Citation V

Initial Pilot Training Manual

SimuFlite

August 2010
NOTICE: This Citation V Initial Pilot Training Manual is to be used
for aircraft familiarization and training purposes only. It is not to be
used as, nor considered a substitute for, the manufacturer’s Pilot
or Maintenance Manual.

Copyright © 2010, CAE SimuFlite, Inc.

All Rights Reserved.

Excerpted materials used in this publication

have been reproduced with permission of
the Cessna Aircraft Company.

Printed in the United States of America.

 Welcome to
SimuFlite
Welcome to SimuFlite!
Our goal is a basic one: to enhance your safety, proficiency
and professionalism within the aviation community. All of
us at SimuFlite know that the success of our company
depends upon our commitment to your needs. We strive
for excellence by focusing on our service to you.
We urge you to participate actively in all training activities.
Through your involvement, interaction, and practice, the
full value of your training will be transferred to the
operational environment. As you apply the techniques
presented through SimuFlite training, they will become
“second nature” to you.
Thank you for choosing SimuFlite. We recognize that you
have a choice of training sources. We trust you will find
us committed to providing responsive, service-oriented
training of the highest quality.
Our best wishes are with you for a most successful and
rewarding training experience.

The Staff of SimuFlite

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 Introduction

This manual is a stand-alone document appropriate for various

levels of training. Its purpose is to serve as an informational Using this
resource and study aid. Manual
The Quick Reference section provides limitations, memory
items from procedural checklists, and other data for quick review.
The Operating Procedures section contains sub-chapters that
provide a pictorial preflight inspection of the aircraft, normal pro-
cedures in an expanded format, standard operating procedures,
maneuvers, and other information for day-to-day operations.
The Flight Planning chapter covers weight and balance and
performance; a sample problem is included.
The Systems section is subdivided by aircraft system. Each
system chapter contains a discussion of components, preflight
and servicing procedures, and abnormal and emergency pro-
cedures. At the beginning of the Systems chapter, a list of
systems is cross referenced to ATA codes to facilitate further
self study, if desired, with the manufacturer’s manuals.
Two graphics are used in this manual to direct your attention to
a specific location in the cockpit. A shaded area locates various
instruments or switches shown in the adjacent text.

The graphic at right (top) represents the cockpit forward panel.

The graphic at right (below) represents the center pedestal.

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This chapter has six major subsections, which are in turn sub-
divided by categories. The General Information subsection Quick
contains information and limitations that pertain to the aircraft
as a whole. The Operational Limits subsection presents func- Reference
tional and structural limits such as:
■ weight
■ speed
Chapter 2
■ takeoff, landing, and enroute operations
■ load factors.
The System Data subsection provides system-by-system limi-
tations and other important data. The Instrument Markings
subsection summarizes system parameters by showing the man-
ufacturer’s markings on specific gages. The AFM Supplements
subsection is a survey of the limits that apply if an aircraft is
modified as indicated by the individual supplements. The System
Data Summaries subsection repeats the charts found at the end
of each system chapter of this manual; these charts summarize
power sources, distribution, controls, and monitors applicable
to the system.
All aircraft limitations are printed in blue.
Some information may be repeated throughout this chapter as
it applies to different categories and subsections.

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 Quick Reference

General Limitations . . . . . . . . . . . . . . . . . . 2-9

Table of
Authorized Operations . . . . . . . . . . . . . . . . . 2-9
Baggage . . . . . . . . . . . . . . . . . . . . . . . . 2-9
Contents
Certification Status . . . . . . . . . . . . . . . . . . . 2-9
Emergency Exit Door Pin . . . . . . . . . . . . . . . . 2-9
Maneuvers . . . . . . . . . . . . . . . . . . . . . . . 2-9
Minimum Crew . . . . . . . . . . . . . . . . . . . . . 2-9
Noise Levels . . . . . . . . . . . . . . . . . . . . . . 2-10
Passenger Seat Position . . . . . . . . . . . . . . . 2-10

Operational Limits . . . . . . . . . . . . . . . . . . 2-11

Weight . . . . . . . . . . . . . . . . . . . . . . . . . 2-11
Maximum Ramp Weight . . . . . . . . . . . . . . . 2-11
Maximum Takeoff Weight . . . . . . . . . . . . . . 2-11
Maximum Landing Weight . . . . . . . . . . . . . . 2-11
Maximum Zero Fuel Weight . . . . . . . . . . . . . 2-11
Center of Gravity Limits . . . . . . . . . . . . . . . . 2-12
Speed Limitations . . . . . . . . . . . . . . . . . . . 2-13
VA, Maximum Maneuvering Speed . . . . . . . . . 2-13
Maximum Maneuvering Speeds . . . . . . . . . . . 2-13
VFE, Maximum Flap Extended Speed . . . . . . . . 2-14
VLE, Maximum Landing Gear Extended Speed . . . 2-14
VLO, Maximum Landing Gear Operating Speed . . . 2-14
VMCA, Minimum Control Speed, Air . . . . . . . . . 2-14
VMCG, Minimum Control Speed, Ground . . . . . . . 2-14
VMO/MMO, Maximum Operating Speed for
Standard Zero Fuel Weight of 11,200 Lbs . . . . . . 2-14
VMO/MMO, Maximum Operating Speed for
Optional Zero Fuel Weight of 12,200 Lbs . . . . . . 2-14
VSB, Maximum Speed Brake Operation Speed . . . 2-14
Autopilot Operation . . . . . . . . . . . . . . . . . 2-14

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 Approach/Landing in Icing Conditions . . . . . . . . . 2-15
Operation in Severe Icing Conditions –
AD 98-04-38 . . . . . . . . . . . . . . . . . . . . . 2-16
Procedures for Exiting the Severe Icing
Environment . . . . . . . . . . . . . . . . . . . . 2-17
Takeoff and Landing Operational Limits . . . . . . . . 2-17
Takeoff/Landing/Enroute Temperature Limitations . . 2-18
Enroute Operational Limits . . . . . . . . . . . . . . 2-19
Load Factors . . . . . . . . . . . . . . . . . . . . . . 2-19
In Flight . . . . . . . . . . . . . . . . . . . . . . . 2-19
Landing . . . . . . . . . . . . . . . . . . . . . . . 2-19
Gravel Runway Modification . . . . . . . . . . . . . 2-19

System Data . . . . . . . . . . . . . . . . . . . . . 2-21

Avionics and Communications . . . . . . . . . . . . 2-21
Angle-of-Attack/Stick Shaker . . . . . . . . . . . . 2-21
Autopilot . . . . . . . . . . . . . . . . . . . . . . . 2-21
OMEGA/VLF . . . . . . . . . . . . . . . . . . . . . 2-21
GPS . . . . . . . . . . . . . . . . . . . . . . . . . 2-22
High Frequency (HF) Automatic Direction
Finder (ADF) System . . . . . . . . . . . . . . . . 2-22
Honeywell SPZ-500 Flight Guidance System
Category II . . . . . . . . . . . . . . . . . . . . . 2-22
Honeywell Single EDZ-605 Electronic
Flight Instrument System (EFIS) . . . . . . . . . . 2-22
Standby Gyro Horizon . . . . . . . . . . . . . . . . 2-23
Electrical and Lighting . . . . . . . . . . . . . . . . . 2-23
Battery Limitations . . . . . . . . . . . . . . . . . . 2-23
Prolonged Ground Operation . . . . . . . . . . . . 2-23
Starter Cycle Limitations . . . . . . . . . . . . . . . 2-24

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 Quick Reference

Fuel . . . . . . . . . . . . . . . . . . . . . . . . . . 2-24
Anti-Ice Additive . . . . . . . . . . . . . . . . . . . 2-24
Fuel Additive . . . . . . . . . . . . . . . . . . . . . 2-24
Approved Fuels . . . . . . . . . . . . . . . . . . . . 2-25
Aviation Gasoline . . . . . . . . . . . . . . . . . . 2-25
Boost Pumps . . . . . . . . . . . . . . . . . . . . . 2-26
Differential . . . . . . . . . . . . . . . . . . . . . . 2-26
Unusable Fuel . . . . . . . . . . . . . . . . . . . . 2-26
Hydraulics . . . . . . . . . . . . . . . . . . . . . . . 2-26
Fluid . . . . . . . . . . . . . . . . . . . . . . . . . 2-26
Ice and Rain Protection . . . . . . . . . . . . . . . . 2-26
Prolonged Ground Operation . . . . . . . . . . . . 2-26
Windshield Ice Protection Fluid . . . . . . . . . . . 2-26
Landing Gear . . . . . . . . . . . . . . . . . . . . . 2-26
Nose Tires . . . . . . . . . . . . . . . . . . . . . . 2-26
Miscellaneous . . . . . . . . . . . . . . . . . . . . . 2-27
Oxygen . . . . . . . . . . . . . . . . . . . . . . . . 2-27
Thrust Reverser . . . . . . . . . . . . . . . . . . . 2-27
Powerplant – Pratt and Whitney
Turbofan JT15D-5A . . . . . . . . . . . . . . . . . 2-28
Approved Oils . . . . . . . . . . . . . . . . . . . . 2-28
Engine Fan Inspection . . . . . . . . . . . . . . . . 2-28
Engine Operating Limits . . . . . . . . . . . . . . . 2-29
Engine Overspeed Limits . . . . . . . . . . . . . . 2-29
Ground Idle Switch . . . . . . . . . . . . . . . . . . 2-29
ITT Limits . . . . . . . . . . . . . . . . . . . . . . 2-30
Minimum Oil Temperature for Starting . . . . . . . . 2-31
Prolonged Ground Operation . . . . . . . . . . . . 2-31
Pneumatic and Pressurization . . . . . . . . . . . . . 2-32
Cabin Pressurization . . . . . . . . . . . . . . . . . 2-32

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 Instrument Markings . . . . . . . . . . . . . . . . . 2-33

AFM Supplements . . . . . . . . . . . . . . . . . . 2-34

Blade Antenna (E-Field) Installation . . . . . . . . . . 2-34
Audio Checklist (with or without Cabin Briefer) . . . . 2-34
Center Panel Mounted Single Clock . . . . . . . . . . 2-34
EROS Oxygen Mask . . . . . . . . . . . . . . . . . 2-34
Flood Cooling System . . . . . . . . . . . . . . . . . 2-35
Global Airborne Flight Information System (AFIS) . . 2-35
Global GNS-X with or without LORAN-C, and/or
VLF/OMEGA and/or GPS Navigation Sensors . . . . 2-36
Global GNS-X with LORAN-C . . . . . . . . . . . . 2-38
Global GNS-X with VLF/OMEGA . . . . . . . . . . 2-38
Global GNS-X with GPS . . . . . . . . . . . . . . . 2-38
Honeywell FMZ-600/800/900 Flight
Management System (FMS) with
VLF/OMEGA Sensor and Optional AFIS . . . . . . . 2-39
Honeywell Dual EDZ-605 Electronic
Flight Instrument System (EFIS) with
Multifunction Display (MFD) . . . . . . . . . . . . . 2-41
Honeywell LSZ-850 Lightning Sensor System . . . . 2-42
Honeywell Primus II SRZ-850 Integrated
Radio System . . . . . . . . . . . . . . . . . . . . 2-43
Honeywell Single EDZ-605 Electronic
Flight Instrument System (EFIS) with
Multifunction Display (MFD) . . . . . . . . . . . . . 2-43
J.E.T. ADI-330 Attitude Indicator . . . . . . . . . . . 2-44
Maximum Design Zero Fuel Weight
of 12,200 LBS . . . . . . . . . . . . . . . . . . . . 2-45
UNS-1A Flight Management System with
CMA-764 VLF/OMEGA Sensor . . . . . . . . . . . 2-45

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 Quick Reference

System Data Summaries . . . . . . . . . . . . . . . 2-49

Air Conditioning System Data . . . . . . . . . . . . . 2-49
Electrical System Data . . . . . . . . . . . . . . . . 2-50
Fire Protection Data . . . . . . . . . . . . . . . . . . 2-51
Flight Controls Systems Data . . . . . . . . . . . . . 2-51
Flaps Data . . . . . . . . . . . . . . . . . . . . . . 2-51
Speedbrakes Data . . . . . . . . . . . . . . . . . . 2-51
Fuel System Data . . . . . . . . . . . . . . . . . . . 2-52
Hydraulic System Data . . . . . . . . . . . . . . . . 2-53
Ice and Rain Protection Data . . . . . . . . . . . . . 2-54
Engine Anti-Ice Data . . . . . . . . . . . . . . . . . 2-54
Windshield Bleed Air Data . . . . . . . . . . . . . . 2-54
Surface Deice Data . . . . . . . . . . . . . . . . . 2-54
Pitot/Static Deice Data . . . . . . . . . . . . . . . . 2-55
Landing Gear System Data . . . . . . . . . . . . . . 2-56
Landing Gear Data . . . . . . . . . . . . . . . . . . 2-56
Brakes/Anti-Skid Data . . . . . . . . . . . . . . . . 2-56
Oxygen Systems Data . . . . . . . . . . . . . . . . . 2-57
Pneumatic Systems Data . . . . . . . . . . . . . . . 2-58
Bleed Air System Data . . . . . . . . . . . . . . . . 2-58
Pressurization System Data . . . . . . . . . . . . . 2-59
Thrust Reverser Systems Data . . . . . . . . . . . . 2-60

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 Quick Reference

Authorized Operations General

■ Day and night
■ VFR and IFR flight Limitations
■ Flight into icing conditions

Baggage
Nose Baggage Compartment . . . . . . . . . . . . . . 350 LBS
Aft Cabin Storage Compartment . . . . . . . . . . . . . 600 LBS
Aft Cabin Storage Floor Load Limit . . . . . . . 120 LBS/SQ FT
Tailcone Baggage Compartment . . . . . . . . 500 LBS TOTAL
Forward Section . . . . . . . . . . . . . . . . . . . . . 300 LBS
Aft Section . . . . . . . . . . . . . . . . . . . . . . . . 200 LBS

Certification Status
FAR, Part 25

Emergency Exit Door Pin

The emergency exit door pin must be removed prior to flight.

Maneuvers
■ No acrobatic maneuvers, including spins
■ No intentional stalls above 25,000 ft

Minimum Crew
Pilot and copilot

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 Noise Levels
■ The Effective Perceived Noise Levels (EPNL) established in
compliance with FAR 36, Amendment 12 (Stage 3) are shown in
Table 2-A.
■ The ICAO Annex 16 Chapter 3 noise values are the same as those
for FAR Part 36, Amendment 12, and were obtained with the
procedures used to establish compliance with FAR Part 36,
Amendment 12. See Table 2-A.
■ The takeoff weight for these figures was 15,900 lbs; the landing
weight was 15,200 lbs.
■ Takeoff and sideline noise levels were established with a climb
speed of 125 KIAS and with 7° flaps. The thrust was reduced at
3,565 ft AGL from the takeoff N1% RPM to 82.7 N1% RPM.
■ Approach noise levels were obtained at 116 KIAS and FULL
flaps.
■ No determination has been made by the FAA that the noise lev-
els in the AFM are or should be acceptable or unacceptable for
operation at, into, or out of any airport.
■ ICAO data applies only after approval of the Civil Aviation
Approval Authority of the country of the aircraft’s registration
(including approval of equivalent procedures used to establish
FAR 36, Amendment 12 compliance).

Passenger Seat Position

For all takeoffs and landings, the following apply.
NOTE: Unobstructed access to
■ Seats must be fully upright and outboard.
the emergency exit is fostered by
moving the adjacent seat aft. ■ The seat just aft of emergency exit must be in most aft position.

Noise Reference FAR 36 ICAO Annex 16, Chapter 3

Takeoff 83.7 68.7

Sideline 94.7 ––

Approach 88.9 80.5

Table 2-A; Effective Perceived Noise Levels (EPNL)

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 Quick Reference

Weight Operational
Maximum Ramp Weight
Weight . . . . . . . . . . . . . . . . . . . . . . . . 16,100 LBS
Limits
Maximum Takeoff Weight
Weight . . . . . . . . . . . . . . . . . . . . . . . . 15,900 LBS
Takeoff weight is limited by the most restrictive of:
■ maximum certified takeoff weight

■ maximum takeoff weight permitted by climb requirements

■ takeoff field length.

Maximum Landing Weight

Maximum Weight . . . . . . . . . . . . . . . . . . 15,200 LBS
Landing weight is limited by the most restrictive of:
■ maximum certified landing weight

■ maximum landing weight permitted by climb requirements or

brake energy limit

■ landing distance.

Maximum Zero Fuel Weight

Standard . . . . . . . . . . . . . . . . . . . . . . . . 11,200 LBS
Optional . . . . . . . . . . . . . . . . . . . . . . . . 12,200 LBS

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 Center of Gravity Limits

2-1

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 Quick Reference

Speed Limitations
VA, Maximum Maneuvering Speed
■ Refer to Figure 2-2.
■ Full application of rudder and aileron controls as well as maneu-
vers that involve angles of attack near the stall should be confined
to speeds below maximum maneuvering speed.
Maximum Maneuvering Speeds

2-2

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 VFE, Maximum Flap Extended Speed
Full Flaps, LAND Position (35°) . . . . . . . . . . . . 173 KIAS
Partial Flaps, T.O., and T.O. & APPR. Position . . . . 200 KIAS
VLE, Maximum Landing Gear Extended Speed
Speed . . . . . . . . . . . . . . . . . . . . . . . . . . 292 KIAS
VLO, Maximum Landing Gear Operating Speed
Extending . . . . . . . . . . . . . . . . . . . . . . . . 250 KIAS
Retracting . . . . . . . . . . . . . . . . . . . . . . . . 200 KIAS
VMCA, Minimum Control Speed, Air
Speed . . . . . . . . . . . . . . . . . . . . . . . . . . . 85 KIAS
VMCG, Minimum Control Speed, Ground
Speed . . . . . . . . . . . . . . . . . . . . . . . . . . . 86 KIAS
VMO/MMO, Maximum Operating Speed for Standard Zero
Fuel Weight of 11,200 Lbs
VMO Below 8,000 Ft . . . . . . . . . . . . . . . . . . 261 KIAS
MMO Between 8,000 and 28,907 Ft . . . . . . . . . . 292 KIAS
MMO Above 28,907 Ft . . . . . . . . . . . 0.755 M INDICATED
The maximum operating limit speeds may not be deliberately
exceeded in any regime of flight (climb, cruise, or descent) unless a
higher speed is authorized for flight test or pilot training.
VMO/MMO, Maximum Operating Speed for Optional Zero
Fuel Weight of 12,200 Lbs
VMO Below 8,000 Ft . . . . . . . . . . . . . . . . . . 261 KIAS
MMO Between 8,000 and 31,400 Ft . . . . . . . . . . 276 KIAS
MMO Above 31,400 Ft . . . . . . . . . . . 0.755 M INDICATED
The maximum operating limit speeds may not be deliberately
exceeded in any regime of flight (climb, cruise, or descent) unless a
higher speed is authorized for flight test or pilot training.
VSB, Maximum Speed Brake Operation Speed
Speed . . . . . . . . . . . . . . . . . . . . . . . . . NO LIMIT
Autopilot Operation
Speed . . . . . . . . . . . . . . . . . . . 292 KIAS OR 0.755 M

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 Quick Reference

Approach/Landing In Icing Conditions WARNING: With any residual

■ When any residual ice is present or can be expected during ice present, do not attempt to fly
approach and landing, VREF and VAPP must be increased. VREF and uncorrected VREF/VAPP speeds.
VAPP, the landing distance, and the maximum weight permitted by Stall speeds increase and stall
brake energy must be corrected per Table 2-B. warning margins decrease.
■ When reconfiguring for approach and landing (i.e., flaps extended
and gear down) with any ice accretion visible on the wing leading
edge, regardless of thickness, activate the surface deice system.
Continue to monitor wing leading edge for any reaccumulation. CAUTION: In icing conditions,
a small amount of residual ice
■ For increased rates of descent icing conditions, Cessna recom- forms on unprotected areas. This
mends use of landing flaps. This allows a higher power setting, if is normal, but can cause an
necessary, to maintain anti-icing capabilities.
increase in stall speeds. When
any amount of residual ice is
Residual Ice – Landing Corrections Factors visible, the stall speeds increase
by 8 kts; the VREF/VAPP speeds
Maximum Landing
Weight Limited By and landing distances and the
Landing Climb or Brake Energy Landing VAPP VREF maximum landing weight per-
Field Anti-Ice Anti-Ice Distance mitted by brake energy must be
Conditions On Off corrected per the chart below.
Field
Elevation
4000 ft or
below; No No No Multiply Add 8 Add 8
downhill Correction Correction by 1.30 knots knots
gradient;
No tailwind

Field
Elevation
Above No Multiply Multiply Add 8 Add 8
4000 ft to Correction by 0.88 by 1.30 knots knots
10,000 ft

All Other No Multiply Multiply Add 8 Add 8

Conditions Correction by 0.73 by 1.30 knots knots

Table 2-B; Residual Ice – Landing Correction Factors

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 Operation in Severe Icing Conditions –
AD 98-04-38
Severe icing may result from environmental conditions outside of
those for which the aircraft is certificated. Flight in freezing rain,
freezing drizzle, or mixed icing conditions (supercooled liquid water
and ice crystals) may result in ice build-up on protected surfaces
exceeding the capability of the ice protection system, or may result
in ice forming aft of the protected surfaces. This ice may not be shed
using the ice protection systems, and may seriously degrade the per-
formance and controllability of the airplane.
During flight, severe icing conditions that exceed those for which the
aircraft is certificated are determined by the following visual cues.
If one or more of these visual cues exists, immediately request pri-
ority handling from ATC to facilitate a route or an altitude change to
exit the icing conditions.
■ Unusually extensive ice accumulation on the airframe and wind-
shield in areas not normally observed to collect ice.
■ Accumulation of ice on the upper surface of the wing aft of the pro-
tected area.
Since the autopilot, when installed and operating, may mask tactile
cues that indicate adverse changes in handling characteristics, use of
the autopilot is prohibited when any of the visual cues specified above
exist, or when unusual lateral trim requirements or autopilot trim
warnings are encountered while the aircraft is in icing conditions.
All wing icing inspection lights must be operative prior to flight into
know or forecast icing conditions at night. This supersedes any relief
provided by the Master Minimum Equipment List (MMEL).
The following weather conditions may be conducive to severe in-
flight icing:
■ Visible rain at temperatures below 0°C ambient air temperature.

■ Droplets that splash or splatter on impact at temperatures below

0°C ambient air temperature.

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 Quick Reference

Procedures for Exiting the Severe Icing Environment

These procedures are applicable to all flight phases from takeoff to
landing. Monitor the ambient air temperature. While severe icing
may form at temperatures as cold as -18°C, increased vigilance is war-
ranted at temperatures around freezing with visible moisture pre-
sent. If the visual cues specified in the Limitations Section of the
AFM for identifying severe icing conditions are observed, accom-
plish the following:
■ Immediately request priority handling from ATC to facilitate a route
or an altitude change to exit the sever icing conditions in order to
avoid extended exposure to flight conditions more severe than those
for which the airplane has been certificated.
■ Avoid abrust and excessive maneuvering that may exacerbate con-
trol difficulties.
■ Do not engage the autopilot.

■ If autopilot is engaged, hold control wheel firmly and disengage

the autopilot.
If an unusual roll response or uncommanded roll control movement
is observed, reduce angle-of-attack.
Do not extend flaps when holding in icing conditions. Operation
with flaps extended can result in a reduced wing angle-of-attack,
with the possibility of ice forming on the upper surface further aft on
the wing than normal, possibly aft of the protected area.
If the flaps are extended, do not retract until airframe is clear of ice.
Report these weather conditions to ATC.

Takeoff and Landing Operational Limits

Maximum Altitude Limit . . . . . . . . . . . . . . . . 14,000 FT
Maximum Ambient Temperature (Figure 2-3) . . . . ISA + 39°C

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 Takeoff/Landing/Enroute Temperature Limitations

* Maximum Enroute Operating Temperature Limit is ISA +39°C ambient temperature

adjusted for Ram Rise or the Indicated Outside Air Temperature from Figure 2-7,
whichever is less.

2-3

Minimum Ambient Temperature . . . . . . . . . . . . . . -54°C

Maximum Crosswind DEMONSTRATED . . . . . . . . 20 KTS
Maximum Tailwind Component . . . . . . . . . . . . . 10 KTS
Maximum Tire Groundspeed . . . . . . . . . . . . . . 165 KTS
Maximum Water/Slush on Runway . . . . . . . . . . . 0.5 INCH
■ The autopilot and yaw damper must be off for takeoff and landing.

■ Vertical navigation system must be off below 500 ft AGL.

Nose Tire Pressure . . . . . . . . . . . . . . . . . . . 120 ±5 PSI

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 Quick Reference

Enroute Operational Limits

Maximum Operating Altitude . . . . . . . . . . . . . 45,000 FT
Maximum Ambient Temperature . . . . . REFER TO Figure 2-3
Minimum Ambient Temperature . . . . . REFER TO Figure 2-3
Generator Load . . . . . . . . . . . . . . . . . . . . 300 AMPS

Load Factors
In Flight
Flaps Up Position (0°) . . . . . -1.52 TO +3.8 G AT 15,900 LBS
Flaps T.O., T.O.
& APPR To LAND
Position (7 to 35°) . . . . . . . . 0.0 TO +2.0 G AT 15,900 LBS
These accelerations limit the angle-of-bank in turns and severity of
pullup maneuvers.
Landing
Flaps T.O. &
APPR To LAND
Position (15 to 35°) . . . . . . . . . . . . +3.5 G AT 15,200 LBS

Gravel Runway Modification CAUTION: Damage to the air-

For aircraft equipped for gravel runway operations, the following craft structure and engines may
limitations apply. occur when operations are
conducted on gravel. Exercise
VLE, Maximum Landing Gear Extended . . . . . . . . 200 KIAS extreme care in maintaining
VLO, Maximum Landing Gear Operating proper approach speeds. High
(Extending or Retracting) . . . . . . . . . . . . . . . . 200 KIAS speeds during touchdown may
■ The anti-skid system must be operational and on for takeoff and
cause gravel damage from the
landing on gravel runways. nose wheel.

■ The nosewheel spin-up system shall be operational and on for

landing on a gravel runway.

■ Single pilot operation, per Exemption 4050A, is prohibited when

operating from gravel runways.

■ The nosewheel steering must be disengaged for takeoff and/or

landing operation on gravel runways.

■ Observe the Caution at right.

Citation V For SimuFlite training only 2-19

March 1999
2-20 For SimuFlite training only Citation V
March 1999
 Quick Reference

Avionics and Communications System Data

Angle-of-Attack/Stick Shaker
■ Use the angle-of-attack (AOA) indicating system as a reference,
but do not use it as a replacement for the airspeed indicator as a
primary instrument.
■ Use the AOA system as a reference for approach speed (1.3 VS1)
at all aircraft weights and center-of-gravity locations at zero, take-
off, takeoff/approach, and landing flap positions.
■ If the stick shaker does not operate during the warning system
test, or the AOA system is otherwise inoperative, it must be repaired
before flight, except when operating the aircraft according to an
approved Minimum Equipment List (MEL).
Autopilot
■ One pilot must remain in his seat with seat belt fastened during all
autopilot operations.
■ Functionally test the autopilot torque monitor and roll monitor
per the procedures in AFM Section III prior to inflight use of the
autopilot.
■ Autopilot operation is prohibited above 14,500 ft if the torque
monitor (AP TORQUE annunciator) does not test per the normal
procedures in AFM, Section III.
■ Autopilot operation is prohibited if the roll monitor annunciator
light illuminates in flight.
■ The autopilot and yaw damper must be off for takeoff and landing.
■ The vertical navigation system must be off below 500 ft AGL.
OMEGA/VLF
■ If the UNS-1 OMEGA/VLF sensor is receiving usable signals
from at least two OMEGA navigation stations, the UNS-1
OMEGA/VLF sensor is demonstrated to be capable of and has
been shown to meet the requirements of:
– VFR/IFR RNAV operation within the conterminous United
States and Alaska in accordance with enroute criteria of AC
20-130 or the criteria of AC 20-101C.
– Flight in the NAT-MNPS Airspace in accordance with AC 120-
33 or AC 91-49.
■ The UNS-1 with OMEGA/VLF only operating is not to be used
as a primary reference during operation into valleys, i.e., between
peaks in mountainous terrain.

Citation V For SimuFlite training only 2-21

March 1999
 GPS
CAUTION: When operating the
UNS-1 GPS/OMEGA/VLF sys- ■ The Global Positioning System (GPS) sensor is approved as a
tem, proximity to buildings, contributing sensor for navigation if:
hangars, power lines, etc., may – The GPS sensor is not the only sensor selected for use.
cause interference with the oper-
ation of the system. Verify posi-
– It is receiving four or more satellites.
tion coordinates prior to departure. High Frequency (HF) Automatic Direction Finder (ADF)
System
The ADF bearing information may be erratic when keying the HF
transmitter. Should this occur, disregard the ADF bearing during
periods of transmission.
Honeywell SPZ-500 Flight Guidance System –
Category II
■ Equipment operation must be according to the Category II Manual.
Category II operation is approved only from the left seat with the
SPZ-500 system mounted for use from the pilot’s instrument panel.
■ The autopilot must be off at 80 ft AGL for approaches with the flaps
in the landing position.
■ The marker beacon audio muting must not be activated at the mid-
dle marker on a Category II approach. If it is activated, the inner
marker audio remains muted because of the short time between
markers.
■ Prior to initiating a coupled Category II approach, the pilot must
verify that the autopilot is functioning.
■ Make Category II approaches in the following configurations
only: flaps in LAND, gear down, and both engines operating.
Honeywell Single EDZ-605 Electronic Flight
Instrument System (EFIS)
■ The Honeywell SPZ-500C Integrated Flight Control System Pilot’s
NOTE: Crew qualification is Manual for the Cessna Citation II/V (Publication EDZ-605/805,
required to conduct Category II number 28-1146-63-00, dated November 1989 or later revision)
approaches. must be immediately available to the flight crew for aircraft
The Honeywell EDZ-605 EFIS equipped with the Honeywell Single EDZ-605/805 Electronic
is compatible with the Flight Instrument System.
Honeywell SPZ-500 Flight
Guidance System.
■ Both flight director- and autopilot-coupled Category II approach-
es are approved for using the Honeywell EDZ-605 EFIS displays.
■ Category II approaches are not approved in the composite mode
(REV selected).
■ Operating in the composite mode (REV selected) is approved only
with the flight director selected.

2-22 For SimuFlite training only Citation V

March 1999
 Quick Reference

■ Limit EFIS ground operation with the pilot’s DISP FAN annun-
ciator light illuminated to 10 minutes or until either EADI HOT
or EHSI HOT annunciator illuminates, whichever occurs first.
■ Do not dispatch if either the EADI HOT, EHSI HOT, or SG HOT
annunciator illuminates.
■ Do not dispatch in instrument meteorological conditions with the
pilot’s DISP FAN annunciator illuminated. Dispatch in visual
meteorological conditions is allowed with the DISP FAN annun-
ciator illuminated if the DISP FAN Illuminated on Ground abnor-
mal procedures are followed.
■ Do not dispatch following a flight where either an EADI HOT or
EHSI HOT annunciator light is illuminated until identifying and
correcting the condition.
■ The pilot’s EADI and EHSI must be installed and operational in
the normal (non-reversionary) mode for takeoff.
■ Verify the EDZ-605 system as operational per a satisfactory pre-
flight test.
■ On unit 0001 to 0074, limit operations on the ground with or with-
out engines operating as shown below.
Below 45°C . . . . . . . . . . . . . . . . . . . . NO LIMIT
Between 45 and 51°C . . . . . . . . . . ONE HOUR LIMIT
Above 51°C . . . . . . . . NO OPERATIONS PERMITTED
Standby Gyro Horizon
Accomplish a satisfactory preflight test on the standby gyro system.

Electrical and Lighting NOTE: If a battery limitation is

Battery Limitations exceeded, accomplish a deep
cycle (including a capacity check)
Engine Starts . . . . . . . . . . . . . . . . THREE PER HOUR to detect possible cell damage.
■ If the BATT O’TEMP light illuminates during ground operation, Three generator-assisted cross
do not take off until after the proper maintenance procedures have starts equal one battery start.
been accomplished. If using a GPU for start, do not
■ Refer to the Note at right. count a battery cycle.
Use of a GPU with voltage in
Prolonged Ground Operation excess of 28V DC or current in
excess of 1,000 amps may
Continuous ground operation of generator above 125 amps at ground damage the starter.
idle (46% turbine speed) or 225 amps at flight idle (52% turbine
speed) is prohibited.

Citation V For SimuFlite training only 2-23

March 1999
 Starter Cycle Limitations
With GPU or Generator-Assisted Cross Start As Starter Power Source
Engine Starts . . . . . . . . . . . . TWO PER 30 MINUTES
Cycles of Operation . . . . . . . TWO WITH A 30-SECOND
. . . . . . . . . . . . . . . . . . . . . . . . .REST PERIOD
. . . . . . . . . . . . . . . . . . . . . .BETWEEN CYCLES
With Battery As Starter Power Source
Engine Starts . . . . . . . . . . . THREE PER 30 MINUTES
Cycles of Operation . . . . . THREE WITH A 30-SECOND
. . . . . . . . . . . . . . . . . . . . . . . . .REST PERIOD
. . . . . . . . . . . . . . . . . . . . . .BETWEEN CYCLES

WARNING: Anti-ice additive

Fuel
is combustible and toxic. It is dan- Anti-Ice Additive
gerous when inhaled and/or
absorbed into the skin. When in
■ Add anti-icing additive (MIL-I-27686E) to all fuels not present-
contact with anti-ice additive, use
ly containing the additive.
appropriate protective equipment ■ Use a minimum anti-icing additive concentration of 0.06% by
(e.g., eye goggles/shield, respira- volume and a maximum of 0.15% by volume.
tor with organic vapor cartridges, ■ Use not less than 20 fluid ounces of additive per 260 gallons of fuel
non-absorbing gloves, additional or more than 20 fluid ounces of additive per 104 gallons of fuel.
skin protection). If anti-ice enters
the eyes, flush with water and ■ Insufficient additive concentrations may result in fuel system
contact a physician immediately icing. Excessive additive may cause fuel tank damage or erro-
neous fuel quantity indications.
Fuel Additive
■ The minimum additive concentration of EGME shall be 0.06%
by volume, and maximum concentration shall be 0.15% by
volume. Fuel, when added to the tank, should have a minimum
concentration of 0.06% by volume.
■ Use not less than 20 fluid ounces of EGME additive per 260
gallons of fuel or more than 20 fluid ounces of EGME additive per
104 gallons of fuel.
■ The minimum additive concentration for DIEGME shall be 0.10%
by volume, and maximum concentration shall be 0.15% by
volume. Fuel, when added to the tank, should have a minimum
concentration of 0.10% by volume.
■ Use not less than 20 fluid ounces of DIEGME additive per 156
gallons of fuel or more than 20 fluid ounces of DIEGME addi-
tive per 104 gallons of fuel.

2-24 For SimuFlite training only Citation V

March 1999
 Quick Reference

Approved Fuels
Condition Per Canadian Pratt and Aviation NOTE: Military JP-4, JP-5, and
Whitney Specification 204 Gasoline JP-8 have anti-icing pre-blended
Per by the refinery.
Jet A, A-1, Jet B and MIL-G-5572
A-2, JP-5, JP-4
and JP-8
CAUTION: Direct anti-ice addi-
Minimum Fuel Temperature -29 -54 -54 tive into the flowing fuel stream
(Takeoff) (°C)
and start the additive flow after
Minimum Fuel Temperature -23 -54 -54 the fuel flow starts; stop the
(Starting) (°C) additive flow before fuel flow
Maximum Fuel Temperature +50 +50 +32
stops. Do not allow concentrat-
(°C) ed additive to contact coated
interior of fuel tank or aircraft
Maximum Altitude (ft) 45,000 45,000 18,000 painted surface.
Maximum Asymmetric 200 200 200
Fuel Differential for Normal
Operations (lbs)

Emergency Asymmetric 600 600 600

Fuel Differential1 (lbs)

Fuel Control Density 0.81 0.79 0.73

Adjustment for Optimum
Engine Acceleration

Table 2-C; Approved Fuels and Associated Conditions

1 Flight characteristic requirements were not demonstrated with unbalanced fuel
above 200 lbs.

Aviation Gasoline
All grades of MIL-G-5572 avgas are permitted for a maximum of 50
hours or 3,500 gallons between overhauls if:
■ The pilot confirms fuel temperature is within limits.

■ The maximum ambient air temperature at takeoff is +32°C.

■ The boost pumps are on. (To crossfeed, turn the boost pump off
on the side opposite the selected tank.)
■ The hours avgas used is entered in the engine logbook. (For record
keeping purposes, assume one hour of engine operation equals
70 gallons of gasoline.)
■ The maximum operating altitude is 18,000 ft.

Citation V For SimuFlite training only 2-25

March 1999
 Boost Pumps
The boost pumps must be on when the low fuel lights illuminate or
when indicated fuel is 185 lbs or less.
Differential
Maximum Asymmetric
Fuel Differential in Flight . . . . . . . . . . . . . . . . 200 LBS
Maximum Asymmetric Fuel
Differential in an Emergency . . . . . . . . . . . . . . 600 LBS
Flight characteristic requirements were not demonstrated with unbal-
anced fuel above 200 lbs.
Unusable Fuel
Fuel remaining in the fuel tanks when the fuel quantity indicator
reads zero is not usable in flight.

Hydraulics
Fluid
Use Skydrol 500A, B, B-4, C or LD -4; or use Hyjet, Hyjet W, III,
or IV.

Ice and Rain Protection

Prolonged Ground Operation
Limit ground operation of pitot/static heat to two minutes to pre-
clude damage to angle-of-attack system.
Windshield Ice Protection Fluid
Use TT-I-735 isopropyl alcohol for windshield anti-ice.

Landing Gear
Nose Tires
■ BF Goodrich/Michelin part number 031-613-8 and Goodyear part
number 184F08-1 and 184F13-5 are the nose tires approved.
■ The nose tire must be inflated to 120 PSI ±5 PSI.

2-26 For SimuFlite training only Citation V

March 1999
 Quick Reference

Miscellaneous
Oxygen
■ The pressure demand sweep-on oxygen mask must be properly
stowed to qualify as a quick donning oxygen mask.
■ Continuous use of the supplemental oxygen system is prohibited
above the following cabin altitudes:
Carrying Passengers . . . . . . . . . . . . . . . . . 25,000 FT
Crew Only . . . . . . . . . . . . . . . . . . . . . . 37,000 FT
■ Headsets, eyeglasses, or hats worn by the crew may interfere with
the quick donning capabilities of the optional oxygen masks.
Thrust Reverser
■ Reduce reverse thrust power to idle reverse at 60 KIAS on land-
ing roll.
■ Limit maximum reverse thrust setting to 86% fan speed for ambi-
ent temperatures at or above -18°C and 79% for ambient temper-
atures below -18°C.
■ Maximum allowable thrust reverser deployed time is 15 minutes
in any one hour period.
■ Limit engine static ground operation to idle power if thrust reversers
are deployed.
■ Do not use thrust reversers during touch and go landings.
■ Verify the thrust reverser(s) as operational using the Before Takeoff
test in the AFM normal procedures.
■ Thrust reverser not approved for use on gravel or sod runways.

Citation V For SimuFlite training only 2-27

March 1999
 Powerplant – Pratt and Whitney Turbofan
JT15D-5A
Approved Oils
■ Use only the following oils:
■ Mobil Jet Oil II

■ Mobil Jet Oil 254

■ Exxon Turbo Oil 2380

■ Castrol 5000

■ Royco Turbine Oil 500

■ Royco Turbine Oil 560

■ Aero Shell Turbine Oil 500

■ Aero Shell Turbine Oil 560.

■ Oils listed for the engine in the latest revision of Pratt and Whitney
Canada Inc. SB Number 7001 are approved for use.
■ For contingency purposes, oil replenishment using any approved
oil brand listed is permitted if either of the following is true.
CAUTION: When changing
■ The total quantity of additional oil does not exceed two U.S.
from an existing lubricant for-
quarts in any 400-hour period.
mulation to a “third generation”
lubricant formulation (e.g., Aero ■ If more than two U.S. quarts of dissimilar oil brands must be

Shell/Royco Turbine Oil 560 or added, drain and flush the complete oil system, then refill with
Mobil Jet 254), the engine man- an approved brand according to Engine Maintenance Manual
ufacturer strongly recommends instructions.
that such a change should only ■ If oils of nonapproved brands or of different viscosities become
be made when an engine is new intermixed, drain and flush the complete oil system, then refill
or freshly overhauled. For addi- with an approved oil according to the Engine Maintenance Manual
tional information on use of third instructions.
generation oils, refer to the
engine manufacturer’s pertinent Engine Fan Inspection
oil service bulletins.
To ensure accurate fan speed thrust indication, inspect fan for dam-
age prior to each flight.

2-28 For SimuFlite training only Citation V

March 1999
 Quick Reference

Engine Operating Limits

Operating Operating Limits

Conditions

Thrust Time Limit ITT Temp. N2% Turbine N1% Fan Oil Pressure Oil Temp.
Setting (Minutes) °C RPM RPM PSIG2 °C

Takeoff 5 700 96 1044 60 to 833 10 to 121

Maximum Continuous 680 96 1044 60 to 83 10 to 121

Continuous

Idle Flight Continuous 580 52 Minimum –– 40 Minimum -40 to 121

Ground Continuous 580 46 Minimum –– 40 Minimum -40 to 121

Start –– Note 1 –– –– –– -40 Minimum

Transient –– 700 96 106 Note 3 -18 to 129

Table 2-D; Engine Operating Limits

1 The maximum start limit is 700°C for two seconds. See Figure 2-5.
2 Normal oil pressures are 60 to 83 PSIG above 52% turbine RPM. Oil pressures
below 60 PSIG are undesirable and should be tolerated only for completion of the
flight, preferably at a reduced power setting. Oil pressures below 40 PSIG are
unsafe and require that either the engine be shut down or a landing be made as
soon as possible using the minimum power required to sustain flight.
3 The maximum transient oil pressure can be 95 PSIG for 90 seconds.
4 Refer to the appropriate thrust setting charts in AFM Section IV for percent FAN
RPM setting.

Citation V For SimuFlite training only 2-29

March 1999
 Engine Overspeed Limits

2-4

Ground Idle Switch

■ The ground idle switch must be in HIGH when conducting touch
and go landings.
■ The ground idle switch must be in HIGH when operating on the
ground with engine anti-ice bleed ON.

2-30 For SimuFlite training only Citation V

March 1999
 Quick Reference

ITT Limits

2-5

Minimum Oil Temperature for Starting

Temperature . . . . . . . . . . . . . . . . . . . . . . . . . -40°C
Prolonged Ground Operation
Limit continuous engine ground static operation, up to and includ-
ing five minutes at takeoff thrust, to ambient temperatures not exceed-
ing ISA +39°C.

Pneumatic and Pressurization

Cabin Pressurization
Normal Cabin Pressurization . . . . . . . - 0.0 to 8.9, ±0.1 PSID

Citation V For SimuFlite training only 2-31

March 1999
2-32 For SimuFlite training only Citation V
March 1999
 Quick Reference

Airspeed Indicator, Standard Zero Fuel Weight of

11,200 LBS Instrument
Red Lines . . . . . . . . . . . . 261 KIAS, 292 KIAS, 0.755 M Markings
Airspeed Indicator, Optional Zero Fuel Weight of
12,200 LBS
Red Lines . . . . . . . . . . . . . . . . . . 276 KIAS, 0.755 MI
Ammeter Indicators
Red Line . . . . . . . . . . . . . . . . . . . . . . . . 300 AMPS
Brake and Gear Pneumatic Pressure Indicator
Wide Red Arc . . . . . . . . . . . . . . . . . ABOVE 2,050 PSI
Wide Green Arc . . . . . . . . . . . . . . . 1,800 TO 2,050 PSI
Yellow Arc . . . . . . . . . . . . . . . . . . 1,600 TO 1,800 PSI
Narrow Red Arc . . . . . . . . . . . . . . . . 0.0 TO 1,600 PSI
Brake Hydraulic Accumulator Pressure Indicator
Light Green Arc . . . . . . . . . . . . . . STATIC PRESSURE
Dark Green Arc . . . . . . . . . . PRESSURIZED PRESSURE
Cabin Differential Pressure Indicator
Red Line . . . . . . . . . . . . . . . . . . . . . . . . . . 8.9 PSI
Green Arc . . . . . . . . . . . . . . . . . . . . . 0.0 TO 8.9 PSI
Fan RPM Indicators
Red Line . . . . . . . . . . . . . . . . . . . . . . . . . . 104.0%
Green Band . . . . . . . . . . . . . . . . . . . . 25 TO 104.0%
ITT Indicators
Red Line . . . . . . . . . . . . . . . . . . . . . . . . . . 700°C
Yellow Band . . . . . . . . . . . . . . . . . . 680°C TO 700°C
Green Band . . . . . . . . . . . . . . . . . . . 150°C TO 680°C
Oil Pressure Indicators
Red Line . . . . . . . . . . . . . . . . . . . . . . . . . . 40 PSI
Yellow Band . . . . . . . . . . . . . . . . . . . . 40 TO 60 PSI
Green Band . . . . . . . . . . . . . . . . . . . . . 60 TO 83 PSI

Citation V For SimuFlite training only 2-33

March 1999
 Oil Temperature Indicators
Red Line . . . . . . . . . . . . . . . . . . . . . . . . . . 121°C
Green Band . . . . . . . . . . . . . . . . . . . . 0°C TO 121°C
Oxygen Pressure Indicator
Red Line . . . . . . . . . . . . . . . . . . . . . . . . 2,000 PSI
Green Arc . . . . . . . . . . . . . . . . . . 1,600 TO 1,800 PSI
Yellow Arc . . . . . . . . . . . . . . . . . . . . 0.0 TO 400 PSI
Turbine RPM Indicators
Red Light and Flashing Readout . . . . . . . . . . . . 96% RPM
Normal Operating . . . . . . . . . . . . . . . . 46 TO 96% RPM

2-34 For SimuFlite training only Citation V

March 1999
 Quick Reference

The following originates in selected AFM supplements. For the sup-

plements applicable to a specific aircraft, refer to the Supplements AFM
section in the AFM for that aircraft.
Supplements
Blade Antenna (E-Field) Installation
Used with Global GNS-X VLF/OMEGA
Flight in P-static conditions (e.g., snow, rain, dust) causes
VLF/OMEGA signal degradation. An associated decrease in posi-
tion certainty may occur during these encounters.

Audio Checklist (with or without

Cabin Briefer)
■ The audio checklist must contain FAA-approved operating pro-
cedures. It is the operator’s responsibility to ensure checklist con-
tents are consistent with current AFM revisions.

Center Panel Mounted Single Clock

■ Single pilot operation is prohibited when a center panel mounted
single clock is installed.

EROS Oxygen Mask

■ Prior to flight, check the EROS oxygen mask and properly stow
it in its receptacle to qualify it as a quick-donning oxygen mask.
■ Continuous use of the supplemental oxygen system is prohibited
above the following cabin altitudes:
Carrying Passengers . . . . . . . . . . . . . . . . . 25,000 FT
Crew Only . . . . . . . . . . . . . . . . . . . . . . 40,000 FT
■ Headsets, eyeglasses or hats worn by the crew may interfere with
the quick-donning capability of the oxygen mask.
■ Unless carefully trimmed, mustaches and/or beards worn by
crewmembers may interfere with proper sealing of the oxygen
mask. Check mask fit and seal while on the ground prior to flight.
■ The oxygen supply has the duration shown in Figure 2-6, follow-
ing page. AFM Supplement 25, pg 5/6.

Citation V For SimuFlite training only 2-35

March 1999
 Oxygen Supply Chart

2-6

Flood Cooling System

■ Do not operate the flood cooling system above 10,000 ft pressure
altitude.
■ Do not use the flood cooling system for cabin heating.

Global Airborne Flight Information

System (AFIS)
■ The Global System, Incorporated, Operator’s Manual GNS-X
(Report Number 1280, dated April 1, 1988, or later revision) and
AFIS Operator’s Manual Supplement for Global-Wulfsburg
Systems GNS-X (Report Number 1278, dated March 1989, revi-
sion 2 or later revision) must be immediately available to the flight
crew for aircraft equipped with the Global AFIS.
■ Navigation predicated on the GNS-X is prohibited until the
navigation program is verified to be version PROG 02.
■ The pilot in command must exercise reasonable and prudent
judgment in the use of the advisory services of AFIS.

2-36 For SimuFlite training only Citation V

March 1999
 Quick Reference

Global GNS-X with or without LORAN-C,

and/or VLF/OMEGA and/or GPS
Navigation Sensors
Global GNS-X Flight Management System
■ The GNS Operator’s Manual, Report No. 1280 with the applica-
ble or later revisions must be available to the flight crew whenever
navigation is predicated on the use of the GNS-X.
Refer to the applicable AFM supplement for pertinent program
and report number information. Program version must be veri-
fied on the system’s initialized page.
■ Whenever operation is predicated on the use of the optional AFIS,
the respective operator’s manual, report number xxxx with the
applicable or later revisions, must be available to the flight crew.
Refer to the applicable AFM supplement for pertinent program
and report number information.
■ The GNS-X is not approved as the sole means of navigation. Other
navigation equipment appropriate to the ground facilities along
the intended route must be installed and operable, as required by
the FARs applicable to the specific type of operation (i.e., VOR,
DME, etc).
■ The GNS-X position information must be checked for accuracy
(reasonableness) prior to use as a means of navigation. The GNS-
X position should be updated when a cross-check with other
approved navigation equipment reveals an error greater than 3
NM along-track or cross-track.
■ Navigation within the national aerospace system shall not be pred-
icated upon the GNS-X during periods of dead reckoning (DR).
■ Following a period of dead reckoning, position should be verified
by visually sighting ground reference points and/or using other
navigation equipment such as NDB, VOR, DME, or radar fix.
■ The GNS-X is not to be used as the primary reference in terminal
areas, during approaches to or departures from airports unless the
VPU sensor is operating and contributing to the position solution.
■ When latitude/longitude transferred from the internal database
(IDB) is displayed on the CDU, the pilot will ensure that it is a rea-
sonable position for the requested identifier.
■ The internal data base (IDB) must be updated to the latest revision
every 28 days; updating to be accomplished with the Global
Wulfsberg Systems update disk or equivalent. Update disks will
be received by mail (to subscribers) or obtained from authorized
Global Wulfsberg installation centers or update centers.

Citation V For SimuFlite training only 2-37

March 1999
 ■ The fuel management mode is for advisory purposes only and it
does not replace the airplane primary fuel flow and fuel quantity
systems.
■ When operating outside the magnetic variation model (north of
70° north latitude or south of 60° south latitude), the pilot must
manually insert magnetic variation.
Global GNS-X WITH LORAN-C
Navigation may not be predicated on LORAN-C when operating
outside the approved LOA boundaries as shown in the GNS-X
Operator’s Manual, or when the LORAN-C sensor is in DR.
Global GNS-X WITH VLF/OMEGA
■ Navigation may not be predicated on VLF/OMEGA when the
VLF/OMEGA sensor is in DR.
■ The GNS-X with only the VLF/OMEGA sensor operating is not
approved for operation into valleys (e.g., between peaks in moun-
tainous terrain).
Global GNS-X WITH GPS
■ The Global Positioning System (GPS) sensor is approved as a
contributing sensor for navigation if one of the following is true:
■ the GPS sensor is not the only sensor selected for use, it receives

four or more satellites, the GDOP is six or less, and the sensor
is not using altitude
■ the system is receiving three or more satellites and the GDOP,

using altitude, is 10 or less.

■ Use manually entered altitude only after failure of the automatic
inputs; update it every five minutes.

2-38 For SimuFlite training only Citation V

March 1999
 Quick Reference

Honeywell FMZ-600/800/900 Flight

Management System (FMS) with VLF/
OMEGA Sensor and Optional AFIS
■ The Honeywell FMS Pilot’s Manual (Report Number 5400-0154
with the following applicable or later revisions) must be imme-
diately available to the flight crew for aircraft equipped with the
Honeywell FMZ-600/800/900 FMS. The AFM supplement is
intended for use with the following FMS program versions:
■ FMS Program Version – NZ-8803, NZ-8804, NZ-9002

■ FMS Pilot’s Manual – F/4-25-88, G/1-27-89, H/1-8-90.

■ The FMZ, with either one VOR/DME or two DMEs available for
navigation, is approved for VFR/IFR RNAV operation within the
contiguous U.S. and Alaska according to the enroute, terminal,
and approach criteria of AC 90-45A and AC 20-130.
■ If the Honeywell OZ-800 VLF/OMEGA sensor is receiving usable
signals from at least two OMEGA navigation stations, the
Honeywell OZ-800 VLF/OMEGA sensor is capable of and meets
the requirements of:
■ VFR/IFR RNAV operation within the contiguous U.S. and

Alaska according to the enroute criteria of AC 20-101C

■ flight in the NAT-MNPS airspace according to AC 120-33 or AC

91-49 with dual independent system installations

■ operation as sole means of lateral navigation according to AC

120-37 in the areas between latitudes 70°N and 55°S with the
exceptions of that area above 45°N bounded by Longitudes
30°E and 120°E extended across the Asian continent.
■ The FMZ is not approved as the sole means of navigation.
Therefore, when the FMZ with the Honeywell OZ-800 VLF/
OMEGA sensor is to be used as the primary means of navigation,
or when coupled to the autopilot, flight director, or HSI, the nav-
igation equipment required by the FAR applicable to the specific
type of operation must be installed and operating.
■ Check the OZ-800 VLF-OMEGA position information for accu-
racy (reasonableness) prior to use as a means of navigation and
under the following conditions:
■ prior to compulsory reporting points during IFR operation when

not under radar surveillance or control

■ at or prior to arrival at each enroute waypoint during

VLF/OMEGA operation along approved RNAV routes

■ prior to requesting off airway routing, and at hourly intervals

thereafter during VLF operation off of approved RNAV routes.

Citation V For SimuFlite training only 2-39

March 1999
 ■ Navigation shall not be predicated on the use of the OZ-800
VLF/OMEGA sensor during periods of sensor DR (annunciated
as DR on VLF/OMEGA status page 2).
■ Following a period of DR (annunciated as NO POSITION SEN-
SORS), visually sight ground reference points and/or by use other
navigation equipment (e.g., RNAV, NDB, VOR/DME, radar fix)
to verify position fix.
■ The OZ-800 VLF/OMEGA sensor is not be used as the primary
reference during any operation in a terminal area or during depar-
tures from or approaches to airports or into valleys (e.g., between
peaks in mountainous terrain).
■ The FMZ is approved for IFR RNAV approach operation if:
■ the pilot shall confirm the tuned navaid used for VOR/DME

updating and/or display

■ both the EHSI and FMZ CDU annunciate the approach mode

(EHSI annunciates APP in blue and FMZ CDU annunciates

APRCH in white).
■ the FMZ is programmed with data from the current published

instrument approach procedures only

■ the FMZ CDU degrade annunciator is not illuminated

■ the last waypoint in the active flight plan must be the destina-

tion airport.
■ The use of manually inserted runway coordinates (place/bear-
ing/distance, place/bearing/place/bearing, or latitude and longi-
tude) is approved for VFR operations only.
■ The FMS data base information must be kept current by month-
ly updates obtained from Honeywell - Commercial Flight Systems
Division.
■ When transferring latitude/longitude and vertical waypoint alti-
tude from the data base, ensure that it is a reasonable position and
correct vertical waypoint altitude for the requested identifier.
■ When using previously stored flight plans and waypoints, verify
them for reasonable position prior to use.
■ During terminal area operation with the degrade annunciator illu-
minated on the FMZ CDU, verify the VOR/DME position.
■ Fuel display parameters are advisory only; they do not replace
the primary fuel quantity and flow indicators.
■ The pilot’s and copilot’s altimeters are the primary altitude refer-
ence during all vertical navigation (VNAV) operations.

2-40 For SimuFlite training only Citation V

March 1999
 Quick Reference

■ Navigation data from the FMZ during periods of DR as indicat-

ed by the illumination of the CDU DR light above the CRT are not
reliable; therefore, their use is prohibited for navigation.
■ FMS vertical navigation data is advisory only and is limited to
FMS CDU and EFIS. It cannot be coupled to the autopilot or
EFIS.
■ The pilot in command must exercise reasonable and prudent judg-
ment in the use of the advisory services of AFIS.

Honeywell Dual EDZ-605 Electronic Flight

Instrument System (EFIS) with
Multifunction Display (MFD)
■ The Honeywell SPZ-500C Integrated Flight Control System Pilot’s
Manual for the Cessna Citation II/V (Publication Number 28- NOTE: Crew qualification is
1146-63-00, dated November 1989 or later revision) must be required to conduct Category II
immediately available to the flight crew for aircraft with the approaches. The Honeywell
Honeywell Dual EDZ-605 EFIS. EDZ-605 EFIS is compatible
with the Sperry SPZ-500 Flight
■ Both flight director- and autopilot-coupled Category II approach- Guidance System.
es are approved using the Honeywell EDZ-605 EFIS displays.
■ Category II approaches are not approved in the composite mode
(REV selected).
■ Operating in the composite mode (REV selected) is approved only
with the flight director (single cue or crosspointer) selected.
■ EFIS ground operation with either the pilot’s or the copilot’s DISP
FAN annunciator light illuminated is limited to 10 minutes or until
either EADI HOT or EHSI HOT annunciator light illuminates,
whichever occurs first.
■ Dispatch is approved with any combination of two of the follow-
ing symbol generators (SG) operational: pilot’s SG, copilot’s SG,
or MFD SG. The hot annunciator light associated with the two
operational symbol generators must be extinguished.
■ Do not dispatch with any EADI HOT or EHSI HOT annunciator
light illuminated.
■ Do not dispatch with both pilot’s and copilot’s DISP FAN annun-
ciator lights illuminated.
■ If the DISP FAN Light Illuminated on Ground abnormal procedures
are followed, dispatch in visual meteorological conditions is
allowed with one EFIS FAN (pilots or copilot’s) annunciator light
illuminated.

Citation V For SimuFlite training only 2-41

March 1999
 ■ Do not dispatch following a flight where either an EADI HOT or
EHSI HOT annunciator light is illuminated, until the condition is
identified and corrected.
■ Both the pilot’s and the copilot’s EADIs and EHSIs must be
installed and operational in the normal (non-reversionary) mode
for takeoff and flight into instrument meteorological conditions
(IMC).
■ The EDZ-605 system must be operational per a satisfactory pre-
flight test as contained in the AFM normal procedures.
■ The aircraft must have an operational, independently powered,
attitude gyro.
■ Use the “T” speed display in the EADI for reference but do not use
it to replace the airspeed indicator as a primary instrument. The
copilot’s “T” speed references the pilot’s pitot/static system.
■ The radar checklist must contain FAA-approved operating pro-
cedures. It is the operator’s responsibility to ensure the checklist
contents are consistent with current AFM revisions.
■ On unit 0001 to 0074 with the Honeywell Dual EDZ-605 EFIS
with MFD, ground operation with or without engines operating is
limited when ambient temperatures are as shown below.
Below 45°C . . . . . . . . . . . . . . . . . . . . NO LIMIT
Between 45°C and 51°C . . . . . . . . . ONE HOUR LIMIT
Above 51°C . . . . . . . . NO OPERATIONS PERMITTED

Honeywell LSZ-850 Lightning Sensor

System
■ The Honeywell LSZ-850 Lightning Sensor System Pilot’s
Operating Manual (Publication Number 28-1146-54-00, dated
June 1988 or later revision) must be immediately available to the
flight crew for aircraft equipped with the Honeywell LSZ-850
Lightning Sensor System.
■ Thunderstorm avoidance must not be solely predicated upon the
use of the LSZ-850 Lightning Sensor System.

2-42 For SimuFlite training only Citation V

March 1999
 Quick Reference

Honeywell Primus II SRZ-850 Integrated

Radio System
■ The Honeywell Pilot’s Operating Handbook, SRZ-850 (Publication
Number 28-1146-50-01, dated June 1988 or later revision) must
be immediately available to the flight crew for aircraft equipped
with the Honeywell Primus II SRZ-850 Integrated Radio System.
■ The aircraft must be equipped with an independent operational
auxiliary COM 1/NAV 2 control display unit.

Honeywell Single EDZ-605 Electronic

Flight Instrument System (EFIS) with
Multifunction Display (MFD)
■ The Honeywell SPZ-500C Integrated Flight Control System Pilot’s
Manual for the Citation II/V (Publication Number 28-1146-63- NOTE: Crew qualification is
00, dated November 1989 or later revision) must be immediate- required to conduct Category II
ly available to the flight crew for aircraft equipped with the single approaches. The Honeywell
EDZ-605 EFIS. EDZ-605 EFIS is compatible
with the Honeywell SPZ-500
■ Both flight director and autopilot-coupled Category II approach- Flight Guidance System.
es are approved using the Honeywell EDZ-605 EFIS displays.
■ Category II approaches are not approved in the composite mode
(REV SELECTED).
■ Operating in the composite mode (REV SELECTED) is approved
only with the flight director selected.
■ Limit EFIS ground operation with the pilot’s DISP FAN annun-
ciator light illuminated to 10 minutes or until either EADI HOT
or EHSI annunciator light illuminates, whichever occurs first.
■ Do not dispatch with any EADI HOT or EHSI HOT annunciator
light illuminated.
■ Do not dispatch in instrument meteorological conditions with the
pilot’s DISP FAN annunciator light illuminated. Dispatch in visu-
al meteorological conditions is allowed with the DISP FAN annun-
ciator illuminated if the DISP FAN Illuminated on Ground abnor-
mal procedures are followed.
■ Do not dispatch following a flight where either an EADI HOT or
EHSI HOT annunciator light is illuminated, until the condition is
identified and corrected.

Citation V For SimuFlite training only 2-43

March 1999
 ■ The pilot’s EADI and EHSI must be installed and operational in
the normal (non-reversionary) mode for takeoff.
■ Verify the EDZ-605 system is operational with a satisfactory pre-
flight test per AFM normal procedures.
■ Use the “T” speed display in the EADI for reference, but do not
use it to replace the airspeed indicator as a primary instrument.
■ The radar checklist must contain FAA-approved operating pro-
cedures. The operator must ensure the checklist contents are con-
sistent with current AFM revisions.
■ On unit 0001 to 0074 with the Honeywell Single EDZ-605 EFIS,
ground operation with engines inoperative is limited to ambient
temperatures below ISA + 34°C as shown below.
Below 45°C . . . . . . . . . . . . . . . . . . . . NO LIMIT
Between 45°C and 51°C . . . . . . . . . ONE HOUR LIMIT
Above 51°C . . . . . . . . NO OPERATIONS PERMITTED

J.E.T. ADI-330 Attitude Indicator

■ Navigation information is limited to ILS localizer and glideslope
or BC localizer (NAV 1 or 2 only).
■ Unless installed in place of the AI-804 standby gyro, the ADI-330
attitude indicator does not meet the instrument panel location
requirements of FAR 121.305(j) for a third attitude instrument.

2-44 For SimuFlite training only Citation V

March 1999
 Quick Reference

Maximum Design Zero Fuel Weight of

12,200 LBS
■ If the aircraft’s maximum zero fuel weight is 12,200 lbs, the
following limits apply.
VMO Below 8,000 Ft . . . . . . . . . . . . . . . . 261 KIAS
MMO Between 8,000 Ft and 31,400 Ft . . . . . . . 276 KIAS
MMO Above 31,400 Ft . . . . . . . . . . . . . . . . 0.755 MI
■ The center-of-gravity envelope is as shown in Figure 2-7. CG
Envel., AFM, Supp 3, pg 4.
■ The airspeed indicator has the following markings.
Red Lines . . . . . . . . . . . . . . . . . 276 KIAS, 0.755 MI

UNS-1A Flight Management System with

CMA-764 VLF/OMEGA Sensor
■ The UNS-1A Operator’s Manual, as applicable to the software
version, must be available to the flight crew whenever navigation
is predicated on the use of the UNS-1A/CMA-764 VLF/OMEGA
sensor.

Software Version Operator’s Manual NOTE: The limitations at right

are specific, and individual limi-
302X UNS-1 Operator’s Manual (Report No. 2409SV302,
change 1, June, 1988 or appropriate later revision) tations apply to the various sen-
sor modes of the UNS-1A FMS.
304X UNS-1 Operator’s Manual (Report No. 2409SCV304, When all sensors are operating
reissue 0, March, 1990 or appropriate later revision) normally, composite navigation
information is output; in this
case, the limitations that apply
■ The UNS-1A, with only the VOR-DME mode operational, is
are a composite of the
approved for VFR/IFR RNAV operation and FMS approach mode
limitations shown.
operation within the contiguous U.S. and Alaska according to the
enroute criteria of AC90-45A.
■ IFR FMS approaches must be performed in the FMS approach
mode only, and according to published area navigation approach
procedures.

Citation V For SimuFlite training only 2-45

March 1999
 Center-of-Gravity Moment Envelope

2-7

2-46 For SimuFlite training only Citation V

March 1999
 Quick Reference

■ If the UNS-1A/CMA-764 VLF/OMEGA sensor receives usable

signals from at least two OMEGA navigation stations, the UNS-
1A/CMA-764 VLF/OMEGA sensor is capable of and meets the
requirements of:
■ VFR/IFR RNAV operation within the contiguous U.S. and

Alaska according to the enroute criteria of AC90-45A or the

criteria of AC20-101B.
■ flight in the NAT-MNPS airspace according to AC120-33 or

AC91-49.
■ The UNS-1A/CMA-764 VLF/OMEGA installation is not approved
as the sole means of navigation. Therefore, the navigation equip-
ment required by the FAR applicable to the specific type of oper-
ation must be installed and operating under either of the follow-
ing conditions:
■ when the CMA-764 VLF/OMEGA is used as a primary means

of navigation
■ when coupled to the autopilot, flight director, or HSI.

■ Check the UNS-1A/CMA-764 VLF/OMEGA position informa-

tion for accuracy (reasonableness) prior to use as a means of nav-
igation and under the following conditions:
■ prior to compulsory reporting points during IFR operation when

not under radar surveillance or control

■ at or prior to arrival at each enroute waypoint during VLF oper-

ation along approved RNAV routes

■ prior to requesting off airway routing and at hourly intervals

thereafter during VLF operation off approved RNAV routes.

■ Do not predicate navigation on the use of the UNS-1A/CMA-764
VLF/OMEGA during periods of DR.
■ Following a period of DR (annunciated as POS UNCERTAIN on
MSG page), visually sight ground reference points and/or by use
other navigation equipment (e.g., RNAV, NDB, VOR/DME, radar
fix) to verify position fix.

Citation V For SimuFlite training only 2-47

March 1999
 ■ Do not use the UNS-1A/CMA-764 VLF OMEGA as a primary
reference during any operation in a terminal area or during depar-
tures from or approaches to airports or into valleys (e.g., between
peaks in mountainous terrain).
■ Do not use the UNS-1A/CMA-764 VLF/OMEGA as a primary
reference during departures from airports.
■ Keep Jeppesen Data Base information current with monthly updates
distributed by Universal Navigation Corporation or Jeppesen
Incorporated.
■ When transferring latitude/longitude the Jeppesen Data Base,
ensure it is a reasonable position for the requested identifier.
■ The UNS-1A-displayed VNAV information is advisory only;
it is not valid below 18,000 ft flight altitude.
■ Without TAS input (B and D TAS), VNAV advisory information
is not available; manual TAS entry is required for IND computa-
tions.
■ Fuel display parameters are advisory only, and do not replace the
primary fuel quantity and flow indicators.

2-48 For SimuFlite training only Citation V

March 1999
 Quick Reference

Air Conditioning System Data System Data

Power Source Air cycle machine (ACM)
Ram air Summaries
Mixing valve (bypass modulating and shutoff valve)
Distribution Overhead duct
Under-floor duct
Control PRESS SOURCE selector
Auto temp rheostat
Manual heat/cool switch
DEFOG fan
OVHD fan
FLOW DISTR selector
Monitor Cabin temperature indicator
Annunciators
AIR DUCT O’HEAT
EMERG PRESS ON
BLD AIR GND

Citation V For SimuFlite training only 2-49

March 1999
 Electrical System Data
Power Source (one) 20 cell, 44 amp-hour battery
(two) Engine-driven starter/generators 28.5V, 300 amp
(two) AC inverters 26V and 115V, 400 Hz AC
Ground power – 28V, 600 to 1,000 amp
Distribution DC Buses
Hot Battery
Battery
Emergency
LH/RH Main
LH/RH Main Bus Extension
LH/RH Crossover
AC Buses
Number One/Two 26V AC
Number One/Two 115V AC
Control DC
Engine generator switches
Battery switch
Engine fire switches
Avionics DC master switch
AC
AC bus magnetic circuit breakers
AC power switch
AC test switch
Monitor DC
Voltmeters and ammeter
L/R GEN OFF annunciators
BATT O’TEMP annunciator
AC
AC FAIL annunciator
INVERTER FAIL 1/2 annunciators
Protection Circuit breakers
Current limiters
Relays
Generator control units

2-50 For SimuFlite training only Citation V

March 1999
 Quick Reference

Fire Protection Data

Power Source Left Main bus Extension
Right Crossover bus
Distribution Each bottle discharges to either engine, as
selected by the ENGINE FIRE PUSH switchlight
Control L/R ENGINE FIRE PUSH switchlights
BOTTLE 1/2 ARMED PUSH switchlights
Monitor L/R ENGINE FIRE PUSH switchlights
L/R F/W SHUTOFF annunciators
Protection Overtemperature/pressure relief valve on each fire
bottle

Flight Controls Systems Data

Flaps Data

Power Source LH Main bus Extension

Hydraulic system
Control Flap preselect handle
Monitor Flap indicator
HYD PRESS ON annunciator

Speedbrakes Data
Power Source LH Main bus Extension
Hydraulic system
Control Speedbrake switches/85% N2 throttle switches
Monitor SPEED BRAKE EXTEND annunciator
HYD PRESS ON annunciator

Citation V For SimuFlite training only 2-51

March 1999
 Fuel System Data
Power Source Fuel
L/R wing tanks
Pressure
L/R ejector pumps
L/R boost pumps
Distribution L/R engine-driven pumps to L/R engines
Control CROSSFEED switch
L/R FUEL BOOST pump switch
L/R ENG FIRE switchlights
L/R throttles
L/R ENGINE START switches
STARTER DISENGAGE switch
Monitor Annunciators
L/R FUEL BOOST ON
L/R FUEL LOW PRESS
L/R FUEL LOW LEVEL
L/R FUEL FLTR BYPASS
L/R FIREWALL SHUTOFF
Fuel quantity gage
Fuel flow gage
Crossfeed INTRANSIT light

2-52 For SimuFlite training only Citation V

March 1999
 Quick Reference

Hydraulic System Data

Power Source Hydraulic fluid in reservoir
Left and right hydraulic pumps as pressure sources
Distribution Left and right engine-driven hydraulic pumps
Control Left and right ENG FIRE switchlights
Landing gear handle
Speedbrake switch
Thrust reverser levers
Flaps lever
System CBs
Monitor LH/RH HYD FLOW LOW annunciators
Flap position indicator
SPEED BRAKE EXTEND annunciator
HYD LOW LEVEL annunciator
LH/RH ENG FIRE switchlights
HYD PRESS ON annunciator
Thrust reverser lights
LH/RH F/W SHUT OFF annunciator
Protection Reservoir pressure relief valve
Bypass relief valve
Firewall valve
Thermal relief

Citation V For SimuFlite training only 2-53

March 1999
 Ice and Rain Protection Data
Engine Anti-Ice Data

Power Source Engine bleed air

LH and RH Main bus Extension
Distribution Each engine
Bleed air to inboard wing leading edge
Bleed air to compressor stator vanes
Bleed air to nose cone, T1 probe
Engine igniters
Control LH/RH ENGINE ANTI-ICE switches
LH/RH ENG anti-ice CBs
Throttle switches: 60% N2
Monitor LH/RH ENG ANTI-ICE annunciators
WING O’HEAT annunciator
Visual scan – WING INSP light
Engine ITT/RPM
Engine ignition lights

Windshield Bleed Air Data

Power Source Bleed air from LH/RH bleed air clusters
Alcohol reservoir
LH Main bus Extension
Distribution LH/RH windshield nozzles
Control W/S BLEED switch
W/S ALCOHOL switch
Manual bleed air control valves
Rain augmenter door handles
Monitor W/S AIR O’HEAT annunciator
Bleed air noise

Surface Deice Data

Power Source Bleed air from LH/RH bleed air clusters
RH Crossover bus
Distribution Upper and lower wing boots
LH/RH horizontal stabilizer boots
Control SURFACE DEICE switch
Monitor SURFACE DEICE annunciator
Visual scan

2-54 For SimuFlite training only Citation V

March 1999
 Quick Reference

Ice and Rain Protection Data (cont.)

Pitot/Static Anti-Ice Data

Power Source LH Main bus Extension and

Emer DC bus (pilot side CB panel)
Distribution AOA heater
LH/RH pitot tube
LH/RH static ports
Control PITOT & STATIC ANTI-ICE switch
Monitor P/S HTR OFF annunciator
AOA HTR FAIL annunciator

Citation V For SimuFlite training only 2-55

March 1999
 Landing Gear System Data
Landing Gear Data

Power Source Hydraulic system

Emergency air bottle
LH Main bus Extension
Distribution Gear actuators
Control Gear control handle
Aux gear controls
T-handle
Blowdown knob
LDG GEAR position on test switch
Landing gear horn silence
Monitor Landing gear down lights
GEAR UNLOCKED light
Landing gear horn
Emergency air press gage (preflight)

Brakes/Anti-Skid Data
Power Source Independent hydraulic system
Emergency air bottle
LH Main bus Extension
Distribution Brake actuators
Control Brake pedals
ANTI-SKID switch
Parking brake handle
Emergency brake handle
ANTI-SKID position on test switch
Monitor POWER BRAKE LOW PRESS annunciator
ANTI-SKID INOP annunciator
Emergency air press gage (preflight)
Brake accumulator press gage (preflight)
Brake pedal feel

2-56 For SimuFlite training only Citation V

March 1999
 Quick Reference

Oxygen Systems Data

Power Source LH Main DC bus
Control Oxygen pressure regulator
Oxygen control valve
Passenger oxygen mask switch
Microphone oxygen mask selector switch
Oxygen selector on mask
Shutoff valve on oxygen cylinder
Altitude pressure switch/solenoid valve
Monitor Oxygen pressure gage
Overboard discharge indicator
Visual scan of passengers/masks
Crew oxygen flow indicator
CAB ALT 10,000 FT annunciator
Pressurization system annunciators/indicators

Citation V For SimuFlite training only 2-57

March 1999
 Pneumatic Systems Data
Bleed Air System Data

Power Source LH/RH engines

Distribution Air cycle machine (ACM)
LH engine to cabin (EMER position)
Wing deice system
Pressurization controller
Door seal
Engine anti-ice system
Windshield bleed air
Control Pressurization source select (bleed air selector) switch
OFF: Bleed air from either engine cluster assembly.
Pressure for windshield bleed air, cabin pressure
control, cabin door seal, wing deice, and engine
anti-ice available
GND: Bleed air from right engine cluster assembly
to ACM through ground valve (18 ppm)
LH: Bleed air from left engine cluster assembly to
ACM through left flow control valve (6 ppm)
NORM: Bleed air from both engines cluster
assembly to ACM through left and right flow
control valves (12 ppm)
RH: Bleed air from right engine cluster assembly
to ACM through right flow control valve (6 ppm)
EMER: Bleed air from left engine cluster assembly
direct to cabin through emergency pressurization
valve from left engine
Monitor Annunciators
W/S AIR O’HEAT
SURFACE DEICE
BLD AIR GND
EMERG PRESS ON
CABIN ALT 10,000 FT
AIR DUCT O’HEAT
ENG ANTI-ICE (LH/RH)
ACM O’PRESS

2-58 For SimuFlite training only Citation V

March 1999
 Quick Reference

Pneumatic Systems Data (cont.)

Pressurization System Data

Power Source Air cycle machine (ACM)

Emergency pressurization from left engine
Distribution Aircraft pressure vessel
Control Pressurization controller
CABIN RATE knob
PRESS SOURCE selector
EMER DUMP valve
Landing gear switch
Throttle switches
Pressure regulator
Control power (28V DC and vacuum)
Monitor CABIN ALT/DIFF PRESS indicator
Cabin rate of change indicator
Annunciators
CABIN ALT 10,000 FT
BLD AIR GND
EMERG PRESS ON
ACM O’PRESS

Citation V For SimuFlite training only 2-59

March 1999
 Thrust Reverser Systems Data
Power Source LH main bus Extension
RH Crossover bus
Hydraulic system
Control Thrust reverser levers
Emergency STOW SW
THRU REV position on rotary TEST switch
LH/RH ENGINE FIRE PUSH switchlights
LH/RH THRUST REVERSER CBs
Monitor Thrust reverser ARM/DEPLOY/UNLOCK indicators
HYD PRESS ON annunciator
MASTER WARNING annunciators

2-60 For SimuFlite training only Citation V

March 1999
This section presents four individual elements of flight opera-
tions: preflight inspection, expanded normal procedures, stan- Operating
dard operating procedures (SOPs), and maneuvers. Although
they are addressed individually in this manual, their smooth Procedures
integration is critical to ensuring safe, efficient operations.
The Preflight Inspection chapter illustrates a step-by-step
exterior inspection of the aircraft. Preflight cockpit and cabin
checks are also discussed.
The Expanded Normal Procedures chapter presents check-
lists for normal phases of flight. Each item, when appropriate,
is expanded to include limitations, cautions, warnings, and light
indications.
The Standard Operating Procedures chapter details Pilot
Flying/Pilot Not Flying callouts and verbal or physical responses.
The Maneuvers chapter pictorially illustrates common and
emergency profiles. Additionally, written descriptions are includ-
ed for most phases of flight with one or both engines operating.

Citation V For SimuFlite training only 3-1

March 1999
3-2 For SimuFlite training only Citation V
March 1999
An essential part of the preparations made before any flight is
the preflight inspection. During this inspection, verify the Preflight
aircraft’s physical readiness. A thorough preflight before the
first flight of the day allows the subsequent inspections to be Inspection
carried out in less time.
No detail should be overlooked during the first preflight of the
day. Abnormal conditions (e.g., low tire pressure) must be Chapter 3A
corrected prior to flight. Even minor discrepancies should be
rectified prior to flight to ensure safety.
The preflight inspection begins inside the aircraft where the
initial cockpit setup and essential functions are verified.
The actual exterior inspection follows; it begins at the left side
of the nose, proceeds clockwise around the aircraft, and ends at
the left wing root. Lastly, the pilot returns to the interior of the
aircraft to check the passenger compartment and cockpit for
flight readiness.

Citation V For SimuFlite training only 3A-1

March 1999
3A-2 For SimuFlite training only Citation V
March 1999
 Preflight Inspection

Cockpit Inspection . . . . . . . . . . . . . . . . . . 3A-5

Table of
Exterior Inspection . . . . . . . . . . . . . . . . . . 3A-7
Contents
General Exterior . . . . . . . . . . . . . . . . . . . . 3A-7
Exterior Inspection Schematic . . . . . . . . . . . . . 3A-8
Left Nose . . . . . . . . . . . . . . . . . . . . . . . 3A-9
Right Nose . . . . . . . . . . . . . . . . . . . . . . 3A-11
Right Wing . . . . . . . . . . . . . . . . . . . . . . 3A-15
Right Nacelle . . . . . . . . . . . . . . . . . . . . . 3A-21
Right Empennage . . . . . . . . . . . . . . . . . . 3A-23
Aft Compartment/Left Empennage . . . . . . . . . . 3A-27
Left Nacelle . . . . . . . . . . . . . . . . . . . . . 3A-31
Left Wing . . . . . . . . . . . . . . . . . . . . . . . 3A-33

Cabin Inspection . . . . . . . . . . . . . . . . . . 3A-39

Cockpit Inspection . . . . . . . . . . . . . . . . . 3A-41

Citation V For SimuFlite training only 3A-3

March 1999
3A-4 For SimuFlite training only Citation V
March 1999
 Preflight Inspection

Control Lock: Ensure that the control lock is unlocked. Control

surfaces need to be free for inspection during the exterior walk- Cockpit
around.
Landing Gear Lever: Check that the lever is down.
Inspection
Elevator Trim: Position the trim tab indicator within takeoff trim
range and ensure during the walkaround that the actual trim tab on
the elevator matches the indication.
Flap Handle: Check that the handle agrees with the flap position
indicator and ensure during the walkaround that the flaps are in the
same position.
Circuit Breakers: Ensure that all circuit breakers on the left and
right CB panels are in.
Generators: Generator switches should be in the OFF position for
an external power start or in GEN for a battery/cross generator start.
All Other Switches: Check that all switches on the instrument
panel are in OFF or NORM as required prior to engine start.
Throttles: Ensure that the throttles are off and cannot be moved
forward.
Battery Switch: Move the Battery Switch to BATT and check
for a minimum of 24V DC.
Parking Brake: Set the brakes by depressing the brake pedals and
pulling the parking brake handle out; this traps the applied pressure
to the wheel brakes. The battery switch must be on to charge the
brake accumulator if the accumulator pressure is below the dark
green arc.
Recognition Lights: Turn the Recognition lights switch to ON
and check illumination on the ground. Turn the switch to OFF. NOTE: For both the Recognition
lights and Landing/Exterior lights
Landing/Exterior Lights: Turn on the lights, check for illumi- checks, another crew member or
nation, and then turn the lights off. ground personnel observing these
Passenger Advisory Lights: Move the advisory light switch light indications outside the
to PASS SAFETY. Check the exit, emergency exit, and seat belt/no aircraft greatly enhances the accu-
racy and speed of the tests.
smoking lights for illumination. Turn the switch to OFF.
Pitot/Static Heat: Turn the pitot/static switch on for 30 seconds
to heat the pitot tubes, static ports, and angle-of-attack probe. The P/S
HTR OFF annunciator extinguishes if the power draw on the system NOTE: The landing and exterior
is correct. Turn the pitot/static switch to OFF. light check may be omitted if
night flight is not anticipated. To
Fuel Quantity: Note the total and individual fuel loads for prop- minimize battery drain if external
er flight planning requirements. power unit is not used, expedite
Battery Switch: Turn to OFF now that checks needing electrical the light and pitot heat check.
power have been completed.

Citation V For SimuFlite training only 3A-5

March 1999
3A-6 For SimuFlite training only Citation V
March 1999
 Preflight Inspection

Exterior
Unfold the preflight inspection diagram on the following page for ease
of reference. Note that each segment of the preflight inspection is Exterior Inspection
identified by letters A through H. Subsequent pages provide
sequenced checklists of each preflight inspection segment. Large
Inspection Walkaround
locator photos identify the general location of each inspection area.
Adjacent photos detail the checklist items. Photos read left to right.
Limitations and specifications are noted if relevant to the checklist.
Before starting the exterior inspection of the aircraft, obtain a flash-
light, screwdriver, and bucket or other suitable container for dis-
posal of fuel samples.

General Exterior
Make a general check for security, condition, and cleanliness of the
aircraft and components. Check particularly for damage, fluid leak-
age, security of access panels, and removal of keys from locks.
Remove all covers from the pitot tubes, static ports, probes, and
engine inlets and exhausts.

Citation V For SimuFlite training only 3A-7 3A-8

March 1999
Exterior Inspection Walkaround

1
C

H
D

2 3A 3B
G
F E

4 5

3A-8 For SimuFlite training only Citation V

March 1999
 Preflight Inspection

A Left Nose
1. Static Ports: Inspect the ports and ensure they are clear and Nose Baggage Limitation
warm. Because it is difficult to feel heat from the static port, run the
Maximum weight for nose
back of a finger from the aircraft skin to the port and back to the
baggage is 350 lbs.
aircraft skin to feel the difference in temperatures.
2. Baggage Door: Ensure that the door is firmly closed and
locked. The baggage door must be key-locked to actuate the door
locked microswitches and extinguish the DOOR NOT LOCKED
annunciator in the cockpit.
3. Nose Gear and Door: On the ground, the two forward gear
doors are closed while the rear door is open to allow a visual inspec-
tion of the nose gear assembly, shimmy damper, and nose gear steer-
ing bellcrank (3A). Check for an oleo strut extension (3B) of approx-
imately three inches for a fully fueled aircraft.
4. Nose Tire: Inspect the chine and tread of the nose tire. Chine
must be intact to meet the water/slush runway operating limitations.
Cord must not be visible on the tire. Check for proper inflation: 120
±5 PSI.
5. Left and Right Pitot Tubes: The tubes are cooling from the
cockpit inspection accomplished earlier. Do not, however, grasp a pitot
tube firmly since severe burns can result. Check the left tube, then
round the nose and check the right. Check that both are clear and
hot.
6. True Airspeed Probe: Check for general condition and
security.
7. Angle of Attack Sensors: Check that probe rotates and is
hot.

Citation V For SimuFlite training only 3A-9

March 1999
 1

2 3 4

5 6

3A-10 For SimuFlite training only Citation V

March 1999
 Preflight Inspection

B Right Nose
1. Baggage Compartment: Open the right baggage door. Raise Windshield Alcohol
the door until it catches and holds in the full open position. A bag- Limitation
gage light is available for night inspection.
Use TT-1-735 Isopropyl Alcohol
2. Windshield Alcohol Sight Gage: Fluid should be visible for windshield anti-ice.
with float ball at the top of the gage.
3. Gear/Brake Pneumatic Pressure Gage: Observe that the
gage is in the green arc. The pressure should read between 1800 and
2050 PSI.
4. Power Brake Accumulator Charge: Note that the gage
reads in the light green arc. The reading may be slightly higher if
the accumulator has been recently powered.
5. Brake Fluid Reservoir Sight Gages: The metal star in the
upper sight gage has a purple tint when the reservoir is full; the ball
should be at the top of the upper sight gage if the accumulator charge
is in the light green arc. See Table 3A-A.

Accumulator Accumulator Status

Discharged Charged

Fluid visible at top Fluid visible at bottom Normal minimum full

of upper gage of upper gage

Fluid visible in Fluid above top of Refill when practical

upper gage bottom gage

Fluid level not visible Fluid at or below top Refill before operation
in upper gage of bottom gage

Table 3A-A; Brake Fluid Reservoir Sight Gage

6. Baggage Door: Ensure that the door is firmly closed and

locked. The baggage door must be key-locked to actuate the door
locked microswitches and extinguish the DOOR NOT LOCKED
annunciator in the cockpit.

Citation V For SimuFlite training only 3A-11

March 1999
 7 8 9

10

3A-12 For SimuFlite training only Citation V

March 1999
 Preflight Inspection

B Right Nose (continued)

7. Overboard Vent Lines: Examine the vacuum, brake reser-
voir, alcohol bottle, and gear vent lines; ensure they are clear.
8. Static Ports: Inspect the ports and ensure they are clear and
warm. Because it is difficult to feel heat from the static port, run the
back of a finger from the aircraft skin to the port and back to the
aircraft skin to feel the difference in temperatures.
9. Angle-of-Attack Sensor: Carefully check the probe for heat-
ing as well as ease and range of rotation. The slots in the probe
should be unobstructed, or the optional vane is free to move.
10. Emergency Exit: Ensure that the door and handle are flush
with the fuselage skin.

Citation V For SimuFlite training only 3A-13

March 1999
 1

2 3 4

5 6 7

3A-14 For SimuFlite training only Citation V

March 1999
 Preflight Inspection

C Right Wing
1. Dorsal Fin Air Inlet: Observe that the inlet is clear.
2. Pylon Tailcone Air Inlet: Note that the pylon tailcone air
inlet is clear.
3. Engine Fan Duct/Fan/Forward T1 Sensor: Inspect the
duct to ensure it is clear. Check for bent blades, nicks, and foreign
objects. If the fan is windmilling, install a rear engine cover or place
your hand on the bullet nose to stop rotation. Observe the condition
of the T1 sensor.
4. Generator Cooling Air Inlet: Ensure the inlet is clear.
5. Wing Inspection Light: Examine the lens for cracks and
integrity.
6. Anti-Ice Bleed Air Cooling Air Inlet: Observe that the inlet
is clear.
7. Heated Leading Edge: Inspect the condition of the leading
edge; ensure vents are clear. Check the attachment of the stall strip.

Citation V For SimuFlite training only 3A-15

March 1999
 8 9 10

11 12 13

14

3A-16 For SimuFlite training only Citation V

March 1999
 Preflight Inspection

C Right Wing (continued)

8. Fuel Quick Drains and Fuel Filter Drain: When taking Fuel Limitation
fuel samples, push straight up on the fuel quick drains (Figure
See Quick Reference section for
3A-1). The drain locks open if the sampler is turned. Drain the fuel
approved fuels.
filter drain and check the fuel samples.
9. Main Gear Door/Wheel/Tire: Examine the door for securi-
ty. Check the wheel hubcap for condition and security of fastening.
Inspect the gear for general security and fluid leakage. The oleo strut
extension should be approximately one to two inches if the aircraft
is fully fueled. Check for tire wear (no cord may show) and inflation
pressure of 130 ±5 PSI.
10. Landing Light: Inspect the landing light for security.
11. Deice Boot: Examine the boot for cuts and cracks that might
prevent inflation. Inspect the surface for signs of delamination. Check
the attachment of the stall strip.
12. Fuel Filler Cap: Inspect the filler cap for security. Ensure
that the locking latch is closed and directed aft (Figure 3A-1).
13. Fuel Tank Vent: Ensure the vent is clear to preclude a neg-
ative pressure buildup that could cause the tank to collapse (Figure
3A-1).
14. Navigation/Strobe/Recognition Lights: Check the
lenses for cracks and integrity.

FUEL FILLER CAP

(TOP OF WING)

1 2
3
4
5
6

FUEL TANK VENT

FILTER DRAIN (BOTTOM OF WING)

3A-1

Citation V For SimuFlite training only 3A-17

March 1999
 15 16A 16B

16C

3A-18 For SimuFlite training only Citation V

March 1999
 Preflight Inspection

C Right Wing (continued)

15. Static Wicks: Look for and inspect the five static wicks: one
on the wing tip, two on the wing trailing edge outboard of the aileron,
and two on the trailing edge of the aileron. If an aileron static wick
is missing, replace it before flight to ensure control surface balance.
16. Aileron/Flap/Speedbrakes: Ensure that the aileron (16A)
has freedom of movement and the hinge points are secure. Check the
flaps (16B) and speedbrakes (16C) for security. Note that the flaps are
in the same position as the indicator in the cockpit.

Citation V For SimuFlite training only 3A-19

March 1999
 1

2 3 4

5 6

3A-20 For SimuFlite training only Citation V

March 1999
 Preflight Inspection

D Right Nacelle
1. Oil Level: Check for the correct level on the dipstick while the Oil Limitations
engine is still hot. If oil level is in doubt, run the engine for two min-
The following oils are approved
utes, shut down, and recheck. Accurate readings can be made up to
for use:
10 minutes after shutdown. Ensure that the filler cap is on securely;
■ Mobil Jet Oil II
fasten the access door.
■ Mobil Jet Oil 254
2. Generator Cooling Air Exhaust: Ensure that the exhaust ■ Exxon Turbo Oil 2380
opening is clear.
■ Castrol 5000
3. Engine Fluid Drain Mast: Examine the engine fluid drain ■ Aero Shell Turbine Oil 500/560
mast and ensure it is clear. Some residual fluid on the drain lines is
■ Royco Turbine Oil 500/560.
considered normal.
In addition, oils listed for the
4. Engine Exhaust/Bypass Ducts: Check for fuel leakage, engine in the latest revision to
damage to turbine blades, cracks, and general security. Pratt & Whitney Canada Inc. SB
5. Aft T1 Sensor: Observe the condition of sensor. 7001 are approved.
6. Thrust Reversers: Ensure the thrust reversers are stowed. Minimum starting oil temperature
is -40°C.
Check for cracks, damage, and general security.
Should it be necessary to replen-
ish oil when oil of the same brand
in the tank is unavailable, comply
with the following:
■ the total quantity of added oil

should not exceed two U.S.

quarts in any 400 hour period
■ if it is necessary to add more

than two U.S. quarts of dissimi-

lar oil brands, drain and flush
complete oil system. Refill with
an approved oil in accordance
with the Engine Maintenance
Manual instructions.
Should oils of nonapproved
brands or of different viscosities
become intermixed, drain and
flush complete oil system. Refill
with an approved oil in
accordance with the Engine
Maintenance Manual instructions.

Citation V For SimuFlite training only 3A-21

March 1999
 1

2 3 4

5 6 7

3A-22 For SimuFlite training only Citation V

March 1999
 Preflight Inspection

E Right Empennage
1. Deice Boot Overboard Vent: Ensure that the vent is clear.
2. Air Cycle Machine Overboard Exhaust: Examine the
exhaust on the lower side of the fuselage aft of the wing trailing
edge and ensure it is clear.
3. Hydraulic Service Door: Check that the drain is clear, the
panel access door is secured, and that no fluid is coming from the
drain.
4. Anti-Ice Bleed Air Cooling Air Exhaust: Ensure that the
exhaust port is clear.
5. Right Horizontal Stabilizer Deice Boot: Inspect the boot
for cuts and cracks that might prevent inflation. Examine the
surface for signs of delamination.
6. Right Elevator/Trim Tab: Examine the elevator for freedom
of movement and the hinge points for security. Ensure the trim tab
position matches the elevator trim position indicator in the cockpit.
7. Tail-mounted Rotating Beacon Light: Inspect the lens for
cracks and security.

Citation V For SimuFlite training only 3A-23

March 1999
 8 9 10

11 12 13

3A-24 For SimuFlite training only Citation V

March 1999
 Preflight Inspection

E Right Empennage (continued)

8. Tail Skid: Observe condition and ensure that it is secure.
9. Rudder/Trim Tab: Ensure that the rudder has freedom of
movement. Examine the hinge points for security. Check that the
trim tab moves in the opposite direction when the rudder is dis-
placed. Note that the trim tab position matches the rudder trim posi-
tion indicator in the cockpit. Check the rudder/aileron interconnect.
Ensure the ailerons move up on the side to which rudder is deflect-
ed and down on the opposite side.
10. Static Wicks: Observe that there are nine static wicks: two on
the trailing edge of each elevator (four total), two on the upper trail-
ing edge of the rudder; one above the rudder; one on the top of the
vertical fin, and one on the stinger. If a rudder or elevator static wick
is missing, replace it before flight to ensure proper control surface
balance.
11. Left Elevator/Trim Tab: Ensure that the elevator has free-
dom of movement. Check the hinge points for security. Note that the
trim tab position matches the elevator trim position indicator in the
cockpit.
12. Left Horizontal Stabilizer Deice Boot: Check the boot
for cuts and cracks that might prevent inflation. Inspect the surface
for signs of delamination.
13. Oxygen Blowout Disc: Ensure that the green disc is in
place. If it is missing, the oxygen bottle may be empty.

Citation V For SimuFlite training only 3A-25

March 1999
 1

2A 2B 3

4 5A 5B

3A-26 For SimuFlite training only Citation V

March 1999
 Preflight Inspection

F Aft Compartment/Left Empennage

1. Hydraulic Fluid Quantity: Check the reservoir sight gage
Hydraulic Limitations
for a reading above the REFILL mark.
Approved fluids include:
2. Fire Bottle Pressure Gages: Note bottle pressure (2A) is ■ Skydrol 500A, B, B-4, C or
correct for the ambient temperature according to the pressure/tem- LD-4
perature chart (2B) between the bottles. A visual inspection of ■ Hyjet, Hyjet W, Hyjet III, or IV
bottle pressure is the only way to determine if the bottles are full.
3. J-Box Circuit Breakers: Ensure that all the visible circuit
breakers are set.
4. Tailcone Baggage Compartment: Inspect the baggage and
engine covers and ensure they are secure. Check that the inner tail-
cone access door is closed and latched firmly.
5. Aft Compartment Light/Access Door: Observe that the aft
compartment light is off (5A). Close the access door and ensure that
the door latches are firmly closed (5B). The aft compartment access
door must be key-locked to actuate the door locked microswitches and
extinguish the DOOR NOT LOCKED annunciator.

Citation V For SimuFlite training only 3A-27

March 1999
 6 7 8

3A-28 For SimuFlite training only Citation V

March 1999
 Preflight Inspection

F Aft Compartment/Left Empennage

(continued)
6. External Power Service Door: Inspect the door for security.
7. Battery Cooling Intake/Vent Lines: Check that the intake
and vent lines are clear.
8. Windshield Heat Exchanger Overboard Exhaust: Ensure
that the exhaust opening is clear.
9. Anti-Ice Bleed Air Cooling Air Exhaust: Examine the
exhaust port and ensure that it is clear.

Citation V For SimuFlite training only 3A-29

March 1999
 1

2 3 4

5 6

3A-30 For SimuFlite training only Citation V

March 1999
 Preflight Inspection

G Left Nacelle
1. Thrust Reversers: Ensure that the thrust reversers are stowed. Oil Limitations
Check for cracks, damage, and general security.
The following oils are approved
2. Aft T1 Sensor: Observe the condition of the sensor. for use:
3. Engine Exhaust/Bypass Ducts: Check for fuel leakage, ■ Mobil Jet Oil II

damage to turbine blades, cracks, and general security. ■ Mobil Jet Oil 254

4. Engine Fluid Drain Mast: Examine the engine fluid drain ■ Exxon Turbo Oil 2380

mast and ensure it is clear. Some residual fluid on the drain lines is ■ Castrol 5000

considered normal. ■ Aero Shell Turbine Oil 500/560

5. Generator Cooling Air Exhaust: Ensure that the exhaust ■ Royco Turbine Oil 500/560.

opening is clear. In addition, oils listed for the

6. Oil Level: Check for the correct level on the dipstick while the engine in the latest revision to
Pratt & Whitney Canada Inc. SB
engine is still hot. If oil level is in doubt, run the engine for two min-
7001 are approved.
utes, shut down, and recheck. Accurate readings can be made up to
10 minutes after shutdown. Ensure that the filler cap is on securely; Minimum starting oil temperature
is -40°C.
fasten the access door.
Should it be necessary to replen-
ish oil when oil of the same brand
in the tank is unavailable, comply
with the following:
■ the total quantity of added oil

should not exceed two U.S.

quarts in any 400 hour period
■ if it is necessary to add more

than two U.S. quarts of dissimi-

lar oil brands, drain and flush
complete oil system. Refill with
an approved oil in accordance
with the Engine Maintenance
Manual instructions.
Should oils of nonapproved
brands or of different viscosities
become intermixed, drain and
flush complete oil system. Refill
with an approved oil in
accordance with the Engine
Maintenance Manual instructions.

Citation V For SimuFlite training only 3A-31

March 1999
 1A

1B 1C 2

3 4 5

3A-32 For SimuFlite training only Citation V

March 1999
 Preflight Inspection

H Left Wing
1. Flap/Speedbrakes/Aileron/Trim Tab: Check the flaps (1A)
and speedbrakes (1B) for security. Note that the flaps are in the same
position as the indicator in the cockpit. Ensure that the aileron (1C)
has freedom of movement and that the hinge points are secure. Note
that the trim tab position matches the aileron trim position in the
cockpit.
2. Static Wicks: Look for and inspect the five static wicks: one
on the wing tip, two on the wing trailing edge outboard of the aileron,
and two on the trailing edge of the aileron. If an aileron static wick
is missing, replace it before flight to ensure control surface balance.
3. Navigation/Strobe/Recognition Lights: Check the lens-
es for cracks and integrity.
4. Fuel Tank Vent: Ensure the vent is clear to preclude a negative
pressure buildup that could cause the tank to collapse (Figure 3A-2)
5. Fuel Filler Cap: Inspect the filler cap for security. Ensure that
the locking latch is closed and directed aft (Figure 3A-2).

FUEL TANK VENT

(BOTTOM OF WING)

FUEL FILLER CAP

(TOP OF WING)

3A-2

Citation V For SimuFlite training only 3A-33

March 1999
 6 7 8

9 10 11

3A-34 For SimuFlite training only Citation V

March 1999
 Preflight Inspection

H Left Wing (continued)

6. Deice Boot: Examine the boot for cuts and cracks that might Fuel Limitations
prevent inflation. Inspect the surface for signs of delamination. Check
See Quick Reference for approved
the attachment of the stall strip.
fuels.
7. Main Gear Door/Wheel/Tire: Examine the door for securi-
ty. Check the wheel hubcap for condition and security of fastening.
Inspect the gear for general security and fluid leakage. The oleo strut
extension should be approximately one to two inches if the aircraft
is fully fueled. Check for tire wear (no cord may show) and inflation
pressure of 130 ±5 PSI.
8. Fuel Filter Drain and Fuel Quick Drains: Drain the fuel fil-
ter drain and check the fuel samples. When taking fuel samples,
push straight up on the fuel quick drains (Figure 3A-3). The drain
locks open if the sampler is turned.
9. Heated Leading Edge: Inspect the condition of the leading
edge; ensure vents are clear. Check the attachment of the stall strip.
10. Anti-Ice Bleed Air Cooling Air Inlet: Ensure the inlet is
clear.
11. Wing Inspection Light: Examine the lens for cracks and
integrity.

2 3 4
1
5
6

FILTER DRAIN

3A-3

Citation V For SimuFlite training only 3A-35

March 1999
 12 13 14

15

3A-36 For SimuFlite training only Citation V

March 1999
 Preflight Inspection

H Left Wing (continued)

12. Generator Cooling Air Inlet: Ensure the inlet is clear.
13. Engine Fan Duct/Fan/Forward T1 Sensor: Inspect the
duct to ensure it is clear. Check for bent blades, nicks, and foreign
objects. If the fan is windmilling, install a rear engine cover or place
your hand on the bullet nose to stop rotation. Observe the condition
of the T1 sensor.
14. Dorsal Fin Air Inlet: Observe that the inlet is clear.
15. Secondary Cabin Door Seal: Check for rips, tears, and
folding.

Citation V For SimuFlite training only 3A-37

March 1999
 1 2 3

4 5 6

3A-38 For SimuFlite training only Citation V

March 1999
 Preflight Inspection

1. Emergency Exit: Check the fit of the door and that the han-
dle is stowed, the guard is in place, and the locking pin is removed. Cabin
2. Passenger Seats: Ensure that the passenger seats are fully
upright and outboard, and positioned aft or forward to prevent block-
Inspection
ing exit doors.
3. Door Entry Lights: Check that the switch on the entry door
post is turned off.
4. Luminescent Exit Placard: Inspect for security.
5. Portable Fire Extinguisher: Note that the cabin fire extin-
guisher is serviced and secured.
6. Documents: Check that airworthiness and registration
certificates and radio station license are on board and displayed.

Citation V For SimuFlite training only 3A-39

March 1999
 2 3 5

3A-40 For SimuFlite training only Citation V

March 1999
 Preflight Inspection

1. FAA Approved Flight Manual: Ensure that this is on board

and available to the pilots. Cockpit
2. Microphones, Headsets, and Oxygen Masks: Check that
these are on board and plugged in. Test the oxygen masks and set the
Inspection
masks to 100%.
3. Oxygen Control Valve: Ensure that it is in the NORMAL
position.
4. Flashlight: Check that it is on board and operational.
5. Portable Fire Extinguisher: Check that the portable fire
extinguisher in the quick-release holder under the copilot seat is ser-
viced and secured. The pressure gage should read in the white arc;
this indicates a 150 PSI charge.
6. Emergency First Aid Kit: Ensure that it is on board the air-
craft and that it is fully equipped.

Citation V For SimuFlite training only 3A-41

March 1999
3A-42 For SimuFlite training only Citation V
March 1999
This section outlines and expands normal operating procedures
and includes applicable cautions and warnings. Also present- Expanded
ed are cold weather operations as well as parking, mooring,
and short-term aircraft storage requirements. Normal
Procedures

Chapter 3B

Citation V For SimuFlite training only 3B-1

March 1999
3B-2 For SimuFlite training only Citation V
March 1999
 Expanded Normal Procedures

Normal Procedures . . . . . . . . . . . . . . . . . 3B-5

Table of
Before Starting Engines . . . . . . . . . . . . . . . . 3B-5
Start . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-8
Contents
Before Taxi . . . . . . . . . . . . . . . . . . . . . . 3B-10
Taxi . . . . . . . . . . . . . . . . . . . . . . . . . . 3B-14
Before Takeoff . . . . . . . . . . . . . . . . . . . . 3B-15
Takeoff . . . . . . . . . . . . . . . . . . . . . . . . 3B-15
Climb . . . . . . . . . . . . . . . . . . . . . . . . . 3B-16
Transition Level . . . . . . . . . . . . . . . . . . . 3B-17
Cruise . . . . . . . . . . . . . . . . . . . . . . . . 3B-17
Descent . . . . . . . . . . . . . . . . . . . . . . . 3B-18
Transition Level . . . . . . . . . . . . . . . . . . . 3B-19
Approach . . . . . . . . . . . . . . . . . . . . . . . 3B-19
Before Landing . . . . . . . . . . . . . . . . . . . . 3B-20
Landing With Thrust Reversers . . . . . . . . . . . 3B-22
After Landing . . . . . . . . . . . . . . . . . . . . . 3B-23
Shutdown . . . . . . . . . . . . . . . . . . . . . . 3B-24
Parking . . . . . . . . . . . . . . . . . . . . . . . . 3B-24
Mooring . . . . . . . . . . . . . . . . . . . . . . . 3B-26
Towing/Taxiing . . . . . . . . . . . . . . . . . . . . 3B-29
Nose Gear Towing . . . . . . . . . . . . . . . . . 3B-29
Main Gear Towing . . . . . . . . . . . . . . . . . 3B-30
Taxiing (Ground Movement) . . . . . . . . . . . . 3B-31

Storage/Restoring . . . . . . . . . . . . . . . . . 3B-33
Storage . . . . . . . . . . . . . . . . . . . . . . . . 3B-33
0-7 Days . . . . . . . . . . . . . . . . . . . . . . 3B-33
7-30 Days . . . . . . . . . . . . . . . . . . . . . 3B-33
Restoring . . . . . . . . . . . . . . . . . . . . . . 3B-34
From 0-30 Days . . . . . . . . . . . . . . . . . . 3B-34

Citation V For SimuFlite training only 3B-3

March 1999
 Cold Weather Operations . . . . . . . . . . . . . 3B-35
Preflight . . . . . . . . . . . . . . . . . . . . . . . 3B-35
EPU Start . . . . . . . . . . . . . . . . . . . . . . 3B-36
Engine Start . . . . . . . . . . . . . . . . . . . . . 3B-36
After Engine Start . . . . . . . . . . . . . . . . . . 3B-36
Taxi/Before Takeoff . . . . . . . . . . . . . . . . . . 3B-37
Takeoff . . . . . . . . . . . . . . . . . . . . . . . . 3B-37
Rejected Takeoff on Slick Runway . . . . . . . . . . 3B-37
In Flight . . . . . . . . . . . . . . . . . . . . . . . 3B-38
Taxi-in and Park . . . . . . . . . . . . . . . . . . . 3B-38
Securing Overnight or for Extended Period
(Aircraft Unattended) . . . . . . . . . . . . . . . . 3B-39
Ground Cooling . . . . . . . . . . . . . . . . . . . 3B-39
Performance Computations, Takeoff . . . . . . . . . 3B-39

3B-4 For SimuFlite training only Citation V

March 1999
 Expanded Normal Procedures

Before Starting Engines Normal

Preflight Inspection . . . . . . . . . . . . . . COMPLETED
Cabin Door . . . . . . . . . . . . . . . . CLOSED/LOCKED
Procedures
■ Check green indicators for proper door pin position.
■ Ensure handle is vertical and in the detent position.
Passenger Briefing . . . . . . . . . . . . . . . COMPLETED
Seat and Rudder Pedals . . . . . . . . . . . . . . ADJUSTED
■ Adjust seat fore and aft; the handle is below the forward center of
the seat.
■ Adjust seat vertically; the handle is on the aisle side forward corner.
■ Adjust seat tilt; the handle is at lower rear, aisle side of the seat.
■ Adjust rudder pedals; depress the tab on inboard side, move the ped-
als fore or aft into one of the three positions, and then release the tab.
Seat Belts, Shoulder Harnesses . . . . . . . . . . SECURED
Adjust the seat belt for a snug fit. Ensure the shoulder harnesses are
latched to the buckle.
Parking Brake . . . . . . . . . . . . . . . . . . . . . . . SET
Depress the brake pedals and pull the parking brake handle to trap
pressure. If the brake accumulator charge did not indicate in the light
or dark green arc on the external walkaround check, the accumulator
must be charged. Turn the Battery switch to BATT shortly before set-
ting the brakes to charge the accumulator.
Control Lock . . . . . . . . . . . . . . . . . . . . . . . . OFF
Rotate the handle clockwise 45° from horizontal and push in to release;
check that controls and throttles are free. Damage to the throttle con-
trol may occur if the throttles are forced past the lock position and the
control lock is engaged.
Crew Oxygen Masks and Pressure . . . . . . . . CHECKED
■ Oxygen pressure gage should read 1,600 to 1,800 PSI.
■ Oxygen mask connection should be secure.
■ OXYGEN CONTROL VALVE should be in NORMAL.
■ Regulator should be set to 100%.
■ Ensure flow by donning the mask, adjusting the fit, and breathing sev-
eral times.
■ Crew masks must be stowed in the quick donning hook.
Circuit Breakers . . . . . . . . . . . . . . . . . . CHECKED
Check that all circuit breakers on the left and right CB panels and the
right sub CB panel are in.

Citation V For SimuFlite training only 3B-5

March 1999
 LH/RH Gyro Slave . . . . . . . . . . . . . . . . . . . AUTO
Check that the left and right gyro slave switches are in AUTO.
All Switches . . . . . . . . . . . . . . . . OFF OR NORMAL
Check that all instrument panel switches are either positioned to off or
normal. In preparation for engine start:
■ Generators switches OFF or GEN, depending on type of start.

■ Radio master switch and inverter switch off to protect against volt-
age variances.
■ Pressurization bleed air may remain in normal.

■ Crossfeed, windshield bleed air valves, and engine synchronization

should be positioned to OFF.
■ Boost pumps and ignition should be in NORM.

Ground Idle Switch . . . . . . . . . . . . . . . . . . . NORM

Landing Gear . . . . . . . . . . . . . . . . . . . . . . DOWN
Check that the landing gear handle is down.
Generators . . . . . . . . . . GEN (OFF FOR EPU START)
■ Generator switches GEN for battery and cross generator starts.
■ Generator switches OFF for an EPU start.
Standby Gyro . . . . . . . . . . . . . . . . . . . TESTED/ON
Standby Gyro check should be accomplished with the battery switch
off to ensure the gyro is receiving emergency power.
Standby Gyro Switch . . . . . . . . . . . . . . . . . . TEST
Place switch in TEST only momentarily. Check that the green light
illuminates.
Standby Gyro Switch . . . . . . . . . . . . . . . . . . . . ON
Check that the STDBY GYRO amber light illuminates.
Standby Gyro Caging Knob . . . . . . . . . . . . UNCAGED
Gyro Horizon . . . . . . . . . . . . . . . . . . . . NO FLAG
Battery Switch . . . . . . . . . . . . . . . . . . . . . . EMER
Check power to emergency bus items. Ensure that at least one avion-
ics item powered from the Emergency DC bus receives power (NAV2,
COM1, Copilot HSI, Dual EFIS Copilots RMI) etc. This step is required
on aircraft 0032 and subsequent.
Battery Switch . . . . . . . . . . BATT/VOLTS CHECKED
Battery voltage for start is 24V DC minimum.
External Power (if applicable) . . . . . . . . . CONNECTED
Once EPU is connected and on, the voltmeter should read 28V DC.

3B-6 For SimuFlite training only Citation V

March 1999
 Expanded Normal Procedures

Warning Test . . . . . . . . . . . . . . . . . . . . CHECKED

■ OFF – The red light above the rotary test switch is off and test sys-
tem inoperative.
■ FIRE WARN – Both engine fire lights on the upper center instrument
panel illuminate.
■ LDG GEAR – Three green (gear safe) lights and one red (UNLOCKED)
light illuminate; the gear horn sounds. Check that the horn can be
silenced by pressing the horn silence button on the landing gear panel.
The horn can be silenced only if flap position is less than 15°. With
the flap handle in the second detent (15°), the horn does not silence.
■ BATT TEMP – BATT O’TEMP annunciator flashes; battery tem-
perature gage shows 160°F to demonstrate circuit integrity. MASTER
WARNING light also illuminates.
■ STICK SHAKER – On the probe-type stick shaker the angle-of-
attack indicator drives to zero and the flag appears. The flag disap-
pears and the indicator moves to 1.0. At approximately 0.82, the
stick shaker activates for a few seconds. This cycle is repeated as long
as the test switch remains in this position.
On the vane-type stick shaker the stick shaker operates. The AOA
meter needle goes past the red area and the EADI fast/slow needles
go past slow. Red indexer lights flash on and off, if installed.
■ T/REV – Six thrust reverser lights illuminate. The MASTER WARN-
ING lights also illuminate.
■ W/S TEMP – The W/S AIR O’HEAT annunciator illuminates when
LOW or HIGH is selected on the windshield bleed air switch.
■ OVERSPEED – The audible overspeed warning sounds.
■ ANTI-SKID – The anti-skid system initiates a self-test: ANTI-SKID
INOP annunciator illuminates and remains on three to four seconds
after the test switch is moved to OFF; the annunciator extinguishes
if the system checks operational. If the system fails the check, the
annunciator remains illuminated.
■ ANNU – All annunciators and the MASTER WARNING lights
illuminate. The turbine speed indicator self-tests. For example, the red
light illuminates and the digital display flashes 8s. When the avionic
power switches are on, the altitude alert horn sounds, and the altitude
alert and autopilot/flight director mode selector panel lights illuminate.
EFIS and FMS lights also illuminate. The MASTER WARNING lights
cannot be reset while the rotary selector switch is in this position.
DOOR NOT LOCKED Annunciator . . . . . . . . . . OUT
Engine Instruments . . . . . . . . . . . . . . . . CHECKED
Check that no OFF flags are visible at the top of each engine instrument.
Fuel Quantity . . . . . . . . . . . . . . . . . . . . CHECKED
Check that proper fuel quantity is indicated on the fuel gages. Maximum
fuel imbalance between tanks is 200 lbs.
Throttles . . . . . . . . . . . . . . . . . . . . . . . . . . OFF

Citation V For SimuFlite training only 3B-7

March 1999
 Start
Center Panel Lights . . . . . . . . FULL BRIGHT (NIGHT)
Rotating Beacon . . . . . . . . . . . . . . . . . . . . . . . ON
Flood Center Panel Lights . . . . . . . . . . FULL BRIGHT
First Engine . . . . . . . . . . . . . . . . . . . . . STARTED
If passengers are aboard and use of ground bleed air is not anticipat-
ed, start the left engine first. Spool up is slightly faster since there is
less line loss with the battery mounted on the left side of the tailcone
compartment.
START Button . . . . . . . . . . . . . . . . . . . PRESSED
Momentarily pressing the START button begins engine rotation by
closing the start relay. The START button illuminates white when the
start relay is engaged. The ignition system then arms for actuation.
The engine instrument floodlights and the associated FUEL BOOST
ON annunciator illuminate. The FUEL LOW PRESS annunciator
extinguishes as boost pump pressure increases.
Throttle . . . . . . . . . . . . . . . . IDLE AT 8 TO 10% N2
Lift the cutoff latch and advance the throttle to IDLE. Fuel flow ini-
tiates and the ignition system activates. The associated ignition light
illuminates.
ITT . . . . . . . . . . . . . . . . . . . . . . . . . CHECKED
Abort start if there is no ITT indication within 10 seconds or the ITT
exceeds 550°C. Maximum ITT for start is 700°C for two seconds.
N1 Speed . . . . . . . . . . . . . . . . . . . . . . CHECKED
Check for an N1 indication between 20 and 25% N2. Abort start if
no N1 indication by 25% N2.
Engine Instruments . . . . . . . . . . CHECKED NORMAL
Monitor the engine instruments during acceleration. Abort start for
abnormal indication.
Annunciators . . . . . . . . . . . . . . . . . . . . CHECKED
The engine start cycle terminates at approximately 38% N2. The
START button light, ignition light, instrument floodlights, and FUEL
BOOST ON annunciator extinguish. The HYD FLOW LOW annun-
ciator extinguishes as hydraulic flow increases during start. The
GEN OFF annunciator extinguishes after the start is terminated, at
approximately 38% N2, if the GEN switch is in GEN.
Should automatic start sequencing not terminate, the FUEL BOOST
ON annunciator, ignition and associated lights remain illuminated.
The starter, however, discontinues cranking because speed sensing,
which governs at approximately 38% N1, does not register. Depress
the STARTER DISENGAGE button to terminate the automatic start
sequence.

3B-8 For SimuFlite training only Citation V

March 1999
 Expanded Normal Procedures

Ground Idle Switch . . . HIGH (Check 52% Turbine RPM)

■ After engine has reached ground idle RPM of approximately 46%
N2, place the ground idle switch in HIGH. The N2 RPM should
increase to 52% (flight idle). Check that it does so.
■ The registering of 52% N2 indicates the ground idle system is work-
ing and that proper RPM, which ensures correct torque on the oper-
ating generator drive, is available for a cross generator start.
■ For a cross generator start – wait until the 52% turbine RPM is
achieved and the generator is on line. Start the second engine. Both
starter buttons illuminate during a cross generator start. A cross gen-
erator start reduces battery heat by eliminated a charging cycle.
■ For an external power start – both generator switches should be off
until start is complete. Do not turn on any electrical equipment until
both GEN OFF lights are extinguished. The Citation V has an over-
current and overvoltage protection system for EPU usage.
Second Engine . . . . . . . . . . . . . . . . . . . . STARTED
For a cross generator start – set no higher than 53% turbine RPM on CAUTION: Turbine speed
the operating engine. greater than 53% on the operat-
ing engine produces a generator
Ground Idle Switch . . . NORM (Check 46% Turbine RPM) output that may damage the gen-
Both engines should idle at 46% turbine RPM. erator drive during the second
If operating in visible moisture and ambient air temperature engine start.
between +10°C and -30°C:
Ground Idle Switch . . . . . . . . . . . . . . . . . . . HIGH
Pitot/Static Heat . . . . . . . . . . . . . . . . . . . . . . ON
Limit Pitot Heat to a maximum of two minutes at a time or damage
to the angle-of-attack vane heater may occur.
LH/RH Engine Anti-Ice Systems . . . . . . . . . . . . . ON
W/S BLEED Air Switch . . . . . . . . . . . . . . LOW or HI
■ If temperature is below -18°C, use the HI position. Ensure that
the windshield bleed air valves are in MAX.
■ For sustained ground operation, the engines should operate for
one out of every four minutes at 65% or above turbine RPM.
Engine Instruments . . . . . . . . . . . . . . . . CHECKED
■ Check that all engine instruments are within normal range.
■ Check that engine annunciators are off (except AC FAIL and BLEED
AIR GND, if selected).
External Power (if applicable) . . . . . . DISCONNECTED
Verify that the power unit is off by confirming a 24V DC battery read-
ing on the voltmeter.

Citation V For SimuFlite training only 3B-9

March 1999
 Generator Switches
(if external power used for start) . . . . . . . . . . . . . GEN
L/R generator annunciators are off, and the ammeters show shared
load. The generators may not come on until the external power is
removed.

Before Taxi
Volt/Ammeters . . . . . . . . . . . . . . . . . . . CHECKED
Voltmeter indicates 28.5V DC and ammeters indicate a shared load.
Right Generator . . . . . . . . . . . . . . . . . . . . . . OFF
Left generator powers the main DC buses; voltmeter shows 28.5V
DC under load.
Voltage Selector . . . . . . . . . . . . . . . . . RIGHT GEN
Shows 28.5V DC without load.
Right Generator . . . . . . . . . . . . . . . . . . . . . . ON
Generators again share the load.
Left Generator . . . . . . . . . . . . . . . . . . . . . . . OFF
Right generator powers the main DC buses; voltmeter shows 28.5V
DC under load.
Voltage Selector . . . . . . . . . . . . . . . . . . LEFT GEN
Shows 28.5V DC without load.
Left Generator . . . . . . . . . . . . . . . . . . . . . . . ON
Check for a shared load on ammeters.
Inverters . . . . . . . . . . . . . . . . . . . . . . . . . . . AC
Avionics Master Switch . . . . . . . . . . . . . . . . . . . ON
The right avionics power switch serves as the master avionics power
switch and supplies DC power to avionics equipment.
External Lights . . . . . . . . . . . . . . . . AS REQUIRED
Passenger Advisory Lights . . . . . . . . . . PASS SAFETY
This position advises passengers to fasten seat belts and to stop smok-
ing for takeoff; it also illuminates cabin exit lights and the baggage
area light.
AntiSkid . . . . . . . . . . . . . . . . . . . . . . . . . . . ON
If the antiskid is turned off prior to or during taxiing, it must be turned
on prior to takeoff. The antiskid self-test sequence must be complet-
ed when the aircraft is stationary. Failure to do so renders the anti-skid
inoperative during takeoff.
Radar . . . . . . . . . . . . . . . . . . . . . . . . STANDBY

3B-10 For SimuFlite training only Citation V

March 1999
 Expanded Normal Procedures

Pressurization/Temperature . . . . . . . . . . . . . . . . SET
■ Normally, cruise altitude plus 1,000 ft is selected on the dial labeled
ACFT. Cabin altitude is displayed on the inner scale labeled CABIN.
Set cabin rate control selector knob pointer in the white arc.
■ Select the temperature control to automatic and at approximately
the 12 o’clock position.
Cabin Fan (if applicable) . . . . . . . . . . . . . . HI OR LO
Anti-Ice/Deice Systems . . . . . CHECKED/AS REQUIRED
CAUTION: Limit ground oper-
Windshield Bleed Air ation of pitot/static heat to two
W/S BLEED Air Switch . . . . . . . . . . . . . . . . . LOW minutes to preclude damage to
W/S BLEED Air Valves . . . . . . . . . . . . . . . . . MAX the angle-of-attack system.
Check for bleed air noise.
Engine Anti-Ice
Turbine (N2) . . . . . . . . . . . . . . SET 65% OR ABOVE CAUTION: Do not operate
L/R Engine Anti-Ice Switches . . . . . . . . . . . . . . . ON deice boots when ambient air
temperature is below -40°C
ITT . . . . . . . . . . . . . . . . . . . . CHECK INCREASE (-40°F).
RPM . . . . . . . . . . . . . . . . . . CHECK DECREASE
Ignition Lights . . . . . . . . . . . . . . . . . . . . . . . ON
Anti-Ice Fail Lights . . . . . . . . . . . . . . . . . . . . OFF
NOTE: When operating in visi-
Should extinguish in two minutes or less with 70% N2 set.
ble moisture and ambient air tem-
L/R Engine Anti-Ice Switches . . . . . . . . . . . . . . OFF perature is between a minimum
Deice System (Boots) of +10°C and -30°C, position
ground idle switch to HIGH; turn
Surface Deice Switch . . . . . . . . . . . . . . . . . . AUTO pitot and static heat and engine
Set power at 60% N2 or above. Below 60% N2, the SURFACE LH and RH anti-ice systems ON.
DEICE annunciator may not illuminate. Momentarily place the If temperature is above -18°C,
SURFACE DEICE switch in AUTO and observe the wing boot turn W/S BLEED air switch to
inflation cycles. The lower wing boots should inflate for six LOW. If temperature is -18°C or
seconds (SURFACE DEICE annunciator illuminated) followed by below, turn W/S BLEED air
a rest period of six seconds. Then the upper wing boots should inflate switch to HI. Check that
for six seconds (SURFACE DEICE annunciator illuminated). The windshield bleed air valves are in
cycle should not repeat. MAX. For sustained ground
Ground Idle Switch . . . . . . . NORM (HIGH as required) operation, the engines should be
operated for one out of every four
Flight Controls . . . . . . . . . . . . . . . . . . . CHECKED minutes at 65% N2 or above.
Check for freedom of movement and correct displacement. Observe the two-minute
limitation on Pitot Heat.

Citation V For SimuFlite training only 3B-11

March 1999
 Flaps . . . . . . . . . . . . . . . . . . CHECKED AND SET
Extend flaps to LAND and check indicator movement. Verify flap trim
interconnect operation between 15° and 25°. Retract flaps to T.O. &
APPR (15°) or T.O. (7°) as appropriate.
If flap trim interconnect does not operate, determine cause and expect
higher than normal elevator forces during flap extension and retraction.
Speedbrakes . . . . . . . . . . . . . . . . . . . . CHECKED
Speedbrakes . . . . . . . . . . . . . . . . . . . . . EXTEND
Check that the HYD PRESS ON annunciator illuminates until speed-
brake are extended; the annunciator then extinguishes. The SPD
BRAKE EXTEND annunciator also illuminates. Observe speed-
brakes on top of the wing.
Speedbrakes . . . . . . . . . . . . . . . . . . . . RETRACT
Check that the HYD PRESS ON annunciator illuminates and then
extinguishes. The SPD BRAKE EXTEND annunciator extinguish-
es. Visually check that the speedbrakes are retracted.
Engine Sync . . . . . . . . . . . . . . . . . . . . . . . . . OFF
Inverter Test . . . . . . . . . . . . . . . . . . COMPLETED
This tests each inverter’s ability to power the complete AC system.
AC TEST Switch . . . . . . . . . . . . . . . . . . . . INV 1
The No. 1 inverter is turned off and the No. 1 INVERTER FAIL
annunciator illuminates. The AC system continues to operate on
power from the No. 2 inverter.
AC TEST Switch . . . . . . . . . . . . . . . . . . . . INV 2
The No. 2 inverter is turned off and the No. 2 INVERTER FAIL
annunciator illuminates. The AC system continues to operate on
power from the No. 1 inverter.
Autopilot . . . . . . . . . . . . . . . . . . . . . . CHECKED
Autopilot . . . . . . . . . . . . . . . . . . . . . . ENGAGE
TEST EACH FLT Button . . . . . PRESS/HOLD FOR 5 SEC
The AP TORQUE and AP ROLL MONITOR annunciators illumi-
nate, and then the autopilot disengages; the AUTOPILOT OFF
annunciator illuminates, and the autopilot warning horn sounds for
one second.

3B-12 For SimuFlite training only Citation V

March 1999
 Expanded Normal Procedures

Autopilot . . . . . . . . . . . . . . . . . . . . . . ENGAGE
The elevator and ailerons should be in the neutral position for this
check.
■ Rotate pitch wheel up and down – control column must move in
direction of pitch wheel movement.
■ Rotate turn knob left and right – control wheel must move in the
direction of the turn knob movement.
■ Engage heading mode on the Flight Director – move heading
cursor to the left or right of the lubber line; control wheel should
follow.
■ Engage ALT mode on the Flight Director – the control column
moves aft as a lower altitude is selected on the pilot’s altimeter and
forward as a higher altitude is selected
■ Disengage ALT mode and reset altimeter.

■ Pull aft on the control wheel and ensure the elevator trim wheel,
after a short delay, starts trimming nose down. The opposite should
occur when pushing forward.
Autopilot . . . . . . . . . . . . . . . . . . . . DISENGAGE
Check both pilot’s and copilot’s AP/TRIM DISC switches and the
Go-Around made; autopilot should disengage and AUTOPILOT
OFF annunciator illuminates for one second. The autopilot must be
re-engaged between the tests.
Trim . . . . . . . . . . . . . . . . . . . CHECKED AND SET
Copilot Electric Trim Switch:
Trim . . . . . . . . . . . . . . . . . . . OPERATE NOSE-UP
Manual Trim Wheel rotates nose up.
AP/TRIM DISC Switch . . . . . . . . . . . . . . . . . PUSH
Check that trim stops.
Trim . . . . . . . . . . . . . . . . OPERATE NOSE-DOWN
Manual Trim Wheel rotates nose down.
AP/TRIM DISC Switch . . . . . . . . . . . . . . . . . PUSH
Check that trim stops.
Pilot Electric Trim: repeat previous check.
Rudder and Aileron Trim . . . . . . . . . SET AT NEUTRAL
Elevator Trim . . . . . . . ENSURE IN TAKEOFF RANGE

Citation V For SimuFlite training only 3B-13

March 1999
 Avionics . . . . . . . . . . . . . . . . . CHECKED AND SET
EFIS Test Button . . . . . . . . . . . . . . . . . . . . PUSH
Verify the following:
■ Radio altimeter test value on pilot display is 50 ft

■ All digit readouts replaced with dashes (except radio altimeter)

■ All flags in view

■ Command cue, if selected, bias from view

■ (optional dual EFIS) the comparator monitor annunciators

illuminate ATT, HDG, and ILS if ILS sources are selected on
both sides
■ Test pass light in upper left corner of EADI illuminates.

Navigation and Communication Radios . . . . . . . . . SET

Transponder . . . . . . . . . . . . . . . . . . . . . . . . SET
Altimeters/Altitude Alerter . . . . . . . . . . . . . . . . SET
Altitude alerter should be set to initial departure altitude.
TOLD Card/Bugs . . . . . . . . . COMPLETED AND SET
Pilot’s airspeed bug should be set to V2; copilot’s airspeed bug to V1.

Taxi
Both pilots should maintain the maximum possible taxiway/ramp
vigilance. At least one pilot should maintain taxi lookout vigilance
CAUTION: If, during taxiing, a while checks are being accomplished.
hard brake pedal/no braking con-
Brakes . . . . . . . . . . . . . . . . . . . . . . . . CHECKED
dition is encountered, operated
the emergency brake system. Thrust Reverser . . . . . . . . . . . . . . . . . . CHECKED
Maintenance is required before T/R Levers . . . . . . . . . . . . . . . . . . IDLE REVERSE
flight. Check that the ARM, UNLOCK, and DEPLOY lights are on.
Emergency Stow Switches . . . . . . . . . . . . . . . STOW
The UNLOCK and DEPLOY lights should extinguish.
CAUTION: Do not attempt to T/R Levers . . . . . . . . . . . . . . . . . . . . . FORWARD
fly aircraft if the thrust reverser The ARM light remains on.
preflight check is unsuccessful. Emergency Stow Switches . . . . . . . . . . . . . NORMAL
The ARM light should extinguish.

3B-14 For SimuFlite training only Citation V

March 1999
 Expanded Normal Procedures

Flight Instruments . . . . . . . . . . . . . . . . . . . CHECK

■ Check that no warning flags are visible.
■ Check that the heading instruments are in agreement (HSIs, RMIs,
and magnetic compass).
■ Check that the VSIs are at zero.
■ Check that the pilot and copilot altimeters are within normal para-
meters
■ Check the flight instruments for correct indications during turns.
■ Ensure that the GYRO SLAVE switches are in AUTO.
Crew Briefing . . . . . . . . . . . . . . . . . . COMPLETED
Refer to the Standard Operating Procedures (SOPs) chapter for detailed
explanation of items on the takeoff briefing.

Before Takeoff
Pressurization Source Select . . . . . . . . . . . . NORMAL
If the source select is left in GND, excessive air extraction occurs on
the right engine, and it will not develop full takeoff thrust.
Temperature Control Selector . . . AUTOMATIC (mid-range)
The ACM overtemperature protection circuit operates only in the
AUTOMATIC mode.
Anti-Skid (when aircraft stopped) . . . . . . . . . . . . . ON
Radar . . . . . . . . . . . . . . . . . . . . ON OR STANDBY
Transponder . . . . . . . . . . . . . . . . . . . . . . . . . ON
Ensure the code is correct and mode C is on.

Takeoff (When Cleared for Takeoff)

Exterior Lights . . . . . . . . . . . . . . . . . . . . . . . ON
For flights 30 minutes before sunset to 30 minutes after sunrise, nav-
igation lights should be on. Do not operate anti-collision lights in con-
ditions of fog, clouds, or haze; reflection of the light beam can cause
disorientation or vertigo.
Pitot Heat . . . . . . . . . . . . . . . . . . . . . . . . . . ON
Limit ground operation of pitot/static heat to two minutes or damage
to the angle-of-attack system may occur.
Anti-Ice System . . . . . . . . . . . . . . . . AS REQUIRED
Ignition . . . . . . . . . . . . . . . . . . . . . . . . . . . . ON
Turning ignition systems on may prevent flameout should an engine
ingestion problem arise during takeoff.

Citation V For SimuFlite training only 3B-15

March 1999
 Landing Lights . . . . . . . . . . . . . . . . . . . . . . . ON
Annunciator Panel/Flight Director . . . . . . . . CHECKED
■ All annunciators should be extinguished except possibly GROUND
IDLE and ENG ICE FAIL if that system is selected with low power.
■ The flight director should be in GO AROUND with Heading and
Altitude Select functions selected.
Engine Instruments . . . . . . . . . . . . . . . . CHECKED
Check that all indications are normal or in the green.

Climb
Landing Gear . . . . . . . . . . . . . . . . . . . . . . . . UP
When a positive rate-of-climb is indicated, pull the gear handle out
and move it to UP to initiate the retraction cycle. This illuminates the
GEAR UNLOCKED and HYD PRESS ON annunciators. Check that
both lights extinguish to indicate the gear is up and locked.
Landing Lights . . . . . . . . . . . . . . . . . . . . . . . OFF
Flaps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . UP
At a comfortable altitude with the wings level and a minimum air-
speed of V2 +10 KIAS, push the flap handle down to clear the detent
and then move to full forward. Check that the position indicator left of
the handle moves to the FLAP UP position.
Yaw Damper . . . . . . . . . . . . . . . . . . . . ENGAGED
Check that YAW DAMPER ENGAGE light is on. Yaw damper
improves aircraft control and passenger comfort.
Climb Power . . . . . . . . . . . . . . . . . . . . . . . . SET
Use indicated temperature and the Climb Thrust Setting chart to deter-
mine climb N1.
Engine Sync . . . . . . . . . . . . . . . . . . . . . . . . . ON
When N1 is set (within 1.5%) or N2 (within 1%), turn the engine syn-
chronize selector switch to FAN or TURB position. Check that the
engine instruments remain within limits. N1 RPM increases with alti-
tude; throttle adjustments may be necessary to maintain specified thrust
setting.
Pressurization . . . . . . . . . . . . . . . . CHECKED/SET
The controller was programmed prior to taxi. Adjust the rate knob to
achieve a comfortable cabin rate-of-climb (usually 300 to 500 fpm).
Observe differential pressure/cabin altitude and cabin vertical speed
gages. A thorough understanding of the DIFF PRESS/CABIN ALT
gage interpretation aids the crew in smooth operation of the pressur-
ization system.

3B-16 For SimuFlite training only Citation V

March 1999
 Expanded Normal Procedures

Ignition . . . . . . . . . . . . . . . . . . . . . . . . NORMAL
When clear of any bird hazard, return IGNITION switches to NORM
when cockpit workload permits.
Anti-Ice Systems . . . . . . . . . . . . . . . AS REQUIRED
Select anti-ice systems as required for climb. Use of engine anti-ice
reduces allowable fan speed and dictates close monitoring of ITT and
RPM limits.
Passenger Advisory Lights . . . . . . . . . . AS REQUIRED
Place the switch to SEAT BELT to keep the cabin advisory light on
and extinguish the NO SMOKING and emergency exit lights. If no
turbulence is anticipated, place the switch to OFF to extinguish advisory
and emergency exit lights.

Transition Level
Altimeters . . . . . . . . . . . . . . . . . . . . . . . . . . SET
Set to 29.92 and crosscheck.
Recognition Lights . . . . . . . . . . . . . . . . . . . . . OFF
Crew Oxygen Masks . . . . . . . . . . . . . . POSITIONED
Above FL250, ensure the mask is in the quick donning position and set
at 100%. Above FL350 with one pilot in the cockpit, that pilot must be
wearing an oxygen mask. Above FL410, at least one pilot must wear
an oxygen mask.
Temperature Select . . . . . . . . . . AUTO (above 31,000 ft)
AUTO above 31,000 ft reduces the possibility of an ACM (air cycle
machine) overheat and normally maintains a comfortable cabin tem-
perature. With low airspeed and high power settings, an air cycle
machine overheat is possible with an excessively cold setting in MAN-
UAL. An approximate indication of airflow warmth into the cabin can
be made by placing a hand over an open crew foot warmer outlet since
the foot warmer is an extension of the same source as the cabin under-
floor ducting.

Cruise
Cruise Power . . . . . . . . . . . . . . . . . . . . . . . . SET
Maintain climb thrust until acceleration to the desired cruise speed is
attained. If engine RPM does not automatically synchronize at desired
cruise setting, turn the engine synchronizer switch to OFF; this allows
the synchronizer actuator to center. Roughly synchronize the engines
with the throttles and turn switch to FAN or TURB.
Engine Instruments . . . . . . . . . . . . . . . . CHECKED

Citation V For SimuFlite training only 3B-17

March 1999
 Fuel Quantity . . . . . . . . . . . . . . . . . . . . CHECKED
Ensure proper consumption rate and balance fuel as required to main-
tain the 200 lb limitation.
Pressurization . . . . . . . . . . . . . . . . . . . CHECKED
Reset cabin altitude and/or rate as required. Maintain the AUTO TEMP
SELECT knob in the 12 to 2 o’clock position for a comfortable cabin
temperature.
Anti-Ice . . . . . . . . . . . . . . . . . . . . AS REQUIRED
CAUTION: Do not operate Check anti-ice system for proper operation prior to entering areas in
deice boots when indicated OAT which icing might be encountered. The engine bleed air anti-ice must
is below -40°C (-40°F). be activated when operating in visible moisture at temperatures of
+10°C to -30°C indicated OAT and any time icing is occurring. The pitot
and static anti-ice is normally operated during flight.

Descent
Defog Fan . . . . . . . . . . . HI (15 minutes before descent)
The DEFOG fan should be turned on and foot warmers closed approx-
imately 15 minutes before descent to reduce condensation on the wind-
shield and cockpit side windows.
Pilot/Copilot Footwarmers . . . . . . . . . . . . . . CLOSED
Closing footwarmers increases the flow of air available for windshield
defogging and also isolates dry conditioned air between the cockpit
side windows to inhibit condensation formation.
Airflow Distribution . . . . . . . . . . . . . . . . . . . . SET
The airflow distribution selector should be biased toward the cockpit
for maximum defog capabilities.
Windshield Bleed Air Switch/Valves . . . . . . . . LO/MAX
Windshield bleed air can be used to externally warm the windshield in
extreme conditions. Normally, the LO position of the windshield bleed
air switch provides adequate temperature.
Pressurization . . . . . . . . . . . . . . . . CHECKED/SET
After beginning descent, set destination field pressure altitude + 200
ft on the CABIN dial of the controller. Monitor the differential pres-
sure/cabin altitude and cabin vertical speed gages. A high cabin
altitude and low differential pressure indicates insufficient rate of
descent; depressurization occurs when cabin and aircraft altitude are
identical. High cabin descent rates may be uncomfortable and may
result in programmed cabin altitude reached well before landing.
Spreading the cabin descent requirement over the majority of the let-
down time realizes optimum comfort.

3B-18 For SimuFlite training only Citation V

March 1999
 Expanded Normal Procedures

Anti-Ice Systems . . . . . . . . . . . . . . . AS REQUIRED

A minimum of 70% N2 or higher is required to keep the engine anti-
ice system operating properly. Pitot/static and engine anti-ice should
be on and operating and windshield bleed air as required when oper-
ating in visible moisture with an OAT of -30°C to +10°C indicated.

Transition Level
Recognition Lights . . . . . . . . . . . . . . . . . . . . . ON
Altimeters . . . . . . . . . . . . . . . . . . . . . . . . . . SET
When cleared below or passing through the transition altitude, set the
reported or landing field barometric pressure on both altimeters.
Crosscheck the settings.

Approach
TOLD Card/Bugs . . . . . . . . . . . . . COMPLETED/SET
■ Complete the APPROACH side of TOLD card.
■ Airspeed bugs should be set to VREF.
Seat Belts/Shoulder Harnesses . . . . . . . . . . . SECURED
Check that the seats are locked in the desired position. Check that the
seat belts and shoulder harnesses are latched to the buckle.
Aft Facing Seat . . . . . . . . UPRIGHT, OUTBOARD, AFT
Ensure there is an unobstructed access to the emergency exit door.
Flight Instruments . . . . . . . . . . . . . . . . . CHECKED
Check vertical gyros; no warning flags should be visible. The compass
headings, altimeters, and airspeed indicators should be in agreement.
For Category II approaches:
Autopilot
TEST EACH FLT Button . . PRESS/HOLD FOR 5 SECONDS
■ The AP TORQUE and AP ROLL MONITOR annunciators illu-
minate followed by the disengagement of the autopilot.
■ The AUTOPILOT OFF annunciator illuminates and autopilot
warning horn sounds.
■ If AP TORQUE or AP ROLL MONITOR annunciators do not
illuminate, high torque is not available or the roll monitor is inop-
erative; an autopilot coupled Category II approach is not approved.
Avionics . . . . . . . . . . . . . . . . . . . . . . . . . . . SET
Tune navigation equipment and identify: courses should be set and
the flight director programmed as required.

Citation V For SimuFlite training only 3B-19

March 1999
 Fuel Crossfeed . . . . . . . . . . . . . . . . . . . . . . . OFF
Check CROSSFEED knob is OFF and INTRANSIT and FUEL
BOOST ON lights extinguished.
Passenger Advisory Lights . . . . . . . . . . PASS SAFETY
Engine Sync . . . . . . . . . . . . . . . . . . . . . . . . . OFF
Crew Briefing . . . . . . . . . . . . . . . . . . COMPLETED
Check the Standard Operating Procedures (SOPs) chapter for a list of
items that should be included in the approach briefing.

Before Landing
Flaps . . . . . . . . . . . . . . . . . . . . . . . . T.O & APPR
Flaps may be extended to T.O. & APPR below 200 KIAS. Check indi-
cator to verify position.
Ignition . . . . . . . . . . . . . . . . . . . . . . . . . . . . ON
This may prevent flameout should an engine problem arise during the
approach and landing phase.
Ground Idle Switch . . . . . . . NORM (HIGH as required)
Use HIGH if ground icing is anticipated or for touch and go landings.
If the GROUND IDLE annunciator illuminates in flight, select HIGH
(flight idle) and ensure that N2 RPM does not decrease below 52%.
Engine acceleration from below 52% N2 RPM may be excessive for
some flight conditions. If conditions are such that icing may be expect-
ed on the ground, select HIGH on the ground idle switch.
Landing Gear . . . . . . . . . . . . . DOWN AND LOCKED
Pull gear handle out and move to DOWN; HYD PRESS ON and GEAR
UNLOCKED annunciators illuminate while the gear is extending.
Three green lights should illuminate and the GEAR UNLOCKED and
HYD PRESS ON annunciators should extinguish.
Landing Lights . . . . . . . . . . . . . . . . . . . . . . . ON
Annunciator Panel . . . . . . . . . . . . . . . . . CHECKED
Flaps . . . . . . . . . . . . . . . . . . . . . . . . . . . LAND
Flaps should be in the LAND position for all normal landings. Flaps may
be extended to LAND below 173 KIAS. Handle must be pushed down
to clear the T.O. & APPR detent. Verify position with indicator.

3B-20 For SimuFlite training only Citation V

March 1999
 Expanded Normal Procedures

Icing Conditions . . . . . . . . . . . . . . . . . . CHECKED

When any residual ice is present or can be expected during approach WARNING: With any residual
and landing, VREF and VAPP must be increased. VREF and VAPP, the ice present, do not attempt to fly
landing distance, and the maximum landing weight permitted by brake uncorrected VREF/VAPP speeds.
energy must be corrected per chart below. Stall speeds increase and stall
warning margins decrease.
Residual Ice – Landing Correction Factors

Maximum Landing
Weight Limited By CAUTION: In icing conditions,
Landing Climb or Brake Energy Landing VAPP VREF a small amount of residual ice
Field Anti-Ice Anti-Ice Distance
forms on unprotected areas. This
Conditions On Off
is normal, but can cause an
Field increase in stall speeds. When
Elevation any amount of residual ice is vis-
4000 ft or ible, the stall speeds increase by
below; No No No Multiply Add 8 Add 8
8 knots; the VREF/VAPP speeds
downhill Correction Correction by 1.30 knots knots
gradient; and landing distances and the
No tailwind maximum landing weight per-
mitted by brake energy must be
Field corrected per chart below.
Elevation
Above No Multiply Multiply Add 8 Add 8
4000 ft to Correction by 0.88 by 1.30 knots knots
10,000 ft
NOTE: When reconfiguring for
All Other No Multiply Multiply Add 8 Add 8 approach and landing (i.e., flaps
Conditions Correction by 0.73 by 1.30 knots knots extended and gear down) with
any ice accretion visible on the
wing leading edge, regardless of
thickness, activate the surface
Pressurization . . . . . . . . . . . . ZERO DIFFERENTIAL deice system. Continue to monitor
Approximately 500 ft above ground level, check that the cabin differential the wing leading edge for any
pressure is near zero. If it is in excess of 0.5 PSI, select a higher cabin reaccumulation.
altitude and adjust RATE so the cabin ascends. Differential pressure
should be at zero for landing; at touchdown, any existing pressure is
dumped. NOTE: For increased rates of
If landing above 12,000 ft pressure altitude, turn the OXYGEN CON- descent icing conditions, Cessna
TROL VALVE to CREW ONLY and bleed air to OFF to preclude recommends use of landing flaps.
passenger mask deployment. This allows a higher power
Autopilot and Yaw Damper . . . . . . . . . . . . . . . . OFF setting, if necessary, to maintain
anti-icing capabilities.
Depress the AP/TRIM DISC switch on either control wheel. With the
yaw damper off, the pilot has complete rudder authority for landing.
Speedbrakes . . . . . . RETRACTED PRIOR TO 50 FEET NOTE: Do not allow turbine
Extended speedbrakes are not approved for landing. (N2) RPM to be less than 52%.

Citation V For SimuFlite training only 3B-21

March 1999
 Landing With Thrust Reversers
Throttles . . . . . . . . . . . . . . . . . . . . . . . . . . IDLE
Suggested crosswind technique involves flying a crab down final
approach and aligning the longitudinal axis of the aircraft to runway
centerline with rudder just before touchdown. The wide expanse of
cockpit visibility makes small crab angles difficult to detect and par-
ticular attention should be devoted to this area to achieve smooth cross-
wind landings.
Eight seconds after touchdown, the engines spool down from flight
idle (approximately 52% N2) to ground idle (46% N2) if the flight idle
switch is in the NORM position. The GROUND IDLE annunciator
illuminates.
Brakes (after touchdown) . . . . . . . . . . . . . . APPLIED
CAUTION: If, during landing, To obtain maximum braking performance from the antiskid system,
a hard brake pedal/no braking apply continuous maximum effort (no modulation) to brake pedals.
condition is encountered, oper-
ate the emergency brake system. Speedbrakes (after touchdown) . . . . . . . . . EXTENDED
Maintenance is required before Touchdown, preceded by a slight flare, should occur on the main
next flight. wheels. Check thrust at idle and extend speedbrakes while lowering the
nose wheel.
Thrust Reversers
(after nose wheel on ground) . . . . . . . . . . DEPLOYED
CAUTION: Do not advance Apply wheel brakes and deploy the thrust reversers. The aircraft pitch-
throttles until the thrust reversers es slightly upward during deployment, therefore, use slight nosedown
unlock lights are extinguished. elevator pressure during thrust reverser deployment especially at high
On aircraft unit 0001 through speeds such as a refused takeoff or no flap landing.
0118 not incorporating SB 560- The nose wheel must be on the ground before actuation of the thrust
78-02, the automatic power reversers to reduce the possibility of pitch-up and lift-off and improve
retard system activates causing directional control. Do not exceed approximately 15 lbs of force on
misrigging of the throttle link- the thrust reverser levers during deployment because the throttle lock-
age systems. This results in only out cams could jam.
partial takeoff power or possibly
Reverser Indicator Lights . . . . CHECK ILLUMINATION
a flameout if the throttle is placed
in idle. The ARM, UNLOCK, and DEPLOY lights should illuminate.
On aircraft unit 0119 and on Reverser Power . . . . . . . . . . . . . . . . AS REQUIRED
0001 through 0118 incorporat- Do not exceed 79% N1 when OAT is below -18°C or 86% N1 at or
ing SB 560-78-02, there is a dan- above -18°C. Once the thrust reversers are deployed, move the levers
ger of the throttle being rapidly aft to a maximum reverser thrust. “Stops” on the levers provide 86%
returned to idle position, which N1 on a -18°C day at sea level so the pilot can keep his attention on the
could cause injury. landing rollout. The factory setting results in lower than 86% N1 at
warmer temperatures and may be reset for higher N1 if temperatures
are predominantly warmer. Do not exceed 86% N1.

3B-22 For SimuFlite training only Citation V

March 1999
 Expanded Normal Procedures

At 60 KIAS:
Thrust Reverser Levers . . . . . . . . . . . IDLE REVERSE
With the thrust reverser levers in the IDLE REVERSE detent, the
reversers are left deployed for aerodynamic drag. Thrust reversing and
braking should commence according to runway length. With excess
runway, braking normally is begun after thrust reverser deceleration is
below 60 knots.
Do not use the thrust reversers for touch and go landings; a full stop
landing must be made once the reversers are selected.
Use caution on runways with small loose grit because it may be ingest-
ed in the engine at idle reverse at low taxi speed.

After Landing
This checklist should be accomplished after the aircraft is clear of
the runway.
Thrust Reversers . . . . . . . . . . . . . . . . . . . STOWED
Flaps . . . . . . . . . . . . . . . . . . . . . . . . . . . . . UP
Check that the HYD PRESS ON annunciator extinguishes after the
flaps are up. Taxiing with flaps extended on a snow or slush covered
taxiway could result in obstruction of the flaps.
Speedbrakes . . . . . . . . . . . . . . . . . . . RETRACTED
Check that the HYD PRESS ON and SPD BRAKE EXTEND annun-
ciators are extinguished.
Ignition . . . . . . . . . . . . . . . . . . . . . . . . . . NORM
Ground Idle Switch . . . . . . . . . . . . . . AS REQUIRED
Pitot/Static Heat/Anti-Ice . . . . . . . . . . . . . . . . . OFF
W/S BLEED AIR may be used as required in falling precipitation.
Turn engine anti-ice ON and operate the engines at or above 65% N2
one out of every four minutes during taxi in visible moisture with a
temperature of +10°C to -30°C. Ensure the pitot/static switch is off.
Exterior Lights . . . . . . . . . . . . . . . . AS REQUIRED
Recognition light life is shortened considerably if used during ground
operations.
Radar . . . . . . . . . . . . . . . . . . . . . . . . . . . . OFF
Transponder . . . . . . . . . . . . . . . . . . . . . STANDBY

Citation V For SimuFlite training only 3B-23

March 1999
 Shutdown
Parking Brake . . . . . . . . . . . . . . . . . . . . . . . SET
Do not set the parking brake if brakes are very hot. This can increase
heat transfer from the brakes to the wheel, causing the fusible plug to
melt and deflate the tire.
Avionics Power Switches . . . . . . . . . . . . . . . . . . OFF
AC power and avionics master switches should be set to OFF.
Exterior Lights . . . . . . . . . . . . . . . . . . . . . . . OFF
Standby Gyro . . . . . . . . . . . . . . . . . . CAGED/OFF
Passenger Advisory Lights . . . . . . . . . . . . . . . . OFF
Overhead and Defog Fans . . . . . . . . . . . . . . . . . OFF
Throttles . . . . . . . . . . . . . . . . . . . . . . . . . . OFF
Allow ITT to stabilize for at least one minute at minimum value. Lifting
the latch and placing the throttle full aft terminates fuel flow to the
engine. A canister collects manifold fuel on shutdown. During the next
flight, this fuel is returned to the fuel cell. Repeated starts for ground
operations cause the canister to overflow through the lower nacelle
after the third shutdown.
Rotating Beacon . . . . . . . . . . . . . . . . . . . . . . OFF
Battery . . . . . . . . . . . . . . . . . . . . . . . . . . . OFF
Move the BATT switch to OFF. Care should be exercised that it is not
placed in EMER. The emergency bus items will drain the battery sig-
nificantly over an extended period.
For deplaning at night, the battery switch may be left in BATT to make
available all cabin lighting until passengers and cabin baggage are
deplaned. Turn the EXTERIOR WING INSP LIGHTS switch to ON
to provide additional illumination in front of the cabin door. An illu-
minated courtesy light switch on the forward door post is wired to the
hot battery bus to turn on the emergency exit lights and one aft baggage
compartment light.
Control Lock . . . . . . . . . . . . . . . . . AS REQUIRED
Engine Covers . . . . . . . . . . . . . . . . . . INSTALLED
In blowing or drifting snow, install engine covers after shutdown as soon
as engines cool sufficiently.

Parking
Normally, the aircraft is parked facing a direction that facilitates ser-
vicing; no concern need be given to the prevailing wind. If not
already accomplished, ensure the following are completed:
Aircraft . . . . . . PARKED ON HARD, LEVEL SURFACE

3B-24 For SimuFlite training only Citation V

March 1999
 Expanded Normal Procedures

Parking Brake and Control Lock . . . . . . . . . . . . . SET

Setting the parking brake is optional. The aircraft may be relocated
without anyone entering the aircraft if the parking brake is not set.
Main Gear . . . . . . . . . . . . . . . . . . . . . CHOCKED
Static Ground Cable . . . . . . . . . . . . . . CONNECTED
Protective Covers
(Figure 3B-1) . . . . . . . . . . INSTALLED AS REQUIRED
Foul Weather Window/Door . . CLOSED AS NECESSARY

Protective Covers

ENGINE EXHAUST COVERS

GENERATOR
INLET COVERS

ENGINE
INLET COVERS

PITOT TUBE COVERS

3B-1

Citation V For SimuFlite training only 3B-25

March 1999
 Mooring
If extended parking plans or impending weather necessitates moor-
ing the aircraft, 3/4-inch ropes (or an equivalent substitute) are
attached to the nose gear and main gear struts. This procedure
requires tie-down eyelets set into the apron; there is no procedure
for mooring at unprepared facilities.
Aircraft . . . . . . . . . . . . . . . . . PARKED ON HARD,
. . . . . . . . . . .LEVEL SURFACE/HEAD INTO WIND
Parking Brake and Control Lock . . . . . . . . . . . . . SET
Setting the parking brake is optional. The aircraft may be relocated
without anyone entering the aircraft if the parking brake is not set.
Main Gear . . . . . . . . . . . . . . . . . . . . . CHOCKED
Static Ground Cable . . . . . . . . . . . . . . CONNECTED
Protective Covers . . . . . . . INSTALLED AS REQUIRED
Ropes (Figure 3B-2) . . . . . . . . . ATTACHED TO NOSE/
. . . . . .MAIN GEAR/SECURED TO PARKING APRON
Foul Weather Window/Door . . CLOSED AS NECESSARY

NOSE GEAR
MAIN GEAR

3B-2

3B-26 For SimuFlite training only Citation V

March 1999
 Expanded Normal Procedures

Taxi Turning Distance

WALL TO WALL 69.90 ft

17.60 ft

21.74 ft

CURB TO CURB 39.40 ft

3B-3

Citation V For SimuFlite training only 3B-27

March 1999
Towbar Turning Distance

WALL TO WALL 52.26 ft

26.13 ft

19.87 ft
8.80 ft

23.75 ft

CURB TO CURB 28.67 ft

3B-4

3B-28 For SimuFlite training only Citation V

March 1999
 Expanded Normal Procedures

Towing/Taxiing
Taxiing the aircraft may be accomplished on hard surfaces as well
as on gravel or sod surfaces. Rudder pedal movement operates the
nose wheel steering system.
On hard surfaces, the aircraft can be towed using a yoke-type tow
bar attached to the nose gear.
While towing or taxiing an aircraft with a flat tire is not recom-
mended, a situation may require it. In such a case, tow or taxi the
aircraft forward enough to clear the immediate area; avoid sharp
turns if towing.
Observe the aircraft taxi turning with brakes and towbar turning dis-
tances depicted on Figure 3B-3 and Figure 3B-4 (previous pages).
Nose Gear Towing
All turns performed during nose gear towing are accomplished
through the tow bar.
Tow Bar (Figure 3B-5) . . . . . PLACED AT NOSE WHEEL
Insert tow bar into nose wheel axle and secure tow bar locking handle.
Tow Bar . . . . . . CONNECTED TO TOWING VEHICLE

NOSE
WHEEL
AXLE

TOWBAR

3B-5

Citation V For SimuFlite training only 3B-29

March 1999
 Pilot’s Seat . . . . . . . . . . . . . . . . . . . . . OCCUPIED
Control Lock . . . . . . . . . . . . . . . . . . DISENGAGED
Parking Brake . . . . . . . . . . . . . . . . . DISENGAGED
If the parking brake is not set, towing can be accomplished without
entering the aircraft.
Chocks/Static Ground Cable/Mooring Ropes . . REMOVED
Wing/Tail Walkers . . . . . . . . STATIONED (OPTIONAL)
In congested areas, wing/tail walkers ensure adequate clearance between
the aircraft and adjacent equipment or structures.
Aircraft . . . . . . . . . . . . . . . . . . . . . . . . . TOWED
Use smooth starts and stops.
When Towing Operation is Complete:
Nosewheel . . . . . . . . . . . . . . . . . . . . . CENTERED
Parking Brake . . . . . . . . . . . ENGAGED (OPTIONAL)
Control Lock . . . . . . . . . . . . . . . . . . . . ENGAGED
Main Gear Wheels . . . . . . . . . . . . . . . . . CHOCKED
Static Ground Cable . . . . . . . . . . . . . . CONNECTED
Tow Bar . . . . . . . . . . . . . . . . . . . . . . REMOVED

Main Gear Towing

Pilot’s Seat . . . . . . . . . . . . . . . . . . . . . OCCUPIED
Main Gear Towing Adapters . . . . . . . . . . INSTALLED
Cables . . . . . . . ATTACHED TO TOWING ADAPTERS/
. . . . . . . . . . . . . . . . . . . . . .TOWING VEHICLE
Use care to prevent crushing of wiring or linkage rods in the area.
Make sure the cables are long enough to clear the aircraft and that the
towing vehicle is on a hard surface.
Chocks/Static Ground Cable/Mooring Ropes . . REMOVED
Parking Brake . . . . . . . . . . . . . . . . . DISENGAGED
Control Lock . . . . . . . . . . . . . . . . . . DISENGAGED
Aircraft . . . . . . . . . . . . . STEERED WITH BRAKES
. . . . . . . . . . . . . . . . . . . .AND RUDDER PEDALS
Using a smooth and even pressure, apply aircraft brakes as required.

3B-30 For SimuFlite training only Citation V

March 1999
 Expanded Normal Procedures

When Towing Operation Is Complete:

Nosewheel . . . . . . . . . . . . . . . . . . . . . CENTERED
Parking Brake . . . . . . . . . . . ENGAGED (OPTIONAL)
Control Lock . . . . . . . . . . . . . . . . . . . . ENGAGED
Main Gear Wheels . . . . . . . . . . . . . . . . . CHOCKED
Static Ground Cable . . . . . . . . . . . . . . CONNECTED
Tow Cables and Towing Adapters . . . . . . . . REMOVED
Taxiing (Ground Movement)
During taxi, the aircraft is controlled via selected engine thrust,
rudder pedal steering, and brakes. Taxiing can be accomplished
with one or both engines operating.
Immediate Area . . . . . . . . . . . . . . . . . . CLEARED
Main Wheels . . . . . . . . . . . . . . . . . . . . CHECKED
Static Ground Cable . . . . . . . . . . . . . . . . REMOVED
Wing Fuel . . . . . . . . . . . . . . . ENSURE BALANCED
The left and right wing tanks should be balanced within 200 pounds.
Pilot Stations . . . . . . . . . . . . . . . BOTH OCCUPIED
Pilot in left seat maneuvers the aircraft. Pilot in right seat acts as an
observer and assists as necessary.
Parking Brake . . . . . . . . . . . . . . . . . . . ENGAGED
Engine(s) . . . . . . . . . . . . . . . . . . . . . . . STARTED
Refer to expanded normal procedure for details.
Wheel Chocks . . . . . . . . . . . . . . . . . . . REMOVED
Parking Brake . . . . . . . . . . . . . . . . . . . RELEASED
Thrust . . . . . APPLIED SUFFICIENT TO START ROLL
Aircraft Brakes . . . . . . . . . APPLIED AS NECESSARY
Use brakes intermittently rather than continuously. Allow the aircraft
to accelerate, then brake to acceptable taxi speed.
Aircraft . . . . . . . . STEERED WITH RUDDER PEDALS
Nose wheel steering is accomplished through the rudder pedals.
Wing Walkers . . . . . . . . . . . . . . . . . . STATIONED
Taxi . . . . . . . . . . . . . . . . . . . . . ACCOMPLISHED
Parking Brake . . . . . . . . . . . . . . . . . . . ENGAGED

Citation V For SimuFlite training only 3B-31

March 1999
 Engine(s) . . . . . . . . . . . . . . . . . . . . . SHUT DOWN
Refer to expanded normal procedure for details.
Main Gear Wheels . . . . . . . . . . . . . . . . . CHOCKED

3B-32 For SimuFlite training only Citation V

March 1999
 Expanded Normal Procedures

Storage Storage and

Aircraft storage practices vary depending on the length of the stor-
age period. There are, however, several general policies to observe. Restoring
If the aircraft is to be stored outside, adhere to parking and moor-
ing requirements. All protective covers should be installed. Cover
the aircraft’s air inlets. Do not set the parking brake. In all other
cases, the following are generally performed.
0–7 Days
If the engines are in a sheltered environment (i.e., not exposed to
excessive humidity or temperature changes), no action need be
taken beyond installing protective covers.
Fuel Crossfeed Line . . . . . . . . . . . . . . . . . DRAINED
Wing Tanks . . . . . . . . . . . . . . . . . . . . . . FILLED
7–30 Days
In addition to the procedures for 0–7 day storage, the following
are performed.
Battery . . . . . . . . . . . . . . . . . . . DISCONNECTED
If the battery is left in the aircraft, regular servicing is required to pre-
vent discharge. If the battery is removed, check its charge regularly.
Wheels . . . . . . . . . . . . ROTATED 100° CLOCKWISE
. . . . . . . . . . . . . . . . . . . . .EVERY TWO WEEKS
Use the valve stem as a reference point.
Engine Desiccant . . . . . . . . INSTALLED IN EXHAUST
■ If engine openings are sealed off, and if relative humidity in the
engine is maintained below 40%, no further action need be taken
concerning the engine. Installing a desiccant ensures the engine’s
relative humidity remains at less than 40%.
■ Keep the desiccant off all engine parts by placing it on racks.
For storage longer than 30 days, refer to Chapter 10 of the Model
560 Maintenance Manual for a detailed list of procedures.

Citation V For SimuFlite training only 3B-33

March 1999
 Restoring
After an aircraft has been stored for a period, it must be restored
to an airworthy state. Based on the length of storage, the follow-
ing are generally performed.

From 0–30 Days

All Protective Covers . . . . . . . . . . . . . . . REMOVED
Engines . . . . . . . . . . . . . . . . . . . . . . INSPECTED
Remove desiccants. Check for foreign objects, and examine overall
condition.
Aircraft . . . . . . . . . . . . . . . CLEAN, IF REQUIRED
Aircraft . . . . . . . . . . . . . . . . . . . . . LUBRICATED
Fuel Water/Sediment . . . . . . . . . . . DRAINED USING
. . . . . . . . . . . . . . . . . . . . . FUEL DRAIN VALVES
If Oil Tank Drained . . . . . . . . . . . . . . FILLED WITH
. . . . . . . . . . . . . . . . . . . . .RECOMMENDED OIL
Brake Cylinders and Reservoir . . . . . . . . . . CHECKED
Shock Struts . . . . . . . . . . . . . . . . . . . . CHECKED
Tires . . . . . . . . . . . . . . . . . . . . . . . . . CHECKED
Hydraulic Reservoir . . . . . . . . . . . . . . . . CHECKED
Battery . . . . . . . . . . . . . . . . . . . . . CONNECTED
Fuel Manual Shutoff Valve . . . . . . OPENED, IF CLOSED

3B-34 For SimuFlite training only Citation V

March 1999
 Expanded Normal Procedures

Preflight Cold Weather

During preflight preparation, inspect areas where surface snow or
frost could change or affect normal system operations. Supplemental Operations
preflight checks should include the following:
All Protection Covers . . . . . . . . . . . . . . . REMOVED
Surface . . CHECKED FREE OF FROST, ICE, AND SNOW
The wing leading edges, all control surfaces, tab surfaces, and control
cavities must be free of ice or snow. Check control cavities for drainage
after snow removal because water puddles may re-freeze in flight.
Generator Inlets . . CLEARED OF INTERNAL ICE/SNOW
NOTE: Visible moisture
Engine Inlets . . . . CLEARED OF INTERNAL ICE/SNOW includes, but is not limited to,
Check that the inlet cowling and tailcone air inlet are free of ice or the following conditions: fog
snow and that the engine fan is free to rotate. with visibility less than one
mile, wet snow, and rain.
Fuel Tank Vents . . CHECKED FREE OF ICE AND SNOW
Check the fuel tank vents; all traces of ice or snow should be removed.
Fuel Drains . . . . . . . . . . . . . ALL WATER DRAINED
Pitot Heads And Static Ports . . . . . . CLEARED OF ICE
Water rundown resulting from snow removal may re-freeze immedi-
ately forward of the static ports. This causes an ice buildup that results
in disturbed airflow over the static ports and can cause erroneous sta-
tic readings even though the static ports themselves are clear.
Landing Gear Doors . . . . . . . . . . . . . . . . CHECKED
Make sure the landing gear doors are unobstructed and free of impact-
ed ice or snow.
Air Conditioning Inlets and Exits . . . . . . . . . CLEARED
Verify that the air inlets and exits are clear of ice or snow.
Aircraft Deicing . . . . . . . . . . . . . . . . COMPLETED
For different deicing fluids, the times of protection (the holdover times)
vary considerably. Furthermore, these times depend to a large extent
on meteorological conditions.
Preliminary Cockpit Preparations . . . . . . COMPLETED
Battery . . . . . . . . . . . . . INSTALLED/PREHEATED/
. . . . . . . . . . . . . . . . . . . . . . .FULLY CHARGED
Minimum voltage is 24V DC.
Engine . . . . . . . . . . . . . . . . . . . . . . PREHEATED
Engine preheating is best accomplished by installing the engine cov-
ers and directing hot air through the oil filler access door.

Citation V For SimuFlite training only 3B-35

March 1999
 EPU Start
If the airplane has been cold soaked below -10°C, an external
power unit and/or preheat procedure should be used for starting.

Engine Start
During cold weather starts, initial oil pressure may be slow in ris-
ing; the oil pressure annunciator may remain illuminated longer
than normal.

After Engine Start

Instruments . . . . . . . . . . . . . . . . . OBSERVED FOR
. . . . . . . . . . . . . . . . . . . .NORMAL OPERATION
The engine instruments should display approximately normal indica-
tions within a short time after reaching idle.
Engine Oil Pressure . . . . . . . . . . . . . . . . CHECKED
During cold weather starts, the oil pressure may temporarily exceed
maximum pressure limits until the oil temperature rises. At low ambi-
ent temperatures, a temporary high pressure above maximum limits
may be tolerated, but takeoff should be delayed until the pressure drops
into normal limits.
Anti-Ice . . . . . . . . . . . . . . . . . . . . AS REQUIRED
During operation from snow-covered runways, engine anti-ice should
be on during taxi and takeoff. Takeoff should be preceded by a static
engine run-up to as high a power level as practical to ensure observa-
tion of stable engine operation prior to brake release.
If severe icing conditions are present, engine anti-ice should be turned
on immediately after engine start. During prolonged ground opera-
tion, periodic engine run-up may be performed to reduce the possi-
bility of ice buildup. For ground operation, the system should be turned
on one minute out of four with N2 set above 65%.
Flight Controls . . . . . . . . . . . . . . . . CHECKED FOR
. . . . . . . . . . . . . . . . .FREEDOM OF MOVEMENT
Accomplish this check when the aircraft has been exposed for an
extended period of time to snow, freezing rain or other conditions that
could restrict flight control movement. Increased control forces can
be expected at low temperatures because of the increased resistance in
cables and the congealed oil in snubbers and bearings. It may be desir-
able to accomplish an additional control check prior to taxi.
Wing Flaps . . . CHECKED FOR NORMAL OPERATION

3B-36 For SimuFlite training only Citation V

March 1999
 Expanded Normal Procedures

Air Conditioning Switches . . . . . . . . . . . . . . SET FOR

. . . . . . . . . . . . . . . . . . .MAXIMUM CABIN HEAT
Temperature Control . . . . . . . . . . . . . MANUAL HOT
Overhead Fan . . . . . . . . . . . . . . . . . . . . . . . . HI
Press Source Selector . . . . . . . . . . . . . . . . . . . GND
Remember to reduce temperature control prior to takeoff.
Windshield Bleed Air . . . . . . . . . . . . . . LOW OR HI
Use windshield bleed air and defog fan to clear the windshield.

Taxi/Before Takeoff
Flaps . . . . . . . . . . . . . . . . . . . T.O. or T.O. & APPR
Extend flaps to the T.O. or the T.O. & APPR setting at this time if they
have been held because of slush or wet snow.
Before Takeoff Checklist . . . . . . . . . . . . COMPLETED
To ensure the aircraft is configured for takeoff, recheck the flap posi-
tion indicator.

Takeoff
If Engine Anti-Ice is used for Takeoff:
Thrust . . . . . . . . . . . . . . . . . ANTI-ICE SETTINGS
V1/Takeoff Field Length . . . . . . . . . . . . . ADJUSTED
. . . . . . . . . . . . . . . .IN ACCORDANCE WITH AFM
If Aircraft Slides on Ice or Snow during Engine Power Check:
Brakes . . . . . . . . . . . . . . . . . . . . . . . RELEASED
Takeoff Roll . . . . . . . . . . . . . . . . . . . . . . BEGUN
Continue engine check during early part of takeoff roll. During take-
offs on icy runways, a lag in nosewheel steering must be expected and
corrections must be anticipated. A light forward pressure on the
control column increases nosewheel steering effectiveness.

Rejected Takeoff on Slick Runway

Maximum Braking . . . . . . . . . . . . . . . . . . . . USED
Speedbrakes . . . . . . . . . . . . . . . . . . . . DEPLOYED
Reverse Thrust . . . . . . . . . . . . . . . . . . MAXIMUM
Rudder . . . . . . . . . . . . . . . . USED FOR PRIMARY
. . . . . . . . . . . . . . . . . .DIRECTIONAL CONTROL

Citation V For SimuFlite training only 3B-37

March 1999
 If a Skid Develops:
Reverse Thrust . . . . . . REDUCED TO IDLE REVERSE
Engine Power . . . . . . . . . . . . . . . . . RETURNED TO
. . . . . . . . . . .FORWARD THRUST AT LOW POWER
. . . . . . . . . . . . . .(IF NECESSARY) TO RETURN TO
. . . . . . . . . . . . . . . . . . . .RUNWAY CENTERLINE
Rudder Pedal Steering . . . . . USED FOR DIRECTIONAL
. . . . . . . . . . . . . . . . . . .CONTROL, IF POSSIBLE
Rudder and differential braking may also be used as necessary for
directional control. Reduce brake pressure if directional control prob-
lems are encountered because of excessive anti-skid cycling.

In Flight
Pitot Heat . . . . . . . . ON FOR DURATION OF FLIGHT
Windshield Bleed Air . . . . . . . . . . . . . . LOW OR HI
This keeps the windshield clear (HI at -18°C [0°F] OAT or below).
Speedbrakes plus increased power settings provide additional bleed
air.
Windshield Alcohol . . . . . . . . . . . . . . . . . . . . . ON
This should be used if windshield bleed air fails. The alcohol lasts
approximately 10 minutes and is distributed to the pilot’s windshield
only.
Engine Anti-Ice . . . . . . . . . . . . . . . . . . . . . . . ON
Use when operating in visible moisture with outside air temperatures
between -30°C and +10°C; use anti-ice thrust settings.
Surface Deice . . . . . . . . . . . . . . . . . . . . . . . AUTO
Use when wing ice buildup is estimated between 1/4 and 1/2 inch.

Taxi-in and Park

Engine Anti-Ice . . . . . . . . . . . . . . . . AS REQUIRED
If severe icing conditions are present, engine anti-icing should be
turned on. During prolonged ground operation, periodic engine run-ups
may be performed to reduce the possibility of ice buildup. For ground
operation, the system should be turned on one minute out of four with
N2 set at 70%.
Windshield Bleed Air . . . . . . . . . . . . . . LOW OR HI
Use windshield bleed air and the defog fan to clear the windshield.

3B-38 For SimuFlite training only Citation V

March 1999
 Expanded Normal Procedures

Securing Overnight or for Extended

Period (Aircraft Unattended)
Wheel Chocks . . . . . . . . . . . . . CHECKED IN PLACE
Parking Brake . . . . . . . . . . . . . . . . . . . . . . . OFF
This eliminates the possibility of the brakes freezing.
Protective Covers . . . . . . . . . . . . . . . . . INSTALLED
Water Storage Containers . . . . . . . . . . . . . DRAINED
Toilets . . . . . . . . . . . . . . . . . . . . . . . . DRAINED
Battery . . . . . . . . . . . . . . . . . . . . . . . REMOVED
If the ni-cad battery will be exposed to temperatures below -18°C
(0°F), the battery should be removed and stored in an area warmer
than -18°C (0°F) but below 40°C (104°F). Subsequent re-installation
of the warm battery enhances starting capability.
Doors . . . . . . . . . . . . . . . . CLOSED AND LOCKED

Ground Cooling
For maximum ground cabin cooling:
Temperature Control . . . . . . . . . . . . . . FULL COLD
Press Source Selector . . . . . . . . . . . . . . . . . . . GND
Overhead Fan . . . . . . . . . . . . . . . . . . . . . . . . . HI
Defog Fan . . . . . . . . . . . . . . . . . . . . . . . . . . . HI
Freon Air Conditioner (if installed) . . . . . . . . . . . . ON

Performance Computations, Takeoff

Make takeoff performance computations in accordance with the
FAA approved AFM.

Citation V For SimuFlite training only 3B-39

March 1999
3B-40 For SimuFlite training only Citation V
March 1999
SimuFlite strongly supports the premise that the disciplined use
of well-developed Standard Operating Procedures (SOP) is Standard
central to safe, professional aircraft operations, especially in
multi-crew, complex, or high performance aircraft. Operating
If your flight department has developed an SOP, we encourage
you to use it during your training. If your flight department
Procedures
does not already have one, we welcome your use of the
SimuFlite SOP.
Experienced corporate pilots carefully developed this SOP for
Chapter 3C
SimuFlite. It is the product of their experience and the result of
a consensus on the way SimuFlite conducts its flight operations.
The procedures described herein are specific to the Citation V
unless manufacturer- or FAA-specified procedures override
them. They apply to the phases of flight as specified, and the
flight crew member designated for each step should accom-
plish it as indicated.

Citation V For SimuFlite training only 3C-1

March 1999
3C-2 For SimuFlite training only Citation V
March 1999
 Standard Operating Procedures

General Information . . . . . . . . . . . . . . . . . 3C-5

Table of
Definitions . . . . . . . . . . . . . . . . . . . . . . . 3C-5
Flow Patterns . . . . . . . . . . . . . . . . . . . . . 3C-5
Contents
Checklists . . . . . . . . . . . . . . . . . . . . . . . 3C-5
Omission of Checklists . . . . . . . . . . . . . . . 3C-5
Challenge/No Response . . . . . . . . . . . . . . 3C-5
Abnormal/Emergency Procedures . . . . . . . . . . 3C-5
Time Critical Situations . . . . . . . . . . . . . . . 3C-6
Rejected Takeoffs . . . . . . . . . . . . . . . . . . 3C-6
Critical Malfunctions in Flight . . . . . . . . . . . . 3C-6
Non-Critical Malfunctions in Flight . . . . . . . . . . 3C-7
Radio Tuning and Communication . . . . . . . . . . 3C-7
Altitude Assignment . . . . . . . . . . . . . . . . . . 3C-7
Pre-Departure Briefings . . . . . . . . . . . . . . . . 3C-7
Advising of Aircraft Configuration Change . . . . . . 3C-7
Transitioning from Instrument
to Visual Conditions . . . . . . . . . . . . . . . . . . 3C-7

Phase of Flight SOPs . . . . . . . . . . . . . . . . 3C-9

Holding Short . . . . . . . . . . . . . . . . . . . . . 3C-9
Takeoff Roll . . . . . . . . . . . . . . . . . . . . . 3C-10
Climb . . . . . . . . . . . . . . . . . . . . . . . . . 3C-11
Cruise . . . . . . . . . . . . . . . . . . . . . . . . 3C-13
Descent . . . . . . . . . . . . . . . . . . . . . . . 3C-14
Approach . . . . . . . . . . . . . . . . . . . . . . . 3C-16
Approach Deviations . . . . . . . . . . . . . . . . . 3C-20
Missed Approach . . . . . . . . . . . . . . . . . . 3C-22
Visual Traffic Patterns . . . . . . . . . . . . . . . . 3C-24
Landing . . . . . . . . . . . . . . . . . . . . . . . 3C-26

Citation V For SimuFlite training only 3C-3

March 1999
3C-4 For SimuFlite training only Citation V
March 1999
 Standard Operating Procedures

Definitions
LH/RH: Pilot Station. Designation of
In all cases, the response by either pilot
is confirmed by the other and any dis- General
seat position for accomplishing a given
task because of proximity to the
agreement is resolved prior to contin-
uing the checklist. Information
respective control/indicator. Regardless After the completion of any checklist,
of PF or PNF role, the pilot in that seat the PNF states “______ checklist is
performs tasks and responds to check- complete.” This allows the PF to main-
list challenges accordingly. tain situational awareness during check-
list phases and prompts the PF to con-
PF: Pilot Flying. The pilot responsible
tinue to the next checklist, if required.
for controlling the flight of the aircraft.
Effective checklists are pertinent and
PIC: Pilot-in-Command. The pilot
concise. Use them the way they are
responsible for the operation and safe-
written: verbatim, smartly, and pro-
ty of an aircraft during flight time.
fessionally.
PNF: Pilot Not Flying. The pilot who
is not controlling the flight of the air- Omission of Checklists
craft. While the PF is responsible for initiat-
ing checklists, the PNF should ask the
Flow Patterns PF whether a checklist should be start-
Flow patterns are an integral part of ed if, in his opinion, a checklist is over-
the SOP. Accomplsih the cockpit setup looked. As an expression of good crew
for each phase of flight with a flow resource management, such prompting
pattern, then refer to the checklist to is appropriate for any flight situation:
verify the setup. Use normal check- training, operations, or checkrides.
lists as “done lists” instead of “do
Challenge/No Response
lists.”
Flow patterns are disciplined proce- If the PNF observes and challenges a
dures; they require pilots who under- flight deviation or critical situation,
stand the aircraft systems/controls and the PF should respond immediately. If
who methodically accomplish the flow the PF does not respond by oral com-
pattern. munication or action, the PNF must
issue a second challenge that is loud
A standardized flow pattern appears and clear. If the PF does not respond
in the Preflight Inspection chapter. after the second challenge, the PNF
must ensure the safety of the aircraft.
Checklists The PNF must announce that he is
Use a challenge-response method to assuming control and then take the
execute any checklist. After the PF ini- necessary actions to return the aircraft
tiates the checklist, the PNF challenges to a safe operating envelope.
by reading the checklist item aloud.
The PF is responsible for verifying that Abnormal/Emergency
the items designated as PF or his seat Procedures
position (i.e., LH or RH) are accom- When any crewmember recognizes an
plished and for responding orally to abnormal or emergency condition, the
the challenge. Items designated on the PIC designates who controls the air-
checklist as PNF or by his seat posi- craft, who performs the tasks, and any
tion are the PNF’s responsibility. The items to be monitored.
PNF confirms the accomplishment of
the item, then responds orally to his Following these designations, the PIC
own challenge. calls for the appropriate checklist.

Citation V For SimuFlite training only 3C-5

March 1999
 The crewmember designated on the ■ RESPOND – Take appropriate
NOTE: “Control” means checklist accomplishes the checklist action.
responsible for flight control of items with the appropriate challenge/ Rejected Takeoffs
the aircraft, whether manual or response.
automatic. The rejected takeoff procedure is a pre-
The pilot designated to fly the aircraft planned maneuver; both crewmembers
(i.e., PF) does not perform tasks that must be aware of and briefed on the
compromise this primary responsibil- types of malfunctions that mandate an
ity, regardless of whether he uses the abort. Assuming the crew trains to a
autopilot or flies manually. firmly established SOP, either crew-
Both pilots must be able to respond to member may call for an abort.
an emergency situation that requires The PF normally commands and exe-
immediate corrective action without cutes the takeoff abort for directional
reference to a checklist. The elements of control problems or catastrophic mal-
an emergency procedure that must be functions. Additionally, any indication
performed without reference to the of the following malfunctions prior to
appropriate checklist are called mem- V1 is cause for an abort:
ory or recall items. Accomplish all other
■ engine failure
abnormal and emergency procedures
while referring to the printed checklist. ■ engine fire
Accomplishing abnormal and emer- ■ thrust reverser deployment
gency checklists differs from accom- ■
plishing normal procedure checklists loss of directional control
in that the pilot reading the checklist In addition to the above, the PF usu-
states both the challenge and the ally executes an abort prior to 70 KIAS
response when challenging each item. for any abnormality observed.
When a checklist procedure calls for When the PNF calls an abort, the PF
the movement or manipulation of con- announces “Abort” or “Continue” and
trols or switches critical to safety of executes the appropriate procedure.
flight (e.g., throttles, engine fire switch-
es, fire bottle discharge switches), the Critical Malfunctions in Flight
pilot performing the action obtains ver- In flight, the observing crewmember
ification from the other pilot that he is positively announces a malfunction.
moving the correct control or switch As time permits, the other crewmem-
prior to initiating the action. ber makes every effort to confirm/iden-
Any checklist action pertaining to a tify the malfunction before initiating
specific control, switch, or equipment any emergency action.
that is duplicated in the cockpit is read If thePNF is the first to observe any
to include its relative position and the indication of a critical failure, he
action required (e.g., “Left Throttle – announces it and simultaneously iden-
OFF; Left Boost Pump – NORMAL”). tifies the malfunction to the PF by
Time Critical Situations pointing to the indicator/annunciator.
After verifying the malfunction, the
When the aircraft, passengers, and/or PF announces his decision and com-
crew are in jeopardy, remember three mands accomplishment of any check-
things. list memory items. The PF monitors
■ FLY THE AIRPLANE – Maintain the PNF during the accomplishment
aircraft control. of those tasks assigned to him.
■ RECOGNIZE CHALLENGE –
Analyze the situation.

3C-6 For SimuFlite training only Citation V

March 1999
 Standard Operating Procedures

Non-Critical Malfunctions tude. The PNF continues to point to the

in Flight altitude alerter until the PF confirms the NOTE: The acronym AWARE
altitude assignment and alerter setting. stands for the following:
Procedures for recognizing and veri- ■ Aircraft status
fying a non-critical malfunction or Pre-Departure Briefings ■ Weather
impending malfunction are the same
as those used for time critical situa- The PIC should conduct a pre-depar- ■ Airport information

tions: use positive oral and graphic ture briefing prior to each flight to
■ Route of flight
communication to identify and direct address potential problems, weather
delays, safety considerations, and oper- ■ Extra.
the proper response. Time, however,
is not as critical and allows a more ational issues. Pre-departure briefings
deliberate response to the malfunction. should include all crewmembers to
Always use the appropriate checklist to enhance team-building and set the tone
accomplish the corrective action. for the flight. The briefing may be for-
mal or informal, but should include
Radio Tuning and some standard items. The acronym
Communcation AWARE works well to ensure no
points are missed. This is also an
The PNF accomplishes navigation and opportunity to brief any takeoff or
communication radio tuning, identifi- departure deviations from the SOP due
cation, and ground communication. to weather or runway conditions.
For navigation radios, the PNF tunes
and identifies all navigation aids. Advising of Aircraft
Before tuning the PF’s radios, he Configuration Change
announces the NAVAID to be set. In
tuning the primary NAVAID, in par- If the PF is about to make an aircraft
ticular, the PNF coordinates with the control or configuration change, he
PF to ensure proper selection sequenc- alerts the PNF to the forthcoming
ing with the autopilot mode. After tun- change (e.g., gear, speedbrake, and flap
ing and identifying the PF’s NAVAID, selections). If time permits, he also
the PNF announces “(Facility) tuned, announces any abrupt flight path
and identified.” changes so there is always mutual
understanding of the intended flight
Monitor NDB audio output anytime path.
the NDB is in use as the NAVAID. Use
the marker beacon audio as backup to Time permitting, a PA announcement
visual annunciation for marker pas- to the passengers precedes maneuvers
sage confirmation. involving unusual deck or roll angles.
In tuning the VHF radios for ATC com- Transitioning from
munication, the PNF places the newly Instrument to Visual
assigned frequency in the head not in Conditions
use (i.e., preselected) at the time of
receipt. After contact on the new fre- If visual meteorological conditions
quency, the PNF retains the previously (VMC) are encountered during an
assigned frequency for a reasonable instrument approach, the PNF nor-
time period. mally continues to make callouts for
the instrument approach being con-
Altitude Assignment ducted. However, the PF may request
The PNF sets the assigned altitude in a changeover to visual traffic pattern
the altitude alerter and points to the callouts.
alerter while orally repeating the alti-

Citation V For SimuFlite training only 3C-7

March 1999
3C-8 For SimuFlite training only Citation V
March 1999
 Standard Operating Procedures

Phase of Flight SOP

Holding Short


P

PN

call: “Before Takeoff Checklist”

action: Complete Before
Takeoff Checklist


call: “Before Takeoff
Checklist Complete”

Takeoff Briefing

action: Brief the Following:

ƒƒ Assigned Runway for Takeoff
ƒƒ Initial Heading/Course
ƒƒ Initial Altitude
ƒƒ Airspeed Limit (If Applicable)
ƒƒ Clearance Limit
ƒƒ Emergency Return Plan
ƒƒ SOP Deviations
Consider the Following:
ƒƒ Impaired Runway Conditions
ƒƒ Weather
ƒƒ Obstacle Clearance
ƒƒ Intrument Departure Procedures
ƒƒ Abort

Cleared for Takeoff



action: Confirm Assigned

Runway for Takeoff
and Check Heading
Indicator Agreement


call: “Assigned Runway
Confirmed, Heading


action: Confirm Assigned Runway Checked”
for Takeoff and Check Heading
Indicator Agreement


call: “Assigned Runway
Confirmed, Heading
Checked”

call: “Takeoff Checklist”

action: Complete Takeoff
Checklist


call: ”Takeoff Checklist
Complete”

Citation V For Training Purposes Only 3C-9

August 2010
 Takeoff Roll


P

PN
Setting Takeoff Power


call: ”Set _____.”

call: “_____ set.”

Initial Airspeed Indication

call: “Airspeed alive.” At 70 KIAS


call: “70 kts crosscheck.”

At V1


call: “V1.”


action: Move hand from
throttles to yoke.
At VR


call: “Rotate.”


action: Rotate to approximately
10° pitch altitude
for takeoff.”

3C-10 For Training Purposes Only Citation V

March 1999
 Standard Operating Procedures

Climb
PF PNF

At Positive Rate of Climb

CALL “Positive rate.”
Only after PNF’s call,
CALL “Gear – UP.”
CALL “Gear selected up.”
When annunciator
illuminates,
“Gear indicates UP.”

After Gear Retraction

ACTION Immediately
accomplish attitude
correlation check.
■ PF’s and PNF’s ADI

displays agree.
■ Pitch and bank

angles are
acceptable
CALL “Attitudes check,”
or if a fault exists, give
a concise statement of
the discrepancy.

At V2 + 10 KIAS and 400 Ft Above Ground Level (Minimum)

CALL “V2 +10 KIAS.”
CALL “Flaps – UP.”
CALL “Flaps selected UP.”
When indicator
shows UP,
“Flaps indicate UP.”

At VENR (Minimum)
CALL “Climb power.”
CALL “Climb power set.”

Citation V For SimuFlite training only 3C-11

March 1999
 Climb (continued)
PF PNF

At 3,000 Ft Above Ground Level and Clear of Traffic

CALL “Climb checklist.”

ACTION Complete Climb
checklist.
CALL “Climb checklist
complete.”

At FL180 (or Transition Level)

CALL “29.92 set.” CALL “29.92 set.”
“Transition Level
checklist.”
ACTION Complete Transition
Level checklist.
CALL “Transition Level
checklist complete.”

At 1,000 Ft Below Assigned Altitude

CALL “1,000 ft to go.”

CALL “Check. 1,000 ft.”

3C-12 For SimuFlite training only Citation V

March 1999
 Standard Operating Procedures

Cruise
PF PNF

At Final Cruise Altitude

CALL “Cruise checklist.”
ACTION Complete Cruise
checklist.
CALL “Cruise checklist
complete.”

Altitude Deviation in Excess of 100 Ft

CALL “Altitude.”
CALL “Correcting.”

Course Deviation in Excess of One Half Dot

CALL “Course.”
CALL “Correcting.”

Citation V For SimuFlite training only 3C-13

March 1999
 Descent
PF PNF

Upon Initial Descent from Cruise

CALL “Descent checklist.”
ACTION Set bugs. ACTION Complete Descent
checklist. Set airspeed
and altimeter bugs.
CALL “Descent checklist
complete.”

At 1,000 Ft Above Assigned Altitude

CALL “1000 ft to go.”

CALL ”Check. 1000 ft.”

At FL180

CALL “Altimeter set ____.” CALL “Altimeter set ____.”

”Transition Level
Checklist.”
ACTION Complete Transition
Level checklist.
CALL ”Transition Level
checklist complete.”

At 10,000 Ft

CALL “10,000 ft.”

CALL “Check.
Speed 250 kts.”

Strive for sterile cockpit – Below 10,000 ft MSL

3C-14 For SimuFlite training only Citation V

March 1999
 Standard Operating Procedures

Descent (continued)
PF PNF

At Appropriate Workload Time

REVIEW REVIEW

■ approach to be executed
■ field elevation
■ appropriate minimum sector altitude(s)
■ inbound leg to FAF, procedure turn direction and altitude
■ configuration (flap setting)
■ approach speed
■ final approach course heading and intercept altitude
■ timing required
■ DH/MDA
■ MAP (non-precision)
■ VDP
■ special procedures (DME step-down, arc, etc.)
■ type of approach lights in use (and radio keying
procedures, if required)
■ missed approach procedures
■ abnormal implications

ACTION Brief:
■ Configuration

■ Approach speed

■ Minimum safe altitude

■ Approach course

■ FAF altitude

■ DA/MDA altitude

■ Field elevation

■ VDP

■ Missed approach

■ Heading

■ Altitude

■ Intentions

■ Abnormal implications

Accomplish as many checklist items as possible. The approach

checklist must be completed prior to the initial approach fix.

Citation V For SimuFlite training only 3C-15

March 1999
 Approach
PF PNF

Prior to Initial Approach Fix

CALL “Approach checklist.”

ACTION Complete Approach
checklist.
CALL “Approach checklist
complete.”

After Level-off on Intermediate Approach Segment

CALL “Flaps – APPROACH.”

CALL “Flaps selected to
APPROACH.”
When flaps indicate
APPROACH,
“Flaps indicate
APPROACH.”

At Initial Convergence of Course Deflection Bar

CALL “Localizer/course CALL “Localizer/course

alive.” alive.”

At Initial Downward Movement of Glideslope Raw Data Indicator

CALL “Glideslope alive.” CALL “Glideslope alive.”

When Annunciators Indicate Localizer Capture

CALL “Localizer/course CALL “Localizer/course

captured.” captured.”

3C-16 For SimuFlite training only Citation V

March 1999
 Standard Operating Procedures

Approach (continued)
PF PNF
At One Dot From Glideslope Intercept (Prior to FAF on
Non-Precision Approach)
CALL “One dot to go.”
CALL “Gear – Down.
Before Landing
checklist.”
CALL “Gear selected
DOWN.”
When gear indicates
DOWN, “Gear
indicates DOWN.”
ACTION Complete Before
Landing checklist
except for full flaps
and autopilot/yaw
damper.

When Annunciator Indicates Glideslope Capture

CALL “Glideslope captured.” CALL “Glideslope captured.”

For two-engine ILS only,
CALL “Flaps – LAND.”
CALL “Flaps selected to
LAND.”
When flaps indicate
LAND,
“Flaps indicate LAND.”

If the VOR on the PNF’s side is used for crosschecks on the

intermediate segment, the PNF’s localizer and glideslope status calls
are accomplished at the time the PNF changes to the ILS frequency.
This should be no later than at completion of the FAF crosscheck, if
required. PNF should tune and identify his NAV radios to the specific
approach and monitor.

Citation V For SimuFlite training only 3C-17

March 1999
 Approach (continued)
PF PNF

At FAF

CALL “Outer marker.” or CALL “Outer marker,” or

“Final fix.” “Final fix.”
ACTION ■ Start timing.
■ Visually crosscheck

that both altimeters

agree.
■ Set missed approach

altitude in altitude
alerter.
■ Check PF and PNF

instruments.
■ Call FAF inbound.

CALL “Altimeters check.

no flags.”

At 1,000 Ft Above Minimum

CALL “1,000 ft to minimums.
No flags.”
CALL “Check. No flags.”

At 500 Ft Above Minimum

CALL “500 ft to minimums.
No flags.”
CALL “Check.”

Abandoned Approach
When within 500 ft AGL, the aircraft must be within the “Approach
Window”:
■ within one dot deflection, both LOC and GS

■ IVSI less than 1,000 fpm

■ IAS with VAP +10 (no less than V

REF or 0.6 AOA, whichever is
less)
■ no flight instrument flags with the landing runway or visual

references not in sight

■ landing configuration, except for full flaps (non-precision or single

engine approaches).
If the aircraft is not within this “window,” a go-around must be
executed.

3C-18 For SimuFlite training only Citation V

March 1999
 Standard Operating Procedures

Approach (continued)
PF PNF

At 200 Ft Above Minimums

CALL “200 ft to go.”

CALL “Check.”

At 100 Ft Above Minimums

CALL “100 ft to go.”

CALL “Check.”

At Minimums (Precision)

CALL “Minimums. Runway

(or visual reference)
____ o’clock.”
ACTION Announce intentions
CALL “Going visual. Land,”
or “Missed approach.”
ACTION As PF goes visual,
PNF transitions to
instruments.

Non-Precision at MDA
CALL “Minimums. ____ time
to go,” or “Minimums.
____ distance to go.”
CALL “Check.”

At Point Where PNF Sights Runway or Visual References

CALL “Runway (or visual

reference) ___
o’clock.”
CALL “Going visual. Land,”
or “Missed approach.”

Citation V For SimuFlite training only 3C-19

March 1999
 Approach Deviations
PF PNF

Abandoned Approach
When within 500 ft AGL, the aircraft must be within the “Approach
Window”:
■ within one dot deflection, both LOC and GS

■ IVSI less than 1,000 fpm

■ IAS with VAP +10 (no less than V

REF or 0.6 AOA, whichever is
less)
■ no flight instrument flags with the landing runway or visual

references not in sight

■ landing configuration, except for full flaps (non-precision or single

engine approaches).
If the aircraft is not within this “window,” a go-around must be
executed.

± One Dot – Glideslope

CALL “One dot (high, low)

and (increasing,
holding, decreasing).”
CALL “Correcting.”
± One Dot – Localizer/VOR

CALL “One dot (right, left)

and (increasing,
holding, decreasing).”
CALL “Correcting.”
± 5° ADF Pointer At or Beyond Midpoint for NDB Approach

CALL “(degrees off course)

degrees (right/left) and
(increasing, holding,
decreasing).”
CALL “Correcting.”

VREF + _______ (Correction for Wind and Configuration)

CALL “VREF plus ____ and

(increasing, holding,
decreasing).”
CALL “Correcting.”

3C-20 For SimuFlite training only Citation V

March 1999
 Standard Operating Procedures

Approach Deviations (continued)

PF PNF

At or Below VREF

CALL “VREF,” or “VREF

minus (knots below
VREF).”
CALL “Correcting.”

Rate of Descent Exceeds 1,000 FPM

CALL “Sink (amount)

hundred and
(increasing, holding,
decreasing).”
CALL “Correcting.”

Citation V For SimuFlite training only 3C-21

March 1999
 Missed Approach
PF PNF

At DA (Precision)

CALL “Minimums. No
runway.”
CALL “Missed approach.”
ACTION Apply power firmly and ACTION Assist PF in setting
positively. power for go-around.
Activate go-around
mode and initially
rotate the nose to
the flight director
go-around attitude.
CALL Flaps – APPROACH.”
CALL “Flaps APPROACH
selected.”
When flaps indicate
APPROACH,
“Flaps indicate
APPROACH.”

At MAP (Non-Precision)

CALL “Missed approach

point. No runway.”
ACTION Apply power firmly and ACTION Assist PF in setting
positively. Activate power for go-around.
go-around mode and
initially rotate the nose
to the flight director
go-around attitude.
CALL “Flaps – APPROACH.”
NOTE: Flaps should already be CALL “Flaps indicate
at APPROACH before calling APPROACH.”
“Flaps – Approach.”

3C-22 For SimuFlite training only Citation V

March 1999
 Standard Operating Procedures

Missed Approach (continued)

PF PNF

At Positive Rate of Climb

CALL “Positive rate.”

CALL “Gear – UP.”
CALL “Gear selected UP.”
When gear indicates
UP, “Gear indicates
UP.”
ACTION Announce heading and
altitude for missed
approach.

At VREF + 10 and 400 Ft Above Ground Level (Minimum)

CALL “Flaps – UP.”

CALL “Flaps selected up.”
When flaps indicate
UP, “Flaps indicate
UP.”

To ATC

CALL “Missed approach.”

At 3,000 Ft Above Ground Level or Level Off (whichever is lower)

CALL “Climb checklist.”

ACTION Complete Climb
checklist.
CALL “Climb checklist
complete.”

Citation V For SimuFlite training only 3C-23

March 1999
 Visual Traffic Patterns
PF PNF

Before Pattern Entry/Downwind (1,500 Ft Above Ground Level)

CALL “Approach checklist.”

ACTION Complete Approach
checklist.
CALL “Approach checklist
complete.”

Downwind

CALL “Flaps – APPROACH.”

CALL “Flaps selected
APPROACH.”
When flaps indicate
APPROACH, “Flaps
indicate APPROACH.”
CALL “Gear – DOWN;
Before Landing
checklist.”
CALL “Gear selected
DOWN.”
When gear indicates
DOWN, “Gear
indicates DOWN.”
ACTION Complete Before
Landing checklist
except for full flaps and
yaw damper.

3C-24 For SimuFlite training only Citation V

March 1999
 Standard Operating Procedures

Visual Traffic Patterns (continued)

PF PNF

At 1,000 Ft Above Ground Level

CALL “1,000 ft AGL.”

CALL “Check. 1,000 ft.”

At 500 Ft Above Ground Level

CALL “500 ft AGL.”

CALL “Check. 500 ft.”

At 200 Ft Above Ground Level

CALL “200 ft AGL.”

CALL “Check. 200 ft.”

Citation V For SimuFlite training only 3C-25

March 1999
 Landing
PF PNF

At Point on Approach When PF Sights Runway or

Visual Reference (Landing Assured)
CALL “Going visual. Land.”
”Flaps – LAND.”
CALL “Flaps selected. Land.”
ACTION Pushes autopilot trim ACTION Continue with:
disconnect switch ■ Speed check

■ Vertical speed check

CALL “Autopilot/yaw damper ■ Callouts

off.” ■ Gear down

verification
■ Flap verification

CALL “Final gear and flaps

recheck.”
“Before Landing
checklist complete.”

At 100 ft Above Ground Level

CALL “100 ft AGL.”

At 50 ft Above Ground Level

CALL “50 ft AGL.”

At Touchdown

ACTION Extend speedbrakes

CALL “Speedbrakes
extended.”

At Thrust Reverser Deployment

CALL “Six lights.”

At Thrust Reverser Idle Speed (60 KIAS)

CALL “60 kts.”

3C-26 For SimuFlite training only Citation V

March 1999
This chapter presents written descriptions of various maneu-
vers and techniques applicable to normal and single engine Maneuvers
operations. The second part of this chapter contains pictorial
examples of selected maneuvers.
Chapter 3D

Citation V For SimuFlite training only 3D-1

March 1999
3D-2 For SimuFlite training only Citation V
March 1999
 Maneuvers

Two Engine Operation . . . . . . . . . . . . . . . . 3D-7

Table of
Taxiing . . . . . . . . . . . . . . . . . . . . . . . . . 3D-7
Before Takeoff . . . . . . . . . . . . . . . . . . . . . 3D-7
Contents
Takeoff (General) . . . . . . . . . . . . . . . . . . . 3D-8
Normal Standing Takeoff . . . . . . . . . . . . . . 3D-8
Rolling Takeoff . . . . . . . . . . . . . . . . . . . . 3D-8
Crosswind Takeoff . . . . . . . . . . . . . . . . . . 3D-8
Takeoff Rotation . . . . . . . . . . . . . . . . . . . . 3D-8
Rejected Takeoff . . . . . . . . . . . . . . . . . . . 3D-8
Initial Climbout . . . . . . . . . . . . . . . . . . . . 3D-8
Climb . . . . . . . . . . . . . . . . . . . . . . . . . 3D-9
Cruise . . . . . . . . . . . . . . . . . . . . . . . . . 3D-9
Thrust Setting . . . . . . . . . . . . . . . . . . . . 3D-9
Cabin Temperature . . . . . . . . . . . . . . . . . 3D-9
Turbulent Air Penetration . . . . . . . . . . . . . . 3D-10
Operation in Icing Conditions . . . . . . . . . . . 3D-10
Inflight Procedures . . . . . . . . . . . . . . . . . . 3D-10
Speedbrake Deployment . . . . . . . . . . . . . . 3D-10
Change of Airspeed . . . . . . . . . . . . . . . . 3D-11
Steep Turns . . . . . . . . . . . . . . . . . . . . 3D-11
Stall Recognition and Recovery . . . . . . . . . . . 3D-11
Approach to Stall . . . . . . . . . . . . . . . . . . 3D-11
Clean Configuration – Flaps and Gear Up . . . . . 3D-11
Approach/Departure Configuration –
15° Flaps and Gear Up or Down . . . . . . . . . 3D-12
Landing Configuration – Full Flaps, Gear Down . . . 3D-12
Unusual Attitudes . . . . . . . . . . . . . . . . . . 3D-12
Recovery from Nose-High Attitude . . . . . . . . . 3D-12
Recovery from Nose-Low Attitude . . . . . . . . . 3D-12
Instrument Procedures . . . . . . . . . . . . . . . . 3D-12
Holding . . . . . . . . . . . . . . . . . . . . . . . 3D-12

Citation V For SimuFlite training only 3D-3

March 1999
 Flight Director . . . . . . . . . . . . . . . . . . . 3D-13
Instrument Approach Considerations . . . . . . . 3D-13
Additional Instrument Systems . . . . . . . . . . . 3D-14
Normal Descent . . . . . . . . . . . . . . . . . . . 3D-14
Condensation Precautions . . . . . . . . . . . . . 3D-14
Pressurization . . . . . . . . . . . . . . . . . . . 3D-14
Anti-Icing . . . . . . . . . . . . . . . . . . . . . . 3D-14
Approach . . . . . . . . . . . . . . . . . . . . . . 3D-14
Emergency Descent . . . . . . . . . . . . . . . . . 3D-15
VFR Traffic Pattern . . . . . . . . . . . . . . . . . 3D-15
Approaches . . . . . . . . . . . . . . . . . . . . . 3D-15
Checklist and Configuration . . . . . . . . . . . . 3D-15
Typical Precision Approach (ILS) . . . . . . . . . 3D-15
Typical Non-Precision Approach . . . . . . . . . . 3D-16
Flap Inoperative Approach and Landing
(Zero Flap Approach and Landing) . . . . . . . . 3D-16
Go-Around/Missed Approach/Balked Landing . . . 3D-17
Go-Around Procedure . . . . . . . . . . . . . . . 3D-17
After a Missed Approach –
Proceeding for Another Approach . . . . . . . . 3D-17
After a Missed Approach – Departing Area . . . . 3D-17
Circling Approach . . . . . . . . . . . . . . . . . 3D-18
Landing . . . . . . . . . . . . . . . . . . . . . . . 3D-18
Thrust Reversers . . . . . . . . . . . . . . . . . . 3D-18
Crosswind . . . . . . . . . . . . . . . . . . . . . 3D-18
Touch-and-Go Landings . . . . . . . . . . . . . . 3D-19
Contaminated Runways . . . . . . . . . . . . . . 3D-19
After Landing . . . . . . . . . . . . . . . . . . . . . 3D-19

Single Engine Operation . . . . . . . . . . . . . . 3D-21

Engine Failure After V1 – Takeoff Continued . . . . 3D-21
Single Engine ILS Approach and Landing . . . . . . 3D-21

3D-4 For SimuFlite training only Citation V

March 1999
 Maneuvers

Single Engine Go-Around/Missed Approach/

Balked Landing . . . . . . . . . . . . . . . . . . . 3D-21

Flight Profiles . . . . . . . . . . . . . . . . . . . . 3D-23

Normal Takeoff . . . . . . . . . . . . . . . . . . . . 3D-25
Steep Turns . . . . . . . . . . . . . . . . . . . . . 3D-27
Approach to Stall at Pattern Altitudes
(Simulator Only) . . . . . . . . . . . . . . . . . . 3D-29
Stall Series at Altitude . . . . . . . . . . . . . . . . 3D-31
Visual Approach . . . . . . . . . . . . . . . . . . . 3D-33
Two Engine ILS Approach and Landing/
Missed Approach/Balked Landing . . . . . . . . . 3D-35
Two Engine Non-Precision Approach and Landing/
Missed Approach/Balked Landing . . . . . . . . . 3D-37
Circling Approach . . . . . . . . . . . . . . . . . . 3D-39
Engine Failure After V1/Takeoff Continued . . . . . 3D-41
Single Engine ILS Approach and Landing/
Missed Approach/Balked Landing . . . . . . . . . 3D-45
Flap Inoperative Approach and Landing
(Zero Flap Approach and Landing) . . . . . . . . 3D-47
Emergency Descent . . . . . . . . . . . . . . . . . 3D-49

Citation V For SimuFlite training only 3D-5

March 1999
3D-6 For SimuFlite training only Citation V
March 1999
 Maneuvers

Taxiing ground personnel to ensure adequate

Two Engine
clearance.
Prior to taxiing the Citation V, com-
plete all items of the Before Taxi
checklist. Obtain clearance from the
When clear of other aircraft after taxi
begins, check both pilot’s and copi-
Operation
appropriate control agency and ensure lot’s brakes as soon as possible Both
both pilots understand the taxi route pilots should maintain good look-out
prior to aircraft movement. Both pilots discipline while taxiing. Avoid tests,
should visually check the area around checks, and paperwork activity that
the aircraft for ground equipment, other compromise necessary visual clearing.
obstructions, and personnel. Taxi speed should be kept to the min-
imum practical for safety and passen- NOTE: Power settings and
Also a visual check should be made airspeeds for maneuvering are
ger comfort.
of the passenger cabin to note that bag- target values and will vary based
gage and equipment are stowed, emer- Items of the Taxi and Before Takeoff
checklists should be accomplished upon aircraft gross weight, den-
gency exit access is clear, galley equip-
when visual clearing is not compro- sity altitude, and environmental
ment and supplies are secure, and that
mised. Whenever it is necessary to stop conditions such as icing.
passengers are seated with seat belts
fastened. If necessary, a verbal or PA aircraft movement with the engine run-
announcement can be made that the ning, maintain firm pressure on the
aircraft is being taxied. brake pedals or set the parking brake.
Plan ahead — be sure that the aircraft
When ready to taxi, release the parking
and its pilots and passengers are ready
brake. The rudder pedals provide
for flight before calling for takeoff
approximately 20° left or right nose
clearance.
wheel travel for takeoff and landing.
Additional castering of the nose wheel
can be achieved by applying differen- Before Takeoff
tial power and braking. For smooth Prior to takeoff, consider the following:
steering, gently move the rudder ped- ■ use of flight director
als only the amount necessary to make
■ thrust application
the required turn.
■ brake release
When applying power to taxi, use care
and good judgment to avoid exhaust ■ runway alignment

blast to other aircraft, personnel, equip- ■ proper use of controls

ment, and buildings. Apply sufficient
■ proper rotation
power to start the aircraft rolling; check
proper operation of the wheel brakes ■ gear retraction

and then reduce power to idle. At ■ thrust reduction to climb power

lighter weights and higher elevations,
■ adherence to airport area speed
the aircraft may accelerate easily; at
idle power, it is easy to generate taxi limits
speeds much higher than desired. If it ■ icing conditions.

is necessary to make a sharp turn after The PF’s takeoff briefing, in accor-
moving from the parking spot, main- dance with SOP, should be clear, con-
tain above idle power until sufficient cise, and pertinent to the specific take-
speed is gained to complete the turn off. Set airspeed bugs according to the
with idle thrust. The additional speed SOP. Navigation aids should be tuned
prevents the aircraft from stopping dur- and identified; the specific courses
ing the turn and then requiring excess should be set. The altitude alerter
thrust to move again. If taxiing in a should be set to the proper altitude.
congested area and close to other air- When cleared for takeoff, complete all
craft, hangars, or other obstacles, use items of the Takeoff checklist.

Citation V For SimuFlite training only 3D-7

March 1999
 Takeoff (General) The PF is responsible for manipulat-
ing the control wheel and uses the con-
The primary instruments for setting ventional aileron into the wind tech-
takeoff thrust are the N1 gages. Obtain nique (i.e., applying full deflection at
the required takeoff power settings the beginning of the takeoff roll and
from the manufacturer’s AFM or from slowly decreasing the deflection as air-
the SimuFlite Citation V Operating speed increases to V1).
Handbook. The manufacturer’s AFM
and Operating Manual state that this Takeoff Rotation
power is set statically for normal take-
offs and that charted takeoff perfor- At VR, smoothly rotate to a takeoff
mance is based on such a setting. pitch attitude of approximately 10° to
12°. Smooth rotation prevents a
Normal Standing Takeoff decrease in airspeed. Early or late rota-
tion degrades takeoff performance.
Refer to profile on page 3D-25.
Hold the brakes firmly and advance Rejected Takeoff
the throttles to approximately 80% N1.
Allow the engines to spool up and sta- For an abort prior to V1, immediately
bilize at nearly equal N1 indications and simultaneously apply wheelbrakes,
before advancing the levers to the retard throttles to IDLE, extend the
desired takeoff N1. When power is set, speedbrakes, and deploy the thrust
check engine instruments and release reversers. When the thrust reversers
the brakes smoothly. are deployed, increase reverse thrust
to help slow the aircraft. Deploying
To optimize coordination, the PNF both thrust reversers for an engine fail-
should monitor the instruments and ure abort is authorized. Consider not
assist with the throttles to enable to the using the thrust reverser of an engine
PF to concentrate on directional con- on fire. Use caution with asymmetrical
trol. At 70 KIAS or greater, the PNF reverse; directional control could be a
cross checks the airspeed indications. problem.
At V1, the PF’s right hand moves to
the control wheel in preparation for Decrease reverser power to idle re-
takeoff rotation. verse by 60 kts. However, maximum
reverse thrust may be used to a full
Rolling Takeoff stop if absolutely necessary. Use
reverse thrust cautiously on wet or slip-
A rolling takeoff may be accomplished
pery runways. Use caution also dur-
when actual runway length adequate-
ing strong crosswind conditions
ly excess takeoff field length and obsta-
because reverse thrust may aggravate
cle clearance is not a factor. Once the
any weather-vaning tendency. Maintain
aircraft is aligned with the runway,
directional control with nosewheel
apply the brakes and advance the throt-
steering to remain on the runway cen-
tles to 80% N1. Release the brakes and
terline.
adjust power to the takeoff N1 setting
prior to 60 KIAS. The AFM takeoff
field length data and takeoff N1 set-
Initial Climbout
tings assume a standing start. Once the vertical speed indicator and
altimeter indicate a positive rate of
Crosswind Takeoff climb, move the landing gear lever to
When required, a crosswind takeoff may UP. Confirm gear has retracted and
be combined with any other takeoff. monitor annunciators and engine instru-
Directional and lateral control through- ments. When the airspeed increases to
out a crosswind takeoff are critical. a minimum of V2 +10 KIAS and the

3D-8 For SimuFlite training only Citation V

March 1999
 Maneuvers

aircraft reaches 400 ft AGL minimum, A thorough understanding of the DIFF

retract the flaps. PRESS/CABIN ALT gage aids the CAUTION: Do not operate
At a minimum speed of VENR, climb crew in smooth operation of the pres- deice boots when indicated OAT
power should be set. The initial set- surization system. is below -40°C (-40°F).
ting is made by reference to the climb
N1. After making the initial setting, Cruise
consult the climb N1 chart for the max-
imum allowable N1 values and adjust Thrust Setting
NOTE: Check anti-ice system
throttles accordingly. Normally, climb power is maintained for proper operation prior to
at level-off until acceleration to the entering areas in which icing
Climb desired cruise Mach, then power is might be encountered.
After setting the climb power and when adjusted to the appropriate setting.
clear of the airport traffic area, both During the climb and acceleration to
pilots complete the Climb checklist. cruise speed, the ITT should be mon-
itored.
Through the climb, the PNF compares
the indicated N1 with the climb N1 If engine RPM does not automatical-
chart. N1 RPM increases with altitude; ly synchronize at the desired cruise
several power adjustments may be nec- setting, turn the engine synchronizer
essary during climb to maintain the switch to OFF; this allows the syn-
specified setting required by the climb chronizer actuator to center. Roughly
charts. If a temperature inversion is synchronize engines with the throttles,
encountered during the climb, closely then turn the synchronizer switch to
monitor the climb N1 setting to stay FAN or TURB.
within the climb N1 limits. For maximum range, the thrust neces-
Using engine anti-ice reduces allow- sary to maintain optimum angle-of-
able fan speed and dictates close mon- attack diminishes with fuel burnoff.
itoring of ITT and RPM limits. Better performance and lower airspeed
requirements as weight decreases
Observe the differential pressure/cabin lessen the necessary thrust.
altitude and cabin vertical speed gages
for proper operation and comfort rate. Cabin Temperature
Periodic checks of time to climb
Monitor the environmental control
remaining, cabin altitude, and rate of
panel to ensure proper comfort level
cabin ascent provide required informa-
for the passengers and crew. Normally,
tion to determine necessary adjustments.
the temperature control selector is in
For example: passing 20,000 ft with a AUTO between the 11 and 1 o’clock
cabin altitude of 4,000 ft and an esti- position. During daylight, the cockpit
mated time remaining of 10 minutes may not be an accurate reference of
to 35,000 ft (8,000 ft cabin altitude) cabin comfort level due to solar heat-
requires a cabin climb rate of 400 fpm ing through cockpit windows.
to attain planned cruise and cabin alti-
Crew foot warmers are an extension
tudes concurrently. With RATE set too
from the same source as the cabin
low, maximum differential pressure
underfloor ducting; use them as a ref-
may occur before cruise altitude; this
erence for temperature adjustments.
takes system control away from the
Place a hand over an open foot warmer
crew. The outflow valve relieves pres-
outlet for an approximate indication
sure as necessary to maintain maxi-
of the temperature of the airflow into
mum differential. A RATE setting too
the cabin.
high may be uncomfortable and results
in the programmed cabin altitude
occurring before the cruise altitude.

Citation V For SimuFlite training only 3D-9

March 1999
 Turbulent Air Penetration on the wing inspection lights illumi-
CAUTION: Do not operate nates the wing leading edges for ice
Although the aircraft is not opera- detection during night operations.
deice boots when indicated OAT
tionally restricted in rough air, flight
is below -40°C (-40°F). All anti-ice/deice systems must be
through severe turbulence should be
avoided if possible. checked and found operational prior
to flights into known icing. Engine
Carefully plan turbulence avoidance anti-ice should be used on the ground
strategy with an understanding of or in the air when the indicated air tem-
mountain wave dynamics, thunder- perature (OAT) is between +10°C and
storm characteristics, and weight ver- -30°C and visible moisture is present.
sus altitude buffet margins. If severe
turbulence is encountered, the follow- Windshield anti-ice should be operat-
ing steps are recommended. ed in LO when OAT is above -18°C
or in HI when OAT is -18°C or below.
1. Maintain airspeed at approximately The deice boots should be activated
180 KIAS. Do not chase airspeed. when ice accumulation on the wings
2. Turn on the ignition system, and is estimated to be to inch thick. An
set thrust to maintain target airspeed. 18 second cycle actuates when the
Change thrust only for extreme air- AUTO/OFF/MANUAL switch is
speed variation. momentarily placed in AUTO. The
3. With the autopilot not engaged, lower deice boots of both wings and
keep control movements moderate and the left horizontal stabilizer boot inflate
smooth. Maintain wings level and for six seconds. The system rests for
desired pitch attitude. Use the attitude six seconds and then the upper wing
indicator as the primary instrument. In boots and right horizontal stabilizer
extreme drafts, large attitude changes boots inflate for the last six seconds.
may occur. Do not make sudden, large The manual position of the switch
control movements. After establishing inflates all boots for as long as the
trim setting for penetration speed, do switch is held in MANUAL.
not change the stabilizer trim. In icing conditions, turn engine anti-
4. Large altitude changes are possible ice switches on and off one at a time;
in severe turbulence. Allow the alti- pause momentarily between moving
tude to vary to maintain the desired each switch. If ice accumulations break
attitude and airspeed. Do not chase away and are ingested by the engines,
altitude or airspeed. pausing reduces the risk of a dual
flameout occurring.
5. Ensure the yaw damper is engaged
to reduce yaw/roll oscillations. If anti-ice is required during takeoff,
turn the anti-ice system on prior to set-
6. If turbulence is penetrated with the ting takeoff power. For proper anti-ice
autopilot on, engage the Soft Ride operation and engine protection, ensure
mode. Turn off the Altitude, Speed, or adherence to the maximum anti-ice N1
Vertical Speed Hold mode. power settings for takeoff, climb, and
7. Turn on the FASTEN SEAT BELT cruise. When power is reduced (e.g.,
sign. for descent), monitor the anti-ice fail
lights to ensure adequate power is
Operation in Icing maintained.
Conditions
The engine and windshield anti-ice Inflight Procedures
systems prevent the accumulation of Speedbrake Deployment
icing; they should be turned on prior to
encountering such conditions. Turning Speedbrakes may be used to expedite
a descent or reduce airspeed. A light

3D-10 For SimuFlite training only Citation V

March 1999
 Maneuvers

buffeting is noticeable with speed- stall (i.e., airframe buffet or stick shak-
brakes extended. er, whichever occurs first). CAUTION: The adjacent dis-
Speedbrakes may be used at any speed The stick shaker on the forward side of cussion is presented only in the
and with flaps in any position, but they the pilot’s control column provides tac- context of recovery training.
must be retracted before landing. There tile warning (i.e., vibration) of impend- Stalls in high performance air-
is no data regarding the effect of speed- ing stall. The angle-of-attack transmit- craft should not be deliberately
brakes on stall speed or VREF. ter activates the stick shaker. At the executed unless they are part of
beginning of the lower speed buffet a supervised pilot training pro-
Change of Airspeed (approximately 8 to 10% above the gram. Safety of flight considera-
Speedbrakes may be used in conjunc- actual stall speed), the stick shaker acti- tions dictate that the utmost cau-
tion with thrust reduction when reduc- vates. If the AOA system is inopera- tion be employed during such
ing airspeed quickly. Reduce thrust to tive, the stick shaker does not operate.
exercises.
the appropriate setting for the desired At the first warning indication, initi-
airspeed, then extend the speedbrakes. ate an immediate recovery. Do not
Upon reaching desired airspeed, retract allow the aircraft to go into full stall.
the speedbrakes. Perform the approach to stall in the
Smoothly coordinate all power and clean, takeoff, and landing configura-
flight control inputs to maintain desired tions. Practice altitude should be no
heading, airspeed, and altitude. Speed- higher than 25,000 ft MSL (AFM lim-
brakes may also be used to control air- itation), and no lower than 10,000 ft
speed during inflight operation of the above terrain.
engine anti-ice when higher-than-nor- Before practicing approaches to stall,
mal engine power settings are required. clear the cockpit area of loose articles;
Steep Turns visually clear the practice area, com-
pute VREF, and set airspeed indicator
Refer to the profile on page 3D-27. bugs. Turn on the ignition system as a
Steep turns (e.g., 45° bank) confirm precaution.
the aerodynamic principle that increas-
ing bank requires increased pitch and Clean Configuration –
power to maintain altitude. Flaps and Gear Up
At intermediate altitudes (e.g., 10,000 While maintaining altitude and heading
ft MSL), practice steep turns at 200 kts. (wings level), retard throttles to idle.
As the aircraft slows, maintain altitude
The initial engine power setting is
with back pressure. Use trim to reduce
about 72% N1. When passing through
30° bank, increase power setting 2 to stick forces. Stop trimming at approx-
4% N1 and pitch attitude approxi- imately 0.6 AOA, and use elevator con-
mately 1°. Trim out back pressure as trol from 0.6 AOA to stall warning. Do
needed. Lead the rollout heading not trim all the way to stall.
approximately 10°, and reduce thrust Increase power to 50% N1 at VREF.
and pitch to the original setting. Observe the angle-of-attack indicator
information; stick shaker actuation
Stall Recognition occurs when the indicator approaches
the red band. At the first evidence of a
and Recovery stall (i.e., airframe buffet, or stick shak-
Refer to the profiles on pages 3D-29 er, whichever occurs first):
and 3D-31. 1. Advance the throttles to maximum
power.
Approach to Stall
2. Maintain pitch attitude (i.e., approx-
The approach to stall should be con- imately 10° nose-up) and wings level.
tinued only to the first evidence of a

Citation V For SimuFlite training only 3D-11

March 1999
 3. As airspeed increases, reduce the 3. After a positive rate of climb is
pitch attitude to maintain altitude. achieved, retract landing gear.
4. Reduce power to maintain desired 4. Accelerate to VREF + 10 KIAS and
airspeed. select flaps up.
Approach/Departure 5. Reduce power as necessary to
maintain desired airspeed and altitude.
Configuration – 15° Flaps
and Gear Up or Down Unusual Attitudes
Establish a level turn using 15° to 30° A number of causes (e.g., jet upset,
bank; retard throttles to idle. As the air- failed attitude references, autopilot
craft slows, maintain altitude with back malfunction, pilot incapacitation) may
pressure. Use trim to reduce stick forces; result in unusual attitudes. Prior to exe-
however, stop trimming at 0.6 AOA. cuting the proper recovery, confirm
Increase power to 50% N1 at VREF. the unusual attitude by cross checking
Observe angle-of-attack indicator the attitude indicators, altimeters, ver-
information; stick shaker actuation tical speed, and airspeed indicators.
occurs when the indicator approaches
the red band. At the first evidence of a Recovery from
stall (i.e., airframe buffet, or stick shak- Nose-High Attitude
er, whichever occurs first): After confirming a nose-high attitude,
1. Advance the throttles to maximum low-airspeed condition exists, apply
power while rolling the wings level. thrust while rolling toward the nearest
2. Maintain pitch attitude (i.e., approx- horizon. Use up to 90° bank, depend-
imately 10° nose-up) and wings level. ing on severity of the condition. When
the nose passes through the horizon,
3. Confirm the flaps are at T.O &
smoothly roll to a wings-level attitude
APPR.
and recover to level flight.
4. As airspeed increases, reduce the
pitch attitude to maintain altitude. Recovery from
5. After reaching VREF + 10, select Nose-Low Attitude
flaps up. After confirming a nose-low attitude
6. Reduce power to maintain desired with airspeed increasing, reduce thrust
airspeed. to idle while simultaneously rolling to
a wings-level attitude. Increase pitch
Landing Configuration – attitude to recover to level or climb-
Full Flaps, Gear Down ing flight. Use speedbrakes, if neces-
sary, to minimize airspeed increase and
While maintaining altitude and head-
altitude loss. Use caution to avoid
ing (wings level), retard throttles to
exceeding G-limits during recovery.
50% N1. At the first evidence of a stall
(i.e., airframe buffet, or stick shaker,
whichever occurs first): Instrument
1. Advance the throttles to maximum Procedures
power. Lower the nose to the horizon Holding
or slightly below.
The maximum holding speeds are:
2. As the aircraft begins to accelerate
■ 14,000 ft MSL and below 230 KIAS
with no buffet, select flaps to T.O. &
APPR and increase the pitch smooth- ■ Above 14,000 MSL 265 KIAS.

ly to 10° to 12° nose-up.

3D-12 For SimuFlite training only Citation V

March 1999
 Maneuvers

If endurance is a factor, determine the Instrument Approach

recommended holding speed for the Considerations
existing flight weight from the Holding
Chart in the SimuFlite Citation V Several factors should be considered
Operating Handbook or the manufac- prior to commencing an approach in
turer’s Performance Manual. a high performance jet aircraft. The
pilot must have a thorough knowledge
Slow to holding speed within three of the destination and alternate weath-
minutes of reaching the holding fix. er conditions before descending out of
Holding pattern recommended entries the high altitude structure. Many
are parallel, teardrop, and direct. weather and traffic advisory sources
Outbound timing begins over or abeam are available, including:
the holding fix, whichever occurs later. ■ Flight Service Stations that may be
If the abeam position cannot be deter- used enroute at any time to obtain the
mined, start timing when the turn to latest destination and alternate weath-
outboard is completed. er conditions
The initial outboard leg is flown for ■ Destination Tower and/or Approach
one or one and one-half minute(s) as Control
appropriate for altitude.
■ ARTCC where controllers can obtain
Inbound leg time at 14,000 ft MSL or information (if requested) pertaining to
below is one minute. Above 14,000 ft traffic delays and whether aircraft are
MSL, the inbound leg time is one and successfully completing approaches
one-half minutes.
■ ATIS.
Timing of subsequent outbound legs
should be adjusted as necessary to If weather is at or near minimums for
achieve proper inbound leg time. For the approaches available, review the
a crosswind correction, double inbound time and fuel requirements to an alter-
drift correction on the outbound leg. nate. To continue the approach to a
landing after arrival at minimums, FAR
Flight Director 91.175 requires that:
The fight director is effective for mak- (c) Operation below DH or MDA.
ing an accurate approach in adverse Where a DH or MDA is applicable, no
weather conditions. If command bars pilot may operate an aircraft, except
are followed precisely, the fight direc- a military aircraft of the United States,
tor computes drift corrections based at any airport below the authorized
on track results. These computations MDA or continue an approach below
command slow and deliberate correc- the authorized DH unless –
tions toward interception of track and (1) The aircraft is continuously in a
glideslope. position from which a descent to a
While following the flight director landing on the intended runway can
commands, remember to cross check be made at a normal rate of descent
the raw data presentations. The flight using normal maneuvers, and for oper-
director is extremely reliable, but the ations conducted under part 121 or
command bar(s) displays computed part 135 unless that descent rate will
(i.e., trend) information only. allow touchdown to occur within the
touchdown zone of the runway of the
Monitor warning lights for indication
intended landing;
of a malfunction. If the computer is
not working properly, erroneous infor- (2) The flight visibility is not less than
mation may be presented. the visibility prescribed in the standard
instrument approach being used; and

Citation V For SimuFlite training only 3D-13

March 1999
 (3) Except for a Category II or Normal Descent
Category III approach where any nec-
essary visual reference requirements Condensation
are specified by the Administrator, at Precautions
least one of the following visual refer- The defog fan should be in HI with the
ence for the intended runway is dis- foot warmers closed up to 15 minutes
tinctly visible and identifiable to the prior to descent. Closing the foot
pilot: warmers circulates dry, conditioned
(i) The approach light system, except air between the side windows to inhib-
that the pilot may not descend below it condensation in these areas.
100 ft above the touchdown zone ele- The HI position of the fan is activated
vation using the approach lights as a prior to initiating descent and if rapid-
reference unless the red terminating ly descending into an area of antici-
bars or the red side row bars are also pated high humidity after cold soak-
distinctly visible and identifiable. ing at altitude. Use windshield bleed
(ii) The threshold. air to externally warm the windshield
(iii) The threshold markings. in extreme conditions.
(iv) The threshold lights. Airflow distribution should be biased
toward the cockpit for maximum defog
(v) The runway end identifier lights.
capabilities.
(vi) The visual approach slope indi-
cator. Pressurization
(vii) The touchdown zone or touch- After initiating descent, set the desti-
down zone markings. nation pressure altitude plus 200 ft in
(viii) The touchdown zone lights. the CABIN dial of the controller; this
ensures cabin depressurization prior
(ix) The runway or runway markings. to touchdown. Adjust the cabin rate
(x) The runway lights. for a 300 to 500 fpm descent.
(d) Landing. No pilot operating an Continue to monitor the differential
aircraft, except a military aircraft of pressure, cabin altitude, and cabin ver-
the United States, may land that air- tical speed gages throughout descent.
craft when the flight visibility is less The most comfortable condition occurs
than the visibility prescribed in the when cabin descent is distributed over
standard instrument approach proce- the majority of the aircraft descent time.
dure being used.
Anti-Icing
Additional Instrument
All anti-ice systems should be on when
Systems operating in visible moisture if the indi-
The following additional equipment is cated outside air temperature is +10°C
available on most aircraft and should to -30°C.
be set according to company SOP:
■ radio altimeter
Approach
■ terrain advisory voice encoding
Double-check landing field informa-
altimeter tion and estimated arrival gross weight;
check runway requirements, determine
■ vertical navigation computer con-
VREF, and set airspeed bugs in accor-
troller (VNAV) dance with the SOP. When descend-
■ long-range navigation equipment. ing through the transition altitude, set
the altimeters to field pressure and
check for agreement.

3D-14 For SimuFlite training only Citation V

March 1999
 Maneuvers

The radar altimeter may be bugged to is VREF +20 +wind factor. Lower the
either the maximum descent altitude gear opposite the touchdown zone or NOTE: Wind factor is half the
or the decision altitude, or it may be about two miles out on base leg or steady wind plus all of the gusts;
bugged as desired in VFR operation straight-in final (but not lower than not to exceed 10 kts.
for terrain proximity warning. traffic pattern altitude).
Complete the Before Landing check-
Emergency list to Flaps – LANDING. Maintain
Descent VREF +20 +wind factor until the turn is
completed on final or descent is start-
Refer to the profile on page 3D-49. ed, if straight-in. Lower flaps to LAND
An emergency descent moves the air- and maintain VREF +10 +wind factor
craft rapidly from a high altitude to a until transition to VREF +wind factor
lower altitude; it is most used in con- over threshold. Disengage yaw damper
junction with a loss of pressurization. prior to landing.
Put on oxygen masks, establish com- The angle-of-attack indicator supple-
munications, disconnect autopilot, ments the airspeed indicator during all
retard throttles to idle, extend speed- maneuvers; include it in the instrument
brakes, and roll into a moderate bank crosscheck. On final approach, use it to
while lowering the nose (initially 25°) confirm the accuracy of VREF compu-
below the horizon. Adjust pitch as nec- tation.
essary to approach, but not to exceed,
VMO/MMO. Approaches
If flying in turbulent air or if structural Checklist and
integrity is questionable, make the Configuration
descent at a lesser and more prudent
speed. The PNF should set the transpon- For instrument approaches where a
der to 7700. procedure turn is flown, initiate the
Before Landing checklist after turning
When conditions permit, the engine
outbound from the approach fix. Flaps
ignition system may be placed on, the
should be lowered to T.O. & APPR
FASTEN SEAT BELT sign and NO
and the airspeed maintained at a min-
SMOKING sign illuminated, oxygen
imum of VREF +20. The target power
availability to passengers checked, and
setting is 60% N1.
ATC contacted for assistance and
instructions. The PNF should monitor If the aircraft is receiving radar vec-
the descent progress, establish the min- tors for an approach, initiate the Before
imum altitude for terrain avoidance, and Landing checklist and aircraft config-
complete the checklists on command. uration changes when abeam the FAF
outbound, or three to five miles before
VFR Traffic Pattern the FAF for a straight-in approach.
Refer to the profile on page 3D-33. At uncontrolled airports, make all
required position/intention reports on
Traffic pattern altitude for jets nor- the appropriate Common Traffic
mally is 1,500 ft AGL. In clean con- Advisory Frequency (CTAF).
figuration, slow to a minimum of VREF
+30. The target power setting is Typical Precision
approximately 60% N1. Approach (ILS)
Initiate the Before Landing checklist Refer to the profile on page 3D-35.
no later than the downwind leg entry
point; lower flaps to T.O. & APPR. An ILS approach is normal when both
The minimum airspeed on downwind engines, the appropriate ILS facilities,

Citation V For SimuFlite training only 3D-15

March 1999
 and airborne equipment are operating 6. With the runway environment in
normally. Accomplish the following. sight, set full flaps and disengage the
1. When established on the localizer autopilot/yaw damper to complete the
inbound to the FAF, ensure flaps are Before Landing checklist. Maintain
set flaps at T.O. & APPR. VAP while intercepting the proper visu-
al glidepath for landing. Cross the land-
2. Maintain airspeed at VREF +20 ing threshold at VREF +wind factor.
+wind factor and initiate the Before
Landing checklist. Flap Inoperative
3. When glideslope indicates one dot Approach and Landing
prior to intercept, lower landing gear. (Zero Flap Approach and
4. At glideslope intercept, start tim- Landing)
ing, begin descent, and extend full Refer to the profile on page 3D-47.
flaps. Complete the Before Landing
checklist to the autopilot/yaw damper. The Citation V has only one approved
flap position for landing (full flaps).
5. Maintain airspeed at VREF +10 Landing with other than full flaps is an
+wind factor (VAP). abnormal condition. A flap inoperative
6. At or before DA, establish visual approach and landing begins with ini-
contact with the runway. tiating the Flap Inoperative Approach
7. Reduce power slightly to ensure and Landing checklist in the SimuFlite
crossing the runway threshold at VREF Citation V Operating Handbook or the
+wind factor. Disengage the autopi- AFM.
lot/yaw damper to complete the Before Two major adjustments to a normal
Landing checklist. approach are the increased landing dis-
tance multiplying factor and the adjust-
Typical Non-Precision ed VREF for the flap setting.
Approach If the flap position indicator does not
Refer to the profile on page 3D-37. agree with the position of the flap han-
When established on the inbound course dle, move the handle to agree with the
to the FAF, perform the following. indicator. Check that the FLAP CON-
TROL CB is set.
1. Set flaps to T.O. & APPR and ini-
tiate the Before Landing checklist. Perform an approach with the flaps
jammed at 15° while maintaining an
2. Adjust airspeed to VREF +20 +wind airspeed of VREF +5; for flaps 7°,
factor; the power setting should be maintain VREF +10, and for flaps 0°
about 60% N1. or unknown, perform the approach
3. Extend landing gear and complete while maintaining VREF +15.
the Before Landing checklist down to During the initial phase of the
full flaps and autopilot/yaw damper approach, maintain a minimum speed
before FAF. of original VREF +30 KIAS (VZF)
4. Upon crossing FAF, start timing, with target power at approximately
notify ATC, and descend to MDA 60% N1 until the flap inoperative
while maintaining airspeed at VAP adjusted VREF is established.
(VREF +10 +wind factor) with power at Speedbrakes may be used during the
about 50% N1. Vertical speed in the approach. Lower the landing gear early
descent should normally be 1,000 to in the approach to help control airspeed.
1,500 fpm.
Once established on final, reduce speed
5. After leveling off at MDA, increase to adjusted VREF +wind factor.
power to hold airspeed at VAP while Depending on flap setting, the power
proceeding to the VDP or MAP.

3D-16 For SimuFlite training only Citation V

March 1999
 Maneuvers

setting may be as low as idle during erate to VENR minimum. Adjust pitch
the descent to maintain the approach attitude and power as necessary.
speed. Ensure that the autopilot/yaw 5. Reduce power to climb N1. At the
damper is off and that the speedbrakes relatively light gross weight at which
are retracted to complete the checklist. missed approaches are normally
During landing, the aircraft has a ten- accomplished, the aircraft accelerates
dency to float in ground effect due to quickly. Pitch and power need to be
the increased airspeed and low drag adjusted accordingly.
configuration. Once touchdown occurs, 6. Set the flight director as required.
use speedbrakes and thrust reversers Use the heading bug and the Heading
as normal to slow the aircraft. Use a mode to fly a desired heading, and a
slight forward pressure on the control navigation mode and the course selec-
wheel when deploying the thrust tor to capture a desired radial/track.
reversers because the nose may pitch After the initial fixed (i.e., 10° nose-
up when the thrust reversers are up) climb attitude is established, vari-
deployed at the higher speed. able climb attitudes may be com-
Go-Around/Missed manded with the touch control steering
button on the control wheel. Use one of
Approach/Balked Landing the vertical modes to capture and main-
Refer to the profiles on pages 3D-35 tain desired climbs or altitudes.
and 3D-37. 7. Confirm the level-off altitude and
Accomplish the go-around/missed heading/course needed for the missed
approach/balked landing at the DA or approach. Comply with the published
MDA with time expired (if applica- missed approach instructions unless
ble) and runway visual reference either other directions are received from
not in sight or not in a position from ATC.
which a normal visual landing
approach can be accomplished. After a Missed Approach –
An approach with a visual descent
Proceeding for Another
point (VDP) positions the aircraft for Approach
a normal glideslope to landing. When Accomplish the following.
an aircraft proceeds beyond the VDP 1. After level-off, complete the Climb
without visual reference to the runway, checklist and maintain VREF +30 KIAS
the probability of a missed approach minimum.
is increased.
2. Review TOLD cards and bugs for
Go-Around Procedure next approach. Brief the approach and
complete the Approach checklist.
Accomplish the following.
Maintain a minimum of VREF +30
1. Apply go-around power. KIAS until established on a portion
2. Push the go-around button; rotate of the next approach and flaps are
to the flight director go-around atti- extended.
tude (i.e, approximately 10° nose-up).
After a Missed Approach –
3. With airspeed at a minimum of Departing Area
VAC, set flaps to T.O. & APPR. Retract
gear at indication of a positive rate of Accomplish the following.
climb on both altimeter and VSI. 1. Accelerate to normal climb speed.
4. When clear of obstacles and at a 2. Complete the Climb checklist.
minimum airspeed of VREF +10 KIAS,
3. Follow normal climbout procedures.
retract the remaining flaps and accel-

Citation V For SimuFlite training only 3D-17

March 1999
 Circling Approach Reduce thrust slowly to idle and raise
NOTE: The maximum the nose slightly from the attitude
demonstrated crosswind for the Refer to the profile on page 3D-39. maintained on final approach. With aft
Citation V is 20 knots. A circling approach is an instrument mounted engines, the nose tends to rise
approach requiring a heading change as thrust is reduced and thus requires
of 30° or more to align the aircraft with little back pressure.
the landing runway. Once visual con- Maintain attitude and allow the aircraft
ditions are reached, the circling to fly onto the runway surface.
approach is a modified version of the
VFR traffic pattern. Upon touchdown, lower the nosewheel
smoothly to the runway, extend the
Turbulence, strong winds, poor visi- speedbrakes, and apply braking as nec-
bility, and low maneuvering altitude essary. Monitor anti-skid for proper
are factors that must be considered functioning. To achieve maximum
when planning a circling approach. benefit from the anti-skid system, do
Plan to use a minimum circling alti- not pump the brakes; instead, apply
tude and distance appropriate to the steady pressure on the pedals. Use
airspeed or approach category. The nosewheel steering and differential
Citation V normally falls into Category braking to maintain directional con-
B for straight in approaches. Due to trol, and deploy the thrust reversers.
the probability of speeds up to 140
KIAS during the approach, it is rec- Thrust Reversers
ommended that circling be accom-
plished with Category C minimums. When thrust reversers are used, extend
Speeds higher than 140 KIAS require speedbrakes upon touchdown and
Category D minimums. lower nosewheel to the runway. Pull
the thrust reverser levers to the
At uncontrolled airports, observe local REVERSE IDLE/DEPLOY position.
traffic direction and restrictions. Check that the ARM, UNLOCK, and
It is recommended that the approach DEPLOY lights illuminate before
be flown with gear down and flaps at smoothly pulling the thrust reverser
T.O. & APPR until arriving at a posi- levers aft to the desired reverse thrust.
tion from which a normal descent for If reverse thrust is not needed, con-
landing can be made. At that time, sider leaving the reverser levers in the
begin descent and select full flaps. idle thrust setting for FOD protection
While maneuvering during the circling and noise abatement.
approach, fly a minimum of VREF +20
+wind factor. When established on Crosswind
final in the landing configuration, fly On the final approach in a crosswind,
at VREF +10 +wind factor until reduc- either the crab approach or the wing-
ing power slightly to cross the runway down method may be used.
threshold at VREF +wind factor.
Do not allow the aircraft to float with
power off prior to touchdown. Fly to
Landing touchdown with little, if any, flare.
Refer to the profiles on pages 3D-35 Deploy speedbrakes on touchdown.
and 3D-37. Follow through the landing roll with
With FULL flaps, cross the threshold ailerons into the wind. The aileron rud-
at 50 ft AGL with a speed of VREF der interconnect can be overridden
+wind factor. with input forces. Use nosewheel steer-
ing and differential braking for direc-
tional control.

3D-18 For SimuFlite training only Citation V

March 1999
 Maneuvers

Touch-and-Go Landings planing speed before using the wheel-

brakes. Hydroplaning speed (Vh), based
If touch-and-go landings are to be prac- on NASA test data, is:
ticed, they should be preplanned and
■ Takeoff
briefed. The thrust reversers and speed-
brakes should not be used on landing. Vh = 9 x √tire pressure
The ground idle switch should be in ■ Landing
HIGH. The PNF should reset the flaps
to T.O. & APPR, set the stabilizer trim Vh = 7.7 x √tire pressure
in the takeoff range, and confirm these The difference in hydroplaning speed
settings with the PF before throttles between takeoff and landing is due to
are advanced to takeoff power. the wheels rolling for takeoff and not
rolling prior to landing.
Contaminated Runways
Normal nose gear tire pressure is 120
Landing on a slick surface requires ±5 PSI and main gear tire pressure is
careful consideration of many factors: 130 ±5 PSI. Assuming the pressures
type of runway surface, approach haz- are 130 and 120 PSI, the following
ards, aircraft weight/speed, wind con- speeds are applicable: for takeoff, the
ditions, temperature, ice, water, and main wheel Vh is 103 KIAS and the
snow. Correction factors can be found nose wheel Vh is 99 KIAS.
in the SimuFlite Citation V Operating For landing, this speed range is
Handbook or in the AFM. Do not rely decreased to 88 KIAS for the main
on thrust reversers to ensure a reduced tires and 85 KIAS for the nose wheel.
stopping distance.
If braking action is started while the
Exercise caution when using differen- tires are hydroplaning, the condition
tial reversing on a slick runway. Be can continue to a much slower speed.
prepared to return to REVERSE/ IDLE
immediately if the aircraft starts to
slide sideways. Strong consideration
After Landing
should be given for only REVERSE/ After clearing the runway, complete
IDLE on slippery surfaces. the After Landing checklist. The
Do not attempt single engine thrust engines should be operated at idle for
reverse on a slick runway. at least one minute prior to shutdown;
taxi time may be included. After the
If there is a possibility of hydroplaning aircraft is parked, complete the Shut-
on surface water, slow below hydro- down checklist.

Citation V For SimuFlite training only 3D-19

March 1999
3D-20 For SimuFlite training only Citation V
March 1999
 Maneuvers

Engine Failure higher than comparable two engine

Single Engine
After V1 – settings.

Takeoff Continued Extend full flaps when runway prox-

imity does not require additional
Operation
Refer to the profile on page 3D-41. engine thrust to overcome flap drag to
With an engine fire or failure indication arrive at the normal touchdown point.
after V1, continue the takeoff. Once the flaps are full down, retard
the operating engine throttle to idle as
Maintain directional control using the the aircraft crosses the runway thresh-
rudder/nosewheel steering, and accel- old at VREF +wind factor.
erate to VR. At VR, rotate the aircraft
to 10° nose-up and climb at V2. If the If rudder trim is used during approach
indication occurs after exceeding V2, to counter asymmetric thrust, zero the
maintain the existing airspeed. Retract rudder trim prior to or during the land-
the landing gear when a positive rate of ing power reduction to prevent unwant-
climb is established. ed yaw. Thrust reduction and flare are
similar to a normal landing. Thrust
When clear of obstacles and at a min- reduction should be slower than nor-
imum of 400 ft AGL, accelerate to V2 mal to counter roll due to yaw effect.
+10 KIAS and retract the flaps. Consequently, slightly less flare than
Increase speed to VENR and continue normal is required to prevent floating.
the climb to the required altitude above
the takeoff field elevation. After touchdown, lower the nose,
extend the speedbrakes, apply wheel
In case of engine fire, consideration braking, and keep the wings level. Use
must be given to fighting the fire once rudder and differential braking.
the flaps are retracted and the aircraft Reverse thrust may be used on the
is accelerating toward VENR. operating engine.
When time and conditions permit,
complete the Engine Failure/Fire Single Engine
checklist and the Climb checklist.
Go-Around/Missed
Single Engine ILS Approach/Balked
Approach and Landing
Landing Refer to the profiles on pages 3D-43
and 3D-45.
Refer to the profile on page 3D-43.
Apply takeoff power on the operating
A single engine inoperative approach
engine and push the flight director go-
is flown essentially the same as an
around button to select the Go-Around
approach with both engines operating.
mode. Rotate to approximately 10°
On final approach, however, do not
nose-up as commanded by the flight
lower full flaps until landing is assured.
director and retract the flaps to T.O &
Up to the final descent point, the air- APPR; normally, the flaps are not
craft is configured normally with the extended beyond the approach setting
previously recommended speeds flown until landing is assured. As thrust is
for each configuration. A single engine increased, apply rudder pressure as
thrust setting is approximately 10% required to counter yaw.

Citation V For SimuFlite training only 3D-21

March 1999
 Maintain the go-around pitch attitude When time permits, the PNF sets the
and minimum airspeed of VAC. Retract PF’s heading bug on the missed
the landing gear when a positive rate of approach heading and selects the
climb is established. Climb to 400 ft requested modes on the flight direc-
AGL (minimum), then retract flaps at tor. At appropriate time, advise ATC
VREF +10 KIAS (minimum). Continue of missed approach and request fur-
accelerating to VENR, set climb power, ther clearance (e.g., another approach,
and continue the climb on the pub- diversion to the alternate airport).
lished missed approach.

3D-22 For SimuFlite training only Citation V

March 1999
 Maneuvers

The following flight profiles illustrate how selected maneuvers are

performed. Each maneuver is broken down into sequential events Flight Profiles
that illustrate appropriate configurations.
■ Normal Takeoff
■ Steep Turns
■ Approach to Stall at Pattern Altitudes (Simulator Only)
■ Stall Series at Altitude
■ Visual Approach
■ Two Engine ILS Approach and Landing/Missed Approach/
Balked Landing
■ Two Engine Non-Precision Approach and Landing/
Missed Approach/Balked Landing
■ Circling Approach
■ Engine Failure after V1/Takeoff Continued
■ Single Engine ILS Approach and Landing/Missed Approach/
Balked Landing
■ Single Engine Non-Precision Approach and Landing/
Missed Approach/Balked Landing
■ Flap Inoperative Approach and Landing (Zero Flap Approach and
Landing)
■ Emergency Descent.

Citation V For SimuFlite training only 3D-23

March 1999
3D-24 For SimuFlite training only Citation V
March 1999
 Maneuvers

Normal Takeoff

1 NON-ROLLING TAKEOFF
BRAKES – HOLD 7 AT V2 + 10 MINIMUM AND 400 FT AGL
TAKEOFF POWER – SET FLAPS – RETRACT
BRAKES – RELEASE CLIMB POWER – SET
TRANSITION TO ENROUTE CLIMB
3 70 KT CROSSCHECK

6 PITCH – MAINTAIN
SPEED – ALLOW TO INCREASE
TO CLIMB SCHEDULE

4 AT VR
AIRCRAFT – ROTATE TO 10°-12°

2 ROLLING TAKEOFF
TAKEOFF POWER – SET BY 60 KIAS

5 AT POSITIVE RATE OF CLIMB

GEAR – UP

Citation V For SimuFlite training only 3D-25

March 1999
 Maneuvers

Normal Takeoff

3D-25 3D-26 For SimuFlite training only Citation V

March 1999
 Maneuvers

Steep Turns

2 SMOOTHLY ROLL INTO 45° BANK

BACK PRESSURE – INCREASE TO MAINTAIN ALTITUDE
POWER – INCREASE TO MAINTAIN 200 KIAS

1 CLEAN CONFIGURATION
SPEED – 200 KIAS
WINGS – LEVEL AT ASSIGNED
ALTITUDE AND HEADING

4 LEAD ROLLOUT TO ASSIGNED HEADING BY

APPROXIMATELY 10°
WINGS – SMOOTHLY ROLL LEVEL
BACK PRESSURE – EASE TO MAINTAIN ALTITUDE
POWER – REDUCE TO MAINTAIN 200 KIAS

THIS MANEUVER MAY BE USED FOR A 180° OR 360° TURN,

AND MAY BE FOLLOWED BY A REVERSAL IN THE OPPOSITE DIRECTION

THE PNF MAY ASSIST AS DIRECTED BY THE PF.

TOLERANCES ARE:
SPEED ± 10 KIAS
ALTITUDE ± 100 FT
BANK ± 5°
ROLLOUT ON HEADING ± 10°

3 MAINTAIN
ALTITUDE – ± 100 FT
SPEED – 200 KIAS ± 10 KTS
ATTITUDE – 45° ± 5° BANK

Citation V For SimuFlite training only 3D-27

March 1999
 Maneuvers

Steep Turns

3D-27 3D-28 For SimuFlite training only Citation V

March 1999
 Maneuvers

Approach to Stall at Pattern Altitudes (Simulator Only)

2 APPROACH CONFIGURATION – BASE LEG 15° TO 30° BANK,
FLAPS - T.O. & APPR, GEAR – UP OR DOWN
1 CLEAN CONFIGURATION – DOWNWIND LEG
THROTTLES – IDLE
THROTTLES – IDLE
BANK ANGLE – 15° TO 30°
INCREASE POWER TO 50% V1 AT VREF
INCREASE POWER TO 50% V1 AT VREF
MAINTAIN PATTERN ALTITUDE AND
MAINTAIN PATTERN ALTITUDE AND SLOW
SLOW TO FIRST INDICATION OF STALL
TO FIRST INDICATION OF STALL
SIMULTANEOUSLY
SIMULTANEOUSLY
- MAXIMUM POWER – APPLY
- MAXIMUM POWER – APPLY
- PITCH ATTITUDE – MAINTAIN 10° UP
- WING – ROLL LEVEL
- FLAPS APPROACH – CONFIRM
- PITCH ATTITUDE – MAINTAIN 10° UP

1 3 LANDING CONFIGURATION – FINAL APPROACH (FULL FLAPS, GEAR DOWN)

THROTTLES – 50% N1
MAINTAIN ILS GLIDESLOPE
SPEED – SLOW TO FIRST INDICATION OF STALL
MAXIMUM POWER – APPLY
PITCH ATTITUDE – LOWER TO HORIZON OR SLIGHTLY BELOW
AS AIRSPEED INCREASES, FLAPS – T.O. & APPR
AIRCRAFT – ROTATE TO 10° TO 12° UP
AT POSITIVE RATE OF CLIMB
- GEAR – UP
- SPEED – VREF + 10 MINIMUM
- FLAPS – UP
NOTE: Approaches to stall are MISSED APPROACH PROCEDURE – PERFORM
performed at pattern altitudes
only in the simulator.

CAUTION: Maximum power

should be used in those stalls
occurring near the ground and in
emergency situations. Stalls prac-
ticed at altitude never require
exceeding engine limitations.

Citation V For SimuFlite training only 3D-29

March 1999
 Maneuvers

Approach to Stall at
Pattern Altitudes
(Simulator Only)

3D-29 3D-30 For SimuFlite training only Citation V

March 1999
 Maneuvers

Stall Series at Atltiude

2 APPROACH CONFIGURATION – 15° TO 30° BANK,
FLAPS - T.O. & APPR, GEAR – UP OR DOWN
1 CLEAN CONFIGURATION
THROTTLES – IDLE
THROTTLES – IDLE
BANK ANGLE – 15° TO 30°
INCREASE POWER TO 50% V1 AT VREF
INCREASE POWER TO 50% V1 AT VREF
MAINTAIN ALTITUDE AND
MAINTAIN ALTITUDE AND SLOW
SLOW TO FIRST INDICATION OF STALL
TO FIRST INDICATION OF STALL
SIMULTANEOUSLY
SIMULTANEOUSLY
- MAXIMUM POWER – APPLY
- MAXIMUM POWER – APPLY
- PITCH ATTITUDE – MAINTAIN 10° UP
- WING – APPROACH – LEVEL
- FLAPS APPROACH – CONFIRM
- PITCH ATTITUDE – MAINTAIN 10° UP

1 3 LANDING CONFIGURATION – (FULL FLAPS, GEAR DOWN)

THROTTLES – 50% N1
MAINTAIN ILS GLIDESLOPE
MAINTAIN ALTITUDE AND SLOW TO FIRST INDICATION OF STALL
MAXIMUM POWER – APPLY
PITCH ATTITUDE – LOWER TO HORIZON OR SLIGHTLY BELOW
AS AIRSPEED INCREASES, FLAPS – T.O. & APPR
AIRCRAFT – ROTATE TO 10° TO 12° UP
AT POSITIVE RATE OF CLIMB
- GEAR – UP
- SPEED – VREF + 10 MINIMUM
- FLAPS – UP
NOTE: Approaches to stall are
performed at pattern altitudes
only in the simulator.

CAUTION: Maximum power

should be used in those stalls
occurring near the ground and in
emergency situations. Stalls prac-
ticed at altitude never require
exceeding engine limitations.

Citation V For SimuFlite training only 3D-31

March 1999
 Maneuvers

Stall Series at
Altitude

3D-31 3D-32 For SimuFlite training only Citation V

March 1999
 Maneuvers

Visual Approach

BASE LEG
3 ABEAM RUNWAY END
4 LANDING GEAR – EXTEND
SINK RATE – DESCEND AT 500 TO 600 FPM BEFORE LANDING CHECKLIST – INITIATE
SPEED – VREF + 20 + WIND FACTOR

2 BEFORE PATTERN ENTRY/DOWNWIND (1,500 FT AGL)

APPROACH CHECKLIST – COMPLETE
FLAPS – T.O. & APPR
SPEED – VREF + 20 + WIND FACTOR

1500 FT AGL

1 BEFORE DESCENT
DESCENT CHECKLIST – COMPLETE
AIRSPEED BUGS – SET TO VREF, VAP, AND VZF
6 THRESHOLD
SPEED –VREF + WIND FACTOR

7 TOUCHDOWN
SPEEDBRAKES – DEPLOY
BRAKES – AS REQUIRED
REVERSE THRUST – AS REQUIRED

5 LANDING ASSURED
FLAPS – FULL DOWN
YAW DAMPER – OFF
SPEED – VAP
BEFORE LANDING CHECKLIST COMPLETE

NOTE: VAP is VREF +10 + wind

factor. Wind factor is one-half
the steady state wind and all of
the gust not to exceed 10 KIAS.

Citation V For SimuFlite training only 3D-33

March 1999
 Maneuvers

Visual Approach

3D-33 3D-34 For SimuFlite training only Citation V

March 1999
 Maneuvers

Two Engine ILS Approach and Landing/Missed Approach/Balked Landing

1 WITHIN THREE MINUTES OF ETA
FLAPS – UP
APPROACH CHECKLIST – COMPLETE
AIRSPEED BUGS – SET TO VREF, VAP, AND VZF
SPEED – REDUCE TO VREF + 40

2 IAF OUTBOUND
TIMING – START
SPEED – VREF + 20 + WIND FACTOR

3 PROCEDURE TURN INBOUND

FLAPS – T.O. & APPR
SPEED – VREF + 20 + WIND FACTOR

8 AT 400 FT
SPEED – VREF + 10 (MINIMUM)
FLAPS – UP

5 MISSED APPROACH
GO AROUND POWER – APPLY
AIRCRAFT – ROTATE TO 10° PITCH
FLAPS – T.O. & APPR 9 ATC – ADVISE

7 POSITIVE RATE OF CLIMB

GEAR – UP
SPEED – VAC

4 ONE DOT BELOW GLIDESLOPE INTERCEPT

GEAR – EXTEND
AT GLIDESLOPE INTERCEPT
TIMING – START
DESCENT – BEGIN
FLAPS – FULL DOWN
BEFORE LANDING CHECKLIST – COMPLETE DOWN
TO AUTOPILOT/YAW DAMPER
SPEED – VAP 6 THRESHOLD
SPEED – VREF + WIND FACTOR

Citation V For SimuFlite training only 3D-35

March 1999
 Maneuvers

Two Engine ILS Approach and

Landing/Missed Approach/
Balked Landing

3D-35 3D-36 For SimuFlite training only Citation V

March 1999
 Maneuvers

Two Engine Non-Precision Approach and Landing/Missed Approach/Balked Landing 1 WITHIN THREE MINUTES OF ETA
FLAPS – UP
APPROACH CHECKLIST – COMPLETE
AIRSPEED BUGS – SET TO VREF , VAP , AND VZF

2 IAF OUTBOUND
TIMING – START
FLAPS – T.O. & APPR
SPEED – VREF + 20 + WIND FACTOR

3 PROCEDURE TURN INBOUND

SPEED – VREF + 20 + WIND FACTOR
EXTEND GEAR PRIOR TO FAF

10 AT 400 FT
SPEED – VREF + 10 (MINIMUM)
FLAPS – UP

8 MISSED APPROACH
GO-AROUND POWER – APPLY
5 DEPARTING MDA AIRCRAFT – ROTATE TO 10° PITCH
SPEED – VAP FLAPS – T.O. & APPR
11 ATC – ADVISE

9 POSITIVE RATE OF CLIMB

GEAR – UP
SPEED – VAP

4 FINAL APPROACH FIX INBOUND

TIMING – START 7 THRESHOLD
SPEED – VAP 6 LANDING ASSURED SPEED – VREF + WIND FACTOR
DESCENT – BEGIN FLAPS – FULL
BEFORE LANDING CHECKLIST – COMPLETE TO FULL SPEED – VREF + WIND FACTOR
FLAPS AND AUTOPILOT/YAW DAMPER – OFF AUTOPILOT/YAW DAMPER – OFF
BEFORE LANDING CHECKLIST – COMPLETE

Citation V For SimuFlite training only 3D-37

March 1999
 Maneuvers

Two Engine Non-Precision

Approach and Landing/
Missed Approach/
Balked Landing

3D-37 3D-38 For SimuFlite training only Citation V

March 1999
 Maneuvers

Circling Approach

1 1

FLY 90° TO RUNWAY

TIMING – START AT
RUNWAY CENTERLINE FOR GO AROUND
AFTER 15 SECONDS, FLY DOWN RUNWAY
HEADING – 30° BANK FLY OVER RUNWAY
AT RUNWAY END, HEADING – 30° BANK
TURN TO DOWNWIND WHEN ESTABLISHED ON CENTERLINE,
1 TURN TO DOWNWIND
HEADING – 30° BANK TURN
TO DOWNWIND

45°

15 SEC
HEADING – TURN 45° TO DOWNWIND 4
USING 30° BANK TURN
TIMING – START 30° 30°
AFTER 30 SECONDS, HEADING –
30

BANK BANK
SE

30° BANK TURN TO DOWNWIND 2 ABEAM LANDING POINT 3

C

15 SEC

NOTES
PROCEDURE IS BASED ON 30° BANK TURNS
MAXIMUM SPEED FOR CATEGORY C IS 140 KIAS
A 300 FT OBSTACLE CLEARANCE IS PROVIDED
AT CIRCLING MINIMUMS
CATEGORY C CIRCLING MINIMUMS OBSTACLE
CLEARANCE IS TO 1.7 NM FROM RUNWAY

RECOMMENDATIONS BASIC CIRCLING PATTERN

FLAPS – APPROACH 1
GEAR – DOWN ENTER PATTERN AS APPROPRIATE
SPEED – VREF + 20 MINIMUM (140 KTS MAXIMUM) 1
FOR AIRCRAFT POSITION
AUTOPILOT – ENGAGED
ALTITUDE HOLD – SELECTED (F/D) 2 START TIMING ABEAM APPROACH END OF RUNWAY
HEADING SELECTED (F/D) FLY OVER RUNWAY
USE FULL DEFLECTION OF TURN KNOB FOR 30° WHEN ESTABLISHED ON CENTERLINE,
3 START TURN TO FINAL, 30° BANK TURN
BANKED TURN OR HEADING BUG HEADING – 30° BANK TURN TO DOWNWIND

4 WITH RUNWAY IN SIGHT AND IN POSITION TO MAKE

A NORMAL APPROACH TO LANDING
FLAPS – LAND
CAUTION: FAR 91.175 re- AUTOPILOT – DISENGAGED
quires immediate execution of DESCENT – BEGIN FROM MDA
SPEED – VAP
the missed approach procedure IF NOT IN A POSITION TO MAKE A NORMAL LANDING, GO-AROUND – PERFORM
when an identifiable part of the AT THRESHOLD, SPEED VREF + WIND FACTOR
airport is not distinctly visible
unless due to normal banking
during approach.

Citation V For SimuFlite training only 3D-39

March 1999
 Maneuvers

Circling Approach

3D-39 3D-40 For SimuFlite training only Citation V

March 1999
 Maneuvers

Engine Failure After V1/Takeoff Continued

8 CLEAR OF OBSTACLES (400 FT AGL MINIMUM)

SPEED – ACCELERATE TO V2 + 10
FLAPS – RETRACT
SPEED – ACCELERATE TO VENR
1 NON-ROLLING TAKEOFF FAILED ENGINE – IDENTIFY
BRAKES – HOLD ENGINE FAILURE CHECKLIST – COMPLETE
TAKEOFF POWER – SET ATC – CALL
BRAKES – RELEASE

3 70 KTS CROSSCHECK

5 AT VR , AIRCRAFT, ROTATE TO 10°

2 ROLLING TAKEOFF
SET TAKEOFF 7 MAINTAIN V2 UNTIL REACHING 400 FT AGL
POWER BY 60 KIAS AND CLEAR OF OBSTACLES

4 V1 DECISION SPEED
ENGINE FAILURE RECOGNIZED

6 POSITIVE RATE OF CLIMB

GEAR – UP

Citation V For SimuFlite training only 3D-41

March 1999
 Maneuvers

Engine Failure
After V1 – Takeoff
Continued

3D-41 3D-42 For SimuFlite training only Citation V

March 1999
 Maneuvers

Single Engine ILS Approach and Landing/Missed Approach/Balked Landing

1 WITHIN THREE MINUTES OF ETA

CONFIGURATION – CLEAN
SINGLE ENGINE APPROACH AND LANDING CHECKLIST – INITIATE
AIRSPEED BUGS – SET TO VREF, VAP, AND VZF
SPEED – REDUCE TO VREF + 40

2 IAF OUTBOUND
TIMING – START
FLAPS – T.O. & APPR
SPEED – VREF + 20 + WIND FACTOR

3 PROCEDURE TURN INBOUND

SPEED – VREF + 20 + WIND FACTOR

9 AT 400 FT (MINIMUM)
SPEED – VREF + 10
FLAPS – UP
SPEED – ACCELERATE TO VENR

7 MISSED APPROACH
GO-AROUND POWER – APPLY
AIRCRAFT – ROTATE TO 10° PITCH 10 ATC – ADVISE
FLAPS – T.O. & APPR

8
7 POSITIVE RATE OF CLIMB
GEAR – UP
SPEED – VAC

4 ONE DOT BELOW GLIDESLOPE 5 LANDING ASSURED

GEAR – EXTEND 6 THRESHOLD
FLAPS – FULL
SPEED – VAP SPEED – VREF + WIND FACTOR

Citation V For SimuFlite training only 3D-43

March 1999
 Maneuvers

Single Engine ILS

Approach and Landing/
Missed Approach/
Balked Landing

3D-43 3D-44 For SimuFlite training only Citation V

March 1999
 Maneuvers

Single Engine Non-Precision Approach and Landing/Missed Approach/Balked Landing

1 WITHIN THREE MINUTES OF ETA
CONFIGURATION – CLEAN
SINGLE ENGINE APPROACH AND LANDING CHECKLIST – INITIATE
AIRSPEED BUGS – SET TO VREF, VAP, AND VZF
SPEED – REDUCE TO VREF + 40

2 IAF OUTBOUND
TIMING – START
FLAPS – T.O. & APPR
SPEED – VREF + 20 + WIND FACTOR

3 PROCEDURE TURN INBOUND

SPEED – VREF + 20 + WIND FACTOR

9 AT 400 FT (MINIMUM)
SPEED – VREF + 10
FLAPS – UP
SPEED – ACCELERATE TO VENR

7 MISSED APPROACH
GO-AROUND POWER – APPLY
AIRCRAFT – ROTATE TO 10° PITCH 10 ATC – ADVISE
FLAPS – T.O. & APPR

7 POSITIVE RATE OF CLIMB

8
GEAR – UP
SPEED – VAC

4 AT FINAL APPROACH FIX INBOUND 5 LANDING ASSURED

GEAR – EXTEND 6 THRESHOLD
FLAPS – FULL
TIMING – START SPEED – VAP SPEED – VREF + WIND FACTOR
DESCENT – BEGIN
SPEED VAP

Citation V For SimuFlite training only 3D-45

March 1999
 Maneuvers

Single Engine Non-Precision

Approach and Landing/
Missed Approach/
Balked Landing

3D-45 3D-46 For SimuFlite training only Citation V

March 1999
 Maneuvers

Flap Inoperative Approach and Landing (Zero Flap Approach and Landing)

2 IAF OUTBOUND
TIMING – START
SPEED – VZF

1 WITHIN THREE MINUTES OF ETA

FLAPS INOPERATIVE APPROACH AND LANDING
3 PROCEDURE TURN INBOUND CHECKLIST – INITIATE
FLAPS VZF AIRSPEED BUGS – SET TO VREF + 15 (ADJUSTED VREF)
EXTEND GEAR PRIOR TO FAF SPEED – REDUCE TO VZF

4 AT FAF 5 THRESHOLD
TIMING – START SPEED – ADJUSTED VREF + WIND FACTOR
DESCENT – BEGIN
FLAP INOPERATIVE APPROACH AND LANDING CHECKLIST – COMPLETE
SPEED – ADJUSTED VREF + WIND FACTOR

NOTE: VZF is VREF +10 KIAS

and is intended as a minimum
speed for maneuvering the
aircraft up to 30° bank with no
flaps.

Citation V For SimuFlite training only 3D-47

March 1999
 Maneuvers

Flap Inoperative
Approach and Landing
(Zero Flap Approach and
Landing)

3D-47 3D-48 For SimuFlite training only Citation V

March 1999
 Maneuvers

Emergency Descent
1 OXYGEN MASKS – DON/100%
EMERGENCY DESCENT – AS REQUIRED
THROTTLES – IDLE
SPEEDBRAKES – EXTEND
MODERATE BANK – INITIATE

3 SPEED – VMO/MMO MAXIMUM

PAX ADVISORY – PASS SAFETY

2 PASSENGER OXYGEN – ENSURE USE

OXYGEN MIC SWITCHES – MIC OXYGEN

4 TIME AND CIRCUMSTANCES PERMITTING

CHECKLIST – CONFIRM ACCOMPLISHED
DESCENT / APPROACH / BEFORE LANDING CHECKLIST – PERFORM

5 LEVEL OFF AS REQUIRED CONSIDERING TERRAIN

NOTE: Use reduced speed if

structural damage has occurred.

NOTE: Attempt to control cabin

altitude with pressurization
controls. If unsuccessful, move
source select to EMER.

Citation V For SimuFlite training only 3D-49

March 1999
3D-50 For SimuFlite training only Citation V
March 1999
Flight planning is one of the most important activities that occurs
prior to each flight. Flight
This chapter instructs you in flight planning and parallels ground-
school instruction.
Planning
An example flight planning problem for the Citation V, with
appropriate charts, appears in this chapter. Italics depict data
for this example.
Chapter 4
Charts needed for any computations appear opposite the pro-
cedural steps. While arm is read in inches and moment is read
in inch-pounds, the notations “inches” or “in/lbs” are omitted.

Citation V For SimuFlite training only 4-1

March 1999
4-2 For SimuFlite training only Citation V
March 1999
 Flight Planning

General Information . . . . . . . . . . . . . . . . . . 4-5

Table of
Trip Planning Data . . . . . . . . . . . . . . . . . . . 4-7
Contents
Departure . . . . . . . . . . . . . . . . . . . . . . . . 4-7
Enroute . . . . . . . . . . . . . . . . . . . . . . . . . 4-7
Arrival . . . . . . . . . . . . . . . . . . . . . . . . . . 4-7
Alternate . . . . . . . . . . . . . . . . . . . . . . . . 4-8

Summary Flight Planning . . . . . . . . . . . . . . . 4-9

Detailed Flight Planning . . . . . . . . . . . . . . . 4-13

Climb Time, Distance and Fuel . . . . . . . . . . . . 4-14
Descent Time, Distance and Fuel . . . . . . . . . . . 4-16
Cruise Distance . . . . . . . . . . . . . . . . . . . . 4-16
Cruise Thrust and Fuel Consumption . . . . . . . . . 4-18
True Airspeed and Fuel Flow . . . . . . . . . . . . . 4-20
Cruise Groundspeed . . . . . . . . . . . . . . . . . 4-20
Time at Cruise . . . . . . . . . . . . . . . . . . . . . 4-20
Cruise Fuel . . . . . . . . . . . . . . . . . . . . . . 4-22
Total Time Enroute . . . . . . . . . . . . . . . . . . 4-24
Fuel Required Enroute . . . . . . . . . . . . . . . . 4-24
Alternate Leg Computations . . . . . . . . . . . . . . 4-24
Reserve Fuel . . . . . . . . . . . . . . . . . . . . . 4-30
Total Fuel and Time . . . . . . . . . . . . . . . . . . 4-30

Weight and Balance . . . . . . . . . . . . . . . . . 4-33

Basic Empty Weight and Moment . . . . . . . . . . . 4-34
Payload Weight and Moment . . . . . . . . . . . . . 4-36
Zero Fuel Weight and Moment . . . . . . . . . . . . 4-40
Fuel Weight and Moment . . . . . . . . . . . . . . . 4-42
Ramp Weight and Moment . . . . . . . . . . . . . . 4-44
Takeoff Weight and Moment . . . . . . . . . . . . . . 4-46
Takeoff CG Limits . . . . . . . . . . . . . . . . . . . 4-48

Citation V For SimuFlite training only 4-3

March 1999
 Landing Weight and Moment . . . . . . . . . . . . . 4-50
Landing CG Limits . . . . . . . . . . . . . . . . . . . 4-52

Definitions . . . . . . . . . . . . . . . . . . . . . . 4-55

Performance . . . . . . . . . . . . . . . . . . . . . 4-61
TOLD Card . . . . . . . . . . . . . . . . . . . . . . 4-61
Aircraft Takeoff Weight . . . . . . . . . . . . . . . . . 4-62
Airport Information . . . . . . . . . . . . . . . . . . . 4-62
Crosswind Component at Takeoff . . . . . . . . . . . 4-62
Maximum Allowable Climb Weight . . . . . . . . . . 4-64
Takeoff Field Length and V Speeds . . . . . . . . . . 4-66
Obstacle Clearance . . . . . . . . . . . . . . . . . . 4-68
Single Engine Takeoff Flight Path Tables . . . . . . . 4-72
Single Engine Takeoff Flight Path . . . . . . . . . . . 4-72
Available vs. Required Field Length . . . . . . . . . . 4-74
Emergency Return . . . . . . . . . . . . . . . . . . . 4-74
Takeoff Thrust Setting . . . . . . . . . . . . . . . . . 4-76
Single Engine Enroute Thrust Setting . . . . . . . . . 4-76
Gross Weight at Destination . . . . . . . . . . . . . . 4-78
Crosswind Component at Destination . . . . . . . . . 4-78
Maximum Allowable Landing Weight . . . . . . . . . 4-80
Landing Distance and Approach Speeds . . . . . . . 4-82
Go-Around/Takeoff Thrust Setting . . . . . . . . . . . 4-84

Supplemental Information . . . . . . . . . . . . . . 4-87

General . . . . . . . . . . . . . . . . . . . . . . . . 4-88
Approach and Landing Gross Climb Gradients . . . . 4-88
Performance for Contaminated Runways . . . . . . . 4-90
Simplified Takeoff Criteria . . . . . . . . . . . . . . . 4-91
Single Engine Enroute Net Climb Gradient – Percent . . 4-92
Single Engine Takeoff Flight Path Distances . . . . . 4-94

4-4 For SimuFlite training only Citation V

March 1999
 Flight Planning

A preflight briefing may be obtained from a DUAT or from a Flight

Service Station by telephone, radio, or personal visit. The briefing General
should consist of weather, airport, enroute NAVAID information,
and NOTAMS.
Information
Normally, plan the trip and compute the weight and balance first.
However, when conditions at the departure airport are near the max-
imum operating limits of the aircraft, determine takeoff performance
data first. This prevents planning a trip and then discovering that
takeoff is impossible with the planned passenger and fuel load.
The performance tables require that the planned altitude and approx-
imate aircraft weight be known. Aircraft weight decreases as fuel
is consumed and can be estimated by scheduling 1,400 lbs for the first
hour, 1,200 lbs for the second, and 1,000 lbs for each subsequent.
In real world situations, the estimated fuel required must be modi-
fied for known delays (e.g., weather, diversions, air traffic flow).
If fuel conservation is more important than time to destination, con-
sult the specific range vs. cruise wind tables in the Citation V
Operating Manual for long range cruise information.
For maximum groundspeed considerations, this chapter assumes the
Citation V maximum thrust setting of 104% N1 for the cruise legs
to the primary and alternate destinations.

Citation V For SimuFlite training only 4-5

March 1999
4-6 For SimuFlite training only Citation V
March 1999
 Flight Planning

The example depicted in this chapter is based on the following data.

Trip Planning
Departure Data
Runway Length . . . . . . . . . . . . . . . . . . . . . 5,300 FT
Runway Gradient . . . . . . . . . . . . . . . . . . . . . . . 0%
Runway Heading . . . . . . . . . . . . . . . . . . . . . . 170°
Ramp Weight . . . . . . . . . . . . . . . . . . . . 16,100 LBS
Anti-Ice . . . . . . . . . . . . . . . . . . . . . . . . . . . OFF
Anti-Skid . . . . . . . . . . . . . . . . . . . . . . . . . . . ON
Takeoff Flaps . . . . . . . . . . . . . . . . . . . . . . 7° or 15°
OAT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20°C
Pressure Altitude . . . . . . . . . . . . . . . . . . . . 2,000 FT
Runway Winds . . . . . . . . . . . . . . . . . . . 190°/15 KTS
Obstacle . . . . . . . . . . . 300 FT HIGH AT 3,640 FT FROM
. . . . . . . . . . . . . . . .DEPARTURE END OF RUNWAY

Enroute
Cruising Altitude . . . . . . . . . . . . . . . . . . . 37,000 FT
Temperature . . . . . . . . . . . . . . . . . . . . . . . . . ISA
Headwind Component . . . . . . . . . . . . . . . . . . 50 KTS
Distance to Destination . . . . . . . . . . . . . . . . . 800 NM

Arrival
Runway Length . . . . . . . . . . . . . . . . . . . . . 6,000 FT
Runway Gradient . . . . . . . . . . . . . . . . . . . . . . +2%
Runway Heading . . . . . . . . . . . . . . . . . . . . . . 140°
Anti-Ice . . . . . . . . . . . . . . . . . . . . . . . . . . . OFF
Anti-Skid . . . . . . . . . . . . . . . . . . . . . . . . . . . ON
OAT . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25°C
Pressure Altitude . . . . . . . . . . . . . . . . . . . . 1,000 FT
Runway Winds . . . . . . . . . . . . . . . . . . . 130°/20 KTS

Citation V For SimuFlite training only 4-7

March 1999
 Alternate
Cruising Altitude . . . . . . . . . . . . . . . . . . . 17,000 FT
Temperature . . . . . . . . . . . . . . . . . . . . . . . . . 15°C
Headwind Component . . . . . . . . . . . . . . . . . . . 0 KTS
Distance From Destination . . . . . . . . . . . . . . . 100 NM

4-8 For SimuFlite training only Citation V

March 1999
 Flight Planning

The example begins with a Summary Flight Planning estimate of

trip time and fuel consumption. Using this estimate, compute the Summary
approximate time, distance, and fuel.
Flight
Planning

Citation V For SimuFlite training only 4-9

March 1999
 Determine estimated trip time and fuel consumption, including the
alternate leg, by using the applicable Maximum Cruise Thrust chart
(Figure 4-1).
For this example, use the chart for 104% N1 at 37,000 ft.

1. Enter the chart from the bottom left at the correct cruise wind
(50 kts).
2. Move to the right to the correct total stage length arc, including
alternate.
For this example, the distance is 900 NM (800 to destination plus
100 to alternate).

3. Move up to the time and fuel angled reference lines to an approx-

imate takeoff weight.
For conservatism, use the 15,900 reference line on each pair of
lines.

4. Move to the left from the fuel line to the edge of the chart to
determine the fuel requirement (3,900 lbs).
This figure does not include the reserve fuel. An average fuel
requirement for 45 min. of reserve is 800 lbs.
3,900 lbs trip fuel + 800 lbs reserve fuel equals a 4700 lbs
estimated fuel requirement.

5. Move to the right from the time line to the edge of the chart to
determine the duration of the flight (2:50).

4-10 For SimuFlite training only Citation V

March 1999
 Flight Planning

Maximum Cruise Thrust

Cruise (104.0% N1)

3,900 LBS

2:50

4-1

Citation V For SimuFlite training only 4-11

March 1999
4-12 For SimuFlite training only Citation V
March 1999
 Flight Planning

Proper detailed planning is required to ensure safe performance.

This section reviews the computations necessary to plan a trip. Detailed
After establishing the approximate fuel and time requirements for the Flight
trip, determine the approximate gross takeoff weight.
In this example, there are no unusual conditions (e.g., distance, Planning
elevation, climb grade requirements, airport ambient tempera-
tures, runway lengths). A takeoff weight of 15,900 lbs is desired;
the additional weight above the basic empty weight, payload, and
minimum fuel requirement is composed of tankered fuel.

Citation V For SimuFlite training only 4-13

March 1999
 Climb Time, Distance and Fuel
Use the appropriate Cruise Climb table (Figure 4-2) to determine
time, distance, and fuel required for the climb to cruising altitude. To
determine the effect of winds aloft on climb distance, see the Wind
Effect on Climb Distance table at the bottom and use the “60/40”
rule; that is, apply 60% of the cruise wind to estimate the average wind
speed in the climb.
1. Enter the appropriate altitude block (37,000 ft) at the correct ISA
value (ISA).
2. Move to the right until the appropriate takeoff weight column is
reached (15,900 lbs). Read the necessary time, distance, fuel
requirement, and rate of climb (21 minutes, 106 NM, 668 lbs,
and 829 fpm, respectively).
3. To determine the effect of wind on climb distance, perform the
following.
a. Enter the Wind Effect on Climb Distance table as closely as
possible to the previously determined climb time.
Because the time is 21 minutes, use the 20 minute row.

b. Move to the right to the appropriate wind column and read the
correction factor. Interpolate as required.
Sixty percent of the cruise headwind is 30 kts; the
interpolated correction factor is 9.6 NM; round this up
to an even 10 NM.

3. Add or subtract the correction factor to/from the previously deter-

mined climb distance.
Since this is a headwind correction, subtracting 10 NM from 106
yields the corrected climb distance, or 96 NM.

4-14 For SimuFlite training only Citation V

March 1999
 Flight Planning

Cruise Climb
250 KIAS/.55 Indicated MACH
Time, Distance, Fuel, and Rate-of-Climb, Anti-Ice Systems – OFF

4-2

Citation V For SimuFlite training only 4-15

March 1999
 Descent Time, Distance and Fuel
Use the Normal Descent table (Figure 4-3) to determine time, distance,
and fuel in descent. Again, apply the “60/40” rule; that is, apply
40% of the cruise wind to estimate the average wind speed in the
descent.
1. Enter the table from the left at the appropriate altitude (37,000 ft).
2. Move to the right to obtain the time and fuel used in the descent
(18.5 minutes and 247 lbs, respectively).
3. Determine the descent distance.
Interpolation for 20 kts (i.e., 40% of the cruise headwind of 50 kts)
yields 94.4 NM. Round this up to an even 95 NM.

Cruise Distance
C = T - (L + D)
where: Determine the distance covered at cruising altitude by subtracting the
C = Cruise Distance climb and descent distances from the total distance.
T = Total Distance The climb distance previously determined is 96 NM. The descent
L = Climb Distance distance previously determined is 95 NM. Thus, the cruise dis-
D = Descent Distance tance is 609 NM (800 minus (96 plus 95) equals 609).

4-16 For SimuFlite training only Citation V

March 1999
 Flight Planning

Normal Descent – 2000 ft per Minute

Anti-Ice Systems – OFF, Speedbrakes Retracted,
Gear and Flaps Up
800 lbs per Hour (400 lbs per Hour per Engine)

4-3

Citation V For SimuFlite training only 4-17

March 1999
 Cruise Thrust and Fuel Consumption
If long range cruise endurance is preferred to higher groundspeed, use
the appropriate Specific Range vs. Cruise Wind chart (Figure 4-4).
Assume for a moment that thrust is not 104% N1, but that you
wish to determine thrust for long range cruise.

1. Enter the table at the top of the figure from the left at the planned
altitude (37,000 ft).
2. Move to the right to the forecast cruise wind column (50 kts head-
wind) and read the thrust setting (89.9).
3. To determine specific fuel consumption at the long range cruise
thrust setting, perform the following.
a. Enter the graph at the bottom of the figure from the bottom at
the headwind component (50 kts headwind).
b. Follow the line up to its intersection with the desired cruise
altitude (37,000 ft).
c. Move to the left to the edge of the chart to read the fuel
consumption.
The specific fuel consumption is 0.355 NM/lb.

4-18 For SimuFlite training only Citation V

March 1999
 Flight Planning

Specific Range vs Cruise Wind

Long Range Cruise, Standard Day
13,500 lbs Average Cruise Weight

.355

4-4

Citation V For SimuFlite training only 4-19

March 1999
 True Airspeed and Fuel Flow
Use the Cruise table (Figure 4-5) to determine the cruise true air-
T–L=C speed and fuel flow. Use the highest listed value closest to the weight
where: of the aircraft at the start of the cruise segment.
T = True Airspeed
L = Climb Fuel 1. Enter the table from the left with the appropriate weight, fan
C = Weight at Start of RPM, and OAT.
Cruise
2. Move right and interpolate the figures for fuel flow and true air-
speed.
Interpolating for 15,232 lbs (15,900 minus 668 equals 15,232)
reveals that the fuel flow is 1,310 lbs/hr and the true airspeed is
409 kts.

Cruise Groundspeed
Determine the groundspeed by adding or subtracting the forecast
T ±C = G cruise wind to/from the true airspeed.
where:
T = True Airspeed Because a 50 knot headwind is forecast, the groundspeed is
C = Cruise Wind 359 kts (409 minus 50 equals 359).
G = Groundspeed
Time at Cruise
C ÷G = T Find the time at cruise by dividing cruise distance by the computed
where: groundspeed.
C = Cruise Distance The cruise distance was previously determined to be 609 NM,
G = Groundspeed and the groundspeed is 359 kts (609 ÷ 359 = 1.69). Thus, the
T = Time at Cruise time at cruise is 1.7 hours (1 hour, 42 minutes).
This is a rough estimate of the time at cruise based on the first
hour’s true airspeed. For a longer trip length, the time at cruise
should be based on the true airspeed for each subsequent hour.
As the aircraft weight decreases the true airspeed increases with
the same power setting.

4-20 For SimuFlite training only Citation V

March 1999
 Flight Planning

Cruise
37,000 Feet
Anti-Ice Systems – OFF, Two Engines

4-5

Citation V For SimuFlite training only 4-21

March 1999
 Cruise Fuel
The amount of fuel required for cruise is the sum of the fuel flow rates
determined for each hour or part of an hour at cruise. Use the Cruise
table (Figure 4-6).
Time at cruise was computed previously as 1.7 hours (1 hour, 42
minutes). Fuel flow for the first hour is 1,310 lbs/hr at an air-
craft weight of 15,232 lbs.

The cruising weight for the second hour is obtained by subtracting

the fuel used during the first hour from the first hour airplane weight.
Thus, an aircraft weight of 15,232 lbs minus 1,310 lbs fuel is
13,922 lbs.

To determine second hour fuel consumption, perform the following.

1. Enter the table at the weight of the aircraft at the beginning of
the second hour (13,922 lbs).
2. Move to the right to read fuel flow, fan RPM, and temperature.
After interpolation, the fuel flow is 1,321 lbs/hr; the fan setting
(104.0%) and temperature (-56°C) remain the same as for the
first hour.

3. Calculate second hour fuel consumption.

When 0.7 hour (i.e., 42 minutes) is multiplied by 1,321, the result
is 924.70 lbs (rounded to 925 lbs).

4. Add the fuel values to determine fuel required at cruise.

Adding the 1,310 lbs for the first hour to the 925 lbs for the final
42 minutes yields a total cruise fuel consumption of 2,235 lbs.

4-22 For SimuFlite training only Citation V

March 1999
 Flight Planning

Cruise
37,000 Feet
Anti-Ice Systems – OFF, Two Engines

4-6

Citation V For SimuFlite training only 4-23

March 1999
 Total Time Enroute
The total time enroute is determined by adding the times for climb,
cruise, and descent.
Climb 0:21.0 hours:minutes
Cruise 1:42.0
Descent + 0:18.5
TOTAL 2:21.5 (rounded to 2:22)

Fuel Required Enroute

The total fuel required enroute is the sum of the fuel needed for
climb, cruise, and descent.
Climb 668 lbs
Cruise 2235
Descent + 247
TOTAL 3150

Alternate Leg Computations

Use the appropriate Normal Descent, Cruise Climb, and 17,000 Feet
Cruise tables to determine the fuel required to the alternate airport.
Assume a zero headwind.

Descent distance is the key to establishing the altitude for cruise and
for beginning the descent to the alternate. Compute the climb and
cruise segments of the alternate leg after determining the altitude
from which the descent begins.
When an alternate is 100 NM or less from the primary destination,
use a “60/40” ratio to determine climb and descent segments enroute
to the alternate. That is, 40% of the distance to the alternate is in the
descent.
The fuel required to the alternate airport is the sum of the fuel require-
ments for the climb, cruise, and descent segments enroute to the
alternate.

4-24 For SimuFlite training only Citation V

March 1999
 Flight Planning

First, refer to the Normal Descent table (Figure 4-7).

1. Using the “60/40” ratio, enter the table in the appropriate wind
column (zero). Read down to the number closest to 40% of the
distance to the alternate.
Because the distance to the alternate is 100 NM, 40% of this
number is 40 NM. Use the figures for 39 NM.

2. Move to the left to the edge of the table to determine the altitude
at which to begin the descent, the descent time, and the descent
fuel. Interpolate as required.
Because an odd altitude is appropriate for the leg, 39 NM and
17,000 ft are chosen. The time required is 8.5 minutes, and the fuel
used is 113 lbs.

Normal Descent – 2000 ft per Minute

Anti-Ice Systems – OFF, Speedbrakes Retracted,
Gear and Flaps Up
800 lbs per Hour (400 lbs per Hour per Engine)

4-7

Citation V For SimuFlite training only 4-25

March 1999
 Alternate Leg Computations (continued)
Next, use the Cruise Climb table (Figure 4-8) to determine the time,
distance, and fuel required for the climb to the alternate cruise alti-
tude (17,000 ft).
1. Enter the table at the appropriate weight for the alternate leg alti-
tude (17,000 ft) and move down to the data block opposite the
correct ISA value (ISA).
Because the 15,900 lbs takeoff weight minus the 3,150 lbs enroute
fuel weight is 12,750 lbs, use the 13,000 lbs column.

2. Read the data block.

The figures for the alternate climb are 6 minutes, 25 NM, and
227 lbs of fuel.

4-26 For SimuFlite training only Citation V

March 1999
 Flight Planning

Cruise Climb
250 KIAS/.55 Indicated MACH
Time, Distance, Fuel, and Rate-of-Climb; Anti-Ice Systems – OFF

4-8

Citation V For SimuFlite training only 4-27

March 1999
 Alternate Leg Computations (continued)
Next, compute the cruise distance to the alternate as you did for the
destination cruise segment: subtract the sum of climb distance and
descent distance from the total distance.
Cruise distance equals 100 NM minus (25 NM plus 39 NM), or
36 NM.

Now, use the appropriate Cruise table (17,000 ft) (Figure 4-9) to
compute cruise time and fuel to the alternate. Interpolate as required.
1. Enter the table from the left with the appropriate aircraft weight
(12,750 lbs) and fan speed (maximum cruise thrust).
2. Read the true airspeed.
Because the airspeed is 370 KTAS for both 13,500 lbs and
12,500 lbs, it is also 370 KTAS for 12,750 lbs.

3. Compute time at cruise by dividing the cruise distance by the

groundspeed.
The 36 NM cruise distance divided by 370 equals a cruise time
of 0.1 hrs, or 6 minutes.

4. Read the fuel flow; interpolate as required.

After interpolation, the fuel flow is 1,646 lbs/hr.

5. Compute fuel consumption for the cruise time.

The cruise time is 0.1 hours, thus 0.1 times 1,646 is 165 lbs.

6. Add fuel and time values for all alternate segments.

Flight Segment Time (minutes) Fuel (lbs)
Climb 6.0 227
Cruise 6.0 165
Descent + 8.5 + 113
TOTAL 20.5 505
or (21)

4-28 For SimuFlite training only Citation V

March 1999
 Flight Planning

Cruise
17,000 Feet
Anti-Ice Systems – OFF, Two Engines

4-9

Citation V For SimuFlite training only 4-29

March 1999
 Reserve Fuel
FAR 91.23(a)(3) required enough reserve fuel to fly for 45 minutes
at normal cruise airspeed after reaching the alternate airport.
Use the Cruise table for 17,000 ft to determine the required fuel
reserves for this problem.

To determine required reserve fuel, perform the following.

1. Enter the appropriate Cruise table (Figure 4-10) at the approxi-
mate aircraft weight (takeoff weight minus total fuel consumed),
correct ISA, and fan RPM.
In this case, use 12,245 lbs (15,900 lbs minus (3,150 plus 505))
at approximately 80% fan RPM.

2. Interpolate the fuel flow for the weight.

At 12,245 lbs, the fuel flow is 1,033 lbs/hr.

3. Compute 75% of the obtained value for the 45 minute reserve.

The reserve fuel is 75% of 1,033 lbs, or 775 lbs.

Another way to estimate reserve fuel is to use a recommended quan-

tity of 800 lbs. This represents a safe and realistic fuel reserve.
For this example, however, use the computed figure of 775 lbs.

Total Fuel and Time

The total fuel required is the sum of the determined requirements
for all trip segments, including start, taxi, and takeoff.
Remember to allow 200 lbs for start, taxi, and takeoff. Time is
the sum of the destination and alternate times.

For this example, the following figures apply.

Segment Fuel (lbs) Time (hrs & min)
Taxi 200
Destination 3150 2:22
Alternate 505 :21
Reserve + 775 + ______
TOTAL FUEL REQUIRED 4630 2:43
The original estimates of the trip time and fuel requirements were
2:50 hrs and 4,700 lbs of fuel.

4-30 For SimuFlite training only Citation V

March 1999
 Flight Planning

Cruise
17,000 Feet
Anti-Ice Systems – OFF, Two Engines

4-10

Citation V For SimuFlite training only 4-31

March 1999
4-32 For SimuFlite training only Citation V
March 1999
 Flight Planning

Precise weight and balance computations are essential elements

of flight planning. Accuracy of these computations helps to Weight and
ensure a safe flight. This section reviews the procedures for
computing weight and balance data.
Balance

Citation V For SimuFlite training only 4-33

March 1999
 Basic Empty Weight and Moment
Basic empty weight is the weight of the aircraft including full oil
and all undrainable fluids. This weight and its moment are noted on
the aircraft weighing form. When the aircraft is altered, refer to the
weight and balance record for the corrected information.
Assume that the basic empty weight is 9,390 lbs with a moment
of 29,051. Record this information on the loading schedule
(Figure 4-11).

4-34 For SimuFlite training only Citation V

March 1999
 Flight Planning

Citation V Loading Schedule

Payload Computations Item Weight Moment/
(lbs) 100

Item ARM Weight Moment/ 1. Basic Empty Weight 9,390 29,051.00

(lbs) 100 * Aircraft CG = 309.04

Occupants 2. Payload

Seat 1 131 3. Zero Fuel Weight

(Sub-total)(Do not exceed
Seat 2 131 maximum zero fuel weight
of 11,200 lbs (standard)
Seat 3
or 12,200 lbs (optional).)
Seat 4 221
4. Fuel Loading
Seat 5 262
Seat 6 262 5. Ramp Weight
Seat 7 308
6. Less Fuel for Taxiing
Seat 9 308
7. ** Takeoff Weight
Seat 10 170 (Do not exceed maximum
takeoff weight of 15,900 lbs)
* Aircraft CG = ***
Toilet 345
8. Less Fuel to Destination

Baggage 9. ** Landing Weight

(Do not exceed maximum
Nose 74.00 landing weight of 15,200 lbs
Tailcone * Aircraft CG = ***
434.00
Tailcone 462.00 MOMENT/100
* Aircraft CG = X 100
WEIGHT

Storage ** Totals must be within approved weight and center-of-gravity

Cabin 348.00 limits. It is the responsibility of the operator to ensure that
the aircraft is loaded properly. The Basic Empty Weight CG
is noted on the Aircraft Weighing Form. If the aircraft has
been altered, refer to the Weight and Balance Record for
Cabinet 155.00 information.
Contents

Payload *** Enter the Center-of-Gravity Limits Envelope Graph to check

if within approved limits (shaded area).

4-11

Citation V For SimuFlite training only 4-35

March 1999
 Payload Weight and Moment
Payload is defined as the crew, passengers, cabinet, contents, and
baggage. The payload moment is the sum of the individual moments
of each of these components. Calculate moment by multiplying a
weight by its arm (i.e., the distance from the reference datum line),
then dividing by 100 to facilitate further computations.
In this example, use the manufacturer-supplied tables in
Figures 4-12 and 4-13.

Assume the crew, passenger, baggage (Figure 4-14), and equipment

load shown in Table 4-A.

Passenger/Item Arm Weight Moment

Crew, Seat 1 131 170 222.70

Crew, Seat 2 131 180 235.80
Passenger, Seat 4 221 150 331.50
Passenger, Seat 5 262 180 471.60
Passenger, Seat 7 308 160 492.80
Passenger, Seat 8 308 170 523.60
Passenger, Seat 10 170 170 289.00

Nose Baggage 74 80 59.20

Aft Tailcone Baggage 462 80 369.60
Aft Cabin Baggage/Aft Zone 348 20 69.60
Refreshment Center 155 40 62.00

Table 4-A; Passenger/Baggage Weights and Moments

4-36 For SimuFlite training only Citation V

March 1999
 Flight Planning

Crew and Passengers Compartments Baggage and Cabinet Compartments

Weight and Moment – Optional Weight and Moment Tables

4-12

4-13

Citation V For SimuFlite training only 4-37

March 1999
 Flight Planning

Weight and Balance Manual Payload Weight and Moment (continued)

Cessna Model 560 1. Add all weights and moments to arrive at total payload figures
(1,580 lbs and 3,599 respectively).
2. Record these figures on the loading schedule (Figure 4-15).

4-14

4-37 4-38 For SimuFlite training only Citation V

March 1999
 Flight Planning

Citation V Loading Schedule

Payload Computations Item Weight Moment/
(lbs) 100

Item ARM Weight Moment/ 1. Basic Empty Weight 9,390 29,051.00

(lbs) 100 * Aircraft CG = 309.04

Occupants 2. Payload 1,580 3,599.00

Seat 1 131 170 222.70 3. Zero Fuel Weight

(Sub-total)(Do not exceed
Seat 2 131 180 235.80 maximum zero fuel weight
of 11,200 lbs (standard)
Seat 3
or 12,200 lbs (optional).)
Seat 4 221 150 331.50
180 471.60 4. Fuel Loading
Seat 5 262
Seat 6 262 180 471.60 5. Ramp Weight
Seat 7 308 160 492.80
6. Less Fuel for Taxiing
Seat 9 308 170 523.60
7. ** Takeoff Weight
Seat 10 170 170 289.00 (Do not exceed maximum
takeoff weight of 15,900 lbs)
* Aircraft CG = ***
Toilet 345
8. Less Fuel to Destination

Baggage 9. ** Landing Weight

(Do not exceed maximum
Nose 74.00 landing weight of 15,200 lbs
Tailcone * Aircraft CG = ***
434.00
Tailcone 462.00 MOMENT/100
* Aircraft CG = X 100
WEIGHT

Storage ** Totals must be within approved weight and center-of-gravity

Cabin 348.00 limits. It is the responsibility of the operator to ensure that
the aircraft is loaded properly. The Basic Empty Weight CG
is noted on the Aircraft Weighing Form. If the aircraft has
been altered, refer to the Weight and Balance Record for
Cabinet 155.00 information.
Contents

Payload *** Enter the Center-of-Gravity Limits Envelope Graph to check

if within approved limits (shaded area).

4-15

Citation V For SimuFlite training only 4-39

March 1999
 Zero Fuel Weight and Moment
The zero fuel weight is the basic empty weight plus the payload.
Zero Fuel Weight Limit
Enter these figures on the loading schedule (Figure 4-16).
11,200 lbs
Because the basic empty weight is 9,390 lbs, adding the 1,580 lb
payload to this figure yields a zero fuel weight of 10,970 lbs.
Adding the moments for the basic empty weight and the payload
yields the zero fuel weight moment of 32,650.

4-40 For SimuFlite training only Citation V

March 1999
 Flight Planning

Citation V Loading Schedule

Payload Computations Item Weight Moment/
(lbs) 100

Item ARM Weight Moment/ 1. Basic Empty Weight 9,390 29,051.00

(lbs) 100 * Aircraft CG = 309.04

Occupants 2. Payload 1,580 3,599.00

Seat 1 131 170 222.70 3. Zero Fuel Weight 10,970 32,650.00

(Sub-total)(Do not exceed
Seat 2 131 180 235.80 maximum zero fuel weight
of 11,200 lbs (standard)
Seat 3
or 12,200 lbs (optional).)
Seat 4 221 150 331.50
180 471.60 4. Fuel Loading
Seat 5 262
Seat 6 262 180 471.60 5. Ramp Weight
Seat 7 308 160 492.80
6. Less Fuel for Taxiing
Seat 9 308 170 523.60
7. ** Takeoff Weight
Seat 10 170 170 289.00 (Do not exceed maximum
takeoff weight of 15,900 lbs)
* Aircraft CG = ***
Toilet 345
8. Less Fuel to Destination

Baggage 9. ** Landing Weight

(Do not exceed maximum
Nose 74.00 80 59.20 landing weight of 15,200 lbs
Tailcone * Aircraft CG = ***
434.00
Tailcone 462.00 80 369.60 MOMENT/100
* Aircraft CG = X 100
WEIGHT

Storage ** Totals must be within approved weight and center-of-gravity

Cabin 348.00 limits. It is the responsibility of the operator to ensure that
20 69.60
the aircraft is loaded properly. The Basic Empty Weight CG
is noted on the Aircraft Weighing Form. If the aircraft has
been altered, refer to the Weight and Balance Record for
Cabinet 155.00 40 62.00 information.
Contents

Payload 1,580 3,599.00 *** Enter the Center-of-Gravity Limits Envelope Graph to check
if within approved limits (shaded area).

4-16

Citation V For SimuFlite training only 4-41

March 1999
 Fuel Weight and Moment
Use the Fuel table (Figure 4-17) to determine the moment for the
Weight Limitations
fuel requirement for the flight.
Maximum Ramp
Weight . . . . . . . 16,100 lbs Because the zero fuel weight is 10,970 lbs and we previously
decided to take on fuel to the maximum possible weight, we now
Maximum Takeoff
Weight . . . . . . . 15,900 lbs
load 5,130 lbs of fuel (16,100 (maximum ramp weight) minus
10,970 (current zero fuel weight) equals 5,130).

1. Enter the table and determine the moment for the fuel load; use
interpolation if necessary.
The fuel load is 5,130; its moment is 15,532.95.

2. Enter this information on the loading schedule (Figure 4-18).

Fuel Loading Weight and Moment Table

5,130 15,523.95
LBS

4-17

4-42 For SimuFlite training only Citation V

March 1999
 Flight Planning

Citation V Loading Schedule

Payload Computations Item Weight Moment/
(lbs) 100

Item ARM Weight Moment/ 1. Basic Empty Weight 9,390 29,051.00

(lbs) 100 * Aircraft CG = 309.04

Occupants 2. Payload 1,580 3,599.00

Seat 1 131 170 222.70 3. Zero Fuel Weight 10,970 32,650.00

(Sub-total)(Do not exceed
Seat 2 131 180 235.80 maximum zero fuel weight
of 11,200 lbs (standard)
Seat 3
or 12,200 lbs (optional).)
Seat 4 221 150 331.50
180 471.60 4. Fuel Loading 5,130 15,532.95
Seat 5 262
Seat 6 262 180 471.60 5. Ramp Weight
Seat 7 308 160 492.80
6. Less Fuel for Taxiing
Seat 9 308 170 523.60
7. ** Takeoff Weight
Seat 10 170 170 289.00 (Do not exceed maximum
takeoff weight of 15,900 lbs)
* Aircraft CG = ***
Toilet 345
8. Less Fuel to Destination

Baggage 9. ** Landing Weight

(Do not exceed maximum
Nose 74.00 80 59.20 landing weight of 15,200 lbs
Tailcone * Aircraft CG = ***
434.00
Tailcone 462.00 80 369.60 MOMENT/100
* Aircraft CG = X 100
WEIGHT

Storage ** Totals must be within approved weight and center-of-gravity

Cabin 348.00 limits. It is the responsibility of the operator to ensure that
20 69.60
the aircraft is loaded properly. The Basic Empty Weight CG
is noted on the Aircraft Weighing Form. If the aircraft has
been altered, refer to the Weight and Balance Record for
Cabinet 155.00 40 62.00 information.
Contents

Payload 1,580 3,599.00 *** Enter the Center-of-Gravity Limits Envelope Graph to check
if within approved limits (shaded area).

4-18

Citation V For SimuFlite training only 4-43

March 1999
 Ramp Weight and Moment
Ramp Weight Limit The ramp weight is the gross weight of the aircraft fully loaded for
16,100 lbs flight. It is the sum of the zero fuel weight and the fuel load. The
ramp moment is the sum of the zero fuel weight moment and the
fuel load moment.
Zero Fuel Weight 10,970 lbs
Fuel Load + 5,130
RAMP WEIGHT 1 6,100

Zero Fuel Weight Moment 32,650.00

Fuel Load Moment + 15,532.95
RAMP WEIGHT MOMENT 48,182.95

Record these values on the loading schedule (Figure 4-19).

4-44 For SimuFlite training only Citation V

March 1999
 Flight Planning

Citation V Loading Schedule

Payload Computations Item Weight Moment/
(lbs) 100

Item ARM Weight Moment/ 1. Basic Empty Weight 9,390 29,051.00

(lbs) 100 * Aircraft CG = 309.04

Occupants 2. Payload 1,580 3,599.00

Seat 1 131 170 222.70 3. Zero Fuel Weight 10,970 32,650.00

(Sub-total)(Do not exceed
Seat 2 131 180 235.80 maximum zero fuel weight
of 11,200 lbs (standard)
Seat 3
or 12,200 lbs (optional).)
Seat 4 221 150 331.50
180 471.60 4. Fuel Loading 5,130 15,532.95
Seat 5 262
Seat 6 262 180 471.60 5. Ramp Weight 16,100 48,182.95
Seat 7 308 160 492.80
6. Less Fuel for Taxiing
Seat 9 308 170 523.60
7. ** Takeoff Weight
Seat 10 170 170 289.00 (Do not exceed maximum
takeoff weight of 15,900 lbs)
* Aircraft CG = ***
Toilet 345
8. Less Fuel to Destination

Baggage 9. ** Landing Weight

(Do not exceed maximum
Nose 74.00 80 59.20 landing weight of 15,200 lbs
Tailcone * Aircraft CG = ***
434.00
Tailcone 462.00 80 369.60 MOMENT/100
* Aircraft CG = X 100
WEIGHT

Storage ** Totals must be within approved weight and center-of-gravity

Cabin 348.00 limits. It is the responsibility of the operator to ensure that
20 69.60
the aircraft is loaded properly. The Basic Empty Weight CG
is noted on the Aircraft Weighing Form. If the aircraft has
been altered, refer to the Weight and Balance Record for
Cabinet 155.00 40 62.00 information.
Contents

Payload 1,580 3,599.00 *** Enter the Center-of-Gravity Limits Envelope Graph to check
if within approved limits (shaded area).

4-19

Citation V For SimuFlite training only 4-45

March 1999
 Takeoff Weight and Moment
The takeoff weight is the ramp weight minus the fuel used during taxi.
Takeoff Weight Limit
Use the Fuel Moment table (Figure 4-20) as before.
15,900 lbs
Assume a 200 lb fuel burn for taxi.
Takeoff weight is limited by the
most restrictive of: 1. Determine the moment for the takeoff fuel load using interpola-
■ maximum certified takeoff
tion.
weight
■ maximum takeoff weight
The takeoff fuel load, 4,930 lbs (5,130 lbs minus 200), has a
moment of 14,915.50.
permitted by climb
requirements
2. Subtract the fuel weight and moment after taxi from the ramp
■ takeoff field length.
weight. The result is the moment for the taxi fuel.
5130 lbs 15,532.95 moment
– 4930 – 14,915.50
200 617.45

3. Determine the takeoff weight and moment by subtracting the taxi

figures from the ramp figures.
The takeoff weight is 15,900 lbs (16,100 minus 200). The takeoff
moment is 47,565.50 (48,182.95 minus 617.45).

4. Record this information on the loading schedule (Figure 4-21).

4-46 For SimuFlite training only Citation V

March 1999
 Flight Planning

Fuel Loading Weight and Moment Table Citation V Loading Schedule

Payload Computations Item Weight Moment/
(lbs) 100

Item ARM Weight Moment/ 1. Basic Empty Weight 9,390 29,051.00

(lbs) 100 * Aircraft CG = 309.04

Occupants 2. Payload 1,580 3,599.00

Seat 1 131 170 222.70 3. Zero Fuel Weight 10,970 32,650.00

(Sub-total)(Do not exceed
Seat 2 131 180 235.80 maximum zero fuel weight
of 11,200 lbs (standard)
Seat 3
or 12,200 lbs (optional).)
Seat 4 221 150 331.50
180 471.60 4. Fuel Loading 5,130 15,532.95
Seat 5 262
Seat 6 262 180 471.60 5. Ramp Weight 16,100 48,182.95
Seat 7 308 160 492.80
6. Less Fuel for Taxiing 200 617.45
Seat 9 308 170 523.60
7. ** Takeoff Weight 15,900 47,565.50
Seat 10 170 170 289.00 (Do not exceed maximum
4,930 14,915.50 takeoff weight of 15,900 lbs)
LBS * Aircraft CG = ***
Toilet 345
8. Less Fuel to Destination

Baggage 9. ** Landing Weight

(Do not exceed maximum
Nose 74.00 80 59.20 landing weight of 15,200 lbs
Tailcone * Aircraft CG = ***
434.00
Tailcone 462.00 MOMENT/100
80 369.60 * Aircraft CG = X 100
WEIGHT

Storage ** Totals must be within approved weight and center-of-gravity

Cabin 348.00 limits. It is the responsibility of the operator to ensure that
20 69.60
the aircraft is loaded properly. The Basic Empty Weight CG
is noted on the Aircraft Weighing Form. If the aircraft has
4-20 been altered, refer to the Weight and Balance Record for
Cabinet 155.00 40 62.00 information.
Contents

Payload 1,580 3,599.00 *** Enter the Center-of-Gravity Limits Envelope Graph to check
if within approved limits (shaded area).

4-21

Citation V For SimuFlite training only 4-47

March 1999
 Flight Planning

Fuel Loading Takeoff CG Limits

Weight and Use the Center of Gravity Envelope (Figure 4-22) to determine
Moment Table whether the calculated takeoff moment is within defined limits.
1. Enter the envelope from the left with the calculated takeoff weight
Loading (15,900 lbs).
Schedule 2. Move to the right to intersect the moment (47,565.50).
3. If the intersection of these values is within the CG envelope
(shaded area), the aircraft is within limits.
The intersection of the values is within takeoff CG limits.

There is a second and much more accurate method to determine

whether the proposed flight is within CG limits.
1. Divide the total moment by the total weight to derive the CG arm.
Because tmoment is divided by 100 to simplify computations,
multiply the result by 100 to derive the actual moment in inches.
When 47,565.50 is divided by 15,900, the result is 299.15.
Multiplying this by 100 results in 299.15 inches.
2. Re-enter the chart from the bottom at this number (299.15) and
move up to intersect the takeoff weight (15,900). If the intersec-
tion falls within the shaded area for maximum takeoff weight,
the proposed fight is within CG limits for takeoff.
Projecting up from 299.15 to the takeoff weight (15,900 lbs)
confirms that the aircraft is within limits for takeoff.
3. Record the takeoff CG arm (299.15) on the loading schedule
(Figure 4-23).

4-47 4-48 For SimuFlite training only Citation V

March 1999
 Flight Planning

Center-of-Gravity Limits Envelope Citation V Loading Schedule

Payload Computations Item Weight Moment/
(lbs) 100

Item ARM Weight Moment/ 1. Basic Empty Weight 9,390 29,051.00

(lbs) 100 * Aircraft CG = 309.04

47,565.50 Occupants 2. Payload 1,580 3,599.00

15,900 LBS Seat 1 131 170 222.70 3. Zero Fuel Weight 10,970 32,650.00
(Sub-total)(Do not exceed
Seat 2 131 180 235.80 maximum zero fuel weight
of 11,200 lbs (standard)
Seat 3
or 12,200 lbs (optional).)
Seat 4 221 150 331.50
180 471.60 4. Fuel Loading 5,130 15,532.95
Seat 5 262
Seat 6 262 180 471.60 5. Ramp Weight 16,100 48,182.95
Seat 7 308 160 492.80
6. Less Fuel for Taxiing 200 617.45
Seat 9 308 170 523.60
7. ** Takeoff Weight 15,900 47,565.50
Seat 10 170 170 289.00 (Do not exceed maximum
takeoff weight of 15,900 lbs)
* Aircraft CG = ***
Toilet 345
8. Less Fuel to Destination

Baggage 9. ** Landing Weight

(Do not exceed maximum
Nose 74.00 80 59.20 landing weight of 15,200 lbs
Tailcone * Aircraft CG = ***
434.00
Tailcone 462.00 MOMENT/100
80 369.60 * Aircraft CG = X 100
WEIGHT

Storage ** Totals must be within approved weight and center-of-gravity

Cabin 348.00 limits. It is the responsibility of the operator to ensure that
20 69.60
the aircraft is loaded properly. The Basic Empty Weight CG
is noted on the Aircraft Weighing Form. If the aircraft has
been altered, refer to the Weight and Balance Record for
Cabinet 155.00 40 62.00 information.
Contents

Payload 1,580 3,599.00 *** Enter the Center-of-Gravity Limits Envelope Graph to check
if within approved limits (shaded area).

4-23

299.15
4-22

Citation V For SimuFlite training only 4-49

March 1999
 Flight Planning

Center-of-Gravity Landing Weight and Moment

Limits Envelope The landing weight is the takeoff weight minus the fuel used to
Landing Weight Limit
reach the destination.
Loading Schedule 15,200 lbs
The landing weight moment is the takeoff weight moment minus
Landing weight is limited by the
most restrictive of: the difference between takeoff fuel moment and landing fuel moment.
■ maximum certified landing
1. Subtract the weight of the fuel consumed (3,150 lbs) from the
weight takeoff weight (15,900 lbs) to determine the landing weight
■ maximum landing weight
(12,750 lbs).
permitted by climb requirements
or brake energy limit 2. Subtract the fuel consumed (3,150 lbs) from the fuel load at
■ landing distance. takeoff (4,930 lbs (5130 - 200)). This is the landing fuel load
(1,780 lbs). Interpolate its moment (Figure 4-24).
After interpolation, the derived moment is 5,337.52.

3. Compute the difference between the takeoff and landing fuel

moments.
Takeoff 4930 lbs 14,915.50 moment
Landing – 1780 – 5337.52
TOTAL 3150 9577.98

4. Determine the landing weight moment by subtracting the dif-

ference between the takeoff and landing fuel moments (9,577.98)
from the takeoff weight moment (47,565.50).
The landing moment is 37,987.52.

5. Record this information on the loading schedule (Figure 4-25).

4-49 4-50 For SimuFlite training only Citation V

March 1999
 Flight Planning

Fuel Loading Weight and Moment Table Citation V Loading Schedule

Payload Computations Item Weight Moment/
(lbs) 100

Item ARM Weight Moment/ 1. Basic Empty Weight 9,390 29,051.00

(lbs) 100 * Aircraft CG = 309.04

Occupants 2. Payload 1,580 3,599.00

Seat 1 131 170 222.70 3. Zero Fuel Weight 10,970 32,650.00

(Sub-total)(Do not exceed
Seat 2 131 180 235.80 maximum zero fuel weight
of 11,200 lbs (standard)
Seat 3
or 12,200 lbs (optional).)
Seat 4 221 150 331.50
180 471.60 4. Fuel Loading 5,130 15,532.95
Seat 5 262
Seat 6 262 180 471.60 5. Ramp Weight 16,100 48,182.95
Seat 7 308 160 492.80
6. Less Fuel for Taxiing 200 617.45
Seat 9 308 170 523.60
1,780 5,337.52 7. ** Takeoff Weight 15,900 47,565.50
LBS Seat 10 170 170 289.00 (Do not exceed maximum
4,930 14.915.50 takeoff weight of 15,900 lbs)
* Aircraft CG = ***
Toilet 345
8. Less Fuel to Destination 3,150 9,577.98

Baggage 9. ** Landing Weight 12,750 37,987.52

(Do not exceed maximum
Nose 74.00 80 59.20 landing weight of 15,200 lbs
Tailcone * Aircraft CG = ***
434.00
Tailcone 462.00 MOMENT/100
80 369.60 * Aircraft CG = X 100
WEIGHT

Storage ** Totals must be within approved weight and center-of-gravity

Cabin 348.00 limits. It is the responsibility of the operator to ensure that
20 69.60
the aircraft is loaded properly. The Basic Empty Weight CG
is noted on the Aircraft Weighing Form. If the aircraft has
been altered, refer to the Weight and Balance Record for
Cabinet 155.00 40 62.00 information.
4-24 Contents

Payload 1,580 3,599.00 *** Enter the Center-of-Gravity Limits Envelope Graph to check
if within approved limits (shaded area).

4-25

Citation V For SimuFlite training only 4-51

March 1999
 Flight Planning

Fuel Loading Landing CG Limits

Weight and Use the Center of Gravity Envelope chart (Figure 4-26) to deter-
Moment Table mine whether the aircraft is within landing CG limits.
1. Enter from the left with the calculated landing weight (12,750 lbs).
Loading Schedule
2. Move across to intersect the moment (37,987.52).
3. Determine if the intersection of these values is within the
envelope.
The intersection falls within the shaded area; the aircraft is with-
in CG limits for landing.
Verify whether the proposed flight is within CG limits using the
more accurate technique.
1. Divide the landing moment (37,987.52) by the landing weight
(12,750 lbs) to derive the CG arm. Because moment is divided
by 100 to simplify computations, multiply the result by 100 to
derive the actual moment in inches.
When 37,987.53 is divided by 12,750, the result is 2.9794.
Multiplying this by 100 results in 297.94 inches.
2. Re-enter the chart from the bottom at this number (297.94) and
move up to intersect the landing weight (12,750 lbs). If the inter-
section falls within the shaded area for maximum takeoff weight,
the proposed flight is within CG limits for takeoff.
Projecting up from 297.94 to 12,750 lbs confirms that the air-
craft is within limits for landing.
3. Record the landing CG arm (297.94) on the loading schedule
(Figure 4-27).

4-51 4-52 For SimuFlite training only Citation V

March 1999
 Flight Planning

Center-of-Gravity Limits Envelope Citation V Loading Schedule

Payload Computations Item Weight Moment/
(lbs) 100

Item ARM Weight Moment/ 1. Basic Empty Weight 9,390 29,051.00

(lbs) 100 * Aircraft CG = 309.04

Occupants 2. Payload 1,580 3,599.00

Seat 1 131 170 222.70 3. Zero Fuel Weight 10,970 32,650.00

(Sub-total)(Do not exceed
Seat 2 131 180 235.80 maximum zero fuel weight
of 11,200 lbs (standard)
Seat 3
or 12,200 lbs (optional).)
Seat 4 221 150 331.50
180 471.60 4. Fuel Loading 5,130 15,532.95
Seat 5 262
Seat 6 262 180 471.60 5. Ramp Weight 16,100 48,182.95
Seat 7 308 160 492.80
6. Less Fuel for Taxiing 200 617.45
Seat 9 308 170 523.60
7. ** Takeoff Weight 15,900 47,565.50
Seat 10 170 170 289.00 (Do not exceed maximum
takeoff weight of 15,900 lbs)
* Aircraft CG = ***
12,750 Toilet 345
LBS 8. Less Fuel to Destination 3,150 9,577.98
37,987.52
Baggage 9. ** Landing Weight 12,750 37,987.52
(Do not exceed maximum
Nose 74.00 80 59.20 landing weight of 15,200 lbs
Tailcone * Aircraft CG = ***
434.00
Tailcone 462.00 MOMENT/100
80 369.60 * Aircraft CG = X 100
WEIGHT

Storage ** Totals must be within approved weight and center-of-gravity

Cabin 348.00 limits. It is the responsibility of the operator to ensure that
20 69.60
the aircraft is loaded properly. The Basic Empty Weight CG
is noted on the Aircraft Weighing Form. If the aircraft has
been altered, refer to the Weight and Balance Record for
Cabinet 155.00 40 62.00 information.
Contents

Payload 1,580 3,599.00 *** Enter the Center-of-Gravity Limits Envelope Graph to check
if within approved limits (shaded area).

4-27

297.94

4-26

Citation V For SimuFlite training only 4-53

March 1999
 Flight Planning

Center-of-Gravity
Limits Envelope

Loading Schedule

4-53 4-54 For SimuFlite training only Citation V

March 1999
 Flight Planning

Before performance can be discussed, certain terms must be clear-

ly understood. To that end, the following definitions are provided. Definitions
Accelerate-go distance – On a takeoff during which an engine fails
at V1 and the takeoff is continued, the horizontal distance from brake
release to the point at which the aircraft attains a height of 35 ft
above the runway surface. The AFM data was obtained under the fol-
lowing conditions:
■ the power was set static to the setting corresponding the AFM
Figure 4-8, then the brakes were released
■ engine failure was recognized at V1
■ the aircraft continued to accelerate to VR at which time positive
rotation to 12° nose-up pitch attitude was made. After reaching
35 ft AGL, pitch attitude was adjusted to +10° nose-up (i.e., flight
director go-around mode). After rotation a slight pull may be nec-
essary to achieve V2
■ the landing gear was retracted when a positive climb rate was
established at or above 35 ft AGL
■ V2 was maintained from the 35-foot point above the runway to
400 ft AGL
■ the aircraft then accelerated to V2 + 10 KIAS at which time the
flaps were retracted and the acceleration continued to VENR. The
climb continued to 1,500 ft AGL
■ the brake anti-skid system was on during all tests.
Accelerate-stop distance – The distance required to accelerate the
aircraft and abort the takeoff due to a failed engine, or other emer-
gency, occurring just prior to V1 with brake application commencing
at V1. The AFM data was obtained under the following conditions:
■ the power was set static to the setting corresponding the AFM
Figure 4-8, then the brakes were released
■ the pilot recognized a reason to stop (e.g., engine failure) just
prior to V1
■ maximum pilot braking effort was started at V1 and continued
until the aircraft came to a full stop
■ both throttles were brought to idle immediately after brake appli-
cation
■ directional control was maintained through the rudder pedals and
differential braking as required.

Citation V For SimuFlite training only 4-55

March 1999
 Altitude – All altitudes used in this manual are pressure altitudes
unless otherwise stated.
Calibrated airspeed (KCAS) – Indicated airspeed (knots) correct-
ed for position error.
CAT II – Category II operation. A straight-in ILS approach to the
runway of an airport under a Category II ILS instrument approach
procedure.
Climb gradient – The ratio of the change in height during a portion
of a climb to the horizontal distance transversed in the same time
interval.
Demonstrated crosswind – The demonstrated crosswind velocity
of 20 kts (measured 30 ft above the runway surface) is the velocity
of the crosswind component for which adequate control of the air-
craft during takeoff and landing was actually demonstrated during cer-
tification tests. This is not limiting. The demonstrated velocity for an
aircraft using thrust reversers is 20 kts.
Engine out accelerate-go distance – On a takeoff during which an
engine fails at V1 and the takeoff is continued, the horizontal distance
from brake release to the point at which the aircraft attains a height
of 35 ft above the runway surface.
Gross climb gradient – The climb gradient that the aircraft can
actually achieve with ideal conditions.
Indicated airspeed (KIAS) – Airspeed indicator readings (knots).
Zero instrument error is assumed.
Indicated outside air temperature (OAT) – The indicated outside
air temperature as read from the pilot’s panel.
ISA – International Standard Atmosphere.
Landing distance – The distance from a point 50 ft above the run-
way surface to the point at which the aircraft would come to a full
stop on the runway. The AFM data was obtained under the follow-
ing conditions:
■ the flaps were set at LAND, both engines were operating, the land-
ing gear was extended, and the anti-skid system was operative
■ landing preceded by a steady 3° angle approach down to the 50-
foot height point with airspeed at VREF
■ thrust setting during approach was selected to maintain the 3°
approach angle

4-56 For SimuFlite training only Citation V

March 1999
 Flight Planning

■ idle thrust was established at the 50-foot height point and throttles
remained in that setting until the aircraft stopped
■ rotation to a three-point attitude after touchdown was accom-
plished at a normal rate
■ hard wheel braking was initiated immediately on nosewheel con-
tact and continued throughout the landing roll
■ The anti-skid system was on during all tests.
Mach number – The ratio of true airspeed to the speed of sound.
Net climb gradient – The gross climb gradient reduced by 0.8%
during the takeoff phase and 1.1% during enroute. This conservatism
is required by FAR 25 for terrain clearance determination to account
for variables encountered in service.
OAT – Outside Air Temperature or Ambient Air Temperature. The
free air static temperature, obtained either from ground meteoro-
logical sources or from inflight temperature indications adjusted for
instrument error and compressibility effects.
Position correction – A correction applied to indicated airspeed or
altitude to eliminate the effect of the location of the static pressure
source on the instrument reading. No position corrections are required
when using AFM performance charts because all airspeeds and alti-
tudes in that section are presented as “indicated” values, except for
stall speeds that are presented as “calibrated” values.
RAT – Ram Air Temperature. The indicated outside air tempera-
ture as read from the pilot’s panel.
Reference zero – The point in the takeoff flight path at which the air-
craft is 35 ft above the takeoff surface and at the end of the takeoff
distance required.
Residual ice – The ice which is not completely removed from the
leading edge stagnation areas of the wing and horizontal stabilizer
by the surface anti-ice/deice system during operation in icing
conditions.
Takeoff field length – The takeoff field length given for each
combination of gross weight, ambient temperature, altitude, wind, and
runway gradients is the greatest of the following:
■ 115% of the two-engine horizontal takeoff distance from start to
a height of 35 ft above runway surface
■ accelerate-stop distance
■ engine-out accelerate-go distance.

Citation V For SimuFlite training only 4-57

March 1999
 No specific identification is made on the charts as to which of these
distances governs a specific case. In all cases considered by the charts,
the field length is governed by either the second or the third condi-
tion because the two-engine takeoff distance is always shorter.
The AFM data was obtained under the following conditions:
■ the power was set static to the setting corresponding the AFM
Figure 4-8, then the brakes were released
■ positive rotation to 12° nose-up pitch attitude was made at VR to
achieve V2. After reaching 35 + 10 by 35 ft AGL
■ the landing gear was retracted when a positive climb rate was
established above 35 ft AGL. After reaching 35 ft, pitch attitude
and flaps were set as desired.
True airspeed – The airspeed (knots) of an aircraft relative to undis-
turbed air.
VA – The maximum speed at which application of full available
aerodynamic control does not overstress the aircraft.
VAPP/VAC – The landing approach airspeed (1.3 VS1) with T.O. &
APPR flaps and landing gear up.
VENR – Single engine enroute climb speed.
VFE – Maximum flap extended speed. The highest speed permissi-
ble with wing flaps in a prescribed extended position.
VLE – Maximum landing gear extended speed. The maximum speed at
which an aircraft can be safely flown with the landing gear extended.
VLO (Extension) – Maximum landing gear extension speed. The
maximum speed at which the landing gear can be safely extended.
VLO (Retraction) – Maximum landing gear retraction speed. The
maximum speed at which the landing gear can be safely retracted.
VMCA – Minimum airspeed in the air at which directional control can
be maintained, when one engine is suddenly made inoperative. VMCA
is a function of engine thrust, which varies with altitude and tem-
perature. The VMCA presented in AFM Section II was determined for
maximum thrust.
VMCG – Minimum airspeed on the ground at which directional con-
trol can be maintained, when one engine is suddenly made inoper-
ative, using only aerodynamic controls. VMCG is a function of engine
thrust, which varies with altitude and temperature. The VMCG pre-
sented in AFM Section II was determined for maximum thrust.
VMO/MMO – Maximum operating limit speed.

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March 1999
 Flight Planning

VR – Rotation speed. The speed at which rotation is initiated during

takeoff to attain the V2 climb speed at or before a height of 35 ft
above runway surface is reached.
VREF – The airspeed equal to the landing 50-foot point speed (1.3
VS0) with full flaps and landing gear extended.
VSB – Maximum operating speed with speedbrakes extended.
VS0 – Stalling speed or the minimum steady flight speed in the land-
ing configuration.
VS1 – Stalling speed or the minimum steady flight speed obtained in
a specific configuration.
V1 – Takeoff decision speed. The distance to continue a takeoff to
35 ft does not exceed the scheduled takeoff field length if recogni-
tion occurred at V1 (accelerate-go). The distance to bring the air-
craft to a full stop (accelerate-stop) will not exceed the scheduled
takeoff field length provided that the brakes are applied at V1.
V2 – Takeoff safety speed. This climb speed is the actual speed at 35
ft above the runway surface as demonstrated in flight during takeoff
with one engine inoperative.
Visible moisture – Visible moisture includes, but is not limited to,
the following: fog with visibility less than one mile, wet snow, and
rain.
Wind – The wind velocities recorded as variables on the charts of
the AFM are understood as the headwind or tailwind components of
the actual winds at 30 ft above the runway surface (tower winds).

Citation V For SimuFlite training only 4-59

March 1999
4-60 For SimuFlite training only Citation V
March 1999
 Flight Planning

To understand performance, it is necessary to be thoroughly famil-

iar with the terms used in describing an aircraft’s efficiency. This
section reviews the definitions and terms used while determining
Performance
performance data, and presents computations required to fill out the
TOLD card. In addition, you are taken step by step through the
process necessary to determine takeoff and landing data and V speeds.

TOLD Card
TAKEOFF CITATION
Use a TakeOff and Landing Data (TOLD) card to record takeoff and ATIS
landing data. It serves as a convenient reference aid in the cockpit.
The Takeoff side of the card provides spaces for the following GROSS WEIGHT
information: V1
FLAPS TRIM
■ ATIS
VR POWER
■ V1 – Takeoff Decision Speed
T/O %
■ VR – Rotation Speed V2 CLIMB (SE) %
FLAP RETRACT CLIMB
■ V2 – Takeoff Safety Speed
VFR RWY RQD. FT.
■ VFR – Flap Retract Speed (minimum) RETURN

■ VENR – Single-Engine Enroute Climb Speed

VENR VREF
CLEARANCE
■ GROSS WEIGHT – Takeoff Weight

■ FLAPS – Takeoff Flap Setting

CALL 1-800-527-2463 FOR ADDITIONAL CARDS

■ T/O POWER – N1 Takeoff Power Setting

■ CLIMB POWER – N1 Maximum Continuous Power Setting

■ RWY RQD – Compute Takeoff Field Length

■ RETURN/VREF – Emergency Return VREF

APPROACH CITATION
ATIS
■ CLEARANCE.

The Approach side of the card provides spaces for the following GROSS WEIGHT
information: VREF
FLAPS
■ ATIS
VAPP
■ VREF – Landing Configuration 50-Foot Point Speed R R
W Q
■ VAP – Approach Target Speed VAP* Y D
FT.
FLAP RETRACT GO AROUND
■ VZF – Zero Flap Maneuvering Speed POWER
VFR
■ RWY RQD – Computed Landing Field Length
VZF (MNVR’G) %
■ GROSS WEIGHT – Landing Weight
■ FLAPS – Landing Flap Setting
■ PWR – Go Around N1 Thrust Setting VREF + 10 KIAS + WIND FACTOR
CALL 1-800-527-2463 FOR ADDITIONAL CARDS
■ VAC – Approach Climb Speed
■ NOTES.
Citation V For SimuFlite training only 4-61
March 1999
Takeoff Weight Limit Aircraft Takeoff Weight
■ 15,900 lbs Enter the gross takeoff weight determined by weight and balance
■ Maximum demonstrated (15,900 lbs) in the GROSS WEIGHT block of the Takeoff side of the
crosswind component is 20 kts. TOLD card.
Because the takeoff flap setting has not been determined, begin
filling out Takeoff TOLD cards for both 7° and 15° flap settings.
TAKEOFF CITATION
ATIS Airport Information
Airport information is obtained from the standard sources.
GROSS WEIGHT

V1 15,900 In this case, use the trip planning data provided and assume a
FLAPS TRIM
forecast runway wind of 190° and 15 kts.
7∂
VR POWER
T/O % Crosswind Component at Takeoff
V2 CLIMB (SE) %
FLAP RETRACT CLIMB Use the Cross Wind Component chart (Figure 4-28) to determine
VFR RWY RQD. FT. the wind component at takeoff.
RETURN

VENR VREF 1. First, determine the angle between the runway heading and the
CLEARANCE forecast wind direction.
With a runway heading of 170° and a forecast wind from 190°, the
resultant angle is 20°.
CALL 1-800-527-2463 FOR ADDITIONAL CARDS

2. Plot the point at which the forecast wind velocity (15 kts) inter-
sects the angular difference between the runway heading and the
forecast wind direction (20°).
APPROACH CITATION
3. Move left to the edge of the chart to obtain the headwind/tail-
ATIS
wind component (14 kts).
4. Move down from the intersection to the bottom of the chart to
GROSS WEIGHT

15,900 obtain the crosswind component (5 kts).

VREF
FLAPS The maximum demonstrated crosswind of 20 kts is not exceeded.
VAPP 15∂
R R
W Q
VAP* Y D
FT.
FLAP RETRACT GO AROUND
POWER
VFR
VZF (MNVR’G) %

VREF + 10 KIAS + WIND FACTOR

CALL 1-800-527-2463 FOR ADDITIONAL CARDS

4-62 For SimuFlite training only Citation V

March 1999
 Flight Planning

Crosswind Component
20°

14 KTS

5 KTS
4-28

Citation V For SimuFlite training only 4-63

March 1999
 Maximum Allowable Climb Weight
Use the appropriate Maximum Takeoff Weight – Pounds Permitted
Takeoff Weight Limit
by Climb Requirements tables (Figure 4-29 and Figure 4-30) to deter-
Takeoff weight is limited by the mine the maximum weight permitted by climb requirements.
most restrictive of:
■ maximum certified takeoff The AFM allows the crew to choose the flap setting for departure.
weight If runway length is not a factor, the climb gradient is greater with 7°
■ maximum takeoff weight takeoff flaps. If runway length is a factor, the takeoff length is short-
permitted by climb er with the flaps set at 15°; however, the climb gradient is less than
requirements for a flaps 7° takeoff.
■ takeoff field length. Use the tables for flaps 7° and 15° and 2,000 ft pressure altitude
to determine the maximum weight permitted by climb require-
ments using either flap setting.

1. Enter the appropriate altitude block (2,000 ft) from the left at the
ambient temperature (20°C).
2. Read the maximum takeoff weight permitted by climb require-
ments.
In this case, the maximum allowable weight is 15,900 lbs for
either flap setting.

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March 1999
 Flight Planning

Maximum Takeoff Weight – Lbs

Permitted by Climb Requirements
Flaps – 15°, Anti-Ice – OFF

4-29

Maximum Takeoff Weight – Lbs

Permitted by Climb Requirements
Flaps – 7°, Anti-Ice – OFF

4-30

Citation V For SimuFlite training only 4-65

March 1999
 Takeoff Field Length and V Speeds
Use the Takeoff Field Length – Feet tables (Figures 4-31 and Figure
4-32) to determine takeoff field length, V1, VR, V2, and VENR.
TAKEOFF CITATION 1. Enter the tables at the correct weight (15,900 lbs) and ambient
ATIS temperature (20°C).
2. Move to the right to the correct headwind column; read V1 and
GROSS WEIGHT
distance. Interpolate as necessary.
V1 101 15,900
FLAPS TRIM For conservatism, use zero wind. For flaps 7°, V1 is 101 KIAS
7∂ and the takeoff distance is 3,870 ft. For flaps 15°, V1 is 97 KIAS
VR 107
POWER
%
and the takeoff distance is 3,640 ft.
T/O
V2 115 CLIMB (SE) %
FLAP RETRACT CLIMB
3. Move further right to the VR/V2 columns. Read the required VR
VFR 125 RWY RQD. 3,870 FT. and V2.
RETURN
For flaps 7°, VR is 107 KIAS and V2 is 115 KIAS. For flaps 15°,
VENR 166 VREF VR is 103 KIAS and V2 is 111 KIAS.
CLEARANCE

4. To determine VFR, add 10 kts to V2.

CALL 1-800-527-2463 FOR ADDITIONAL CARDS
For flaps 7°, VFR is 125 KIAS. For flaps 15°, VFR is 121.

5. Read the value for VENR at the top of the weight block (166 KIAS
for both).
6. Enter the correct V speeds and required runway on the TOLD
TAKEOFF CITATION card.
ATIS
7. Check the Takeoff Correction Factors tables (Figure 4-33 and
Figure 4-34) for any needed adjustments.
GROSS WEIGHT

V1 97 15,900 In this case, no adjustments are required for runway gradient or

FLAPS TRIM anti-ice systems.
15∂
VR 103
POWER
T/O %
V2 111 CLIMB (SE) %
FLAP RETRACT CLIMB

VFR 121 RWY RQD. 3,640FT.

RETURN

VENR 166 VREF

CLEARANCE

CALL 1-800-527-2463 FOR ADDITIONAL CARDS

4-66 For SimuFlite training only Citation V

March 1999
 Flight Planning

Takeoff Field Length – Feet Takeoff Correction Factors – Flaps 15°

Flaps – 15°, 2000 Feet
CORRECTION FACTORS – RUNWAY GRADIENT

SHADED AREA NONSHADED AREA

RUNWAY
GRADIENT MULTIPLY MULTIPLY
V1* DISTANCE V1* DISTANCE
BY BY

2% UPHILL ADD 4 KNOTS 1.7 ADD 4 KNOTS 1.7

1% UPHILL ADD 2 KNOTS 1.2 ADD 2 KNOTS 1.2

1% DOWNHILL SUBTRACT 1 KNOT 1.0 NO CORRECTION 1.0

2% DOWNHILL SUBTRACT 2 KNOTS 1.0 NO CORRECTION 1.0

* IF THE ADJUSTED V1 IS GREATER THAN VR, THE VALUE OF VR MUST BE USED FOR V1.

CORRECTION FACTORS – ANTI-ICE ON

4-31 V1 – KIAS NO CORRECTION
TAKEOFF FIELD LENGTH – FEET MULTIPLY DISTANCE BY 1.10
4-33

Takeoff Field Length – Feet

Takeoff Correction Factors – Flaps 7°
Flaps – 7°, 2000 Feet
CORRECTION FACTORS – RUNWAY GRADIENT

SHADED AREA NONSHADED AREA

RUNWAY
GRADIENT MULTIPLY MULTIPLY
V1* DISTANCE V1* DISTANCE
BY BY

2% UPHILL ADD 4 KNOTS 1.7** ADD 4 KNOTS 1.7**

1% UPHILL ADD 2 KNOTS 1.2** ADD 2 KNOTS 1.2**

1% DOWNHILL SUBTRACT 3 KNOTS 1.0 NO CORRECTION 1.0

2% DOWNHILL SUBTRACT 8 KNOTS 1.0 NO CORRECTION 1.0

* IF THE ADJUSTED V1 IS GREATER THAN VR, THE VALUE OF VR MUST BE USED FOR V1.
** TAKEOFFS PROHIBITED FOR CORRECTED TAKEOFF FIELD LENGTHS GREATER THAN 11,000 FEET.

4-32 CORRECTION FACTORS – ANTI-ICE ON

V1 – KIAS NO CORRECTION
TAKEOFF FIELD LENGTH – FEET MULTIPLY DISTANCE BY 1.10
4-34

Citation V For SimuFlite training only 4-67

March 1999
 Flight Planning

Takeoff Field Obstacle Clearance

Length Assume that an engine fails at V1 and the takeoff is to be contin-
ued. Whether an obstacle in the flight path can be cleared becomes
Takeoff critical.
Correction In this example, an obstacle 300 ft high is 3,640 ft from the depar-
Factors ture end of the runway.
To determine obstacle clearance, climb gradients must be addressed.
Climb gradients are the mathematical relationship of vertical dis-
tance (i.e., “rise”) divided by the horizontal distance (i.e., “run”)
and expressed in percent.
For example, if an aircraft climbs 50 ft for every 1,000 ft cov-
ered horizontally, it has a 5.0% climb gradient.
50 x 100 = 5.0%
1000

The point at which the aircraft reaches 35 ft above the runway sur-
face is called “reference zero.” Reference zero has two factors: its
distance from brake release (i.e., takeoff field length) and its height
above the runway surface (i.e., 35 ft) (Figure 4-35 and Figure 4-36).
Shortly after reaching reference zero, the gear completes retraction.
The first segment of climb is from reference zero to gear retraction.
The second segment is from gear retraction to 400 ft AGL. If the pilot
holds V2 during the first and second climb segments, the manufac-
turer guarantees a known climb gradient (net climb gradient) for
that weight, pressure altitude, and temperature.
1. Compare the known climb gradient with the required climb gra-
dient. The required gradient is that needed to clear the obstacle.
With an engine failure at V1 and flaps 7°, the aircraft is 35 ft AGL
at 3,870 ft from brake release (i.e., takeoff field length), which is
1,430 ft from the end of the runway (5,300 minus 3,870 equals 1,430).
With an engine failure at V1 and flaps 15°, the aircraft is 35 ft AGL
at 3,640 ft from brake release (i.e., takeoff field length), which is
1,660 ft from the end of the runway (5,300 minus 3,640 equals 1,660).
Because the obstacle is 3,640 ft from the departure end of the run-
way, the distance to the obstacle from reference zero is 5,070 ft with
7° flaps (3,640 plus 1,430 equals 5,070). With flaps 15°, the obsta-
cle is 5,300 ft from reference zero (3,640 plus 1,660 equals 5,300).
The height of the obstacle is 300 ft. The aircraft is already at
35 ft AGL when it begins its climb segments; therefore, the air-
craft needs to climb an additional 265 ft (300 minus 35 equals
265) to clear the obstacle (i.e., 265 ft is the rise distance).

4-67 4-68 For SimuFlite training only Citation V

March 1999
 Flight Planning

Single Engine Takeoff Flight Path, Flaps – 15° Single Engine Takeoff Flight Path, Flaps – 7°

4-35

4-36

Citation V For SimuFlite training only 4-69

March 1999
 Flight Planning

Single Engine Obstacle Clearance (continued)

Takeoff Flight 2. Use the climb gradient equation to determine the required climb
Path gradient.
For flaps 7°, the climb gradient is 5.2%; for flaps 15°, the gra-
dient is 5.0%.
3. Use the Second Segment Takeoff Climb Gradient – Percent tables
(Figure 4-37 and Figure 4-38) to determine if the aircraft is with-
in the maximum takeoff weight for obstacle clearance. Use the
obstacle clearance diagram (Figure 4-39) for a graphic example
of this problem.
a. Enter the Second Segment Takeoff Net Climb Gradient –
Percent tables (flaps 7° and 15°) for the appropriate takeoff
weight (15,900 lbs) from the left at the pressure altitude
(2,000 ft), then move right one column and select the
correct temperature row (20°).
b. In this row, move to the wind (zero) and read the net climb
gradient (5.3% for flaps 7° and 4.3% for flaps 15°).
In this example, the Second Segment Takeoff Net Climb
Gradient – Percent charts determine that a flaps 7° takeoff
must be used to clear the 300 ft obstacle 3,640 ft from the end
of a 5,300 ft runway.
A flaps 15° takeoff results in a 4.3% net climb gradient when a
5% gradient is required. A flaps 7° takeoff results in a 5.3% net
climb gradient, when a 5.2% gradient is required.
If a flap 15° takeoff is required due to a short runway, the aircraft
weight must be reduced to 15,000 lbs to clear the obstacle.

4-69 4-70 For SimuFlite training only Citation V

March 1999
 Flight Planning

Second Segment Takeoff Net Climb Gradient – Percent Second Segment Takeoff Net Climb Gradient – Percent
Flaps – 7° Flaps – 15°

4-37 4-38

Citation V For SimuFlite training only 4-71

March 1999
 Flight Planning

Obstacle Clearance Single Engine Takeoff Flight Path Tables

An alternate method to determine obstacle clearance is to use the
Single Engine Takeoff Flight Path tables (Figure 4-40 and 4-41).
These tabulated charts in the AFM compare known obstacle heights
with various climb gradients to determine the computed distance to
clear an obstruction. Each distance is from reference zero; the pilot
must compute the “run” distance to the obstacle to compare it with
the charted value. Use the full obstacle height because the charts
were created with the 35 ft reference zero as part of the equation.
The chart is limited to 50 ft increments in obstacle heights and is not
computed for net climb gradients less than 2%. FAR 25 requires an
aircraft capability of at least a 2.4% climb gradient for takeoff. The
2.4% climb gradient is a gross climb gradient. Net climb gradients are
gross climb gradients reduced by 0.8% in the second segment climb.

Single Engine Takeoff Flight Path

Use the Single Engine Takeoff Flight Path charts to determine
the climb gradient required to clear the obstacle with one engine
operating.
The 300 ft obstacle is 5,300 ft from reference zero (flaps 15°)
and 5,070 feet from reference zero (flaps 7°).
Enter the chart at either the obstacle’s height (300 ft) or the com-
puted required gradient.
4-39 For flaps 7°, a 300 ft obstacle is cleared in 6,030 ft with a 5%
gradient.
For flaps 15°, a 300 ft obstacle is cleared in 5,790 ft with a 5%
gradient.
This confirms the calculated figures using the rise over run for-
mula that the obstacle clearance gradient would have to be
greater than 5% to clear the obstacle
Initially, it appears that the flaps 15° aircraft performs better than
the flaps 7° aircraft (5,790 vs 6,030 ft); however, for both aircraft
to achieve a 5% climb gradient, as required in this example, the
flaps 15° aircraft must be approximately 15,000 lbs and the flaps
7° aircraft must be over 15,900 lbs (see the Net Climb Gradient
charts (Figures 4-37, 4-38, and 4-39), on the previous page).
When comparing flap setting performance for both aircraft at
the same weight and atmospheric conditions, remember that a
flaps 15° aircraft always has a shorter takeoff field length; a
flaps 7° aircraft always has a greater climb gradient.

4-71 4-72 For SimuFlite training only Citation V

March 1999
 Flight Planning

Single Engine Takeoff Flight Path, Flaps – 15°

First and Second Segment, 35 ft to 400 ft

4-40

Single Engine Takeoff Flight Path, Flaps – 7°

First and Second Segment, 35 ft to 400 ft

4-41

Citation V For SimuFlite training only 4-73

March 1999
 Available vs. Required Field Length
Use the values previously computed to determine if the required
runway length is within the length of the available runway.
TAKEOFF CITATION
The available runway length is 5,300 ft. This distance exceeds
ATIS
the required runway length of 3,870 ft for the selected takeoff
flap setting of 7°.
GROSS WEIGHT

V1 101 15,900
FLAPS TRIM Emergency Return
7∂
VR 107
POWER If an emergency return to the departure airport becomes necessary,
V2
T/O %
determine VREF from the Landing Distance – Feet table (Figure 4-42).
115 CLIMB (SE) %
FLAP RETRACT CLIMB
For operation under FAR Part 135 the landing distance is
VFR 125 RWY RQD. 3,870 FT. divided by 0.6 to derive a factored landing distance for safety
RETURN
considerations.
VENR 166 VREF 106
CLEARANCE
For an emergency return at just under 15,900 lbs the VREF is
108 KIAS with a landing distance of 3,260 ft. Factored by 0.6
the distance is 5,433 ft.
CALL 1-800-527-2463 FOR ADDITIONAL CARDS
If a fuel burn of 700 lbs was used from takeoff to landing, to get
to the maximum landing weight of 15,200 lbs, the VREF would
be 106 KIAS with a landing distance of 3,100 ft. Factored by 0.6
the distance is 5,166 ft.
From this example using the FAR Part 135 factor, the aircraft
landing weight must be lowered to meet the 5,300 ft of runway
length.
Enter the Emergency Return VREF of 106 on the TOLD card.

4-74 For SimuFlite training only Citation V

March 1999
 Flight Planning

Landing Distance – Feet

Actual Distance, Flaps – Full, 2000 Feet

4-42

Citation V For SimuFlite training only 4-75

March 1999
 Takeoff Thrust Setting
Consult the Takeoff/Go Around Thrust Setting chart (Figure 4-43) to
Takeoff Thrust Limits
determine the correct takeoff thrust setting.
N2 . . . . . . . . . . . . . 96%
N1 . . . . . . . . . . . . 104% 1. Enter the chart from the bottom at the ambient temperature (20°C)
and move up to the appropriate anti-ice/environmental system
line or the pressure altitude line, whichever is first.
Maximum Continuous Thrust
The Environmental Systems On Anti-Ice Systems Off line
Limits occurs first.
N2 . . . . . . . . . . . . . 96% 2. Move left to the edge of the chart and read the fan (N1) speed
N1 . . . . . . . . . . . . 104% required (98.7%). This becomes the takeoff thrust limit for these
conditions.
3. Enter this information on the TOLD card.

Single Engine Enroute Thrust Setting

TAKEOFF CITATION Use the Maximum Continuous Thrust Setting chart (Figure 4-44) to
ATIS determine the single-engine climb thrust setting.
1. Enter the chart from the bottom at the ram air temperature (20°C)
GROSS WEIGHT and move up to the appropriate anti-ice/environmental systems line
V1 101 15,900 or pressure altitude line, whichever is first.
FLAPS TRIM
7∂ 2. Move left to the edge of the chart and read the fan speed (95.8%).
VR 107
POWER
T/O
98.7 %
3. Enter this information on the TOLD card.
V2 115 CLIMB (SE)95.8 %
FLAP RETRACT CLIMB

VFR 125 RWY RQD. 3,870 FT.

RETURN

VENR 166 VREF 106

CLEARANCE

CALL 1-800-527-2463 FOR ADDITIONAL CARDS

4-76 For SimuFlite training only Citation V

March 1999
 Flight Planning

Takeoff/Go Around Thrust Setting Maximum Continuous Thrust Setting

Enroute Climb

98.7%

95.8%

4-43 4-44

Citation V For SimuFlite training only 4-77

March 1999
 Flight Planning

Takeoff/Go Gross Weight at Destination

Around Thrust Consult the Weight and Balance form to obtain the landing weight
Landing Weight Limit
Setting of the aircraft at the destination airport (12,750 lbs). Enter this weight
15,200 lbs in the GROSS WEIGHT box of the TOLD card.
Landing weight is limited by the
Maximum most restrictive of:
Continuous ■ maximum certified landing
Crosswind Component at Destination
Thrust Setting weight Use the Crosswind Component chart (Figure 4-45) to determine the
■ maximum landing weight wind component.
permitted by climb requirements
or brake energy limit 1. Determine the angle between the runway heading and the forecast
■ landing distance
wind direction.
■ maximum demonstrated cross- With a runway heading of 140° and a forecast wind from 130°, the
wind component is 20 kts. resultant angle is 10°.
2. Plot the point at which the forecast wind velocity (20 kts) inter-
sects the difference between the runway heading and the fore-
cast wind direction (10°).
APPROACH CITATION
ATIS
3. Move left to the edge of the chart to obtain the headwind/tail-
wind component (20 kts).

GROSS WEIGHT
4. Move down to the bottom of the chart from the intersection to
VREF 12,750 obtain the crosswind component.
FLAPS The crosswind component is less than 3.5 kts; for conservatism,
VAPP FULL use 4 kts.
R R
W Q
VAP* Y D
FT.
The maximum demonstrated crosswind component of 20 kts is
FLAP RETRACT GO AROUND not exceeded.
POWER
VFR
VZF (MNVR’G) %

VREF + 10 KIAS + WIND FACTOR

CALL 1-800-527-2463 FOR ADDITIONAL CARDS

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 Flight Planning

Crosswind Component

10°

20 KTS

3.5 KTS

4-45

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March 1999
 NOTE: Refer to Quick Reference
Maximum Allowable Landing Weight
chapter for approach and landing in
icing conditions limitations.
Use the Maximum Landing Weight Permitted by Climb Requirements
or Brake Energy Limits – Pounds table (Figure 4-46) to determine the
maximum allowable landing weight under these conditions.
1. Enter the appropriate altitude block (1,000 ft) from the left at the
correct ambient temperature (25°C).
2. Move to the right to the appropriate wind component block
(20 kts) and, within the block, to the appropriate runway gradient
column (2%). Read the maximum landing weight permitted
(15,200 lbs).
The landing weight for this example, 12,750 lbs, is well below
the 15,200 lb limit.

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 Flight Planning

Maximum Landing Weight Permitted by Climb

Requirements or Brake Energy Limits – Lbs
Flaps – 15°

4-46

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March 1999
 Landing Distance and Approach Speeds
Use the Landing Distance – Feet table (Figure 4-47) to determine
L = R VREF, VAPP, and runway length required by weight and ambient
0.6 conditions.
where: Interpolate between the 13,000 and 12,500 lb blocks because the
landing weight is 12,750 lbs.
L = Landing Distance
0.6 = FAR 135 Factor 1. Enter the block for the appropriate altitude (1,000 ft) and weight
(13,000 and 12,500 lbs) from the left at the correct ambient tem-
R = Minimum Runway perature (25°C).
Distance Required prior
to Dispatch 2. Move to the right until the applicable wind component is reached
(zero). Use zero for added safety.
3. Read the landing distance; interpolate if necessary.
The landing distance is 2,555 ft.
APPROACH CITATION
ATIS
4. To meet FAR 135 requirements or for an additional margin of
safety, divide the landing distance by 0.6.
2555 ft = 4258 ft
GROSS WEIGHT
97 12,750 0.6
VREF
FLAPS
5. From the top of the block, just below the weight, read VREF and
VAPP 104 FULL VAPP/VAC speeds. Interpolate as necessary.
R R

112
W Q
For 12,500 lbs, VREF is 96 KIAS and VAPP/VAC is 103 KIAS. For
VAP* Y D
4,285 FT.
FLAP RETRACT GO AROUND
13,000 lbs, VREF is 98 KIAS and VAPP/VAC is 105 KIAS.
VFR
POWER
Interpolation for 12,750 lbs determines that VREF is 97 KIAS and
127 VAPP/VAC is 104 KIAS.
VZF (MNVR’G) %
6. The formula for VAP, the approach target speed, is VREF + 10
KIAS + (the headwind component and 1/2 of the gust factor). The
total adjustment may not exceed VREF + 20 KIAS.
VREF + 10 KIAS + WIND FACTOR
CALL 1-800-527-2463 FOR ADDITIONAL CARDS 97 + 10 + 5 + 0 = 112 KIAS

7. Find VZF by adding 30 kts to VREF.

97 + 30 = 127

8. Enter this information on the TOLD card.

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 Flight Planning

Landing Distance – Feet

Actual Distance, Flaps – FULL, 1000 Feet

4-47

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 Go-Around/Takeoff Thrust Setting
Use the Takeoff/Go Around Thrust Setting chart (Figure 4-48) to
determine the takeoff thrust setting.
APPROACH CITATION
ATIS
1. Enter the graph from the bottom at the correct ambient tempera-
ture (25°C). Move up to the appropriate anti-ice/environmental
systems or the pressure altitude line, whichever is first.
GROSS WEIGHT

VREF 97 12,750 2. Move left to the edge of the chart to obtain the fan speed
FLAPS (98.1% N1).
VAPP 104 FULL
R R
3. Enter this number in the GO AROUND PWR block of the TOLD
VAP* 112
W Q
Y D card.
4,285 FT.
FLAP RETRACT GO AROUND
POWER
VFR
127
VZF (MNVR’G) 98.1 %

VREF + 10 KIAS + WIND FACTOR

CALL 1-800-527-2463 FOR ADDITIONAL CARDS

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 Flight Planning

Takeoff/Go Around Thrust Setting

98.1%

25°C
4-48

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March 1999
4-86 For SimuFlite training only Citation V
March 1999
 Flight Planning

This section provides information for specific instances outside

normal routines addressed elsewhere in this manual. These procedures Supplemental
provide an awareness and understanding of situations that may be
encountered and should not be confused with normal performance
Information
procedures.

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March 1999
 General
There is no relationship between the information computed in the pre-
vious example and the information exhibited below. Consider each
chart independently and based on the conditions given for the chart.

Approach and Landing Gross Climb

Gradients
These tables present the percentage actual climb gradient of the air-
craft given the conditions at the top of each chart. This information
was derived from flight test data.
The difference between the approach and landing tables is stated in
the conditions at the top of each chart.
The following procedure applies to both the Approach Gross Climb
Gradient-Percent table (Figure 4-49) and Landing Gross Climb
Gradient-Percent table (Figure 4-50).
Assume the following conditions.
Landing Field Altitude . . . . . . . . . . . . . . . 3,000 FT
OAT . . . . . . . . . . . . . . . . . . . . . . . . . . . -10°C
Landing Weight . . . . . . . . . . . . . . . . . . 14,500 LBS
Wind Component . . . . . . . . . . . . . 20 KT HEADWIND
Anti-Ice . . . . . . . . . . . . . . . . . . . . . . . . . . ON

1. Enter table from the left at the landing field altitude (3,000 ft).
2. Following the appropriate temperature (-10°C), move to the
right to the appropriate landing weight and headwind column
(14,500 lbs and 20 kt headwind).
3. Read the indicated climb gradients.
The approach climb gradient initially is 8.9%; because the anti-
ice systems are on, this is reduced by 1 to 7.9%.
The landing climb gradient initially is 23; because the anti-ice
systems are on, this is reduced by 3 to 20%.

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 Flight Planning

Approach Gross Climb Gradient – Percent Landing Gross Climb Gradient – Percent
Flaps – 15° Flaps – Full

4-49 4-50

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 Flight Planning

Approach Gross Performance for Contaminated Runways

Climb Gradient – Use the figures in Table 4-B to identify changes in takeoff perfor-
Contaminated Runway
Percent Limitations mance when operating on contaminated runways. Observe the
■
following AFM notes.
Maximum Takeoff Weight –
Landing Gross 15,900 lbs 1. Avoid takeoffs and landings with actual headwinds exceeding
Climb Gradient – ■ Maximum Landing Weight – 20 kts or actual tailwinds exceeding 10 kts.
Percent 15,200 lbs
2. Normal takeoff V1 speeds and takeoff distances with 15° flaps
are determined from AFM Section IV, Figure 4-17 for takeoff.
Takeoff field lengths for 7° flaps or anti-ice systems on should not
be used.
3. The published limiting maximum tailwind component for this
aircraft is 10 kts; however, Cessna does not recommend landings
on precipitation-covered runways with any tailwind component.
If a tailwind landing cannot be avoided, add 0.25 to all landing fac-
tors. Normal landing distances are determined from AFM Section
IV, Figure 4-32.
4. Takeoffs with tailwinds in water, add 0.10 to factor.
5. Takeoffs with tailwinds should not be attempted in snow.

Runway Depth of Takeoff Landing

Condition Precipitation

Multiply Multiply
Takeoff Field Landing
Length By Distance By

Wet Less than 0.01 inch 1.45 1.40

Water

Water Less than 0.4 inch 2.40 2.20

Water
Slush Less than 0.5 inch 2.35 2.05

Loose, Dry 1.0 inch 3.50 2.50

Snow
2.0 inches 2.80 2.10

Ice Unknown 2.60

Table 4-B; Performance for Contaminated Runways

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 Flight Planning

Simplified Takeoff Criteria

Cessna supplies simplified takeoff data that applies to the majority
of situations where runway length is greater than that required for this
aircraft. The tabulated data gives more exact performance criteria
through a range of conditions that include all but the most extreme
cases.
Most takeoffs result in field length margins that allow the use of a
single set of values for takeoff speeds and power settings. If the
following conditions are met, the simplified procedures may be used:
■ no obstacle in the flight path
■ anti-ice systems off
■ flaps set 15°
■ takeoff field length available of 4,700 ft or longer
■ no tailwind
■ no runway gradient.
Simplified criteria utilizes the values shown in Table 4-C.

Weight 15,900 lbs or less 15,000 lbs or less 13,500 lbs or

less
Altitude of 2,000 ft or below 4,000 ft or below 6,000 ft or below
Airport

Ambient Between 10 and Between 12 and Between 9 and

Temperature 41°C 39°C 42°C

V1 99 KIAS 97 KIAS 92 KIAS

VR 103 KIAS 100 KIAS 95 KIAS

V2 111 KIAS 108 KIAS 102 KIAS

Single Engine 166 KIAS 163 KIAS 158 KIAS

Climb Speed

Takeoff Fan 96.2% 96.5% 96.0%

Single Engine 92.7% 93.0% 92.6%

Climb Fan

Table 4-C; Simplified Takeoff Criteria

When conditions are other than those specified in the simplified

criteria, the appropriate tabulated data must be referred to.

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 Single Engine Enroute Net Climb
Gradient – Percent
If the aircraft is in a single-engine condition maintaining the approach
VENR, the Single Engine Enroute Net Climb Gradient - Percent
charts (Figure 4-51 and Figure 4-52) show the net climb gradient for
the aircraft.

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 Flight Planning

Single Engine Enroute Net Climb Gradient – Percent Single Engine Enroute Net Climb Gradient – Percent
Flaps – Up Flaps – Up

4-52

4-51

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 Flight Planning

Single Engine Single Engine Takeoff Flight Path Distances

Enroute Net The data presented in these tables determines the horizontal distance
Climb Gradient – in the climb from 35 to 1,500 ft; the AFM has tables for all takeoff
Percent altitudes. All of these tables function in the same manner.
Review the flight segments (Figures 4-53 and 4-54) at this time. Each
has a distinctive configuration as shown in Table 4-D.

Flight Number of Thrust Flaps Gear

Segment Operating Setting
Engines (Degrees)

First Segment 1 Takeoff 7° (T.O.) or Down

Takeoff Climb 15° (T.O. &
Appr.)

Second 1 Takeoff 7° (T.O.) or Up

Segment 15° (T.O. &
Takeoff Climb Appr.)
Final 1 Maximum 0° Up
Segment Continuous
Climb

Enroute Climb 1 Maximum 0° Up

Continuous

Approach 1 Takeoff 15° (T.O. & Up

Climb Appr.)

Landing Climb 2 Takeoff 35° (Land) Down

Table 4-D; Takeoff Flight Segments

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 Flight Planning

Single Engine Takeoff Flight Path, Flaps – 15° Single Engine Takeoff Flight Path, Flaps – 7°

4-53 4-54

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March 1999
 Flight Planning

Single Engine With the flight segments in mind, the flight path distances for this fol-
lowing example can be determined from the Single Engine Takeoff
Takeoff Flight Flight Path Distances table (Figure 4-55 and Figure 4-56).
Path, Flaps – 15° Assume the following conditions.
Ambient Temperature . . . . . . . . . . . . . . . . . . . 5°C
Single Engine
Takeoff Flight Flaps . . . . . . . . . . . . . . . . . . . . . . . . . 7° or 15°
Path, Flaps – 7° Pressure Altitude . . . . . . . . . . . . . . . . SEA LEVEL
Gross Takeoff Weight . . . . . . . . . . . . . . 15,000 LBS
Wind . . . . . . . . . . . . . . . . . . . . . . . . . . NONE
Anti-Ice Systems . . . . . . . . . . . . . . . . . . . . . OFF

1. Enter the table from the left at the gross weight (15,000 lbs).
Move right to the next column to the row for the ambient tem-
perature (5°C).
2. Move right to the block for the applicable wind component (zero
wind). Read the applicable distances.
For flaps 7°, the cumulative distances are shown below.
First and Second Segments . . . . . . . . . . . . . 5,227 FT
Third Segment (Including First
and Second Segments) . . . . . . . . . . . . . . . 14,173 FT
Final Segment (Including First, Second
and Third Segments . . . . . . . . . . . . . . . . 31,607 FT
For flaps 15°, the cumulative distances are shown below.
First and Second Segments . . . . . . . . . . . . . 6,035 FT
Third Segment (Including First
and Second Segments) . . . . . . . . . . . . . . . 17,088 FT
Final Segment (Including First, Second,
and Third Segments) . . . . . . . . . . . . . . . . 34,522 FT

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 Flight Planning

Single Engine Takeoff Flight Path Distances Single Engine Takeoff Flight Path Distances
Flaps – 15°, Sea Level Flaps – 7°, Sea Level

4-55 4-56

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 Flight Planning

Single Engine
Takeoff Flight
Path Distances

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Several chapters contain multiple systems to facilitate a more
coherent presentation of information. The systems covered are Systems
listed below in alphabetical order opposite the chapter in which
they are located. ATA codes are noted in parentheses.

SYSTEM (ATA Code) CHAPTER

Aircraft Structure (51) . . . . . . . . AIRCRAFT OVERVIEW
Air Conditioning (21) . . . . . . . . . . . . . . . PNEUMATIC
APU (49) . . . . . . . . . . . . . . . . . . . . POWERPLANT
Brakes (32) . . . . . . . . . . . . . . . . . . LANDING GEAR
Communications (23) . . . . . . . . . . . . . . . . AVIONICS
Dimensions and Areas (6) . . . . . . AIRCRAFT OVERVIEW
Doors (52) . . . . . . . . . . . . . . AIRCRAFT OVERVIEW
Electrical (24) . . . . . . . . . . . . . . . . . . ELECTRICAL
Equipment/Furnishings (25) . . . . . AIRCRAFT OVERVIEW
Engine (71) . . . . . . . . . . . . . . . . . . . POWERPLANT
Engine Controls (76) . . . . . . . . . . . . . . POWERPLANT
Engine Fuel and Control (73) . . . . . . . . . POWERPLANT
Engine Indicating (77) . . . . . . . . . . . . . POWERPLANT
Fire Protection (26) . . . . . . . . . . . . . . . . . . . . . FIRE
Flight Controls (27) . . . . . . . . . . . FLIGHT CONTROLS
Fuel (28) . . . . . . . . . . . . . . . . . . . . . . . . . . FUEL
Fuselage (53) . . . . . . . . . . . . . AIRCRAFT OVERVIEW
Hydraulics (29) . . . . . . . . . . . . . . . . . HYDRAULICS
Ice and Rain Protection (30) . . . . . . . . . . ICE AND RAIN
Ignition (74) . . . . . . . . . . . . . . . . . . POWERPLANT
Landing Gear (32) . . . . . LANDING GEAR AND BRAKES
Lighting (33) . . . . . . . . . . . . . . . . . . . ELECTRICAL
Navigation (34) . . . . . . . . . . . . . . . . . . . AVIONICS
Oil (79) . . . . . . . . . . . . . . . . . . . . . POWERPLANT
Oxygen (35) . . . . . . . . . . . . . . . . MISCELLANEOUS
Pitot/Static (34) . . . . . . . . . . . . . . . . . . . AVIONICS
Pneumatic (36) . . . . . . . . . . . . . . . . . . PNEUMATIC
Pressurization (21) . . . . . . . . . . . . . . . . PNEUMATIC
Stabilizers (55) . . . . . . . . . . . . AIRCRAFT OVERVIEW
Stall Warning (27) . . . . . . . . . . . . FLIGHT CONTROLS

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 Thrust Reversers (78) . . . . . . . . . . . MISCELLANEOUS
Warning Lights (33) . . . . . . . . . . . . MISCELLANEOUS
Windows (56) . . . . . . . . . . . . . AIRCRAFT OVERVIEW
Wings (57) . . . . . . . . . . . . . . AIRCRAFT OVERVIEW

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This section presents an overview of the Cessna Citation 560
aircraft. It includes major features, airframe structures, dimen- Aircraft
sions, and danger areas, as well as a list of service bulletins
referenced in this manual. Overview
This manual references the manufacturer’s serial numbers and,
where system differences warrant, it publishes separate data
and schematics. Chapter 5A

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5A-2 For SimuFlite training only Citation V
March 1999
 Aircraft Overview

Airframe Description . . . . . . . . . . . . . . . . . 5A-5

Table of
Engine . . . . . . . . . . . . . . . . . . . . . . . . . 5A-5
Firewall . . . . . . . . . . . . . . . . . . . . . . . 5A-5
Contents
Fuselage . . . . . . . . . . . . . . . . . . . . . . . 5A-5
Nose Section . . . . . . . . . . . . . . . . . . . . 5A-5
Nose Baggage Compartment . . . . . . . . . . . 5A-5
Pressurized Compartment . . . . . . . . . . . . . . 5A-6
Cockpit . . . . . . . . . . . . . . . . . . . . . . . 5A-6
Entrance Door . . . . . . . . . . . . . . . . . . . 5A-7
Optional Cargo Door . . . . . . . . . . . . . . . . 5A-9
Emergency Exit . . . . . . . . . . . . . . . . . . 5A-10
Passenger Compartment . . . . . . . . . . . . . 5A-11
Refreshment Centers . . . . . . . . . . . . . . . 5A-11
Aft Cabin Baggage Compartment . . . . . . . . 5A-12
Waste Facilities . . . . . . . . . . . . . . . . . . 5A-12
Cabin Windows . . . . . . . . . . . . . . . . . . 5A-12
Tailcone . . . . . . . . . . . . . . . . . . . . . . 5A-12
Forward Tailcone Compartment . . . . . . . . . 5A-13
Tailcone Baggage Compartment . . . . . . . . . 5A-13
Wing . . . . . . . . . . . . . . . . . . . . . . . . . 5A-14
Fuel Tanks . . . . . . . . . . . . . . . . . . . . . 5A-14
Empennage . . . . . . . . . . . . . . . . . . . . . 5A-15

Aircraft Dimensions . . . . . . . . . . . . . . . . 5A-17

Danger Areas . . . . . . . . . . . . . . . . . . . . 5A-19

Radar . . . . . . . . . . . . . . . . . . . . . . . . 5A-19
Engine Inlet Air Draw . . . . . . . . . . . . . . . . 5A-19
Engine Exhaust Plume . . . . . . . . . . . . . . . . 5A-19

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March 1999
Aircraft Features

DORSAL FIN
AIR INLET
AVIONICS EQUIPMENT ACCESS
RUDDER SERVO
AND TRIM TAB

AFT COMPARTMENT/BAGGAGE
ACCESS DOOR (LEFT SIDE ONLY)
FOUL WEATHER WINDOW
NOSE LANDING GEAR (LEFT SIDE ONLY)
MAIN LANDING GEAR

SPEEDBRAKE (UPPER AND LOWER)

EMERGENCY EXIT

ELEVATOR TRIM

CABIN DOOR

AILERON TRIM TAB (LEFT SIDE ONLY)

NAVIGATION/ ANTI-COLLISION LIGHT

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March 1999
 Aircraft Overview

The Cessna Citation V is certified in

accordance with FAR Part 25 airwor-
Firewall
Stainless steel outboard ribs of the
Airframe
thiness standards for two pilots, IFR,
VFR, day, night, and flight into known
pylon frame form a vapor barrier
between the engine and pylon. Fire-
Description
icing conditions. The low-wing, twin-
proof sealant is in the upper and lower
engine turbofan aircraft is certified to
pylon skin. In addition, seals separate
an operating altitude of 45,000 ft. The
the pylons from the thrust reverser fair-
standard model without the increased
ings. Fuel and hydraulic lines mount
zero fuel weight has airspeed limits of
into the firewall seal plates to provide
261 KIAS below 8,000 ft, 292 KIAS
a disconnect area for those lines.
from 8,000 to 28,890 ft, and 0.755
Mach from 28,890 to 45,000 ft. The
maximum takeoff weight is 15,900 lbs.
Fuselage
Low take-off and landing speeds per- The fuselage (Figure 5A-1) is an all
mit operation at small and unimproved metal, semi-monocoque structure con-
airports. sisting of a nose section, a pressurized
The aircraft has hydraulically powered compartment, and a tailcone.
landing gear, flap, speedbrake, and
Nose Section
thrust reverser operation through an
open center hydraulic system. A sepa- The unpressurized nose section con-
rate closed hydraulic system operates tains a 16-cubic foot baggage com-
the main gear wheelbrakes. In addition, partment (less space in aircraft with
a backup pneumatic system can operate additional avionics options), an avion-
landing gear extension and braking. ics compartment accessible through the
removable radome, the nose gear wheel
Engine well, the power brake accumulator
gage, and various reservoirs.
Two Pratt and Whitney Canada Inc.
JT15D-5A turbofan engines on the rear Nose Baggage Compartment
fuselage each produce 2,900 lbs of
Two top-hinged, swing-up doors
thrust and incorporate ice protection
(Figure 5A-2), one on each side of the
and fire detection/extinguishing
nose section, are key-locked and latched
systems. Conventional piggy-back con-
at the bottom. When either nose com-
trols on the throttles operate target-type
partment door is fully open, a door
thrust reversers (see the Miscellaneous
stop/latch on the aft door hinge engages
chapter).

5A-2

5A-1

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March 1999
 to hold the door open. To close the door, doors out of the up and latched posi-
push the release pad on the latch and tion removes power from the light.
lower the door.
Pressurized Compartment
On the bottom of each door, two latch-
es (forward and aft) and a center key- The pressurized compartment com-
lock hold the door closed. The nose prises the cockpit and passenger area.
compartment door key-locks connect The main entry door is on the left fuse-
to the DOOR NOT LOCKED circuit, lage between the cockpit and passenger
along with those in the aft baggage and cabin.
cabin entrance door. If any one or more
Cockpit
of the doors in the circuit is not locked,
the DOOR NOT LOCKED annuncia- The two-crew cockpit (Figure 5A-4)
tor illuminates. has dual controls and a conventional
Two microswitches, one in each door instrument panel with center pedestal.
stop/latch, and a lighted manual rock- Both pilot seats have a five-point
er switch control the nose baggage restraint system and are adjustable for-
compartment light (Figure 5A-3) at the ward, aft, vertically, and tilted.
top center of the baggage area. The Hot The cockpit windows include the left
Battery bus powers the light. When and right windshields, pilot and copilot
either door is fully open, the door’s side windows, and the pilot’s foul
stop/latch engages a microswitch that weather window.
sends power to the manual rocker Each windshield (Figure 5A-5) com-
switch. A small light in the rocker prises three laminated layers: an outer
switch lens illuminates so the switch pane of stretched acrylic 0.414 inch
can be seen easily in the dark. thick; a vinyl layer 0.05 inch thick; and
Moving the manual switch to ON illu- an inner pane of stretched acrylic 0.414
minates the baggage light. Moving the inch thick. The pilot and copilot side
manual switch to OFF or releasing both windows, aft of the windshields on each

5A-3

5A-4

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March 1999
 Aircraft Overview

side of the fuselage, consist of a 0.45- door. Bleed air from the pneumatic sys-
inch thick prestressed Plexiglas outer tem inflates the door seal. A male fit-
pane, an air space, and a 0.25-inch thick ting mounted on the fuselage in the
pre-stressed Plexiglas inner pane. hinge area mates with a female fitting
The pilot’s foul weather window (Fig- in the hinge when the door is closed.
ure 5A-6), on the left side between the As the lower forward latch pin enters its
windshield and the pilot’s side window, fuselage door frame socket, a bleed air
provides ventilation and access for a valve in the forward door frame opens
ground heating/cooling hose. It also to allow air to enter the primary seal. As
enables the pilot to communicate with the door latch retracts, the bleed air
ground crew personnel. The foul weath- valve dumps the air in the primary seal
er window consists of three laminated to allow the door to open. A pressure
layers: an outer panel of stretched switch illuminates the DOOR SEAL
acrylic 0.276 inch thick; a vinyl layer annunciator when pressure is less than
0.05 inch thick; and an inner panel of 23 PSI.
stretched acrylic 0.092 inch thick. The A secondary (passive) door seal is
window opens inward by releasing a installed around the outer edge of the
latch at the top and lowering the pane on door under the primary seal. Cabin pres-
its bottom-mounted hinge. sure extrudes the secondary seal into
Entrance Door position. With loss of the primary seal
(i.e., DOOR SEAL light illuminated),
The flush-fitting entrance door on the the aircraft is limited to 41,000 ft.
left forward side of the aircraft (Figure
5A-7, following page) incorporates the
When the engines are operating, pneu-
forward cabin window. The door key- matic air inflates a small rain seal
locks from the outside and attaches to mounted on the fuselage at the hinge
the fuselage by a single hinge. to seal the area around the hinge.
A primary inflatable door seal of mold- Handles inside and outside the door
ed rubber is installed in a retainer that share a common shaft for opening and
forms a continuous ring around the closing the door. The handles remain

5A-6

5A-5

Citation V For SimuFlite training only 5A-7

March 1999
 stowed by spring tension. A series of 12 The door key-locks from the outside. A
locking pins that extend from the door hinged double-step stair is separate
into corresponding fuselage door frame from the door, mounted at the bottom
sockets hold the door closed; there are of the door frame. To extend the stair,
four sets of locking pins/sockets on pull out and down, then unfold the sec-
each side (forward and aft) of the door ond step to provide convenient entry.
and two sets each at the top and bot- To open the door from inside, depress
tom. The lower forward fuselage sock- the trigger on the handle to release the
et houses the bleed air valve that handle from the striker plate. Rotate
inflates the primary door seal and a the handle counterclockwise. The lock-
microswitch that connects to the ing linkage retracts the locking pins
DOOR NOT LOCKED circuit. from the fuselage door frame sockets,
Visual inspection of the linkage can be and the bleed air valve in the fuselage
made through five inspection windows. forward door frame dumps air from the
A small round viewing port is next to inflatable door seal. The door warning
each of the four top and bottom latch- microswitch illuminates the DOOR
es; a green dot on a white background NOT LOCKED annunciator, while the
is present in each of the four ports when door seal pressure switch illuminates
the latches are secure. A fifth rectan- the DOOR SEAL annunciator. Push the
gular window is next to the base of the door outward to the full open position
door handle; two green bars are in line until a spring-loaded door stop holds
with each other when the door is safe. the door in the open position and return
the handle to the stowed position.

5A-7

5A-8 For SimuFlite training only Citation V

March 1999
 Aircraft Overview

To open the door from outside the air- in place of the standard passenger/crew
craft, push the end of the handle entrance door. The cargo door consists
marked PUSH until the opposite end of two doors, an upper and lower, that
of the handle clears the socket. Rotate attach to the fuselage with hinge pins.
the handle clockwise and pull the door The locking mechanism for each door
outward as described above. operates independently.
To close the door, press the door stop The upper door has a gas spring exten-
lever to release the door stop lug. Use der on each side of the door. The exten-
the door-assist chain to pull the door ders automatically open the door after
closed from inside the aircraft. Close it is unlocked. A locking mechanism
the door and rotate the inside handle in the forward extender holds the door
clockwise (outside handle counter- open during gusty conditions.
clockwise) until the door lock pins A gas spring snubber on the forward
engage the fuselage door sockets. side of the lower door allows the lower
Return the handle to its stowed posi- door to open slowly. Two cables, one
tion. Check the door warning annun- on each side of the door, support the
ciators in the cockpit. lower door when it is fully open. The
Optional Cargo Door cables connect at the fuselage, then
route inside the door where they con-
An optional 35.12-inch cargo entrance nect to reels that take up the cable slack
door (Figure 5A-8) may be installed when the door is closed.

UPPER
GAS SPRING CABIN
EXTENDERS DOOR

LATCH
FITTINGS

INFLATABLE
DOOR SEAL

DOOR CABLE

GAS
SPRING
FOLD OUT SNUBBER
STEPS

LOWER CORNER
CABIN SEAL
DOOR

5A-8

Citation V For SimuFlite training only 5A-9

March 1999
 Steps built into the lower door auto- back of each fitting is visible only if
matically fold out and back again via a the fitting is locked. In addition, a win-
cable and spring as the door opens and dow next to the upper door handle dis-
closes. The cable attaches to the fuse- plays aligned reference marks and the
lage door frame, runs down through words “DOOR CLOSED” when the
the door, through eye bolts in the step, door handle is the closed position.
then connects to a spring at the bottom Five microswitches, four in the upper
of the door. door and one in the lower door, are tied
The locking mechanism for each door to the DOOR NOT LOCKED electri-
operates the same, although there are cal circuit. If the upper and lower doors
some mechanical differences. The upper are not properly closed and latched, the
door locking mechanism consists of a DOOR NOT LOCKED annunciator
cable, cable pulleys, push rods, latch illuminates.
pins, bellcranks, cable drums, and latch
fittings. The lower door locking mech- Emergency Exit
anism consists of two cables, cable pul- The plug-type emergency exit door
leys, cable drums, and latch fittings. (Figure 5A-9) on the right forward side
As the door handle on either door of the cabin (opposite entry door) is
moves from one position to the other installed from inside the aircraft and
(open or closed), the cable(s) inside the opens inward for quick egress. It is lock-
door rotates the latch fitting around the able from the inside only (Figure
latch post on the fuselage door frame. 5A-10). The emergency exit door incor-
When both doors are closed and the porates the right forward cabin window.
upper handle is rotated, arms in the A latch pin on the top and two retainers
lower door and pins in the upper door at the bottom of the door hold it in place.
lock together. Cabin air pressure acting against the
The cargo doors have two seals: a door in flight forces the plug door firm-
pneumatically inflated seal around the ly against the fuselage door frame. A
fuselage door frame, and a standard rubber bulb-type seal around the edge of
rubber seal on the lower door that pro- the door provides a pressure seal. Holes
vides a positive seal between the doors. through the inner surface of the seal and
in the door frame allow cabin air pres-
Inspection holes around the upper and sure to inflate the seal. The flush out-
lower door frame allow visual inspec- side handle shares a common latch cam
tion of the latch fittings. A cam on the assembly with the inside handle; both

5A-9 5A-10

5A-10 For SimuFlite training only Citation V

March 1999
 Aircraft Overview

handles are at the top of the door. A portable fire extinguisher mounts in
To open the emergency exit door, rotate a red quick-release bracket behind the
either handle to disengage the latch pin. aft left passenger seat.
Pull the door inward until it clears the Refreshment Centers
door frame, then raise the bottom of
the door to clear the retainers from the Three refreshments centers are avail-
stop blocks. Remove the door. After able for installation in the Citation V. A
several pressure cycles, the door standard 28-inch refreshment center
becomes firmly seated and is difficult, (Figure 5A-12) is forward of the entry
but not impossible, to remove. door, an optional 8-inch refreshment
center may be on the right forward of
Remove the locking pin inside the han- the emergency exit door, or an option-
dle prior to flight. This allows ground al midship refreshment center may be
personnel to enter the aircraft in res- installed.
cue operations. The emergency exit is
not connected to the DOOR NOT The standard 28-inch refreshment cen-
LOCKED warning circuit. ter contains three decanters, two heat-
ed liquid containers, a drip pan with an
Passenger Compartment heated overboard drain, and two cup
The passenger compartment (Figure dispensers enclosed by tambour doors.
5A-11) can be fitted to seat up to 11 Cold cathode tubes light the enclosed
passengers, although configurations of area automatically when the tambour
six to eight seats are typical. doors are raised. Above the lighted
work area an overhead compartment
The center and forward seats may face provides additional storage. A lighted
forward or aft, depending on configu- 8-inch clothes closet is in the forward
ration. Foldout tables are available for section of the unit. The lower section of
the middle seats. The seats are the unit includes an ice chest with heat-
adjustable forward-aft, left-right, and ed overboard drain, storage space and
in reclining positions. Adjustable arm- a trash container.
rests, seat belts, and shoulder harness-
es are standard. Each passenger posi- The optional 8-inch refreshment center
tion has individual seat lighting, a includes a heated liquid container, cold
ventilation outlet, and a drop-down water container, and drip pan with heat-
oxygen mask for emergency use (see ed overboard drain.
the Miscellaneous chapter for a detailed Another optional midship refreshment
description of the oxygen system). center contains an ice chest, storage

5A-11 5A-12

Citation V For SimuFlite training only 5A-11

March 1999
 drawers, drip pan, a one-gallon hot ity section consists of a sink with hot
tank, a 0.9-gallon cold water tank, and and cold water and a towel dispenser.
an overboard drain.
Cabin Windows
On the standard or optional models,
the REFRESH CTR CB provides 28V Fourteen cabin windows (Figure
DC for lighting and to heat the liquid 5A-13), seven on each side of the air-
containers and the overboard drip drain. craft, include the windows in the cabin
entrance and emergency exit doors. The
Aft Cabin Baggage windows are constructed identically.
Compartment Each window consists of three lami-
An aft cabin baggage compartment nated layers: an outer panel of stretched
stows up to 600 lbs of baggage. If acrylic 0.187 inch thick, a vinyl layer
installed, an optional aft divider and 0.05 inch thick, and an inner panel of
sliding doors separate the aft cabin bag- stretched acrylic 0.094 inch thick.
gage area from the passenger section
Tailcone
and reduce aircraft noise in the cabin.
The tailcone area houses the major
Waste Facilities components of the electrical, hydraulic,
An electrically operated, recirculating environmental, engine fire extinguish-
flush toilet is behind a privacy divider ing, and flight control systems, as well
in the right aft cabin baggage compart- as a baggage compartment.
ment. The self-contained unit operates
Forward Tailcone
on the aircraft’s 28V DC power. The Compartment
installation includes a relief tube, seat
cushion, lap belt, and shoulder harness, A small door on the forward bulkhead
as well as a life vest in the seat base of the tailcone baggage compartment
assembly. (Figure 5A-14) provides access to sys-
An optional vanity may be in the par- tems equipment. To reach the equip-
titioned toilet area. The upper vanity ment access door, first enter through
section includes a water tank storage the drop-down tailcone baggage door,
area, mirror, and light. The lower van- lift the carpet on the forward bulkhead

5A-13

5A-12 For SimuFlite training only Citation V

March 1999
 Aircraft Overview

of the tailcone baggage compartment, Tailcone Baggage

release the access door latches, and Compartment
raise the door.
The tailcone baggage compartment is
The aircraft battery and electrical junc- aft of the equipment compartment. The
tion boxes are in the forward tailcone tailcone baggage compartment has car-
area, along with major electrical buses, peted wall and floor panels and pro-
current limiters, circuit breakers, and vides storage space for baggage as well
anti-ice control boxes. as the engine covers. Nets secure the
The hydraulic reservoir, shutoff valves, baggage in place. The tailcone equip-
system valves, and plumbing are also ment access door is on the tailcone bag-
in the forward tailcone. Servicing is gage compartment forward bulkhead
normally accomplished with a (see the access door description above).
hydraulic “mule” via the hydraulic The drop-down tailcone baggage door
access panel on the lower aft fuselage. (Figure 5A-15) on the left fuselage has
In addition, the environmental system, a piano-type hinge at the bottom, with
air cycle machine (ACM), optional two latches (forward and aft) and a key-
flood cooling, and associated plumbing lock at the top. If the door is not prop-
are in the forward tailcone area. erly closed and locked, a microswitch in
the key-lock illuminates the DOOR
Two 125-cubic inch engine fire extin-
NOT LOCKED annunciator. When the
guishing bottles are below the
door is closed and locked, the annun-
hydraulic reservoir forward of the
ciator extinguishes.
ACM. The bottles contain CBrF3 nitro-
gen pressurized to about 600 PSI at A two-position ON/OFF light switch
70°F (21.1°C). Pressure gages are on mounted on the tailcone access for-
the bottles; a correction tempera- ward door frame controls the tailcone
ture/pressure table is between the bot- compartment and aft baggage lights.
tles for visual inspection during pre- The lights receive power from the Hot
flight (see Fire Protection chapter). Battery bus. A microswitch in the
doorframe removes power from the

5A-15

5A-14

Citation V For SimuFlite training only 5A-13

March 1999
 lights when the tailcone baggage door attached to the bleed air panels on the
is closed. inboard leading edge as well as the
deice boot.
Wing A cable-operated outboard aileron and
The fully-cantilevered wing (Figure two hydraulically operated inboard
5A-16) is a multiple spar, integral fuel flaps are on each wing trailing edge.
tank structure covered by aluminum Hydraulically operated speedbrakes
alloy skin panels that provide electri- are on the upper and lower surfaces
cal conductivity for lightning strike. of the wing, just forward of the out-
The stub wing assembly is the fuselage ermost flaps.
wing carry-through and extends beyond
the fuselage. Each wing attaches to the
Fuel Tanks
stub wing with upper and lower main An integral tank in each wing supplies
and rear spar fittings. fuel to its respective engine or cross-
The wing fixed leading edge supports feeds to the opposite engine, as select-
two wing ice protection systems: the ed. Each 431 gallon tank has an over-
outboard leading edge pneumatically wing fueling port (Figure 5A-17).
inflated deice boots and an inboard lead- Except for the area above the main
ing edge bleed air anti-ice system (see landing gear wheel well, the integral
Ice and Rain chapter). Stall strips are fuel tank includes all the wing area for-

5A-16

5A-17 5A-18

5A-14 For SimuFlite training only Citation V

March 1999
 Aircraft Overview

ward of the rear spar. Liquid-tight ribs porates the elevators; its leading edge
at the inboard and outboard ends of the has pneumatically inflated deice boots
wing complete the boundaries of the (see Ice and Rain chapter).
fuel tanks. The chemically-treated fuel A structural honeycomb panel at the
tank interior surface is coated with lower end of the vertical stabilizer pro-
epoxy primer for corrosion resistance. vides access to elevator and rudder
Plates in the lower wing skin panels control cables and components, avion-
provide access to the fuel cells. ics flux detector cable assembly con-
nectors, and deice boot line assemblies.
Empennage A saddle assembly at the lower leading
The conventional empennage (Figure edge of the vertical stabilizer provides
5A-18) consists of a fixed horizontal a smooth contour between the dorsal
stabilizer, elevators, elevator trim tabs, fin and vertical stabilizer. The locator
vertical stabilizer, rudder, and rudder beacon antenna mounts on the saddle
trim tab. assembly, while the optional emergency
The single-assembly horizontal locator transmitter (ELT) is in the
stabilizer is mounted on the vertical dorsal fin. The rudder is on the vertical
stabilizer above the fuselage tail sec- stabilizer trailing edge (see Flight
tion. The horizontal stabilizer incor- Controls chapter).

Citation V For SimuFlite training only 5A-15

March 1999
5A-16 For SimuFlite training only Citation V
March 1999
 Aircraft Overview

Radome to Rudder . . . . . . . . .
Nose Gear to Main Gear . . . . . .
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
.
. . 48.90 ft
. . 19.90 ft
Aircraft
Main Gear to Main Gear . . . . . . . . . . . . . . . . . 17.60 ft Dimensions
Wing Tip to Wing Tip . . . . . . . . . . . . . . . . . . 52.50 ft
Horizontal Stabilizer, Tip to Tip . . . . . . . . . . . . . 21.47 ft
Ground to Top of Vertical Stabilizer . . . . . . . . . . . 15.00 ft NOTE: See the Expanded Normal
Cabin Interior Diameter . . . . . . . . . . . . . . . . 59 inches Operations Chapter under Taxiing
Cabin Interior Height . . . . . . . . . . . . . . . . . . 57 inches and Towing for turning radii.

52.50 FT

21.47 FT

1.25 FT

17.60 FT

7.55 FT

15.00 FT

19.90 FT

48.90 FT

Citation V For SimuFlite training only 5A-17

March 1999
5A-18 For SimuFlite training only Citation V
March 1999
 Aircraft Overview

The primary dangers (Figure 5A-19)

around the aircraft are weather radar
Ingestion of small articles (e.g., keys
and pebbles) can cause considerable Danger Areas
emissions from the radome, the engine damage to the engine.
inlet air draw, and the exhaust plume.
Engine Exhaust
Radar Plume
If the weather radar is operating out of The engine exhaust danger area extends
the forced standby mode, emissions are 160 ft from the rear of the engine; as
hazardous up to a 9-ft radius of the the distance from the nacelle increases,
radome. The safest practice is not to the danger decreases. The engine
operate the radar on the ground except exhaust hazards lie in plume tempera-
for brief system tests or to scan the ture and velocity. Exhaust gases imme-
departure path immediately prior to diately exiting the nacelle average
takeoff. During ground tests, advise 1,000°F (537°C) at 600 kts. This
ground personnel to remain well clear decreases to 84°F (29°C) at 15 kts at a
of the radome. distance of 160 ft.
Engine Inlet Air Advise ground personnel of imminent
engine starts. Do not start an engine
Draw without verifying that the immediate
The engine inlet air draw is hazardous area behind and in front of the aircraft
25 ft in front of the engine nacelle. The is clear of ground personnel, small arti-
draw increases close to the nacelle. cles, and sensitive equipment.

25 FT RADIUS
84 °F
1,000 °F

9 FT RADIUS

160 FT

5A-19

Citation V For SimuFlite training only 5A-19

March 1999
5A-20 For SimuFlite training only Citation V
March 1999
This chapter provides a brief overview of the following:
Avionics
■ pitot/static system and instruments
■ communication equipment
■ navigation equipment Chapter 5B
■ flight control systems.
Besides these specific areas, this chapter includes instrumenta-
tion not addressed in other chapters.
Cockpit panel art at the front of the chapter provides a ready
reference to locate instruments and equipment addressed in this
discussion.
For a detailed description of and operating procedures for a
particular piece of equipment, refer to the applicable pilot’s guide.

Citation V For SimuFlite training only 5B-1

March 1999
5B-2 For SimuFlite training only Citation V
March 1999
 Avionics

Cockpit Instrument Panel Schematic . . . . . . . . 5B-8

Table of
Pedestal Schematic . . . . . . . . . . . . . . . . . 5B-9
Pitot/Static System Schematic . . . . . . . . . . . 5B-10
Contents
Flight Data Systems . . . . . . . . . . . . . . . . 5B-11
Pitot/Static System . . . . . . . . . . . . . . . . . . 5B-11
Air Data Computer . . . . . . . . . . . . . . . . . 5B-11
Pitot/Static Instruments . . . . . . . . . . . . . . . 5B-11
Altimeters . . . . . . . . . . . . . . . . . . . . . . 5B-12
Mach/Airspeed Indicator . . . . . . . . . . . . . . 5B-12
True Airspeed System . . . . . . . . . . . . . . . 5B-13
Instantaneous Vertical Speed Indicator . . . . . . 5B-13
Overspeed Warning . . . . . . . . . . . . . . . . 5B-13
Altitude Alerting and Reporting . . . . . . . . . . . 5B-14
AOA System . . . . . . . . . . . . . . . . . . . . . 5B-14

Avionics Equipment . . . . . . . . . . . . . . . . 5B-17

Communications . . . . . . . . . . . . . . . . . . . 5B-17
Audio Control Panels . . . . . . . . . . . . . . . . 5B-17
VHF Communications . . . . . . . . . . . . . . . 5B-18
Collins CTL-22 . . . . . . . . . . . . . . . . . . 5B-18
Honeywell RMU-850 . . . . . . . . . . . . . . . 5B-18
Clearance Deliver Unit . . . . . . . . . . . . . . . 5B-19
HF Communications . . . . . . . . . . . . . . . . 5B-19
Cockpit Voice Recorder . . . . . . . . . . . . . . 5B-19
Emergency Locator Transmitter . . . . . . . . . . 5B-20
Static Discharging . . . . . . . . . . . . . . . . . 5B-20
Navigation . . . . . . . . . . . . . . . . . . . . . . 5B-20
VHF Navigation . . . . . . . . . . . . . . . . . . 5B-21
Instrument Landing System . . . . . . . . . . . . 5B-21
Marker Beacon . . . . . . . . . . . . . . . . . . . 5B-21
Automatic Direction Finder . . . . . . . . . . . . . 5B-22

Citation V For SimuFlite training only 5B-3

March 1999
 Radio Magnetic Indicator . . . . . . . . . . . . . . 5B-22
Flight Management System . . . . . . . . . . . . 5B-22
Flight Data Recorder . . . . . . . . . . . . . . . . 5B-24
Pulse Equipment . . . . . . . . . . . . . . . . . . . 5B-24
Transponder . . . . . . . . . . . . . . . . . . . . 5B-24
Distance Measuring Equipment . . . . . . . . . . 5B-25
Radio Altimeter . . . . . . . . . . . . . . . . . . . 5B-25
Weather Radar . . . . . . . . . . . . . . . . . . . 5B-26

Other Instruments . . . . . . . . . . . . . . . . . 5B-27

Digital Clock . . . . . . . . . . . . . . . . . . . . . 5B-27
Outside Air Temperature . . . . . . . . . . . . . . . 5B-27
Standby Attitude Indicator . . . . . . . . . . . . . . 5B-27
Magnetic Compass . . . . . . . . . . . . . . . . . 5B-28

Flight Control Systems . . . . . . . . . . . . . . . 5B-29

Autopilot . . . . . . . . . . . . . . . . . . . . . . . 5B-29
Autopilot Controller . . . . . . . . . . . . . . . . . 5B-29
Autopilot Computer . . . . . . . . . . . . . . . . . 5B-31
Servos . . . . . . . . . . . . . . . . . . . . . . . 5B-31
Operation . . . . . . . . . . . . . . . . . . . . . . 5B-32
Flight Director . . . . . . . . . . . . . . . . . . . . 5B-32
Flight Director Computer . . . . . . . . . . . . . . 5B-32
Mode Selector . . . . . . . . . . . . . . . . . . . 5B-33
Accelerometers . . . . . . . . . . . . . . . . . . . 5B-33
Flight Instruments . . . . . . . . . . . . . . . . . . 5B-33
Copilot’s Attitude Indicator . . . . . . . . . . . . . 5B-34
Copilot’s Horizontal Situation Indicator . . . . . . . 5B-34

5B-4 For SimuFlite training only Citation V

March 1999
 Avionics

EFIS . . . . . . . . . . . . . . . . . . . . . . . . 5B-35
EADI . . . . . . . . . . . . . . . . . . . . . . . . 5B-35
EHSI . . . . . . . . . . . . . . . . . . . . . . . . 5B-36
Display Controller . . . . . . . . . . . . . . . . . 5B-37
Multifunction Display Unit . . . . . . . . . . . . . 5B-37
MFD Controller . . . . . . . . . . . . . . . . . . 5B-38
Symbol Generator . . . . . . . . . . . . . . . . 5B-38
Course and Heading Controller . . . . . . . . . . 5B-38
Vertical Navigation . . . . . . . . . . . . . . . . . 5B-38
Attitude and Heading Reference . . . . . . . . . . . 5B-39
Vertical Gyro . . . . . . . . . . . . . . . . . . . . 5B-39
Directional Gyros . . . . . . . . . . . . . . . . . . 5B-39
Rate Gyros . . . . . . . . . . . . . . . . . . . . . 5B-40

Preflight and Procedures . . . . . . . . . . . . . 5B-41

Preflight . . . . . . . . . . . . . . . . . . . . . . . 5B-41
Abnormal Procedures . . . . . . . . . . . . . . . . 5B-41
Autopilot Torque Failure . . . . . . . . . . . . . . 5B-41
Autopilot Roll Monitor Failure . . . . . . . . . . . 5B-41
Pitot/Static Heater Failure . . . . . . . . . . . . . 5B-41
Angle-of-Attack Probe Heater Failure . . . . . . . 5B-41
Erratic Instrument Indications . . . . . . . . . . . 5B-41
EFIS Display Failure . . . . . . . . . . . . . . . . 5B-42
EFIS Heading Failure . . . . . . . . . . . . . . . 5B-42
EFIS Attitude Failure . . . . . . . . . . . . . . . . 5B-42
EFIS Cooling Fan Failure . . . . . . . . . . . . . 5B-42
EADI/EHSI Overtemperature . . . . . . . . . . . . 5B-42
Symbol Generator Overheat . . . . . . . . . . . . 5B-43

Citation V For SimuFlite training only 5B-5

March 1999
 Symbol Generator Failure . . . . . . . . . . . . . 5B-43
Single Radio Management Unit Failure . . . . . .5B-43
Dual Radio Management Unit Failure . . . . . . . 5B-43
NOSE COMP O’TEMP Annunciator . . . . . . . . 5B-43
Angle-of-Attack Indicator Flagged . . . . . . . . . 5B-44
Emergency Procedures . . . . . . . . . . . . . . . 5B-44
EFIS Symbol Generator Failure . . . . . . . . . . 5B-44
EFIS Symbol Generator Overheat . . . . . . . . . 5B-44
Autopilot Hardover . . . . . . . . . . . . . . . . . 5B-44
Red Gun Display Failure . . . . . . . . . . . . . . 5B-44

5B-6 For SimuFlite training only Citation V

March 1999
 Avionics

Cockpit
Instrument
Panel

Citation V For SimuFlite training only 5B-7 5B-8

March 1999
 Cockpit Instrument Panel

S EMER EMER S
ARM ARM
T T
O BOTTLE 1 LH RH BOTTLE 2 O
UNLOCK W ARMED ENG ENG ARMED W UNLOCK
PUSH FIRE FIRE PUSH
DEPLOY S S DEPLOY
W NORMAL NORMAL W

9 7 8
EADI DISP MSG SX DR
HOT FAN OUTER MIDDLE INNER
AUTOPILOT MASTER
OFF AP EHSI SG
HOT HOT
WPT
WARNING MASTER
ROLL
MONITOR 2000 3000 100 WARNING
AP TORQUE RESET 110 O O O RESET
800 O 1800 F
O 120
F F F O F O
M877 F F F F O F O F
100 F U 1600 U I 100 I
I 80 I
CHRONOMETER
E E 2500 L L
100 FEET 1400 L L
L L 80
1200 P 60 P
ALT F 90 F F F FEET
300 320 0 I 700 I F F T T R R 100 1
0
1 A 80 A L
1000
L
U 2000 U
E
60 E
300 320 0 0 0 0 0 0 0 10
GMT LT FT ET
280 M 40 HDG GS 9 N N T T 800 E E E 40 E 280 M 40 9 1
A O O M M S S A C TOTAL HOURS
VG2 70 T T 40
0 560 L L
260 C W W O
LOC 20 20 600 P P 260 C

5
SELECT CONTROL H 60 S S D

5
H 60
4
8 2 60 1500 20 4 8
E
SEL CTL 240 AOA 600 L 400 R Q Q 20 30 30 2
80 50 L R L R 240 20 20
200 T T
0 5 60
80
DAVTRON KNOTS F 10 10 40 Y Y 0 0 KNOTS 10 10 ALT
220 MB ALT IN HG LBS/HR 1000 ¡C PSI 220
100 7 3 30 100 1000 FT
3
200 G 1013 29.92 L 20 R L
500
R L R 7 100 FT
29.92 M877
200 20 20
IN HG CHRONOMETER
6 4 % RPM 400 500 30 30
10 -0+ 10 TURBINE 4
120
1.0 180
5 180 120 6 90
20 20 160 140
S 10 10 300 160 140 5 60 ¡F 120
+ 30 60 ANGLE .8 SE BARO
_0 30 INDET 200 SE
INDET
30 OF X LL
30 90 ATTAC O 0 X PU TO CT 30 150
40 40 F .6 ER
E GMT LT FT ET
120 F CMD 20 20 ¡C % RPM LBS OFF 0 180 SELECT CONTROL
50
60
50 250 KTS DH CLI MB
140 B A T T SEL CTL
.0 .4
60 OAT
F
60 200 DH 140 RA T E M P
C .2 100 FT
DAVTRON
STAEL ALT
33 0 2 AC BATT CAB OIL PRESS FUEL LOW FUEL LOW HYD FLOW ENG
3 1
FAI O’TEMP ALT WARN LEVEL PRESS LOW ANTI-ICE TO DI VE 33 0 3 2
4 LH RH LH RH LH RH LH RH LH RH
NM 1 4
30

INST VERT FRM SET

30
5 INST VERT EMER COMM 1 COMM ADF NA
6

EMER COMM 1 5

6
COMM ADF NA VANG
SPEED GEN
OFF
INVERTER
FAIL
EMERG
PRESS ON POWER
FUEL FLTR FUEL
BY PASS BOOST ON
HYD LOW
LEVEL
WING
O- DEG SPEED
BRAKE PULL PASS
PASS DIM
6
27

TO
0 BLD AIR
6

27
SPKR
9

LOW PRESS HYD PRESS

LH RH 1 2 LH RH LH RH LH RH CAGE 0

9
SPKR GND ON
N 1 2 1 2 1 2
1 2 1 2 1 2
33 3 2
2 CRS MAG1 VOR1 FT MIN DOOR AIR DUCT ANTI SKID P/S HTR F/W NOSE COMP W/S AIR
FT MIN
12

COMM AUTO IDENT DME SPKR

24

12
SPEED SEAL O’HEAT INOP OFF SHUTOFF O’TEMP O-HEAT

24
COMM AUTO IDENT DME SPKR SURFACE 5 FMS
000 H 10.0 NM x 1000 4 BRAKE
DOOR NOT DE-ICE
x 1000 4 1

30
ACM
1 EXTEND GROUND AOA HTR

6
1 LH RH LH RH OFF
LOCKED O’PRESS 1
R

OFF 15 IDLE FAIL VOLUME

+I ILS 1

R
21 15
VO

VOLUME
¡ 21

VO
18
SEL VOICE 1 2 HDPH AD N
2 XFR XFR XFR
AD
18
2 H S
SEL VOICE 1 2 HDPH

W
H S
3 D P

E
F
D P
MKR
33 F
P K MKR
P K H R
H R FD OFF HDG NAV APR BC MEM MEM MEM MEM MEM MEM

30

12
24
SQ SQ PASS SPKR MASTER 1 2 MUTE
PASS SPKR MASTER 1 2 MUTE OFF ON ARM CAP ARM CAP ARM CAP ON OFF COM ON OFF COM ADF TONE ADF 15
WX OFF STO OFF STO ANT STO 21
W

E
G S
VOLTAGE SEL -2¡ DIM ALT ALT SEL VNAV VS IAS OFF
DH
BAT ON ARM CAP ARM CAP ON ON TEST TEST TEST
24

12
LH RH D.C. 200 200 VOR1 ACT ACT ACT DH 10
GEN VOLT D.C. AMP
GEN 20 30 D.C. AMP
300 300 ABS
100 100
21
15
ALT 0
VOR2 S CH SEL PWR 1 5 OXYGEN 15 PHONE COCKPIT VOICE RECORDER ERASE
NAV
TEST
10 40 0 400 0 400
CH SEL PWR HDG GSPD 25 2 HOLD
022 100 KTS AUTOTUNE XFR XFR XFR
AFT FWD
5 SEC
TEST FA FAN COMP 20
MODE 3
MFD SG A/C N HI ON 0 PSI 20
OFF
FIRE DC POWER AVIONIC POWER DTRK MAG1 360 HDG FMS1 HOT MEM MEM MEM MEM MEM MEM 15 RAD ALT x100
ANNU WAR 315 130 NM MFD MFD XFR F O
DIM

4
L GEN BAT R GEN INV1 AC ON FAN HOT F 10
ANTI +I ON HLD NA ON HLD NA ON AL AT A 5
LDG ¡ F

PR
N

T
ES S
S TO TE
SKID GEAR T N OFF STO OFF STO STBY IDENT
3 OFF MEM MEM
OVER BATT E 33 OFF LO LO
OFF OFF
SPEED TEMP S TEST
TEST TEST TEST ADF
T ADF TONE
W/S TEMP STICK ACT ACT PRE
SHAKER RESET EMER RESET INV2 OFF OFF HDG
V ANT STO 70 80
T/REV OFF CABIN
NORMAL VOR1
FUEL BOOST ENGINE START IGNITION
TEMP 90
LH RH
HDG
REV
50 TEST 60
VOR2 MAP ACT
LH ON RH LH ON RH WX NORM EMER
PLAN
O
HDG
GSPD
50 100
F STARTER 319 260 KTS
DISENGAGE INC MFD
¡F
F FULL GSPD VOR RCL SKP
MAP WX ET NA FMS R
ARC TIP
N SG HSI
NORM NORM G
NORM NORM
CROSSFEED ANTI ICE / DEICE APT DAT PAG ENT
DEC
NAV ADF 1 NAV ADF 2
PITOT & W/S BLEED W/S ALCOHOL ENGINE SURFACE MODE DIM
OFF FMS 1 OFF
TEST

STATIC HI ON LH RH AUTO
IN TRANSIT
OFF OFF LDG GEAR
O O O
ADI DH HSI WX
F F F UP DN
BRG DIM TST DIM DIM BRG CLOSED
F F F NOSE
LH OFF RH TEMPERATURE CONTROL
TAN TAN OFF LOW OFF XDF XDF MANUAL 15 MANUAL AUTOMATIC
ANTI- LH RH 2 WINDSHIELD BLEED AIR FA RH
LH GYRO SLAVE EXTERIOR LIGHTS 20 HOT
VERT GYRO LEFT CENTER RIGHT EL
HORN SKID 10 1 4
FLOOD LTS ON STDBY PASS LIGHTS 4 5 LH RH OVHD DEFOG VERT GYRO GYRO SLAVE
HI MAN LH RECOG ANTI NA WING SILENCE ON 3 6 25
.5 CABIN HI HI HI MAN LH
ON GYRO SAFETY BEACON B LANDING ON 6 OFF
GEAR 2 7 0
O O C UNLOCKED 6 O
5
1 8 30 FT MIN COLD F OPEN
F F N 9 .5 x1000
HOT
F
LH RH F F 4 MANUAL
/ PUSH DIFF 35 1 MANUAL
ENG ENG PRESS 2 NORMAL
OFF OFF OFF T 0 45
40 COLD OFF MAX OFF MAX LOW LOW NORM AUTO RH
TEST DOWN OFF CABIN ALT
NORM AUTO RH COLL INSP SEA OFF L LH RH OFF STATIC PITOT
PANEL LIGHT CONTROL x1000 FT
LH RH
T E TAS COMP
PARK BRAKE-PULL M
CONTROL LOCK PULL FLOOD COOLING E
EMER BRAKE-PULL 7 6 PULL RAIN
HIGH ON 8 R GND EMER
N

AL
0 35 X

BI
UNLOCK T4

CA

T-F
GND D OFF

10
AC

T
IDLE U PRESS SOURCE

00
AUX GEAR M
CONTROL P
NORM OFF

1.PULL & TURN INC CABIN

TEE HANDLE CO R CKPT CABIN
45¡ CW TO
CABIN RATE NTROLLE
UNLOCK
DEPRESSURIZE CABIN AIR
FLOW DISTR
2. PULL ROUND BEFORE LANDING PRESS-ENVIRON
KNOB TO BLOW
GEAR DOWN

5B-8 For SimuFlite training only Citation V

March 1999
 Avionics

Pedestal
LH FULL RH

FLAPS 0¡
UP

T.O 7¡

TRIM T
NOSE H
T.O &
DOWN R APPR
15¡
O
T
T
T L
O
E
NOSE

LAND 35¡

ENGINE SYNC
OFF
OFF
FA TURB MUST BE OFF
RETRACT FOR TAKEOFF
& LANDING

EXTEND

1 2 3

4 5 6
7 8 9 O
N

HOLD 0 BACK B
R
T
ENTER

M NA PLAN A B C D E F G
S
G
D FPL H I J K L M N #
APCH O P Q R S T U
TUNE V W X Y Z – SP

COURSE HEADING
PULL PULL
DIR SYNC

RCT STA TGT SECT

ON FP WX GMAP
STBY TEST
+
OFF 0 15
PRESET -
GAIN MODE TIL
BANK SOFT
LIMIT RIDE TURN
DESCEND
P LOW ON
I T
R DN TEST
T I EACH
C M
UP FLT
H
CLIMB ENGAGE ENGAGE

AP YD

L R

L R

Citation V For SimuFlite training only 5B-9

March 1999
Pitot/Static System

MACH/AIRSPEED ALTIMETER IVSI IVSI ALTIMETER MACH/AIRSPEED

100 FEET

300 320 0 2 2 10 0
FE E
T 300 320
0
1 0 0
280
40 9 1 4 1 4 C 9 MILLIBARS 1 280
40

0 5 00
O
260 5 INST VERT 5 INST VERT D 1018.5 260
5

5
60 E 60
4 8 2 SPEED SPEED 8 1018 2 4
240 240
6 6
220 KNOTS
80

MB ALT IN
3
0
5 FT MIN
0
5 FT MIN
0 5 00 ALT

3
220
KNOTS
80

100 7 HG 7 3007 100

200
1018 30.07 1 x 1000 4 1 x 1000 4 IN HG
200
6 4 2 2 6 4
180 120
5 5 180 120
SET SET
INDEX 160 140 INDEX 160 140

LEFT RIGHT
PITOT PITOT
TUBE TUBE
CABIN
PRESSURE
DIFFERENTIAL
GAGE

15
20
10
4 5
3 6 25

2 7
5 8 30
1
9

0
DIFF
PRESS 40
45
35
PITOT
CABIN AL
ALTT
x1000 FT
VALVE

GEAR
WARNING TAS
MACH AIRSPEED
ADC AIRSPEED SWITCH
SWITCH

STATIC
VALVE

TOP TOP

XPDR VNAV FMS FDR

PILOT'S PITOT

COPILOT'S PITOT
PILOT'S STATIC

COPILOT's STATIC

5B-10 For SimuFlite training only Citation V

March 1999
 Avionics

Flight data systems include the pitot/

static system, pitot/static instruments,
instruments. The pilot and copilot sta-
tic lines do not connect, but they do Flight Data
air data computer, and the angle-of-
attack system.
cross over within the aircraft so that
the pilot’s and copilot’s static sources Systems
are on both sides of the aircraft.
Pitot/Static System The pilot’s static sources supply the
Two electrically heated pitot tubes pilot’s altimeter, Mach/airspeed indi-
(Figure 5B-1) on either side of the air- cator, instantaneous vertical speed
craft nose supply ram air pressure. The indicator (IVSI), Mach/airspeed limit
Left Main bus Extension provides 28V warning switch, air data computer
DC for anti-icing of the pilot’s pitot/ (ADC) and FMS. The copilot’s static
static system and the Emergency bus sources supply the copilot’s altimeter,
provides 28V DC for the copilot’s pitot/ Mach/ airspeed indicator, IVSI, cabin
static anti-icing system. The PITOT & differential pressure gage, and land-
STATIC switch (Figure 5B-2) on the ing gear airspeed warning switch.
lower left instrument panel controls
power to the pitot tube, static port, and
Air Data Computer
angle-of-attack probe heating elements. An air data computer (ADC) receives
Two annunciators, LH P/S HEATER ram and static air pressure data from
OFF and RH P/S HEATER OFF, illu- the pilot’s pitot and static sources. The
minate with power off to the system or ADC, in turn, computes standard air
a heating element failure. data functions and provides electrical
The left pitot tube supplies the pilot’s driving signals to the pilot’s altime-
Mach/airspeed indicator, airspeed ter, transponder for altitude encoding,
Mach warning switch, the air data vertical navigation system (VNAV)
computer (ADC) and FMS. The right for altitude alerting, and the flight data
pitot tube supplies the copilot’s recorder and flight management sys-
Mach/airspeed indicator and the land- tem (FMS).
ing gear airspeed warning switch.
Sources of static air come from two
Pitot/Static
dual static ports (Figure 5B-3) on Instruments
either side of the aircraft nose. Each Pitot/static instruments are those
dual port contains a static air pressure instruments that use ram and/or static
source for the pilot’s and copilot’s pressure data from the pitot tubes and

5B-1 5B-2 5B-3

Citation V For SimuFlite training only 5B-11

March 1999
 static ports or information derived An amber warning light on the upper
from the ADC. These include: right edge of the pilot’s altimeter pro-
■ altimeters vides a visual altitude warning from the
vertical navigation system (VNAV).
■ Mach/airspeed indicators
The light illuminates once the aircraft
■ instantaneous vertical speed indi- reaches 1,000 ft of the preselected alti-
cators (IVSIs) tude and extinguishes when the aircraft
■ landing gear airspeed warning is within 250 ft of this altitude. The light
(refer to Landing Gear Chapter) illuminates again if the aircraft deviates
more than 250 ft from the preselected
■ overspeed warning
altitude.
■ altitude alerting and reporting.
An optional copilot barometric encod-
Altimeters ing altimeter provides a second source
of altitude encoding information. An
The pilot’s altimeter (Figure 5B-4) is OFF flag appears in the upper left cor-
an ADC instrument that provides ner of the indicator if the encoder loses
encoded altitude information to the power; the altimeter continues to func-
transponder. The copilot’s altimeter tion normally. A green switchlight,
(Figure 5B-5) is a barometric pres- labeled XPDR ENC ALT PRI/SEC
sure-driven instrument. Both the below the VNAV controller (Figure
pilot’s and copilot’s altimeters show 5B-6), selects the encoded altitude
aircraft altitude with a moving point- source. Normally, the transponder uses
er and drum display. information from the pilot’s altimeter
The altimeters provide an accurate (i.e., primary (PRI)). If the primary
indication of aircraft altitude from 0 source fails, pressing the switch selects
to 50,000 ft. A pointer, moving over a the copilot’s altimeter as the encoding
scale graduated at 20 ft intervals, source and illuminates SEC. Pressing
makes one revolution every 1,000 ft. the switch a second time returns to the
Digital drum displays indicate altitude primary (PRI) encoding source. An
Airspeed Limitations in ten-thousand, thousand, hundred, internal electric vibrator prevents stick-
MMO and 20 ft increments. ing of the altimeter needle. Power is
(above 28,907 ft) . . . 0.755 MI supplied to the vibrator through the 2
Both altimeters can be set to local
amp RH ALT CB on the left circuit
VMO barometric pressure or airport altitude.
breaker panel.
(8,000 to 28,907 ft) . 292 KIAS A BARO setting knob on the lower
left edge of each altimeter allows the Mach/Airspeed Indicator
VMO
setting of barometric pressure in either
(below 8,000 ft) . . . 261 KIAS The pilot’s and copilot’s Mach/air-
millibars or inches of mercury.
speed indicators (Figure 5B-7) are

5B-4 5B-5 5B-6

5B-12 For SimuFlite training only Citation V

March 1999
 Avionics

identical; both receive uncorrected (Figure 5B-8) on the lower right

information from their respective instrument panel. The transducer
pitot/static systems. Both display air- converts the pressure data into electri-
speed in knots with a pointer moving cal signals for the true airspeed com-
over a fixed scale and the Mach num- puter. After receiving pressure and
ber with a rotating scale. temperature data, the computer sup-
Two cutouts in the face of the indica- plies a corrected airspeed signal to the
tor denote VMO. The inner cutout at flight guidance system.
261 kts is red below 8,000 ft and black Instantaneous Vertical
above 9,600 ft. The outer cutout is
black below 8,000 ft and red at 9,600
Speed Indicator
ft and above. A single red radial line Identical instantaneous vertical speed
at Mach 0.755 denotes MMO above indicators (Figure 5B-9) for the pilot
28,907 ft. On aircraft with the option- and copilot use accelerometers to pro-
al 12,200 lbs zero fuel weight (ZFW), vide an instantaneous display of air-
the outer cutout is at 276 kts. craft vertical speed (up or down). Both
An adjustment knob on the lower left indicators have a pointer moving over
edge of the instrument allows the set- a fixed scale that indicates aircraft ver-
ting of a reference airspeed index (bug). tical velocity from 0 to 6,000 feet-per-
minute up or down. The instruments
True Airspeed System receive static air pressure information
from their respective static systems.
The digital true airspeed system uses
pitot, static, and ambient air tempera- Overspeed Warning
ture to compute true airspeed and
Mach number corrected for com- The overspeed warning system uses
pressibility and ram rise. The system ram and static air pressure to trigger an
supplies true airspeed data in an aural warning of an overspeed condi-
ARINC 429 format for the flight guid- tion. The system senses maximum air-
ance systems. The system consists of speed with ram and static pressure
a transducer behind the pilot’s instru- from the pilot’s pitot/static system to
ment panel, a temperature probe (ther- actuate a Mach/airspeed warning horn
mistor) on the forward fuselage, a true in the cockpit. Once the aircraft reach-
airspeed computer, and pitot/static es a set airspeed, contacts close the
source control switches. Mach/airspeed warning switch and the
horn sounds. The warning horn sounds
The transducer receives pitot/static data at approximately 262 kts below 8,000
pressure through the guarded PITOT ft, at approximately 292 kts between
and STATIC TAS COMP switches

5B-7 5B-8 5B-9

Citation V For SimuFlite training only 5B-13

March 1999
 8,000 and 28,907 ft, and at approxi- altitude requires placing the VNAV
AOA Limitations mately Mach 0.755 above 28,907 ft. computer controller selector switch in
The angle-of-attack system may On aircraft with the 12,200 lbs ZFW, ALT SEL and setting a reference alti-
be used as a reference system but the warning horn sounds at approxi- tude on the digital display with the
does not replace the airspeed mately 276 kts above 8,000 ft. SET knob. The system allows setting
indicator as a primary instrument. System testing is through the rotary an altitude within 10 ft from 0 to
TEST switch on the left instrument 10,000 ft altitude and within 100 ft
panel. Placing the switch in OVER- above 10,000 ft.
SPEED tests the system wiring and
components and sounds the warning AOA System
horn. The angle-of-attack system can be
used as a reference for approach speed
Altitude Alerting and (VS1) at all airplane weights and cen-
Reporting ter-of-gravity locations at zero, take-
The Honeywell altitude alerting and off, takeoff/approach, and landing flap
reporting system consists of an air data positions.
computer (ADC), vertical navigation Angle-of-attack (AOA) systems pro-
(VNAV) computer controller (Figure vide a visual indication of wing lift-
5B-10), altimeter, and instantaneous loading to allow the pilot to maintain
vertical speed indicator (IVSI). All a safe margin above stall speed. AOA
these components rely on the pitot/sta- systems on the Citation V include sys-
tic system for ram and static air pres- tems manufactured by Teledyne and
sure data. Safe Flight. The Teledyne system con-
The system provides both visual and sists of a computer, transmitter, probe,
aural indications when the aircraft flap position sensor, indicator, and an
reaches a preselected altitude by illu- indexer. The Safe Flight system con-
minating an amber light on the pilot’s sists of an angle-of-airflow sensor, sig-
altimeter and sounding an alerting nal summing unit, vane heater moni-
tone. The altitude alerting and report- tor, indicator, and an indexer.
ing system also provides accurate The vane or conical probe on the right
encoded altitude information for the forward fuselage (Figure 5B-11or 11A)
transponder. On some aircraft, the rotates (moves) to align itself with the
copilot’s altimeter provides a second airstream. A transducer attached to the
source for altitude encoding. vane/probe provides an electrical sig-
Altitude selection is through the nal to the summing unit/computer.
VNAV computer controller on the The summing/unit computer uses this
center instrument panel. Selecting an signal and one from the flap position

5B-10 5B-11 5B-11A

5B-14 For SimuFlite training only Citation V

March 1999
 Avionics

indicator to drive the AOA indicator The red arc from 0.85 to 1.0 is the
(Figure 5B-12) on the pilot’s instrument warning zone and continues to a full
panel, the indexer on the glareshield stall. The stick shaker activates at
(Figure 5B-13), and the fast/slow dis- approximately 0.79 to 0.88 indicated.
plays (Figure 5B-14) on the electronic A heads-up approach indexer on the
attitude director indicator(s). glareshield has a green circle, a red
The AOA indicator, marked from 0 chevron, and a yellow chevron. The
to 1.0, has colored arcs that indicate indexer provides a visual display of the
the approach and landing, warning, aircraft’s angle-of-attack. The vane-
and caution of angles-of-attack. actuated indexer activates once the nose
The Teledyne indicator has a green gear is down and locked and weight is
arc from 0 to 0.55 that indicates a nor- off the main gear squat switches; a
mal angle-of-attack. A white arc from timer provides a 20 second delay to
0.55 to 0.65 indicates the approach prevent illumination during takeoff.
and landing operating range; 0.6 indi- The red chevron points down to indi-
cates the optimum landing approach cate that pushing the nose down cor-
speed (VAPP/VREF). A yellow arc from rects a high AOA; the yellow chevron
0.65 to 0.75 is a caution zone where points up to indicate that pulling the
the aircraft is approaching a critical nose up corrects a low AOA. Five dif-
angle-of-attack. The red arc from 0.75 ferent combinations of the illuminat-
to 1.0 is the warning zone; it starts just ed circle and/or chevron indicate the
prior to stick shaker activation and various angle-of-attack conditions.
continues to a full stall. At approxi- The system provides:
mately 0.81 to 0.84 indicated, the stick ■ AOA high – red top chevron
shaker activates. illuminated
The Safe Flight indicator has white, ■ AOA slightly high – red top chevron
yellow, and red arcs that indicate the and green circle illuminated
approach and landing, warning, and
■ AOA correct – green circle
caution zones. The uncolored area
from 0.1 to 0.57 indicates the normal illuminated
operating range. A narrow white arc ■ AOA slightly low – green circle

from 0.57 to 0.63 indicates the and yellow bottom chevron illumi-
approach and landing range; 0.6 indi- nated
cates the optimum approach landing ■ AOA low – yellow bottom chevron
speed (VAPP/VREF). The yellow arc illuminated.
from 0.63 to 0.85 is a caution zone
indicating a critical angle-of-attack. At very high angles-of-attack, the red
top chevron flashes as an additional

5B-12 5B-13 5B-14

Citation V For SimuFlite training only 5B-15

March 1999
 warning to push the nose down and The AOA system uses 28V DC from
reduce the aircraft’s angle-of-attack. the Left Main bus Extension for oper-
Placing the rotary TEST switch in ation and anti-icing. The PITOT &
STICK SHAKER tests the angle-of- STATIC ANTI-ICE switch controls
attack system. Simultaneously the anti-icing for the AOA vane. If AOA
AOA indicator needle moves past 1.0 system loses power or fails, the indi-
(full stall), the red chevron illuminates cator pointer deflects and stows at 1.0
on the indexer, and the EADI needle(s) on the indicator; a red X appears on
move to SLOW; this indicates that the the EADI(s) fast/slow display.
AOA system is operating normally.
Without an operating inverter, the
AOA EADI test is not possible.

5B-16 For SimuFlite training only Citation V

March 1999
 Avionics

Avionics equipment on the Citation V

includes communication, navigation,
nications radios or the passenger
address system. A concentric volume Avionics
and pulse equipment.
Please refer to the Operating Manual,
knob controls the volume of outputs
to the cockpit speakers and the head- Equipment
AFM Supplements, and applicable phones. The microphone selector knob
Pilot’s Guides for specific operating has an emergency position that
procedures and limitations. bypasses the audio control panel
amplifiers and connects the micro-
Communications phone directly to the transmitter and
the output directly to the headset.
Communications equipment includes:
The Honeywell AV-850 audio control
■ audio panels panel (Figure 5B-16) is distinctive in
■ very high frequency (VHF) that it employs a digital audio trans-
communications mission technique. Instead of separate
■ high frequency (HF) communications
wires carrying audio signals from the
remote units, the system uses a twist-
■ radio-telephone ed pair of wires. The digital design of
■ passenger address and entertain- the unit reduces interference and
ment unnecessary noise.
■ cockpit voice recorder Each audio source selector on the AV-
■ emergency locator transmitter (ELT)
850 independently selects and con-
trols the volume of an audio source.
■ static discharging.
With the knob in, the audio output is
off. With it out, the control enables
Audio Control Panels receiver output and allows volume
The pilot’s and copilot’s audio con- control. A master volume control
trol panels (Figure 5B-15) control adjusts the volume of the combined
audio source selection and microphone outputs of the various audio sources
output selection; buttons or switches to the headphones or cabin speakers.
select the radio(s) connected to the The AV-850 audio panel has an emer-
headphones or cockpit speakers. A gency mode that automatically acti-
microphone selector knob connects vates during a power or audio system
the microphone in use to the commu-

5B-16

5B-15

Citation V For SimuFlite training only 5B-17

March 1999
 failure. The system bypasses all cir- quencies: one active, one preset, and
cuitry to connect the communication six stored in memory.
radio directly to microphone and head- Frequency tuning is through two con-
phones; pressing the EMER button centric knobs; the inner knob changes
performs the same function. the frequency in 50 kHz steps and the
VHF Communications outer in 1 MHz steps. Once tuned,
pressing the STO (store) button twice
The standard VHF communications stores the frequency in memory.
installation on the Citation V is a Collins Pressing the MEM button replaces the
VHF-22A transceiver controlled by a preset frequency with a stored fre-
Collins CTL-22 control head. The quency. Pressing the XFR button
Honeywell Primus II is an option; it swaps the active frequency with the
consists of a RCZ-850 transceiver con- preset frequency.
trolled by a radio management unit Additional controls on the control
(RMU). The Honeywell RM-850 RMU head include an ON/OFF switch, vol-
controls the VHF communications, ume control, and squelch control.
VHF navigation, automatic direction
finding, and transponder. Honeywell RMU-850
The optional clearance delivery unit The RM-850 RMU (Figure 5B-18)
(CDU) controls a single VHF com- provides frequency tuning, storage,
munication and navigation radio. and selection for VHF communica-
Typical VHF transceivers provide air- tion, VHF receiver, ATC, and ADF
to-air, air-to-ground, and ground-to- systems. The system uses a color cath-
ground communications in the 118 to ode ray tube (CRT) to display fre-
135.975 MHz frequency range. quencies for the various communica-
Frequency spacing of 25 kHz provides tion and navigation systems.
720 discrete channels. The Honeywell The RMU display has five windows
RCZ-850 can be bootstrapped to pro- that group data associated with a par-
vide coverage from 118 to 152 MHz; ticular system function. Each window
this provides an additional 720 dis- (COM, NAV, ATC, ADF, and MLS)
crete channels for a total of 1,440. provides control of both frequency
Collins CTL-22 and operating mode of the particular
function.
The Collins CTL-22 (Figure 5B-17) Each window has two lines: active fre-
uses two digital displays to show the quency and preset frequency, or fre-
active and preset frequency. The quency and operating mode.
system can use and store eight fre-

5B-17 5B-18 5B-19

5B-18 For SimuFlite training only Citation V

March 1999
 Avionics

Transfer and line select keys are on munications without turning on the
the left and right sides of the display. RMUs (and thereby draining battery
A transfer key alternates between the power) or as an emergency radio con-
active and preset frequency with each trol unit if both RMUs fail.
successive press. Pressing a line select The CDU has a liquid-crystal display
key highlights the selected item with (LCD) that displays the COM 1 and
a yellow box and connects the select- NAV 2 frequencies, the radio being
ed frequency with the TUNE knobs tuned, and annunciators for emergency
on the bottom of the RMU for fre- mode, squelch, transmit, and naviga-
quency tuning. Pressing the line select tion radio audio operation. Below the
key either recalls stored frequencies display are the transfer key and
or selects operating modes, depend- switches for squelch and NAV audio.
ing on the function selected. A MODES knob selects the unit oper-
Additional buttons on the RMU ating mode. Two concentric knobs on
include: the right of the unit tune the radios.
■ SQ – squelch
HF Communications
■ DIM – display brightness adjustment
High frequency radios allow very long
■ 1/2 – transfer control to opposite
range communication in the 2 to 29
RMU MHz frequency range. Spacing of 100
■ STO – store selected frequency Hz provides 280,000 discrete chan-
■ ID – transponder identification/ nels. The system operates in the lower
response mode sideband (LSB), upper sideband
(USB), and amplitude modulation
■ PGE – display page selection
(AM) modes. Through the control
■ TST – test selected (highlighted) head, the pilot can tune, store, and use
system up to 99 distinct frequencies.
■ DME – de-slave DME from active Typical systems consist of a trans-
VOR frequency. ceiver, control head (Figure 5B-20),
power amplifier/antenna coupler, and
Clearance Delivery Unit a long wire antenna.
The optional Honeywell CD-850
CDU (Figure 5B-19) on the center
Cockpit Voice Recorder
instrument panel is an auxiliary NAV The cockpit voice recorder (CVR)
1/ COM 2 radio control head for the records the last 30 minutes of con-
integrated radio system. The CDU can versation between the crew, aural
be used before engine start for com- warnings, and communications over

5B-20

Citation V For SimuFlite training only 5B-19

March 1999
 the public address system. The sys- the dorsal fin with a remote control on
tem also records audio signals from the right instrument panel. An impact
the pilot’s and copilot’s audio systems. switch activates the transmitter if the
The system consists of a cockpit- switch senses a force of approximate-
mounted control and monitoring panel ly five Gs along the longitudinal axis of
(Figure 5B-21), isolation amplifiers, the aircraft. The guarded switch on the
and a recorder unit in the tailcone instrument panel either arms, activates,
(Figure 5B-22). The recorder unit is in or turns off the ELT.
an international orange colored box
that protects it from fire and impact. Static Discharging
The CVR operates whenever the bat- Static discharger wicks (Figure
tery switch is in BATT. 5B-23) on the wing tips, horizontal
The CVR control panel on the right and vertical stabilizers, control sur-
instrument panel has a sensitive micro- faces, and tailcone minimize the
phone that filters out extraneous noise effects of lightning strikes on the air-
(i.e. engine noise). An indicating meter craft structure and static charges on
on the unit allows system monitoring avionics equipment. The dischargers
by observing needle movement while bleed off accumulated static charges
talking. A headphone jack allows the to the atmosphere and direct lightning
monitoring of CVR operation while strikes away from the aircraft.
the unit is operating. The ERASE but- During the preflight inspection, check
ton only works on the ground; press- the presence, security, and condition of
ing it for two seconds erases the entire each static wick.
tape. The TEST button initiates a test
of the control and recorder units. Navigation
The recorder unit contains an endless
Navigation equipment on the Citation
(loop) magnetic tape. Access to the
V includes:
recorder unit is through the tailcone
■ VHF navigation (VHF NAV)
access door.
■ instrument landing system (ILS)
Emergency Locator
■ marker beacon
Transmitter
■ automatic direction finder (ADF)
The optional emergency locator trans-
■ radio magnetic indicator (RMI)
mitter (ELT) transmits a sweeping tone
on 121.5 MHz and 243.0 MHz as an ■ flight management system (FMS)
aid in locating a downed aircraft. It con- ■ flight data recorder.
sists of a transmitter and antennas in

5B-21 5B-22 5B-23

5B-20 For SimuFlite training only Citation V

March 1999
 Avionics

VHF Navigation receive and display horizontal (later-

al) and vertical (glideslope) guidance
The VHF navigation receivers oper- information on the attitude director
ate in the 108.00 to 117.95 MHz fre- and horizontal situation indicators.
quency range; 50 MHz channel spac- Together, this allows the flight crew
ing provides 200 distinct frequencies. to follow an established instrument
As part of the instrument landing sys- approach profile.
tem (ILS), the system also provides
40 localizer (LOC) channels in the Tuning a localizer frequency auto-
108.10 to 111.95 MHz range and 40 matically selects the paired glideslope
glideslope channels in the ultra-high channel. As an aid in identifying the
frequency (UHF) band from 329.15 selected localizer, each localizer trans-
to 335.00 MHz frequency range. This mits a distinct four-letter Morse code
provides 40 paired channels for the identifier.
instrument landing system (ILS).
Marker Beacon
VHF navigation receiver tuning and
control is through a separate control The marker beacon receiver operates
head or through the radio management on 75 MHz. The system provides visu-
unit (Honeywell Primus II). The Collins al and aural indications of outer, mid-
CTL-32 has two digital readouts dle, and inner beacon passage.
(Figure 5B-24) that display the active Depending on the flight director sys-
and preset frequency. Aircraft with the tem installed, the electronic attitude
Honeywell Primus II system use radio director indicator(s) has colored sym-
management units (RMUs) to control bols that illuminate during station pas-
the VHF receivers (Figure 5B-25). sage. On aircraft equipped with only
the pilot’s EFIS (Electronic Flight
VHF receivers provide information to Information System), the copilot’s visu-
the attitude director indicators (ADIs), al indication is on three colored annun-
horizontal situation indicators (HSIs), ciators above the copilot’s attitude indi-
radio magnetic indicators (RMIs), cator. On both displays, blue denotes
flight management system (FMS), and the outer marker, amber the middle
other navigation equipment. marker, and white the inner marker.
Instrument Landing Each marker beacon station uses a dif-
System ferent dot-dash sequence and tone to
identify itself. The outer marker, locat-
The instrument landing system (ILS) ed from four to seven miles from the
combines outputs from the VHF NAV runway threshold, has a 400 Hz two
(localizer) and glideslope receivers to dash per second tone. The middle

5B-24 5B-25

Citation V For SimuFlite training only 5B-21

March 1999
 marker, located 3,000 to 6,000 ft from tion for the radio magnetic indicator
the runway threshold, has a 1300 Hz and the horizontal situation indicator;
tone with 95 alternating dots and dash- for easier identification of Morse code
es per minute. The inner marker, has station identifiers, BFO adds a 1000
a 3000 Hz tone with six dots per sec- Hz tone. VOICE mode increases the
ond. As the aircraft passes each receiver bandwidth to improve audio
beacon during an approach, the respec- quality.
tive annunciators flash and the marker
identifying tone sounds through the Radio Magnetic Indicator
speakers or headphones. Two radio magnetic indicators (RMIs)
display magnetic ADF and VOR head-
Automatic Direction Finder ing information with two pointers on a
Automatic direction finder (ADF) sys- calibrated servo-driven compass card
tems (Figure 5B-26) operate in the (Figure 5B-27). A single bar pointer
100.0 to 1799.5 kHz frequency range displays either No. 1 VOR or No. 1
with 0.5 kHz frequency spacing. ADF information and a double bar
These systems provide relative bear- pointer displays either No. 2 VOR or
ing readings to low frequency hom- No. 2 ADF information. Pushbuttons
ing stations, radio beacons, and AM on the lower corners of the instrument
broadcast stations. Typical systems select either VOR or ADF for each
consist of a receiver, control head, and pointer.
a loop and sense antenna. Some instal- The copilot’s RMI, driven by the pilot’s
lations employ dual ADF systems. C-14D directional gyro, slaves to the
The system provides four selectable copilot’s compass system. The pilot’s
operating modes: RMI, driven by the copilot’s C-14D
■ ANT – antenna directional gyro, slaves to the pilot’s
compass system.
■ ADF – automatic direction finder
Please refer to the Flight Control
■ BFO (TONE) – beat frequency
System section, this chapter, for a dis-
oscillation cussion of the directional gyros (com-
■ VOICE – voice audio. pass systems).
In ANT, the system only receives and
Flight Management
does not provide bearing information;
this mode is usable for frequency tun- System
ing in that it prevents random move- The standard Global GNS-X and
ment of the ADF needle. Both ADF optional Honeywell FMZ-600 flight
and BFO compute bearing informa- management systems (FMS) use nav-

5B-26 5B-27

5B-22 For SimuFlite training only Citation V

March 1999
 Avionics

igation inputs from the navigation (dual EFIS) radios as it needs to for
radios and internal sensors to calculate navigation purposes. A NAV /AUTO-
and update aircraft position. These sys- TUNE (GNS-X) or NAV MANUAL/
tems provide a best computed position NAV AUTO/ARM (FMZ-600) switch
based on the accuracy and reliability (Figure 5B-29) on the center instru-
of the input sensors (i.e., signal quali- ment controls this function. Use of
ty, distance to navigation aid, etc). this switch(es) prevents unintentional
Typical systems consist of a naviga- VHF NAV tuning by the FMS. The
tion management unit or computer, FMS also does not autotune the VHF
control display unit (Figure 5B-28), NAV radio(s) when the EFIS system
LORAN-C receiver, VLF/Omega sen- controls the radio.
sor, and antennas. The GNS-X also has the capability to
The system receives bearing and dis- tune the Primus II VHF COM, VHF
tance information from the VHF NAV NAV, ADF, and transponders.
receivers, air data information from the Frequency selection is through the
true airspeed computer, and long range CDU. After tuning the radio, the CDU
navigation information from a frequency blanks as there is not a data-
LORAN-C receiver and its own VLF/ bus between the Primus II system and
Omega sensor. the FMS. The FMZ-600 cannot auto-
The FMS provides position information tune the Honeywell Primus II radio sys-
and navigation commands through the tem; it can only use VHF NAV and
control display unit (CDU), horizontal DME information manually tuned by
situation indicator(s), and/or multifunc- the flight crew.
tion display (if installed). It will provide Amber annunciators on the pilot’s
roll commands to the autopilot. instrument panel repeat information
Through an internal database, the sys- presented on the CDU by the GNS-X
tem can automatically tune navigation and FMZ-600 flight management sys-
aids along the flight path. The data- tems. This includes:
base also contains navigation aid, air- ■ BATT – illuminates when the
port information, and waypoints; it can GNS-X internal battery is power-
also store user entered waypoints. A ing the system
connector on the lower right instru- ■ SX – illuminates after program-
ment panel allows the connection of a ming in a parallel track to the leg
data loader or data transfer unit for shown on the CDU
database updates.
■ WPT – illuminates when the air-
The GNS-X and FMZ-600 can auto- craft is within 30 seconds of the
tune the VHF NAV 1 and VHF NAV 2 next leg change

5B-28 5B-29

Citation V For SimuFlite training only 5B-23

March 1999
 ■ MSG – illuminates when the VOR- The UFDR has an inverter that con-
TAC positioning unit (VPU) is no verts 28V DC from the aircraft elec-
longer supplying navigation or trical system into 115V, 400 Hz AC.
position update data; the CDU An inertia switch removes 28V DC to
explains the message the inverter and turns the UFDR off if
■ DR – illuminates when the system the switch experiences a force of five
is in dead reckoning mode. gravities (5 Gs). Pressing the reset
switch on the UFDR restores power to
The GNS-X also provides fuel man- the inverter. The amber RECORDER
agement functions. The system has PWR FAIL annunciator illuminates to
the ability, after entering aircraft indicate power loss to the recorder or
weight and fuel load before takeoff, that the recorder is not running.
to monitor fuel use and provide an
updated aircraft gross weight. An underwater locating beacon on the
recorder assists in locating the recorder
Flight Data Recorder if the aircraft ditches. The battery pow-
ered beacon begins transmitting on
The optional universal flight data
37.5 kHz once the beacon submerses
recorder (UFDR) records the last 25
in either salt or fresh water. The design
hours of aircraft flight data on mag-
of the beacon allows it to continue
netic tape in a digital format. The
transmitting for 30 days at depths up
UFDR consists of a remotely mount-
to 20,000 ft.
ed recorder unit (Figure 5B-30) and
an accelerometer. The UFDR in the
tailcone is in an international orange
Pulse Equipment
painted steel container built to with- Pulse equipment includes:
stand impact and fire. ■ transponder
The system records aircraft altitude, ■ distance measuring equipment
airspeed, vertical acceleration, and
■ radio altimeter
heading with a time reference. Aircraft
altitude and airspeed come from the ■ weather radar.

air data computer (ADC), heading

information is from the pilot’s RMI, Transponder
and vertical acceleration is from an Standard installation includes two
accelerometer. Converters for air- 4096 Mode A transponders in the nose
speed, course altitude, and fine alti- avionics compartment. Each transpon-
tude convert DC signals from the der functions in Mode C (automatic
ADC into synchro signals that the altitude reporting). The optional
UFDR can record. Honeywell Primus II Remote Radio

5B-30 5B-31 5B-32

5B-24 For SimuFlite training only Citation V

March 1999
 Avionics

System also has Mode S (collision that provides information. Either NAV
avoidance). Each system consists of a 1 or NAV 2 illuminates to indicate the
transceiver, control head (Figure active radio. On the EHSI(s), press-
5B-31), and a transmit/receive anten- ing the NAV button on the respective
na. The system transmits on 1090 MHz EFIS display controller selects the
and receives on 1030 MHz. The pilot’s NAV radio information displayed.
or copilot’s (if installed) encoding
altimeter provides aircraft altitude infor- Radio Altimeter
mation to the transponder system for The Collins ALT-55B radio altimeter
transmission to ATC radar facilities. system provides accurate altitude from
20 to 2,500 ft during approach and land-
Distance Measuring ing. The system consists of a trans-
Equipment ceiver, transmit and receive antenna,
Distance measuring equipment (DME) and indicators (Figure 5B-33). The
computes and provides slant range dis- transceiver transmits a varying 4250 to
tance between the aircraft and a VOR- 4350 MHz (4.3 GHz ±50 MHz) signal
TAC navigation facility. The system toward the ground, receives the bounced
transmits in the 1025 to 1150 MHz signal, and computes altitude by com-
frequency range and receives in the puting the time between transmission
962 to 1213 MHz range. Most sys- and reception of the radio signal.
tems also can tune military TACAN The radio altimeter operates continu-
stations for distance measuring infor- ously when there is power on the air-
mation. craft. As the aircraft descends to 2,500
Pairing of DME channels with VOR ft, the system begins providing alti-
frequencies provides automatic selec- tude information to the pilot’s EADI
tion of DME channels by the VHF and the copilot’s ADI (dual flight
navigation receiver. directors). Between 200 and 2,500 ft,
DME information can be displayed the indicator provides altitude in 10-ft
either on indicators (Figure 5B-32) or increments; below 200 ft it provides
on the electronic horizontal situation altitude in 5-ft increments. The radio
indicator(s). The indicator has two dig- altimeter may fluctuate as much as 50
ital displays that present distance to ft when taxiing over ice or snow due
station in nautical miles (nm) and, by to the radio signal reflective proper-
pressing the SEL button: ground ties of these surfaces.
speed, time-to-station, or station iden- Additional aural and visual indications
tification. Pressing the CH button on provided by the radio altimeter include
the indicator selects the NAV radio a selectable decision height (DH) light

5B-33 5B-34

Citation V For SimuFlite training only 5B-25

March 1999
 and warning horn, and a rising run- ■ RCT – selects rain echo attenua-
way symbol. The EFIS display con- tion compensation technique to
troller sets the altitude when the amber attempt to display weather masked
DH light illuminates in the EADI. by heavy precipitation
Once the aircraft reaches 100 ft above ■ TGT – selects target alerting; alerts
the preselected decision height, a crew to high intensity weather by
white box appears near the radio alti- displaying a yellow “T” to indicate
tude indication on the EADI. Once the red level or worse targets within
aircraft reaches the decision height, 7-1/2 degrees either side of the air-
the amber DH symbol appears. craft nose
Weather Radar ■ MODE SWITCH – selects either
ON, OFF or STANDBY
The Primus 650 WX radar system con-
■ GAIN – adjusts the sensitivity of
sists of an antenna, receiver-transmitter,
display, and system controls (Figure the radar receiver
5B-34, previous page). The system ■ TILT – adjusts antenna tilt up or
operates in the X band to provide color- down
coded weather location and ground ■ BRT – adjusts display brightness
mapping capabilities. In the weather or
■ SCT – selects either 120° or 60°
ground mapping mode, the system has
a maximum range of 300 nm. scanning area
■ AZ – enables or disables azimuth
The system displays storm intensity
in five levels with color. With weak line display
or no returns, the display is dark. As ■ RANGE – adjusts through two
storm intensity increases, color pro- buttons range either up or down
gresses from green to yellow to red, between 5 and 300 nm.
and then to magenta; green represents The radar system uses a landing gear
light rainfall and magenta represents squat switch to prevent unintentional
the heaviest rainfall. The system sens- operation on the ground. The radar
es precipitation, not clouds. In ground remains in forced standby mode until
mapping mode, the system uses black, weight is off the wheels. Simultaneously
cyan, yellow, and magenta to repre- pressing both RANGE buttons takes
sent reflected signal intensity. Cyan the radar out of forced standby mode;
denotes the least reflective surface and the system operates normally.
magenta represents the most reflec-
tive surface. When operating the radar on the
ground, exercise extreme caution to
Controls arranged on the left, right, prevent injury to ground personnel,
and bottom of the display (Figure fuel ignition, or damage to the radar
5B-34, previous page) select system system. Keep personnel from entering
mode, scanning range, antenna tilt, the radar scanning area and within 15
receiver gain (sensitivity), scanning ft of the radome. Do not operate the
angle, display brightness, and target radar during refueling or in the prox-
alerting. Counterclockwise around the imity of explosive materials. Point the
display this includes: aircraft nose away from obstructions
■ WX – selects weather mapping and large metal buildings to prevent
mode damage to the system by reflected
■ MAP – selects mapping mode radar energy.

5B-26 For SimuFlite training only Citation V

March 1999
 Avionics

This section describes flight instru-

ments unrelated to the pitot/static,
tons exits the setting mode and starts
the clock. To select the other clock Other
flight control, autopilot, navigation,
and communications systems. These
modes, press the SEL button; the
UTC, LT, FT, or ET light illuminates Instruments
include: to indicate the selected mode.
■ Davtron digital clock

■ outside air temperature gage

Outside Air
■ standby attitude indicator
Temperature
■ magnetic compass. An outside air temperature (OAT) gage
(Figure 5B-36) is on the left instru-
Digital Clock ment panel. The gage displays tem-
perature uncorrected for ram rise in
Two Davtron M877 digital clocks Celsius and Fahrenheit. The gage uses
(Figure 5B-35) for the pilot and copi- 28V DC for operation and 5V AC for
lot can display local time (LT), lighting. A temperature sensor in the
Coordinated Universal time (UTC), air inlet duct under the dorsal fin scoop
flight time (FT), and elapsed time senses outside temperature.
(ET). Each clock has two function but-
tons, SEL (select) and CTL (control), Standby Attitude
that select the time mode displayed,
start the mode, and reset the selected Indicator
mode. The normal operating mode is A J.E.T. two-inch standby attitude
the display of either local time or indicator (Figure 5B-37) is on the
Coordinated Universal Time (UTC). copilot’s instrument panel. The indi-
Local time is in a 12-hour format; cator provides a standby source of air-
UTC is in a 24-hour format. A landing craft roll and pitch information. A
gear squat switch enables the flight PULL TO CAGE knob is on the right
time mode once weight is off the land- corner of the indicator. Pulling the
ing gear. knob out and rotating it clockwise
Pressing both buttons puts the clock cages the gyro. An OFF flag appears
into the time-setting mode. The tens during a power loss, if there is an open
of hours digit flashes and can be set winding in the motor, or if the the unit Standby Gyro Limitation
by pressing the CTL button to change is caged. Normally, the unit uses 28V
DC from the LH Main bus Extension. A satisfactory preflight test must
the time. Pressing the SEL button
If bus power fails or drops below a set be accomplished on the standby
steps the clock through the individual
digits. Once set, pressing the SEL but- voltage, an emergency battery pack in gyro system.

5B-35 5B-36 5B-37

Citation V For SimuFlite training only 5B-27

March 1999
 the nose avionics compartment pro- TEST performs a battery capacity and
vides power for approximately 30 min- circuit test; illumination of the green
utes of gyro operation. The battery GYRO TEST light indicates an ade-
pack, consisting of lead-acid batteries, quate charge and proper operation of
also provides power for lighting of the the circuits.
copilot’s airspeed indicator, altimeter,
VSI, and attitude indicator. The air- Magnetic Compass
craft’s electrical system provides charg-
A standard, liquid-filled magnetic
ing current for the battery pack.
compass (Figure 5B-38) on the wind-
The STDBY GYRO switch on the shield center post provides magnetic
lower edge of the pilot’s instrument heading information. The compass
panel controls and tests the standby contains provisions for maintenance
gyro horizon. To the left of the switch personnel to adjust the unit to com-
is an amber POWER ON and a green pensate for aircraft generated mag-
GYRO TEST light. With the switch netic fields. A correction card near the
in STDBY GYRO, the attitude indi- compass provides a record of recent
cator operates and the aircraft electri- adjustments to the compass and com-
cal system provides power for opera- pass deviation errors.
tion. If the electrical system fails, the
Heading information from the com-
batteries automatically power the indi-
pass is only accurate in straight and
cator and the amber POWER ON light
level, unaccelerated flight.
illuminates. Placing the switch in

5B-38

5B-28 For SimuFlite training only Citation V

March 1999
 Avionics

The Citation V has a Honeywell SPZ-

500C automatic flight control system
to fly the aircraft through the autopi-
lot controller pitch wheel, turn knob, Flight Control
(AFCS). The system provides three
axes (roll, pitch, and yaw) automatic
or touch control steering. Use of the
pitch wheel or turn knob uncouples Systems
flight control with elevator trim, yaw the flight director, canceling the ver-
damping, pitch-lift compensation, and tical or lateral flight director modes.
Q-switching.
The system combines the following Autopilot
sub-systems into a completely inte- The autopilot system provides aircraft
grated automatic flight control system: attitude stabilization across the pitch,
■ autopilot system roll, and yaw axes. The primary func-
■ flight director system
tion of the autopilot system is to
reduce pilot workload and fatigue dur-
■ flight instrumentation systems
ing long flights. The system allows
■ attitude and heading reference complete control of the aircraft with
systems minimal control inputs provided
■ air data system. through the autopilot controller.
The system provides three options to Components of the autopilot include:
the flight crew: manual operation, ■ SP-200 autopilot computer

automatic operation, or semi-auto- ■ PC-500 autopilot controller

matic operation. Manual operation
■ SM-200 servo motors
allows the pilot to fly the aircraft guid- NOTE: Please refer to the applic-
ed by cues from the flight director ■ controls and indicators. able pilot’s manuals and the
instrumentation. Automatic operation Aircraft Flight Manual for a thor-
flies the aircraft through the autopilot Autopilot Controller ough discussion and operating
coupled to the flight director; the pilot The PC-500 autopilot controller procedures for the particular
only monitors system operation. Semi- (Figure 5B-39) on the pedestal has pieces of the flight control system.
automatic operation allows the pilot controls for the yaw damper and

Autopilot Limitations
One pilot must remain in his seat
with the seat belt fastened during
all autopilot operations.
The autopilot torque monitor and
roll monitor should be functionally
tested per the Normal Procedures
in the AFM prior to inflight use.
Autopilot operation is prohibited
above 14,500 ft if the torque moni-
tor (AP TORQUE annunciator)
does not test per the Normal
Procedures in the AFM.
Autopilot operation is prohibited if
the roll monitor annunciator
illuminates inflight.

5B-39

Citation V For SimuFlite training only 5B-29

March 1999
 autopilot. A PITCH wheel and TURN Disengaging the autopilot with the
knob allows pilot control of the air- AP/TRIM DISC button, electric trim-
craft through the autopilot in the pitch ming (Figure 5B-40), or pressing a go-
and roll axes. Additional controls around button (Figure 5B-41) illumi-
include bank limit and soft ride switch- nates the amber AUTOPILOT OFF
es. The soft ride, bank limit, yaw lights and sounds the autopilot warn-
damper engage, and autopilot engage ing horn for one second. If the autopi-
buttons illuminate when selected. lot disengages for any other reason, the
Pressing the YD button engages the AUTOPILOT OFF light illuminates
yaw damper; the button illuminates continuously, and the warning horn
with ENGAGE to indicate engage- sounds for one second. Pressing the
ment. Use of the yaw damper without AP/TRIM DISC, elevator trim, or go-
the autopilot increases aircraft stabil- around button extinguishes the light.
ity in the yaw axis and results in Illumination of the UP or DN light on
increased passenger comfort. The yaw the controller indicates the application
damper must be disengaged with the of a sustained force by the elevator
red AP TRIM DISC button on the servo to maintain aircraft attitude; both
pilot’s or copilot’s control wheels. the indicators should be off before
Pressing the AP button engages the engaging the autopilot. The TURN
autopilot and yaw damper simultane- knob also must be centered in the
ously. With no flight director modes detent before autopilot engagement.
selected, engaging the autopilot brings Rotating the TURN knob either left
the aircraft to a wings level attitude. or right out of the detent results in a
With a flight director mode selected roll command. The roll angle is pro-
before autopilot engagement, engag- portional to and in the direction of
ing the autopilot automatically cou- knob rotation. The farther the knob is
ples the system to the flight director. turned, the greater the roll angle.
The AP button illuminates with Rotation of the knob automatically
ENGAGE to indicate engagement. cancels any selected flight director lat-
The autopilot normally disengages eral mode.
with: The PITCH wheel provides an aircraft
■ pressing the AP/TRIM DISC pitch movement in the direction and
button on either control wheel in proportion to wheel movement.
Movement of the wheel cancels only
■ using the elevator trim switches
the altitude hold and altitude capture
■ depressing a go-around button on flight director modes. During an
the throttles. autopilot glideslope capture, move-

5B-40 5B-41 5B-42

5B-30 For SimuFlite training only Citation V

March 1999
 Avionics

ment of the wheel has no effect on air- flap position control switch. Roll axis
craft pitch angle. Movement of the pitch signals come from the vertical gyro,
wheel with a flight director vertical directional gyro, flight director com-
mode (VS or IAS) selected changes the puter, and autopilot controller. Yaw axis
EFIS displayed vertical mode reference; signals come from the directional gyro
the autopilot responds to these changes. that provides heading information.
As a safety feature, perform an autopi- Altitude information comes from the
lot system test before autopilot air data computer (ADC) or a baro-
engagement; the autopilot will engage metric switch in the copilot’s static
without this test. system. The autopilot uses the altitude
The BANK LIMIT and SOFT RIDE information to enable or disable torque
buttons limit the operation of the switching. Torque switching provides
autopilot. Pressing BANK LIMIT two distinct rates of servo torque,
reduces the maximum autopilot com- depending on aircraft altitude. With
manded bank angle to 12° in the flight the SPZ-500C autopilot, torque
director heading select mode. The switching occurs at 14,500 ft.
LOW light illuminates to indicate acti- The computer takes these signals and
vation of bank limiting. Pressing the compares the aircraft’s actual attitude to
SOFT RIDE button provides decreased the desired attitude. The computer then
pitch and roll commands from the commands the servos that move the
autopilot in response to sudden atti- flight controls to reposition the aircraft
tude changes caused by flight through to match the desired aircraft attitude.
turbulent air; the button illuminates
with ON to indicate activation. Use Servos
SOFT RIDE only for moderate to Servos for the ailerons, elevator, and
heavy turbulence. rudder consist of a DC motor-
On aircraft with dual flight direc- tachometer, clutch assembly, synchro,
tors, a reversionary switch on the cen- and power gear train. Signals from the
ter instrument panel selects the flight autopilot computer drive the servo
director system providing altitude, motor through cables to position the
heading, and flight director command control surface. A feedback signal pro-
inputs to the autopilot. The split AP duced by the motor-tachometer relays
XFR FD-1/AP XFR FD-2 switchlight control surface position information
(Figure 5B-42) selects between the to the computer. Once the control sur-
pilot’s (No. 1) and copilot’s (No. 2) face reaches the commanded position,
flight directors. The switch illuminates the computer signals the motor to stop.
green to indicate the system in use. During autopilot engagement, the elec-
tromagnetic clutch assembly connects
Autopilot Computer the servo motor to its output shaft.
The SP-200 autopilot computer in the With the autopilot disengaged, the out-
nose avionics compartment receives put shaft moves freely.
pitch, roll, and yaw signals from the A servo amplifier within the comput-
navigation sensors, gyros, autopilot er provides torque limiting for the ser-
controller, and the flight director com- vos. The limiter regulates the power
puter. The computer takes these sig- supplied by the autopilot computer to
nals, performs computations, and the servo drive motor.
sends driving signals to the aileron, The crew can overpower the autopilot
elevator, and rudder servos. through control wheel movement. If a
Pitch axis signals come from flight mechanical failure occurs, the servo’s
director computer, airspeed sensor, clutch slips to allow normal control
autopilot controller, vertical gyro, and movement.

Citation V For SimuFlite training only 5B-31

March 1999
 Operation light do not illuminate and the autopi-
lot remains engaged, the autopilot
The autopilot can provide pitch, roll, requires maintenance.
and heading hold without inputs from
the flight director. The autopilot main-
tains an established pitch angle once it
Flight Director
engages. Normally, without touch con- The flight director system processes
trol steering (TCS) inputs, the autopi- signals from various sensors and equip-
lot rolls the aircraft to a wings level ment to provide horizontal and verti-
attitude while the aircraft is in a bank. cal guidance signals for display on the
If the aircraft is in a bank greater than electronic attitude director indicator(s).
6°, the autopilot holds the heading indi- This includes the radio guidance, iner-
cated as it rolls the aircraft to a wings tial navigation, area navigation, and
level attitude. The autopilot also holds vertical navigation modes of operation.
the heading existing at lateral mode The standard installation is a single
disengagement. flight director system with a two-tube
With touch control steering, the pilot electronic flight instrument system
can maintain a bank greater than 6° (EFIS). EFIS consists of an electronic
during autopilot engagement by press- attitude director indicator (EADI) and
ing the TCS button, engaging the electronic horizontal situation indica-
autopilot, and then releasing the TCS tor (EHSI). The copilot has an attitude
button; the autopilot maintains the indicator and a horizontal situation indi-
established bank. TCS also allows cator (HSI). One option adds a multi-
maneuvering without disengaging any function display (MFD) that replaces
selected flight director modes. If the the radar indicator; this is the three-tube
autopilot is operating in indicated air- EFIS. A second option adds a flight
speed, vertical speed, or altitude hold director and EFIS for the copilot.
mode, pressing the TCS button on the A standard flight director system
control wheel allows maneuvering of consists of:
the aircraft to a new reference. After
■ flight director computer
releasing the TCS button, the autopi-
lot maintains the new airspeed, verti- ■ flight director mode selector

cal speed, or altitude reference. ■ accelerometer

Before autopilot engagement, AFM ■ roll-rate gyro.

limitations require autopilot testing dur-
ing the TAXI checklist. Pressing the Flight Director Computer
TEST EACH FLIGHT button on the
The flight director computer processes
autopilot controller before takeoff with
aircraft position information from the
the autopilot engaged initiates the
navigation sensors, vertical and direc-
torque monitor (current monitor) test
tional gyros, and air data system to pro-
sequence. The system simulates torque
vide pitch and roll cues to the EADI
limiter failure to verify operation of the
and commands to the autopilot system.
autopilot disengagement system. While
holding the button, the AP TORQUE The unit contains the necessary elec-
lights and ROLL MONITOR light tronics to provide the following:
should illuminate and the autopilot ■ heading, roll, and pitch attitude hold
should disengage within two seconds. modes
The AUTOPILOT OFF lights illumi- ■ heading select, vertical navigation,
nate, and the warning horn sounds for and long range navigation modes
approximately one second. If the AP
■ VOR (NAV), VOR approach, local-
TORQUE lights or ROLL MONITOR
izer (NAV), ILS approach modes

5B-32 For SimuFlite training only Citation V

March 1999
 Avionics

■ vertical speed, indicated speed, and ■ BC – follow the localizer back course
altitude hold modes ■ ALT – maintain current altitude or
■ back course, altitude preselect, and altitude indicated on vertical navi-
vertical navigation modes. gation computer/controller
■ ALTSEL – arm and capture alti-
Mode Selector
tude selected through vertical nav-
The flight director has an MS-205 igation computer/controller
mode selector panel (Figure 5B-43) on ■ VNAV – arm and capture a VOR/
the center instrument panel; dual flight DME derived vertical profile (ver-
director installations have two mode tical navigation)
selectors. The panel contains latching
pushbuttons that select all of the flight ■ VS – maintain current vertical speed
director modes except go-around; a or allow selection of a new vertical
pushbutton on each throttle lever selects speed by PITCH wheel movement
this mode. Each button illuminates amber ■ IAS – maintain current indicated
to indicate mode arming and green to airspeed or allow selection of a new
indicate mode capture. Depending on vertical airspeed through PITCH
the mode selected, the EADI also indi- wheel movement.
cates mode selection and status.
Accelerometers
Pressing a selector button commands
the flight director computer to follow An accelerometer provides aircraft
signals from navigation aids, accept acceleration information to the flight
pilot inputs from the system con- director system. The unit is self-con-
trollers, or to maintain the established tained, closed-loop, forced-balanced
aircraft attitude: device that provide a linear output pro-
■ FD OFF – remove command bars portional to the aircraft’s acceleration
from view on the EADI while main- to the flight director computer.
taining flight director operation
■ HDG – follow inputs from the
Flight Instruments
heading bug on the horizontal situ- There are three flight instrument
ation indicator options:
■ NAV – capture and track selected ■ two-tube EFIS; single flight director

navigation aid signals from either a ■ three-tube EFIS; single flight director
VOR, ILS, or LOC ■ five-tube EFIS; dual flight directors.
■ APR – capture and track vertical
The single flight director installations
and lateral deviation signals from
have either a two-tube or three-tube
a VOR or ILS

5B-43

Citation V For SimuFlite training only 5B-33

March 1999
 EFIS for the pilot and mechanical Copilot’s Horizontal
instruments for the copilot. The three- Situation Indicator
tube EFIS replaces the weather radar
display with a multifunction display The copilot’s horizontal situation indi-
unit (MFD). The copilot’s attitude indi- cator (HSI), on aircraft with a single
cator and horizontal situation indicator flight director, displays compass head-
do not integrate with the flight director ing, glideslope and localizer devia-
system. The five-tube EFIS installation tion, and aircraft position relative to
has electronic flight instruments for the a VOR radial (Figure 5B-45). The
pilot and copilot; both sets of instru- directional gyro (compass system)
ments integrate with a flight director. drives the rotating compass card.
An aircraft symbol shows aircraft
Copilot’s Attitude position and heading in relation to an
Indicator azimuth card, lateral deviation bar,
The copilot’s attitude indicator (Figure and selected heading. The azimuth
5B-44), on aircraft with a single flight card displays heading information
director, is an electrically-driven instru- obtained from a gyro-stabilized mag-
ment with an internal gyroscope that netic compass (copilot’s C-14D direc-
drives a sphere-type display. A con- tional gyro). Heading is read on the
ventional liquid-filled inclinometer is card beneath the lubber line at the top
at the bottom of the indicator. The AI center of the indicator. If the compass
provides 360° of roll information. The system is OFF, if it provides invalid
brown-earth, blue-sky sphere has pitch heading information, or if the HSI
attitude reference marks above and loses power, a HDG flag appears at
below the horizon line. Roll angle the top of the display.
marks are to the left and right of the The HSI displays glideslope deviation
wings level mark. with a pointer on the right side of the
Pulling the PULL TO ERECT knob compass card. Dots above and below
out cages the gyro. If the gyro loses the center mark indicate the deviation
power, a red GYRO flag appears on above and below the glideslope. If the
the left side. ILS signal is unusable or unreliable,
a VERT flag appears over the pointer.
The AI operates on 28V DC supplied A course deviation bar in the center
by the EMERGENCY bus through the of the instrument displays the devia-
2-amp ADI 2 circuit breaker. It oper- tion, left or right, from the VOR radi-
ates when the battery switch and al or the localizer. If the navigation
avionics master switch are ON, and signal is lost, unreliable, or invalid, a
when the battery switch is in EMER. NAV flag appears on the left side.

5B-44 5B-45

5B-34 For SimuFlite training only Citation V

March 1999
 Avionics

EFIS Degree lines on the sphere display air-

craft pitch from 5 to 90° either nose-
Electronic flight instrument systems up or nose-down. Roll attitude marks
(EFIS) replace the conventional on the top of the display show aircraft
mechanical attitude director indicator bank angle (roll) from 0 to 60°.
(ADI) and horizontal situation indi-
cator (HSI) with displays that dupli- The symbol generator electronically
cate these instruments electronically. presents the flight director steering
commands with a single cue “V-bar.”
The system uses displays that are high- When not in use, the steering cues dis-
resolution, black-matrix, shadow-mask appear by selecting FD OFF or when
color cathode ray tubes (CRTs) driven no lateral mode is used.
by a symbol generator and operated
through a display controller and an Besides the normal display, the EADI
instrument control. displays:
■ radio altitude from the Collins ALT-
The only difference between the EFIS
displays is the addition of an incli- 55 radio altimeter
nometer on the bottom of the EADI. ■ decision height set through the dis-
The displays are interchangeable and play controller
can display both EADI and EHSI ■ flight director mode annunciators
information on one display if the other
■ marker beacon passage
should fail.
■ rising runway symbol beginning at
EADI 200 ft AGL
The electronic attitude director indi- ■ glideslope deviation
cator (Figure 5B-46) is a color CRT ■ expanded localizer deviation
driven by a symbol generator. The
■ rate-of-turn
EADI presents a fixed aircraft sym-
bol superimposed against a moving ■ vertical navigation display.
blue-sky, brown-earth true-sphere to The EADI displays digital warnings
indicate aircraft attitude. A fast/slow instead of flags to represent system
pointer for the angle-of-attack system failures. If the flight director fails, an
is to the left of the sphere, a glides- amber FD warning appears and the
lope indicator is on the right, and an flight director steering commands dis-
inclinometer is at the bottom. Besides appear. If the vertical gyro fails, a red
the normal aircraft attitude informa- ATT FAIL message appears instead
tion, the EADI displays system status of the airplane symbol and the sphere
and mode selection information turns solid blue. If the glideslope,
around the sphere.

5B-46

Citation V For SimuFlite training only 5B-35

March 1999
 expanded localizer, fast/slow com- plays only 90° of compass coordi-
mand, or rate-of-turn indicator fails, nates. The advantage of a partial com-
a large red X appears instead of the pass format is the ability to display
system pointer. If an internal failure additional information such as weath-
occurs, the entire display blanks. If er radar returns, navigation aid posi-
the symbol generator or any part of tion, range rings, and wind vector.
its circuitry fails, a large red X appears Format selection is through the dis-
across the entire display with an SG play controller.
FAIL annunciation. System failures displayed on the EHSI
The display also repeats the flight include heading, course, azimuth and
director modes selected on the display. vertical deviation failure. If the com-
pass system fails, a red HDG FAIL
EHSI message appears at the top of the dis-
The electronic horizontal situation play while the bearing and course
indicator (Figure 5B-47) presents the pointers and annunciators disappear.
same information as its mechanical If a course, vertical, glideslope or ele-
counterpart. The EHSI displays: vation deviation system fails, a red X
appears over the affected scale. A red
■ aircraft position and heading with
V (vertical), G (glideslope), or E (ele-
respect to magnetic or true north
vation) appears near the affected scale.
■ selected heading and selected
If either the EADI or EHSI display
course
fails, the operating display can pre-
■ DME slant range from 0 to 399.9 sent a combined EADI/EHSI display.
nm or 0 to 3,999 nm Turning the failed display’s dimming
■ vertical deviation from glideslope control completely off signals the sym-
■ TO/FROM indication from the
bol generator to present a combined
display on the operating unit. The
selected VOR navigation aid
phase of flight, navigation radio tun-
■ bearing and track information ing, selected flight director mode, or
■ ground speed in knots other flight director function deter-
■ navigation source selected.
mines the information displayed on
the single operating display. Normally,
The EHSI operates in either a full- this information includes the EADI
compass or partial-compass format. sphere with flight director steering
Full-compass format displays a 360° commands and a horizontal compass
compass rose; partial-compass dis- display.

5B-47 5B-48

5B-36 For SimuFlite training only Citation V

March 1999
 Avionics

Display Controller 1 selects the No. 1 ADF; NAV 1 selects

VOR No. 1, and FMS displays FMS
The DC-811 display controller (Figure
information on the HSI bearing point-
5B-48) selects navigation sources,
er. The BRG ◊ knob operates the same
bearing data, and display format for
as the BRG O knob; it selects the same
the EADI and EHSI. Buttons on the
information from the No. 2 navigation
controller include:
receivers. The ADI DIM and HSI DIM
■ FULL/ARC – selects either full arc knobs vary the brightness of the dis-
(360°) or partial compass (90°) for plays. Turning either knob fully coun-
the EHSI terclockwise turns the display OFF and
■ MAP – changes EHSI to partial presents a composite EADI/EHSI dis-
compass format and displays nav- play on the operating display. The WX
igation aid information DIM knob controls the brightness of
■ WX – superimposes weather radar
weather radar returns on the EHSI in
partial compass format.
information on the partial compass
display Multifunction Display Unit
■ GS/TTG – selects ground speed
The multifunction display unit (MFD)
(GS) and time-to-go (TTG) for dis- (Figure 5B-49) primarily functions as
play on the EHSI a weather radar indicator. It also func-
■ ET – elapsed time selection and dis- tions as a navigation data and check-
play list display and provides backup to the
■ NAV – selects between NAV 1 and EHSI and the EFIS symbol genera-
NAV 2 or between ILS 1 and ILS 2 tor(s). If one of the EHSIs fail, the
receivers for the EHSI MFD can replace it. If one of the sym-
bol generators fail, the MFD symbol
■ FMS – selects flight management
generator can replace the failed unit
system (FMS) information for dis- with no degradation of capabilities.
play on the EHSI
The MFD as a navigation display
■ TEST – pressing the TEST button
operates in map mode, map mode with
on the ground initiates a system weather, and flight plan mode. These
test. If the system passes, TEST modes display:
appears on the displays; if it fails,
■ actual aircraft heading
FD FAIL appears.
■ tuned VOR/DME symbols
The BRG O knob selects the naviga-
tion source driving the HSI bearing ■ aircraft symbols
pointer. OFF removes the pointer; ADF

5B-49

Citation V For SimuFlite training only 5B-37

March 1999
 ■ selected navigation course on the EADI and EHSI. Navigation
■ distance to a waypoint and guidance equipment provide pitch
and roll information, heading, glides-
■ waypoint and waypoint data
lope, localizer (LOC), course devia-
■ crosstrack deviation tion, bearing (NAV and ADF), aircraft
■ desired track line position, airspeed, and ground speed.
The symbol generator processes this
■ aircraft depiction as related to a
information, then displays it on the
desired track. respective displays.
Combining the map mode with the
weather radar presents navigation infor- Course and Heading Controller
mation overlayed with radar returns. A remote course and heading con-
This assists in navigating around poten- troller (Figure 5B-51) for each option
tially hazardous weather. of the flight director system allows
Checklist mode allows the user to the setting of course and heading
access multiple checklists for normal information on the horizontal situa-
and emergency procedures. tion indicator. A COURSE knob sets
the desired navigation course used
MFD Controller with the VOR for navigation. Pulling
The MFD controller (Figure 5B-50) the knob out and then releasing it auto-
allows selection of MFD modes and for- matically slews the course pointer on
mat. Through the controller, the pilot the HSI/EHSI to center the course
can select map, plan, and/or weather deviation to the tuned VOR facility.
mode. The controller also allows the The HEADING knob sets the head-
display and scrolling of normal and ing bug on the HSI/EHSI.
emergency checklists. If a symbol gen-
erator or EHSI fails, reversionary switch-
Vertical Navigation
ing is available through the controller. As part of the flight control system,
the vertical navigation system (VNAV)
Symbol Generator provides inputs for vertical navigation
The symbol generators are the core of and altitude preselect modes. The
the EFIS installation. The symbol gen- VNAV system also functions as part
erator processes information from the of the altitude alerting system (see
navigation receivers, gyros, weather Altitude Alerting and Reporting).
radar, and guidance systems to pro- The system consists of a vertical nav-
vide an integrated display of naviga- igation computer/controller (Figure
tion and aircraft guidance information 5B-52) that provides a data input selec-

5B-50 5B-51 5B-52

5B-38 For SimuFlite training only Citation V

March 1999
 Avionics

tor, mode selector, and data display. The Attitude and

computer/controller allows entry of data
in the following order: Heading Reference
■ preselected altitude for altitude The attitude and heading reference
alerting and waypoint selection systems provide pitch, roll, and stabi-
■ elevation of navigation station lized heading reference information
to the flight control, flight instrument,
■ nautical miles waypoint is before
and weather radar systems.
the station (-bias)
The system consists of:
■ nautical miles waypoint is after the
■ VG-14A vertical gyros
station (+bias)
■ C-14D directional gyros
■ angle to waypoint during a descent
or climb. ■ RG-204 rate gyro.

After selecting a mode with the selec- Vertical Gyro

tor knob, data entry is through the SET
knob. The Honeywell VG-14A vertical gyro
provides pitch and roll attitude infor-
The system integrates with the air data
mation to the electronic attitude direc-
system and navigation receivers for
tor indicator (EADI), flight director,
position information and with the flight
autopilot, and weather radar. The
director system for mode control. The
weather radar uses pitch and roll infor-
VNAV system, in turn, provides pitch
mation for antenna stabilization.
guidance and path deviation indica-
tions to the ADI and HSI to allow the The gyro consists of an electrically-
crew to guide the aircraft to cross an driven gyro rotor spinning on a vertical
aimpoint at a desired altitude. axis, pitch and roll synchros, torquers,
gravity erection system, and monitor-
Before performing vertical navigation,
ing circuits. Gimbals within the unit
the crew selects and tunes the VOR-
limit the amount of movement around
TAC station that provides azimuth and
the pitch and roll axes. The gyro is free
distance measuring information. Once
to move 80° up and down and 360°
selected, arming requires selection of
around the roll axis. Pitch and roll tor-
the desired waypoint and waypoint
quers erect the rotor on its vertical axis
data and selection of the VNAV mode
to prevent gyro tumble.
through the flight director mode con-
trol panel. The crew then sets the way- As the aircraft moves about its pitch
point elevation by selecting ALT and and roll axes, the pitch and roll syn-
entering the waypoint altitude through chros react to the movement of the
the SET knob. gyro gimbal. As the gimbal moves, the
synchros provide electrical signals pro-
The system then requires station ele- portional to the gyro displacement.
vation by selecting STAEL and enter-
ing the appropriate altitude and the Directional Gyros
distance before or after the station by
either selecting TO or FROM. Once The Citation V has dual Honeywell C-
altitude, station elevation, and distance 14D compass systems for the pilot and
are entered, the system computes the copilot. Both provide 360° of magnetic
angle necessary to reach the station at heading information to the horizontal
the desired altitude. The VNAV sys- situation indicators (HSIs), radio mag-
tem then supplies the necessary sig- netic indicators (RMIs), autopilot, flight
nals to the flight director system to director, and other guidance systems.
compute and display the required Each system consists of a directional
steering commands to reach the way- gyro, flux valve, remote compensator,
point at the desired altitude. and control switch.

Citation V For SimuFlite training only 5B-39

March 1999
 Each directional gyro has an electrical- The GYRO MAN/AUTO switches
ly-driven gyro and a base assembly with (Figure 5B-53) control the C-14D com-
an inverter. On aircraft with single EFIS, pass systems. In AUTO, the gyro auto-
the pilot’s compass system receives 28V matically aligns itself to the magnetic
DC from the RH Crossover bus; the heading indicated by the flux valve. In
copilot’s compass system receives MAN, manual gyro alignment is through
power from the Emergency bus. On air- the two-position (LH/RH) SLAVE
craft with dual EFIS, the pilot’s com- switch. Placing the SLAVE switch in
pass system receives power from the either LH or RH moves the copilot’s HSI
Emergency bus and the copilot’s sys- compass card left or right at a rate of
tem receives power from the RH Main approximately 30° per minute. This pro-
bus Extension. If the main AC inverters vides a short-term heading reference
fail, on aircraft equipped with two or when magnetic information is unreliable.
three tube EFIS systems, the copilot’s
compass system inverter provides 26V Rate Gyros
AC power to No. 2 navigation radio Rate gyros provide rate-of-turn signals
(NAV 2) and the copilot’s HSI. On to the flight director and autopilot com-
aircraft equipped with five tube EFIS puters. The flight director drives the
systems, the pilot’s compass system rate-of-turn display on the attitude direc-
inverter provides 26V AC power to the tor indicator. The gyro consists of an
copilot’s RMI. electrically-driven rotor that is free to
The flux valve senses the strength and move on its horizontal axis; a torsion
direction of the earth’s magnetic field bar restricts movement around the roll
and converts it into electrical signals axis. As the aircraft turns, it tries to
for gyro compensation. The compen- move the gyro in the direction of the
sating signal aligns the gyro with mag- aircraft turn. The gyro resists this move-
netic north. The remote compensator ment and reacts by rolling on its longi-
uses adjustable permanent magnets to tudinal axis 90° to the applied force. As
counteract the effect of magnetic fields the gyro moves, a transmitter synchro
on the flux valves created by direct on the other end of the torsion bar pro-
current and ferrous materials in the duces an electrical signal proportional
aircraft. to the rate-of-turn.

5B-53

5B-40 For SimuFlite training only Citation V

March 1999
 Avionics

Preflight Autopilot Torque Failure

During the preflight inspection, Illumination of the AP TORQUE light Preflight and
in flight indicates failure of a torque
remove the pitot/static and angle-of-
attack probe covers. Check the pitot sensing device, failure of torque Procedures
probes and static ports for freedom switching, or improper torque appli-
from obstructions. Carefully check the cation. With the torque limiter failed,
angle-of-attack probe for freedom of the autopilot must remain off during
movement and that the slots are free all flights conducted above 14,500 ft.
from obstructions. Check the pres-
ence, condition, and security of the
Autopilot Roll Monitor
static dischargers. Check the commu- Failure
nication and navigation antennas for Illumination of the AP ROLL MON-
security and damage. ITOR light in flight indicates failure
of the roll monitor. Disengage the
Abnormal autopilot.
Procedures Pitot/Static Heater Failure
Abnormal procedures for the avion- Illumination of either the LH or RH
ics systems include (please refer to P/S HTR OFF annunciator indicates a
your SimuFlite Citation V Operating failure in the pitot/static heating sys-
Handbook for a detailed checklist): tem due to power loss (PITOT & STA-
■ autopilot torque failure TIC anti-ice switch off) or equipment
■ autopilot roll monitor failure malfunction. The pilot’s pitot/static
heating system receives 28V DC
■ pitot/static heater failure
power the Left Main bus Extension,
■ copilot’s pitot/static instruments and the copilot’s system receives
indications erratic power from the Emergency bus. Check
■ angle-of-attack probe heater failure the anti-ice switches and circuit break-
ers to locate the failed system. If the
■ EFIS display failure
pilot’s system fails, the autopilot alti-
■ EFIS heading failure tude hold function is inoperative.
■ EFIS attitude failure (dual EFIS)
Angle-of-Attack Probe
■ EFIS cooling fan failure
Heater Failure
■ EADI/EHSI overtemperature
Illumination of the AOA HTR FAIL
■ EFIS symbol generator overheat
light indicates failure of the angle-of-
(dual EFIS) attack probe heater. Check the PITOT
■ EFIS symbol generator failure (dual & STATIC switch and the AOA circuit
EFIS) breaker. If the system is still inopera-
■ single radio management unit fail-
tive, avoid areas of known or forecast
ure (Primus II radios) icing. If the AOA probe ices over, the
stick shaker is inoperative.
■ dual radio management unit failure
(Primus II radios) Erratic Instrument
■ NOSE COMP O’TEMP annuncia- Indications
tor (aircraft 0075 and subsequent) If one or more of the copilot’s pitot/
■ angle-of-attack and/or slow/fast static instruments (Mach/airspeed indi-
indicator flagged. cator, IVSI, and altimeter) behave errat-
ically, there is a leak in the copilot’s

Citation V For SimuFlite training only 5B-41

March 1999
 pitot/static system. Lift the guards on the and course pointers should reappear.
TAS COMP/PITOT-STATIC switches Continue the flight using the secondary
on the copilot’s instrument panel; close heading source. On aircraft with dual
both switches to isolate the TAS trans- EFIS, verify that both EHSIs display
ducer from the pitot/static system. Pull the same heading source.
the TAS CB on the right CB panel.
Loss of the TAS requires manual entry EFIS Attitude Failure
of airspeed into the flight management On aircraft with dual EFIS, appear-
system. ance of a red ATT FAIL message and
a solid blue sphere with loss of the
EFIS Display Failure pitch and roll scales indicates a loss of
Blanking of an EADI or EHSI indi- attitude information to the EADI. Press
cates a display failure. The EFIS system the ATT REV button on the failed side
can provide a combined EADI/ EHSI to select the opposite attitude source.
on the operating display. Turning the The sphere, pitch scale and roll scale
affected display DIM knob to OFF should reappear. The ATT FAIL mes-
(fully counterclockwise) selects the sage should disappear. Verify that the
composite EADI/EHSI display on the EADI displays the selected attitude
operating display. source (ATT 1 or ATT 2).
Operation in the composite mode is EFIS Cooling Fan Failure
approved only with the flight director
selected. Takeoff with either display Illumination of the yellow DISP FAN
inoperative is prohibited; both the light indicates failure of an EFIS cool-
EADI and EHSI must be operational ing fan.
before takeoff. If the fan fails on the ground, remove
On aircraft with an MFD, the MFD power from the system by pulling the
can replace a failed EHSI. If an EADI applicable EFIS, EADI, and EHSI cir-
or EHSI fails, turning the affected dis- cuit breakers to prevent overheating and
play’s DIM knob to OFF displays a possible failure of the system. Limit
composite EADI/EHSI display on the ground operating time to 10 minutes
operating display. Turning the MFD with the light on. Before takeoff, reset
mode selector button to HSI moves the pulled circuit breakers and monitor
the EHSI display to the MFD display. the EADI HOT and EHSI HOT lights.
In this mode, all operation of the If the fan fails in flight, monitor the
EADI and EHSI are unchanged EADI HOT and EHSI HOT lights. In
(except for dimming). All normal flight, there is adequate air flow to
functions of the MFD are inoperative. prevent system overheating. Reduce
Dispatch is not approved in this mode. cockpit temperature to prevent system
overheating.
EFIS Heading Failure Observe system operating limitations
Appearance of a red HDG FAIL mes- with a DISP FAN light illuminated.
sage and loss of bearing and course
pointers on the EHSI indicates a loss EADI/EHSI
of heading information to the EHSI. Overtemperature
Pressing the HDG REV button selects If the EADI HOT or EHSI HOT lights
between the pilot’s and copilot’s com- illuminate in flight, select MANUAL
pass system. The HDG REV button COLD on the cockpit temperature
illuminates and MAG 1 or MAG 2 selector to provide additional system
illuminates on the EHSI (depending on cooling. Use the display controller
the side selected); the HDG FAIL mes- DIM knob to reduce display bright-
sage should disappear and the bearing

5B-42 For SimuFlite training only Citation V

March 1999
 Avionics

ness. If this fails to extinguish the EADI If the MFD symbol generator fails with
or EHSI HOT light, pull the affected a failed EADI/EHSI symbol generator,
display’s circuit breakers. Enable the push the SG Normal/SG Rev button on
composite display by turning the affect- the failed side to have the operating
ed display DIM knob to OFF. side’s symbol generator replace the
Observe the system operating limita- failed one; both sides of the system are
tions with an EFIS overtemperature using one symbol generator.
condition. Single Radio Management
Symbol Generator Unit Failure
Overheat On aircraft with Primus II radios:
On aircraft with dual EFIS, indications If a radio management unit (RMU)
of a symbol generator overheat due to fails, pushing the 1/2 (cross-side) key
cooling fan failure are SG HOT or MG transfers the entire RMU operation
HOT messages on the EFIS displays. and display to the opposite system.
The legend color changes from white
If the SG HOT light appears on the to magenta to indicate control of the
pilot’s or copilot’s display, use the MFD opposite side.
mode selector knob to select the MFD
symbol generator. Verify that MG Dual Radio Management
appears on the affected side’s EADI. Unit Failure
Pull the affected EFIS circuit breakers.
On aircraft with Primus II radios:
If the MG HOT light appears on the If both Honeywell RMUs fail, select
MFD, disengage the autopilot by press- emergency mode through the clear-
ing the AP/TRIM DISC button. Pull ance delivery unit (CDU); the CDU
the MFD SYM GEN and MFD DISP inhibits communication and naviga-
circuit breakers to remove power to the tion radio tuning through either RMU.
MFD display and symbol generator. The EMRG annunciator on the CDU
Do not engage the autopilot altitude display indicates emergency mode
hold mode with the MFD SYM GEN selection. Tune the COMM and NAV
circuit breaker pulled. radio. During emergency mode oper-
If the MFD symbol generator was replac- ation, the NAV or channeled DME is
ing a failed EADI/EHSI symbol gener- no longer tuned and the DME display
ator, select the opposite symbol genera- is invalid.
tor (cross-side operation) and verify that
SG 1 or SG 2 appears on both EADIs. NOSE COMP O’TEMP
Annunciator
Symbol Generator Failure
On aircraft 0075 and subsequent,
If a symbol generator failure occurs excessive temperature in the nose
on aircraft with a single EFIS instal- avionics compartment illuminates the
lation, refer to Emergency Procedures. NOSE COMP O’TEMP annunciator.
On aircraft with dual EFIS, an oper- Ground operation of avionics equip-
ating symbol generator can replace a ment for longer than one hour with an
failed unit through reversionary switch- air temperature greater than 45°C may
ing. Indications of a failed symbol gen- result in avionics compartment over-
erator are a red SG FAIL on a display heating. High avionics compartment
or loss of an EADI, EHSI, or MFD. temperatures can result in EFIS sys-
Selecting the SG (left or right) on the tem failures and damage to sensitive
MFD controller replaces the failed avionics equipment. Restrict ground
symbol generator with the MFD sym- operating times to prevent compart-
bol generator; the MFD blanks. ment overheating.

Citation V For SimuFlite training only 5B-43

March 1999
 If the NOSE COMP O’HEAT annun- EFIS Symbol
ciator illuminates on the ground, deter- Generator Overheat
mine and correct the fault before
flight. Allow the nose compartment An SG HOT message on the EFIS
to cool before continuing ground oper- displays indicates symbol generator
ation or flight. overheat due to fan failure. Continued
use of the system with the fan failed
If the annunciator illuminates in flight, may result in a system failure due to
land as soon as practical. Determine overheating. Pull the associated EFIS
the cause before the next flight. Usually, (DC), EADI, and EHSI circuit break-
the annunciator does not illuminate ers to remove power from the system.
unless there is an avionics system
failure. Continue the flight with the copilot’s
flight director system and the stand-
Angle-of-Attack by gyro horizon.
Indicator Flagged
Autopilot Hardover
Indications of an angle-of-attack sys-
tem failure are an OFF flag on the A sudden nose-down pitching with
indicator or a red X over the slow/fast the autopilot engaged denotes an
indicator on the EADI(s). The slow/ autopilot hardover. Immediately dis-
fast indicator may be unreliable and engage the autopilot by pressing the
the stick shaker inoperative. AP/ TRIM DISC button on either con-
trol wheel. Regain aircraft control.
Emergency During certification, demonstrated
maximum altitude losses were:
Procedures ■ cruise – 500 ft at 45,000 ft cruis-
Emergency procedures for the avion- ing altitude
ics systems include (please refer to
■ climb – 350 ft at 17,000 ft
your SimuFlite Citation V Operating
Handbook for a detailed checklist): ■ ILS approach – 28 ft.

■ EFIS symbol generator failure (sin- Refer to the Aircraft Flight Manual
gle EFIS) for the Autopilot Glideslope Deviation
■ EFIS symbol generator overheat
Profile.
■ autopilot hardover. Red Gun Display Failure
WARNING: Following a red
gun failure in a display unit, the EFIS Symbol The color in the EFIS display tubes is
Generator Failure produced by a combination of three
red warning annunciators are not
color guns inside the tube. Failure of
visible. Either a large red X and SG FAIL on the red gun changes sky color in the
the EADI or blanking of both the EADI from dark blue to dull dark blue
EADI and EHSI indicates a symbol and the ground from brown to green.
generator failure. Loss of the symbol In the EHSI, the compass rose turns
generator renders the entire EFIS sys- from white to blue. Operation with a
tem inoperative. Continue the flight red gun failure should be undertaken
with the copilot’s flight director sys- with caution because the red warning
tem and the standby gyro horizon. annunciators are not visible.

5B-44 For SimuFlite training only Citation V

March 1999
This section describes the electrical systems and lighting
systems of the Citation V aircraft. Electrical
The Citation V electrical system includes direct current (DC) that
powers the majority of the aircraft systems, and alternating
Systems
current (AC) that powers avionics equipment.
Two engine-driven starter/generators provide DC power to their
respective distribution buses for primary aircraft electrical power.
Chapter 5C
The battery supplies secondary DC power through the Hot
Battery bus, which also serves as the input for ground power.
Two static inverters convert DC to AC. Four distribution buses
carry power from the inverters.
Lighting on the Citation V includes cockpit, passenger cabin,
emergency, and exterior lights.

Citation V For SimuFlite training only 5C-1

March 1999
5C-2 For SimuFlite training only Citation V
March 1999
 Electrical System

DC System Schematic . . . . . . . . . . . . . . . . 5C-6

Table of
DC Circuit Breakers Schematic . . . . . . . . . . . 5C-7
DC System . . . . . . . . . . . . . . . . . . . . . . 5C-9
Contents
Starter/Generators . . . . . . . . . . . . . . . . . . 5C-9
Generator Control Units . . . . . . . . . . . . . . 5C-10
Generator Control Switches . . . . . . . . . . . . 5C-10
Ammeters . . . . . . . . . . . . . . . . . . . . . 5C-10
GEN OFF Annunciators . . . . . . . . . . . . . . 5C-10
Battery . . . . . . . . . . . . . . . . . . . . . . . . 5C-11
Battery Switch . . . . . . . . . . . . . . . . . . . 5C-11
Voltmeter . . . . . . . . . . . . . . . . . . . . . . 5C-12
Battery Overheat Warning . . . . . . . . . . . . . 5C-12
Battery Temperature Gage . . . . . . . . . . . . . 5C-13
Ground Power . . . . . . . . . . . . . . . . . . . . 5C-13
GPU Protection . . . . . . . . . . . . . . . . . . 5C-13
Junction Boxes . . . . . . . . . . . . . . . . . . . 5C-14
DC Distribution Buses . . . . . . . . . . . . . . . . 5C-14
Hot Battery Bus . . . . . . . . . . . . . . . . . . 5C-14
Emergency Bus . . . . . . . . . . . . . . . . . . 5C-15
Battery Bus . . . . . . . . . . . . . . . . . . . . . 5C-15
Main Buses . . . . . . . . . . . . . . . . . . . . . 5C-16
Main Bus Extensions . . . . . . . . . . . . . . . . 5C-16
Crossover Buses . . . . . . . . . . . . . . . . . . 5C-16
DC Relays . . . . . . . . . . . . . . . . . . . . . . 5C-16
Start Relay . . . . . . . . . . . . . . . . . . . . . 5C-17
Emergency Power Relay . . . . . . . . . . . . . . 5C-17
Battery Relay . . . . . . . . . . . . . . . . . . . . 5C-17
Battery Disable Relay . . . . . . . . . . . . . . . 5C-17
Power Relay . . . . . . . . . . . . . . . . . . . . 5C-18
Battery Disconnect Relay . . . . . . . . . . . . . 5C-18

Citation V For SimuFlite training only 5C-3

March 1999
 DC Protection . . . . . . . . . . . . . . . . . . . . 5C-18
Generator Control Units . . . . . . . . . . . . . . 5C-18
Circuit Breakers . . . . . . . . . . . . . . . . . . 5C-18
Current Limiters . . . . . . . . . . . . . . . . . . 5C-19
DC Operation . . . . . . . . . . . . . . . . . . . . 5C-19
GPU Start . . . . . . . . . . . . . . . . . . . . . 5C-19
Battery Start . . . . . . . . . . . . . . . . . . . . 5C-20
Cross Generator Assisted Start . . . . . . . . . . 5C-21
Terminating Abnormal Starts . . . . . . . . . . . . 5C-23

DC Buses . . . . . . . . . . . . . . . . . . . . . . 5C-25

AC System Schematic . . . . . . . . . . . . . . . 5C-28

AC Circuit Breakers Schematic . . . . . . . . . . 5C-28
AC System . . . . . . . . . . . . . . . . . . . . . 5C-29
Components . . . . . . . . . . . . . . . . . . . . . 5C-29
Inverters . . . . . . . . . . . . . . . . . . . . . . 5C-29
Inverter Control . . . . . . . . . . . . . . . . . . 5C-29
Switching Relays . . . . . . . . . . . . . . . . . 5C-30
Failure Annunciators . . . . . . . . . . . . . . . 5C-30
AC Buses . . . . . . . . . . . . . . . . . . . . . . 5C-31
AC Protection . . . . . . . . . . . . . . . . . . . . 5C-31
Circuit Breakers . . . . . . . . . . . . . . . . . . 5C-31

Preflight and Procedures . . . . . . . . . . . . . 5C-33

Preflight . . . . . . . . . . . . . . . . . . . . . . . 5C-33
Abnormal Procedures . . . . . . . . . . . . . . . . 5C-33
Single Generator Failure . . . . . . . . . . . . . . 5C-33
Single Inverter Failure . . . . . . . . . . . . . . . 5C-33
Emergency Procedures . . . . . . . . . . . . . . . 5C-34
Dual Inverter Failure . . . . . . . . . . . . . . . . 5C-34
AC Fail and AC Power Distribution Failure . . . . 5C-34
Battery Overheat . . . . . . . . . . . . . . . . . . 5C-34

5C-4 For SimuFlite training only Citation V

March 1999
 Electrical System

Electrical Fire or Smoke . . . . . . . . . . . . . . 5C-35 DC Electrical

Loss of Both Generators . . . . . . . . . . . . . . 5C-36 System
Generator Reset Decision Tree . . . . . . . . . . 5C-37

Data Summary . . . . . . . . . . . . . . . . . . . 5C-39

Electrical System Data . . . . . . . . . . . . . . . . 5C-39

Lighting . . . . . . . . . . . . . . . . . . . . . . . 5C-41
Interior Lighting . . . . . . . . . . . . . . . . . . . 5C-41
Cockpit Lighting . . . . . . . . . . . . . . . . . . 5C-41
Primary Lighting . . . . . . . . . . . . . . . . . 5C-41
Secondary Lighting . . . . . . . . . . . . . . . . 5C-42
Passenger Compartment . . . . . . . . . . . . . 5C-42
Fluorescent Lighting . . . . . . . . . . . . . . . 5C-42
Reading Lights . . . . . . . . . . . . . . . . . . 5C-43
Entrance Light . . . . . . . . . . . . . . . . . . 5C-43
Footwell . . . . . . . . . . . . . . . . . . . . . . 5C-43
Passenger Advisory . . . . . . . . . . . . . . . 5C-43
Baggage Compartment . . . . . . . . . . . . . . 5C-44
Exterior Lighting . . . . . . . . . . . . . . . . . . . 5C-44
Navigation Lights . . . . . . . . . . . . . . . . . . 5C-45
Anti-Collision Lights . . . . . . . . . . . . . . . . 5C-45
Rotating Beacon . . . . . . . . . . . . . . . . . . 5C-45
Recognition Lights . . . . . . . . . . . . . . . . . 5C-45
Landing/Taxi Lights . . . . . . . . . . . . . . . . . 5C-46
Wing Inspection . . . . . . . . . . . . . . . . . . 5C-46
Emergency Lighting . . . . . . . . . . . . . . . . . 5C-46

Citation V For SimuFlite training only 5C-5 5C-6

March 1999
 DC Electrical System

RIGHT CB PANEL 2 4
6 LH MAIN BUS EXTENSION RH CROSSOVER BUS
EMER BUS AFIS VHF TRANSMITTER ITT-LH ALCOHOL PUMP
ANTI-COLLISION LIGHT LANDING GEAR CONTROL ALT VIBRATOR-RH
AOA SYSTEM COMPUTER/IND LANDING GEAR WARNING CABIN PRESS EMERGENCY
LEFT CB PANEL AOA SYSTEM HEATER LH BUS NO. 1, 2, 3 DIGITAL CLOCK-RH
4 RH CROSSOVER BUS 5 BATTERY OVERTEMP SYSTEM OUTSIDE AIR TEMP DIVIDER FANS
LH CROSSOVER BUS BATTER TEMP GAGE OIL PRESSURE-LH ENG/WING ANTI-ICE VALVE-RH
CABIN FAN OIL TEMP-LH EQUIPMENT COOLING
CABIN PRESS NORM PANEL LIGHTS-5V-LH FIRE DETECT-LH
CABIN TEMP IND PITCH TRIM F/W SHUTOFF
INVERTER 1 INVERTER 2 COCKPIT VOICE RECORDER PITOT/STATIC HEAT-LH FLIGHT HOUR METER
15A 35A 35A 15A DEFOG FAN ROTATING BEACON FUEL BOOST PUMP-
2 LEFT MAIN RH MAIN DIGITAL CLOCK-LH SKID CONTROL LH (SECONDARY)
EL INVERTER SPEED BRAKE IGNITER-LH (PRIMARY)
EXTENSION BUS EXTENSION BUS ENGINE/WING ANTI-ICE VALVE-LH STANDBY GYRO OVERSPEED WARNING
3 ENGINE SYNCHRONIZATION START POWER (PRIMARY) PANEL LIGHTS-5V-CENTER
75A 75A 75A 75A 75A 75A FLIGHT DATA RECORDER T/R POWER AND CONTROL-LH PANEL LIGHTS-5V RH
FIRE DETECT-RH TAIL NAV LIGHT START LIGHTS-RH
F/W SHUTOFF-RH TEMP CONTROL SURFACE DEICE BOOTS
FLAP CONTROL VALVE TURBINE SPEED-LH T/R POWER AND CONTROL-RH
FUEL FLOW-LH W/S BLEED AIR CONTROL WARNING LIGHTS 2
AFT J-BOX 80A 80A 80A 80A 80A 80A FUEL BOOST PUMP- W/S BLEED AIR POWER
225A 225A RH (SECONDARY) WARNING LIGHTS 1
FUEL QUANTITY-LH WING INSPECTION LIGHT-LH 1
LH MAIN BUS BATTERY BUS RH MAIN BUS HOT BATTERY BUS
IGNITER-RH (PRIMARY) WING NAV LIGHT
INVERTER 1 BATT & EMER RELAY
EMER LIGHTING BATTERY PACK
BATTERY EXT LIGHTS

*
IGNITION POWER (START)
BATTERY DISABLE RELAY LH CIRCUIT LOCATOR BEACON
RELAY (CROSS GEN BREAKER NOSE AND BAGGAGE LIGHTS
TAIL CONE LIGHT
200 START) 200 PANEL VOLT METER
100 300 100 300

400 0 400
0
D.C. AMP
BATT D.C. AMP

OFF
EMER
20 30 EMER PWR L/R GEN OFF
6
EMERGENCY BUS (28V DC)
LH 40
RELAY RH RH CIRCUIT
10 AUDIO PANEL (UNIT 0032 & COPILOT'S ATTITUDE IND (SINGLE EFIS)
POWER D.C.
VOLTS POWER Subsequent; prior aircraft with DIRECTIONAL GYRO NO. 1
BREAKER
RELAY GEN RELAY PANEL
OFF SB 560-23-05) LH & RH SPEED IND
COCKPIT FLOOD LIGHTS NAV 2
LH RH COMM 1 NO. 2 VHF NAV REPEATER
VOLTAGE SEL COPILOT'S PITOT/STATIC HEATERS (DUAL EFIS)
COPILOT'S HSI (SINGLE EFIS) VOLTMETER
BATT COPILOT'S RMI (DUAL EFIS)
LH RH
GEN GEN
ENGINE START ENGINE START 3
RH MAIN BUS EXTENSION
L GEN LH DISENGAGE DISENGAGE RH R GEN
ADF 2 INVERTER 2
START START AUDIO 2 (0001-0031 w/o ITT-RH 5 LH CROSSOVER BUS
OFF GCU DISC DISC GCU OFF SB560-23-05) MFD CONTROLLER
RESET RESET COMM 2 MFD DISPLAY ADF 1 EFIS 1 DISPLAY CONTROLLER
LH START RH START DME 2 MFD SYMBOL GEN AFIS VHF RECEIVER EHSI 1
AUDIO 1 (0001-0031 w/o FD 1
RELAY RELAY EADI 2 (DUAL EFIS) OIL PRESS-RH
SB560-23-05 FMS (GNS-X)
EFIS 1 OIL TEMP-RH
LH RH EFIS 2 (DUAL EFIS) RADAR AUTOPILOT WARNING HF COMM 3
GEN HOT BATTERY BUS GEN EFIS 2 CONTROLLER RH BUS NO. 1, 2, 3 AUTOPILOT SERVO NAV 1
(DUAL EFIS RH INST VIBRATOR AUTOPLIOT CONTROL RADIO ALT

28.5V 32.5V 28.5V EHSI 2 (DUAL EFIS) RMI 2 (SINGLE EFIS) DIRECTIONAL GYRO 1 RMI 1
FD NO. 2 (DUAL FD) DME 1 TRANSPONDER 1
300A GROUND BATT 24V 300A TRANSPONDER 2
EADI 1 VLF
OVER VOLTAGE FLITEPHONE TAS HEATER
POWER OVER CURRENT FUEL REMAINING TAS SYSTEM EFIS 1 DISP CONT
RELAY BATTERY FUEL FLOW-RH TURBINE SPEED-RH
DISCONNECT FUEL QUANTITY-RH
1200A
* REFER TO FIGURE 5C-5A
FOR UNITS 001 THRU 031 28.0V
RELAY
1000A (EXTERNAL POWER START)

5C-6 For SimuFlite training only Citation V

March 1999
 Electrical System

DC Circuit Breakers LH CB PANEL

LH FAN LH LH TURB LH FUEL LH FUEL LH OIL LH OIL NORM EMER LH LH LH FW LH FIRE LH

SPEED ITT SPEED FLOW QTY TEMP PRESS PRESS PRESS IGN BOOST SHUTOFF DET BOV
2 2 2 2 2 2 2 5 5 7.5 15 7.5 2 5

ENGINE INSTRUMENTS

STDBY LH LH PITOT LH W/S BLEED CABIN RH RH RH FW RH FIRE RH

GYRO OAT CLOCK STATIC ENG AIR TEMP FAN TEMP IGN BOOST SHUTOFF DET BOV

5 2 2 7.5 5 5 20 5 7.5 15 7.5 2 5

ENVIRONMENTAL ENGINE-FUEL

RH RH RH PITOT RH W/S ANTI BATT LDG WARN WARN OVER VOICE

ALT FLT/HR CLOCK STATIC ENG BLEED AIR NAV COLL TEMP GEAR LTS 1 LTS 2 SPEED RECORDER

2 2 2 7.5 5 5 5 7.5 2 2 5 5 2 5

FLIGHT INSTRUMENTS WARNING

AC AOA SURFACE W/S WING ROTATING SKID LH THRUST ANGLE OF ENGINE NOSE FLIGHT
INVERTER HTR DEICE ALCOHOL INSP BEACON CONTROL REVERSER ATTACK SYNC WHL RPM RECORDER
LH BUS
15 10 5 5 5 5 20 7.5 5 5 2 5
NO 1 NO 2 NO 3
NO. 1 ANTI-ICE
75 75 75
RH CB LH EL LH CENTER RH EQUIP RH THRUST FLAP GEAR PITCH SPEED
PANEL START PANEL PANEL PANEL PANEL COOL REVERSER CONTROL CONTROL TRIM BRAKE
35 7.5 1 5 5 5 7.5 7.5 5 5 5 5

DC POWER LH BUS LIGHTS SYSTEMS

RH CB PANEL

COMM NAV XPDR DME ADF AUDIO RH FAN RH RH TURB RH FUEL RH FUEL RH OIL RH OIL
WARN
1 1 1 1 1 1 SPEED ITT SPEED FLOW QTY TEMP PRESS
2
7.5 3 3 3 3 3 3 2 2 2 2 2 2 2

ENGINE INSTRUMENTS

COMM NAV XPDR DME ADF AUDIO FD EFIS EFIS 1 EHSI EADI DG RMI MFD MFD
2 2 2 2 2 2 1 1 CONT 1 1 1 1 SYM GEN DISP
2 3
7.5 3 3 3 3 3 3 5 1 5 5 5 2 7.5 5

COMM FD EFIS EFIS 2 EHSI EADI DG RMI PHONE RADAR

FMS VLF AFIS AP TAS
3 2 2 CONT 2 2 2 2
3
5 5 5 7.5 7.5 2 3 5 1 5 5 5 2 5 7.5

AVIONICS DC
NAV RMI EFIS VG FD AIR RAD RH AC
AP START
1 1 1 1 1 DATA ALT INVERTER
1 2 2 1 1 1 2 2 7.5 15
RH BUS
NO 2
NO 1 NO 2 NO 3
HSI NAV RMI EFIS VG FD RADAR ADI FLOOD LH CB
75 75 75
2 2 2 2 2 2 2 PANEL
1
2 1 2 2 1 1 1 2 5 35

AC FLT INSTR AVIONICS DC PWR-RH BUS

1 NOT INSTALLED WITH DUAL EFIS

EMERGENCY BUS LH CROSSOVER BUS
LH MAIN BUS EXTENSION RH CROSSOVER BUS 2 0001-0031 w/o SB 560-23-05
AUDIO 1 - LH CROSSOVER BUS
RH MAIN BUS EXTENSION AUDIO 2 - RH MAIN BUS EXTENSION

3 DUAL EFIS INSTALLED

DG 1 - EMER BUS
RMI 2 - EMER BUS

Citation V For SimuFlite training only 5C-7

March 1999
5C-8 For SimuFlite training only Citation V
March 1999
 Electrical System

Two 28V DC, 300 amp starter/gener-

ators are the main source of DC power
power side of the system connect to
the Main bus side through the battery DC System
for the aircraft. A 24V, 44 amp-hour relay. The copilot’s essential instru-
battery supplies power for engine start- ments receive power from the
ing and emergencies. A receptacle Emergency bus through an emergency
below the left engine pylon allows the relay. The battery switch controls the
connection of a 28V DC ground power battery relay and the emergency relay.
unit (GPU). During normal operation, all the buses
Major DC system components include: connect together and the generators
■ two starter/generators with associat- provide power to operate the systems
ed generator control units, control and charge the aircraft battery.
switches, ammeters, and annunciators
■ one battery with associated gages,
Starter/Generators
switches, and annunciators The primary sources of DC power are
■ power junction box (J box)
the two 28.5V DC, 300 amp engine-
driven starter/generators driven by the
■ nine DC buses
engine accessory gearboxes (Figure
■ various DC relays. 5C-1). The starter/generators serve a
The DC buses distribute power from dual function; as motors for engine
the battery, generators, or ground starting, and as generators operating
power unit (GPU) to equipment and in parallel to provide power to the DC
systems requiring DC power. The sys- bus distribution system.
tem is divided into essentially two The generators cross-tie through two
halves. One half contains the Main DC current limiters and the Battery bus. If
buses that connect together to a com- one generator goes off-line, the remain-
mon Battery bus through current lim- ing generator powers both of the Main
iters; the generators supply power to DC buses. The overload capability of
the Main buses through power relays. each generator is 50% for two minutes
The battery and GPU connect to a Hot (450 amps). Either generator is capa-
Battery bus. The battery and ground ble of supplying the entire normal DC
electrical requirements of the aircraft.

5C-1

Citation V For SimuFlite training only 5C-9

March 1999
 The starter/generator, driven by the (refer to DC schematic to trace flow
accessory gearbox, has a shear section to Main bus).
in its drive shaft. If the starter/ gener- The center OFF position disconnects a
ator mechanically fails (i.e., bearing generator from its bus without de-
failure), the shear section fails at 1,500 exciting it by opening its power relay.
inch-pounds of torque to prevent dam-
age to the engine accessory gearbox. The momentary RESET position resets
a generator field relay that tripped from
Each starter/generator has an internal overvoltage, feeder fault, or ENG
fan that cools the unit during ground FIRE switch actuation. Positioning the
operation; in flight, ram air cools the switch to RESET also may be neces-
starter/generator. An air inlet duct on sary following a windmilling airstart.
the forward engine cowling directs air
to the starter/generator. An exhaust Ammeters
duct on the lower cowling directs cool-
The left and right ammeters (Figure
ing air overboard.
5C-3) on the left switch panel display
Generator Control Units the load in amps on each generator.
The ammeters are identical and have a
The generator control units (GCUs) in red radial line at 300 amps that indi-
the tailcone provide voltage regula- cates the maximum generator load.
tion, feeder fault, ground fault, over-
voltage protection, and load sharing. If a generator failure or shutdown
occurs, the load on the affected gen-
Refer to the Protection section in this erator drops to zero and the load on
chapter for further details. the operating generator increases to
Generator Control meet the demand because the buses
are tied together.
Switches
Each generator has a three-position GEN OFF Annunciators
(GEN/OFF/RESET) control switch on The amber LH/RH GEN OFF annun-
the left instrument panel (Figure ciators on the annunciator panel illu-
5C-2). minate when a generator power relay
In GEN, regulation, protection, and opens to disconnect the generator from
Main bus connection are automatic the electrical system.
through the GCU; the generator con- A failure of both generators illuminates
nects to its Main bus once the genera- the LH and RH GEN OFF annuncia-
tor reaches operating speed and begins tors and triggers the red MASTER
supplying power at the correct voltage WARNING lights.

5C-2 5C-3

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March 1999
 Electrical System

Battery switch panel between the left and right

generator switches (Figure 5C-5). The Battery Limitations
The 24V, 44 amp-hour nickel-cadmi- battery switch provides the ground to Do not exceed three engine starts
um battery in the tailcone (Figure control both the battery relay and the per hour.
5C-4) is the secondary source of DC emergency relay. If battery limitation is exceeded, a
power for the aircraft and provides
Control of the battery relay is the same deep cycle including a capacity
engine starting and emergency power.
for all Citation Vs. However, the con- check must be accomplished to
It connects to the aircraft electrical sys-
trol of the emergency relay changes detect possible cell damage. Refer
tem through a quick disconnect that
after unit 0031. to Chapter 24 of the Maintenance
allows the crew to disconnect the bat-
tery during long term parking. The battery relay controls the power Manual for procedure.
line between the Hot Battery bus and Three generator assisted cross
When connected, a ground power unit
the Battery bus. With the relay closed, starts are equal to one battery start.
can charge the battery. The entire DC
power can flow from the Hot Battery
system or only the Emergency bus may
power sources (battery or GPU) to the
be powered by the battery during elec-
Battery bus and then to the Main DC
trical system abnormal or emergency
Buses for distribution. During gener-
situations.
ator operation power flows from the
The battery consists of 20 cells (1.3V Battery bus to the Hot Battery bus to
per cell) connected with busbars in a charge the battery and power the Hot
stainless steel box; the busbars connect Battery bus items. With the battery
to the positive and negative terminals. relay open, the Hot Battery bus and
A steel vent line from the battery case Battery bus are disconnected.
to an opening in the aircraft fuselage
The generators can only charge the
releases battery gases to the atmosphere.
battery with the switch in BATT.
During engine starts using external
The battery relay is normally open
power, the battery disconnect relay
(powered closed). The BATT position
opens to disconnect the battery’s neg-
of the BATTERY switch supplies a
ative terminal from the airframe
ground to close the relay and connect
ground; therefore, external starts do
the two buses. With the switch in OFF
not use battery power and do not count
or EMER, there is no ground and the
in the battery start limitations. At the
spring-loaded relay remains open to
end of the start cycle, the relay closes
disconnect the Battery bus from the
to connect the battery to ground.
Hot Battery bus.
Battery Switch The emergency relay is a normally
open (powered closed) relay that con-
The three-position (BATT/OFF/
trols power to the Emergency bus.
EMER) battery switch is on the left

5C-4 5C-5

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March 1999
 On units 0001 through 0031, the Hot (Figure 5C-7) is in the BATT position.
Battery Limitations Battery bus supplies power to the When the BATTERY and VOLTAGE
If the BATT O’TEMP annunciator Emergency bus when the relay closes; SEL switches are in BATT, the volt-
illuminates during ground no power is available to the Emergency meter displays the highest voltage on
operations, the battery requires bus with the relay open (Figure 5C-5A). the Hot Battery bus (i.e., battery, gen-
maintenance before takeoff. With the BATTERY switch in BATT or erator, or external power). Voltmeter
EMER, the switch supplies a ground to range is 10 to 40 volts. With the VOLT-
close the emergency relay. With the AGE SEL switch in the LH GEN or
switch in OFF, there is no ground and RH GEN position, the voltage output
the emergency relay opens. of the respective generator is displayed.
On units 0032 and subsequent, depend- Battery Overheat Warning
ing on battery switch position, the emer-
gency relay connects the Emergency bus The battery overheat warning system
power line to either the Battery bus or consists of a temperature sensor and a
the Hot Battery bus. With the emergency temperature module. The sensor mea-
relay open, the Emergency bus connects sures the temperature at the center of
to the Battery bus. With the relay closed, the battery. If battery temperature reach-
the Emergency bus receives power from es 145 to 160°F (63 to 71°C), the tem-
the Hot Battery bus. With the BAT- perature module illuminates the red
TERY switch in EMER, the switch sup- BATT O’TEMP annunciator and acti-
plies a ground to close the emergency vates the MASTER WARNING lights.
relay. In this position, the emergency If battery temperature exceeds 160°F
relay connects the Hot Battery bus to (71°C), the BATT O’TEMP annuncia-
the Emergency bus. With the BAT- tor flashes approximately three times
TERY switch in either BATT or OFF, per second.
there is no ground and the relay opens to After an illumination of the BATT
connect the Emergency bus to the O’TEMP annunciator, the battery
Battery bus. requires removal and reconditioning
per the aircraft Maintenance Manual.
Voltmeter
Selecting BATT TEMP on the rotary
The voltmeter is on the left switch panel TEST switch tests the battery overheat
next to the ammeters. The voltmeter warning system by simulating a 160°F
(Figure 5C-6) indicates the voltage of battery temperature condition. Flashing
the Hot Battery bus when the battery of the BATT O’TEMP annunciator indi-
switch is in BATT or EMER and the cates system integrity and proper oper-
spring-loaded VOLTAGE SEL switch

EMER PWR RELAY

BATTERY
DISABLE
RELAY
BATT (X-GEN
O STARTS)
F
B
F A
EMER
T
T
E
R
BATTERY Y
RELAY B
U
S
5C-5A 5C-6 5C-7

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March 1999
 Electrical System

ation of the circuits. This test also drives power source to the Hot Battery bus.
the battery temperature gage to 160°F. Connecting a GPU to the aircraft clos-
es the relay and supplies power to the
Battery Temperature Gage Hot Battery bus. Placing the battery
In addition to the BATT O’TEMP switch in BATT closes the battery relay
annunciator, an optional temperature to supply power to the rest of the elec-
gage on the right meter panel provides trical system. Once a generator begins
a continuous indication of battery tem- supplying power to a Main bus, the
perature (Figure 5C-8). The range is 0 external power relay opens to prevent
to 180°F with a yellow band from 145 DC power flow from the electrical sys-
to 160°F and a red band from 160 to tem to the GPU.
180°F. The gage operates on 28V and The GPU should be disconnected
uses a second temperature sensor when not operating to prevent aircraft
between the battery cells for its reading. battery drain and damage due to
reverse current.
Ground Power
GPU Protection
Where available, a ground power unit
(GPU) can provide power operation of An overvoltage/overcurrent sensor pro-
systems, engine starting, and battery tects the starters from excessive cur-
charging. rent flow and the aircraft electrical sys-
tem from excessive voltage from a
A GPU connects to the aircraft through
GPU. The sensor consists of a detector
the receptacle on the left side of the
and a current sensor.
fuselage below the engine nacelle
(Figure 5C-9). The aircraft requires a The detector monitors voltage and cur-
GPU capable of producing 28V DC rent supplied by the GPU to the air-
with a capacity of 600 to 1,000 amps. craft. If voltage exceeds 32.5V for
The aircraft requires a GPU with a soft more than 200 milliseconds, the detec-
start capability to prevent overtorquing tor isolates the GPU from the aircraft’s
of the starter/generator drive shaft at electrical system. Before reapplying
the beginning of the start cycle. A 1,000 ground power, unplug the GPU to reset
amp GPU has sufficient capacity to the sensor.
provide 50% of the allowable starter A current sensor and the detector pro-
torque (1,500 in-lbs). vide overcurrent monitoring. If GPU
The external power relay in the aft J current is 1,200 ±100 amps for two
box connects the 28V DC ground seconds, the system disconnects GPU

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March 1999
 power to the aircraft. Unplug the GPU DC Distribution
to reset the system.
Buses
Junction Boxes
Buses provide connection and distrib-
Junction boxes in the tailcone and nose ution points for the DC and AC power.
avionics compartment contain compo- The battery switch controls operation of
nents that control the electrical system, the nine bus system through the Hot
thrust reversers, and supply informa- Battery bus that connects the DC sys-
tion from one system or component to tem with the battery. During emergency
another. conditions with the battery switch in
The power junction box in the tailcone EMER, only essential items receive
houses circuit breakers, relays, trans- power (Refer to specific numbers in
formers and current limiters (Figure Figure 5C-11 for this discussion).
5C-10). The power junction box is the Refer to DC Buses after the Data
central distribution for DC power from Summary for a listing of the systems
the generators, battery, and ground supplied by the LH/RH Main, LH/RH
power. Relays control the operation of Main Bus Extension, and LH/RH
the DC electrical system; they include: Crossover buses.
■ battery relay
Hot Battery Bus
■ left and right power relays
The Hot Battery bus (1) connects
■ left and right start relays
directly to the battery and receives
■ external power relay power when either the battery or
■ emergency power relay. ground power is powering the aircraft
or from the aircraft generators when
Table 5C-A provides a list of CBs or
the battery switch is in BATT. The Hot
items powered from the junction box;
Battery bus powers:
these CBs are inaccessible in flight.
■ voltmeter – battery switch in BATT
The Hot Battery bus, Battery bus, LH
or EMER
and RH Main buses are also in the
■ ignition for start
power junction box.

Left Side Right Side

Aft/Fwd Compartment Lights Ammeter – RH
Ammeter – LH Battery Charge (Secondary)
Battery Charge (Primary) Bus Sense – RH
Bus Sense – LH Cabin Lights
Chimes Flood Cooling
Fuel Boost Pump – RH (Primary) Flush Toilet
Generator Off Monitor – LH Flush Toilet Drain Heater
Indirect Light System Fuel Boost Pump – LH (Primary)
Landing Light – LH Generator Off Monitor – RH
Oxygen/Seat Belt Light Landing Light – RH
Power Relay – LH Power Relay – RH
5C-10 Recognition Light – LH Razor Outlet
Refreshment Bar Recognition Light – RH
Refreshment Center Start Light – RH
Start Light – LH Stereo
Tail Light Wing Inspection Light – RH
Tailcone Pressurization Valve Voltmeter (Secondary)
Voltmeter (Primary)
Table 5C-A; Power Junction Box Circuit Breakers

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 Electrical System

■ emergency lighting battery pack Emergency Bus

■ locator beacon (if installed)
The Emergency bus (4), behind the
■ exit lights copilot CB panel, receives power
■ nose and tail baggage compartment through the emergency power relay.
lights The Emergency bus supplies 28V DC
power to selected instruments and sys-
■ tailcone light
tems. With the battery switch in EMER,
■ battery relay and emergency relay the battery powers the bus to provide at
closing. least 30 minutes of power to selected
A ground power unit can be connect- instruments and systems. See Table
ed to provide DC power and charge 5C-B, following page, for a listing of
the battery. The Hot Battery bus in the these items.
J box connects to the Battery bus On units 0001 through 0031, turning
through the battery relay. With the bat- the battery switch OFF removes power
tery switch in EMER, the Hot Battery from the Emergency bus. Even with a
bus supplies power through the emer- generator on-line power cannot reach
gency power relay to the Emergency the bus through the emergency power
bus. The left and right start relays con- relay.
nect their respective starter/generators On units 0032 and subsequent, with
to the Hot Battery bus during engine the battery switch in BATT or OFF,
starts. With the battery switch in OFF, the emergency power relay remains
any of the above items will deplete the relaxed (open). The relay directs power
battery if left on overnight. from the operating generator through
the Battery bus to the Emergency bus.

Battery Connected – Battery Switch Off

TO CB 4 TO CB 4
PANELS 1 PANELS
EMER EMER
BATT BATT
EMERGENCY OR EMERGENCY OR
POWER OFF EMER POWER EMER OFF
3 RELAY 3 RELAY
LH MAIN RH MAIN

2 2
BATT BUS BATT BUS

POWER POWER
RELAY RELAY
BATTERY BATTERY
RELAY RELAY
START START
RELAY RELAY
1 1
GEN HOT BATT GEN HOT BATT

1 UNIT 001-031
GPU BATT
EXTERNAL
POWER BATTERY
RELAY DISCONNECT
RELAY

5C-11

Citation V For SimuFlite training only 5C-15

March 1999
 Aircraft equipped with the optional for individual CBs and receive power
Honeywell Primus II Integrated Radio from their respective Main bus.
System use the clearance delivery unit Normally, the pilot’s items, number one
(CDU) to operate COMM 1 and NAV2 inverter, LH instruments, and systems
whenever the battery switch is in (refer to DC Schematic for listing)
EMER and no generators are on-line. receive power from the LH Main bus
extension. Copilot items (dual EFIS),
Battery Bus number two inverter, RH instruments,
The Battery bus (2), which receives and avionics receive power from the
power from the battery or GPU via the RH Main bus extension (refer to DC
Hot Battery bus through the closed bat- Schematic for listing). The LH Main
tery relay, powers the LH and RH Main Bus Extension is behind the pilot’s CB
buses if both generators are off-line. In panel, and the RH Main Bus Extension
addition, if one generator goes off-line, is behind the copilot’s CB panel.
the Battery bus provides a cross-tie
between the LH and RH Main buses to Crossover Buses
allow an operating generator to power Two Crossover buses serve to group
both Main buses. With either genera- like items together on one CB panel.
tor on-line, the Battery bus provides an Essentially, system items are on the
electrical path for the generator to left CB panel and some receive power
charge the battery through the battery from the RH Crossover bus via the RH
relay and power the Emergency bus. Main Bus Extension. Some instru-
The Battery bus is in the tailcone J box. ments and avionics on the right CB
panel receive power from the LH
Main Buses Crossover Bus Extension via the LH
The LH and RH Main buses (3) receive Main Bus Extension.
power directly from their respective A 35 amp CB connects each Bus
starter/generator or through the Battery Extension to its opposite Crossover
bus. Bus feeds (cables) then distribute bus. The Right Crossover bus is behind
power to the cockpit CB panels to the the left CB panel and the Left
LH Main Bus Extension and the RH crossover bus is behind the right CB
Main Bus Extension. The LH and RH panel (refer to DC schematic).
Main buses are in the J box.
Main Bus Extensions DC Relays
Refer to specific numbers in Figure
The LH and RH Main Bus Extensions
5C-12 for the DC relays.
are the final distribution power points

Standard Single EFIS Optional Dual EFIS

Audio Panels (unit 0032 and Audio Panels (unit 0032 and
subsequent; prior aircraft with subsequent; prior aircraft with
SB560-23-05) SB560-23-05)
Cockpit Flood Lights Cockpit Flood Lights
COMM No. 1 COMM No. 1
Copilot’s Pitot/Static Heaters Copilot’s Pitot/Static Heaters
Copilot’s HSI Copilot’s RMI
Copilot’s Attitude Indicator Directional Gyro No. 1
Directional Gyro No. 2 LH & RH Fan Speed Indicators
LH & RH Fan Speed Indicator NAV No. 2
NAV No. 2 No. 2 VHF NAV Repeater
Voltmeter Voltmeter
Table 5C-B; Emergency Bus Powered Equipment

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March 1999
 Electrical System

Start Relay battery switch in EMER energizes the

relay to carry power from the Hot
To energize the left or right start relay Battery bus to the Emergency bus.
(1), press either engine start button. Placing the battery switch in OFF does
Once the start relay closes, power flows not isolate the Emergency bus if a gen-
to the engine starter; the starter begins erator is on-line. Power flows from the
turning. Power to close the solenoid Battery bus through the emergency
start relay comes from the aft J box only power relay to the Emergency bus.
when the battery switch is in BATT.
On units 0001 through 0031, the nor-
A white light in each starter button indi- mally open relay energizes closed to
cates closing of the start relay contacts. connect the Hot Battery bus with the
The speed sensor in the GCU removes Emergency bus when the battery switch
power from the start relay once the is in BATT or EMER. Placing the bat-
engine reaches approximately 38% N2. tery switch in OFF isolates the
Emergency Power Relay Emergency bus from all power sources.

On units 0032 and subsequent, the Battery Relay

two-way emergency power relay (2) Positioning the battery switch to BATT
determines which source (battery or closes the battery relay (3) to power
generators) provides electrical power the entire DC system. Positioning the
to the Emergency bus. With the battery battery switch to EMER opens the bat-
switch in BATT or OFF, the emergency tery relay and closes the emergency
power relay is unpowered to a position power relay, thus disconnecting the
that supplies power from the Battery LH Main DC bus, RH Main DC bus,
bus to the Emergency bus. Placing the

Battery Switch BATT

TO CB 4 TO CB 4
PANELS 1 PANELS
EMER EMER
BATT BATT
EMERGENCY OR EMERGENCY OR
POWER OFF EMER POWER EMER OFF
3 RELAY 3 RELAY
LH MAIN RH MAIN

2 2
BATT BUS BATT BUS

POWER POWER
RELAY RELAY
BATTERY BATTERY
RELAY RELAY
START START
RELAY RELAY
1 1
GEN HOT BATT GEN HOT BATT

1 UNIT 001-031
GPU BATT
EXTERNAL
POWER BATTERY
RELAY DISCONNECT
RELAY

5C-12

Citation V For SimuFlite training only 5C-17

March 1999
 and the Battery bus from the Hot age/overcurrent sensor detects exces-
Battery bus and connecting the Hot sive GPU voltage or amperage.
Battery bus to the Emergency bus. This Moving the battery switch to BATT
also prevents the generators from connects GPU power from the Hot
charging the battery. Battery bus to the rest of the electrical
Battery Disable Relay system.

The battery disable relay (not shown) DC Protection

protects the current limiters during a
ground cross generator start by open- Three major items provide protection
ing the battery relay. This blocks the for the electrical system:
power from taking an alternative route ■ generator control units (GCUs)

from the Main Bus Extension across ■ circuit breakers (CBs)

the current limiter to the Battery bus
■ current limiters.
and then to the Hot Battery bus and
the starter. With the battery relay open, Generator Control Units
power from one starter/generator goes
directly through the Hot Battery bus A GCU for each starter/generator pro-
to the second starter/generator. A squat vides voltage regulation, overvoltage
switch on the left main gear prevents protection, undervoltage protection,
cross starts in flight. reverse current sensing, and starter
control.
Power Relay In the generator function each GCU
The generator control unit opens the provides the following:
power relay (4) to remove a failed gen- ■ As a voltage regulator, the GCU
erator from the system and isolate the monitors generator output so it can
generator from its associated Main bus. regulate the generator to meet the
Placing the generator switch in the OFF demands of the system.
position also opens the power relay. The
■ For generator paralleling, the GCU
respective amber GEN OFF annuncia-
reads the differences between the
tor illuminates once the relay opens.
generator voltage and the equalizer
Battery Disconnect Relay bus and feeds the change to the reg-
ulator to shift the output to match.
The battery grounds to the airframe In addition, an equalizing relay
through the battery disconnect relay allows the power relay to isolate
(5). This relay opens to disconnect the completely a tripped generator.
battery from the electrical system when
■ Once a generator initially excites,
a GPU is supplying power during an
engine start. the GCU begins sensing for a
ground or feeder fault. If a feeder
External Power Relay fault occurs, the field relay trips to
shut the generator down.
The external power relay (6) is between
■ If the GCU voltage regulator fails,
the external power receptacle and the
Hot Battery bus. Connecting a GPU to the field relay trips to protect the gen-
the aircraft closes the relay; the GPU erator from an overvoltage condition.
supplies power to the Hot Battery bus. If the GCU senses an overvoltage or
Once a generator begins supplying feeder fault, it takes the generator off-
power, the relay opens to prevent the line by opening the field relay. The
flow of current from the electrical sys- GCU senses the resulting undervolt-
tem to the GPU. The external power age and opens the power relay to iso-
relay also opens when the overvolt- late the generator from its Main bus.

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March 1999
 Electrical System

Pressing the ENGINE FIRE switch Complete diagrams of the cockpit CB

also opens the field relay. An open panels are found at the beginning of
field relay may be reset by placing the the DC Systems section.
generator control switch in the momen-
tary RESET position. Current Limiters
The current limiters in the tailcone J
Circuit Breakers box provide the same function as CBs;
Circuit breakers (CBs) provide over- however, they break power to a bus
load protection for specific aircraft sys- rather than to a specific component.
tems and components. Each CB has a The primary current limiters have a
rating that varies with the maximum 225 amp rating and are between the
load of the protected system or com- Battery bus and the Main buses to pre-
ponent. An amperage rating, clearly vent excessive current flow back to the
marked on each CB, indicates the Battery and Hot Battery buses or a
amperage that the CB opens to protect ground fault on one of the buses.
that circuit or component. The opening Secondary current limiters, rated to
of a CB is automatic. open at 80 amps, connect the Main
Panels on both sides of the cockpit buses to each of the three parallel feed-
contain all the electrical system CBs ers to the Main Bus Extensions. Current
accessible in flight. They may be man- limiters actuate automatically; but they
ually reset or pulled by the flight crew cannot be reset. Once a current limiter
because each is a push-to-reset type. opens, it must be replaced.
Those in the tailcone J box are inac-
cessible in flight.

Citation V For SimuFlite training only 5C-19

March 1999
 DC Operation nal power. The external power relay
opens when a generator comes on-line.
After engine start with the generators Generator power always preempts
on-line, the electrical system requires external power.
little more than monitoring unless an
After positioning the switches correct-
abnormal or emergency condition
ly, either engine may be started first.
develops. There are, however, three
For example, press the RH ENGINE
different methods to start the engine:
START button momentarily. This action
■ GPU start simultaneously activates the following:
■ battery start ■ closes the RH start relay – start but-

■ cross generator assist start. ton illuminates indicating start relay

is closed
GPU Start ■ opens the battery disconnect relay

Ground power connects to the aircraft ■ activates RH fuel boost pump – RH

through the receptacle on the left side BOOST PUMP light illuminates;
of the fuselage. Switch positions are RH FUEL PRESS LOW light extin-
the following: guishes
■ both generator switches – OFF ■ arms the right engine ignition sys-
■ ignition switches – NORM tem – ignition light illuminates when
■ battery switch – BATT
the throttle is moved to IDLE detent
■ illuminates the engine instrument
■ boost pump switches – NORM or
OFF. flood light.
Leave the generator switches OFF to Power flows from the GPU to the Hot
ensure that both engines start on exter- Battery bus and then to the right starter/

GPU Start of Right Engine

TO CB 4 TO CB 4
PANELS 1 PANELS
EMER EMER
BATT BATT
EMERGENCY OR EMERGENCY OR
POWER OFF EMER POWER EMER OFF
3 RELAY 3 RELAY
LH MAIN RH MAIN

2 2
BATT BUS BATT BUS

POWER POWER
RELAY RELAY
BATTERY BATTERY
RELAY RELAY
START
RELAY START
1 RELAY 1
GEN HOT BATT GEN HOT BATT

1 UNIT 001-031
GPU BATT
BATTERY
EXTERNAL DISCONNECT
POWER RELAY
RELAY
5C-13

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March 1999
 Electrical System

generator (Figure 5C-13). The GCU flows back to the Hot Battery bus
automatically terminates the start through the Battery bus to charge the CAUTION: Turbine speed
sequence at approximately 38% N2 by battery. Power is also routed through greater than 53% on the operating
disengaging the start relay. the emergency power relay to the engine will produce a generator
Extinguishing of the RH FUEL Emergency bus (Figure 5C-14). output that may damage the gen-
BOOST ON annunciator, ignition light, Battery Start erator drive during the second
engine instrument light, and the RH engine start.
start button light confirm termination A battery start, which is similar to a GPU
of the start sequence. The battery dis- start, begins by checking the voltmeter
connect relay closes to reconnect the for a minimum battery voltage of 24V.
battery to its ground. The RH GEN OFF Switch positions are the following:
annunciator remains illuminated until ■ both generator switches – ON
the generator switch is placed in GEN.
■ ignition switches – NORM
Start the second engine using the same
■ battery switch – BATT
procedure after the first engine stabi-
lizes at idle. Once the second engine ■ boost pump switches – NORM/OFF.

is running and the start cycle termi- With both generator switches ON, the
nates, disconnect the GPU and place GCU can close the power relay and
both generator switches in GEN. Both bring the generator on-line once gen-
GEN OFF annunciators should extin- erator voltage is equal to or greater
guish and both ammeters should show than Main bus voltage. Either engine
load sharing by the generators. may be started first. In extreme con-
With the generator switches ON, the ditions, starting the left engine first
power relays close and power flows reduces the load on the battery slight-
from each generator to the Main and ly because the battery cable is shorter
Extension buses. Generator power also to the left starter/generator.

Normal Operation – Engine Running - Generator On

TO CB 4 TO CB 4
PANELS 1 PANELS
EMER EMER
BATT BATT
EMERGENCY OR EMERGENCY OR
POWER OFF EMER POWER EMER OFF
3 RELAY 3 RELAY
LH MAIN RH MAIN

2 2
BATT BUS BATT BUS

POWER POWER
RELAY RELAY
BATTERY BATTERY
RELAY RELAY
START START
RELAY RELAY
1 1
GEN HOT BATT GEN HOT BATT

1 UNIT 001-031
GPU BATT
EXTERNAL
POWER BATTERY
RELAY DISCONNECT
RELAY
5C-14

Citation V For SimuFlite training only 5C-21

March 1999
 Press the RH ENGINE START button light, engine instrument flood light,
momentarily. This action simultan- and the RH start button light confirm
eously activates the following: start sequence termination.
■ closes the RH start relay – start but- The RH GEN OFF annunciator extin-
ton illuminates indicating start relay guishes when the generator voltage
is closed exceeds the battery voltage.
■ activates the RH fuel boost pump –
Cross Generator
RH FUEL BOOST ON annuncia-
tor illuminates and RH FUEL LOW
Assisted Start
PRESS annunciator extinguishes A cross generator start uses the output
■ arms the right engine ignition sys- of the operating generator to assist the
tem – ignition light illuminates when battery in starting the second engine.
the throttle is moved to IDLE detent The generator assisted start circuit is
disabled in flight by the left main land-
■ illuminates the engine instrument
ing gear squat switch. Set all switches
flood light. in the same positions as a normal battery
Power flows from the battery to the start. However, the start sequence clos-
Hot Battery bus and then to the right es both start relays to allow the operat-
starter/generator (Figure 5C-15). The ing generator to assist in the engine start.
GCU automatically terminates the start Before starting the second engine, place
sequence at approximately 38% N2 by the GND IDLE switch in the HIGH
disengaging the start relay. position. This increases the idle RPM
Extinguishing of the RH FUEL on the operating engine to 52% N2 to
BOOST ON annunciator, ignition ensure that proper torque and genera-

Battery Start of Right Engine

TO CB 4 TO CB 4
PANELS 1 PANELS
EMER EMER
BATT BATT
EMERGENCY OR EMERGENCY OR
POWER OFF EMER POWER EMER OFF
3 RELAY 3 RELAY
LH MAIN RH MAIN

2 2
BATT BUS BATT BUS

POWER POWER
RELAY RELAY
BATTERY BATTERY
RELAY RELAY
START
RELAY START
1 RELAY 1
GEN HOT BATT GEN HOT BATT

1 UNIT 001-031
GPU BATT
EXTERNAL
POWER BATTERY
RELAY DISCONNECT
RELAY
5C-15

5C-22 For SimuFlite training only Citation V

March 1999
 Electrical System

tor output is available for a generator On the ground, the generator assisted
assisted start. In addition, this is a func- start circuit provides power to open the
tional test of the ground idle system. battery disable relay. The open relay
Press the LH ENGINE START button removes electrical power from the bat-
momentarily. This action (Figure tery relay causing it to open. This pre-
5C-16) simultaneously activates the vents the load of the operating genera-
following: tor from using the Main bus as an
alternate route of power for the cross
■ closes both the RH and LH start
generator start. The high starting amper-
relays – both start buttons illumi- age flowing from the Main bus through
nate indicating that both start relays the Battery bus, then back to the Hot
are closed Battery bus would damage the 225 amp
■ battery disable relay opens and current limiter. At start termination, the
removes ground from the battery battery disable relay is de-energized to
relay that opens close and supplies power to close the
■ activates left fuel boost pump – LH battery relay (see the DC Electrical
FUEL BOOST ON annunciator Schematic, front of this chapter).
illuminates and LH FUEL LOW Power flows from the operating gen-
PRESS annunciator extinguishes erator to the Hot Battery bus and then
■ arms the left engine ignition system to the left starter/generator. The GCU
– ignition light illuminates when the automatically terminates the start
throttle is moved to IDLE detent sequence at approximately 38% N2 by
disengaging both start relays.
■ illuminates the engine instrument
flood light. Extinguishing of the LH FUEL
BOOST ON annunciator, ignition light,

Cross Generator Start of Left Engine

TO CB 4 TO CB 4
PANELS 1 PANELS
EMER EMER
BATT BATT
EMERGENCY OR EMERGENCY OR
POWER OFF EMER POWER EMER OFF
3 RELAY 3 RELAY
LH MAIN RH MAIN

2 2
BATT BUS BATT BUS

POWER POWER
RELAY RELAY
BATTERY BATTERY
RELAY RELAY

START START
RELAY 1 RELAY 1
GEN HOT BATT GEN HOT BATT

1 UNIT 001-031
GPU BATT
EXTERNAL
POWER BATTERY
RELAY DISCONNECT
RELAY
5C-16

Citation V For SimuFlite training only 5C-23

March 1999
 engine instrument flood light, and the DISENGAGE switch between the start
RH and LH start button lights confirm buttons. The following occurs:
termination of the start sequence. Once ■ start relay opens – light in start but-
the GCU senses that both generators ton extinguishes
are operating in parallel, the second
■ boost pump turns off – FUEL
generator comes on-line when its power
relay closes. The LH GEN OFF annun- BOOST ON annunciator extin-
ciator then extinguishes. guishes
■ ignition system de-arms or turns off
Terminating Abnormal – ignition light extinguishes if the
Starts throttle has been moved to idle detent
The automatic start sequence may be ■ engine instrument flood light extin-

terminated by pushing the STARTER guishes.

5C-24 For SimuFlite training only Citation V

March 1999
 Electrical System

LH Main Bus (Aft J Box)

Aft/Fwd Compartment Lights Power Relay – LH DC Buses
Ammeter – LH Recognition Light – LH
Battery Charge (Primary) Refreshment Bar
Bus Sense – LH Refreshment Center
Chimes Start Light – LH
Fuel Boost Pump – RH (Primary) Tail Light
Generator Off Monitor – LH Tailcone Pressurization Valve
Indirect Light System Voltmeter (Primary)
Landing Light – LH
Oxygen/Seat Belt Lights

RH Main Bus (Aft J Box)

Ammeter – RH Landing Light – RH
Battery Charge (Secondary) Power Relay – RH
Bus Sense – RH Razor Outlet
Cabin Lights Recognition Light – RH
Flood Cooling Start Light – RH
Flush Toilet Stereo
Flush Toilet Drain Heater Wing Inspection Light – RH
Fuel Boost Pump – LH (Primary) Voltmeter (Secondary)
Generator Off Monitor – RH

LH Main Bus Extension

AFIS VHF Transmitter ITT – LH
Anti-Collision Light Landing Gear Control
AOA System Computer/Indicator Landing Gear Warning
AOA System Heater LH Bus No. 1, 2, 3
Battery Overtemp System Outside Air Temperature
Battery Temperature Gage Oil Pressure – LH
Cabin Fan Oil Temperature – LH
Cabin Pressure Normal Panel Lights – 5V – LH
Cabin Temperature Indicator Pitch Trim
Cockpit Voice Recorder Pitot/Static Heat – LH
Defog Fan Rotating Beacon
Digital Clock – LH Skid Control
EL Inverter Speed Brake
Engine/Wing Anti-Ice Valve – LH Standby Gyro
Engine Synchronization Start Power (Primary)
Flight Data Recorder Thrust Reverser Power and
Fire Detection – RH Control – LH
Firewall Shutoff – RH Tail Navigation Light
Flap Control Valve Temperature Control
Fuel Flow – LH Turbine Speed – LH
Fuel Boost Pump – RH W/S Bleed Air Control
(Secondary) W/S Bleed Air Power
Fuel Quantity – LH Warning Lights 1
Igniter – RH (Primary) Wing Inspection Light – LH
Inverter 1 Wing Navigation Light

Citation V For SimuFlite training only 5C-25

March 1999
 RH Crossover Bus
Alcohol Pump Fuel Boost Pump – LH
Altimeter Vibrator – RH (Secondary)
Cabin Pressure Emergency Igniter – LH (Primary)
Digital Clock – RH Overspeed Warning
Divider Fans Panel Lights – 5V – Center
Engine/Wing Anti-Ice Valve – RH Panel Lights – 5V – RH
Equipment Cooling Start Light (RH)
Fire Detection – LH Surface Deice Boots
Firewall Shutoff – LH Thrust Reverser Power and
Flight Hour Meter Control – RH
Warning Light 2

RH Main Bus Extension

ADF 2 Inverter 2
Audio 2 (0001-0031 w/o ITT – RH
SB560-23-05) MFD Controller
COMM 2 MFD Display
DME 2 MFD Symbol Generator
EADI 2 (dual EFIS) Oil Pressure – RH
EFIS 1 Oil Temperature – RH
EFIS 2 (dual EFIS) Radar
EFIS 2 Controller (dual EFIS) RH Bus No. 1, 2, 3
EHSI 2 (dual EFIS) RH Instrument Vibrator
Flight Director No. 2 (dual FD) RMI 2 (single EFIS)
Flitephone True Airspeed System Heater
Fuel Remaining Transponder 2
Fuel Flow – RH True Airspeed System
Fuel Quantity – RH Turbine Speed – RH

LH Crossover Bus
ADF 1 EHSI 1
AFIS VHF Receiver Flight Director 1
Audio 1 (0001-0031 w/o FMS (GNS-X)
SB560-23-05) HF COMM 3
Autopilot Warning NAV 1
Autopilot Servo Radio Altimeter
Autopilot Control RMI 1
Directional Gyro 1 Transponder 1
DME 1 VLE
EADI 1
EFIS 1 Display Controller

5C-26 For SimuFlite training only Citation V

March 1999
 Electrical System

AC Electrical
System
AC Circuit
Breakers

Citation V For SimuFlite training only 5C-27 5C-28

March 1999
AC Electrical System AC Circuit Breakers

ADC EFIS 2 (OPT)

AUTOPILOT EFIS 1 HSI 2 FD 2 (OPT)
FD 1 NAV 1 NAV 2 VG 2
RADAR RMI 1 RMI 2 COMM NAV XPDR DME ADF AUDIO RH FAN RH RH TURB RH FUEL RH FUEL RH OIL RH OIL
WARN
VG 1 1 1 1 1 1 1 SPEED ITT SPEED FLOW QTY TEMP PRESS
MASTER
7.5 3 3 3 3 3 3 2 2 2 2 2 2 2
WARNING
115V AC #1 26V AC #1 RESET 26V AC #2 115V AC #2 ENGINE INSTRUMENTS

COMM NAV XPDR DME ADF AUDIO FD EFIS EFIS 1 EHSI EADI DG RMI MFD MFD
2 2 2 2 2 2 1 1 CONT 1 1 1 1 SYM GEN DISP

AC 7.5 3 3 3 3 3 3 5 1 5 5 5 2 7.5 5
FAIL

COMM FD EFIS EFIS 2 EHSI EADI DG RMI PHONE RADAR

FMS VLF AFIS AP TAS
3 2 2 CONT 2 2 2 2

5 5 5 7.5 7.5 2 3 5 1 5 5 5 2 5 7.5

AVIONICS DC

NAV RMI EFIS VG FD AIR RAD RH AC

AP START
1 1 1 1 1 DATA ALT INVERTER
1 2 2 1 1 1 2 2 7.5 15
RH BUS
NO 2
NO 1 NO 2 NO 3
HSI NAV RMI EFIS VG FD RADAR ADI FLOOD LH CB
2 2 2 2 2 2 2
75 75 75 PANEL
AC FAIL AC FAIL
(ALLOW (HOLD) 2 1 2 2 1 1 1 2 5 35
RESET)
AC FLT INSTR AVIONICS DC PWR-RH BUS

26V AC BUS NO. 1 & 2

115 VAC 26VAC

115V AC BUS NO. 1 & 2

115V, 400Hz INVERTER 115V, 400Hz BUS 1 BUS 2 BUS 1 BUS 2
FAIL 5 5 10 10
INVERTER 26V, 400Hz AC 26V, 400Hz AC INVERTER
1 1 2 2 AC BUS

AC INV AC INV
NO.1 NO.2

T
28V DC INV 1
E
AC 28V DC
LH MAIN BUS EXT S RH MAIN BUS EXT
INV 2 T OFF

5C-28 For SimuFlite training only Citation V

March 1999
 Electrical System

The Citation V AC electrical system

uses inverters to convert direct current
INVERTER NO. 1 CB and the No. 2
inverter receives power from the RH AC System
(DC) into alternating current (AC) for Main bus Extension through the 15
avionics and navigation equipment. A amp AC INVERTER NO. 2 CB. Each
split-bus distribution system supplies inverter operates independently from
inverter power through CBs to the var- the other. The two inverters intercon-
ious systems requiring AC power. nect to ensure phase synchronization
The two inverters operate simultane- between themselves.
ously and independently to supply their Inverter Control
separate distribution systems. If an
inverter fails, the operating inverter Three AVIONICS POWER switches
can meet the aircraft’s AC electrical (Figure 5C-18) on the left switch panel
demands; however, both inverters must control the inverters and avionics
operate for takeoff. equipment. The left AC switch (INV
1/TEST/INV 2) is a test switch, the
Components middle switch (AC/OFF) controls DC
power to both inverters, and the right
The AC system consists of: switch (ON/OFF) is an avionics mas-
■ two 375 volt-amp (VA) inverters ter switch.
■ two 115V AC buses The AC test switch spring-loads to the
■ two 26V AC buses
center (OFF) position. Placing the
switch in either INV 1 or INV 2 turns
■ system controls and indicators
off the respective inverter by remov-
■ switching relays ing DC power. With the switch in either
■ protection devices.
position, the operating inverter assumes
the loads for both sides of the AC elec-
Inverters trical system.
The inverters (Figure 5C-17) convert The AC/OFF switch controls the DC
28V DC into 115V and 26V, 400 Hz power source to the inverters. With the
single-phase AC power. Each unit is switch in AC, Main bus DC power
a transistorized static inverter (no mov- flows to both inverters. The inverters
ing parts) that receives DC power from then power their respective buses. In
separate sources. The No. 1 inverter OFF, the switch removes power to the
receives power from the LH Main bus DC grounding circuit that connects
Extension through the 15 amp AC each inverter to its Main DC bus and
removes power to the AC buses.

5C-17 5C-18

Citation V For SimuFlite training only 5C-29

March 1999
 The ON/OFF switch is an avionics mas- switch position turns off the operating
ter switch. In ON, the switch removes inverter and causes a complete loss of
power from the normally closed avion- AC power while the switch is in TEST.
ics bus relays. These relays connect the
avionics buses to their respective DC Failure Annunciators
buses. The switch provides DC power The AC inverter system has two
to the avionics bus relays in OFF, there- INVERTER FAIL annunciators and one
by disconnecting the avionics from their AC FAIL annunciator. The fault mon-
DC power sources. With the battery itor continuously monitors inverter out-
switch in EMER and neither generator put and frequency.
connected to the Main DC bus, the If an inverter trips off-line due to DC
emergency avionics bus is powered power loss, overvoltage, or undervolt-
when the DC master switch is in either age, the respective amber 1/2 INVERT-
ON or OFF. The avionics master switch ER FAIL annunciator illuminates. The
does not have to be ON for the invert- operating inverter automatically
ers to operate. assumes the load of the failed inverter.
Switching Relays Illumination of either or both INVERT-
ER FAIL annunciators also illuminates
Each side of the AC electrical system the red AC FAIL annunciator. The AC
has a pair of relays that automatically FAIL annunciator, in turn, triggers the
connect an operating inverter to a failed master warning system that illuminates
inverter’s AC bus(es). Normally, the the MASTER WARNING lights.
26V and 115V AC inverter outputs pro- Pressing the MASTER WARNING
vide power to the relay solenoids. The switch extinguishes the AC FAIL
solenoids maintain the connection annunciator. The failed inverter’s
between the inverter and its 26V and INVERTER FAIL annunciator remains
115V buses. If an inverter AC output illuminated to indicate the malfunc-
fails, the solenoid opens to a second tioning unit. With both inverters failed,
set of contacts that connects the oppo- the AC FAIL annunciator remains illu-
site inverter to the failed inverter’s bus. minated after pushing the MASTER
If the operating inverter fails to assume WARNING RESET.
the loads, the automatic inverter Power loss, overvoltage, and under-
switching system is defective. If one voltage on the 115V AC buses trigger
inverter failure annunciator illuminates the AC FAIL annunciator. If an invert-
in flight, selecting the opposite test er fails or malfunctions, there is usually

26V AC Bus No. 1 26V AC Bus No. 2

EFIS 1 EFIS 2*
NAV 1 HSI 2
RMI 1 NAV 2
RMI 2
115V AC Bus No. 1 115V AC Bus No. 2
Air Data Computer Flight Director 2*
Autopilot Vertical Gyro 2
Flight Director 1
Radar
Vertical Gyro 1

Table 5C-C; AC Buses

* Dual EFIS/Flight Directors

5C-30 For SimuFlite training only Citation V

March 1999
 Electrical System

a power drop or surge on the 115V AC Refer to Table 5C-C for a list of sys-
bus before the inverter drops off-line tems powered by the inverters.
and the operating inverter assumes the
load. The AC FAIL annunciator illu- AC Protection
minates to indicate this power loss.
Various circuit breakers (CBs) and a
Illumination of the AC FAIL annun- fault monitoring system protect the
ciator also triggers the master warn- inverters, the AC buses, and the DC
ing system. Pressing the MASTER buses supplying inverter power from
WARNING switch extinguishes the electrical faults.
AC FAIL annunciator unless there is
total AC failure. Circuit Breakers
Each side of the split-bus AC system
AC Buses employs three CBs: one between each
Each inverter connects to a 26V and a inverter output and the respective AC
115V AC bus through a magnetic cir- bus, and one between the DC bus and
cuit breaker (CB). With both inverters the inverter DC input. Separate CBs
operating normally, the No. 1 inverter connect and protect individual pieces
supplies power to 115V AC bus No. of avionics and navigation equipment
1 through the 5 amp 115 VAC BUS 1 to the AC buses. The right and left CB
CB and the 26V AC bus No. 1 through panels have the 15 amp inverter DC
the 10 amp 26 VAC BUS 1 CB. The supply CBs, and the right CB panel
No. 2 inverter supplies power to the contains the individual CBs for the
115V AC bus No. 2 through the 5 amp avionics equipment. The right CB sub-
115 VAC BUS 2 CB and the 26V AC panel contains the left and right 10 amp
bus No.2 through the 10 amp 26 VAC 26V and 5 amp 115V AC bus CBs.
BUS 2 CB. The four magnetic AC bus CBs are
If an inverter fails, the automatic between the inverter outputs and the
switching system automatically con- AC buses. If a bus short occurs, the CB
nects the operating inverter to the failed opens to remove the bus from the
inverter’s buses. respective inverter output. When the
Circuit breakers between the invert- CB opens, a secondary circuit illumi-
er’s 26V and 115V AC outputs pro- nates the AC FAIL annunciator; the
tect the inverters from bus faults. With AC FAIL annunciator triggers the reset-
a circuit breaker pulled, the protected table MASTER WARNING lights.
bus does not receive power. The four Before resetting one of these CBs, con-
circuit breakers are on the right sub sider the possibility of a short existing
circuit breaker panel. on the respective AC bus.

Citation V For SimuFlite training only 5C-31

March 1999
5C-32 For SimuFlite training only Citation V
March 1999
 Electrical System

Preflight Please refer to your SimuFlite Citation

Preflight and
V Operating Handbook for complete
During the cockpit preflight check,
verify battery voltage on the voltmeter;
the minimum acceptable battery read-
procedure.
Single Generator Failure
Procedures
ing is 24V. Check emergency bus
items. During the Before Start Check, With a LH or RH GEN OFF annunci-
test the battery temperature warning ator illuminated, the initial action is to
system by selecting BATT TEMP with reduce, if required, the electrical load
the rotary TEST switch. A 160°F indi- on the operating generator. One gen-
cation on the battery temperature gage, erator is usually capable of supplying
flashing of the BATT O’TEMP annun- all electrical requirements during flight.
ciator, and illumination of the MAS- If possible after reducing the load,
TER WARNING light indicate proper direct attention toward the failed gen-
system operation. erator while checking and resetting
switches and CBs (Figure 5C-18A).
Check the operation of all the system
annunciator lights prior to flight. Test After reducing the load, check the gen-
the annunciators by selecting ANNU erator switches to ensure that they are
on the rotary TEST switch; all of the in GEN; check for any disengaged
annunciators should illuminate. CBs. Select the voltmeter to read the
voltage output of both generators
Before testing the inverters, the air- before attempting a generator reset. If
craft requires a source of DC power the voltage output on the failed gen-
from either a GPU or the generators erator is less than 10V, the field relay
to keep from running down the bat- is open and might be resettable; if its
tery. Test the operation of the invert- voltage output is 10V or more, but less
ers and the automatic switching sys- than normal (28.5V), the field relay is
tem by placing the AC power switch in closed and the generator may not reset.
AC. Place the inverter test switch in The voltage output of the operating
INV 1 and observe the illumination of generator should be normal (28.5V).
the INVERTER FAIL 1 annunciator.
After resetting the failed generator,
The No. 2 inverter should still power check that its voltage output is normal
all AC buses. Select INV 2 with the test (28.5V) before moving its switch to
switch. The No. 1 inverter should GEN. If the generator does not reset,
assume the AC loads and the INVERT- turn it off and continue the flight with
ER FAIL 2 light should illuminate. The the one operating generator.
switch should spring-load to the center
OFF position. During the test, the AC Single Inverter Failure
FAIL and MASTER WARNING lights
illuminate when each inverter fails. The Illumination of the MASTER WARN-
AC FAIL annunciator can be reset by ING lights and the 1 or 2 INVERTER
pushing the MASTER WARNING FAIL and AC FAIL annunciators indi-
switchlight. cate an inverter failure. Check the AC
Inverter No. 1 and No. 2 CBs and ver-
Abnormal ify that they are set. Reset the MAS-
TER WARNING lights; the AC FAIL
Procedures light extinguishes. The INVERTER
Abnormal procedures include: FAIL annunciator continues to illumi-
nate to identify the failed inverter.
■ single generator failure
Continue the flight with the operating
■ single inverter failure. inverter powering the AC electrical

Citation V For SimuFlite training only 5C-33

March 1999
 system. The flight director usually dis- there is an AC power distribution fail-
engages with an inverter failure and ure. Check the right sub-circuit break-
may be re-engaged. er panel for any open CBs.
If any of the CBs are open, do not reset
Emergency them unless they were unintentionally
Procedures tripped by the crew. Resetting an open
AC circuit breaker may cause addi-
Electrical system emergency proce- tional system damage. Depending on
dures include: the bus lost, different avionics systems
■ double inverter failure are inoperative. Refer to the AC
■ AC system failure and AC power Schematic for a listing of systems pow-
distribution failure ered by the AC buses.
■ battery overheat Battery Overheat
■ electrical fire or smoke
Illumination of the BATT O’TEMP
■ loss of both generators. annunciator and an excessive temper-
Please refer to your SimuFlite Citation ature indication on the battery temper-
V Operating Handbook for complete ature gage indicates a battery overtem-
procedure. perature condition. Positioning the
battery switch to EMER should open
Dual Inverter Failure the battery relay to isolate the generators
from the battery and terminate charg-
Illumination of the 1 and 2 INVERT- ing. Check for an amperage drop to
ER FAIL annunciators, the AC FAIL ensure that charging has stopped. With
annunciator and the MASTER WARN- the battery switch in EMER, the No. 2
ING lights indicate a complete AC radio magnetic indicator (RMI) nee-
power failure. dle(s) may fluctuate if the No. 2 navi-
Attempt to restore at least one inverter gation receiver is providing navigation
by resetting and/or pulling the inverter information. If the battery voltage is
CBs. Land as soon as practical. If one volt less than generator voltage in
unable to reset, place the battery switch 30 seconds to two minutes, monitor the
in EMER to provide a source of AC BATT O’TEMP annunciator and bat-
power from the copilot’s C-14D direc- tery temperature gage for a change.
tional gyro inverter to the copilot’s hor- If the BATT O’TEMP annunciator
izontal situation indicator and NAV 2 in extinguishes and battery temperature
aircraft equipped with a single elec- decreases, land as soon as possible.
tronic flight instrument system (EFIS).
The pilot’s C-14D directional gyro If there is no decrease in battery amper-
inverter provides AC power to the copi- age, turn the floodlights to full intensi-
lot’s RMI and NAV 2 radio in aircraft ty (at night) and turn both generator
equipped with the optional dual EFIS. switches OFF. If the battery relay is not
The pilot’s EFIS and copilot’s EFIS (if stuck with the generators off, the BATT
installed) will be inoperative. Land as O’TEMP annunciator extinguishes
soon as practical (within 30 minutes). immediately along with a loss of all
main DC power. With the switch in
AC Fail and AC Power EMER and the generators off, the emer-
Distribution Failure gency power relay closes to supply bat-
tery power from the Hot Battery bus to
If the AC FAIL annunciator does not the Emergency bus. A fully charged
extinguish with both INVERTER battery can power the Emergency bus
FAIL annunciators extinguished and for at least 30 minutes (see Table 5C-D).
the MASTER WARNING light reset,

5C-34 For SimuFlite training only Citation V

March 1999
 Electrical System

Aircraft with the optional Honeywell Battery bus; the other buses still receive
Primus II radio system use the clear- power. Land as soon as practical. After
ance delivery unit (CDU) to operate landing, check the battery to determine
the COMM 1 and NAV 2 radios with if there is cell damage.
the battery switch in EMER and the If there is no DC power loss, the bat-
generators OFF. tery relay is stuck closed. Turn both
Should a landing be necessary with the of the manual windshield bleed air
battery switch in EMER and both gen- valves OFF. Pull the DC power LH
erators OFF, the following systems are and RH BUS NO. 1, 2, 3 CBs. Land as
inoperative: soon as practical. Once the landing is
■ normal landing gear extension and assured, reset the CBs to obtain power
indication; extend landing gear with for the landing gear and flap systems.
the blow down system Extend the landing gear and flaps;
maintain airspeed at VREF.
■ flaps; perform a flaps inoperative
landing After landing, the battery requires
maintenance to determine if there is
■ power brakes/antiskid; emergency
cell damage.
braking system only
■ engine anti-ice valves; anti-ice Electrical Fire or Smoke
valves in the open position After donning oxygen masks, select
■ outside air temperature gage 100% oxygen, verify that the INPH but-
■ all engine instruments except N1 ton is out (optional Primus II radios),
tachometers. and place the microphone switch in MIC
OXY MASK. Ensure that the PRESS
If the battery relay is not stuck and nor- SOURCE SELECTOR is in NORM
mal DC power is lost, place both gen- and the cockpit divider door is open.
erator switches in GEN and turn the
battery switch OFF. With the battery With a known source, pull the appro-
switches OFF: priate circuit breaker(s) to isolate the
faulty circuit.
On units 0001 through 0031, the
Emergency bus is not powered. With an unknown source, turn the
cockpit floodlights to full bright (at
On units 0032 and subsequent, the night), place the battery switch in
Emergency bus continues to receive EMER, and turn both generators OFF.
power from the generators via the

Standard Single EFIS Optional Dual EFIS

Audio Panels (unit 0032 and Audio Panels (unit 0032 and
subsequent; prior aircraft with subsequent; prior aircraft with
SB560-23-05) SB560-23-05)
Cockpit Flood LIghts Cockpit Flood Lights
COMM No. 1 COMM No. 1
Copilot’s Pitot/Static Heaters Copilot’s Pitot/Static Heaters
Copilot’s Attitude Indicator Copilot’s RMI
Directional Gyro No. 2 Directional Gyro No. 1
LH & RH Fan Speed Indicator LH & RH Fan Speed Indicators
NAV No. 2 Nav No. 2 VHF NAV Repeater
Voltmeter Voltmeter

Table 5C-D; Emergency Bus Powered Equipment

Citation V For SimuFlite training only 5C-35

March 1999
 With the battery switch in EMER and LAND flaps. Maintain airspeed at
generator switches OFF, the battery VREF. Pull the LH No. 1, 2, and 3 cir-
provides at least 30 minutes of power cuit breakers after extending the land-
to the emergency bus items (Table ing gear and flaps. The anti-skid system
5C-D, previous page). is inoperative; multiply landing dis-
Aircraft with the optional Honeywell tance by 1.3%. Braking is available
Primus II radio system use the clear- until accumulator discharge; use the
ance delivery unit (CDU) to operate emergency braking system.
the COMM 1 and NAV 2 radios with Loss of Both Generators
the battery switch in EMER and the
generators OFF. With a dual generator failure, direct
With the battery switch in EMER and primary action toward resetting them.
the generator switches OFF, the If only one generator resets, decrease
following systems are inoperative: electrical load as much as possible; the
same as with SINGLE GENERATOR
■ landing gear normal operation and
FAILURE. Limit operating generator
indication; use the blow-down system load to 300 amps (Figure 5C-18A).
■ flaps; perform a flaps inoperative
If neither generator comes on-line, the
landing battery is the only source of electrical
■ anti-skid/power braking system; use power. Turn the flood lights to full
emergency braking bright (night). Position the battery
■ engine anti-ice valves; valves in switch to EMER to isolate the Hot
open position Battery and Emergency buses from the
aircraft electrical system. The battery
■ outside air temperature gage
provides approximately 30 minutes of
■ engine instruments; fan speed indi- power to the Emergency bus (Table
cators (N1) only. 5C-D, previous page).
On units 0001 through 0031, select Aircraft with the optional Primus II
EMER COMM and HDPH. The flight radios use the CDU to operate the
crew must use the headsets; the inter- COMM 1 and NAV 2 radios with the
com between the crew members is battery switch in EMER and the gen-
inoperative. erators inoperative.
Turn both windshield bleed air manu- The following items are unavailable
al valves off to prevent an excessive or inoperative with only Emergency
amount of hot air from reaching wind- bus power:
shield; the windshield bleed air shut- ■ landing gear normal operation and
off valve fails open. Pull the LH and indication; use the blow-down system
RH No. 1, 2, and 3 circuit breakers.
■ flaps; perform a flaps inoperative
Place the battery switch in BATT and
the generator switches in GEN. Turn landing
all electrical switches OFF except for ■ anti-skid and braking system; use

emergency items. emergency braking

If the severity of smoke warrants, ini- ■ engine anti-ice valves; valves in

tiate smoke removal procedures. Land open position

as soon as possible. ■ outside air temperature gage

When the landing is assured, reset the ■ engine instruments; fan speed (N1)
LH No. 1, 2, and 3 circuit breakers. indicators only.
Lower the landing gear and select

5C-36 For SimuFlite training only Citation V

March 1999
 Electrical System

Turn both windshield bleed air valves Once the landing is assured, select
to OFF. The windshield bleed air shut- BATT with the battery switch to sup-
off valve fail open with an electrical ply power to the landing gear and flap
power loss. Land as soon as possible. systems. Extend the landing gear and
select the appropriate amount of flaps.
Maintain VREF for the landing.

Generator Reset Decision Tree

MAY BE RESET CANNOT BE RESET

OVERVOLTAGE REVERSE CURRENT

FEEDER FAULT UNDERVOLTAGE
FIRE SW PRESS

TRIPS TRIPS

FIELD RELAY POWER RELAY

L/R GEN OFF

Volts -
READ

Greater Than 10V DC Normal 28.5 V DC Less Than 10V DC

But Less Than 28.5V DC

Field Relay Other GEN Volts - READ Field Relay

NOT TRIPPED TRIPPED

GEN SW- OFF If Not Normal GEN SW -

(RESET NOT FUNCTIONAL) GEN SW -OFF RESET, THEN OFF

Volts -
READ

Less Than 10V DC Normal 28.5V DC

GEN SW - LEAVE OFF GEN SW - ON

5C-18A (RESET NOT FUNCTIONAL)

Citation V For SimuFlite training only 5C-37

March 1999
5C-38 For SimuFlite training only Citation V
March 1999
 Electrical System

Electrical System Data

Data
Power Source (one) 20 cell, 44 amp-hour battery
(two) Engine-driven starter/generators 28.5V, 300 amp Summary
(two) AC inverters 26V and 115V, 400 Hz AC
Ground power – 28V, 600 to 1,000 amp
Distribution DC Buses
Hot Battery
Battery
Emergency
LH/RH Main
LH/RH Main Bus Extension
LH/RH Crossover
AC Buses
Number One/Two 26V AC
Number One/Two 115V AC
Control DC
Engine generator switches
Battery switch
Engine fire switches
Avionics DC master switch
AC
AC bus magnetic circuit breakers
AC power switch
AC test switch
Monitor DC
Voltmeters and ammeter
L/R GEN OFF annunciators
BATT O’TEMP annunciator
AC
AC FAIL annunciator
INVERTER FAIL 1/2 annunciators
Protection Circuit breakers
Current limiters
Relays
Generator control units

Citation V For SimuFlite training only 5C-39

March 1999
5C-40 For SimuFlite training only Citation V
March 1999
 Electrical System

Lighting on the Citation V includes

interior, exterior, and emergency light-
cockpit flood lights, engine instrument
flood lights, and two map lights. Lighting
ing. Interior lighting includes systems The majority of the cockpit lighting
for the cockpit, passenger compart- controls are on the lower edge of the
ment, baggage compartments, tailcone pilot’s instrument panel. These con-
compartment, and nose baggage com- trols allow the adjustment of the lights
partment. Exterior lighting includes from OFF to full intensity. The map
the rotating beacon, navigation, land- light controls are on the forward edge
ing, wing inspection, recognition, and of the side consoles.
tail flood lights. An emergency light-
ing system provides interior lighting The lights operate on DC and AC
if a complete electrical failure occurs power from the aircraft’s normal elec-
or if the aircraft experiences a 5G lon- trical system. During a dual generator
gitudinal acceleration (crash). failure, the Emergency bus and the
standby horizon batteries provide
In addition to the interior, exterior, and power for the cockpit flood lights,
emergency lighting systems, the stand- engine instrument flood lights, and
by gyro horizon’s battery pack pro- instrument lighting.
vides power for lighting of selected
instruments in the cockpit. Primary Lighting
The majority of the instruments have
Interior Lighting internal lighting or use post lights.
Interior lighting includes: Electroluminescent panels illuminate
■ cockpit
the switches and control positions.
Each electroluminescent panel con-
■ passenger compartment
sists of a layer of phosphor sandwiched
■ baggage compartment. between two electrodes and encased
in plastic. An electric current flowing
Cockpit Lighting through the electrodes causes the phos-
Primary and secondary lighting sys- phor to glow.
tems illuminate the cockpit instruments Internal instrument lighting uses 5V
and panels. These include internal DC supplied by three inverters: one
lighting of the instruments, electrolu- each for the left, center, and right instru-
minescent panel lighting, and post ment panel lights. The LH PANEL cir-
lights. Secondary lighting includes the cuit breaker supplies 28V DC from the

5C-19 5C-20

Citation V For SimuFlite training only 5C-41

March 1999
 LH Main bus Extension and the RH power from the emergency battery pack
PANEL and CENTER PANEL circuit (see Emergency Lighting).
breakers supply 28V DC from the RH A map light (Figure 5C-22) on the left
Crossover bus. An ON/OFF switch and right overhead panel shine down
supplies power to the inverters and the to illuminate the pilot and copilot sta-
LEFT, CENTER, and RIGHT rheostats tions. Each light receives power from
(Figure 5C-19, previous page) vary the RH Crossover bus through the RH
light intensity. PANEL circuit breaker. A rheostat
The post lights use either 5V or 28V (Figure 5C-23) on the forward left and
DC. The internal and post lighting sys- right side consoles individually controls
tems illuminate the instrument panels, the lights from off to full intensity.
side consoles, and the pedestal. Optional supplemental lighting con-
Electroluminescent panels are on the sists of a cathode light on the under-
instrument panel, circuit breaker pan- side left and right sides of the
els, side consoles, and the control glareshield. The lights powered by an
pedestal. The EL PANEL circuit break- inverter shine down to illuminate the
er supplies 28V DC from the LH Main left and right instrument panels.
Bus Extension to the inverter through
the ON/OFF switch. The inverter then Passenger Compartment
converts 28V DC into 40 to 60V AC, Passenger lighting includes:
400 Hz for the panels. The EL rheo-
■ indirect fluorescent
stat controls the brightness of the pan-
els from off to full intensity. ■ reading

■ entrance
Secondary Lighting
■ footwell
Secondary lighting includes flood,
■ passenger advisory.
map, and supplemental lighting.
Two flood lights on the overhead panel Fluorescent Lighting
(Figure 5C-20, previous page) and one
Fluorescent lights on the overhead pan-
under the fire warning panel illuminate
els provide indirect lighting of the pas-
the cockpit and engine instruments. The
senger compartment. The system con-
FLOOD circuit breaker supplies 28V
sists of two inverters, 10 power
DC controlled by the FLOOD switch
supplies and 10 fluorescent tubes. The
(Figure 5C-21) from the Emergency
system uses 28V DC supplied by the
bus to the lights. When either engine is
LH Main DC bus through the INDI-
in the start cycle, the lights receive
RECT LIGHT circuit breaker (tailcone

5C-21 5C-22 5C-23

5C-42 For SimuFlite training only Citation V

March 1999
 Electrical System

J box). The inverters convert 28V DC receive 28V DC supplied by the RH

into 190V AC for the power supplies. Main bus through the CABIN LIGHTS
Each power supply provides 300 to circuit breaker (tailcone J box).
480 VRMS to its fluorescent tube. The
OFF-BRIGHT-DIM switch forward Footwell
of the entrance door controls the lights A footwell light is on the forward side
(Figure 5C-24). of the toilet area. It receives 28V DC
Selecting BRIGHT illuminates the through the CABIN LIGHTS circuit
lights at full intensity. With the switch breaker (tailcone J box) from the RH
in DIM, the lights illuminate at full Main DC bus. The light operates when
brightness for five seconds then dim. power is on the aircraft.
The fluorescent lighting system also Passenger Advisory
responds to DC input voltage. If the
input voltage to the aft inverter drops Passenger advisory lighting consists of
below 20V, the lights automatically three No Smoking/Fasten Seat Belt
shift to bright. Once voltage exceeds signs (Figure 5C-26) and two exit
20V DC, they return to their original- lights (Figure 5C-27, following page).
ly selected intensity. Operate the lights The No Smoking/Fasten Seat Belt signs
in BRIGHT until a consistent source are on the forward and aft overhead
of 28V DC exists on the aircraft. console and in the toilet area opposite
the belted toilet seat. The exit lights are
Reading Lights above the main entrance door and the
Individual reading lights (Figure emergency exit.
5C-25) for the passengers are above The No Smoking/Fasten Seat Belt signs
the seats. Each light is adjustable and receive 24V DC through the OXY
has an on/off push switch next to it. SEAT BELT circuit breaker (tailcone
The lights receive 28V DC from the J box) from the LH Main DC bus. The
RH Main DC bus through the CABIN PASS SAFETY switch (Figure 5C-28,
LIGHT circuit breaker (tailcone J box). following page) on the pilot’s switch
panel controls the signs. Placing the
Entrance Light
switch in the up position (PASS SAFE-
A light on either side of the overhead TY) illuminates the “No Smoking” and
panel illuminates the main and emer- “Fasten Seat Belt” symbols of the sign
gency exit areas. An illuminated switch and sounds the passenger chime.
forward of the main entrance door and Placing the switch in the SEAT BELT
the passenger safety switch in the cock- position illuminates the “Fasten Seat
pit control both lights. Both lights Belt” symbol and sounds the chime.

5C-24 5C-25 5C-26

Citation V For SimuFlite training only 5C-43

March 1999
 The exit signs do not require power for bus. A manually operated ON/OFF
operation. They are self-illuminating; switch and a microswitch in each door
they glow. release latch control the light. With the
switch ON, opening either door illu-
Baggage Compartment minates the light; closing the door
Baggage compartment lighting pro- extinguishes the light.
vides general lighting of the tailcone,
aft baggage compartment, and the nose Exterior Lighting
baggage compartment. Exterior lighting consists of:
The tailcone and aft baggage com- ■ navigation
partment lights receive 28V DC
■ anti-collision
through the AFT/FWD COMP LT cir-
cuit breaker from the Hot Battery bus. ■ rotating beacon

A manually operated ON/OFF switch ■ recognition

on the door frame and a door-operated
■ landing/taxi
microswitch control the lights. The
tailcone light is detachable for use in ■ wing inspection.

the tailcone area. Closing the door Controls for the exterior lights are on
extinguishes the lights. the pilot’s switch panel (Figure
The nose baggage compartment light 5C-29) and lower left instrument panel
receives 28V DC from the Hot Battery (Figure 5C-30).

5C-27 5C-28

5C-29 5C-30

5C-44 For SimuFlite training only Citation V

March 1999
 Electrical System

Navigation Lights Rotating Beacon

The navigation lights (Figure 5C-31) A red rotating beacon is on the top of
consist of a green light in the right the vertical stabilizer. The beacon
wing tip, a red light in the left wing receives 28V DC from the RH Main
tip, and a white light in the tailcone DC bus through the ROTATING BEA-
stinger. Each light operates on 28V DC CON circuit breaker (tailcone J box).
supplied by the LH Main bus The BEACON switch on the pilot’s
Extension through the NAV circuit switch panel controls it.
breaker. The NAV switch on the lower
left instrument panel controls the lights. Recognition Lights
A white halogen recognition light is
Anti-Collision Lights on each wing tip (Figure 5C-32). The
The anti-collision lights consist of a right light receives 28V DC power
strobe light in the left and right wing from the LH Main DC bus through the
tip with the navigation light. A power RT RECOG LT circuit breaker (tail-
supply supplied with 28V DC from cone J box); the left light receives
the LH Main bus Extension through power from the RH Main DC bus
the ANTI COLL circuit breaker pow- through the LT RECOG LT circuit
ers the strobes. The ANTI COLL breaker (tailcone J box). The RECOG
switch on the lower left instrument switch on the lower left instrument
panel controls the strobes. panel controls the lights.

5C-31 5C-32

5C-33

Citation V For SimuFlite training only 5C-45

March 1999
 Landing/Taxi Lights Emergency
A spare high-intensity light on each Lighting
main gear drag brace (Figure 5C-33) is
a dual-function landing and taxi light. The emergency lighting system con-
The lights only illuminate when the sists of a flood light above the main
landing gear doors are open. The RH and emergency exits and an alternate
Main DC bus powers the right light engine instrument flood light on the
and the LH Main DC bus powers the bottom side of the fire tray. The stand-
left light. The LH LANDING LT and by gyro horizon battery pack also pro-
RH LANDING LT circuit breakers vides power for copilot’s instrument
(tailcone J box) provide circuit protec- lighting during electrical failures.
tion. The LH and RH LANDING A dry 20-cell, nickel-cadmium battery
switches control the lights. Turning the pack above the overhead trim panel
switches ON illuminates the lights once has an inertia switch that activates with
the main gear doors open. a 5G longitudinal force. Normally, the
Hot Battery bus provides power for
Wing Inspection battery charging and operation of the
A wing inspection light (Figure lights. During engine starts, the bat-
5C-34) on the left and right sides of tery pack also illuminates the engine
the fuselage forward of the wing lead- instrument flood light.
ing edge illuminates the wing leading The standby gyro horizon’s battery
edge for inspection purposes. The right pack during an electrical failure pro-
light receives 28V DC supplied by the vides power for lighting of the:
RH Main DC bus through a circuit ■ standby gyro horizon
breaker in the tailcone J box; the left
■ copilot’s attitude indicator (single
light receives 28V DC from the LH
Main bus Extension supplied through EFIS)
the WING INSP circuit breaker on the ■ dual fan tachometer
left circuit breaker panel. The WING ■ copilot’s VSI
INSP switch on the pilot’s switch panel
■ copilot’s airspeed indicator
controls both lights.
■ copilot’s altimeter.

5C-34

5C-46 For SimuFlite training only Citation V

March 1999
This chapter describes the Citation V fire protection system,
which consists of an engine fire detection/extinguishing sys- Fire
tem and manually operated fire extinguishers for use inside the
pressure vessel. Protection
The fire detection system senses and indicates the presence of
an engine overheat condition or fire and provides a visual warn-
ing in the cockpit. Chapter 5D
British-certified aircraft have an aft baggage compartment smoke
detector with an accompanying cockpit annunciator and a fire
alarm bell that sounds if an engine fire or overheat condition
is detected. Australian and German-certified aircraft just have
a fire alarm bell.
The engine fire extinguishing system includes two fire extin-
guisher bottles in the tailcone. Both bottles can be discharged
to either engine. Two portable fire extinguisher bottles are
stowed in the cockpit and cabin area.

Citation V For SimuFlite training only 5D-1

March 1999
5D-2 For SimuFlite training only Citation V
March 1999
 Fire Protection

Fire Protection
Fire Protection System Schematic . . . . . . . . . 5D-4
Table of System
Fire Detection Components Schematic . . . . . . 5D-4
Engine Fire Detection System . . . . . . . . . . . 5D-5
Contents Fire Detection
Fire Detection Loop . . . . . . . . . . . . . . . . . . 5D-5 Components
Fire Control Unit . . . . . . . . . . . . . . . . . . . . 5D-5
ENGINE FIRE PUSH Switchlights
(Warning Function) . . . . . . . . . . . . . . . . . 5D-6
Fire Alarm Bell (British, German,
and Australian-Certified Aircraft) . . . . . . . . . . 5D-6
Aft Baggage Compartment Smoke Detector
(British-Certified Aircraft) . . . . . . . . . . . . . . 5D-6

Engine Fire Extinguishing . . . . . . . . . . . . . . 5D-7

ENGINE FIRE PUSH Switchlights
(Extinguishing System Function) . . . . . . . . . . 5D-7
Firewall Shutoff Valves . . . . . . . . . . . . . . . 5D-7
Fire Extinguisher Bottles . . . . . . . . . . . . . . 5D-8
BOTTLE ARMED PUSH Switchlights . . . . . . . . 5D-8
Portable Fire Extinguishers . . . . . . . . . . . . . . 5D-9

Preflight and Procedures . . . . . . . . . . . . . 5D-11

Preflight . . . . . . . . . . . . . . . . . . . . . . . 5D-11
FIRE WARN Test . . . . . . . . . . . . . . . . . . 5D-11
Abnormal Procedure . . . . . . . . . . . . . . . . . 5D-11
Engine Fire During Shutdown . . . . . . . . . . . 5D-11
Emergency Procedures . . . . . . . . . . . . . . . 5D-11
Engine Fire in Flight . . . . . . . . . . . . . . . . 5D-12
Engine Fire During Takeoff . . . . . . . . . . . . . 5D-12

Data Summary . . . . . . . . . . . . . . . . . . . 5D-13

Fire Protection Data . . . . . . . . . . . . . . . . . 5D-13

Citation V For SimuFlite training only 5D-3 5D-4

March 1999
Fire Protection System Fire Detection Components

RH CROSSOVER BUS LH T/R RH T/R LH MAIN BUS EXT

ISOLATE ISOLATE 1 2
HYDRAULIC FIREWALL
SHUTOFF VALVES

FORWARD FUEL FIREWALL

1 SHUTOFF VALVES
ARM S
T EMER EMER S ARM WING
LH ENG RH ENG T
UNLOCK
O BOTTLE 1
ARMED FIRE FIRE
BOTTLE 2
ARMED
O
UNLOCK
SPAR
W W
S
PUSH PUSH PUSH PUSH
DEPLOY S
W NORMAL NORMAL W DEPLOY

LH GEN RH GEN
FIELD LH FUEL RH FUEL FIELD
F/W S.O. F/W S.O.
LH HYD F/W RH HYD
SHUT OFF
F/W S.O. F/W S.O.
LH RH
FIRE DET
REAR PRESS CONTROL FUEL
BKHD UNITS FIREWALL
SHUTOFF
VALVE
FIRE CONTROL
BOTTLE 1
3 UNITS

HYDRAULIC
2 THRUST REVERSER 3
RIGHT HYDRAULIC RESERVOIR ISOLATION VALVES
FIREWALL SHUTOFF
VALVE

BOTTLE 2
PRESS RIGHT
GAGE THERMAL
RELIEF/ FIRE ENGINE
FIRE PUMP
DETECT REFILL DETECT
LOOP LOOP

FIRE BOTTLES:
125 CU. IN. CBrF3

LEFT HYDRAULIC LEFT

FIREWALL SHUTOFF ENGINE
VALVE PUMP FIRE CONTROL UNIT

5D-4 For SimuFlite training only Citation V

March 1999
 Fire Protection

The fire detection system identifies a

fire or overheat condition in either the
cable has a fire-proof connector at each
end. Engine Fire
left or right engine nacelle. Comp-
onents of the system include a detec-
The sensor loop connects to its respec-
tive control unit in the tailcone on the
Detection
tion sensor cable (loop), a fire control
unit, and an ENGINE FIRE PUSH
aft pressure bulkhead. Normal DC
power supplies the system through the System
switchlight for each engine. 2-amp LH and RH FIRE DET CBs.
If an engine fire or overheat condition The LH Main bus Extension feeds the
occurs, the fire detection loop directs right FIRE DET CB, while the RH
a signal to the fire control unit. The Crossover bus feeds the left CB via
control unit illuminates the corre- the RH Main bus Extension.
sponding red ENGINE FIRE PUSH As the temperature increases, the semi-
switchlight on the glareshield. When conductor resistance decreases. At
pressed, the same switchlight arms the approximately 530°F (277°C), elec-
extinguishing system. tric current from the inner wire flows
to the outer casing to complete a circuit
Fire Detection Loop at the control unit.
Each engine nacelle houses a 188-
cubic inch, heat-sensing loop (Figure Fire Control Unit
5D-1) of flexible, stainless steel cable. Two fire detection control units, one
The cable contains a 28V DC ener- for each engine, are in the aft fuselage
gized inner wire centered in a highly section, accessible through the tailcone
compacted, metal oxide semiconduc- baggage door.
tor material. The hermetically sealed

5D-1

Citation V For SimuFlite training only 5D-5

March 1999
 Each control unit monitors electrical the fire test. However, if a short is pre-
resistance between the cable inner wire sent during the test, the short discrim-
and outer case on the respective engine inating circuit activates and the sys-
fire detector loop. As resistance in the tem test fails to give proper indications.
cable drops below 200 ohms, the fire
alarm circuit activates. The unit mon- ENGINE FIRE PUSH
itors a second and much lower resis-
tance level to activate a short discrim-
Switchlights
inating lockout circuit. (Warning Function)
The control unit recognizes the differ- When a control unit detects a fire or
ence between an actual fire/overheat overheat condition in the sensor cable,
and a short by monitoring the time it it energizes the corresponding guard-
takes for the resistance to fall. The unit ed ENGINE FIRE PUSH switchlight
interprets an instantaneous drop in on the glareshield (see Engine Fire
resistance below the short discrimi- Extinguishing System, ENGINE FIRE
nating lockout circuit level as a short PUSH switchlights in this chapter).
and blocks the signal from reaching
the ENGINE FIRE PUSH switchlight. Fire Alarm Bell
During an actual fire or overheat sit- (British, German, and
uation, the resistance in the cable Australian-Certified
decreases at a much slower rate. As
the resistance decreases below the fire
Aircraft)
alarm point, the fire relay actuates to The circuitry between the detection
illuminate the fire light; simultane- system sensors and the ENGINE FIRE
ously, the relay disables the short dis- PUSH switchlight also ties in to a fire
criminating lockout circuit. If resis- alarm bell that sounds on British,
tance continues to fall, the short German, and Australian-certified air-
discriminating circuit activates, but its craft only. The alarm bell, with its time
signal has no effect because its output delay relay, is mounted in the left side
is disconnected. panel. Illumination of the ENGINE
To test the integrity of the fire warning FIRE PUSH switchlight activates the
system, select the FIRE WARN posi- bell, which sounds for two seconds on
tion of the rotary test switch. The con- Australian-certified aircraft and 4.25
trol unit has an internal test resistor that seconds on British and German-
is applied to the center wire of the sen- certified aircraft, then automatically
sor cable. During the test, the sum of shuts off.
the test resistance plus the sensor cable
center conductor resistance is less than Aft Cabin Baggage
the fire alarm point. This actuates the Smoke Detector
system and verifies the continuity of
the sensor cable, fire bridge and ampli- (British-Certified Aircraft)
fier, fire relay, and the fire warning indi- On British-certified aircraft only, a
cator (see FIRE WARN TEST in the smoke detector in the aft cabin baggage
Preflight section of this chapter). compartment activates a red SMOKE
The sum of the test resistance and the DETECT annunciator in the event of
cable resistance falls between the fire smoke in that area. The WARN LTS 1
alarm point and the short discriminat- CB on the left circuit breaker panel
ing alarm point so that the short dis- powers the smoke detector.
criminating circuit is not tested during

5D-6 For SimuFlite training only Citation V

March 1999
 Fire Protection

The engine fire extinguishing system

consists of two fire extinguisher
■ The white BOTTLE 1 and 2
ARMED PUSH switchlights (fire Engine Fire
bottles, deployment tubes, cockpit
discharge controls, and associated elec-
extinguisher discharge switches)
illuminate to indicate the bottles are Extinguishing
trical circuits. armed to discharge, when pushed,
to the appropriate engine. System
ENGINE FIRE PUSH ■ The system removes electrical

Switchlights power from the normally closed

thrust reverser isolation valve on the
(Extinguishing System corresponding engine; this causes
Function) the valve to remain closed (see
The red ENGINE FIRE PUSH switch- Miscellaneous chapter).
lights (Figure 5D-2) are warning The extinguishing system can be dis-
annunciators for the detection system as armed by depressing the ENGINE
well as arming mechanisms for the FIRE PUSH switchlight a second time.
extinguishing system. This disarms the fire bottles and opens
To arm the extinguishing system, lift the fuel and hydraulic firewall shutoff
the transparent plastic guard cover and valves. If the engine is to be restarted,
press the illuminated ENGINE FIRE the de-energized starter/ generator field
PUSH switchlight; the following occurs. relay must be reset with the generator
control switch.
■ The fuel and hydraulic firewall shut-
off valves on the affected engine Firewall Shutoff Valves
close.
Four electrically motorized firewall
■ The L or R F/W SHUTOFF annun-
shutoff valves, one fuel and one
ciator illuminates in response to hydraulic valve for each engine, close
both valves closing. when the corresponding ENGINE
■ With fuel cut off, the engine shuts FIRE PUSH switchlight is initially
down. pressed; they open if the switchlight is
■ The associated engine starter/gen- pressed a second time. The valves open
erator field relay trips. or close in approximately one second.

5D-2

Citation V For SimuFlite training only 5D-7

March 1999
 A fuel firewall shutoff valve in the dry cartridge assembly is a plumbing net-
bay area of each wing isolates the pri- work that directs the extinguishing
mary fuel line to its respective engine. agent to an engine. Each bottle can be
The left and right hydraulic firewall discharged to either engine, depend-
shutoff valves are between the ing on which cartridge is activated.
hydraulic reservoir and the suction side The ENGINE FIRE PUSH switchlight
of each hydraulic pump. selects the correct cartridge.
The left or right ENGINE FIRE PUSH Each bottle contains monobromotri-
switchlight directs 28V DC via a 7.5- fluormethane (CBrF3) pressurized to
amp FW SHUTOFF CB on the left cir- 600 ±75 PSI at 70°F (21°C). The extin-
cuit breaker panel to close the fuel and guishing agent has no damaging effects
hydraulic shutoff valves on the affect- on any of the engine components;
ed engine’s side. The LH or RH F/W therefore, it is not necessary to clean or
SHUTOFF annunciator illuminates replace the components after the
when both valves are closed on the bottles are fired.
corresponding engine. Relief valves protect the fire bottles
The LH Main bus Extension feeds the from overheat/overpressure. An abnor-
RH F/W SHUTOFF CB, while the RH mal tailcone temperature (approxi-
Crossover bus feeds the LH F/W mately 210°F or 99°C) may cause the
SHUT-OFF CB. bottles to automatically discharge
through the relief valves. The relief
Fire Extinguisher Bottles valve is a filler fitting on each bottle
Two 125-cubic inch spherical fire that melts to channel the extinguish-
extinguisher bottles (Figure 5D-3) in ing agent into the tailcone.
the tailcone are cross-plumbed so that Pressure gages are on the bottom of
either bottle can be discharged into each fire extinguisher bottle. During
either engine. the preflight inspection, compare the
Each fire bottle has two fire extin- gage reading to the temperature chart
guisher cartridges. Attached to each between the bottles to determine if the
bottle requires servicing.

5D-3

5D-8 For SimuFlite training only Citation V

March 1999
 Fire Protection

BOTTLE ARMED PUSH necessarily mean the bottle contents

Switchlights have discharged; it means only that the
switchlight has been pushed. The only
Pressing the ENGINE FIRE PUSH way to verify the bottle has discharged
switchlight applies power to the dis- is to observe the pressure gage on the
charge controls, selects the appropriate bottle during post/preflight.
cartridge on each fire bottle, and illu-
minates the 1 and 2 BOTTLE ARMED For example, if both BOTTLE
PUSH switchlights (Figure 5D-4). ARMED lights are pushed, the lights
extinguish and the bottles should dis-
To discharge a fire extinguisher into charge. If for some reason the
the affected engine, push either the 1 or ENGINE FIRE PUSH switchlight is
2 BOTTLE ARMED PUSH switch- pushed to reset the system and then
light. This directs 28V DC to the cor- pushed again, the BOTTLE ARMED
responding fire extinguisher cartridge. switchlights re-illuminate even though
The resulting explosive pressure breaks the fire bottles are empty.
the end of the cartridge housing assem-
bly, and the fire extinguishing agent
discharges through the swivel into the
Portable Fire
distribution network. Extinguishers
Deployment tubes route the extin- Two portable fire extinguishers are
guishing agent from the bottles to the mounted in red quick-release brack-
selected fire area. In addition, nine ets: one under the copilot’s seat
holes in the front flange of the upper (Figure 5D-5) for easy access to the
thrust reverser door allow extinguish- crew in flight, and a second extin-
ing agent to enter the area between the guisher in the passenger cabin behind
upper thrust reverser door and the duct the aft left seat (Figure 5D-6).
assembly. Each portable extinguisher is rated for
Pushing the BOTTLE ARMED PUSH class B and C fires and contains 4 lbs
switchlight extinguishes its respective of Halon 1211, nitrogen-charged to
light. An extinguished light does not 125 PSI.

5D-4 5D-5 5D-6

Citation V For SimuFlite training only 5D-9

March 1999
5D-10 For SimuFlite training only Citation V
March 1999
 Fire Protection

Preflight Abnormal Procedure Preflight and

During the exterior preflight inspec- The following is a brief explanation
tion, check the pressure gage on each
fire bottle for an indication of 600 ±75
of what happens during an abnormal
procedure that pertains to the fire pro-
Procedures
PSI at 70°F (21°C). Consult the cor- tection system. Please refer to your
rection table (Figure 5D-7) between SimuFlite Citation V Operating Hand-
the bottles if the temperature is other book for a detailed checklist.
than 70°F.
Engine Fire
During the interior preflight inspection, NOTE: Weigh the portable fire
verify pressures of 125 PSI on portable
During Shutdown
extinguishers at least once every
fire extinguishers in the cockpit and A high or sustained reading on the ITT six months. If weight falls below
cabin. Check that the expiration date gage during shutdown indicates a 4-1/2 lbs, ensure than
on the bottle placard is current. continuing combustion in the engine maintenance personnel test for
burner can. leakage and recharge the bottle.
FIRE WARN Test
Check that the affected engine’s throt- After use, recharge immediately.
Test the fire detection system prior to tle is in the OFF position to ensure fuel
flight by rotating the rotary TEST flow is cut off to the engine. Press the
switch (Figure 5D-8) on the left panel START pushbutton momentarily to
to the FIRE WARN position. Check start the engine rotating; this will con- CAUTION: When operating
that both ENGINE FIRE PUSH tinue air flow through the engine to with emergency power, engine
switchlights illuminate; this verifies assist in the removal of any residual fire protection is not available.
the following: fuel and lower the internal tempera-
■ sensor loop continuity (the loop is not ture. After 15 seconds, press the
broken, kinked, open, or shorted) STARTER DISENGAGE pushbutton.
■ operative control unit circuitry and NOTE: If the engine fire during
relays
Emergency shutdown continues, press the
■ power to the switchlights. Procedures appropriate ENGINE FIRE
PUSH switchlight to shut off fuel
If one of the switchlights does not illu- The following is a brief explanation of at the firewall valve. Then push
minate, there may be a short in the what happens during emergency pro- either BOTTLE ARMED PUSH
circuitry, the bulbs in the switchlight cedures that pertain to the fire protection switchlight to discharge an extin-
may be burned out, or a sensor cable system. Please refer to your SimuFlite
guisher bottle into the engine.
on the associated engine may be Citation V Operating Handbook for spe-
disconnected. cific detailed checklists.

5D-7 5D-8

Citation V For SimuFlite training only 5D-11

March 1999
 Engine Fire In Flight Handbook or AFM checklist for non-
memory items. If fire persists, land as
If an ENGINE FIRE PUSH switch- soon as possible.
light illuminates, move the throttle on
the corresponding engine to IDLE. If Engine Fire
the ENGINE FIRE PUSH switchlight During Takeoff
extinguishes, land as soon as practical
without taking further action. If an engine fire occurs during take-
off, maintain aircraft control as a first
However, if the ENGINE FIRE PUSH priority. After establishing a positive
switchlight remains illuminated, push rate of climb, retract the landing gear
the ENGINE FIRE PUSH switchlight and maintain a minimum air-speed of
to shut down the engine and arm the V2 until the aircraft reaches at least
extinguishing system; then press a 400 ft AGL. Accelerate to V2 +10 and
BOTTLE ARMED PUSH switch- retract the flaps. Continue accelerat-
light to discharge a fire extinguisher ing to VENR and accomplish the check-
bottle into the affected engine. Refer list items for Engine Fire in Flight
to the SimuFlite Citation V Operating above.

5D-12 For SimuFlite training only Citation V

March 1999
 Fire Protection

Fire Protection Data Data

Power Source Left Main bus Extension
Right Crossover bus
Summary
Distribution Each bottle discharges to either engine, as selected
by the ENGINE FIRE PUSH switchlight
Control L/R ENGINE FIRE PUSH switchlights
BOTTLE 1/2 ARMED PUSH switchlights
Monitor L/R ENGINE FIRE PUSH switchlights
L/R F/W SHUTOFF annunciators
Bottle pressure gages
Protection Overtemperature/pressure relief valve on each
fire bottle

Citation V For SimuFlite training only 5D-13

March 1999
5D-14 For SimuFlite training only Citation V
March 1999
The primary flight controls consist of the ailerons, elevators,
and rudder. The mechanically driven flight controls operate Flight
through input received from the cockpit via direct cable.
The electrically controlled and hydraulically operated secondary
Controls
flight controls consist of flaps and speedbrakes.
Trim tabs on the left aileron, elevators, and rudder provide for roll,
yaw, and pitch trim respectively. In addition to the primary and
Chapter 5E
secondary flight controls, stall warning devices complete the
flight controls system.

Citation V For SimuFlite training only 5E-1

March 1999
5E-2 For SimuFlite training only Citation V
March 1999
 Flight Controls

Primary Flight Controls Schematic . . . . . . . . . 5E-4

Table of
Primary Flight Controls . . . . . . . . . . . . . . . 5E-5
Ailerons . . . . . . . . . . . . . . . . . . . . . . . . 5E-5
Contents
Roll Trim . . . . . . . . . . . . . . . . . . . . . . . 5E-5
Elevators . . . . . . . . . . . . . . . . . . . . . . . 5E-5
Manual Pitch Trim . . . . . . . . . . . . . . . . . . 5E-6
Electric Pitch Trim System . . . . . . . . . . . . . . 5E-6
Rudder . . . . . . . . . . . . . . . . . . . . . . . . 5E-7
Autopilot/Yaw Damper . . . . . . . . . . . . . . . . 5E-7
Rudder Trim . . . . . . . . . . . . . . . . . . . . . 5E-8
Rudder/Aileron Interconnect . . . . . . . . . . . . . 5E-8
Control Lock . . . . . . . . . . . . . . . . . . . . . . 5E-8

Secondary Flight Controls Schematic . . . . . . 5E-10

Flaps System Schematic . . . . . . . . . . . . . . 5E-10
Speedbrakes System Schematic . . . . . . . . . 5E-12
Secondary Flight Controls . . . . . . . . . . . . . 5E-13
Flaps . . . . . . . . . . . . . . . . . . . . . . . . . 5E-13
Speedbrakes . . . . . . . . . . . . . . . . . . . . . 5E-14
Speedbrake Switch . . . . . . . . . . . . . . . . . 5E-14
Stick Shaker . . . . . . . . . . . . . . . . . . . . . 5E-15

Preflight and Procedures . . . . . . . . . . . . . 5E-17

Preflight . . . . . . . . . . . . . . . . . . . . . . . 5E-17
Abnormal Procedures . . . . . . . . . . . . . . . . 5E-17
Electric Elevator Trim Failure . . . . . . . . . . . . 5E-17
Electric Trim Inoperative . . . . . . . . . . . . . . 5E-17
Jammed Elevator Trim Tab . . . . . . . . . . . . . 5E-17
Flaps Inoperative Approach and Landing . . . . . 5E-17

Data Summaries . . . . . . . . . . . . . . . . . . 5E-19

Flaps Data . . . . . . . . . . . . . . . . . . . . . . 5E-19
Speedbrakes Data . . . . . . . . . . . . . . . . . . 5E-19

Citation V For SimuFlite training only 5E-3

March 1999
Primary Flight Controls

ELEVATOR TRIM TABS

RUDDER

AILERONS RUDDER
SERVO &
TRIM TAB

ELEVATORS

AILERON TRIM TAB

5E-4 For SimuFlite training only Citation V

March 1999
 Flight Controls

Primary flight controls allow control

of the aircraft through the pitch, roll,
Rotation of either control wheel moves
ailerons inversely to one another Primary Flight
and yaw axes. The control columns
and rudder pedals in the cockpit con-
through a synchronized interconnect-
ing cable system (i.e., the aileron on Controls
trol the primary flight controls via one wing moves up as the aileron on
cables and bellcranks. the opposite wing moves down). Full
range of travel for the ailerons is 19 ±1°
Ailerons up and 15 ±1° down.
Ailerons (Figure 5E-1) on the outboard Roll Trim
trailing edge of each wing provide roll
control of the aircraft mechanically A trim tab connected to the trailing
through the control wheel or electron- edge of the left aileron provides roll
ically through the autopilot servo. The trim. A trim tab control knob (Figure
autopilot aileron servo has an override 5E-2) on the center control pedestal
function so the pilot can physically below the autopilot control head
overpower the servo by manually rotat- mechanically operates the trim tab via
ing the control wheel. a cable and an actuator. The rotation
of the knob transmits mechanical
Rotation of the control wheel trans- action to the aileron trim tab actuator
mits aileron directional information to screws that adjust the trim tab up or
the sector assembly and causes it to down.
rotate. The sector assembly is below
the cabin floor aft of the rear carry- An indicator on the pedestal displays
through spar. Rotation of the sector aileron trim position. The indicator dis-
assembly moves the wing cable assem- plays the amount of trim away from
bly to the aileron actuators. The actu- neutral.
ators move the aileron to the desired
position. Elevators
The autopilot actuates the aileron sys- The elevators (Figure 5E-3) on the
tem by rotating the aileron sector trailing edge of the horizontal stabi-
assembly that moves the wing cable lizer provide pitch control of the air-
assembly, aileron actuators, and craft. The elevators are controlled by
ailerons. In addition, autopilot move- two sources, either manually with the
ment of the sector assembly moves the control columns, or electrically with
control wheels in the cockpit. the autopilot servo.

5E-1 5E-2 5E-3

Citation V For SimuFlite training only 5E-5

March 1999
 Movement of the control wheel fore or Electric Pitch Trim System
aft moves the elevators mechanically.
To move the control surfaces, the ele- A flap/elevator trim mixer system per-
vator system uses a bellcrank assem- mits electrical actuation of the eleva-
bly in the aft section of the tailcone, a tor trim control system. The mixer
sector assembly below the flight com- assembly actuates with one or more of
partment floor, and cable assemblies. the following inputs:
Movement of the control columns fore ■ trim switch on the control wheel

or aft mechanically moves the eleva- selected up or down

tor through the sector assembly and ■ autopilot trim system
bellcrank assembly via the cable assem-
■ flap movement in the range of 15°
blies. A four-cable, direct connect
assembly from the control column to to 25°
the tail section drives the elevators ■ elevator trim adjustment when flaps

through a range of 22 +0.5° -1° up and move between 15° and 25°.
15 ±1° down. An electric motor on the trim mixer
The autopilot elevator servo cables assembly engages when one or more of
apply force on the elevator cables to the above conditions occur. The motor
reposition the elevator bellcrank; this drives the sector in the appropriate
deflects the elevator. The elevator cable direction using a chain drive. The chain
assemblies contain cable clevis fittings drive moves the elevator trim cables,
that provide attachment points for the which rotate the actuator screws. The
autopilot elevator servo cables. screws adjust the trim tab. The LH
The autopilot servo includes an over- Main bus Extension supplies electri-
ride function that allows the crew to cal power for the electric motor.
overpower the servo through move- A split trim switch on the left side of
ment of the control column (for a the pilot’s control wheel (Figure 5E-5)
detailed description of the autopilot, or one on the right side of the copilot’s
see Avionics chapter). control wheel provides electrical con-
trol of the motor. The switch completes
Manual Pitch Trim an electrical circuit to the electric trim
A pitch trim tab on each elevator allows motor that positions the tab. The pilot’s
for minor adjustments to the elevator. trim switch interrupts and overrides
The tabs provide a balanced and sta- the copilot’s switch.
ble condition of flight. A manual ele- The AP/TRIM disconnect switch on
vator trim wheel (Figure 5E-4) on the the pilot’s and copilot’s yoke electri-
center pedestal transmits directional cally disables the electric trim if a mal-
information to the tabs. The system function in electric trim should occur.
uses cable assemblies to transmit the Pulling the 5 amp PITCH TRIM CB
information to the elevator trim tab on the left circuit breaker panel removes
actuators. Rotation of the actuator power from the electric trim motor.
screws deflects the trim tab to a full The autopilot trim system signals the
range of travel of 4.5 +1° -0° up and mixer assembly to position the eleva-
16 +0.5° -0° down. tor trim tabs. This neutralizes the
A mechanical pointer driven by the control forces on the autopilot servo
trim wheel on the center pedestal indi- and allows the elevator to remain in
cates NOSE UP or NOSE DOWN trim through the full range of auto-
trim. A takeoff (TO) marker indicates pilot control.
the proper trim for takeoff.

5E-6 For SimuFlite training only Citation V

March 1999
 Flight Controls

Included into the flap system is an rudder bellcrank transmits movement

automatic elevator trim adjustment. information to a torque tube that
The automatic elevator trim adjust- deflects the rudder.
ment compensates for rapid pitch
changes encountered when the flaps Autopilot/Yaw Damper
are cycled up or down. The autopilot electrically controls the
The flap-trim rate moves faster than rudder through autopilot rudder servo
the electric trim rate. Flap-trim and inputs to the system. The autopilot
electric trim rate inputs are additive. computer receives signals from the
Actuation of the control wheel trim directional gyro and supplies input
switch with the flaps traveling between information to the autopilot rudder
15° to 25° results in an increase or servo for rudder deflection and yaw
decrease in trim tab movement rate as dampening. The servo cables apply
required by the two inputs. force to the fuselage cable assemblies
Manually trimming or momentarily to reposition the rudder bellcrank that
pressing the AP/TRIM DISC switch attaches to the rudder torque tube; this
(Figure 5E-5) overrides the flap-trim deflects the rudder.
system momentarily. The autopilot rudder servo includes an
override function to allow the crew to
Rudder overpower the autopilot by manually
depressing the rudder pedals.
The rudder (Figure 5E-6) on the trailing
edge of the vertical stabilizer provides The yaw damper provides automatic
effective yaw control of the aircraft at control of transient yaw motion and is
all flight speeds. A direct connect cable a function of the autopilot. Engaging
system from both sets of rudder pedals the autopilot automatically engages the
to the tail section drives the aircraft rud- yaw damper. The damper may be sep-
der. Full range of motion is 22 +1° -0° arately engaged by pressing the YD
either left or right of center. ENGAGE switch on the autopilot con-
trol head (Figure 5E-7, following page).
Torque tube and bridge assemblies
connect the pilot’s and copilot’s rud- Disengage the yaw damper by pressing
der pedal sets together. The connec- either the AP/TRIM disconnect switch
tion provides corresponding rudder or the go-around button. Disengaging
pedal movement between the sets. the autopilot with the electric trim
Depressing the rudder pedals transmits switch does not disengage the yaw
directional information through fuse- damper.
lage cables to the rudder bellcrank. The

5E-4 5E-5 5E-6

Citation V For SimuFlite training only 5E-7

March 1999
 Rudder Trim springs and cable clamps. Movement
NOTE: Neutralize the primary of the rudder pedals moves the ailerons
flight controls before engaging The rudder trim tab provides for aero- in proportion to the rudder position.
the CONTROL LOCK. dynamic adjustment of the rudder This provides lateral stability as
through manual deflection of the tab. required by certification standards. The
The rudder trim tab system consists of system can be overridden by manual
a trim control assembly in the control pilot input to the rudder pedals and
CAUTION: Due to maximum pedestal, cable assemblies in the fuse- control wheel when cross control is
nose wheel deflection decreased lage, a trim tab on the trailing edge of required (i.e., crosswind landing or
to 60°, do not tow the aircraft the rudder, and a rudder trim tab actu- sideslip).
with the lock engaged. ator on the trailing edge of the verti-
cal stabilizer. The trim tab deflects 10
±1° to the left and right.
Control Lock
Rotating the rudder trim control wheel The control lock (Figure 5E-9) secures
(Figure 5E-8) repositions the fuselage the ailerons, elevators, and rudder in
Yaw Damper Limitation
cable assemblies routed to the trim tab a streamlined neutral position. In addi-
The yaw damper may be used for tion, the lock secures the throttles in
actuator. The actuator displaces the rud-
all phases of flight except takeoff the cutoff position. The control lock
der trim tab that moves the rudder due
and landing. system consists of a control handle,
to aerodynamic forces acting on the tab.
cable assemblies, push rods, and bell-
In addition, the rudder trim tab is a cranks.
servo boost tab. The servo boost tab
boosts the rudder, displacing it from To engage the control lock, align the
the neutral position. The servo moves nose wheel, neutralize all flight con-
the trim tab approximately 1/2° in the trols, and move both throttles to CUT-
opposite direction for each degree of OFF. Pull the T-handle below the
rudder movement. instrument panel on the left side straight
aft to the detent. To release the flight
The cockpit indicator on the center controls and throttles, rotate the CON-
pedestal has neutral, full left, and full TROL LOCK T-handle 45° clockwise
right positions. and push it in to the unlocked position
(return to the horizontal position).
Rudder/Aileron Check flight controls for proper move-
Interconnect ment before flight.
The rudder system interconnects to the
aileron system via an arrangement of

5E-7 5E-8 5E-9

5E-8 For SimuFlite training only Citation V

March 1999
 Flight Controls

Secondary Flight
Controls
Flaps System

Citation V For SimuFlite training only 5E-9 5E-10

March 1999
Secondary Flight Controls Flaps System
FLAPS EXTENDING

FLAP EXTENDING FLAP EXTENDING

WING FLAPS

SPEEDBRAKES
(UPPER & LOWER FLAP
WING SURFACES) SOLENOID
VALVE

PRESSURE RETURN

FLAPS RETRACTING

FLAP RETRACTING FLAP RETRACTING

PRESSURE
RETURN

FLAP
SOLENOID
VALVE

PRESSURE RETURN

5E-10 For SimuFlite training only Citation V

March 1999
 Flight Controls

Citation V For SimuFlite training only 5E-11

March 1999
Speedbrakes System
LH ACTUATOR (EXTENDED) RH ACTUATOR (EXTENDED)

LH ACTUATOR (RETRACTED) RH ACTUATOR (RETRACTED)

SPEED
BRAKE SPD BRAKE
EXTEND EXTENDED

LH MAIN BUS EXT

UP LIMIT SW DOWN LIMIT

5 SWITCH
EXTEND RETRACT

SOLENOID VLV RETRACTED

EXTENDED

THERMAL
RELIEF VALVE

SAFETY VALVE

RETURN PRESSURE
INLET PRESSURE
HYDRAULIC
BYPASS
VALVE

RETRACT
(TYPICAL) POWER LEVER
SWITCHES
INLET PRESSURE
ABOVE 85% N2
RETURN FLOW EXTEND

EXTEND HOLDING
RETRACT RELAY L RH

(IN TAILCONE)

5E-12 For SimuFlite training only Citation V

March 1999
 Flight Controls

The secondary flight controls consist of

wing flaps and speedbrakes.
er moves along a slot next to the flap
lever to indicate the actual flap posi- Secondary
tion. In addition, the pointer removes
Flaps electrical power from the actuation Flight Controls
system once the flaps reach the prese-
The flaps (Figure 5E-10) on the trail-
lected position. A flap preselect control
ing edge of each wing are inboard of
cable under the cabin floor attaches to
the ailerons. With the flaps partially
both the preselect assembly and a
extended, an increase in lift occurs.
crossover cable. Components that con-
Full flap extension increases drag,
trol the flap actuation system include
which helps reduce speed.
up and down position switches and an
The flaps, which are constructed of arm and cam unit.
graphite composite laminates, consist
The flap actuation system consists of
of two segments on each wing.
an electrically operated flap control
The electrically controlled and hydrauli- valve and hydraulic actuators. In addi-
cally actuated flaps travel between 0° tion, the system includes four bell-
to 35°. The flap preselect lever (Figure cranks in each wing that connect to
5E-11) can be positioned to any flap pushrods. The pushrods operate the
position between UP and FULL down. inboard and outboard flaps.
Mechanical detents in the preselect
The flap control valve is a three-posi-
mechanism mark the T.O. (7°) and T.O. Flaps Limitations
tion, four-way solenoid operated valve.
& APPR (15°) flap position. To repo- Flaps extended speeds (VFE)
Up and down position switches on the
sition the flap lever from the detent posi-
preselect lever control electrical power T.O., T.O.
tion, press the lever down and slide it
to the valve. When flaps are selected, & APPR (15°) . . . . 200 KIAS
forward or aft. To select any position
the hydraulic bypass valve closes and
with the flap lever and avoid the detents, LAND (35°) . . . . . 173 KIAS
the HYD PRESS ON annunciator illu-
press and hold the lever down and slide
minates. The energized flap valve
it to the desired setting.
shifts to allow hydraulic fluid to enter
A flap autotrim system retrims the the extend or retract ports, depending
elevator trim tabs to compensate for on the preselect lever position. The
rapid pitch changes that result from flap pressure entering the actuator moves
movement between 15° to 25° (see Flap/ an internal piston. Fluid from the oppo-
Elevator Trim System this chapter). site side of the piston returns to the
The flap preselect system consists of hydraulic system through the valve.
the preselect lever and a pointer assem- The actuator rotates a bellcrank, which,
bly on the control pedestal. The point- in turn, transmits directional informa-

5E-10 5E-11

Citation V For SimuFlite training only 5E-13

March 1999
 tion to the remaining wing bellcranks without exceeding maximum operat-
via pushrods; the flaps travel to the pre- ing speeds (i.e., VMO/MMO) along with
selected position. Reaching the prese- spoiling lift during the landing roll. The
lected position, the arm and cam unit of angular travel for the upper speedbrake
the pointer assembly opens the up or panel is 60 ±2°; the lower speedbrake
down position switch to de-energize panels close to a faired position when
the valve. The valve then shifts to block the upper panels close. The speedbrakes
all four ports; this maintains fluid pres- are on the upper and lower surface of
sure to retain flap selection. The each wing forward of the flaps.
hydraulic system returns to an open- Mechanical ball studs and two spring
center condition. retainers hold the lower speedbrake
Right and left limit switches prevent panels up. Improper adjustment or
flap overrun. The switches open to loose retainers may allow the lower
remove power from the control valve panel to droop open; this activates the
upon completion of full up or full down limit switch and causes the hydraulic
flap travel. This prevents damage to system to pressurize and reposition the
the flaps. speedbrakes to retract.
Selecting the flaps past T.O. & APPR Speedbrake Switch
without the gear down and locked
sounds the gear warning horn. The A SPEED BRAKE switch on the throt-
HORN SILENCE button cannot tle pedestal with EXTEND/RETRACT
silence the horn in this condition. positions controls the speedbrakes
The flap control valve requires elec- (Figure 5E-13). Moving the switch to
trical power to move out of the block EXTEND closes the hydraulic bypass
position. With loss of electrical power valve to pressurize the hydraulic sys-
the flaps remain in the last position, tem; the HYD PRESS ON annuncia-
even if the flap preselect lever is tor illuminates. The speedbrake sole-
moved to a new position. noid valve moves to the extend position
to direct hydraulic fluid flow at 1,500
PSI to the extend side of the speed-
Speedbrakes brake actuators.
The hydraulically actuated and elec- When the speedbrakes are fully
trically controlled speedbrakes (Figure extended, an UP limit switch closes
5E-12) provide increased descent rates and the white SPEED BRAKE

5E-12 5E-13

5E-14 For SimuFlite training only Citation V

March 1999
 Flight Controls

EXTEND annunciator illuminates. The This allows hydraulic fluid pressure

hydraulic bypass valve reopens to to flow to the retract side of the speed-
relieve pressure, and the HYD PRESS brake actuator. The SPEED BRAKE
ON annunciator extinguishes. The EXTEND annunciator extinguishes
solenoid valve returns to neutral to as the speedbrakes retract. Finally, the
block all hydraulic fluid lines to the hydraulic system depressurizes.
speedbrake actuator and maintain
speedbrakes in the extended position. Stick Shaker
If an electrical failure occurs with the A stick shaker on the forward side of
speedbrakes extended, a spring-loaded the pilot’s control column warns of an
safety valve opens to provide a return impending stall (Figure 5E-14). The
path for the hydraulic fluid. This stick shaker uses an electric motor with
allows the speedbrakes to blow down. rotating weights to induce a vibration
If an electrical failure occurs with the feel to the control column.
speedbrakes retracted, they do not
Vibration occurs at approximately 8%
extend. The LH Main bus Extension
to 10% above the actual stall speed.
supplies electrical power for speed-
Stick shaker activation normally
brake operation through the SPEED
occurs before stall buffet except in the
BRAKE CB.
clean configuration where it is near-
If a dual hydraulic pump failure occurs ly the same and either could occur
with the speedbrakes extended, move first.
the switch to RETRACT to energize
The stick shaker receives inputs and
the speedbrake control valve. This
power from the angle-of-attack sys-
allows the speedbrakes to trail.
tem (see Avionics chapter). To test the
Override microswitches in the tailcone shaker system before flight, use the
prevent speedbrake extension above rotary switch on the pilot’s switch
85% N2. Advancing the throttle levers panel (Figure 5E-15).
beyond 85% N2 automatically retracts
The LH Main bus Extension supplies
the speedbrakes.
electrical power for the stick shaker
Once the speedbrake switch is in motor through the ANGLE OF
RETRACT, the hydraulic system pres- ATTACK CB.
surizes and the speedbrake solenoid
valve moves to the retract position.

5E-14 5E-15

Citation V For SimuFlite training only 5E-15

March 1999
5E-16 For SimuFlite training only Citation V
March 1999
 Flight Controls

Preflight Jammed Elevator

Preflight and
During the external inspection, check Trim Tab
all control surfaces for freedom of
movement and general security. In
In the event of a jammed elevator trim
tab during cruise, maintain trimmed
Procedures
addition, check the rudder trim tab to speed as long as practical. On the
ensure it moves in the opposite direc- approach and landing phases of flight,
tion of rudder displacement. Check all do not extend the flaps; use the flap
visible hydraulic lines and components inoperative approach and landing pro-
for leaks. cedures. Do not use the autopilot.
During cockpit inspection, ensure the With a jammed elevator trim tab during
flap lever and position indicator agree a takeoff or go-around procedure, NOTE: Do not use the autopilot
with the actual flap position, the ele- reduce power as necessary to maintain with the electric trim
vator trim is within the takeoff range, 120 KIAS or less. Maintain the current inoperative. The autopilot is
the aileron trim tab is centered, the rud- flap position. Minimum airspeeds are: unable to trim out servo torque.
der trim tab is centered, and the control ■ LAND flaps (35°) – VREF Disengaging the autopilot with
lock is unlocked. light servo torque may result in a
■ T.O. and APPR (15°) – VAC (VAPP)
significant pitch change.
Abnormal +5 KIAS
■ T.O. (7°) – VREF +15 KIAS.
Procedures
There are no emergency procedures Flaps Inoperative
associated with the flight controls sys- Approach and Landing
tem. The following is a discussion of The Citation V is approved for only a
abnormal procedures for the system. FULL flap landing so the only data
For list of specific procedural steps, available is for a full flap landing con-
please refer to your SimuFlite Citation figuration. Any position other than
V Operating Handbook. FULL down is considered abnormal.
Electric Elevator Trim The flaps inoperative approach and
Failure landing check list is similar to the nor-
mal landing check except for adjust-
If an electric elevator trim runaway ment of the landing distance. To adjust
occurs, press and hold the AP/TRIM the landing distance, take the FULL
disconnect switch on the pilot’s or flap landing distance and multiply it
copilot’s control wheel. The switch by 1.2.
removes electrical power and stops
movement of the electric trim. Remove Airspeeds adjustments are made based
electrical power to the elevator trim on the amount of flaps available; all
system by pulling the PITCH TRIM speeds are adjusted from the original
CB. After removing power to the elec- FULL flap VREF.
tric trim motor, release the AP/TRIM ■ flaps at 15°, adjust speed to VREF
disconnect switch. Continue the flight +5
using manual elevator trim. ■ flaps at 7°, adjust speed to VREF +10

Electric Trim Inoperative ■ flaps at 0° or unknown, adjust speed

to VREF +15.
With a failure of the electric trim, ver-
ify the PITCH TRIM CB is set. With The new speed is an adjusted VREF.
the CB set, continue the flight with Use this new speed in the same man-
manual elevator trim actuation. ner as the original VREF.

Citation V For SimuFlite training only 5E-17

March 1999
5E-18 For SimuFlite training only Citation V
March 1999
 Flight Controls

Flaps Data Data

Power Source LH Main bus Extension
Hydraulic system Summaries
Control Flap preselect handle
Monitor Flap indicator
HYD PRESS ON annunciator

Speedbrakes Data
Power Source LH Main bus Extension
Hydraulic system
Control Speedbrake switches/85% N2 throttle switches
Monitor SPEED BRAKE EXTEND annunciator
HYD PRESS ON annunciator

Citation V For SimuFlite training only 5E-19

March 1999
5E-20 For SimuFlite training only Citation V
March 1999
The Citation fuel system is essentially two distinct, yet iden-
tical systems: one for the left engine and one for the right. Fuel
Wing tanks store the fuel for use by the engines; each wing
holds approximately 431 U.S. gallons of usable fuel for a total System
aircraft capacity of 862 U.S. gallons, or approximately 5,814
lbs total usable fuel.
Each wing tank supplies its respective engine; crossfeed oper- Chapter 5F
ations allow both engines to receive fuel from one wing tank.
Normal operation of the system is automatic and self-sustain-
ing after engine start; the system requires no pilot action other
than monitoring.

Citation V For SimuFlite training only 5F-1

March 1999
5F-2 For SimuFlite training only Citation V
March 1999
 Fuel System

Fuel System Schematic . . . . . . . . . . . . . . . 5F-6

Table of
Fuel Storage . . . . . . . . . . . . . . . . . . . . . 5F-7
Fuel Tanks . . . . . . . . . . . . . . . . . . . . . . . 5F-7
Contents
Vent System . . . . . . . . . . . . . . . . . . . . . . 5F-7
Fuel Probes . . . . . . . . . . . . . . . . . . . . . . 5F-8
Fuel Quantity Gage . . . . . . . . . . . . . . . . . . 5F-8
Low Level Warning . . . . . . . . . . . . . . . . . . 5F-8
Fuel Flow Indication . . . . . . . . . . . . . . . . . . 5F-8
Fuel Pressure Switch . . . . . . . . . . . . . . . . . 5F-9
Filler Assembly . . . . . . . . . . . . . . . . . . . . 5F-9
Drain Valves . . . . . . . . . . . . . . . . . . . . . . 5F-9
Manual Shutoff Valve . . . . . . . . . . . . . . . . 5F-10

Fuel Distribution . . . . . . . . . . . . . . . . . . 5F-11

Motive Flow Valve . . . . . . . . . . . . . . . . . . 5F-11
Motive Flow Shutoff Valve . . . . . . . . . . . . . . 5F-11
Primary Ejector Pump . . . . . . . . . . . . . . . . 5F-11
Transfer Ejector Pumps . . . . . . . . . . . . . . . 5F-12
Electric Boost Pump . . . . . . . . . . . . . . . . . 5F-12
Boost Pump Switch . . . . . . . . . . . . . . . . . 5F-12
FUEL BOOST ON Annunciators . . . . . . . . . . 5F-13
Fuel Filters and Bypass . . . . . . . . . . . . . . . 5F-13
Fuel Firewall Shutoff Valves . . . . . . . . . . . . . 5F-13
Crossfeed Switch and Valves . . . . . . . . . . . . 5F-13
Crossfeed IN TRANSIT Light . . . . . . . . . . . . 5F-14
Normal Operations . . . . . . . . . . . . . . . . . . 5F-14
Crossfeed Operations . . . . . . . . . . . . . . . . 5F-15

Servicing and Procedures . . . . . . . . . . . . . 5F-17

Preflight Inspection . . . . . . . . . . . . . . . . . . 5F-17
Servicing . . . . . . . . . . . . . . . . . . . . . . . 5F-17
Additive Procedures . . . . . . . . . . . . . . . . . 5F-17

Citation V For SimuFlite training only 5F-3

March 1999
 Fueling Procedures . . . . . . . . . . . . . . . . . 5F-17
Defueling Procedures . . . . . . . . . . . . . . . . 5F-17
Suction Method . . . . . . . . . . . . . . . . . . . 5F-17
Force Method . . . . . . . . . . . . . . . . . . . . 5F-17
Transfer (Crossfeed) Method . . . . . . . . . . . . 5F-17
Abnormal Procedures . . . . . . . . . . . . . . . . 5F-18
Low Fuel Pressure . . . . . . . . . . . . . . . . . 5F-18
Low Fuel Quantity . . . . . . . . . . . . . . . . . 5F-18
Fuel Filter Bypass . . . . . . . . . . . . . . . . . 5F-18
Fuel Boost Pump ON . . . . . . . . . . . . . . . . 5F-18

Data Summary . . . . . . . . . . . . . . . . . . . 5F-19

Fuel System Data . . . . . . . . . . . . . . . . . . 5F-19

5F-4 For SimuFlite training only Citation V

March 1999
 Fuel System

Fuel System

Citation V For SimuFlite training only 5F-5 5F-6

March 1999
 Fuel System

TRANS PUMPS
MANUAL
PRI EJECT PUMP SHUTOFF
VALVE
X-FEED
IN TRANSIT RIGHT
WING
LEFT LH RH TANK
WING TANK OFF TANK
FUEL TANK
PROBES (6)

LH RH
BOOST ENG ENG
BOOST

FUEL LOW
LEVEL
FUEL LOW
LH RH FUEL FLTR LEVEL
BYPASS
LH RH
LH RH

X-FEED LINE FILTER

RH CROSSOVER BUS LH MAIN BUS EXT
MOTIVE FLOW S.O. VLV
LH ENGINE RH ENGINE
FIRE FIRE
PUSH PUSH FUEL BOOST
3000 O
F FW S.O. VLV LH RH
2700 F ON
O
2400 F
F 2100
F PRESS SW F
U U FUEL LOW
PRESS NORM NORM
E 1800 E
L 1500 L
LH RH
Q 1200 Q MOTIVE FLOW VLV
T T
900
Y Y
600
300 FUEL
L R 2000 O
ENG PUMP & FCU F FUEL BOOST ON
0
1800 F
F F
LBS
U 1600 U PRIMARY EJECTOR OR LH RH
E E BOOST PUMP FLOW
1400
FLOW XMTR L L
1200 TRANSFER EJECTOR FLOW
F F
L 1000 L ENGINE-DRIVEN PUMP
OIL COOLER O 800 O MOTIVE FLOW
W W
600 ENGINE-DRIVEN PUMP
FLOW DIV L 400
R HIGH PRESS FLOW
200
0
EPA RETURN
AUTOMATIC S.O. LH ENGINE LBS/HR
RH ENGINE

5F-6 For SimuFlite training only Citation V

March 1999
 Fuel System

The fuel storage system consists of one

integral wet-wing fuel tank in each
ity feed fuel through the flapper valves
and three transfer ejector pumps that Fuel Storage
wing and a fuel vent system that main- pick up fuel from the low tank areas
tains a positive internal tank pressure near the wing root forward of the
within the structural limitations of the sump. In addition to the sump, each
wing. The vent system also provides tank includes:
an escape for fuel and vapor when a ■ a vent system
temperature rise causes fuel in the tank
■ fuel quantity probes and associated
to expand.
gages and annunciators
Fuel Tanks ■ filler assembly

■ drain valves
The fuel tank in each wing includes
all the wing area forward of the rear ■ manual shutoff valve.
spar except the inboard leading edge
and the area above the main landing Vent System
gear well.
The fuel vent system in each wing
Liquid-tight ribs at each end of the ensures positive internal tank pressure
wing form the boundaries of the tank. to prevent tank cavitation and allows
Holes in the ribs and spars permit fuel excess pressure to port overboard. The
movement within the tanks. Baffle system incorporates an air scoop
plates in the outboard ribs prevent (Figure 5F-1), a vent float valve, a
rapid outboard shift of fuel during surge tank, two check valves, and a
wing-low attitudes. vent line.
A sump area in the root of each wing By design, the air scoop, float valve,
contains a boost pump, a primary ejec- and two check valves remain unblocked
tor pump, and quick drains for water during inflight ice accumulation. The
and sediment removal. The sump air scoop and float valves on the under-
ensures boost pump and ejector pump side of the outboard wingtip share a
operation for a minimum of five sec- common line to the surge tank. The
onds during negative G maneuvers (not surge tank in the outboard wing con-
exceeding -.5 G). tains an anti-siphon hole. The vent line
Two flapper check valves are on top extends from the surge tank to an open-
of the fuel sump cover that forms a ended position near the sump area. The
reservoir of fuel around the pumps. check valves connect to the vent line:
The fuel sumps are kept full by grav- one midway in the line, and the other

5F-1

Citation V For SimuFlite training only 5F-7

March 1999
 near the surge tank. The lines, tank, 3,000 lbs. The gages and probes oper-
Unusable Fuel Limitation and vent provide venting for the wing ate on 28V DC power from the LH
Fuel remaining in the fuel tanks in all normal attitudes of pitch and roll. and RH Main Bus Extension buses on
when the fuel quantity indicator the pilot’s and copilot’s CB panels.
reads zero is not usable in flight. Fuel Probes
A series of six fuel probes and a tem-
Low Level Warning
perature compensator in each wing tank The low level warning system is inde-
supply information to the vertical scale pendent of the quantity indicating sys-
fuel quantity gage in the cockpit. tem. A nylon float switch in each wing
The fuel tank probes have concentric provides a signal to its respective
metal tubes that act as plates of a FUEL LOW LEVEL annunciator. As
capacitor. The value of the capacitance fuel quantity decreases to 185 lbs or
at each probe is proportional to the less, the float switch closes the circuit
height of the dielectric medium (fuel) to its associated light. During opera-
between the plates; that is, the higher tions with minimum fuel loads (above
the fuel is at a given probe, the greater 185 lbs), it is possible for the lights to
the signal the probe produces. The illuminate momentarily if the aircraft
probes in each tank are connected in encounters turbulence, rough runway
parallel; the capacitance values of each or taxi surfaces, or during turning.
probe in a tank combine to form a total
signal from that tank. Fuel Flow Indication
A temperature compensator in the A fuel flow transmitter on the right
sump of each tank adjusts for the side of each engine is in the fuel line
dielectric change of the fuel as tem- between the oil cooler and the fuel
perature changes. The result is that the control unit. The transmitter measures
compensator corrects for fuel density. the fuel flow rate between 145 and
1,800 lbs/hr.
Fuel Quantity Gage A dual vertical tape display indicator
The fuel quantity gages (one for each on the top center of the cockpit instru-
tank) on the top center of the cockpit ment panel allows the flight crew to
instrument panel are vertical scale monitor the fuel flow of both engines
servo-type instruments. The indicator on one indicator (Figure 5F-3). The
converts the total fuel tank capacitance fuel flow indicator’s display range is
signal into a fuel weight indication for from 0 to 2,000 lbs/hr. Internal lighting
each wing tank (Figure 5F-2). The is 5V DC, while power to operate the
indicator’s display range is from 0 to gage is 28V DC through the LH and

5F-2 5F-3

5F-8 For SimuFlite training only Citation V

March 1999
 Fuel System

RH FUEL FLOW CB on their respec- the pressure switch to activate the boost
tive LH and RH Main Bus Extension pump. In a low pressure situation, the
buses (left and right CB panel). When pressure switch relay supplies power
power to the gage is removed, the ver- to energize the boost pump relay.
tical tape drops and displays an OFF The boost pump relay energizes a hold-
warning at the top of the vertical tape. ing relay that supplies power to the
The dual-channel indicator receives boost pump from a CB on the aft J-
signals from its corresponding engine box; the FUEL BOOST ON annunci-
fuel flow transmitter. ator then illuminates. As pressure rises
The indicator is disabled below 10% above 4.65 PSI, the boost pump
turbine speed, or with the throttle in remains energized even though the
CUTOFF. The fuel flow indicating pressure switch has de-activated and
system also provides electrical power removed power from the FUEL LOW
and signals to the flight management PRESS annunciator. To de-energize
system. the holding relay, turn off the boost
pump switch. Returning the switch to
Fuel Pressure NORM rearms the system.
Switch Filler Assembly
The fuel pressure switch is in a hous-
A flush-mounted fuel filler assembly
ing that includes a diaphragm mechan-
is on the upper surface of each wing
ically linked to a microswitch on the
near the outboard end for normal fuel
engine pylon firewall. The switch con-
servicing.
nects to a sensing port in the fuel line
to the engine-driven pump and pro- The assembly includes an adapter,
vides two functions. First, the switch standpipe (to allow for fuel expansion),
monitors the fuel pressure to the cap, and chain to attach the cap to the
engine-driven pump and illuminates adapter (Figure 5F-4). Flush-fitting
the FUEL LOW PRESS annunciator handles fit inside the recessed cap,
if the pressure decreases below 4.65 which has markings to indicate open
PSI (+.15, -.25 PSI). It also completes and closed positions. Remove the cap
an electrical circuit to operate the fuel by lifting the handle and rotating the
boost pump. handle counterclockwise. An optional
locking fuel filler cap is interchange-
Activation of the boost pump is semi-
able with the original. The locking cap
automatic. The boost pump switches
has a chrome cover to protect the lock
must be in NORM, and the respective
from weather.
throttle must be out of CUTOFF for

5F-4 5F-5

Citation V For SimuFlite training only 5F-9

March 1999
 Drain Valves counterclockwise locks it in the OPEN
CAUTION: Use caution when position.
draining fuel. Push straight up Six poppet-type drain valves are on
on the poppet valve to prevent the lower surface of each wing from Manual Shutoff Valve
it from locking OPEN. the gear area to the fuselage (Figure
5F-5, previous page). The semi-flush, A manually-operated ball valve attach-
externally mounted valves allow sed- es to the inboard rib of each wing. The
iment, moisture, and/or residual fuel valve shuts off the fuel supply at the
to be drained from the tanks. wing to allow maintenance on the sys-
tem downstream without defueling.
The poppet head has a phillips screw- The valve is spring-loaded in the
driver-type slot that helps control pop- OPEN position and detented in the
pet movement while draining. Pushing CLOSED position.
the poppet valve up and rotating it

5F-10 For SimuFlite training only Citation V

March 1999
 Fuel System

The fuel distribution system compris- Motive Flow Fuel

es the normal engine feed system, Shutoff Valve
which supplies fuel to the engine from
its respective tank, and the engine The motive flow shutoff valve is in the Distribution
crossfeed system, which supplies fuel dry bay area of the wing fuel system
to an engine from the opposite tank. between the motive flow valve and the
Components of the distribution sys- primary ejector. The shutoff valve is
tem include: a normally-open, electrically-operat-
■ motive flow valves
ed solenoid. The valve closes only dur-
ing crossfeed operations on the tank
■ primary ejector pumps
that is not supplying the fuel.
■ transfer ejector pumps

■ electric boost pumps and associated Primary Ejector

switches and lights Pump
■ filters and associated annunciators
The primary ejector pump in the sump
■ firewall shutoff valves and associ- area of each tank supplies high vol-
ated annunciators ume, low pressure fuel from the wing
■ crossfeed switch, crossfeed valves,
tank to the engine-driven pump.
INTRANSIT light, and motive flow High pressure fuel flows from the
shutoff valve for crossfeed operation. engine-driven pump through the
motive flow valve and motive flow
Motive Flow Valve shutoff valve to the primary ejector
pump. This pump incorporates a ven-
The motive flow valve is in the high
turi (Figure 5F-6) to produce a low-
pressure bypass line from the engine- pressure area at the pump inlet. The
driven pump; it regulates the engine- suction created draws a large volume
driven pump’s high pressure fuel out- of fuel from the sump and outputs it
put for motive flow. The motive flow at a lower pressure (approximately 30
fuel flows back to the tank to operate PSI) back to the engine-driven pump.
the primary ejector pump. Normal reg-
ulated motive flow pressure is 200 to The primary ejector pump operates
250 PSI. any time the engine is running and

MOTIVE
OUTPUT FLOW

INLET

SUCTION
5F-6

Citation V For SimuFlite training only 5F-11

March 1999
 has no moving parts; it must be com- action of the transfer ejectors is to move
pletely submerged to operate. fuel with very low pressure output.

Transfer Ejector Electric Boost

Pumps Pump
Three transfer ejector pumps in each The electric boost pump is in the sump
wing tank scavenge fuel from low area of each wing tank and provides fuel
areas of the tank and direct the fuel to pressure during engine start and cross-
the sump area to ensure a constant sup- feed operation; it also serves as a back-
ply of fuel. up for the primary ejector pump. The
The pumps are continuously operating, electric boost pump output pressure is
ejector-types that use output pressure less than the normal output of the pri-
fuel from the electric boost pump or mary ejector (approximately 25 PSI).
the primary ejector fuel pump as the Boost Pump Switch
motive flow source. The motive flow
pressure is relatively low (approxi- The FUEL BOOST switches, one for
mately 30 PSI); the result of the venturi each side, are on the lower left corner

ELECTRICAL SWITCH
OUTLET (FUEL FILTER BYPASS
PORT ANNUNCIATOR)

FUEL FILTER
MANUAL HEAD AND BYPASS
SHUTOFF
VALVE
INLET
PORT

5F-7

FUEL FILTER
ELEMENT

FILTER
BOWL

5F-8

5F-12 For SimuFlite training only Citation V

March 1999
 Fuel System

of the pilot’s instrument panel. Each can be drained by depressing a man-

switch has three positions: NORM/ ual drain valve on the bottom of the
OFF/ON (Figure 5F-7). filter.
In the OFF position, the pumps are de- The filter contains a bypass valve and
energized except during engine start an electrical switch that activates when
and crossfeed. During these phases the the filter becomes obstructed. The
pumps operate automatically through bypass valve measures the input and
separate circuitry. output pressure of the filter and opens
In the NORM position, the pumps when the pressure differential reach-
operate automatically during engine es 4.5 PSI to 5.0 PSI. The electrical
start and crossfeed as well as when fuel switch on the bypass operates the
output from the primary ejector pump FUEL FLTR BYPASS annunciator. It
is insufficient. The NORM position is an early warning system that indi-
arms the boost pumps to operate when cates an impending bypass and acti-
the fuel pressure switch senses low vates at 3.5 PSID to 4.0 PSID or 1
pressure and the throttle is out of CUT- PSID before actual filter bypass.
OFF. The switch should be in NORM
during a normal flight profile. If the Fuel Firewall
throttle is in CUTOFF, the pumps are Shutoff Valves
disabled in both OFF and NORM. The
pumps operate continuously with the The fuel firewall shutoff valves are in
switch in ON, regardless of throttle the dry bay area of the wing (Figure
position. 5F-10, following page). They are
sometimes called the motorized shut-
FUEL BOOST ON off valves since they are motor-oper-
Annunciators ated ball valves and are not on the fire-
wall. The valves operate in response
The LH/RH FUEL BOOST ON annun- to pushing the L or R ENG FIRE
ciators illuminate when the respective PUSH switchlights; they close or open
boost pump receives power. in approximately one second.
With the FUEL BOOST switch in The LH/RH F/W SHUTOFF annun-
NORM or OFF, the annunciator illu- ciators illuminate when the fuel and
minates during engine start or cross- hydraulic firewall shutoff valves to the
feed operation. The annunciator also associated engine are closed.
illuminates if low fuel pressure exists
and the switch is in NORM. Crossfeed Switch
The annunciators illuminate continu-
ously with the FUEL BOOST switch
and Valves
in the ON position. Two crossfeed valves, one in each dry
bay area, control fuel access to the
Fuel Filters crossfeed fuel line that passes through
the pressure vessel to connect to the
and Bypass feedlines of the engines. In normal
Fuel filters are in the dry bay area of operation, pressurized fuel is not pre-
the wing, downstream of the manual sent in the crossfeed line because of
shutoff valve, and prior to the fuel fire- the valves’ locations.
wall valves (Figure 5F-8). The in-line The crossfeed valves are motor-oper-
fuel filters are tee-type filters with a ated ball valves that open or close in
throw-away paper element. A sump less than 1.5 seconds in response to
area in the lower part of the filter bowl inputs from the crossfeed switch. The

Citation V For SimuFlite training only 5F-13

March 1999
 valves remain open as long as the TRANSIT light that indicates the cross-
switch is in crossfeed. feed valves are cycling from closed to
The CROSSFEED switch on the lower open or vice versa.
left-hand corner of the cockpit instru- With the switch in LH tank or RH tank,
ment panel initiates the crossfeed oper- the light illuminates immediately as the
ation by sending electrical signals to valves begin to open. The light extin-
the various components used during guishes when the valves are fully open
crossfeed operation (Figure 5F-9). (approximately 1.5 seconds).
Switch positions are LH TANK/OFF/ When crossfeed is complete, move the
RH TANK. For left tank to right engine crossfeed switch to OFF; the light illu-
crossfeed, place the switch in LH minates as soon as the valves begin to
TANK. This activates the left boost close. The light extinguishes when the
pump for crossfeeding. Simultaneously, valves are fully closed (approximate-
the crossfeed valves receive signals to ly 1.5 seconds).
open.
Three seconds after the crossfeed Normal Operations
valves open, the motive flow shutoff During normal operation, each engine
valve closes. This action shuts down receives fuel from its respective tank.
the primary ejector pump in the right During engine start, the electric boost
wing and prevents right wing ejector pump supplies fuel pressure to the
pressure from overriding left wing engine-driven fuel pump.
boost pump pressure and crossfeeding
Depressing an engine START button
in the wrong direction.
energizes the boost pump, which
moves fuel through the manual shutoff
Crossfeed valve, fuel filter, and firewall shutoff
IN TRANSIT Light valve to the engine-driven fuel pump.
Immediately above the CROSSFEED The FUEL BOOST ON annunciator
switch on the lower left corner of the illuminates. The FUEL PRESS LOW
instrument panel is the green IN annunciator extinguishes as fuel pres-
sure increases above 4.65 PSI.

FORWARD
WING SPAR

5F-9
FUEL FIREWALL
SHUTOFF VALVE

5F-10

5F-14 For SimuFlite training only Citation V

March 1999
 Fuel System

When the engine starts, an N2 RPM ground operations. This not only tests
sensor switch in the generator control the program logic and exercises the Asymmetric Fuel Limitations
unit terminates the engine start valves, but also replaces the old fuel Maximum asymmetric fuel
sequence. The boost pump is de-ener- in the crossfeed line. differential for normal operations
gized (FUEL BOOST ON annuncia- To initiate crossfeed operations, posi- is 200 lbs.
tors extinguish), and high-pressure fuel tion the CROSSFEED switch to either Maximum asymmetric fuel
(motive flow) from the engine-driven LH or RH TANK. The boost pump in differential for emergency opera-
fuel pump and the motive flow valve the tank selected automatically acti- tions is 600 lbs.
operates the primary ejector fuel pump vates (FUEL BOOST ON annuncia-
that supplies fuel to the engine. Fuel Flight characteristics were not
tor illuminates), and both crossfeed demonstrated with unbalanced
also flows to the transfer ejector valves open (IN TRANSIT light illu-
pumps; these pumps transfer fuel from fuel above 200 lbs.
minates). The IN TRANSIT light
the tank to the sump. extinguishes when the crossfeed valves
The crossfeed valves remain closed are fully open. Three seconds later, the
during normal operations. motive flow shutoff valve on the
engine receiving crossfeed fuel auto-
Crossfeed matically closes.
Operations To terminate crossfeed operations, posi-
tion the CROSSFEED switch to OFF.
Crossfeed allows both engines to feed The motive flow shutoff valve auto-
from either the LH or RH fuel tank or matically opens to re-establish motive
to allow one engine to feed from the flow to the primary ejector pump.
opposite side during single engine
operation. Routine flight does not nor- Approximately three seconds later, the
mally require crossfeed operations; crossfeed valves close (IN TRANSIT
crossfeed is normally used to balance light illuminates) and the boost pump
the fuel load during single engine oper- shuts off (FUEL BOOST ON annun-
ations or when refueling produces an ciator extinguishes). The IN TRAN-
uneven load. SIT light extinguishes when the cross-
feed valves are fully closed.
It is a good practice to test the cross-
feed system occasionally during

Citation V For SimuFlite training only 5F-15

March 1999
5F-16 For SimuFlite training only Citation V
March 1999
 Fuel System

Preflight Inspection Defueling Servicing and

During preflight inspection, check the Procedures
following: There are three methods for draining Procedures
■ fuel gages for service requirements fuel tanks: suction, force-draining, or
■ excessive fuel leakage transfer. External power is required at
some point in the process of the last two
■ fuel sample at each sump drain
methods to operate the boost pump. The
■ fuel cap closed and locked main difference among the three is the
■ fuel tank vent unobstructed. connect point for the defueling unit.
The last two methods require removal
Servicing of residual fuel through the sump
An overwing port near the outboard drains on the lower wing surface near
end of each wing provides fueling the inboard end of each wing.
access. Follow approved grounding Suction Method
procedures for the aircraft and fuel Fuel Limitation
truck during refueling. Incorporate an 1. Remove filler cap and insert suction See Quick Reference chapter for
anti-icing additive with the fuel. line. list of approved fuels.
Prolonged aircraft storage may result 2. Remove required amount of fuel
in the buildup of water that leaches out and then remove suction line.
the anti-icing additive. Excessive water 3. Replace the filler cap.
accumulation in the fuel tank sumps
may be an indication of this. Force Method
If anti-ice additive enters the eyes, 1. Remove lower engine cowl and dis-
flush with water and contact a physi- connect engine fuel supply line at
cian immediately. fuel control.
2. Attach the suction/fuel line.
Additive Procedures 3. BATT switch ON; GPU connected;
Insert the fuel nozzle into the fuel filler FUEL BOOST pump ON.
first and follow with the fuel additive
4. Remove required fuel
nozzle. Direct additive into the flowing
fuel stream and begin the flow after 5. FUEL BOOST pump OFF.
fuel. Stop additive flow before stop- 6. Remove suction/fuel line and con-
ping the fuel flow. nect fuel supply line.
Maintain additive concentration range 7. Replace engine cowl.
in accordance with instructions in the
AFM. Transfer (Crossfeed)
Method
Fueling Procedures 1. BATT switch ON; GPU connected;
1. Determine required fuel. FUEL BOOST pump NORM.
2. Ground fuel supply unit, fuel sup- 2. Throttles in CUTOFF.
ply unit to aircraft, and fuel nozzle 3. CROSSFEED switch in LH or RH
to aircraft. TANK.
3. Remove filler cap and add required 4. When FUEL LOW PRESS annun-
fuel. ciator illuminates, CROSSFEED
4. Replace filler cap and remove switch to OFF.
ground wires. 5. Disconnect GPU and turn BATT
switch to OFF.

Citation V For SimuFlite training only 5F-17

March 1999
 Abnormal can pump the last bit of fuel. Consider
crossfeeding if the fuel level is ade-
Procedures quate on the other side. Land as soon
The following section provides a brief as practical.
discussion of what happens to the fuel
Fuel Filter Bypass
system during abnormal conditions.
Illumination of the FUEL FLTR
Low Fuel Pressure BYPASS annunciator indicates a
The FUEL LOW PRESS annunciator clogged fuel filter. If the annunciator
indicates fuel pressure is low. The 5 illuminates during high altitude flight,
PSI pressure switch should have acti- ice formation across the filter may have
vated the appropriate boost pump; caused the malfunction. Consider the
check that the fuel boost pump CB is possibility of partial or total loss of both
in and the switch is ON to ensure the engines if the bypass valves clog. Check
pump receives power. the sump tank and filter quick drains
after landing for water accumulation.
Crossfeeding may be necessary; it
allows the boost pump in the opposite Fuel Boost Pump ON
tank to pressurize the entire fuel sys-
tem. The annunciator extinguishes Illumination of the FUEL BOOST ON
once the system pressurizes again. annunciator indicates the pump is
receiving power. During an engine start
Low Fuel Quantity or crossfeed operation, this illumina-
tion is normal and requires no action.
The FUEL LOW LEVEL annuncia- If the FUEL BOOST ON annunciator
tor indicates a quantity below 185 lbs illuminates in flight, the low fuel pres-
of usable fuel in the tank. Ensure the sure switch may have activated the
boost pump is on and the fuel boost boost pump. Turn the affected boost
pump CB is in to provide uninterrupt- pump ON, and then to NORMAL. If
ed fuel flow to the engines. Although the light remains on, the pressure
this is not a pressure low situation, the switch activated the system and a low
boost pump is physically lower in the pressure condition remains.
sump area than the ejector pump and

5F-18 For SimuFlite training only Citation V

March 1999
 Fuel System

Fuel Systems Data Data

Power Source Fuel
L/R wing tanks Summary
Pressure
L/R ejector pumps
L/R boost pumps
Distribution L/R engine-driven pumps to L/R engines
Control CROSSFEED switch
L/R FUEL BOOST pump switch
L/R ENG FIRE switchlights
L/R throttles
L/R ENGINE START switches
STARTER DISENGAGE switch
Monitor Annunciators
L/R FUEL BOOST ON
L/R FUEL LOW PRESS
L/R FUEL LOW LEVEL
L/R FUEL FLTR BYPASS
L/R FIREWALL SHUTOFF
Fuel quantity gage
Fuel flow gage
Crossfeed IN TRANSIT light

Citation V For SimuFlite training only 5F-19

March 1999
5F-20 For SimuFlite training only Citation V
March 1999
The Citation V utilizes an open center (on demand) hydraulic
system. Hydraulic fluid continuously circulates through the sys- Hydraulic
tem at approximately 60 PSI because of the resistance in the
lines. When pressure is required, a bypass valve closes and the System
pressure builds up to 1,500 PSI. When the pressure is no longer
needed, the bypass valve opens and returns the system to its
low pressure flow. A relief valve regulates the pressure between
1,350 and 1,500 PSI when the system is activated. Chapter 5G
An open center system requires less hydraulic fluid than a con-
stant pressure system because of reduced heat build-up. This
decreases component wear and hydraulic leakage.
Fluid used in the system is a phosphate ester base hydraulic
fluid (Skydrol/Hyjet).
Systems using hydraulic pressure include:
■ landing gear (refer to Landing Gear chapter)
■ thrust reversers (refer to Miscellaneous chapter)
■ speedbrakes (refer to Flight Controls chapter)
■ flaps (refer to Flight Controls chapter).

Citation V For SimuFlite training only 5G-1

March 1999
5G-2 For SimuFlite training only Citation V
March 1999
 Hydraulic System

Hydraulic
Hydraulic System Schematic . . . . . . . . . . . . 5G-4
Table of System
Hydraulic System Components . . . . . . . . . . . 5G-4
Hydraulic System . . . . . . . . . . . . . . . . . . 5G-5
Contents Hydraulic
Reservoir . . . . . . . . . . . . . . . . . . . . . . . 5G-5 System
Firewall Shutoff Valves . . . . . . . . . . . . . . . . 5G-5
Components
Pumps . . . . . . . . . . . . . . . . . . . . . . . . . 5G-6
Filters . . . . . . . . . . . . . . . . . . . . . . . . . 5G-6
Flow Switch Check Valves . . . . . . . . . . . . . . 5G-7
Open Center Bypass Valve and Relief Valve . . . . . 5G-7
Hydraulic Pressure Indicating System . . . . . . . . 5G-9

Preflight and Procedures . . . . . . . . . . . . . . 5G-9

Servicing and Preflight . . . . . . . . . . . . . . . . 5G-9
Abnormal Procedures . . . . . . . . . . . . . . . . . 5G-9
Low Hydraulic Fluid Level . . . . . . . . . . . . . . 5G-9
Hydraulic Flow Low . . . . . . . . . . . . . . . . . 5G-9
Hydraulic System Remains Pressurized . . . . . . 5G-9

Data Summary . . . . . . . . . . . . . . . . . . . 5G-11

Citation V For SimuFlite training only 5G-3 5G-4

March 1999
Hydraulic System Hydraulic System Components
RH CROSSOVER BUS LH MAIN BUS EXT RIGHT PRESSURE
LINE FILTER

LH ENGINE RH ENGINE HYDRAULIC

FIRE FIRE RESERVOIR THRUST REVERSER
PUSH PUSH ISOLATION VALVE

RELIEF AND RIGHT HYDRAULIC FLOW SWITCHES

BLEED VALVE FIREWALL SHUTOFF CHECK VALVES
VALVE

HYD LOW RH
LH LEVEL
ENG ENG LEFT PRESSURE
PUMP PUMP LINE FILTER

F/W RESERVOIR F/W RETURN LINE

SHUTOFF SHUTOFF FILTER

HYD FLOW
BOOT STRAP LOW FLOW SW
CHK VLV
LH RH
FILTER FILTER RIGHT
ENGINE
PUMP

CHK PRESS
VLV PRESS RET
SW
LANDING GEAR DRAIN

HYD GROUND SERV

PRESS ON
THRUST CHK VLV
REVERSERS

SPEEDBRAKES RELIEF GROUND

SYSTEM
BYPASS VALVE FILTER SERVICE
VALVE (1,350 - CONNECTIONS
(NORMALLY 1,500 PSI)
FLAPS OPEN) LEFT
ENGINE
PUMP
HYDRAULIC
LEFT HYDRAULIC SUB-PANEL
FIREWALL SHUTOFF ASSEMBLY
VALVE
SUPPLY

CIRCULATION PRESS

ACTUATION PRESS

RETURN

5G-4 For SimuFlite training only Citation V

March 1999
 Hydraulic System

The Citation V’s open center hydraulic

system was chosen for its light weight,
the system is not pressurized or the
pumps are not operating, the large area Hydraulic
low component wear, and reduced
leakage. The majority of the system
piston is spring-loaded to maintain 2.7
to 4.0 PSI on the fluid. System
components are located in the upper A plunger type sight gage attached to
level of the tailcone area. the reservoir provides a visual indica-
The system components include: tion of the reservoir fluid level during
■ reservoir preflight (Figure 5G-2). The sight gage
has three indication marks:
■ firewall shutoff valves
■ REFILL – 0.2 gallons
■ pumps
■ FULL – 0.5 gallons
■ filters
■ OVERFILL – 0.6 gallons.
■ flow switch check valves
A microswitch attached to the sight
■ open center bypass valve and relief
gage activates when the plunger drops
valve below the REFILL indication. The
■ hydraulic pressure indicating microswitch energizes the HYD LOW
system. LEVEL annunciator.
A pressure relief valve in the low pres-
Reservoir sure area of the reservoir begins to open
The hydraulic reservoir is in the tail- at 40 PSI and fully opens at 60 PSI.
cone (Figure 5G-1). The entire system With the reservoir filled to capacity,
contains only 3.9 gallons of hydraulic the relief valve opens to drain excess
fluid with the reservoir holding a max- fluid. Manual operation of the relief
imum of 0.6 gallons. valve allows system air pressure to be
bled off for maintenance purposes.
The reservoir prevents cavitation of
the engine-driven pumps by main-
taining a positive fluid pressure input
Firewall Shutoff
to the pumps. During system opera- Valves
tion, 1,500 PSI pressure is applied to a
Hydraulic firewall shutoff valves
small piston on the reservoir. The small
(Figure 5G-3) are installed between
piston attaches to a large area piston
the hydraulic reservoir and the suction
120 times larger than the small piston.
side of each engine-driven pump to cut
The resultant output pressure of the
off the supply of hydraulic fluid from
hydraulic fluid is 15 to 16 PSI. When
the reservoir. An electric motor operates

SHUTOFF
VALVE

5G-1 5G-2 5G-3

Citation V For SimuFlite training only 5G-5

March 1999
 the valves in response to pushing the The pumps are constant displacement
respective ENGINE FIRE PUSH type pumps driven by a splined shaft
switchlight. The respective LH or RH mounted on the accessory gearbox of
F/W SHUTOFF annunciator illumi- each engine.
nates when the fuel and hydraulic fire- The pumps are designed to operate
wall valves close. Once the firewall without fluid flow for indefinite peri-
valves are closed, activating the ENG ods of time before failure; this condi-
FIRE PUSH switchlight a second time tion would exist on a windmilling
reopens them. engine shut down with the ENG FIRE
A thermal relief valve incorporated PUSH switchlight.
into the hydraulic firewall shutoff valve
relieves pressures in excess of 75 PSI Filters
between the closed valve and the
Three replaceable cartridge-type fil-
engine-driven pump.
ters in the hydraulic system have a fil-
tration capability of 5 to 15 microns.
Pumps One filter on the pressure side of each
The pressure to operate the hydraulic engine-driven pump removes impuri-
system comes from two engine-driven ties from the fluid before it reaches the
pumps (one on each engine) (Figure hydraulic items. One large filter is in
5G-4). Either pump can supply enough the hydraulic return line of the reser-
pressure to operate the system. voir (Figure 5G-5). The two filters on

CHECK VALVE

REDUCER

FILTER ASSEMBLY

5G-4

FILTER ELEMENT LINE ASSEMBLY

TO GROUND
DISCONNECT

FILTER BOWL

5G-5

5G-6 For SimuFlite training only Citation V

March 1999
 Hydraulic System

the output side of the pumps have a prevent fluid flow from one engine-
three-gallon per minute (GPM) capac- driven pump to the other during engine
ity while the return line filter is rated at shutdown; they also monitor the fluid
12 GPM. flow in their respective lines. If
A bypass valve is incorporated into hydraulic fluid flow drops to less than
each filter. The valve opens when a 0.45 GPM ±0.10 GPM, the respective
pressure differential of 100 PSI is flow switch check valve closes a circuit
sensed between the input and output to illuminate the appropriate LH or RH
side of the filter. There is no indication HYD FLOW LOW annunciator. The
in the cockpit of hydraulic filter bypass. flow switch opens the circuit to the
annunciator as the pump’s output
Flow Switch increases above 1.33 GPM.

Check Valves Open Center

The output flow of each pump comes Bypass Valve and
together at a T-fitting and enters a com-
mon pressure line to the hydraulic sub- Relief Valve
systems. Downstream of the left and With the selection of one or more
right hydraulic filters and prior to the hydraulic subsystems, a bypass valve
common T fitting are two flow switch/ in the return line closes to allow sys-
check valves (Figure 5G-6). The dual tem pressure to build. Electrical power
purpose valves act as a check valve to to close the normally open bypass valve

FROM
FILTER
RIGHT FLOW SWITCH
CHECK VALVE

LEFT FLOW SWITCH

CHECK VALVE

SPECIAL FITTING
(T - FITTING)

FROM
FILTER

TO HYDRAULIC
PANEL
5G-6

Citation V For SimuFlite training only 5G-7

March 1999
 comes from the circuit breaker(s) of the Hydraulic Pressure
Thrust Reverser Limitation subsystem being activated. A relief
Maximum allowable thrust valve, which operates in parallel with Indicating System
reverser deployed time is 15 the bypass valve, begins to open at A pressure switch on the hydraulic
minutes in any 1-hour period. 1,350 PSI and is fully open at 1,500 PSI panel assembly informs the flight crew
to maintain pressure when the system is when the system is pressurized; it illu-
in operation. The bypass valve remains minates the HYD PRESS ON annun-
closed until the selected hydraulic ciator. Activation of one of the hydraulic
system completes the cycle. The bypass subsystems closes the bypass valve and
valve then opens and depressurizes the the system pressurizes. As the pressure
hydraulic system. increases above 185 PSI, the pressure
With selection of the thrust reversers, switch closes the electrical circuit to the
the same sequence occurs except the HYD PRESS ON annunciator.
system maintains pressurization while As pressure decreases to 155 ±5 PSI,
the thrust reversers remain deployed. the pressure switch opens the circuit
A 15-minute limitation on thrust and the annunciator extinguishes.
reverser use prevents the system from
overheating.

5G-8 For SimuFlite training only Citation V

March 1999
 Hydraulic System

Servicing and If both pumps fail, the speedbrakes,

Preflight and
Preflight flaps and thrust reversers are inopera-

Check the fluid level in the reservoir

tive. The landing gear must be low-
ered by emergency procedures. Procedures
during the preflight inspection by visu-
ally checking the reservoir plunger. Hydraulic System
Check all visible hydraulic lines and Remains Pressurized
components for leaks. If the HYD PRESS ON annunciator
remains illuminated after the speed- Hydraulic System Limitations
Abnormal brakes, landing gear, or flaps have been Approved Fluids:
Procedures operated, or the thrust reversers have ■ Skydrol (500 A, 500 B, 500 B-
been normally stowed, the bypass 4, 500 C, or LD-4)
There are no emergency procedures
valve has failed in the closed position.
associated with the hydraulic system. ■ Hyjet (Hyjet W, III, IV).

The following is a discussion of abnor- Pull the subsystem CBs one at a time
mal procedures for the system. Please to isolate the system that is signaling
refer to your SimuFlite Citation V the bypass valve to remain closed and
Operating Handbook for a detailed maintain pressure on the hydraulic sys- CAUTION: Skydrol hydraulic
checklist. tem. Pulling the correct CB removes fluid, when heated to approxi-
electrical power from the affected con- mately 270°, decomposes into
Low Hydraulic Fluid Level trol valve; this removes power to the acids and other products and
hydraulic open center bypass valve.
Illumination of the amber HYD LOW could cause damage to metal
The valve then opens and relieves the
LEVEL annunciator indicates fluid structure.
hydraulic pressure.
level in the hydraulic reservoir is at
0.2 gallons or less. The speedbrakes, If the system remains pressurized after
flaps, and thrust reversers may not this procedure, the bypass valve has
operate. In addition, the landing gear failed. Land as soon as practical. Reset WARNING: Release system air
may require emergency extension pro- gear, flaps, speedbrakes, and thrust
pressure before servicing the
cedures. Land as soon as practical. reversers CBs before landing.
reservoir.
In some situations, the bypass valve
Hydraulic Flow Low may fail open and pressure is not avail-
The LH/RH HYD FLOW LOW annun- able to the systems. The HYD PRESS
ciator illuminates to advise the crew that ON annunciator does not illuminate.
In these cases, land as soon as practi- NOTE: Moving the flap lever
the left and/or right hydraulic pump flow
rate is below normal (less than 0.35 cal. The speedbrakes, flaps, and thrust allows the flaps to float in a trail
GPM). During single engine operation, reversers are not available; the land- configuration.
illumination of this annunciator is nor- ing gear will not operate using normal
mal. Cycling time of hydraulically actu- procedures.
ated functions increases. NOTE: Pulling the RH or LH
THRUST REVERSER CB deac-
tivates the opposite reverser
emergency stow system.

Citation V For SimuFlite training only 5G-9

March 1999
5G-10 For SimuFlite training only Citation V
March 1999
 Hydraulic System

Hydraulic System Data Data

Power Source Hydraulic fluid in reservoir
Left and right hydraulic pumps as pressure sources Summary
Distribution Left and right engine-driven hydraulic pumps
Control Left and right ENG FIRE switchlights
Landing gear handle
Speedbrake switch
Thrust reverser levers
Flaps lever
System CBs
Monitor LH/RH HYD FLOW LOW annunciators
Flap position indicator
SPEED BRAKE EXTEND annunciator
HYD LOW LEVEL annunciator
LH/RH ENG FIRE switchlights
HYD PRESS ON annunciator
Thrust reverser lights
LH/RH F/W SHUT OFF annunciator
Protection Reservoir pressure relief valve
Bypass relief valve
Firewall valve
Thermal relief

Citation V For SimuFlite training only 5G-11

March 1999
5G-12 For SimuFlite training only Citation V
March 1999
The Citation V utilizes four methods to protect aircraft surfaces
from ice and rain. Ice and Rain
■ Engine bleed air prevents ice formation on the engine bullet
nose cone, temperature probe, engine inlet lip, stator vanes,
Protection
windshield, and a section of the inboard wing.
■ Electric heating elements protect the pitot tubes, static ports,
and the angle-of-attack probe.
Chapter 5H
■ Pneumatic deice boots remove ice from the outboard wing
sections and horizontal stabilizers; the vertical stabilizer lead-
ing edge has no protection.
■ Isopropyl alcohol provides a backup for the pilot’s wind-
shield bleed air system.

Citation V For SimuFlite training only 5H-1

March 1999
5H-2 For SimuFlite training only Citation V
March 1999
 Ice and Rain Protection

Engine Anti-Ice System Schematic . . . . . . . . . 5H-6

Table of
Surface-Ice System Schematic . . . . . . . . . . . 5H-6
Wing Anti-Ice System Schematic . . . . . . . . . . 5H-8
Contents
Wing Anti-Ice Components Schematic . . . . . . . 5H-8
Windshield Bleed Air Anti-Ice System
Schematic . . . . . . . . . . . . . . . . . . . . . 5H-10
Ice Protection . . . . . . . . . . . . . . . . . . . . 5H-11
Ice Detection . . . . . . . . . . . . . . . . . . . . . 5H-11
Bleed Air Anti-Ice Systems . . . . . . . . . . . . . . 5H-11
Engine Anti-Ice . . . . . . . . . . . . . . . . . . . 5H-11
ENGINE ANTI-ICE Switches . . . . . . . . . . . 5H-12
Inboard Wing Leading Edge Anti-Ice . . . . . . . . 5H-13
Surface Deice System . . . . . . . . . . . . . . . . 5H-14
Boot Cycle . . . . . . . . . . . . . . . . . . . . . 5H-15
SURFACE DEICE Annunciator . . . . . . . . . . 5H-15
Windshield Bleed Air Anti-Ice System . . . . . . . . 5H-16
Windshield Bleed Air Controls . . . . . . . . . . . 5H-16
Temperature Controller . . . . . . . . . . . . . . . 5H-16
Windshield Overheat Annunciator . . . . . . . . . 5H-17
Rain Removal . . . . . . . . . . . . . . . . . . . 5H-17

Miscellaneous Systems . . . . . . . . . . . . . . 5H-19

Pitot/Static Anti-Ice System . . . . . . . . . . . . . 5H-19
Pitot/Static Controls . . . . . . . . . . . . . . . . 5H-19
Pitot/Static Annunciator . . . . . . . . . . . . . . 5H-20
Alcohol Anti-Ice System . . . . . . . . . . . . . . . 5H-20
Heated Drains . . . . . . . . . . . . . . . . . . . . 5H-20

Preflight and Procedures . . . . . . . . . . . . . 5H-21

Preflight Inspection . . . . . . . . . . . . . . . . . 5H-21
Servicing . . . . . . . . . . . . . . . . . . . . . . . 5H-21
Normal Operations . . . . . . . . . . . . . . . . . . 5H-21

Citation V For SimuFlite training only 5H-3

March 1999
 Abnormal Procedures . . . . . . . . . . . . . . . . 5H-21
Engine Anti-Ice Failure . . . . . . . . . . . . . . . 5H-22
Wing Bleed Air Overheat . . . . . . . . . . . . . . 5H-22
Windshield Bleed Air Failure . . . . . . . . . . . . 5H-22
Windshield Air Overheat . . . . . . . . . . . . . . 5H-22
Pitot/Static Failure . . . . . . . . . . . . . . . . . 5H-22
AOA HTR Failure . . . . . . . . . . . . . . . . . 5H-22

Data Summaries . . . . . . . . . . . . . . . . . . 5H-23

Engine Anti-Ice Data . . . . . . . . . . . . . . . . . 5H-23
Windshield Bleed Air Data . . . . . . . . . . . . . . 5H-23
Surface Deice Data . . . . . . . . . . . . . . . . . 5H-23
Pitot/Static Deice Data . . . . . . . . . . . . . . . . 5H-24

5H-4 For SimuFlite training only Citation V

March 1999
 Ice and Rain Protection

Engine Anti-Ice
System
Surface-Ice
System

Citation V For SimuFlite training only 5H-5 5H-6

March 1999
Engine Anti-Ice System Surface-Ice System

SURFACE
AUTO
ENGINE TO PNEUMATIC O
BLEED AIR SYSTEM F
F
60% N2 THROTTLE SURFACE
SWITCH (AFTER 5 SEC) DE-ICE MANUAL

>60% N2 NAC/GEN ENGINE

T1 BLEED AIR
INLET
<60% N2

ENGINE ENG LH UPPER 23 PSI RH UPPER

ON NOSE WING BOOT PRESS WING BOOT
LH RH IGN REG
O SYS CONE ENG LH LOWER RH LOWER
OFF
F ANTI-ICE WING BOOT
F
WING BOOT
XFD XFD XFD LH RH

LOW TEMP
SW (220° F)

CONTROL VALVE CONTROL VALVE

AND EJECTOR AND EJECTOR
TO STATORS

INBOARD WING
ANTI-ICE VLV

BLEED AIR COOLING

XFEED VLV TUBE LH HORIZ RH HORIZ
BOOT PRESSURE BOOT
SWITCH
LOW TEMP 21 PSI
SWITCH (300° F)

5H-6 For SimuFlite training only Citation V

March 1999
 Ice and Rain Protection

Wing Anti-Ice
System
Wing Anti-Ice
Components

Citation V For SimuFlite training only 5H-7 5H-8

March 1999
Wing Anti-Ice System Wing Anti-Ice Components
CHECK VALVE

LEFT WING DIFFUSER LEFT WING RIGHT WING RIGHT WING DIFFUSER
TEMP SENSOR TEMP SENSOR FROM RIGHT
ENGINE W/S ANTI-ICE

CHECK VALVE
LOW TEMP SW (300° F) LOW TEMP SW (300° F)
RAM AIR IN RAM AIR IN

VACUUM EJECTOR
ANTI-ICE REGULATOR
VALVE AND SHUTOFF

FROM LEFT ENGINE

COOLING COOLING
TUBE TUBE
DEICE SERVICE AIR 23 PSI
BOOTS PRESSURE
REGULATOR
DOOR SEAL
CROSSFEED
VALVE
LH/RH RAM AIR EXHAUST
MODULATING
TEMP CONTROL ANTI-ICE AIR
TEMP SENSING LINE FLEX JOINT MODULATING TEMP
RAM -AIR SHROUD CONTROL VALVE
(COOLING DUCT)
RAM AIR EXH RAM AIR EXH
LOW TEMP
SWITCH (300° F)
ANTI-ICE TEMP OVERTEMP PROTECTION
BLEED AIR CONTROL SENSOR SWITCHES (160° F)
X-FEED VLV
TO DOOR TO SURFACE
SEAL DE-ICE BOOTS
LH WING ANTI-ICE VALVE 23 RH WING ANTI-ICE VALVE
PSI PRESS REG

ANTI-ICE
TO VAC EJECT SUPPLY LINE

OVERTEMP PROTECTION DIFFUSER

SWITCH (160° F) (PICCOLO) TUBE DEICE BOOT
LEFT ENG BLEED RIGHT ENG BLEED
LEADING EDGE
RAM-AIR ENTRY ASSEMBLY

TO WINDSHIELD ANTI-ICE
HOT BLEED AIR

TEMPERATURE-CONTROLLED
ENGINE BLEED AIR BLEED AIR

REGULATED BLEED AIR 23 PSI SERVICE AIR

COOLING AIR RAM AIR

5H-8 For SimuFlite training only Citation V

March 1999
 Ice and Rain Protection

Citation V For SimuFlite training only 5H-9

March 1999
Windshield Bleed Air Anti-Ice System

FWD PRESSURE
LH MANUAL W/S BLEED AIR BULKHEAD RH MANUAL W/S BLEED AIR
CONTROL CONTROL

LH MAIN BUS EXT RH CROSSOVER BUS

TEMP TRANS
5A W/S BLEED
AIR CONTROL
W/S BLEED 5A 5A WARN LITE 2
AIR POWER

280° F TEMPERATURE
W/S BLEED (138° C) TRANSMITTER
HI (295° F, 146° C)
TEMPERATURE TEMPERATURE
O CONTROLLER WARNING
F
F
LOW
260° F AFT PRESSURE
(127° C) BULKHEAD
OVERHEAT
295° F
(146° C)
TEMP/
PRESSURE W/S AIR
SWITCH O'HEAT
RELAY
TEMPERATURE PRESSURE
TRANSMITTER SWITCH
5 PSI

RAM AIR EXHAUST RAM AIR IN

HEAT EXCHANGER
AIR
CONTROL
VALVE
W/S BLEED AIR VALVE

LH ENG BLEED AIR RH ENG BLEED AIR

SUPPLY
REGULATED AIR
RAM AIR

5H-10 For SimuFlite training only Citation V

March 1999
 Ice and Rain Protection

Ice Detection ous engine components. Use the anti-

Ice Protection
ice systems when operating in visible
During day flight, visually note ice by
moisture with an indicated OAT
observing ice formation on the lower
between +10°C and -30°C.
corners of the windshield or on the
wing leading edge (Figure 5H-1). The bleed air flows continuously
through the engine bullet nose cone
During night flight, use the wing
and T1 probe forward of the fan in the
inspection lights to illuminate the lead-
engine inlet anytime an engine is oper-
ing edges of both wings. The WING
ating. Solenoid valves control bleed air
INSP switch in the cockpit controls
flow to the air inlet lip, inner stator
the lights (Figure 5H-2).
vanes, and inboard wing leading edges.
If ice builds up on the engine, using NOTE: Visible moisture
engine anti-ice can cause ice ingestion. Engine Anti-Ice includes, but is not limited to, the
If ice build-up occurs, climb or descend Stainless steel tubing transfers the following conditions: fog with
into a clear area. If unable to climb or bleed air from the engine to the air inlet visibility less than one mile, wet
descend, consider turning one system duct and engine stator. The aerody- snow, and rain.
on at a time. namic air inlet assembly is a welded
and riveted aluminum duct with space
Bleed Air Anti-Ice between the inner and outer skin to
Systems provide for bleed air flow. With the
system operating, bleed air travels
Engine bleed air supplies the heat nec-
through the duct and exits from the air
essary to prevent ice formation on the
inlet assembly forward of the engine
inboard wing leading edge and vari-
inlet adapter. The engine stator anti-

W/S ALCOHOL

PITOT PROBES (2) WINDSHIELD BLEED AIR

STATIC PORTS (4) RAIN REMOVAL/ANTI-ICE
ELECTRICALLY
ANTI-ICED
ANGLE-OF-ATTACK
PROBE ELECTRICALLY
INBOARD WING ANTI-ICED
BLEED AIR ANTI-ICED

OUTBOARD WING
DE-ICE BOOTS

5H-2

HORIZONTAL
STABILIZER
ENGINE INLET/ DE-ICE BOOTS
INNER STATOR
VANES BLEED
AIR ANTI-ICED

5H-1

Citation V For SimuFlite training only 5H-11

March 1999
 ice system is part of the engine instal- thermal time delay relay and then to
lation except for the electrical con- the valves. After five seconds, the ther-
nection that powers its control valve. mal relay opens to remove power from
The normally-open (electrically closed) the valves, causing them to open.
solenoid valves control the air flow During normal cruise operation with
into the air inlet lip and engine stator. power set above 60% N2 for longer
Electrical power holds the valves than five seconds, activation of the
closed until anti-ice is selected; the ENGINE ANTI-ICE de-energizes the
valves open only with the removal of solenoid valves immediately. Reducing
power and at least 4 PSIG bleed air the power below 60% N2 causes the
pressure available. The valves regu- solenoid valves to energize to the
late bleed air from 11 to 14 PSIG for closed position; this resets the thermal
anti-ice protection. time delay relay. The five-second time
delay allows the engines to accelerate
ENGINE ANTI-ICE Switches quickly during a go-around situation
LH and RH ENGINE ANTI-ICE without robbing power from the
switches on the pilot’s lower left instru- engines for anti-ice.
ment panel electrically control the LH and RH ENGINE ANTI-ICE
engine and inboard wing anti-ice sys- annunciators monitor the engine anti-
tems (Figure 5H-3). With the three- ice system. The corresponding ENGINE
position ON/OFF/XFD switch in OFF, ANTI-ICE annunciator illuminates
the wing anti-ice and engine bleed when the system is ON and any of the
valves receive power to keep them following are true:
closed. Turning the switch ON removes
■ engine inlet lip temperature is below
power from the wing anti-ice valve (see
220°F (±8°F)
Inboard Wing Leading Edge Anti-Ice,
this chapter), sends power to the corre- ■ engine stator anti-ice valve fails to

sponding engine ignition system, and open

routes power to a 60% N2 switch on the ■ wing bleed air temperature is below
respective throttle. 300°F.
With a throttle setting below 60% N2, Proper engine anti-ice protection may
power is continuously supplied to the require power settings as high as 75%
inlet lip and stator valves; this keeps N2. Allow some time for the inlet and
them closed. With power above 60% wing temperature sensors to heat after
N2, electrical power routes through a the system is on.

5H-3

5H-12 For SimuFlite training only Citation V

March 1999
 Ice and Rain Protection

Inboard Wing Leading heat shield spacer divides the assembly

Edge Anti-Ice into a forward hot section and a rear-
ward air space in front of a fuel cell
The inboard wing leading edge portion closeout spacer. In the front of the lead-
of each engine anti-ice system includes ing edge assembly, a piccolo tube, riv-
a 61-inch diffuser panel that extends eted and bonded to a liner, forms an
from the wing fairing to the inboard inner liner assembly. Riveted together,
edge of the deice boot. Control and the inner liner and the outer skin form
monitoring of the system comes from a single assembly. Air disseminated
a low temperature sensor, three high from the piccolo tube heats the lead-
temperature sensors, and a duct tem- ing edge and exhausts through two
perature sensor and modulator valve. scoops on the lower surface near the
The duct temperature sensor and mod- aft edge of the assembly.
ulator valve operate together to regulate
leading edge duct air temperature. Temperature adjustment of the left and
right anti-ice bleed air comes from con-
The anti-ice system uses engine bleed trolling the ram air flow through the
air extracted from the left and right left and right cooling tubes in the wing
bleed air clusters. The air from each root fairing area. Ram air enters the
side flows through one-way check cooling tube from a scoop in the for-
valves to the “christmas tree.” From ward wing root fairing, moves aft in
the “christmas tree,” the bleed air the tube to a modulating valve, and
routes to various systems. A tap into then exhausts overboard at the trailing
the pneumatic lines prior to the one- edge of the wing root fairing.
way check valves provides hot bleed
air for the wing anti-ice. Pneumatic air from the “christmas
tree” regulated to 23 PSI routes for-
The air travels forward to its respec- ward to the left and right pressure oper-
tive left and right inboard wing anti- ated modulating valves of the cooling
ice valve. The wing anti-ice valves are tubes. From the modulating valve,
normally open (electrically closed) plumbing routes the air forward to a
valves controlled by the correspond- temperature-controlled pneumatic
ing left and right ENGINE ANTI-ICE valve threaded into a boss in the probe
switches. Moving the three-position weld assembly of the forward wing
ON/OFF/XFD switch to ON removes root fairing.
power from the valves so they open.
If the temperature of the wing anti-ice
From the wing anti-ice valves, the bleed air is too cold at the boss, the
bleed air lines move internally along pneumatic valve opens to port 23 PSI
the left and right fuselage to a point air overboard. This decreases the pres-
near the trailing edge of the wing. At sure at the modulating valve; the spring
this position, the lines exit the fuse- loaded pressure valve closes. Blockage
lage into a cavity between the fuselage of ram air through the cooling tube
and the wing root fairing. The bleed causes the temperature of the anti-ice
air line enters a ram air shroud (cool- bleed air to increase. As the tempera-
ing tube) and runs forward to the lead- ture of the bleed air increases, the
ing edge of the wing. The bleed air line pneumatic temperature sensor closes
exits the cooling tube and enters the to block the 23 PSI air from porting
leading edge assembly. overboard. This increases the back
The leading edge assembly consists of pressure on the modulating valve,
an aluminum outer skin with com- which opens to allow ram air flow
partmental spacers bonded inside. A through the cooling tube; this cools the
anti-ice bleed air line.

Citation V For SimuFlite training only 5H-13

March 1999
 A separate boss of the probe weld In case of an engine failure, both wings
NOTE: Manual activation of the assembly contains a 300°F bleed air can be heated by the operating engine.
boots does not interrupt the auto- temperature switch. The switch illu- A crossfeed line forward of the left
matic cycle. Holding the switch in minates its respective ENG ANTI-ICE and right wing anti-ice valves connect
MANUAL inflates all the boots. annunciator when the corresponding with the individual left and right anti-
Releasing the switch allows the switch is ON and the bleed air tem- ice lines. A normally closed (electri-
boots to deflate. perature at the wing root boss has not cally opened) crossfeed valve sepa-
reached 300°F. rates the two air sources under normal
Three 160°F overtemperature switch- conditions. When an engine fails,
es on each side protect the system from move the engine anti-ice switch for
CAUTION: Do not operate the an overheat condition. One tempera- that engine to XFD. Power travels to
surface deice below -40°F. ture switch attaches to the fuel closeout the respective wing anti-ice valve, sig-
panel at the midpoint of the leading naling it to close, and to the crossfeed
edge assembly. Another switch mounts valve, which opens. With the cross-
on the fuselage adjacent to the coupling feed valve open, bleed air from the
CAUTION: In icing conditions, at the probe weld assembly. The third operating engine enters the wing anti-
a small amount of residual ice switch mounts beneath the wing root ice lines of the opposite side.
forms on unprotected areas. This fairing approximately midwing, adja-
is normal, but can cause an cent to a joint in the cooling tube. Surface Deice
increase in stall speeds. When The three switches work electrically System
any amount of ice is visible on in parallel. If any switch senses 160°F A pneumatic deice system breaks up
the wing leading edge, increase or greater, the WING O’HEAT annun- ice accumulation on the horizontal sta-
by 8 kts the VREF and VAC ciator for that side illuminates and the bilizer leading edge (Figure 5H-4) and
speeds, In addition, correct the wing anti-ice valve for that wing elec- the outboard wing leading edge (Figure
landing distance and maximum trically closes. 5H-5). Bleed air from the pneumatic
landing weight permitted by After the wing cools, the WING “christmas tree,” regulated to 23 PSI,
brake energy using the Residual O’HEAT annunciator extinguishes and flows through two control valves to
Ice Landing Correction Factors the wing anti-ice valve opens to return rubber boots bonded to the leading edge
Table (in the AFM). bleed air to the wing. Operation of the of the airfoils. The wing boots have
system on the ground in hot weather two chambers (upper and lower) that
with power above 70% N1 may cause inflate separately during normal infla-
the system to overheat due to inade- tion cycles. The system receives power
quate cooling (a normal indication for from a 5 amp SURFACE DEICE CB
the conditions). on the left CB panel. A SURFACE

5H-4

5H-5

5H-14 For SimuFlite training only Citation V

March 1999
 Ice and Rain Protection

DEICE switch and timer control the The interruption of power to the con-
system. A white SURFACE DEICE trol valves directs the bleed air over-
annunciator monitors the inflation board. In addition, the control valves
cycles. create a vacuum to evacuate the boots
The two control valves are in the tail- and lines during the interruption of
cone. One valve supplies pneumatic power to the control valves. This vac-
pressure to the lower wing boots and uum holds the boots deflated.
left horizontal stabilizer; the other sup- MANUAL causes both control valves
plies pressure to the upper wing boots to open and inflates all boots simulta-
and right horizontal stabilizer. neously. Holding the switch in MAN-
Use surface deice when ice buildup is UAL maintains boot inflation.
1/4 to 1/2 inch thick. Early activation Releasing the switch from MANUAL
of the boots may result in ice bridging allows the boots to deflate.
on the wing, rendering the boots inef- There is no reset function associated
fective. Late activation (if accumula- with the surface deice switch. The
tion is more than 1/2 inch) may not SURFACE DEICE CB is the only
clear the ice. means of resetting the boot system.
Boot Cycle SURFACE DEICE
The SURFACE DEICE switch (Figure Annunciator
5H-6) controls the deice boots. The Two pressure switches, mounted on
switch has three positions, AUTO/OFF/ the aft side of the vertical stabilizer
MANUAL; OFF is the spring-loaded forward spar below the horizontal sta-
center switch position. bilizer, monitor boot inflation. One
Placing the switch momentarily in switch is in the left horizontal stabi-
AUTO starts a timer circuit for an lizer deice line and the other is in the
18-second cycle. During the first 6 sec- right horizontal stabilizer deice line.
onds, the lower wing and left horizon- The white SURFACE DEICE annun-
tal stabilizer control valve opens and ciator illuminates when either switch
inflates the boots. Neither valve receives senses a pressure of 21 PSI or greater.
power during the 7 to 12-second cycle. During the auto deice cycle, the SUR-
During the 13 to 18-second cycle, the FACE DEICE annunciator extin-
upper wing and right horizontal stabi- guishes when the middle 6-second rest
lizer control valve receives power. cycle activates.

5H-6

Citation V For SimuFlite training only 5H-15

March 1999
 Windshield Bleed of ram air across the heat exchanger.
Windshield Bleed Air With the WINDSHIELD BLEED AIR
Limitation Air Anti-Ice System knobs, the pilot can manually manip-
■ Windshield bleed air required to ulate the nozzle assemblies and con-
The bleed air anti-ice system provides
prevent ice on windshield when windshield anti-ice under normal con- trol the amount of bleed air that flows
operating in visible moisture with ditions as well as external windshield onto the windshield (Figure 5H-7).
OAT 10°C to -30°C. defog and rain removal.
■ W/S BLEED AIR switch to LO
Windshield Bleed Air
Engine bleed air from the pneumatic Controls
(127°C) when OAT is above “christmas tree” supplies hot air for
-18°C. the system. A normally open (electri- The W/S BLEED switch has three posi-
■ W/S BLEED AIR switch to HI cally closed) solenoid control valve in tions: HI/OFF/LOW (Figure 5H-8).
(138°C) when OAT is -18°C or the bleed air line prior to the heat The WINDSHIELD BLEED AIR man-
exchanger controls the bleed air. The ual controls (Figure 5H-9) for the left
below.
W/S BLEED AIR switch on the pilot’s and right windshield nozzle assemblies
■ If W/S AIR O’HEAT annuncia-
lower left instrument panel controls are on the copilot’s lower left panel
tor (146°C) remains ON for 60 next to the OVHD and DEFOG fan
the power to the valve. The heat
seconds, close the manual valves. exchanger reduces the engine bleed air switches.
temperature for windshield use. Regardless of the position of the W/S
Once the system activates, a tempera- BLEED switch, the manual controls
ture controller senses and automati- must be open to allow bleed air to dis-
cally maintains air temperature to the perse across the exterior surface of the
windshield by regulating the amount windshield. Leave the manual valves

5H-7 5H-8

5H-9 5H-10

5H-16 For SimuFlite training only Citation V

March 1999
 Ice and Rain Protection

in OFF to block inadvertent air flow To enhance defogging when descend-

to the windshield (i.e., electrical failure ing into a warm, moist area, turn the
of the control solenoid). windshield bleed air and defog fan on
When conditions require windshield 15 minutes before descent (see
anti-icing, turn the windshield bleed Pneumatic chapter).
valves on, and position the windshield Windshield Overheat
bleed switch in LOW if the indicated
OAT is above -18°C or in HI if the
Annunciator
indicated OAT is -18°C or below. With the W/S BLEED switch in either
the HI or LOW position, illumination
Temperature Controller of the amber W/S AIR O’HEAT
The temperature controller receives annunciator advises that the bleed air
input from the W/S BLEED switch to the windshield exceeds 146°C. If
position and from two temperature the W/S BLEED switch is in the OFF
transmitters in the windshield bleed air position, a 5 PSI pressure switch in the
line. The controller drives the ram air bleed air line, downstream of the heat
control actuator to modulate the amount exchanger, monitors the system. Any
of ram air permitted to pass through pressure greater than 5 PSI with the
the heat exchanger to achieve the switch OFF illuminates the W/S AIR
required bleed air temperature. In O’HEAT annunciator and may indi-
LOW, the temperature controller reg- cate a leak in the control valve.
ulates the air temperature to 127 ±5°C.
In HI, the air is maintained at 138 ±5°C.
Rain Removal
An additional temperature transmitter The windshield bleed air anti-ice sys-
in the bleed air line automatically clos- tem also provides rain removal.
es the bleed air shutoff valve and illu- Augmenter doors mounted on the exte-
minates the W/S AIR O’HEAT annun- rior of the windshield bleed air dif-
ciator if bleed air temperature exceeds fusers change the airflow pattern to
146°C. deflect the rain and improve forward
visibility. The PULL RAIN handle
When the W/S BLEED AIR switch is under the WINDSHIELD BLEED
OFF, a time delay relay causes the tem- AIR control knobs opens the aug-
perature controller to drive the ram air menter doors (Figure 5H-10).
control actuator to the closed position.
After 60 seconds, power is removed Open the doors before selecting wind-
from the temperature controller. The shield bleed air. The WINDSHIELD
windshield bleed air valve closes when BLEED AIR knobs must be open to
the W/S BLEED AIR switch is OFF. MAX and the W/S BLEED switch
In the event of an electrical failure, the positioned to LOW for effective rain
bleed air shutoff valve fails to the open removal.
position. During low power settings, closing the
Regulate the volume of air to each copilot’s manual valve provides more
windshield with the manual controls. bleed air to the pilot’s windshield.

Citation V For SimuFlite training only 5H-17

March 1999
5H-18 For SimuFlite training only Citation V
March 1999
 Ice and Rain Protection

Pitot/Static Anti-Ice The angle-of-attack (AOA) probe

Miscellaneous
System (Figure 5H-13) or vane (Figure 5H-15)

Integral heating elements provide anti-

heater receives power any time the
PITOT & STATIC switch is ON. Systems
icing for the pitot tubes, static ports Check all of the pitot/static system com-
and angle-of-attack probe. If the pitot ponents during the exterior preflight
tube(s) and/or static port(s) become inspection.
restricted or blocked from ice forma-
tion, unreliable or complete failure of Pitot/Static Controls
flight instruments and other pitot-sta-
When the PITOT & STATIC ANTI-
tic supported systems occurs (refer to
ICE switch is ON, current sensors on
Avionics chapter for information on
the left flight compartment side con-
pitot/static system).
sole (forward of the equipment panel)
Two pitot tubes are on the forward connect to their respective pitot/static
fuselage, one on each side of the nose heater system (Figure 5H-14). The
section (Figure 5H-11). Four static sensors detect a defective heating ele-
ports, two on each side of the aircraft, ment by monitoring the drop in amper-
are just below the flight compartment age to the system. One pitot heater or
windows (Figure 5H-12). All six of both static port heaters of the respec-
the components incorporate electrical tive system (pilot’s or copilot’s) acti-
heating elements controlled by the vate the current sensor and illuminate
PITOT & STATIC ANTI-ICE switch the corresponding LH or RH P/S HTR
on the left switch panel. OFF annunciator.

5H-11 5H-12

5H-13 5H-14 5H-15

Citation V For SimuFlite training only 5H-19

March 1999
 Pitot/Static Annunciator BLEED switch) to ON actuates an
Alcohol Limitations electric pump that sprays alcohol on
■ Use TT-I-735 isopropyl alcohol The LH/RH P/S HTR OFF annuncia- the windshield (Figure 5H-17).
for windshield anti-ice. tor illuminates to advise the pilot that
the PITOT & STATIC ANTI ICE The two-quart reservoir in the nose
■ Backup alcohol system is baggage compartment provides approx-
switch is OFF or a heating element(s)
sufficient for 10 minutes. has malfunctioned. The AOA HTR imately 10 minutes of continuous
FAIL annunciator illuminates in the anti-ice protection. The alcohol anti-
event of a probe heater failure; it does ice system may be turned on or off at
not monitor the case heater. With the any time.
AOA HTR CB engaged and electrical
power applied to the aircraft, the case Heated Drains
heater receives power. Electrically heated drains prevent ice
formation and blockage of the refresh-
Alcohol Anti-Ice ment center and/or relief tubes.
System Electrical power for the drains come
from the shaver/toilet circuit through
An isopropyl alcohol-based fluid pro- the SHAVER/TOILET CB of the
vides a backup for bleed air anti-ice Right Power Distribution J Box. With
protection on the pilot’s windshield the SHAVER/TOILET CB in and
(Figure 5H-16). Positioning the W/S power applied to the aircraft, the drains
ALCOHOL switch (next to the W/S are heated.

5H-16 5H-17

5H-20 For SimuFlite training only Citation V

March 1999
 Ice and Rain Protection

Preflight Inspection cases, during takeoff, use of engine

Preflight and
anti-ice may require a reduced power
Inspect the pitot tubes and static vents
for obstructions during the preflight
inspection. Check the windshield alco-
setting, an increased V1 speed, and a
longer takeoff distance. Climb and Procedures
cruise power settings are also reduced.
hol reservoir sight gage to ensure fluid
is visible. Inspect all deice boots and On descent and approach into icing
deice boot overboard vents. conditions, keep engine turbine speed
(N2) above 75% for more efficient
Accomplish a self-test of the tempera-
bleed air anti-icing (use speedbrakes
ture monitor system during the warning
as required). Compute go-around
system check. Select the W/S TEMP
power using engine anti-ice; consider
position on the rotary test switch
runway conditions for landing field
(Figure 5H-18) and then turn the W/S
length.
BLEED AIR switch to either HI or
LOW. Verify proper system operation After landing, keep all aircraft surfaces
by observing illumination of the W/S clean for the next takeoff. During an
AIR O’HEAT annunciator. approach in heavy icing, check the
engine inlets and the top of the wings
Servicing for any ice accumulation.
Circumstances may require the pilot to Encountering wet snow or slush on
service the alcohol reservoir. Remove the runway requires checking the
the liner and reservoir filler plug. Add wheel wells for freedom from conta-
alcohol to bring the fluid level to the mination before the next flight. The
neck of the plug. Filling to above the wells may become packed with wet
sight gage provides a reserve supply to snow or slush.
perform the preflight inspection or
operational checks without replenish- Abnormal
ing the reservoir (Figure 5H-19). Procedures
The following section provides a brief
Normal Operations discussion of what happens to the ice
If operating in visible moisture during and rain protection system during
taxi with OAT of +10°C to -30°C, abnormal conditions. There are no
ensure removal of all ice or frost from emergency procedures. Please refer to
the aircraft surfaces and operate the the SimuFlite Citation V Operating
engine anti-ice systems one out of Handbook for specific procedures.
every four minutes at 65% N2. In a few

5H-18 5H-19

Citation V For SimuFlite training only 5H-21

March 1999
 Engine Anti-Ice Failure switch OFF indicates a pressure
buildup in the system. Open one of the
Continuous illumination of the amber manual control valves to relieve the
ENG ANTI-ICE annunciator indicates pressure. If the annunciator does not
bleed air is not effectively anti-icing. extinguish, leave the control valves in
Increase throttle to increase bleed air OFF to prevent heat damage to the
anti-icing. If the ENGINE ANTI-ICE windshields.
annunciator does not extinguish, leave
the icing environment. Pitot/Static Failure
Wing Bleed Air Overheat Illumination of either the LH or RH
P/S HTR OFF annunciator indicates a
With continuous illumination of the possible malfunction of a heating ele-
LH or RH WING O’HEAT annuncia- ment in the pitot/static system. Ensure
tor, decrease power on the affected side. anti-ice switches are in their proper
If the annunciator does not extinguish, position and CBs are in. Determine the
leave the icing environment and turn inoperative system, if possible.
the ENGINE ANTI-ICE switches OFF.
If the pilot’s pitot/static system fails,
Windshield Bleed Air the autopilot altitude hold function is
Failure inoperative. If both pitot/static systems
fail, the angle-of-attack vane is avail-
If windshield bleed air fails, activate able for airspeed reference. The copi-
the backup alcohol system. Leave the lot’s pitot/static system receives power
icing environment as soon as possible from the Emergency bus.
before the 10-minute alcohol supply
is gone. For an emergency altitude reference,
dump cabin pressure and use the cabin
Windshield Air Overheat altimeter for a rough determination of
aircraft altitude.
With a momentary illumination of the
W/S AIR O’HEAT annunciator, reduce AOA HTR Failure
the bleed air reaching the windshield
by adjusting the WINDSHIELD Illumination of the AOA HTR FAIL
BLEED AIR manual controls. With a annunciator indicates a probe heater
continuous illumination, turn the sys- failure. Check that the pitot and static
tem off, close the manual valves, and switch is ON. Ensure the AOA CB is
activate the alcohol backup system. in and leave the icing environment. If
the AOA probe heater fails and the
Illumination of the W/S AIR O’HEAT AOA becomes iced, the stick shaker
annunciator with the W/S BLEED may not function.

5H-22 For SimuFlite training only Citation V

March 1999
 Ice and Rain Protection

Engine Anti-Ice Data Data

Power Source Engine bleed air
LH and RH Main bus Extension Summaries
Distribution Each engine
Bleed air to inboard wing leading edge
Bleed air to compressor stator vanes
Bleed air to nose cone, T1 probe
Engine ignitors
Control LH/RH ENGINE ANTI-ICE switches
LH/RH ENG anti-ice CBs
Throttle switches: 60% N2
Monitor LH/RH ENG ANTI-ICE annunciators
WING O’HEAT annunciator
Visual scan – WING INSP light
Engine ITT/RPM
Engine ignition lights

Windshield Bleed Air Data

Power Source Bleed air from LH/RH bleed air clusters
Alcohol reservoir
LH Main bus Extension
Distribution LH/RH windshield nozzles
Control W/S BLEED switch
W/S ALCOHOL switch
Manual bleed air control valves
Rain augmenter door handles
Monitor W/S AIR O’HEAT annunciator
Bleed air noise

Surface Deice Data

Power Source Bleed air from LH/RH bleed air clusters
RH Crossover bus
Distribution Upper and lower wing boots
LH/RH horizontal stabilizer boots
Control SURFACE DEICE switch
Monitor SURFACE DEICE annunciator
Visual scan

Citation V For SimuFlite training only 5H-23

March 1999
 Pitot/Static Deice Data
Power Source LH Main bus Extension (pilot side)
Emer DC bus (copilot side)
Distribution AOA heater
LH/RH pitot tube
LH/RH static ports
Control PITOT & STATIC ANTI-ICE switch
Monitor P/S HTR OFF annunciator
AOA HTR FAIL annunciator

5H-24 For SimuFlite training only Citation V

March 1999
The Citation V landing gear system is a standard tricycle design
electrically controlled and hydraulically actuated. Each gear Landing
assembly has a single tire; the nose tire is chined to deflect
water and slush on the runway away from the aircraft’s engines. Gear
The system is engineered for design simplicity and low main-
tenance with a very small drag penalty. Systems
A back-up pneumatic blow-down system provides gear exten-
sion if the primary system fails.
The Citation V brake system has pedal-operated multiple disc
Chapter 5I
brakes on the main gear wheels; the system is mechanically
controlled and hydraulically actuated. A back-up pneumatic
system supplies pressure for braking if the primary system fails.
An anti-skid system provides maximum braking efficiency on
all runway surfaces.

Citation V For SimuFlite training only 5I-1

March 1999
5I-2 For SimuFlite training only Citation V
March 1999
 Landing Gear Systems

Landing Gear System Schematic . . . . . . . . . . 5I-6

Table of
Landing Gear System . . . . . . . . . . . . . . . . . 5I-7
Main Gear . . . . . . . . . . . . . . . . . . . . . . . 5I-7
Contents
Strut and Links . . . . . . . . . . . . . . . . . . . . 5I-7
Hydraulic Actuator . . . . . . . . . . . . . . . . . . 5I-8
Main Gear Door . . . . . . . . . . . . . . . . . . . 5I-8
Squat Switches . . . . . . . . . . . . . . . . . . . . 5I-8
Nose Gear . . . . . . . . . . . . . . . . . . . . . . . 5I-9
Nose Gear Assembly . . . . . . . . . . . . . . . . 5I-10
Nose Gear Doors . . . . . . . . . . . . . . . . . . 5I-10
Hydraulic Actuator . . . . . . . . . . . . . . . . . . 5I-10
Shimmy Damper . . . . . . . . . . . . . . . . . . 5I-10
Position and Warning System . . . . . . . . . . . . 5I-10
Visual Indications . . . . . . . . . . . . . . . . . . 5I-10
Audible Indication . . . . . . . . . . . . . . . . . . 5I-11
Supplemental Gear . . . . . . . . . . . . . . . . . . 5I-11
Operation . . . . . . . . . . . . . . . . . . . . . . . 5I-11
Landing Gear Control Handle . . . . . . . . . . . . 5I-11
Retraction . . . . . . . . . . . . . . . . . . . . . . 5I-11
Extension . . . . . . . . . . . . . . . . . . . . . . 5I-12
Emergency Extension . . . . . . . . . . . . . . . . 5I-12
Auxiliary Gear Controls . . . . . . . . . . . . . . . 5I-12
Emergency Air Bottle . . . . . . . . . . . . . . . . 5I-12
Nosewheel Steering . . . . . . . . . . . . . . . . . 5I-13
Nosewheel Spin Up System (Optional) . . . . . . . . 5I-13
Nosewheel Steering Disengage . . . . . . . . . . . 5I-14

Brake System Schematic . . . . . . . . . . . . . . 5I-16

Brake System . . . . . . . . . . . . . . . . . . . . 5I-17
Anti-Skid/Power Brake System . . . . . . . . . . . . 5I-17
Wheel Brakes . . . . . . . . . . . . . . . . . . . . . 5I-17

Citation V For SimuFlite training only 5I-3

March 1999
 Electric Hydraulic Pump . . . . . . . . . . . . . . . 5I-18
Master Brake Cylinders . . . . . . . . . . . . . . . . 5I-18
Parking Brake/Valve . . . . . . . . . . . . . . . . . 5I-18
Brake Annunciators . . . . . . . . . . . . . . . . . . 5I-19
Emergency Braking . . . . . . . . . . . . . . . . . . 5I-19

Servicing and Procedures . . . . . . . . . . . . . 5I-21

Preflight Inspection . . . . . . . . . . . . . . . . . . 5I-21
Servicing . . . . . . . . . . . . . . . . . . . . . . . 5I-21
Accumulator Preload . . . . . . . . . . . . . . . . 5I-21
Tire Inflation . . . . . . . . . . . . . . . . . . . . . 5I-21
Strut Inflation . . . . . . . . . . . . . . . . . . . . 5I-21
Abnormal Procedures . . . . . . . . . . . . . . . . . 5I-22
Landing Gear Does Not Extend . . . . . . . . . . . 5I-22
Anti-Skid System Failure . . . . . . . . . . . . . . 5I-22
Power Brake System Failure . . . . . . . . . . . . 5I-22
Wheel Brake Failure . . . . . . . . . . . . . . . . 5I-22

Data Summaries . . . . . . . . . . . . . . . . . . . 5I-23

Landing Gear Data . . . . . . . . . . . . . . . . . . 5I-23
Brakes/Anti-Skid Data . . . . . . . . . . . . . . . . 5I-23

5I-4 For SimuFlite training only Citation V

March 1999
 Landing Gear Systems

Landing Gear
System

Citation V For SimuFlite training only 5I-5 5I-6

March 1999
 Landing Gear System

LDG GEAR
EMERGENGY
GEAR UP NOSE
RELEASE LH RH
CABLE PULL
ANTI-
HORN SKID
SILENCE ON GEAR
UNLOCKED

FROM DUMP VLV

EMERG AIR (1,800 TO 2,050 PSI) HYD SYS PUSH
UPLOCK CABLES
DOWN OFF

U D GEAR
RESTRICTOR P N CONTROL
LG WARN
GEAR CONTROL
LH MAIN BUS EXT

EXTEND PRESSURE

FLUID RETURN

AIR PRESSURE

ELECTRICAL

CABLE

DOWN SW
DOWN SW

DOWN SW

RH
NOSE
LH

5I-6 For SimuFlite training only Citation V

March 1999
 Landing Gear Systems

The landing gear completes a retrac-

tion or extension cycle in less than six
returns to its original unpressurized
condition. Landing Gear
seconds. It can be extended up to 250
KIAS (VLO) and retracted up to 200 Main Gear System
KIAS (VLO). Once extended, airspeed
The main gear assembly mounted on
is limited to 292 KIAS (VLE).
support fittings in the wings consists
The control and indicating system of of a wheel and brake attached to an
the landing gear is through the landing air-oil shock strut (Figure 5I-1). Other
gear control panel on the pilot’s lower components include:
right instrument panel (Figure 5I-2).
■ torque links
The landing gear control handle, when
selected UP or DOWN, signals the ■ single wheel with a multiple disc

hydraulic system to pressurize, and brake (refer to Brakes, this chapter)

positions the landing gear control valve ■ gear door
to the corresponding position. Hydraulic ■ hydraulic actuator
fluid flows through the control valve to
the hydraulic actuators of each gear. ■ squat switches on both main gear.

The gear SAFE and UNSAFE indica-

tions are on the landing gear control Strut and Links
panel. With the gear in a safe UP or Each strut has an isolation piston with
DOWN position, the hydraulic system hydraulic fluid on one side and a nitro-

TRUNNION
MICRO SWITCH BRAKE LINE

SWIVEL FITTING
ACTUATOR
OIL FILL

DRAG BRACE
FIL
LIN ONS
GI NSTRUCTI

DOOR LINKAGE 5I-2

STRUT

SQUAT SWITCH
BRAKE LINES

TORQUE LINKS AXLE

BRAKE ASSY
5I-1

Citation V For SimuFlite training only 5I-7

March 1999
 gen pressure charge on the other side brace to hold the gear in the extend
so the piston can float within the strut position. Hydraulic pressure does not
for shock absorption during taxi, take- hold the gear in the down and locked
off, and landing (Figure 5I-3). A drag position. In the retract cycle, the
brace mounted between the aft trun- hydraulic pressure at the retract port
nion and the lower part of the strut of the actuator forces the sleeve pis-
cylinder assists in absorbing forward ton spring to release the locking ring
and aft forces on the strut. Torque links from the actuator groove. This unlocks
connect the strut cylinder to the pis- the actuator for retraction.
ton and axle assembly to hold the
wheel in alignment. Main Gear Door
With the gear in the retracted position,
Hydraulic Actuator the main gear door covers the wheel
An actuator attached directly to the well, except for a portion of the tire.
strut assembly hydraulically retracts The door hinges at the outboard edge
and extends the main gear. As the actu- of the wheel well and is secured to the
ator piston rod retracts, the gear moves strut with a linkage (Figure 5I-4).
to the retract position; as the rod When the gear retracts or extends, the
extends, the gear moves to the extend door follows.
position.
Squat Switches
With the gear in the retract position, a
mechanical uplock connects with an The squat switches connect to several
uplock roller on the main gear to hold aircraft systems to provide verification
the gear in the retract position as the of the aircraft’s ground status. The left
hydraulic pressure is removed from squat switch does not allow the gear to
the actuator. retract while the aircraft is on the
In the extend position of the actuator, ground. A mechanical compression or
a spring-loaded sleeve piston presses a decompression of the landing gear
locking ring into a groove in the actu- assembly on landing or takeoff closes or
ator rod. In the locked position, the opens the squat switch. When the left
actuator physically becomes one long wheel leaves the ground, the squat

5I-3

5I-4

5I-8 For SimuFlite training only Citation V

March 1999
 Landing Gear Systems

switch signals the landing gear control ■ stick shaker and stick shaker test
handle solenoid lock to retract. Once ■ thrust reversers.
this occurs, the movement of the control
handle can be used for retract selection. Information provided by the right squat
switch controls the following:
Information provided by the left squat
■ stick shaker test
switch controls the following:
■ thrust reversers
■ ground idle
■ anti-skid.
■ radar forced standby (ground)

■ outflow valves ground mode (OPEN

below 80% N2, TAKEOFF mode
Nose Gear
above 80% N2) The nose gear operation is similar to
■ emergency pressure (inoperative on
the main gear. In the UP position,
the ground) uplock hooks hold the nose gear in
place once hydraulic pressure is
■ bleed air ground – right engine
removed. When the gear is in the
(inoperative in flight) DOWN position, an internal actuator
■ cross generator start disable (in flight) lock holds the nose gear extended
■ flight hour meter and digital clocks
without hydraulic pressure. Torque
links connect the strut cylinder to the
■ landing gear handle solenoid and
piston and fork assembly and hold the
anti-skid wheel in alignment.

STEERING GEARS STEERING

UNIVERSAL
JOINT STEERING
ARM
SUPPORT
BRACKET HYDRAULIC
ACTUATOR
SHIMMY DAMPER

FORWARD
DOOR
LINKAGE ACTUATOR
FIL
LIN
G INS UCTIONS
TR
ATTACH FITTING

TRUNNION AFT DOOR LINKAGE

TORQUE LINKS

5I-5

Citation V For SimuFlite training only 5I-9

March 1999
 Nose Gear Assembly Hydraulic Actuator
The shock strut is an air/hydraulic The actuator piston rod on the nose gear
chamber with an isolation piston to operates opposite of the main gear. As
separate the halves. The nose gear the actuator piston rod retracts, the gear
assembly automatically centers during moves to the extend position; as the rod
retraction (Figure 5I-5, previous page). extends, it pushes the gear to the retract
Other components of the assembly position. Once down and locked,
include: hydraulic pressure is not necessary to
■ gear doors hold the nose gear in position because
of its internal downlock. Hydraulic pres-
■ hydraulic actuator
sure, however, is required to unlock the
■ shimmy damper internal downlock for retraction.
■ nosewheel steering.
Shimmy Damper
The nose gear is also used to steer and
tow the aircraft. A steering gear assem- The shimmy damper mounted on top
bly mounted on top of the strut cylin- of the strut connects to a bracket above
der provides a tow turning radius stop the trunnion. Rapid movements of the
(refer to Nosewheel Steering section, nosewheel are dampened by moving
this chapter). a piston through an oil filled cylinder.
Access to the shimmy damper is
Nose Gear Doors through an access plate in the nose
baggage compartment floor.
Three doors mechanically connected to
the nose gear follow the gear during
retraction and extension. The doors
Position and
cover the nose wheel well when the Warning System
nose gear is in the retracted position. Visual and audio indications indicate
Two doors cover the forward portion normal and abnormal landing gear
of the wheel well and a single door cov- positions. Additional aircraft protec-
ers the aft portion. The two forward tion is through a LDG GEAR CB and
doors close when the gear is fully the gear handle lock solenoid that pre-
retracted or fully extended; the aft door vents landing gear retraction with the
closes only when the gear is retracted aircraft on the ground.
(Figure 5I-6).

5I-6 5I-7

5I-10 For SimuFlite training only Citation V

March 1999
 Landing Gear Systems

Visual Indications Operation

Three green safe lights and a red Landing Gear Control
GEAR UNLOCK light are adjacent to Handle
the gear control handle on the bottom
left of the center instrument panel The control handle on the bottom left
(Figure 5I-7). The green lights, labeled of the center instrument panel selects
NOSE, LH, and RH, illuminate when landing gear retraction (UP) or exten-
the respective gear is down and locked. sion (DOWN). The handle must be
pulled out of its detent before it can be
The red GEAR UNLOCK light indi-
moved. A spring-loaded solenoid lock
cates an unsafe gear condition—either
pin holds the handle in the DOWN
a gear assembly is in transit, or it is
position and prevents inadvertent
not locked. The light illuminates when
movement of the handle on the ground.
the gear handle moves out of the UP
detent and remains illuminated until Movement of the handle sends an elec-
all three gear assemblies are down and trical signal to close the hydraulic
locked. At retraction, the light illumi- bypass valve. Simultaneously, the land-
nates when any downlock is released; ing gear control valve moves to the
it remains illuminated until all three retract or extend position as selected.
gear assemblies are up and locked. DC power for the gear handle, warning
Normal visual indications with the gear lights, warning horn, and locking sole-
down and safe is three green lights illu- noid is through the GEAR CONTROL
minated. When the gear retracts, all CB on the LH Main Extension bus. If
lights should extinguish. the solenoid fails or electrical power
is lost, the gear handle cannot be
Audible Indication moved to the UP position, However,
the handle can be moved to the
An audible warning horn sounds if the
DOWN position in flight.
airspeed is below 150 KIAS, either
throttle is retarded below approximately Retraction
70% N2, and the gear is not down and
locked. Silence the warning horn by With weight off the left squat switch,
depressing the horn silence switch next the electric solenoid lock pin under the
to the gear handle. The horn resets if gear handle retracts electrically to allow
the throttle is advanced above 70% N2. the pilot to place the landing gear han-
dle in the UP position.
The horn also sounds if the flaps extend
beyond the 15° position, and the gear is With the gear handle moved to UP, the
not down and locked. In this situation, hydraulic bypass valve closes (refer to
the horn cannot be silenced until the Hydraulics chapter) to pressurize the
flaps are retracted back to 15° or all hydraulic system, the HYD PRESS ON
three gear assemblies are down and annunciator illuminates, and the red
locked. GEAR UNLOCK warning light illu-
minates as the first gear assembly
Supplementary Gear unlocks.
The landing gear control valve electri-
A tail skid extends below the tailcone
cally positions to direct hydraulic pres-
skin. It consists of a steel weld assem-
sure to the retract side of each gear
bly attached to bulkheads. The tail skid
actuator. The internal down-locks of
prevents the tail assembly from scrap-
the actuators are hydraulically
ing the ground during over-rotation.
removed; the gears begin to retract.

Citation V For SimuFlite training only 5I-11

March 1999
 Once the landing gear fully retracts, ■ emergency air bottle
the spring-loaded uplock mechanism ■ dump valve.
holds the landing gear retracted. When
all three uplock mechanisms have trig- Auxiliary Gear Controls
gered their uplock switches, the circuit
Below the landing gear control han-
is interrupted. The hydraulic bypass
dle is a T-handle that mechanically
valve then opens to return the system to
releases all uplocks so the landing gear
an open center state, and the HYD
can free-fall into position (Figure
PRESS ON light extinguishes; the
5I-8).
landing gear control valve moves to
center, and the GEAR UNLOCK light Place the landing gear handle DOWN
extinguishes. for proper indication when the gear is
in place. Pull the red T-handle out and
Extension rotate 45° clockwise to lock the han-
With the gear handle moved to dle; this action moves three cables that
DOWN, the hydraulic bypass valve retract the mechanical uplock hooks.
closes (refer to Hydraulics chapter) to The landing gear control valve in its
increase pressure and route hydraulic unpowered center position or powered
fluid to the landing gear system; the DOWN position allows hydraulic fluid
HYD PRESS ON annunciator illumi- on the UP side of the actuators to return
nates, and the red GEAR UNLOCK to the system; this prevents hydraulic
warning light illuminates as the gear lock. The gear free falls to the down
assembly unlocks. position. It may be necessary to use the
rudder to yaw the aircraft (first in one
The landing gear control valve electri- direction and then the other) to extend
cally positions to route hydraulic fluid the main gear actuators fully. The nose
to the uplocks to release them before gear extends aft, and aerodynamic loads
directing pressure to the down side of force the gear to lock.
the gear actuators. Fluid then flows to
the extend side of the actuating cylin- Emergency Air Bottle
der. As the gear reaches its full exten-
Rotating the red T-handle also unlocks
sion, the internal downlocks of the actu-
the red collar-type blowdown knob
ator engage.
behind the T-handle. The knob con-
The gear down microswitches open the trols the emergency air bottle on the
hydraulic bypass valve to return the right side of the forward pressure bulk-
system to low pressure (HYD PRESS head. The bottle should have 1,800 to
ON light extinguishes) when all three 2,050 PSI air pressure.
gear are down and locked. The GEAR
Pulling the knob routes air pressure
UNLOCK warning light then extin-
from the emergency air bottle through
guishes, and the three green gear lights
auxiliary extension lines to the extend
illuminate.
side of the landing gear actuators. As
Emergency Extension the air bottle pressure flows to the gear
actuators, a dump valve in the landing
If the hydraulic system fails or an elec- gear hydraulic return line opens to pro-
trical malfunction exists in the land- vide an alternate path for the return
ing gear system, the gear uplocks can flow of hydraulic fluid to the reservoir.
be manually released for gear free fall. This also prevents hydraulic retraction
Three additional components are uti- of the gear once the emergency air bot-
lized in the emergency extension tle has been activated.
process:
The main gear actuators have three
■ auxiliary gear controls
separate internal chambers for retrac-

5I-12 For SimuFlite training only Citation V

March 1999
 Landing Gear Systems

tion and extension. The retraction side Nosewheel Steering

of the actuator has one chamber for CAUTION: Do not attempt flight
hydraulic fluid to retract the gear. The Nosewheel steering is actuated by with sheared bolts. Violent nose
extend side has one chamber for nor- deflection of the rudder pedals. Cables wheel shimmy may occur and
mal hydraulic extension and a sepa- connect the rudder pedals to the nose- the nose gear may not center
rate chamber for pneumatic air pres- wheel steering unit. A bellcrank on the when retracted after takeoff.
sure from the blow-down bottle. left side of the wheel well transmits
steering control inputs from the cables
The nose gear has two chambers: one
to a spring-loaded rod (bungee). The
for extension and one for retraction.
bungee inputs to a steering arm con-
Under normal conditions hydraulic
nected to a universal joint.
fluid is used in both. For an emergency
extension, the pneumatic air uses the During gear retraction, the universal
same extension chamber as the joint centers the nose gear and dis-
hydraulic system. The nose gear actu- connects the rudder pedals from the
ator requires servicing when a blow- nosewheel steering gears to allow rud-
down bottle has been activated. der movement without turning the nose
wheel in the wheel well. With nose
When the landing gear is down, the
gear extended, the universal joint con-
microswitches engage and all three
nects to the steering gears for steering
green lights illuminate. The red GEAR
control of the strut. A shimmy damper
UNLOCK light extinguishes.
interconnects with the front steering
The locking rings in the actuators hold gear to decrease steering oscillations.
the landing gear in the extended posi-
Rudder pedal steering is limited to
tion. Leave the blowdown knob in the
approximately 20° either side of cen-
extend position after actuation. Once
ter by stops on the rudder pedals. Use
the air bottle has been utilized, hydraulic
of differential thrust and brakes
operation of the gear is not possible.
increases the deflection. The bungee
Maintenance action is required after an
allows the nosewheel to turn past the
emergency extension to restore normal
limits of the control cable.
operation of the landing gear.
During towing operations, the nose
A sight gage for the emergency air bot-
wheel may be moved up to 95° either
tle is visible in the right forward bag-
side of center before engaging the stop
gage compartment for preflight inspec-
bolts on the trunnion. Turning the nose
tion (Figure 5I-9).
wheel past the bolt limit will damage

5I-9

5I-8

Citation V For SimuFlite training only 5I-13

March 1999
 the system and result in loss of nose- wheel speed and provides nosewheel
NOTE: Do not use the wheel steering capability. The flight con- RPM information to indicator lights
nosewheel spin-up system on trol lock limits the nose wheel to 60° in the cockpit.
takeoff. Nosewheel steering must and should not be used when towing. To operate the system, move the W/S
be disengaged for takeoff and/or BLEED switch to LOW or HIGH
landing from gravel runways. Nosewheel Spin Up (-18°C limitation still applies) and turn
System (Optional) the left and right WINDSHIELD
BLEED AIR valve controls to OFF.
An optional nosewheel spin-up sys- The pilot’s windshield bleed air may
tem allows the aircraft to operate from be used for anti-ice or rain removal,
gravel runways. Rotation of the nose- but nosewheel spin-up time increases.
wheel prior to touchdown reduces Position the nosewheel spin-up con-
gravel spray when the nosewheel con- trol knob on the lower right instrument
tacts the runway surface. panel to ON and maintain engine
The system uses a combination of power above 52% N2 for adequate
engine bleed air from the windshield bleed air supply.
bleed air system and ram air from a Within 90 seconds, the N/W RPM
scoop attached to the spin-up assembly MIN indicator on the upper left panel
to rotate a turbine (Figure 5I-10). The illuminates green as wheel speed
turbine attaches to the nose wheel. increases into the 1,600 to 2,400 RPM
A magnetic pickup in the turbine cover (±50 RPM) range. Maintain wheel
assembly and steel clips on the turbine speed in this range by adjusting the
convert the mechanical motion of the NOSE WHEEL SPIN UP control. If
nosewheel into AC voltage. The volt- the red N/W RPM MAX indicator illu-
age is directly proportional to the minates, an overspeed is indicated.

BLEED AIR
LINE

MAGNETIC
STRAP PICKUP

RAM AIR
SCOOP

5I-10

5I-14 For SimuFlite training only Citation V

March 1999
 Landing Gear Systems

Nosewheel Steering pin and allows the rudder to move in

response to the rudder pedals while the
Brake System
Disengage
nosewheel steering bungee holds the
Aircraft equipped for gravel runway nosewheel centered.
operation also include a nosewheel
steering disengage system. A nosewheel At low speeds, nosewheel castering
steering GRAVEL HANDLE on the loads override the bungee and allow
right side of the center pedestal engages directional control with differential
or disengages nosewheel steering. braking. Disengage the nosewheel
steering for takeoff and landings on a
In this system, the rudder pedals move gravel runway.
two concentric shafts that supply input
to the rudder and to the nosewheel After touchdown, position the NOSE
steering when a locking pin is in place. WHEEL SPIN UP control to OFF and
The GRAVEL HANDLE removes the re-engage the nosewheel steering when
the aircraft slows to a safe speed.

Citation V For SimuFlite training only 5I-15 5I-16

March 1999
 Brake System
PRESS IND (900 - 1,300 PSI)

LDG GEAR
DRAIN UP NOSE
PUMP
LH RH
ACCU CHG VLV ANTI-
R HORN SKID
REL VLV SILENCE ON GEAR
H UNLOCKED

M
A
PUSH
I FILTER
N DOWN OFF
ACCUM PRECHG
(675 PSI) CHK VLV
B L R L R
U
S PILOT'S PILOT'S COPILOT'S COPILOT'S
MAST MAST MAST CYL MAST CYL
E CYL CYL RESS WARN SW (750 PSI)
REL VLV 900 PSI
X
T
750 PSI
GEAR CONTROL

POWER EXT
WARN LIGHTS 2 BRAKE
5 LOW PRESS
RET

EXT
900 PSI
POWER BRAKE RET
WARN LIGHTS 1 5 AND
ANTI-SKID VALVE MOTOR SW
ANTI-SKID
INOP 1,300 PSI
LH MAIN BUS EXT R PARKING
L BRAKE
SERVO VALVE EMERGENCY BRAKE EMERG GEAR AND BRAKE
H VALVE L (1,800 TO 2,050 PSI)
SKID
20 CONTROL
ANTI-
M SKID
A ON
I
N
OFF
B
U GEN
S GEN
E ELECTRONIC
X CONTROL
T BOX

SUPPLY POWER BRAKE PRESSURE

LEFT WHEEL RIGHT WHEEL
RETURN MASTER CYLINDER PRESSURE

METERED PRESSURE

5I-16 For SimuFlite training only Citation V

March 1999
 Landing Gear Systems

Each main landing gear wheel has

hydraulically or pneumatically oper-
The control module averages the volt-
ages from the left and right wheels to Brake System
ated multiple disc brakes. The pneu- create a reference voltage. Any sig-
matic braking system provides a back- nificant variation between either wheel
up if hydraulic braking fails. During speed voltage and the reference voltage
normal braking, an anti-skid system produces an error signal in the control
provides maximum braking efficien- module.
cy and prevents skidding on wet, dry, The control module activates the
or icy runways. power brake and anti-skid valve, which
A parking brake system uses trapped controls the amount of braking applied
hydraulic pressure to lock the brakes. to each wheel.
The brake system receives electrical The anti-skid control module incor-
power through a 20 amp CB on the porates test circuitry that continuous- NOTE: To ensure proper brak-
left CB panel labeled SKID CON- ly monitors the anti-skid system. If a ing, it is necessary for the pilot
TROL. The brake anti-skid system and fault is detected, the ANTISKID INOP to apply maximum effort to the
the power brake hydraulic pump annunciator on the control panel illu- brake pedals throughout the
receive power from this CB which, minates. If hydraulic pressure in the braking run. When the system
when disengaged, disables the power power system drops below 750 PSI, anticipates a skid and releases
brake system. Braking is then avail- the POWER BRAKE LOW PRESS the applied brake pressure, any
able only by means of the backup annunciator illuminates. attempt by the pilot to modulate
pneumatic braking system. When the control module produces an braking can result in an interrup-
error signal, the servo valve segment of tion of the applied brake signal;
Anti-Skid/Power the power brake and anti-skid valve this may increase braking
Brake System energizes. The servo valve modulates distance significantly.
the braking effort applied by the pilot
The anti-skid system provides power to maintain the reference and wheel
assisted braking with protection against speed voltages within skid limits; this
skidding. Components are as follows: prevents skidding. When the airplane NOTE: If the anti-skid is turned
■ two wheel speed generators speed drops below approximately 12 off prior to or during taxiing, it
knots, the anti-skid function disen- must be turned on prior to
■ a power brake relay/anti-skid valve
gages. takeoff. The anti-skid self-test
■ a control module
sequence must be completed
A switch on the instrument panel
■ a reservoir
allows the pilot to select anti-skid ON when the aircraft is stationary.
■ an accumulator or OFF. When the switch is in the ON Failure to do so renders the anti-
position, the anti-skid function is oper- skid inoperative during takeoff.
■ an electrically-driven hydraulic
pump ational when the gear handle is down.
With the control switch in the OFF
■ a hydraulic filter
position, the ANTISKID INOP annun-
■ pressure and control switches ciator illuminates. In this situation, the
■ indicator lights. pilot has power braking available with-
out the anti-skid function. If the power
A wheel speed generator is bolted to
brake system fails, braking is avail-
each main gear axle. As the wheel
able only through the backup pneu-
turns, the generator creates a 36Hz sig-
matic system.
nal for each wheel revolution. These
signals are sent to the control module
as a variable frequency. The control
Wheel Brakes
module converts these signals to a DC The wheel brake assembly mounts on
voltage that is directly proportional to the landing gear axle. It is a multiple
wheel speed. disc design, consisting of three rotating

Citation V For SimuFlite training only 5I-17

March 1999
 and two stationary discs. A piston greater force has control. Depressing
housing contains multiple brake pis- the pedals forces fluid from the outlet
tons, return springs, and an inlet and ports of the master cylinders into the
bleeder port. Cavities within the brake respective brake signal inlet ports.
assembly connect the pistons together Pressure going to the brakes is pro-
to provide simultaneous brake actua- portional to the pressure exerted by the
tion with equalized brake pressure. brake pedals from the outlet side.
Brake application forces the pistons A power brake valve receives the brake
outward, pressing the stationary and fluid from the master cylinders and
rotating discs together; friction between electric pump to provide metered brake
the discs creates braking action. pressure.
Releasing the brakes opens a check
Electric Hydraulic valve to allow fluid to move freely in
Pump either direction. Two relief valves in
the hydraulic lines and the SKID CON-
Hydraulic power for the anti-skid sys- TROL CB protect the system. One
tem is from an electrically-driven relief valve is upstream of the system
hydraulic pump in the left nose of the filter and opens to reroute fluid back
airplane. The pump operates only to the pump if pressure exceeds 1,350
when the gear is down; an accumula- PSI. The second relief valve connects
tor in the system maintains system to the parking brake valve and opens
pressure when the pump is not run- at approximately 1,000 PSI to prevent
ning. The pump is controlled by a pres- damage to the system from excess pres-
sure switch that opens to turn off the sure caused by fluid expansion.
pump when the pressure approaches
1,300 PSI and closes to engage the The brakes can operate with or without
pump when the system pressure drops anti-skid protection.
to 900 PSI.
A filter in the fluid line downstream
Parking
from the pump traps any foreign parti- Brake/Valve
cles in the fluid. A check valve prevents The parking brake is part of the nor-
fluid from returning to the reservoir mal brake system and uses control-
through the pump when the pump is off. lable check valves to prevent hydraulic
A switch on the landing gear control fluid return after brakes are set.
enables the power brake system. When Parking brakes are set by depressing
the landing gear is down, the switch the toe brakes while pulling out the
closes to provide a ground for the power parking brake handle under the lower
brake hydraulic pump motor. The motor left side of the instrument panel (Figure
then runs in response to signals from 5I-11). The handle mechanically links
the motor pressure switch. When the to the parking brake valve downstream
landing gear squat switch senses the from the power brake valve in the
aircraft is on the ground, it signals the hydraulic lines. The parking brake
anti-skid electronic control box, which valve closes to trap fluid in the system.
enables the anti-skid system.
Do not set the parking brake if the
Master Brake brakes are hot. With the brakes hot and
parking brake set, cool-down time
Cylinders increases; heat transfer could cause the
Four master brake cylinders are paired parking brake thermal relief valve to
in series so if both pilots apply brakes open, or the thermal relief plugs in the
simultaneously, the one applying the wheel could melt and deflate the tire.

5I-18 For SimuFlite training only Citation V

March 1999
 Landing Gear Systems

Brake Annunciators Control of the emergency braking sys-

tem is by a lever with a red knob to CAUTION: If the power system
The ANTISKID INOP annunciator the left of the AUX GEAR CON- fails, braking is available through
illuminates when the ANTI-SKID TROL T-handle; the lever is labeled the emergency backup pneumatic
switch is in OFF to advise the pilot he EMER BRAKE PULL (Figure 5I-12). system only (refer to Emergency
has power braking but no anti-skid pro- The lever mechanically actuates the Braking, this chapter).
tection. The annunciator, electrically emergency brake valve. This valve
connected to the anti-skid control box meters air pressure through shuttle
and power brake valve, also illumi- valves to the brake assemblies in direct
nates when the landing gear control proportion to the amount of lever
handle is in DOWN and the power movement. Pulling the lever aft applies CAUTION: Do not apply nor-
braking or anti-skid systems fail. equal pressure to both main landing mal braking while using the
The POWER BRAKE LOW PRESS gear brake assemblies. Releasing the pneumatic brakes. Depressing
annunciator illuminates if hydraulic back pressure on the lever allows it to the pedals repositions the shut-
pressure in the power brake system move forward to relieve the pressure. tle valves in the brake lines to
drops below 750 PSI. The annunciator Air pressure to the brakes may be mod- open; this allows high pressure
is electrically connected to the brake ulated to provide any braking rate air from the brake housing to
system pressure warning switch and desired, but differential braking and enter the brake hydraulic reser-
resets if pressure rises above 900 PSI. anti-skid are not available. Pull the voir and possibly rupture it.
handle slowly to achieve the desired
The annunciator operates only when
rate of deceleration; pulling it rapidly
the landing gear is selected DOWN
could cause a skid. Maintain pressure
because the electric hydraulic pump
until the aircraft stops.
operates only when the gear is down. NOTE: After stopping and
With the pump off, pressure may drop The emergency air bottle, when fully clearing the runway using emer-
to the accumulator pre-charge level of charged, contains sufficient pressure gency braking, shut down the
675 PSI. for several full brake applications even engines and have the aircraft
if landing gear emergency extension towed to the ramp. There is no
Emergency Braking was used. warning in the cockpit that the
If the hydraulic brake system fails, a For the most efficient use of the sys- air bottle is depleted.
pneumatic brake system is available. tem, apply sufficient air pressure to the
The pneumatic pressure required for brakes to obtain the desired decelera-
emergency braking is in the emer- tion rate; maintain that pressure until
gency air bottle; this is the same emer- the airplane stops. When the handle is
gency air utilized in emergency gear released, residual air pressure from the
extension. brakes exhausts overboard.

5I-11 5I-12

Citation V For SimuFlite training only 5I-19

March 1999
5I-20 For SimuFlite training only Citation V
March 1999
 Landing Gear Systems

Preflight Inspection positioned in ANTI-SKID, the ANTI

Servicing and
SKID INOP annunciator illuminates.
During the exterior preflight inspec-
tion, accomplish the following checks
of the landing gear and brake system: Servicing Procedures
■ conduct visual inspection of nose Service the emergency gear and brake
gear assembly, tires for wear, and bottle with high pressure nitrogen and
tire pressure (120 ±5 PSI) refill the bottle to 2,050 PSI when the
pressure gage reads below 1,800 PSI.
■ ensure emergency air bottle pres-
Accomplish servicing through a charg-
sure is 1,800 to 2,050 PSI by check-
ing valve on the bottle behind the right
ing gage in baggage compartment
baggage compartment aft liner.
■ check the brake fluid reservoir in
the nose baggage compartment Accumulator Preload
■ check the brake fluid accumulator Brake accumulator pressure is 675 ±25
pressure. A reading in the light PSI.
green arc indicates a precharge pres-
sure of 675 PSI and the dark green Tire Inflation
arc indicates an operating pressure Main wheel tires are inflated to 130
of 900 to 1,300 PSI ±5 PSI; nosewheel tire inflation is 120
■ check main gear and doors for gen- ±5 PSI.
eral security, fluid leaks, strut exten-
sion, main gear tire pressure (130 Strut Inflation
±5 PSI). A data plate on the strut contains infor-
During the cockpit Before Start check- mation concerning the proper amount
list, test the landing gear with the rotary of visible chromed surface on the
TEST switch on the left side of the lower portion of the strut.
pilot’s left instrument panel (Figure Main gear struts are inflated to display
5I-13). When positioned in LDG one to two inches of chromed surface;
GEAR, all landing gear lights illumi- nose gear struts are inflated to display
nate if they are working properly. When five inches.

5I-13

Citation V For SimuFlite training only 5I-21

March 1999
 Abnormal tem failure. Ensure that the ANTI-
SKID switch is ON and that the anti-
Procedures skid CB is engaged.
There are no emergency procedures If the light remains illuminated, turn
associated with landing gear and the anti-skid system off and adjust
brakes. Refer to the SimuFlite Citation performance data for anti-skid off
V Operating Handbook for specific operations.
procedures.
Power Brake System
Landing Gear Failure
Does Not Extend The power brake pressure and anti-
When the red GEAR UNLOCK light skid annunciators illuminate together
remains illuminated and one or more of to indicate a brake system failure. If
the three green GEAR DOWN lights the power brake annunciator remains
fail to illuminate, hydraulic (especial- illuminated after resetting the SKID
ly if the HYD PRESS ON annunciator CONTROL CB, use emergency brak-
also illuminates) or electrical problems ing for landing.
are interfering with gear extension. Remember that anti-skid does not
Ensure that the control handle is cor- function during emergency braking.
rectly positioned and the GEAR CON- Do not depress the brake pedals while
TROL CB is engaged; extend the gear the emergency air brakes are actuat-
using the manual system. Pull the red ed; high pressure air bypasses the shut-
T-handle to release the gear uplocks. tle valve and could rupture the brake
Yaw the aircraft to achieve three green fluid reservoir if normal braking is
gear down lights. Pull the blow-down applied.
bottle knob to apply pneumatic pres-
sure to the down side of the actuators. Wheel Brake Failure
A wheel brake failure is treated the
Anti-Skid System Failure same as a power brake system failure
The ANTI SKID INOP annunciator in that the emergency braking system
illuminates to indicate an anti-skid sys- should be utilized.

5I-22 For SimuFlite training only Citation V

March 1999
 Landing Gear Systems

Landing Gear Data Data

Power Source Hydraulic system
Emergency air bottle
LH Main bus extension
Summaries
Distribution Gear actuators
Control Gear control handle
Aux gear controls
T-handle
Blowdown knob
LDG GEAR position on test switch
Landing gear horn silence
Monitor Landing gear down lights
GEAR UNLOCKED light
Landing gear horn
Emergency air press gage (preflight)

Brakes/Anti-Skid Data
Power Source Independent hydraulic system
Emergency air bottle
LH Main bus extension
Distribution Gear actuators
Control Brake pedals
ANTI-SKID switch
Parking brake handle
Emergency brake handle
ANTI-SKID position on test switch
Monitor POWER BRAKE LOW PRESS annunciator
ANTI-SKID INOP annunciator
Emergency air press gage (preflight)
Brake accumulator press gage (preflight)
Brake pedal feel

Citation V For SimuFlite training only 5I-23

March 1999
5I-24 For SimuFlite training only Citation V
March 1999
This chapter discusses the thrust reverser system, emergency
equipment on the aircraft, the master warning system, and the Miscellaneous
emergency oxygen system.
The thrust reverser system features external target-type reversers
Systems
that direct exhaust gases forward to provide deceleration force
to assist in braking.
Emergency equipment on the aircraft includes life vests, first aid
Chapter 5J
and survival kits, therapeutic oxygen, and portable fire extin-
guishers.
The master warning system provides warning of an aircraft
equipment malfunction, indication of an unsafe operational
condition that requires immediate attention, or indication of
normal system operation.
The oxygen system provides an emergency air supply to the
crew through quick-donning pressure demand masks. Passenger
masks deploy automatically from overhead compartments if
the cabin altitude exceeds 13,500 ±600 ft, or the flight crew
can manually deploy the passenger masks.

Citation V For SimuFlite training only 5J-1

March 1999
5J-2 For SimuFlite training only Citation V
March 1999
 Miscellaneous Systems

Thrust Reverser Schematics . . . . . . . . . . . . 5J-6

Table of
Thrust Reverser System . . . . . . . . . . . . . . . 5J-7
Hydraulic Actuating System . . . . . . . . . . . . . . 5J-7
Contents
Actuators . . . . . . . . . . . . . . . . . . . . . . . 5J-7
Isolation Valve . . . . . . . . . . . . . . . . . . . . 5J-8
Control Valve . . . . . . . . . . . . . . . . . . . . . 5J-8
Pressure Switch . . . . . . . . . . . . . . . . . . . 5J-8
Thrust Reverser Levers . . . . . . . . . . . . . . . . 5J-8
Throttle Feedback System . . . . . . . . . . . . . . 5J-9
Squat Switches . . . . . . . . . . . . . . . . . . . 5J-9
Deployment . . . . . . . . . . . . . . . . . . . . . 5J-10
Stowing . . . . . . . . . . . . . . . . . . . . . . . 5J-10
Emergency Stowing . . . . . . . . . . . . . . . . 5J-10
Preflight Inspection . . . . . . . . . . . . . . . . . . 5J-11
Abnormal Procedures . . . . . . . . . . . . . . . . 5J-11
ARM Light On in Flight . . . . . . . . . . . . . . . 5J-11
Single Engine Thrust Reversing . . . . . . . . . . 5J-11
Emergency Procedures . . . . . . . . . . . . . . . 5J-11
Inadvertent Deployment During Takeoff . . . . . . 5J-11
Inadvertent Deployment in Flight . . . . . . . . . . 5J-12
UNLOCK Light On in Flight . . . . . . . . . . . . . 5J-12

Data Summary . . . . . . . . . . . . . . . . . . . . 5J-13

Thrust Reverser Systems Data . . . . . . . . . . . . 5J-13

Emergency Equipment . . . . . . . . . . . . . . . 5J-15

Cabin Equipment . . . . . . . . . . . . . . . . . . . 5J-15
Cockpit Equipment . . . . . . . . . . . . . . . . . . 5J-15

Citation V For SimuFlite training only 5J-3

March 1999
 Oxygen System Schematic . . . . . . . . . . . . . 5J-16
Oxygen System . . . . . . . . . . . . . . . . . . . 5J-17
Oxygen Cylinder . . . . . . . . . . . . . . . . . . . 5J-17
Pressure Gage . . . . . . . . . . . . . . . . . . . 5J-17
Discharge Indicator . . . . . . . . . . . . . . . . . 5J-17
Protection – Shuttle Valve and Flow Fuse . . . . . . 5J-18
Control . . . . . . . . . . . . . . . . . . . . . . . . 5J-18
Passenger Oxygen . . . . . . . . . . . . . . . . . . 5J-19
Crew Oxygen Mask . . . . . . . . . . . . . . . . . 5J-19
Crew Mask Microphone . . . . . . . . . . . . . . 5J-20
Preflight Inspection . . . . . . . . . . . . . . . . . . 5J-21

Data Summary . . . . . . . . . . . . . . . . . . . . 5J-23

Oxygen Systems Data . . . . . . . . . . . . . . . . 5J-23

Warning Systems . . . . . . . . . . . . . . . . . . 5J-25

Rotary TEST Switch . . . . . . . . . . . . . . . . . 5J-25
Stick Shaker . . . . . . . . . . . . . . . . . . . . . 5J-25
Probe-type . . . . . . . . . . . . . . . . . . . . . 5J-25
Vane-type . . . . . . . . . . . . . . . . . . . . . . 5J-25

Annunciator Panels . . . . . . . . . . . . . . . . . 5J-28

Annunciator Cross Reference . . . . . . . . . . . 5J-29

5J-4 For SimuFlite training only Citation V

March 1999
 Miscellaneous Systems

Thrust Reverser
System

Citation V For SimuFlite training only 5J-5 5J-6

March 1999
 Thrust Reverser System

STOW
SOLENOID
200 PSI CONTROL VALVE DEPLOY
PRESSURE SWITCH STOW POSITION SHOWN
SPRING
HYD LOW
LEVEL
ISOLATION VALVE
HYD
PRESS ON

1,500 PSI
LH MAIN BUS EXT

DEPLOY
LH T/R

7 1/2
DEPLOY SOLENOID
RH CROSSOVER BUS RETURN

WARN LITE 1
5
RH T/R
7 1/2
RESTRICTOR
S S
ARM T EMER EMER T ARM
O BOTTLE 1 LH ENG RH ENG BOTTLE 2 O AIR
UNLOCK W ARMED FIRE FIRE ARMED W UNLOCK
PUSH PUSH PUSH PUSH GND
DEPLOY S NORM NORM S DEPLOY
W W

SQUAT SWITCH

STOWED

TO LH STOW TO LH T/R
STOW

STOW
UNLOCK LIMIT SW

DEPLOYED
RIGHT THROTTLE DEPLOYED
STOW/DEPLOY SWITCH DEPLOY POSITION SHOWN RIGHT THRUST
REVERSER

DEPLOY LIMIT SWITCH

THROTTLE LOCK SOLENOID STOWED OR

TRANSIT

5J-6 For SimuFlite training only Citation V

March 1999
 Miscellaneous Systems

External target-type thrust reversers The high pressure line (1,500 PSI)
(Figure 5J-1) use two vertical doors, routes to the thrust reverser isolation Thrust
(i.e., buckets) to direct exhaust gases and control valves in the tailcone equip-
forward for extra deceleration force ment section. A restrictor between the Reverser
during landing ground roll. The stow ports of the actuator and the stow
hydraulically actuated and electrically
controlled reversers mount to the aft
port of the control valve limits the
hydraulic return flow during the deploy
System
engine fan nozzle and stow into the cycle. The restrictor prevents the actu-
external aircraft contours to form the ator piston from overrunning the
aft portion of the engine nacelle. hydraulic supply flow at aircraft for-
The pilot activates the thrust reverser ward deploy speeds up to 120 KIAS.
system by operating thrust reverser From the aft equipment section, the
levers that mount “piggy-back” on hydraulic lines run through the pylon
engine throttle levers. The reversers to the engine nacelle.
can be deployed only when primary The control valve of the thrust revers-
throttle levers are in IDLE position and er system routes high pressure to one
the aircraft is on the ground. Landing of two lines, depending on whether the
gear squat switches activate at touch- deploy or stow sequence is activated.
down to complete the electrical circuit The low pressure or return line is
necessary for reverser deployment. always the opposite line. From the con-
The LH Main bus Extension and trol valve, the return line bypasses the
RH Crossover bus power the left and isolation valve and returns to the air-
right thrust reverser control circuits craft hydraulic system through a check
respectively. valve, which prevents surge pressures
from blocking the return line.
Hydraulic Actuating Actuators
System
Each thrust reverser has two actuators,
The aircraft hydraulic system provides one mounted on each side of an alu-
pressure to the thrust reverser actua- minum support beam on the inboard
tors. Each engine’s hydraulic pump and outboard sides of the engine
connects by two hydraulic lines to an exhaust. The actuators transform
open center bypass valve. hydraulic energy into mechanical

5J-1

Citation V For SimuFlite training only 5J-7

March 1999
 power to open and close the thrust An electrical signal from the thrust
Thrust Reverser Limitations reverser doors through a linkage sys- reverser levers actuates both the iso-
Reverse thrust power must be tem on a sliding track. An overcenter lation and control valves.
reduced to the idle reverse detent feature in the linkage locks the revers- The isolation valve is wired through
position at 60 KIAS on landing er in the stowed position. the LH/RH ENGINE FIRE PUSH
roll. Two hydraulic lines connect to ports switchlights. If either switchlight is
Maximum allowable thrust that are used for high pressure and pressed, electrical power is removed
reverser deployed time is 15 return lines, depending on the mode from the isolation valve, which closes
minutes in any one hour period. of operation selected in the cockpit. to prevent thrust reverser deployment
Thrust reversing on sod/dirt, or The control valve in the aft tailcone on the affected engine.
gravel runways is not allowed. provides inputs to the actuators through
two flex hoses. Control Valve
Engine static ground operation is
Mechanically actuated limit switches The control valve, which is down-
limited to IDLE power.
connect to the actuators to provide stream from the isolation valve, has
Use of thrust reversers is prohibit- cockpit indications. When the actua- two independently energized sole-
ed during touch and go landings. tors move, the stow limit microswitch noids: one for the deployment cycle,
The maximum demonstrated closes and the UNLOCK light illumi- the other for stowing. The solenoids
crosswind component for use of nates. Once the doors are fully open, route hydraulic pressure to one or the
thrust reversers is 20 kts. the deploy limit microswitch closes and other side of the control valve. This
the DEPLOY indicator illuminates. moves the control valve. The control
Maximum reverse thrust setting is
Normal operating pressure for the actu- valve directs hydraulic pressure to the
limited to 79% N1 for ambient
ators is 1,500 PSI with a fluid tem- actuator deploy or stow ports, as select-
temperature below -18°C.
perature range of -65°F to +225°F. ed. When one port is pressurized, the
When this high pressure is applied to other port vents the low hydraulic pres-
the actuator, an extend force of 920 sure return. Both solenoids are never
pounds is available; the stroke is 5.215 energized at the same time. The deploy
±.015 inches. solenoid can be energized only when
either squat switch indicates the air-
Isolation Valve craft is on the ground.
The isolation valve isolates the con- Pressure Switch
trol valve and the actuators from the
aircraft hydraulic system while thrust When the isolation valve opens, a pres-
reversers are not in operation by block- sure switch between the isolation and
ing off the high pressure inlet port. control valves illuminates the amber
With the thrust reverser throttle levers ARM light to indicate at least 200 PSI
selected to DEPLOY, 28V DC flows to hydraulic pressure within the thrust
the isolation valve solenoid; this opens reverser system. When the pressure
the valve to direct hydraulic pressure drops to 100 PSI, the pressure switch
to the control valve. The solenoid opens and the ARM light extinguishes.
remains energized while the system is
in the deploy mode. Thrust Reverser
With the thrust reverser lever in STOW Levers
position, the stow switches close to The piggy-back thrust reverser levers
energize the isolation valve during the on the throttles control thrust reverser
stow cycle. Upon completion of the deploy and stow functions (Figure
stow cycle, the stow switches open to 5J-2). With throttles at IDLE, the levers
remove power from the isolation valve pull up to the DEPLOY position or
solenoid. The isolation valve then clos- move down to the STOW position.
es to isolate the control valve system
from hydraulic supply pressure. When the levers are positioned, a
microswitch (either the deploy or stow

5J-8 For SimuFlite training only Citation V

March 1999
 Miscellaneous Systems

microswitch, depending upon position) of the thrust reverser. See Emergency

relays an electrical signal to the Procedures, this section. WARNING: Do not use throttle
hydraulic bypass valve, the isolation If thrust reversers are inadvertently friction or manually restrain the
valve, and the deploy or stow solenoid deployed, an automatic throttle retard- throttle levers during takeoff.
of the thrust reverser control valve to ing device brings the throttle to approx- Inadvertent thrust reverser deploy-
start the selected cycle. imately idle thrust, depending on the ment can cause a dangerous
During normal operation, ensure the amount of throttle friction applied. asymmetrical thrust condition.
engine throttles are at IDLE before After the device activates, it is possi-
activation of thrust reverser system. ble to advance the throttle lever; this
creates corresponding reverse thrust.
Throttle Feedback System
If inadvertent deployment occurs, the
A throttle feedback system ensures pilot can manually override the retrac-
engine thrust remains at an idle level tion mechanism with the throttle levers.
during inadvertent translation of the Subsequently reducing the throttle
thrust reverser system. lever to idle can cause engine flameout
The feedback system consists of a if mechanical damage results from the
positive action mechanical linkage deployment.
connected directly to the fuel flow
regulator. During the thrust reverser
Squat Switches
transition stage, a flexible cable drives Squat switches on the left and right
a spring-loaded throttle control lever main landing gear (see the Landing
cam that positions the aircraft throttle Gear chapter) prevent actuation of the
linkage to the IDLE position. This pre- thrust reversers in flight. Compression
vents the linkage from advancing the of either landing gear when the aircraft
power setting until the deploy cycle is is on the ground mechanically actu-
complete. ates the switch and allows the deploy
Normally feedback control is not need- solenoid to operate.
ed in the landing phase when engines If either landing gear squat switch fails
are at idle for thrust reverser operation. or if a reverser is in an armed condition
If a thrust reverser deploys inadvertently without actuation of either squat
during takeoff or while in flight, the switch, the MASTER WARNING
feedback system physically moves the lights illuminate.
power lever toward idle, which reduces The stow solenoid does not require
the thrust on the affected engine to the squat switch input to operate.
idle position during the translation cycle

5J-2

Citation V For SimuFlite training only 5J-9

March 1999
 Deployment ■ With the isolation valve powered, the
NOTE: The DEPLOY indicator stow solenoid actuates and the con-
should illuminate in less than 1.5 The reversers are designed for deploy-
trol valve opens to apply hydraulic
seconds after the hydraulic ment during landing ground roll only.
pressure to the stow ports of the
UNLOCK indicator illuminates. For deployment, move the throttle
reverser actuators.
levers to IDLE and the thrust reverser
An erroneous sequencing or a ■ The actuators start to extend; the
levers to the DEPLOY position to acti-
delay in the illumination of the deploy limit switch releases, and the
vate the deploy switches under the
thrust reverser indicators indicates DEPLOY light extinguishes.
throttle quadrant. During deployment
a failure in the thrust reverser sys- the following occurs. ■ When the actuators fully extend to
tem. Either or both conditions
■ The throttle lock solenoid prevents the stow position, the stow limit
require a maintenance check.
thrust reverser lever movement switch depresses; the thrust revers-
beyond idle reverse. er is fully stowed and locked.
■ An electrical signal closes the ■ The UNLOCK light extinguishes.
hydraulic open center bypass valve ■ With power removed from the iso-
to pressurize the hydraulic system; lation valve, the control valve clos-
the HYD PRESS ON annunciator es and the thrust reverser hydraulic
illuminates. system depressurizes.
■ The isolation valve opens to direct ■ The hydraulic system returns to open
hydraulic pressure to the thrust revers- center. The ARM light and HYD
er system; the ARM light illuminates PRESS ON annunciator extinguish.
when the system is pressurized.
■ If either landing gear squat switch
Emergency Stowing
actuates, the deploy solenoid on the Emergency stow switches (Figure
thrust reverser control valve is ener- 5J-4) adjacent to the reverser lights on
gized and the control valve opens the glareshield initiate an emergency
to apply hydraulic pressure to the stow sequence.
deploy port of the actuators. Moving the emergency stow switch to
■ Movement of the actuators and EMER bypasses the normal system
reversing mechanism unlocks the and provides an override command to
doors; the UNLOCK light illumi- stow by relaying an electrical signal
nates as soon as the reverser doors directly to the isolation valve, hydraulic
are unlocked and deployment begins. bypass valve, and stow solenoid of the
■ The actuators move approximately control valve.
five inches. When the thrust revers- After the reverser stows, the DEPLOY
er doors are fully deployed they con- and UNLOCK annunciators extinguish;
tact the deploy limit switch. This however, the ARM and HYD PRESS
illuminates the deploy light (Figure ON annunciators remain illuminated
5J-3) and energizes the throttle lock until the emergency stow switch is
solenoid to permit thrust reverser moved from EMER to NORM.
lever movement to increase power Power to operate the emergency stow
for reverse thrust. system comes from the opposite sys-
Stowing tem’s normal circuit breaker (i.e., the
RH TR CB feeds the left emergency
Moving the thrust reverser control stow, and the LH TR CB feeds the
lever forward to the STOW position right emergency stow.
actuates the stow switch in the throttle
quadrant; the following occurs.

5J-10 For SimuFlite training only Citation V

March 1999
 Miscellaneous Systems

Preflight Inspection Single Engine

Thrust Reversing WARNING: Do not attempt to
During preflight inspection, check the fly the aircraft if the thrust
thrust reversers for cracks, damage, and To utilize single engine reverse thrust, reversers fail the preflight test.
general security. Ensure the thrust reduce both throttles to IDLE and
reverser lock-out tools are not inserted. apply the brakes. Extend the speed-
In the cockpit, test the system with the brakes to aid in slowing the aircraft;
rotary TEST switch on the left panel. after the nosewheel is on the ground,
The thrust reverser lights and the deploy the desired thrust reverser.
MASTER WARNING lights illumi- Check illumination of the ARM,
nate with the test switch positioned to UNLOCK and DEPLOY lights for
THRU REV. proper operation, and apply reverser
power as required. Move the thrust
During taxi, deploy the thrust reversers reverser levers to IDLE REVERSE
normally. Actuate the emergency stow when the aircraft reaches 60 KIAS.
switch on both sides and observe the
thrust reversers stow. Return the piggy-
backs to stow and deactivate the emer-
Emergency
gency stow switches. Procedures
The following is a discussion of
Abnormal emergency procedures for the thrust
Procedures reverser system.
The following is a discussion of abnor- Inadvertent Deployment
mal procedures for the thrust reverser During Takeoff
system. Please refer to the SimuFlite
Citation V Operating Hand-book for If deployment occurs below V1, reject
specific procedures. the takeoff. Apply the brakes, move the
throttles to IDLE, extend the speed-
ARM Light On in Flight brakes, and deploy both thrust reversers.
The ARM light indicates the isolation If deployment occurs above V1, con-
valve is open and the hydraulic sys- tinue the takeoff and actuate the emer-
tem is pressurized. If the ARM light gency stow switch on the affected
illuminates during flight, ensure that engine. After establishing a positive
the thrust reverser levers are down and rate of climb, retract the landing gear.
that the emergency stow switch is in Do not exceed 125 KIAS until the
NORM. thrust reverser stows. At 400 ft and

5J-3 5J-4

Citation V For SimuFlite training only 5J-11

March 1999
 V2 +10, retract the flaps and accelerate. not exceed 200 KIAS. Land as soon as
Do not exceed 200 KIAS after the practical. If reverser does not stow, shut
thrust reverser stows. Land as soon as down the affected engine; perform a
practical. If the thrust reverser will not single engine approach maintaining 150
stow, shut down the engine and main- KIAS or below during the approach.
tain 150 KIAS or less during the Land as soon as possible.
approach.
UNLOCK Light On in
Inadvertent Deployment Flight
in Flight Illumination of the UNLOCK light
Check illumination of the ARM, during flight indicates the thrust revers-
UNLOCK, and DEPLOY annuncia- er is not fully stowed and the reverser
tors. Check that the throttle feedback doors are unlocked. The emergency
system has retarded the affected throt- stow switch should be immediately
tle to IDLE. Position the emergency activated. Ensure that the thrust revers-
stow switch to EMER. Reduce airspeed er levers are down. If the light does
to 125 KIAS (115 KIAS with flaps not extinguish, maintain 200 KIAS or
extended). After the reverser stows, do below. Land as soon as practical.

5J-12 For SimuFlite training only Citation V

March 1999
 Miscellaneous Systems

Thrust Reverser Systems Data Data

Power Source LH Main bus Extension
RH Crossover bus Summary
Hydraulic system
Control Thrust reverser levers
Emergency STOW SW
THRU REV position on rotary TEST switch
LH/RH ENGINE FIRE PUSH switchlights
LH/RH THRUST REVERSER CBs
Monitor Thrust reverser ARM/DEPLOY/UNLOCK indicators
HYD PRESS ON annunciator
MASTER WARNING annunciators

Citation V For SimuFlite training only 5J-13

March 1999
5J-14 For SimuFlite training only Citation V
March 1999
 Miscellaneous Systems

Citation V emergency equipment

includes life vests, optional portable
■ Therapeutic Oxygen (optional) – If
installed, a portable therapeutic oxy- Emergency
therapeutic oxygen, a first aid kit, a sur-
vival kit, and portable fire extinguishers.
gen container is on the forward side
of the right aft cabin partition. Equipment
Cabin Equipment Cockpit Equipment
Cabin emergency equipment includes: Cockpit emergency equipment may
■ Life Vests – A life vest ( Figure
include:
5J-5) is stowed in each passenger ■ Fire Extinguishers – Two portable
and crew seat pocket and in the base fire extinguishers are stowed in red
of toilet seat assembly. quick-release brackets for use inside
■ First Aid Kit – An approved first
the pressure vessel. One extin-
aid kit is normally stowed in the guisher is beneath the copilot’s seat
map case on the forward left side of and the other is in the cabin aft of
the cabin (Figure 5J-6). left rear seat (see the Fire Protection
System chapter of this manual).
■ Survival Kit – If onboard, a univer-
■ Fire Axe (optional) – A fire axe is
sal survival kit in the cabin provides
food, equipment, and four sleeping stowed in a bracket on the forward
bags in the event of a downed air- side of the left map compartment.
craft. A supplemental kit in the tail- ■ Smoke Goggles (optional) – Smoke
cone baggage compartment contains goggles are normally stowed in the
an axe, winter clothing, and snow aft pouch of each crew seat.
shoes.

5J-5 5J-6

Citation V For SimuFlite training only 5J-15

March 1999
Oxygen System

OXYGEN/SEATBELT
ALTITUDE PRESSURE SWITCH
LH MAIN DC BUS (AFT J BOX) 5 SOLENOID VALVE
(OPENS AT 13,500' ± 600',
CLOSES AT 8,000' )

TO PASSENGER DISTRIBUTION

PILOT'S FACE MASK CREW ONLY CONTROL VALVE

TO COPILOT'S FACE MASK

MANUAL
DROP VALVE

MANUAL DROP
NORMAL CREW ONLY

SHUTTLE VALVE

FLOW FUSE
CYLINDER PRESSURE GAUGE
OVERBOARD SHUTOFF
DISCHARGE INDICATOR VALVE KNOB

PRESSURE
REGULATOR

CREW NORMAL MANUAL

ONLY DROP
1,600
1,800
PSI
FILLER VALVE AND
PROTECTIVE CAP

OXYGEN CYLINDER OXYGEN CONTROL VALVE

(REAR MOUNTED)
( (

HIGH PRESSURE

LOW PRESSURE

5J-16 For SimuFlite training only Citation V

March 1999
 Miscellaneous Systems

Oxygen is available to the crew at all

times and to the passengers either man-
vicing through the filler port. The shut-
off valve, which is open, and pressure Oxygen
ually through cockpit control or auto-
matically if cabin altitude exceeds
regulator, which meters line pressure to
70 ±10 PSI, control the flow of oxy- System
13,500 ±600 ft. gen to the distribution system.
The pressurization system normally A fully charged bottle provides approx-
maintains an 8,000 ft cabin altitude; imately one hour of oxygen for the
therefore, supplemental oxygen is used crew and six passengers. To compute
primarily as an emergency oxygen duration for the actual number of peo-
supply. ple aboard, assume consumption at a
rate of 4.3 liters per minute per occu-
Oxygen Cylinder pant and a usable full bottle output of
1,750 liters (see Table 5J-A).
A 76-cubic ft oxygen cylinder (Figure
5J-7) is mounted in the tailcone, just Pressure Gage
aft of the tailcone access door.
The oxygen bottle pressure gage on
The oxygen bottle assembly contains the right instrument panel (Figure
a pressure reducing valve, shutoff 5J-8) indicates cylinder pressure when
valve, and provisions for external ser- the system is charged, regardless of

WARNING: Smoking is pro-

hibited during oxygen use. In
addition, certain fatty materials
such as oil, grease, soap, lipstick,
and lip balm are serious fire
hazards when in contact with
oxygen.
5J-7 5J-8

Available Time in Hours:Minutes

NOTE: For the table at left,
Cabin 1 2 2 2 2 2 2 cockpit masks are assumed to be
Altitude Cockpit Cockpit Cockpit Cockpit Cockpit Cockpit Cockpit at the normal setting at 20,000 ft
2 Cabin 4 Cabin 6 Cabin 8 Cabin 10 Cabin cabin altitude with a respiratory
8,000 17:20 8:40 2:47 1:39 1:11 0:55 0:45 rate of 10 liters per minute (body
10,000 18:18 9:09 2:50 1:40 1:11 0:55 0:45 temperature pressure saturated)
and at 100% setting at and above
15,000 20:35 10:17 2:56 1:42 1:12 0:56 0:45
25,000 ft.
20,000 23:32 11:46 3:03 1:45 1:13 0:56 0:46
25,000 10:59 5:44 2:21 1:29 1:05 0:52 0:42
30,000 14:19 7:09
35,000 19:22 9:41
37,000 23:32 11:46
Table 5J-A; Oxygen Supply Chart

Citation V For SimuFlite training only 5J-17

March 1999
 the position of the shutoff valve on the A completely discharged oxygen cylin-
cylinder. Markings on the gage indi- der must be removed and overhauled
cate the following pressure: as if it were a damaged cylinder.
■ Yellow – 0 to 400 PSI

■ Green – 1,600 to 1,800 PSI

Protection – Shuttle
■ Red – 2,000 PSI
Valve and Flow Fuse
■ Blowout Relief – 2,500 PSI. Two separate lines carry oxygen from
the tailcone-mounted cylinder to the
A fully serviced system should read cockpit. To protect the system against
in the green arc between 1,600 and oxygen loss if one line develops a leak,
1,800 PSI. Have the system serviced a flow fuse at the aft intersection of the
when the gage indicates out of the two lines senses a too-high rate of flow
green arc or pressure drops below 400 (i.e., an overboard leak) and closes.
PSI. If the bottle depletes to empty, it This isolates the affected line and
must be purged. allows the oxygen to continue in the
Discharge Indicator opposite line.
As the oxygen flows past a shuttle valve
An overpressure vent line connects
at the forward intersection of the two
with a flush-mounted green disc on the
lines, pressure from the operating line
lower left side of the aft fuselage. If
closes the shuttle valve to the affected
cylinder pressure exceeds 2,500 PSI,
line and prevents oxygen from escaping
the disc ruptures and the oxygen
overboard from that point as well.
discharges overboard through the vent
line. There is no cockpit indication of either
the valve or fuse actuation; mainte-
A missing green disc (Figure 5J-9)
nance personnel should inspect the lines
indicates the oxygen cylinder has dis-
periodically.
charged due to over-temperature/ pres-
sure. If this occurs, the cylinder must
be removed and overhauled by an
Control
FAA-approved overhaul station. The OXYGEN CONTROL VALVE on
Do not allow the cylinder to become the left console regulates flow of oxy-
completely discharged because cont- gen to the passenger compartment. The
aminants may be drawn into the regu- oxygen control valve selector has three
lator by negative pressures in the cylin- positions: CREW ONLY/NORMAL/
der caused by temperature variations. MANUAL DROP (Figure 5J-10).

5J-9 5J-10

5J-18 For SimuFlite training only Citation V

March 1999
 Miscellaneous Systems

NORMAL directs oxygen to both the each passenger seat. The masks deploy
cockpit and passenger compartment. automatically or manually, depending WARNING: The passenger
With the OXYGEN selector in NOR- on the oxygen selector switch posi- oxygen system is not suitable for
MAL, low pressure oxygen (70 PSI) tion. After the mask drops, the pas- continuous operation above
is available to a solenoid valve on the senger must pull the lanyard to release 25,000 ft. The crew oxygen sys-
oxygen selector and to crewmembers oxygen flow. tem is not suitable for continuous
through side console outlets. The sole- With the oxygen selector in the NOR- operation above 37,000 ft.
noid valve normally is spring-loaded MAL position, when the cabin returns
to block oxygen flow to the passenger to 8,000 ft cabin altitude, the solenoid
system. In the event of cabin decom- valve on the control valve closes and
pression (when the cabin altitude passenger oxygen automatically shuts
exceeds 13,500 ±600 ft), an altitude- NOTE: Use of the EMER posi-
off. If required, passengers can con-
sensing pressure switch opens the sole- tion on the crew oxygen mask
tinue to receive oxygen below 8,000
noid valve to release oxygen to the pas- ft cabin altitude if the crew selects quickly diminishes the oxygen
senger distribution system. MANUAL DROP. If the masks were supply.
If electric power is not available, the manually dropped, passenger oxygen
solenoid valve cannot be opened; in continues to flow until the crew selects
this case, select MANUAL DROP. CREW ONLY.
CREW ONLY directs oxygen to the
cockpit only (at the pilots’ side con- Crew Oxygen Mask
sole oxygen mask outlets). Select A quick-donning, sweep-on flight crew
CREW ONLY prior to landing at field mask (Figure 5J-12) stows beside each
elevations above 12,000 ft to prevent pilot’s seat and has a quick-disconnect
inadvertent deployment of passenger oxygen line that plugs into a connec-
oxygen masks. tor on each side console. A three-posi-
MANUAL DROP deploys passenger tion regulator control/ selector knob
oxygen masks. If the passenger masks (EMER/100%/NORM) on the mask
fail to deploy automatically, if the elec- provides the following selections.
trical system fails, or if passenger oxy- ■ With EMER selected on the regula-
gen is required below a cabin altitude tor control knob, the regulator sup-
of 13,500 ±600 ft (e.g., smoke, fumes), plies 100% oxygen at 2 to 4 inches
select MANUAL DROP to route oxy- H2O positive to the mask (pressure
gen through the manual control valve. breathing). Select EMER when using
mask for smoke protection or when
Passenger Oxygen increased oxygen flow is desired.
Passenger masks are stowed in over- ■ In the 100% position, the regulator
head containers (Figure 5J-11) above supplies 100% oxygen on demand,

5J-11 5J-12

Citation V For SimuFlite training only 5J-19

March 1999
 regardless of altitude. Leave the selec- OXY MASK/MIC HEADSET) for
tor in the 100% position so that the audio transmission and reception.
mask is always ready for emergency ■ Each crewmember has an indepen-
use at high altitudes and for smoke dently operated audio control panel
protection. that includes pushbutton selections,
In the NORM position, the regulator or a rotary switch for PASS SPKR
automatically supplies an air/oxygen (passenger speaker), COMM 1,
mixture (diluter demand). Selecting COMM 2, or EMER/COMM 1 for
NORM for oxygen use below 20,000 transmission only (see the Avionics
ft extends the oxygen duration. chapter for a description of the audio
An oxygen pressure detector in the control panel).
oxygen mask hose indicates a green ■ Each control wheel has a push-to-
band in the transparent portion of the talk MIC/OFF/INPH switch (Figure
supply line to verify oxygen pressure. 5J-14) to select either external radio
If the band is red, no oxygen pressure transmission or interphone use via
is present. the oxygen mask microphone.
Crew Mask Microphone To use the intercom (via overhead
cockpit speakers) between crewmem-
The crew mask has a microphone bers, select MIC OXY MASK on the
(Figure 5J-13) for internal communi- side console, then press the push-to-
cation and external radio transmission. talk switch to the INPH position; the
To transmit during oxygen use, the MIC OXY mode overrides the spring
crew uses the following controls. action to give the pilots an open micro-
■ A two-position microphone selec- phone until the selection is manually
tor switch on each crewmember’s changed. Transmission through one
side console independently selects crewmember’s oxygen mask micro-
either the oxygen mask microphone phone is heard through the other
or headset microphone (i.e., MIC crewmember’s overhead speaker.

5J-13 5J-14

5J-20 For SimuFlite training only Citation V

March 1999
 Miscellaneous Systems

During external radio communication, Preflight Inspection

the intercom is inoperative. On air-
craft unit 0001 through 0031 not During the exterior preflight inspec-
modified by SB560-23-05, with tion, check the oxygen bottle over-
EMER/COMM 1 selected, the over- pressure port to ensure that the green
head cockpit speakers are inoperative. disc has not ruptured. Verify that the
oxygen gage indicates between 1,600
To transmit over passenger speakers
to 1,800 PSI.
during oxygen usage, select MIC OXY
MASK, INPH, and PASS SPKR; in Prior to flight, instruct passengers on
this mode, external radio transmission the use of passenger oxygen masks.
is inoperative.

Citation V For SimuFlite training only 5J-21

March 1999
5J-22 For SimuFlite training only Citation V
March 1999
 Miscellaneous Systems

Oxygen Systems Data Data

Power Source
Control
LH Main DC bus
Oxygen pressure regulator
Summary
Oxygen control valve
Passenger oxygen mask switch
Microphone oxygen mask selector switch
Oxygen selector on mask
Shutoff valve on oxygen cylinder
Altitude pressure switch/solenoid valve
Monitor Oxygen pressure gage
Overboard discharge indicator
Visual scan of passengers/masks
Crew oxygen flow indicator
CAB ALT 10,000 FT annunciator
Pressurization system annunciators/indicators

Citation V For SimuFlite training only 5J-23

March 1999
5J-24 For SimuFlite training only Citation V
March 1999
 Miscellaneous Systems

The MASTER WARNING lights and Rotary TEST Switch Warning

annunciator panel provide visual indi-
To test the audio and other systems,
cation of various conditions and/or
functions of selected systems. Red
use the rotary TEST switch (Figure
5J-15) on the left instrument panel.
Systems
lights indicate a malfunction that
When the switch is rotated through
requires immediate corrective action;
each position, the associated system
amber lights indicate either a mal-
warnings occur. The TEST switch
function that requires attention, but not
checks various annunciators, lights,
necessarily an immediate action, or
and audio warnings before takeoff.
normal system operation. The adviso-
Refer to Table 5J-B, following page.
ry white lights notify the pilot of the
system’s operational status.
Stick Shaker
Annunciators dim automatically when
the PANEL LIGHT CONTROL toggle Probe-type
switch is ON (see Lighting in the The angle-of-attack indicator flags and
Electrical chapter). drives to zero. Then the indicator flag
A red MASTER WARNING light illu- pulls from view and the indicator nee-
minates when a red annunciator illu- dle drives to 1.0. As the needle passes
minates, inadvertent thrust reverser approximately 0.82, the stick shaker
deployment occurs, or both amber activates for a few seconds. This cycle
GEN OFF annunciators illuminate (see is repeated as long as the test switch
the Electrical chapter). The MASTER remains in the stick shaker position. If
WARNING lights incorporate a reset the cycle is incomplete when the test
switch actuated by pressing the switch is moved to another position, the
MASTER WARNING RESET light. cycle continues until it is completed.
The light remains illuminated and flash-
ing until reset, even if the malfunction Vane-type
is corrected. Pressing a MASTER The stick shaker operates. The angle-
WARNING light normally does not of-attack meter needle moves past the
extinguish the associated annunciator. red area and the EADI fast/slow nee-
Landing gear warning and VMO/MMO dles go past slow. The red indexer
overspeed audio sounds are also part of (optional) light flashes on and off.
the warning system.

5J-15

Citation V For SimuFlite training only 5J-25

March 1999
 Switch Position Effect

OFF The red light above the rotary test switch is off and test
system inoperative.

FIRE Both engine fire lights on the upper center instrument

WARN panel illuminate.

LDG Three green (gear safe) lights and one red (UNLOCKED)
GEAR light illuminate; the gear horn sounds. Check that the horn
can be silenced by pressing the horn silence button on the
landing gear panel. The horn can be silenced only if flap
position is less than 15°. With the flap handle in the
second detent (15°), the horn does not silence.

BATT BATT O’TEMP annunciator flashes; battery

TEMP temperature gage shows 160°F to demonstrate circuit
integrity.

STICK On the probe-type stick shaker, the angle-of-attack

SHAKER indicator drives to zero and the flap appears. The flap dis-
appears and the indicator moves to 1.0. At approximately
0.82, the stick shaker activates for a few seconds. This
cycle is repeated as long as the test switch remains in this
position. On the vane-type stick shaker, the stick shaker
operates. The AOA meter needle goes past the red area
and the EADI fast/slow needles go past slow. Red indexer
lights flash on and off, if installed.

T/REV Six thrust reverser lights illuminate. The MASTER

WARNING lights also illuminate.

W/S TEMP The W/S AIR O’HEAT annunciator illuminates when LOW
or HIGH is selected on the windshield bleed air switch.

OVER The audible overspeed warning sounds.

SPEED

ANTI The anti-skid system initiates a self-test: ANTI-SKID INOP

SKID annunciator illuminates and remains on three to four sec-
onds after the test switch is moved to OFF; the annunciator
extinguishes if the system checks operational. If the system
fails the check, the annunciator remains illuminated.

ANNU All annunciators and the MASTER WARNING lights illumi-

nate. The turbine speed indicator self-tests. For example,
the red light illuminates and the digital display flashes 8s.
When the avionic power switches are on, the altitude alert
horn sounds, and the altitude alert and autopilot/ flight
director mode selector panel lights illuminate. EFIS and
FMS lights also illuminate. The MASTER WARNING lights
cannot be reset while the rotary selector switch is in this
position.

Table 5J-B; Rotary Test Switch Positions

5J-26 For SimuFlite training only Citation V

March 1999
 Miscellaneous Systems

Annunciator
Panels

Citation V For SimuFlite training only 5J-27 5J-28

March 1999
 Annunciator Panels

A 1 2 3 2 1 G
S S D A
ARM T EMER EMER T ARM
O BOTTLE 1 LH ENG RH ENG BOTTLE 2 O B
UNLOCK W ARMED FIRE FIRE ARMED W UNLOCK

S NORM
PUSH PUSH PUSH PUSH C
DEPLOY NORM S DEPLOY
W W

B
4 5 6 7 8 9 10 11

AC BATT CAB ALT OIL PRESS FUEL LOW FUEL LOW HYD FLOW ENG
FAIL O’TEMP 10000 FT WARN LEVEL PRESS LOW ANTI-ICE

LH RH LH RH LH RH LH RH LH RH

12 13 14 15 16 17 18 19

GEN INVERTER EMERG FUEL FLTR FUEL HYD LOW WING

OFF FAIL PRESS ON POWER BY PASS BOOST ON LEVEL O-HEAT
BRAKE
BLD AIR LOW PRESS HYD PRESS
LH RH 1 2 GND LH RH LH RH ON LH RH

H E
20 21 22 23 24 25 26 27 28
F
DOOR AIR DUCT ANTI SKID P/S HTR F/W NOSE COMP W/S AIR
SPEED SEAL O’HEAT INOP OFF SHUTOFF O’TEMP O-HEAT SURFACE
BRAKE DE-ICE
EXTEND DOOR NOT ACM GROUND AOA HTR
LOCKED O’PRESS IDLE LH RH LH RH FAIL

C D
29 30 31 32 33 34 35

MFD MFD EADI DISP

AUTOPILOT AP SG HOT HOT FAN MSG DR
OFF ROLL HOT
AP TORQUE MONITOR EHSI MFD EHSI SG WPT BATT
HOT FAN HOT HOT H
40 TEST

OFF
FIRE
ANNU WARN
E 36
F 37 38
G 39 ANTI LDG
SKID GEAR
OVER BATT
NOSE MASTER SPEED TEMP
GEAR
SG FAIL
UNLOCKED WARNING W/S TEMP
STICK
LH RH RESET SHAKER
T/REV

5J-28 For SimuFlite training only Citation V

March 1999
 Miscellaneous Systems

Major annunciators, with brief explanations, are listed alpha-

betically to correspond with alphabetical designations on the Annunciator
illustration on the facing page. Numbers in the second column
of the table also correspond to numbered annunciators on the
Cross
illustration. Specific information about what causes each Reference
annunciator to illuminate is in the appropriate system chapter.

ARM/ 1 The amber ARM light indicates the isolation valve is

UNLOCK/ open and the system is pressurized. The amber
DEPLOY UNLOCK light indicates the T/R is not fully stowed
and that reverser doors are unlocked. The white
DEPLOY light indicates the thrust reverser is fully
deployed.

BOTTLE 2 The white BOTTLE ARMED PUSH switchlights

ARMED indicate the engine fire extinguisher cartridges are
PUSH armed to fire into the engine selected by the ENGINE
1/2 FIRE PUSH switchlight. Push the BOTTLE ARMED
PUSH switchlight to release extinguishing agent.

ENGINE FIRE 3 An illuminated red ENGINE FIRE PUSH switchlight

PUSH indicates the corresponding fire detection loop has
LH/RH detected a temperature of approximately 500°F.
Pushing the switchlight closes the fuel and hydraulic
firewall shutoff valves, arms both extinguisher bottles
(illuminates the BOTTLE 1/2 ARMED PUSH switch-
lights), trips the generator field relay, and removes
power from the thrust reverser isolation valve.

AC 4 The red alternating current fail annunciator

FAIL illuminates if AC power bus voltage is above 130V
AC or below 90V AC. The master warning light also
illuminates.

ACM 22 The amber air cycle machine overpressure annunci-

O PRESS ator illuminates if the ACM pressure exceeds 42 PSI.

AIR DUCT 22 The amber air duct overheat annunciator illuminates

O’HEAT if the ventilation duct temperature exceeds safe
limits (315°F).

ANTI-SKID 23 The amber anti-skid inoperative annunciator

INOP illuminates if the anti-skid system is inoperative or
the switch is off.

AOA HTR 27 The amber angle-of-attack heater fail annunciator

FAIL illuminates if the probe heating element is operative
or if the pitot heat switch is off.

Citation V For SimuFlite training only 5J-29

March 1999
 B (continued)

BATT 5 The red battery overtemperature annunciator

O’TEMP illuminates steadily at temperatures above 145°F;
it flashes at temperatures above 160°F. The master
warning light also illuminates.

BLD AIR 14 The amber bleed air ground annunciator illuminates

GND when a high flow rate of bleed air from the right
engine is selected for ground operation of the ACM;
the pressurization source selector is in the GND
position.

CAB ALT 6 The red cabin altitude annunciator illuminates if cabin

10,000 FT pressure altitude exceeds 10,000 ft. The master
warning light also illuminates.

DOOR NOT 21 The amber DOOR NOT LOCKED annunciator illumi-

LOCKED nates if any one or more of the following are not
locked:
■ cabin entrance door

■ tailcone baggage door lock

■ nose baggage door lock.

DOOR 21 The amber DOOR SEAL annunciator illuminates if

SEAL the cabin entrance door seal is not inflated.

EMER 14 The amber emergency pressurization on annuncia-

PRESS ON tor illuminates if emergency pressurization is either
manually selected or automatically activated by an
air cycle machine (ACM) overheat.

ENG 11 The amber engine anti-ice annunciator illuminates if

ANTI-ICE the respective engine cowl leading edge is below
LH/RH 220° ±8°F, the engine stator valve is not open, or the
inboard wing bleed air temperature is below 300°F.

F/W 25 The amber firewall shutoff annunciator illuminates if

SHUTOFF the respective fuel and hydraulic firewall shutoff
LH/RH valves are closed.

FUEL 17 The amber fuel boost on light illuminates if electric

BOOST ON power is applied to the respective fuel boost pump.
LH/RH

HYD LOW 18 The amber hydraulic low level annunciator illuminates

LEVEL if the fluid in the hydraulic reservoir is 0.2 gallons or
less.

FUEL FLTR 16 The amber fuel filter bypass annunciator illuminates

BYPASS when the respective fuel filter is contaminated and
LH/RH an actual or impending bypass exists.

FUEL LOW 9 The amber fuel low pressure annunciator illuminates if

PRESSURE fuel pressure in the respective system is below 5 PSI.

5J-30 For SimuFlite training only Citation V

March 1999
 Miscellaneous Systems

B (continued)

FUEL LOW 8 The amber fuel low level annunciator illuminates if

LEVEL the fuel quantity in the respective tank is below
LH/RH approximately 185 lbs.

GEN 12 The amber generator off annunciator illuminates if

OFF the corresponding generator power relay is open and
LH/RH the generator is not connected to the aircraft bus.
The master warning light illuminates when both left
and right generator OFF lights illuminate.

GROUND 23 With the aircraft on the ground, the amber ground

IDLE idle annunciator illuminates if the ground idle switch
is in the normal (NORM) position.

HYD FLOW 10 The amber hydraulic flow low annunciator illuminates

LOW if the respective hydraulic system flow is below
approximately 0.35 to 0.55 gallons per minute.

HYD PRESS 18 The amber hydraulic pressure on annunciator illumi-

ON nates if the hydraulic system is pressurized.

INVERTER 13 The amber inverter fail annunciator illuminates if

FAIL either the No. 1 or No. 2 inverter output voltage is
1/2 above 130V AC or below 90V AC. If either inverter
fails, the AC FAIL annunciator also illuminates.
Resetting the master warning switch extinguishes
the AC FAIL annunciator unless both inverter fail
annunciators are illuminated.

NOSE COMP 26 The amber nose compartment overtemperature

O TEMP annunciator illuminates if the temperature in the nose
avionics compartment is excessively high. High
temperatures in that are can damage sensitive
avionics equipment (aircraft 560-075 and sub.).

OIL PRESS 7 A red oil pressure warning annunciator illuminates if

WARN the oil pressure is below safe limits (40 PSI) in the
LH/RH corresponding engine. The master warning light also
illuminates.

P/S HTR 24 The amber pitot static heater off annunciator

OFF illuminates if the corresponding pitot heat switch is
LH/RH off or power has been lost to the pitot tube or one or
both static ports in that system.

POWER 15 The amber power brake low pressure annunciator

BRAKE illuminates if the power brake hydraulic pressure is
LOW PRESS low (less than 750 PSI).

SPEED 20 The white advisory speedbrakes extended annuncia-

BRAKE tor illuminates when the left and right speedbrakes
EXTEND are fully extended.

Citation V For SimuFlite training only 5J-31

March 1999
 B (continued)

SURFACE 28 The white surface deice annunciator illuminates if

DEICE inflation pressure in the surface deice boots exceeds
21 PSI.

W/S AIR 27 The amber windshield air overheat annunciator illu-

O HEAT minates if the temperature of bleed air flowing onto
the windshield exceeds safe temperature limits, or,
with the windshield heat switch OFF, air pressure in
the line exceeds 5 PSI.

WING 19 The amber wing overheat annunciator illuminates if

O’ HEAT the corresponding wing temperature sensors detect
LH/RH heat above 160°F in their location.

AP ROLL 30 The amber AP ROLL MONITOR annunciator illumi-

MONITOR nates when the autopilot roll rate monitor detects an
autopilot excessive roll rate command.

AP TORQUE 29 The amber AP TORQUE annunciator illuminates

(steady illumination) if the autopilot torque fails to the
high torque mode. Momentary illumination indicates
autopilot is manually disengaged by the AP/TRIM
DISC button on the control wheel, the go-around
button, or the electrical elevator trim switch.

AUTOPILOT 29 Steady illumination of the amber AUTOPILOT OFF

OFF annunciator indicates the autopilot is disengaged
due to one or more of the following:
Abnormal – AC or DC power loss, vertical or
directional gyro failure, or torque limiter failure
Normal – A/P test button pushed, vertical gyro
switch (HI) or LH/RH gyro switch actuation.
Momentary illumination indicates the A/P trim
disconnect switch is pushed.

BATT 35 The amber battery annunciator illuminates if the

GNS-X RPU is being powered by its own internal
battery. This annunciator is present only when the
optional VLF/OMEGA is installed.

DR 35 The amber DR light illuminates when the system is

in dead reckoning.

DISP 33 The amber DISP FAN annunciator illuminates if the

FAN EADI/EHSI cooling fan fails. Continued operation
with the fan failed may cause display failure(s) due
to overheating.

5J-32 For SimuFlite training only Citation V

March 1999
 Miscellaneous Systems

D (continued)

EADI 33 The amber EADI HOT annunciator illuminates if the

HOT EFIS electronic display indicator is overheated due
to EADI/EHSI cooling fan failure. Continued
operation may result in display failure due to
overheating.

EHSI 33 The amber EHSI HOT annunciator illuminates if the

HOT EHSI display indicator is overheated due to
EADI/EHSI cooling fan failure. Continued operation
may cause display failure due to overheating.

MFD FAN 32 The amber MFD FAN annunciator illuminates if the

MFD cooling fan fails.

MFD HOT 32 The amber MFD HOT annunciator illuminates if the

multifunction display is overtemperature.

MFD SG HOT 31 (EFIS) The amber multi-function display signal

generator annunciator illuminates if the MFD signal
generator is overheated.

MSG 34 Illumination of the amber message annunciator

indicates the VORTAC positioning unit (VPU) is not
supplying navigation or position update information
to the flight management system.

SG HOT 33 The amber SG HOT annunciator illuminates if a

symbol generator fan fails. Continued operation with
the fan failed may cause a symbol generator
overheat and failure.

WPT 34 The amber waypoint annunciator illuminates when

the aircraft is within 30 seconds of the next leg
change.

SG FAIL 36 The red SG FAIL annunciator illuminates if the EFIS

symbol generator fails. The red X and SG FAIL
annunciator illuminate if the EADI symbol generator
fails.
On single EFIS, the EFIS system is inoperative
unless an MFD SG is available.
On dual EFIS, SG 1 or SG 2 illuminated indicates the
cross-side symbol generator is driving the affected
displays.

Citation V For SimuFlite training only 5J-33

March 1999
 F

GEAR 38 The red GEAR UNLOCKED annunciator illuminates

UNLOCKED with the gear handle down to indicate one or more
gear assemblies are not down and locked. It
illuminates with the gear handle up to indicate at
least one gear assembly has unlocked from the
down position, but is not yet up and locked.

NOSE 37 The green NOSE LH/RH annunciators illuminate on

LH/RH gear extension as each gear microswitch (nose gear,
left and right main gear) makes contact with its
respective gear down microswitch; this indicates the
gear is in the down and locked position.

MASTER 39 The red master warning switchlight flashes when a

WARNING red annunciator illuminates, inadvertent thrust
RESET reverser deployment occurs, or both amber GEN
OFF annunciators illuminate. It continues to flash
until reset, even if the malfunction is corrected.

ROTARY 40 When the red light above the rotary TEST switch is
TEST SWITCH off, the test system is inoperative and the TEST
LIGHT switch is in OFF. See Table 5J-B, page 5J-26 for
explanations of test positions.

5J-34 For SimuFlite training only Citation V

March 1999
This chapter describes the aircraft systems that extract, distribute,
and control engine bleed air. The combination of these systems Pneumatic
in this chapter presents the flow of the engine bleed air and its use
throughout the aircraft. For the onboard oxygen system, see the Systems
Miscellaneous chapter.
The bleed air system extracts engine bleed air and transfers it to
various other systems (e.g., air conditioning, ice and rain pro- Chapter 5K
tection, and pressurization). For a description of the ice and
rain protection system, see the Ice and Rain chapter.
The air conditioning system routes engine bleed air through
an air cycle machine (ACM) for temperature adjustment, then
distributes this conditioned air throughout the aircraft for cabin
pressurization. Temperature controls and various valves regu-
late the air’s temperature and distribution.
Conditioned bleed air enters the cabin through outlets in the
floor and ceiling. Two pressurization system outflow valves
regulate the flow of air out of the pressure vessel to maintain a
comfortable cabin pressure.

Citation V For SimuFlite training only 5K-1

March 1999
5K-2 For SimuFlite training only Citation V
March 1999
 Pneumatic Systems

Pneumatic/Air Conditioning System Schematic . . 5K-6

Table of
Bleed Air System . . . . . . . . . . . . . . . . . . . 5K-7
Extraction and Collection . . . . . . . . . . . . . . . 5K-7
Contents
Tailcone Distribution . . . . . . . . . . . . . . . . . . 5K-7
Cluster Assemblies . . . . . . . . . . . . . . . . . 5K-7
Bleed Air Shutoff (Flow Control) Valves . . . . . . . 5K-8
Ground Shutoff and Pressure Regulating Valve . . . 5K-9
Emergency Pressurization Valve . . . . . . . . . . 5K-10
PRESS SOURCE Selector Switch . . . . . . . . . 5K-10
Pneumatic Bleed Air . . . . . . . . . . . . . . . . . 5K-10
Bleed Air Pressure Regulator and Relief Valve . . 5K-11
Entrance Door Seal System . . . . . . . . . . . . 5K-11

Air Conditioning System . . . . . . . . . . . . . . 5K-13

Air Cycle Machine . . . . . . . . . . . . . . . . . . 5K-13
Overtemperature Protection . . . . . . . . . . . . 5K-15
Air Duct Overheat Annunciator . . . . . . . . . . . 5K-15
Bypass Mixing Valve . . . . . . . . . . . . . . . . 5K-15
Temperature Control . . . . . . . . . . . . . . . . 5K-15
Temperature Sensors . . . . . . . . . . . . . . . 5K-15
Air Distribution . . . . . . . . . . . . . . . . . . . . 5K-16
Mixing Tube . . . . . . . . . . . . . . . . . . . . 5K-16
Flow Divider . . . . . . . . . . . . . . . . . . . . 5K-16
Ventilation Outlets . . . . . . . . . . . . . . . . . 5K-16
Overhead Air Blower . . . . . . . . . . . . . . . . 5K-16
Defog Fan . . . . . . . . . . . . . . . . . . . . . 5K-16
Fresh (Ram) Air . . . . . . . . . . . . . . . . . . 5K-17
Flood Cooling . . . . . . . . . . . . . . . . . . . . 5K-17
Avionics Bay Cooling System . . . . . . . . . . . . 5K-17
Instrument Panel Cooling . . . . . . . . . . . . . . 5K-17

Citation V For SimuFlite training only 5K-3

March 1999
 Pressurization System Schematic . . . . . . . . . 5K-18
Pressurization System . . . . . . . . . . . . . . . 5K-19
Components . . . . . . . . . . . . . . . . . . . . . 5K-19
Vacuum Jet Pump and Regulator . . . . . . . . . 5K-20
Pressurization Controller . . . . . . . . . . . . . . 5K-20
Auxiliary Volume Tank . . . . . . . . . . . . . . . 5K-21
Pneumatic Relay . . . . . . . . . . . . . . . . . . 5K-21
Altitude Limit Valves . . . . . . . . . . . . . . . . 5K-21
Solenoid Air Valves . . . . . . . . . . . . . . . . . 5K-21
Air Filter . . . . . . . . . . . . . . . . . . . . . . 5K-22
Cabin Altimeter and Cabin Rate of
Change Indicators . . . . . . . . . . . . . . . . 5K-22
Cabin Altitude Annunciator . . . . . . . . . . . . . 5K-22
Emergency Dump Valve . . . . . . . . . . . . . . 5K-22
Emergency Pressurization . . . . . . . . . . . . . . 5K-22
Emergency Pressurization Annunciator . . . . . . . 5K-23
Mixing Tube . . . . . . . . . . . . . . . . . . . . 5K-23
Tailcone Pressurization . . . . . . . . . . . . . . . 5K-23

Servicing and Procedures . . . . . . . . . . . . . 5K-25

Preflight Inspection . . . . . . . . . . . . . . . . . . 5K-25
Abnormal Procedures . . . . . . . . . . . . . . . . 5K-25
Air Duct Overheat . . . . . . . . . . . . . . . . . 5K-25
Emergency Pressurization On . . . . . . . . . . . 5K-25
Air Cycle Machine Overpressure . . . . . . . . . . 5K-25
Vacuum System Failure . . . . . . . . . . . . . . 5K-26
Overpressurization . . . . . . . . . . . . . . . . . 5K-26
Emergency Procedures . . . . . . . . . . . . . . . 5K-26
Rapid Decompression (CAB ALT 10,000 FT
Annunciator Illuminated) . . . . . . . . . . . . . 5K-26
Emergency Descent . . . . . . . . . . . . . . . . 5K-26
Environmental System Smoke or Odor . . . . . . 5K-26

5K-4 For SimuFlite training only Citation V

March 1999
 Pneumatic Systems

Smoke Removal . . . . . . . . . . . . . . . . . . 5K-26 Pneumatic/

Data Summaries . . . . . . . . . . . . . . . . . . 5K-27 Air Conditioning
System
Bleed Air System Data . . . . . . . . . . . . . . . . 5K-27
Air Conditioning System Data . . . . . . . . . . . . 5K-28
Pressurization System Data . . . . . . . . . . . . . 5K-28

Citation V For SimuFlite training only 5K-5 5K-6

March 1999
 Pneumatic/Air Conditioning System

RIGHT FLOW CONTROL GROUND CONTROL

LH MAIN BUS EXT RH MAIN DC BUS (AFT J BOX) SHUTOFF VALVE VALVE (18 LBS/MIN) ACM
(6 LBS / MIN) O'PRESS

20A CABIN AIR FLOOD

20A COOLING RIGHT PRIMARY
BLEED AIR OVERPRESSURE BLEED AIR
GND
1 ENGINE BLEED AIR 2 SYSTEMS BLEED AIR CLUSTER SWITCH
FLOOD COOLING
■ CONE ANTI-ICE ■ DE-ICE BOOTS OVHD DEFOG ON
■ T1 PROBE ANTI-ICE ■ W/S BLEED AIR
FLOOD LH MAIN BUS EXT
■ INLET ANTI-ICE HI HI
COOLING
■ PRESS CONTROLLER NORM
VALVE
■ STATOR ANTI-ICE ■ DOOR SEAL O CPT CABIN PRESS TEMP
■ STATOR COOLING ■ WING ANTI-ICE F OFF 1 RH CROSSOVER BUS
F 5A 5A
■ TURBINE COOLING ■ 23 PSI CONTROLLER LOW LOW FLOOD
■ OIL SEALS
AIR
COOLING EMER
FLOW DISTR 5A PRESS
■ FUEL SYSTEM
AFT PRESSURE NORM

BULKHEAD FRESH
AIR RT WING
ANTI-ICE LH RH

GND EMER
OFF

AUXILIARY SECONDARY RAM AIR

PLENUM OVERPRESSURE SWITCH
OVERHEAD
BLOWER
3

BYPASS MIXING
MIXING 2 VALVE
CHECK TUBE
WEMAC VALVE SYSTEMS COOLING TURBINE (T) AND
CABIN TEMP SENSOR BLEED AIR AMBIENT COOLING COMPRESSOR (C)
WINDSHIELD & BLOWER ASSEMBLY OUTLETS OVERHEAD AIR IN
DEFOG T C
COCKPIT VENT (ABOVE INTERIOR (WEMAC) DUCT BLOWER BLEED AIR
FOOT WARMER HEADLINER) DEFOG FAN PRECOOLER
WEMAC CABIN AIR
DUCT
O'HEAT OVERBOARD
SENSOR PLENUM
WATER
OUTFLOW SEPARATOR
RESTRICTOR VALVES
WATER TO
ASPIRATOR
AT AMBIENT
100°F (38°C) LT WING AIR INLET
THERMOSWITCH ANTI-ICE AMBIENT
AIR OUT
SIDE WINDOW PASSENGER ARMREST UNDER-FLOOR MAIN PLENUM FLOW DIVIDER LEFT EMERG
PRESS ON
DEFOG FOOTWARMER WARMER DUCT BLEED AIR
CLUSTER ACM OVERHEAT
SENSOR
3 450°F AIR DUCT
RAM AIR TEMPERATURE CONTROL O'HEAT
MANUAL AUTOMATIC
COLD AIR HOT OVERTEMP
1 PROTECTION 315°F (157°C)
BLEED AIR CIRCUIT
OFF
HEAT-EXCHANGER COOLED BLEED AIR 410°F (210°C)
MANUAL COLD
CONDITIONED AIR COLD HOT HEAT EXCHANGER
MANUAL

EMERGENCY LEFT FLOW CONTROL

PRESSURIZATION VALVE SHUTOFF VALVE (6 LBS / MIN)

5K-6 For SimuFlite training only Citation V

March 1999
 Pneumatic Systems

Extraction Tailcone Bleed Air

and Collection Distribution
Extraction of bleed air occurs during Major components of the distribution System
engine(s) operation from the four and system include:
eight o’clock positions of each engine’s ■ cluster assemblies
high pressure (HP) compressor. The
■ flow control shutoff valves
transfer tubes are on the front of the
bypass duct. The transfer tube’s design ■ ground shutoff valve

restricts the airflow from the HP com- ■ emergency pressurization valve

pressor to prevent excessive bleed air
■ check valves
draw from the engines.
■ PRESS SOURCE selector switch
Elbow assemblies collect the engine
bleed air from the transfer tubes and ■ bleed air pressure regulator.

route it to the tailcone bleed air distri-

bution system (Figure 5K-1). See the Cluster Assemblies
Powerplant chapter for details. Engine bleed air passes through the
The distribution of bleed air occurs via nacelles into each engine’s respective
the nacelle pneumatic system through cluster assembly manifold. The cluster
the pylon into the tailcone, where it is assembly is in the tailcone where the
available for use in the respective air- distribution of air occurs. Each assem-
craft systems. bly provides three individual outlets

NORMALLY
OPEN

BLEED AIR
CLUSTER
MOTOR VALVE

NORMALLY
CLOSED

TRANSFER
ELBOW TUBES
ASSEMBLY

5K-1

Citation V For SimuFlite training only 5K-7

March 1999
 for bleed air to one of the following valve prevents one engine’s bleed air
systems: from entering the lower RPM engine’s
■ air conditioning and pressurization through the shutoff valve. In addition,
the check valve prevents the flow of
■ emergency pressurization (left
bleed air from an operating engine into
engine) and bleed air ground (right a non-operating engine. A normally
engine) open solenoid controls the valve; bleed
■ deice boots, anti-ice, rain removal, air pressure activates the valve.
and other pneumatic systems. Bleed air passes through the shutoff
Bleed Air Shutoff valves into a tee and enters the refrig-
eration unit and pressurization system
(Flow Control) Valves (Figure 5K-2).
Air leaving the cluster assembly out- The PRESS SOURCE rotary switch
let to the air conditioning system pass- (Figure 5K-3) labeled OFF, GND, LH,
es through its respective bleed air shut- NORMAL, RH and EMER is in the
off valve. The valve allows a minimum cockpit. The switch electrically controls
flow rate of 6.0 ppm to the air condi- the shutoff valves. In flight, shutoff
tioning system. Each valve combines valves usually remain open. Electrical
the functions of a flow control shutoff power moves the shutoff valves to the
valve and a check valve. The check

GROUND S.O.
RH WING AND PRESS
ANTI-ICE REG. CLOSED

TO AIR CYCLE
CHECK MACHINE (ACM)
VLV

LH WING FLOW CONTROL

ANTI-ICE S.O. VALVES
(OPEN WHEN BLD
AIR PRESSURE SENSED)
EMERG
PRESS
CLOSED

5K-2

5K-8 For SimuFlite training only Citation V

March 1999
 Pneumatic Systems

closed position. The valve returns to the Ground Shutoff and

open position if an electrical failure Pressure Regulating Valve
occurs with the switch selected to a posi-
tion other than NORMAL. The motorized ground shutoff and
pressure regulating valve is part of an
A sensor and relays automatically close automatic ground mode shutoff sys-
the shutoff valves if an overheat con- tem. This system prevents the over-
dition occurs in the ACM. The sensor pressurization of the ACM.
illuminates the EMERG PRESS ON
annunciator. The sensor and relays On the ground only, selecting GND on
switch the left and right shutoff valves the PRESS SOURCE switch opens the
from open to closed and open the valve and closes both bleed air shutoff
emergency pressurization valve of the valves. The right engine supplies the
left cluster during flight. The left squat bleed air for the air conditioning system.
switch prevents emergency pressur- Prior to the ground shutoff valve, a pri-
ization activation on the ground. mary pressure switch closes the ground
The LH Main bus Extension supplies shutoff valve if a pressure of 32 to 38
power to the rotary switch for shutoff PSIG (approximately 72% right engine
valve operation through the 5 amp N2) occurs. With the shutoff and pres-
NORM PRESS CB. sure regulating valve energized, the
BLD AIR GND annunciator illumi-

Switch Function
Position

OFF Closes (energizes) both flow control shutoff valves, to stop

airflow to the ACM.

GND With the right engine operating, the ground shutoff and pressure-
regulating valve is open (energized). This allows approximately
18 ppm of bleed air flow through the ACM to ventilate the cabin.
Both flow control shutoff valves are closed (energized) and the
BLEED AIR GND annunciator illuminates. This completes the
circuit for the primary and secondary overpressure switches.
5K-3
LH Opens (de-energizes) the left flow control shutoff valve and closes
(energizes) the right flow control shutoff valve. This allows bleed
air to flow only from the left engine to the ACM.

NORMAL To operate the ACM, both flow control shutoff valves open (de-
energize). This allows an airflow of 6 ±0.5 ppm from each engine
(combined for 12 ppm).

RH Opens (de-energizes) the right flow control and shutoff valve and
closes (energizes) the left flow control and shutoff valve. This
allows bleed air to flow only from the right engine to the ACM.

EMER Both flow control and shutoff valves close (energizes), the
ground shutoff valve closes, and the emergency pressurization
valve opens (energizes). This routes left engine bleed air to a
mixing tube in the cabin air distribution system for emergency
cabin pressurization. The EMERG PRESS ON annunciator illu-
minates when the emergency pressurization valve opens.

Table 5K-A; PRESS SOURCE Switch Selection

Citation V For SimuFlite training only 5K-9

March 1999
 nates to advise that bleed air for ground and de-energizes an assortment of
operation of the air conditioner is orig- valves. The six-position PRESS
inating from the right engine. If a fail- SOURCE switch is on the center tilt
ure of the primary pressure switch instrument panel. For switch position
should occur, a secondary pressure and function see Table 5K-A, previous
switch downstream of the valve closes page.
the valve at a pressure range from 36 to
42 PSIG (74% right engine N2) and Pneumatic Bleed Air
illuminates the ACM O’PRESS annun-
Pneumatic bleed air is unconditioned
ciator. The illumination of the annun-
hot bleed air extracted from the left and
ciator notifies the crew that the prima-
right bleed air clusters. This bleed air
ry switch has failed.
feeds the following pneumatic systems:
Emergency ■ left and right wing anti-ice
Pressurization Valve ■ wing anti-ice control

See Pressurization section of this chap- ■ pressure control (vacuum)

ter for the description of the emergency ■ deice boots
pressurization valve.
■ windshield bleed air
PRESS SOURCE ■ door seal.
Selector Switch
Air tapped from the bleed air clusters,
The PRESS SOURCE selector switch travels to a set of tee fittings (“christ-
(Figure 5K-3, previous page) energizes mas tree”) for distribution. Each air

RIGHT WING
ANTI-ICE

2
SERVICE BLEED AIR
CHRISTMAS TREE
ROUTING TO: 2
DOOR SEAL 1
DEICE BOOTS
W/S BLEED AIR ENGINE BLEED AIR
PRESS CONTROL OIL SEALS
WING A/I & TEMP TURB COOL
CONTROL CONE A/I
T1 A/I
STATOR A/I, COOL
FUEL SYS
LEFT WING INLET A/I
ANTI-ICE

5K-4

5K-10 For SimuFlite training only Citation V

March 1999
 Pneumatic Systems

source passes through a check valve. One output port of the “tree” supplies
The check valve prevents the escape bleed air to the cabin door inflatable
of air through an inoperative engine. seal. Another output supplies the de-ice
Prior to the check valves, a tee fitting boots. A third outlet supplies air for
extracts the hot bleed air for wing anti- control of the cooling tubes of the wing
ice. The respective engine feeds its own anti-ice.
wing anti-ice air source. However, a
crossfeed capability exists. Bleed air Entrance Door Seal
from both engines continues past the System
check valves and combines at the A check valve, a spring-loaded nor-
“christmas tree.” The tree distributes mally closed valve, an inflatable door
the air to various systems. The wind- seal, and bleed air lines and fittings
shield bleed air and pressure control comprise the entrance door seal sys-
use unregulated air for operation (see tem. The inflatable pressure seal is a
Ice and Rain chapter). The windshield molded rubber continuous ring fitted
bleed air uses the high temperature air inside the door frame. In addition, an
for anti-ice. The pressure control uses extruded rubber weather seal encloses
the high pressure air to generate a vac- the outer edge of the door. A rain seal
uum source through a vacuum regula- covers the door seal at the hinge area.
tor (venturi). The remaining air pass-
es through a bleed air pressure regulator A bleed air line from the pressure
and relief valve. regulator routes through the upper aft
pressure bulkhead to the door seal valve
Service Bleed Air at the forward door frame. Locking the
Pressure Regulator and door allows the lower forward door lock
Relief Valve pin to actuate the door valve, which
inflates the door seal with bleed air.
The pressure regulator maintains bleed
A check valve in the input line to the
air pressure for the entrance door seal
door seal valve prevents an immediate
system, wing de-ice boots, and the
deflation of the door seal that could
wing anti-ice temperature control.
lead to rapid decompression. Unlocking
Bleed air passes through various con-
the door opens a spring-loaded valve.
trols to regulate it to 23 ±1 PSIG with
This allows trapped air in the seal to
relief set at 27 ±1 PSIG. The regulat-
escape through a vent in the valve body,
ed air flows into further branches of
which frees the door to open.
the “christmas tree” (Figure 5K-4).

Citation V For SimuFlite training only 5K-11

March 1999
5K-12 For SimuFlite training only Citation V
March 1999
 Pneumatic Systems

Conditioned air is the result of bleed air

routed through the ACM. Overhead and
unit, a fan housing and water aspirator,
a 410°F overtemp protection circuit, Air
under-floor ducting in the cabin and
cockpit distribute the conditioned air.
and a 450°F ACM overheat sensor.
The ACM combines a radial flow
Conditioning
Placement of the PRESS SOURCE
switch determines the source of the
compressor and a radial flow turbine
on a common shaft. An extension of System
bleed air. In addition, the switch deter- the common shaft drives an external
mines the configuration of the bleed cooling fan. The fan draws air from
air valves (energized/de-energized), the tailcone across the precooler, pri-
which route the engine bleed air into mary, and secondary heat exchangers.
the ACM. A fan in the ACM draws The heat exchangers cool the air flow-
cooling air across the heat exchangers. ing into the compressor and turbine.
The cooling air passing across the heat
Air Cycle Machine exchangers carries the heat and
exhausts overboard through a vent on
The ACM (Figure 5K-5) utilizes engine
the under side of the tailcone. The
bleed air received from either the flow
ACM compressor compresses the
control shutoff valves or the ground
bleed air from the precooler and pri-
shutoff valve to provide conditioned air
mary heat exchanger. The air then
to the cabin. The ACM is on the right
passes to the secondary heat exchang-
side of the forward tailcone.
er and into the ACM turbine where the
Components include a turbine and air expands rapidly to produce cold air.
compressor, a bypass mixing valve,
The bypass mixing valve controls air
three heat exchangers joined as one
temperature by modulating the amount

BLEED AIR IN PRIMARY STAGE

TO HEAT EXCHANGER
AND COMPRESSOR
CONDITIONED
AIR DUCT BYPASS
HOT AIR
DUCT
HEAT
EXCHANGER
COOLING
AIR IN

HEAT EXCHANGER
COOLING AIR OUT
THROUGH EXHAUST
VENT
WATER
SEPARATOR
SECONDARY STAGE
AIR OUT WATER OF HEAT EXCHANGER
TO CABIN TUBE TO TURBINE

5K-5

Citation V For SimuFlite training only 5K-13

March 1999
 of hot bleed air that bypasses the ACM water separator. The water separator
and mixes directly with conditioned controls the humidity of conditioned
air. The bypass valve’s position air entering the cabin by removing
depends on the TEMPERATURE moisture from the conditioned air. The
CONTROL switch selection. moisture is routed from the separator
Engine bleed air passes through the and sprayed across the heat exchangers
precooler and primary heat exchangers to aid in cooling. Eventually the water
to lower the air’s temperature prior to exhausts overboard (Figure 5K-7) with
the compression cycle. The air con- the cooling air that flow across the heat
tinues through the ACM compressor exchangers.
where the air’s temperature and pres- If the TEMPERATURE CONTROL
sure increase. The air passes through switch is in AUTOMATIC, the mois-
the secondary heat exchanger to lower ture in the water separator should not
the temperature while pressure remains freeze. With the TEMPERATURE
high (refer to the Bleed Air/Air CONTROL switch in MANUAL, the
Conditioning schematic). moisture may freeze and raise the pres-
The cooled high pressure air then pass- sure differential across the water sep-
es through the expansion turbine and arator condenser assembly. If the dif-
into a cabin supply duct. The expand- ferential reaches 4.6 PSI, the poppet
ing air and dropping pressure cause a relief valve opens to direct air through
substantial drop in the temperature of the water separator without passing
the air exiting the ACM. through the condenser, that extracts
moisture from the air.
A small portion of bleed air from the
heat exchanger routes to the aspirator of Bleed air bypassing the ACM provides
the fan housing. Here the high veloci- cabin heat. An electrically operated
ty of the air creates a vacuum on the valve controls the bypass air that mixes
tube leading to the water separator. This with the ACM’s conditioned air to reg-
removes the water collected by the ulate cabin temperature.

TEMPERATURE CONTROL INLET

MANUAL AUTOMATIC
HOT

OFF

MANUAL COLD
COLD HOT
MANUAL

5K-6 WATER

DRAIN
TUBE VORTEX VANE
OUTLET
POPPET
RELIEF
VALVE

WATER

COALESCER
5K-7

5K-14 For SimuFlite training only Citation V

March 1999
 Pneumatic Systems

Overtemperature Bypass Mixing Valve

Protection Regulating the flow of hot bleed air
During high altitude operations, with a through the bypass mixing valve con-
low airspeed and high power settings, trols the conditioned air’s temperature.
it is possible to exceed the cooling The temperature control system elec-
capacity of the ACM. A 450°F ACM trically controls the bypass mixing
overheat sensor between the com- valve.
pressor and the secondary heat
exchanger monitors the temperature Temperature Control
in the ACM. If an overheat condition The TEMPERATURE CONTROL
should occur, the ACM shuts down by switch (Figure 5K-6) controls the cabin
closing the left and right flow control temperature either automatically or
valves and automatically opening the manually. Electrical control of the tem-
emergency pressurization valve of the perature occurs during the selection of
left bleed air cluster. AUTOMATIC. A temperature sensor
To prevent an ACM shutdown, an over- in the distribution duct combines with
temperature protection circuit/ switch a cabin temperature sensor and the
biases the temperature controller when position of the temperature control
the ACM compressor discharge reach- switch to modulate the bypass mixing
es 410°F. The bias prompts the tem- valve for proper temperature control.
perature controller to react and simulate In the MANUAL mode, the switch is
a warmer setting selection. This a momentary contact switch that directs
decreases the amount of air through the 28V to move the bypass valve. When
ACM by increasing the air flow manually controlled, the valve travels
through the bypass until the overtem- from full open to full closed in approx-
perature condition no longer exists imately 11 seconds and vice versa.
(below 410°F). With the decrease in The temperature sensor of the auto-
temperature, the bias automatically matic system is inoperative with the
returns the ACM to the temperature selection of MANUAL. With the tem-
controller setting. perature sensor inoperative, ice for-
The ACM overtemperature protection mation in the water separator and/or
functions only with AUTOMATIC ducting is possible with maximum
temperature control selected. For oper- cooling requested and outside humid-
ations above 31,000 ft, use only the ity above 40%.
AUTOMATIC temperature mode.
Selecting MANUAL temperature
Temperature Sensors
mode may cause the ACM to shut The various air conditioning system
down, which in turn generates an auto- sensors include a cabin temperature
matic EMER PRESS ON condition. sensor and a duct temperature sensor.
Air Duct Overheat The temperature sensor in the overhead
of the cabin detects the temperature of
Annunciator cabin ambient air and varies the auto-
The amber AIR DUCT O’HEAT matic control system accordingly. A
annunciator advises the pilot that the fan in the top of the cabin draws cabin
temperature in the duct leading to the air across the temperature sensor.
cabin exceeds 315 ±8°F. The air duct The automatic temperature control
overheat sensor is downstream from balances the selected input with the
the water separator. duct temperature sensor and cabin tem-

Citation V For SimuFlite training only 5K-15

March 1999
 perature sensor inputs to control ACM a motor-operated damper valve to
output. The temperature control range block air from the duct and open an
is 65°F to 85°F. air port beneath the cabin floor. Cabin
air is drawn from under the cabin floor
Air Distribution area into the overhead air duct.
The air distribution system directs and A control lever on either side of the
controls the flow of hot, cold, and/or cockpit controls the foot warmer out-
fresh air through the cabin and cockpit let valve. The lever also controls the
(refer to Bleed Air/Air Conditioning respective side window defog valve.
schematic). The defog valve directs the maximum
air flow to side windows for defog-
The ACM supplies the conditioned air.
ging with the foot warmers closed.
Fresh (unconditioned) air enters the
system from the ram air ducting in the The passenger foot warmer manifolds
tailcone (see the Fresh Air section, this are along the outboard cabin walls at
chapter). floor level. The flow divider directs
conditioned air to the foot warmers
Mixing Tube through the above-floor air distribu-
tion ducting.
During normal operation, conditioned
air from the ACM enters the mixing Conditioned air from the under-floor
tube in the air ducting under the aft air distribution system supplies the foot
cabin floor and continues throughout warmers just forward of the main
the air distribution system. entrance door.
The passenger armrest warmer mani-
Flow Divider folds are along the outboard cabin
The flow divider in the under-floor walls. Air holes under the armrests
ducting provides adjustable volumes receive conditioned air from the same
of conditioned air to the cabin above- source as the foot warmers.
floor air distribution and the cockpit
under-floor air distribution. The AIR Overhead Air Blower
FLOW DISTR switch in the cockpit The OVHD fan switch on the tilt panel
has five detents to control the flow controls the overhead air blower, which
divider. Each detent selection increas- increases the airflow through the over-
es or decreases the amount of air flow head outlets (WEMAC). The switch
into the cockpit or cabin depending on consists of HI/OFF/LOW positions.
switch position. Normally, because of When operating, the blower draws air
the “green house” effect of the win- from the ACM main distribution duct
dows, the crew positions the switch one if the temperature is below 100°F or
detent left of center to allow the cock- from the under-floor area when tem-
pit to receive the larger volume of air. perature in the duct is above 100°F.
Ventilation Outlets Defog Fan
The cabin overhead and cockpit con- The defog fan under the floor in the
tain fresh air vents (WEMACS). The conditioned air ducting to the cockpit
WEMACS receive air from the ACM increases airflow to the cockpit foot
distribution system if the temperature warmers. In addition, the fan supplies
is less than 100°F. A 100°F thermo- conditioned air to the windshield and
switch monitors the temperature in the side window defog outlets. The cock-
duct. As the temperature increases pit switch has three positions: HI, OFF,
above 100°F, the thermoswitch signals and LOW.

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March 1999
 Pneumatic Systems

Fresh (Ram) Air anytime the battery switch is ON. The

system consists of two fans: one fan
Pressurization
The cabin and tailcone receive fresh on the left and one on the right side of System
air through flush scoops in the dorsal the forward divider. The system cir-
fin. An air duct adjacent to the scoop culates air using the two fans to pull
and cabin ventilation junction box dis- air from the avionics bay into the bag-
tributes the fresh air. A check valve in gage area. This cools the avionics
the fresh air line to the cabin prevents equipment in the avionics bay. The RH
reverse air flow when the pressure in Crossover bus supplies 28V DC power
the cabin is greater than the ram air to the fans. The EQUIP COOL CB
pressure. A drain hole in the duct protects the system.
attachment connection at the aft pres-
sure bulkhead allows moisture in the A thermal switch monitors the avion-
duct to escape. ics bay temperature. Should the tem-
perature in the bay drop to 80°F to
Flood Cooling 85°F, the thermal switch opens to inter-
rupt power to both fans. A tempera-
The optional flood cooling system ture above 95 ±5°F closes the switch to
cools the cabin during ground or low supply power to the fans for operation.
altitude operation. The system supplies
a high volume flow of conditioned Instrument Panel
bleed air to flood the cabin for quick
and efficient cooling. The FLOOD Cooling
COOLING ON/OFF switch is on the The cooling system consists of an
environmental section of the tilt panel. exhaust fan and an optional tube-axial
When the switch is ON, all ACM con- blower. The system circulates and
ditioned air diverts to the flood cooling exhausts air on the forward side of the
system, which mixes fresh conditioned instrument panel to cool the panel.
air with grill or upper recirculated air The box-type exhaust fan on the defog
and routes it to the cabin outlet in the deck is to the left of the airplane cen-
aft pressure bulkhead. A check valve terline. The fan exhausts air from the
prevents reverse flow when the cabin forward side of the instrument panel.
is pressurized and the system is OFF.
The optional tube-axial type blower is
Use the system only during ground forward of the instrument panel. To
operation or in flight below 10,000 ft. prevent foreign object ingestion, the
Do not use the system for cabin heat- blower uses a mesh screen over the
ing. Using the system for heating or intake and exhaust. The blower circu-
above 10,000 ft may cause the tem- lates air over the instrument panel.
perature and/or pressure limitations of
With the temperature behind the instru-
the tailcone duct to be exceeded.
ment panel reaching 98°F, a thermal
switch automatically supplies 28V DC
Avionics Bay to operate the cooling fan. The EQUIP
Cooling System COOL CB protects the instrument cool-
ing system. The system is operational
The avionics bay cooling system oper- anytime power is on the main buses.
ates automatically and continuously

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March 1999
 Pressurization System

MAN S.O.
VLV NO. 2
BREATHER
LH/RH DC BUSES CAB ALT LIMIT
PNEU RELAY ASSY
(13,000' ± 1,500)

SOLENOID VLV "B" (N. C.) 28V DC

VLV
SOL AIR
VLV "A" (N.O.) TEST PORT OUTFLOW
VLV
CAB
AIR CAB ALT LIMIT
ELEC REG TEST (13,000' ± 1,500)
SYS VAC PORTS CONTROL

AUX VOL TANK

SOLENOID AIRCRAFT
4-WAY RESTRICT VALVE "C" (N.C.)
TAILCONE
EMER
DUMP
VLV (N.C.)
CONTROL

BLEED AIR
CABIN
ALT SEL CABIN RATE VACUUM
OUTFLOW
VALVE TEST
R/H OUTFLOW CONTROL
OUTFLOW CONTROL <80% REF PRESS BLEED AIR
8 7 6
FILTER DISCH PRESS
N >85%
AL
AC BI

FLT
X

FILTER
CA

L/H CABIN PRESSURE

GRND MAN S.O.
CABIN
CONTROLLER
<80% VLV NO. 3
GEAR VACUUM
>85% CHK VLV PUMP REG
FACE VIEW
(3.75 TO
THROTTLE SW 4.75" HG)
CABIN
FORWARD ATMOSPHERE
REG VAC CONTROL MAN S.O. VLV NO. 1

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March 1999
 Pneumatic Systems

As a constant source of conditioned

bleed air flows into the aircraft, out-
valve opens when cabin pressure ver-
sus outside pressure air exceeds 8.9 Pressurization
flow valves maintain cabin pressure
and differential to a maximum of 8.9
PSID. When the valve opens, air with-
in the outflow valve body flows over- System
±0.1 PSID. board and creates a vacuum. The vac-
Two spring-loaded closed outflow uum opens the outflow valve to allow
valves control the airflow out of the cabin pressure to decrease.
aircraft. When closed, the valves
restrict cabin air outflow. Cabin alti- Components
tude descends as the bleed air system Components of the pressurization sys-
continues to pump bleed air into the tem, in addition to the outflow valves,
cabin. The outflow valves are under include:
the cabin floor on the aft pressure bulk-
■ vacuum jet pump and regulator
head (Figure 5K-8).
■ pressurization controller
Vacuum pressure created from bleed
air through a jet pump counters the ■ auxiliary volume tank

spring action of the outflow valves. ■ pneumatic relay

The pressurization controller (CABIN ■ altitude limit valves
CONTROLLER) uses vacuum pres-
sure to signal a bleed air relay. The ■ solenoid air valves

bleed air relay amplifies the signal and ■ air filter

applies the proper amount of vacuum ■ cabin altimeter and cabin rate of
to the outflow valves. The vacuum
change indicator
opens the outflow valves, which allows
■ cabin altitude annunciator
cabin air to escape and cabin altitude
to ascend. A maximum differential ■ emergency dump valve
pressure relief valve on the outflow ■ various controls and indicators.

AFT
PRESSURE
BULKHEAD
TEST ALTITUDE LIMIT
PORTS CONTROL
TO AUX
VOLUME RIGHT
CHAMBER OUTFLOW
VALVE
STATIC LEFT
PORT OUTFLOW
TO VALVE
PNEUMATIC
RELAY

TO SOLENOID
VALVE (RELAY
DISABLE)
SHUTOFF
VALVE 2

SHUTOFF
VALVE 3
AIR FILTER STATIC
ASSEMBLY PORT
5K-8

Citation V For SimuFlite training only 5K-19

March 1999
 Vacuum Jet Pump panel meters the vacuum (i.e., control)
and Regulator air to the pneumatic relay which in turn
modulates the outflow valves to
When either or both engines are oper- achieve the desired cabin altitude and
ating, the bleed air system directs bleed rate of climb.
air through an air ejector to create a
vacuum at an air inlet port (refer to the A moveable diaphragm separates two
Pressurization schematic). The air chambers of the controller. One of the
passes through a filter that removes chambers senses cabin pressure while
foreign particles before reaching the the other senses the reference ambient
jet pump and regulator. The regulator pressure trapped prior to takeoff. As
takes the filtered air and regulates the cabin altitude increases, the pressure
vacuum pressure between 3.75 and differences between the two chambers
4.75 inches of mercury. cause the diaphragm to route vacuum
to the pneumatic relay, which ampli-
Vacuum is supplied to the pressuriza- fies the signal to control the outflow
tion controller, the pneumatic relay, valves. Cabin pressure increases or
and the EMER DUMP switch. decreases until equilibrium between
The pressurization controller regulates the two chambers occurs.
through trapped reference air a pneu- Select the desired cabin altitude by
matic relay below the floor panels in rotating the cabin altitude selector
the aft cabin area. The relay varies the knob. This applies a spring bias to the
amount of vacuum applied to the out- diaphragm and changes the pressure
flow valves. differential to achieve equilibrium
If the vacuum system fails, cabin pres- between the two chambers.
sure goes to maximum differential The CABIN RATE knob controls the
because vacuum or control air is not rate at which the cabin climbs or
available to manipulate the outflow descends. The knob operates a valve
valves. that bleeds air between the two sealed
chambers. The valve determines the
Pressurization Controller rate at which the spring pressure is
The pressurization controller selector applied to the moveable diaphragm with
(Figure 5K-9) on the lower center the selection of a new cabin altitude.

5K-9

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March 1999
 Pneumatic Systems

Auxiliary Volume Tank pressure chamber in the controller. The

valve opens to negate the system ref- NOTE: Control of cabin pressure
The auxiliary volume tank is a small erence air; this prevents the controller is automatic during taxi, takeoff
sealed chamber with a pneumatic tube from operating the pneumatic relay. roll, and up to aircraft cruise alti-
connected to the control port of the tude. The cabin pressure control
pneumatic relay assembly. The tank Solenoid air valve “C” between the
system requires no adjustments.
connects to a pneumatic tube contain- pneumatic relay and the cabin altitude
ing a breather plug and a test port num- limit valves is normally closed. When
ber 1 fitting, which connects to the the valve is open, vacuum from the jet
pneumatic relay assembly. The auxil- pump enters the control line and the
iary volume tank acts as a system shock outflow valves are held open to prevent
absorber (pneumatic accumulator) and the cabin from pressurizing.
provides a greater accuracy of cabin Throttle position and the left squat
rate of change control. switch control the operation of the sole-
noid valves (see the Pressurization
Pneumatic Relay schematic and pneumatic relay descrip-
The pneumatic relay is a signal boost- tion). The solenoid valves operate only
er that receives control information on the ground for pre-pressurization on
from the pressurization controller. The the takeoff roll as well as to prevent
relay amplifies the signal from the pressurization during ground operation.
pressurization controller and relays it During ground operation, the landing
to the outflow valves for position infor- gear squat switch provides a ground for
mation. The pneumatic relay is below the throttle switches. With throttle posi-
the floor panels in the aft cabin area. tion at less than 80%, both the solenoid
air valve “B” and solenoid air valve “C”
Altitude Limit Valves receive electrical power from the LH/
The cabin altitude limit valves prevent RH DC buses. Both valves, which are
cabin altitude from exceeding 13,000 normally closed, open. The pneumatic
±1,500 ft. The valves open to allow relay loses the reference signal from the
ambient cabin air directly into the vac- pressure controller and a vacuum port
uum lines to reduce the vacuum. The opens in the outflow valve line so that
outflow valves then close under their full vacuum reaches the outflow valves.
own spring tension to re-establish The vacuum produced by the vacuum
cabin pressure. jet pump and regulator overcomes the
internal spring in the outflow valves
Solenoid Air Valves and moves the valves to full open.
Three solenoid air valves activate elec- As the throttles increase above 85%, the
trically when the aircraft is on the throttle switches provide power to close
ground. the normally open solenoid air valve
Solenoid air valve “A” on the pres- “A” and simultaneously removes power
surization controller, when activated, from solenoid valves“B” and “C.” As
controls the ground differential pres- “B” and “C” valves close, reference sig-
sure control; it causes the controller to nals to the pneumatic relay are restored,
descend the cabin 60 ft below field and raw vacuum to the outflow valve is
altitude and renders the cabin rate of replaced by regulated vacuum from the
change function inoperative. pneumatic relay. Closing the solenoid
air valve “A” interrupts the cabin rate
The solenoid air valve “B” is down- controller’s signal and replaces it with a
stream of the auxiliary volume tank and 60 ft below field elevation command.
is normally closed. With the solenoid This permits the outflow valves to ini-
air valve “B” closed, the volume tank tiate cabin pressurization and to avoid
provides additional volume to the rate a sudden bump at aircraft rotation.

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March 1999
 At takeoff, the landing gear squat vates the annunciator at 10,000 ±350 ft
switch opens to remove power to the on increasing altitude. The annuncia-
solenoid air valve “A”. This allows the tor extinguishes on decreasing altitude
pressurization controller to take over prior to reaching 8,450 ft.
control of the outflow valves and cabin
pressurization. Emergency Dump Valve
The guarded EMER DUMP lever
Air Filter (Figure 5K-11) on the tilt panel allows
An air filter assembly prevents conta- the crew to rapidly dump cabin pres-
minated air from entering the pressur- sure if required. The guard covers the
ization system. The assembly consists lever to prevent accidental operation.
of a boot, a filter element, and an outlet The dump valve routes vacuum direct-
connection. ly to the outflow valves to force the
valves open; this releases cabin pres-
Cabin Altimeter and Cabin sure and allows the cabin altitude to
Rate of Change Indicators equalize with aircraft altitude up to
The CABIN ALT and CABIN rate of approximately 13,000 ±1,500 ft. Select
change indicators (Figure 5K-10) are the PRESS SOURCE selector to OFF
adjacent to the pressurization controller to obtain complete cabin depressuriza-
on the tilt panel. The cabin altimeter tion at altitudes above 13,000 ±1,500
presents the existing cabin altitude on ft. The cabin altitude limit valves pre-
the outer scale and pressure differential vent a system failure or inadvertent pilot
on the inner scale. action from depressurizing the cabin
above 13,000 ±1,500 ft with bleed air
A red line on the pressure differential available in the distribution system.
scale indicates the maximum differ-
ential of 8.9 PSI. Emergency
The cabin rate of change indicator
shows the rate at which the cabin alti-
Pressurization
tude ascends or descends. If the pilot selects EMER pressuriza-
tion on the PRESS SOURCE selector
Cabin Altitude Annunciator switch or an overheat condition occurs
The red CABIN ALT 10,000 FT annun- in the ACM, the emergency pressur-
ciator illuminates when cabin altitude ization valve, off the left bleed air
exceeds 10,000 ft; illumination of the cluster, is automatically powered open.
annunciator also triggers the MASTER Both bleed air flow control and shut-
WARNING lights. A barometric pres- off valves are powered closed. On the
sure switch on the left side console acti- ground, the left squat switch input pre-

5K-10 5K-11

5K-22 For SimuFlite training only Citation V

March 1999
 Pneumatic Systems

vents the emergency pressurization ual selection of the emergency pres-

valve from opening. surization or with automatic activation
The emergency pressurization valve due to an ACM overheat. If emergency
routes left engine hot bleed air through pressurization is selected on the ground,
the mixing tube into the cabin where it manually or automatically, the annun-
mixes with the recirculating air to ciator illuminates but the emergency
maintain cabin pressurization during pressurization valve does not open.
emergencies in flight. Mixing Tube
The pilot manually opens the emer-
gency pressurization valve through the A multi-function mixing tube is under
selection of EMER on the PRESS the aft cabin flooring. During emer-
SOURCE selector switch. gency cabin pressurization operation,
bleed air from the emergency pressur-
The overheat sensor in the ACM acti- ization valve ejects through a nozzle
vates emergency pressurization if the into the mixing tube. The mixing tube
temperature exceeds 450°F between through a check valve assembly pulls
the compressor and the secondary heat cabin air into it to mix with the emer-
exchanger. gency pressurization bleed air. This
If ACM temperature reduces within lowers the bleed air temperature as it
12 seconds, the system automatically enters the air distribution system.
returns to normal operation. If the tem- Cabin pressurization remains without
perature does not reduce within 12 the benefit of the ACM.
seconds, rotate the PRESS SOURCE
selector switch to the EMER position Tailcone
to reset the lockout relay. Then rotate
the PRESS SOURCE selector switch
Pressurization
to LH, RH, or NORMAL to return the An air outlet valve eliminates buildup
valves to their normal positions. of pressurized air in the tailcone.
Consider selecting the LH or RH posi- The ram air scoop under the right
tion to reduce the amount of air flow pylon and the scoop on both sides of
through the ACM. the dorsal fin provide pressurized air
for the tailcone.
Emergency A solenoid operates the air outlet valve,
Pressurization which opens when the aircraft is on
Annunciator the ground. A time delay relay allows
a 60-second delay before the tailcone
The EMERG PRESS ON annunciator pressurization valve opens. As the air-
illuminates to notify the crew of an craft leaves the ground, the touchdown
emergency pressurization situation. The relay actuates to drive the valve closed.
annunciator illuminates through man-

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March 1999
5K-24 For SimuFlite training only Citation V
March 1999
 Pneumatic Systems

Preflight Inspection Emergency

Servicing and
Accomplish the preflight of the pneu- Pressurization On
matic, air conditioning, and pressur-
ization systems in accordance with the
The amber EMERG PRESS ON
annunciator illuminates if the ACM
Procedures
Preflight chapter of this manual. shuts down, fails, or the NORM pres-
Prior to flight, check that the ACM oil surization CB disengages. The NORM
level is above the fill line (Figure PRESS CB is on the pilot’s CB panel.
5K-12). View the ACM oil container During an ACM overheat, a time delay
through the tailcone equipment access relay locks the system into emergency
door on the forward bulkhead of the pressurization if the ACM tempera-
aft baggage compartment; use a flash- ture remains too high for 12 seconds or
light if necessary. more. If the ACM cools sufficiently
in less than 12 seconds, the system
Abnormal automatically returns to the previous-
Procedures ly selected mode.
Follow procedural steps to cool the
This section discusses what happens
ACM by allowing more hot bleed air
within the pneumatic, pressurization,
into the cabin; this decreases the quan-
and air conditioning systems during
tity of hot bleed air cooled by the ACM.
abnormal situations. For a list of spe-
cific procedural steps, please refer to High altitude climbs at low airspeeds
the SimuFlite Citation V Operating with low temperature selected on the
Handbook. controls may cause overheating of the
ACM. The combination of an airspeed
Air Duct Overheat increase and warmer cabin tempera-
ture selected corrects the problem.
Cooling down the ventilation duct dom-
inates this procedure. Manually manip- Air Cycle Machine
ulate the temperature controls to force Overpressure
cooler air through the duct and lower
the temperature. After the AIR DUCT If the ACM O’PRESS annunciator
O’HEAT annunciator extinguishes, illuminates while the aircraft is on the
return cabin temperature controls to the ground, seek maintenance to correct
AUTOMATIC mode. the malfunction before the next flight.

5K-12

Citation V For SimuFlite training only 5K-25

March 1999
 If the annunciator illuminates while the Do not dump cabin pressure at high
aircraft is airborne, reduce the amount altitudes; descend before dumping.
of bleed air entering the ACM by
selecting RH on the PRESS SOURCE Emergency
selector. Pull the EMER PRESS CB.
Next pull and reset the NORM PRESS
Procedures
CB. Finally, reset the EMER PRESS The following discusses what happens
CB. The sequence of pulling and reset- within the bleed air, air-conditioning,
ting the CBs prevents emergency pres- or pressurization systems during emer-
surization activation. If the system does gency situations. For a list of specific
not reset, reduce power on the corre- procedural steps, refer to SimuFlite’s
sponding engine to 80%. Citation V Operating Handbook.
Vacuum System Failure Rapid Decompression
An indication of vacuum system fail- (CABIN ALT 10,000 FT
ure is that the cabin pressure goes to Annunciator Illuminated)
maximum differential. Excess cabin Ensure that the crew and passengers
pressure vents overboard with activa- are using oxygen masks with 100%
tion of the max PSID relief valve on oxygen; then execute emergency
the outflow valves. descent according to SOP.
Select LH or RH on the PRESS
SOURCE selector switch and reduce Emergency Descent
power on the corresponding engine to If unable to maintain cabin pressure,
below 70% N1. This reduces airflow reduce engine throttles to idle, extend
into the aircraft, which reduces cabin the speedbrakes, roll into a moderate
differential pressure. The EMER DUMP bank, lower the nose, and execute an
valve is inoperative. emergency descent according to the
Prior to landing, turn the PRESS SOP.
SOURCE selector switch to OFF to
allow cabin pressure to bleed over- Environmental System
board. After landing and before Smoke or Odor
deplaning, rotate the cabin door latch After donning oxygen masks, turn off
slowly; this allows any remaining the OVHD and DEFOG fans to isolate
pressure to escape past the door seal the source of smoke or odor. The pri-
and prevents discomfort to passengers mary action is selecting a single source
and crew. (LH or RH) for bleed air and allowing
time for system purging. If smoke con-
Overpressurization tinues, select EMER to maintain cabin
Decrease cabin pressure by increasing pressurization without use of the ACM.
the cabin altitude and rate of climb. Land as soon as possible.
Also reduce bleed air into the ACM
by selecting a single source to control Smoke Removal
input into the ACM through the engine If smoke is intense, utilize crew and
throttle. passenger oxygen masks at 100% oxy-
With less bleed air entering the ACM, gen. Descend and use the EMER
cabin pressure should reduce. If pres- DUMP valve at a safe altitude to clear
sure remains high, dump the cabin air the cabin.
via the EMER DUMP valve.

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March 1999
 Pneumatic Systems

Bleed Air System Data

Data
Power Source LH/RH engines
Distribution Air cycle machine (ACM) Summaries
LH engine to cabin (EMER position)
Wing deice system
Pressurization controller
Door seal
Engine anti-ice system
Windshield bleed air
Control Pressurization source select (bleed air selector) switch
OFF: Bleed air from either engine cluster assembly.
Pressure for windshield bleed air, cabin pressure
control, cabin door seal, wing deice, and engine
anti-ice available.
GND: Bleed air from right engine cluster assembly
to ACM through ground valve (18 ppm)
LH: Bleed air from left engine cluster assembly to
ACM through left flow control valve (6 ppm)
NORM: Bleed air from both engines cluster assem-
bly to ACM through left and right flow control valves
(12 ppm)
RH: Bleed air from right engine cluster assembly to
ACM through right flow control valve (6 ppm)
EMER: Bleed air from left engine cluster assembly
direct to cabin through emergency pressurization
valve from left engine
Monitor Annunciators
W/S AIR O’HEAT
SURFACE DEICE
BLD AIR GND
EMERG PRESS ON
CABIN ALT 10,000 FT
AIR DUCT O’HEAT
ENG ANTI-ICE (LH/RH)
ACM O’PRESS

Citation V For SimuFlite training only 5K-27

March 1999
 Air Conditioning System Data
Power Source Air cycle machine (ACM)
Ram air
Mixing valve (bypass modulating and shutoff valve)
Distribution Overhead duct
Under-floor duct
Control PRESS SOURCE selector
Auto temp rheostat
Manual heat/cool switch
DEFOG fan
OVHD fan
FLOW DISTR selector
Monitor Cabin temperature indicator
Annunciators
AIR DUCT O’HEAT
EMERG PRESS ON
BLD AIR GND

Pressurization System Data

Power Source Air cycle machine (ACM)
Emergency pressurization from left engine
Distribution Aircraft pressure vessel
Control Pressurization controller
CABIN RATE knob
PRESS SOURCE selector
EMER DUMP valve
Landing gear switch
Throttle switches
Pressure regulator
Control power (28V DC and vacuum)
Monitor CABIN ALT/DIFF PRESS indicator
Cabin rate of change indicator
Annunciators
CABIN ALT 10,000 FT
BLD AIR GND
EMERG PRESS ON
ACM O’PRESS

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March 1999
The Powerplant chapter contains information on the following:
Powerplant
■ powerplant and engine indicating
■ lubrication system
■ fuel and fuel control Chapter 5L
■ ignition
■ bleed air
■ power control.
Two Pratt & Whitney Canada JT15D-5A turbofan engines power
the Citation V. The JT15D-5A is a lightweight, twin-spool
turbofan engine that produces 2,900 lbs of static thrust.

Citation V For SimuFlite training only 5L-1

March 1999
5L-2 For SimuFlite training only Citation V
March 1999
 Powerplant

Pratt & Whitney JT15D-5A

Turbofan Engine Schematic . . . . . . . . . . . . 5L-6 Table of
Turbofan Engines . . . . . . . . . . . . . . . . . . 5L-7 Contents
Components . . . . . . . . . . . . . . . . . . . . . . 5L-7
Low Pressure Compressor Rotor . . . . . . . . . . 5L-7
Low Pressure Compressor Case . . . . . . . . . . 5L-7
High Pressure Compressor Assembly . . . . . . . . 5L-8
Compressor Bleed System (Bleed Off Valve) . . . 5L-8
Combustion Section . . . . . . . . . . . . . . . . . 5L-8
Turbine Section . . . . . . . . . . . . . . . . . . . 5L-9
Exhaust Duct . . . . . . . . . . . . . . . . . . . . . 5L-9
Accessory Gearbox . . . . . . . . . . . . . . . . . 5L-9
Engine Indicating . . . . . . . . . . . . . . . . . . 5L-9
Fan Speed (N1) . . . . . . . . . . . . . . . . . . 5L-9
Turbine Speed (N2) . . . . . . . . . . . . . . . . 5L-10
Interturbine Temperature (ITT) . . . . . . . . . . 5L-10

Lubrication System Schematic . . . . . . . . . . 5L-12

Ignition System Schematic . . . . . . . . . . . . . 5L-12
Powerplant Systems . . . . . . . . . . . . . . . . 5L-13
Lubrication . . . . . . . . . . . . . . . . . . . . . . 5L-13
Components . . . . . . . . . . . . . . . . . . . . 5L-13
Oil Tank . . . . . . . . . . . . . . . . . . . . . . 5L-13
Oil Pump . . . . . . . . . . . . . . . . . . . . . 5L-14
Oil Cooler . . . . . . . . . . . . . . . . . . . . . 5L-14
Oil Filter . . . . . . . . . . . . . . . . . . . . . . 5L-14
Transfer Tubes and Passages . . . . . . . . . . 5L-14
Centrifugal Breather . . . . . . . . . . . . . . . 5L-14
Oil Indicating . . . . . . . . . . . . . . . . . . . . 5L-14
Oil Pressure Gage . . . . . . . . . . . . . . . . 5L-14
Low Oil Pressure Warning . . . . . . . . . . . . 5L-15
Oil Temperature Gage . . . . . . . . . . . . . . 5L-15

Citation V For SimuFlite training only 5L-3

March 1999
 Fuel and Fuel Control . . . . . . . . . . . . . . . . 5L-15
Components . . . . . . . . . . . . . . . . . . . . 5L-16
Fuel Pump . . . . . . . . . . . . . . . . . . . . 5L-16
Fuel Control Unit . . . . . . . . . . . . . . . . . 5L-17
Oil Cooler . . . . . . . . . . . . . . . . . . . . . 5L-18
Flow Divider Valve . . . . . . . . . . . . . . . . 5L-18
Motive Flow Valve . . . . . . . . . . . . . . . . . 5L-18
Fuel Manifold Assembly . . . . . . . . . . . . . 5L-18
Adapter Assemblies . . . . . . . . . . . . . . . . 5L-18
Emergency Fuel Shutoff . . . . . . . . . . . . . 5L-19
Fuel Indicating . . . . . . . . . . . . . . . . . . . 5L-19
Fuel Flow . . . . . . . . . . . . . . . . . . . . . 5L-19
Fuel Pressure Switch . . . . . . . . . . . . . . . 5L-19
Ignition System . . . . . . . . . . . . . . . . . . . . 5L-19
Ignition Exciters . . . . . . . . . . . . . . . . . . . 5L-19
Ignition Cables . . . . . . . . . . . . . . . . . . . 5L-20
Igniter Plugs . . . . . . . . . . . . . . . . . . . . 5L-20
Control and Indication . . . . . . . . . . . . . . . 5L-20
Start Control Relay and Throttle Lever
Cutoff Switch . . . . . . . . . . . . . . . . . . 5L-20
Ignition Switches . . . . . . . . . . . . . . . . . 5L-20
Engine Anti-Ice Switches . . . . . . . . . . . . . 5L-20
Ignition Lights . . . . . . . . . . . . . . . . . . . 5L-21
Bleed Air . . . . . . . . . . . . . . . . . . . . . . . 5L-21
Bearing Sealing . . . . . . . . . . . . . . . . . . . 5L-21
Engine Cooling . . . . . . . . . . . . . . . . . . . 5L-21
Engine Anti-Icing . . . . . . . . . . . . . . . . . . 5L-21
Engine Inlet Anti-Icing . . . . . . . . . . . . . . . 5L-22
Engine Controls . . . . . . . . . . . . . . . . . . . 5L-22
Throttle levers . . . . . . . . . . . . . . . . . . . 5L-22

5L-4 For SimuFlite training only Citation V

March 1999
 Powerplant

Engine Synchronizer . . . . . . . . . . . . . . . . 5L-22 Pratt & Whitney

Ground Idle . . . . . . . . . . . . . . . . . . . . . 5L-23 JT15D-5A
Turbofan Engine
Preflight and Procedures . . . . . . . . . . . . . . 5L-25
Preflight . . . . . . . . . . . . . . . . . . . . . . . 5L-25
Servicing . . . . . . . . . . . . . . . . . . . . . . . 5L-25
Abnormal Procedures . . . . . . . . . . . . . . . . 5L-25
False Engine Start (Engine Does Not Light) . . . . 5L-26
Engine Starter Will Not Disengage . . . . . . . . . 5L-26
Engine Fire During Ground Shutdown . . . . . . . 5L-26
Low Oil Pressure . . . . . . . . . . . . . . . . . . 5L-26
OIL PRESS WARN Illuminated . . . . . . . . . . 5L-26
OIL PRESS WARN Not Illuminated . . . . . . . . 5L-26
Firewall Shutoff Valve Closed . . . . . . . . . . . 5L-26
Emergency Procedures . . . . . . . . . . . . . . . 5L-26
Engine Failure or Fire During Takeoff . . . . . . . 5L-27
Engine Failure/Precautionary Shutdown . . . . . . 5L-27
Engine Failure During Coupled Approach . . . . . 5L-27
Emergency Restart – One Engine . . . . . . . . . 5L-27
Starter Assisted . . . . . . . . . . . . . . . . . . 5L-27
Windmilling with Airspeed above 200 KIAS . . . . 5L-28
Emergency Restart – Two Engines . . . . . . . . . 5L-28
Engine Fire . . . . . . . . . . . . . . . . . . . . . 5L-28

Engine Limitations . . . . . . . . . . . . . . . . . 5L-29

Inter-Turbine Temperature Limits . . . . . . . . . . . 5L-29
Engine Overspeed Limits . . . . . . . . . . . . . . 5L-30
Engine Operating Limits . . . . . . . . . . . . . . . 5L-31

Citation V For SimuFlite training only 5L-5 5L-6

March 1999
Pratt & Whitney JT15D-5A Turbofan Engine

5L-6 For SimuFlite training only Citation V

March 1999
 Powerplant

Turbine engines accelerate and com- Components Turbofan

press air to produce thrust. Thrust can be
Major components of the engine are:
generated two ways: a small volume of
air accelerated to a very high velocity or ■ low pressure compressor rotor Engines
a large volume of air accelerated to a assemblies
lower velocity. ■ low pressure compressor case and

The turbofan engine uses both meth- stator assemblies.

ods. It compresses a small portion of ■ high pressure compressor assembly
the incoming air, mixes it with fuel, ■ combustion section
ignites and burns the fuel/air mixture,
■ turbine section
and exhausts the hot combustion gases.
The fan, in turn, accelerates a large ■ exhaust duct
volume of air at a lower velocity and ■ accessory gearbox.
bypasses it around the engine core.
The relationship of the mass of Low Pressure
bypassed air to the mass of air going Compressor Rotor
through the combustion chamber is the The low pressure compressor rotor
bypass ratio. The Pratt & Whitney of assemblies rotate within the low pres-
Canada JT15D-5A is a medium- sure compressor case, which consists
bypass engine (2.1:1) with a static of a first-stage front fan with a boost-
thrust rating of 2,900 lbs at sea level. er stage disk. The front fan consists of
The turbofan engine is essentially two 19 titanium airfoil-shaped fan blades
separate powerplants. The N1 section mounted on a disk. A dual-walled
uses a fan to move a large volume of compressor inlet cone assembly is on
cool, low velocity air through the the front of the fan disk. The booster
engine bypass duct. This section is stage consists of 50 titanium blades
responsible for providing approxi- mounted on a disk.
mately two-thirds of the total thrust.
Together the front fan and booster disk
The N2 section produces energy in the accelerate air rearward into the engine.
form of high velocity hot air through Approximately two-thirds of the intake
a high pressure turbine. The high pres- air flows to a bypass duct that routes air
sure turbine extracts energy to drive its around the engine. This airflow pass-
compressor and accessory gearbox. The es around the engine core, then com-
low pressure turbine extracts energy to bines with the gas stream exiting the
drive the fan and axial compressor. gas generator at the exhaust duct; the
The remaining energy flows into the remaining one-third enters an axial
exhaust section where it joins with the compressor.
bypass airflow to provide thrust.
Low Pressure
A concentric drive shaft system sup- Compressor Case
ports the fan and turbine rotors. The
inner drive shaft connects the fan and The low pressure compressor case
the booster stage at the front of the consists of a stainless steel housing
engine to the two rear low pressure tur- that contains the front inner, rear inner,
bines. The outer drive shaft connects and bypass stator assemblies. Each sta-
the centrifugal compressor (impeller) tor assembly consists of airfoil-shaped
and the high pressure turbine. guide vanes that direct, straighten, and
divide the airflow exiting the low pres-
sure compressor.

Citation V For SimuFlite training only 5L-7

March 1999
 The front inner stator assembly con- The system consists of a bleed off
sists of 52 guide vanes that direct the valve (BOV), actuator, linkage, bleed
airflow from the front fan to the boost- off valve control (BVC), electrical har-
er stage. The rear inner stator assembly ness, and a BVC-controlled solenoid.
consists of 47 guide vanes that direct The BVC on the engine nacelle uses
the airflow through the intermediate compressor speed (N2), altitude (Po),
casing to the high pressure compres- and gas generator acceleration and
sor. The bypass stator assembly con- deceleration rate inputs to command
sists of 60 guide vanes that direct air- the opening and closing of the BOV.
flow through the bypass ducts. The annular (ring shaped) bleed off
High Pressure valve on the intermediate case sur-
rounds the compressor impeller. The
Compressor Assembly BVC, in response to engine parame-
From the lower pressure compressor, ters, energizes the solenoid. The sole-
low pressure high velocity air enters noid, in turn, releases pressure fuel from
the high pressure compressor assem- the engine-driven fuel pump into the
bly, which consists of an impeller with actuator to move a piston. Movement
an impeller shroud and a diffuser of the piston, translated through a link-
assembly. A stubshaft riding in the No. age, opens the BOV.
2 engine bearing supports the front of The left and right engine BVC operate
the impeller. The high pressure rotor on 28V DC supplied by the aircraft
shaft and the No. 3 bearing support the electrical system. The left BVC
rear of the impeller. receives power from the LH Main bus
The rotating impeller accelerates the Extension through the LH BOV circuit
airflow from the low pressure com- breaker; the right BVC receives power
pressor to create kinetic energy before from the RH Crossover bus through
it directs the air flow toward the dif- the RH BOV circuit breaker. If a BVC
fuser. The diffuser then converts this loses power, the compressor bleed sys-
kinetic energy into high pressure ener- tem fails in the closed position. With
gy at a low velocity before directing the system inoperative, do not adjust
the airflow to the combustion section. the throttle rapidly at high altitudes.
A compressor bleed valve system Combustion Section
releases high pressure compressor bleed
air from the impeller to prevent engine The combustion section is at the rear
surging and compressor stalling. of the gas generator case. It consists of
a combustion liner with small and large
Compressor Bleed System exit ducts. Flowing aft from the dif-
(Bleed Off Valve) fuser, high pressure air flows into the
During rapid engine acceleration and combustion liner. The combustion liner
deceleration, air pressure tends to build turns the air flow 180° and directs it
within the high pressure compressor. forward to the exit ducts. Fuel nozzles
This pressure build-up results in engine introduce fuel into the combustion sec-
surging and compressor stalling usu- tion where the air and fuel mix; the
ally at high altitudes. To prevent this, air/fuel mixture ignites. The hot expand-
a compressor bleed valve (bleed off ing gases flow forward into the exit duct
valve) system releases excess high where they make another 180° turn
pressure compressor bleed air from the before reaching the high pressure tur-
impeller into the bypass duct. bine stator vanes in the turbine section.

5L-8 For SimuFlite training only Citation V

March 1999
 Powerplant

Turbine Section sure and scavenge oil for the No. 4

engine bearing. A T6 temperature probe Fan Speed (N1) Limitations
The turbine section consists of a single- is in the leading edge of each strut. Takeoff (5 minutes) . . . 104%
stage high pressure turbine and a two-
stage low pressure turbine. The high A housing and cover assembly in the Maximum Continuous . . 104%
and low pressure turbines turn on sep- center of the ducts contains the No. 4 Transient . . . . . . . . . 106%
arate concentric shafts. The high pres- engine bearing. The overspeed control
sure turbine drives the impeller through actuating rod extends through the cover
the outer shaft and the low pressure assembly.
turbine drives the low pressure com-
Accessory Gearbox
pressor (front fan and booster stage)
through the inner shaft. A vertical tower shaft drives the acces-
The high pressure turbine inlet guide sory gearbox on the bottom of the
vanes direct the gas stream from the intermediate engine casing. The tower
combustion section at a correct angle shaft has bevel gears on either end;
to efficiently drive the high pressure one end meshes with a bevel gear on
turbine. The high pressure turbine con- the high pressure compressor rotor
sists of 71 airfoil-shaped blades mount- shaft and the other meshes with a bevel
ed to a disk. As the high velocity gas gear on the starter/generator drive shaft
stream passes and expands through the in the accessory gearbox.
turbine, the turbine rotates and extracts The accessory gearbox drives the:
energy to drive the impeller. ■ starter/generator

After passing through the high pres- ■ fuel pump

sure turbine, the gas stream reaches
■ hydraulic pump
the two low pressure turbines. Each
turbine stage consists of inlet guide ■ oil pump

vanes followed by a turbine disk. The ■ N2 tachometer/generator.

guide vanes straighten the gas flow
before it reaches the turbine disks. The Engine Indicating
high velocity gases flow through the
Engine indicating includes:
first and second-stages to rotate the
■ fan speed (N1) indicator
turbines. The low pressure turbines, in
turn, drive the low pressure compres- ■ turbine speed (N2) indicator
sor through a shaft. After exiting the ■ interturbine temperature (ITT).
turbine section, the gas stream enters
the exhaust duct. Fan Speed (N1)
Exhaust Duct The N1 speed sensor probe, driven by
the low pressure compressor shaft,
The exhaust duct consists of a front drives a tachometer/generator that sup-
and rear turbine exhaust duct inner plies a pulsating voltage. The voltage’s
assembly and a front and rear outer frequency is proportional to the rota-
bypass duct assembly. The thrust tional speed of the fan. This signal
reverser attaches to the rear of the fan drives the cockpit fan speed indicator.
duct assembly. Hot gases from the tur-
bine section flow through the front and The fan speed indicator (Figure 5L-1,
rear turbine exhaust duct assemblies following page) is a dual-reading
where they mix with cold air flow vertical tape display with two digital
directed by the bypass duct assemblies. displays that show fan speed as a per-
centage of maximum RPM. Each tape
Six evenly spaced hollow struts sup- displays fan speed from 20 to 110%
port the inner and outer assemblies. Two RPM; the digital displays show fan
of the struts provide a passage for pres- speed from 0 to 110% RPM. A green

Citation V For SimuFlite training only 5L-9

March 1999
 colored band on the tape indicates the right indicator. If either indicator loses
Turbine Speed (N2) normal operating fan speed from 25 power, the respective display blanks.
Limitations to 104.0% RPM; a red radial line at
Takeoff (5 minutes) . . . . 96% 104.0% RPM indicates the maximum Interturbine Temperature (ITT)
Maximum Continuous . . 96% operating speed. The ITT gage (Figure 5L-3) provides
The transducer and indicator operate a synthetic or calculated indication of
Flight Idle (minimum) . . . 52%
on 28V DC; the left and right fan speed engine interturbine temperature (ITT).
Ground Idle (minimum) . . 46% A series of chromel/alumel tempera-
indicating system receives power from
Transient . . . . . . . . . . 96% the Emergency bus through the LH ture probes and thermocouples sense
and RH FAN SPEED circuit breakers. temperature in the bypass duct, low
If the left or right indicator loses power, compressor case, and the exhaust
the tape and display blank. nozzle.
ITT Limitations
T1 thermocouples in the inlet and
Takeoff (5 minutes) . . . 700°C Turbine Speed (N2)
exhaust area of the bypass duct pro-
Maximum Continuous . . 680°C A tachometer/generator driven by the vide an average temperature reading
Flight Idle . . . . . . . . 580°C accessory gearbox provides a pulsating equal to three times the temperature
voltage with a frequency proportional rise through the bypass duct. T6 ther-
Ground Idle . . . . . . . 580°C to turbine speed. This signal drives the mocouples in the leading edges of the
Starting (2 seconds) . . . 700°C cockpit indicator. exhaust case nozzle struts connect in
Transient . . . . . . . . . 700°C The turbine speed indicator (Figure parallel to provide an average tem-
5L-2) is a dual-reading digital display perature reading of the gas stream exit-
that indicates turbine speed as a per- ing the turbine section. The ITT indi-
centage of maximum RPM. Each cating system provides a simulated
three-digit digital displays shows N2% temperature by adding the tempera-
RPM to 0.1%. At the maximum tur- ture rise sensed by the T1 thermo-
bine speeds of 96% N2 RPM, the dig- couples to the T6 temperature.
its and a red light flash as warning. The dual-reading vertical scale ITT
The tachometer/generator and indica- gage on the top left center of the instru-
tor operate on 28V DC supplied by the ment panel displays left and right
normal electrical system. The LH Main engine temperature. The vertical tape
bus Extension through the LH TURB displays ITT from 150 to 800°C. A
SPEED circuit breaker powers the left green colored band from 150 to 680°
engine indicator; the RH Main bus indicates the normal operating range
Extension through the RH TURB of the engine. A yellow band starts at
SPEED circuit breaker powers the the maximum continuous ITT of 680°

5L-1 5L-2 5L-3

5L-10 For SimuFlite training only Citation V

March 1999
 Powerplant

and continues to 700°C. A red radial through the LH ITT circuit breaker; the
line at 700°C denotes the maximum right engine indicator receives power
Lubrication
takeoff and start ITT. from the RH Main bus Extension System
The indicator operates on 28V DC sup- through the RH ITT circuit breaker. If
plied by the normal electrical system. the indicator loses 28V DC, a red OFF Ignition
flag appears above the respective tape
The left engine ITT tape receives power
display. System
from the LH Main bus Extension

Citation V For SimuFlite training only 5L-11 5L-12

March 1999
Lubrication System Ignition System
BEARINGS
1,2,3,31/2
TO BEARINGS1,#2,#3
IGNITER

RESERVOIR

CUTOFF

START IGNITION
CIRCUIT BREAKER RFI FILTER
B NO. 2

GRAVITY FEED TO GEAR BOX

PRESSURE A EXCITER
RELIEF VALVE PUMP ASSEMBLY THROTTLE LEVER SWITCH FUSE

OIL PRESSURE
COOLER ELEMENT START CONTROL RELAY

IGNITION LIGHT
RH/LH MAIN DC BUSES #4 BEARING

TEMP IGNITION
SCAVENGE
SWITCH
PRESS
PRESS

ELEMENT
(SPDT)

ON
OIL IGNITION FAIL
FILTER SYSTEM DRAIN OFF LIGHT RELAY

NORM
TO BEARINGS
1,#2,#3 IGNITION CIRCUIT
BREAKER (COCKPIT)
SCAVENGE ELEMENT ON
TO BEARINGS 4 AND 31/2 OFF RFI FILTER
FUSE
ON
A
ENGINE/WING NO. 1
ENGINE/WING
ANTI-ICE (COCKPIT) OFF TO ANTI-ICE VALVES B EXCITER
O
O 100 O O O 120
F O ENGINE
I F I I F I
L 80 F L L
100
L ANTI-ICE
OIL PRESS P P T
80
T
SWITCH
60
WARN R R E 60 E (DPDT)
E E M M
S 40 S P 40 P
LH RH S S
20 L 20 R
L R
0
0 °C
PSI
IGNITER

SUPPLY

PUMP PRESSURE

RETURN

5L-12 For SimuFlite training only Citation V

March 1999
 Powerplant

Powerplant systems include:

■ lubrication
of the pump draws oil from the sumps
to the oil tank. Powerplant
■ fuel and fuel control The breather system equalizes excess
air from the lubrication system to the
Systems
■ ignition
atmosphere. This aids in lubrication
■ bleed air and anti-icing by preventing the build-up of air in the
■ engine controls. engine that could cause pump cavita-
tion and excessive pressure.
Lubrication Engine oil indicating system is dis-
cussed with the oil system.
The lubrication system supplies oil
under constant pressure to the engine Components
and accessory gearbox for lubrication,
cooling, and cleaning. The lubrication The oil system consists of:
system consists of three interrelated ■ oil tank
systems: pressure oil, scavenge oil, and ■ oil pump
breather.
■ fuel-to-oil cooler
The pressure system uses an engine-
■ oil filter
driven pump to pressurize oil from a
tank. From the pump, oil passes ■ transfer tubes and passages
through a fuel-to-oil cooler that cools ■ centrifugal breather.
the oil before it reaches a filter that
removes impurities. After passing Oil Tank
through the filter, internal and external The oil tank has a capacity of 2.03 U.S.
transfer tubes and passages route oil to gallons (7.58 liters) of which 1.20 gal-
the engine bearings, bevel and spur lons (4.54 liters) are usable. The tank
gears, and the accessory gearbox. is an integral part of the intermediate
The scavenge system returns oil from casing and has an oil filler neck with a
the engine to the oil tank. Oil drains dipstick and cap assembly (Figure
from the bearings into the accessory 5L-4). Marks on the dipstick indicate
gearbox and from the No. 4 bearing the maximum hot quantity (MAX
into a second sump. The scavenge side HOT) and maximum cold quantity

5L-4

Citation V For SimuFlite training only 5L-13

March 1999
 (MAX COLD). Additional marks indi- Opening and closing of the pressure
Oil Pressure Limitations cate the number of quarts to be added relief valve maintains a fairly constant
Takeoff . . . . . 60 to 83 PSIG to reach the full mark. oil pressure by returning excess oil to
Maximum the oil pump inlet.
Oil Pump
Continuous . . . 60 to 83 PSIG Transfer Tubes and Passages
An engine-driven oil pump on the inter-
Idle (minimum) . . . . 40 PSIG
mediate casing pressurizes oil from the External transfer tubes and internal
Transient . . . . . . . . 95 PSIG oil tank for circulation through the passages carry oil from the oil pump to
engine and returns oil from the engine the engine bearings and accessory
to the oil tank. The three-element oil gearbox. Additional tubes and passages
pump consists of a single-element pres- carry oil from the bearings to the acces-
sure side and a dual-element scavenge sory gearbox and bearing sumps back
side driven by a common shaft. The to the scavenge side of the oil pump.
pump uses the intermeshing of gears
to draw oil from the oil tank (pressure Centrifugal Breather
pump) and the accessory gearbox The centrifugal breather uses a gear-
(scavenge pump). box-driven impeller to separate air and
oil. The oil particles adhere to the
Oil Cooler
impeller; as the impeller rotates, it
An oil cooler on the filter housing cir- slings the particles outward. The rel-
culates oil and fuel through separate atively oil free air passes through the
passages to cool hot oil flowing from hollow impeller shaft to a breather
the pump to the filter. Hot oil flowing adapter where a tube carries it over-
through the cooler transfers heat to the board for venting.
fuel. Fuel under pressure comes from
the fuel control unit (FCU). If the oil Oil Indicating
passages clog, a bypass valve opens Oil indicating systems include:
to route oil around the cooler.
■ oil pressure gage

Oil Filter ■ low oil pressure warning

The oil filter and housing are on the ■ oil temperature gage.
accessory gearbox. The housing con-
tains a removable filter element, two Oil Pressure Gage
check valves, a pressure relief valve, An oil pressure transmitter on the for-
and a bypass valve. The housing also ward cowling bulkhead assembly of
contains bosses for the oil temperature each engine senses oil pressure from a
and pressure transmitters and a mount boss on the oil filter housing. Each
for the oil cooler. transmitter electrically drives the oil
The removable oil filter has a 10-micron pressure gage. The oil pressure gage
filter element that can be removed and (Figure 5L-5) is a dual-reading vertical
cleaned. If the filter begins to clog, oil scale indicator that displays oil pres-
pressure overcomes spring pressure to sure for each engine from 0 to 100 PSI.
open the bypass valve; oil flow bypass- Colored bands and a radial line indi-
es the filter. cate the minimum, warning, and nor-
mal operating pressures. A red radial
Two check valves in the filter housing
line at 40 PSI indicates the minimum
prevent the gravity flow of oil from
oil pressure at engine idle. A yellow
the engine. Without these check valves,
band from 40 to 60 PSI indicates the
oil tank draining would be necessary
warning zone; pressures below 60 PSI
when changing the oil filter or remov-
are undesirable. A green band from 60
ing transfer tubes.
to 83 PSI is the normal operating range.

5L-14 For SimuFlite training only Citation V

March 1999
 Powerplant

The left engine oil pressure gage and A red radial line at 121°C indicates the
transmitter receives 28V DC from the maximum operating oil temperature. Oil Temperature Limitations
LH Main bus Extension through the The transmitter and gage operate on Takeoff . . . . . . . 10 to 121°C
LH OIL PRESS circuit breaker; the 28V DC. The left indicating system Maximum
right engine system receives power receives power from the LH Main bus Continuous . . . . 10 to 121°C
from the RH Main bus Extension Extension through the LH OIL TEMP Idle . . . . . . . . -40 to 121°C
through the RH OIL PRESS circuit circuit breaker; the right system receives
breaker. If the system loses power, a Starting (min) . . . . . . -40°C
power from the RH Main bus Extension
red OFF flag appears above the affect- through the RH OIL TEMP circuit Transient . . . . . -18 to 129°C
ed scale. breaker. If the system loses DC power,
Low Oil Pressure Warning a red OFF flag appears above the affect-
ed scale.
An oil pressure switch on each
engine’s oil filter housing and annun- Fuel and Fuel
ciators in the cockpit warn of low oil
pressure conditions. If an engine’s oil Control
pressure drops to approximately 35 The fuel and fuel control system sup-
PSIG, the pressure switch illuminates plies metered fuel under pressure to
the respective red OIL PRESS WARN atomize and mix fuel with air in the
annunciator; illumination of the annun- combustion chamber. The system reg-
ciator also triggers the MASTER ulates the flow of fuel to the engine
WARNING lights. The annunciator under all operating conditions and
extinguishes once oil pressure reaches engine speeds.
approximately 47 PSIG.
Fuel under pressure from the wing fuel
The low oil pressure warning system system enters the engine fuel system
receives 28V DC from the Main buses; from the firewall shutoff valve in the
there is no circuit protection. wing sump area. From the shutoff valve
Oil Temperature Gage a flexible fuel line carries fuel aft to a
special T-fitting at the engine firewall.
A temperature transmitter on each oil The T-fitting connects to a rigid line
filter housing senses oil temperature that carries fuel to the engine fuel pump.
for display on the cockpit indicators.
The engine fuel pump receives fuel at
The dual-reading vertical scale tem-
approximately 20 to 30 PSI and
perature gage (Figure 5L-6) displays
increases fuel pressure to approxi-
engine oil temperature from 0 to 130°C.
mately 500 to 700 PSI. A fuel filter
A green colored band from 0 to 121°C
downstream of the fuel pump removes
indicates the normal temperature range.
impurities. After travelling through the

5L-5 5L-6

Citation V For SimuFlite training only 5L-15

March 1999
 filter, fuel enters the fuel control unit ■ adapter assemblies
(FCU). The FCU regulates engine N2 ■ emergency fuel shutoff.
speed by varying fuel flow.
From the FCU, fuel flows under pres- Fuel Pump
sure through the flow meter and oil The accessory gearbox-driven fuel
cooler before reaching the flow divider pump contains a centrifugal boost stage
valve. The flow divider valve splits and a spur-gear positive displacement
the fuel flow into two streams: prima- unit. The pump also has two fuel fil-
ry and secondary. The primary and sec- ters: a 74 micron inlet screen and a 10
ondary fuel streams travel through the micron discharge filter. If the screen
dual manifold to the fuel nozzles. The clogs, fuel pressure unseats the screen
fuel nozzles provide an atomized spray to allow fuel flow past it. If the filter
of fuel into the combustion chamber. clogs, the bypass valve opens to allow
An emergency fuel shutoff system uses fuel flow around the filter.
a mechanical linkage to cut fuel to the From the aircraft fuel system, fuel
engine and shut down the engine if a enters the impeller boost stage of the
mechanical failure occurs within the engine-driven fuel pump. The boost
engine. stage increases fuel pressure to prevent
cavitation of the positive displacement
Components unit (gear cavity). Fuel passes through
The engine fuel system consists of: the inlet filter screen, then enters the
■ fuel pump
gear cavity. Intermeshing gears with-
in the cavity, combined with the boost
■ fuel control unit (FCU)
stage, increase fuel pressure from 20
■ oil cooler to 30 PSI at the fuel pump inlet to 500
■ flow divider valve
to 700 PSI at the outlet depending on
engine speed. After passing through
■ motive flow valve
the gear cavity, the discharge filter
■ fuel surge reservoir (Figure 5L-7) removes impurities from the fuel before
■ fuel manifold assembly
allowing it to travel to the FCU.

BLEED
AIR
SHUTOFF

CONTROL
FROM
FUEL
TANK ENGINE FUEL FUEL
FUEL OIL FUEL FUEL
DRIVEN CONTROL FLOW COOLER SURGE DIV. NOZZLES
PUMP UNIT RESERVOIR
AUTOMATIC
FUEL
MOTIVE TO SHUTOFF
FLOW FLOW
VALVE METER
IGNITION STEP
MODULATOR
MOTIVE FLOW
TO FUEL TANK

5L-7

5L-16 For SimuFlite training only Citation V

March 1999
 Powerplant

Excess fuel from the pump reroutes sure to regulate three bellows within
back to the fuel pump bypass port to the section. Movement of the bellows in
mix with fuel from the centrifugal response to pressure changes controls
boost stage. the position of the metering valve to
regulate fuel flow. During engine oper-
Fuel Control Unit ation, pressures increase and decrease in
The hydro-mechanical FCU receives response to engine acceleration and
fuel from the pump at approximately deceleration. As pressures build, the
500 to 700 PSI. Through a splined cou- bellows open the metering valve to
pling, the fuel pump transmits a speed increase fuel flow. As the engine decel-
signal proportional to engine com- erates, pressures decrease to close the
pressor turbine speed (N2) to the gov- metering valve. The action of this sec-
erning section of the FCU. tion controls the operation of the meter-
The FCU controls engine N2 speed by ing valve, which provides efficient fuel
varying fuel flow to the engine; scheduling for engine operation.
increasing fuel flow increases engine The power input and speed governing
speed and vice versa. The FCU meters section uses a flyweight and springs
fuel for engine starting, acceleration, to operate three pneumatic orifices. It
governing, altitude compensation, operates one orifice in response to the
deceleration, idle, bleed off valve acti- flyweight force and two orifices in
vation, and engine shutdown. response to the difference between a
The FCU has three separate sections; selected engine speed and actual
it consists of: engine speed. This section regulates
pressures supplied to the fuel com-
■ fuel metering section
puting section in response to engine
■ computing (air) section speed and desired engine speed.
■ power input and speed governing A temperature sensor (T2) in the
section bypass duct monitors ambient air tem-
■ temperature sensor perature. The sensor reacts to ambient
air temperature changes by bleeding
■ dual idle solenoid.
off more or less Px air to the FCU.
The fuel metering section contains a The dual idle reset solenoid in the FCU
mechanically actuated metering valve provides two engine idle settings: ground
that varies fuel flow by increasing or idle and flight idle. The two-position
decreasing the area of an orifice. The (NORM/HIGH) GND IDLE switch
action of the metering valve and a (Figure 5L-8, next page), a delay timer,
bypass valve establish fuel flow to the and the landing gear squat switches con-
engine. The bypass valve routes excess trol the solenoid. With the switch in
fuel back to the fuel pump to maintain NORM, the solenoid activates eight sec-
a constant fuel pressure differential onds after the squat switches activate
across the metering valve. The fuel (weight on wheels) to reduce engine idle
metering section also uses a high pres- to 46% N2 RPM. The lower idle RPM
sure relief valve to prevent excess pres- reduces thrust from the engine to slow
sure in the FCU by rerouting fuel back taxi speeds. With the switch in HIGH
to the fuel pump. Fuel leaving the or weight off the wheels, the ground idle
metering valve is referred to as metered system deactivates to provide a normal
fuel (P2). engine idle speed of not less than 52%
The fuel computing (air) section uses N2 RPM. Once the ground idle system
acceleration bellows (Px), governor bel- activates, the GROUND IDLE annun-
lows (Py), and compressor air (Pc) pres- ciator illuminates.

Citation V For SimuFlite training only 5L-17

March 1999
 Oil Cooler to the manifolds. With the shutoff lever
closed, the shutoff and windmill bypass
The oil cooler uses fuel from the FCU
valve routes fuel back to the fuel pump.
to cool engine oil before it reaches the
engine bearings and accessory gear- Motive Flow Valve
box. Oil and fuel flow through sepa-
rate passages within the cooler where A motive flow valve on the bracket
the hot oil exchanges heat with the cold that supports the flow divider valve
fuel. A bypass valve within the cooler uses excess fuel pressure (P1) to oper-
bypasses oil around the passages if the ate a wing sump primary ejector fuel
cooler clogs. pump. The motive flow valve supplies
high pressure, low volume fuel at
Flow Divider Valve approximately 200 to 250 PSI to the
ejector. The ejector fuel pump supplies
After passing through the oil cooler,
high volume low pressure fuel to the
metered fuel enters the flow divider
engine during normal operation.
valve. The flow divider valve splits
the fuel flow into two; one for the pri- Fuel Manifold Assembly
mary manifold and the other for the
secondary manifold. The divider valve The fuel manifold assembly delivers
also controls fuel pressure to the pri- fuel from the flow divider valve to the
mary manifold during engine start, primary and secondary passages of the
dumps fuel from the manifolds during fuel nozzles. The manifold consists of
engine shut down, and ensures that fuel 12 dual-orifice (primary and sec-
does not reach the manifolds until it ondary) adapter assemblies connect-
reaches a minimum pressure. ed by transfer tubes. Two bolts on each
adapter assembly secure the entire
During engine start the flow divider manifold assembly to the low pressure
valve provides fuel only to the prima- turbine stator support assembly.
ry manifold; any excess fuel is routed
back to the fuel pump inlet. As engine Adapter Assemblies
speed increases toward idle, the divider Each of the 12 adapter assemblies con-
valve begins supplying fuel to the sec- tains a dual-orifice fuel nozzle with a
ondary manifold until both the prima- swirl-type tip and a sheath. Each fuel
ry and secondary manifolds receive nozzle provides a finely atomized
the same fuel flow. spray of fuel into the combustion
A manually operated shutoff lever on chamber. Compressor discharge air
the flow divider valve allows mainte- flowing around the sheath assists in
nance personnel to shut FCU fuel flow fuel atomization and nozzle cooling.

5L-8 5L-9

5L-18 For SimuFlite training only Citation V

March 1999
 Powerplant

Emergency Fuel Shutoff illuminates a FUEL LOW PRESS

annunciator and activates a boost pump
The emergency fuel shutoff system pre-
when the fuel pressure drops below
vents engine overspeed by shutting off
4.65 PSIG.
the fuel flow to the engine. The sys-
tem consists of a mechanical linkage If fuel pressure falls below 4.65 PSIG
between the No. 4 bearing housing and with the throttle cutoff switch out of
the fuel inlet and shutoff valve assem- cutoff, the respective FUEL LOW
bly on the fuel manifold. Axial dis- PRESS annunciator illuminates. If the
placement aft of the low pressure tur- boost pump switch is in NORMAL,
bine shaft activates a plunger in the the boost pump automatically activates
shutoff valve piston assembly. The to increase fuel pressure to the engine.
shutoff valve piston assembly activates
the fuel inlet and emergency fuel shut- Ignition System
off valve to stop fuel flow to the engine. During the engine start cycle, the igni-
Fuel Indicating tion system provides high-energy dis-
charges for fuel/air mixture ignition.
The fuel indicating system consists of: The system consists of:
■ fuel flow transmitters ■ ignition exciters

■ fuel flow gages ■ ignition cables

■ fuel pressure switches and annun- ■ igniter plugs.

ciators.
During the engine start cycle, the Hot
Fuel Flow Battery bus supplies power for the left
and right engine ignition systems. With
A fuel flow transmitter in the fuel line the ignition switches or engine anti-ice
between the oil cooler and FCU on each switches ON, the RH Crossover bus
engine senses fuel flow from 145 to and the LH Main Bus Extension from
1800 pounds-per-hour (PPH). The the LH and RH IGN circuit breakers
transmitters drive a dual-reading verti- supply 28V DC to the left and right
cal scale fuel flow gage (Figure 5L-9) engine ignition exciters. The LH and
on the center instrument panel. The RH IGN circuit breakers on the LH cir-
gage displays fuel flow for each engine cuit breaker panel must be closed before
from 0 to 2,000 PPH. Throttle cutoff the ignition or engine anti-ice switch-
switches disable the indicators below es can operate the ignition system.
10% N2 RPM.
The transmitters and fuel flow gage Ignition Exciters
operate on 28V DC supplied by the air- Each engine has two ignition exciters:
craft electrical system. The left engine one on the outer bypass duct and one
transmitter and indicator receive power on the intermediate case. Each exciter
from the LH Main bus Extension powers one igniter plug. The exciters
through the LH FUEL FLOW circuit operate on 9 to 30V DC.
breaker; the right engine system receives
power from the RH Main bus Extension The exciters are capacitance discharge
through the RH FUEL FLOW circuit units that use a step-up transformer to
breaker. If either system loses 28V DC generate and a capacitor to store a four
power, an OFF flag appears above the joule charge at approximately 3,000 volts
affected scale. for the igniters. Once the exciter creates
a sufficient charge, it discharges to fire
Fuel Pressure Switch the igniter. The design of the ignition
system ensures that if one exciter fails,
A fuel pressure switch in each engine’s
the other exciter continues to operate.
fuel line performs a dual function: it

Citation V For SimuFlite training only 5L-19

March 1999
 Ignition Cables unit to energize the start control relay;
the start control relay closes.
High tension leads carry the electrical
output from the ignition exciters to the Advancing the throttle lever out of
igniter plugs. Each consists of an elec- IDLE CUT-OFF closes the throttle
trical lead in a flexible metal braiding. lever cutoff switch. Power then flows
from the Hot Battery bus through the
Igniter Plugs START IGNITION POWER circuit
breaker (aft J box), throttle lever cut-
The igniter plugs are at the 5 and 7
off switch, and the closed start control
o’clock positions on the bypass duct.
relay to the ignition exciters; the sys-
From the bypass duct, the igniter plugs
tem operates. The ignition light illu-
protrude through the gas generator case
minates to indicate the flow of power
and into the combustion chamber.
to both ignition exciters.
Control and Indication Once the engine reaches approximately
40% N2 RPM, the starter/ generator
Controls for the ignition system include:
speed sensor signals the generator con-
■ start control relay
trol unit to terminate the start sequence.
■ throttle lever cutoff switch The generator control unit opens the
■ ignition switches
start control relay cutting power to the
starter/generator.
■ engine anti-ice switches.
Ignition Switches
A green ignition light above each igni-
tion switch provides a visual indication The two-position (ON/NORM) igni-
of proper system operation. tion switches (Figure 5L-10) on the
pilot’s switch panel supply power to
Start Control Relay and the ignition systems. In ON, the
Throttle Lever Cutoff Switch
switches provide 28V DC to the sys-
The start control relay, controlled by tem from the LH and RH IGN circuit
the generator control unit, supplies breakers; the igniters operate contin-
power from the Hot Battery bus uously. Placing the switch in NORM
through the throttle lever cutoff switch (off), interrupts power to the igniters.
during engine start cycle. The normal operating position for the
Pressing a starter switch supplies bat- switches is NORM.
tery power to the auxiliary start relay. Engine Anti-Ice Switches
The auxiliary start relay closes to per-
mit power from the generator control Positioning the engine anti-ice switch-
es (Figure 5L-11) ON activates the

5L-10 5L-11

5L-20 For SimuFlite training only Citation V

March 1999
 Powerplant

engine inlet and inboard wing leading Bearing Sealing

edge anti-ice system as well as the
engine ignition system. Compressor discharge air flowing
through the bearing labyrinth seals
With the engine anti-icing switches ON contains oil within the bearing com-
and the ignition switches in NORM or partments. Each labyrinth seal consists
OFF, the ignition system operates con- of a runner and a labyrinth that form a
tinuously. Power flows from the LH seal to control the pressure gradient
and RH IGN circuit breakers through within each bearing compartment. The
relays controlled by the engine anti-ice bleed air flowing through the seals also
switches to the ignition exciters. assists in oil scavenging by forcing
Activation of continuous engine igni- excess oil through the passages that
tion during icing conditions with the carries oil to the bearing sumps.
engine anti-icing switches ON prevents
engine flameout due ice ingestion. Engine Cooling
Ignition Lights Compressor discharge air cools the
high pressure turbine disk, second-stage
Fuses between the power source and
and third-stage turbine, and the high
ignition exciter control the green igni-
pressure turbine stator vanes. After
tion lights above each ignition switch.
cooling the turbine disks, the air dis-
During normal operation, both lights
charges into the main gas stream
illuminate. If one of the fuses open with
through the engine. The stator vanes
the ignition system operating, the light
receive compressor discharge air for
extinguishes. Check the respective IGN
cooling through small holes in the small
circuit breaker. The design of the ignition
exit duct. The air flows through the
system ensures that if one exciter fails,
holes and enters the stator vanes
the other system continues to operate.
through the outer stator vane ring. Air
flowing through the vanes provides
Bleed Air cooling before being discharged
The engine bleed air system uses high through holes in the vane trailing edges.
pressure bleed air to provide:
Engine Anti-Icing
■ engine bearing sealing

■ engine cooling
The engine bullet nose cone and T1
thermocouple are continuously anti-
■ engine and airframe ice and rain iced by engine bleed air whenever the
protection (anti-icing) engine is operating. Each receives
■ engine fuel system operation bleed air from different sources.
■ air conditioning, pressurization, and Compressor discharge air travelling
temperature control down a hollow low pressure com-
■ pressurization control
pressor shaft enters the double wall
inlet cone where it provides anti-icing
■ door sealing.
protection. The bleed air travels rear-
Two taps on the outer bypass duct pro- ward where holes in the inner wall
vide high pressure compressor (Pc) allow it to discharge into the engine
bleed air for airframe systems (see Air bullet nose cone cavity. From the cone
Conditioning and Pressurization); this cavity, it flows to the No. 1 bearing
discussion concerns bleed air used by labyrinth seal.
the engine. A tube assembly and solenoid shutoff
valve carry bleed air from the gas gen-

Citation V For SimuFlite training only 5L-21

March 1999
 erator case to the intermediate casing. Turning the engine anti-ice switch OFF
From the casing, the air travels through or positioning the throttle lever below
the outer strut into the inner bore of the 60% N2 position applies power to
the intermediate case. An annular the solenoid valves; the valves close
(ring-shaped) anti-icing air manifold to stop air flow to the engine inlet duct
then routes the air to the hollow stator and inner stator vanes.
vanes. After travelling through the A temperature sensor monitors bleed
vanes, the air discharges around the air temperature in the inlet duct.
leading rotary shroud where it mixes Normally, the switch opens at approx-
with the primary air stream. imately 220°F to extinguish its respec-
An external transfer tube on the engine tive ENG ANTI-ICE annunciator when
bypass duct routes warm high pressure the system is operating. If its tempera-
compressor air from the gas generator ture falls to approximately 200°F, the
case for T1 thermocouple anti-icing. The sensor closes to illuminate the annun-
air then travels through the probe where ciator. Additionally, the inner stator anti-
it provides leading edge anti-icing. After ice valve position switch extinguishes its
exiting the probe, the air mixes with air respective ENG ANTI-ICE annunciator
travelling through the engine. when the solenoid valve is open.
Engine Inlet Anti-Icing Engine Controls
The engine inlet anti-icing system uses Engine controls includes:
compressor bleed air to warm the
■ throttle levers
engine inlet duct.
■ engine synchronizer
A stainless steel duct carries com-
pressor bleed air from the engine to an ■ ground idle system.

electrically-closed anti-ice solenoid

valve. The solenoid valve controls Throttle Levers
bleed air flow to the engine inlet duct. The throttle levers (Figure 5L-12)
The anti-ice solenoid valve closes with manually connect through cables and
the application of DC power and opens bellcranks to their respective FCU.
once power is removed. If there is a Each throttle lever controls the FCU
power interruption and pressure exists from the fuel cutoff to full throttle posi-
at the valve, the valve opens to provide tions. Throttle travel ranges from CUT-
inlet anti-icing. The LH and RH OFF (full aft) through IDLE to FULL
ENGINE anti-ice switches on the left forward or maximum thrust. A
instrument panel control the engine mechanical CUTOFF stop prevents
inlet anti-icing system. Turning a inadvertent selection of this position. A
switch on with engine power above the latch on the throttle must be raised
60% N2 position supplies power to a before it can be moved to or from the
thermal delay relay heating element. CUTOFF position.
After five seconds, the relay opens to
remove power to the solenoid valve; Engine Synchronizer
the valve opens to supply bleed air to The engine synchronizer provides auto-
the engine inlet anti-icing system. matic N1 or N2 RPM matching of the
Air flowing through the valve enters right (slave) engine to the left (master)
the inlet manifold where it heats the engine. A turbine out-of-sync condition
double-walled engine inlet duct. It then is generally more noticeable in the cock-
travels through the anti-icing air ducts pit; a fan out-of-sync condition is
where it flows rearward and mixes generally more noticeable in the pas-
with the engine air flow. senger cabin.

5L-22 For SimuFlite training only Citation V

March 1999
 Powerplant

The synchronizer consists of a syn- not match, the control box activates the
chronizer control box, speed setting speed setting actuator. Driven by the Ground Idle Limitations
actuator, a trimming assembly, and a actuator, the trimming and rotary shaft The ground idle switch must be in
flexible rotary shaft assembly. Mono- assemblies adjust the slave engine’s HIGH position when conducting
poles on the engine fans and turbines FCU to match the engine’s turbine or touch and go landings.
provide an alternating current to the fan speed to the master engine.
control box that corresponds to the fan The actuator has a 1.5% N1 RPM or
and turbine speeds of both engines. 1.0% N2 RPM range of authority. This
The control box in the control pedestal limited range of authority prevents the
operates on 28V DC supplied by the slave engine from losing more than a
LH Main bus Extension through the fixed amount of power as it tries to syn-
ENG SYNC circuit breaker. A three- chronize with a failing master engine.
position (FAN/OFF/TURB) rotary If the system fails to synchronize the
switch (Figure 5L-13) below the throt- engines, the actuator has reached its
tle levers controls the synchronizer. full range of travel; turn the system
The FAN position synchronizes engine OFF and manually synchronize the
N1 RPM; TURB synchronizes engine engines with the throttle levers.
N2 RPM, and OFF deactivates the sys-
tem and drives the speed setting actu- Ground Idle
ator to its center range before stopping. The ground idle switch must be in
An indicator light next to the switch HIGH position when operating on the
illuminates during system operation. ground with engine anti-ice bleed ON.
After manually synchronizing the The ground idle system reduces engine
engines, select either FAN or TURB speed from a flight idle of 52% N2
to activate the synchronizer. The system RPM to 46% N2 RPM. Reducing the
then compares the slave engine to the engine idle speed slows taxi speed and
master engine. If the slave engine does results in decreased brake wear.

5L-13

5L-12

Citation V For SimuFlite training only 5L-23

March 1999
 The two-position (NORM/HIGH) second delay to reduce engine speed
GND IDLE switch (Figure 5L-14) con- to 46% N2 RPM. With the switch in
trols the ground idle system. In NORM, HIGH or weight off the wheels, the
the ground idle system arms. Once the system deactivates and resets to the
landing gear squat switch closes normal engine flight idle speed.
(weight on wheels), the dual idle reset The amber GROUND IDLE annunci-
solenoids (ground idle solenoids) in the ator illuminates when the GND IDLE
fuel control units activate after an eight switch is in NORM.

5L-14

5L-24 For SimuFlite training only Citation V

March 1999
 Powerplant

Preflight made only when an engine is new or

Preflight and
freshly overhauled. For additional
During the preflight inspection, visual-
ly inspect the engine inlet for foreign
objects. Inspect the fan blades for any
information regarding the use of third
generation oils in the JT15D-5A, refer Procedures
to the engine manufacturer’s pertinent
evidence of foreign object damage (e.g.,
service bulletins.
nicks, cracks, dents). Verify that the
generator cooling air exhaust and fuel Should it be necessary to replenish oil
drain masts are clear. Check the condi- when oil of the same brand in the tank
tion of the T1 probe in the engine inlet. is unavailable, observe the following.
Check the engine exhaust and bypass ■ Total quantity added of a different

ducts for condition and foreign objects. brand of oil does not exceed two
Check for fuel leakage, damaged tur- U.S. quarts in any 400-hour period.
bine blades, cracks and general secu- ■ If it is necessary to add more than
rity of the entire area. Check that the oil two quarts of a different brand, drain
filler cap and access door are secure. and flush the oil system and refill
Check engine oil level within 10 with an approved oil.
minutes of engine shutdown. After 10 Should oil of unapproved brands or of
minutes oil may begin draining from different viscosities become inter-
the oil tank into the accessory gear- mixed, drain and flush the engine oil
box. The oil remains in the engine but system and refill with an approved oil
gives a false reading on the dipstick. in accordance with the JT15D-5A
If the oil level reads low after 10 min- Engine Maintenance Manual.
utes of engine shutdown, run the
To replenish oil:
engine for a minimum of two minutes
and then shut down and recheck. 1. Open the oil tank access door and
unlock the filler cap.
Servicing 2. Withdraw the dipstick and check
The oil tank holds 2.03 U.S. gallons the oil level.
(7.68 liters) with a usable quantity of 3. Refill the oil tank and insert the
1.20 U.S. gallons (4.54 liters). Maximum filler cap.
oil consumption is 0.5 pounds per hour 4. Lock the dipstick and then close
over a 10-hour period. the access door.
Approved types include the following
and those listed in P&W SB No. 7001: Abnormal
■ Exxon Turbo Oil 2380 Procedures
■ Aero Shell/Royco Turbine Oil 500 The following is a discussion of abnor-
■ Aero Shell/Royco Turbine Oil 560 mal procedures for the system. Please
(third generation oil) refer to your Simuflite Citation V
■ Mobil Jet II
operating handbook for a detailed
checklist. Abnormal procedures for
■ Mobil Jet 254 (third generation oil)
the powerplant include:
■ Castrol 5000. ■ false engine start

When changing from an existing lubri- ■ engine starter will not disengage
cant formulation to a third generation ■ engine fire during ground shutdown
lubricant formulation (Aero Shell/
Royco Turbine Oil 560 or Mobil Jet ■ low oil pressure

254), Pratt & Whitney Canada strong- ■ firewall shutoff valve closed
ly recommends that such a change be ■ ground idle light on in flight.

Citation V For SimuFlite training only 5L-25

March 1999
 False Engine Start OIL PRESS WARN Illuminated
(Engine Does Not Light) If oil pressure is above 60 PSI, land as
False engine start indications include soon as practical.
no increase in ITT and no continued If pressure is between 40 and 60 PSI,
increase in N2. Move the associated reduce power on the affected engine. If
throttle to CUTOFF and wait 15 sec- the pressure stabilizes after reducing
onds before pressing the STARTER power, maintain the reduced power
DISENGAGE pushbutton to clear setting for the remainder of the flight
excess fuel from the engine. and land as soon as practical.
Engine Starter Will Not If oil pressure is below 40 PSI, shut
down the engine by placing the throt-
Disengage
tle lever in CUTOFF. Land as soon as
If the starter fails to disengage once practical.
the engine reaches approximately 40%
N2 RPM, press the starter disengage OIL PRESS WARN Not
button. Illuminated
If the start button light remains illumi- Normal oil pressure should be indi-
nated, the start relay is stuck. Turn the cating with 52% N2 and above. If the
battery switch OFF to remove power oil pressure is between 40 and 60 PSI,
to the starter. Disconnect the battery in reduce power. If the pressure stabilizes
the tailcone to isolate the battery from after reducing power, maintain the
the electrical system. Bring the throttle reduced power setting for the remain-
lever(s) to the CUTOFF position to shut der of the flight and land as soon as
down the engine(s). practical.
If oil pressure is below 40 PSI, land
Engine Fire During as soon as practical.
Ground Shutdown
A high or sustained interturbine tem-
Firewall Shutoff Valve
perature during ground shutdown is a Closed
positive indication of an internal engine Illumination of the LH or RH F/W
fire. Check that the throttle lever is in SHUTOFF annunciator indicates the
the CUTOFF position. Momentarily closing of both the fuel and hydraulic
press the start button to motor the firewall shutoff valves. Pressing an
engine. Motoring the engine assists in ENGINE FIRE PUSH switchlight
clearing fuel from the engine and may closes the valves and isolates the thrust
blow the fire out. Press the starter dis- reversers. It also trips the field excita-
engage button after 15 seconds of tion on the associated generator and
engine motoring. arms both fire extinguisher bottles. To
open firewall shutoff valves and reset
Low Oil Pressure the system, push the ENGINE FIRE
A decreasing oil pressure and the illu- PUSH switchlight a second time.
mination of a red OIL PRESS WARN
annunciator signal an unacceptable oil Emergency
pressure. The red MASTER WARN- Procedures
ING lights also illuminate. The only
crew action is to accomplish a pre- Please refer to your Simuflite Citation
cautionary engine shutdown. V operating handbook for a detailed
checklist for emergency procedures.
The OIL PRESS WARN annunciator
is an independent system from the oil
pressure indicator.

5L-26 For SimuFlite training only Citation V

March 1999
 Powerplant

Emergency procedures for the power- Shut down the affected engine by mov-
plant include: ing the throttle to CUTOFF and turn-
■ engine failure or fire during takeoff ing the ignition system and generator
OFF. Reduce the electrical load as
■ engine failure/precautionary shut-
required to prevent overloading the
down operating generator. Turn the engine
■ engine failure during coupled synchronizer OFF. Use the fuel cross-
approach feed system to maintain a balanced
■ emergency restart – one engine fuel load; do not exceed a 200 lb
asymmetrical fuel load. If engine anti-
■ emergency restart – two engines
icing is on, select ANTI-ICE XFD on
■ engine fire. the failed engine.
Engine Failure or Fire If no fire hazard exists, leave the fire-
wall shutoff valve open and turn the
During Takeoff
fuel boost pump ON to prevent dam-
If speed is below V1 during an engine age to the engine-driven fuel pump.
failure or fire during takeoff, abort the
takeoff. Apply brakes as required to Engine Failure During
slow the aircraft, move the throttle to Coupled Approach
IDLE, extend the speedbrakes, and If an engine fails during a coupled
deploy the thrust reverser on the unaf- approach, increase power on the oper-
fected engine. Verify reverser opera- ating engine to maintain approach speed
tion by observing the illumination of and correct rate of descent. Disengage
the ARM, UNLOCK, and DEPLOY the autopilot and yaw damper. Add rud-
lights. To obtain maximum braking der trim toward the operating engine to
action, apply continuous pressure to compensate for aircraft yaw away from
the brake pedals; do not modulate the operating engine. Maintain airspeed
brake pressure. If there is an engine at VREF +10 KIAS before raising the
fire, perform an engine fire procedure. flaps to T.O. & APPR.
If an engine failure occurs above V1, Continue with the procedure by mov-
continue the takeoff. After establish- ing the affected engine’s throttle lever
ing a positive rate of climb, retract the to CUTOFF. If the failure occurred due
landing gear. At 400 ft, retract the flaps to an engine fire, perform the engine
at V2 +10. Accelerate to VENR (single fire procedure.
engine enroute climb speed). If there is
an engine fire, perform the engine fire Emergency Restart –
procedure. If there is an engine fail- One Engine
ure, accomplish an engine shut down.
There are two methods for an emer-
Engine Failure/ gency engine restart: one with starter
Precautionary Shutdown assist, and one by windmilling the
engine.
Several conditions such as abnormal-
ly high or low oil pressure, abnormal Starter Assisted
or rising ITT, engine vibration, fan/tur-
Place the throttle lever in CUTOFF,
bine RPM fluctuations, abnormal oil
the generator switch to GEN, and
temperature, or erratic fuel flow could
check that the firewall shutoff valves
necessitate an engine shutdown in
are open (LH or RH F/W SHUTOFF
flight.
annunciator extinguished). Position the

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March 1999
 ignition switch to ON; press the START Emergency Restart –
button momentarily. With weight off Two Engines
the wheels, the cross generator start sys-
tem disables to prevent damage to the Select both engine ignition and boost
operating engine generator. pump switches to ON. Automatic
sequencing and selection of these func-
Move the throttle lever to IDLE once tions does not occur when the start but-
the engine reaches 8 to 10% N2. ton is not used. Leave both throttle
Monitor engine instruments to prevent levers in IDLE.
exceeding 700°C (maximum start ITT).
Place the ignition switch to NORM. If altitude allows, increase airspeed to
200 KIAS to improve the possibility
Windmilling with Airspeed of an immediate start.
above 200 KIAS
If no start occurs in 10 seconds or accel-
Place the throttle in CUTOFF. Ensure eration to 200 KIAS is not feasible,
that the firewall shutoff valves are open momentarily press either START but-
(F/W SHUTOFF annunciator extin- ton and attempt a starter assist restart.
guished). Turn the respective ignition
switch and boost pump ON. When per- Engine Fire
forming a windmilling airstart, the If an ENGINE FIRE PUSH switch-
automatic sequencing and selection of light illuminates, immediately bring
these functions does not occur because the affected throttle lever to IDLE.
the button is not used. At airspeeds
above 200 KIAS, advance the throttle If the switchlight remains illuminated,
to IDLE. Monitor the engine instru- lift the cover on the illuminated switch
ments as the engine begins spooling and push it. Pushing the switch closes
up to idle speed. the firewall shutoff valves to cutoff
fuel and hydraulics. It also isolates the
After the engine stabilizes, position thrust reverser, trips the generator field,
the boost pump and ignition switches and arms the fire extinguisher bottles.
to NORM. It may be necessary to Push one of the illuminated fire bottle
select the associated generator RESET lights to discharge an extinguisher to
position momentarily to reinstate the the engine.
generator. Monitor ITT to prevent
exceeding the 700°C temperature limit Place the ignition switch in NORM.
for two seconds. Select CUTOFF with the affected
engine throttle lever. Reduce the elec-
During a windmilling airstart at high trical load to prevent overloading the
altitudes below 200 KIAS the engine operating generator. Turn the boost
may rumble and have an ITT above pump OFF and select ANTI-ICE XFD
500°C. This occurs as the engine on the affected engine. If required, land
accelerates between 30 and 42% N2 as soon as practical.
RPM because of the fuel-air mixture
instability. If the fire warning light remains illu-
minated after 30 seconds, push the sec-
ond fire bottle light. Land as soon as
possible.

5L-28 For SimuFlite training only Citation V

March 1999
 Powerplant

Airstart Envelope
(Maximum Altitude JP4/JET B: 30,000 ft)

5L-15

Citation V For SimuFlite training only 5L-29

March 1999
5L-30 For SimuFlite training only Citation V
March 1999
 Powerplant

Inter-Turbine Temperature Limits

Starting Conditions Only Engine
Limitations

All Conditions Except Starting

If the inter-turbine temperature (ITT) exceeds 700°C during takeoff, or if 680°C is
exceeded at any other time other than takeoff, refer to Chapter 77 of the Maintenance
Manual and Engine Maintenance Manual.

5L-16

Citation V For SimuFlite training only 5L-31

March 1999
 Engine Overspeed Limits

5L-17

5L-32 For SimuFlite training only Citation V

March 1999
 Powerplant

Engine Operating Limits

Thrust Setting Operating Limits

Time Limit ITT (°C) % Turbine % Fan RPM Oil Pressure Oil Temperature
(minutes) RPM (N2) (N1) (PSIG)2 (°C)

Takeoff 5 700 96 1044 60 to 833 10 to 121

Maximum Continuous 680 96 1044 60 to 83 10 to 121

Continuous

Idle Flight Continuous 580 52 (min) –– 40 (min) -40 to 121

Ground Continuous 580 52 (min) –– 40 (min) -40 to 121

Start –– see note 1 –– –– –– -40 (min)

Transient –– 700 96 106 see note 3 -18 to 129

1. The maximum start limit is 700°C for 2 seconds; refer to Figure 5L-16.
2. Normal oil pressures are 60 to 83 PSIG above 52% Turbine RPM. Oil pressures
below 60 PSIG are undesirable and should be tolerated only for the completion of
the flight, preferably at reduced power settings. Oil pressures below 40 PSIG are
unsafe and require that either the engine be shut down or a landing be made as
soon as possible, using the minimum power required to sustain flight.
3. The maximum transient oil pressure can be 95 PSIG for 90 seconds.
4. Refer to the appropriate thrust setting charts in the AFM for % Fan RPM setting.

Citation V For SimuFlite training only 5L-33

March 1999
5L-34 For SimuFlite training only Citation V
March 1999