PILOTS MANUAL

Learjet 60XR

This Pilots Manual provides information supplemental to the Learjet 60XR FAA Approved Airplane Flight Manual. In the event any information herein conflicts with information in the FAA Approved Airplane Flight Manual, the FAA Approved Airplane Flight Manual shall take precedence.

PM-133

May 2007

Subject: Learjet 60XR Pilots Manual Change 1

The following summary describes the changes that are incorporated with this change.

FRONT MATTER
Introduction Updated LOEP.

SECTION III HYDRAULICS & LANDING GEAR

Emergency Air System Corrected Emergency Braking was Emergency Brakes.

SECTION IV ELECTRICAL & LIGHTING

Table of Contents Battery Overheat Warning System (Page 4-9) BAT 60 AND BAT 71 Lights (Page 4-10) BAT TEMP Display (Page 4-10) Electrical Page Display Updated TOC Removed deleted paragraph entries. Removed text Ni-Cad battery not an option on the 60XR.

Removed text Ni-Cad battery not an option on the 60XR.

Removed text Ni-Cad battery not an option on the 60XR.

Removed Battery Temperature Indication not a feature with Lead Acid Battery. Revised beacon strobe operation when modified by SB-60-33-7.

Anti-Collision Beacon Strobe Lights

SECTION V FLIGHT SYSTEMS & AVONICS

Table of Contents Air Data Computers Updated TOC Removed deleted paragraph entries. Added ADC 2 when in EMER BUS operation. Added bullet item Cabin Pressurization. Removed text not a feature on the 60xr. Compatible only when used with the B.F. Goodrich GH-3000 standby instrument.

Remote Air Data Computer (Page 5-28)

SECTION VIII FLIGHT CHARACTERISTICS & OPERATIONAL PLANNING

Table of Contents Updated TOC Temperature was Tempature.

Climb Performance Two Engines

Corrected 23,500 was 25,500.

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Highlights-1

Instructions

LEGEND

- A ADD SHEET - D DESTROY SHEET - ALL OTHER SHEETS REVISED

A / BLANK Hydraulics 3-3 / 3-4 Electrical & Lighting IV-1 / IV-2 4-9/ 4-10 4-13 / 4-14 4-21 / 4-22 Flight Control Systems & Avionics V-1 / V-2 5-25 / 5-26 5-27 / 5-28 Flight Characteristics & Operational Planning VIII-1 / VIII-2 8-23 / 8-24

11/2009

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LIST OF EFFECTIVE PAGES

Use this List of Effective Pages to determine the current status of the Pilots Manual. Pages affected by the current change are indicated by an asterisk (*) immediately preceding the page number.

Dates of issue for Original and Changed pages are: Original ....................................... O ................................ May 2007 Change .........................................1 ..................... November 2009

Page

Change

Page

Change

Title.............................................O * A ................................................. 1 i and ii ........................................O General Description I-1................................................O 1-1 thru 1-21 ..............................O Engines & Fuel II-1 thru II-3...............................O 2-1 thru 2-51 ..............................O Hydraulics & Landing Gear III-1 .............................................O 3-1 thru 3-3 ................................O * 3-4 ............................................... 1 3-5 thru 3-19 ..............................O Electrical & Lighting * IV-1 ............................................. 1 IV-2 .............................................O 4-1 thru 4-8 ................................O * 4-9 and 4-10 ............................... 1 4-11 thru 4-13 ............................O * 4-14 ............................................. 1 4-15 thru 4-21 ............................O * 4-22 ............................................. 1 4-23 thru 4-37 ............................O Flight Control Systems & Avionics V-1 .............................................O * V-2 ............................................. 1 V-3 and V-4................................O 5-1 thru 5-24 ..............................O * 5-25 and 5-26 ............................. 1 5-27 .............................................O * 5-28 ............................................. 1 5-29 thru 5-58 ............................O
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Anti-Ice & Environmental VI-1 thru VI-3..............................O 6-1 thru 6-46 ..............................O Interior Equipment VII-1 and VII-2..........................O 7-1 thru 7-35 ..............................O Flight Characteristics & Operational Planning * VIII-1 ......................................... 1 VIII-2 ..........................................O 8-1 thru 8-23 ..............................O * 8-24 ............................................. 1 8-25 thru 8-104 ..........................O

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INTRODUCTION
The information in this manual is intended to augment the information in the Learjet 60XR FAA Approved Airplane Flight Manual and in no manner supersedes any Flight Manual limitations, procedures, or performance data. In the event that any information in this manual should conflict with that in the FAA Approved Airplane Flight Manual, the FAA Approved Airplane Flight Manual shall take precedence.

THE MANUAL
Sections I through VII of this manual are intended to provide the operator of the Learjet 60XR with a basic description of the aircraft operating systems from the cockpit controls and indicators to the actuating mechanisms in the systems. No attempt has been made to establish a specific standard aircraft due to the numerous customer options. Therefore, the illustrations and descriptions within this manual are for a typical aircraft and may not match a specific aircraft. Specific serialization is shown only when more than one version of the same system is incorporated into production on a nonretrofit basis. Section VIII of this manual contains tabular performance and fuel consumption data derived from the Flight Manual and flight testing. This data may be used by the operator for flight planning.

REVISING THE MANUAL

Periodically, Numbered Changes may be issued against this manual. Pages included in Numbered Changes supersede like numbered pages in the Pilots Manual. Each page of a Numbered Change will contain a Change number located at the lower inside margin of the page. Portions of the text affected by the change are indicated by a vertical bar at the outer margin of the page. The vertical bars may not appear on pages that contain graphs or tables. Additionally, when a changed page occurs as the result of a rearrangement of material due to a change on a previous page, no vertical bar will appear.

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REVISING THE MANUAL (CONT)

The List of Effective Pages provides the user with a guide to establish the current effective date of each page in the Pilots Manual and may be used as an instruction sheet for incorporating the latest Numbered Change into the Pilots Manual. Information included in the List of Effective Pages states the current Change number for each page and the dates of Original issue and Numbered Changes. An asterisk (*) next to a page number indicates the page was changed, added, or deleted by the current change.

ADDRESSES
Your comments and suggestions concerning this manual are solicited and should be forwarded to: Learjet, Inc. P.O. Box 7707 Wichita, Kansas 67277-7707 Attn: Technical Publications

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SECTION I GENERAL DESCRIPTION

TABLE OF CONTENTS Aircraft General Description.................................................................... 1-1 Airplane Three-View (Figure 1-1)........................................................ 1-2 General Arrangement - Exterior (Figure 1-2) .................................... 1-3 Cabin Entry Door ...................................................................................... 1-5 ENTRY DOOR Light ............................................................................. 1-5 Cabin Door Operation........................................................................... 1-6 Opening Cabin Door (From Outside) (Figure 1-3)....................... 1-6 Closing Cabin Door (From Inside) (Figure 1-4)............................ 1-7 Opening Cabin Door (From Inside) (Figure 1-5).......................... 1-8 Closing Cabin Door (From Outside) (Figure 1-6) ........................ 1-9 Emergency Exit/Baggage Door............................................................. 1-10 AFT CAB DOOR Light........................................................................ 1-10 Emergency Exit/Baggage Door Operation ...................................... 1-10 Emergency Exit/Baggage Door Operation (From Inside) (Figure 1-7).............................................................. 1-11 Emergency Exit/Baggage Door Operation (From Outside) (Figure 1-8)........................................................... 1-12 External Doors ......................................................................................... 1-13 EXT DOORS Light ............................................................................... 1-13 Tailcone Baggage Compartment........................................................ 1-13 Turning Radius (Figure 1-9) ................................................................... 1-14 Danger Areas (Figure 1-10) .................................................................... 1-15 Pedestal (Typical) (Figure 1-11) ............................................................. 1-16 Instrument Panel (Typical) (Figure 1-12) ............................................. 1-17 Pilots Circuit Breaker Panel Layout (Figure 1-13) ............................. 1-19 Copilots Circuit Breaker Panel Layout (Figure 1-14) ........................ 1-21

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SECTION I GENERAL DESCRIPTION

AIRCRAFT GENERAL DESCRIPTION
The Learjet 60XR aircraft, manufactured by Learjet, Inc., is an all metal, pressurized, low-wing, turbofan-powered monoplane. The high-aspect ratio, fully cantilevered, swept-back wings with winglets are of conventional riveted construction except for the upper section of the winglets, which is full-depth honeycomb core bonded to the outer skin. The fuselage is of area rule design and semi-monocoque construction. Two inverted V ventral fins (delta fins) are fitted to the aft section of the tailcone to provide the aircraft with favorable stall recovery characteristics and additional lateral/directional stability. Thrust is provided by two pod-mounted PW305A turbofan engines manufactured by Pratt and Whitney Canada, Inc. Independent fuel systems supply fuel to the engines with fuel storage available in wing and fuselage tanks. Enginedriven hydraulic pumps supply hydraulic power for braking, extending and retracting the landing gear, wing flaps, and spoilers. The landing gear system is a fully retractable tricycle-type gear with dual maingear wheels, anti-skid braking, and nose-wheel steering. The flight controls are manually controlled through cables, bellcranks, pulleys, and push-pull tubes. Lateral and directional trim is accomplished by means of electrically-actuated trim tabs installed on the left aileron and on the rudder. Longitudinal trim is accomplished by changing the angle of incidence of the horizontal stabilizer with an electrically-operated linear actuator. Aircraft air conditioning systems provides heating, cooling, and pressurization for the crew, passenger, and cabin baggage compartments.

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NOTE: All dimensions shown for aircraft in static position.

14 ft 7 in (4.44 m)

56 ft 2 in (17.12 m) 58 ft 8 in (17.89 m) 14 ft 8 in (4.48 m)

8 ft 3 in (2.51 m) 43 ft 10 in (13.35 m)

AIRPLANE THREE-VIEW Figure 1-1

1-2

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1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15. 16. 17. 18. 19. 20. 21. 22. 23.

Fwd Avionics Compartment Stall Warning Vane Pitot-Static Tube Lower Cabin Entry Door Upper Cabin Entry Door Inboard Wing Stall Fence Boundary Layer Energizers Outboard Wing Stall Fence Wing Fuel Filler Cap Wing Navigation Light Aft Baggage Compartment Door Tailcone Compartment Access Door Rudder Elevator Tail Navigation Light Tail Strobe/Beacon Recognition Light Ram Air Inlet Fuselage Fuel Cell Flap Spoiler/Spoileron Aileron Emergency Exit/Baggage Door

GENERAL ARRANGEMENT - EXTERIOR Figure 1-2

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CABIN ENTRY DOOR

The cabin door consists of an upper portion that forms a canopy when open and a lower portion with integral steps. The upper portion has gas-charged struts (gas springs) installed to assist in door opening. A latch, when over centered, retains the door in the open position. A door release handle, located on the aft door frame, mechanically releases the latch to allow the upper door to close. The gas-charged struts soften door opening and closing movements. The lower portion of the door incorporates a torsion bar system to provide closing assistance. Cables attached to take-up reels are installed on the forward and aft lower door structure to aid in closing and prevent damage if the door is inadvertently allowed to drop open. A self-contained hydraulic damper is also attached to the lower door as an additional protection against dropping the door. Each door half has a locking handle which, when rotated, drives a series of locking pins into the fuselage structure and through interlocking arms secure the halves together. When the pins are engaged, the door becomes a rigid structural member. There is a secondary safety latch installation on the lower door separate from the door locking system. This installation will hold the lower door against the door frame seal, and align the locking pins with the pin holes. When the lower door is unlocked, the safety latch will keep the door from falling open. This latch may be operated from either inside or outside the aircraft. A key lock is provided on the upper door to secure the aircraft from the outside. Rotating the key lock will move a locking bar over the inside upper door handle, preventing it from rotating to the open position. ENTRY DOOR LIGHT A red ENTRY DOOR warning light is installed on the glareshield annunciator panel to provide the crew with visual indication of cabin door security. The light will illuminate and flash to indicate that one or more of the locking pins is not fully engaged or that the key lock is in the locked position. The light will illuminate steady when the entry door is full open and power is on the aircraft. If all pins are fully engaged, and the locking bar is recessed, the most probable cause for illumination is a switch malfunction or misalignment.

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CABIN DOOR OPERATION To open the cabin door from the outside: 1. Insert key in key lock and rotate. The key lock will retract the upper door handle locking bar. 2. Insert finger in the handle finger pull door and pull out handle halves. Rotate the handle halves clockwise to the stop. 3. Raise upper door to the full open position. 4. Reach inside and rotate lower door locking handle to OPEN position. 5. Release safety catch, located on forward side of middle step, from the inside, or outside using exterior button. 6. Gently lower door to the full down position.

OPENING CABIN DOOR (FROM OUTSIDE) Figure 1-3 1-6

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CABIN DOOR OPERATION (CONT) To close cabin door from inside: 1. Raise lower door, using forward cable knob, until safety latch fully engages. 2. Rotate lower door locking handle to the locked position. 3. Release upper door with door release handle on aft door frame. 4. With the upper door locking handle in OPEN position, pull door tightly against door seal and rotate locking handle to the locked position. (If preparing for flight, check ENTRY DOOR warning light extinguished.)

CLOSING CABIN DOOR (FROM INSIDE) Figure 1-4

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CABIN DOOR OPERATION (CONT) To open cabin door from the inside: 1. 2. 3. 4. 5. Lift upper door locking handle to the OPEN position. Push upper door outward and up to the full open position. Rotate lower door locking handle to OPEN position. Release safety latch, located on forward side of middle step. Gently lower the lower door to full down position using the forward cable knob.

OPENING CABIN DOOR (FROM INSIDE) Figure 1-5

1-8

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CABIN DOOR OPERATION (CONT) To close and lock cabin door from the outside: 1. Raise lower door until the safety latch fully engages. 2. Reach inside and rotate lower door locking handle to the locked position. 3. Release upper door with door release handle on aft door frame. 4. With upper door locking handle in the OPEN position, gently lower upper door and push tightly against door frame. 5. Rotate exterior handle halves counterclockwise to the stop and ensure each half recesses into door structure. 6. Insert key in key lock and rotate. This will extend the upper door locking bar over the locking handle.

CLOSING CABIN DOOR (FROM OUTSIDE) Figure 1-6

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EMERGENCY EXIT/BAGGAGE DOOR

The emergency exit/baggage door, located on the aft right side of the cabin, serves a dual function. It provides egress from the cabin during emergencies and access from the outside to the aft cabin baggage area. The door is attached to the airframe by hinges at the top and secured by locking pins at the side and lower edge. The door structure incorporates a window similar to those installed in the cabin. Gas-charged struts (gas springs) are installed to assist in door opening and closing and to hold the door open when fully extended. For security on the ground, the inner door latching handle has a red streamered locking pin installed through a hole in the handle to restrict movement. This pin must be removed before every flight. AFT CAB DOOR LIGHT To provide cockpit visual indication as to the flight status of the emergency exit/baggage door, a red AFT CAB DOOR warning light is installed on the glareshield annunciator panel. The light will illuminate and flash if the locking pins are not fully engaged, the handle mechanism is not in the latched position, or the red streamered locking pin has not been removed for flight. The light will illuminate steady when the handle is at the full open position. If all components are found to be properly positioned, a switch malfunction or misalignment is the probable cause for illumination. EMERGENCY EXIT/BAGGAGE DOOR OPERATION To open emergency exit/baggage door from the inside: 1. Remove red streamered locking pin. 2. Rotate locking handle to the OPEN position. 3. Push door outward and up to the full open position. To close the emergency exit/baggage door from the inside: 1. With the door locking handle in the OPEN position, gently lower the door. 2. Pull door tight against door seal and rotate the locking handle to the locked position. 3. If preparing for flight, no further action is required except to check AFT CAB DOOR warning light extinguished. If securing door on the ground, rotate pin cover knob and insert red streamered locking pin.

1-10

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EMERGENCY EXIT/BAGGAGE DOOR OPERATION (FROM INSIDE) Figure 1-7

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EMERGENCY EXIT/BAGGAGE DOOR OPERATION (CONTD) To open emergency exit/baggage door from the outside: 1. Insert finger in the handle finger pull door and pull out handle halves. Rotate the handle halves clockwise to the stop. 2. Raise door upward to the full open position.
NOTE

Stand clear if there is a chance the cabin is still pressurized.

EMERGENCY EXIT/BAGGAGE DOOR OPERATION (FROM OUTSIDE) Figure 1-8 To close the emergency exit/baggage door from the outside: 1. With the door locking handle in the OPEN position, gently lower the door and push tightly against door frame. 2. Rotate exterior handle halves counterclockwise to the stop and ensure each half recesses into door structure. 3. If preparing for flight, no further action is required except to check AFT CAB DOOR warning light extinguished.

1-12

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EXTERNAL DOORS
External doors are installed to provide for baggage loading and maintenance access. The nose area forward of the cockpit is accessible through four doors two on the left side and two on the right side. The tailcone is accessible through the tailcone access door and aft baggage door, both located on the left side. Two doors provide access to the single-point pressure refueling system. These doors are located side by side on the right side of the fuselage beneath the right engine. Access to the external servicing provisions for the toilet is through a door on the underside of the fuselage below the toilet. EXT DOORS LIGHT Illumination of the red EXT DOORS warning light, located on the glareshield annunciator panel, indicates the tailcone access door and/ or the aft baggage door is not properly closed and latched. The primary purpose of the light is to indicate a door open condition prior to takeoff. If the doors were properly latched prior to takeoff and the light illuminates in flight, the most probable cause is a switch failure. TAILCONE BAGGAGE COMPARTMENT The tailcone baggage compartment is accessed through a door located under the left engine pylon. A slight pressure differential (0.25 psi) is maintained to prevent fluids from entering the compartment. The pressure is provided by ram air entering the dorsal inlet. An outflow valve, located on the top of the baggage compartment, controls the pressure.

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36.5 feet (11.1 meters)

(8 29. .9 3 m fee et t er s)

NOSE WHEEL

WING TIP

NOTE

Turning radius expressed above is based upon 60 nose wheel travel (full-authority/low-speed steering). Limited authority steering provides 24 of nose wheel travel. Turning radius will increase accordingly. TURNING RADIUS Figure 1-9

1-14

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2 feet (0.6 m)

WEATHER RADAR

35 feet (10.7 m)

35 feet (10.7 m)

ENGINE INTAKE

Engine danger area shown for takeoff RPM.

ENGINE EXHAUST
1000F (538C)

75F (24C) 240 feet (73.2 m)

DANGER AREAS Figure 1-10

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22 21

2 3 4

20 19 17 16 18

15

14 13

12

1. Thrust Levers 2. Thrust Reverser Levers 3. APR Switch (Recessed) 4. Flap Switch 5. Check List Switch 6. Copilot Course (CRS) Control 7. Copilot Cursor Control Panel (CCP) 8. Control Display Unit #2 6 (CDU) 9. Fuel Control Panel 7 10. Dual HF Comm Control Head 8 11. AIRSHOW Flight Deck Controller 12. Emergency Exit Lights Control Panel 13. Radio Tuning Unit #2 (RTU) 9 14. Trim Control Panel 15. Control Display Unit #1 (CDU) 16. Pilot Cursor Control Panel (CCP) 10 17. Pilot Course (CRS) Control 11 18. Nose Steer Switch 19. Parking Brake Handle 20. Engine Sync Switches 21. Emergency Brake Handle 22. Spoiler Lever

PEDESTAL (TYPICAL) Figure 1-11

1-16

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5 4

10

11 12

3 2

13 14

1
1. 2. 3. 4. 5. 6. 7. 8. 9. 10.

20
11. 12. 13. 14. 15. 16. 17. 18. 19. 20.

19

18

17

16

15

Pilots Switch Panel Pilots Audio Control Panel Pilots EFIS Control Panel Pilots Flight Instruments (PFD & MFD) Pilots Display Control Panel (DCP) Electronic Standby Instrument System (ESIS) Flight Control Panel (FCP) Heading Speed Altitude Panel (HSA) Annunciator Panel Fuel Quantity Indicator

Copilots Display Control Panel (DCP) Copilots Flight Instruments (MFD & PFD) Copilots EFIS Control Panel Copilots Audio Control Panel Copilots Switch Panel Cockpit Voice Recorder Control Panel Landing Gear Control Panel Radio Tuning Unit #1 (RTU) Center Switch Panel ELT Control Switch

INSTRUMENT PANEL (TYPICAL) Figure 1-12

16-125B

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ELECTRICAL LIGHTS

L EMER BUS CONT FLOOD LTS WARN LTS

NAV LTS

L INV

EMER BAT 1

L INSTR LTS

L EL LTS

L AC BUS EMER BAT 3

CENTER PANEL PED LTS CHART HOLDERS

TRIM-FLT CONT

BAT TEMP L DC BUS 1 PRI PITCH TRIM

L STALL WARN

CABIN PWR BUS L DC BUS 2 ROLL TRIM

WHEEL MASTER

L GEN

L DC BUS 3

YAW TRIM

SQUAT SW
ANTI-ICE

Denotes AC circuit breakers

L DC BUS 4
AFCS

Denotes DC circuit breakers

L PITOT HEAT L WSHLD DEFOG L IAPS AP 1 L STALL VANE HEAT L WSHLD DEFOG PITCH SERVO ROLL-YAW SERVO
ENVIRONMENT

L NAC HEAT L ICE DETECT LIGHT ICE DETECTOR

Denotes unused circuit breaker positions

MACH TRIM
FUEL

Denotes circuit breakers on the emergency bus

L BLEED AIR TEMP CONTROL IND FUS TANK XFR PUMP OXYGEN VALVE BLEED AIR OV HT FUEL QTY PWR 1 L STBY SCAV PUMP L JET PUMP XFR VALVE COOL CONTROL
INSTRUMENTS

CABIN PRESS IND MANUAL TEMP CONTROL

L FUEL FLOW

XFLO VALVE
L ENGINE

AHS 1

PFD 1

MFD 1

L FIRE DETECT

L FW SOV

ADC 1

LO SPD WARN1

L FIRE EXT

L START

L CLOCK
AVIONICS

EFIS CONTROL 1

L ENG CH A

L ENG CH B

AUDIO 1

RTU 1

L AVIONICS MASTER

L IGN CH A

L IGN CH B

COMM 1

NAV 1

ATC 1

ENGINE SYNC

ADF 1

DME 1

PFD 1 HEAT

L TR CONT

L TR AUTO STOW

HF 1

GPS 1

MFD 1 HEAT

L OIL PRESS

L ENGINE VIB MON

AVIONICS

FMS DISPLAY 1

DISPLAY CONTROL 1

FSU 1

DATA LINK RADIO

IAPS TEMP

MFD CONTROL 1

DCU 1

EDC 1

FDR

PHONE

SELCAL

RADIO ALT

XM WEATHER

TAWS

TAWS

SATCOM

CABIN

CABIN

READ LTS

TABLE LTS

AISLE LTS

CABIN LTS

STEREO

VANITY DRN

WATER HEATER

TOILET

TOILET SERVICE

VIDEO

PILOTS CIRCUIT BREAKER PANEL LAYOUT Figure 1-13 1-19/1-20 (Blank) 1-19

Pilots Manual

GALLEY DRN

HOT CUP

MICRO WAVE

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LIGHTS ELECTRICAL

EMER LTS WING INSP LT

WARN LTS

R EMER BUS CONT

EMER BUS TIE

BEACONSTROBE LTS R EL LTS R INSTR LTS FLASH LTS

TRIM-FLT CONT

EMER BAT 2

AC BUS TIE

LOGO LT

PULSE RECOG LT

R DC BUS 1

NOSE STEER R STALL WARN

SEC PITCH TRIM

R DC BUS 2

O P E N

DC BUS 1 TIE DC BUS 2 TIE

NOSE STEER SPOILER

FLAPS

R DC BUS 3

DC BUS 3 TIE

RUDDER PEDAL ADJUST SPOILERON

HYDRAULICS

Denotes AC circuit breakers

TRIM-FLAP -SPOILER IND R DC BUS 4 HYDRAULIC PRESS IND GEAR R AC BUS
AFCS

Denotes DC circuit breakers

AIR PRESS IND ANTI SKID AP 2 SYSTEM TEST
ANTI-ICE FUEL

R INV

R GEN

Denotes unused circuit breaker positions

R NAC HEAT R WSHLD DEFOG R PITOT -STALLTAT HEAT R WSHLD DEFOG STANDBY PITOT HEAT R ICE DETECT LIGHT STAB HEAT WSHLD HEAT R STALL VANE HEAT WING HEAT ALCOHOL SYSTEM
ENVIRONMENT

R IAPS

Denotes circuit breakers on the emergency bus

FUS TANK AUX PUMP R JET PUMP -XFR VALVE R FUEL FLOW
R ENGINE

FUEL QTY PWR 2 R STBY -SCAV PUMP

TAT PROBE HEAT

R FW SOV

R FIRE DETECT

CABIN PRESS SYS

R BLEED AIR AUTO TEMP CONT

R START

R FIRE EXT

CREW FAN

CABIN AIR

R ENG CH B

R ENG CH A

AUX CREW HEAT

CABIN FAN
INSTRUMENTS

AUX CABIN HEAT

R IGN CH B

R IGN CH A

MFD 2

PFD 2

AHS 2

ENGINE DIAGNOSTIC SYSTEM

LO SPD WARN 2

ADC 2

R TR AUTO STOW

R TR CONT

EFIS CONTROL 2

R CLOCK
AVIONICS

STATIC SOURCE

R ENGINE VIB MON

AVIONICS

R OIL PRESS

R AVIONICS MASTER

RTU 2

AUDIO 2

MFD 2 HEAT

PFD 2 HEAT

ATC 2

NAV 2

COMM 2

NOSE FAN

INSTR PANEL FANS

TCAS

DME 2

ADF 2

RADAR

CVR

GPS 2

HF 2

ELT NAV

ELT

FSU 2

DISPLAY CONTROL 2

FMS DISPLAY 2

DATA LINK

EDC 2

DCU 2
CABIN

MFD CONTROL 2

STORM SCOPE
CABIN

HOUR METER

110 VAC OUTLETS

220 VAC OUTLETS

PASS SPKR

CABIN DISPLAY

CABIN FIRE DETECT 110 VAC INV 220 VAC INV PASS INFO CABIN AUDIO ENTRY LTS

COPILOTS CIRCUIT BREAKER PANEL LAYOUT Figure 1-14

1-21/1-22 (Blank) 1-21

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PASS CONTROL

PASS AUDIO

AFT BAG LT

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SECTION II ENGINES & FUEL

TABLE OF CONTENTS Engines........................................................................................................ 2-1 Engine Fuel and Control System ......................................................... 2-1 Engine Control Logic Diagram (Figure 2-1)....................................... 2-2 Thrust Levers.......................................................................................... 2-3 Engine-Driven Fuel Pump.................................................................... 2-3 Hydro-Mechanical Fuel Control Unit ................................................. 2-3 Full Authority Digital Electronic Control (FADEC) ......................... 2-4 ENG CMPTR Switches..................................................................... 2-4 ENG CMPTR Lights ......................................................................... 2-5 Variable Inlet Guide Vanes and Variable Stator Vanes ..................... 2-5 Surge Bleed Control............................................................................... 2-5 Automatic Performance Reserve (APR)................................................. 2-6 APR Switch ............................................................................................. 2-6 APR ARM Indicator .............................................................................. 2-6 APR ON Indicator ................................................................................. 2-7 Engine Synchronizer ................................................................................. 2-7 ENG SYNC Switches ............................................................................. 2-7 ENG SYNC Indicators........................................................................... 2-8 Ground Idle System .................................................................................. 2-8 Engine Oil System ..................................................................................... 2-8 Pressure System ..................................................................................... 2-8 Engine Oil System Schematic (Figure 2-2) ......................................... 2-9 Scavenge System .................................................................................. 2-10 Breather System.................................................................................... 2-10 Engine Ignition System........................................................................... 2-10 IGNITION Switches ............................................................................ 2-11 Ignition Lights ...................................................................................... 2-11 Engine Indicating System (EIS) ............................................................. 2-11 EIS Engine Page (Figure 2-3) .............................................................. 2-11 N1 Indicators......................................................................................... 2-12 ITT Indicators ....................................................................................... 2-12 N2 Indicators......................................................................................... 2-13 FUEL FLOW (FF) Indicators .............................................................. 2-13 Engine Oil Indicators (Pressure and Temperature) ........................ 2-13 Oil Pressure Lights............................................................................... 2-13 Engine Chip Lights .............................................................................. 2-14 ENG FILTERS Light............................................................................. 2-14 ENG VIB Lights.................................................................................... 2-14 PM-133 II-1

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TABLE OF CONTENTS (Cont) Engine Diagnostic System (EDS) ...................................................... 2-14 Engine Diagnostic System (Figure 2-4) ............................................ 2-15 Engine Diagnostic Unit (EDU) ..................................................... 2-16 Isolation Units................................................................................. 2-16 Control Display Unit (CDU) ......................................................... 2-16 EDS FAULT Annunciator .............................................................. 2-16 EDS Record Switch......................................................................... 2-16 Engine Fire Detection System ............................................................... 2-17 SYSTEM TEST Switch Fire Detection Function ......................... 2-17 ENG FIRE PULL Light ....................................................................... 2-17 Engine Fire Extinguishing System........................................................ 2-18 ENG FIRE PULL Handle and ENG EXT ARMED Lights ............. 2-18 Engine Fire Extinguishing System (Figure 2-5)............................... 2-19 Fire Extinguisher Discharge Indicators ............................................ 2-20 Thrust Reverser System ......................................................................... 2-20 Deploy................................................................................................... 2-21 Stow ....................................................................................................... 2-21 Auto Stow ............................................................................................. 2-22 Thrust Reverser Assembly ................................................................. 2-22 Thrust Reverser System Schematic (Figure 2-6) ............................. 2-23 Thrust Reverser Lever ........................................................................ 2-25 Throttle Balk Solenoid ........................................................................ 2-25 Hydraulic Control Unit (HCU) ......................................................... 2-26 Thrust Reverser Relay Box ................................................................. 2-26 Aircraft Fuel System ............................................................................... 2-27 Wing Tanks ........................................................................................... 2-27 Fuselage Tank....................................................................................... 2-27 Fuel Control Panel Switches and Annunciators ............................. 2-27 Fuel Control Panel (Figure 2-7) ......................................................... 2-28 JET PUMP Switches ....................................................................... 2-28 Fuel System Schematic (Figure 2-8) .................................................. 2-29 STBY PUMP Switches.................................................................... 2-31 XFLO VALVE Switch ..................................................................... 2-31 NORM XFR Switch ........................................................................ 2-32 AUX XFR Switch ............................................................................ 2-33 GRVTY XFR Switch........................................................................ 2-34 FILL Switch ..................................................................................... 2-34 Fuselage Tank Switch Priority ...................................................... 2-35 Fuselage Tank Full Light ............................................................... 2-35 Fuselage Tank Empty Light .......................................................... 2-35 LO FUEL PRESS Lights ................................................................. 2-35

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TABLE OF CONTENTS (Cont) Fuel Gaging System............................................................................. 2-36 Fuel Quantity Indicator.................................................................. 2-36 Fuel Quantity Probes...................................................................... 2-36 Total Quantity Indicator (SPPR) ................................................... 2-36 Fuel System Glareshield Lights ......................................................... 2-37 FUEL PRESS Lights ........................................................................ 2-37 LOW FUEL Light ............................................................................ 2-37 FUEL SYS Light............................................................................... 2-37 Ram Air Fuel Vent System.................................................................. 2-38 Single-Point Pressure Refuel (SPPR) System ................................... 2-38 Single-Point Refuel System Schematic (Figure 2-9) ........................ 2-39 WING and FUS PRECHECK Valves............................................ 2-41 SPPR BATT Switch ......................................................................... 2-41 Refuel Selector Switch .................................................................... 2-42 FUS FULL Light .............................................................................. 2-42 VENT OPEN Light ......................................................................... 2-42 Fuel Drains (Figure 2-10) .................................................................... 2-43 Fuel Anti-Icing Additive..................................................................... 2-44 Refueling ............................................................................................... 2-44 Auxiliary Power Unit (APU) ................................................................. 2-45 APU Control Panel .............................................................................. 2-46 APU Control Panel (Figure 2-11) ....................................................... 2-46 APU AMPS Indicator ..................................................................... 2-46 APU FIRE ......................................................................................... 2-46 APU FAULT/STOP Switch ........................................................... 2-47 APU RUNNING/START Switch.................................................. 2-47 APU MASTER Switch .................................................................... 2-47 APU ON Indicator .......................................................................... 2-47 APU SYSTEM TEST Switch........................................................... 2-47 APU Relay Panel.................................................................................. 2-48 APU Relay Panel (Figure 2-12) .......................................................... 2-48 FIRE DET BITE Indicator............................................................... 2-48 GEN FAULT BITE Indicator.......................................................... 2-48 FAULT Reset Switch ....................................................................... 2-48 APU BITE Annunciator Box............................................................... 2-49 APU BITE Annunciator Box (Figure 2-13) ....................................... 2-49 APU Generator..................................................................................... 2-49 APU Operating Procedures................................................................ 2-50 APU Pre-Start Check ...................................................................... 2-50 APU Start-Up .................................................................................. 2-50 APU Shutdown ............................................................................... 2-51 APU Shutdown Features (Automatic)......................................... 2-51
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SECTION II ENGINES & FUEL

ENGINES
The Learjet 60XR is powered by two PW305A Pratt and Whitney twospool, front-fan engines. Each engine is rated at 4600 pounds thrust at sea level. A spinner and an axial-flow fan, located at the forward end of the engine, are driven by the low-pressure rotor. The low-pressure rotor consists of an axial-flow fan (low-pressure compressor) and a three-stage low-pressure axial turbine, mounted on a common shaft. The highpressure rotor consists of a high-pressure compressor (four axial stages and a single centrifugal stage) and a two-stage high-pressure axial turbine, mounted on a common shaft. The rotor shafts are concentric, so that the low-pressure rotor shaft passes through the high-pressure rotor shaft. The high-pressure rotor drives the accessory gearbox through a driveshaft geared to the N2 rotor shaft. An annular duct serves to bypass fan air for direct thrust and also diverts a portion of the fan air to the high-pressure compressor. The bypass ratio (bypass flow to core flow) is 4.55:1. Air from the low-pressure compressor flows through variable inlet guide vanes and first-stage variable stator vanes to the high-pressure compressor and is discharged into the annular combustor. Combustion products flow through the high- and low-pressure turbines and are discharged axially through the exhaust duct to provide additional thrust. ENGINE FUEL AND CONTROL SYSTEM The engine fuel and control system pressurizes fuel routed to the engine from the aircraft fuel system, meters fuel flow, and delivers atomized fuel to the combustion section of the engine. The system also supplies high-pressure motive-flow fuel to the aircraft fuel system for jet pump operation. The major components of the system are the thrust levers, the engine-driven fuel pump, the hydro-mechanical fuel control unit (HFCU), the full authority digital electronic control (FADEC), variable inlet guide vanes, variable stator vanes, and the surge bleed control.

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Thrust Lever (TLA)

Aircraft Air Data Computer Cockpit Displays Pilot Select and Aircraft Discretes Engine Trims 28V DC Power Surge Bleed Control

Full Authority Digital Electronic Control (FADEC)

N1 Inlet Flight Conditions

P3

N2 T4.5

IGV BOV

Metered Fuel (Wf) IGV Position Demand Wf Fuel Demand Torquemotor Valves Servo Pressure Metering Valve Overspeed Trip Fuel Shutoffs Hydro-mechanical Fuel Control Unit (HFCU) Pump Fuel In Standby Shutdown

ENGINE CONTROL LOGIC DIAGRAM Figure 2-1

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THRUST LEVERS Two thrust levers (one for each engine) are located on the upper portion of the pedestal, and operate in a conventional manner with the full forward position being maximum thrust. Stops at the IDLE position prevent inadvertent reduction of the thrust levers to CUT-OFF. The IDLE stops can be released by lifting a finger lift on the outboard side of each thrust lever. Detents are provided for CUT-OFF, IDLE, maximum cruise (MCR), maximum continuous thrust (MCT), takeoff (TO), and automatic performance reserve (APR). Each thrust lever is mechanically linked to a rotary variable differential transformer (RVDT) position transducer. The RVDT provides dual electrical signals to the FADEC which correspond to the thrust lever angle (TLA). A switch, which actuates in the CUT-OFF position, provides a discrete signal to the FADEC to initiate the normal shutdown sequence. ENGINE-DRIVEN FUEL PUMP The engine-driven fuel pump provides high-pressure fuel to the engine fuel control system as well as motive-flow fuel for operation of the aircraft jet pumps. The pump consists of a low-pressure pump element, high-pressure pump element, relief valve, and motive flow provisions. The pump itself is housed in the hydro-mechanical fuel control unit. Fuel from the low-pressure element passes through a filter before it enters the high-pressure element. In the event the pressure differential across the fuel filter increases to a preset level, the impending bypass indicator will actuate and the white ENG FILTERS light will illuminate. If the pressure differential continues to increase, due to clogging, the filter bypass valve will open to allow fuel to bypass the filter. HYDRO-MECHANICAL FUEL CONTROL UNIT (HFCU) The HFCU mounts to the permanent magnet alternator on the aft side of the accessory gearbox. The HFCUs main function is to control fuel flow to the engines fuel nozzles. Fuel flow is regulated in response to commands from the FADEC which computes the necessary settings for the existing conditions. The HFCU also provides servo pressure to the variable guide vane actuator, houses the engine-driven fuel pump, and provides fuel pressure regulation.

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FULL AUTHORITY DIGITAL ELECTRONIC CONTROL (FADEC) There are two FADECs installed, one on each engine. Each FADEC has two channels (A and B), each fully capable of controlling the engine. During normal operation (ENG CMPTR switch in AUTO), the most capable channel is automatically selected to control the engine. FADEC functions include: Thrust Management Overspeed Protection Bleed-Off Valve Control Automatic Performance Reserve Inlet Guide Vane & Inlet Stator Vane Control Igniter Operation Surge Protection Fault Detection N1 Bug Setting Engine Synchronization Starting & Shutdown Control Digital ITT

The crew is able to control the engine through the FADEC by changing the TLA input to change desired thrust level. The FADEC receives input from several engine sensors and the aircrafts air data computers and together with the TLA input it determines the appropriate signals to send to the HFCU, the inlet guide vane and stator vane actuator, and the bleed-off valve solenoid to achieve the desired engine operation. The aircrafts air data computers provide inlet static pressure (PAMB) and Mach number as primary signals to the FADEC. PAMB and Mach number are also measured by the FADEC transducer but used only as a backup to the air data computer signals. Sensors on the engine provide inlet total temperature (TT0) signals to the FADEC. A TT0 signal is provided by the air data computer, but used only as a backup to the engine sensor signals. Electrical power is supplied by an engine-driven permanent-magnet alternator. Backup power and power for starting is provided through the ENG CH A and ENG CH B circuit breakers on the pilots and copilots circuit breaker panels. Backup power is available to channel A during EMER BUS mode. ENG CMPTR SWITCHES Two switches, one for each engine, on the center switch panel labeled ENG CMPTR CH. A/AUTO/CH. B enable the flight crew to select the FADEC channel (A or B) to be used to control the engine. Normally, the switches are left in the AUTO position which allows the FADEC to automatically select the most capable channel. During abnormal situations, the crew may use this switch to force the desired channel to take control of the engine.

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ENG CMPTR LIGHTS Two ENG CMPTR lights are provided for each engine and reside in the annunciator panel. One light is white and one is amber. Illumination of a white light indicates a minor malfunction in one or both channels of the associated FADEC. Illumination of an amber light indicates a major malfunction in one channel of the associated FADEC. Illumination of both the white and amber lights indicates a malfunction in both channels of the associated FADEC. Dispatch is not permitted with any white or amber light illuminated. VARIABLE INLET GUIDE VANES AND VARIABLE STATOR VANES The engine is equipped with variable inlet guide vanes to direct air into the first stage axial compressor and variable stator vanes to direct air into the second stage axial compressor. This feature permits peak compressor efficiency throughout various operating conditions. A variable guide vane actuator is used to simultaneously position the guide vanes and stator vanes. The FADEC computes the desired vane position and commands the HFCU to provide servo pressures (fuel) to the actuator which positions the vanes. A rotary variable differential transformer (RVDT) position transducer, mounted on the actuator, sends an electrical feedback signal to the FADEC. SURGE BLEED CONTROL Each engine has a surge bleed control system which allows surge free operation throughout various operating conditions and improves engine starting characteristics. The system consists of a solenoid control valve and three bleed-off valves (BOV). Two valves bleed compressor air from station 2.5 while the third valve bleeds air from station 2.8. BOV position is controlled by the FADEC via the solenoid control valve. Compressor discharge air (P3) is used to provide servo pressure to close the bleed-off valves. The solenoid control valve applies P3 pressure to the BOVs to close them and vents P3 pressure to open them. In the event a solenoid control valve fails, the bleed-off valves will go to the open position.

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AUTOMATIC PERFORMANCE RESERVE (APR)

The APR system provides for an automatic change from the takeoff N1 rating to the APR rating for the operative engine in the event of loss of thrust from one engine during takeoff. The amount of thrust change will depend on ambient conditions. Since the engines installed on the Learjet 60XR are flat rated, the difference between takeoff and APR thrust will be very small under some ambient conditions. The system consists of an APR switch on the forward pedestal, APR ARM and APR ON indicators which display on the EIS Engine Page normally displayed on the pilots MFD, and associated aircraft wiring. To detect loss of thrust, the FADEC continuously monitors the opposite engines N1 and N2 signals. Loss of thrust is defined by the FADECs as meeting one or more of the following criteria: The N1 of one engine differs more than 15% from the N1 of the other engine. The N2 of one engine differs more than 7.5% from the N2 of the other engine. The N1 of one engine differs more than 4% from the N1 of the other engine and N1 is decreasing at a rate greater than 5% per second. The N2 of one engine differs more than 2% from the N2 of the other engine and N2 is decreasing at a rate greater than 2% per second. APR SWITCH APR system automatic operation is pilot controlled through the APR ARM-OFF switch located on the right side of the pedestal adjacent to the thrust levers. The switch is recessed to prevent inadvertent APR activation. The switch has two positions: OFF and ARM. For automatic operation the switch is set to ARM. When ARM is selected, the APR ARM indicator on the EIS will illuminate provided no faults exist which affect the APR function. When a loss of thrust is detected by one of the FADECs, an uptrim of the operative engine is commanded. The FADEC checks that the change to the appropriate APR N1 setting has been triggered and if it has, the APR ON indicator on the EIS will illuminate. Should automatic activation of APR fail to occur, APR thrust can be manually obtained by setting the thrust lever to the APR detent. In this case, the APR ON indicator on the EIS will not illuminate. Once invoked, the APR thrust schedule will remain active until the APR switch is set to OFF. APR ARM INDICATOR The green ARM indicator on the EIS will illuminate when the APR switch is in the ARM position provided no faults exist which affect the APR function. 2-6
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APR ON INDICATOR If APR is activated automatically by the FADEC, the amber APR ON indicator on the EIS will illuminate once APR thrust has been achieved. The APR ON indicator will not illuminate if APR thrust is obtained manually using the thrust lever detent.

ENGINE SYNCHRONIZER
The engine synchronizer system consists of two ENG SYNC switches, an amber or green SYNC indicator on the EIS Engine Page, and engine synchronizer circuits within the FADECs. During flight, the engine synchronizer, if selected, will maintain the two engines N1 or N2 in sync with each other. The engine synchronizer must not be used during takeoff, landing, or single-engine operations. Engine synchronization is not available on the ground or whenever APR is armed. Electrical power for the engine synchronizer is 28 VDC supplied through the ENGINE SYNC circuit breaker on the pilots circuit breaker panel. Synchronization is accomplished by maintaining the speed of the slave engine in sync with the speed of the master engine. The master engine is determined and so designated during installation. The following criteria must be satisfied before the system will operate: The ENG SYNC switch is set to SYNC. The difference between the N1 speed of each engine is no more than 5%. Thrust levers are in the range from IDLE to MCT. Thrust reversers are stowed. APR is disarmed. Deviating from any of these criteria will cancel engine synchronization. The system will raise flight idle of the master engine by a maximum of 1% N1 when activated. ENG SYNC SWITCHES Two ENG SYNC switches are installed on the pedestal immediately below the thrust levers. The ENG SYNC control switch is labeled SYNCOFF and the ENG SYNC selector switch is labeled N1-N2. When moved to the SYNC position, the control switch will activate the engine synchronizer and remove N1 Indicator compensation; therefore, the N1 and N1 bug presentations will reflect actual N1 speed. When SYNC is selected, N1 or N2 synchronization is selected by moving the ENG SYNC selector switch to N1 or N2 as desired.
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ENG SYNC INDICATORS The green SYNC indicator on the EIS will illuminate when the SYNCOFF switch is in the SYNC position. The amber SYNC indicator on the EIS and the amber ENG SYNC light on the glareshield will illuminate when the nose gear is not up and the SYNC-OFF switch is in the SYNC position.

GROUND IDLE SYSTEM

The ground idle system provides reduced engine idle speeds for ground operations. When the thrust lever is in the IDLE detent and the squat switch is in the ground mode, idle speed is reduced from approximately 65% N2 (flight idle) to approximately 52% N2 (ground idle). In flight, the idle speed setting is selected to ensure adequate transient response to full takeoff power. The system incorporates a 10-second delay after touchdown before ground idle is activated.

ENGINE OIL SYSTEM

The engine oil system provides lubrication and cooling for the mainshaft bearings, all accessory drive gears and all accessory bearings. The system consists of a pressure system, a scavenge system, and a breather system. PRESSURE SYSTEM The oil tank is an integral part of the engine intermediate case. Oil is drawn from the tank by a gear-type pressure pump. Pump output is directed through a pressure adjusting valve which bleeds excess pressure back to the pump inlet. From there, oil passes through an oil filter and fuel/oil heat exchanger before being routed to the mainshaft bearings, accessory drive gears, and accessory bearings. A cold-start valve diverts oil from the pump outlet into the accessory gearbox sump if pressure exceeds 200 psi during cold weather operation. The oil filter incorporates a bypass valve allowing oil to bypass the filter should it become clogged. An impending bypass indicator provides both a pop-up type visual indicator and an electrical signal to activate the ENG FILTERS light in the cockpit. To avoid false indications at engine start-up with cold oil, a thermal lockout inhibits the impending bypass indication if oil temperature is below 38 C (100 F).

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An anti-siphon device is incorporated to prevent oil from being siphoned out of the oil tank following engine shutdown. The device contains a small hole drilled through to the expansion space at the top of the oil tank. This breaks the siphon action caused by the oil tank level being higher than the main bearing oil jets.
SIPHON BREAK

Oil Tank

#1 Bearing

#2 Bearing

#3 Bearing

#4 Bearing

Breather Sump L ENG CHIP L OIL PRESS

FUEL/OIL HEAT EXCHANGER

L ENG FILTERS

OIL SUPPLY LINE OIL PRESSURE LINE OIL SCAVENGE LINE OIL BYPASS LINE ELECTRICAL

OIL FILTER BYPASS VALVE CHIP DETECTOR IMPENDING BYPASS INDICATOR

OIL TEMPERATURE SENSOR OIL PRESSURE SENSOR OIL PRESSURE SWITCH PRESSURE PUMP SCAVENGE PUMP PRESSURE ADJUSTING VALVE

COLD START VALVE STRAINER

ENGINE OIL SYSTEM SCHEMATIC Figure 2-2

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SCAVENGE SYSTEM The scavenge system incorporates three gear-type scavenge pumps installed in the accessory gearbox. Oil from the number 1 and 2 bearing compartments drains by gravity into the accessory gearbox sump. Oil from number 3 and 4 bearings is pumped by scavenge pumps into the accessory gearbox sump. Scavenge flow from all bearing compartments is aided by pressurizing airflow through the labyrinth air seals. Bypass valves are incorporated around the number 3 and 4 bearing scavenge pumps to prevent pressure build-up in the scavenge lines at higher bearing cavity pressure conditions. Oil collected in the accessory gearbox sump is pumped to the top of the oil tank by a separate scavenge pump. BREATHER SYSTEM Air from the bearing compartments, accessory gearbox, and oil tank is vented overboard through an impeller-type centrifugal air/oil separator installed in the accessory gearbox.

ENGINE IGNITION SYSTEM

Each engine ignition system consists of an IGNITION switch, a green annunciator, two ignition exciter units, two shielded cables, two igniter plugs, and associated aircraft wiring. The ignition exciter unit is a solidstate, high-voltage unit which provides a spark rate of 1 to 4 sparks per second at an output of 24,000 to 35,000 volts. The igniter plugs are mounted at four and five oclock positions in the combustion chamber case. The plugs are operated by separate cables and spark when pulsed by the ignition exciter units. During the start cycle, the ignition system is automatically energized by the FADEC when the thrust levers are placed in the IDLE position and N2 is above approximately 6%. The ignition system is automatically de-energized by the FADEC at approximately 40% N2. At pressure altitudes below 20,000 feet and TLA at or above IDLE, the FADEC will sequence the ignition system on should N2 speed fall below 40%. This feature provides for an immediate relight when the aircraft is below 20,000 feet. The ignition system may be operated continuously through the corresponding IGNITION switch. The ignition system light will be illuminated whenever the associated ignition system is operating either continuously (IGNITION On) or automatically (FADEC control). The ignition system is powered by 28 VDC from the L and R IGN CH A and IGN CH B circuit breakers on the pilots and copilots circuit breaker panels. The ignition system is operative during EMER BUS mode.

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IGNITION SWITCHES The IGNITION switches, located on the center switch panel, are used to obtain continuous engine ignition. The switch controlling the left engine ignition system is labeled L-OFF. The switch controlling the right engine ignition system is labeled R-OFF. When an IGNITION switch is placed in the On (L or R as applicable) position, 28 VDC from the corresponding L or R IGN CH A and IGN CH B circuit breakers is applied to the corresponding ignition exciter units. IGNITION LIGHTS Green lights above each IGNITION switch are installed to indicate ignition system operation. The corresponding light will be illuminated when the associated ignition system is operating either continuously (IGNITION On) or automatically (FADEC control).

ENGINE INDICATING SYSTEM (EIS)

The EIS Engine Page consists of full time displays, normally on the pilots MFD, of N1, ITT, N2, Fuel Flow, Oil Pressure, and Oil Temperature. The EIS Engine Page can be displayed on any Adaptive Flight Display (AFD) by pressing the SYS button on the respective DCP or pressing a line select key (LSK). Unless in reversionary mode, EIS pages normally displayed on the MFDs when selected to a different EIS page will redisplay after 20 seconds. The EIS Engine Page information is also available on the RTU STBY DISPLAY page.

EIS ENGINE PAGE Figure 2-3

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N1 INDICATORS There is a N1 indicator for each engine. Each indicator utilizes both a digital display and an arc-sweep display with a pointer to indicate N1. The N1 pointer shares the same sweep display as the ITT indicator for each engine. The digital display shows the fan speed to the nearest tenth of a percent. Each indicator also has a trapezoid-shaped N1 bug driven by a signal from the associated FADEC. The N1 bug represents the speed the engine should achieve given the ambient conditions, thrust lever setting, flap setting, and squat switch position. N1 is an indication of engine speed plus compensation. The FADEC takes into consideration its inputs to calculate and transmit the proper N1 bug settings for the ambient conditions. While airborne with the flaps up, the N1 bugs will show the proper N1 for the selected throttle detent or, if the throttles are in between detents, the next higher setting. While on the ground, or inflight with flaps 3 or lower, the N1 bugs will show takeoff power. On the ground with the thrust reversers deployed, the N1 bugs will show the maximum reverse N1 for the current conditions. Each engine FADEC has an externally mounted trim plug which provides trim compensation to the N1 signal. This trim plug will ensure consistent N1 indications for a specific paired throttle position. When ENG SYNC is On, compensation is removed. Each engine is also equipped with two induction-type speed sensors at the aft end of the low-pressure rotor. A toothed wheel is attached to the low-pressure shaft rotating adjacent to the stationary speed sensors. As the toothed wheel turns, its teeth cause the frequency output of the speed sensors to change proportionally. The frequency of the output signal represents the speed of the rotating N1 group. One sensor provides output signals to the N1 indicator, and channel A of the FADEC while the other sensor provides output signals to channel B of the FADEC and the opposite engines FADEC (used for APR and engine synchronizer). ITT INDICATORS There is an ITT indicator for each engine. Each indicator utilizes digital display and an arc-sweep display with a pointer to indicate ITT. The ITT pointer shares the same sweep display as the N1 indicator for each engine. The digital display shows the turbine temperature to the nearest degree. Interstage turbine temperature for each engine is sensed by Chromel-Alumel parallel wired thermocouples positioned between the high- and low-pressure turbine sections at engine station 4.5. The thermocouples provide an average T4.5 signal to the FADEC. The ITT indicator is driven by a signal from the FADEC.

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N2 INDICATORS There is a digital N2 display for each engine. The display shows the turbine speed to the nearest tenth of a percent. Each engine is equipped with two induction-type speed sensors installed on the right side of the accessory gearbox. The gearshaft teeth on the centrifugal impeller (within the accessory gearbox) rotate adjacent to the stationary speed sensors. As the gearshaft turns, its teeth cause the frequency output of the speed sensors to change proportionally. Since the accessory gearbox is driven by the N2 spool, the frequency of the output signal represents the speed of the rotating N2 group. One sensor provides output signals to the N2 indicator, and channel A of the FADEC while the other sensor provides output signals to channel B of the FADEC and the opposite engines FADEC (used for APR and engine synchronizer). FUEL FLOW (FF) INDICATION There is a digital fuel flow (FF) display for each engines fuel burn rate. The digital display indicates fuel flow to the nearest 10 pounds per hour. A fuel-flow transmitter (flowmeter) for each engine measures fuel flow by means of a rotary vane installed in the engine fuel supply line between the hydro-mechanical fuel control unit and the fuel dump valve. As fuel flows through the flowmeter, an amplitude-modulated constant-frequency sine wave signal is generated and applied to the fuel flow signal. The analog signal is converted to a digital signal of fuel burn rate (pounds per hour) for display. The Fuel Flow indicating system also provides a signal to the flight management system for each pound of fuel burned. ENGINE OIL INDICATIONS (Pressure and Temperature) There are two digital engine OIL displays for each engine one for pressure and one for temperature. The pressure ranges from 0 to 220 psi. The temperature ranges from -50C to 150C. A resistance-type temperature sensor located in an oil pressure line on each engine provides the temperature information. A pressure transducer which senses the pressure differential between the oil scavenge line and the oil pressure line on each engine provides the pressure information. OIL PRESSURE LIGHTS Red L OIL PRESS and R OIL PRESS warning lights are installed in the glareshield annunciator panel. In the event that either engines oil pressure drops below approximately 20 psi, a pressure switch connected to the oil pressure line and oil scavenge line of the affected engine will cause the applicable light to illuminate. Also, the applicable light will be illuminated whenever electrical power is on the aircraft and the corresponding engine is not operating.
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ENGINE CHIP LIGHTS Illumination of either amber L ENG CHIP or R ENG CHIP light indicates the presence of contaminants and debris in the corresponding engines oil system. The lights are activated by a magnetic chip detector installed in the scavenge oil passage of each engines accessory gear box. ENG FILTERS LIGHT Illumination of a white ENG FILTERS light on the glareshield annunciator panel indicates one or more of the following conditions: Impending bypass of the respective engine fuel filter Impending bypass of the respective engine oil filter Impending bypass of the respective airframe-mounted fuel filter The airframe-mounted fuel filter circuit is wired through the squat switch and may cause the ENG FILTERS light to illuminate only if the aircraft is on the ground. The engine fuel filter circuit is not wired through the squat switch and may cause the ENG FILTERS light to illuminate either in flight or on the ground. A maintenance panel, installed in the tailcone, is utilized by maintenance personnel to determine the specific filter causing the ENG FILTERS light to illuminate and to reset the system after the corrective action has been taken. ENG VIB LIGHTS Illumination of either amber L ENG VIB or R ENG VIB light indicates an abnormally high level of vibration in the associated engine. The lights are activated by a signal conditioning box located in the tailcone. A transducer installed on a mounting pad of each engines intermediate case provides the trigger to initiate an engine vibration caution. ENGINE DIAGNOSTIC SYSTEM (EDS) An EDS is installed to provide engine fault recording and trend monitoring. The system periodically records engine parameters and allows the crew to request that conditions be recorded at anytime. Normal use of the system entails downloading data from the EDS and submitting to Pratt and Whitney Canada for analysis on a monthly basis. The data may be downloaded at any time to assist in diagnosing engine problems which may be encountered. The EDS is intended for maintenance functions only and not for in-flight monitoring or diagnosis by the flight crew. 2-14
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The system consists of an Engine Diagnostic Unit (EDU), two isolation units (one for each engine), a Control Display Unit (CDU), a white EDS FAULT annunciator and an EDS RECORD switch on the center switch panel. The system is powered by 28 VDC through the ENGINE DIAGNOSTIC SYSTEM circuit breaker on the copilots circuit breaker panel.

FADEC

ISOLATION UNIT
AIRCRAFT EQUIPMENT

CDU EDU
ENT

564456 STREET CITY, STATE COUNTRY

GROUND SUPPORT EQUIPMENT

DATA CARD POST OR COURIER TO P&WC ANALYSIS AT P&WC

DATA TRANSFER BY OPERATOR TO P&WC VIA MODEM

ANALYSIS ON SITE

ENGINE DIAGNOSTIC SYSTEM Figure 2-4

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ENGINE DIAGNOSTIC UNIT (EDU) The EDU contains the memory used to store the collected data for each engine. The units capacity allows approximately 200 hours of data storage. The unit is installed in the tailcone. On the back of the EDU is a green, an amber, and a red light. The green light illuminates to indicate the EDS is powered. The red light illuminates to indicate the EDS has failed the self test. ISOLATION UNITS The isolation units are installed in the tailcone and provide protection for the FADECs in case of a fault in the engine diagnostic system. CONTROL DISPLAY UNIT (CDU) The CDU contains the display, control keys and connections necessary to control the system and download data. The CDU incorporates provisions to interface the system with a personal computer and provisions to download data onto a solid state data card. EDS FAULT ANNUNCIATOR The white EDS FAULT annunciator is located in the glareshield annunciator panel. Illumination of the light indicates one of the following: The EDS is off. The EDU Built In Test Equipment (BITE) has detected a system failure. The EDU memory is 85% full. The system has detected an engine condition which is out of acceptable parameters. EDS RECORD SWITCH The EDS RECORD switch is located on the center switch panel. The purpose of the switch is to allow the flight crew to initiate data collection by the EDS. When the switch is actuated, the engine parameters existing four minutes prior to and one minute after switch actuation will be recorded in the EDU memory.

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ENGINE FIRE DETECTION SYSTEM

Three heat-sensing elements connected in series are located in each engine nacelle to detect an engine fire. One element is located around the accessory gearbox; one is located around the engine tailcone; and another around the engine firewall. The fire detection system is controlled by two fire-detect control boxes located in the tailcone. In the event of an engine fire, the control box(es) will sense a resistance change in the sensing elements and flash the applicable ENG FIRE PULL light. The FIRE indicator on the EIS will illuminate inside the appropriate N1/ITT analog display. Electrical power for the system is 28 VDC supplied through the L and R FIRE DETECT circuit breakers on the pilots and copilots circuit breaker panels respectively. The fire detect system is operative during EMER BUS mode. SYSTEM TEST SWITCH FIRE DETECTION FUNCTION The rotary-type SYSTEM TEST switch on the instrument panel is used to test the fire detection system. Rotating the switch to FIRE DET and depressing the switch TEST button will connect a resistance into both fire detect system circuits. This resistance, simulating an engine fire, will cause both ENG FIRE PULL lights to illuminate and flash. It also tests and lights the ENG EXT ARMED lights. This test function also tests the tailcone bleed air overheat system. Depressing the TEST button will cause both red BLEED AIR L and BLEED AIR R lights to illuminate and the FIRE indicator on the EIS to illuminate. These tests check the heat-sensing elements for continuity. ENG FIRE PULL LIGHT A red ENG FIRE PULL warning light is part of a T-handle installed on the glareshield to warn the crew of a fire in the associated engine nacelle. In the event of an engine fire, the associated ENG FIRE PULL light will illuminate and flash. Operation of the T-handle is explained under ENGINE FIRE EXTINGUISHING SYSTEM.

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ENGINE FIRE EXTINGUISHING SYSTEM

The engine fire extinguishing system components include: two spherical extinguishing agent containers, an ENG FIRE PULL T-handle for each engine, two amber ENG EXT ARMED light/switches, a hydraulic shutoff valve for each engine, a fuel shutoff valve for each engine, a thermal discharge indicator, a manual discharge indicator, and associated wiring and plumbing. The system also utilizes the pneumatic system bleed-air shutoff valves. The system is plumbed to provide the contents of either or both extinguishing agent containers to either engine nacelle. Two-way check valves are installed to prevent extinguishing agent flow between containers. The extinguishing agent, Halon 1301 (bromotrifluoromethane [CF3Br]), is stored under pressure in the extinguisher containers and a pressure gage on each container is visible from inside the tailcone. Halon 1301 is non-toxic at normal temperatures and is non-corrosive. As Halon 1301 is non-corrosive, no special cleaning of the engine or nacelle area is required in the event the system has been used. The system operates on 28 VDC supplied through the L and R FIRE EXT circuit breakers on the pilots and copilots circuit breaker panels respectively. The fire extinguishing system is operative during EMER BUS mode. ENG FIRE PULL HANDLE AND ENG EXT ARMED LIGHTS The engine fire extinguishing system is operated through the ENG FIRE PULL T-handles and the ENG EXT ARMED lights located on either end of the glareshield annunciator panel. The ENG EXT ARMED lights are combination light/switches. When the ENG FIRE PULL T-handle is pulled, the associated engine fuel, hydraulic, and bleed-air shutoff valves will close to isolate the affected engine. The associated thrust reverser isolation valve will also close, shutting off hydraulic fluid to the associated thrust reverser. A solenoid valve in the HFCU shuts off fuel to the engine causing immediate shutdown, and both ENG EXT ARMED lights will illuminate. Illumination of the ENG EXT ARMED lights indicates that the fire extinguishing system is armed. Depressing an illuminated ENG EXT ARMED light will discharge the contents of an extinguisher bottle into the affected engine nacelle. Depressing the second ENG EXT ARMED light will discharge the contents of the other extinguisher bottle into the affected nacelle.

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ENG EXT ARMED ENG EXT ARMED

ENG FIRE PULL ENG FIRE PULL

BLEED AIR SHUTOFF VALVE

BLEED AIR SHUTOFF VALVE

PRESSURE GAUGE

PRESSURE GAUGE

#1 CONTAINER TWO-WAY CHECK VALVES

FUEL SHUTOFF VALVE

#2 CONTAINER FUEL SHUTOFF VALVE

RELIEF VALVE

RELIEF VALVE

FIRE EXTINGUISHING SYSTEM Figure 2-5

RED YELLOW THERMAL DISCHARGE INDICATOR MANUAL DISCHARGE INDICATOR

LH NACELLE

HYDRAULIC SHUTOFF VALVE

HYDRAULIC SHUTOFF VALVE

RH NACELLE

HFCU

HFCU

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FIRE EXTINGUISHER DISCHARGE INDICATORS Two disk-type indicators are flush-mounted in the fuselage under the left engine pylon. If the contents of either or both containers have been discharged into the engine nacelles, the yellow disk will be ruptured. If the contents of either or both containers have been discharged overboard as the result of an overheat condition causing excessive pressure within the containers, the red disk will be ruptured. If both disks are intact, the system has not been discharged. The indicators are readily accessible for visual inspection and must be checked for condition prior to each flight.

THRUST REVERSER SYSTEM

Each engine is equipped with an independent, electrically controlled, hydraulically actuated, target-type thrust reverser. The thrust reverser system consists of a thrust reverser assembly installation on each engine, thrust reverser levers on the main thrust levers, a throttle balk solenoid, associated hydraulic plumbing and associated electrical wiring. Each thrust reverser assembly installation consists of an upper and lower target-type door, four-bar door linkage, an inboard and outboard door actuator, two secondary latches, four stow switches and one deploy switch. A hydraulic control unit (HCU) for each thrust reverser is installed in the tailcone. The HCU controls the hydraulic flow to the associated thrust reverser in response to electrical inputs. Hydraulic power for thrust reverser operation is supplied by a combination of engine driven hydraulic pump flow and a thrust reverser accumulator. Pressure from the auxiliary hydraulic pump is not available to the thrust reverser system. The thrust reverser accumulator is plumbed primarily to power thrust reverser operations but assists the main system accumulator for landing gear, flap and brake operation. Refer to Section III for more details on the thrust reverser hydraulic system. Electrical power for thrust reverser control and auto stow functions is 28 VDC supplied through the L and R TR CONT and the L and R TR AUTO STOW circuit breakers on the pilots and copilots circuit breaker panels. The WARN LTS circuit breakers supply electrical power for FADEC discrete signals and a redundant power source for the annunciator circuits. The status of the thrust reversers is indicated on the EIS Engine Page in the lower portion of the N1/ITT analog display.

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DEPLOY In order to arm a thrust reverser, both squat switches must be in the ground mode (aircraft weight on wheels), and the applicable thrust lever must be in the IDLE detent. When the prerequisite conditions are met, a signal from the applicable thrust reverser relay box will open the applicable isolation valve (within the HCU) allowing hydraulic pressure to be available for thrust reverser deployment. The presence of hydraulic pressure will actuate a pressure switch and illuminate the green REV indicator on the EIS. Lifting the thrust reverser lever to the DEPLOY detent will signal the applicable HCU to apply hydraulic pressure to the secondary latch actuators and deploy port of the thrust reverser actuators (inboard and outboard). When the secondary latches are released, the secondary latch stow switches send a signal to illuminate the amber UNL indicator on the EIS. Once the thrust reverser doors move out of the stowed position, the primary latch stow switches send a discrete signal to the on-side FADEC to limit engine thrust to idle. When the doors reach the fully deployed position, the deploy switch sends a signal to illuminate the white DEP indicator on the EIS and a discrete signal is sent to the on-side FADEC to allow engine thrust to increase above idle. The N1 bug will reposition indicating the FADEC is utilizing the reverse thrust schedule. A throttle balk solenoid prevents either thrust reverser lever from moving significantly above reverse idle until both thrust reversers are fully deployed. Once the deploy switches on both thrust reversers are actuated, the solenoid is energized allowing the thrust reverser levers to move into the reverse thrust range. STOW To stow the thrust reverser, the thrust reverser lever is moved into the STOW position. The thrust reverser relay box will signal the applicable HCU to apply hydraulic pressure to the stow port of the thrust reverser actuators (inboard and outboard). Once the thrust reverser doors move out of the deployed position, the deploy switch sends a signal to illuminate the amber UNL indicator on the EIS and a discrete signal is sent to the on-side FADEC to limit engine thrust to idle. When the doors reach the stowed position, the primary latch stow switches send a discrete signal to the on-side FADEC to restore engine thrust. The upper and lower doors trip their respective spring-loaded secondary latches as they reach the stowed and locked position. At this point, the secondary latch stow switches send a signal to remove the amber UNL indicator from the EIS.

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AUTO STOW The thrust reverser doors are mechanically secured in the stowed position by a four-bar overcenter door linkage (primary latch). Should an uncommanded unlock condition be sensed by the primary latch stow switches, an auto stow sequence will be initiated and the UNL indicator on the EIS will illuminate (amber on the ground or red in flight). The thrust reverser relay box will command the HCU to open the isolation valve and apply hydraulic pressure to the stow port of the thrust reverser actuators (inboard and outboard). A primary latch unlock condition will result in a discrete signal being sent to the on-side FADEC to limit thrust to flight idle, regardless of throttle position, until the thrust reverser is returned to the stowed position. An unlock condition sensed by the secondary latch stow switches will illuminate the UNL indicator on the EIS (amber on the ground or red in flight) but will not initiate the auto stow sequence. THRUST REVERSER ASSEMBLY Each engine is equipped with a thrust reverser assembly attached to the engine outer fan duct. When stowed, the thrust reverser fairs with the nacelle and forms the engine afterbody. Each upper and lower door is attached to the support structure by a four-bar linkage. Two links are idler links and two are driver links. The driver links connect to the inboard and outboard actuators with an overcenter link. After stowing the doors, the actuators continue to drive the overcenter links to an overcenter position. This provides a mechanical latch to keep the doors stowed. This overcenter mechanism is referred to as the primary latch. In addition to the primary latch, each thrust reverser door is held in the stowed position by a secondary latch. A latch plate on each door engages the spring-loaded secondary latch mechanism securing the door in the stowed and locked position. During the deployment sequence, each secondary latch is released by hydraulic pressure from the deploy line. Each assembly is equipped with two primary latch stow switches, two secondary latch stow switches, and one deploy switch. The primary latch stow switches are used to detect the extreme aft (locked) position of the inboard and outboard actuators. The secondary latch stow switches are used to detect the engagement of the secondary latch with the thrust reverser doors. The deploy switch is actuated by one of the idler links and detects the fully deployed position. These switches provide signals to sequence the thrust reverser operation, control the thrust reverser annunciators, control the throttle balk solenoid and initiate the auto stow sequence. 2-22
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AIRCRAFT HYDRAULIC SYSTEM

Pressure

THRUST REVERSER RELAY BOX

FADEC

Return

Deploy Light
Deploy Switch Stow Switch

Arm

HYDRAULIC CONTROL UNIT

Arm Light Stow Deploy

Thrust Reverser Actuators

DEPLOY

STOW

Unlock Light

THRUST REVERSER SYSTEM SCHEMATIC Figure 2-6 2-23/2-24 (Blank) 2-23

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THRUST REVERSER LEVER A thrust reverser lever is mounted piggy-back fashion on each main thrust lever. The thrust reverser lever cannot be moved out of the STOW position unless the associated main thrust lever is at the IDLE stop. Similarly, the main thrust lever cannot be moved from the IDLE position when the associated thrust reverser lever is in the DEPLOY and reverse thrust range. Moving the main thrust lever to IDLE actuates a switch in the throttle quadrant to signal the system to arm if the aircraft is on the ground. Another switch in the throttle quadrant is actuated by the thrust reverser lever and signals the system to stow or deploy the associated thrust reverser. When both thrust reversers are fully deployed, the thrust reverser levers are allowed to move beyond the DEPLOY detent into the reverse thrust range. Moving the thrust reverser lever above reverse idle allows the engine to spool up providing the desired amount of reverse thrust. The FADEC will schedule reverse thrust as a function of airspeed (provided by ADC 1 and 2), decreasing thrust as the airplane slows down. If airspeed data is not provided to the FADEC, the maximum reverse thrust available will be 65% N1. THROTTLE BALK SOLENOID A throttle balk solenoid is installed in the pedestal to mechanically prevent either thrust reverser lever from moving into the reverse range until both thrust reversers are fully deployed. When the solenoid is deenergized, a spring-loaded lockout mechanism allows the thrust reverser levers to move between the STOW and DEPLOY positions only. When energized, the solenoid will overcome the spring-loaded lockout mechanism allowing the thrust reverser levers to move beyond the DEPLOY position into the reverse thrust range

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HYDRAULIC CONTROL UNIT (HCU) The HCU functions as a shutoff valve to isolate the thrust reverser system from the aircrafts hydraulic system and also as a selector valve directing hydraulic fluid to stow and deploy the thrust reverser doors as commanded. The HCU incorporates both a mechanical and an electrical isolation valve. The mechanical valve may be manually closed and secured with a locking pin thereby deactivating the thrust reversers. The electrical valve is closed until the conditions for arming are satisfied or the auto stow sequence is initiated. The electrical signals to operate the HCU come from the applicable thrust reverser relay box. When the left or right ENG FIRE PULL T-handle is pulled, the associated isolation valve will close, shutting off hydraulic fluid to the associated thrust reverser. A pressure switch, in the HCU, senses hydraulic pressure availability to the selector valve. When pressure is present, the switch will illuminate the REV indicator on the EIS (green on the ground and amber in flight). Each HCU incorporates a check valve in the hydraulic return port which allows free flow from the HCU to the aircrafts hydraulic return system but no flow in the reverse direction. THRUST REVERSER RELAY BOX Two thrust reverser relay boxes are installed in the tailcone. One box controls the left thrust reverser system and the other controls the right. Inputs to each relay box are provided from: left and right squat switches, arming switch (throttle quadrant), stow/deploy switch (throttle quadrant), stow switches (thrust reverser assembly), deploy switch (thrust reverser assembly), and pressure switch (HCU). From the input signals the relay box determines the appropriate output signals including: arm thrust reverser (open isolation valve in the HCU), deploy thrust reverser, stow thrust reverser, initiate auto stow, limit engine thrust to idle (discrete signal to FADEC), restore engine thrust to normal (discrete signal to FADEC), enable thrust reverser levers (throttle balk solenoid), annunciate thrust reverser conditions and indicate to the takeoff monitor whether the thrust reverser is locked or unlocked.

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AIRCRAFT FUEL SYSTEM

The aircraft fuel system consists of two wing tanks, a fuselage fuel tank, a fuel supply system, a fuel quantity indicating system, a fuel transfer system and a fuel vent system. Fuel fillers are located outboard near each wing tip. A single-point pressure refuel (SPPR) system is also installed. WING TANKS The wing is divided by a center bulkhead into two separate fuel-tight compartments which serve as fuel tanks. Each tank extends from the center bulkhead outboard to the wing tip rib, thus providing a separate fuel supply for each engine. A tank crossflow valve is installed to permit fuel transfer between wing tanks. Center bulkhead relief valves prevent wing tank overpressurization during fuel crossflow operations. Flapper-type check valves, located in the various wing ribs, allow free fuel flow inboard but restrict outboard fuel flow. A jet pump and an electric standby pump are mounted in each wing tank near the center bulkhead to supply fuel under pressure to the respective engine fuel system. An electric scavenge pump, located in the forward inboard section of each wing tank, is used to transfer fuel to the section containing the main fuel pumps and is operated by the low-fuel float switch. Three jet-type transfer pumps, located along the aft portion of each wing tank, transfer fuel to the section containing the main fuel pumps. A filler cap, located in the outer section of the wing tank, is used for fuel servicing. FUSELAGE TANK The fuselage tank, installed in the aft fuselage, consists of two interconnected bladder-type cells. The fuselage tank is provided with two transfer pumps, a float switch, a fuel quantity probe, and single-point pressure refuel provisions. The fuselage tank can be refueled by pumping wing fuel with the wing tank standby pumps through both transfer lines or by using the single-point pressure refuel system. Fuel can be transferred to the wing tanks by normal fuel transfer, auxiliary fuel transfer, rapid fuel transfer or gravity transfer. During the normal fuel transfer, the left fuselage tank transfer pump will pump fuel into both wing tanks. During the auxiliary fuel transfer, the right fuselage tank transfer pump will pump fuel into both wing tanks. During rapid fuel transfer, both the normal and auxiliary fuel transfer modes are energized. During gravity transfer, fuel will flow to both wing tanks through both transfer lines. FUEL CONTROL PANEL SWITCHES AND ANNUNCIATORS The fuel control panel incorporates all the necessary switches to maintain proper fuel management and to fuel the aircraft.
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FUEL SYSTEM
L WING
ON OFF JET PUMP STBY PUMP XFLO VALVE

R WING
ON STBY PUMP OFF JET PUMP

LO FUEL PRESS

N X O F RR M

ON

ON

A X U F X R

LO FUEL PRESS

L ENG

ON GRVTY XFR

FULL EMPTY

ON FILL

R ENG

FUSELAGE
FUEL CONTROL PANEL Figure 2-7 JET PUMP SWITCHES The JET PUMP switches, on the fuel control panel, control the motive flow valves. The switches are an alternate action type. Selecting On, opens the corresponding motive flow valve and allows high-pressure fuel from the corresponding engine-driven fuel pump to flow to the corresponding jet pumps. Selecting OFF, closes the corresponding motive flow valve and renders the associated jet pumps inoperative. When OFF is selected, an OFF annunciation (on the switch) will illuminate and the Master CAUT lights will flash (Master CAUT will not illuminate during engine start). If a motive flow valve is neither open nor closed, the corresponding OFF annunciator will flash. The motive flow valves operate on 28 VDC supplied through the L and R JET PUMPXFR VALVE circuit breakers on the pilots and copilots circuit breaker panels. Loss of power to the motive flow valve causes the valve to remain in its last position. Motive flow valves are operative during EMER BUS mode.

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LOW FUEL
XFLO VALVE

FUEL PROBE FLOAT SWITCH FILLER SQUAT SWITCH RELAY TRANSFER PUMP SCAVENGE PUMP ENGINE FUEL PUMP STANDBY PUMP JET PUMP PRESS RELIEF VALVE FUEL FILTER CROSSFLOW VALVE
S S S L FUEL PRESS ENG FIRE PULL ENG FIRE PULL R FUEL PRESS L ENG FILTERS R ENG FILTERS

SHUTOFF VALVE MOTIVE FLOW VALVE TRANSFER VALVE WING FLOAT RELAY PRESSURE SWITCH RELIEF VALVE CHECK VALVE HIGH PRESSURE FUEL LOW PRESSURE FUEL ELECTRICAL
GEN OFF START

OPEN

M T

OPEN

NX OF RR M

T
ON

ON GRVTY XFR ON

AX U F XR

OPEN

ON FILL

FULL EMPTY

GEN
OFF JET PUMP OFF JET PUMP

OFF START

ON STBY PUMP

ON STBY PUMP

FUEL SYSTEM SCHEMATIC Figure 2-8

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STBY PUMP SWITCHES The STBY PUMP switches, on the fuel control panel, control the operation of the standby electric pumps. The switches are an alternate action type. The switches normally remain Off except in the event of a jet pump failure or during fuel crossflow. Regardless of switch position, the standby pumps are automatically de-energized during fuselage fuel transfer operations. The standby pumps are automatically energized when the fuselage tank FILL function is selected or the STARTGEN switch is set to START. An ON annunciation (on the switch) will illuminate whenever power is applied to the corresponding standby pump. The green FUEL SYS light, on the glareshield annunciator panel, will also illuminate whenever a standby pump is on. The standby pumps operate on 28 VDC supplied through the L and R STBY-SCAV PUMP circuit breakers on the pilots and copilots circuit breaker panels. XFLO VALVE SWITCH The XFLO VALVE switch, on the fuel control panel, controls the crossflow valve. The switch is an alternate action type. Selecting Open, opens the crossflow valve allowing fuel to flow between the wing tanks. Whenever the crossflow valve is open, a horizontal bar (on the switch) will illuminate to annunciate the valves open status. The green FUEL SYS light will also illuminate whenever the crossflow valve is fully opened. If the crossflow valve is neither open nor closed, the horizontal bar will flash. The crossflow valve is opened automatically when filling the fuselage tank from the wings and during fuselage fuel transfer operations. To balance wing fuel, the XFLO VALVE switch should be set to Open and the heavy side STBY PUMP switch set to ON. The standby pump on the light side should be OFF. The standby pump will continue to operate until the STBY PUMP switch is set to Off. The crossflow valve allows all usable wing fuel aboard the aircraft to be available to either engine. The switch should be set to Off except when correcting an out-of-balance condition. The crossflow valve operates on 28 VDC supplied through the XFLO VALVE circuit breaker on the pilots circuit breaker panel. Loss of power to the crossflow valve causes the valve to remain in its last position. The crossflow valve is operative during EMER BUS mode.

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NORM XFR SWITCH The NORM XFR switch, on the fuel control panel, is used to operate the normal (left) fuel transfer system. The switch is an alternate action type. When NORM XFR is selected, the left transfer pump is energized, the left transfer valve will open, both standby pumps will be rendered inoperative, and the crossflow valve will open. Fuel will then be pumped from the fuselage tank to the wing tanks until the wing float switches actuate to de-energize the transfer pump and close the transfer valve (the crossflow valve will remain open). If the fuselage tank should empty before the wing float switches shut down the left transfer system, a pressure switch in the fuselage tank transfer line will illuminate the EMPTY light. The green FUEL SYS light will illuminate when NORM XFR is selected and flash whenever the EMPTY light illuminates. Setting the switch to Off will extinguish the EMPTY light (if illuminated), close the left transfer valve, de-energize the left transfer pump, enable the standby pumps, and close the crossflow valve. Whenever the left transfer valve is open, a vertical bar (on the switch) will illuminate to annunciate the valves open status. If the transfer valve is neither open nor closed, the vertical bar will flash. An ON annunciation (on the switch) will illuminate whenever power is applied to the left transfer pump. The left fuel transfer valve operates on 28 VDC supplied through the L JET PUMP-XFR VALVE circuit breaker on the pilots circuit breaker panel. Loss of power to the left transfer valve causes the valve to remain in its last position. The left transfer pump operates on 28 VDC supplied through the FUS TANK XFR PUMP circuit breaker on the pilots circuit breaker panel. Both the valve and pump are operative during EMER BUS mode.

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AUX XFR SWITCH The AUX XFR switch, on the fuel control panel, operates the auxiliary (right) fuel transfer system which provides an alternate transfer system in the event the normal system fails or, when used in conjunction with the normal system, allows rapid transfer of fuselage fuel if desired. The switch is an alternate action type. When AUX XFR is selected, the right fuselage transfer pump is energized, the right transfer valve will open, both standby pumps will be rendered inoperative, and the crossflow valve will open. Fuel will then be pumped from the fuselage tank into the wing tanks. The switch should be set to Off when either the EMPTY light illuminates or the wing tanks become full. The green FUEL SYS light will illuminate when AUX XFR is selected and flash whenever the EMPTY light illuminates. Setting the switch to Off will close the right transfer valve, de-energize the right transfer pump, close the crossflow valve, enable the standby pumps, and extinguish the EMPTY light, if illuminated. Actuation of the wing float switches has no effect on the auxiliary (right) fuel transfer system. Therefore, if the switch is not set to OFF when the wing tanks are full, fuel will continue to circulate between the fuselage and wing tanks through the wing expansion and fuel transfer lines. When the fuselage tank is emptied, a pressure switch in the right transfer line will actuate to illuminate the EMPTY light. Whenever the right transfer valve is open, a vertical bar (on the switch) will illuminate to annunciate the valves open status. If the transfer valve is neither open nor closed, the vertical bar will flash. An ON annunciation (on the switch) will illuminate whenever power is applied to the right transfer pump. The right fuel transfer valve operates on 28 VDC supplied through the R JET PUMP-XFR VALVE circuit breaker on the copilots circuit breaker panel. Loss of power to the right transfer valve causes the valve to remain in its last position. The right transfer pump operates on 28 VDC supplied through the FUS TANK AUX PUMP circuit breaker on the copilots circuit breaker panel. Both the valve and pump are operative during EMER BUS mode.

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GRVTY XFR SWITCH The GRVTY XFR switch, on the fuel control panel, can be used to transfer fuselage fuel without using the transfer pumps. The switch is an alternate action type. When GRVTY XFR is selected, both transfer valves will open, the crossflow valve will open, and both standby pumps will be rendered inoperative. Fuel will then gravity flow from the fuselage tank to the wing tanks until the wings are full or the wing and fuselage tank heads are equal. When using this method to transfer fuel, approximately 350 pounds (159 kilograms) of fuel will remain in the fuselage tank and the EMPTY light will be inoperative. To assure all possible fuel has been transferred, reference must be made to the fuel quantity indicator. The switch should be set to Off when all fuel possible has been transferred and during approach and landing. The green FUEL SYS light and an ON annunciation (on the switch) will illuminate whenever gravity transfer is selected. Gravity transfer is operative during EMER BUS mode. FILL SWITCH The FILL switch, on the fuel control panel, is used to operate the fuselage tank fill system. The switch is an alternate action type and must be held approximately 3 seconds to select the FILL function. When FILL is selected, both wing tank standby pumps are energized, both left and right transfer valves are opened via the fuselage tank float switch, and the crossflow valve will open. Fuel will then be pumped into the fuselage tank from the wing tanks until the switch is turned Off or the fuselage tank float switch actuates to close the transfer valves, shut down the standby pumps, and illuminate the FULL light. Placing the switch in the Off position will extinguish the FULL light and close the crossflow valve. The green FUEL SYS light and an ON annunciation (on the switch) will illuminate whenever fuselage tank fill is selected. If FILL is selected and the left wing float switch trips the LOW FUEL light or the squat switch goes to the air mode, the fuselage tank fill function will be automatically deselected. The FILL function may be subsequently reselected, if desired.

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FUSELAGE TANK SWITCH PRIORITY The FUSELAGE Tank switches are listed below in their order of priority (highest to lowest). If the FUSELAGE Tank switches are positioned to contradictory positions, the function with the highest priority will override conflicting functions. 1. NORM XFR and AUX XFR switches (both have same priority) 2. FILL switch 3. GRVTY XFR switch FUSELAGE TANK FULL LIGHT The FUSELAGE FULL light, on the fuel control panel, is installed to indicate a fuselage tank full condition during fuselage tank fill operations. The light is illuminated through actuation of the fuselage tank float switch. During normal fuselage tank fill operations, actuation of the float switch will illuminate the FULL light, close the transfer valves, and shut down the standby pumps. The FILL switch must be set to Off to extinguish the light. FUSELAGE TANK EMPTY LIGHT The FUSELAGE EMPTY light, on the fuel control panel, is installed to indicate a fuselage tank empty condition during fuel transfer. The light is operated by pressure switches in the left and right fuselage fuel transfer lines. As the fuselage tank empties during transfer operations, the pressure switches sense a loss of pressure in the transfer line and complete circuits to illuminate the EMPTY light. Either pressure switch can illuminate the light. Setting the NORM XFR and/or AUX XFR switch (as applicable) to Off will extinguish the light. LO FUEL PRESS LIGHTS The two LO FUEL PRESS lights, on the fuel control panel, repeat the L and R FUEL PRESS annunciators on the glareshield panel. See FUEL SYSTEM GLARESHIELD LIGHTS, this section.

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FUEL GAGING SYSTEM The fuel gaging system consist of a fuel quantity indicator installed in the cockpit, fuel quantity probes located in the various fuel tanks, and an optional total quantity indicator located near the single point pressure refueling controls. The fuel gaging system operates on 28 VDC supplied through the FUEL QTY PWR 1 and FUEL QTY PWR 2 circuit breakers on the pilots and copilots circuit breaker panels. The fuel gaging system is operative during EMER BUS mode. FUEL QUANTITY INDICATOR The fuel quantity indicator, on the instrument panel, indicates fuel quantity in pounds (or optionally kilograms) of fuel. The indicator has four digital readouts one for the left wing tank, one for the right wing tank, one for the fuselage tank, and one which shows the total of the other three summed together. Inputs from the attitude heading reference system are used to correct the fuel quantity indication for aircraft pitch attitude. The indicator incorporates a feature to alert the crew of a fuel imbalance between the left and right wing tanks. Should a fuel imbalance of 500 pounds, (200 pounds if flaps are 8 or lower) or more occur, the fuel quantity reading representing the heavy wing and the IMB annunciator, on the fuel quantity indicator, will flash. The flashing annunciations may be cancelled by depressing and releasing the mute switch in the right thrust lever. FUEL QUANTITY PROBES Fuel quantity is sensed by four capacitance-type fuel quantity probes in each wing tank and a capacitance-type fuel quantity probe in the fuselage fuel tank. The left inboard fuel quantity probe incorporates a fuel temperature compensator which compensates for fuel density changes due to temperature. TOTAL QUANTITY INDICATOR (SPPR) The optional total quantity indicator, located with the single point pressure refueling controls, indicates total fuel quantity in pounds of fuel. The system may also be configured to indicate kilograms of fuel. The indicator has a digital readout which repeats the total indication shown on the cockpit indicator. Refueling personnel can use the indicator to determine the total fuel load without reference to the cockpit indicator.

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FUEL SYSTEM GLARESHIELD LIGHTS FUEL PRESS LIGHTS The red L FUEL PRESS and R FUEL PRESS warning lights in the glareshield annunciator panel are installed to alert the pilot of a low fuel pressure condition. The FUEL PRESS lights are energized by a pressure switch installed in each engine fuel supply line between the aircraft fuel filter and the engine-driven fuel pump. When fuel supply pressure drops to 2.75 psi or below, the pressure switch closes to illuminate the respective light. At 3.75 psi, the switch will reopen. Should the light illuminate, the standby pumps should be used to supply engine fuel. The fuel control panel incorporates two LO FUEL PRESS lights which illuminate in conjunction with the associated glareshield warning light. LOW FUEL LIGHT The amber LOW FUEL caution light in the glareshield annunciator panel will illuminate when the fuel quantity in either wing tank decreases to approximately 410 pounds (186 kilograms) of fuel with the aircraft in a level attitude. The light is operated by a low wing fuel float switch installed in each wing tank. Either float switch may cause the light to illuminate. FUEL SYS LIGHT The green FUEL SYS light in the glareshield annunciator panel will illuminate whenever a fuel transfer function is selected on the fuel control panel. The following conditions cause the light to illuminate: Crossflow valve is fully opened Either transfer valve (left or right) is open NORM, AUX, or GRVTY XFR is selected FILL is selected Either standby pump is on The following conditions cause the light to flash: The fuselage EMPTY light is illuminated The fuselage FULL light is illuminated

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RAM AIR FUEL VENT SYSTEM The fuel vent system provides ram air pressure to all interconnected components of the fuel system to ensure positive pressure during all flight conditions. Flush mounted underwing scoops (inboard) admit pressure to the fuselage vent system, and a separate set of underwing scoops (outboard) admit pressure for the wing vent systems. The fuselage vent line is connected to a sump that has a moisture drain valve. Each wing tank vent system has a sump with a moisture drain valve located next to the wing vent underwing scoops. Overpressurization due to thermal expansion in the wing tanks is relieved through the left and right expansion lines to the fuselage tank. Overpressurization of the fuselage tank, should the vent and expansion lines be clogged, is relieved overboard through a pair of pressure relief valves and a separate vent line. SINGLE-POINT PRESSURE REFUEL (SPPR) SYSTEM The single-point pressure refueling (SPPR) system allows the entire fuel system to be serviced through a fuel servicing adapter located on the right side of the aircraft below the engine pylon. An SPPR control panel is located immediately forward of the refuel adapter. The SPPR incorporates a precheck system which allows the operator to check the operation of the system vent and shutoff valves before commencing refuel operations. The major system components are the refuel adapter, the control panel, a vent valve, a shutoff valve and pilot valve for each tank (both wings and fuselage), solenoid valve for the fuselage tank, two precheck valves, and associated plumbing and wiring. The control panel is located on the right fuselage below the engine pylon. Electrical power to operate the system indicator lights and solenoid valve is 28 VDC supplied from the #2 battery through the BATT ON-OFF switch on the refuel control panel. The vent valve is installed to prevent system overpressurization in the event of a shutoff valve failure. Operation of the valve is checked during the precheck sequence. The valve automatically opens whenever fuel pressure is applied to the system. When the valve reaches the full open position, a switch in the valve completes a circuit to illuminate the VENT OPEN light on the SPPR control panel.

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VENT OPEN

FUS FULL

G
VENT VALVE

A
1

TOTAL
1 2 Vent open. Vent open and wing float switches (full).
FUS PILOT VALVE FUS FLOAT SWITCH SOLENOID SHUTOFF

PARTIAL
2
REFUEL ADAPTER

FUS PRECHECK VALVE

FUSELAGE TANK

WING PRECHECK VALVE

FUS FUEL SHUTOFF

WING FLOAT SWITCH WING PILOT VALVE WING FUEL SHUTOFF WING FUEL SHUTOFF WING PILOT VALVE

WING FLOAT SWITCH

L WING TANK

R WING TANK

SINGLE-POINT REFUEL SYSTEM SCHEMATIC Figure 2-9

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Each shutoff valve is controlled by the associated pilot valve located at the high point in each tank. When refueling pressure is applied to the system through the refuel adapter, pressurized fuel is applied to each shutoff valve. This pressure is applied to both sides of the valve poppet. If the pilot valve is open (associated tank not full), some of the pressure acting to hold the valve closed will be vented through the pilot valve and the pressure acting to unseat the poppet will drive the valve open against the spring tension. When the tank fills, the pilot valve will close, fuel pressure on both sides of the shutoff valve poppet will equalize, and spring tension will drive the valve closed. The solenoid valve for the fuselage tank is located between the tank pilot valve and shutoff valve in the vent line. This valve is normally closed and must be energized open in order to open the shutoff valve for filling the tank. The valve is used to isolate the fuselage tank if filling that tank is not desired. WING AND FUS PRECHECK VALVES The WING and FUS PRECHECK valves are used to check operation of the system vent valve and individual shutoff valves before full refueling procedures are commenced. System precheck is accomplished with the Refuel Selector switch set to TOTAL in order to check all shutoff valves. When the WING and FUS PRECHECK valves are set to OPEN (grips vertical) and refuel pressure is applied to the refuel adapter, fuel will be admitted to the precheck lines and to the tank fill lines. The shutoff valves will open and fuel will flow into all tanks. The fuel in the precheck lines will empty into a float basin at each pilot valve. When the basin fills the pilot valve float will close the pilot valve, which causes the associated shutoff valve to close terminating fuel flow. The vent valve should open when fuel flow is initiated. Fuel flow should stop within 10 to 20 seconds. SPPR BATT SWITCH The BATT ON-OFF switch, on the refuel control panel, allows operation of the single-point pressure refuel system without the need to enter the cockpit in order to energize aircraft power. When the switch is set to ON, DC power from the aircrafts #2 battery is applied to the SPPR control circuits.

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REFUEL SELECTOR SWITCH The Refuel Selector switch, on the SPPR fuel control panel, is used to select the tank(s) to be filled during refueling. The switch has two positions: TOTAL and PARTIAL. The TOTAL position of the Refuel Selector switch is used to fill the wing and fuselage tanks simultaneously. When TOTAL is selected and refueling pressure is applied (vent valve opens), circuits are completed to open the fuselage tank solenoid valve. When the solenoid valve opens the fuselage tank shutoff valve will open to admit fuel into the fuselage tank. The PARTIAL position of the Refuel Selector switch is used to fill the wings first and then the fuselage. This is useful when full wings and less than full fuselage fuel is desired. When PARTIAL is selected and the vent valve opens, the fuselage tank solenoid valve will be controlled by the wing high-level float switches. When the wings are full, the wing high-level float switches complete the circuit to open the fuselage tank solenoid valve. When the solenoid valve opens, the fuselage tank shutoff valve will open and admit fuel to the fuselage tank. FUS FULL LIGHT The amber FUS FULL light, on the refuel control panel, will illuminate whenever the fuselage tank float switch actuates. The light illuminates to alert the operator that refuel operations should have automatically terminated. If fuel flow continues with the light illuminated, fueling operations should be immediately terminated. VENT OPEN LIGHT The green VENT OPEN light, on the refuel control panel, will illuminate whenever the fuselage tank vent valve opens. The light is operated by a microswitch in the valve. The circuit for the fuselage tank solenoid valve is wired through this switch to prevent filling the fuselage tank until the vent valve opens.

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16 15 14 13 17

1 2 3 4

12 11 10 9 8 7

1. 2. 3. 4. 5. 6. 7. 8. 9.

Left Wing Scavenge Pump 10. Fuselage Tank Sump Left Wing Sump 11. Right Wing Transfer Line Left Engine Fuel 12. Right Wing Expansion Line Left Wing Vent (sump) 13. Right Wing Vent (sump) Left Wing Expansion Line 14. Right Engine Fuel Left Wing Transfer Line 15. Right Wing Sump Fuel Vent (fuselage) 16. Right Wing Scavenge Pump Left Fuel Filter 17. Fuel Crossover Right Fuel Filter FUEL DRAINS Figure 2-10 2-43

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FUEL ANTI-ICING ADDITIVE Anti-icing additive is not a requirement. However, for microbial protection, it is recommended that anti-icing additive be used at least once a week for aircraft in regular use and whenever a fueled aircraft will be out of service for a week or more. Refer to the Airplane Flight Manual for the recommended concentration and the proper method of blending anti-icing additive. REFUELING The aircraft may be refueled through filler caps on each wing tip or through the single-point pressure refuel adapter on the right fuselage below the engine pylon. Bonding jacks are located on the underside of each wing near the fuel filler and behind the SPPR control panel door. Refer to the Airplane Flight Manual for approved fuels and proper refueling procedures.

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AUXILIARY POWER UNIT (APU)

The Auxiliary Power Unit (APU), located in the rear equipment bay, is a self-contained, single stage gas turbine unit that can be operated continuously up to an ambient temperature of 130 F (54 C). The APU provides electric power for ground operations of the aircraft electrical system, independent of the aircraft main engines. It is restricted to ground operations only. The starting, acceleration and operation of the engine is controlled by an integral system of automatic and coordinated pneumatic and electromechanical controls. The APU engine is comprised of three major sections: the accessory section, compressor section and turbine section. Engine power for the auxiliary power unit is developed through compression of ambient air by a single entry, radial, outward-flow, centrifugal compressor. The compressed air, when mixed with fuel and ignited, drives a radial inwardflow turbine rotor. The APU control panel (located above the copilots circuit breaker panel) contains all the primary controls to operate the APU. There is also an APU Relay Panel and APU BITE (Built-In-Test-Equipment) box (primarily for maintenance use), located in the APU compartment, which displays the fault codes associated with the APU. The engine is controlled and serviced by four systems: the engine fuel system, lubrication system, electrical system and indicating system. Fuel for the APU flows from the left wing fuel tank, through the APU boost pump, a shutoff valve and a fuel filter prior to reaching the APU. The APU uses approximately 40 pounds of fuel per hour. Running out of fuel in the left wing fuel tank will introduce air in the APU fuel lines which will cavitate the APU and prevent it from restarting immediately. The APU gearbox serves as an oil sump for the APU self-contained lubrication system. The APU Electronic Sequence Unit (ESU) is a fully automatic system that directs delivery of the correct amount of fuel regardless of ambient conditions and load requirements, as well as properly sequencing control of fuel and ignition during starting. The ESU also monitors engine parameters during operation and automatically shuts down the APU in the event a parameter is not within operational limits. A weight-on-wheels input prevents operation of the APU while airborne.

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APU CONTROL PANEL The APU control panel, located above the copilots circuit breaker panel, houses the necessary controls for operation and monitoring. APU fire detection/extinguishing controls are also located on the APU control panel.
AMPS (350 MAX)

APU
FIRE ON
FAULT STOP

SYSTEM

RUNNING

5
FIRE

10
APU

10
GEN

TEST

MASTER

START

APU CONTROL PANEL Figure 2-11 APU AMPS INDICATOR The AMPS indicator is a digital display indicating the amperage output of the APU generator (shows zero during start). Display will flash when current is at or above 400 amps. APU FIRE This switch/indicator is used to show an APU system fire or overheat (800F at a single point in the fire loop or 375F within overall length of the fire loop) and activate the APU fire extinguishing system. Should there be a fire/overheat in the APU, as detected by the fire loop, the FIRE switch/indicator will indicate FIRE (red), the aircraft Master WARN light will illuminate, and the APU fire warning horn will sound. The fire detection/extinguishing system will automatically shut down the APU by closing the fuel shutoff valve, and activate the fire extinguisher within 20 seconds. Depressing the FIRE switch/indicator will also shut down the APU and discharge the APU fire extinguishing bottle.

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APU FAULT/STOP SWITCH This switch/indicator is a momentary, two cell, lighted switch. The lower portion is labeled STOP (white) and during normal operation this switch is used to shut down the APU by sending an overspeed signal to the Electronic Sequence Unit of the APU. A normal shutdown will not cause the FAULT half of the switch to illuminate. The top portion of this switch is labeled FAULT (amber) and shows a malfunction in the APU system. The APU will automatically shut down if a fault is sensed. The FAULT indicator circuit is latched and is cleared by the FAULT RESET switch on the APU relay box, located near the APU. APU RUNNING/START SWITCH This switch/indicator is a momentary, two cell, lighted switch. Depressing this switch initiates the APU start sequence. The lower portion is labeled START (white) and is illuminated whenever the MASTER Switch is on to identify the switch. The top portion is labeled RUNNING (green) and is illuminated when the APU is running and supplying or ready to supply power to the aircraft. APU MASTER SWITCH The APU MASTER switch is used to power up the APU control circuits from the aircraft normal electrical system. The legend is daylight readable and illuminated white when the aircraft NAV light switch is on. APU ON INDICATOR The APU ON (green) indicator illuminates when the MASTER switch is on. APU SYSTEM TEST SWITCH The APU SYSTEM TEST switch tests the integrity of the APU fire loop/ extinguishing system. Depressing this switch will also test all annunciator lights on the APU control panel, sound the APU fire horn, close the APU fuel shutoff valve and illuminate the aircraft Master WARN/ CAUT lights. Depressing this switch while the APU is running will close the APU fuel shutoff valve and shut down the APU.

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APU RELAY PANEL The APU relay panel is located in the rear equipment bay, next to the APU. The panel contains circuit breakers and relays which interface to the APU control panel and system components for starting and operating the APU. The relay panel also contains two magnetic latching BITE indicators to display generator faults or overheat faults.
NORM K2 10 FUEL J5 GEN FAULT FIRE DET RESET FAULT RESET 15 GEN K4 K5 K6 K7 K9 K11 K14 K16 K15 K17

5
POR

K1

K3

K8

K10

K12

K13

APU RELAY PANEL Figure 2-12 FIRE DET BITE INDICATOR The white FIRE DET indicator shows a fire or overheat condition has been detected. GEN FAULT BITE INDICATOR The white GEN FAULT indicator shows a generator fault has been detected by the ESU. FAULT RESET SWITCH This switch has two positions, NORM and RESET. The switch is spring loaded to remain in the NORM position for normal APU operations. Selecting the RESET position resets the FIRE DET and the GEN FAULT BITE indicators.

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APU BITE ANNUNCIATOR BOX The BITE annunciator box, located in the APU compartment, will display any fault codes (BITE indication) encountered. An indicator activated white shows a malfunction.

PROCESSOR FAIL

OVERSPEED

OVERTEMP NO 1

LOW OIL PRESS

TIME OUT

APU BITE ANNUNCIATOR BOX Figure 2-13

APU GENERATOR Refer to Section IV, ELECTRICAL & LIGHTING, for information on the APU generator.

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APU OPERATING PROCEDURES APU PRE-START CHECK This check should be accomplished in addition to the Preflight Inspection in Section II of the FAA approved Airplane Flight Manual. 1. APU Oil Level Check. 2. Check APU area for indications of oil or fuel leaks. 3. FUEL, GEN, & POR (Point of Regulation) Circuit Breakers (APU Relay Panel) Set. 4. APU Inlet & Exhaust Clear. 5. FIRE, APU, & GEN Circuit Breakers (APU Control Panel) Set. 6. BATTERY 1 & BATTERY 2 Switches On. 7. GPU (if desired) Connect. 8. Verify 18 volts minimum are available for starting the APU. 9. Left Wing Fuel Quantity Check. 10. APU MASTER Switch Press. Verify ON, START, STOP and AMPS indicator all illuminate. 11. APU SYSTEM TEST Switch Press. APU fire horn sounds, APU FIRE warning switch, all APU annunciator lights illuminate and the digital AMPS indicator displays all 8s. APU START-UP To start the APU: 1. BCN/STROBE Switch BCN. 2. APU START Switch Press (momentarily). An automatic start sequence is initiated and the following events will occur: - The APU engine start relay receives starting power from the aircraft batteries or external power. - At 5% RPM the APU fuel shutoff valve opens. - At 65% RPM the starter is de-energized. - At 98% RPM + 20 seconds the green RUNNING annunciator illuminates indicating the APU is ready to provide electrical power. If external ground power is not being used, the APU generator will automatically go on-line and the AMPS indicator will indicate the APU generator load.

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APU SHUTDOWN To shut down the APU: 1. APU STOP Switch Press (momentarily). An automatic shutdown sequence is initiated. Verify that the green RUNNING light goes off. 2. APU MASTER Switch Press. The APU ON annunciator will extinguish. 3. BCN/STROBE Switch Off. 4. BATTERY Switches Off. APU SHUTDOWN FEATURES (Automatic) During APU operation, the ESU monitors engine speed, temperature, oil pressure and electrical surge conditions. The ESU contains circuitry which will automatically send a signal to the APU Relay Panel which in turn will close the fuel shutoff valve and shut down the APU under the following conditions: - Overspeed - Underspeed - Over temperature - Low oil pressure - Loss of EGT signal to the APU ESU - Loss of RPM - High oil temperature - APU fire indication - Low fire bottle pressure - Generator malfunction

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SECTION III HYDRAULICS & LANDING GEAR

TABLE OF CONTENTS Hydraulic System ...................................................................................... 3-1 HYD PUMP Switch ............................................................................... 3-2 HYDR PRESS Lights.............................................................................. 3-2 Hydraulic System Schematic (Figure 3-1) .......................................... 3-3 HYD PRESS Indicator ........................................................................... 3-4 Emergency Air System ............................................................................. 3-4 Emergency Air Pressure Indicator ...................................................... 3-4 Landing Gear System................................................................................ 3-5 Landing Gear Selector Switch.............................................................. 3-5 Landing Gear Extension/Retraction Schematic (Figure 3-2) .......... 3-6 Landing Gear Position Indicators........................................................ 3-8 Landing Gear Warning System............................................................ 3-9 Landing Gear Alternate Extension.................................................... 3-10 Gear Blow Down............................................................................. 3-11 Gear Free Fall................................................................................... 3-12 Nose Wheel Steering System ................................................................. 3-13 STEER ON Light .................................................................................. 3-14 NOSE STEER/ARM Switch ............................................................... 3-14 Control Wheel Master Switch Nose Steering Function............. 3-14 Wheel Brake System................................................................................ 3-15 Parking Brake ....................................................................................... 3-15 PARK BRAKE Light ............................................................................ 3-15 Wheel Brake System Schematic (Figure 3-3).................................... 3-16 Emergency Braking ............................................................................. 3-17 Anti-Skid System ..................................................................................... 3-18 ANTI-SKID Lights ............................................................................... 3-19 ANTI-SKID Switch .............................................................................. 3-19

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SECTION III HYDRAULICS & LANDING GEAR

HYDRAULIC SYSTEM
The aircraft hydraulic system supplies hydraulic pressure for operation of the aircraft landing gear, brake, flap, spoiler and thrust reverser systems. Hydraulic fluid is supplied from the hydraulic reservoir through shutoff valves to the engine-driven hydraulic pumps for distribution to the required systems upon demand. The engine-driven, variable-volume hydraulic pumps will normally maintain system pressure between 1400 and 1550 psi. A pressure relief valve installed between the high-pressure and return lines will open to relieve pressure in excess of 1750 psi. Reservoir pressure is maintained at approximately 20 psi by bleed air supplied through a pressure regulator. Reservoir pressure in excess of 20 psi is relieved overboard by a pressure relief valve and a vacuum relief valve prevents negative pressure in the reservoir. Two precharged (850 psi) hydraulic accumulators are installed to absorb pressure surges. Both accumulator indicators are located under the right engine behind a transparent sight panel. The right-hand accumulator is plumbed for the brakes, landing gear and flaps; the left-hand accumulator is plumbed primarily to power thrust reverser operations but assists the main system accumulator for landing gear, flap and brake operation. Two high-pressure filters and one return filter prevent hydraulic fluid contamination. The return filter incorporates a bypass valve which will open in the event it becomes clogged. Both the highpressure and return filter incorporate an overpressure bypass button. An auxiliary hydraulic pump is installed to provide system pressure in the event of a malfunction or during engine-off ground operations. The thrust reverser hydraulic system incorporates a mechanically controlled isolation valve that will shut off hydraulic fluid to the thrust reverser system if it senses that hydraulic pressure in the main hydraulic system has dropped below approximately 150 psi. This prevents thrust reverser activation in the unlikely event of engine-driven pump failure. A one-way check valve downstream of the thrust reverser system ensures that fluid does not back-up from the main system.

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Two motor-driven firewall shutoff valves can stop hydraulic fluid flow to the engine-driven hydraulic pumps in the event of an emergency or engine fire. Each shutoff valve is operated by the corresponding ENG FIRE PULL T-handle on the glareshield. (Refer to ENGINE FIRE EXTINGUISHING). The valves operate on 28 VDC supplied through the L and R FW SOV circuit breaker on the pilots and copilots circuit breaker panels respectively. Loss of power causes the shutoff valves to remain in their last position. The firewall shutoff valves are operative during EMER BUS mode. The system is serviced through a ground service access located below the right engine pylon. The service access includes quick-disconnect ports for pressure and return lines, an air valve for accumulator charging, and a direct-reading accumulator pressure gage. HYD PUMP SWITCH The auxiliary hydraulic pump is controlled by the HYD PUMP switch located on the center switch panel. When the switch is placed in the On (HYD PUMP) position, the auxiliary hydraulic pump is cycled by a pressure sensing switch plumbed into the high-pressure side of the system. The pressure switch will energize the auxiliary hydraulic pump if system pressure drops below approximately 1000 psi and then de-energize the pump when system pressure rises above approximately 1100 psi. The auxiliary hydraulic pump is plumbed to provide hydraulic pressure for the landing gear, wheel brake, and flap systems only and will not supply pressure for operation of the spoilers or thrust reversers. The auxiliary hydraulic pump operates on 28 VDC supplied through a current limiter and is available when EMER BUS is selected. Refer to Airplane Flight Manual for hydraulic pump limitations. HYDR PRESS LIGHTS Illumination of the amber L and R HYDR PRESS lights on the glareshield annunciator panel indicate low hydraulic system pressure from either the left or right engine-driven pump respectively. The lights are operated by the hydraulic pump pressure switches that sense hydraulic pressure provided by each engine-driven pump. The L or R HYDR PRESS light will illuminate when hydraulic system pressure drops below approximately 150 (50) psi in the engine-driven hydraulic pump line.

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REGULATED BLEED AIR 20 PSI

L
ENG FIRE PULL

R
ENG FIRE PULL

VACUUM RELIEF VALVE

RESERVOIR RELIEF VALVE L HYDR PRESS

HYDRAULIC RESERVOIR

EXTERNAL PRESSURE

R HYDR PRESS PRESSURE FILL VALVE

EXTERNAL RETURN

CHARGE VALVE

SPOILER SYSTEM

ACCUMULATOR ACCUMULATOR
P S I X 1 0 0 0

2 1
0

H Y D P R E S S

HYD PUMP

CHARGE VALVE

OFF ISOLATION VALVE

1750 PSI

FLAP SYSTEM

LANDING GEAR DOOR SELECTOR VALVE

BRAKE SYSTEM

LANDING GEAR SELECTOR VALVE

THRUST REVERSER SYSTEM

AIR CHECK VALVE SUPPLY PRESSURE PRESSURE FILL RETURN

FIREWALL SHUTOFF VALVE

ENGINE DRIVEN PUMP

RELIEF VALVE

FILTER

AUXILIARY PUMP

PRESSURE GAGE FILTER

GROUND SERVICE QUICK DISCONNECT

ISOLATON VALVE PILOT PRESSURE ELECTRICAL

TRANSDUCER RESTRICTOR

HYDRAULIC SYSTEM SCHEMATIC Figure 3-1

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HYD PRESS INDICATOR The HYD PRESS indicator is a vertical-scale instrument and is located on the center switch panel adjacent to the auxiliary hydraulic pump and anti-skid switches. The indicator face consists of a vertical scale marked from 0 to 2000 psi in 500 psi increments and a pointer at the right margin of the instrument. The instrument is operated by a pressure transducer plumbed to the high-pressure side of the hydraulic system in the gear, flap and brake part of the circuit. The indicator operates on 28 VDC supplied through the HYDRAULIC PRESS IND circuit breaker on the copilots circuit breaker panel. Refer to Airplane Flight Manual for instrument limit markings.

EMERGENCY AIR SYSTEM

Two emergency air bottles (3000 psi) are installed to provide alternate gear extension and emergency braking in the event of an electrical or hydraulic system failure. One bottle provides air pressure to operate the emergency gear extension blow down system and the other bottle provides air pressure to operate the emergency brakes and emergency gear extension free fall systems. One emergency air bottle is installed behind the left wing/fuselage fairing, and the other is installed behind the right wing/fuselage fairing. Refer to LANDING GEAR ALTERNATE EXTENSION and EMERGENCY BRAKING for system operation. EMERGENCY AIR PRESSURE INDICATOR The emergency air pressure indicator is a vertical scale, dual-reading instrument and is located on the center switch panel adjacent to the hydraulic pressure indicator. The indicator face consists of a center scale reading from 0 to 4000 psi in 500 psi increments and two pointers on opposite margins of the scale. The left margin is labeled GEAR AIR and the right margin is labeled BRAKE AIR. The indicator pointers are operated by transducers plumbed to the corresponding emergency air bottles. The GEAR AIR pointer indicates the state of charge for the air bottle operating the alternate gear extension blow down system and the BRAKE AIR pointer indicates the state of charge for the air bottle operating the emergency braking and alternate gear extension free fall systems. The indicator operates on 28 VDC supplied through the AIR PRESS IND circuit breaker on the copilots circuit breaker panel. Refer to Airplane Flight Manual for instrument limit markings.

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LANDING GEAR SYSTEM

The landing gear is hydraulically retractable, tricycle gear with airhydraulic shock strut-type nose and main gear. The main gear has dual wheels and brakes on each strut. Each main gear wheel is equipped with two fusible plugs which will melt and release tire pressure in the event wheel temperature reaches 390F. The brake system incorporates four power-boosted disc-type brakes with an integral anti-skid system. The nose gear utilizes a chined tire to prevent splashing into the engine inlet. Nose wheel steering is electrically controlled by the rudder pedals. Hydraulic pressure for gear retraction and extension is transmitted by a system of tubing, hoses, and actuating cylinders, and is electrically controlled by limit switches and solenoid valves. Alternate extension can be accomplished pneumatically in case of hydraulic or electrical system failure. Two doors enclose each main gear after retraction. The inboard doors are hydraulically operated and the outboard doors are mechanically operated by linkage connected to the main gear struts. The nose gear doors operate mechanically with linkage attached to the nose gear shock strut. LANDING GEAR SELECTOR SWITCH The LANDING GEAR switch, located on the center instrument panel, is a lever-lock type switch and must be pulled aft before selecting the UP or DN position. The switch controls the position of the gear selector valve and the door selector valve through gear and door position switches. Electrical power for the control circuits is 28 VDC supplied through the GEAR circuit breaker on the copilots circuit breaker panel. The landing gear control circuits are operative during EMER BUS mode.

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TO BRAKES

EMERGENCY AIR BOTTLE (GEAR FREE FALL AND BRAKES)

CHARGE VALVE EMERGENCY AIR BOTTLE (GEAR BLOW DOWN)

CHARGE VALVE

LEFT DOOR UPLOCK ACTUATORS

RIGHT DOOR UPLOCK ACTUATORS

S
LEFT MAINGEAR INBD DOOR RIGHT MAIN GEAR INBD DOOR

GEAR SELECTOR VALVE

DOOR CONTROL VALVE

UP

DN TO BRAKES

NOSE GEAR UPLOCK ACTUATORS SQUAT SWITCH SQUAT SWITCH

NOSE GEAR DOORS

HYDRAULIC PRESSURE

SHUTTLE VALVE

RELIEF VALVE

HYDRAULIC RETURN ELECTRICAL GEAR DOWN PRESSURE

EMERGENCY AIR CONTROL VALVE

FILTER CHECK VALVE

GEAR UP PRESSURE

SHUNT VALVE

EMERGENCY AIR PRESSURE

LANDING GEAR EXTENSION/RETRACTION SCHEMATIC Figure 3-2 3-6

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Landing gear retraction cycle: When the LANDING GEAR switch is placed in the UP position and the squat switches are in the air mode, the following sequence of events will occur: 1. 28 VDC will be applied to the open solenoid of the door selector valve and hydraulic pressure will be applied to both inboard main gear door uplock actuators and door actuators. 2. When the inboard main gear doors open, door open switches will complete a circuit from the LANDING GEAR switch to the up solenoid of the gear selector valve. Hydraulic pressure will be applied to the main and nose gear actuators and the gear will retract. 3. When the main gear retract, gear up switches will complete a circuit from the LANDING GEAR switch to the close solenoid of the door selector valve. Hydraulic pressure will be applied to the inboard main gear doors actuators to raise the gear doors. Additionally, a gear down safety switch will complete a circuit to the up solenoid of the gear selector valve to maintain continuous hydraulic pressure in the gear actuators. 4. The gear doors are latched by uplatch actuator spring tension. Landing gear extension cycle: When the LANDING GEAR switch is placed in the DN position the following sequence of events will occur: 1. 28 VDC will be applied to the open solenoid of the door selector valve and hydraulic pressure will be applied to both inboard main gear door uplock actuators and door actuators. 2. When the main gear doors open, door open switches will complete a circuit from the LANDING GEAR switch to the down solenoid of the gear selector valve. Hydraulic pressure will be applied to the main and nose gear actuators and the gear will extend. 3. When the main gear are full down, gear down switches will complete a circuit from the LANDING GEAR switch to the close solenoid of the door selector valve. Hydraulic pressure will be applied to the inboard main gear door actuators to raise the gear doors. Additionally, a gear down safety switch will complete a circuit to the down solenoid of the gear selector valve to maintain continuous hydraulic pressure in the gear actuators. 4. The gear doors are latched by uplatch actuator spring tension.

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LANDING GEAR POSITION INDICATORS The landing gear position display, located on the EIS Flight Page, consists of gear indications arranged in a triangular pattern. The indicators are green, red, amber, or white in color. The location of each indicator in the triangular arrangement corresponds to the location of the gear on the aircraft. A DN (green) indication signifies the corresponding gear is down and locked. An unsafe (red rectangle) signifies that the corresponding gear is not in the down and locked position. A door unsafe (white or amber rectangle) displayed along with the DN (green) indication, signifies that the corresponding main gear door is open. During the gear retraction sequence, the unsafe (white rectangle) indicators will display when the sequence is initiated, remain displayed throughout the retraction cycle, and then extinguish when the nose gear is up and locked and the main gear inboard doors close. During the gear extension sequence, the unsafe (white rectangle) indicators will display when the sequence is initiated, remain displayed throughout the extension cycle, and then extinguish when the nose gear is down and locked and the main gear inboard doors close. The indicators are operated by the same switches that control the landing gear extension and retraction cycles. Refer to Airplane Flight Manual for detailed information on the landing gear position indicators. The indicators may be tested with the landing gear retracted by using the GEAR function of the system test switch. When the system test switch is pressed, the landing gear unsafe indicators on the EIS Flight Page will display, the mute light will illuminate on the landing gear switch panel and the landing gear warning horn will sound. If the landing gear is down, only the landing gear warning horn will sound.

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LANDING GEAR WARNING SYSTEM A landing gear warning system is installed to warn the operator of potentially unsafe flight conditions with the landing gear retracted. The system consists of the landing gear warning horn, a thrust lever position switch, and flap position switches. The warning system also uses the landing gear position switches and unsafe indicators. The ADCs (air data computers) provide the airspeed/altitude trip signal. Depending upon the flight condition encountered, one of two distinct warnings will be given as follows: Warning horn sounds and three red gear unsafe indicators display This indicates that the landing gear is not down, airspeed is below approximately 170 KIAS, altitude is below approximately 16,300 feet, and at least one thrust lever is below the 60% N1 position. When the horn sounds under these conditions, the horn can be silenced by depressing the MUTE switch on the LANDING GEAR control panel or depressing the MUTE button in the right thrust lever handle. Whenever the warning horn has been muted, the amber MUTE light on the LANDING GEAR control panel will illuminate. The unsafe indicators will continue to display until either the landing gear is extended or one of the above conditions is corrected. Warning horn only sounds Normally, sounding of the warning horn without a corresponding unsafe indicator being displayed signifies that the landing gear is not down and the flaps are lowered beyond 25. When the horn sounds because the flaps are lowered, the horn cannot be silenced by either mute switch. The horn will continue to sound until either the landing gear is extended or the flaps are retracted.

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LANDING GEAR ALTERNATE EXTENSION In the event of a main hydraulic system failure or an electrical system malfunction, the landing gear can be extended pneumatically. Pneumatic gear extension can be accomplished by using either the alternate gear blow down system or the alternate gear free fall system. However, to ensure adequate emergency air supply for emergency braking (hydraulic system failure) or to ensure hydraulic pressure can be regained (electrical malfunction), it is recommended that blow down be selected first. If an attempt to blow down the gear is unsuccessful, alternate gear free fall should be selected. Air pressure to operate the blow down system is supplied by the GEAR AIR emergency air bottle and is controlled by the EMERGENCY BLOW DOWN GEAR lever on the right side of the pedestal. Air pressure to operate the free fall system is supplied by the BRAKE AIR emergency air bottle and is controlled by the EMERGENCY FREE FALL GEAR lever on the right side of the pedestal forward of the blow down lever. Whenever alternate gear extension is to be selected, the LANDING GEAR selector switch should be placed in the DOWN position and the GEAR circuit breaker on the copilots circuit breaker panel should be pulled. This will prevent inadvertent gear retraction in the event electrical power to the system is regained.

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GEAR BLOW DOWN When the EMERGENCY BLOW DOWN GEAR lever on the right side of the pedestal is pushed full down (until lever latches), air pressure from the GEAR AIR emergency air bottle is admitted to the blow down system through the lever actuated blow down valve. Since the air pressure is greater than the landing gear system hydraulic pressure, shuttle valves in the landing gear system will reposition to admit air pressure to the landing gear system inboard main gear door and door uplock actuators, the main gear actuators, the nose gear uplock and gear actuators, the gear control valve, and the door control valve. The gear and door selector valves are positioned to down to prevent inadvertent gear retraction. When the landing gear is down and locked, the three green DN indicators will display. The two main gear door unsafe indicators will remain displayed after gear extension due to the inboard main gear doors remaining open. When emergency gear blow down is selected, it is not required that the EMERGENCY BLOW DOWN GEAR lever be returned to the up position prior to landing. However, the lever must be returned to the up position prior to servicing either the GEAR AIR bottle or the hydraulic system. The EMERGENCY BLOW DOWN GEAR lever is returned to the up position by lifting the lever release (small metal tab available through a small hole immediately forward of the lever) and pulling the lever to the full up (latched) position.

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GEAR FREE FALL When the EMERGENCY FREE FALL GEAR lever on the right side of the pedestal is pushed full down (until lever latches), air pressure from the BRAKE AIR and free fall emergency air bottle is admitted to the free fall system through the lever actuated free fall valve. The air pressure is directly applied to an uplock actuator for each inboard main gear door, a nose gear uplock actuator, the door selector valve, the gear selector valve, and a hydraulic pressure shunt. The uplock actuators open the gear doors and release the nose gear uplock allowing the gear to free fall. The gear and door selector valves are positioned to down to prevent inadvertent gear retraction. The hydraulic pressure shunt diverts hydraulic system pressure to a hydraulic return line. Full gear extension should occur within 30 seconds with a complete loss of hydraulic pressure. When the landing gear is down and locked, the three green DN indicators will display. The two main gear door unsafe indicators will remain displayed after extension due to the inboard main gear doors remaining open. When emergency gear free fall is selected, the EMERGENCY FREE FALL GEAR lever must be returned to the up position in order to retain BRAKE AIR bottle pressure for emergency braking (hydraulic system failure) or in order to allow the hydraulic shunt to reposition, allowing the hydraulic system to regain pressure (electrical malfunction). The EMERGENCY FREE FALL GEAR lever is returned to the up position by lifting the lever release (small metal tab available through the small hole immediately forward of the lever) and pulling the lever to the full up (latched) position.

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NOSE WHEEL STEERING SYSTEM

The digital nose wheel steering system is a steer by wire system that receives pilot commands through dual rudder pedal position and dual rudder pedal force sensors. The computer processes information from the rudder pedal position and force sensors and three anti-skid wheel speed generators and steering authority is modified as a function of aircraft ground speed. For low speed ground operations 60 of steering authority either side of neutral is available. At low speed and large rudder pedal deflection the nose wheel displacement will be large for high maneuverability. Once a rudder pedal has reached its stop, further nose wheel displacement is generated by additional force being applied to that rudder pedal. As ground speed increases, the maximum wheel deflection is reduced to zero. At 90 knots 28 VDC is removed and the system disengages. Above 90 knots the nose wheel is allowed to castor. Nose wheel steering engage circuits are controlled through the momentary-action pedestal-mounted NOSE STEER/ARM switch and the Control Wheel Master Switches (MSW). When the squat switches are in the ground mode, depressing and releasing the NOSE STEER/ARM switch will activate the computer when AC and DC power are available, the nose gear is down and locked, and no faults are detected by the system monitor. When the system is active the STEER ON annunciator on the glareshield and the ARM annunciator on the NOSE STEER/ ARM switch will illuminate. At 90 knots, when the system disengages, the glareshield STEER ON annunciator will extinguish. When the nose gear is no longer in the down and locked position, the ARM annunciator on the NOSE STEER/ARM switch will extinguish, however; the computer is still powered and system monitor circuitry remains active. When the nose gear is down and locked for landing the ARM annunciator on the NOSE STEER/ARM switch will illuminate provided no faults have been detected. After touchdown, when ground speed decreases to 90 knots, the STEER ON light on the glareshield will illuminate and steering authority will increase as ground speed decreases. If the system cannot be armed, limited authority steering (24 either side of neutral) is available by depressing and holding either MSW. It should be noted that in some instances, even though a fault has been detected, the system will continue to function normally until shutdown. After that, however; it will not be possible to operate the system with full steering authority until the fault has been corrected. If the system cannot be accessed by either MSW, sufficient control is still available by differential braking.

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The nose wheel steering system is powered by 28 VDC supplied through the NOSE STEER circuit breaker and 115 VAC supplied through the NOSE STEER circuit breaker in the TRIM-FLT CONT group on the copilots circuit breaker panel. STEER ON LIGHT The green STEER ON light on the glareshield annunciator panel illuminates to indicate the nose wheel steering system is capable of responding to rudder pedal inputs. NOSE STEER/ARM SWITCH Normally, the NOSE STEER switch is used to activate nose steering circuits for taxi operations. Momentarily depressing the NOSE STEER switch will activate the system and the ARM annunciator will illuminate. When nose steering has been activated, the system can be disengaged by depressing then releasing either the pilots or copilots Control Wheel Master Switch (MSW) or by depressing the NOSE STEER switch a second time. The disconnect tone will sound. CONTROL WHEEL MASTER SWITCH NOSE STEERING FUNCTION Depressing and holding either Control Wheel Master Switch (MSW) will engage the nose wheel steering system. While the MSW is held, the nose steering system will operate normally and the STEER ON annunciator will be illuminated. When the MSW is released, the nose wheel steering system will disconnect. The STEER ON annunciator will extinguish. In the event that nose wheel steering will not arm, the MSW can be depressed and held for limited authority steering, under some fault conditions.

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WHEEL BRAKE SYSTEM

The primary brake system utilizes hydraulic system pressure for power boost. Hydraulic pressure from the nose gear down line is metered to the disc-type wheel brakes by the power brake valves. The valves are controlled by the rudder pedal toe brakes through mechanical linkage. Two shuttle valves in the pressure lines prevent fluid feedback between the pilots and copilots pedals. Four additional shuttle valves connect the pneumatic system to the brake system for emergency braking. Hydraulic fuses, located in the main gear wheel wells, will close to prevent pressure loss if fluid flow exceeds normal brake actuation rate. Snubbing of the main gear wheels is accomplished during retraction by means of hydraulic back pressure in the brake lines caused by a restrictor in the return line. An integral anti-skid system is installed to effect maximum braking efficiency. When parking, it is advisable to have the wheels chocked prior to releasing brakes. PARKING BRAKE The parking brake handle is labeled PARKING BRAKE and is located on the pedestal below the thrust levers. The handle is mechanically connected to the parking brake valve through which all pressure from the primary brake system must pass. The parking brake system is actuated by pressing and holding the toe brakes (hydraulic system pressurized) then pulling the parking brake handle which closes the parking brake valve, locking pressure against the wheel brakes. Pulling the parking brake handle also closes the solenoid shutoff valve on the antiskid system to prevent leakage through the anti-skid valve. Returning the parking brake handle to the off position releases the brakes. The anti-skid system is inoperative when the parking brake is engaged. PARK BRAKE LIGHT An amber PARK BRAKE light, on the pilots subpanel, immediately above the ANTI-SKID lights, is installed to alert the operator that the parking brake may be engaged. The light is operated by a switch attached to the parking brake valve and will be illuminated whenever power is on the aircraft and the PARKING BRAKE handle is not full in.

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20-27B

WHEEL BRAKE SYSTEM SCHEMATIC Figure 3-3

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EMERGENCY BRAKING In the event of a main hydraulic system failure, the wheel brakes can be applied pneumatically. Emergency (pneumatic) braking is initiated and controlled through the red EMER BRAKE handle located on the pedestal to the left of the thrust levers. Emergency braking is initiated by pulling the handle out of the recess and pushing down. As the EMER BRAKE handle is pushed down, air pressure from the BRAKE AIR emergency air bottle is directed to the wheel brake shuttle valves through the lever actuated emergency brake valve. If the emergency air pressure is greater than the brake system pressure, the wheel brake shuttle valves will reposition to admit air pressure to apply the brakes. As the brake handle is released, excess air will be vented overboard and the brakes will release. Because the emergency air lines are plumbed into the hydraulic brake system between the anti-skid control valves and the wheel brakes, anti-skid protection is not available when using emergency brakes. Also, the parking brake will be inoperative when using emergency air pressure.

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ANTI-SKID SYSTEM
An anti-skid system is integrated into the hydraulic brake system to provide maximum braking efficiency under all runway surface conditions without skidding the tires. The system consists of the ANTI-SKID control switch, anti-skid control box, two anti-skid control valves, monitoring lights, four wheel-speed transducers (one in each main wheel axle), and associated aircraft wiring. Each anti-skid control valve is a dual unit capable of individually modulating brake pressure for both associated brakes. As the transducers are driven by the main wheels, a frequency proportional to the wheel speed is induced and forwarded to the control box. The control box converts the wheel-speed frequency to an analog signal and compares the analog to a reference representing the normal deceleration limits. Should the wheel speed deviate from the normal deceleration limits, the control box will signal the affected wheels control valve to reduce braking pressure on the affected wheel. Braking pressure is reduced by bypassing some of the hydraulic system pressure into a return line by means of a servo controlled valve in the control valve. As the wheel speed increases, normal braking pressure is restored. To ensure full manual control of the hydraulic braking system and to prevent pressure loss when the parking brake is set, a solenoidoperated shutoff valve at each control valve return port is de-energized closed when the ANTI-SKID switch is OFF or the parking brake is set. Electrical power for the anti-skid system control circuits is 28 VDC supplied through the ANTI-SKID circuit breaker in the hydraulics group on the copilots circuit breaker panel.

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ANTI-SKID LIGHTS Four amber ANTI-SKID lights on the pilots subpanel provide a continuous cockpit indication of the anti-skid system control circuits. The two lights labeled L represent control circuits for the left main gear brakes and the two lights labeled R represent control circuits for the right main gear brakes. The anti-skid control box continuously monitors the system circuits and will illuminate the applicable light(s) should any of the following conditions arise: loss of input power, open and short transducer circuits, open or short control valve circuits, and failure of control box circuits. Also, the lights will be illuminated any time the gear is down and locked, power is on the aircraft, and the ANTI-SKID switch is off. ANTI-SKID SWITCH The ANTI-SKID switch is located on the center switch panel and has two positions: On (ANTI-SKID) and OFF. When the switch is in the On (ANTI-SKID) position, 28 VDC is applied to the anti-skid system control circuits. Normally, the switch remains in the On (ANTI-SKID) position for all operations.

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SECTION IV ELECTRICAL & LIGHTING

TABLE OF CONTENTS DC Power Distribution ............................................................................. 4-1 BATTERY Switches................................................................................ 4-3 START/GEN Switches .......................................................................... 4-3 Start Lights.............................................................................................. 4-4 GEN RESET Switches............................................................................ 4-4 GEN Lights ............................................................................................. 4-4 DC Generation and Start (Figure 4-1) ................................................. 4-5 DC Power Distribution (Figure 4-2).................................................... 4-7 CUR LIM Light....................................................................................... 4-9 DC Circuit Breakers ............................................................................... 4-9 External Power Receptacle ................................................................. 4-10 AC Power and Distribution (Figure 4-3).............................................. 4-11 AC Power Distribution ........................................................................... 4-13 INVERTER Switches............................................................................ 4-13 AC Circuit Breakers............................................................................. 4-13 Electrical Page Display ........................................................................... 4-14 EIS Electrical Page (Figure 4-4).............................................................. 4-14 Automatic Load Shedding System ....................................................... 4-15 Emergency Bus System........................................................................... 4-15 Cabin Power Control Switch.............................................................. 4-15 EMER BUS Switch ............................................................................... 4-16 Emergency Bus System (Figure 4-5).................................................. 4-17 Avionics Power System .......................................................................... 4-19 Avionics Master Switch....................................................................... 4-19 Auxiliary Power Unit (APU) Generator .............................................. 4-19 Emergency Power System...................................................................... 4-20 EMER BAT Switch ............................................................................... 4-20 Exterior Lighting ..................................................................................... 4-21 Landing/Taxi Lights ........................................................................... 4-21 Navigation Lights ................................................................................ 4-21 Tail Logo Lights.................................................................................... 4-22 Anti-Collision (Beacon/Strobe) Lights ............................................. 4-22 Recognition Light................................................................................. 4-23 Wing Inspection Light......................................................................... 4-23 Exterior Convenience Lights ............................................................. 4-23

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TABLE OF CONTENTS (Cont) Cockpit Lighting ..................................................................................... 4-24 Instrument Panel Floodlights ............................................................ 4-24 Instrument Lights ................................................................................ 4-24 Two-Stage Lighting ............................................................................. 4-25 Switch Panel Lighting......................................................................... 4-26 Adaptive Flight Display Lighting ..................................................... 4-26 Map Reading Lights............................................................................ 4-27 Lighted Chart Holders........................................................................ 4-27 Dome Lights ......................................................................................... 4-27 Passenger Compartment Lighting........................................................ 4-28 Aisle Lights........................................................................................... 4-28 Passenger Reading Lights .................................................................. 4-28 Overhead Lights .................................................................................. 4-28 Entry Door Switch Panel (Figure 4-6)............................................... 4-29 Cabin Control Switch Panel (Figure 4-7).......................................... 4-29 Passenger Control Switch Panel (Figure 4-8) .................................. 4-29 Lavatory Switch Panel (Figure 4-9)................................................... 4-30 Entry Light ........................................................................................... 4-30 Lavatory Lights.................................................................................... 4-30 Baggage Compartment Light ............................................................ 4-31 No Smoking and Fasten Seat Belt Signs........................................... 4-31 Cargo and Servicing Compartment Lighting ..................................... 4-32 Tailcone Baggage Lights ..................................................................... 4-32 Tailcone Maintenance Light ............................................................... 4-32 Illuminated Exit Sign System ................................................................ 4-33 Emergency Exit Lights Battery Units ............................................... 4-33 Egress Light Assemblies..................................................................... 4-33 Emergency Exit Lighting.................................................................... 4-33 EMERGENCY EXIT LIGHTS Control Panel ................................... 4-34 EMERGENCY EXIT LIGHTS Control Panel (Figure 4-10)....... 4-34 Control Switch ................................................................................ 4-34 TEST Switch .................................................................................... 4-35 Annunciators................................................................................... 4-35 Master Caution/Warning and Annunciator Panel Lights ................ 4-36

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SECTION IV ELECTRICAL & LIGHTING

DC POWER DISTRIBUTION
Primary electrical power for aircraft and avionics systems requiring DC power is supplied by two engine-driven, 30-volt, 400-ampere starter/ generators. Secondary DC electrical power is supplied by two 24-volt Concorde Lead Acid batteries. An external power receptacle is installed for engine start and stationary ground operations. A generator control unit (GCU) is installed for each starter/generator. The GCUs contain circuits to maintain generator output at approximately 28 VDC throughout varying engine speeds and loads. The GCUs also contain circuits to equalize generator load during parallel operation, provide overvoltage protection, and provide current limiting during ground operations and during generator-assisted cross starts. During normal operation, the generators supply all aircraft DC power requirements. Regulated 28 VDC output from the generators is applied to the respective generator buses. The voltage on the generator buses is applied to the battery charging bus through 275-amp current limiters. Battery charge is maintained from the battery charging bus through the battery relays and battery buses. The DC BUS 2 and 3 buses in the circuit breaker panels are powered from the respective generator buses through 50-amp current limiters. The DC BUS 4 buses in the circuit breaker panels are powered from the battery charging bus through 40amp current limiters. The battery bus in the pilots circuit breaker panel is powered from the #1 battery through a 20-amp current limiter. The battery bus in the copilots circuit breaker panel is powered from the #2 battery through a 10-amp current limiter. The DC BUS 1 buses in the circuit breaker panels are powered from the respective generator bus through an overload sensor and a control relay. A CABIN PWR BUS is installed in the pilots circuit breaker panel. The CABIN PWR BUS is powered from the battery charging bus through a 100-amp current limiter, an overload sensor, and a control relay. The inverters are powered through overload sensors and control relays. Additionally, aircraft systems producing heavy loads; such as resistance heaters, freon compressor, large lamps, inverters, blowers, heavy-duty motors, and heavyduty pumps, are supplied power through current limiters connected to either the battery charging bus or generator buses.
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Overload sensors are installed between the DC BUS 1 buses and the associated generator bus. The overload sensors are installed to protect the DC BUS 1 feeder circuits from an overload. Basically, each overload sensor is a 70-amp circuit breaker mechanically connected to a switch. Should an overload condition occur, the circuit breaker will reposition the switch to de-energize a power relay, thereby disconnecting the DC BUS 1 bus. Additionally, the switch will apply a ground to trip the affected L or R DC BUS 1 circuit breaker. When the overload sensor circuit breaker cools, the switch will reset; however, the power relay will not re-energize due to the open L or R DC BUS 1 circuit breaker. When the malfunction has been corrected and the affected L or R DC BUS 1 circuit breaker reset, the power relay will re-energize and power to the DC BUS 1 bus will be restored. An overload sensor is installed between the CABIN PWR bus and the battery charging bus. The overload sensor is installed to protect the CABIN PWR BUS feeder circuit from an overload. Operation of the CABIN PWR BUS overload sensor is the same as that described for the DC BUS 1 overload sensors. The generators will not come on-line if an operating ground power unit is connected to the aircraft. A cross start relay box is installed which enables an operating generator to assist in providing power to start the opposite engine. If one generator is on-line and a start of the opposite engine is initiated, the cross start relay circuits will cause both left and right starter relays to close. In effect, this will bypass both battery charging bus 275-amp current limiters and the output of the operating generator will supplement the aircraft batteries in providing power for the starter. An airstart relay box is installed which prevents the primary flight displays from blanking and ensures certain equipment, necessary for a successful start, has adequate voltage during airstarts. During an airstart, the #2 battery is isolated from the battery charging bus and its power is dedicated to the following loads: L & R STBY-SCAV PUMP L & R ENG CH A (FADEC) L & R ENG CH B (FADEC) L & R START MFD 1 & 2 DCP 1 & 2 L & R JET PUMP-XFR VALVE L & R IGN CH A L & R IGN CH B AHS 1 & 2 PFD 1 & 2

When the aircraft is on the ground, operation of the airstart circuits is inhibited and both batteries will be available to power the starter.

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An emergency bus system is installed to operate selected equipment from the aircraft batteries for the maximum duration in the event of a dual generator failure. When the emergency buses are selected, the battery charging bus is isolated from the batteries and the equipment connected to the emergency buses will be powered from the aircraft batteries. BATTERY SWITCHES The aircraft batteries are controlled through the BATTERY 1 and 2 switches on the pilots switch panel. The #1 battery is wired directly to the battery bus in the pilots circuit breaker panel and the #2 battery is wired directly to the battery bus in the copilots circuit breaker panel. When either BATTERY switch is placed in the On position, the corresponding battery relay closes to connect the respective battery bus to the battery charging bus if the EMER BUS switch is in the NORMAL position. When the BATTERY switch is placed in the OFF position, the battery relay is de-energized and the respective battery bus is isolated from the battery charging bus. The battery relays will also be de-energized whenever the EMER BUS switch is in the EMER BUS position. START/GEN SWITCHES The starter/generators are controlled through the START/L GEN and START/R GEN switches on the pilots switch panel. Additionally, the START position of each switch is used to control various functions required for the starting sequence. These functions are described below. Each switch has three positions: START, OFF, and GEN. Prior to initiating the starting sequence, the associated thrust lever should be placed in the IDLE detent. START position: With the BATTERY switches On, DC power from the L and R START circuit breakers is applied to the left and right START/ GEN switches. When a START/GEN switch is set to START, DC power from the corresponding START circuit breaker is applied to close the corresponding starter relay, activate the corresponding standby pump, cause the corresponding motive flow valve to close, shutdown the cooling, auxiliary heating, and stabilizer heat systems, and energize the FADEC start sequence relay (supplies a discrete start signal to the FADEC). When the starter relay closes, the starter will begin to spool the engine and the START light will illuminate. When N2 reaches approximately 6%, the FADEC automatically activates the ignition system and turns on fuel flow to the engine. When N2 reaches approximately 40%, the ignition will automatically terminate. When N2 reaches approximately 45%, a speed sensor in the starter/generator will cause
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power to be removed from the starter relay (starter will be de-energized and the START light will extinguish) and from the FADEC start sequence relay (discrete start signal to FADEC will be removed and the corresponding motive flow valve will open). When the switch is moved out of the START position, the corresponding standby pump will shut down. If the associated thrust lever is not in the IDLE detent, ignition and fuel flow will not occur as stated above. GEN position: During the engine start sequence, when engine RPM reaches idle speed, the START/GEN switch should be set to GEN. When GEN is selected, the corresponding generation circuits will be activated. The generator will not come on-line with a GPU connected. Additionally, the cooling and auxiliary heating systems, and stabilizer heat system cutout relays will be reset. The generation circuits activate and control the corresponding generator through the generator control unit. START LIGHTS Amber lights adjacent to each START/GEN switch are installed to indicate starter operation. The corresponding light will be illuminated whenever the associated starter is energized. GEN RESET SWITCHES The GEN RESET buttons are located on the pilots switch panel adjacent to the START/GEN switches. Should a generator fault occur, the corresponding generator control unit will de-energize the affected generator field circuit and open the generator relay isolating the generator from the respective generator bus. Momentarily depressing the applicable GEN RESET button will reset the generator by closing the affected generator field circuit and closing the generator relay. The GEN RESET buttons have no effect with the corresponding START/GEN switch OFF or the corresponding START and/or GEN circuit breaker open. GEN LIGHTS Amber L GEN and R GEN annunciator lights are installed in the glareshield annunciator panel. The lights are controlled by the corresponding generator control circuits and will illuminate whenever the corresponding generator has failed or is off line. The light will also illuminate whenever the corresponding START/GEN switch is in either START or OFF and at least one BATTERY switch is On.

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L GEN NORM L GEN OFF EMER BUS RESET L START L STANDBY PUMP NORMAL OFF BATTERY 1 #1 BATTERY RELAY L JET PUMP Close L GEN LINE RELAY L GEN BUS CROSS START RELAY BOX #1 BATTERY L FADEC (start sequence) L START L STBY PUMP L MOTIVE FLOW VALVE L START L GENERATOR CONTROL UNIT SPEED SENSOR GEN INTERPOLE FIELD RESET GEN BUS 28V GEN BUS SENSE START IN LINE RELAY GEN START OUT EQUALIZER BUS STARTER/GEN
VAC

00.0
CURRENT SENSOR

VDC AMPS

000

000

#1 BAT BUS

1
EXTERNAL POWER RECEPTACLE

GPU/GEN LOCKOUT

L START RELAY

VAC

00.0 000
CUR LIM

VDC AMPS

000

APU SYSTEM (if installed) (certified for ground operation only)

EXT POWER RELAY

28 VDC OUT

R FADEC (start sequence)

R START RELAY

APU Control Inputs

R GEN BUS

EXTERNAL POWER OVER-VOLTAGE CUTOUT CIRCUIT R JET PUMP #2 BAT BUS #2 BATTERY RELAY #2 BATTERY BATTERY 2 OFF R STANDBY PUMP R START OFF R GEN R START R GEN RESET R STBY PUMP Close R GENERATOR CONTROL UNIT R START R GEN LINE RELAY

AMPS(350 MAX)

APU

FIRE

5
FIRE

10
APU

10
GEN

SYSTEM TEST

MASTER

R MOTIVE FLOW VALVE

EQUALIZER BUS START OUT GEN LINE RELAY START IN GEN BUS SENSE GEN BUS 28V RESET FIELD GEN INTERPOLE SPEED SENSOR

CURRENT SENSOR

VAC

00.0
STARTER/GEN

VDC AMPS

000

000

Ground when either generator is on-line.

NORM

NOTE: Airstart circuit not shown.

DC GENERATION AND START Figure 4-1

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PILOTS CB PANEL
1 L STARTER/GENERATOR

L GEN RELAY POWER RELAY L START RELAY

OVERLOAD SENSOR L DC EMER BUS

L DC BUS 1 L DC BUS 2 L DC BUS 3 L DC BUS 4 L DC BUS 2

L DC BUS 1

L GEN BUS

L DC BUS 3 L BATTERY BUS L DC BUS 4 CABIN PWR BUS CABIN PWR BUS

#1 BATTERY

#1 BATTERY RELAY #2 BATTERY RELAY

BATTERY CHARGING BUS

OVERLOAD SENSOR POWER RELAY

#2 BATTERY

R DC BUS 4 R DC BUS 4

R BATTERY BUS

DC BUS 3 TIE

DC BUS 2 TIE

DC BUS 1 TIE

R GEN BUS

R DC BUS 3 R DC BUS 3 R DC BUS 2 R DC BUS 2 R DC BUS 1 R DC BUS 1 R DC EMER BUS

R START RELAY

POWER RELAY

EMER BUS TIE

R GEN RELAY R STARTER/GENERATOR

OVERLOAD SENSOR

COPILOTS CB PANEL

Controlled by EMER BUS Switch. See figure 4-5 for schematic of EMER BUS system.

DC POWER DISTRIBUTION Figure 4-2 4-7/4-8 (Blank) 4-7

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CUR LIM LIGHT The amber CUR LIM annunciator light, on the glareshield annunciator panel, is installed to indicate the continuity of the 275-amp current limiters. The 275-amp current limiters connect the battery charging bus to the generator buses. Failure of both 275-amp current limiters will cause the equipment connected to the battery charging bus to be powered from the aircrafts batteries only. The light is illuminated by sensors wired across the current limiter terminals. A failure of either current limiter will cause the respective sensor to illuminate the CUR LIM light. DC CIRCUIT BREAKERS The aircraft DC electrical circuits are protected by push-to-reset, thermal-type circuit breakers. Most DC circuit breakers are located on the pilots and copilots circuit breaker panels. The L and R DC BUS 1, DC BUS 2, and DC BUS 3 buses may be interconnected through the DC BUS 1 TIE, DC BUS 2 TIE, and DC BUS 3 TIE circuit breaker/switches on the copilots circuit breaker panel. Normally the L and R DC buses are not tied together. If it is desired to tie a L DC BUS and R DC BUS together, the appropriate DC BUS TIE circuit breaker/switch must be in the up (closed) position. The DC BUS 1 circuit breaker on each circuit breaker panel controls power to the associated DC BUS 1 bus through control relays. Circuit breakers are grouped together into system types (e.g. ELECTRICAL, LIGHTS, AVIONICS). Power to operate the emergency bus system is supplied from the batteries through the respective EMER BUS CONT circuit breaker (see figure 4-5). The circuit breakers for equipment powered during EMER BUS mode are denoted by red rings on the overlay.

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EXTERNAL POWER RECEPTACLE External power may be connected to the aircraft DC electrical distribution system through a standard receptacle located on the right fuselage below the pylon. To start an engine or operate aircraft systems using external power at least one BATTERY switch must be in the On position; however, the generators will not come on-line with an external power source connected. External power over-voltage protection circuits will open the external power relay and disconnect external power from the aircraft DC distribution system in the event the external power source exceeds approximately 32 volts. External power source amperage must be limited to a maximum of 1500 amps as specified on the placard above the external power receptacle.

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000

VAC VDC AMPS

000

L INVERTER POWER RELAY

FAULT SIGNAL 115 VAC

L AC BUS

L INV

L INV

BATTERY CHARGING BUS

WSHLD DEFOG OFF R INV

R INV

115 VAC for WS DEFOG

WINDSHIELD SWITCHING LOGIC Both Inverters ON Left inverter powers left WS Right inverter powers right WS Only One Inverter ON Operating inverter powers both WS

115 VAC for WS DEFOG

L GEN BUS

PILOTS CB PANEL

AC BUS TIE

R GEN BUS

PHASE LOCK OFF POWER RELAY R AC BUS

R INVERTER

115 VAC FAULT SIGNAL

COPILOTS CB PANEL

R 115 VAC BUS

L 115 VAC BUS

000

VAC VDC AMPS

000

AC POWER AND DISTRIBUTION Figure 4-3

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AC POWER DISTRIBUTION
Electrical power for aircraft and avionics systems requiring AC power is supplied through two 115-volt, 400-Hz, 1500VA, solid-state inverters. During normal operation, the left and right inverter output voltages are applied to the left and right AC buses respectively. The left and right AC buses may be interconnected through the AC BUS TIE circuit breaker/switch on the copilots circuit breaker panel. Each AC bus is intended to be powered by only one inverter. Therefore, the AC BUS TIE switch should only be closed after removing power from one of the buses and setting the respective INVERTER switch to OFF. If both INVERTER switches are On, a relay in the copilots circuit breaker panel will prevent the AC BUS TIE from functioning (electrically). An inverter relay box controls 28 VDC input to the inverters and provides isolation between the inverter output and AC bus should an inverter fault occur. A phase lock function within the right inverter keeps the output of each inverter in-phase. Input power to operate the left and right inverters is 28 VDC supplied through 100-amp current limiters connected to the left and right generator buses respectively. INVERTER SWITCHES Operation of the left and right inverters is controlled through the two INVERTER switches on the pilots switch panel. The switch controlling the left inverter is labeled L-OFF and the switch controlling the right inverter is labeled R-OFF. When either switch is moved to the On (L or R) position, the associated power relay is energized to supply input power to the associated inverter. When one switch is On and the other is OFF, a relay in the inverter relay box is energized isolating the inoperative inverter from its associated AC bus. The inverter control circuits operate on 28 VDC supplied through the L INV and R INV circuit breakers on the pilots and copilots circuit breaker panels respectively. AC CIRCUIT BREAKERS The aircraft AC electrical circuits are protected by push-to-reset magnetic-type circuit breakers. AC circuit breakers are denoted by a white ring on the panel overlay. The copilots circuit breaker panel also contains the AC BUS TIE circuit breaker/switch which is used to tie the L AC BUS and R AC BUS together in the abnormal situation of single inverter operation. Circuit breakers are grouped together into system types (e.g. ELECTRICAL, AFCS, AVIONICS).

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ELECTRICAL PAGE DISPLAY

The EIS Electrical Page is used to monitor left and right AC bus voltage, left and right DC generator load and the DC charging bus voltage. Digital displays are used for voltage and amperage readouts. Each parameter being monitored is divided into Normal, Caution and Warning ranges. Whenever any parameter goes from the normal range to the caution range, the digital readout will display in amber and flash for five seconds. If the parameter progresses into the warning range, the digital readout will display in red and flash for five seconds. The amber or red digital readout will remain until the affected parameter returns to the normal range. Caution and warning annunciations are inhibited during starter engagement. An amber boxed C located adjacent to the VAC display indicates that the inverter is out of phase.

EIS ELECTRICAL PAGE Figure 4-4 Voltage and amperage parameters are shown in the following table: NORMAL
AC Voltage 110 <= VAC <= 130

CAUTION
90 <= VAC <= 109 OR 131 <= VAC <= 134 18.0 <= VDC < 22.0 OR 29.5 < VDC <= 31.5 330 <= DCA <= 400 N/A 325 <= DCA <= 400

WARNING
VAC < 90 OR VAC > 134 VDC < 18.0 OR VDC > 31.5 DCA > 400 DCA > 400 DCA > 400

DC Voltage DC Amperage On The Ground Up To 31,000 Feet From 31,001 Feet To 46,000 Feet From 46,001 Feet To 51,000 Feet Or Loss Of Air Data Information High Capacity Generator

22.0 <= VDC <= 29.5

DCA <= 325 DCA <= 400 DCA <= 325

DCA <= 300

300 <= DCA <= 400

DCA > 400

DCA <= 400

N/A

DCA > 400

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AUTOMATIC LOAD SHEDDING SYSTEM

An automatic electrical load-shedding system is installed to automatically reduce generator loading in the event of a single generator failure. The system is only active during flight (weight not on wheels). Should either L or R GEN light illuminate in flight, the following loads will automatically shut down to reduce the load on the operating generator: CABIN PWR BUS Loads Air Conditioning System Cockpit Floorboard Heater System Baggage Compartment Heater System

If the generator is brought back on-line, these loads will be regained.

EMERGENCY BUS SYSTEM

An emergency bus system is installed to provide 28 VDC to selected systems in the event of a dual generator system failure or to quickly deenergize and isolate all nonessential equipment in the event of electrical smoke or fire. The system uses the aircrafts batteries to supply DC power to the DC equipment on the emergency bus. All emergency bus circuit breakers are denoted by a red ring on the panel overlay. The EMER BUS TIE is located on the copilots circuit breaker panel. The emergency bus system control circuits operate on 28 VDC supplied by the batteries through the EMER BUS CONT circuit breakers in the pilots and copilots circuit breaker panel. CABIN POWER CONTROL SWITCH The cabin power control switch system adds a CABIN PWR OFF switch inline with the CABIN PWR BUS circuit breaker. This allows the pilot to quickly and efficiently load shed all cabin power systems by selecting the CABIN PWR switch to the OFF position. When the cabin power switch is selected off it will disable all of the cabin entertainment equipment, ordinance signs and standard cabin lighting. Cabin Downwash Lighting will still be available and if not already on can be selected ON from the Master Control Switch Panel or the Cabin Control Switch Panel located in the LH FWD closet. Also, selecting CABIN PWR OFF is one means of reducing generator loads when required by abnormal procedures in the FAA Approved Airplane Flight Manual. During single-generator operation, the aircraft load shed will automatically cause the CABIN PWR to go to the OFF mode.

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EMER BUS SWITCH The EMER BUS switch on the pilots switch panel is used to select the power source for the emergency buses. The switch has two positions EMER BUS and NORMAL. When the EMER BUS switch is in the NORMAL position, the emergency bus system relays will be de-energized and equipment on the emergency buses will be powered from the normal electrical system. DC equipment on the emergency buses will be powered through the associated DC BUS 1, 2, or 3. When the switch is in the EMER BUS position, the battery relays will be de-energized, the emergency bus system relays will be energized, and equipment on the emergency buses will be powered through the emergency bus system. When the battery relays are de-energized, the aircraft batteries are completely isolated from the battery charging bus and the normal DC power distribution system. When EMER BUS is selected, electrical power will be distributed as follows: 1. DC power for the primary pitch trim motor will be switched from the battery charging bus to the #1 aircraft battery. 2. DC power for the auxiliary hydraulic pump will be switched from the battery charging bus to the #2 aircraft battery. 3. DC power to heat the standby pitot-static probe will be switched from the battery charging bus to the #2 aircraft battery. 4. DC powered equipment on the emergency buses will be switched from the associated DC BUS 1 to the aircraft batteries. 5. The DC voltmeter will display the voltage of both batteries (EMER BUS TIE must be closed).
NOTE

The conditions just described assume that both BATTERY switches are in the On position. If only the BATTERY 1 switch is On, the auxiliary hydraulic pump will not be available, heat for the standby pitot-static probe will not be available, and the DC voltmeter will display the voltage of the #1 battery. All other conditions will be as described. If only the BATTERY 2 switch is On, Primary Pitch Trim will not be available and the DC voltmeter will display the voltage of the #2 battery. All other conditions will be as described.

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PILOTS CB PANEL

ELEC PWR MON L DC BUS 1 L IGN CH A L DC BUS 2 #1 BATTERY RELAY L DC BUS 3 L IGN CH B L EMER BUS CONT FUEL QTY PWR 1 FUS TANK XFR PUMP L JET PUMP-XFR VALVE XFLO VALVE L EMER DC BUS L FIRE DETECT L FIRE EXT L FW SOV L ENG CH A FLOOD LTS CENTER PANEL-PED LTS WARN LTS PRI PITCH TRIM L STALL WARN BATTERY CHARGING BUS EMER BUS

ATC 1 MFD 1 DCU 1 ADF 1 GPS 1 AUDIO 1 COMM 1 FMS DISPLAY 1 FSU 1 ADC 1 FDR MFD CONTROL 1 DISPLAY CONTROL 1

2
#1 BATTERY

1 1 2

WHEEL MASTER BLEED AIR OV HT CABIN PRESS IND AHS 1

NORMAL

STAB ACT AUX HYD PUMP

RTU 1 NAV 1

DC VOLTMETER STANDBY PITOT HEAT

R EMER BUS CONT EMER BUS TIE

COPILOTS CB PANEL

1
R DC BUS 1

1
R DC BUS 2 R DC BUS 3 R IGN CH B R IGN CH A #2 BATTERY RELAY

CVR FUEL QTY PWR 2 FUS TANK AUX PUMP R JET PUMP-XFR VALVE R FIRE DETECT R FIRE EXT R FW SOV ELT NAV DCU 2 ADC 2

R EMER DC BUS

#2 BATTERY

R ENG CH A WARN LTS SEC PITCH TRIM FLAPS TRIM-FLAP-SPOILER INDICATOR R STALL WARN SPOILER GEAR CABIN PRESS SYS AHS 2 AUDIO 2 CABIN FIRE DETECT PASS SPKR

1 2 3

Ground to activate relays supplied through EMER BUS position of EMER BUS switch and "On" position of BATTERY switches. Ground to activate relays supplied through EMER BUS position of EMER BUS switch. With EMER BUS switch in EMER BUS position and: a. Both BATTERY switches "On" Voltmeter will display voltage of both batteries (EMER BUS TIE must be closed). b. Only one BATTERY switch "On" Voltmeter will display voltage of the battery whose respective BATTERY switch is "On".

EMERGENCY BUS SYSTEM Figure 4-5 4-17/4-18 (Blank) 4-17

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AVIONICS POWER SYSTEM

An avionics power system is installed to allow selected DC powered avionics systems to be powered up through the use of two master switches. The system consists of a LEFT MASTER and RIGHT MASTER switch, and a control relay in each circuit breaker panel. The control relays operate on 28 VDC supplied through the corresponding AVIONICS MASTER circuit breaker in the associated circuit breaker panel. The AVIONICS MASTER switches have no effect when EMER BUS is selected and the generators are off-line. AVIONICS MASTER SWITCH The LEFT MASTER switch is installed in the pilots switch panel and the RIGHT MASTER switch is installed in the copilots switch panel. These two switches allow the crew to turn groups of avionic equipment off and on with only two switches. Refer to the Airplane Flight Manual for a listing of equipment controlled by the MASTER switches. The actual equipment affected may vary with customized wiring options.

AUXILIARY POWER UNIT (APU) GENERATOR

The APU generator provides 28 volts DC electrical power to the aircraft battery charging bus. The generator is controlled by a Generator Control Unit (GCU). The APU is only certified for ground use. After starting the APU using the APU control panel on the copilots circuit breaker panel, the green APU RUNNING annunciator will illuminate indicating that the APU system is ready to supply power to the aircraft. Refer to Auxiliary Power Unit in Section II of this manual.

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EMERGENCY POWER SYSTEM

The aircraft is equipped with either a dual or triple emergency power system to supply electrical power to selected equipment in the event of a normal electrical power system failure. Operating time of equipment powered by the emergency power supply is presented in the Airplane Flight Manual. Power for the emergency power system is supplied by two emergency power supply units located in the right, aft, nose avionics compartment. Each emergency power supply unit contains a 12-cell lead-acid battery to provide electrical power. The emergency power supply batteries are trickle charged from the aircraft normal electrical system through the EMER BAT circuit breakers on the pilots and copilots circuit breaker panels. If the normal electrical system has failed, EMER BAT 1 power supply will provide electrical power for the Electronic Standby Indicating System (ESIS) and lighting for the compass RTU 1, and CDU; EMER BAT 2 will supply electrical power for NAV 1, RTU 1, Data Concentrator Units (DCU 1 & 2), Attitude Heading Reference System (AHS 1 & 2), and air data computers (ADC 1 & 2); if a third emergency backup battery is installed, EMER BAT 3 will supply emergency power to COMM 1, AUDIO 1, FMS Display 1, and GPS 1/ADF 1 (Either or). The system is controlled through the EMER BAT 1, EMER BAT 2, and EMER BAT 3 switches on the pilots switch panel. Amber EMR PWR 1, EMR PWR 2 and EMR PWR 3 annunciators on the center instrument panel will illuminate whenever electrical power from the associated emergency power supply is being used. EMER BAT SWITCH The EMER BAT switches have two positions: On (EMER BAT 1, 2, 3) and OFF. With a switch in the On position, electrical power from the corresponding emergency power supply battery is available to supply emergency power should the normal electrical system fail. Normally, electrical power from the emergency power supply batteries is not used because 28 VDC from the normal electrical system is balanced against it. In the event of a failure of the normal electrical system, the balanced condition is removed and electrical power from the emergency power supplies is used.

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EXTERIOR LIGHTING
LANDING/TAXI LIGHTS A landing/taxi light is installed on each main landing gear. The lights are controlled by the LDG LT switches on the center switch panel. The LDG LT switches have three positions: On (L and R), TAXI, and OFF. The landing light control circuits are wired through the main gear down-and-locked switches; therefore, the landing lights are inoperative when the landing gear is not down and locked. When the LDG LT switches are placed in the On position, control circuits apply full 28 VDC to the landing lights and the lights will illuminate full bright. When the LDG LT switches are in the TAXI position, resistors shunt the lamp input power to 21 VDC and the lights are dimmed. In order to extend the service life of the lamps, it is recommended that the lights be used as sparingly as possible in the LDG LT mode. The lamps and control circuits are supplied electrical power through 20-amp current limiters. Some aircraft are equipped with a pulsating landing light option which is used in conjunction with the pulsating recognition light. On these aircraft, a pulse controller unit controls the landing lights by delivering pulsating DC current at approximately 45 cycles per minute. The effect of this pulsating current is to cause the bulbs brightness to continually vary between approximately 40% and 100% of full bright. The pulsating feature is activated when the RECOG light switch is set to the PULSE position, the applicable LDG LT switch is OFF and the landing gear is down and locked. When the LDG LT switch is positioned to On or TAXI, the landing/taxi lights will illuminate steadily. NAVIGATION LIGHTS Navigation lights are installed in the forward portion of the wing tips and in the vertical stabilizer upper aft fairing (bullet). The lights are controlled through the NAV switch in the LIGHTS group on the center switch panel. When the NAV light switch is placed in the On (NAV) position, the navigation lights will illuminate. Additionally, setting the NAV light switch to On (NAV) activates two-stage dimming and certain cockpit lights are automatically dimmed. Refer to TWO-STAGE DIMMING, this section. Electrical power for the navigation lights is 28 VDC supplied through the NAV LTS circuit breaker on the pilots circuit breaker panel.

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TAIL LOGO LIGHTS (OPTIONAL) Optional tail logo lights may be installed in the horizontal stabilizer on either side of the vertical stabilizer. These lights are used to illuminate both sides of the vertical stabilizer. The lights are controlled through the NAV switch in the LIGHTS group on the center switch panel. Aircraft with NAV LOGO-NAV-OFF Switch: When the NAV light switch is placed in the NAV LOGO position, the tail logo lights and navigation lights will illuminate. To use the navigation lights without the tail logo lights, select the NAV position of the switch. Electrical power for the tail logo lights is 28 VDC supplied through a 15-amp current limiter. Power for the control circuit is 28 VDC supplied through the LOGO LT circuit breaker on the copilots circuit breaker panel. ANTI-COLLISION (BEACON/STROBE) LIGHTS Anti-collision lights are mounted on top of the vertical stabilizer and on the bottom of the fuselage. Each light incorporates two flashtubes one with an aviation red filter and one with a clear filter. The lights are controlled through the BCN/STROBE light switch in the LIGHTS group on the center switch panel. On aircraft not modified by SB-60-33-7 (Modification of Strobe Light Switch), when the switch is placed in the BCN/STROBE position, the red flashtube in each light will flash if the aircrafts weight is on the wheels or the clear flashtube will flash if the aircrafts weight is not on the wheels. On aircraft modified by SB-60-33-7 (Modification of Strobe Light Switch), when the switch is placed in the STROBE position, the white flashtube in each light will flash whether or not the aircrafts weight is on the wheels.
.

When the switch is placed in the BCN/STROBE position, the red flashtube in each light will flash if the aircrafts weight is on the wheels or the clear flashtube will flash if the aircrafts weight is not on the wheels. When the switch is placed in the BCN position, the red flashtube in each light will flash whether or not the aircrafts weight is on the wheels. Therefore, when the clear strobe light is not desired in flight, the switch must be set to BCN or OFF. Each flashtube pulses at a rate of approximately 50 pulses per minute. The lights operate on 28 VDC supplied through the 7.5-amp BEACON-STROBE LTS circuit breaker on the copilots circuit breaker panel.

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RECOGNITION LIGHT A recognition light is installed on the upper, leading edge of the vertical stabilizer. The light is controlled through the RECOG light switch in the LIGHTS group on the center switch panel. When the switch is placed in the on (RECOG) position, control circuits apply full 28 VDC from the battery charging bus to illuminate the light. For greatest lamp life, it is recommended that the recognition light be turned OFF at altitudes of 18,000 feet or above. The recognition light operates on 28 VDC supplied through a 20-amp current limiter. Some aircraft are equipped with a pulsating recognition light option. On these aircraft, the RECOG light switch has a middle position labeled PULSE and a pulse controller unit. When the switch is placed in the PULSE position, 28 VDC from the PULSE RECOG LT circuit breaker is applied to the pulse controller unit which in turn lights the recognition light by delivering pulsating DC current at approximately 45 cycles per minute. The effect of this pulsating current is to cause the bulbs brightness to continually vary between approximately 40% and 100% of full bright. This feature results in enhanced aircraft recognition and improved bulb life. Also, the landing lights will pulse alternately with the recognition light if the landing gear is down and locked and the LDG LT switches are OFF. On aircraft with a pulsating recognition light, a PULSE RECOG LT circuit breaker on the copilots circuit breaker panel supplies 28 VDC to the pulse controller unit. WING INSPECTION LIGHT For a description of the wing inspection light, refer to Section VI, ANTI-ICE AND ENVIRONMENTAL. EXTERIOR CONVENIENCE LIGHTS Exterior convenience lights consist of a light on the underside of each engine pylon. The lights will illuminate the area around the tailcone baggage compartment and the single-point pressure refueling access. The lights are controlled by the entry light switch located near the entry door and are inoperative when the aircraft is in flight.

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COCKPIT LIGHTING
INSTRUMENT PANEL FLOODLIGHTS Lights are installed in the glareshield assembly to provide flood illumination of the instrument panel. The lights are controlled and dimmed through the FLOOD rheostat switch on the pilots switch panel. Electrical power is 28 VDC supplied through the FLOOD LTS circuit breaker on the pilots circuit breaker panel. Instrument panel floodlights are operative during EMER BUS mode. INSTRUMENT LIGHTS Lighting is installed for the pilots indicators, copilots indicators, center instrument panel indicators, pedestal indicators, and magnetic compass. Electrical power is 28 VDC supplied through the L and R INSTR LTS circuit breakers and the CENTER PANEL-PED LTS circuit breaker on the pilots and copilots circuit breaker panels. The lights are controlled and dimmed by the INSTR and CENTER PNL/PEDESTAL rheostat switches on the pilots switch panel and the INSTR rheostat switch on the copilots switch panel. Pilots INSTR dimmer switch: The pilots INSTR dimmer switch provides variable dimming for the following: Oxygen pressure indicator Pilots clock Pilots angle-of-attack indicator Copilots INSTR dimmer switch: The copilots INSTR dimmer switch provides dimming for the following: Pressurization panel Copilots clock Copilots angle-of-attack indicator APU control panel

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CENTER PNL/PEDESTAL dimmer switch: The CENTER PNL/PEDESTAL dimmer switch on the pilots switch panel provides dimming for the following: Autopilot panel Fuel quantity indicator ESIS HYD PRESS indicator Magnetic compass GEAR & BRAKE AIR indicator WING TEMP indicator NOSE STEER switch Fuel control panel HF control head Trim switch panel AIRSHOW Flight Deck Controller RTU, CDU and CCP panels Cabin pressure indicator Two master instrument light switches may be installed. They consist of two INSTR LIGHTS MASTER switches and the associated aircraft wiring. One master switch is located in the L INSTR LIGHTS group on the pilots switch panel and the other is located in the R INSTR LIGHTS group on the copilots switch panel. The WING INSP LIGHT switch, normally located on the copilots switch panel, may be relocated to a position on the instrument panel. The INSTR LIGHTS MASTER switches allow certain cockpit lighting to be turned on and off using one switch instead of multiple switches. The following lighting groups are controlled by the INSTR LIGHTS MASTER switches: L INSTR LIGHTS EL PNL CB PNL INSTR CENTER PNL/PEDESTAL R INSTR LIGHTS EL PNL INSTR CB PNL

The individual controls are used to select the brightness level of the affected instrument lights and the master switch is used to turn the lighting groups off and on as desired. TWO-STAGE LIGHTING Certain lights are automatically dimmed when the NAV light switch is set to NAV. When the NAV light switch is set to OFF, full 28 VDC is applied to the lights allowing them to illuminate at full brightness. When the NAV light switch is set to NAV, the voltage applied to the lights is reduced to approximately 14 VDC reducing their brightness. The lights dimmed by the two-stage dimmers are: Autopilot controller ANTI-SKID lights IGNITION lights Fuel control panel lights START lights PARK BRAKE light
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EFIS reversionary mode lights Pressurization FAULT/MANUAL light Pressurization EMER DEPRESS light CVR TEST & CVR ERASE switches NOSE STEER ARM annunciator

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SWITCH PANEL LIGHTING Electroluminescent panel lighting is provided for the pilots and copilots switch panels, the center switch panel, audio control panels, MIC/ PHONE jack panels, the pressurization control panel, anti-skid panel, system test switch panel, landing gear control panel, rudder pedal adjust panels, Display Control Panels (DCP), Cursor Control Panels (CCP), and circuit breaker panels. The panels are supplied 115 VAC through the L and R EL LTS circuit breakers on the pilots and copilots circuit breaker panels. The lights are controlled and dimmed through the EL PNL and CB PNL rheostat switches on the pilots and copilots switch panels. Pilots EL PNL and CB PNL dimmer switches: The pilots EL PNL dimmer switch controls the electroluminescent lighting of the pilots inboard and outboard switch panels, the center switch panel, the pilots audio control panel, the pilots rudder pedal adjust panel, the anti-skid panel, the system test switch panel, the landing gear control panel, the pilots DCP and CCP panels, throttle quadrant overlay, and the engine synchronizer switch panel. The pilots CB PNL dimmer switch controls the electroluminescent lighting of the pilots circuit breaker panel, and MIC/PHONE jack panel. Copilots EL PNL and CB PNL dimmer switches: The copilots EL PNL dimmer switch controls the electroluminescent lighting of the copilots switch panel, the pressurization control panel, the copilots audio control panel, the copilots DCP and CCP panels, and the copilots rudder pedal adjust panel. The copilots CB PNL dimmer switch controls the electroluminescent lighting of the copilots circuit breaker panel, MIC/PHONE jack panel and the APU control panel. ADAPTIVE FLIGHT DISPLAY (AFD) LIGHTING The brightness of the AFD tubes is controlled by two DISPLAY dimmer controls one on the pilots switch panel and one on the copilots switch panel. Each DISPLAY dimmer is used to adjust the brightness of the on-side outboard display, primary flight display (PFD) and the onside inboard display, multi-function display (MFD). The CDU screen lighting is controlled by the BRT Knob.

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MAP READING LIGHTS Map reading lights are located on the left and right cockpit sidewalls above the circuit breaker panels. Each lamp is mounted on a flexible conduit and is controlled by a rheostat switch located on the base of the assembly. The lights operate on 28 VDC supplied through the L and R INSTR LTS circuit breakers on the pilots and copilots circuit breaker panels. LIGHTED CHART HOLDERS A Lighted chart holder is located on each control wheel. Lighting is controlled by a control knob located on each chart holder. When the control knob is rotated fully counterclockwise the light is off. Rotating the knob clockwise will cause the light to come on and brighten as the knob is rotated. Chart holder lighting is powered by 28 VDC through the CHART HLDRS circuit breaker on the copilots circuit breaker panel. DOME LIGHTS Dome lights are installed in the cockpit overhead panel. These lights are used to illuminate the entire cockpit area. The lights are controlled by two separate electrical circuits. A rocker switch next to each light has three positions ON-off-REMOTE. If a BATTERY switch is on, setting a Dome Light switch to ON will illuminate the associated dome light. Rotating the associated OVHD dimmer control (pilots and copilots switch panel) will vary the brightness of the dome light. The ON position of the Dome Light switch is powered by 28 VDC through the R INSTR LTS circuit breaker on the copilots circuit breaker panel. When a Dome Light switch is placed in the REMOTE position, the associated dome light is controlled by the dome light function of the membrane switch panel, located near the entry door. The REMOTE position does not require a BATTERY switch to be on. The REMOTE position of the Dome Light switch is powered by 28 VDC supplied through the ENTRY LTS circuit breaker on the copilots circuit breaker panel.

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PASSENGER COMPARTMENT LIGHTING

The passenger compartment lighting consists of aisle lights, passenger reading lights, overhead lights, entry lights, NO SMOKING/FASTEN SEAT BELTS signs, lavatory lights, cabin baggage compartment lights, and the cove cabinet lights. AISLE LIGHTS Aisle lights are installed on each side of the center aisle to provide foot path lighting. The lights are controlled by the aisle light function of the Cabin Touch Screen located on the upper inboard portion of the left forward closet and the Master Control unit. The lights operate on 28 VDC supplied through the AISLE LTS circuit breaker on the pilots circuit breaker panel. PASSENGER READING LIGHTS Passenger reading lights are installed in the convenience panels above the seats on each side of the cabin. Some convenience panels consist of an eyeball-type air outlet and a reading light while others consist of a two-light assembly referred to as table lights. Each light includes an integral, directionally-adjustable lens. The lights are controlled through a CMS touch screen switch panel (READ LIGHTS and TABLE LIGHTS) in the armrest adjacent to each seat location. The lights operate on 28 VDC supplied through the READ LTS and TABLE LTS circuit breakers on the pilots circuit breaker panel. OVERHEAD LIGHTS General cabin lighting is provided by lights recessed in the cabin convenience panel. The lights operate on 28 VDC supplied through the CABIN LTS circuit breaker on the pilots circuit breaker panel. The lights are controlled through Cabin Touch Screen located on the upper inboard portion of the left forward closet and the Master Control unit. The switch panel provides on/off, bright and dim functions. In the event of cabin depressurization, the lights will automatically illuminate full bright if the cabin altitude reaches approximately 14,500 feet. Refer to OXYGEN SYSTEM for a description of emergency operation of the overhead lights.

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Pressing the upper switch will toggle the ON/OFF state of the cockpit dome light if the dome light switch in the cockpit is in the remote position. Pressing the middle switch will toggle the ON/OFF state of the entryway lights. Pressing the lower switch will toggle the On/Off state of the baggage light, vanity light, and lavatory reading light.

Entry Door Switch Panel (Located outboard on the aft side of the left forward cabinet) Figure 4-6
Pressing the Lighting position on the Cabin Control Switch Panel will cause the Cabin Control Switch Panel to advance to the lighting control panel. The lighting control panel toggles the following lights On/Off: Vanity/Bag Worksurface (Galley) DN Wash (Cabin Downwash) Club Accent Spot Light

Cabin Control Switch Panel (Located on the inboard top side of the left forward cabinet) Figure 4-7
Pressing the Reading Light position on a Passenger Control Switch Panel will cause the Reading Light for that seat to toggle On/Off: Pressing the Table Light position on a Passenger Control Switch Panel will cause the Table Light for that seat to toggle On/Off:

Passenger Control Switch Panel (Located in the armrest adjacent to passenger seats) Figure 4-8
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The Lavatory Switch Panel toggles the following lights On/Off: Lavatory/Vanity Lavatory Downwash Baggage Reading

Lavatory Switch Panel (Located in the lavatory wall) Figure 4-9 ENTRY LIGHT A cabin entry lights consist of a light in the top section of the door and a light on the bottom of the left forward cabinet. The lights are controlled by the entry light function of the entry door switch panel, located near the entry door. The lights circuits are wired to the right battery bus through the ENTRY LTS circuit breaker on the copilots circuit breaker panel. Therefore, the light is operable regardless of BATTERY switch position. The aircraft has a timer function that turns the cabin entry lights off after approximately 60 minutes after the upper cabin door is closed. LAVATORY LIGHTS The lavatory is illuminated by lights recessed in the lavatory convenience panel, a reading light in the RH overhead convenience panel, a vanity light assembly installed over the vanity cabinet, and vanity mirror lights. The reading, downwash lights and vanity/lavatory light are controlled with a membrane switch panel located on the RH lavatory wall. The reading light operates on 28 VDC supplied through the READ LTS circuit breaker on the pilots circuit breaker panel. The vanity/lavatory light operate on 28 VDC supplied through the ENTRY LTS circuit breaker on the pilots circuit breaker panel. The downwash lights operates on 28 VDC supplied through the CABIN LTS circuit breaker on the pilots circuit breaker panel.

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BAGGAGE COMPARTMENT LIGHT Overhead lights are installed in the cabin baggage compartment to provide illumination of the compartment. The lights are controlled by the entry light function of the membrane switch panel, located near the entry door or through a membrane-type baggage light switch located in the aft lavatory. The lights circuits are wired to the right battery bus through the ENTRY LTS circuit breaker on the copilots circuit breaker panel. Therefore, the light is operable regardless of BATTERY switch position. The aircraft has a timer function that turns the cabin entry lights off after approximately 60 minutes after the upper cabin door is closed. NO SMOKING AND FASTEN SEAT BELT SIGNS No smoking and fasten seat belt signs are installed in the cabin headliner immediately aft of the crew compartment and in the aft cabin. When illuminated, the sign displays symbolic representations for no smoking and fasten seat belts. Illumination of the sign is controlled through the NO SMOKING FASTEN SEAT BELT-OFF-FASTEN SEAT BELT switch on the center switch panel. When the switch is set to NO SMOKING FASTEN SEAT BELT, both symbols will illuminate and a chime will sound. When the switch is set to FASTEN SEAT BELT, only the fasten-seat-belt symbols will illuminate and the tone will sound. Additionally, a RETURN TO SEAT sign is installed in the lavatory. The RETURN TO SEAT sign will be illuminated whenever the fasten seat belt sign is illuminated. Electrical power to illuminate the signs is 28 VDC supplied through the PASS INFO circuit breaker on the copilots circuit breaker panel. The chime is generated by the passenger speaker amplifier and broadcast through the passenger speakers. When the CABIN PWR switch is selected OFF, the illuminated NO SMOKING/FASTEN SEAT BELT sign is disabled. Some aircraft have a no smoking cabin. In these aircraft, the no smoking portion of the no smoking and fasten seat belt signs is illuminated anytime one of the BATTERY switches is on. A two-position FASTEN SEAT BELT-OFF switch replaces the three-position NO SMOKING FASTEN SEAT BELT-OFF-FASTEN SEAT BELT switch on the center switch panel.

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CARGO AND SERVICING COMPARTMENT LIGHTING

TAILCONE BAGGAGE LIGHTS Two lights are installed along the LH side of the tailcone baggage compartment to provide illumination of the compartment. A door-actuated switch and BAGGAGE LIGHTS - OFF toggle switch are installed. The toggle and door-activated switches are wired in series to the light assemblies; therefore, the baggage access door must be open and the toggle switch set to BAGGAGE LIGHTS to illuminate the lights. When the toggle switch is set to OFF, the lights will extinguish regardless of the door position. The lights will operate regardless of BATTERY switch position. TAILCONE MAINTENANCE LIGHT A tailcone maintenance light is installed in the tailcone equipment compartment to provide illumination of the compartment. The system consists of a light assembly, a MAINT LIGHTS - OFF toggle switch and a door-actuated switch. The toggle switch and door-actuated switch are wired in series to the light assembly; therefore, the tailcone access door must be open and the toggle switch set to the MAINT LIGHTS position to illuminate the light. When the toggle switch is set to OFF, the light will extinguish regardless of the access door position. When the access door is closed, the light will extinguish regardless of the toggle switch position. The maintenance light operates on 28 VDC supplied from the #1 battery through a current limiter.

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ILLUMINATED EXIT SIGN SYSTEM

The Learjet 60XR aircraft comes standard with six illuminated exit signs, one located above the entry door, two in the LH FWD cabinet, one in the RH AFT partition, one in the lavatory toilet shroud and one above the emergency door. The illuminated exit signs system provides exit sign lighting in the event of a normal electrical system failure. The system also includes two emergency battery units, two egress light assemblies (located in the aircraft exit doors) an illuminated exit sign control panel in the cockpit and associated aircraft wiring. The batteries are charged through the EMER LTS circuit breaker on the copilot's circuit breaker panel. If armed, the system will automatically activate whenever R DC BUS 4 loses normal electrical power. Therefore, the system will automatically activate during EMER BUS mode. EMERGENCY EXIT LIGHTS BATTERY UNITS The battery units, used in the illuminated exit sign system, are rechargeable, 24-volt, and maintenance-free. Each battery unit incorporates a relay that when activated will connect the battery to the lights utilized for emergency illumination of the exit signs. The relay will remain latched in this position until a signal to reset is received. Therefore, once activated the illuminated exit sign system will remain activated even though control wiring may become severed. One battery is located in the forward part of the cabin while the other is located in the aft part of the cabin. Either battery is capable of powering the entire illuminated exit sign system by itself, thus allowing all illuminated exit signs to activate even with a vertical transverse separation of the cabin. EGRESS LIGHT ASSEMBLIES An egress light assembly is installed in the upper cabin door and the emergency escape/baggage door. When activated, these lights provide illumination of the emergency exits. Each light assembly includes a momentary push button switch. If the system is armed but not activated, pressing either push button switch will manually activate the system. EMERGENCY EXIT LIGHTING (OPTIONAL) The optional emergency exit lighting is supplemental to and works in conjunction with the illuminated exit sign system. The additional lighting provided by this option consists of the three cabin table lights, galley work surface light and the cabin aisle lights. These lights are utilized to provide cabin lighting for emergency egress.
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EMERGENCY EXIT LIGHTS CONTROL PANEL The EMERGENCY EXIT LIGHTS control panel, in the cockpit, provides control, testing, and indicating functions for the illuminated exit signs, egress lights and the optional emergency exit lighting. The panel includes: one control switch (ON-ARMED-OFF/RESET), one test switch (TEST BAT 1-NORM-TEST BAT 2), one white ON annunciator, and one amber NOT ARMED annunciator.

EMERGENCY EXIT LIGHTS

ON ARMED NOT ARMED OFF/RESET BAT 2 TEST BAT 1 NORM

EMERGENCY EXIT LIGHTS CONTROL PANEL Figure 4-10 CONTROL SWITCH Functions of the control switch are shown in the following table:

SWITCH POSITION

SYSTEM RESPONSE
The relays in both battery units will reset to off and all emergency exit lighting will go out. Pressing one of the push button switches at either exit will activate the system while held. Upon release, the system will reset to off. Arms the system to automatically activate should normal electrical power be lost. Selecting ARMED prior to powering up the aircraft will cause the system to activate immediately. Pressing one of the push button switches at either exit will manually activate the system. To manually activate the system, hold switch momentarily to ON and release. The switch will spring back to the ARMED position and the system will remain activated.
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TEST SWITCH The test switch is a three-position switch spring loaded to the NORM position. The test switch is used to verify each battery unit is capable of powering all the emergency exit lighting by itself. To test system: 1. Aircraft BATTERY Switches On. 2. EMERGENCY EXIT LIGHTS Switch ARMED. 3. TEST Switch BAT 1 and hold. All six illuminated exit signs and both egress lights will illuminate. ON annunciator will also illuminate. If the optional emergency exit lighting is installed then the cabin table lights, galley work surface light and the cabin aisle lights will also illuminate. 4. TEST Switch BAT 2 and hold. All six illuminated exit signs and both egress lights will illuminate. ON annunciator will also illuminate. If the optional emergency exit lighting is installed then the cabin table lights, galley work surface light and the cabin aisle lights will also illuminate. 5. TEST Switch Release to NORM. Emergency exit lighting will reset to off and the ON annunciator will extinguish. ANNUNCIATORS Meaning of the ON and NOT ARMED lights is shown in the following table:

ANNUNCIATION
ON

MEANS
The system is activated either manually or automatically. Also annunciates during test. The aircraft is powered up and the system is not yet armed. Also annunciates whenever the system has been automatically activated. Illumination of NOT ARMED will trip the Master CAUT lights.

NOT ARMED

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MASTER CAUTION/WARNING AND ANNUNCIATOR PANEL LIGHTS

Master WARN/CAUT lights on the pilots and copilots instrument panels and annunciator panel cockpit warning lights give a visual indication of various systems operating conditions. The annunciator panel lights are white (advisory), green (normal), amber (caution) and red (warning). The annunciator panel cockpit warning lights may be tested by pressing the test switch on either side of the panel. During the first 3 seconds of the lamp test, the two bulbs in each light will alternately illuminate. Thereafter, all the bulbs will illuminate until the test switch is released. Photoelectric cells, outboard of each ENG FIRE PULL switch, automatically dim the annunciator panel lights to a level corresponding to existing light in the cockpit or to a minimum preset level for a totally dark cockpit. Other cockpit annunciator lights are dimmed when the NAV lights are on. If an annunciator light illuminates and the condition is corrected, the light will extinguish. If the condition recurs, the light will again illuminate. Illumination of any red cockpit annunciator will cause both Master WARN lights to illuminate and flash. Depressing the Master WARN/ CAUT light will extinguish the Master WARN light even though the annunciator light may be flashing (ENTRY DOOR, AFT CAB DOOR, L or R STALL, CABIN FIRE, or either ENG FIRE PULL). Illumination of any amber cockpit annunciator, except starter engaged lights (during ground operations), will cause both Master CAUT lights to illuminate and flash unless the master caution feature has been inhibited. Depressing the Master WARN/CAUT light will extinguish the Master CAUT light even though the annunciator light may be illuminated. The annunciator light will remain on as long as the condition exists. When the aircraft is on the ground, the master caution feature may be inhibited by depressing and holding either Master WARN/CAUT light until the Master CAUT light illuminates steadily. Approximately 10 seconds after takeoff, the master caution feature will revert to the normal (uninhibited) mode.

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Most white annunciators may be extinguished in flight by depressing either Master WARN/CAUT light. Depressing either warning lights Test switch will cause the annunciators to illuminate again. A white ENG CMPTR light accompanied by an amber ENG CMPTR light may not be extinguished. Any white annunciators which were extinguished in flight will again illuminate shortly after touchdown. When an EIS page is not displayed and parameters on that page are out of tolerance, there will be an amber or red flag in the lower left of the currently displayed page indicating the page with the out of tolerance indication. This indication is in addition to the Master WARN/CAUT light.

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SECTION V FLIGHT SYSTEMS & AVIONICS

TABLE OF CONTENTS Flight Controls ........................................................................................... 5-1 Aileron and Elevator ............................................................................. 5-1 Rudder..................................................................................................... 5-1 Pedal Adjust Switches...................................................................... 5-1 Controls Gust Lock ................................................................................ 5-2 Controls Gust Lock (Figure 5-1)........................................................... 5-2 Flaps......................................................................................................... 5-3 Flap Selector Switch ......................................................................... 5-3 Flight Control Page (Figure 5-2)........................................................... 5-3 Flap Position Indicator .......................................................................... 5-4 Spoilers .................................................................................................... 5-4 SPOILER Lever.................................................................................. 5-6 SPOILER EXT Light.......................................................................... 5-6 SPOILER ARM Light........................................................................ 5-7 SPOILER MON Light ....................................................................... 5-7 System Test Switch Spoiler Reset Function.............................. 5-7 Trim Systems .............................................................................................. 5-8 Mach Trim ............................................................................................... 5-8 Pitch Trim Selector Switch Mach Trim Function ..................... 5-9 MACH TRIM Light .......................................................................... 5-9 System Test Switch Mach Trim Function ................................. 5-9 Trim Control Panel (Figure 5-3) ......................................................... 5-10 Pitch Trim .............................................................................................. 5-10 Pitch Trim System Block Diagram (Figure 5-4)................................ 5-11 Pitch Trim Selector Switch ............................................................. 5-12 Control Wheel Trim Switches Pitch Function........................ 5-12 NOSE DN-OFF-NOSE UP Switch ................................................ 5-12 Control Wheel Master Switches Pitch Trim Function .......... 5-13 PITCH TRIM Light ......................................................................... 5-13 T. O. Trim Light ............................................................................... 5-13 System Test Switch Trim Overspeed Function ...................... 5-13 Pitch Trim Indicator........................................................................ 5-14 Trim-In-Motion Audio Clicker...................................................... 5-14 Roll Trim................................................................................................ 5-15 Control Wheel Trim Switches Roll Function.......................... 5-15 Control Wheel Master Switches Roll Trim ............................. 5-15 Aileron Trim Indicator ................................................................... 5-15
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TABLE OF CONTENTS (Cont) Yaw Trim ............................................................................................... 5-16 Rudder Trim Switch ....................................................................... 5-16 Rudder Trim Indicator................................................................... 5-16 Control Wheel Master Switches Yaw Trim ............................ 5-16 Warning Systems..................................................................................... 5-17 Stall Warning System .......................................................................... 5-17 Stall Warning System Block Diagram (Figure 5-5) ......................... 5-18 Angle-of-Attack Indicators ........................................................... 5-19 Low-Speed Awareness Cues.............................................................. 5-19 STALL Warning Lights .................................................................. 5-19 System Test Switch Stall Warning Function .......................... 5-19 Overspeed Warning System.......................................................... 5-20 System Test Switch Overspeed Warning Function............... 5-20 Takeoff Warning System..................................................................... 5-20 Enhanced Ground Proximity Warning System with Windshear Detection (EGPWS/WS) .................................... 5-21 Traffic Alert and Collision Avoidance System (TCAS) .................. 5-22 Air Data Systems..................................................................................... 5-23 Primary Pitot-Static System ............................................................... 5-23 Primary Pitot-Static System Schematic (Figure 5-6) ....................... 5-24 Static Source Switch ............................................................................ 5-24 Standby Pitot-Static System ............................................................... 5-25 Standby Pitot-Static System Schematic (Figure 5-7) ....................... 5-25 Air Data Computers ............................................................................ 5-25 ADC/ADC Transfer Switch............................................................... 5-26 Attitude Heading System ...................................................................... 5-27 Heading Control Switches ................................................................. 5-28 AHS/AHS Reversionary Mode......................................................... 5-28 Magnetic Compass .............................................................................. 5-28 Electronic Standby Instrument System (ESIS) ................................ 5-29 Electronic Standby Instrument System (Figure 5-8) ....................... 5-29 Electronic Flight Instrument System (EFIS) ........................................ 5-30 Primary Flight Display (PFD)............................................................ 5-31 Multifunction Display (MFD)............................................................ 5-31 EFIS Control Panel .............................................................................. 5-32 Display Control Panel (DCP)............................................................. 5-32 Heading, Speed, Altitude Panel (HSA)............................................ 5-33 Course (CRS) Control Knobs ............................................................. 5-33 Cursor Control Panel (CCP) .............................................................. 5-33

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TABLE OF CONTENTS (Cont) Communications ..................................................................................... 5-34 VHF Communications ........................................................................ 5-34 HF Communications ........................................................................... 5-34 SELCAL ................................................................................................. 5-35 Audio Control System......................................................................... 5-35 Audio Control Panel....................................................................... 5-35 Audio Control Panel (Figure 5-9) ................................................. 5-36 MIC SELECT Switch....................................................................... 5-36 NORM MIC/OXY MIC Switch .................................................... 5-36 Volume Controls ............................................................................. 5-37 Radio Monitor Switches................................................................. 5-37 BOTH/VOICE/IDENT Switch..................................................... 5-38 Marker Beacon HI/LO Switch...................................................... 5-38 Audio Control Flight Operation .............................................. 5-38 Cabin Briefing System......................................................................... 5-39 Airshow Cabin Video Information System................................. 5-39 Navigation ................................................................................................ 5-40 VHF Navigation................................................................................... 5-40 Marker Beacon Display.................................................................. 5-41 Distance Measuring Equipment (DME) ...................................... 5-41 Automatic Direction Finding (ADF) ................................................. 5-41 ATC Transponders ............................................................................... 5-41 Radio Altimeter.................................................................................... 5-42 Flight Control System (FCS) .................................................................. 5-42 Autopilot/Flight Director System ..................................................... 5-42 Flight Control Panel (FCP) ............................................................ 5-43 Self-Test ............................................................................................ 5-43 Autopilot Engage Functions ......................................................... 5-44 Autopilot/Flight Guidance Mode Selection ............................... 5-44 FCP Annunciators........................................................................... 5-46 Control Wheel Master Switches Autopilot Function............ 5-47 Pitch Trim Selector Switch Autopilot Function ..................... 5-47 Control Wheel Trim Switches Autopilot/Flight Director Function ......................................... 5-47 NOSE DN-OFF-NOSE UP Switch Autopilot Function ........ 5-48 SYNC Switches................................................................................ 5-48 FD CLEAR Switches....................................................................... 5-48

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TABLE OF CONTENTS (Cont) Yaw Damper System ........................................................................... 5-48 Yaw Damper Control ..................................................................... 5-48 Control Wheel Master Switches Yaw Damper Function ..... 5-49 Rudder Boost System.......................................................................... 5-49 Rudder Boost Switch...................................................................... 5-49 Flight Management System (FMS) ....................................................... 5-50 Weather Radar ......................................................................................... 5-51 Miscellaneous .......................................................................................... 5-52 Cockpit Voice Recorder (CVR) .......................................................... 5-52 Flight Data Recorder (FDR) ............................................................... 5-53 Clocks .................................................................................................... 5-53 Hourmeter Aircraft ........................................................................ 5-54 Emergency Locator Transmitter ........................................................ 5-54 Transmitter and Antenna .............................................................. 5-54 Remote Switch ................................................................................ 5-54 Lightning detection System (LDS) (Optional)................................. 5-55 XM Satellite Weather (Optional) ....................................................... 5-56 Universal Weather (Optional)............................................................ 5-56 Navigation Source ............................................................................... 5-57 3D Map Format (Optional) ................................................................ 5-57 E-Charts (Optional) ............................................................................. 5-57 Jeppesen Display Charts (Optional) ................................................. 5-58 NOTAMS (Optional) ........................................................................... 5-58

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SECTION V FLIGHT SYSTEMS & AVIONICS

FLIGHT CONTROLS
The primary flight controls (ailerons, elevator, and rudder) are mechanically operated through the control columns, control wheels, and rudder pedals. The flaps and spoilers are hydraulically operated and electrically controlled. Aircraft trim systems (pitch, roll, and yaw) are electrically operated and controlled. AILERON AND ELEVATOR Movement of the control columns and control wheels is mechanically translated into elevator and aileron control surface movement through systems of cables, pulleys, and push-pull rods. In addition to aileron control, the control wheels incorporate switches that control normal trim, pitch-axis interrupt, autopilot and yaw damper disconnect, flight director clear, flight director sync, microphone keying, and nose wheel steering engage and disengage circuits. Control wheel switch functions are discussed under the applicable system. RUDDER Rudder pedal movement is mechanically translated into rudder control surface movement through a system of cables, pulleys, and bellcranks. Nose wheel steering, when engaged, is electronically controlled by the pedals and braking may be accomplished by depressing the upper portion of the pedals. PEDAL ADJUST SWITCHES The pilots and copilots rudder pedals are individually adjustable through the PEDAL ADJUST switches on the pilots and copilots outboard switch panels. Each switch has three positions: FWD, OFF, and AFT. When either switch is held to the FWD or AFT position, an electrically controlled actuator will move the corresponding rudder pedals in the desired direction. The rudder pedal adjust system operates on 28 VDC supplied through the RUDDER PEDAL ADJUST circuit breaker on the copilots circuit breaker panel.

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CONTROLS GUST LOCK A controls gust lock is provided to help prevent wind gust damage to the movable control surfaces. When installed, the lock provides security by holding full rudder, full aileron, and full down elevator.

CONTROLS GUST LOCK Figure 5-1

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FLAPS The hydraulically-actuated, electrically-controlled flap system provides flap settings of UP (0), 8, 20, and DN (40). The single-slotted flaps are attached to the rear wing spar with tracks, rollers, and hinges. The flap selector switch controls a solenoid-operated hydraulic control valve that meters hydraulic pressure to the flap actuators. The actuators mechanically rotate sectors attached to the flaps through adjustable push-pull tubes. Interconnecting cables and pulleys synchronize flap movement throughout the range of travel. A flap position switch is mechanically connected to each flap sector. These switches provide flap position information to the landing gear warning, stall warning, spoiler warning, trim-in-motion warning, spoileron, and autopilot systems. A flap limit switch is mechanically connected to each sector to automatically maintain flap position at the selected setting. Overtravel, when the flaps are fully extended, is mechanically prevented. The flap control system operates on 28 VDC supplied through the FLAPS circuit breaker on the copilots circuit breaker panel. The flaps are operative during the EMER BUS mode. FLAP SELECTOR SWITCH The flap selector switch is located on the right side of the pedestal near the thrust levers. The switch has four positions: UP, 8, 20, and DN. The switch handle is shaped like an airfoil. When 8 or 20 flaps is selected, 28 VDC is directed to the applicable (up or down) solenoid of the flap control valve. The flap control valve will meter hydraulic pressure to the flap actuators and move the flaps in the desired direction. As the flaps approach within 1 of the selected setting, the applicable flap limit switch will remove power from the flap control valve solenoid and flap travel will stop. When UP is selected, 28 VDC is directed to the up solenoid of the flap control valve and the flaps will move in the up direction. When DN is selected, 28 VDC is directed to the down solenoid of the flap control valve and the flaps will move in the down direction. When the flaps reach full extension, the down pressure will remain to maintain the flaps full down.

SELCAL VHF 1 VHF 2

FLIGHT CONTROL PAGE Figure 5-2

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FLAP POSITION INDICATOR The FLAPS indicator, located on the EIS Flight Display Page, provides the crew with visual indication of flap position. The indicator face consists of a scale, which has markings for UP (0), 8, 20, and DN (40), and a pointer on the left of the scale. A potentiometer connected to the left flap sector transmits the flap position signal to the indicator. The indicator operates on 28 VDC supplied through the TRIM-FLAP-SPOILER INDICATOR circuit breaker on the copilots circuit breaker panel. The flap position indicator is operative during the EMER BUS mode. SPOILERS The spoilers, located on the upper surface of the wings forward of the flaps, may be extended symmetrically for use as spoilers or asymmetrically for aileron augmentation when the flaps are extended. The spoilers are electrically controlled and hydraulically actuated either by a control switch (Normal Spoiler Mode), by the wing flap position switches (Spoileron), or automatically during ground operations when the thrust levers are pulled to idle (Autospoilers). Autospoilers: The autospoiler mode is used to automatically extend the spoilers on landing or in the event of an aborted takeoff. When the SPOILER lever is set to ARM, the system will be armed (SPOILER ARM light will illuminate) to automatically extend both spoilers when one of the following conditions are met.

Flight Phase
Aborted Takeoff

Autospoilers will deploy when:

Aircraft accelerates to 40 knots or greater groundspeed and the thrust levers are brought to IDLE per the ABORTED TAKEOFF procedure. Spoilers will remain deployed unless a thrust lever is advanced above IDLE. Either of the following occurs: 1. Both squat switches indicate an on ground condition and both thrust levers are in IDLE (one may be in CUTOFF) or 2. A wheel speed of 40 knots or greater is attained at touchdown and both thrust levers are in IDLE (one may be in CUTOFF). Spoilers will remain deployed unless a thrust lever is advanced above IDLE.

Landing

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Once spoilers are deployed, the deploy signal will latch and cycling the squat switches will not stow the spoilers. Advancing one or both throttles will release the latch and stow the spoilers. Normal spoiler extension and retraction will override the autospoiler logic. Flap position has no effect on autospoiler operation and autospoilers are not operational when EXT or RET is selected. Autospoiler control circuits operate on 28 VDC supplied through the SPOILER circuit breaker on the copilots circuit breaker panel. Autospoilers are operative during the EMER BUS mode. Normal Spoiler Mode: During the spoiler mode, the spoilers are symmetrically extended and retracted through the SPOILER lever on the forward pedestal. In flight, the spoilers may be extended to any desired position by placing the SPOILER lever in any position between ARM and EXT. Detents for approximately 10 and 20 positions are provided between the ARM and EXT positions of the lever. On the ground, the spoilers will extend fully whenever any partial extension is selected. The SPOILER indicator, located on the EIS Flight Display Page, provides the crew with visual indication of spoiler position. The spoiler mode, when selected, will override the aileron augmentation (spoileron) mode, if aileron augmentation is engaged. When the spoiler lever is positioned for spoiler extension, a computer-amplifier will command a selector valve and two servo valves to the extend position. These valves will apply hydraulic pressure to the spoiler actuators and cause the spoilers to extend. As the spoilers unseat and extend through 1, the SPOILER EXT light will illuminate and the computer will close a restrictor bypass to restrict hydraulic flow into the return line. The spoilers will fully extend in approximately 5 to 7 seconds. Full extension is approximately 45. However, during flight, a pressure relief allows the spoilers to blow down to a lesser extension angle. When RET is selected, the computer-amplifier will command the servo valves closed and the selector valve to retract. The selector valve will then apply hydraulic pressure to the spoiler actuators and cause the spoilers to retract. When retracted, the spoilers are secured by an internal locking mechanism in the actuators. The spoilers will fully retract in approximately 4 seconds. A monitor circuit will automatically retract both spoilers and illuminate the SPOILER MON light should a malfunction occur. Spoiler mode control circuits operate on 28 VDC supplied through the SPOILER circuit breaker on the copilots circuit breaker panel. The spoilers are operative during EMER BUS mode.

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Spoileron Mode: During the spoileron (aileron augmentation) mode, the spoilers are independently raised and lowered in a one-to-one ratio with the upgoing aileron to improve lateral control with the flaps full down. Aileron augmentation is automatically engaged when the flaps are lowered beyond 25 and the SPOILER lever is in the RET or ARM position. During the spoileron mode, the computer-amplifier continuously monitors aileron position through follow-ups on the aileron sectors. As the ailerons move, the computer-amplifier actuates the spoiler selector and servo valves to control spoiler movement. As one aileron moves up, the servo valves are positioned so that the spoiler on the same wing moves up with the aileron while the opposite spoiler remains retracted. A limit switch for each spoiler limits spoiler extension to approximately 15. A monitor circuit will automatically retract both spoilers and illuminate the SPOILER MON light should a malfunction occur. The spoileron mode operates on 115 VAC supplied through the SPOILERON circuit breaker on the copilots circuit breaker panel. SPOILER LEVER Symmetric extension and retraction of the spoilers is controlled through the SPOILER lever located on the left side of the pedestal adjacent to the thrust levers. The lever has five positions: RET, ARM, two partial extension detents and EXT. When the switch is set to EXT, both spoilers will extend and the SPOILER EXT light will illuminate. When the lever is set to ARM, the autospoiler system will be armed for automatic spoiler extension and the SPOILER ARM light will illuminate. When the lever is set to RET, both spoilers will retract. The spoilers may be extended partially by placing the spoiler lever between ARM and EXT. When on the ground, the spoilers will extend fully when the spoiler lever is in any position between ARM and EXT. SPOILER EXT LIGHT The SPOILER EXT light, located on the glareshield annunciator panel, will illuminate steady whenever the flaps are UP and the spoilers are extended. The light will flash if the spoilers are extended and the flaps are beyond 3. The light is operated by a 1-up position switch for each spoiler. The light will illuminate if either 1-up switch is actuated except during spoileron mode.

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SPOILER ARM LIGHT The SPOILER ARM light, on the glareshield annunciator panel, will illuminate whenever the autospoiler mode is armed and remains illuminated when autospoilers are extended. The light will not illuminate and the autospoiler system will not arm (SPOILER ARM light will not come on), or will disarm (SPOILER ARM light will go out), if the squat switches are in an asymmetric condition for more than approximately 2 minutes. SPOILER MON LIGHT The amber SPOILER MON light, located on the glareshield annunciator panel, will illuminate whenever monitor circuits in the computeramplifier detect a malfunction during the spoileron mode or unequal spoiler extension during the spoiler mode. Should the monitor detect a malfunction during aileron augmentation, the monitor will automatically disengage the spoileron mode and the spoilers will immediately retract. If the monitor has disabled aileron augmentation or the SPOILERON circuit breaker is pulled, normal spoiler mode operation will not be available in flight; however, the spoilers will be available for ground operation. The autospoilers will also be operational but should not be armed if the SPOILERON circuit breaker is open. During the spoiler mode, the SPOILER MON light will illuminate and both spoilers will retract in the event of unequal spoiler extension where the difference is 6 or more. Additionally, the SPOILER MON light will also illuminate if either of the autospoiler dual logic circuits fail. SYSTEM TEST SWITCH SPOILER RESET FUNCTION The rotary-type system test switch, located on the center instrument panel, is used to test the spoiler system. During flight, the SPOILER RESET position is used to reset the spoiler/ spoileron system in the event of a malfunction. Should the monitor disable spoiler/spoileron mode (SPOILER MON light illuminated) and the fault clears, the system may be enabled by momentarily placing the system test switch in the SPOILER RESET position. If the system is reset, the SPOILER MON light will extinguish. If the spoiler/spoileron system cannot be reset, the SPOILER MON light will remain illuminated and normal spoiler or spoilerons will not be available in flight. During ground operations, the switch is used during the spoileron and autospoiler test sequence to verify system operation. Placing the system test switch in the SPOILER RESET position and depressing the PRESS TEST button in the center of the switch will simulate a malfunction.
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TRIM SYSTEMS
MACH TRIM The Mach trim system provides automatic pitch trim in response to Mach changes to increase longitudinal stability and counteract the center-of-lift movement at speeds above approximately 0.70 MI if the autopilot is disengaged or inoperative. The system consists of a computer, a pitch trim followup, the MACH TRIM annunciator light, and associated aircraft wiring. The Mach trim computer receives Mach data from the air data computers. The Mach trim system utilizes the primary motor of the horizontal-stabilizer pitch-trim actuator to affect trim changes. The Mach trim computer operates on 115 VAC supplied through the MACH TRIM circuit breaker and 28 VDC supplied through the PRI PITCH TRIM circuit breaker on the pilots circuit breaker panel. The Mach trim system is inoperative during EMER BUS mode. During flight, with the autopilot disengaged or inoperative, the Mach trim system will automatically engage at approximately 0.70 MI. As the aircraft Mach number changes, the change is sensed by the air data computers and transmitted to the Mach trim computer. If the aircraft is not retrimmed to compensate for the Mach change, the Mach trim computer will command the appropriate pitch trim change (nose up for increased Mach and nose down for decreased Mach) through the horizontal-stabilizer pitch-trim actuator. A followup on the horizontal stabilizer will transmit a horizontal stabilizer position signal to the Mach trim computer. Stabilizer trim motion will cease as the followup stabilizer position signal cancels the pitch trim signal from the Mach trim computer. Monitors are installed to disengage Mach trim in the event of a malfunction. If a monitor disengages Mach trim and Mach is above 0.77 MI, the overspeed warning horn will sound. The Mach trim system is resynchronized whenever either pilot manually trims the aircraft and a synchronous standby mode is maintained if the autopilot is engaged. In flight, Mach trim monitor may also be reset through the SYSTEM TEST switch on the center instrument panel.

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PITCH TRIM SELECTOR SWITCH MACH TRIM FUNCTION The Mach trim system utilizes the primary motor of the horizontal stabilizer pitch trim actuator to increase longitudinal stability. If the PITCH TRIM selector switch on the pedestal is in the PRI position, Mach trim will automatically engage at approximately 0.70 MI if the autopilot is disengaged or inoperative. Mach trim will not engage or will disengage when the PITCH TRIM selector switch is moved to the OFF or SEC position. If the PITCH TRIM selector switch is in OFF or SEC, the Mach trim monitor will remain active and will illuminate the MACH TRIM light and cause the overspeed warning horn to sound at or above 0.77 MI if the monitor detects a sufficient Mach/horizontal stabilizer position error. MACH TRIM LIGHT The amber MACH TRIM annunciator light, located on the glareshield annunciator panel, will illuminate whenever the Mach trim monitor or Mach monitor has disengaged the Mach trim system. Whenever the Mach trim system is disengaged and Mach is above 0.77 MI, the overspeed warning horn will sound if the autopilot is inoperative or not engaged. The Mach trim monitor continuously monitors input signals and power to the Mach trim computer. In the event of loss of power to the Mach trim computer or primary pitch trim system, loss of input signals to the Mach trim computer, or a Mach/horizontal stabilizer position error, the Mach trim monitor will disengage Mach trim and illuminate the MACH TRIM light. SYSTEM TEST SWITCH MACH TRIM FUNCTION The rotary-type SYSTEM TEST switch on the center instrument panel is used to test the Mach trim system and the Mach trim monitor while the aircraft is on the ground. In flight, the switch is used to resynchronize the system if the Mach trim monitor has disengaged the system. The test function is initiated by rotating the switch to MACH TRIM and then depressing the switch PRESS TEST button. When the aircraft is on the ground and the test sequence is initiated, the test switch inserts a signal that causes the horizontal stabilizer to trim in the nose-up direction. Since there is no corresponding airspeed change, the Mach trim monitor senses a Mach/horizontal stabilizer position error, disengages Mach trim, and illuminates the MACH TRIM light. In flight, depressing the PRESS TEST button will resynchronize the Mach trim system to the horizontal stabilizer position and Mach existing when the PRESS TEST button was depressed.
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PITCH TRIM PRI O F F SEC NOSE UP NOSE DN O F F

RUDDER TRIM NOSE OFF NOSE LEFT RIGHT

TRIM CONTROL PANEL Figure 5-3 PITCH TRIM Pitch trim is accomplished by repositioning the horizontal stabilizer to the desired trim setting through actuation of the horizontal stabilizer pitch trim actuator. The actuator is a dual-motor, screwjack-type actuator. The primary motor is operated by the aircraft primary pitch trim system and the Mach trim system. The secondary motor is operated by the aircraft secondary pitch trim system and the autopilot. A speed controller in the primary pitch trim system changes primary pitch trim rate as a function of horizontal stabilizer trim position. The speed controller allows high trim rates when the aircraft is trimmed for takeoff or approach and low trim rates when the aircraft is trimmed for cruise. A trim speed monitor is incorporated into the speed controller to alert the crew of a trim speed error. The primary and secondary pitch trim systems are electrically independent and mode selection is made through a selector switch. Primary pitch trim is pilot controlled through trim switches on each control wheel. Secondary pitch trim is pilot controlled through a switch on the pedestal. Emergency interrupt is provided for both systems through the Control Wheel Master switches (MSW). The ELEV trim indicator, located on the EIS Flight Display Page, provides the crew with visual indication of horizontal stabilizer position. Primary pitch trim control circuits operate on 28 VDC supplied through the PRI PITCH TRIM circuit breaker on the pilots circuit breaker panel. Secondary pitch trim control circuits operate on 28 VDC supplied through the SEC PITCH TRIM circuit breaker on the copilots circuit breaker panel. Both the primary and secondary pitch trim systems are operative during EMER BUS mode.

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3 FLAP SWITCH

PILOTS CONTROL WHEEL TRIM SWITCH

UP-DN

COPILOTS CONTROL WHEEL TRIM SWITCH

UP-DN

PEDESTAL NOSE DN-OFFNOSE UP SWITCH

UP-DN

MON

TRIM SW PANEL MACH TRIM COMPUTER

UP-DN SYNC (PILOT AUTHORITY) (PITCH TRIM SEL) (AUTOPILOT DISENGAGE) PRI UP-DN SEC UP-DN

MACH TRIM ANNUNCIATOR

AUTOPILOT COMPUTER
UP-DN

PRIMARY TRIM CONTROL PITCH TRIM ANNUNCIATOR

MACH MON OVERSPEED MON DISABLE UP-DN SPEED CONTROL DISABLE

DISENGAGE

CONTROL WHEEL MASTER SWITCH

(MSW)

POSITION

RATE

CONTROL WHEEL MASTER SWITCH

(MSW)

P R I M A R Y

HORIZONTAL STABILIZER PITCH TRIM ACTUATOR

POSITION POSITION MOTION

S E C O N D A R Y

AUDIO CLICKER

POSITION

RATE SWITCH
(LO SPD WATCH)

T.O. TRIM LIGHT

PITCH TRIM INDICATOR

TRIM-IN-MOTION DETECTOR

MECHANICAL ELECTRICAL

PITCH TRIM SYSTEM BLOCK DIAGRAM Figure 5-4

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PITCH TRIM SELECTOR SWITCH The PITCH TRIM selector switch, located on the pedestal trim control panel, provides primary and secondary mode selection for the aircraft trim systems. The switch has three positions: PRI, OFF, and SEC. When the switch is set to PRI, a ground path is provided for the primary pitch trim system control circuits and trim changes are accomplished through the control wheel trim switches. When the switch is set to SEC, a ground path is provided for the secondary pitch trim system control circuits and trim changes are accomplished through the pedestal NOSE DN-OFF-NOSE UP switch. When the switch is set to the OFF position, both pitch trim electrical control circuits are isolated from the aircraft electrical system. The Mach trim system is inoperative with the PITCH TRIM selector switch in the OFF or SEC positions. The autopilot is inoperative with the PITCH TRIM selector switch in the OFF position. CONTROL WHEEL TRIM SWITCHES PITCH FUNCTION Each control wheel trim switch is a dual-function (trim and trim arming) switch which controls primary pitch trim and roll trim. One switch is located on the outboard horn of each control wheel. Each switch has four positions: LWD, RWD, NOSE UP, and NOSE DN. The trim arming button on top of the switch must be depressed for trim motion to occur. With the PITCH TRIM selector switch in the PRI position, actuation of either switch to NOSE UP or NOSE DN will signal the primary motor in the horizontal stabilizer pitch trim actuator to move the stabilizer in the appropriate direction. Actuation of the pilots switch will override actuation of the copilots switch. Actuation of either switch to any of the four positions (LWD, RWD, NOSE UP, or NOSE DN) will disengage the autopilot. Actuation of either switch to NOSE UP or NOSE DN will resynchronize the Mach trim computer. NOSE DN-OFF-NOSE UP SWITCH The NOSE DN-OFF-NOSE UP switch, located on the pedestal trim control panel, controls secondary pitch trim. The switch is spring loaded to the center (OFF) position. With the PITCH TRIM selector switch in the SEC position, actuation of the NOSE DN-OFF-NOSE UP switch to NOSE DN or NOSE UP will signal the secondary motor of the horizontal stabilizer pitch trim actuator to move the stabilizer in the appropriate direction. Actuation of secondary pitch trim will disengage the autopilot. The Mach trim system is inoperative when using secondary pitch trim. With the PITCH TRIM selector switch in the PRI or OFF position, this switch has no effect. 5-12
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CONTROL WHEEL MASTER SWITCHES PITCH TRIM FUNCTION A Control Wheel Master Switch (MSW) is located beneath the control wheel trim switch on the outboard horn of each control wheel. In addition to the switches other functions, either Control Wheel Master Switch (MSW), when depressed, will inhibit primary or secondary pitch trim. If the Control Wheel Master Switch is used to inhibit primary pitch trim, primary pitch trim cannot be reactivated until the Control Wheel Master Switch is released and the trim input is removed. Therefore, during the preflight check of the primary pitch trim system, it is necessary to release the control wheel trim switch as well as the Control Wheel Master Switch (MSW) to reset the system. Secondary pitch trim, however, will be inhibited only as long as the Control Wheel Master Switch (MSW) is held. PITCH TRIM LIGHT An amber PITCH TRIM annunciator light, located on the glareshield annunciator panel, is installed to alert the crew of primary pitch trim system malfunctions during flight. Additionally, the PITCH TRIM light will illuminate whenever either Control Wheel Master Switch (MSW) is depressed. T. O. TRIM LIGHT An amber T. O. TRIM annunciator light, located on the glareshield annunciator panel, is installed to alert the crew that the PITCH TRIM indicator pointer is not within the T. O. segment when the aircraft is on the ground. The light will be extinguished whenever the indicator pointer is set within the T. O. segment. The light is disabled during flight operations. SYSTEM TEST SWITCH TRIM OVERSPEED FUNCTION The rotary-type SYSTEM TEST switch, located on the pilots instrument panel, is used to test the trim speed monitor. Prior to beginning the trim speed monitor test, the pitch trim must be set on the high trim rate (N UP) side of the index on the PITCH TRIM indicator. The monitor test is initiated by rotating the SYSTEM TEST switch to TRIM OVSP, initiating primary pitch trim through either control wheel trim switch, and then depressing the switch PRESS TEST button. When the PRESS TEST button is depressed, a false low trim rate range horizontal stabilizer position signal is applied to the trim speed monitor. With the trim speed monitor in the low trim rate watch mode, running the primary pitch trim at the high trim rate will cause the trim speed monitor to illuminate the PITCH TRIM light.

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PITCH TRIM INDICATOR The ELEV indicator, located on the EIS Flight Display Page, provides the crew with visual indication of the horizontal stabilizer trim position. There is a pointer on the right side of the vertical scale with a digital readout of horizontal stabilizer trim position. The position pointer is green when on the ground and the pointer is within the T.O. segment. The position pointer is white when on the ground and the pointer is not within the T.O. segment. In air mode, the ELEV pointer is always green, regardless of position. The indicator range is from 1 to 12 of horizontal stabilizer travel. ND and NU markings indicate the direction of trim travel for airplane nose down and airplane nose up respectively. The T.O. (takeoff) segment from 5.7 to 8.75 is marked with a thick line. A triangle at the 6.5 position, separates the high and low trim rate ranges. At pitch trim settings on the NU side of the triangle, the trim speed controller will be in the high trim rate (low airspeed) mode. At pitch trim settings on the ND side of the triangle, the trim speed controller will be in the low trim rate (high airspeed) mode. The pitch trim indicator receives horizontal stabilizer position inputs from a potentiometer installed in the horizontal stabilizer pitch trim actuator. The system operates on 28 VDC supplied through the TRIM-FLAP-SPOILER INDICATOR circuit breaker on the copilots circuit breaker panel. TRIM-IN-MOTION AUDIO CLICKER A trim-in-motion audio clicker system is installed to alert the crew of horizontal stabilizer movement. The system will annunciate continuous movement of the horizontal stabilizer by producing a series of audible clicks through the headsets and cockpit speakers. The system consists of a potentiometer in the horizontal stabilizer pitch trim actuator, a trim-in-motion detector box and associated aircraft wiring. As the horizontal stabilizer actuator drives the stabilizer, the output signal from the potentiometer is altered. The change in potentiometer signal is sensed by the detector box. After approximately 1/4 second of continuous stabilizer movement, the detector box will produce the speaker and headset clicks. The trim-in-motion audio clicker system is wired through the flap position switches and will not sound if the flaps are lowered beyond 3. The trim-in-motion audible clicker may or may not sound during autopilot trim due to the duration of the trim inputs. Power for system operation is 28 VDC supplied from the WARN LTS circuit breakers on the pilots and copilots circuit breaker panels through the warning lights control box. These circuit breakers are powered during EMER BUS mode.

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ROLL TRIM Roll trim is accomplished by positioning the aileron trim tab on the inboard trailing edge of the left aileron through actuation of the roll trim actuator. The roll trim actuator is an electrically-operated, rotary-type actuator connected to the aileron trim tab by a push-pull rod. The system is controlled through the pilots and copilots control wheel trim switches. The AIL indicator, located on the EIS Flight Display Page, provides the crew with visual indication of the roll trim setting. The roll trim system operates on 28 VDC supplied through the ROLL TRIM circuit breaker on the pilots circuit breaker panel. CONTROL WHEEL TRIM SWITCHES ROLL FUNCTION Each control wheel trim switch is a dual-function (trim and trim arming) switch which controls roll trim and primary pitch trim. One switch is located on the outboard horn of each control wheel. Each switch has four positions: LWD, RWD, NOSE UP, and NOSE DN. The arming button on top of the switch must be depressed for trim motion to occur. Actuation of either control wheel trim switch to LWD or RWD will signal the aileron trim tab actuator to move the tab as required to lower the appropriate wing. Actuation of the pilots switch will override actuation of the copilots switch. Actuation of either switch to any of the four positions (LWD, RWD, NOSE-UP, or NOSE-DN) will disengage the autopilot if the trim arming button is depressed. CONTROL WHEEL MASTER SWITCHES ROLL TRIM A Control Wheel Master Switch (MSW) is located beneath the control wheel trim switch on the outboard horn of each control wheel. In addition to the switches other functions, either Control Wheel Master Switch (MSW), when depressed, will inhibit roll trim. The roll trim is inhibited only as long as the Control Wheel Master Switch (MSW) is held. AILERON TRIM INDICATOR Aileron trim information is provided by the AIL indication on the EIS Flight Display Page. Two semi-circular scales and pointers present the trim tab position in terms of left wing down and right wing down. The scale markings represent increments of trim tab travel. The aileron trim indicator receives inputs from a potentiometer in the roll trim actuator. The system operates on 28 VDC supplied through the TRIM-FLAPSPOILER INDICATOR circuit breaker on the copilots circuit breaker panel.
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YAW TRIM Yaw trim is accomplished by positioning the rudder trim tab on the lower trailing edge of the rudder through actuation of the yaw trim actuator. The yaw trim actuator is an electrically-operated, rotary-type actuator connected to the rudder trim tab by two push-pull rods. Yaw trim is pilot controlled through the RUDDER TRIM switch on the pedestal. The RUDDER indicator, located on the EIS Flight Display Page, provides the crew with visual indication of the yaw trim setting. The yaw trim system operates on 28 VDC supplied through the YAW TRIM circuit breaker on the pilots circuit breaker panel. RUDDER TRIM SWITCH Yaw trim is pilot controlled through the RUDDER TRIM switch located on the pedestal trim control panel. The switch has three positions: NOSE LEFT, OFF, and NOSE RIGHT. The switch knob is split and both halves must be rotated simultaneously to initiate yaw trim motion. When the switch is released, both halves will return to the center OFF position. Actuation of the RUDDER TRIM switch to NOSE LEFT or NOSE RIGHT will signal the yaw trim actuator to move the rudder trim tab in the appropriate direction. RUDDER TRIM INDICATOR Rudder trim tab position indication is provided by the RUDDER indication on the EIS Flight Display Page. A horizontal scale and pointer indicates the direction (L or R) of yaw trim. The scale markings represent increments of rudder trim tab travel. The rudder trim indicator receives inputs from a potentiometer in the rudder trim actuator. The system operates on 28 VDC supplied through the TRIM-FLAP-SPOILER INDICATOR circuit breaker on the copilots circuit breaker panel. The RUDDER TRIM indicator will be operative during the EMER BUS mode. CONTROL WHEEL MASTER SWITCHES YAW TRIM A Control Wheel Master Switch (MSW) is located beneath the control wheel trim switch on the outboard horn of each control wheel. In addition to the switches other functions, either Control Wheel Master Switch (MSW), when depressed, will inhibit yaw trim. The yaw trim is inhibited only as long as the Control Wheel Master Switch (MSW) is held.

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WARNING SYSTEMS
STALL WARNING SYSTEM A stall warning system is installed to provide the crew with visual and tactile warning of an impending stall. The major components of the stall warning system consist of the following: left and right stall vanes on the forward fuselage, a two-channel computer-amplifier, flap position switches for each flap, two 18,100-foot altitude switches, a stick shaker for each crew position, an angle-of-attack indicator for each crew position, L and R STALL warning lights, and associated aircraft wiring. The flap position switches provide bias information to the computer-amplifier which will decrease stall warning speeds as the flaps go from 0 to 40. Above approximately 18,100 feet pressure altitude, the altitude switches bias the system to increase stall warning speeds approximately 15 knots. The stick shakers are eccentric weights driven by an electric motor and actuation is evidenced by a high-frequency vibration of the control columns. The left and right systems are completely independent and utilize separate electronics, stall vanes, altitude switches, shaker motors, and flap switches. The stall warning system operates on 28 VDC supplied through the L and R STALL WARN circuit breakers on the pilots and copilots circuit breaker panels respectively. The stick shaker and STALL warning light circuits are wired through the squat switches; therefore, the stick shaker and STALL warning lights are deactivated when the squat switches are in the ground mode. The stick shaker and STALL warning lights will be deactivated for 3 to 5 seconds after lift-off. The angle-of-attack indicators remain active both on the ground and inflight, however the angle of attack displays are not available on the PFD while on the ground. The stall warning systems may be tested on the ground using the rotarytype systems test switch, located on the center instrument panel. During flight, the stall warning vanes align with the local airstream and transducers produce a voltage proportional to airplane angle of attack. The transducer signals are transmitted to the appropriate computeramplifier channel along with flap position information from the flap position switches and altitude information from the altitude switches. The angle-of-attack indicator pointers will enter the amber segment, the L and R STALL lights will illuminate and flash, and the stick shakers will actuate when the angle of attack increases to an angle corresponding to an airspeed at least 7% above the stall speed published in the Airplane Flight Manual.

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LEFT FLAP POSITION SWITCH

ADC 1

ADC 2

RIGHT FLAP POSITION SWITCH

LEFT STALL WARNING VANE

RIGHT STALL WARNING VANE

L E F T

C H A N N E L

STALL WARNING COMPUTER

R I G H T

C H A N N E L

SHAKER

LEFT SQUAT SWITCH

SHAKER

PILOTS ANGLE-OF-ATTACK INDICATOR

COPILOTS ANGLE-OF-ATTACK INDICATOR

RIGHT SQUAT SWITCH

LEFT SHAKER

AUTOPILOT/ SHAKER INTERFACE

RIGHT SHAKER

LEFT STALL ANNUNCIATOR

RIGHT STALL ANNUNCIATOR

L PFD

R PFD

STALL WARNING SYSTEM BLOCK DIAGRAM Figure 5-5

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ANGLE-OF-ATTACK INDICATORS The angle-of-attack indicators, located on the pilots and copilots instrument panels, translate signals from the stall warning computer-amplifier into a visual indication of angle-of-attack. These indicators present normalized angle-of-attack information for all flap settings on a scale from 1.0 (max lift) to 0 (zero lift). The left stall warning system utilizes the pilots angle-of-attack indicator and the right stall warning system utilizes the copilots angle-of-attack indicator. Each indicator face is divided into three segments as follows: green -safe, amber -caution/shaker, and red -warning. Low-Speed Awareness Cues The PFD Airspeed displays receive information from the stall warning computer and display the following types of Low-Speed Awareness cues: Impending Stall Speed reference cue (ISS) which is represented by the top of the red bar on the airspeed cue and .82 AOA. Reference Approach Speed cue (RAS) which is represented by the 1.3Vs green line on the airspeed cue and .6 AOA. Airspeed Trend Vector on the airspeed cue.
WARNING

Low-Speed awareness cues serve as an approximation of stall speed and do not replace the actual stall warning system.

STALL WARNING LIGHTS The red L and R STALL warning lights, located in the glareshield annunciator panel, are installed to indicate impending stall or a system malfunction. During flight operations, the lights will illuminate and flash when the shaker is actuated. The lights are pulsed at the same frequency and duration as the shakers; therefore, the flash frequency will increase as the angle-of-attack increases from initial shaker actuation. At or just prior to the angle-of-attack pointer entering the red segment, the flash frequency is sufficient to cause the lights to appear steady. SYSTEM TEST SWITCH STALL WARNING FUNCTION The rotary-type system test switch, located on the center instrument panel, is used to test the left and right stall warning systems. Each system is individually tested through the L STALL and R STALL positions of the system test switch. The test is initiated by rotating the system test switch to L or R STALL (as applicable) and then depressing the switch PRESS TEST button. When the test sequence is initiated, the correPM-133

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sponding angle-of-attack indicator pointer will begin to sweep from the green segment toward the red segment. As the pointer passes the green-amber margin, the stick-shaker will actuate, Master WARN lights will illuminate, and the applicable STALL light will begin to flash. Shaker actuation is made evident by high frequency vibration of the control column. OVERSPEED WARNING SYSTEM The overspeed warning system provides an audible overspeed warning in the event aircraft speed exceeds a Mach or airspeed limit. The overspeed warning horn is activated by the air data computers when the position of the airspeed and the maximum allowable airspeed coincide. 28 VDC for system circuits is supplied through the WARN LTS circuit breakers on the pilots and copilots circuit breaker panels and will be powered during emergency bus operations. The overspeed warning horn will sound under any of the following conditions: 1. Airspeed exceeds VMO. 2. Mach exceeds MMO. SYSTEM TEST SWITCH OVERSPEED WARNING FUNCTION The rotary-type system test switch, located on the center instrument panel, is used to test the overspeed warning system. The test sequence is initiated by rotating the system test switch to OVSP and then depressing the switch PRESS TEST button. The overspeed warning will sound three times, each separated by a brief pause. The third warning horn will continue until the TEST button is released. TAKEOFF WARNING SYSTEM The takeoff configuration monitor system consists of a monitor box, throttle quadrant switch and various system switches (provide the input signals to the monitor box). The system is active when the aircraft is on the ground (right squat switch in ground mode). A takeoff monitor aural warning will sound during ground operations when the right thrust lever is advanced to the MCR position or above and one or more of the following conditions exist: 1. 2. 3. 4. 5. 5-20 Thrust reverser unlocked or deployed. Flaps not set for takeoff. Spoilers not retracted. Pitch trim not in a safe position for takeoff. Parking brake not released.
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ENHANCED GROUND PROXIMITY WARNING SYSTEM WITH WINDSHEAR DETECTION (EGPWS/WS) The Enhanced Ground Proximity Warning System with Windshear Detection (EGPWS/WS) provides the pilot with aural and visual warning of potentially dangerous flight paths relative to ground and windshear. The system automatically and continuously monitors the airplanes flight path with respect to terrain when the aircraft is below 2450 feet radio altitude (altitude AGL). If the airplanes projected flight path would imminently result in terrain impact, the system issues appropriate visual and voice warnings. Warnings are issued for excessive sink rate, excessive terrain closure rate, descent after takeoff or missed approach, proximity to terrain with flaps and/or gear up, descent below glideslope, and descent below decision height (DH) or minimum descent altitude (MDA). The system computes windshear and alerts the crew of windshear of sufficient magnitude to be hazardous to the aircraft. Windshear alerts are given for increasing headwind/decreasing tailwind and/or updraft. Windshear warnings are given for decreasing headwind/increasing tailwind and/or down-draft. The system consists of the EGPWS/WS computer, annunciators on the AFDs, INHIBIT/OVRD switches on the instrument panel for G/S INH, TERR, and TAES FLAP, and associated aircraft wiring. Voice warnings are made through the cockpit speakers and the headphones. Voice warnings generated by the EGPWS will have priority over voice warnings generated by the TCAS. The system receives inputs from the either air data computer, either AHRS, both stall warning vanes, radio altimeter, both nav receivers, nose gear down and locked switch, and the left flap 8, 20 and 40 switch. The system operates on 28 VDC supplied through the EGPWS circuit breaker on the pilots circuit breaker panel. Refer to the Collins Pro Line 21 Avionics System with IFIS for the Learjet 60XR Operators Guide (Collins P/N 523-0807841, edition 1, dated April 24, 2006 or later applicable version) and the Learjet 60XR FAA Approved Airplane Flight Manual (FM-133) for additional information.

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TRAFFIC ALERT AND COLLISION AVOIDANCE SYSTEM (TCAS) The Traffic Alert and Collision Avoidance System (TCAS) provides the pilot with aural and visual indications of potentially dangerous flight paths relative to other aircraft in the vicinity. The system uses the transponder to interrogate other transponder-equipped aircraft and determine their bearing, range, and altitude. With this information, the TCAS processor can generate advisories to prevent or correct traffic conflicts. The TCAS consists of a receiver/transmitter/processor, two directional antennas, and associated aircraft wiring. Power for system operation is 28 VDC supplied through the TCAS circuit breaker on the copilots circuit breaker panel. Advisories are issued to the crew via the aircraft audio system and integrated displays (PFDs and MFDs). Aural advisories generated by the ground proximity/windshear warning system (if installed) will have priority over aural advisories generated by the TCAS. Refer to the Collins Pro Line 21 Avionics System with IFIS for the Learjet 60XR Operators Guide (Collins P/N 523-0807841, edition 1, dated April 24, 2006 or later applicable version) and the Learjet 60XR FAA Approved Airplane Flight Manual (FM-133) for additional information.

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AIR DATA SYSTEMS

Air data for instruments and equipment requiring flight environment air data for display or operation is provided by two separate air data systems. The dual primary air data system consists of the primary pitot-static system, two air data computers, a total temperature probe and reversionary mode switch/annunciators. A separate standby pitotstatic system is installed to provide flight environment air data for display on the standby instruments. PRIMARY PITOT-STATIC SYSTEM Pitot and static pressure for the air data computers and other using systems is obtained from the two primary pitot-static probes. One probe is located on each side of the nose compartment. Each probe contains a pitot (impact pressure) port and two static pressure ports. The probes also contain electrical heating elements controlled by the L and R PITOT HEAT switches. Four drain valves, located near the nose gear doors, are installed at the system low spots to drain moisture from the system. The pilots pitot system is completely independent of the copilots pitot system and utilizes the left pitot-static probe as the source of pitot pressure. The copilots system utilizes the right pitot-static probe to obtain pitot pressure. The pilots and copilots systems each utilize a separate static source on each of the probes. A solenoid-operated shutoff valve is installed in each static source line to ensure accurate static pressure in the event one probe becomes clogged or unreliable. The shutoff valves are controlled through the STATIC SOURCE switch on the pilots switch panel and operate on 28 VDC supplied through the STATIC SOURCE circuit breaker on the copilots circuit breaker panel. The pilots pitot source supplies pitot pressure for ADC 1 air data computer. The copilots pitot source supplies pitot pressure for ADC 2 air data computer. Each pitot-static probe contains two static sources. One static source on each probe is interconnected with a static source on the opposite probe to supply static pressure to ADC 1. The other static source on each probe is interconnected with a static source on the opposite probe to supply static pressure to ADC 2. In the event a static source becomes clogged or unreliable, the affected pitot-static probes static sources can be isolated, allowing all equipment to be operated from static sources on the opposite probe.

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ADC 2 Pilot Pitot/Static Probe

Copilot Pitot/Static Probe ADC 1

PITOT STATIC 1 SHUTOFF VALVES STATIC 2 SHUTOFF VALVES

PITOT STATIC 1

STATIC 2

ADC 1 PITOT PRESSURE ADC 1 STATIC PRESSURE

ADC 2 PITOT PRESSURE ADC 2 STATIC PRESSURE

PRIMARY PITOT-STATIC SYSTEM SCHEMATIC Figure 5-6 STATIC SOURCE SWITCH The STATIC SOURCE switch controls solenoid-operated shutoff valves, in the static plumbing, to ensure accurate static pressure sensing in the event one of the pitot-static probes become inoperable or unreliable. The STATIC SOURCE switch, located on the pilots switch panel, has three positions: L, BOTH, and R. When the switch is in the BOTH position all four shutoff valves are de-energized open and static pressure for the air data instruments and equipment is available from static ports in both pitot-static probes. Normally, the switch is in the BOTH position for all operations. When the switch is set to L or R, the shutoff valves for the opposite pitot-static probe are energized closed, and static pressure will be supplied by the selected pitot-static probe only.

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STANDBY PITOT-STATIC SYSTEM The standby pitot-static system is independent of the primary system and supplies pitot-static pressure to the standby Mach/airspeed indicator and the standby altimeter. The standby pitot-static probe is located on the right side of the nose compartment. This probe contains a pitot (impact pressure) port and two static pressure ports. The standby pitot-static probe contains an electrical heating element controlled by the R PITOT HEAT switch. Two drain valves, located near the nose gear doors, are installed at the system low spot to drain moisture from the system.

STANDBY MACH/AIRSPEED INDICATOR

STANDBY PITOT-STATIC SYSTEM SCHEMATIC Figure 5-7 AIR DATA COMPUTERS

Two digital air data computers receive pitot and static pressures from the primary pitot-static system and temperature data from the total temperature probe for computation of the flight environment. The computed results of the sensor inputs are converted to electrical signals and transmitted to the associated cockpit displays. Additional outputs from the air data computers are transmitted to the integrated avionics processor system (IAPS) for distribution to other systems that require air data for proper operation. The following table summarizes the various outputs under normal conditions. The air data computers operate on 28 VDC through the ADC circuit breakers on the pilots and copilots circuit breaker panels. ADC 1 and ADC 2 are operative during EMER BUS operations.

;; ;;
Standby Pitot/Static Probe STANDBY STATIC PRESSURE STANDBY PITOT PRESSURE
PITOT STATIC 1 STATIC 2

STANDBY ALTIMETER

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ADC 1 Pilots Instruments (EFIS) Airspeed/Mach Altitude/Vertical speed Altitude Alert L Stall Warning Gear Warning Overspeed profile AHS (TAS) L FCS Mach Trim ATC 1 (encoded altitude) FMS 1 FADEC SAT TAS Cabin Pressurization

ADC 2 Copilots Instruments (EFIS) Airspeed/Mach Altitude/Vertical Speed Altitude Alert R Stall Warning Gear Warning Overspeed profile AHS (TAS) R FCS Mach Trim ATC 2(encoded altitude) FMS 2 FADEC SAT TAS Cabin Pressurization

Refer to the Collins Pro Line 21 Avionics System with IFIS for the Learjet 60XR Operators Guide (Collins P/N 523-0807841, edition 1, dated April 24, 2006 or later applicable version) and the Learjet 60XR FAA Approved Airplane Flight Manual (FM-133) for additional operational information and a complete description of the air data system interfaces and instruments. ADC/ADC TRANSFER SWITCH The ADC/ADC transfer switches on the EFIS CONTROL panels are used to select the ADC source for display on the on-side display. Onside ADC is the normal selection indicated by a green annunciation of the switch. Reversionary (cross-side) selection is indicated by an amber annunciation on the switch. ADC reversion on either side will also cause the following annunciations: ADC # (# = system supplying air data [1 or 2]) on both PFDs.

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ATTITUDE HEADING SYSTEM

Aircraft avionics displays and equipment requiring attitude or heading information are supplied that information from the dual, independent Collins Attitude Heading Systems (AHS 1 and AHS 2). Each system consists of an attitude heading computer with internal compensator, a magnetic flux sensor in the associated wing tip, two HEADING control switches, and associated aircraft wiring. The attitude heading computer is composed of inertial instruments, electronics, interface hardware, processing and memory circuits to provide attitude and heading information to other aircraft systems. One magnetic slaving unit is located in each wing tip and is used to sense the earths magnetic field. The HEADING SLAVE-FREE switch allows the crew to select either Free or Slaved Magnetic Heading mode. The system has two operating modes, normal and basic. During normal operation, a true airspeed input is supplied by the air data system to improve accuracy. If the true airspeed input is lost, the system will continue to operate in the basic mode. AHS operation is automatic and both systems will initialize when battery power is applied to the aircraft. During the nominal 70 second alignment, the system determines its orientation with the local vertical and magnetic North and performs a series of self-test and calibration functions. The AHS 1 and 2 systems are powered by 28 VDC AHS 1 and AHS 2 circuit breakers on the pilots and copilots circuit breaker panels. Both AHS 1 and AHS 2 will be powered during EMER BUS operations. In the event of a power loss, approximately 11 minutes of back-up power (28 VDC) will be supplied to AHS 1 and AHS 2 by EMER BAT 2. This feature makes it unnecessary to reinitialize the system should a momentary power loss be experienced. Should one of the systems fail, the functions of the failed system may be assumed by the remaining system using the AHS/AHS reversionary mode. Attitude/heading data is provided for the following using systems: EFIS Displays attitude and heading displays Flight Management System heading data Flight Control System attitude, heading and acceleration data Fuel Quantity System attitude, heading and acceleration data TCAS System attitude, heading and acceleration data EGPWS System attitude, heading and acceleration data Weather Radar pitch and roll data for antenna stabilization Lightning Detection System (if installed) pitch and roll data for heading stabilization

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HEADING CONTROL SWITCHES The HEADING control switches, located in the AVIONICS group on the pilots and copilots switch panels, are used to control the heading output of the associated AHS. The switches on the pilots side control AHS 1 while the switches on the copilots side control AHS 2. The SLAVE-FREE switch provides slaving mode selection for the associated AHS heading output. When the switch is set to SLAVE, the associated AHS heading output will be referenced to its magnetic slaving unit and the associated compass cards will reflect this slaved alignment. When the switch is set to FREE, the associated AHS heading output will not be referenced to its magnetic slaving unit. The SLAVE L-R switch provides for manual slewing of the associated compass cards. Small heading splits can usually be cleared by cycling the SLAVE-FREE switch to FREE and then back to SLAVE while the aircraft is in straight and level, unaccelerated flight. AHS/AHS REVERSIONARY MODE The AHS/AHS switches on the EFIS CONTROL panels are used to select the attitude heading system for the respective EFIS display and flight director. On-side AHS is the normal selection indicated by green annunciation on the switch. Reversionary (cross-side) selection is indicated by an amber annunciation on the switch. AHS reversion on either side will also cause the following annunciations: ATT # (# = system supplying attitude data [1 or 2]) on both PFDs and MAG # (# = system supplying heading data [1 or 2]) above each compass card.

MAGNETIC COMPASS A direct-reading magnetic compass is installed on the windshield center post. The liquid filled compass contains a horizontal drum dial and a lubber line. The drum has a 360 scale graduated in 5 increments. Numerical markings appear at 30 intervals except that 0, 90, 180 and 270 are labeled N, E, S, and W respectively. N/S and E/W compensator screws are located under the cover plate. A compass steering correction card is located near the compass.

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ELECTRONIC STANDBY INSTRUMENT SYSTEM (ESIS) The ESIS is located on the center instrument panel. This indicator is a L3 Communications Avionics Systems solid state, graphic display standby indicator system. The system consists of a self-sensing single box unit and is powered by 28 VDC supplied by EMER BAT 1. This single LCD indicator provides the pilot and copilot with pitch and roll, slip/skid indications, altitude, airspeed, Mach number, dual baro-set, and VMO/MMO indications. Localizer and glideslope deviation is provided if NAV 1 is tuned to an ILS. It is designed to mimic the primary EFIS system. For a more detailed description of this system, refer to the current L3 Communications Avionics Systems Electronic Standby Instrument System Pilots Guide (P/N TP-560).

ELECTRONIC STANDBY INSTRUMENT SYSTEM Figure 5-8

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ELECTRONIC FLIGHT INSTRUMENT SYSTEM (EFIS)

The EFIS is a Collins 4-panel composite color display system. The system consists of a primary flight display (PFD) and a multifunction display (MFD) on each pilots instrument pane, heading, speed, altitude panel (HSA), one course heading panel (CHP), two cursor control panels (CCP), two EFIS Control panels (ECP), and two Control Display Units (CDU). Cooling for the PFDs and MFDs is provided by fans integral to each display unit and an avionics cooling fan. Failure of the avionics cooling fan is indicated by illumination of the white INSTR FAN annunciator on the glareshield annunciator panel. The system is powered by 28 VDC from the following circuit breakers: PFD 1 & 2, MFD 1 & 2, and EFIS CONTROL 1 & 2. The EFIS is used to display airplane altitude, airspeed/Mach, vertical speed, air temperature, attitude data, navigational data, flight director commands, mode annunciators, weather, checklists, warnings, and diagnostic messages. This description covers the system in a general manner and is intended for familiarization only. Refer to the Collins Pro Line 21 Avionics System with IFIS for the Learjet 60XR Operators Guide (Collins P/N 5230807841, edition 1, dated April 24, 2006 or later applicable version), the Learjet 60XR FAA Approved Airplane Flight Manual (FM-133), and the Collins FMS 5000 Operators Guide for additional operational information and a complete description of the EFIS interfaces and instruments.

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PRIMARY FLIGHT DISPLAY (PFD) The PFD on each side displays attitude, primary air data and lateral navigation display elements. The PFDs provide the following information:
Pitch and Roll Attitude Mode Annunciations Vertical Speed Baro Corrected Altitude Altitude Preselect Temperature DME Data Marker Beacon TCAS RAs Flight Director Commands Heading, Course & Bearing Airspeed Radio Altitude Reporting Altitude, MDA or DH Set VNAV Deviation Warning Annunciations & Flags Glideslope and Localizer Deviation

MULTIFUNCTION DISPLAY (MFD) The MFD on each side brings together numerous displays to show a map-like presentation of the airplanes horizontal navigation situation. The MFDs provide the following information:
Heading Source Annunciations Course Deviation Selected Heading Bearing Pointer Wind DME Data Warning Annunciations & Flags VNAV Deviation Selected Course/Desired Track Weather Radar

In addition, the MFD is capable of displaying the following information:

Checklists Maintenance Diagnostics Avionics Status Sensor Status Approach Charts Airways Flight Plan Map Nearby Nav Aids, Airports, etc. Performance and Progress TCAS TFC Display Graphical Weather Geographical Data

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EFIS CONTROL PANEL An EFIS control panel is installed on both the pilots and copilots instrument panel. Each panel controls its respective EFIS. Each switch is an alternate action switch. On-side selection is indicated by a green annunciation and cross-side or reversionary mode selection is indicated by an amber annunciation.
AHS AHS

This switch selects the attitude heading system for the respective EFIS display, flight director and other systems requiring attitude or heading data. The switch is used to recover attitude and heading data if the on-side AHS fails. Whenever cross-side AHS data is selected, the pitch, roll, and heading comparators will be disabled, and all equipment normally sourced by the on-side AHS will be sourced by the cross-side AHS.

ADC ADC

This switch selects the air data system for the respective EFIS display, flight director and other systems requiring air data. The switch is used to recover air data if the onside ADC fails. This reversionary mode selection switch is used to recover data on the MFD. When actuated in the REV mode, the adjacent PFD functions will be assumed by the MFD. This would be used if a PFD tube fails. This reversionary mode selection switch is used to recover data on the PFD. When actuated in the REV mode, the adjacent MFD functions will be assumed by the PFD. This would be used if a MFD tube fails. This switch is only located on the copilots panel. The switch displays the engine indication display on the copilots PFD.

REV MFD

REV PFD

ON ENG

DISPLAY CONTROL PANEL (DCP) Two DCPs (one on the pilots instrument panel and one on the copilots instrument panel) provide PFD and MFD display control. The DCP is used to select control menus on the PFD and to adjust the display range on the PFD and MFD. The DCP provides dedicated controls for the Air Data System and Weather Radar System. For a detailed description of the DCP refer to the Collins Pro Line 21 Avionics System with IFIS for the Learjet 60XR Operators Guide (Collins P/N 523-0807841, edition 1, dated April 24, 2006 or later applicable version). 5-32
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HEADING, SPEED, ALTITUDE PANEL (HSA) The single HSA is located below the FCP on the glareshield and provides for heading selection, speed/vertical speed selection, as well as altitude pre-select inputs. The HDG knob is used to change the selected heading indicated by the heading bug on both PFDs and MFDs simultaneously. Pressing the inset PUSH SYNC switch in the center of the HDG knob will synchronize the heading bug on all of the large displays to the current airplane heading as read under the lubber line on the pilots PFD. COURSE (CRS) CONTROL KNOBS Two course (CRS) knobs are located on the pedestal forward of the cursor control panels. they are used to change the active selected course on the on-side PFD/MFD when VOR is the active NAV sensor. When FMS is the active NAV sensor and in the SEL CRS mode, these knobs change the course angle to the TO waypoint. Pressing the center PUSH DIRECT switch on either CRS knob will zero the course deviation and establish a course directly to the active NAV sensor. CURSOR CONTROL PANEL (CCP) Two Cursor Control Panels, located on the pedestal forward of the CDUs, operate MFD menus and select display formats. The Cursor Control Panel (CCP) is used to select and control the optional Integrated Flight Information System (IFIS) functions by MFD on-screen menus and to adjust the orientation of the optional FMS 3D Map. Dedicated controls are provided for chart selection, a joystick for panning and zooming charts, quick MFD format access keys, and MFD menu controls. Three quick access keys are used to store and then recall display format configurations for the MFD. CONTROL DISPLAY UNIT (CDU) Dual Collins CDUs are installed in the pedestal to control the PFDs, MFDs, and FMS. The CDUs also provide an additional method (other than the RTUs) for tuning NAV/COM radios and entering transponder codes. The CDU uses a combination of displayed menus, line-keys, full alphanumeric keypad, control knobs and dedicated control keys. In most cases, the CDUs can be operated simultaneously or independently. For instance, the pilot may change or edit the flight plan while the copilot manages NAV/COM frequencies. Neither CDU has priority over the other. If both CDUs tune the same radio, the most recent change is the one that will be used. The pilot should note that there are some functions that cannot be done simultaneously. 5-33

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COMMUNICATIONS
VHF COMMUNICATIONS Dual VHF communications transceivers are installed to provide AM voice communication capability. The VHF COMMs are capable of tuning 8.33Khz steps. The transceivers are SELCAL compatible with analog audio interfaces. Tuning is accomplished via the Radio Tuning Units (RTU) or via the Control Display Units (CDU). The CDUs have similar radio management functions but differ on RTU failure procedures. (Refer to AFM for detailed malfunction information). The design of the system is such that all radio management functions are channeled through the RTUs, regardless of their origin. The center instrument panel RTU normally tunes COMM 1 and the pedestal RTU normally tunes COMM 2. If an RTU fails, the remaining RTU is capable of tuning both COMM 1 and COMM 2. Power for the system is 28 VDC supplied through the COMM 1 and COMM 2 circuit breakers on the pilots and copilots circuit breaker panels. COMM 1 is powered during EMER BUS operations. The above information is presented in a general manner and is intended for familiarization only. For a detailed description and operation of the VHF communications system refer to the Collins Pro Line 21 Avionics System with IFIS for the Learjet 60XR Operators Guide (Collins P/N 523-0807841, edition 1, dated April 24, 2006 or later applicable version). HF COMMUNICATIONS An HF (high frequency) communication system is installed to provide long range communication capability. The system operates on any 0.1 kHz frequency between 2.0 and 29.9999 MHz. The system consists of a control/display unit (pedestal), a remote power amplifier and antenna coupler, remote receiver/transmitter, and antenna. System power is 28 VDC supplied through current limiters and controlled by a remote control circuit breaker. The remote control circuit breaker is controlled by the HF 1 circuit breaker on the pilots circuit breaker panel. The HF receiver is SELCAL compatible. The above information is presented in a general manner and is intended for familiarization only. For a detailed description and operation of the HF communications system refer to the appropriate HF operators manual. 5-34
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SELCAL The SELCAL system permits the selective calling of individual aircraft over normal radio communications circuits linking the ground station with the aircraft. The SELCAL system is integrated into the communication systems to relieve the flight crew from continuously monitoring communications frequencies during flights of extended duration. The system consists of a decoder unit and the SELCAL indication on the EIS Flight Page. The system is powered by 28 VDC through the SELCAL circuit breaker on the pilots circuit breaker panel. When a call is received, an indication in the SELCAL area of the flight display will show and an intermittent aural tone will sound. When the mic button is momentarily depressed, the aural tone will cease. The SELCAL system can be tested by turning the system test switch to the SELCAL position and pushing the knob to start the test. The SELCAL alert tone will sound and each of the SELCAL enabled radios indicators (VHF 1, VHF 2, HF 1, HF 2) will be displayed. AUDIO CONTROL SYSTEM The audio control system is used to select the desired audio inputs for broadcast through the speakers or headphones. The audio control system is also used to select the desired transmitter to which microphone inputs will be directed. A separate audio control system is provided for pilot and copilot. Each system consists of an audio amplifier and audio control panel. The audio control system operates on 28 VDC supplied through the L and R Audio circuit breakers on the pilots and copilots circuit breaker panels respectively. The audio control systems will operate during EMER BUS mode. AUDIO CONTROL PANEL An audio control panel is installed at the outboard end of the pilots and copilots instrument panels. Each panel provides the controls necessary to direct audio signals and adjust volume levels. Each panel is used in conjunction with the on-side microphone, headphone and cockpit speaker.

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INPH

PHONE

SPKR

PASS MIC SELECT VHF 2 1 1 HF 2 PASS

VHF1

MASTER VOLUME VHF2 HF1 HF2

NAV1

NAV2

ADF1

ADF2

BOTH

DME1

DME2

MKR1

IDENT MKR2 HI

V O I C E

NORM MIC OXY MIC

LO

AUDIO CONTROL PANEL Figure 5-9 MIC SELECT SWITCH The MIC SELECT Switch is a multi-position rotary-type switch labeled VHF 1, VHF 2, HF 1, and HF 2, and PASS. This switch provides the proper microphone audio inputs for the respective functions. VHF 1, VHF 2, HF 1 and HF 2 Positions When any of these positions are selected, microphone inputs are provided for the respective transceiver. Microphone must be keyed to transmit. PASS Position When this position is selected, the pilot or copilot, utilizing this function, may speak to the passengers through the passenger speaker. Microphone must be keyed to transmit. PASS should not be selected on both audio control panels simultaneously as degradation of the volume level may result. NORM MIC/OXY MIC SWITCH NORM MIC Position When the switch is in this position, voice transmissions are accomplished with the headset microphone or handheld microphone. OXY MIC Position When the switch is in this position, voice transmissions are accomplished with the oxygen mask microphone. Both cockpit speakers, phone and interphone function (see VOLUME CONTROLS) will be active. The microphone must be keyed to transmit to the passengers or via a communications radio. 5-36
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VOLUME CONTROLS The volume controls consist of four MASTER VOLUME (INPH, PHONE, SPKR and PASS) controls. Each control is rotated to regulate the overall volume level to the applicable output device. The INPH and SPKR controls have a push-ON/push-OFF function. In the ON position, the control knob will protrude further than in the OFF position. Also, the controls will illuminate in the ON position. INPH Volume This control regulates the volume level of the crew interphone system. The interphone employs a voice-activated hot microphone. SPKR Volume This control regulates the volume level of the on-side cockpit speaker audio. PHONE Volume This control regulates the volume level of the on-side headphone audio. PASS Volume This control regulates the volume level of the passenger speaker audio. RADIO MONITOR SWITCHES Each control has a push-ON/push-OFF function and a volume control which is rotated to regulate the volume level of individual audio inputs. In the ON position, the control knob will protrude further than in the OFF position. Also, the control will illuminate in the ON position. Radio monitor switches on the audio control panel are labeled and perform the following functions: VHF 1 and VHF 2 Switches When in the ON position, provide audio from the VHF 1 and VHF 2 transceivers respectively. HF 1 and HF 2 Switches When in the ON position, provide audio from the HF 1 and HF 2 (if installed) transceiver respectively. NAV 1 and NAV 2 Switches When in the ON position, provide audio from the NAV 1 and NAV 2 receivers respectively. ADF 1 and ADF 2 Switches When in the ON position, provide audio from the ADF 1 and ADF 2 (if installed) receiver. DME 1 and DME 2 Switches When in the ON position, provide audio from the DME 1 and DME 2 receivers respectively. MKR 1 and MKR 2 Switches When in the ON position, provide audio from the MKR 1 and MKR 2 receivers respectively.

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BOTH/VOICE/IDENT SWITCH This switch controls the audio filtering for the NAV and ADF receivers. BOTH Position When the switch is in this position, both the station identifier and voice transmissions will be heard. The BOTH position is the normal position. VOICE Position When the switch is in this position, only the voice transmissions will be heard. IDENT Position When the switch is in this position, only the station identifier will be heard. MARKER BEACON HI/LO SWITCH The HI/LO switch on the pilots audio control panel controls the #1 marker beacon receiver and the HI/LO switch on the copilots audio control panel controls the #2 marker beacon receiver. HI Position When the switch is in this position, the marker beacon receiver sensitivity is increased. LO Position When the switch is in this position, the marker beacon receiver sensitivity is decreased. AUDIO CONTROL FLIGHT OPERATION 1. Applicable MASTER VOLUME Controls Set to the ON position and rotate to a comfortable listening level. 2. Applicable Radio Monitor Switches Set to the ON position and rotate to a comfortable listening level. The VHF 1 and VHF 2 volume controls do not affect sidetone levels. The HF 1 and HF 2 volume controls will affect the sidetone level since the audio and sidetone utilize a common line from the transceivers. 3. MIC SELECT Switch Rotate to desired position.

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CABIN BRIEFING SYSTEM One of the optional systems (Airshow 410 or Airshow 4000 Cabin Video Information System) may be installed. Either system is designed to give passengers a recorded briefing for various phases of flight. AIRSHOW CABIN VIDEO INFORMATION SYSTEM An optional Airshow Cabin Video Information System may be installed. The system includes a serial mouse, video monitor and a flight deck controller. The Airshow system is selected for display from the cabin control switch panel located on the inboard upper side of the forward left-hand cabinet or from the master control switch panel, located in the cabin armrest. The passenger briefing feature consists of three messages, (TAKEOFF, LANDING and TURBULENCE). To access these briefings, scroll through the menu and select Time To Destination (TTD), select SEL BRF from the sub-menu if using the optional flight deck controller, or by selecting SEL BRF from the INFO MENU if using the serial mouse. After selecting the desired briefing, the message will be heard through the overhead cabin speakers and in each passenger headphone. The briefing will override any other audio source except for paging. To cancel a briefing scroll to CANCEL or reselect the same briefing.

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NAVIGATION
The navigation system includes the radios and controls used for VOR/ ILS navigation, DME, ADF navigation, ATC transponder operation and radio altitude measurement. Tuning of all these functions except the radio altimeter is accomplished via the Radio Tuning Units (RTU) on the center instrument panel or via the Control Display Units (CDU) in the pedestal. The design of the system though is such that all navigation radio management functions are channeled through the RTUs regardless of their origin. The left RTU normally tunes NAV 1, ADF 1, ATC 1, etc. and the right RTU normally tunes the #2 radios. If an RTU fails, the remaining RTU is capable of tuning both #1 and #2 systems. Power for the RTUs is 28 VDC supplied through the RTU 1 and RTU 2 circuit breakers on the pilots and copilots circuit breaker panels. RTU 1 will be operative during EMER BUS operations. The radio altimeter will be discussed later. Navigation information is presented in a general manner and is intended for familiarization only. For a detailed description and operation of the navigation system refer to the Collins Pro Line 21 Avionics System with IFIS for the Learjet 60XR Operators Guide (Collins P/N 5230807841, edition 1, dated April 24, 2006 or later applicable version). VHF NAVIGATION Dual VHF navigation receivers and controls are installed to provide the crew with VOR bearing, VOR audio, localizer deviation, glideslope deviation, marker beacon passage identification and marker beacon audio. The receivers are capable of tuning the entire navigation and glideslope frequency range. The NAV 1 and NAV 2 circuit breakers on the pilots and copilots circuit breaker panels supply 28 VDC to power the VHF navigation receivers. NAV 1 will be powered during EMER BUS operations.

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MARKER BEACON DISPLAY Marker beacon passage, displayed on the PFD, is indicated by a cyan box with OM for outer marker, a yellow box with MM for middle marker, or white box with IM for inner marker. All marker beacon annunciations flash when they are displayed. DISTANCE MEASURING EQUIPMENT (DME) Dual DME transceivers are installed to provide distance, time-tostation, ground speed, and station ident information for use by other units in the avionics system. Each DME can track as many as three stations at the same time. Channel 1 of each DME is paired with a VOR frequency and tuned via the RTU or CDU for direct display by the crew. Channels 2 and 3 are used by the Flight Management System for multisensor navigation and are automatically tuned by the FMS. DME Hold can be activated on the RTU to hold the current DME frequency and allow the navigation receiver to be independently retuned. 28 VDC power for the DME receivers is supplied by the DME 1 and DME 2 circuit breakers on the pilots and copilots circuit breaker panels. AUTOMATIC DIRECTION FINDING (ADF) An ADF system is installed to provide aural reception of signals from a selected ground station and indicate relative bearing to that station. The system operates in the normal ADF frequency range and is tuned via the RTU or CDU for direct display by the crew. Functions such as BFO ON or OFF are controlled by the RTU. The ADF 1 circuit breaker is located on the pilots circuit breaker panel to supply 28 VDC to the ADF receiver. ADF 1 will be operative during EMER BUS operations. ATC TRANSPONDERS Two ATC transponders are installed to provide identification (ModeA), altitude (Mode-C), and select (Mode-S) reporting for the ATC radar beacon system. The traditional 4096 Mode-A codes are available and altitude reporting is selectable. The Mode-S data link feature is used for TCAS operation. The TDRs are equipped for Mode-S and Flight ID which includes Enhanced Surveillance. Code selection may be accomplished from the RTU or CDU. Other functions such as STBY mode, ID (ident) and turning off and on altitude reporting are controlled by the RTU. Power for the transponders is 28 VDC supplied by the ATC 1 and ATC 2 circuit breakers on the pilots and copilots circuit breaker panels. Identification and altitude reporting will be provided by ATC 1 during EMER BUS operations.
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RADIO ALTIMETER A radio altimeter is installed to give the pilot and copilot a direct radio height measurement from 0 to 2,500 feet AGL. The radio altitude is automatically displayed in green digits on both PFDs when the radio altitude is below 2,500 feet AGL. Changes in altitude are displayed by the radio altimeter in 50-foot increments when the altitude is above 1,000 and in 10-foot increments when the altitude is below 1,000 feet. No tuning is required and there are no operating controls that affect the radio altimeter. During a radio altimeter test, selected from the RTU, a fixed value of 50 feet will be displayed on both PFDs. The RADIO ALT circuit breaker on the pilots circuit breaker panel supplies 28 VDC power to the radio altimeter.

FLIGHT CONTROL SYSTEM (FCS)

The FCS provides 3-axis autopilot/yaw damper, dual flight director, rudder boost and automatic pitch trim functions. The FCS contains two flight control computers and three primary servos and is controlled by a glareshield-mounted Flight Control Panel (FCP). Each side of the dual system (pilot and copilot) operates the same and both work together to drive the servos and the pitch trim system. The following information is presented in a general manner and is intended for familiarization only. Refer to the Collins Pro Line 21 Avionics System with IFIS for the Learjet 60XR Operators Guide (Collins P/ N 523-0807841, edition 1, dated April 24, 2006 or later and the FAA Approved Airplane Flight Manual for further information on the Flight Control System. AUTOPILOT/FLIGHT DIRECTOR SYSTEM The autopilot/flight director system provides automatic flight control and guidance for climb, cruise, descent and approach. The system provides dual channel flight guidance, and either channel can be coupled to the autopilot. Mode selection and annunciation for each flight guidance channel and engage controls for autopilot and yaw damper are provided through the glareshield-mounted FCP. Mode and system status annunciation is also provided on the appropriate cockpit displays. The system provides dual-channel flight guidance in the pitch and roll axis. Dual-channel yaw axis outputs are used for yaw damping. Pitch and roll axis change, when commanded by the autopilot, is affected through autopilot elevator and aileron servos. The autopilot also provides pitch trim commands to the secondary trim system motor of the horizontal stabilizer pitch trim actuator. Autopilot pitch authority is 5-42
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limited to 10 nose down and 20 nose up and roll authority is limited to 32 for lateral command, 27 bank for heading or course capture, and 15 for course tracking and roll rate is limited to 5 per second. Pilot inputs to the autopilot/flight director system are accomplished through the FCP, control wheel switches and the course heading panels. The pilots flight guidance system operates on 28 VDC supplied through the AP 1 and the FD 1 circuit breakers on the pilots circuit breaker panel. The copilots flight guidance system operates on 28 VDC supplied through the AP 2 and the FD 2 circuit breakers on the copilots circuit breaker panel. The autopilot system operates on 28 VDC supplied through the AP 1 and AP 2 circuit breakers. The autopilot/flight guidance system is active whenever power is on the aircraft and both avionics master switches are on. The autopilot may be coupled to either the pilots or copilots flight guidance channel using the AP XFR and AP ENG switches on the FCP. When the autopilot is engaged, the associated or on-side PFD will display steering information from the on-side flight guidance channel. Whenever the autopilot is engaged, the on-side PFD command bars will display the steering command and the on-side instruments may be used to monitor autopilot performance. When the autopilot is not engaged, the PFD attitude display can be used to manually fly the airplane in response to steering commands from the on-side flight guidance channel (provided a vertical or lateral mode is selected). FLIGHT CONTROL PANEL (FCP) Autopilot/flight guidance mode selection and autopilot engagement functions are accomplished through the glareshield-mounted FCP. The controller contains three groupings of buttons. The center grouping provides the autopilot selection and engage buttons as well as autopilot status annunciators. The grouping on the left provides mode selection for the pilots flight guidance channel and the grouping on the right provides mode selection for the copilots flight guidance channel. SELF-TEST The system initiates a self-test sequence when the system is powered up (LEFT and RIGHT AVIONICS MASTER Switches ON). If the selftest sequence is not successfully completed, the autopilot will not engage and an FD flag will be displayed on the PFDs.

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AUTOPILOT ENGAGE FUNCTIONS AP XFR The AP XFR is a momentary push-on/push-off button which is used to select the flight guidance channel to be coupled with the autopilot. A green triangle, on the FCP, will illuminate and point to the side which will couple to the autopilot, when engaged. AP The AP button is a momentary push-on/push-off button which is used to couple the autopilot to the selected flight guidance channel. If the autopilot passed the power-up self-test, the autopilot will engage and the green light will illuminate and a green AP or AP (as appropriate) annunciation will appear on the primary flight displays. An electrical interlock in the FCP automatically engages the yaw damper whenever the autopilot is engaged. Thereafter, the yaw damper may be independently disengaged. YD The YD button is a momentary push-on/push-off button which is used to engage the yaw damper. When engaged, the indicator above the YD button illuminates. The yaw damper can be disengaged by depressing the YD button a second time or by depressing the Control Wheel Master (MSW) switch. TURB The TURB button is a momentary push-on/push-off button which is used to select the autopilot turbulence mode. When TURB is selected, the autopilot will provide softer responses in the pitch and roll axis for flying through turbulence. TURB is not available during flight director only operation and is locked out in APPR mode. AUTOPILOT/FLIGHT GUIDANCE MODE SELECTION All mode selection buttons are the momentary push-on/push-off type. A light above the mode selector button will illuminate if all conditions for the mode are satisfied. Any selected mode can be cancelled by selecting an incompatible mode, depressing the mode selector button a second time, or depressing the FD CLEAR button. Mode selection and operation is identical for the left and right channels. Attitude Hold When the flight director is operating and no vertical mode is selected, pitch attitude hold will automatically be active. When the flight director is operating and no lateral mode is selected, roll attitude hold will automatically be active. Although active, the roll attitude hold cannot be entered without the autopilot first being engaged in the roll mode and then disconnected. These modes are used to maintain a reference pitch and bank angle. The reference angles may be established by manually flying the aircraft to the desired pitch and bank 5-44
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angle and depressing the SYNC button (on the control wheel). When the SYNC button is released, the flight director will generate commands to maintain the existing pitch and roll attitude. If the bank angle is less than 5, the flight director will command heading hold. The reference values may be changed using the vertical and lateral command function of the control wheel trim switches. HDG (heading) When HDG is selected, autopilot/flight director commands are generated to maneuver the airplane as necessary to fly a heading by position of the heading bug on the PFD. 1/2 BANK When 1/2 BANK is selected, the flight director reduces its maximum roll attitude command to one-half of the normal limit. 1/ 2 BANK may be engaged in conjunction with any lateral mode except Approach. 1/2 BANK is automatically selected when the airplanes pressure altitude is at or above 41,500 feet. 1/2 BANK automatically clears when the airplane descends below this altitude. NAV (navigation) The NAV mode provides flight director commands to capture and track the navigational course set on the PFD. APPR (approach) The APPR mode provides flight director commands to capture and track the navigational course set on the PFD with approach accuracy. During ILS approaches, commands to capture and track the glideslope will be generated after the localizer has been captured. ALT (altitude hold) The ALT mode provides flight director commands to track the indicated altitude present at the time of mode engagement. VS (vertical speed hold) The VS mode provides flight director commands to maintain the vertical speed selected. In the absence of a preselected vertical speed, flight director commands will be generated to maintain the vertical speed present at the time of engagement. VNAV (vertical navigation) VNAV allows the pilot to program the FMS to provide vertical guidance in descent planning or to meet altitude crossing restrictions. FLC (Flight Level Change) FLC provides commands to acquire and track an IAS or Mach reference airspeed while taking into account the need to climb or descend to bring the aircraft to the active reference altitude (Preselect Altitude or Flight Plan Target Altitude).

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Go-Around The go-around (GA) mode is a flight director only mode and is selected by depressing the GO-AROUND button in the left thrust lever knob. When GA is selected, the autopilot will disengage, selected lateral and vertical modes will be cancelled, and a fixed 9 nose-up, heading hold steering command will be presented on the PFD. FCP ANNUNCIATORS The FCP incorporates annunciators to provide the status of the rudder boost and automatic pitch trim systems and an annunciator to indicate which flight director is selected. TRIM (pitch trim) The red TRIM annunciator will illuminate when an automatic pitch trim failure has been detected. The autopilot cannot be engaged while the red TRIM light is illuminated. If already engaged and the light illuminated, the autopilot will remain engaged until manually disengaged. RB (rudder boost) Two separate RB annunciators, one green and one amber, are installed. Illumination of the green RB annunciator indicates the rudder boost system is active. Illumination of the amber RB annunciator indicates a rudder boost system failure or that the RUDDER BOOST switch is off. LEFT & RIGHT ARROWS (autopilot transfer arrows) The left or right green arrow illuminates to indicate which flight director is selected. When the autopilot is engaged, the arrow points to the coupled flight director. If the autopilot s disengaged, a white arrow points to the selected flight director.
NOTE

In ILS approach and go-around modes, both FGCs are used independently to provide steering commands to their on-side PFD and both left and right arrows will illuminate.

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CONTROL WHEEL MASTER SWITCHES AUTOPILOT FUNCTION The Control Wheel Master Switches (MSW), located on the outboard horn of the pilots and copilots control wheels, may be used to disengage the autopilot. Depressing either the pilots or copilots MSW will disengage the autopilot. When the autopilot disengages, the green light above the AP button on the FCP will extinguish and the autopilot disengage tone will sound. For a pilot initiated autopilot disconnect, the AP annunciation flashes amber for 5 seconds, then self-clears. If a monitored disengagement occurs, the autopilot disconnect is a red AP and red transfer arrow for 5 seconds, then steady and will clear when the AP or MSW button is pressed, or the autopilot is re-engaged. When the autopilot is disengaged using the MSW, the flight director will remain active and will display steering information from the flight guidance computer, if a vertical or lateral mode is selected. PITCH TRIM SELECTOR SWITCH AUTOPILOT FUNCTION When the autopilot is engaged, the autopilot maintains aircraft pitch trim through the secondary motor of the horizontal stabilizer pitch trim actuator if the PITCH TRIM selector switch on the pedestal is in the PRI or SEC position. The autopilot will not engage or will disengage if the PITCH TRIM selector switch is moved to the OFF position. CONTROL WHEEL TRIM SWITCHES AUTOPILOT/FLIGHT DIRECTOR FUNCTION When either Control Wheel Trim switch (arming button depressed) is moved to any of the four positions (LWD, RWD, NOSE UP or NOSE DN), an aircraft trim input is made and the autopilot will disengage. If the arming button is not depressed, the on-side switch may be used to input lateral commands (LWD and RWD) and vertical commands (NOSE UP and NOSE DN) to the autopilot. Using this feature causes active modes (except GS) in the applicable axis to disengage and revert to the attitude hold mode. Armed modes are not effected. The control wheel trim switch has no effect on the flight director.

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NOSE DN-OFF-NOSE UP SWITCH AUTOPILOT FUNCTION The NOSE DN-OFF-NOSE UP switch, located on the pedestal trim control panel, may be used to disengage the autopilot or to make trim adjustments with the autopilot pitch and roll axes inhibited. With the PITCH TRIM selector switch in the SEC position, actuation of secondary pitch trim through the NOSE DN-OFF-NOSE UP switch will disengage the autopilot, extinguish the green light above the AP button, and sound the autopilot disengage tone. When the autopilot is disengaged through the NOSE DN-OFF-NOSE UP switch, the flight director will remain active and will display steering information from the flight guidance computer. SYNC SWITCHES The SYNC switches in the control wheels are normally used with the on-side flight director to change a vertical mode (except GS, LVL CHG and ALTS) reference values without reselecting the mode. The only lateral mode in which SYNC switches are active is roll attitude hold (ROLL). FD CLEAR SWITCHES Depressing the FD CLEAR switch in either control wheel will remove the command bars and cancel any selected vertical or lateral mode from the on-side flight director. Depressing the FD CLEAR if the autopilot is coupled to the on-side flight director will remove the command bars and must be depressed to redisplay the command bars. YAW DAMPER SYSTEM The yaw damper augments aircraft stability by opposing uncommanded motion about the yaw axis and provides turn coordination. The yaw damper is provided by the yaw axis of the autopilot/flight guidance system. The yaw damper operates independent of the autopilot. YAW DAMPER CONTROL The yaw damper button and annunciator are located on the FCP. The yaw damper engages when the autopilot is engaged, or by depressing the YD button on the FCP. When the yaw damper is engaged, the green light above the YD button will be illuminated. If the yaw damper is already engaged, depressing the YD button will disengage the yaw damper.

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CONTROL WHEEL MASTER SWITCHES - YAW DAMPER FUNCTION The Control Wheel Master Switches (MSW), located on the outboard horn of the pilots and copilots control wheels, may be used to disengage the yaw damper. Depressing either the pilots or copilots Control Wheel Master Switch (MSW) will disengage the yaw damper. When the yaw damper is disengaged through pilot action, the yaw damper disengage tone will sound, and an amber YD annunciator on the EFIS will flash for 5 seconds, then extinguish. The green indicator light above the YD button on the FCP will also extinguish. RUDDER BOOST SYSTEM The rudder boost system is installed to provide reduced rudder pedal force, increased directional control effectiveness and improved takeoff performance. With the rudder boost on, minimum control speedground (VMCG), takeoff speeds and distances are all lower. Rudder boost is a function of the autopilot. In addition to the autopilot, the system consists of a yaw force interface box, force sensors, flap position switch, RUDDER BOOST Switch, and associated aircraft wiring. The yaw damper servo provides the boost to assist the pilot in moving the rudder in the desired direction. The rudder boost system is supplied 28 VDC through the FD 1 circuit breaker on the pilots circuit breaker panel. Normally the RUDDER BOOST Switch, on the pilots switch panel, is left on at all times. With flaps lowered more than 3, applying approximately 50 pounds of force to either rudder pedal will cause the yaw servo to automatically engage and apply force to the rudder in the same direction as the pilot. As pilot input force is increased, the servo force will also increase up to the maximum yaw servo force. When the rudder boost engages, the green RB annunciator, on the FCP, illuminates to indicate rudder boost is active. If the yaw damper is on when the rudder boost engages, the system will make a smooth transition from yaw damper to rudder boost. A failure of the system is indicated by illumination of the amber RB annunciator on the FCP. Self-test of the system is initiated during system power-up. RUDDER BOOST SWITCH Arming of the rudder boost system is controlled by the RUDDER BOOST Switch located on the pilots switch panel. When the switch is set to ON, the system will be armed. Setting the switch to OFF will disarm the system and the amber RB annunciator, on the FCP, will illuminate.
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FLIGHT MANAGEMENT SYSTEM (FMS)

The Learjet 60XR is equipped with a dual Collins FMS-5000 flight management system. The FMS is an integrated navigation management system that provides the pilot with centralized control for the airplanes navigation sensors, computer based flight planning, and fuel management. FMS capabilities include VFR/IFR RNAV operation, direct-to functions, VNAV, approach, and fuel management. The system also receives true airspeed and altitude information from the air data computer and fuel flow data from the fuel flow sensors. The FMS provides worldwide point-to-point and great circle navigation. The FMS uses sensor data from GPS, VOR/DME navaids, and air data systems, along with the active flight plan and its own database information. The sensor data is used by the FMS to determine the present position, direction, and speed. GPS can be used as the primary means of navigation in oceanic and remote areas if a pre-departure verification of GPS navigation availability over the entire planned route is performed before each flight. The FMS contains a subscription data base which has the appropriate navaids and airports. The FMS scans for DME signals which, according to its data base present position, are expected to be received. The outputs of the two DMEs, three channels for each DME allowing up to six DMEs to be scanned. As navigation station signals are received, their identifiers are decoded for station verification. If at least three properly positioned DME signals are received, the airplane position can be determined. When less than three DMEs are available, then VOR radial and DME distance is used. The fuel management function of the FMS allows the pilot to plan fuel requirements while on the ground. Pilot-supplied data and inputs from the airplanes fuel flow sensors give the FMS the necessary information to calculate and display significant real-time fuel management information throughout the flight. For a detailed description and operation of the FMS, Refer to the Collins Pro Line 21 Avionics System with IFIS for the Learjet 60XR Operators Guide (Collins P/N 523-0807841, edition 1, dated April 24, 2006 or later applicable version).

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WEATHER RADAR
A weather radar system is installed to give the pilot a pictorial representation of the safest possible flight path during adverse weather conditions. The single unit X-Band weather radar provides data from atmospheric moisture and ground features. The resulting radar pictures may be displayed on any of the AFDs. Terrain mapping is possible with the radar, and with practice, the pilot will be able to identify coastlines, large rivers and lakes, mountainous areas and cities. As the radar system becomes more familiar, it may be used to verify position, track, ground speed, altitude and attitude as well as for weather avoidance. The radar can be operated in a split mode or sync mode. In the split mode, both pilots have the option of placing the radar in different mode and range settings on alternate sweeps. This gives the appearance of two independent radars. In the sync mode, both sides show the same radar display. Some installations include the capability to detect precipitation related turbulence. Control of the weather radar is accomplished from the pilots and copilots Display Control Panels (DCP) and the line select keys on the PFD/ MFD. Primary stabilization for the radar is obtained from the left Attitude Heading System (AHS). If the left AHS fails, stabilization is automatically obtained from the right AHS. For a detailed description and operation of the weather radar system refer to the Collins Pro Line 21 Avionics System with IFIS for the Learjet 60XR Operators Guide (Collins P/N 523-0807841, edition 1, dated April 24, 2006 or later applicable version).

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MISCELLANEOUS
COCKPIT VOICE RECORDER (CVR) A cockpit voice recording system is installed to record all cockpit voice, radio communication, aural annunciation, and aural navigation signals for the last 30 minutes of operation. System components consist of a TEST switch, an ERASE switch, a pass indicator, a fail indicator, a headphone jack, a microphone and a voice recorder unit. The ERASE switch, TEST switch, pass indicator, fail indicator and HEADPHONE jack are installed on the copilots switch panel. The area microphone, installed in the center of the instrument panel, picks up all cockpit audio. The microphone incorporates electronic background noise suppression. The voice recorder unit converts audio input to digital format. The digital format audio is stored in a crash-survivable solid-state memory. The digital storage unit has a maximum recording interval of 30 minutes. After 30 minutes of continuous recording, the recorder automatically starts recording over the previously stored audio data. The CVR TEST switch is pressed and held for at least 2 seconds to initiate the automatic self-test. During the self-test the PASS and FAIL annunciators will flash alternately for approximately 15 seconds. At the end of a successful self-test the PASS annunciator will illuminate steady for approximately 10 seconds. If the self-test fails, the FAIL annunciator will come on either steady or flashing. The pattern of flashes is an indication to maintenance personnel as to the nature of the failure. Squat switch, parking brake and anti-skid ON interlock switching control the bulk erasure function. Voice recorder system power is 28 VDC supplied through the CVR circuit breaker on the copilots circuit breaker panel. The CVR will be operative during EMER BUS operations. There is an optional 120 minute capacity CVR available. The only difference between the standard and optional CVR is the recording time.

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FLIGHT DATA RECORDER (FDR) (Optional) The flight data recorder will record pertinent flight profile data. A white FDR FAIL annunciator is installed in the warning lights annunciator panel to annunciate system malfunctions. The system is powered by 28 VDC through the FDR circuit breaker on the pilots circuit breaker panel and is powered by the Emergency Bus. The Flight Data Recorder is recording whenever power is applied to the aircraft. There are no controls or switches associated with the FDR and operation is completely automatic. Upon power application to the aircraft, the system will perform a selftest. When the BATTERY switches are set to on, the FDR FAIL annunciator will illuminate briefly, then extinguish. The test will continue for another 60 seconds. The light should not come back on during the test. CLOCKS Each instrument panel is equipped with a multi-function chronometer to display GMT, local time (LT), flight time (FT), and elapsed time (ET). Power for the chronometers is 28 VDC supplied through the L and R CLOCK circuit breakers on the pilots and copilots circuit breaker panels. The SEL button selects what is to be displayed and the CTL button controls what is being displayed. Pressing SEL sequentially selects GMT, LT, FT or ET for display. FT starts counting when the squat switches transition to the air mode and stops counting when they transition back to ground mode. The CTL button resets FT back to zero when held down for 3 seconds. ET is started and reset when the CTL button is pushed momentarily. Depressing the SEL and CTL buttons simultaneously enters the set mode and GMT or LT can be set. The CTL button is then pressed to increment the flashing digit to the desired value. Pressing the SEL button then enters that value and toggles to the next digit to be set.

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HOURMETER AIRCRAFT An hourmeter is installed to measure aircraft accumulated time. The hourmeter is located behind the carpeted access panel on the step behind the cockpit or in the copilots circuit breaker panel. It is wired to the right squat switch and will measure accumulated time as soon as the aircraft lifts off. The hourmeter receives 28 VDC through the HOUR METER circuit breaker on the copilots circuit breaker panel. EMERGENCY LOCATOR TRANSMITTER The Emergency Locator Transmitter (ELT) transmits distress signals assisting rescue personnel in locating a downed aircraft. The ELT consists of a transmitter, antenna, and remote switch. TRANSMITTER AND ANTENNA The transmitter and antenna are installed in the vertical stabilizer. Power for the transmitter is provided by an internal battery. The transmitter will automatically activate under emergency conditions or may be manually activated using the cockpit switch. REMOTE SWITCH A remote switch is installed in the cockpit to allow manual activation and resetting of the ELT transmitter without accessing the transmitter itself.

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LIGHTNING DETECTION SYSTEM (LDS) (OPTIONAL) The LDS, also called the L3 Communications Stormscope Series II Weather Mapping System, is a passive system; that is, it does not transmit energy. Instead, the LDS detects electrical discharges (lightning) through passive reception of their energy and displays them as a moving map on an adaptive flight display (AFD). Since the LDS does not plot water droplets like regular weather radar, it is not subject to attenuation. The LDS will automatically position thunderstorm information relative to aircraft heading. The LDS system includes an antenna, LDS processor, and associated aircraft wiring. Operator control inputs include inputs from the line select keys on the AFDs, DCPs, reversionary switching, and other remote-mounted controls. Data collection and distribution is provided by the IAPS. The LDS processor calculates lightning azimuth and range, and generates lightning symbology, operating, and fault message for display on the AFDs. The LDS uses built-in test equipment to verify proper operation and to generate fault messages for display on the AFDs. Displayed electromagnetic discharges associated with thunderstorm activity appear as lighting bolts on the display. The lighting bolts are color coded to identify different levels of intensity. The lighting bolts are removed from the screen after 2 minutes. When changing from one range display to another, no loss of data will occur since electrical discharge information is acquired and stored on all ranges simultaneously. The LDS should never be used to attempt thunderstorm penetration. Thunderstorm avoidance must not be solely predicated upon the use of the LDS.

CAUTION

The LDS receives 28 VDC through the STORMSCOPE circuit breaker on the copilots circuit breaker panel. The preceding information on the LDS is meant as a familiarization only to the LDS. For a detailed description and operation of the LDS refer to the Collins Pro Line 21 Avionics System with IFIS for the Learjet 60XR Operators Guide (Collins P/N 523-0807841, edition 1, dated April 24, 2006 or later applicable version).

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XM Satellite Weather (Optional) The optional XM Satellite data link receiver is part of the optional Integrated Flight Information System (IFIS). The XM Satellite Receiver provides a constant stream of graphical and textual weather data from the XM Satellite Radio weather service to the FSU. The Graphical Weather (GWX-3000) format provides the ability to show GWX images or reports on the MFD. The GWX images are provided by Baron Services by a satellite Information Service Provider (XM Satellite Weather Service). The GWX image can be a textual weather report/forecast or a graphical image. Weather reports include Significant Meteorological (SIGMET) and Airman Meteorological (AIRMET) advisories and Aviation Routine Weather Reports (METAR). Weather forecasts are Terminal Area Forecast (TAF). Observation images include NEXRAD and Echo Tops. Universal Weather (Optional) The GWX format provides the ability to show one GWX image at a time on the MFD. New GWX images are requested by the pilot with controls on the CDU. Refer to the Rockwell Collins Corporate Datalink System CMU-4000/RIU-40X0 Operator Guide, Collins Part Number 5230790499, for detailed information on using the CDU to request and view GWX images. The GWX images are uplinked with VHF datalink system from the Information Service Provider (Universal Weather). A list of saved and available GWX images shows on the MFD when requested by the pilot. Using controls on the CCP, the pilot selects the desired GWX image to show on the MFD. The GWX image can be a forecast or an observation image. Forecast images include WINDS ALOFT, ICING, and TURBULENCE. Observation images include NEXRAD, TOPS/MOVEMENT, and Weather (WX) DEPICTION. Each datalinked GWX image is paired with a corresponding geopolitical background image. A title/time banner shows for each GWX image.

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NAVIGATION SOURCE The Navigation (NAV) or Map source legend shows along the left side of the MFD when the compass Arc, Rose, FMS Plan Map, PPOS map, or 3D Map is the active format on the MFD. The NAV source field is four lines of text that show along the left side of the MFD when the active NAV source and the Map source are the same FMS and the compass Arc or Rose is the active format on the MFD. The active NAV source is selected from the NAV SOURCE menu on the PFD. When the MFD Plan Map, PPOS Map, or 3D Map are selected for display on the PFD, the NAV source. The Map source is set to FMS1 or FMS2 with the MAP menu on the MFD. 3D MAP FORMAT (Optional) The 3D Map is an optional, advanced FMS feature which provides lateral, vertical, and performance-predicted flight plan information in a single, three-dimensional (3D) format on the MFD. The map data is a combination of what would typically be presented as two separate map formats a vertical profile and a plan map. The 3D Map has an adjustable viewing orientation which is used to customize the viewing angle. The 3D Map allows predicted flight path views that are referenced from the ground (such as a vertical profile view), referenced directly over a map center position (such as a Plan Map view), or referenced from an intermediate point in between. E-CHARTS (Optional) The E-Chart format provides the ability to show an electronic version of a conventional paper instrument chart on the MFD. The E-Charts are linked automatically by the FMS when a flight plan is entered and can also be selected manually by the pilot. The available charts are listed on the Chart Main Index. Controls for chart selection are on the CCP. When aircraft position data is available, a moving aircraft symbol shows on E-Charts that are geographic-referenced. A non-geographicreferenced chart has a magenta aircraft symbol with a circle and slash on the top right hand corner of the chart.

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JEPPESEN CHART DISPLAY (Optional) The selected Jeppesen E-Chart shows when selected by the pilot. Controls on the CCP are used to select a chart for display, pan around the chart, zoom in or out on the chart, and change the orientation of the chart. A moving aircraft symbol shows on the chart when the chart is geographically-referenced, as determined by the Jeppesen database and the aircraft position is within the geographically-referenced part of the chart. NOTAMS (Optional) The Chart NOTAMS menu shows the chart NOTAMS available for the selected airport. The page is broken into two fields, the NOTAM summary and NOTAM details. When more than one NOTAM is available for the selected airport, the selected NOTAM and total number of NOTAMs shows in the summary field. The selected NOTAM readout is also a data entry field that allows the user to select another NOTAM for viewing. The NOTAM type, effectivity, begin date, and end date show in the summery field. The NOTAM text as defined in the Jeppesen charts database shows in the details field.

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SECTION VI ANTI-ICE & ENVIRONMENTAL

TABLE OF CONTENTS Bleed Air Supply ....................................................................................... 6-1 BLEED AIR Switches............................................................................. 6-1 Bleed Air Supply Schematic (Figure 6-1) ........................................... 6-2 CABIN AIR Light................................................................................... 6-3 BLEED AIR Warning Light................................................................... 6-3 Anti-Ice Systems ........................................................................................ 6-4 Rosemount Ice Detector System ......................................................... 6-4 Ice Detect Lights..................................................................................... 6-5 Wing Inspection Light........................................................................... 6-5 Engine and Nacelle Inlet Anti-Ice ....................................................... 6-6 NAC HEAT Switches ....................................................................... 6-6 NAC HT Lights ................................................................................. 6-7 Wing Anti-Ice ........................................................................................ 6-7 Wing Anti-Ice System (Figure 6-2) ................................................. 6-8 STAB WING HEAT Switch Wing Heat Function .................... 6-9 WING TEMP Indicator .................................................................... 6-9 WING HT Light .............................................................................. 6-10 Horizontal Stabilizer Anti-Ice ............................................................ 6-10 STAB WING HEAT Switch Stabilizer Heat Function........... 6-10 STAB HT Light ................................................................................ 6-11 Stabilizer Heat Self Test.................................................................. 6-11 Windshield Anti-Ice............................................................................. 6-12 WSHLD HEAT Switch ................................................................... 6-12 Windshield Anti-Ice System (Figure 6-3) .................................... 6-13 WSHLD HT Light ........................................................................... 6-14 WSHLD OV HT Light .................................................................... 6-14 Windshield Defog ................................................................................ 6-15 Windshield Defog System (Figure 6-4)........................................ 6-15 WSHLD DEFOG Switch ................................................................ 6-16 L and R WS DEFOG Annunciators .............................................. 6-16 Windshield Anti-Ice Alcohol System ........................................... 6-17 WSHLD ALC Switch...................................................................... 6-17 ALC LOW Caution Light............................................................... 6-17 Alcohol Anti-Ice System (Figure 6-5)........................................... 6-18 Pitot-Static and Stall Warning Anti-Ice............................................. 6-19 PITOT HEAT Switches................................................................... 6-19 PITOT HT Light .............................................................................. 6-20 L, R and STBY Pitot Heat Lights .................................................. 6-20 PM-133 VI-1

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TABLE OF CONTENTS (Cont) Oxygen System........................................................................................ 6-21 Oxygen Storage and Pressure Regulation ....................................... 6-21 Oxygen System Schematic Single Forward Cylinder (Figure 6-6).......................................... 6-22 Single Aft Cylinder (Figure 6-6A) ................................................ 6-23 Dual Cylinders (Figure 6-6B) ........................................................ 6-24 Oxygen Pressure Indicator................................................................. 6-25 Oxygen System Cockpit Controls ..................................................... 6-26 Passenger Masks ................................................................................. 6-27 Passenger Mask (Figure 6-7) ......................................................... 6-27 Crew Masks Scott ATO.................................................................. 6-28 Crew Mask Scott ATO (Figure 6-8) ......................................... 6-28 Pressurization System ............................................................................ 6-29 Pressurization System Schematic (Figure 6-10) .............................. 6-30 Normal Pressurization........................................................................ 6-31 Emergency Pressurization.................................................................. 6-32 Pressurization Controls and Indicators............................................ 6-33 MODE Switch ................................................................................. 6-33 MAN ALT Control ......................................................................... 6-33 EMER DEPRESS Switch ................................................................ 6-34 LDG ALT Selector........................................................................... 6-34 High Altitude Pressurization Mode ............................................ 6-34 Pressurization Indicator ................................................................ 6-35 PRESS SYS Light............................................................................. 6-35 EMER PRESS Light ........................................................................ 6-36 BLEED AIR Switches EMER Function ................................... 6-36 Cabin Altitude Warning Horn and Mute Function................... 6-36 CABIN ALT HI Light ..................................................................... 6-36 SYSTEM TEST Switch CABIN ALT Function ....................... 6-36 Air Conditioning and Heating .............................................................. 6-37 Primary Heating and Cooling Bleed Air..................................... 6-37 Air Distribution Schematic (Figure 6-11) .................................... 6-38 Temperature Control Schematic (Figure 6-12) ........................... 6-39 CAB AIR Switch ............................................................................. 6-40 Crew AUTO-MAN Switch ............................................................ 6-40 Cabin AUTO-CABIN-MAN Switch............................................. 6-40 Crew and Cabin COLD-HOT Selector Switches........................ 6-41 TEMP CONT Indicator.................................................................. 6-41 CAB TEMP Indicator ..................................................................... 6-41

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TABLE OF CONTENTS (Cont) R-134A Cooling System ...................................................................... 6-42 Refrigerant Cooling System (Figure 6-13) ........................................ 6-43 Cabin Climate Switches ...................................................................... 6-44 COOL-OFF Switch.......................................................................... 6-44 CABIN FAN Switch........................................................................ 6-44 CREW FAN Switch ......................................................................... 6-44 Hourmeter Compressor ................................................................ 6-44 Auxiliary Heating System.................................................................. 6-45 Cabin Auxiliary Heat ..................................................................... 6-45 Cockpit Floorboard Heaters .......................................................... 6-45 AUX HT Switch............................................................................... 6-46 Tailcone Baggage Compartment Heater System ................................ 6-46

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SECTION VI ANTI-ICE & ENVIRONMENTAL

BLEED AIR SUPPLY
Engine bleed air is used extensively for anti-icing and cabin environmental control. The source of this air is low- and high-pressure ports on each engine compressor. From the engine compressor, the bleed air is mixed and regulated in the mixing/regulating valve mounted on each engine. The bleed air is then ducted from the engines into the tailcone where it is available for several using systems. Shutoff valves and check valves are installed in the tailcone plumbing to control the bleed air from the left and right engines. In addition to the plumbing, the system includes BLEED AIR switches and an overheat warning system. BLEED AIR SWITCHES The L and R BLEED AIR switches, located in the BLEED AIR group on the copilots switch panel, control the respective left and right bleed-air shutoff valves and left and right emergency pressurization valves. Each BLEED AIR switch has three positions: EMER, ON and OFF. When a BLEED AIR switch is in the ON position, the respective bleed-air shutoff valve will open and the emergency pressurization valve will be closed. When a BLEED AIR switch is set to OFF, the respective bleedair shutoff valve will be energized to the closed position. When a BLEED AIR switch is set to EMER, the respective bleed-air shutoff valve will close and the emergency pressurization valve will be energized open and the high-stage bleed air will be shut off. The bleed-air shutoff valve will close automatically whenever emergency pressurization is activated or the ENG FIRE PULL T-handle is pulled on the respective side. The bleed-air shutoff valves control bleed-air flow to the cabin air distribution and temperature control systems, wing antiice system, and windshield anti-ice system. Bleed air for nacelle, engine anti-icing, and windshield alcohol tank pressurization is still available with the shutoff valves closed. The bleed-air shutoff valves and emergency pressurization valves operate on 28 VDC supplied through the L and R BLEED AIR circuit breakers on the pilots and copilots circuit breaker panels.

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BLEED AIR

EMER

O N CABIN AIR DISTRIBUTION OFF COCKPIT AIR DISTRIBUTION CABIN AIR DISTRIBUTION

EMERGENCY PRESSURIZATION VALVE

BLEED-AIR SHUT-OFF VALVE

BLEED-AIR SHUT-OFF VALVE

EMERGENCY PRESSURIZATION VALVE

WING ANTI-ICE LOW PRESS ENGINE BLEED HIGH PRESS BLEED AIR MIX VALVE WINDSHIELD ANTI-ICE

GROUND SERVICE

PRESSURIZATION JET PUMP & ALCOHOL ANTI-ICE

BLEED AIR MIX VALVE

LOW PRESS ENGINE BLEED HIGH PRESS

WING ANTI-ICE BYPASS CIRCUIT TEMPERATURE CONTROL SYSTEM (SERVO AIR)

WING ANTI-ICE BYPASS CIRCUIT

NACELLE INLET ANTI-ICE

HYDRAULIC SYSTEM

NACELLE INLET ANTI-ICE

CABIN AIR ON

FLOW CONTROL VALVE

CREW OFF

CABIN TEMPERATURE CONTROL VALVE TEMPERATURE CONTROL VALVE

COLD

HOT

COLD

HOT

RAM AIR

RAM AIR

RAM AIR PLENUM

HEAT EXCHANGER

BLEED AIR

CONDITIONED AIR

BLEED AIR SUPPLY SCHEMATIC Figure 6-1

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CABIN AIR LIGHT A white CABIN AIR advisory light indicates that either the L BLEED AIR, R BLEED AIR or CAB AIR switches are in the off position. BLEED AIR WARNING LIGHT Engine pylon, bleed-air duct, and tailcone overheat indication is provided by the red BLEED AIR L and BLEED AIR R warning lights. Each light is operated by thermoswitches installed in the pylon structure and in the bleed-air ducting. Activation of either thermoswitch will illuminate the associated light. The thermoswitch in the pylon structure will cause the associated light to illuminate if the pylon structure temperature reaches approximately 250F. The thermoswitch in the pylon bleed-air ducting will cause the associated light to illuminate if the duct temperature reaches approximately 600F. In addition to the thermoswitches, a tailcone sensing element is installed to detect elevated tailcone temperatures caused by a leak in the bleed-air ducting. If both the BLEED AIR L and BLEED AIR R warning lights illuminate simultaneously, the tailcone overheat sensor has tripped the lights. The lights operate on 28 VDC supplied through the WARN LTS circuit breakers on the pilots and copilots circuit breaker panels. The tailcone overheat detection system operates on 28 VDC supplied through the BLEED AIR OV HT circuit breaker on the pilots circuit breaker panel. Warning lights and tailcone overheat detection is operative during EMER BUS mode

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ANTI-ICE SYSTEMS
Aircraft anti-ice protection is provided through the use of electrically heated anti-ice systems, engine bleed-air heated anti-ice systems, and an alcohol anti-ice system. Electrically heated systems include the pitot-static probes, total air temperature probe, engine inlet air temperature/pressure sensors, stall warning vanes, and horizontal stabilizer leading edge. Electrically-heated windshields provide defogging for the windshield interior. Engine bleed air is utilized to provide anti-icing for the wing leading edge, windshield, nacelle inlets, lowpressure compressor inner stator, and engine fan spinners. The alcohol system is installed to provide backup anti-ice protection for the pilots windshield in event of normal anti-icing system malfunction. ROSEMOUNT ICE DETECTOR SYSTEM (OPTIONAL) The optional Rosemount Ice Detector system is installed to detect an icing condition and notifies the pilots by illumination of the amber or white ICE DET lights, in the glareshield annunciator panel, and both Master CAUT lights. A self-test of the Rosemount Ice Detector system is conducted every time aircraft power is turned on, and the ICE DETECTOR circuit breaker is engaged. The ice detector system self-test will show a failed self-test if the amber ICE DET light and both Master CAUT lights are illuminated. The Rosemount Ice Detection System provides an additional means of ice detection and should not be used as the only source of ice detection. The Rosemount Ice Detector System receives 28 VDC through the ICE DETECTOR circuit breaker on the pilots circuit breaker panel. When the Rosemount Ice Detector probe detects an icing condition, and the STAB WING HEAT switch is Off, the amber ICE DET light located in the glareshield annunciator panel, and both Master CAUT lights will illuminate. Probe de-icing is done automatically by the Rosemount system itself. Selecting the STAB WING HEAT switch On will inhibit the amber ICE DET light and enable the white ICE DET light. The ICE DET white light is an advisory light which will illuminate only when icing is detected while the STAB WING HEAT switch is On. Illumination of the ICE DET amber light with the STAB WING HEAT switch On indicates a failure of the Rosemount Ice Detection system.

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ICE DETECT LIGHTS Two ice detect lights are installed on the forward glareshield to indicate ice or moisture formation on the windshield during night operations. These lights are illuminated whenever the BATTERY switches are On. When particles of ice or moisture form, light refraction results in the appearance of two red areas, approximately 1-1/2 inches (38 mm) in diameter, on the windshield. The light on the pilots side is located in a position covered by the windshield anti-ice airstream. The copilots light is positioned outside the airstream; therefore, the copilots windshield must be monitored whenever windshield anti-ice system is in operation. The red areas indicate ice encounters when the SAT is below freezing and moisture encounters when the SAT is above freezing. The lights are supplied 28 VDC through the L and R ICE DETECT LIGHT circuit breakers on the pilots and copilots circuit breaker panels respectively. WING INSPECTION LIGHT The wing inspection light, located on the right forward fuselage, may be used to visually inspect the right wing leading edge for ice accumulation during night operations. The light is illuminated by depressing the WING INSP LIGHT momentary switch. The switch is located on the copilots switch panel. The light illuminates a black dot on the outboard wing leading edge to enhance visual detection of ice accumulation. Power is supplied through the WING INSP LT circuit breaker on the copilots circuit breaker panel.

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ENGINE AND NACELLE INLET ANTI-ICE The engine and nacelle inlet anti-ice system provides anti-ice protection for the engine fan spinners, low pressure compressor inner stator, nacelle inlets, and the engine inlet air temperature and pressure sensors. The fan spinners, low pressure compressor inner stator, and nacelle inlets are anti-iced by engine bleed air. The fan spinners are continually heated by bleed air flowing between their double-wall construction. The low pressure compressor inner stator and nacelle inlet are heated by bleed air when the associated NAC HEAT switch is on. The engine air temperature (TT0) and pressure (PT) sensors are anti-iced by integral electrical heating elements. Each engine anti-ice system is independently operated and consists of TT0/PT sensor heating elements, a nacelle inlet anti-ice control valve (controls flow to the nacelle inlet lip), an engine anti-ice control valve (controls flow to the low-pressure compressor inner stator), a pressure switch, a control switch, a NAC HT light, and associated aircraft wiring and bleed-air plumbing. Control circuits are powered by 28 VDC supplied through the L and R NAC HEAT circuit breakers on the pilots and copilots circuit breakerpanels respectively. NAC HEAT SWITCHES The left and right engine and nacelle inlet anti-ice systems are independently controlled through the NAC HEAT switches in the ANTI-ICE group on the center switch panel. Each NAC HEAT switch has two positions: On (L or R) and OFF. When a NAC HEAT switch is placed in the On (L or R) position, the associated TT0/PT sensor elements will be energized and the associated engine and nacelle inlet anti-ice control valves will open. Engine bleed air will flow through the open valves to the low pressure compressor inner stator and nacelle inlet lip. Since the control valves are energized closed, engine and nacelle inlet anti-ice protection will still be available in the event of an electrical system failure.

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NAC HT LIGHTS The amber L and R NAC HT lights on the glareshield annunciator panel provide the crew with visual indication of an engine or nacelle inlet anti-ice system malfunction. The lights are operated by a pressure switch in the associated nacelle inlet bleed air plumbing and a proximity switch built into the engine anti-ice control valve. Illumination of a NAC HT light when the associated NAC HEAT switch is in the On position, indicates that insufficient pressure is being applied to the nacelle inlet or the engine anti-ice control valve has failed to open. Illumination of a NAC HT light, when the associated NAC HEAT switch is in the OFF position, indicates that bleed-air pressure is being applied to the nacelle anti-ice system due to a malfunction of the nacelle anti-ice control valve. The green NAC HT light on the glareshield annunciator panel provides the crew with visual indication that either nacelle heat switch is On. WING ANTI-ICE The wing anti-ice system utilizes engine bleed air directed through diffuser tubes in each wing leading edge. The heated air is distributed to the wing root and leading edge and then allowed to exit into the center wing/wheel well area. The system consists of wing diffuser tubes, a WING HT caution light, two thermoswitches (one underheat sensor and one overheat sensor), a wing temperature sensor, an anti-ice shutoff and pressure regulator valve, a bleed air bypass valve on each engine, a wing temperature indicator, a system switch, and associated aircraft wiring. Electrical power for system operation is 28 VDC supplied through the WING HEAT circuit breaker on the copilots circuit breaker panel.

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Bleed Air Mix Valve LP Bleed Air (from engine) Temperature Sensor HP Bleed Air (from engine)
STAB WING HEAT

Underheat Thermoswitch
WING HT

Overheat Thermoswitch

Bleed Air Bypass Valve

OFF

Bleed Air Shutoff Valve HP Bleed Air (from LH engine) Shutoff & Pressure Regulator Mixed Bleed Air (from LH engine)

To Other Systems

W I N G

T E M P

When STAB WING HEAT Switch is ON, connection will be made.

WING ANTI-ICE SYSTEM Figure 6-2

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STAB WING HEAT SWITCH WING HEAT FUNCTION The wing anti-ice system is controlled through the STAB WING HEAT switch located in the ANTI-ICE group on the center switch panel. The switch has two positions: On (STAB WING HEAT) and OFF. When the STAB WING HEAT switch is set On, the anti-ice shutoff and pressure regulator valve control solenoid will close allowing pressure to build within the valve reference chambers. The building pressure will open a butterfly valve in the bleed-air airstream and allow heated air to flow through the ducting into the wing diffuser tubes. The valve will maintain a regulated 15 (2.5) psi bleed airflow providing the butterfly remains open. In the event of an electrical system failure, the valve will shut off the bleed-air flow and wing anti-ice protection will not be available. Two sources of bleed air are used for wing anti-ice. In addition to the normal bleed-air supply (mixed low- and high-pressure), bypass circuits are activated which makes hotter bleed air from the engines high pressure ports available for wing anti-icing. A temperature sensor will deactivate the bypass circuit if the respective high-pressure duct becomes too hot. When the STAB WING HEAT switch is set to OFF, the bypass circuits are deactivated. Additionally, the bypass circuit is deactivated if the respective BLEED AIR switch is not ON or the respective ENG FIRE PULL T-handle is pulled. WING TEMP INDICATOR The WING TEMP indicator, located on the center switch panel in the ANTI-ICE group, is installed to provide a visual indication of the wing leading edge temperature. The indicator receives input signals from the wing temperature sensor installed on the inner surface of the left wing leading edge. The indicator face is divided into three colored segments: blue, green, and red. If the indicator pointer is in the blue segment, wing leading edge temperature is cold enough for moisture to freeze on the surface. If the indicator pointer is in the green segment, wing leading edge temperature is warm enough that moisture will not freeze on the surface. If the indicator pointer is in the red segment, the wing leading edge is approaching an overheat condition and corrective action must be taken. The wing anti-ice system should be energized whenever flying through visible moisture and the WING TEMP indicator pointer is in the blue segment.

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WING HT LIGHT The amber WING HT light, on the glareshield annunciator panel, will illuminate to indicate the wing anti-ice system is not maintaining the temperature of the leading edge in the normal operating range. In the event that the wing leading edge heats to 215F (102C), the overheat thermoswitch located on the inner skin of the right wing leading edge will cause the light to illuminate. If the wing leading edge temperature cools to 55F (13C) and the STAB WING HEAT switch is on, the underheat thermoswitch located on inner skin of the right wing leading edge will cause the light to illuminate. The light will illuminate upon initial activation of the wing anti-ice system if the wing temperature is below the set point of the underheat thermoswitch. As the temperature of the wing leading edge rises, the light should extinguish. HORIZONTAL STABILIZER ANTI-ICE The horizontal stabilizer anti-ice system utilizes sequenced electrical heating elements along the horizontal stabilizer leading edge. The system consists of an electrically heated blanket bonded to each half of the horizontal stabilizer leading edge, three remote control circuit breakers (RCCB), a heat controller, a caution light, a system switch, and associated aircraft wiring. Control circuits operate on 28 VDC supplied through the STAB HEAT circuit breaker on the copilots circuit breaker panel. Electrical power for the heating elements is 28 VDC supplied through three 50-amp current limiters. STAB WING HEAT SWITCH STABILIZER HEAT FUNCTION The horizontal stabilizer anti-ice system is controlled through the STAB WING HEAT switch located in the ANTI-ICE group on the center switch panel. The switch has two positions: On (STAB WING HEAT) and OFF. When the aircraft is in flight and the STAB WING HEAT switch is On, 28 VDC is supplied through the three RCCBs to the heat controller. The heat controller distributes intermittent electrical power to the individual heating elements in a forward-to-aft sequence of 15 seconds duration each. Approximately 3 minutes are required to complete a full cycle. The center, or parting elements, are supplied with continuous electrical power. At least one engine generator must be operating to enable the heat controller circuits. The controller circuits are biased by starter engaged and weight-on-wheels signals; therefore, the system is inoperative when the squat switch is in the ground mode and during engine start.

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STAB HT LIGHT The amber STAB HT light, located on the glareshield annunciator panel will illuminate when any of the following conditions exist: On the ground STAB HEAT circuit breaker is pulled. STAB WING HEAT switch is On. In flight STAB HEAT circuit breaker is pulled. The STAB WING HEAT switch is On and any one heating element fails (remaining elements will continue to function normally). During flight, illumination of the STAB HT light indicates system failure. During ground operation, the STAB HT light should illuminate whenever the STAB WING HEAT switch is On. STABILIZER HEAT SELF TEST A self test may be conducted with the aircraft on the ground and a generator on-line. Under these conditions, when the STAB WING HEAT switch is turned on the following events should happen: 1. The STAB HT light will illuminate. 2. The generator load will increase approximately 120 amps total for 2 to 3 seconds and then decrease to the STAB HEAT off value. 3. The STAB HT light will remain illuminated indicating the system is functioning normally. The following events indicate a failure of the system: 1. STAB HT light does not illuminate when STAB WING HEAT switch is turned on. Turn STAB WING HEAT switch off. 2. Load does not decrease within 5 seconds. Turn STAB WING HEAT switch off. 3. STAB HT light flashes approximately 3 times per second. One or more heating elements are not within their operating tolerance (element failure). Turning STAB WING HEAT switch off will cancel the flashing. The STAB WING HEAT switch must be off for 3 minutes allowing the system to reset before another self test attempt can be made.

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WINDSHIELD ANTI-ICE Primary windshield anti-icing is accomplished by directing conditioned engine bleed air through ducting and control valves to external outlet nozzles forward of the windshield. The windshield anti-ice system consists of a shutoff valve, an anti-ice modulating valve, two lowlimit overheat thermoswitches, two high-limit overheat thermoswitches, a green WSHLD HT light, an amber WSHLD OV HT caution light, a ram air modulating valve, an anti-ice duct temperature sensor, an anti-ice heat exchanger, two outlet nozzle assemblies, a system control switch, and associated aircraft wiring and bleed-air ducting. Electrical power to the control circuits is 28 VDC supplied through the WSHLD HEAT circuit breaker on the copilots circuit breaker panel. WSHLD HEAT SWITCH The windshield anti-ice system is controlled through the WSHLD HEAT switch in the ANTI-ICE grouping on the center switch panel. The switch has three positions: WSHLD HEAT (On), HOLD, and OFF. When power is applied to the aircraft, or the BATTERY switches are set On, the windshield anti-ice shutoff valve is energized to the open position. When open, the shutoff valve allows engine bleed air to the antiice modulating valve downstream. When the WSHLD HEAT switch is placed in the On position, a circuit is completed to the anti-ice modulating valve and WSHLD HT indicator light. The anti-ice modulating valve will move toward full open until the valve is fully open or the WSHLD HEAT switch is set to HOLD. When the switch is in the HOLD position, the anti-ice modulating valve will remain in its last attained position, and allow bleed air to the anti-ice heat exchanger. When the WSHLD HEAT switch is set to OFF, the anti-ice modulating valve will move towards the closed position until the valve is fully closed or the WSHLD HEAT switch is set to HOLD. The anti-ice modulating valve will fully open or close in approximately 15 seconds. The anti-ice heat exchanger cools the bleed air with ram air regulated by a ram air modulating valve. This valve is controlled by the downstream anti-ice duct temperature sensor and regulates the anti-ice bleed air temperature by varying the amount of ram air allowed into the heat exchanger.

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WSHLD OV HT

WSHLD HT

WSHLD HEAT

WSHLD HEAT HOLD OFF WING ANTI-ICE SYSTEM

ENGINE BLEED AIR

ENGINE BLEED AIR

ANTI-ICE SHUTOFF VALVE

ANTI-ICE MOD VALVE

ENVIRONMENTAL SYSTEM HEAT EXCHANGER

WINDSHIELD ANTI-ICE HEAT EXCHANGER RAM AIR MODULATING VALVE

RAM AIR

HIGH TEMPERATURE LIMIT THERMOSWITCH LOW TEMPERATURE LIMIT THERMOSWITCH SQUAT SWITCH RELAY (makes connection when aircraft is on the ground)

Anti-Ice Shutoff Valve is normally closed

(must be energized open) WSHLD HT light out

Electrical ground on this wire turns Electrical ground on this wire turns
WSHLD OV HT light on

WINDSHIELD ANTI-ICE SYSTEM Figure 6-3

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WSHLD HT LIGHT The green WSHLD HT light, located on the glareshield annunciator panel, provides the crew with a visual indication of windshield heat operation. The light is extinguished when the WSHLD HEAT switch is set to OFF. The light will illuminate when the WSHLD HEAT switch is moved out of the OFF position and remain illuminated until either the switch is set to OFF or an overheat thermoswitch trips shutting airflow off and extinguishing the green WSHLD HT light. WSHLD OV HT LIGHT Illumination of the amber WSHLD OV HT caution light, on the glareshield annunciator panel, indicates that the bleed air temperature in one or both of the windshield outlet nozzles has reached the respective low- or high-limit thermoswitch settings and the windshield antiice system has been shutdown by either the low- or high-limit thermoswitches. During ground operations, the light is controlled by the low-limit switches. In flight, the light is controlled by the high-limit switches. If the bleed air temperature in either outlet nozzle reaches 250F (121C) during ground operation, the low-limit overheat thermoswitches will close the anti-ice shutoff valve and illuminate the WSHLD OV HT caution light. If the outlet nozzle bleed air temperature in either nozzle reaches 347F (175C) in flight, the high-limit overheat thermoswitches will perform the same function. When the nozzle bleed air temperature drops to 240F (115C) during ground operations, or 311F (155C) in flight, the overheat thermoswitches will reset allowing the anti-ice shutoff valve to open and extinguish the WSHLD OV HT caution light. To avoid a false WSHLD OV HT indication upon landing, the low-limit overheat thermoswitch circuitry is disabled for 10 seconds after touchdown, after which normal functioning will resume.

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WINDSHIELD DEFOG Windshield internal defogging is accomplished using electrically heated windshield panels. The system is designed so that it may be activated before takeoff and remain on until shutdown. The system consists of two windshield panels with integral heaters, windshield heat control unit, system switch, L and R WS DEFOG annunciators, and associated aircraft wiring. The system utilizes the 115 VAC output from the inverter system to power the integral heaters. The control circuit receives 28 VDC through the L WSHLD DEFOG and R WSHLD DEFOG circuit breakers on the pilots and copilots circuit breaker panels. The 115 VAC input to the system is provided through the L and R WSHLD DEFOG circuit breakers on the pilots and copilots circuit breaker panels. .
LH WINDSHIELD TEMP CONTROL SENSOR OVER-TEMP SENSOR RH WINDSHIELD TEMP CONTROL SENSOR OVER-TEMP SENSOR

HEATER

HEATER

L WS DEFOG

R WS DEFOG

WINDSHIELD HEAT CONTROL UNIT

L WSHLD R WSHLD DEFOG DEFOG

WSHLD DEFOG NORM

R WSHLD DEFOG

L WSHLD DEFOG

*Aircraft with three

115 VAC IN position switch

*LOW
OFF

WINDSHIELD DEFOG SYSTEM Figure 6-4

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WSHLD DEFOG SWITCH The windshield defog system is controlled through the WSHLD DEFOG switch in the ANTI-ICE group on the center switch panel. The switch positions are OFF, LOW and NORM. With the WSHLD DEFOG switch set to LOW or NORM, the integral heaters will be supplied 115 volts AC power from the inverter system via the windshield heat control unit. When the switch is set to LOW, operating temperature range of the windshield is 90-97F (32-36C). When the WSHLD DEFOG switch is set to NORM, operating temperature range of the windshield is 105-120F (41-49C).
NOTE

Normally, the left inverter will power the left windshield panel while the right inverter will power the right windshield panel. However, either inverter is capable of powering both windshield panels. Should one inverter switch be in the on position and the other in the off position, switching will occur allowing the operative inverter to power both windshield panels.

Normal system operation is indicated by illumination of the L and R WS DEFOG annunciators when the system is activated (windshield temperature below 85F [29C]). When the windshield is heated above 85F (29C), the annunciators will extinguish. L AND R WS DEFOG ANNUNCIATORS Illumination of a WS DEFOG annunciator, located on the glareshield annunciator panel, indicates an over-temperature condition, undertemperature condition or loss of AC or DC power. Temperature sensors are attached to each windshield panel which provide temperature data to the windshield heat control unit. Should the temperature of the windshield drop below 85F (29C), the applicable WS DEFOG annunciator will illuminate to alert the crew. Should the temperature of the windshield increase above 150F (66C), the applicable WS DEFOG annunciator will illuminate and the affected windshield will be deactivated. When the windshield cools to the normal operating range, the system will reactivate and the WS DEFOG annunciator will extinguish. Electrical faults detected by the system monitor will cause the affected WS DEFOG annunciator to illuminate.

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WINDSHIELD ANTI-ICE ALCOHOL SYSTEM The alcohol anti-ice system is utilized for windshield anti-icing in the event of a windshield heating system malfunction. Alcohol anti-icing is accomplished by directing methyl alcohol over the pilots windshield surface through an external outlet in the windshield heat outlet nozzle assembly. The system consists of a 2.35 gallon alcohol reservoir, a float switch, a filter, a relief valve, a three-way control valve, a bleed air shutoff and pressure regulator valve, a system switch, an amber ALC LOW caution light and associated aircraft wiring. The pressure relief valve is installed to prevent system overpressurization by venting system pressure greater than 2.6 psi above ambient, and bleed system pressure when the system is off. The system control circuits operate on 28 VDC supplied through the ALCOHOL SYSTEM circuit breaker on the copilots circuit breaker panel. WSHLD ALC SWITCH The windshield alcohol anti-ice system is controlled by the WSHLD ALC switch in the ANTI-ICE group on the center switch panel. The switch has two positions: WSHLD ALC (On) and OFF. When the switch is set to WSHLD ALC, circuits are completed to open the shutoff and pressure regulator valve and position the three-way control valve for alcohol flow to the windshield. The alcohol reservoir, pressurized to approximately 2.4 psi above ambient through the shutoff and pressure regulator valve, supplies alcohol to the windshield outlet through a filter and the three-way control valve. When the switch is set to OFF, the shutoff and pressure regulator valve will close, the three-way valve will reposition to cut off flow and system pressure will bleed off through the pressure relief valve. ALC LOW CAUTION LIGHT Illumination of the amber ALC LOW light, located on the glareshield annunciator panel, indicates the alcohol supply in the reservoir is low. The reservoir float switch will illuminate the light through a relay when in the full down position. When the relay is energized, a holding circuit is also energized to prevent the light from flickering due to the bobbing motion of the float. The holding circuit is de-energized when the BATTERY switches are set to OFF and the alcohol reservoir is filled. A completely filled reservoir will supply the windshield alcohol anti-ice system with approximately 45 minutes of alcohol flow.

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WSHLD ALC

OFF

BLEED AIR

ALCOHOL RESERVOIR

TO ENVIRONMENTAL SYSTEM

ALC LOW

BLEED AIR

OVERBOARD VENT BLEED AIR PRESSURE ALCOHOL SUPPLY ELECTRICAL FILTER FLOAT SWITCH

PRESSURE RELIEF VALVE CHECK VALVE ALCOHOL PRESSURE REGULATOR & SHUTOFF VALVE THREE-WAY VALVE

ALCOHOL ANTI-ICE SYSTEM Figure 6-5

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PITOT-STATIC AND STALL WARNING ANTI-ICE Anti-ice protection for the pitot-static probes, total temperature probe, stall warning vanes, and the pressurization static port is accomplished by energizing integral electrical heating elements in each component. The independent pitot-static probe, total temperature probe, and stall warning vane anti-ice systems consist of control switches, probe heaters, vane heaters, and pitot heat monitors. Both left, right and standby systems utilize the same PITOT HT light. The pressurization static port heater is part of the right system. The pitot-static probe heating elements receive 28 VDC through their respective L PITOT HEAT, R PITOT-STALL-TAT HEAT, and STANDBY PITOT HEAT circuit breakers on the pilots and copilots circuit breaker panels. The total temperature probe heating element receives 28 VDC through the TAT PROBE HEAT circuit breaker on the copilots circuit breaker panel. Total temperature probe heat is only enabled when the squat switch is in the air mode. The pressurization static port heating element receives 28 VDC through the R PITOT-STALL-TAT HEAT circuit breaker on the copilots circuit breaker panel. The stall warning vane heating elements receive 28 VDC through the respective L and R STALL VANE HEAT circuit breakers on the pilots and copilots circuit breaker panels. An optional Triple Pitot Heat Indication System may be installed. The system does not change the anti-ice protection for the pitot-static probes, stall warning vane, or total temperature probe. It does add specific warning annunciators in the event of failure of either left, right, or standby pitot-static heat system. The annunciators are installed on the center instrument panel, below the PITOT HEAT placard. PITOT HEAT SWITCHES The pitot-static heat systems are controlled through the PITOT HEAT switches in the ANTI-ICE group on the center switch panel. Each switch has two positions: On (L or R) and OFF. When the L and R PITOT HEAT switches are set to On (L and R), power is supplied to each pitot-static probe heater, each stall warning vane heater, the total temperature probe heater (aircraft in flight), and the pressurization static port heater. The standby pitot-static probe, pressurization static port, and the total temperature probe heat are activated through the R PITOT HEAT switch.

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PITOT HT LIGHT A pitot heat monitor system is installed to alert the pilot if insufficient current is being applied to any of the pitot-static probe heating elements (left, right and standby). Each monitor is basically a relay which maintains an open circuit for the PITOT HT light as long as sufficient current is being applied to the associated pitot-static probe heating element. In the event of a malfunction in or loss of power to the associated pitot-static probe heating element, the relay will release and complete the PITOT HT light circuit. Illumination of the amber PITOT HT light, in the glareshield annunciator panel, indicates a malfunction in either the left, right or standby pitot-static heat system, or that at least one PITOT HEAT switch is OFF. L, R AND STBY PITOT HEAT LIGHTS In the event of a malfunction in the pitot-static heat system, the applicable amber L, R, or STBY annunciator, and both Master CAUT lights will illuminate and flash. Additional pitot-static heat system failures will cause the applicable individual L, R, or STBY annunciator to illuminate and both Master CAUT lights to illuminate and flash. When the aircraft is powered from the EMER BUS, the L and R pitot heat annunciators will illuminate to notify pilots that only the standby pitot heat is operational.

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OXYGEN SYSTEM
The aircraft oxygen system provides oxygen service for the crew and passengers. The system consists of the crew and passenger distribution systems, a high-pressure oxygen storage cylinder, a shutoff valve and pressure regulator assembly, an oxygen pressure transducer, an oxygen pressure indicator, an overboard discharge relief valve and indicator, a passenger oxygen control valve, lanyard actuated passenger mask oxygen valves, and crew and passenger oxygen masks. Electrical power to operate the passenger oxygen control valve and oxygen indicator is supplied through the OXYGEN VALVE circuit breaker on the pilots circuit breaker panel. Oxygen is available to the crew at all times and can be made available to the passengers either automatically above 14,500 (250) feet cabin altitude, or manually at all altitudes through the use of the cockpit controls on the pilots circuit breaker panel. The oxygen system is designed for use during emergency descent to a cabin altitude not requiring oxygen and is not to be used for extended periods of flight at cabin altitudes requiring oxygen or as a substitute for the normal pressurization system. Smoking is prohibited when oxygen is in use. OXYGEN STORAGE AND PRESSURE REGULATION Several oxygen storage cylinder arrangements are used: Single cylinder in the nose compartment (40 or 77 cubic feet) Single cylinder in the vertical stabilizer (77 cubic feet) Dual cylinders one in the nose compartment (40 or 77 cubic feet) and one in the vertical stabilizer (77 cubic feet) The shutoff and pressure regulator assembly forms an integral part of the storage cylinder and provides for pressure regulation, pressure indication, and servicing. Oxygen pressure for the passenger and crew distribution systems is regulated to a pressure of 60 to 80 psi. The shutoff and pressure regulator assembly also incorporates a burst disc pressure relief valve to discharge the oxygen cylinder contents overboard in the event that cylinder pressure reaches 2700 to 3000 psi. Should the cylinder contents be discharged overboard, the green overboard discharge indicator will be ruptured or missing. Storage cylinders mounted in the nose compartment have the overboard discharge indicator located on the lower left side of the nose section. Storage cylinders mounted in the vertical stabilizer have the overboard discharge indicator located on the left side at the base of the vertical stabilizer.

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FILLER VALVE TRANSDUCER

FORWARD OXYGEN CYLINDER

P S I

O X Y P R E S S

OVERBOARD DISCHARGE INDICATOR RELIEF VALVE

X 1 0 0 0

CREW MASK

QUICK-DISCONNECT VALVE

QUICK-DISCONNECT VALVE

CREW MASK

AUTO
F OF

DE

PL

PASSENGER MASK AUTOMATIC DEPLOY @ 14,500 FEET (CABIN ALTITUDE)

PASSENGER OXYGEN

PASSENGER MASK ASSEMBLY

(with single forward cylinder) OXYGEN SYSTEM SCHEMATIC Figure 6-6 6-22
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PRESSURE SWITCH

FILLER VALVE SEAL VALVE AFT OXYGEN CYLINDER

TRANSDUCER
AFT
P S I X 1 0 0 0

O X Y P R E S S

RELIEF VALVE

OVERBOARD DISCHARGE INDICATOR

CREW MASK

QUICKDISCONNECT VALVE

QUICKDISCONNECT VALVE

CREW MASK

AUTO

F OF

DE PL OY

PASSENGER MASK AUTOMATIC DEPLOY @ 14,500 FEET (CABIN ALTITUDE)

PASSENGER OXYGEN

PASSENGER MASK ASSEMBLY

F6006000006601

(with single aft cylinder) OXYGEN SYSTEM SCHEMATIC Figure 6-6A

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FILLER VAL VE PRESSURE SWITCH SEAL VAL VE

FILLER VAL VE

FORWARD OXYGEN CYLINDER

FWD
P S I X 1 0 0 0

AFT OXYGEN CYLINDER

O X Y P R E S S

TRANSDUCER

TRANSDUCER
AFT

RELIEF VAL VE OVERBOARD DISCHARGE INDICATOR

P S I X 1 0 0 0

O X Y P R E S S

RELIEF VAL VE

OVERBOARD DISCHARGE INDICATOR

CREW MASK QUICK-DISCONNECT VAL VE QUICK-DISCONNECT VAL VE

CREW MASK

AUTO

F OF

DE PL OY

PASSENGER MASK AUTOMATIC DEPLOY @14,500 FEET (CABIN ALTITUDE)

PASSENGER OXYGEN

PASSENGER MASK ASSEMBL Y

F60-060000-001-01

(with dual cylinders) OXYGEN SYSTEM SCHEMATIC Figure 6-6B 6-24

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OXYGEN PRESSURE INDICATOR The vertical-scale oxygen pressure indicator is located on the pilots circuit breaker panel. The indicator face is marked from 0 to 2000 psi in 250 psi increments and is controlled by an electric transducer plumbed to the high-pressure side of the shutoff and pressure regulator assembly. The oxygen supply system may be a single cylinder or dual cylinder system. The pressure indicator is located on the pilots circuit breaker panel. In aircraft with dual systems, a second pressure indicator is added to the pilots circuit breaker panel to allow determination of the oxygen pressure in each oxygen cylinder. The transducer for the aft oxygen system is wired through a pressure switch to the aft pressure indicator. The pressure switch senses loss of pressure in the aft oxygen tube. If the aft cylinder is pressurized but the supply tube is not (for example; due to blockage) the indicator will read zero. Since pressure will vary due to temperature the fore and aft cylinder may not indicate the same during flight.

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OXYGEN SYSTEM COCKPIT CONTROLS The oxygen system cockpit controls consist of one control valve, labeled PASSENGER OXYGEN OFF-AUTO-DEPLOY, located on the pilots circuit breaker panel. The control valve controls oxygen availability to the passenger oxygen distribution system and provides automatic or manual mode selection. Oxygen is available to the crew oxygen distribution system at all times when the oxygen cylinder shutoff valve is open. Control positions and system functions are as follows: 1. With the PASSENGER OXYGEN valve in the AUTO position, oxygen is available to the passenger distribution system and the passenger masks will deploy automatically in the event cabin altitude climbs to 14,500 feet. Should the cabin altitude reach 14,500 (250) feet, an electrical signal from the pressurization indicator will cause the solenoid valve (integral with the PASSENGER OXYGEN valve) to open, the passenger oxygen masks will deploy, and the cabin overhead lights will illuminate to provide maximum visibility for donning masks. Normally, the control should be in this position. 2. With the PASSENGER OXYGEN valve in the DEPLOY position, oxygen is available to the passenger distribution system and the passenger masks will deploy. Setting the PASSENGER OXYGEN valve to the DEPLOY position will manually open the PASSENGER OXYGEN valve and allow oxygen pressure to deploy the passenger masks. This position can be used to deploy the passenger masks at any cabin altitude and must be used if electrical power is unavailable. 3. With the PASSENGER OXYGEN valve in the OFF position, oxygen will not be available to the passenger distribution system regardless of cabin altitude. This position can be used when oxygen is required for the crew members only.

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PASSENGER MASKS The passenger oxygen masks are stowed in compartments in the convenience panels above the passenger seats. Whenever the compartment doors open automatically (PASSENGER OXYGEN-AUTO) or manually (PASSENGER OXYGEN-DEPLOY) the passenger oxygen masks will fall free and oxygen will be available for passenger use. Passengers should don masks and pull the mask lanyard to initiate oxygen flow. An orifice incorporated in the mask tubing connections will provide a constant flow rate of 4.5 liters per minute. A green area of the reservoir bag inflates when oxygen is flowing. Should the doors be inadvertently opened from the cockpit, pressure must be bled from the system by pulling one of the mask lanyards before the masks can be restowed. The compartment doors can be opened manually for mask cleaning and servicing per Maintenance Manual instructions.

OXYGEN VALVE (LANYARD OPERATED) ELASTIC STRAP DOOR (OPEN) OXYGEN MASK
FL ATED N EEN I N OK G R OXYGE

LANYARD

RESERVOIR BAG

GREEN INFLATED OXYGEN OK OXYGEN MASK

F6006000000201

PASSENGER MASK Figure 6-7

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CREW MASKS Scott ATO The flight crew oxygen masks are stowed in accessible stowage boxes just aft of the pilots and copilots circuit breaker panels or in storage cups just aft of the pilot and copilot on the bulkhead. The mask regulators provide for normal, 100% oxygen, and emergency operation (refer to the Airplane Flight Manual for detailed operational procedures). Each mask incorporates a microphone controlled by the NORM MIC/ OXY MIC switch on the respective audio control panel. When the OXY MIC is in use, a voice-activated hot interphone exists for crew member communication. An optional oxygen pressure detector may be located in the oxygen line. If sufficient pressure is available in the line, the detector shows green. .
INFLATABLE HARNESS

MASK

N
100% PUSH

MASK REGULATOR
EMERGENCY

OXYGEN LINE

PRESS TO TEST

MICROPHONE LINE
F6006000006801

CREW MASK SCOTT ATO Figure 6-8

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PRESSURIZATION SYSTEM
Cabin pressurization is provided by conditioned air entering the cabin through the air distribution ducts and controlled by modulating the amount of air exhausted from the cabin. The pressurization system consists of a cabin primary outflow valve, a cabin secondary outflow valve, an electronic pressurization controller, a LDG ALT selector, a MAN ALT control valve with rate control, a MODE switch, an EMER DEPRESS switch, a pressurization vacuum jet pump, a vacuum regulator, a pressurization indicator, two emergency pressurization valves, two emergency pressurization aneroid switches, an amber PRESS SYS caution light, an amber EMER PRESS caution light, and an aural warning system. All system controls are located in the PRESSURIZATION group on the copilots switch panel. The pressurization indicators are located directly above the system controls. Power for the control circuits is 28 VDC supplied through the CABIN PRESS SYS circuit breaker on the copilots circuit breaker panel. Power for the pressurization indicator is 28 VDC supplied through the CABIN PRESS IND circuit breaker on the pilots circuit breaker panel. Automatic and manual pressurization modes are available during EMER BUS mode. The pressurization indicator is operative during EMER BUS mode.

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PRESSURIZATION SYSTEM SCHEMATIC Figure 6-10 6-30

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NORMAL PRESSURIZATION Normal pressurization is controlled by regulating control pressure to the cabin primary and secondary outflow valves. The control pressure may be regulated automatically by the electronic pressurization controller or manually by the MAN ALT control knob. A pressurization vacuum jet pump provides vacuum (servo pressure) to operate the outflow valves. MANUAL mode operation is completely independent of the aircraft electrical system. If the cabin-to-ambient differential pressure reaches 9.7 psid, the positive pressure relief metering section of the outflow valves will cause the outflow valves to open and maintain a 9.7 psi differential. The outflow valves incorporate a cabin altitude limiter which limits cabin altitude to approximately 13,700 (500) feet should the system fail to maintain the normal cabin altitude. Should the cabin altitude reach approximately 13,700 (500) feet, the altitude limiters will vent cabin pressure to the outflow valve control chambers causing the valves to close. Should a rapid descent cause a negative pressure in the cabin, both the primary and secondary outflow valves will open to vent ambient atmospheric pressure to the cabin. When the system is in the automatic mode, the electronic controller maintains cabin pressure based on air data from the aircrafts air data computers, landing field elevation selected on the LDG ALT selector, position of the thrust levers, position of the landing gear squat switch, and the systems preprogrammed climb and descent schedules. The electronic controller features built-in test equipment which performs fault detection and annunciation routines during ground and flight operation. Should a fault be detected, the FAULT annunciator on the mode switch will illuminate and the system will automatically revert to manual mode. Depressing the mode switch will extinguish the FAULT annunciator and illuminate the MANUAL annunciator. When the system is in the manual or fault modes, the crew maintains the desired cabin pressure using the MAN ALT and MAN RATE controls to position the outflow valves. Moving the MAN ALT control to UP or DN controls the outflow valves directly causing them to open or close as appropriate until the MAN ALT control is moved to the center position. The desired cabin altitude is then controlled by the crew by reference to the pressurization indicator. The rate at which the outflow valves will respond to MAN ALT control movement is controlled by rotating the MAN RATE knob from MIN to MAX as desired.

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EMERGENCY PRESSURIZATION In the event of normal cabin airflow malfunction, emergency pressurization is provided by routing low pressure engine bleed air directly into the cabin through the emergency pressurization valves. Emergency pressurization is accomplished automatically by opening the emergency pressurization valves in response to signals from the aneroid switches when the cabin altitude increases to 9500 (250) feet or manually by setting the BLEED AIR switches to EMER. When the aircraft is below 25,000 feet pressure altitude and the system is in automatic mode with a takeoff or landing field elevation greater than 8000 feet specified, the aneroid switches will not trigger the emergency pressurization unless the cabin altitude increases to 14,500 (250) feet. Emergency pressurization is provided by two independent circuits left and right. If triggered automatically, the left and right circuits will activate approximately at the same time in response to the aneroid switch signals. If triggered manually, the left and right circuits may be activated separately. When emergency pressurization is triggered the following events occur: Emergency pressurization valve opens The bleed-air mix valve goes to the low-pressure bleed port The bleed-air shutoff valve closes The wing anti-ice bypass circuit is deactivated The EMER PRESS annunciator illuminates The result is that engine low-pressure bleed air is ducted directly into the cabin air overhead and floor diffusers. This bypasses all bleed-air plumbing in the tailcone area and will stabilize cabin altitude if the pressurization failure has occurred in that area. The emergency pressurization valves are energized to the open position and de-energized for normal bleed-air flow. Each valve is independent of the other and, whenever both valves are open, temperature control and bleed air for wing and windshield anti-ice will be unavailable. Operating power for emergency valve actuation is 28 VDC supplied through the L and R BLEED AIR circuit breakers on the pilots and copilots circuit breaker panels.

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PRESSURIZATION CONTROLS AND INDICATORS MODE SWITCH The MODE switch is an alternate-action switch located on the copilots switch panel. The switch is used to toggle the pressurization system between the automatic and manual modes. Upon initial power-up, the system will be in automatic mode if no faults were revealed in the selftest. If a fault is detected, the system will revert to manual and the FAULT annunciator (part of the MODE switch) will illuminate. To switch from automatic to manual mode and vice versa, the MODE switch is depressed and released. When manual mode is selected, the MANUAL annunciator (part of the MODE switch) will be illuminated. MAN ALT CONTROL The MAN ALT control is a 3-position valve located on the copilots switch panel. The control is used to direct either regulated vacuum or cabin pressure to the outflow valves positioning them so that the desired cabin altitude results. Moving the control to the UP detent applies regulated vacuum to the outflow valves causing them to move toward the open position and increasing cabin altitude. Moving the control to the DN detent applies cabin pressure to the outflow valves causing them to move toward the close position and decreasing cabin altitude. When the control is in the center position, the outflow valves remain in their last attained position stabilizing the cabin altitude. Incorporated into the MAN ALT control valve is a MAN RATE control. The MAN RATE control is an adjustable needle valve which restricts the passage between the MAN ALT valve and the outflow valves. The rate at which the outflow valves react to the MAN ALT control is adjusted by varying this restriction.

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EMER DEPRESS SWITCH The EMER DEPRESS switch is an alternate-action switch located on the copilots switch panel. A guard is installed over the switch to prevent inadvertent actuation. The switch is used to depressurize the cabin and increase cabin airflow for smoke and fume evacuation. The EMER DEPRESS function is available in both automatic and manual modes. When EMER DEPRESS is selected, the outflow valves receive a signal to move toward the full open position. The cabin altitude will ascend to the aircraft altitude or 13,700 (500) feet (cabin altitude limiter), whichever is less. When EMER DEPRESS mode is selected, the EMER DEPRESS annunciator (part of the EMER DEPRESS switch) will be illuminated. To de-select this mode, depress and release the EMER DEPRESS switch. LDG ALT SELECTOR The LDG ALT selector is located on the copilots switch panel. The selector consists of a circular instrument graduated from -1000 to 14,000 feet in 500-foot increments and a setting knob used by the crew to select the landing field elevation. As the setting knob is moved, the needle on the instrument moves to show the selected landing altitude. The selected landing field elevation signal is supplied to the pressurization controller for use in determining the appropriate cabin climb and descent profile. The elevation of the destination airport is selected on the LDG ALT selector prior to takeoff and checked again prior to descent. The LDG ALT selector has no effect in manual mode. HIGH ALTITUDE PRESSURIZATION MODE When the aircraft is going to takeoff or land at a field elevation greater than 8000 feet, the system changes to high altitude pressurization mode. This increases the warning elevation to 14,500 (250) feet cabin altitude when the aircraft is below 25,000 feet pressure altitude.

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PRESSURIZATION INDICATOR The pressurization indicator consists of a circular CABIN ALT instrument graduated from -1000 to 20,000 feet, a circular CABIN RATE instrument graduated from 2000 feet per minute down to 2000 feet per minute up, and a digital readout to display differential pressure. All three components of the indicator require electrical power. If power to the indicator is lost, the CABIN ALT and CABIN RATE needles will go to the OFF position and the DIFF PRESS display will go blank. The DIFF PRESS readout is capable of displaying differential pressure from 0.0 to 9.9 psid. If the differential pressure exceeds the maximum of 9.8 psid, the display will flash. If the differential pressure exceeds 0.5 psid negative, the DIFF PRESS readout will flash 0.5. The indicator provides outputs for the following: 8750 (250) feet cabin altitude Illuminates PRESS SYS caution light if in the manual mode. Activates cabin altitude aural warning horn and red CABIN ALT HI light at: 10,100 (250) feet cabin altitude whenever the aircraft is above 25,000 feet pressure altitude. 10,100 (250) feet cabin altitude if the aircraft is below 25,000 feet pressure altitude and the system detects takeoff or landing at a field elevation less than 8000 feet. 14,500 (250) feet cabin altitude if the aircraft is below 25,000 feet pressure altitude and the system detects takeoff or landing at a field elevation greater than 8000 feet. 14,500 (250) feet cabin altitude Activates automatic deployment of passenger oxygen masks and turns on cabin overhead lighting. Differential pressure exceeds - 0.5 or + 9.8 psid Illuminates PRESS SYS caution light. PRESS SYS LIGHT The amber PRESS SYS caution light, on the glareshield annunciator panel, illuminates to annunciate the following conditions: Differential pressure has exceeded the limit (- 0.5 to + 9.8 psid). In automatic mode cabin altitude exceeds: 14,500 (250) feet if the aircraft is below 25,000 feet pressure altitude and the system detects takeoff or landing at a field elevation greater than 8000 feet. 8600 (200) feet for all other conditions. In manual mode cabin altitude exceeds 8750 (250) feet. The pressurization system detects a fault.
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EMER PRESS LIGHT The amber EMER PRESS caution light, on the glareshield annunciator panel, illuminates to annunciate the following conditions: The emergency pressurization has activated on one or both sides. If emergency pressurization has not activated, an electrical fault exists which may prevent activation of emergency airflow. BLEED AIR SWITCHES EMER FUNCTION The L and R BLEED AIR switches may be used to manually activate emergency pressurization. When a BLEED AIR switch is set to EMER, the respective bleed-air shutoff valve will close and emergency pressurization valve will be energized open and the high-stage bleed air will be shut off. To reset the emergency pressurization valve, reduce power on the respective engine and set the BLEED AIR switch to OFF. CABIN ALTITUDE WARNING HORN and MUTE FUNCTION A cabin altitude aural warning horn will sound to alert the crew to a problem with the cabin pressurization system. The horn is controlled by an output from the cabin pressurization indicator which activates the warning horn circuit (see pressurization indicator). The cabin altitude warning horn circuit is tested through the SYSTEM TEST switch on the instrument panel. The MUTE switch, on right thrust lever knob, may be used to interrupt the horn for approximately 60 seconds in the event the horn sounds. CABIN ALT HI LIGHT A red CABIN ALT HI light will illuminate in conjunction with the cabin altitude warning horn. SYSTEM TEST SWITCH CABIN ALT FUNCTION The rotary-type SYSTEM TEST switch on the instrument panel is used to test the cabin altitude warning system. Rotating the switch to CABIN ALT and depressing the switch TEST button will provide a ground simulating the 10,100-foot trigger signal.

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AIR CONDITIONING AND HEATING

Primary heating and cooling is accomplished by controlling the temperature of the bleed air entering the independently controlled cockpit and cabin air distribution systems. An R-134A vapor cycle cooling system is installed to provide additional cooling. An auxiliary (electrical) heating system is installed to provide additional heating for the cabin, if desired. PRIMARY HEATING AND COOLING-BLEED AIR Cockpit and cabin temperature is regulated by controlling the temperature of the pressurization bleed air entering the cockpit and cabin air distribution systems. With the BLEED AIR switches ON and the CAB AIR switch ON, engine bleed air is admitted to the ram air heat exchanger through a flow control valve. The bleed air is cooled in the heat exchanger by ram air entering the dorsal inlet, passing through the exchanger, and then exiting overboard. The conditioned bleed air then passes out of the exchanger into the cockpit and cabin air distribution ducts. The temperature of the conditioned air is controlled by the temperature control valve on each distribution system duct. These valves bypass some of the bleed air around the heat exchanger and mix it directly with the conditioned air exiting the heat exchanger. Temperature control valve position, thus, temperature regulation, is pneumatically controlled by the electrically operated temperature control system. Whenever either cabin or cockpit temperature AUTOMAN switch is set to AUTO, the respective system temperature controller will automatically maintain the temperature set with the (CREW or CABIN) COLD-HOT selector. The cabin temperature AUTO-MAN switch also has a CABIN position which allows the temperature to be set using a temperature control panel in the cabin area. The controllers maintain the selected temperature by comparing input signals from various temperature sensors and then electrically controlling the torque motors that provide pneumatic pressure (servo air) to the temperature control valves. Duct temperature sensors are installed in each system to close the temperature control valves and light the DUCT OV HT caution light whenever excessively high duct temperatures are sensed. The cockpit and cabin air temperature sensors have small blowers that draw air past the sensing elements to assure rapid sensing of temperature changes.

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FOOTWARMER

FOOTWARMER

PEDESTAL DIFFUSER PEDESTAL OUTLETS

COCKPIT EVAPORATOR

OVERHEAD OUTLET

OVERHEAD OUTLET

SHOULDER OUTLET

SHOULDER OUTLET

COCKPIT CABIN

FLOOR DIFFUSER

OVERHEAD DIFFUSER FLOOR DIFFUSER OVERHEAD OUTLETS

OVERHEAD DIFFUSER

AUX HEATER

AUX HEATER

CABIN EVAPORATOR

CABIN BLOWER CABIN TAILCONE

FROM HEAT EXCHANGER CHECK VALVE

FROM FROM HEAT HEAT EXCHANGER EXCHANGER VARIABLE OPENING AIR OUTLET

F6006000000301

AIR DISTRIBUTION SCHEMATIC Figure 6-11 6-38

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TORQUE MOTOR SKIN TEMPERATURE SENSOR
CABIN COLD
AUTO C A B I N MAN

HEAT EXCHANGER

CABIN TEMPERATURE CONTROL

HOT

CABIN TEMP CONTROL VALVE

CABIN TEMPERATURE SENSOR

CABIN FAN

CABIN
TORQUE MOTOR

COCKPIT
SKIN TEMPERATURE SENSOR
CREW

TEMP CONTROL

CREW 4 CABIN 4

COCKPIT TEMP CONTROL VALVE

COCKPIT TEMPERATURE CONTROL

COLD

HOT

AUTO

TEMPERATURE CONTROL SCHEMATIC Figure 6-12

COCKPIT TEMPERATURE SENSOR SUPPLY PRESSURE REGULATOR RAM AIR CONDITIONED AIR BLEED AIR CHECK VALVE

MAN

COCKPIT FAN

ELECTRICAL

Pilots Manual

HIGH-PRESSURE BLEED AIR (SERVO AIR)

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Whenever MAN mode is selected with either system AUTO-MAN switch, temperature control valve position is controlled by rotating the CREW or CABIN COLD-HOT selector switch. The rheostat type switch will vary the input current to the affected torque motor to pneumatically position the temperature control valve. Duct overheat protection is provided in this mode also. Power for the temperature control circuits is 28 VDC supplied through the AUTO TEMP CONT circuit breaker on the copilots circuit breaker panel (AUTO mode), and the MANUAL TEMP CONTROL circuit breaker on the pilots circuit breaker panel (MAN mode). CAB AIR SWITCH The CAB AIR switch, on the copilots switch panel, controls the flow control valve. With the BLEED AIR switches ON, setting the CAB AIR switch ON will de-energize the flow control valve controlling solenoid and allow system pressure to the valves controlling chambers. Internal pressures will position the valve shutoff sleeve, controlling bleed-air flow to the heat exchanger. Setting the CAB AIR switch OFF will energize the valve control solenoid which will shutoff control pressure and allow the valve shutoff sleeve to block bleed-air flow. CREW AUTO-MAN SWITCH An AUTO-MAN mode switch is located below the CREW COLD-HOT selector on the copilots switch panel. The switch provides automatic or manual mode operation for the cockpit temperature control system. When AUTO is selected, the cockpit temperature controller will automatically position the cockpit temperature control valve (through inputs to the torque motor) to maintain the temperature set on the CREW COLD-HOT selector. When MAN is selected, cockpit temperature control valve position is controlled directly from the CREW COLD-HOT selector. CABIN AUTO-CABIN-MAN SWITCH An AUTO-CABIN-MAN switch is located below the CABIN COLDHOT selector on the copilots switch panel. The switch provides automatic, automatic remote, and manual mode selection for the cabin temperature control system. When AUTO is selected, the cabin temperature control will automatically position the cabin temperature control valve (through inputs to the torque motor) to maintain the temperature set on the CABIN COLD-HOT selector above the AUTOMAN switch. The CABIN mode operates identical to AUTO except that the temperature is set using a remote temperature selector in the cabin. When MAN is selected, cabin temperature control valve position is controlled directly from the CABIN COLD-HOT selector on the copilots switch panel. 6-40
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CREW AND CABIN COLD-HOT SELECTOR SWITCHES A CREW COLD-HOT and a CABIN COLD-HOT selector switch are located on the copilots switch panel and a remote temperature selector is located in the cabin. In system AUTO mode, these switches are used to select the desired system temperature to be maintained automatically by the temperature controllers. In MAN mode, these rheostat type switches directly vary the current input to the pneumatic torque motors which position the temperature control valves. Rotating the switches clockwise from COLD to HOT is equivalent to selecting temperatures ranging from 60F (16C) to 90F (32C). When CABIN is selected on the cabin AUTO-CABIN-MAN switch, a remote selector switch in the cabin can be used to select the desired cabin temperature. TEMP CONTROL INDICATOR A TEMP CONTROL indication, located on the EIS Electrical Page, provides the crew with a visual indication of the position of the crew and cabin temperature control valves. The indication ranges from 0 at full cold to 9 at full hot. Each TEMP CONTROL indication is controlled by an externally mounted potentiometer on each temperature control valve. The potentiometers are mechanically linked to the duct airflow control flappers. They operate on 28 VDC supplied through the TEMP CONTROL IND circuit breaker on the pilots circuit breaker panel. CAB TEMP INDICATOR The CAB TEMP indication, located on the EIS Electrical Page, provides the crew with indication of cabin temperature in C.

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R-134A COOLING SYSTEM The R-134A vapor cycle cooling system is installed for cockpit and cabin cooling during ground operations, inflight cooling, and cabin dehumidification. On the ground, power must be supplied by an engine generator, APU or ground power unit. In flight, the air conditioning system must be powered by both engine generators. When the COOLOFF switch is set to COOL, power is supplied to the compressor motor and the system refrigerant is compressed and circulated under high pressure through a receiver/dehydrator (dryer) to the cockpit and cabin evaporators. A cockpit blower, located below the cockpit floor, and a cabin blower, located in the aft cabin overhead, circulate air through the system evaporators to provide cooling. Also, pressurization bleed air is used to provide airflow through the cabin evaporator. The system is protected against overpressure conditions by two separate safety devices. The first is a binary high/low pressure switch located on the compressor discharge port. This switch will open at approximately 350 psig and will interrupt power to the compressor control circuit. This in turn will de-energize the compressor motor relay and remove power to the compressor motor. The system pressure will then drop. The switch will also interrupt power to the compressor control circuit under low pressure conditions. This low pressure switch may shut down the compressor if the average refrigerant temperature between the cabin and tailcone is 35F (1.7C) or less. The second overpressure safety device is a fuse plug located on the receiver /dehydrator bottle. This plug will vent the system refrigerant safely overboard in the event of a system pressure in excess of 425 psig. The compressor motor is automatically cut out during engine start, STAB WING HEAT operation, and inflight when only one generator is operating. When the aircraft is on external power, the compressor motor is powered by 28 VDC supplied through a 175-amp current limiter connected to the battery charging bus and a power contactor. When the generators are operating, the compressor motor is powered by 28 VDC supplied through two power contactors and two 175-amp current limiters connected to the generator buses. A fault isolator will remove power from the compressor motor should a fault occur which causes the compressor load to become unequally shared between the generators (except during single generator operation on the ground).

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System control circuits, including the cabin blowers, are powered by 28 VDC supplied through the COOL CONTROL circuit breaker on the pilots circuit breaker panel. The cabin blowers are powered by 28 VDC through a 50-amp current limiter. Speed control circuits for the cabin blowers are powered through the CABIN FAN circuit breaker on the copilots circuit breaker panel. The cockpit blower (including speed control circuit) is powered by 28 VDC through the CREW FAN circuit breaker on the copilots circuit breaker panel.

REFRIGERANT COOLING SYSTEM Figure 6-13

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CABIN CLIMATE SWITCHES COOL-OFF SWITCH The COOL-OFF switch, located in the CABIN CLIMATE group on the copilots switch panel, controls the freon cooling system. When set to COOL, the switch allows power to the freon compressor motor and cabin and cockpit blower circuits. If either the CREW or CABIN FAN switch is off when the switch is set to COOL, the respective blower, cockpit or cabin, will run at minimum speed. Blower speed may be increased by rotating the CREW or CABIN FAN switch, as applicable, in a clockwise direction until the desired speed is reached. CABIN FAN SWITCH Cabin blower speed is controlled during cooling and supplemental air circulation modes by the rheostat-type CABIN FAN switch located in the CABIN CLIMATE group on the copilots switch panel. Rotating the switch clockwise out of the off detent position will turn on the cabin blowers and blower speed will increase with further clockwise movement. Power must be supplied by an engine generator, ground power unit or APU. During pressurized flight (CAB AIR switch ON), cabin cooling is accomplished by pressurization airflow through the cabin evaporator. CREW FAN SWITCH The rheostat-type CREW FAN switch is located in the CABIN CLIMATE group on the copilots switch panel. The switch controls the cockpit blower which is available for all ground and inflight cooling or air circulation modes. When the cooling system is in operation, the blower will force air through the cockpit evaporator to provide cooling or circulate air when the air circulation mode is selected. Air circulated by the cockpit blower is exhausted through the cockpit and cabin overhead eyeball outlets when they are rotated to the open position. HOURMETER COMPRESSOR An hourmeter may be installed in the tailcone compartment to measure accumulated compressor usage time. The hourmeter is activated whenever the compressor motor is running. There is no separate circuit breaker installed with this installation.

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AUXILIARY HEATING SYSTEM An auxiliary heating system is installed to provide additional cabin and cockpit heating when desired. The COOL-OFF switch must be set to the OFF position in order to operate the cabin auxiliary heater. Power must be supplied by an engine generator, APU, or ground power unit. The AUX HT switch, on the copilots switch panel, is used to control the system. The auxiliary heater control circuit is wired through the start cutout relay; therefore, the system is inoperable during engine start. CABIN AUXILIARY HEAT The cabin auxiliary heat is provided by two heater assemblies located in the cabin left and right overhead diffusers. The system utilizes the cabin blower to provide air circulation. The heater assemblies incorporate several thermostatic controls to cycle the heaters at approximately 170 F. The thermostatic controls of each heater are connected in series to each other; therefore, cycling of each heater occurs simultaneously. The cabin blower will start when either heater warms to approximately 75 F. An overheat monitor is installed to monitor the temperature of both heaters. If either heater exceeds approximately 300 F or a switching failure occurs, both heaters will be disabled. Maintenance action is required when the overheat monitor disables the system. Each heater incorporates a thermofuse which will melt and disconnect electrical power to that heater should an overheat condition occur. The system control circuit operates on 28 VDC supplied through the AUX CABINCREW HEAT circuit breaker on the copilots circuit breaker panel. The heater assemblies are supplied 28 VDC through two 50-amp current limiters. Operation of the cabin heaters is only available if the CAB AIR switch is OFF. During pressurized flight (CAB AIR switch ON), cabin heating is accomplished by pressurization airflow. COCKPIT FLOORBOARD HEATERS The cockpit floorboard heater system provides direct contact heat for crew foot warming. There are four heaters, one located beneath each rudder pedal. Each heater contains two heater blankets and a temperature limiting circuit which controls temperature between 100F and 130F independently of the other three heaters. When the temperature of a heater reaches 103F, a relay will remove power to the two heater blankets causing them to cool. The cockpit floorboard heater is controlled through the use of the AUX HT switch. The system control circuit operates on 28 VDC supplied through the AUX CABIN-CREW HEAT circuit breaker on the copilots circuit breaker panel.

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AUX HT SWITCH The auxiliary heating system is controlled through the use of the AUX HT switch located in the CABIN CLIMATE group on the copilots switch panel. The switch has three positions: OFF, CREW and CAB & CREW. With the switch in the CAB & CREW position, the cabin heaters and blower will energize to provide cabin heat and the cockpit floorboard heaters (if applicable) will energize to provide cockpit heat. With the switch in the CREW position, only the cockpit floorboard heaters will be energized. TAILCONE BAGGAGE COMPARTMENT HEATER SYSTEM Tailcone baggage compartment heat is provided to keep the tailcone baggage compartment temperature between 35F and 50F. The BAGGAGE HEAT switch is located in the tailcone baggage compartment and is normally left in the ON position at all times. There is also a baggage heat switch located on the copilots circuit breaker panel. The tailcone baggage heater elements are activated when either external power is connected, or at least one engine-driven generator is powering the electrical system, and the tailcone baggage heater switch is in the ON position. The tailcone baggage heaters are powered by 28 VDC through a 50-amp current limiter.

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SECTION VII INTERIOR EQUIPMENT

TABLE OF CONTENTS Cockpit Description .................................................................................. 7-1 Cockpit Seats .......................................................................................... 7-2 Cockpit Seat (Figure 7-1).................................................................. 7-4 General Arrangement Cockpit (Figure 7-2) ............................. 7-5 Cabin Description...................................................................................... 7-7 Passenger Seats ...................................................................................... 7-7 Passenger Seat (Figure 7-3).............................................................. 7-8 Emergency Equipment ............................................................................. 7-9 Cabin Baggage Compartment Smoke Detection............................... 7-9 Smoke Goggles....................................................................................... 7-9 Hand Fire Extinguisher......................................................................... 7-9 Protective Breathing Equipment........................................................ 7-10 Normal Operation .......................................................................... 7-10 Abnormal Condition of Operation............................................... 7-14 Flotation Equipment............................................................................ 7-15 Miscellaneous Equipment ...................................................................... 7-16 Crew Compartment............................................................................. 7-16 Flashlights ........................................................................................ 7-16 Crew Work Table............................................................................. 7-17 Checklist Holder ............................................................................. 7-17 Sunvisor............................................................................................ 7-17 Passenger Compartment ........................................................................ 7-18 Cabinets, Drawers & Tables ............................................................... 7-18 Galley Cabinet ................................................................................. 7-18 Galley Switch Panel (Figure 7-4) .................................................. 7-20 Galley Cabinet (Figure 7-5) ........................................................... 7-21 Forward Left-Hand Cabinet.......................................................... 7-22 Forward Left-Hand Cabinet (Figure 7-6) .................................... 7-22 Cabin Control Switch Panel (Figure 7-7)..................................... 7-23 Entry Switch Panel (Figure 7-8) .................................................... 7-23 Forward Right-Hand Cabinet ....................................................... 7-24 Forward Right-Hand Cabinet (Figure 7-9).................................. 7-24 Pyramid Cabinets .......................................................................... 7-25 Pyramid Cabinets (Figure 7-10) .................................................... 7-25 Sidewall Storage Boxes .................................................................. 7-26 Executive Tables .............................................................................. 7-26 Executive Table Installation (Figure 7-11) ................................... 7-26
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TABLE OF CONTENTS (Cont) Passenger Entertainment System...................................................... 7-27 Stereo System .................................................................................. 7-27 CD and DVD Players (Figure 7-12).............................................. 7-28 Master Control Switch Panel (Figure 7-13)................................. 7-28 Passenger Control Switch Panel (Figure 7-14) ........................... 7-29 Video System................................................................................... 7-29 Airshow System.............................................................................. 7-30 Airshow Flight Deck Controller (Figure 7-15) ........................... 7-30 Remote Cabin Temperature Control................................................. 7-31 Cabin Temperature Control Panel (Figure 7-16) ............................. 7-31 Iridium SATCOM System .................................................................. 7-32 Dataport ................................................................................................ 7-32 AC Outlets ............................................................................................ 7-32 Window Shades ................................................................................... 7-32 Gasper Outlets ..................................................................................... 7-32 Cabin Baggage Compartment ........................................................... 7-33 Lavatory/Vanity .................................................................................. 7-33 Vanity (Figure 7-17)........................................................................ 7-34 Vanity Switch Panel (Figure 7-18) ................................................ 7-35 Toilet ................................................................................................. 7-35

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SECTION VII INTERIOR EQUIPMENT

COCKPIT DESCRIPTION
The instrument panel is readable by either crew member and the pedestal is accessible and readable by either crew member. Circuit breaker panels are located on the cockpit sidewalls. A magnetic compass is installed on the windshield center post. No switches (except dome light switches), instruments, or placards are located overhead. The pilots and copilots seats are adjustable forward, aft, and vertically. Life vest storage, in some installations, is provided behind each crew seat. On other installations, the life vests are installed in a pouch assembly added to the front of the crew seats. The pilots and copilots rudder pedals are adjustable forward and aft. A curtain, located behind the crew, may be closed for privacy or to darken the cockpit. A handheld fire extinguisher is installed on the bulkhead behind each crew station at approximately shoulder height. A certificate holder is located just aft of the pilots station. Air outlets are installed in each sidewall just aft of the armrest, in each kickplate adjacent to the outboard rudder pedals, on the front side of the center pedestal, and in the headliner above each crew station. An ashtray and drink holder is installed on each side just forward of the circuit breaker panels. Storage is provided as follows: pouches installed on the underside of the glareshield on each side, pouches attached to the lower part of each circuit breaker panel, Jeppesen-size manual holders located at the forward lower edge of each circuit breaker panel, checklist holders located on the side of the pedestal at each crew station, and storage compartments attached to each sidewall outboard of each crew seat. Oxygen masks will be stored in a stowage cup just aft of the pilot and copilots seat or in an accessible compartment just aft of the pilots and copilot s circuit breaker panel. A crew member PBE (protective breathing equipment) is stored in a box accessible to the crew (typically on the aft end of the pedestal). Map lights are installed in each sidewall above the circuit breaker panels and dome lights are installed in the headliner on each side. A work table is installed above the circuit breaker panels at each crew station. Each table hinges enabling it to be stowed against the sidewall when not in use. Sunvisors are installed in tracks at the upper edge of the windshield at each crew station and pull-out extensions are available at the outboard corners of the glareshield. An assist handle, installed overhead, provides a handhold for improved cockpit access.

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COCKPIT SEATS The cockpit seats (figure 7-1) are comprised of two basic structures; the upper structure containing the controls to adjust the headrest, recline, and lumbar support and the base structure containing the controls to adjust the thigh pad, seat height and seat horizontal position. The seat belt system inertia reel is attached to the rear of the seat back. The seat belt reel lock is located on the outboard side of the seat, below and to the rear of the armrest. To lock the seat belt reels, push the reel lock handle down. For automatic reel control, move the reel handle up. The lap and crotch strap are mounted on the seat pan. Seat height adjustment is accomplished by pressing a button on the height lock handle on the outboard side of the seat. When the button is pressed and handle pulled up, the seat will raise. When the button is pressed and the handle pushed down, the seat will lower. Release the button at the desired height to lock the seat into place. Seat tracking is made with the track handle on the inboard side of the seat. Moving the handle aft will allow the seat to be moved forward and aft as desired. Release the track handle to lock the seat track into place. The headrest may be adjusted for angle by moving the headrest to the right and rotating it to one of eight possible lock positions. The back cushion/lumbar support adjustment is controlled by two handwheels, one on each side of the seat. The handwheel on the outboard side of the seat controls the up/down movement, the inboard handwheel controls the in/out movement. Full up/down movement of the back cushion is obtained within 3 1/2 turns of the handwheel and full in/out movement of the back cushion is obtained within 2 3/4 turns of the handwheel. The armrests are padded and can be individually adjusted. Each armrest has an adjusting knob at the forward end of the arm. When either knob is turned counterclockwise, the armrest will lower. When either knob is turned clockwise, the armrest raises. The armrests can be folded back and pushed in towards the seat spine to facilitate entry and exit to the seat. Slide the armrest out and rotate down for use.

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Thigh support pad adjustment is accomplished by turning the thigh pad adjusting handwheel located on the inboard, center section of the seat pan. Rotate the knob forward to raise the thigh pads, and rotate it backward to lower them. When the seat occupant uses the foot controls, thus putting pressure on the thigh pads, tension springs within the linkages are overridden allowing either thigh pad to be pushed downwards. When the thigh pad pressure is released the thigh pads return to their pre-set position. The recline control lever is located on the outboard side of the seat below the lumbar support adjustment. Seats may be reclined to a maximum of 35.

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NOTE: Pilots seat shown. On the

copilots seat, seat height, recline control, inertia reel, track lock, and thigh pad controls are on the opposite side.

COCKPIT SEAT (TYPICAL) Figure 7-1

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1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. PM-133

Pedestal & Throttle Quadrant Ankle Air Outlet Pilots Circuit Breaker Panel Copilots Circuit Breaker Panel Pilots JEPP Storage Cabinet Copilots JEPP Storage Cabinet Shoulder Air Outlet Oxygen Controls & Mic/Phone Jack Panel Foldout Work Table Map Light Assist Handle Sunvisor Dome Light

14. 15. 16. 17. 18. 19. 20. 21. 22. 23. 24. 25.

Overhead Air Outlet Pilots Control Column & Wheel Copilots Control Column & Wheel Instrument Panel Magnetic Compass Annunciator Panel Copilots Mic/Phone Jack Panel Cockpit Phone Cockpit Speakers Flashlight Pedestal Air Outlet APU Control Panel

GENERAL ARRANGEMENT - COCKPIT Figure 7-2

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Pilots Manual

CABIN DESCRIPTION The aircraft cabin is divided into three areas: the passenger area, the lavatory, and the cabin baggage compartment. Access to the baggage compartment may be accomplished through the cabin or through the emergency exit/baggage door on the right side of the fuselage. The lavatory is located in the aft cabin immediately forward of the baggage compartment. Individual reading lights, air outlets, and passenger oxygen masks are located in the overhead convenience panels above the seats. PASSENGER SEATS Lap belts are included in each passenger seat (figure 7-3). Optional shoulder harnesses for three-point latching is available. Passenger seats do not have break-over backs. A life vest is stowed in a pocket under each seat bottom. Access is through a panel on the front of the seat above the storage drawer. Passenger seats can be swiveled 360 but normal aircraft installation is limited to 180. Seats have lateral tracking on the seat base which allows them to be positioned as far outboard as possible for take-off and landing, thus maintaining maximum aisle clearance. Seat tracking or swivel is accomplished by lifting on the inboard release handle on the inboard armrest. Optional floortracking is accomplished by lifting on the release handle near the base of the seat. Passenger seat backs may be reclined to a maximum of 30 with a mechanical button on the outboard armrest. The optional berthing position is available which allows the seat to go full flat. Seats certified for aft facing take-offs and landing will be equipped with hidden bread board headrests which can be pulled up for use or stowed into the top of the seat. Inboard armrests may be moved down by pulling up slightly on the armrest and allowing it to lower. Outboard armrests have an optional feature to be stowed as well. Armrest(s) may be raised and locked into place by pulling the armrest up until it clicks into place. Armrests may be either up or down for take-off and landing. Do not sit on the armrests since this could cause damage to the internal latching device.

CAUTION

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Storage drawers may be located below each seat and are accessed by pulling the knob on the drawer. These drawers are held shut by friction latches at the back of the drawer. Passenger seats may be equipped with a recliner-style. When desired, the footrest can be pulled out for use. Fire blocking of seat cushions is an optional feature to meet FAR Part 25 requirements. Passenger seats may include an optional mechanical lumbar support adjustment knob on the outboard side of the seat back. Rotating the knob forward moves the lumbar support outward thus providing lower back support.

PASSENGER SEAT (TYPICAL) Figure 7-3 7-8

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EMERGENCY EQUIPMENT
CABIN BAGGAGE COMPARTMENT SMOKE DETECTION A baggage area smoke detection system is installed to provide the crew with visual warning of a possible fire in the cabin baggage compartment. The system receives power from the 3-amp CABIN FIRE DETECT circuit breaker on the copilots circuit breaker panel. If the smoke detector, located in the aft cabin baggage area, senses smoke in the aft cabin baggage or lavatory area, a signal is transmitted to an amplifier which will illuminate the red CABIN FIRE light on the glareshield annunciator panel. When the smoke clears, the light will extinguish. The cabin smoke detection system is operative during EMER BUS mode. Self test of the smoke detector is accomplished by pressing the annunciator light test switch. Illumination of the CABIN FIRE light indicates a successful self test. SMOKE GOGGLES Smoke goggles are provided for each crew member and are stowed in sidewall compartments just below the flashlight holder. The goggles must be donned should smoke or fumes be present in the aircraft. Refer to the AFM for the specific procedures. HAND FIRE EXTINGUISHER Halon 1211 fire extinguishers are installed for cockpit and cabin fire protection. The fire extinguishers, in some installations, are attached to the bulkheads just behind each crew station at approximately shoulder height. On other installations, the fire extinguishers may be attached just aft of the pedestal in the cockpit area. A fire extinguisher is also located next to the lavatory seat under the arm rest. The extinguishers incorporate a pressure gage which indicates the state of propellant charge. If properly charged, the indicator needle will be within the green segment. When an extinguisher has been manually discharged, the indicator will be in the red area. This provides the crew with visual indication that the bottle has been partially or totally discharged. The bottle takes approximately 10 seconds to fully discharge. The extinguishers are rechargeable.

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PROTECTIVE BREATHING EQUIPMENT Protective breathing equipment (PBE) is available for a crew member to use in fighting cabin fires. The PBE is designed to protect the users eyes and respiratory system from the harmful atmosphere which may be generated by a cabin fire. The PBE is a hood with a visor which is placed over the head and seals around the neck. An oxygen-generating canister provides breathing oxygen for the user. The PBE is vacuum sealed in a bag and stored in a box accessible to the crew. The PBE is a throw-away unit that must be replaced whenever the vacuum seal has been broken. It is imperative that the vacuum seal be maintained since the oxygen-generating chemicals react with moisture. Duration of oxygen production is nominally 15 minutes depending upon the work rate and size of the user. Useful life of a sealed PBE is 10 years from date of manufacture. NORMAL OPERATION Donning the PBE: There are two available carriers for the PBE. A portable container stored in a cabinet behind the cockpit or a mounted container (normally mounted to the aft side of the pedestal). 1. Removing mask from container. a. To open the portable container, lift the single latch on the cover and lift. Remove sealed bag from the container. b. On the mounted container, grasp the red access handle on the protective container firmly and pull forcible to disengage the cover. When the cover is removed from the container, immediately drop it. (The vacuum sealed bag does not need to be removed from the container to open.) The packaged unit may be removed from the stowage container prior to opening and carried to a remote location for use. 2. To remove the PBE from the vacuum sealed bag, locate the red I.D. tag and pull sharply to tear open the vacuum sealed bag. Reach into the opened vacuum-sealed bag and firmly grasp the PBE. Pull the PBE straight out of the bag. If necessary hold the bag with the opposite hand. 3. Place both hands inside the neckseal opening with palms facing each other and PBE visor facing downward with the oxygengenerating canister resting on the tip of the hands. 4. With the head bent forward, guide the PBE neckseal over the top of the head and down over the face using the hands to shield the face and glasses from the oronasal mask cone.

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5. With both hands, grasp the adjustment straps at the lower corners of the visor and pull outward sharply to actuate the starter candle. Within 1-5 seconds, a rushing noise of oxygen entering the hood will be heard and inflation will be evident.
WARNING

Human hair is highly flammable. Hair that protrudes through the neckseal could ignite if brought into direct contact with flame.

6. With the straps still in hand and head bent forward, pull backward to secure the oronasal mask cone high on the nose for a tight seal. 7. If wearing glasses, you may adjust their position to rest on tip of the oronasal mask cone by moving the sides of the frame through the hood fabric. Do not attempt to adjust through the neckseal as this will result in infiltration of the surrounding atmosphere into the interior of the hood. 8. When the neckseal is positioned at the neck and the oxygengenerating canister is resting on the nape of the neck, remove the hands, checking to see that clothing is not trapped in the seal and hair does not protrude between the seal and the neck. Pull the protective neck shield down to cover the collar and upper shoulder area.

STEP 1
Grasp red access handle and pull forcibly to disengage the cover. Locate red I.D. tag and pull sharply to tear open the vacuum-sealed bag. PM-133

STEP 2
Pull PBE out of the vacuum-sealed bag and shake hood open.

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STEP 3
Place both hands inside the neckseal opening with palms facing each other and PBE visor facing downward with the canister resting on tip of hands.

STEP 4
With the head bent forward, guide the PBE neckseal over the tip of the head and down over the face using the hands to shield the face and glasses from oronasal mask cone.

STEP 5
With both hands, grasp the adjustment straps at the lower corners of the visor and pull outward sharply to actuate the starter candle.

STEP 6
With the straps still in hand and head bent forward, pull backward to secure the oronasal mask cone high on the nose for a tight seal.

STEP 7
If wearing glasses, you may adjust their position to rest on top of the oronasal mask cone by moving the sides of the frame through the hood fabric. Do not attempt to adjust through the neckseal as this will result in infiltration of the surrounding atmosphere into the interior of the hood.

STEP 8
When the neckseal is positioned at the neck and the canister is resting on the nape of the neck, remove the hands, checking to see that clothing is not trapped in the seal and hair does not protrude between the seal and the neck. Pull the protective neck shield down to cover the collar and upper shoulder area. PM-133

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Following actuation, the hood will inflate over a 15-20 second period. After this period, the starter candle will cease flowing and the only sound will be slight rustling of the fabric on each inhalation and exhalation. Dependent upon breathing rate, there will be a slight exhalation resistance as the exhaled breath is forced through the oxygen-generating canister. Inhalation resistance will be almost unrecognizable since inhalation is directly from the interior of the hood through a diaphragm type check valve located at the base of the oronasal mask. The visor should remain clear of fogging or misting. Heat is produced by both the chemical air regeneration process and transfer of body heat during the rebreathing cycle. Heat build-up within the hood is normal and is dependent upon the amount of work performed. There should be no irritating or strong unusual odors within the hood. Operational duration is variable dependent upon the amount of work performed by the user. If the PBE is worn to exhaustion of the chemical regeneration system, this will be evidenced by a gradual reduction in the expended volume of the hood until the point that the hood is collapsed tightly around the head at the end of a full inhalation. Additionally, there will be a rapid buildup of heat and moisture in the hood as the canister looses its effectiveness. At this point, the wearer should immediately retire to a safe breathing area clear of flame and toxic fumes and remove the device. Removing the PBE: 1. Go to a safe area away from immediate contact with fire or open flame and/or toxic fumes. 2. With both hands, reach for the two lower corners of the visor area and push forward on the metal tabs of the adjustment strap buckles to release the strap tension. 3. Place both hands under the neckseal in forward area and pull up, guiding the oronasal cone and neckseal over the face/ glasses until the PBE is clear of the head. 4. Place the expended PBE in a safe place to cool away from fire or exposure to water. Disposal: The expended PBE still contains unreacted oxidizing material and strong alkali materials. At the completion of flight, it must be turned over to maintenance for authorized disposal.

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ABNORMAL CONDITION OF OPERATION

CAUTION

This device produces oxygen which will vigorously accelerate combustion. Do not intentionally expose the device to direct flame contact, or remove in the immediate presence of fire or flame. Due to oxygen saturation of the hair, do not smoke or become exposed to fire or flame immediately after removing.

Users should be trained to recognize abnormal conditions which could signify malfunction or failure of the equipment to properly operate. Failure of the starter candle: If the starter candle fails to actuate when the adjustment strap is pulled, an additional sharp pull on the strap may be sufficient to dislodge the lanyard pin and actuate the device. If the device still fails to actuate, the hood will continue to function, although the initial purge capability is lost. Sticking the fingers into the neckseal to allow a large lung inhalation may be required to enable sufficient breathing volume until the chemical regeneration system begins producing a surplus of oxygen. Inadequate oronasal mask seal: Absence of a tight seal of the oronasal cone to the face may result in excess leakage of the exhaled breath into the hood, short circuiting the oxygen-generating canister. This condition may result in a build-up of CO2 within the rebreathing volume in the hood. Excessive CO2 is normally indicated by breathing distress such as rapid and labored breathing accompanied by a general feeling of insufficient ability to get ones breath, although there is no restriction to breathing. Presence of moisture or fogging on the visor and the sensation of air escaping from the mask, particularly around the nose and eyes are indications of a lack of proper fit. Adjustment of the mask straps and mask position to minimize leakage should rapidly alleviate the problem. If the perception of breathing distress persists, the user should quickly go to a safe area and remove the PBE and don alternate breathing equipment if required.

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Loss of infiltration seal: The smoke and toxic fumes generated by the combustion of most aircraft cabin interior materials has many strong irritants. The continued presence of strong irritation odors inside the hood resulting in eye and respiratory tract discomfort is a good indicator of the lack of an effective infiltration seal. Verify that the seal is in contact with the skin or the neck and does not have clothing or jewelry trapped in the seal, or hair protruding between the seal and the neck. If the condition persists, or there is evidence of a tear in the neckseal, the user should quickly go to a safe area and remove the PBE and don alternate breathing equipment if required. FLOTATION EQUIPMENT Pilots and copilots life vests are either stowed in a pocket on the pilots and copilots seat back or in a pouch assembly on the front of the pilots and copilots seats. Life vests in the passenger cabin are stowed in a compartment under each passenger/cabin seat. There is also a life vest stowed in the armrest next to the aft lavatory toilet seat. The life vests are inflated by pulling the red CO2 release tabs.

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MISCELLANEOUS EQUIPMENT
CREW COMPARTMENT FLASHLIGHTS Flashlights are located on the Jeppesen storage units next to the pilots and copilots seats. The rechargeable flashlights are waterproof, flame retardant, and floatable. The rechargeable flashlights must be properly placed in the retention bracket to ensure their recharging. Ensure the D ring is properly secured into the flashlight end cap. Place the head end of the light against the retaining disc at the top end of the bracket with the switch toward the bracket and the small red LED light facing out. Once the head of the flashlight is positioned, snap the butt of the flashlight into the clips at the bottom of the bracket. When the flashlight is recharging, the LED light should be on. To remove the flashlight from the bracket, grasp and pull the lower end of the light out of the bracket clips. Do not install the flashlight into the recharging base while the flashlight is still turned on since recharging and lamp life would be significantly reduced. The lamp inside the flashlight may need to be changed after approximately 20 hours of service. To change the lamp, unscrew the head of the light and remove the lens cap and reflector assembly. Remove the lamp from the reflector by unscrewing the threaded plastic retainer. Insert the new lamp and replace the retainer. Be sure to reinstall the spacer/washers to retain its highly focused lighting ability. Do not touch the shiny surface of the reflector or the glass portions of the lamp. If the reflector surface requires cleaning, use only a soft, dry cloth. The useful life per charge of the flashlight is approximately 45 minutes and requires about 16 hours to recharge after a full battery depletion. Leaving the flashlight on constant charge in extreme temperatures (below 30F and above 100F) could affect the useful life of the battery pack. The flashlights recharge only when an aircraft battery switch(es) is turned on. The power source for the recharging base, if installed, is 28 VDC from the FLASH LTS circuit breaker on the copilots circuit breaker panel.

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CREW WORK TABLE A fold down work table, with hinged leaf, is located in the outboard panel adjacent to each pilots seat. The table is folded out of its compartment by the available finger hold at the top edge of the panel compartment. Unfold its leaf for use. To stow the table fold the leaf up and push the table back into its compartment. CHECKLIST HOLDER A one-piece checklist holder is installed on the floor on each side of the forward pedestal. It can hold the checklist and prevent it from becoming displaced during flight. SUNVISOR Each pilot has a sunvisor located at the upper edge of the windshield. Each sunvisor is hinged so that it can be folded down and slid along its track as desired. Some aircraft may have pull-out extensions available at the outboard corners of the glareshield.

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PASSENGER COMPARTMENT
CABINETS, DRAWERS & TABLES Standard and optional cabinets, drawers and tables may be built into the passenger compartment. Due to the wide variety of options available, the following descriptions and figures show only the most common accessories. Power for the cabinet kicker lights and cabin aisle lights is 28 VDC from the AISLE LTS circuit breaker on the pilots circuit breaker panel. GALLEY CABINET The galley cabinet (figure 7-5) has storage cabinets and drawers accessible through press-to-open buttons on the cabinet doors and drawers. There is a galley work light controlled by the galley work light switch located on the galley switch panel. Power for the galley work light is 28 VDC from the TABLE LTS circuit breaker on the pilots circuit breaker panel. Internal galley lights are actuated by micro-switches in the cabinet doors. Power for the Internal galley lights is 28 VDC from the CABIN LTS circuit breaker on the pilots circuit breaker panel. Top galley cabinet contains one 1.5 gallon (5.71) or two .66 gallon (2.51l) vented, stainless steel, removable liquid dispenser containers. This insulated container incorporates a heating element along the bottom and is automatically plugged into a power source when installed in the cabinet. An over-temperature sensor and a thermostat is built in, which will keep even small amounts of liquid warm without burning the container. The lighted On/Off liquid warmer switch(es) are located on the galley switch panel. With at least one battery switch on, and a warmer switch pressed ON, the switch will illuminate and the warmer will keep already hot liquids between 150 and 170F. Power to these warmers is 28 VDC from the HOT CUP circuit breaker on the pilots circuit breaker panel. The liquid warmer container(s) can also be controlled from the Cabin Control Switch Panel. When aircraft power is cycled the hot liquid container(s) will turn off and the switch(es) will have to be selected to on when power is restored to the aircraft.

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The container is removed by opening the top cabinet doors and pulling down the dispenser button panel located in the upper section of the cabinet. The dispenser button panel is held into place with ball-catches. Remove the dispenser by pulling it straight out from the cabinet. The containers can be drained through the screw on/off cap on the top of the unit, by pressing the spigot and allowing fluids to drain, or unscrewing the outside spigot ring and removing the spigot. The container is filled through the top cap. To reinstall the container, ensure the cap is screwed on tightly, and push the container completely into the cabinet, thus connecting the heating element to its power source. Flip the dispenser button panel over the spigot outlets before closing the top and middle cabinet doors. The warmers are not able to heat cold liquids to very warm temperatures. Before installing the dispenser in the airplane, and to aid in sustaining hot liquids, it is recommended that very hot water be poured into the container. Install the lid and allow the container to preheat for approximately 15 minutes. Drain the hot water and add whatever hot beverage is desired. If desired, cold liquids may be available by not turning on the applicable warmer. To serve liquids from the dispenser, position a cup under the desired liquid dispenser. Press the dispenser button which, in turn presses the spigot drain. A drip pan below the dispenser outlets will catch small amounts of overflow. The top galley cabinet also contains door-mounted glass storage racks, two disposable cup holders mounted horizontally immediately above the liquid dispensers and a large general storage area below the hot liquid container(s). Slide-out drawers for storage and a divided ice drawer are located in the lower galley cabinet. Drainage for the ice drawer and the galley drip pan is provided through a drain valve on the underside of the cabinet. To open the drain press the drain position on the galley switch panel. The water will drain out through the forward cabinet drain mast. The drain mast is heated to prevent ice build up around the drain hole. The drain will only remain open while the switch is depressed. Power to the galley drain is 28 VDC from the GALLEY DRN circuit breaker on the pilots circuit breaker panel.

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The middle compartment is available for storage or an optional microwave oven. Power for the microwave oven is 28 VDC from a 50 amp current limiter located in the tailcone. The MICROWAVE circuit breaker on the pilots circuit breaker panel controls a relay which will remove power from the microwave oven. The left compartment is available for storage or an optional warming oven. A lighted On/Off warming oven switch is located on the galley switch panel. With the warming oven switch pressed on (illuminated) power is sent to the warming oven. Power for the warming oven is 28 VDC from the OVEN circuit breaker on the pilots circuit breaker panel. A wine storage unit in this cabinet is located at the center outboard edge of the worktop. There is a pull-out trash container and a pull-out work surface on the forward side of the galley. No cigarettes, matches, or otherwise flammable materials, should be discarded in the trash container.

GALLEY SWITCH PANEL (TYPICAL) Figure 7-4

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Galley Switch Panel General Storage Glass Storage General Storage Drawer Divided Ice Drawer Clean & Dirty Ice Compartments Microwave (Optional) Pocket Door General Storage (Standard) General Soda/Water Bottle Storage

Dual Cup Dispenser

Dual Hot Liquid Containers (Optional) Single Hot Liquid Container (Standard)

Power Outlet 110V

Wine Bottle Storage Pull-out Work Surface and Pull-out Trash Container

GALLEY CABINET (TYPICAL) Figure 7-5

Work Surface

General Storage Drawer

General Storage/Catering Tray Storage

Pilots Manual

Storage Compartment (Standard) Warming Oven (Optional)

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FORWARD LEFT-HAND CABINET The forward left-hand cabinet (figure 7-6) has mini liquor storage, PBE storage and a closet with a coat rod accessible through press-to-open buttons on the cabinet doors. On the inboard upper side of the cabinet is the cabin control switch panel (figure 7-7) and on the aft side of the cabinet is the entry switch panel (figure 7-8).

Cabin Control Switch Panel PBE Storage

Mini Liquor Storage

Entry Switch Panel

FORWARD LEFT-HAND CABINET Figure 7-6

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CABIN CONTROL SWITCH PANEL Figure 7-7

ENTRY SWITCH PANEL Figure 7-8

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FORWARD RIGHT-HAND CABINET The forward right-hand cabinet (figure 7-9) has a closet accessible through press-to-open button on the cabinet door. On the aft side of the cabinet is the infrared eye which receives commands from the remote control and the optional 15.1 inch Liquid Crystal Display (LCD) video monitor.
Infrared Eye

Optional 15.1" Monitor

FORWARD RIGHT-HAND CABINET Figure 7-9 7-24

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PYRAMID CABINETS Optional pyramid cabinets (figure 7-10) may be located behind the individual cabin seats against the forward and/or aft bulkhead. Access is by pressing the button at the top, center section of the door/drawer panel. The cabinet door opens outward for miscellaneous storage.

PYRAMID CABINETS (TYPICAL) Figure 7-10 7-25

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SIDEWALL STORAGE BOXES Headphones, as well as other items, may be stored in the outboard sidewall storage boxes located along the cabin armrests. EXECUTIVE TABLES Pull-out executive tables (figure 7-11) are available in the sidewall between the aft and forward facing seat locations. The table is tilted away from the wall, pulled up and then the leaf unfolded for use.

EXECUTIVE TABLE INSTALLATION (TYPICAL) Figure 7-11

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PASSENGER ENTERTAINMENT SYSTEM STEREO SYSTEM An audio signal is supplied to speakers on both sides of the cabin and to individual passenger switch panel headphone jack from a ten disc CD changer located in the vanity (figure 7-12). There is a master control switch panel, located in the cabin armrest (figure 7-13), which incorporates lighting, cabin speaker, audio select, video select (if installed) and remote cabin temperature controls. There are also passenger control switch panels, located in the cabin armrests adjacent to the passenger seats (figure 7-14), which incorporate lighting, headphone volume control, audio select controls, and a headphone jack. Press the Cabin Audio position on the master control switch panel or the cabin control switch panel to change to the cabin audio control panel. The cabin audio control panel is used to select the desired audio source (e.g., CD, DVD), turn the cabin speakers on and off, and to control the volume, bass and treble settings for the speakers. Each passenger location has a passenger control switch panel that may be used to select individual audio source, volume, bass and treble settings for use with headphones. Power for the stereo system is 28 VDC from the STEREO circuit breaker on the pilots circuit breaker panel.Power for the video system is 28 VDC from the VIDEO circuit breaker on the pilots circuit breaker panel. Power to operate the audio distribution module and audio digital selectors is 28 VDC from the CABIN AUDIO circuit breaker on the copilots circuit breaker panel. Power for the passenger speakers is 28 VDC from the PASS SPKR circuit breaker on the copilots circuit breaker panel. Keying the passenger address or passenger briefing system will automatically override any cabin stereo channel, including overhead speakers that have been turned off by the cabin control switch panel or the cabin master control switch panel. Passenger address and passenger briefings are transmitted over cabin speakers and headphone jacks.

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CD Controller DVD Player

10 Disc CD Changer

CD and DVD PLAYERS Figure 7-12

MASTER CONTROL SWITCH PANEL Figure 7-13 7-28

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PASSENGER CONTROL SWITCH PANEL Figure 7-14 VIDEO SYSTEM Optional 15.1 inch Liquid Crystal Display (LCD) video monitors may be installed in conjunction with a single or dual DVD player installed in the vanity (figure 7-12) and/or an Airshow system. The optional monitors are installed in either the forward right-hand cabinet facing aft and/or the aft right-hand partition facing forward. The video monitors and the DVD player receive 28 VDC from a VIDEO circuit breaker on the pilots circuit breaker panel. Press the Cabin Video position on the master control switch panel or the cabin control switch panel to change to the cabin video control panel. The cabin video control panel is used to select the desired video source (e.g., DVD1, DVD2, AIRSHOW) and turn the cabin LCD video monitors on and off. Press the Cabin audio position on the master control switch panel or the cabin control switch panel to change to the cabin audio control panel. The cabin audio control panel is used to select the audio source corresponding to the selected video source.

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AIRSHOW SYSTEM An optional Airshow system may be installed which allows passengers to be informed of flight status without interrupting the pilots, in addition to other pertinent inflight information. The unit interfaces with FMS-1 and can display customized modes of operation. The Airshow system receives 28 VDC from the PASS INFO circuit breaker on the copilots circuit breaker panel. For additional information, reference the Airshow Operators Manual. Pressing the Video position on the Cabin Control Switch Panel (located on the inboard top side of the left forward cabinet) or the Cabin Master Control Switch Panel (located in the cabin armrest) will cause that Switch Panel to change to the Video control panel. From this control panel the monitors are switched on and off and the video source for each monitor is selected. Pressing the Airshow Mode position on the Cabin Control Switch Panel or the Master Control Switch Panel will cause that Switch Panel to change to the AIRSHOW control panel. The various modes of the Airshow display are accessed from the Airshow control panel. The Airshow has an optional Flight Deck Controller (figure 7-15) which has an display with a push button SELECT switch and a SCROLL knob. The controller can be used to enter time to destination, Greenwich Mean Time, and the destination airport identifier. For a detailed description of the Airshow system refer to the current Airshow operators manual.

SCROLL

SELECT

AIRSHOW FLIGHT DECK CONTROLLER Figure 7-15

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REMOTE CABIN TEMPERATURE CONTROL A remote cabin temperature control (figure 7-16) is located on the Cabin Control Switch Panel (located on the inboard top side of the left forward cabinet) and on the Cabin Master Control Switch Panel (located in the cabin armrest). When the AUTO-CABIN-MAN switch located below the CABIN HOTCOLD selector on the copilots switch panel is set to CABIN, control for cabin temperature is given to the cabin control switch panel. Pressing the Cabin temp position on the Cabin Control Switch Panel or the Master Control Switch Panel will cause that Switch Panel to change to the Cabin Temperature control panel. The temperature control panel consists of a bar graph with C at one end and H at the other. The Temp (up) and Temp (down) position are used to raise and lower the setting.

CABIN TEMPERATURE CONTROL PANEL Figure 7-16

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IRIDIUM SATCOM SYSTEM (OPTIONAL) The ICS-100 Iridium SATCOM is a single channel system and the ICS200 Iridium SATCOM is a dual channel system. The SATCOM system consists of a transceiver, handsets, and low profile top mounted antenna. The SATCOM system provides features such as air to air, air to ground, ground to air, call transfer, extension to extension calling, and three party conferencing. The system uses the Iridium Low Earth Orbit (LEO) satellite constellation for global voice and data communications services including the polar regions. A customer selected service provider is identified on the Subscriber Identity Module (SIM) card installed in the transceiver. Power to the Iridium SATCOM system is through a SATCOM circuit breaker on the pilots circuit breaker panel. Refer to the Iridium SATCOM users manual for more detailed instructions on the use of the Iridium SATCOM system. DATAPORT A dataport may be installed in the cabin. The dataport is used in conjunction with the flight phone system to communicate to the internet for e-mails, etc. AC OUTLETS 110 VAC 60 Hz outlets are located inside the storage box at each passenger seat location, for the three place divan there are two outlets located on either side of the center storage compartment in the armrest ledge, and for the two place divan there is one outlet located between the two storage compartments in the armrest ledge. They receive 110 VAC through an AC OUTLETS circuit breaker on the copilots circuit breaker panel. An aneroid switch will disconnect power to the outlets if the cabin altitude should reach 9500 (250) feet. Power will be restored if normal cabin altitude is regained. The maximum load for each outlet is 220 Watts. The optional 220 VAC 50 Hz outlets replace the 110 VAC 60 Hz outlets. WINDOW SHADES Window shades are installed in all passenger compartment windows. The shades can be lowered or raised to any level. The shades are translucent and will not totally block out light. GASPER OUTLETS Individual gasper, or air outlets, are available in the cockpit and in the cabin convenience panels. These outlets may be turned to approximately 40 around its center to direct air flow as desired. Rotate the conical port counterclockwise to open and clockwise to close. 7-32
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CABIN BAGGAGE COMPARTMENT The door to the aft cabin baggage compartment is located in the lavatory. It is a bi-fold door with a recessed, pull-type latch to open and close. When the door is closed and the latch pushed fully in, bolts in the door will engage into the top, bottom, and outboard side of the door jamb thus securing the door. The maximum weight for the cabin baggage compartment is placarded. The cabin baggage compartment door in the Vanity may be accessed through the emergency exit/ baggage door. LAVATORY/VANITY The lavatory is equipped with a toilet and a vanity consisting of a sink, faucet, potable water tank, soap dispenser, tissue holder, trash container, AC outlet, swing out lighted mirror, and storage drawers. The lavatory is separated from the passenger cabin with a sliding door that is stowed and latched on the left-side of the bulkhead. The door is latched open with a recessed latch on the aft-side of the door to a catch in the aft-side of the bulkhead wall. A magnetic strip along the door edge allows the door to be closed but cannot be locked shut. The potable water tank, pump, and heater are located under the sink. The tank itself is in the lavatory aft cabinet below the sink and holds approximately 1.7 gallons (6.4 liters). It is equipped with a quick disconnect shutoff for easy removal and installation. To remove the potable water tank, press the disconnect lever on the plumbing connection and pull it apart from the tubing. Pull the tank straight out from the cabinet. It is recommended that the potable water tank be removed from the aircraft during extended cold weather to prevent the water in the tank from freezing and damaging the tank. For more information on the servicing of the potable water tank, reference the GROUND HANDLING, SERVICING AND EMERGENCY INFORMATION manual. The heater is part of the potable water tank and disconnects electrically when the tank is removed from the cabinet. The potable water tank heater turns on when DC power is applied to the airplane. It increases water temperature to 100F (38C). The water heater receives 28 VDC from the WATER HEATER circuit breaker on the pilots circuit breaker panel. The switch for the water faucet is to the left of the faucet on the lavatory wall. When the switch is pressed a timer starts and the water pump is turned on. Only warm water from the potable water tank is available from the faucet. The water pump receives 28 VDC power through the VANITY DRAIN circuit breaker on the pilots circuit breaker panel.
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The sink is drained by pressing the DRAIN switch located on the vanity switch panel. A green LED on the switch will illuminate while the switch is pressed. The LED will extinguish when the switch is released. The drain switch receives 28 VDC from the VANITY DRAIN circuit breaker on the pilots circuit breaker panel. The water is drained through a heated drain mast on the bottom of the aircraft. The heater is activated through a squat switch and prevents ice from forming on the drain mast.

Lavatory Light Accordian Shade

Window

Mirror

Countertop Ledge

110V Outlet

Vanity Switch Panel Soap Dispenser

Faucet Tissue General Storage Drawer Toilet Paper

Trash Container

Heated Water Container

VANITY Figure 7-17

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VANITY SWITCH PANEL Figure 7-18 TOILET A flushing toilet is installed in the lavatory. This unit features a twocompartment design isolating the flushing fluid from the waste. Raising the lid opens the sealed valve at the bottom of the bowl. Closing the lid automatically flushes the toilet. Length of the flush cycle is controlled automatically. Two electric pumps are used in this unit. The flushing pump circulates the flushing fluid during the flush cycle. The macerator/pump unloads the waste from the toilet during servicing only.
CAUTION

Use only biodegradable toilet paper such as that used in recreational vehicles. Do not use the toilet to dispose of other paper products, cigarettes, sanitary napkins, coffee grounds, etc. The macerator/pump will become clogged with these items making external servicing of the toilet impossible.

Servicing of the toilet is accomplished using servicing ports located on the aircraft exterior. The macerator/pump is used to pump the waste from the toilet while fresh flushing fluid is pumped into the toilet from the servicing equipment. Refer to Chapter 12 in the maintenance manual for servicing instructions. Power to operate the flushing circuit is 28 VDC from the 5-amp TOILET circuit breaker on the pilots circuit breaker panel. Power to operate the servicing circuit is 28 VDC from the 10-amp TOILET SERVICE circuit breaker on the pilots circuit breaker panel. The TOILET SERVICE circuit breaker is powered from the left battery bus; therefore, servicing can be accomplished without turning the battery switches on.

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SECTION VIII FLIGHT CHARACTERISTICS & OPERATIONAL PLANNING

TABLE OF CONTENTS General Flight Characteristics ............................................................... 8-1 Operational Planning .............................................................................. 8-3 Operational Planning Form (Figure 8-1) ............................................ 8-4 Temperature Conversion (Figure 8-2)................................................. 8-5 Linear Conversions (Figure 8-3) .......................................................... 8-6 Volume Conversions (Figure 8-4) ........................................................ 8-7 Weight Conversions (Figure 8-5) ......................................................... 8-8 Relation of Temperature (C) to ISA (Figure 8-6).............................. 8-9 Speed/Temperature Conversion (Figure 8-7).................................. 8-10 Climb Performance ................................................................................ 8-11 Climb Power Setting............................................................................ 8-11 Climb Performance Schedule............................................................. 8-11 Maximum Continuous Thrust for Climb (N1) (Figure 8-8)........... 8-12 Climb Performance Two Engine (Figure 8-9)(12 Sheets) ............ 8-13 Cruise Performance ................................................................................ 8-25 Normal Cruise ...................................................................................... 8-25 Maximum Specific Range ................................................................... 8-25 Maximum Range Cruise Two Engines .......................................... 8-25 Long Range Cruise Two Engines.................................................... 8-25 High Speed Cruise............................................................................... 8-25 Maximum Range Descent One Engine ......................................... 8-26 Long Range Cruise One Engine ..................................................... 8-26 Normal Cruise (Figure 8-10)(10 Sheets)............................................ 8-27 Maximum Specific Range (Figure 8-11)............................................ 8-37 Maximum Range Cruise Two Engines (Figure 8-12)(19 Sheets) 8-38 Long Range Cruise Two Engines (Figure 8-13)(19 Sheets) ......... 8-57 High Speed Cruise (Figure 8-14)(19 Sheets)..................................... 8-76 Maximum Range Descent One Engine (Figure 8-15) .................. 8-95 Long Range Cruise One Engine (Figure 8-16)(5 Sheets) ............. 8-96

hange 1
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TABLE OF CONTENTS (Cont) Descent and Holding Performance .................................................. 8-101 Descent Performance Schedule ....................................................... 8-101 Holding Operations .......................................................................... 8-101 Descent Performance Schedule Normal (Figure 8-17) ..................................................................... 8-102 High Speed (Figure 8-18) .............................................................. 8-103 Holding Operations (Figure 8-19)................................................... 8-104

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SECTION VIII FLIGHT CHARACTERISTICS & OPERATIONAL PLANNING

GENERAL FLIGHT CHARACTERISTICS
Taxi operations can be conducted using one or both engines. If nosewheel steering is inoperative or when taxiing on a slick or icy surface, it is recommended that taxiing be conducted using both engines to preclude aggravating the problem with asymmetric thrust. The digital nose-wheel steering system provides excellent taxi maneuverability. At low ground speeds, nose wheel travel is approximately 60 either side of neutral. The steering authority tapers off as ground speed increases and is reduced to zero at approximately 80 knots. At 90 knots, the system will automatically disengage. The rudder is effective for directional control above 45 KIAS. The two pod-mounted PW305A engines, manufactured by Pratt and Whitney Canada, Inc., are rated at 4600 pounds thrust at sea level. The time required to accelerate these engines from idle RPM to maximum thrust RPM is approximately seven (7) seconds. The engine thrust and acceleration characteristics complement the Learjet 60XR airframe so that outstanding performance, flexibility, and safety margins are available in all flight regimes. Single-engine performance offers an example of these capabilities in that the sea-level single-engine rate of climb at 23,100 pounds is approximately 1,340 feet per minute and the singleengine service ceiling is approximately 31,000 feet at a cruise weight of 19,000 pounds. Although the flight control systems are manual, stick forces are light to moderate throughout the flight envelope. Stability is good at all airspeeds and airplane configurations. Aircraft responsiveness and flight control authority are very good throughout the flight envelope. A yaw damper is employed to damp lateral oscillations caused by turbulent air; however, it is not required for dispatch. Trim changes due to use of the landing gear, flaps and power are slight; however, a trim change is required when spoilers are extended or retracted.

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GENERAL FLIGHT CHARACTERISTICS (Cont)

The dual stall warning system provides an excellent indication of impending airplane stall. Additionally, the airplane exhibits an aerodynamic stall warning buffet in all configurations. The shaker actuates at least 7% above the stall speed published in the Airplane Flight Manual. The shaker system produces a high-frequency, low-amplitude vibration transmitted to the control columns. As the shakers actuate, the red low-speed awareness cue reaches the center of the airspeed display on the EFIS, the angle-of-attack indicator needle enters the yellow arc and the stall warning lights illuminate and flash. Recovery is easily accomplished by lowering the nose of the airplane while simultaneously advancing power as necessary to accelerate out of the stall regime. Good aircraft response, to elevator inputs, occurs throughout the aircraft operating envelope. The spoiler system provides an effective means of increasing normal rates of descent and may be used as a drag device to achieve rapid airspeed deceleration. The spoilers are used just after touchdown to spoil the lift for more effective braking action and to increase drag for minimum landing roll. Aileron augmentation is accomplished by the spoiler system when the SPOILER switch is in the RET or ARM position and the flaps are lowered beyond 25.

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OPERATIONAL PLANNING
The charts and tables on the following pages contain performance data for climb, cruise, descent and holding. Takeoff and landing performance data is presented in tabular form in the FAA Approved Flight Manual. Fuel consumption information is presented based on flight test data and average engine characteristics. The following conditions are to be assumed when extracting data from this section: WEIGHT All weights presented in this section are to be understood as the gross weight of the airplane in pounds. For flight planning, the climb weight used is the gross weight of the airplane at the start of climb, the cruise weight used is the mid-weight between the start cruise weight and the end cruise weight and the descent weight used is assumed to be 16,000 pounds. All altitudes presented in this section are to be understood as pressure altitude in feet. OAT Outside Air Temperature. For presentation in this section, Temperature is to be understood as OAT unless otherwise specified. SAT Static Air Temperature obtained from inflight indications. SAT is equivalent to OAT. RAT Ram Air Temperature obtained from inflight measurement (includes compression rise). FUEL FLOW The fuel flows presented are for two engines except where single-engine performance is specified. The wing flap position for various flight conditions is as follows: Climb.................................................. UP-0 Enroute............................................... UP-0 Holding .............................................. UP-0

ALTITUDE

TEMPERATURE

FLAPS

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OPERATIONAL PLANNING FORM

WEIGHT
ZERO FUEL WEIGHT FUEL LOAD RAMP WEIGHT WARM UP & TAKEOFF
Altitude=
Table 1:

TIME

DISTANCE

FUEL

START CLIMB WEIGHT CLIMB END CLIMB WEIGHT

Altitude=

START CRUISE WEIGHT CRUISE END CRUISE WEIGHT Altitude= START CLIMB WEIGHT CLIMB END CLIMB WEIGHT Altitude= START CRUISE WEIGHT CRUISE END CRUISE WEIGHT Altitude= START CLIMB WEIGHT CLIMB END CLIMB WEIGHT Altitude= START CRUISE WEIGHT CRUISE END CRUISE WEIGHT Altitude= START DESCENT WEIGHT DESCENT END DESCENT WEIGHT Altitude= RESERVES ZERO FUEL WEIGHT

Total 8-4 Figure 8-1

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TEMPERATURE CONVERSION
To convert from Celsius to Fahrenheit, find, in bold face columns, the number representing the Celsius temperature to be converted. The equivalent Fahrenheit temperature is read in the adjacent column headed F. To convert from Fahrenheit to Celsius, find, in bold face columns, the number representing the Fahrenheit temperature to be converted. The equivalent Celsius temperature is read in the adjacent column headed C.
C -73.3 -72.8 -72.2 -71.7 -71.1 -70.6 -70.0 -69.4 -68.9 -68.3 -67.8 -67.2 -66.7 -66.1 -65.6 -65.0 -64.4 -63.9 -63.3 -62.8 -62.2 -61.7 -61.1 -60.6 -60.0 -59.4 -58.9 -58.3 -57.8 -57.2 -56.7 -56.1 -55.6 -55.0 -54.4 -53.9 -53.3 -52.8 -52.2 -51.7 -51.1 -50.6 -50.0 -49.4 -48.9 -48.3 -47.8 -47.2 -46.7 -46.1 F -58.0 -56.2 -54.4 -52.6 -50.8 -49.0 -47.2 -45.4 -43.6 -41.8 -40.0 -38.2 -36.4 -34.6 -32.8 -31.0 -29.2 -27.4 -25.6 -23.8 -22.0 -20.2 -18.4 -16.6 -14.8 -13.0 -11.2 -9.4 -7.6 -5.8 -4.0 -2.2 -0.4 1.4 3.2 5.0 6.8 8.6 10.4 12.2 14.0 15.8 17.6 19.4 21.2 23.0 24.8 26.6 28.4 30.2 -50 -49 -48 -47 -46 -45 -44 -43 -42 -41 -40 -39 -38 -37 -36 -35 -34 -33 -32 -31 -30 -29 -28 -27 -26 -25 -24 -23 -22 -21 -20 -19 -18 -17 -16 -15 -14 -13 -12 -11 -10 -9 -8 -7 -6 -5 -4 -3 -2 -1 C -45.6 -45.0 -44.4 -43.9 -43.3 -42.8 -42.2 -41.7 -41.1 -40.6 -40.0 -39.4 -38.9 -38.3 -37.8 -37.2 -36.7 -36.1 -35.6 -35.0 -34.4 -33.9 -33.3 -32.8 -32.2 -31.7 -31.1 -30.6 -30.0 -29.4 -28.9 -28.3 -27.8 -27.2 -26.7 -26.1 -25.6 -25.0 -24.4 -23.9 -23.3 -22.8 -22.2 -21.7 -21.1 -20.6 -20.0 -19.4 -18.9 -18.3 F 32.0 33.8 35.6 37.4 39.2 41.0 42.8 44.6 46.4 48.2 50.0 51.8 53.6 55.4 57.2 59.0 60.8 62.6 64.4 66.2 68.0 69.8 71.6 73.4 75.2 77.0 78.8 80.6 82.4 84.2 86.0 87.8 89.6 91.4 93.2 95.0 96.8 98.6 100.4 102.2 104.0 105.8 107.6 109.4 111.2 113.0 114.8 116.6 118.4 120.2 0 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48 49 C -17.8 -17.2 -16.7 -16.1 -15.6 -15.0 -14.4 -13.9 -13.3 -12.8 -12.2 -11.7 -11.1 -10.6 -10.0 -9.4 -8.9 -8.3 -7.8 -7.2 -6.7 -6.1 -5.6 -5.0 -4.4 -3.9 -3.3 -2.8 -2.2 -1.7 -1.1 -0.6 0.0 0.6 1.1 1.7 2.2 2.8 3.3 3.9 4.4 5.0 5.6 6.1 6.7 7.2 7.8 8.3 8.9 9.4 F 122.0 123.8 125.6 127.4 129.2 131.0 132.8 134.6 136.4 138.2 140.0 141.8 143.6 145.4 147.2 149.0 150.8 152.6 154.4 156.2 158.0 159.8 161.6 163.4 165.2 167.0 168.8 170.6 172.4 174.2 176.0 177.8 179.6 181.4 183.2 185.0 186.8 188.6 190.4 192.2 194.0 195.8 197.6 199.4 201.2 203.0 204.8 206.6 208.4 210.2 50 51 52 53 54 55 56 57 58 59 60 61 62 63 64 65 66 67 68 69 70 71 72 73 74 75 76 77 78 79 80 81 82 83 84 85 86 87 88 89 90 91 92 93 94 95 96 97 98 99 C 10.0 10.6 11.1 11.7 12.2 12.8 13.3 13.9 14.4 15.0 15.6 16.1 16.7 17.2 17.8 18.3 18.9 19.4 20.0 20.6 21.1 21.7 22.2 22.8 23.3 23.9 24.4 25.0 25.6 26.1 26.7 27.2 27.8 28.3 28.9 29.4 30.0 30.6 31.1 31.7 32.2 32.8 33.3 33.9 34.4 35.0 35.6 36.1 36.7 37.2 F 212.0 213.8 215.6 217.4 219.2 221.0 222.8 224.6 226.4 228.2 230.0 231.8 233.6 235.4 237.2 239.0 240.8 242.6 244.4 246.2 248.0 249.8 251.6 253.4 255.2 257.0 258.8 260.6 262.4 264.2 266.0 267.8 269.6 271.4 273.2 275.0 276.8 278.6 280.4 282.2 284.0 285.8 287.6 289.4 291.2 293.0 294.8 296.6 298.4 300.2 100 101 102 103 104 105 106 107 108 109 110 111 112 113 114 115 116 117 118 119 120 121 122 123 124 125 126 127 128 129 130 131 132 133 134 135 136 137 138 139 140 141 142 143 144 145 146 147 148 149 C 37.8 38.3 38.9 39.4 40.0 40.6 41.1 41.7 42.2 42.8 43.3 43.9 44.4 45.0 45.6 46.1 46.7 47.2 47.8 48.3 48.9 49.4 50.0 50.6 51.1 51.7 52.2 52.8 53.3 53.9 54.4 55.0 55.6 56.1 56.7 57.2 57.8 58.3 58.9 59.4 60.0 60.6 61.1 61.7 62.2 62.8 63.3 63.9 64.4 65.0

F -148.0 -100 -146.2 -99 -144.4 -98 -142.6 -97 -140.8 -96 -139.0 -137.2 -135.4 -133.6 -131.8 -130.0 -128.2 -126.4 -124.6 -122.8 -121.0 -119.2 -117.4 -115.6 -113.8 -112.0 -110.2 -108.4 -106.6 -104.8 -103.0 -101.2 -99.4 -97.6 -95.8 -94.0 -92.2 -90.4 -88.6 -86.8 -85.0 -83.2 -81.4 -79.6 -77.8 -76.0 -74.2 -72.4 -70.6 -68.8 -67.0 -65.2 -63.4 -61.6 -59.8 -95 -94 -93 -92 -91 -90 -89 -88 -87 -86 -85 -84 -83 -82 -81 -80 -79 -78 -77 -76 -75 -74 -73 -72 -71 -70 -69 -68 -67 -66 -65 -64 -63 -62 -61 -60 -59 -58 -57 -56 -55 -54 -53 -52 -51

Figure 8-2
PM-133

8-5

Pilots Manual

LINEAR CONVERSIONS
To convert from meters to feet, find, in the bold face columns, the number of meters to be converted. The equivalent number of feet is read in the adjacent column headed FEET. To convert from feet to meters, find, in the bold face columns, the number of feet to be converted. The equivalent number of meters is read in the adjacent column headed METERS.
METERS 304.8 335.3 365.8 396.2 426.7 457.2 487.7 518.2 548.6 579.1 609.6 640.1 670.6 701.0 731.5 762.0 792.5 823.0 853.5 883.9 914.4 944.9 975.4 1005.9 1036.3 1066.8 1097.3 1127.8 1158.3 1188.7 1219.2 1249.7 1280.2 1310.7 1000 1100 1200 1300 1400 1500 1600 1700 1800 1900 2000 2100 2200 2300 2400 2500 2600 2700 2800 2900 3000 3100 3200 3300 3400 3500 3600 3700 3800 3900 4000 4100 4200 4300 FEET 3280.8 3608.9 3937.0 4265.0 4593.1 4921.2 5249.3 5577.4 5905.4 6233.5 6561.6 6889.7 7217.8 7545.8 7873.9 8202.0 8530.1 8858.2 9186.2 9514.3 9842.4 10170.5 10498.6 10826.6 11154.7 11482.8 11810.9 12139.0 12467.0 12795.1 13123.2 13451.3 13779.4 14107.4 METERS 1341.1 1371.6 1402.1 1432.6 1463.1 1493.5 1524.0 1554.5 1585.0 1615.5 1645.9 1676.4 1706.9 1737.4 1767.9 1798.3 1828.8 1859.3 1889.8 1920.3 1950.7 1981.2 2011.7 2042.2 2072.7 2103.1 2133.6 2164.1 2194.6 2225.1 2255.5 2286.0 2316.5 2347.0 4400 4500 4600 4700 4800 4900 5000 5100 5200 5300 5400 5500 5600 5700 5800 5900 6000 6100 6200 6300 6400 6500 6600 6700 6800 6900 7000 7100 7200 7300 7400 7500 7600 7700 FEET 14435.5 14763.6 15091.7 15419.8 15747.8 16075.9 16404.0 16732.1 17060.2 17388.2 17716.3 18044.4 18372.5 18700.6 19028.6 19356.7 19684.8 20012.9 20341.0 20669.0 20997.1 21325.2 21653.3 21981.4 22309.4 22637.5 22965.6 23293.7 23621.8 23949.8 24277.9 24606.0 24934.1 25262.2 METERS 2377.5 2407.9 2438.4 2468.9 2499.4 2529.9 2560.4 2590.8 2621.3 2651.8 2682.3 2712.8 2743.2 2773.7 2804.2 2834.7 2865.2 2895.6 2926.1 2956.6 2987.1 3017.6 3048.0 3352.8 3657.6 3962.4 4267.3 4572.1 4876.9 5181.7 5486.5 5791.3 6096.1 6400.9 7800 7900 8000 8100 8200 8300 8400 8500 8600 8700 8800 8900 9000 9100 9200 9300 9400 9500 9600 9700 9800 9900 10000 11000 12000 13000 14000 15000 16000 17000 18000 19000 20000 21000 FEET 25590.2 25918.3 26246.4 26574.5 26902.6 27230.6 27558.7 27886.8 28214.9 28543.0 28871.0 29199.1 29527.2 29855.3 30183.4 30511.4 30839.5 31167.6 31495.7 31823.8 32151.8 32479.9 32808.0 36088.8 39369.6 42650.4 45931.2 49212.0 52492.8 55773.6 59054.4 62335.2 65616.0 68896.8

Figure 8-3

8-6

PM-133

Pilots Manual

VOLUME CONVERSIONS
To convert from liters to gallons, find, in the bold face columns, the number of liters to be converted. The equivalent number of gallons is read in the adjacent column headed GALLONS. To convert from gallons to liters, find, in the bold face columns, the number of gallons to be converted. The equivalent number of liters is read in the adjacent column headed LITERS.
LITERS 18.9 37.9 75.7 113.6 151.4 189.3 227.1 265.0 302.8 340.7 378.5 416.4 454.2 492.1 529.9 567.8 605.6 643.5 681.3 719.2 757.0 794.9 832.7 870.6 908.4 946.3 984.1 1022.0 1059.8 1097.7 1135.5 1173.4 1211.2 1249.1 1286.9 1324.8 1362.6 1400.5 1438.3 5 10 20 30 40 50 60 70 80 90 100 110 120 130 140 150 160 170 180 190 200 210 220 230 240 250 260 270 280 290 300 310 320 330 340 350 360 370 380 GALLONS 1.3 2.6 5.3 7.9 10.6 13.2 15.9 18.5 21.1 23.8 26.4 29.1 31.7 34.3 37.0 39.6 42.3 44.9 47.6 50.2 52.8 55.5 58.1 60.8 63.4 66.1 68.7 71.3 74.0 76.6 79.3 81.9 84.5 87.2 89.8 92.5 95.1 97.8 100.4 LITERS 1476.2 1514.0 1551.9 1589.7 1627.6 1665.4 1703.3 1741.1 1779.0 1816.8 1854.7 1892.5 1930.4 1968.2 2006.1 2043.9 2081.8 2119.6 2157.5 2195.3 2233.2 2271.0 2308.9 2346.7 2384.6 2422.4 2460.3 2498.1 2536.0 2573.8 2611.7 2649.5 2687.4 2725.2 2763.1 2800.9 2838.8 2876.6 2914.5 390 400 410 420 430 440 450 460 470 480 490 500 510 520 530 540 550 560 570 580 590 600 610 620 630 640 650 660 670 680 690 700 710 720 730 740 750 760 770 GALLONS 103.0 105.7 108.3 111.0 113.6 116.2 118.9 121.5 124.2 126.8 129.5 132.1 134.7 137.4 140.0 142.7 145.3 148.0 150.6 153.2 155.9 158.5 161.2 163.8 166.4 169.1 171.7 174.4 177.0 179.7 182.3 184.9 187.6 190.2 192.9 195.5 198.2 200.8 203.4 LITERS 2952.3 2990.2 3028.0 3065.9 3103.7 3141.6 3179.4 3217.3 3255.1 3293.0 3330.8 3368.7 3406.5 3444.4 3482.2 3520.1 3557.9 3595.8 3633.6 3671.5 3709.3 3747.2 3785.0 4163.5 4542.0 4920.5 5299.0 5677.5 6056.0 6434.5 6813.0 7191.5 7570.0 7948.5 8327.0 8705.5 9084.0 9462.5 9841.0 780 790 800 810 820 830 840 850 860 870 880 890 900 910 920 930 940 950 960 970 980 990 1000 1100 1200 1300 1400 1500 1600 1700 1800 1900 2000 2100 2200 2300 2400 2500 2600 GALLONS 206.1 208.7 211.4 214.0 216.6 219.3 221.9 224.6 227.2 229.9 232.5 235.1 237.8 240.4 243.1 245.7 248.3 251.0 253.6 256.3 258.9 261.6 264.2 290.6 317.0 343.5 369.9 396.3 422.7 449.1 475.6 502.0 528.4 554.8 581.2 607.7 634.1 660.5 686.9

Figure 8-4 8-7

PM-133

Pilots Manual

WEIGHT CONVERSIONS
To convert from kilograms to pounds, find, in the bold face columns, the number of kilograms to be converted. The equivalent number of pounds is read in the adjacent column headed POUNDS. To convert from pounds to kilograms, find, in the bold face columns, the number of pounds to be converted. The equivalent number of kilograms is read in the adjacent column headed KILOGRAMS.
KILOGRAMS 4.5 9.1 13.6 18.1 22.7 27.2 31.8 36.3 40.8 45.4 49.9 54.4 59.0 63.5 68.0 72.6 77.1 81.6 86.2 90.7 95.3 99.8 104.3 108.9 113.4 117.9 122.5 127.0 131.5 136.1 140.6 145.2 149.7 154.2 158.8 163.3 167.8 172.4 176.9 181.4 186.0 190.5 195.0 199.6 204.1 10 20 30 40 50 60 70 80 90 100 110 120 130 140 150 160 170 180 190 200 210 220 230 240 250 260 270 280 290 300 310 320 330 340 350 360 370 380 390 400 410 420 430 440 450 POUNDS KILOGRAMS 22.0 44.1 66.1 88.2 110.2 132.3 154.3 176.4 198.4 220.5 242.5 264.6 286.6 308.6 330.7 352.7 374.8 396.8 418.9 440.9 463.0 485.0 507.1 529.1 551.1 573.2 595.2 617.3 639.3 661.4 683.4 705.5 727.5 749.6 771.6 793.7 815.7 837.7 859.8 881.8 903.9 925.9 948.0 970.0 992.1 208.7 213.2 217.7 222.3 226.8 231.3 235.9 240.4 244.9 249.5 254.0 258.6 263.1 267.6 272.2 276.7 281.2 285.8 290.3 294.8 299.4 303.9 308.4 313.0 317.5 322.1 326.6 331.1 335.7 340.2 344.7 349.3 353.8 358.3 362.9 367.4 371.9 376.5 381.0 385.6 390.1 394.6 399.2 403.7 408.2 460 470 480 490 500 510 520 530 540 550 560 570 580 590 600 610 620 630 640 650 660 670 680 690 700 710 720 730 740 750 760 770 780 790 800 810 820 830 840 850 860 870 880 890 900 POUNDS 1014.1 1036.2 1058.2 1080.3 1102.3 1124.3 1146.4 1168.4 1190.5 1212.5 1234.6 1256.6 1278.7 1300.7 1322.8 1344.8 1366.9 1388.9 1410.9 1433.0 1455.0 1477.1 1499.1 1521.2 1543.2 1565.3 1587.3 1609.4 1631.4 1653.4 1675.5 1697.5 1719.6 1741.6 1763.7 1785.7 1807.8 1829.8 1851.9 1873.9 1896.0 1918.0 1940.0 1962.1 1984.1 KILOGRAMS 412.8 417.3 421.8 426.4 430.9 435.5 440.0 444.5 449.1 453.6 499.0 544.3 589.7 635.0 680.4 907.2 1134.0 1360.8 1587.6 1814.4 2041.2 2268.0 2494.8 2721.6 2948.4 3175.2 3402.0 3628.8 3855.6 4082.4 4309.2 4536.0 4989.6 5443.2 5896.8 6350.4 6804.0 7257.6 7711.1 8164.7 8618.3 9071.9 9525.5 9979.1 10432.7 910 920 930 940 950 960 970 980 990 1000 1100 1200 1300 1400 1500 2000 2500 3000 3500 4000 4500 5000 5500 6000 6500 7000 7500 8000 8500 9000 9500 10000 11000 12000 13000 14000 15000 16000 17000 18000 19000 20000 21000 22000 23000 POUNDS 2006.2 2028.2 2050.3 2072.3 2094.4 2116.4 2138.5 2160.5 2182.6 2204.6 2425.1 2645.5 2866.0 3086.4 3306.9 4409.2 5511.5 6613.8 7716.1 8818.4 9920.7 11023.0 12125.3 13227.6 14329.9 15432.2 16534.5 17636.8 18739.1 19841.4 20943.7 22046.0 24250.6 26455.2 28659.8 30864.4 33069.0 35273.6 37478.2 39682.8 41887.4 44092.0 46296.6 48501.2 50705.8

Figure 8-5 8-8

PM-133

Pilots Manual

RELATION OF TEMPERATURE (C) TO ISA

-50C 51
-106.5

-40C
-96.5

-30C
-86.5

-20C
-76.5

-10C
-66.5

ISA
-56.5

+10C
-46.5

+20C
-36.5

+30C
-26.5

37 35 33 31 30 29 28 27 ALTITUDE ~ 1000 FT 26 25 24 23 22 21 20 19 18 16 14 12 10 8 6 4 2 S.L.

-106.5 -104.2 -100.3 -96.3 -94.4 -92.4 -90.4 -88.4 -86.5 -84.5 -82.5 -80.5 -78.6 -76.6 -74.6 -72.6 -70.6 -66.7 -62.7 -58.8 -54.8 -50.8 -46.9 -42.9 -39.0 -35.0

-96.5 -94.2 -90.3 -86.3 -84.4 -82.4 -80.4 -78.4 -76.5 -74.5 -72.5 -70.5 -68.6 -66.6 -64.6 -62.6 -60.6 -56.7 -52.7 -48.8 -44.8 -40.8 -36.9 -32.9 -29.0 -25.0

-86.5 -84.2 -80.3 -76.3 -74.4 -72.4 -70.4 -68.4 -66.5 -64.5 -62.5 -60.5 -58.6 -56.6 -54.6 -52.6 -50.6 -46.7 -42.7 -38.8 -34.8 -30.8 -26.9 -22.9 -19.0 -15.0

-76.5 -74.2 -70.3 -66.3 -64.4 -62.4 -60.4 -58.4 -56.5 -54.5 -52.5 -50.5 -48.6 -46.6 -44.6 -42.6 -40.6 -36.7 -32.7 -28.8 -24.8 -20.8 -16.9 -12.9 -9.0 -5.0

-66.5 -64.2 -60.3 -56.3 -54.4 -52.4 -50.4 -48.4 -46.5 -44.5 -42.5 -40.5 -38.6 -36.6 -34.6 -32.6 -30.6 -26.7 -22.7 -18.8 -14.8 -10.8 -6.9 -2.9 1.0 5.0

-56.5 -54.2 -50.3 -46.3 -44.4 -42.4 -40.4 -38.4 -36.5 -34.5 -32.5 -30.5 -28.6 -26.6 -24.6 -22.6 -20.6 -16.7 -12.7 -8.8 -4.8 -0.8 3.1 7.1 11.0 15.0

-46.5 -44.2 -40.3 -36.3 -34.4 -32.4 -30.4 -28.4 -26.5 -24.5 -22.5 -20.5 -18.6 -16.6 -14.6 -12.6 -10.6 -6.7 -2.7 1.2 5.2 9.2 13.1 17.1 21.0 25.0

-36.5 -34.2 -30.3 -26.3 -24.4 -22.4 -20.4 -18.4 -16.5 -14.5 -12.5 -10.5 -8.6 -6.6 -4.6 -2.6 -0.6 3.3 7.3 11.2 15.2 19.2 23.1 27.1 31.0 35.0

-26.5 -24.2 -20.3 -16.3 -14.4 -12.4 -10.4 -8.4 -6.5 -4.5 -2.5 -0.5 1.4 3.4 5.4 7.4 9.4 13.3 17.3 21.2 25.2 29.2 33.1 37.1 41.0 45.0

-50C

-40C

-30C

-20C

-10C

ISA

+10C

+20C

+30C

Figure 8-6

PM-133

8-9

Pilots Manual

SPEED/TEMPERATURE CONVERSION
MACH TRUE .60 0 -5 -10 -15 -20 -25 SAT KTAS SAT KTAS SAT KTAS SAT KTAS SAT KTAS SAT KTAS SAT KTAS SAT KTAS SAT KTAS SAT KTAS SAT KTAS SAT KTAS SAT KTAS SAT KTAS SAT KTAS SAT KTAS SAT KTAS SAT KTAS -18 374 -23 370 -27 367 -32 363 -37 359 -41 356 -46 352 -51 349 -55 345 -60 341 -65 338 -69 334 -74 330 -79 326 -83 323 -88 318 -93 314 -97 310 .62 -19 385 -24 381 -28 378 -33 374 -38 371 -42 367 -47 363 -52 359 -56 356 -61 352 -66 348 -70 344 -75 340 -80 336 -84 332 -89 328 -94 323 -98 320 .64 -20 397 -25 293 -30 389 -34 386 -39 382 -43 378 -48 374 -53 370 -57 366 -62 362 -67 358 -71 355 -76 350 -80 347 -85 342 -90 338 -94 334 -99 329 .66 -21 408 -26 404 -31 400 -35 397 -40 393 -45 389 -49 385 -54 381 -58 377 -63 373 -68 369 -72 365 -77 360 -81 357 -86 352 -91 347 -95 343 -100 339 .68 -23 419 -27 416 -32 412 -36 408 -41 404 -46 399 -50 396 -55 391 -59 387 -64 383 -69 379 -73 375 -78 370 -82 366 -87 362 -91 358 -96 353 -101 348 .70 -24 431 -28 427 -33 423 -38 418 -42 415 -47 410 -51 406 -56 402 -60 398 -65 394 -70 389 -74 385 -79 380 -83 376 -88 371 -92 367 -97 362 -101 358 .72 -25 442 -30 438 -34 434 -39 429 -43 425 -48 421 -52 417 -57 413 -62 408 -66 404 -71 399 -75 395 -80 390 -84 386 -89 381 -93 377 -98 371 -102 367 .74 -26 453 -31 449 -36 445 -40 440 -45 436 -49 432 -54 427 -58 423 -63 418 -67 414 -72 409 -76 405 -81 400 -85 396 -90 390 -94 386 -99 381 -103 376 .76 -28 464 -32 460 -37 455 -41 451 -46 447 -50 443 -55 438 -59 434 -64 429 -68 424 -73 419 -77 415 -82 410 -86 405 -91 400 -95 395 -100 390 -104 385 .78 -29 475 -34 470 -38 466 -42 462 -47 457 -51 453 -56 448 -60 444 -65 439 -69 434 -74 429 -78 425 -83 419 -87 415 -92 409 -96 405 -101 399 -105 394 .80 -30 486 -35 481 -39 477 -44 474 -48 468 -53 463 -57 458 -62 453 -66 449 -70 444 -75 439 -79 435 -84 429 -88 424 -93 419 -97 414 -102 408 -106 403 .82 -32 496 -36 492 -41 487 -45 483 -49 478 -54 473 -58 468 -63 463 -67 459 -72 453 -76 449 -80 444 -85 438 -89 434 -94 428 -98 423 -102 418 -107 412

C
RAM AIR TEMPERATURE (RAT)
-30 -35 -40 -45 -50 -55 -60 -65 -70 -75 -80 -85

Figure 8-7

8-10

PM-133

Pilots Manual

CLIMB PERFORMANCE
CLIMB POWER SETTING Figure 8-8 presents the climb maximum continuous thrust settings. At the start of the climb, the thrust levers are moved to the Maximum Continuous Thrust (MCT) position. When airborne with the flaps up, the FADEC will determine the proper maximum continuous thrust N1 and position the N1 bug to that value. The N1 needle should align with the N1 bug.

CLIMB PERFORMANCE SCHEDULE Figure 8-9 shows time, distance and fuel used to climb from sea level to altitude for standard and off-standard days at various weights. The climb weight used is the start-of-climb weight. Subtraction of performance values for two altitudes results in the time, distance and fuel required for climb between the two altitudes. The climb speed schedule presented with each table is based upon an operational climb schedule to optimize fuel consumption and approximates the best rate-of-climb speeds. The climb speeds given are 250 KIAS up to 32,000 feet and 0.70 MI above 32,000 feet. Climb thrust is maximum continuous thrust (MCT).

PM-123

8-11

Pilots Manual

MAXIMUM CONTINUOUS THRUST FOR CLIMB (N1) ALL ENGINE

ALTITUDE - 1000 FEET
S.L. 55 50 45 40 35 30 25
88.23 86.07 88.34 86.19 89.07 86.94 89.80 87.69 90.51 88.41 91.22 89.14 91.99 89.92 92.77 90.72 92.55 91.50 91.74 91.74 90.93 90.93 90.11 90.11 89.28 89.28 88.44 88.44 87.60 87.60 86.75 86.75 85.88 85.88 85.02 85.02 84.14 84.14 83.25 83.25 82.35 82.35 81.44 81.44 80.53 80.53 79.60 79.60 89.11 86.84 89.78 87.53 90.48 88.24 91.18 88.96 91.90 89.70 92.63 90.45 93.37 91.20 94.10 91.96 94.83 92.71 95.56 93.45 94.95 94.20 94.06 94.06 93.16 93.16 92.25 92.25 91.34 91.34 90.41 90.41 89.48 89.48 88.53 88.53 87.58 87.58 86.61 86.61 85.64 85.64 84.65 84.65 90.42 87.95 91.12 88.67 91.81 89.39 92.49 90.08 93.16 90.78 93.88 91.52 94.60 92.25 95.31 92.99 96.02 93.72 96.73 94.45 97.44 95.19 98.10 95.86 97.31 96.50 96.32 96.32 95.33 95.33 94.32 94.32 93.31 93.31 92.28 92.28 91.24 91.24 90.19 90.19 91.80 89.16 92.43 89.82 93.03 90.44 93.67 91.11 94.37 91.82 95.06 92.54 95.75 93.26 96.45 93.97 97.14 94.69 97.82 95.39 98.40 96.00 99.02 96.64 100.21 97.86 92.96 90.12 93.49 90.67 94.10 91.30 94.71 91.94 95.38 92.64 96.07 93.35 96.76 94.07 97.45 94.79 98.06 95.43 98.64 96.03 99.57 96.99 93.73 90.64 94.20 91.14 94.69 91.65 95.39 92.38 96.09 93.11 96.79 93.84 97.48 94.56 98.07 95.18 98.66 95.80 99.87 97.04 94.34 91.02 94.83 91.54 95.48 92.22 96.13 92.90 96.78 93.59 97.41 94.25 98.01 94.88 98.78 95.68 95.08 91.54 95.69 92.18 96.30 92.83 96.90 93.47 97.51 94.11 98.11 94.75 99.30 95.98 95.79 91.99 96.40 92.64 97.01 93.29 97.62 93.94 98.82 95.17 94.29 90.27 94.91 90.93 95.52 91.58 96.72 92.83 98.00 94.14 98.36 94.55 98.57 94.81 98.79 95.07 99.00 95.33 93.13 88.90 93.75 89.57 94.37 90.23 95.57 91.48 96.85 92.80 97.22 93.22 97.44 93.48 97.66 93.75 97.88 94.02 93.12 88.86 93.74 89.52 94.36 90.19 95.57 91.44 96.85 92.77 97.21 93.18 97.44 93.45 97.66 93.72 97.88 93.98
60-096

10

15

20

25

30

35

40

45

50

51

STATIC AIR TEMPERATURE C

20 15 10 5 0 -5 -10 -15 -20 -25 -30 -35 -40 -45 -50 -55 -60 -65 -70

101.41 100.79 99.09 98.23

101.68 101.52 101.10 100.03 99.38 99.00 98.30 96.97

101.35 101.71 101.39 101.05 100.57 100.08 99.59 99.21 98.62 98.02 97.28 96.48 100.21 101.90 101.59 101.24 100.92 100.44 99.80 99.43 98.85 98.25 97.66 96.87 99.05 99.05 101.00 101.62 101.44 101.12 100.65 99.65 99.07 98.49 97.90 97.12 100.42 101.18 101.32 100.86 99.30 98.72 98.15 97.38 99.96 98.95 100.36 100.36 98.39 97.63

XX.XX ANTI-ICE OFF XX.XX FULL ANTI-ICE ON

8-12

Figure 8-8

SPEED SCHEDULE 250 KIAS up to 32,000 ft .70 MI above 32,000 ft

PM-123

PRESSURE ALTITUDE 1000 FEET

PM-133 ISA +10C ISA +15C ISA +20C Fuel Time Dist Fuel Time Dist Fuel Time Dist Fuel Lb Min. N.M. Lb Min. N.M. Lb Min. N.M. Lb 509.6 416.5 19.9 128.9 516.3 379.2 14.5 91.7 440.7 18.4 118.6 502.3 24.2 159.8 592.7 353.3 12.2 75.7 403.7 14.5 91.5 445.0 17.2 110.8 495.1 332.4 10.7 65.6 377.2 12.4 77.2 410.8 14.4 91.3 451.3 314.4 9.6 58.3 355.6 11.0 67.6 385.0 12.7 79.2 420.5 297.9 8.8 52.4 336.5 10.0 60.5 363.2 11.4 70.3 395.4 282.6 8.1 47.6 319.0 9.1 54.6 343.5 10.4 63.2 373.1 268.5 7.5 43.6 303.0 8.4 49.8 325.8 9.6 57.5 353.2 254.5 6.9 39.8 286.9 7.8 45.3 308.0 8.8 52.1 333.3 237.3 6.3 35.2 266.6 7.0 40.0 285.5 7.9 45.7 308.0 219.9 5.6 30.9 246.2 6.3 35.0 263.1 7.1 39.8 282.8 203.3 5.1 27.2 227.0 5.6 30.7 242.1 6.3 34.7 259.5 187.4 4.6 24.0 208.7 5.1 26.9 222.2 5.6 30.3 237.5 171.9 4.1 21.0 191.0 4.5 23.5 202.9 5.0 26.4 216.3 156.6 3.6 18.3 173.6 4.0 20.4 184.1 4.4 22.9 196.0 141.6 3.2 15.9 156.6 3.5 17.7 165.8 3.9 19.7 176.3 126.7 2.8 13.6 139.8 3.1 15.2 147.9 3.4 16.9 157.1 111.8 2.4 11.6 123.3 2.7 12.9 130.3 2.9 14.3 138.2 97.1 2.1 9.7 106.8 2.3 10.7 112.9 2.5 11.9 119.5 82.3 1.7 7.9 90.4 1.9 8.8 95.5 2.1 9.7 101.0 67.5 1.4 6.2 74.1 1.5 6.9 78.2 1.6 7.7 82.6 52.8 1.0 4.7 57.8 1.1 5.2 60.9 1.3 5.7 64.2 38.0 0.7 3.3 41.4 0.8 3.6 43.6 0.9 4.0 45.9 23.0 0.4 1.9 25.0 0.5 2.1 26.2 0.5 2.3 27.5 7.8 0.1 0.6 8.4 0.2 0.7 8.8 0.2 0.7 9.2

WEIGHT 14,000 LB

CLIMB PERFORMANCE TWO ENGINE

Figure 8-9 (Sheet 1 of 12)

51 49 47 45 43 41 39 37 35 33 31 29 27 25 23 21 19 17 15 13 11 9 7 5 3 1

ISA -10C ISA Time Dist Fuel Time Dist Min. N.M. Lb Min. N.M. 14.4 89.4 418.7 20.4 130.5 11.2 68.0 371.2 13.6 85.0 9.4 56.7 343.0 11.2 68.8 8.3 49.0 321.4 9.7 58.8 7.4 43.2 303.3 8.6 51.6 6.7 38.6 287.2 7.8 46.0 6.1 34.7 272.4 7.1 41.5 5.6 31.4 258.4 6.5 37.6 5.2 28.7 245.5 6.0 34.4 4.8 26.1 232.8 5.6 31.4 4.3 23.3 217.4 5.1 27.9 3.9 20.6 201.7 4.6 24.7 3.5 18.2 186.7 4.2 21.8 3.2 16.1 172.4 3.7 19.3 2.9 14.2 158.3 3.4 17.0 2.6 12.5 144.5 3.0 14.9 2.3 10.9 130.9 2.7 13.0 2.0 9.5 117.4 2.3 11.2 1.8 8.1 103.9 2.0 9.5 1.5 6.9 90.4 1.7 8.0 1.3 5.7 76.9 1.4 6.5 1.0 4.6 63.4 1.2 5.2 0.8 3.5 49.7 0.9 3.9 0.6 2.5 35.9 0.6 2.7 0.4 1.5 21.9 0.4 1.6 0.1 0.5 7.4 0.1 0.5

Pilots Manual

8-13

CLIMB SPEED:

250 KIAS up to 32,000 feet. 0.70 MI above 32,000 feet.

PRESSURE ALTITUDE 1000 FEET

8-14
Fuel Lb

Pilots Manual

WEIGHT 15,000 LB

ISA +10C Time Dist Fuel Min. N.M. Lb

ISA +15C Time Dist Fuel Min. N.M. Lb

ISA +20C Time Dist Fuel Min. N.M. Lb

CLIMB PERFORMANCE TWO ENGINE

Figure 8-9 (Sheet 2 of 12)

51 49 47 45 43 41 39 37 35 33 31 29 27 25 23 21 19 17 15 13 11 9 7 5 3 1 475.7 419.5 386.8 362.0 341.1 322.6 305.5 290.0 274.8 256.0 237.2 219.3 202.1 185.2 168.7 152.5 136.4 120.5 104.5 88.6 72.7 56.8 40.9 24.8 8.3 16.9 107.2 496.7 13.6 84.9 445.2 11.8 72.4 412.4 10.5 63.7 386.9 9.5 57.1 365.2 8.7 51.7 345.5 8.1 47.2 327.9 7.5 43.1 310.3 6.8 38.1 288.1 6.1 33.4 266.0 5.5 29.4 245.2 4.9 25.9 225.3 4.4 22.7 206.1 3.9 19.8 187.3 3.4 17.1 168.9 3.0 14.7 150.8 2.6 12.5 132.9 2.2 10.4 115.1 1.8 8.5 97.5 1.5 6.7 79.9 1.1 5.1 62.3 0.8 3.5 44.6 0.5 2.0 26.9 0.2 0.7 9.0 20.2 130.5 561.6 16.2 102.6 498.8 14.0 87.4 460.5 12.5 77.0 430.9 11.3 68.9 405.5 10.4 62.5 383.3 9.6 56.6 361.3 8.6 49.5 333.5 7.7 43.1 306.1 6.8 37.6 280.7 6.1 32.8 256.8 5.4 28.5 233.8 4.8 24.7 211.7 4.2 21.3 190.3 3.7 18.2 169.5 3.2 15.4 149.1 2.7 12.9 129.0 2.2 10.5 109.0 1.8 8.3 89.1 1.4 6.2 69.3 0.9 4.3 49.5 0.6 2.5 29.7 0.2 0.8 9.9 23.5 153.3 598.5 16.5 104.7 496.4 13.7 85.9 451.3 12.1 74.3 420.1 10.9 66.0 394.9 9.9 59.4 372.7 9.2 54.1 353.0 8.4 49.1 333.4 7.6 43.3 308.8 6.8 37.8 284.4 6.1 33.1 261.6 5.5 29.1 240.0 4.9 25.4 219.0 4.3 22.1 198.7 3.8 19.1 178.9 3.3 16.4 159.6 2.9 13.9 140.5 2.4 11.6 121.7 2.0 9.4 102.9 1.6 7.5 84.3 1.2 5.6 65.6 0.9 3.9 47.0 0.5 2.2 28.2 0.2 0.7 9.4

ISA -10C ISA Time Dist Fuel Time Dist Min. N.M. Lb Min. N.M. 20.0 125.5 518.6 12.9 78.8 415.4 16.3 102.3 10.5 63.4 377.1 12.6 77.8 9.1 54.0 350.8 10.7 65.2 8.1 47.2 329.5 9.4 56.6 7.2 41.9 311.2 8.5 50.2 6.6 37.7 294.5 7.7 45.0 6.0 34.0 279.0 7.1 40.7 5.6 31.0 264.9 6.5 37.2 5.1 28.2 251.1 6.1 33.9 4.7 25.1 234.3 5.5 30.2 4.2 22.2 217.4 5.0 26.6 3.8 19.6 201.2 4.5 23.5 3.4 17.3 185.7 4.0 20.8 3.1 15.3 170.5 3.6 18.3 2.8 13.5 155.6 3.2 16.0 2.5 11.8 140.9 2.9 14.0 2.2 10.2 126.4 2.5 12.1 1.9 8.8 111.8 2.2 10.3 1.6 7.4 97.3 1.9 8.6 1.4 6.1 82.8 1.5 7.1 1.1 4.9 68.2 1.2 5.6 0.9 3.8 53.5 1.0 4.2 0.6 2.7 38.7 0.7 2.9 0.4 1.6 23.6 0.4 1.7 0.1 0.5 8.0 0.1 0.6

PM-133

CLIMB SPEED:

250 KIAS up to 32,000 feet. 0.70 MI above 32,000 feet.

PRESSURE ALTITUDE 1000 FEET

PM-133 ISA Dist N.M. Fuel Lb ISA +10C Time Dist Fuel Min. N.M. Lb ISA +15C Time Dist Fuel Min. N.M. Lb ISA +20C Time Dist Fuel Min. N.M. Lb 22.1 141.1 585.3 14.4 89.3 467.0 11.9 72.4 423.5 10.3 62.1 393.5 9.2 54.6 369.2 8.3 48.8 348.2 7.6 44.0 329.2 7.0 40.1 312.2 6.5 36.5 295.5 5.9 32.4 275.2 5.3 28.6 254.8 4.8 25.3 235.5 4.3 22.3 217.0 3.9 19.7 198.9 3.5 17.2 181.1 3.1 15.0 163.6 2.7 12.9 146.4 2.3 11.0 129.2 2.0 9.2 112.1 1.7 7.6 95.0 1.3 6.0 78.0 1.0 4.5 60.9 0.7 3.2 43.8 0.4 1.9 26.5 0.1 0.6 8.9 20.6 132.2 576.0 15.3 96.0 492.4 13.0 80.0 450.6 11.5 69.7 420.2 10.4 62.1 395.2 9.5 55.9 373.2 8.7 51.0 353.6 8.1 46.5 334.3 7.3 41.1 310.2 6.6 36.0 286.2 5.9 31.7 263.7 5.3 27.8 242.3 4.7 24.4 221.5 4.2 21.2 201.2 3.7 18.4 181.4 3.2 15.8 161.9 2.8 13.4 142.7 2.4 11.2 123.6 2.0 9.1 104.6 1.6 7.2 85.7 1.2 5.4 66.8 0.9 3.8 47.9 0.5 2.2 28.8 0.2 0.7 9.7 19.2 122.6 560.1 15.3 96.0 496.5 13.2 81.7 457.9 11.8 72.0 428.5 10.8 64.5 403.3 9.9 58.5 381.3 9.1 53.1 359.7 8.2 46.7 332.9 7.3 40.8 306.3 6.6 35.7 281.7 5.9 31.3 258.2 5.2 27.3 235.6 4.6 23.7 213.7 4.1 20.5 192.3 3.6 17.6 171.5 3.1 14.9 151.0 2.6 12.4 130.7 2.2 10.1 110.5 1.7 8.0 90.5 1.3 6.0 70.5 0.9 4.2 50.4 0.5 2.4 30.3 0.2 0.8 10.1 24.9 163.0 658.4 18.3 116.4 553.9 15.5 96.8 504.3 13.7 84.3 469.0 12.3 75.0 439.7 11.3 67.8 414.7 10.4 61.2 390.3 9.3 53.5 359.9 8.3 46.5 330.1 7.4 40.5 302.5 6.6 35.3 276.6 5.8 30.7 251.7 5.2 26.6 227.8 4.5 22.9 204.7 3.9 19.6 182.3 3.4 16.6 160.3 2.9 13.8 138.6 2.4 11.3 117.1 1.9 8.9 95.7 1.5 6.6 74.4 1.0 4.6 53.1 0.6 2.6 31.9 0.2 0.8 10.6

WEIGHT 16,000 LB

Time Min.

ISA -10C Dist Fuel N.M. Lb

Time Min.

CLIMB PERFORMANCE TWO ENGINE

Figure 8-9 (Sheet 3 of 12)

51 49 47 45 43 41 39 37 35 33 31 29 27 25 23 21 19 17 15 13 11 9 7 5 3 1

15.5 11.8 10.0 8.8 7.9 7.1 6.5 6.0 5.5 5.0 4.5 4.1 3.7 3.3 3.0 2.6 2.3 2.0 1.7 1.5 1.2 0.9 0.7 0.4 0.1

95.4 71.5 59.6 51.6 45.5 40.7 36.7 33.3 30.3 27.0 23.8 21.1 18.6 16.4 14.4 12.6 10.9 9.4 7.9 6.6 5.3 4.0 2.9 1.7 0.6

474.7 415.6 382.4 357.3 336.2 317.5 300.3 284.8 269.7 251.6 233.4 215.9 199.2 182.9 166.9 151.1 135.5 119.9 104.3 88.7 73.1 57.3 41.4 25.3 8.6

Pilots Manual

8-15

CLIMB SPEED:

250 KIAS up to 32,000 feet. 0.70 MI above 32,000 feet.

WEIGHT 17,000 LB Fuel Lb

PRESSURE ALTITUDE 1000 FEET

8-16

Pilots Manual

Time Min.

ISA -10C Dist Fuel N.M. Lb ISA Dist N.M.

Time Min.

ISA +10C Time Dist Fuel Min. N.M. Lb

ISA +15C Time Dist Fuel Min. N.M. Lb

ISA +20C Time Dist Fuel Min. N.M. Lb

CLIMB PERFORMANCE TWO ENGINE

Figure 8-9 (Sheet 4 of 12)

51 49 47 45 43 41 39 37 35 33 31 29 27 25 23 21 19 17 15 13 11 9 7 5 3 1 17.0 106.0 529.2 13.2 81.0 464.9 11.3 68.2 427.7 10.0 59.4 399.3 9.0 52.8 375.4 8.2 47.4 354.1 7.6 43.1 335.3 7.0 39.3 317.1 6.3 34.8 295.1 5.7 30.7 273.0 5.2 27.1 252.2 4.6 23.9 232.3 4.2 21.0 212.8 3.7 18.4 193.7 3.3 16.0 175.0 2.9 13.8 156.5 2.5 11.8 138.1 2.1 9.9 119.8 1.8 8.1 101.5 1.4 6.4 83.3 1.1 4.8 65.0 0.8 3.4 46.8 0.5 2.0 28.3 0.2 0.7 9.5 17.5 110.3 549.5 14.4 88.8 493.3 12.5 76.3 456.5 11.2 67.5 427.5 10.2 60.6 402.5 9.4 55.1 380.8 8.7 50.1 359.6 7.9 44.2 333.3 7.0 38.7 307.2 6.3 34.0 282.9 5.7 29.9 259.7 5.1 26.1 237.4 4.5 22.8 215.6 4.0 19.7 194.2 3.5 16.9 173.3 3.0 14.4 152.7 2.5 12.0 132.2 2.1 9.8 111.9 1.7 7.7 91.6 1.3 5.8 71.4 0.9 4.0 51.2 0.5 2.3 30.8 0.2 0.8 10.3 23.6 152.4 653.7 17.2 108.4 549.0 14.6 90.1 499.7 12.9 78.6 464.8 11.7 70.0 436.0 10.7 63.4 411.4 9.8 57.3 387.6 8.8 50.3 358.2 7.9 43.9 329.2 7.1 38.4 302.5 6.3 33.6 277.1 5.6 29.3 252.7 5.0 25.4 229.1 4.4 22.0 206.1 3.8 18.8 183.7 3.3 16.0 161.7 2.8 13.3 139.9 2.3 10.8 118.3 1.9 8.6 96.8 1.4 6.4 75.4 1.0 4.4 53.9 0.6 2.6 32.4 0.2 0.8 10.8 21.0 134.8 622.0 17.1 107.8 553.9 15.0 92.6 510.8 13.4 81.8 476.8 12.3 73.6 448.6 11.2 66.3 421.4 10.0 57.8 387.9 8.9 50.1 355.2 7.9 43.6 325.3 7.1 38.0 297.1 6.3 33.0 270.2 5.5 28.6 244.4 4.9 24.6 219.6 4.2 21.0 195.4 3.6 17.8 171.8 3.1 14.8 148.5 2.6 12.1 125.4 2.0 9.5 102.4 1.6 7.1 79.6 1.1 4.9 56.8 0.6 2.8 34.1 0.2 0.9 11.4

21.7 135.6 593.8 13.5 82.0 461.7 11.1 66.1 417.3 9.6 56.4 387.1 8.5 49.4 362.7 7.7 44.0 341.6 7.0 39.5 322.5 6.4 35.8 305.5 5.9 32.6 289.1 5.4 28.9 269.5 4.9 25.5 249.8 4.4 22.5 231.0 4.0 19.9 213.1 3.6 17.6 195.6 3.2 15.4 178.4 2.8 13.5 161.5 2.5 11.7 144.7 2.2 10.0 128.1 1.9 8.5 111.4 1.6 7.0 94.8 1.3 5.6 78.1 1.0 4.3 61.2 0.7 3.1 44.3 0.4 1.8 27.0 0.1 0.6 9.2

PM-133

CLIMB SPEED:

250 KIAS up to 32,000 feet. 0.70 MI above 32,000 feet.

PRESSURE ALTITUDE 1000 FEET

PM-133 ISA Dist N.M. Fuel Lb ISA +10C Time Dist Fuel Min. N.M. Lb ISA +15C Time Dist Fuel Min. N.M. Lb ISA +20C Time Dist Fuel Min. N.M. Lb 22.3 141.0 639.8 14.9 91.8 513.7 12.4 75.2 465.4 10.9 64.7 431.7 9.7 57.1 404.3 8.8 51.1 380.4 8.1 46.4 359.7 7.5 42.1 339.8 6.8 37.3 315.8 6.1 32.9 292.0 5.5 29.0 269.6 5.0 25.6 248.1 4.4 22.5 227.2 4.0 19.7 206.8 3.5 17.1 186.8 3.1 14.8 167.0 2.7 12.6 147.3 2.3 10.5 127.7 1.9 8.6 108.2 1.5 6.8 88.8 1.2 5.2 69.3 0.8 3.6 49.9 0.5 2.1 30.2 0.2 0.7 10.2 20.7 131.9 627.0 16.0 99.3 542.0 13.7 83.8 496.2 12.2 73.4 462.3 11.1 65.6 433.9 10.2 59.5 409.8 9.4 54.0 386.4 8.4 47.5 357.7 7.6 41.5 329.3 6.8 36.4 302.9 6.1 32.0 278.0 5.4 28.0 253.9 4.8 24.4 230.4 4.2 21.1 207.6 3.7 18.1 185.1 3.2 15.3 163.0 2.7 12.8 141.1 2.2 10.4 119.4 1.8 8.2 97.8 1.4 6.2 76.2 1.0 4.3 54.5 0.6 2.5 32.9 0.2 0.8 11.0 19.6 124.5 612.9 16.1 99.9 546.7 14.1 86.0 504.6 12.6 76.1 471.3 11.6 68.6 443.7 10.6 61.9 417.3 9.5 54.2 384.9 8.5 47.2 353.4 7.6 41.2 324.3 6.8 36.0 296.9 6.0 31.4 270.6 5.3 27.2 245.1 4.7 23.5 220.4 4.1 20.1 196.4 3.5 17.1 172.8 3.0 14.2 149.5 2.5 11.6 126.4 2.0 9.1 103.4 1.5 6.9 80.4 1.1 4.7 57.5 0.6 2.7 34.6 0.2 0.9 11.5 25.1 162.6 715.9 19.2 121.2 611.6 16.4 102.0 557.4 14.6 89.3 517.4 13.3 80.0 485.2 12.1 71.8 454.9 10.8 62.4 417.8 9.6 54.0 381.9 8.5 46.9 349.3 7.6 40.8 318.8 6.7 35.4 289.7 5.9 30.6 261.8 5.2 26.4 235.1 4.5 22.5 209.2 3.9 19.0 183.8 3.3 15.8 158.7 2.7 12.9 134.0 2.2 10.1 109.5 1.7 7.6 85.0 1.2 5.2 60.7 0.7 3.0 36.4 0.2 1.0 12.1

WEIGHT 18,000 LB

Time Min.

ISA -10C Dist Fuel N.M. Lb

Time Min.

CLIMB PERFORMANCE TWO ENGINE

Figure 8-9 (Sheet 5 of 12)

51 49 47 45 43 41 39 37 35 33 31 29 27 25 23 21 19 17 15 13 11 9 7 5 3 1

16.0 12.3 10.5 9.2 8.3 7.5 6.9 6.4 5.8 5.2 4.7 4.2 3.8 3.4 3.0 2.7 2.3 2.0 1.7 1.4 1.1 0.8 0.5 0.2

97.8 73.7 61.8 53.6 47.5 42.5 38.4 34.9 30.9 27.3 24.1 21.2 18.7 16.4 14.4 12.5 10.7 9.0 7.5 6.0 4.6 3.3 2.0 0.7

523.5 456.8 419.4 391.0 367.1 345.8 327.1 309.2 288.0 266.8 246.6 227.3 208.6 190.2 172.1 154.3 136.5 118.7 101.0 83.1 65.2 47.1 28.7 9.8

Pilots Manual

8-17

CLIMB SPEED:

250 KIAS up to 32,000 feet. 0.70 MI above 32,000 feet.

WEIGHT 19,000 LB Fuel Lb

PRESSURE ALTITUDE 1000 FEET

8-18

Pilots Manual

Time Min.

ISA -10C Dist Fuel N.M. Lb ISA Dist N.M.

Time Min.

ISA +10C Time Dist Fuel Min. N.M. Lb

ISA +15C Time Dist Fuel Min. N.M. Lb

ISA +20C Time Dist Fuel Min. N.M. Lb

CLIMB PERFORMANCE TWO ENGINE

Figure 8-9 (Sheet 6 of 12)

51 49 47 45 43 41 39 37 35 33 31 29 27 25 23 21 19 17 15 13 11 9 7 5 3 1 17.3 107.0 576.4 13.7 83.3 507.7 11.8 70.6 466.7 10.5 61.8 435.1 9.5 55.1 408.3 8.7 49.8 385.3 8.0 45.2 363.6 7.3 39.9 337.5 6.5 35.1 311.7 5.9 30.9 287.6 5.3 27.3 264.5 4.7 24.0 242.2 4.2 21.0 220.3 3.7 18.2 198.9 3.3 15.7 177.7 2.8 13.4 156.8 2.4 11.2 135.9 2.0 9.2 115.1 1.6 7.3 94.4 1.2 5.5 73.7 0.9 3.8 53.0 0.5 2.3 32.1 0.2 0.7 10.8 23.2 148.3 699.4 17.9 111.5 600.5 15.4 94.3 548.5 13.7 82.7 509.6 12.5 74.3 478.4 11.4 66.9 448.9 10.2 58.3 413.3 9.1 50.7 378.8 8.1 44.2 347.2 7.2 38.6 317.6 6.4 33.6 289.2 5.7 29.1 261.8 5.0 25.1 235.4 4.4 21.5 209.6 3.8 18.2 184.3 3.2 15.2 159.4 2.6 12.4 134.7 2.1 9.7 110.1 1.6 7.3 85.7 1.1 5.1 61.2 0.7 2.9 36.8 0.2 0.9 12.3 28.1 181.3 778.3 18.0 112.5 599.7 15.1 92.3 540.4 13.3 79.9 500.0 11.9 71.0 467.6 11.0 64.2 440.6 10.1 58.1 414.8 9.1 51.0 383.3 8.1 44.5 352.4 7.2 39.0 323.9 6.5 34.2 297.0 5.8 29.9 271.1 5.1 26.0 245.9 4.5 22.5 221.4 3.9 19.3 197.4 3.4 16.3 173.7 2.9 13.6 150.3 2.4 11.1 127.1 1.9 8.8 104.1 1.5 6.6 81.1 1.0 4.6 58.0 0.6 2.7 34.9 0.2 0.9 11.7 21.7 138.2 681.2 18.1 113.0 609.9 16.0 97.7 562.0 14.4 87.0 525.0 13.1 77.8 490.9 11.7 67.4 449.7 10.3 58.1 410.2 9.1 50.4 374.6 8.1 43.7 341.5 7.2 37.9 310.1 6.3 32.8 280.1 5.6 28.2 251.3 4.8 24.1 223.4 4.2 20.3 196.2 3.5 16.9 169.4 2.9 13.7 143.0 2.3 10.8 116.7 1.8 8.1 90.7 1.2 5.6 64.7 0.7 3.2 38.8 0.2 1.0 12.9

21.4 132.5 639.1 13.9 83.5 503.6 11.5 67.9 454.8 10.0 58.3 421.3 8.9 51.2 394.1 8.1 45.6 370.3 7.4 41.2 349.7 6.8 37.3 330.2 6.2 33.0 307.3 5.5 29.1 284.3 5.0 25.7 262.7 4.5 22.6 242.0 4.0 19.9 222.0 3.6 17.5 202.4 3.2 15.3 183.1 2.8 13.2 164.0 2.5 11.4 145.1 2.1 9.6 126.2 1.8 7.9 107.3 1.4 6.4 88.4 1.1 4.9 69.3 0.8 3.5 50.1 0.5 2.1 30.5 0.2 0.7 10.4

PM-133

CLIMB SPEED:

250 KIAS up to 32,000 feet. 0.70 MI above 32,000 feet.

PRESSURE ALTITUDE 1000 FEET

PM-133 ISA Dist N.M. Fuel Lb ISA +10C Time Dist Fuel Min. N.M. Lb ISA +15C Time Dist Fuel Min. N.M. Lb ISA +20C Time Dist Fuel Min. N.M. Lb 21.5 134.5 675.6 15.3 93.2 556.8 12.9 77.2 505.1 11.4 67.0 468.2 10.2 59.3 437.8 9.4 53.5 412.4 8.6 48.4 388.5 7.8 42.7 360.1 7.0 37.5 332.3 6.3 33.0 306.3 5.6 29.0 281.6 5.0 25.5 257.6 4.5 22.3 234.3 4.0 19.4 211.4 3.5 16.7 188.9 3.0 14.2 166.5 2.6 11.9 144.3 2.1 9.7 122.2 1.7 7.7 100.2 1.3 5.8 78.2 0.9 4.1 56.3 0.6 2.4 34.1 0.2 0.8 11.5 20.8 130.7 673.8 16.6 102.2 590.1 14.4 87.3 541.3 12.9 76.9 503.8 11.8 69.2 473.4 10.8 62.5 444.9 9.7 54.7 410.3 8.7 47.7 376.6 7.7 41.7 345.8 6.9 36.5 316.8 6.2 31.9 289.0 5.5 27.7 262.0 4.8 24.0 235.7 4.2 20.5 210.1 3.6 17.4 184.8 3.1 14.5 159.9 2.5 11.8 135.2 2.0 9.3 110.6 1.6 7.0 86.1 1.1 4.9 61.7 0.6 2.8 37.1 0.2 0.9 12.4 30.7 199.3 860.3 20.1 125.8 664.0 16.9 103.8 597.4 14.9 90.1 551.4 13.5 80.5 515.7 12.3 72.2 482.8 11.0 62.8 443.4 9.8 54.3 405.6 8.7 47.3 371.3 7.7 41.3 339.3 6.9 35.9 308.7 6.1 31.1 279.3 5.3 26.8 250.9 4.6 22.9 223.3 4.0 19.4 196.3 3.4 16.1 169.7 2.8 13.1 143.3 2.3 10.4 117.2 1.7 7.8 91.1 1.2 5.4 65.1 0.7 3.1 39.1 0.2 1.0 13.0 25.3 161.7 771.2 20.1 126.0 670.3 17.5 107.3 611.4 15.7 94.8 568.5 14.2 84.4 529.8 12.6 72.8 483.8 11.1 62.5 440.3 9.8 54.1 401.4 8.7 46.9 365.5 7.7 40.6 331.5 6.8 35.0 299.1 5.9 30.1 268.2 5.2 25.7 238.3 4.4 21.7 209.2 3.8 18.0 180.6 3.1 14.6 152.3 2.5 11.5 124.3 1.9 8.6 96.5 1.3 5.9 68.8 0.8 3.4 41.2 0.3 1.1 13.7

WEIGHT 20,000 LB

Time Min.

ISA -10C Dist Fuel N.M. Lb

Time Min.

CLIMB PERFORMANCE TWO ENGINE

Figure 8-9 (Sheet 7 of 12)

51 49 47 45 43 41 39 37 35 33 31 29 27 25 23 21 19 17 15 13 11 9 7 5 3 1

16.0 12.6 10.8 9.6 8.6 7.9 7.3 6.6 5.9 5.3 4.8 4.3 3.8 3.4 3.0 2.6 2.2 1.9 1.5 1.2 0.9 0.5 0.2

97.3 75.0 63.3 55.3 49.0 44.1 39.9 35.3 31.0 27.3 24.1 21.2 18.6 16.2 14.1 12.1 10.2 8.4 6.8 5.2 3.7 2.2 0.7

564.3 494.7 454.0 422.8 396.1 373.4 352.1 327.2 302.5 279.3 257.2 235.8 214.9 194.4 174.1 154.0 133.9 113.8 93.7 73.5 53.1 32.4 11.0

Pilots Manual

8-19

CLIMB SPEED:

250 KIAS up to 32,000 feet. 0.70 MI above 32,000 feet.

WEIGHT 21,000 LB Fuel Lb

PRESSURE ALTITUDE 1000 FEET

8-20

Pilots Manual

Time Min.

ISA -10C Dist Fuel N.M. Lb ISA Dist N.M.

Time Min.

ISA +10C Time Dist Fuel Min. N.M. Lb

ISA +15C Time Dist Fuel Min. N.M. Lb

ISA +20C Time Dist Fuel Min. N.M. Lb

CLIMB PERFORMANCE TWO ENGINE

Figure 8-9 (Sheet 8 of 12)

51 49 47 45 43 41 39 37 35 33 31 29 27 25 23 21 19 17 15 13 11 9 7 5 3 1 17.3 106.4 617.5 14.1 84.8 547.8 12.3 72.6 503.9 11.0 63.9 469.3 10.0 57.4 440.9 9.2 51.8 414.7 8.3 45.6 383.8 7.4 39.9 353.6 6.7 35.1 325.6 6.0 30.9 299.2 5.4 27.1 273.6 4.8 23.7 248.7 4.2 20.6 224.3 3.7 17.7 200.3 3.2 15.1 176.5 2.7 12.6 152.9 2.3 10.3 129.5 1.8 8.2 106.1 1.4 6.2 82.8 1.0 4.3 59.5 0.6 2.5 36.0 0.2 0.8 12.1 23.0 145.2 744.3 18.6 114.9 652.9 16.2 98.5 597.5 14.6 87.3 556.4 13.3 78.0 519.3 11.8 67.5 475.5 10.4 58.3 434.0 9.3 50.6 396.7 8.3 44.1 362.1 7.3 38.3 329.1 6.5 33.1 297.5 5.7 28.5 267.1 4.9 24.4 237.6 4.3 20.6 208.7 3.6 17.2 180.3 3.0 14.0 152.3 2.4 11.0 124.4 1.8 8.3 96.7 1.3 5.7 69.1 0.7 3.3 41.5 0.2 1.1 13.8 31.2 22.6 19.2 17.1 15.4 13.5 11.9 10.5 9.3 8.2 7.2 6.3 5.5 4.7 4.0 3.3 2.6 2.0 1.4 0.8 0.3 202.5 142.1 118.3 103.6 91.7 78.6 67.3 58.0 50.2 43.4 37.4 32.1 27.4 23.1 19.2 15.6 12.2 9.2 6.3 3.6 1.2 25.7 164.0 793.3 18.5 114.2 648.0 15.7 95.5 586.9 14.0 83.4 543.2 12.7 74.8 508.8 11.6 67.3 477.0 10.4 58.7 438.9 9.3 51.0 402.1 8.3 44.5 368.7 7.4 38.9 337.5 6.6 34.0 307.6 5.8 29.5 278.7 5.1 25.5 250.7 4.5 21.8 223.3 3.9 18.5 196.3 3.3 15.4 169.8 2.7 12.5 143.5 2.2 9.9 117.4 1.7 7.4 91.4 1.2 5.1 65.4 0.7 3.0 39.3 0.2 1.0 13.2 911.9 741.9 667.2 616.5 572.2 520.6 472.3 429.7 390.7 354.0 319.2 286.0 253.9 222.7 192.1 162.0 132.1 102.5 73.1 43.8 14.6

20.0 122.8 662.3 14.0 83.8 541.1 11.8 69.1 489.7 10.3 59.7 453.4 9.3 52.6 423.4 8.4 47.2 398.2 7.7 42.6 375.0 7.0 37.6 348.1 6.3 33.0 321.4 5.6 29.0 296.5 5.1 25.5 272.9 4.6 22.5 250.1 4.1 19.7 227.8 3.6 17.2 206.0 3.2 14.9 184.4 2.8 12.8 163.1 2.4 10.8 141.8 2.0 8.9 120.5 1.6 7.2 99.2 1.3 5.5 77.8 0.9 3.9 56.2 0.6 2.3 34.3 0.2 0.8 11.6

PM-133

CLIMB SPEED:

250 KIAS up to 32,000 feet. 0.70 MI above 32,000 feet.

PRESSURE ALTITUDE 1000 FEET

PM-133 ISA Dist N.M. Fuel Lb ISA +10C Time Dist Fuel Min. N.M. Lb ISA +15C Time Dist Fuel Min. N.M. Lb ISA +20C Time Dist Fuel Min. N.M. Lb 20.5 127.0 703.4 15.6 93.8 596.6 13.3 78.9 542.9 11.8 68.9 503.2 10.7 61.6 471.4 9.8 55.4 442.4 8.8 48.7 408.7 7.9 42.5 376.0 7.1 37.3 345.9 6.4 32.8 317.6 5.7 28.8 290.3 5.1 25.1 263.7 4.5 21.8 237.7 3.9 18.8 212.2 3.4 16.0 187.0 2.9 13.4 161.9 2.4 10.9 137.0 1.9 8.6 112.3 1.5 6.5 87.6 1.0 4.5 63.0 0.6 2.7 38.1 0.2 0.9 12.8 20.9 129.9 719.1 17.2 105.0 637.9 15.2 90.7 586.2 13.7 80.8 547.0 12.5 72.4 511.4 11.1 63.0 469.3 9.9 54.5 429.1 8.8 47.5 392.9 7.9 41.5 359.3 7.0 36.1 327.2 6.2 31.4 296.2 5.4 27.1 266.2 4.7 23.2 237.0 4.1 19.6 208.3 3.5 16.3 180.0 2.9 13.3 152.1 2.3 10.5 124.4 1.8 7.9 96.8 1.2 5.4 69.2 0.7 3.2 41.6 0.2 1.0 13.9 27.7 176.8 863.3 20.7 128.3 717.3 17.7 107.9 648.8 15.9 95.0 600.8 14.3 84.3 558.8 12.7 72.7 509.9 11.2 62.5 464.1 9.9 54.1 423.5 8.8 47.1 386.1 7.8 40.8 350.6 6.9 35.3 316.7 6.0 30.4 284.1 5.3 25.9 252.5 4.5 21.9 221.7 3.8 18.2 191.4 3.2 14.8 161.6 2.5 11.7 132.0 1.9 8.8 102.6 1.4 6.0 73.2 0.8 3.5 43.9 0.3 1.1 14.6 25.8 163.1 830.9 21.2 131.3 731.2 18.7 113.6 670.1 16.7 99.8 618.8 14.6 85.0 560.4 12.8 72.4 506.7 11.3 62.2 459.9 10.0 53.7 417.6 8.8 46.3 377.8 7.7 39.9 340.3 6.7 34.2 304.6 5.9 29.2 270.3 5.0 24.6 236.9 4.2 20.4 204.2 3.5 16.6 172.1 2.8 13.0 140.3 2.1 9.7 108.8 1.5 6.7 77.5 0.9 3.9 46.4 0.3 1.2 15.4

WEIGHT 22,000 LB

Time Min.

ISA -10C Dist Fuel N.M. Lb

Time Min.

CLIMB PERFORMANCE TWO ENGINE

Figure 8-9 (Sheet 9 of 12)

51 49 47 45 43 41 39 37 35 33 31 29 27 25 23 21 19 17 15 13 11 9 7 5 3 1

15.9 12.9 11.2 9.9 9.0 8.2 7.4 6.7 6.0 5.4 4.8 4.3 3.8 3.4 2.9 2.5 2.1 1.7 1.3 1.0 0.6 0.2

95.5 75.6 64.5 56.5 50.5 45.4 40.0 35.0 30.8 27.1 23.8 20.9 18.2 15.8 13.5 11.4 9.4 7.6 5.8 4.1 2.5 0.8

598.7 529.4 486.4 452.4 424.5 399.0 369.8 341.1 314.3 289.2 264.8 241.2 218.0 195.1 172.4 149.9 127.4 104.8 82.2 59.4 36.2 12.3

Pilots Manual

8-21

CLIMB SPEED:

250 KIAS up to 32,000 feet. 0.70 MI above 32,000 feet.

WEIGHT 22,750 LB Fuel Lb

PRESSURE ALTITUDE 1000 FEET

8-22

Pilots Manual

Time Min.

ISA -10C Dist Fuel N.M. Lb ISA Dist N.M.

Time Min.

ISA +10C Time Dist Fuel Min. N.M. Lb

ISA +15C Time Dist Fuel Min. N.M. Lb

ISA +20C Time Dist Fuel Min. N.M. Lb

CLIMB PERFORMANCE TWO ENGINE

Figure 8-9 (Sheet 10 of 12)

51 49 47 45 43 41 39 37 35 33 31 29 27 25 23 21 19 17 15 13 11 9 7 5 3 1 25.2 157.7 817.7 16.8 102.0 639.1 14.2 84.1 574.6 12.5 72.9 530.2 11.3 65.0 495.5 10.3 58.3 464.3 9.3 51.1 428.2 8.3 44.6 393.5 7.4 39.1 361.7 6.6 34.3 331.9 5.9 30.1 303.2 5.3 26.3 275.3 4.7 22.8 248.1 4.1 19.6 221.4 3.5 16.7 195.0 3.0 13.9 168.8 2.5 11.4 142.9 2.0 9.0 117.0 1.5 6.8 91.3 1.1 4.7 65.6 0.6 2.8 39.7 0.2 0.9 13.3 23.4 146.3 788.9 18.5 113.2 680.8 16.1 96.7 621.2 14.5 85.7 577.7 13.2 76.6 538.8 11.7 66.4 493.4 10.4 57.3 450.3 9.2 49.8 411.8 8.2 43.5 376.3 7.3 37.9 342.4 6.5 32.8 309.9 5.7 28.3 278.3 5.0 24.2 247.7 4.3 20.5 217.6 3.6 17.0 188.0 3.0 13.9 158.7 2.4 10.9 129.8 1.8 8.2 101.0 1.3 5.7 72.2 0.8 3.3 43.4 0.2 1.1 14.5 22.6 140.6 774.2 19.0 116.0 691.4 16.9 101.3 637.2 15.2 89.5 590.6 13.4 76.8 537.4 11.8 65.8 488.0 10.4 56.9 444.7 9.3 49.4 405.0 8.2 42.8 367.5 7.2 37.0 331.6 6.3 31.8 297.3 5.5 27.1 264.1 4.7 22.9 231.8 4.0 19.1 200.1 3.3 15.5 168.8 2.6 12.2 137.8 2.0 9.2 107.1 1.4 6.3 76.4 0.8 3.6 45.8 0.3 1.2 15.3 29.0 23.0 20.0 17.8 15.5 13.5 11.9 10.5 9.2 8.1 7.1 6.1 5.3 4.4 3.7 2.9 2.2 1.5 0.9 0.3 185.0 142.7 122.1 106.6 90.2 76.5 65.6 56.5 48.7 41.9 35.9 30.5 25.7 21.3 17.3 13.6 10.2 7.0 4.0 1.3 918.1 786.0 714.7 657.0 592.4 534.1 483.9 438.8 396.6 356.9 319.2 283.1 248.0 213.7 180.0 146.7 113.7 81.0 48.5 16.1

17.9 108.2 656.1 13.8 81.4 562.9 11.8 68.4 512.9 10.5 59.6 475.4 9.5 53.1 445.1 8.6 47.6 417.8 7.8 41.9 386.7 7.0 36.6 356.3 6.2 32.1 328.2 5.6 28.3 301.7 5.0 24.9 276.3 4.5 21.8 251.4 4.0 19.0 227.2 3.5 16.4 203.3 3.1 14.1 179.7 2.6 11.9 156.1 2.2 9.8 132.7 1.8 7.9 109.2 1.4 6.0 85.6 1.0 4.3 61.8 0.6 2.6 37.7 0.2 0.9 12.8

PM-133

CLIMB SPEED:

250 KIAS up to 32,000 feet. 0.70 MI above 32,000 feet.

PRESSURE ALTITUDE 1000 FEET

PM-133 ISA Dist N.M. Fuel Lb ISA +10C Time Dist Fuel Min. N.M. Lb ISA +15C Time Dist Fuel Min. N.M. Lb ISA +20C Time Dist Fuel Min. N.M. Lb 27.6 173.0 872.0 17.3 105.2 654.8 14.5 86.0 585.8 12.7 74.3 539.5 11.5 66.1 503.8 10.5 59.3 471.8 9.4 51.9 434.9 8.4 45.3 399.5 7.5 39.7 367.1 6.7 34.8 336.7 6.0 30.5 307.6 5.4 26.6 279.2 4.7 23.1 251.6 4.1 19.9 224.5 3.6 16.9 197.7 3.0 14.1 171.2 2.5 11.5 144.8 2.0 9.1 118.6 1.6 6.9 92.5 1.1 4.8 66.5 0.7 2.8 40.2 0.2 0.9 13.5 24.5 153.5 817.7 19.0 116.2 696.3 16.5 98.8 633.5 14.8 87.5 588.4 13.4 78.0 548.4 11.9 67.5 501.7 10.6 58.3 457.6 9.4 50.6 418.3 8.4 44.2 382.1 7.4 38.4 347.6 6.6 33.3 314.5 5.8 28.7 282.5 5.0 24.6 251.3 4.3 20.8 220.8 3.7 17.3 190.7 3.0 14.1 161.0 2.4 11.1 131.6 1.9 8.3 102.4 1.3 5.8 73.2 0.8 3.3 44.0 0.3 1.1 14.7 23.3 145.3 795.4 19.5 118.9 706.6 17.2 103.5 649.9 15.5 91.4 601.7 13.6 78.3 546.9 12.0 67.0 496.3 10.6 57.9 451.9 9.4 50.2 411.5 8.3 43.5 373.2 7.3 37.6 336.8 6.4 32.3 301.8 5.6 27.5 268.1 4.8 23.3 235.3 4.1 19.3 203.0 3.4 15.7 171.3 2.7 12.4 139.8 2.0 9.3 108.6 1.4 6.4 77.5 0.8 3.7 46.5 0.3 1.2 15.5 30.0 23.6 20.5 18.2 15.8 13.7 12.1 10.6 9.4 8.2 7.2 6.2 5.3 4.5 3.7 3.0 2.3 1.6 0.9 0.3 191.5 147.0 125.1 109.0 92.0 77.9 66.7 57.5 49.5 42.5 36.5 31.0 26.1 21.7 17.6 13.8 10.3 7.1 4.1 1.3 947.5 806.0 730.6 670.4 603.6 543.6 492.2 446.0 403.0 362.5 324.2 287.4 251.8 216.9 182.7 148.9 115.4 82.2 49.2 16.3

WEIGHT 23,000 LB

Time Min.

ISA -10C Dist Fuel N.M. Lb

Time Min.

CLIMB PERFORMANCE TWO ENGINE

Figure 8-9 (Sheet 11 of 12)

51 49 47 45 43 41 39 37 35 33 31 29 27 25 23 21 19 17 15 13 11 9 7 5 3 1

18.8 113.8 680.2 14.1 83.5 574.9 12.0 69.8 522.2 10.6 60.7 483.3 9.6 54.0 452.2 8.8 48.4 424.2 7.9 42.5 392.5 7.1 37.2 361.5 6.3 32.6 332.9 5.7 28.7 306.0 5.1 25.2 280.1 4.6 22.1 254.9 4.0 19.2 230.3 3.6 16.7 206.1 3.1 14.3 182.1 2.7 12.0 158.3 2.2 9.9 134.5 1.8 8.0 110.7 1.4 6.1 86.7 1.0 4.3 62.7 0.6 2.6 38.2 0.2 0.9 13.0

Pilots Manual

8-23

CLIMB SPEED:

250 KIAS up to 32,000 feet. 0.70 MI above 32,000 feet.

WEIGHT 23,500 LB Fuel Lb

PRESSURE ALTITUDE 1000 FEET

8-24

ISA -10 C Time Dist Fuel Min. N.M. Lb ISA Dist N.M.

Pilots Manual

Time Min.

ISA +10 C Time Dist Fuel Min. N.M. Lb

ISA +15 C Time Dist Fuel Min. N.M. Lb

ISA +20 C Time Dist Fuel Min. N.M. Lb

CLIMB PERFORMANCE TWO ENGINE

Figure 8-9 (Sheet 12 of 12)

51 49 47 45 43 41 39 37 35 33 31 29 27 25 23 21 19 17 15 13 11 9 7 5 3 1 18.5 112.3 689.5 15.1 89.9 609.2 13.2 77.3 558.9 11.9 68.6 520.9 10.8 61.4 487.1 9.7 53.6 448.5 8.7 46.7 411.7 7.8 40.9 378.0 6.9 35.9 346.6 6.2 31.4 316.5 5.5 27.4 287.2 4.9 23.8 258.8 4.3 20.4 230.8 3.7 17.4 203.2 3.1 14.5 175.9 2.6 11.8 148.8 2.1 9.4 121.9 1.6 7.1 95.0 1.1 4.9 68.3 0.7 2.9 41.3 0.2 1.0 13.9 24.9 156.0 842.5 20.4 125.0 738.6 18.0 108.2 676.6 16.1 95.2 624.8 14.1 81.2 566.5 12.4 69.4 513.2 11.0 59.9 466.8 9.7 51.9 424.7 8.6 44.9 385.0 7.6 38.7 347.2 6.6 33.3 311.1 5.8 28.4 276.2 4.9 24.0 242.3 4.2 19.9 209.0 3.5 16.2 176.3 2.8 12.7 143.9 2.1 9.5 111.7 1.5 6.6 79.7 0.9 3.8 47.8 0.3 1.2 15.0 33.9 25.0 21.5 19.0 16.4 14.3 12.5 11.0 9.7 8.5 7.4 6.4 5.5 4.6 3.8 3.1 2.3 1.6 0.9 0.3 27.3 172.6 891.3 20.0 122.7 729.3 17.2 103.3 659.1 15.4 91.1 610.6 13.9 81.0 568.0 12.3 69.9 518.7 10.9 60.2 472.5 9.7 52.3 431.6 8.6 45.6 394.0 7.7 39.6 358.3 6.8 34.4 324.0 5.9 29.6 290.9 5.2 25.3 258.7 4.5 21.4 227.2 3.8 17.8 196.2 3.1 14.5 165.6 2.5 11.4 135.3 1.9 8.6 105.2 1.3 5.9 75.2 0.8 3.4 45.2 0.3 1.1 15.1 218.4 1043.0 156.2 848.9 131.7 764.1 114.0 698.5 95.8 626.8 80.9 563.1 69.1 509.1 59.4 461.0 51.1 416.2 43.9 374.1 37.6 334.4 32.0 296.4 26.9 259.5 22.3 223.5 18.1 188.1 14.2 153.3 10.6 118.8 7.3 84.6 4.2 50.6 1.3 16.8

21.2 129.4 743.8 14.9 88.1 600.6 12.5 72.7 541.4 11.0 62.9 499.6 9.9 55.8 466.6 9.1 50.0 437.4 8.1 43.8 404.2 7.3 38.3 372.1 6.5 33.6 342.4 5.9 29.5 314.6 5.2 25.9 288.0 4.7 22.7 262.0 4.1 19.8 236.7 3.7 17.1 211.7 3.2 14.6 187.1 2.7 12.4 162.5 2.3 10.2 138.1 1.9 8.2 113.6 1.5 6.3 89.1 1.0 4.5 64.3 0.6 2.7 39.2 0.2 0.9 13.3

PM-133 Change 1

CLIMB SPEED:

250 KIAS up to 32,000 feet. 0.70 MI above 32,000 feet.

Pilots Manual

CRUISE PERFORMANCE
The cruise performance on the following pages is based on flight test data and represents the average delivered aircraft. NORMAL CRUISE The Normal Cruise tables (Figure 8-10) provide fuel flows and true airspeed for constant 0.76 MI cruise at weights from 14,000 to 23,000 pounds. Engine power is adjusted to maintain constant Mach as weight decreases. Standard and off-standard day temperatures provide interpolation factors. MAXIMUM SPECIFIC RANGE Figure 8-11 presents a graphic description of the range capability at ISA as a function of weight and altitude. The data is based upon two engine, maximum-range cruise at ISA. In general, the cruise altitude selected should be near the maximum nautical miles per pound fuel for a given aircraft weight. MAXIMUM-RANGE CRUISE - TWO ENGINES The Maximum-Range Cruise - Two-Engine tables (Figure 8-12) provide fuel flow, indicated Mach or airspeed, and true airspeed for 100% maximum range cruise at weights from 14,000 to 23,000 pounds. Standard and off-standard day temperatures provide interpolation factors. LONG-RANGE CRUISE - TWO ENGINES The Long-Range Cruise - Two-Engine tables (Figure 8-13) provide fuel flow, indicated Mach or airspeed, and true airspeed for 99% maximum range cruise at weights from 14,000 to 23,000 pounds. Standard and offstandard day temperatures provide interpolation factors. HIGH-SPEED CRUISE The High Speed Cruise tables (Figure 8-14) provide fuel flows, indicated Mach or airspeed, and true airspeed for a MMO/VMO or VMAX cruise at weights from 14,000 to 23,000 pounds. Power for maximum speed cruise is for the limiting condition (MMO/VMO, or maximum cruise power). Standard and off-standard day temperatures provide interpolation factors.
PM-133

8-25

Pilots Manual

MAXIMUM RANGE DESCENT - ONE ENGINE Figure 8-15 shows the descent speed schedule for a maximum range descent to an altitude at or below the single-engine service ceiling for the aircraft gross weight. LONG-RANGE CRUISE - ONE ENGINE The Long-Range Cruise - One Engine tables (Figure 8-16) provide fuel flows, indicated Mach or airspeed and true airspeed for 99% maximum range cruise at weights from 14,000 to 23,000 pounds. Standard nd offstandard day temperatures provide interpolation factors.

8-26

PM-133

Pilots Manual

NORMAL CRUISE
WEIGHT 14,000 LB Mach .76 MI KTAS 51 Fuel - Lb/Hr KTAS 49 Fuel - Lb/Hr KTAS 47 Fuel - Lb/Hr KTAS 45 Fuel - Lb/Hr KTAS 43 Fuel - Lb/Hr KTAS 41 Fuel - Lb/Hr KTAS 39 Fuel - Lb/Hr KTAS 37 Fuel - Lb/Hr KTAS 35 Fuel - Lb/Hr KTAS 30 Fuel - Lb/Hr KTAS 25 Fuel - Lb/Hr ISA -10 425 878 425 860 425 862 425 881 425 912 425 953 425 1005 425 1069 427 1152 437 1452 447 1830 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

435 885 435 887 435 907 435 937 435 978 435 1031 435 1096 437 1187 447 1484 456 1883

445 910 445 931 445 961 445 1004 445 1058 445 1126 447 1215 456 1526 466 1930

450 942 450 974 450 1018 450 1071 450 1138 452 1230 461 1545 470 1946

454 954 455 986 455 1031 455 1086 455 1154 457 1248 466 1563 475 1969

ALTITUDE 1000 FEET

WEIGHT 14,500 LB Mach .76 MI KTAS 51 Fuel - Lb/Hr KTAS 49 Fuel - Lb/Hr KTAS 47 Fuel - Lb/Hr KTAS 45 Fuel - Lb/Hr KTAS 43 Fuel - Lb/Hr KTAS 41 Fuel - Lb/Hr KTAS 39 Fuel - Lb/Hr KTAS 37 Fuel - Lb/Hr KTAS 35 Fuel - Lb/Hr KTAS 30 Fuel - Lb/Hr KTAS 25 Fuel - Lb/Hr

ISA -10

TEMPERATURE C ISA ISA +10 ISA +15

ISA +20

425 889 425 886 425 900 425 927 425 966 425 1017 425 1078 427 1161 437 1456 447 1831

435 914 435 912 435 927 435 954 435 993 435 1044 435 1107 437 1196 447 1487 456 1885

445 935 445 951 445 979 445 1018 445 1071 445 1136 447 1224 456 1530 466 1931

450 962 450 991 450 1033 450 1085 450 1149 452 1240 461 1550 470 1948

ALTITUDE 1000 FEET

455 1003 455 1046 455 1100 455 1166 457 1258 466 1567 475 1970

PM-133

Figure 8-10 (Sheet 1 of 10)

8-27

Pilots Manual

NORMAL CRUISE
WEIGHT 15,000 LB Mach .76 MI KTAS 51 Fuel - Lb/Hr KTAS 49 Fuel - Lb/Hr KTAS 47 Fuel - Lb/Hr KTAS 45 Fuel - Lb/Hr KTAS 43 Fuel - Lb/Hr KTAS 41 Fuel - Lb/Hr KTAS 39 Fuel - Lb/Hr KTAS 37 Fuel - Lb/Hr KTAS 35 Fuel - Lb/Hr KTAS 30 Fuel - Lb/Hr KTAS 25 Fuel - Lb/Hr WEIGHT 15,500 LB Mach .76 MI KTAS 51 Fuel - Lb/Hr KTAS 49 Fuel - Lb/Hr KTAS 47 Fuel - Lb/Hr KTAS 45 Fuel - Lb/Hr KTAS 43 Fuel - Lb/Hr KTAS 41 Fuel - Lb/Hr KTAS 39 Fuel - Lb/Hr KTAS 37 Fuel - Lb/Hr KTAS 35 Fuel - Lb/Hr KTAS 30 Fuel - Lb/Hr KTAS 25 Fuel - Lb/Hr ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

425 920 425 911 425 921 425 945 425 981 425 1030 425 1090 427 1169 437 1460 447 1833

435 946 435 938 435 947 435 972 435 1009 435 1058 435 1119 437 1206 447 1491 456 1886

445 972 445 998 445 1034 445 1085 445 1148 447 1233 456 1534 466 1933

450 985 450 1010 450 1049 450 1100 450 1162 452 1250 461 1554 470 1949

ALTITUDE 1000 FEET

454 1023 455 1062 455 1114 455 1179 457 1268 466 1571 475 1972

ISA -10

TEMPERATURE C ISA ISA +10 ISA +15

ISA +20

425 956 425 939 425 945 425 962 425 995 425 1043 425 1101 427 1179 437 1464 447 1834

435 965 435 970 435 991 435 1024 435 1073 435 1132 437 1216 447 1495 456 1887

445 995 445 1018 445 1051 445 1099 445 1161 447 1243 456 1538 466 1934

ALTITUDE 1000 FEET

450 1029 450 1065 450 1115 450 1176 452 1261 461 1558 470 1951

454 1043 455 1078 455 1129 455 1192 457 1279 466 1576 475 1973

8-28

Figure 8-10 (Sheet 2 of 10)

PM-133

Pilots Manual

NORMAL CRUISE
WEIGHT 16,000 LB Mach .76 MI KTAS 51 Fuel - Lb/Hr KTAS 49 Fuel - Lb/Hr KTAS 47 Fuel - Lb/Hr KTAS 45 Fuel - Lb/Hr KTAS 43 Fuel - Lb/Hr KTAS 41 Fuel - Lb/Hr KTAS 39 Fuel - Lb/Hr KTAS 37 Fuel - Lb/Hr KTAS 35 Fuel - Lb/Hr KTAS 30 Fuel - Lb/Hr KTAS 25 Fuel - Lb/Hr WEIGHT 16,500 LB Mach .76 MI KTAS 51 Fuel - Lb/Hr KTAS 49 Fuel - Lb/Hr KTAS 47 Fuel - Lb/Hr KTAS 45 Fuel - Lb/Hr KTAS 43 Fuel - Lb/Hr KTAS 41 Fuel - Lb/Hr KTAS 39 Fuel - Lb/Hr KTAS 37 Fuel - Lb/Hr KTAS 35 Fuel - Lb/Hr KTAS 30 Fuel - Lb/Hr KTAS 25 Fuel - Lb/Hr ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

425 996 425 967 425 969 425 982 425 1011 425 1056 425 1114 427 1191 437 1468 447 1835

435 994 435 995 435 1011 435 1041 435 1087 435 1146 437 1227 447 1499 456 1889

445 1020 445 1039 445 1069 445 1115 445 1174 447 1255 456 1543 466 1935

ALTITUDE 1000 FEET

450 1050 450 1082 450 1130 450 1190 452 1273 461 1563 470 1952

454 1064 455 1095 455 1145 455 1206 457 1290 466 1580 475 1975

ISA -10

TEMPERATURE C ISA ISA +10 ISA +15

ISA +20

425 997 425 993 425 1004 425 1027 425 1070 425 1126 427 1204 437 1472 447 1837

435 1026 435 1021 435 1031 435 1059 435 1103 435 1160 437 1239 447 1503 456 1891

445 1048 445 1059 445 1088 445 1131 445 1188 447 1268 456 1548 466 1937

ALTITUDE 1000 FEET

450 1071 450 1100 450 1146 450 1204 452 1286 461 1567 470 1954

454 1085 454 1114 455 1160 455 1221 457 1303 466 1585 475 1977

PM-133

Figure 8-10 (Sheet 3 of 10)

8-29

Pilots Manual

NORMAL CRUISE
WEIGHT 17,000 LB Mach .76 MI KTAS 51 Fuel - Lb/Hr KTAS 49 Fuel - Lb/Hr KTAS 47 Fuel - Lb/Hr KTAS 45 Fuel - Lb/Hr KTAS 43 Fuel - Lb/Hr KTAS 41 Fuel - Lb/Hr KTAS 39 Fuel - Lb/Hr KTAS 37 Fuel - Lb/Hr KTAS 35 Fuel - Lb/Hr KTAS 30 Fuel - Lb/Hr KTAS 25 Fuel - Lb/Hr WEIGHT 17,500 LB Mach .76 MI KTAS 51 Fuel - Lb/Hr KTAS 49 Fuel - Lb/Hr KTAS 47 Fuel - Lb/Hr KTAS 45 Fuel - Lb/Hr KTAS 43 Fuel - Lb/Hr KTAS 41 Fuel - Lb/Hr KTAS 39 Fuel - Lb/Hr KTAS 37 Fuel - Lb/Hr KTAS 35 Fuel - Lb/Hr KTAS 30 Fuel - Lb/Hr KTAS 25 Fuel - Lb/Hr ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

425 1033 425 1020 425 1027 425 1046 425 1085 425 1139 427 1217 437 1479 447 1839

435 1048 435 1052 435 1078 435 1118 435 1174 437 1252 447 1510 456 1893

ALTITUDE 1000 FEET

445 1081 445 1108 445 1148 445 1203 447 1282 456 1555 466 1940

450 1094 450 1119 450 1162 450 1219 452 1299 461 1575 470 1956

454 1134 455 1177 455 1236 457 1317 466 1592 475 1980

ISA -10

TEMPERATURE C ISA ISA +10 ISA +15

ISA +20

425 1071 425 1047 425 1050 425 1068 425 1100 425 1153 427 1231 437 1488 447 1843

435 1077 435 1076 435 1097 435 1134 435 1188 437 1265 447 1519 456 1897

ALTITUDE 1000 FEET

445 1105 445 1129 445 1165 445 1218 447 1296 456 1565 466 1944

450 1119 450 1139 450 1179 450 1234 452 1313 461 1584 470 1960

454 1155 455 1193 455 1251 457 1330 466 1601 475 1984

8-30

Figure 8-10 (Sheet 4 of 10)

PM-133

Pilots Manual

NORMAL CRUISE
WEIGHT 18,000 LB Mach .76 MI KTAS 51 Fuel - Lb/Hr KTAS 49 Fuel - Lb/Hr KTAS 47 Fuel - Lb/Hr KTAS 45 Fuel - Lb/Hr KTAS 43 Fuel - Lb/Hr KTAS 41 Fuel - Lb/Hr KTAS 39 Fuel - Lb/Hr KTAS 37 Fuel - Lb/Hr KTAS 35 Fuel - Lb/Hr KTAS 30 Fuel - Lb/Hr KTAS 25 Fuel - Lb/Hr ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

425 1114 425 1076 425 1074 425 1090 425 1120 425 1169 427 1245 437 1498 447 1846

435 1106 435 1102 435 1117 435 1152 435 1203 437 1278 447 1528 456 1902

ALTITUDE 1000 FEET

445 1130 445 1150 445 1184 445 1235 447 1311 456 1575 466 1948

450 1161 450 1197 450 1250 452 1327 461 1594 470 1964

454 1176 454 1212 455 1266 457 1345 466 1611 475 1988

WEIGHT 18,500 LB Mach .76 MI KTAS 51 Fuel - Lb/Hr KTAS 49 Fuel - Lb/Hr KTAS 47 Fuel - Lb/Hr KTAS 45 Fuel - Lb/Hr KTAS 43 Fuel - Lb/Hr KTAS 41 Fuel - Lb/Hr KTAS 39 Fuel - Lb/Hr KTAS 37 Fuel - Lb/Hr KTAS 35 Fuel - Lb/Hr KTAS 30 Fuel - Lb/Hr KTAS 25 Fuel - Lb/Hr

ISA -10

TEMPERATURE C ISA ISA +10 ISA +15

ISA +20

425 1111 425 1100 425 1112 425 1140 425 1185 427 1260 437 1508 447 1850

435 1143 435 1129 435 1137 435 1170 435 1218 437 1292 447 1538 456 1906

ALTITUDE 1000 FEET

445 1157 445 1171 445 1204 445 1251 447 1326 456 1585 466 1952

450 1183 450 1216 450 1266 452 1342 461 1605 470 1968

454 1198 454 1231 455 1282 457 1359 466 1621 475 1992

PM-133

Figure 8-10 (Sheet 5 of 10)

8-31

Pilots Manual

NORMAL CRUISE
WEIGHT 19,000 LB Mach .76 MI KTAS 51 Fuel - Lb/Hr KTAS 49 Fuel - Lb/Hr KTAS 47 Fuel - Lb/Hr KTAS 45 Fuel - Lb/Hr KTAS 43 Fuel - Lb/Hr KTAS 41 Fuel - Lb/Hr KTAS 39 Fuel - Lb/Hr KTAS 37 Fuel - Lb/Hr KTAS 35 Fuel - Lb/Hr KTAS 30 Fuel - Lb/Hr KTAS 25 Fuel - Lb/Hr ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

425 1148 425 1126 425 1134 425 1160 425 1202 427 1275 437 1517 447 1854

435 1180 435 1157 435 1161 435 1189 435 1234 437 1306 447 1549 456 1910

ALTITUDE 1000 FEET

445 1187 445 1192 445 1224 445 1268 447 1341 456 1596 466 1956

450 1206 450 1235 450 1282 452 1357 461 1615 470 1972

454 1252 455 1298 457 1374 466 1632 475 1997

WEIGHT 19,500 LB Mach .76 MI KTAS 51 Fuel - Lb/Hr KTAS 49 Fuel - Lb/Hr KTAS 47 Fuel - Lb/Hr KTAS 45 Fuel - Lb/Hr KTAS 43 Fuel - Lb/Hr KTAS 41 Fuel - Lb/Hr KTAS 39 Fuel - Lb/Hr KTAS 37 Fuel - Lb/Hr KTAS 35 Fuel - Lb/Hr KTAS 30 Fuel - Lb/Hr KTAS 25 Fuel - Lb/Hr

ISA -10

TEMPERATURE C ISA ISA +10 ISA +15

ISA +20

425 1187 425 1153 425 1158 425 1181 425 1220 427 1290 437 1528 447 1858

ALTITUDE 1000 FEET

435 1186 435 1186 435 1208 435 1250 437 1320 447 1560 456 1914

445 1216 445 1244 445 1286 447 1357 456 1606 466 1960

450 1232 450 1255 450 1299 452 1372 461 1626 470 1976

454 1273 454 1316 457 1389 466 1642 475 2001

8-32

Figure 8-10 (Sheet 6 of 10)

PM-133

Pilots Manual

NORMAL CRUISE
WEIGHT 20,000 LB Mach .76 MI KTAS 51 Fuel - Lb/Hr KTAS 49 Fuel - Lb/Hr KTAS 47 Fuel - Lb/Hr KTAS 45 Fuel - Lb/Hr KTAS 43 Fuel - Lb/Hr KTAS 41 Fuel - Lb/Hr KTAS 39 Fuel - Lb/Hr KTAS 37 Fuel - Lb/Hr KTAS 35 Fuel - Lb/Hr KTAS 30 Fuel - Lb/Hr KTAS 25 Fuel - Lb/Hr ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

425 1235 425 1184 425 1182 425 1203 425 1239 427 1306 437 1538 447 1862

ALTITUDE 1000 FEET

435 1217 435 1211 435 1228 435 1268 437 1335 447 1571 456 1919

445 1243 445 1265 445 1305 447 1373 456 1617 466 1965

450 1259 450 1277 450 1318 452 1388 461 1637 470 1981

454 1294 454 1334 457 1405 466 1653 475 2006

WEIGHT 20,500 LB Mach .76 MI KTAS 51 Fuel - Lb/Hr KTAS 49 Fuel - Lb/Hr KTAS 47 Fuel - Lb/Hr KTAS 45 Fuel - Lb/Hr KTAS 43 Fuel - Lb/Hr KTAS 41 Fuel - Lb/Hr KTAS 39 Fuel - Lb/Hr KTAS 37 Fuel - Lb/Hr KTAS 35 Fuel - Lb/Hr KTAS 30 Fuel - Lb/Hr KTAS 25 Fuel - Lb/Hr

ISA -10

TEMPERATURE C ISA ISA +10 ISA +15

ISA +20

ALTITUDE 1000 FEET

425 1220 425 1207 425 1224 425 1258 427 1322 437 1550 447 1866

435 1253 435 1239 435 1250 435 1286 437 1350 447 1584 456 1924

445 1271 445 1286 445 1324 447 1389 456 1630 466 1969

450 1300 450 1337 452 1404 461 1649 471 1985

454 1316 454 1354 457 1420 466 1665 475 2011

PM-133

Figure 8-10 (Sheet 7 of 10)

8-33

Pilots Manual

NORMAL CRUISE
WEIGHT 21,000 LB Mach .76 MI KTAS 51 Fuel - Lb/Hr KTAS 49 Fuel - Lb/Hr KTAS 47 Fuel - Lb/Hr KTAS 45 Fuel - Lb/Hr KTAS 43 Fuel - Lb/Hr KTAS 41 Fuel - Lb/Hr KTAS 39 Fuel - Lb/Hr KTAS 37 Fuel - Lb/Hr KTAS 35 Fuel - Lb/Hr KTAS 30 Fuel - Lb/Hr KTAS 25 Fuel - Lb/Hr ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

ALTITUDE 1000 FEET

425 1257 425 1233 425 1246 425 1278 427 1338 437 1562 447 1871

435 1292 435 1267 435 1274 435 1304 437 1365 447 1597 456 1929

445 1299 445 1308 445 1344 447 1406 457 1643 466 1975

450 1324 450 1356 452 1420 461 1663 471 1991

454 1375 457 1437 466 1678 475 2017

WEIGHT 21,500 LB Mach .76 MI KTAS 51 Fuel - Lb/Hr KTAS 49 Fuel - Lb/Hr KTAS 47 Fuel - Lb/Hr KTAS 45 Fuel - Lb/Hr KTAS 43 Fuel - Lb/Hr KTAS 41 Fuel - Lb/Hr KTAS 39 Fuel - Lb/Hr KTAS 37 Fuel - Lb/Hr KTAS 35 Fuel - Lb/Hr KTAS 30 Fuel - Lb/Hr KTAS 25 Fuel - Lb/Hr

ISA -10

TEMPERATURE C ISA ISA +10 ISA +15

ISA +20

ALTITUDE 1000 FEET

425 1295 425 1260 425 1269 425 1299 427 1356 437 1575 447 1881

435 1332 435 1296 435 1299 435 1323 437 1382 447 1611 456 1939

445 1329 445 1331 445 1365 447 1424 457 1657 466 1984

450 1348 450 1376 452 1438 461 1676 471 2000

454 1396 457 1455 466 1692 475 2027

8-34

Figure 8-10 (Sheet 8 of 10)

PM-133

Pilots Manual

NORMAL CRUISE
WEIGHT 22,000 LB Mach .76 MI KTAS 51 Fuel - Lb/Hr KTAS 49 Fuel - Lb/Hr KTAS 47 Fuel - Lb/Hr KTAS 45 Fuel - Lb/Hr KTAS 43 Fuel - Lb/Hr KTAS 41 Fuel - Lb/Hr KTAS 39 Fuel - Lb/Hr KTAS 37 Fuel - Lb/Hr KTAS 35 Fuel - Lb/Hr KTAS 30 Fuel - Lb/Hr KTAS 25 Fuel - Lb/Hr ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

ALTITUDE 1000 FEET

425 1341 425 1291 425 1292 425 1319 427 1374 437 1588 447 1891

435 1327 435 1325 435 1346 437 1399 447 1625 456 1949

445 1358 445 1385 447 1443 457 1670 466 1994

450 1375 450 1398 452 1456 461 1690 471 2011

454 1417 457 1475 466 1707 475 2037

WEIGHT 22,500 LB Mach .76 MI KTAS 51 Fuel - Lb/Hr KTAS 49 Fuel - Lb/Hr KTAS 47 Fuel - Lb/Hr KTAS 45 Fuel - Lb/Hr KTAS 43 Fuel - Lb/Hr KTAS 41 Fuel - Lb/Hr KTAS 39 Fuel - Lb/Hr KTAS 37 Fuel - Lb/Hr KTAS 35 Fuel - Lb/Hr KTAS 30 Fuel - Lb/Hr KTAS 25 Fuel - Lb/Hr

ISA -10

TEMPERATURE C ISA ISA +10 ISA +15

ISA +20

ALTITUDE 1000 FEET

425 1392 425 1327 425 1317 425 1340 427 1393 437 1601 447 1901

435 1363 435 1352 435 1369 437 1420 447 1640 456 1959

445 1386 445 1406 447 1462 457 1685 466 2004

450 1404 450 1422 452 1475 461 1704 471 2021

454 1439 457 1495 466 1722 475 2047

PM-133

Figure 8-10 (Sheet 9 of 10)

8-35

Pilots Manual

NORMAL CRUISE
WEIGHT 23,000 LB Mach .76 MI KTAS 51 Fuel - Lb/Hr KTAS 49 Fuel - Lb/Hr KTAS 47 Fuel - Lb/Hr KTAS 45 Fuel - Lb/Hr KTAS 43 Fuel - Lb/Hr KTAS 41 Fuel - Lb/Hr KTAS 39 Fuel - Lb/Hr KTAS 37 Fuel - Lb/Hr KTAS 35 Fuel - Lb/Hr KTAS 30 Fuel - Lb/Hr KTAS 25 Fuel - Lb/Hr ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

ALTITUDE 1000 FEET

425 1364 425 1343 425 1361 427 1412 437 1614 447 1911

435 1400 435 1380 435 1393 437 1442 447 1655 456 1969

445 1414 445 1427 447 1481 457 1699 466 2013

450 1445 452 1494 461 1718 471 2032

455 1463 457 1515 466 1737 475 2058

8-36

Figure 8-10 (Sheet 10 of 10)

PM-133

Pilots Manual

MAXIMUM SPECIFIC RANGE

Figure 8-11 8-37

PM-133

Pilots Manual

MAXIMUM RANGE CRUISE

TWO ENGINE
WEIGHT 14,000 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 .733 409 820 .708 395 763 .703 392 751 .687 383 743 .677 377 746 .666 371 750 .642 357 740 .606 337 713 .579 324 713 .506 289 705 191 273 753 190 250 791 200 243 888 207 233 960 204 213 972 202 196 1016 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.711 406 790 .704 402 776 .673 384 750 .654 373 740 .633 361 730 .607 346 716 .601 342 729 .573 328 728 .504 295 724 191 279 776 191 256 821 204 253 921 203 232 948 198 211 965 199 197 1028

.713 417 814 .704 411 796 .675 394 772 .654 382 759 .641 374 758 .611 356 740 .604 352 753 .585 342 763 .510 305 755 191 284 801 195 267 856 199 251 909 198 231 941 194 210 972 196 198 1041

.704 415 805 .671 396 779 .654 386 769 .635 374 760 .616 363 754 .604 356 761 .585 346 772 .516 311 774 191 287 810 196 271 865 197 251 906 195 229 938 192 210 979 195 198 1052

.704 420 816 .682 407 800 .654 390 778 .634 378 770 .607 362 753 .598 356 762 .585 350 780 .521 317 790 191 290 818 195 272 866 195 251 906 192 228 938 191 211 990 194 199 1066 PM-133

8-38

Figure 8-12 (Sheet 1 of 19)

Pilots Manual

MAXIMUM RANGE CRUISE

TWO ENGINE
WEIGHT 14,500 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET PM-133 ISA -10 .732 409 856 .720 402 808 .704 392 776 .675 376 751 .662 369 747 .654 364 750 .631 351 740 .604 336 725 .593 332 744 .518 296 736 195 278 775 194 255 815 205 249 921 209 235 976 206 215 989 204 198 1037 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.720 411 830 .704 402 799 .687 392 785 .657 375 763 .642 366 759 .617 352 745 .604 344 749 .583 334 756 .514 300 752 194 283 802 194 260 844 206 255 937 205 235 967 200 213 985 202 200 1052

.704 411 819 .687 401 808 .658 384 783 .644 376 779 .621 362 770 .604 352 768 .596 349 792 .522 312 786 195 289 827 198 271 877 201 253 928 200 233 960 196 212 994 198 200 1065

.704 415 828 .688 406 820 .659 389 795 .643 379 788 .618 364 774 .604 356 776 .596 352 800 .527 318 803 195 292 836 198 274 884 199 253 925 196 232 959 194 212 1003 197 200 1077

.688 410 829 .658 393 804 .644 384 800 .617 368 783 .604 360 786 .581 347 789 .531 324 819 195 295 843 197 274 883 197 253 925 194 230 959 193 213 1015 196 201 1091

Figure 8-12 (Sheet 2 of 19)

8-39

Pilots Manual

MAXIMUM RANGE CRUISE

TWO ENGINE
WEIGHT 15,000 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.731 408 855 .706 394 803 .695 388 795 .663 369 767 .648 361 761 .632 352 757 .604 336 740 .596 333 761 .529 303 766 197 281 797 197 259 841 208 252 941 210 237 993 207 217 1008 206 200 1060

.723 413 868 .708 404 828 .694 396 816 .668 381 795 .650 371 787 .626 357 774 .604 344 765 .594 340 784 .525 307 783 198 288 829 197 265 868 208 258 956 207 237 987 202 215 1005 204 202 1074

.704 411 844 .694 405 837 .667 389 814 .651 380 807 .634 370 802 .604 352 784 .604 354 817 .533 319 818 198 294 853 200 274 897 205 258 956 201 235 980 198 214 1018 200 202 1088

.695 410 850 .669 395 828 .651 384 817 .626 369 803 .604 356 791 .604 358 825 .537 324 833 198 297 862 200 276 903 201 256 944 198 234 980 197 215 1028 199 202 1101

.695 415 859 .669 399 837 .654 390 832 .626 373 814 .604 360 803 .591 354 819 .540 329 846 198 300 869 198 277 902 199 256 945 196 233 984 195 215 1039 198 203 1115 PM-133

8-40

Figure 8-12 (Sheet 3 of 19)

Pilots Manual

MAXIMUM RANGE CRUISE

TWO ENGINE
WEIGHT 15,500 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET PM-133 ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.733 409 891 .710 396 833 .703 392 827 .673 375 800 .654 364 786 .633 353 776 .608 338 760 .601 336 783 .531 304 783 200 286 824 200 263 867 210 255 960 212 239 1011 210 219 1029 207 202 1079

.723 413 904 .710 405 858 .703 402 851 .682 389 832 .654 373 810 .635 362 803 .610 348 789 .604 346 812 .536 314 814 201 293 856 201 269 893 210 260 972 209 239 1006 204 217 1026 206 204 1096

.711 415 883 .703 411 873 .683 399 855 .654 382 829 .635 370 821 .613 357 812 .604 354 832 .545 325 848 201 299 879 203 277 918 206 260 973 203 237 1000 200 217 1041 202 204 1113

.703 415 882 .672 397 853 .654 386 841 .634 374 832 .607 358 813 .601 356 836 .547 330 862 201 302 888 202 279 922 203 258 965 200 236 1001 199 217 1053 201 205 1127

.703 420 892 .682 407 876 .655 390 852 .636 379 845 .609 363 829 .601 360 849 .542 330 863 201 305 895 201 280 923 201 258 965 199 237 1014 199 219 1071 200 205 1142

Figure 8-12 (Sheet 4 of 19)

8-41

Pilots Manual

MAXIMUM RANGE CRUISE

TWO ENGINE
WEIGHT 16,000 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.726 405 919 .720 402 877 .704 392 849 .689 384 839 .655 365 807 .641 357 804 .616 343 788 .604 338 802 .540 309 812 204 290 851 204 268 893 212 257 978 215 241 1032 211 221 1047 209 204 1101

.720 412 903 .704 402 874 .685 391 857 .659 376 835 .643 367 832 .619 353 818 .604 346 828 .547 320 844 204 298 882 204 273 918 211 262 989 210 241 1025 205 218 1047 208 206 1118

.721 421 927 .704 411 898 .688 401 882 .657 384 853 .643 375 849 .624 364 842 .604 354 847 .556 332 880 205 304 905 205 280 939 208 262 990 205 239 1020 202 219 1064 205 206 1137

.703 415 905 .689 407 896 .657 387 865 .642 378 861 .616 363 843 .604 358 856 .557 336 891 205 307 914 204 282 941 204 260 983 202 239 1025 202 221 1083 204 207 1152

.703 420 916 .689 411 906 .658 392 877 .645 384 876 .618 368 859 .604 361 870 .546 333 884 205 310 921 204 284 948 203 261 987 202 241 1043 202 222 1102 203 207 1168 PM-133

8-42

Figure 8-12 (Sheet 5 of 19)

Pilots Manual

MAXIMUM RANGE CRUISE

TWO ENGINE
WEIGHT 16,500 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET PM-133 ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.732 409 967 .723 404 914 .711 397 882 .695 387 866 .667 372 842 .649 361 832 .623 347 814 .604 338 817 .550 315 840 207 295 877 207 272 918 214 260 998 217 244 1050 213 223 1067 211 205 1121

.723 413 940 .710 406 907 .695 397 891 .671 383 870 .652 372 860 .627 358 847 .604 346 844 .557 326 874 207 301 905 207 277 943 213 264 1007 212 243 1043 207 221 1068 209 207 1141

.711 415 932 .692 404 909 .666 389 885 .649 379 877 .625 364 861 .605 354 862 .566 338 909 208 309 932 207 283 960 209 264 1007 207 241 1040 205 222 1094 206 208 1161

.704 415 931 .695 410 925 .666 393 897 .649 383 890 .624 368 873 .604 358 873 .556 336 903 208 312 940 207 286 966 206 263 1002 205 242 1053 205 224 1114 206 209 1178

.695 415 935 .668 398 911 .654 390 907 .627 374 890 .604 361 887 .555 339 915 207 313 943 207 288 975 206 265 1015 205 244 1074 205 226 1135 205 210 1195

Figure 8-12 (Sheet 6 of 19)

8-43

Pilots Manual

MAXIMUM RANGE CRUISE

TWO ENGINE
WEIGHT 17,000 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.723 404 948 .708 395 904 .704 392 897 .682 380 880 .654 364 856 .630 351 841 .607 340 838 .559 320 869 207 295 891 210 276 944 216 262 1017 219 246 1069 215 224 1086 213 207 1142

.723 413 975 .708 404 929 .704 402 924 .672 383 890 .654 373 881 .636 363 876 .611 350 871 .568 332 905 209 303 925 210 282 969 215 266 1024 214 245 1063 209 222 1090 211 209 1163

.711 415 959 .700 409 943 .674 394 917 .654 381 902 .631 368 888 .607 356 882 .574 343 936 211 313 959 210 287 985 211 266 1025 209 244 1063 208 225 1124 208 210 1186

.709 419 968 .704 415 958 .669 395 923 .654 386 915 .632 372 903 .607 359 895 .558 337 922 211 316 967 210 290 991 209 267 1029 209 246 1082 208 227 1145 208 211 1204

.704 420 969 .681 406 951 .654 390 926 .635 379 920 .610 365 914 .564 344 945 207 314 959 210 293 1001 209 269 1043 209 248 1105 208 229 1167 208 213 1227 PM-133

8-44

Figure 8-12 (Sheet 7 of 19)

Pilots Manual

MAXIMUM RANGE CRUISE

TWO ENGINE
WEIGHT 17,500 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET PM-133 ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.723 404 984 .718 401 947 .704 392 919 .689 384 909 .654 364 875 .637 355 869 .614 344 865 .568 325 897 207 295 907 213 280 970 218 265 1036 220 248 1087 217 226 1107 214 209 1164

.723 414 1011 .718 410 974 .704 402 947 .685 391 930 .661 377 910 .644 367 905 .620 355 900 .578 338 935 212 309 954 213 286 993 216 268 1043 216 247 1082 211 225 1115 213 211 1187

.718 420 1000 .703 411 972 .686 401 954 .654 382 922 .638 372 917 .614 360 910 .571 341 945 214 318 985 213 291 1010 212 268 1045 212 247 1092 211 229 1154 211 212 1213

.704 415 981 .687 405 969 .654 386 935 .639 377 932 .614 364 925 .567 342 952 212 318 984 213 294 1018 212 270 1056 212 249 1111 211 231 1176 211 214 1236

.703 420 993 .688 410 981 .654 390 947 .644 384 950 .618 370 943 .573 350 975 210 319 989 213 297 1028 212 273 1071 212 251 1135 211 233 1199 211 216 1260

Figure 8-12 (Sheet 8 of 19)

8-45

Pilots Manual

MAXIMUM RANGE CRUISE

TWO ENGINE
WEIGHT 18,000 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.732 409 1034 .723 403 985 .709 396 948 .694 387 934 .663 370 907 .644 359 896 .621 348 892 .577 330 926 211 301 937 216 284 996 220 267 1055 222 250 1106 218 228 1126 216 210 1186

.723 413 1013 .709 405 978 .693 396 962 .664 379 932 .652 372 933 .628 360 929 .588 344 965 216 314 983 216 290 1018 218 270 1060 217 249 1101 214 228 1143 215 213 1210

.723 423 1041 .707 413 1003 .692 404 983 .661 386 952 .645 376 945 .621 364 939 .573 343 964 217 322 1012 216 295 1035 215 272 1070 215 251 1119 214 232 1185 214 215 1244

.704 415 1006 .694 409 1000 .664 392 969 .647 382 962 .622 368 955 .576 348 983 213 320 1005 216 298 1044 215 274 1084 215 253 1142 214 234 1208 214 217 1269

.704 420 1019 .693 414 1010 .663 396 981 .651 388 981 .626 375 973 .582 355 1005 214 324 1019 216 301 1058 215 277 1100 215 255 1167 214 236 1231 214 219 1294 PM-133

8-46

Figure 8-12 (Sheet 9 of 19)

Pilots Manual

MAXIMUM RANGE CRUISE

TWO ENGINE
WEIGHT 18,500 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET PM-133 ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.722 403 1063 .723 404 1019 .704 393 968 .701 391 964 .673 375 940 .651 363 924 .627 351 919 .585 335 954 215 306 968 219 288 1021 222 269 1075 225 253 1130 220 230 1147 218 212 1208

.723 413 1047 .713 407 1008 .701 400 995 .668 382 960 .654 373 953 .636 364 957 .598 350 995 219 318 1012 219 294 1044 219 272 1079 219 251 1122 217 231 1172 217 215 1234

.713 417 1037 .693 405 1009 .669 390 984 .651 380 973 .627 368 967 .581 348 994 220 326 1035 219 300 1061 218 275 1097 218 254 1149 217 235 1217 217 218 1276

.710 419 1042 .704 415 1036 .666 393 993 .654 386 991 .630 373 985 .584 353 1013 217 325 1035 219 302 1071 218 278 1112 217 256 1171 217 237 1241 217 220 1301

.701 418 1043 .673 401 1016 .653 389 1002 .633 379 1003 .591 360 1035 217 329 1049 219 305 1086 218 280 1126 217 258 1198 217 239 1264 217 222 1328

Figure 8-12 (Sheet 10 of 19)

8-47

Pilots Manual

MAXIMUM RANGE CRUISE

TWO ENGINE
WEIGHT 19,000 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.723 404 1053 .713 398 1009 .704 392 988 .685 382 975 .654 364 946 .634 355 946 .594 340 982 219 311 998 222 292 1047 224 272 1095 226 254 1147 222 232 1167 220 214 1230

.723 413 1083 .713 408 1035 .704 402 1021 .683 390 1002 .654 373 971 .643 368 984 .604 354 1020 223 323 1041 222 298 1069 221 274 1098 221 253 1142 220 234 1201 220 217 1262

.713 417 1063 .703 411 1048 .683 399 1025 .654 381 996 .634 372 996 .589 352 1023 219 324 1046 222 304 1087 221 279 1124 221 257 1177 220 238 1248 220 221 1308

.713 421 1076 .704 415 1058 .686 405 1045 .654 386 1011 .637 377 1015 .592 358 1044 220 329 1064 222 306 1099 221 282 1140 221 260 1202 220 240 1273 220 223 1334

.703 420 1070 .684 408 1054 .654 390 1023 .641 384 1033 .599 366 1066 220 333 1079 222 309 1114 221 284 1158 225 267 1251 220 242 1299 220 225 1362 PM-133

8-48

Figure 8-12 (Sheet 11 of 19)

Pilots Manual

MAXIMUM RANGE CRUISE

TWO ENGINE
WEIGHT 19,500 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.728 406 1097 .722 403 1052 .704 392 1009 .691 385 1003 .654 365 965 .641 359 973 .602 345 1011 222 316 1028 225 295 1073 226 274 1115 228 256 1164 223 233 1188 222 216 1259

.728 416 1128 .722 413 1080 .705 403 1046 .690 394 1033 .654 373 992 .640 367 998 .604 354 1034 226 328 1070 225 302 1094 224 278 1124 224 256 1168 223 237 1230 222 220 1293

.722 422 1109 .704 411 1073 .688 402 1054 .654 382 1017 .640 375 1025 .596 356 1052 222 329 1076 225 308 1113 224 283 1152 223 261 1206 223 241 1280 222 224 1341

.704 415 1081 .691 408 1074 .661 390 1041 .644 382 1045 .601 363 1074 223 334 1094 225 310 1128 224 285 1169 223 263 1230 223 243 1306 222 226 1366

.704 420 1095 .690 412 1085 .656 391 1046 .643 385 1054 .604 369 1090 224 338 1110 225 313 1143 224 288 1186 227 270 1275 239 263 1422 222 228 1394

PM-133

Figure 8-12 (Sheet 12 of 19)

8-49

Pilots Manual

MAXIMUM RANGE CRUISE

TWO ENGINE
WEIGHT 20,000 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.732 409 1143 .723 403 1085 .703 392 1035 .695 387 1028 .662 369 996 .647 362 1001 .604 346 1029 226 321 1058 228 299 1098 228 277 1135 229 257 1182 226 236 1214 225 219 1289

.723 413 1114 .712 407 1080 .695 397 1061 .662 378 1025 .646 371 1027 .604 354 1048 230 333 1098 228 306 1120 227 281 1150 226 259 1195 226 240 1260 225 223 1323

.723 423 1144 .710 414 1108 .680 397 1064 .662 387 1050 .646 379 1053 .603 361 1081 225 334 1107 228 311 1141 227 286 1179 225 263 1233 226 244 1312 225 227 1372

.704 415 1106 .695 410 1102 .662 391 1063 .651 386 1075 .604 365 1096 226 338 1124 228 314 1156 227 289 1198 224 264 1251 226 247 1338 225 229 1400

.706 421 1124 .695 415 1113 .664 396 1080 .648 388 1082 .604 369 1105 227 343 1140 228 317 1172 227 292 1214 229 272 1299 241 266 1444 225 231 1429 PM-133

8-50

Figure 8-12 (Sheet 13 of 19)

Pilots Manual

MAXIMUM RANGE CRUISE

TWO ENGINE
WEIGHT 20,500 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.722 403 1173 .723 403 1117 .707 394 1068 .704 392 1060 .671 374 1028 .653 366 1028 .604 346 1044 230 326 1087 231 303 1124 230 279 1156 230 259 1200 229 239 1241 228 222 1317

.723 413 1148 .713 407 1105 .704 402 1096 .671 383 1059 .653 374 1056 .604 354 1065 232 337 1124 231 309 1145 230 284 1175 229 263 1222 228 243 1290 228 226 1353

.723 423 1179 .713 417 1139 .695 406 1111 .670 391 1083 .653 383 1083 .604 361 1099 229 338 1137 230 313 1163 230 290 1207 227 265 1257 228 247 1344 228 230 1405

.710 419 1145 .704 415 1137 .670 395 1097 .654 388 1098 .604 365 1112 229 343 1153 230 316 1178 230 293 1226 229 270 1293 243 265 1450 228 232 1434

.711 425 1160 .704 420 1149 .672 401 1114 .653 391 1111 .604 369 1121 230 347 1170 230 319 1193 230 295 1242 232 275 1324 244 268 1466 228 233 1464

PM-133

Figure 8-12 (Sheet 14 of 19)

8-51

Pilots Manual

MAXIMUM RANGE CRUISE

TWO ENGINE
WEIGHT 21,000 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.723 403 1152 .715 399 1108 .699 390 1077 .684 381 1068 .654 366 1046 .605 346 1061 233 331 1117 234 306 1151 233 282 1180 232 261 1220 231 242 1269 231 224 1347

.723 413 1184 .715 408 1135 .703 402 1118 .684 391 1101 .654 375 1077 .604 354 1082 232 337 1138 234 313 1171 233 288 1202 232 266 1251 231 246 1321 231 228 1384

.715 418 1168 .703 411 1150 .683 399 1124 .654 383 1103 .604 361 1116 232 343 1167 230 313 1177 233 293 1237 229 267 1282 231 250 1377 231 232 1439

.715 422 1180 .703 415 1159 .684 404 1142 .654 387 1116 .605 366 1130 232 347 1183 230 316 1192 233 296 1255 232 273 1319 245 267 1469 231 234 1468

.703 420 1172 .684 408 1156 .654 391 1132 .605 369 1139 233 352 1201 230 319 1208 233 299 1271 234 278 1348 245 270 1487 269 275 1746 PM-133

8-52

Figure 8-12 (Sheet 15 of 19)

Pilots Manual

MAXIMUM RANGE CRUISE

TWO ENGINE
WEIGHT 21,500 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.728 407 1198 .723 403 1151 .703 392 1109 .690 384 1095 .655 367 1067 .612 351 1089 237 336 1147 237 310 1177 236 286 1206 235 264 1246 234 244 1295 233 227 1378

.728 416 1231 .723 413 1181 .704 402 1141 .690 394 1130 .655 376 1099 .604 354 1100 230 334 1143 237 316 1198 236 291 1229 235 269 1279 234 249 1352 233 231 1416

.723 422 1212 .704 411 1175 .678 396 1137 .655 384 1125 .610 365 1144 235 347 1197 231 315 1199 235 297 1265 235 274 1327 245 265 1467 233 235 1471

.723 427 1228 .704 415 1182 .690 407 1175 .655 388 1139 .612 370 1160 235 352 1213 231 319 1215 235 300 1283 235 277 1347 245 268 1483 233 237 1501

.704 420 1198 .689 411 1186 .655 392 1155 .612 374 1170 236 356 1231 232 323 1235 236 302 1299 236 281 1372 247 272 1508 254 260 1651

PM-133

Figure 8-12 (Sheet 16 of 19)

8-53

Pilots Manual

MAXIMUM RANGE CRUISE

TWO ENGINE
WEIGHT 22,000 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.732 409 1245 .723 403 1183 .703 392 1135 .685 382 1107 .663 371 1098 .619 355 1117 240 341 1177 239 314 1203 238 289 1231 238 267 1272 237 247 1324 236 230 1408

.732 419 1280 .723 413 1214 .710 405 1174 .695 397 1158 .662 380 1131 .609 357 1128 233 338 1174 237 317 1212 238 295 1255 237 272 1307 237 252 1383 236 234 1446

.723 423 1246 .708 414 1208 .679 396 1162 .662 388 1157 .616 369 1174 238 352 1226 234 320 1231 238 300 1293 238 277 1356 246 266 1483 236 238 1505

.703 415 1206 .695 410 1203 .662 393 1174 .619 374 1191 238 356 1242 235 323 1248 238 303 1311 238 280 1374 247 270 1504 270 274 1758

.704 420 1224 .695 414 1216 .662 397 1189 .610 373 1186 239 361 1261 235 327 1267 238 306 1327 239 284 1397 249 274 1526 256 262 1670 PM-133

8-54

Figure 8-12 (Sheet 17 of 19)

Pilots Manual

MAXIMUM RANGE CRUISE

TWO ENGINE
WEIGHT 22,500 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.722 403 1268 .723 404 1215 .706 394 1165 .693 387 1145 .671 376 1130 .626 359 1146 244 345 1206 242 317 1229 241 292 1256 240 270 1299 239 250 1352 239 232 1439

.723 413 1249 .711 406 1199 .702 401 1191 .670 384 1165 .615 360 1158 236 342 1204 237 317 1226 240 297 1278 238 273 1329 238 254 1409 239 236 1479

.723 423 1281 .713 417 1242 .694 405 1212 .670 393 1191 .623 373 1205 241 356 1257 238 324 1263 241 304 1322 240 280 1383 248 268 1505 239 241 1539

.707 417 1240 .703 415 1238 .669 397 1209 .626 378 1222 241 360 1273 238 328 1280 241 307 1340 240 283 1401 249 272 1525 257 261 1677

.706 422 1255 .702 419 1251 .669 401 1223 .616 376 1216 242 365 1291 239 332 1299 241 309 1355 241 286 1421 250 275 1545 257 264 1691

PM-133

Figure 8-12 (Sheet 18 of 19)

8-55

Pilots Manual

MAXIMUM RANGE CRUISE

TWO ENGINE
WEIGHT 23,000 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.726 405 1323 .723 404 1249 .713 398 1204 .699 390 1179 .668 374 1147 .625 358 1161 247 349 1234 245 321 1256 244 296 1282 243 273 1325 242 253 1381 241 235 1469

.723 413 1284 .713 407 1231 .703 402 1215 .683 392 1208 .621 364 1187 239 347 1235 240 321 1257 243 301 1308 240 275 1352 240 256 1435 241 239 1511

.723 423 1317 .713 417 1267 .703 411 1252 .668 392 1209 .628 376 1234 244 360 1287 241 329 1294 244 307 1350 243 283 1410 250 270 1528 259 261 1683

.713 421 1280 .703 415 1262 .686 407 1261 .624 377 1237 244 364 1302 241 332 1313 244 310 1368 243 286 1428 251 274 1546 259 263 1696

.713 426 1295 .702 419 1273 .676 405 1255 .621 380 1247 245 370 1322 242 336 1331 244 313 1383 243 289 1445 252 277 1565 259 265 1710 PM-133

8-56

Figure 8-12 (Sheet 19 of 19)

Pilots Manual

LONG RANGE CRUISE

TWO ENGINE
WEIGHT 14,000 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 .746 416 843 .734 409 800 .721 402 779 .712 397 779 .700 390 779 .684 381 777 .661 368 769 .629 350 747 .610 341 758 .541 309 762 205 291 812 205 269 858 221 268 990 224 251 1048 220 230 1059 218 212 1109 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.735 420 826 .723 413 806 .713 407 803 .696 397 796 .673 384 785 .646 368 769 .623 355 764 .606 347 777 .538 314 780 202 294 826 211 282 911 221 273 1004 220 251 1036 215 228 1055 216 214 1125

.726 424 832 .724 423 828 .714 417 826 .695 406 815 .677 395 809 .657 383 804 .630 367 793 .612 358 807 .544 325 813 206 305 868 214 292 944 218 274 1002 214 250 1028 210 227 1063 213 214 1141

.724 428 838 .714 421 836 .697 411 828 .674 397 815 .650 383 803 .629 370 800 .612 362 816 .547 330 828 208 311 886 213 294 948 214 272 993 211 248 1025 209 228 1073 212 215 1153

.714 426 846 .698 416 839 .676 403 829 .646 385 808 .623 371 801 .607 363 817 .550 335 843 208 315 896 211 294 946 212 271 990 208 247 1025 207 228 1084 211 215 1166

PM-133

Figure 8-13 (Sheet 1 of 19)

8-57

Pilots Manual

LONG RANGE CRUISE

TWO ENGINE
WEIGHT 14,500 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 .744 416 876 .738 412 837 .728 406 811 .716 399 805 .699 390 796 .682 380 790 .662 368 784 .634 353 769 .614 343 778 .550 314 789 205 292 825 207 272 878 224 271 1012 227 254 1070 222 232 1080 220 214 1132 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.739 422 861 .729 416 836 .715 408 827 .705 402 826 .683 389 815 .657 374 802 .630 359 789 .613 351 803 .549 321 811 206 299 856 214 286 937 223 276 1024 222 254 1057 217 230 1077 219 216 1150

.729 426 857 .717 418 851 .703 410 845 .682 398 834 .664 387 831 .638 372 820 .619 362 831 .555 331 844 209 310 894 217 296 967 220 277 1024 216 252 1049 213 230 1088 215 217 1167

.727 430 865 .718 424 865 .704 415 858 .682 402 845 .657 387 831 .633 373 821 .617 365 837 .558 336 858 210 315 910 216 297 969 216 275 1013 213 250 1048 211 230 1099 214 217 1180

.717 428 872 .704 420 868 .687 409 862 .658 392 843 .630 375 828 .612 366 839 .559 341 871 211 319 920 213 297 966 213 274 1010 210 249 1049 210 231 1111 213 218 1195

8-58

Figure 8-13 (Sheet 2 of 19)

PM-133

Pilots Manual

LONG RANGE CRUISE

TWO ENGINE
WEIGHT 15,000 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.743 415 878 .733 409 843 .721 402 834 .708 394 827 .688 383 815 .665 370 805 .639 355 790 .618 346 797 .557 318 814 208 296 847 211 276 905 226 274 1033 229 257 1090 225 234 1101 223 216 1157

.743 425 902 .732 418 866 .720 411 856 .711 406 855 .691 394 845 .667 380 833 .638 364 816 .620 355 827 .560 327 843 209 304 883 217 290 962 225 278 1043 224 257 1079 219 232 1099 221 219 1176

.730 426 884 .720 420 878 .709 414 875 .689 402 862 .669 390 855 .648 378 849 .625 366 854 .566 338 876 212 314 919 219 299 989 221 279 1041 218 254 1070 215 232 1112 218 219 1193

.722 426 892 .709 418 886 .689 406 874 .666 392 862 .637 375 842 .619 366 854 .568 342 889 213 319 934 218 300 989 218 277 1032 215 253 1070 213 233 1125 217 220 1208

.720 430 899 .710 423 898 .693 413 891 .668 398 877 .638 380 856 .619 370 866 .564 344 893 213 322 943 215 299 985 215 276 1029 212 251 1073 212 233 1137 215 220 1223

PM-133

Figure 8-13 (Sheet 3 of 19)

8-59

Pilots Manual

LONG RANGE CRUISE

TWO ENGINE
WEIGHT 15,500 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.746 416 916 .733 409 870 .726 405 862 .715 399 858 .696 388 845 .673 375 833 .645 359 815 .624 349 820 .566 323 843 210 299 872 214 281 933 229 277 1055 231 260 1112 227 237 1123 225 218 1180

.743 425 935 .733 419 896 .725 414 886 .713 407 879 .699 399 875 .676 386 864 .649 370 848 .627 359 852 .570 333 874 212 308 909 220 294 987 228 281 1063 226 259 1101 221 234 1122 223 221 1199

.735 430 923 .725 424 909 .715 418 905 .696 406 891 .674 393 879 .653 381 873 .632 370 878 .576 344 907 214 318 944 222 302 1012 223 281 1060 220 256 1092 217 235 1138 220 221 1219

.725 428 919 .714 422 916 .696 411 904 .674 397 892 .646 380 873 .623 369 876 .576 347 916 216 323 958 220 303 1010 219 279 1051 217 255 1093 216 235 1152 219 222 1235

.724 432 928 .715 427 928 .698 416 918 .678 404 911 .649 386 891 .625 374 892 .570 347 916 215 325 962 217 302 1006 217 278 1049 214 254 1097 214 236 1166 217 223 1250

8-60

Figure 8-13 (Sheet 4 of 19)

PM-133

Pilots Manual

LONG RANGE CRUISE

TWO ENGINE
WEIGHT 16,000 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.746 416 954 .737 411 907 .732 408 893 .719 401 884 .704 393 876 .681 379 862 .655 364 846 .630 352 844 .575 329 873 213 303 897 218 285 962 231 280 1077 234 262 1132 229 239 1147 227 220 1203

.737 421 933 .730 417 917 .716 409 905 .707 403 905 .685 391 895 .660 376 881 .634 363 878 .581 340 906 215 312 935 223 298 1013 229 283 1082 228 261 1119 223 237 1145 225 223 1224

.731 427 942 .717 419 929 .703 410 922 .681 397 909 .658 383 896 .636 372 899 .587 351 938 217 322 969 224 306 1034 225 283 1079 222 259 1114 219 237 1164 222 223 1245

.729 430 948 .719 424 944 .703 415 934 .681 402 923 .656 387 907 .629 372 901 .580 350 937 218 327 982 222 306 1032 221 281 1071 219 257 1117 218 238 1178 221 224 1262

.718 428 954 .704 420 948 .688 410 944 .659 393 925 .632 378 919 .579 353 947 216 326 978 219 305 1027 218 280 1070 216 257 1125 217 239 1195 220 225 1280

PM-133

Figure 8-13 (Sheet 5 of 19)

8-61

Pilots Manual

LONG RANGE CRUISE

TWO ENGINE
WEIGHT 16,500 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.743 415 988 .742 414 948 .732 408 917 .722 403 909 .710 396 905 .688 383 891 .662 369 874 .636 356 869 .585 334 903 216 307 921 221 290 990 234 283 1098 236 265 1153 231 241 1166 228 222 1225

.742 424 975 .734 420 948 .721 412 934 .709 405 928 .694 396 926 .669 382 913 .642 367 905 .591 346 937 218 316 961 226 302 1038 231 285 1101 230 263 1139 224 239 1168 227 225 1249

.734 429 972 .718 419 953 .709 414 952 .687 401 938 .663 387 923 .639 374 921 .593 354 962 220 326 995 226 309 1057 226 285 1097 224 261 1136 221 239 1189 224 225 1271

.730 431 976 .722 426 972 .708 418 963 .689 406 953 .664 392 939 .636 376 929 .584 353 959 220 329 1002 224 309 1054 222 283 1091 221 260 1142 220 240 1206 223 226 1291

.720 430 979 .710 423 978 .692 412 970 .668 398 958 .639 382 947 .589 359 979 217 328 997 221 307 1048 220 282 1092 218 259 1151 219 241 1223 222 227 1307

8-62

Figure 8-13 (Sheet 6 of 19)

PM-133

Pilots Manual

LONG RANGE CRUISE

TWO ENGINE
WEIGHT 17,000 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.745 416 988 .733 409 945 .726 405 935 .715 399 933 .695 387 920 .669 373 903 .645 361 899 .594 340 933 219 311 948 225 294 1018 236 285 1119 238 267 1174 233 243 1188 230 224 1248

.745 426 1016 .734 419 973 .725 414 963 .711 406 953 .700 399 953 .678 387 944 .652 373 938 .601 351 967 221 321 990 229 306 1064 233 288 1121 231 265 1160 226 241 1192 229 227 1274

.735 430 1002 .724 423 986 .714 417 981 .693 404 967 .670 391 952 .645 377 946 .596 356 982 224 331 1024 229 312 1081 228 287 1116 225 263 1158 223 242 1217 226 227 1298

.725 428 998 .713 421 994 .695 410 983 .672 396 971 .646 382 962 .593 358 990 221 331 1022 226 311 1075 224 285 1111 223 262 1167 223 243 1236 225 229 1318

.723 432 1007 .714 426 1007 .696 415 996 .678 404 991 .649 388 982 .597 364 1010 220 333 1027 223 310 1071 222 285 1115 221 262 1178 222 244 1254 224 229 1337

PM-133

Figure 8-13 (Sheet 7 of 19)

8-63

Pilots Manual

LONG RANGE CRUISE

TWO ENGINE
WEIGHT 17,500 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.746 416 1025 .737 411 982 .730 407 963 .719 401 958 .703 392 950 .676 377 932 .653 366 930 .603 345 963 222 316 978 228 299 1046 238 288 1140 240 269 1195 234 245 1209 232 226 1272

.741 424 1042 .737 421 1009 .730 417 993 .716 409 983 .702 401 976 .687 392 975 .661 378 969 .609 356 995 225 327 1020 232 310 1088 235 290 1140 233 266 1179 228 243 1215 231 229 1299

.737 431 1036 .729 426 1019 .716 418 1006 .699 408 996 .677 395 982 .653 382 978 .601 359 1004 226 335 1050 231 315 1102 229 288 1136 227 265 1180 226 244 1245 228 230 1326

.727 430 1025 .718 424 1024 .701 413 1012 .680 401 1002 .655 388 996 .602 364 1022 223 334 1045 227 313 1094 226 287 1134 225 265 1193 225 245 1264 227 231 1347

.727 434 1038 .717 428 1034 .702 419 1026 .684 408 1021 .659 394 1015 .606 370 1041 224 338 1058 225 313 1095 224 288 1141 223 265 1208 224 247 1285 227 232 1369

8-64

Figure 8-13 (Sheet 8 of 19)

PM-133

Pilots Manual

LONG RANGE CRUISE

TWO ENGINE
WEIGHT 18,000 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.745 416 1063 .740 413 1020 .728 406 984 .721 402 982 .708 395 978 .683 380 960 .661 370 959 .610 349 990 226 321 1009 231 303 1072 240 291 1160 242 271 1214 236 247 1231 234 227 1295

.740 423 1048 .735 420 1024 .720 411 1011 .705 402 999 .690 394 998 .669 383 999 .615 360 1020 229 332 1051 234 313 1110 236 292 1159 234 268 1199 230 245 1240 233 231 1325

.734 429 1053 .717 419 1028 .706 412 1027 .683 398 1011 .660 387 1009 .607 363 1031 228 337 1070 232 317 1120 231 291 1157 229 267 1203 228 247 1273 230 232 1355

.732 432 1058 .721 426 1050 .705 416 1040 .688 406 1034 .664 393 1029 .609 368 1050 227 339 1075 229 315 1115 228 290 1157 227 267 1219 227 248 1295 230 233 1376

.720 430 1060 .708 422 1057 .688 410 1046 .667 399 1048 .612 373 1068 227 343 1090 228 317 1126 228 293 1177 229 272 1254 227 250 1316 229 234 1399

PM-133

Figure 8-13 (Sheet 9 of 19)

8-65

Pilots Manual

LONG RANGE CRUISE

TWO ENGINE
WEIGHT 18,500 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.742 415 1100 .744 415 1059 .732 408 1016 .724 404 1006 .713 398 1006 .689 384 988 .667 373 987 .617 353 1016 230 326 1041 234 306 1098 242 293 1182 243 273 1232 238 248 1252 236 229 1320

.744 425 1089 .735 420 1050 .724 413 1039 .711 406 1031 .692 395 1018 .673 386 1023 .620 363 1043 232 337 1080 236 316 1133 238 294 1179 236 270 1220 232 247 1265 235 233 1350

.737 430 1082 .721 421 1059 .711 415 1057 .689 402 1040 .667 391 1039 .613 366 1058 230 340 1092 233 318 1137 232 293 1178 231 269 1227 230 249 1302 233 234 1384

.734 433 1088 .724 428 1077 .710 419 1069 .694 409 1063 .671 398 1061 .615 371 1077 230 344 1107 231 318 1137 230 293 1184 229 270 1245 230 251 1325 232 235 1406

.722 431 1086 .712 425 1086 .691 412 1071 .673 403 1077 .618 377 1094 230 348 1121 231 321 1155 233 299 1215 240 285 1333 247 272 1451 231 237 1431

8-66

Figure 8-13 (Sheet 10 of 19)

PM-133

Pilots Manual

LONG RANGE CRUISE

TWO ENGINE
WEIGHT 19,000 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.745 416 1097 .735 410 1050 .726 405 1030 .717 400 1031 .695 387 1016 .674 377 1016 .623 357 1042 233 331 1071 236 309 1123 244 296 1203 244 274 1250 240 250 1275 238 231 1343

.745 426 1128 .735 420 1077 .728 416 1067 .715 408 1060 .695 396 1042 .675 387 1043 .622 364 1060 236 341 1110 238 319 1156 240 296 1200 237 272 1239 234 249 1292 237 235 1376

.735 430 1107 .726 424 1093 .714 417 1084 .694 405 1068 .673 394 1068 .618 370 1085 233 344 1121 234 319 1154 234 295 1200 232 271 1251 233 252 1332 235 236 1413

.734 434 1119 .727 430 1105 .716 423 1103 .697 411 1088 .679 402 1093 .621 375 1105 233 348 1138 233 321 1163 234 297 1216 233 274 1279 232 253 1356 234 238 1437

.664 397 1043 .725 433 1114 .716 427 1114 .696 415 1100 .677 405 1102 .623 380 1119 234 353 1153 234 325 1181 236 303 1246 248 294 1392 258 283 1533 234 239 1464

PM-133

Figure 8-13 (Sheet 11 of 19)

8-67

Pilots Manual

LONG RANGE CRUISE

TWO ENGINE
WEIGHT 19,500 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.745 416 1135 .738 412 1086 .724 404 1049 .719 401 1055 .703 392 1047 .681 381 1044 .629 360 1067 237 336 1102 239 313 1148 246 298 1226 246 276 1269 242 252 1297 239 233 1368

.746 426 1168 .738 422 1114 .733 419 1099 .718 410 1087 .702 400 1075 .679 389 1069 .624 365 1078 239 346 1139 241 322 1179 242 298 1221 239 274 1260 237 252 1318 239 237 1403

.738 432 1145 .731 427 1126 .714 417 1105 .700 409 1100 .680 398 1099 .624 373 1112 236 349 1153 235 321 1174 237 298 1227 234 273 1277 235 254 1363 237 239 1442

.729 431 1132 .719 424 1129 .701 413 1115 .686 406 1124 .626 378 1131 236 353 1169 236 324 1190 238 302 1250 241 283 1341 253 276 1496 237 240 1469

.729 435 1147 .718 428 1141 .703 419 1133 .680 407 1126 .628 383 1144 237 357 1186 236 328 1208 239 307 1277 250 297 1418 261 287 1565 236 242 1497

8-68

Figure 8-13 (Sheet 12 of 19)

PM-133

Pilots Manual

LONG RANGE CRUISE

TWO ENGINE
WEIGHT 20,000 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.744 416 1173 .742 414 1126 .729 407 1084 .721 402 1078 .708 395 1075 .687 385 1073 .635 363 1092 240 340 1132 242 316 1173 248 301 1247 247 277 1287 244 254 1321 242 235 1395

.742 424 1155 .735 420 1127 .721 412 1113 .707 404 1106 .685 393 1099 .625 366 1096 240 348 1160 243 324 1202 244 301 1243 241 275 1282 239 254 1345 241 239 1431

.742 434 1186 .735 430 1160 .717 419 1132 .706 412 1131 .686 402 1128 .629 376 1138 239 353 1184 238 324 1200 240 302 1255 236 276 1304 237 257 1393 239 241 1473

.732 432 1163 .721 426 1156 .705 416 1144 .688 408 1147 .632 382 1158 240 358 1200 238 327 1217 241 306 1282 248 292 1400 262 286 1566 239 242 1499

.733 438 1180 .720 430 1166 .707 422 1163 .686 411 1157 .633 387 1170 240 362 1217 239 331 1236 242 311 1307 253 300 1445 263 289 1588 261 267 1672

PM-133

Figure 8-13 (Sheet 13 of 19)

8-69

Pilots Manual

LONG RANGE CRUISE

TWO ENGINE
WEIGHT 20,500 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.745 416 1223 .744 416 1163 .733 409 1119 .720 401 1096 .713 398 1104 .693 388 1101 .641 367 1119 244 345 1163 245 320 1200 251 303 1269 249 279 1307 245 256 1344 243 237 1420

.744 425 1195 .734 420 1150 .724 414 1140 .712 406 1135 .690 396 1128 .630 369 1121 242 350 1179 245 327 1224 246 303 1266 243 278 1306 241 256 1374 243 241 1459

.738 431 1192 .722 422 1167 .709 414 1158 .691 405 1158 .635 379 1166 243 358 1216 240 328 1227 243 305 1286 243 284 1359 256 277 1519 242 243 1504

.735 434 1198 .724 427 1181 .711 420 1177 .690 409 1169 .638 385 1186 243 362 1231 240 331 1244 244 310 1313 251 295 1427 264 288 1589 241 245 1531

.722 431 1192 .712 424 1192 .690 414 1186 .637 389 1194 244 367 1248 241 335 1263 245 314 1337 255 303 1471 265 291 1609 272 279 1765

8-70

Figure 8-13 (Sheet 14 of 19)

PM-133

Pilots Manual

LONG RANGE CRUISE

TWO ENGINE
WEIGHT 21,000 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.745 416 1200 .735 410 1153 .721 402 1123 .716 399 1129 .698 391 1129 .649 371 1149 247 350 1193 248 324 1228 253 306 1291 250 281 1327 247 258 1367 245 239 1447

.745 426 1234 .735 420 1180 .728 416 1169 .715 408 1162 .696 399 1158 .634 371 1146 243 352 1200 246 329 1244 247 305 1287 244 280 1330 243 259 1403 245 243 1487

.737 431 1217 .726 424 1200 .710 415 1180 .696 408 1187 .641 383 1195 246 363 1247 243 331 1257 246 309 1317 250 291 1410 265 287 1593 244 246 1536

.735 434 1226 .727 430 1211 .715 422 1207 .692 410 1193 .645 390 1216 246 367 1262 243 335 1274 247 314 1343 254 298 1456 266 290 1612 263 267 1690

.725 433 1221 .715 427 1220 .696 417 1217 .640 391 1217 247 371 1280 244 339 1294 248 318 1367 258 306 1497 267 293 1631 281 287 1841

PM-133

Figure 8-13 (Sheet 15 of 19)

8-71

Pilots Manual

LONG RANGE CRUISE

TWO ENGINE
WEIGHT 21,500 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.745 416 1239 .738 412 1188 .725 405 1155 .714 398 1146 .702 393 1155 .655 375 1177 250 354 1222 251 328 1256 255 308 1313 252 282 1346 249 260 1391 247 241 1475

.745 426 1273 .738 422 1218 .733 419 1200 .718 410 1188 .702 403 1190 .641 375 1178 246 355 1229 247 330 1263 249 307 1309 246 282 1356 245 261 1431 248 245 1517

.738 432 1251 .731 427 1233 .712 416 1206 .700 410 1214 .649 388 1230 249 367 1278 246 336 1289 249 313 1349 253 295 1444 267 289 1617 246 248 1568

.729 431 1238 .718 424 1234 .699 414 1228 .652 394 1249 249 371 1294 247 339 1307 250 318 1374 257 302 1484 268 292 1635 275 279 1783

.729 435 1253 .717 428 1246 .702 420 1249 .642 392 1240 250 376 1312 247 343 1327 251 322 1397 260 309 1524 269 295 1652 277 283 1814

8-72

Figure 8-13 (Sheet 16 of 19)

PM-133

Pilots Manual

LONG RANGE CRUISE

TWO ENGINE
WEIGHT 22,000 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.745 416 1280 .742 414 1228 .729 407 1189 .715 398 1168 .702 393 1174 .658 377 1199 253 358 1248 254 332 1285 256 310 1334 253 284 1368 251 262 1415 249 243 1502

.745 426 1316 .741 424 1258 .734 419 1227 .720 411 1213 .707 406 1221 .648 379 1212 249 360 1261 249 332 1284 251 310 1332 248 284 1382 248 263 1461 250 247 1547

.742 434 1293 .735 429 1266 .717 419 1240 .702 411 1238 .657 393 1264 251 371 1306 250 340 1322 252 317 1380 256 298 1474 270 292 1642 264 266 1700

.730 431 1265 .720 425 1260 .706 418 1264 .656 397 1276 252 375 1322 250 344 1340 253 321 1405 259 305 1512 270 295 1660 283 287 1859

.732 437 1287 .719 429 1273 .707 423 1281 .648 396 1272 252 379 1338 251 348 1360 254 326 1427 263 312 1551 271 298 1676 279 285 1837

PM-133

Figure 8-13 (Sheet 17 of 19)

8-73

Pilots Manual

LONG RANGE CRUISE

TWO ENGINE
WEIGHT 22,500 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.744 416 1320 .744 416 1264 .734 409 1224 .718 401 1199 .701 393 1193 .660 378 1220 254 360 1270 257 335 1313 258 312 1354 255 286 1391 252 264 1440 252 245 1532

.744 425 1299 .732 418 1248 .723 413 1239 .712 408 1250 .655 384 1246 251 363 1290 252 336 1315 253 313 1361 252 289 1419 258 274 1538 252 249 1575

.744 435 1332 .738 431 1298 .721 421 1272 .702 412 1261 .664 397 1297 253 374 1333 253 344 1354 255 320 1408 259 302 1504 272 294 1668 277 279 1803

.734 434 1302 .722 427 1287 .710 421 1296 .659 399 1301 254 378 1348 253 348 1374 256 324 1433 262 308 1540 272 297 1683 281 285 1848

.721 431 1299 .710 426 1311 .655 400 1306 253 381 1359 254 352 1392 256 328 1453 265 314 1577 273 300 1699 280 287 1859

8-74

Figure 8-13 (Sheet 18 of 19)

PM-133

Pilots Manual

LONG RANGE CRUISE

TWO ENGINE
WEIGHT 23,000 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.744 416 1370 .745 416 1300 .735 410 1255 .722 403 1231 .706 396 1225 .662 379 1241 255 360 1286 260 339 1341 259 314 1374 257 288 1414 254 265 1465 254 247 1562

.745 426 1338 .734 419 1281 .726 415 1267 .715 410 1277 .661 387 1276 254 366 1318 255 340 1346 256 316 1387 256 294 1458 267 283 1607 254 252 1606

.738 431 1325 .725 423 1304 .707 415 1294 .668 400 1325 256 377 1359 256 348 1387 257 323 1435 262 305 1533 274 297 1693 283 285 1858

.735 434 1333 .725 428 1315 .715 424 1327 .662 400 1325 256 381 1375 256 352 1406 258 327 1459 264 311 1569 275 299 1708 283 287 1869

.723 431 1325 .714 428 1340 .661 404 1339 254 382 1379 257 356 1425 259 331 1480 268 317 1604 274 302 1721 282 289 1881

PM-133

Figure 8-13 (Sheet 19 of 19)

8-75

Pilots Manual

HIGH SPEED CRUISE

WEIGHT 14,000 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind KTAS 45 Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS KTAS 5 Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 .763 427 887 .780 437 910 .780 437 914 .780 437 933 .780 437 961 .790 443 1053 .800 449 1178 .810 456 1346 .810 458 1459 .810 469 1825 330 462 1997 340 440 2054 340 408 2057 340 380 2100 300 312 1729 300 291 1815 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.768 440 912 .780 447 962 .780 447 980 .780 447 1013 .790 454 1107 .800 460 1233 .810 466 1410 .810 469 1531 .810 479 1912 330 472 2098 340 449 2156 340 416 2163 340 387 2189 300 318 1799 300 296 1886

.726 424 832 .768 450 937 .780 457 1029 .780 457 1065 .790 464 1164 .800 470 1297 .810 477 1479 .810 479 1602 .810 490 2001 330 481 2183 340 458 2247 340 424 2257 340 394 2280 300 323 1861 300 301 1959

.745 440 880 .778 461 1012 .780 462 1088 .790 469 1190 .800 475 1326 .810 482 1513 .810 485 1638 .810 495 2047 330 486 2233 340 462 2287 340 428 2298 340 398 2324 300 326 1894 300 304 1992

.707 422 821 .764 457 967 .780 467 1115 .790 474 1220 .800 481 1359 .810 487 1549 .810 490 1677 .805 497 1951 330 491 2280 340 467 2338 340 432 2340 340 401 2368 300 329 1930 300 306 2027

Figure 8-14 (Sheet 1 of 19) 8-76

PM-133

Pilots Manual

HIGH SPEED CRUISE

WEIGHT 14,500 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind KTAS 45 Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS KTAS 5 Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 .744 416 876 .780 437 941 .780 437 939 .780 437 955 .780 437 981 .790 443 1071 .800 449 1194 .810 456 1363 .810 458 1470 .810 469 1831 330 462 1999 340 440 2055 340 408 2058 340 380 2102 300 312 1731 300 291 1818 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.756 432 903 .780 447 988 .780 447 1004 .780 447 1031 .790 454 1124 .800 460 1248 .810 466 1429 .810 469 1543 .810 479 1917 330 472 2100 340 449 2157 340 416 2164 340 387 2191 300 318 1802 300 296 1888

.760 445 936 .780 457 1052 .780 457 1084 .790 464 1183 .800 470 1314 .810 477 1497 .810 479 1613 .810 490 2007 330 481 2186 340 458 2248 340 424 2258 340 394 2282 300 324 1863 300 301 1961

.730 431 871 .772 458 1008 .780 462 1108 .790 469 1208 .800 475 1342 .810 482 1532 .810 485 1650 .810 495 2052 330 486 2236 340 462 2288 340 428 2299 340 398 2326 300 326 1896 300 304 1995

.755 451 960 .780 467 1135 .790 474 1239 .800 481 1376 .810 487 1568 .810 490 1689 .805 496 1950 330 491 2283 340 467 2339 340 432 2341 340 401 2371 300 329 1932 300 306 2029

Figure 8-14 (Sheet 2 of 19)

PM-133

8-77

Pilots Manual

HIGH SPEED CRUISE

WEIGHT 15,000 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS KTAS S.L. Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.780 437 977 .780 437 965 .780 437 977 .780 437 1000 .790 443 1090 .800 449 1212 .810 456 1380 .810 458 1480 .810 469 1837 330 462 2002 340 440 2057 340 408 2060 340 380 2103 300 312 1734 300 291 1820

.737 421 890 .780 447 1017 .780 447 1028 .780 447 1050 .790 453 1142 .800 460 1265 .810 466 1449 .810 469 1555 .810 479 1923 330 472 2103 340 449 2159 340 416 2166 340 387 2193 300 318 1804 300 296 1891

.751 439 934 .780 457 1079 .780 457 1103 .790 464 1201 .800 470 1332 .810 477 1515 .810 479 1624 .810 490 2013 330 481 2189 340 458 2250 340 424 2260 340 394 2284 300 324 1866 300 301 1964

.704 416 854 .765 453 1004 .780 462 1129 .790 469 1228 .800 475 1361 .810 482 1553 .810 485 1663 .810 495 2058 330 486 2239 340 462 2290 340 428 2301 340 398 2328 300 326 1899 300 304 1997

.745 445 952 .779 466 1100 .790 474 1258 .800 481 1395 .810 487 1588 .810 490 1701 .804 496 1949 330 491 2286 340 467 2341 340 432 2343 340 401 2373 300 329 1935 300 306 2032

Figure 8-14 (Sheet 3 of 19) 8-78

PM-133

Pilots Manual

HIGH SPEED CRUISE

WEIGHT 15,500 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS KTAS S.L. Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.774 433 1011 .780 437 993 .780 437 1001 .780 437 1021 .790 443 1110 .800 449 1230 .810 456 1398 .810 458 1493 .810 469 1844 330 462 2004 340 440 2058 340 408 2061 340 380 2105 300 312 1736 300 291 1822

.778 446 1037 .780 447 1053 .780 447 1070 .790 453 1160 .800 460 1286 .810 466 1470 .810 469 1570 .810 479 1929 330 472 2105 340 449 2160 340 416 2168 340 387 2195 300 318 1806 300 296 1893

.739 432 932 .778 456 1070 .780 457 1124 .790 464 1221 .800 470 1351 .810 477 1534 .810 479 1638 .810 490 2020 330 481 2192 340 458 2251 340 424 2262 340 394 2287 300 324 1868 300 301 1966

.757 448 999 .780 462 1151 .790 469 1249 .800 475 1382 .810 482 1574 .810 485 1679 .810 495 2065 330 486 2241 340 462 2292 340 428 2302 340 398 2330 300 326 1901 300 304 2000

.731 436 942 .774 463 1096 .790 474 1279 .800 481 1415 .808 486 1522 .810 490 1717 .803 495 1947 330 491 2289 340 467 2343 340 432 2345 340 401 2375 300 329 1937 300 306 2035

Figure 8-14 (Sheet 4 of 19)

PM-133

8-79

Pilots Manual

HIGH SPEED CRUISE

WEIGHT 16,000 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS KTAS S.L. Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.761 426 1002 .780 437 1024 .780 437 1026 .780 437 1043 .790 443 1131 .800 449 1250 .810 456 1417 .810 458 1509 .810 469 1854 330 462 2007 340 440 2060 340 408 2063 340 380 2107 300 312 1740 300 291 1825

.769 440 1028 .780 447 1079 .780 447 1094 .790 453 1183 .800 460 1308 .810 466 1491 .810 469 1589 .810 479 1938 330 472 2108 340 449 2162 340 416 2169 340 387 2198 300 318 1810 300 296 1896

.721 422 929 .771 452 1068 .780 457 1148 .790 464 1242 .800 470 1370 .810 477 1557 .810 479 1656 .810 490 2030 330 481 2195 340 458 2253 340 424 2263 340 394 2289 300 324 1872 300 301 1970

.747 441 993 .780 462 1176 .790 469 1272 .800 475 1404 .810 482 1595 .810 485 1698 .810 495 2075 330 486 2244 340 462 2294 340 428 2304 340 398 2333 300 326 1905 300 304 2004

.712 425 928 .767 459 1091 .787 472 1228 .799 480 1368 .807 485 1519 .809 489 1654 .803 495 1946 330 491 2292 340 467 2345 340 432 2346 340 401 2377 300 329 1941 300 306 2038

Figure 8-14 (Sheet 5 of 19) 8-80

PM-133

Pilots Manual

HIGH SPEED CRUISE

WEIGHT 16,500 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS KTAS S.L. Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.743 415 988 .780 437 1060 .780 437 1051 .780 437 1065 .790 443 1152 .800 449 1269 .810 456 1436 .810 458 1526 .810 469 1864 330 462 2010 340 440 2062 340 408 2064 340 380 2109 300 312 1744 300 291 1829

.757 433 1017 .780 447 1107 .780 447 1118 .790 453 1207 .800 460 1330 .810 466 1513 .810 469 1608 .810 479 1948 330 472 2111 340 449 2164 340 416 2171 340 387 2200 300 318 1814 300 296 1901

.765 448 1066 .780 457 1172 .790 464 1266 .800 470 1393 .810 477 1581 .810 479 1677 .810 490 2040 330 481 2198 340 458 2255 340 424 2265 340 394 2291 300 324 1877 300 301 1974

.734 433 985 .777 461 1148 .790 469 1295 .800 475 1426 .810 482 1618 .810 485 1718 .810 495 2022 330 486 2247 340 462 2296 340 428 2306 340 398 2335 300 326 1910 300 304 2008

.760 455 1085 .783 469 1225 .797 478 1366 .804 484 1516 .807 488 1651 .802 494 1945 330 491 2295 340 467 2347 340 432 2348 340 401 2380 300 329 1946 300 306 2043

Figure 8-14 (Sheet 6 of 19)

PM-133

8-81

Pilots Manual

HIGH SPEED CRUISE

WEIGHT 17,000 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.780 437 1099 .780 437 1079 .780 437 1088 .790 443 1174 .800 449 1290 .810 456 1458 .810 458 1543 .810 469 1874 330 462 2013 340 440 2064 340 408 2066 340 380 2111 300 312 1748 300 291 1833

.739 423 1001 .780 447 1136 .780 447 1143 .790 453 1232 .800 460 1354 .810 466 1538 .810 469 1628 .810 479 1958 330 472 2114 340 449 2166 340 416 2173 340 387 2202 300 318 1819 300 296 1905

.756 442 1064 .780 457 1198 .790 464 1291 .800 470 1417 .810 477 1608 .810 479 1699 .810 490 2050 330 481 2201 340 458 2257 340 424 2267 340 394 2294 300 324 1881 300 301 1979

.712 420 972 .771 457 1144 .790 469 1323 .800 475 1449 .810 482 1646 .810 485 1739 .809 494 2020 330 486 2251 340 462 2298 340 428 2308 340 398 2337 300 326 1914 300 304 2013

.752 449 1079 .779 467 1221 .794 476 1363 .802 482 1512 .806 487 1648 .801 494 1943 330 491 2299 340 467 2349 340 432 2350 340 401 2382 300 329 1951 300 306 2048

Figure 8-14 (Sheet 7 of 19) 8-82

PM-133

Pilots Manual

HIGH SPEED CRUISE

WEIGHT 17,500 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.771 432 1126 .780 437 1109 .780 437 1111 .790 443 1196 .800 449 1312 .810 456 1481 .810 458 1561 .810 469 1885 330 462 2016 340 440 2066 340 408 2068 340 380 2113 300 312 1752 300 291 1838

.779 447 1165 .780 447 1169 .790 453 1256 .800 460 1379 .810 466 1565 .810 469 1649 .810 479 1968 330 472 2117 340 449 2168 340 417 2175 340 387 2205 300 318 1823 300 296 1909

.746 436 1061 .780 457 1226 .790 464 1317 .800 470 1444 .810 477 1638 .810 479 1721 .810 490 2061 330 481 2204 340 458 2258 340 424 2269 340 394 2296 300 324 1886 300 301 1983

.765 453 1139 .788 468 1293 .800 475 1479 .809 482 1611 .810 485 1761 .808 493 2018 330 486 2254 340 462 2300 340 428 2310 340 398 2340 300 327 1919 300 304 2017

.740 442 1070 .774 464 1218 .790 474 1360 .800 481 1509 .804 485 1645 .801 493 1942 330 491 2302 340 467 2351 340 432 2352 340 401 2384 300 329 1956 300 306 2053

Figure 8-14 (Sheet 8 of 19)

PM-133

8-83

Pilots Manual

HIGH SPEED CRUISE

WEIGHT 18,000 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.760 425 1116 .780 437 1144 .780 437 1136 .790 443 1219 .800 449 1336 .810 456 1509 .810 458 1581 .810 469 1896 330 462 2019 340 440 2068 340 408 2070 340 380 2116 300 312 1757 300 291 1842

.771 442 1156 .780 447 1196 .790 453 1282 .800 460 1406 .810 466 1593 .810 469 1670 .810 479 1979 330 472 2120 340 449 2170 340 417 2177 340 387 2207 300 318 1828 300 296 1914

.732 428 1057 .778 456 1220 .790 464 1344 .800 470 1473 .810 477 1668 .810 479 1744 .810 490 2072 330 481 2208 340 458 2260 340 425 2271 340 394 2298 300 324 1890 300 302 1988

.757 448 1134 .784 465 1289 .799 475 1445 .807 480 1609 .810 485 1785 .807 493 2017 330 486 2257 340 462 2302 340 428 2312 340 398 2342 300 327 1924 300 304 2022

.727 434 1060 .769 460 1213 .787 472 1357 .799 480 1507 .802 484 1642 .800 493 1941 330 491 2306 340 467 2353 340 432 2354 340 401 2387 300 329 1961 300 307 2058

Figure 8-14 (Sheet 9 of 19) 8-84

PM-133

Pilots Manual

HIGH SPEED CRUISE

WEIGHT 18,500 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.742 415 1100 .780 437 1182 .780 437 1163 .790 443 1243 .800 449 1360 .810 456 1538 .810 458 1605 .810 469 1908 330 462 2023 340 440 2070 340 408 2072 340 380 2118 300 312 1762 300 291 1846

.760 435 1145 .780 447 1224 .790 453 1308 .800 460 1434 .810 466 1622 .810 469 1693 .810 479 1990 330 472 2124 340 449 2172 340 417 2179 340 387 2210 300 318 1833 300 296 1918

.773 453 1217 .790 464 1371 .800 470 1502 .810 477 1698 .810 479 1770 .810 490 2084 330 481 2212 340 458 2262 340 425 2273 340 394 2301 300 324 1895 300 302 1993

.747 441 1127 .780 462 1286 .796 473 1443 .805 479 1606 .809 484 1741 .807 492 2015 330 486 2262 340 462 2304 340 428 2314 340 398 2345 300 327 1929 300 304 2027

.701 418 1042 .762 456 1208 .783 469 1354 .796 478 1504 .800 483 1639 .799 492 1940 330 491 2311 340 467 2355 340 432 2356 340 401 2389 300 329 1966 300 307 2063

Figure 8-14 (Sheet 10 of 19)

PM-133

8-85

Pilots Manual

HIGH SPEED CRUISE

WEIGHT 19,000 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.780 437 1224 .780 437 1191 .790 443 1267 .800 449 1387 .810 456 1568 .810 458 1631 .810 469 1919 330 462 2027 340 440 2072 340 408 2074 340 380 2120 300 312 1766 300 291 1851

.745 426 1129 .780 447 1253 .790 453 1335 .800 460 1462 .810 466 1651 .810 469 1719 .810 479 2001 330 472 2128 340 449 2174 340 417 2181 340 387 2212 300 318 1837 300 296 1923

.765 448 1213 .790 464 1400 .800 470 1532 .810 477 1730 .810 479 1798 .810 490 2096 330 481 2217 340 458 2264 340 425 2275 340 394 2304 300 324 1900 300 302 1998

.734 434 1119 .775 459 1283 .793 471 1440 .803 478 1603 .807 483 1739 .806 492 2014 330 486 2266 340 462 2307 340 428 2316 340 398 2347 300 327 1934 300 304 2032

.755 451 1203 .780 467 1351 .793 476 1501 .799 482 1636 .798 492 1938 330 491 2315 340 467 2357 340 432 2358 340 401 2392 300 329 1971 300 307 2068

Figure 8-14 (Sheet 11 of 19) 8-86

PM-133

Pilots Manual

HIGH SPEED CRUISE

WEIGHT 19,500 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.774 433 1277 .780 437 1225 .790 443 1294 .800 449 1416 .810 456 1599 .810 458 1659 .810 469 1934 330 462 2031 340 440 2074 340 408 2076 340 380 2122 300 312 1771 300 291 1856

.780 447 1288 .790 453 1362 .800 460 1491 .810 466 1682 .810 469 1746 .810 479 2015 330 472 2132 340 449 2176 340 417 2183 340 387 2215 300 318 1842 300 297 1928

.758 443 1209 .787 462 1375 .800 470 1564 .810 477 1718 .810 479 1827 .810 490 2111 330 481 2221 340 458 2267 340 425 2277 340 394 2306 300 324 1905 300 302 2003

.705 416 1101 .770 456 1279 .790 469 1437 .801 476 1600 .805 481 1736 .804 491 2011 330 486 2271 340 462 2309 340 428 2318 340 398 2350 300 327 1939 300 304 2038

.746 446 1195 .775 464 1347 .790 474 1498 .796 480 1633 .796 490 1935 330 491 2320 340 467 2360 340 432 2360 340 401 2395 300 329 1976 300 307 2074

Figure 8-14 (Sheet 12 of 19)

PM-133

8-87

Pilots Manual

HIGH SPEED CRUISE

WEIGHT 20,000 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.765 428 1266 .780 437 1260 .790 443 1326 .800 449 1449 .810 455 1636 .810 458 1687 .810 469 1952 330 462 2035 340 440 2078 340 408 2080 340 380 2125 300 312 1776 300 291 1860

.777 445 1307 .790 453 1394 .800 460 1525 .810 466 1718 .810 469 1773 .810 479 2032 330 472 2136 340 449 2180 340 417 2187 340 387 2218 300 318 1847 300 297 1933

.747 437 1203 .783 459 1372 .799 469 1540 .807 475 1714 .810 479 1857 .810 490 2130 330 481 2226 340 458 2270 340 425 2281 340 394 2310 300 324 1911 300 302 2008

.763 452 1274 .786 466 1434 .799 475 1597 .803 480 1733 .803 490 2009 330 486 2275 340 462 2313 340 429 2322 340 398 2353 300 327 1944 300 304 2043

.735 439 1188 .770 461 1343 .787 472 1495 .794 479 1631 .795 489 1933 330 491 2325 340 467 2364 340 432 2364 340 401 2398 300 329 1982 300 307 2079

Figure 8-14 (Sheet 13 of 19) 8-88

PM-133

Pilots Manual

HIGH SPEED CRUISE

WEIGHT 20,500 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.752 420 1251 .780 437 1296 .790 443 1358 .800 449 1484 .810 455 1673 .810 458 1715 .810 469 1971 330 462 2040 340 440 2081 340 408 2084 340 380 2129 300 312 1781 300 291 1865

.769 441 1298 .790 453 1427 .800 460 1559 .810 466 1757 .810 469 1801 .810 479 2051 330 472 2141 340 449 2183 340 417 2191 340 387 2223 300 318 1853 300 297 1938

.733 428 1196 .778 456 1369 .795 467 1537 .805 473 1709 .810 479 1847 .809 489 2096 330 481 2231 340 458 2274 340 425 2285 340 394 2314 300 324 1916 300 302 2013

.755 446 1268 .782 464 1431 .797 473 1595 .802 479 1730 .802 489 2006 330 486 2280 340 462 2317 340 429 2326 340 398 2358 300 327 1950 300 304 2048

.719 429 1175 .765 457 1339 .783 470 1492 .791 477 1628 .793 488 1930 330 491 2238 340 467 2368 340 432 2369 340 401 2403 300 330 1988 300 307 2085

Figure 8-14 (Sheet 14 of 19)

PM-133

8-89

Pilots Manual

HIGH SPEED CRUISE

WEIGHT 21,000 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.780 437 1338 .790 443 1391 .800 449 1520 .810 455 1711 .810 458 1744 .810 469 1990 330 462 2044 340 440 2084 340 408 2089 340 380 2134 300 313 1786 300 291 1870

.758 434 1285 .790 453 1462 .800 460 1596 .810 466 1798 .810 469 1830 .810 479 2072 330 472 2145 340 449 2186 340 417 2196 340 387 2228 300 318 1858 300 297 1943

.772 452 1364 .793 466 1535 .803 472 1706 .808 478 1843 .808 488 2093 330 481 2235 340 458 2277 340 425 2290 340 394 2319 300 324 1921 300 302 2018

.746 441 1261 .778 461 1428 .794 471 1592 .800 478 1728 .801 488 2004 330 486 2285 340 462 2320 340 429 2330 340 398 2363 300 327 1955 300 304 2054

.758 453 1333 .780 467 1489 .788 475 1625 .791 487 1927 330 491 2237 340 467 2371 340 432 2373 340 401 2408 300 330 1993 300 307 2091

Figure 8-14 (Sheet 15 of 19) 8-90

PM-133

Pilots Manual

HIGH SPEED CRUISE

WEIGHT 21,500 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.780 437 1397 .790 443 1425 .800 449 1556 .810 455 1751 .810 458 1776 .810 469 2009 330 462 2049 340 440 2088 340 408 2093 340 380 2138 300 313 1792 300 291 1875

.743 425 1267 .787 451 1466 .800 460 1634 .809 466 1827 .810 469 1863 .810 479 2093 330 472 2150 340 449 2190 340 417 2200 340 387 2232 300 318 1863 300 297 1949

.765 448 1359 .790 464 1532 .801 471 1703 .806 477 1839 .807 487 2091 330 481 2240 340 458 2281 340 425 2294 340 394 2324 300 324 1927 300 302 2024

.733 433 1252 .773 458 1424 .790 469 1588 .798 476 1725 .799 487 2001 330 486 2290 340 463 2324 340 429 2335 340 398 2368 300 327 1961 300 304 2060

.752 449 1328 .776 465 1486 .785 473 1622 .790 486 1925 329 490 2236 340 467 2375 340 432 2378 340 401 2413 300 330 1999 300 307 2097

Figure 8-14 (Sheet 16 of 19)

PM-133

8-91

Pilots Manual

HIGH SPEED CRUISE

WEIGHT 22,000 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.773 433 1439 .790 443 1473 .800 449 1593 .810 455 1792 .810 458 1812 .810 469 2029 330 462 2053 340 440 2092 340 409 2098 340 380 2142 300 313 1797 300 291 1881

.781 448 1458 .799 459 1640 .807 465 1823 .810 469 1901 .810 479 2115 330 472 2154 340 449 2194 340 417 2205 340 387 2237 300 318 1869 300 297 1954

.758 443 1353 .786 461 1528 .799 470 1700 .804 475 1835 .805 486 2087 330 481 2246 340 458 2284 340 425 2299 340 394 2329 300 324 1933 300 302 2030

.700 413 1231 .768 455 1419 .788 467 1586 .795 475 1722 .798 486 1999 330 486 2295 340 463 2328 340 429 2340 340 398 2373 300 327 1967 300 304 2066

.742 443 1320 .772 462 1482 .782 471 1618 .788 484 1922 329 490 2235 340 467 2379 340 432 2382 340 401 2418 300 330 2005 300 307 2103

Figure 8-14 (Sheet 17 of 19) 8-92

PM-133

Pilots Manual

HIGH SPEED CRUISE

WEIGHT 22,500 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.765 428 1427 .790 443 1531 .800 449 1649 .810 455 1854 .810 458 1848 .810 469 2050 330 462 2060 340 440 2095 340 409 2103 340 380 2147 300 313 1803 300 292 1886

.775 444 1450 .795 457 1635 .804 463 1818 .810 469 1940 .810 479 2137 330 472 2161 340 449 2197 340 417 2210 340 387 2243 300 318 1874 300 297 1959

.747 437 1346 .782 459 1524 .797 468 1697 .802 474 1832 .804 485 2084 330 481 2253 340 458 2288 340 425 2304 340 394 2334 300 324 1938 300 302 2035

.762 451 1414 .784 465 1582 .792 473 1718 .796 485 1996 330 486 2303 340 463 2332 340 429 2344 340 398 2378 300 327 1973 300 304 2072

.732 437 1312 .767 459 1477 .779 469 1615 .786 483 1918 329 489 2234 340 467 2383 340 432 2387 340 402 2423 300 330 2012 300 307 2109

Figure 8-14 (Sheet 18 of 19)

PM-133

8-93

Pilots Manual

HIGH SPEED CRUISE

WEIGHT 23,000 LB Mach Ind 51 KTAS Fuel - Lb/Hr Mach Ind 49 KTAS Fuel - Lb/Hr Mach Ind 47 KTAS Fuel - Lb/Hr Mach Ind 45 KTAS Fuel - Lb/Hr Mach Ind 43 KTAS Fuel - Lb/Hr Mach Ind 41 KTAS Fuel - Lb/Hr Mach Ind 39 KTAS Fuel - Lb/Hr Mach Ind 37 KTAS Fuel - Lb/Hr Mach Ind 35 KTAS Fuel - Lb/Hr Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr KIAS S.L. KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET ISA -10 TEMPERATURE C ISA ISA +10 ISA +15 ISA +20

.753 421 1410 .790 443 1631 .800 449 1710 .810 455 1924 .810 458 1886 .810 469 2070 330 462 2071 340 440 2099 340 409 2107 340 380 2151 300 313 1808 300 292 1891

.766 439 1438 .792 455 1630 .803 461 1815 .809 468 1970 .810 479 2160 330 472 2172 340 449 2201 340 417 2214 340 387 2248 300 318 1880 300 297 1965

.732 428 1335 .778 456 1521 .794 466 1694 .800 473 1829 .802 484 2082 330 481 2265 340 458 2292 340 425 2309 340 395 2340 300 324 1944 300 302 2041

.754 446 1407 .781 463 1579 .790 471 1716 .794 483 1993 330 486 2315 340 463 2336 340 429 2349 340 398 2384 300 327 1979 300 304 2078

.715 427 1299 .762 456 1472 .776 467 1611 .783 481 1914 328 488 2232 340 467 2388 340 433 2392 340 402 2429 300 330 2018 300 307 2115

Figure 8-14 (Sheet 19 of 19) 8-94

PM-133

Pilots Manual

MAXIMUM RANGE DESCENT - ONE ENGINE

ALTITUDE ~ FT 51,000 to 49,000 49,000 to 29,000 29,000 to 21,000 21,000 and below

DESCENT SPEED 0.70 MI 170 KIAS 0.45 MI 200 KIAS

NOTE: This table represents the minimum sink-rate speed above the single-engine service ceiling and approximates the best rate-of-climb speed below the single-engine service ceiling.

Figure 8-15

PM-133

8-95

Pilots Manual

LONG RANGE CRUISE

ONE ENGINE
TEMPERATURE C ISA ISA +10 ISA +15 .533 .534 .538 312 319 324 722 761 782 198 196 198 289 292 297 728 753 776 199 200 199 267 273 274 739 773 783 202 202 202 250 254 258 760 794 814 202 205 208 231 239 244 787 838 867 209 214 216 222 232 235 859 919 939

WEIGHT 14,000 LB Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET

ISA -10 .536 306 695 198 282 693 199 261 703 200 243 724 200 225 742 203 212 794

ISA +20 .533 325 791 198 300 795 199 277 802 202 260 829 211 250 897 216 237 952

WEIGHT 15,000 LB Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET

ISA -10 .558 319 758 205 292 750 205 270 756 207 251 774 207 233 796 211 220 847

TEMPERATURE C ISA ISA +10 ISA +15 .545 .559 .552 319 333 333 774 835 840 205 204 205 298 303 308 783 818 839 206 205 205 277 280 283 794 821 842 208 208 207 257 262 264 810 848 861 209 212 214 239 247 252 842 893 921 216 221 222 230 239 242 911 971 989

ISA +20 .554 337 862 205 311 858 206 287 862 206 265 874 213 253 931 219 241 987

Figure 8-16 (Sheet 1 of 5) 8-96

PM-133

Pilots Manual

LONG RANGE CRUISE

ONE ENGINE
TEMPERATURE C ISA ISA +10 ISA +15 .567 .568 .570 332 339 344 849 886 910 208 211 211 303 313 317 827 881 899 212 212 212 284 289 292 845 881 901 213 212 212 264 268 270 859 892 912 216 218 217 247 255 255 898 945 956 223 228 224 237 246 245 964 1022 1020

WEIGHT 16,000 LB Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET

ISA -10 .566 324 801 212 301 806 212 278 809 213 258 824 214 241 850 218 227 900

ISA +20 .548 334 886 210 318 912 212 296 922 212 273 930 215 255 966 219 241 1011

WEIGHT 17,000 LB Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET

ISA -10 .584 334 868 217 309 858 219 287 864 219 266 876 221 248 905 224 234 953

TEMPERATURE C ISA ISA +10 ISA +15 .588 .584 .561 344 349 339 921 955 929 216 217 216 314 322 323 898 942 953 216 218 218 290 298 301 892 941 962 219 218 219 271 275 278 912 949 971 223 221 221 255 258 260 953 984 1000 229 229 225 244 248 245 1016 1052 1045

ISA +20 .503 307 873 217 327 978 219 304 982 219 281 991 221 263 1025 220 242 1038

Figure 8-16 (Sheet 2 of 5)

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LONG RANGE CRUISE

ONE ENGINE
TEMPERATURE C ISA ISA +10 ISA +15 .600 .570 .505 351 341 305 980 965 907 224 222 223 325 329 333 968 998 1022 223 224 224 299 306 309 955 1001 1021 224 225 225 277 283 286 961 1008 1031 228 226 226 261 263 267 1002 1032 1057 236 230 227 251 249 248 1069 1080 1082

WEIGHT 18,000 LB Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET

ISA -10 .603 345 938 220 312 900 225 295 918 226 274 929 227 255 957 231 241 1006

ISA +20

215 326 1004 223 311 1037 225 289 1052 227 269 1080 227 250 1100

WEIGHT 19,000 LB Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET

ISA -10 .610 349 991 227 322 970 227 298 957 230 279 977 233 261 1006 237 248 1058

TEMPERATURE C ISA ISA +10 ISA +15 .609 356 1036 232 230 221 336 340 331 1038 1073 1046 230 230 229 308 314 315 1018 1059 1077 230 231 231 284 291 294 1019 1067 1090 231 231 232 264 270 273 1042 1089 1112 238 234 233 253 253 255 1105 1123 1141

ISA +20

202 306 989 230 319 1102 231 296 1110 232 275 1135 233 256 1161

Figure 8-16 (Sheet 3 of 5) 8-98

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LONG RANGE CRUISE

ONE ENGINE
TEMPERATURE C ISA ISA +10 ISA +15 .559 327 1000 234 228 207 340 338 311 1084 1097 1031 236 234 235 315 320 324 1077 1115 1142 236 236 236 291 298 300 1077 1126 1146 236 237 237 271 276 279 1097 1146 1171 240 240 239 255 259 261 1135 1180 1202

WEIGHT 20,000 LB Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET

ISA -10 .613 351 1043 234 332 1040 233 305 1015 234 283 1020 238 267 1056 244 254 1111

ISA +20

231 321 1140 235 302 1166 238 282 1195 239 263 1224

WEIGHT 21,000 LB Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET

ISA -10 .569 326 1020 241 342 1110 240 314 1083 240 291 1081 241 271 1100 247 258 1153

TEMPERATURE C ISA ISA +10 ISA +15

ISA +20

240 348 1151 241 323 1136 241 298 1134 242 277 1154 244 260 1184

216 321 1085 240 328 1179 242 304 1184 243 283 1205 245 266 1239

237 326 1184 240 306 1201 243 286 1231 245 268 1261

222 310 1130 241 309 1230 243 288 1253 245 270 1285

Figure 8-16 (Sheet 4 of 5)

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LONG RANGE CRUISE

ONE ENGINE
TEMPERATURE C ISA ISA +10 ISA +15

WEIGHT 22,000 LB Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET

ISA -10

ISA +20

247 349 1172 247 323 1152 246 298 1141 247 277 1156 249 260 1185

242 351 1199 245 327 1187 247 305 1194 248 284 1212 249 265 1237

244 332 1232 246 309 1236 249 290 1266 250 271 1291

229 316 1175 246 313 1266 248 292 1287 250 273 1316

245 314 1280 248 294 1311 250 275 1342

WEIGHT 23,000 LB Mach Ind 30 KTAS Fuel - Lb/Hr KIAS 25 KTAS Fuel - Lb/Hr KIAS 20 KTAS Fuel - Lb/Hr KIAS 15 KTAS Fuel - Lb/Hr KIAS 10 KTAS Fuel - Lb/Hr KIAS 5 KTAS Fuel - Lb/Hr ALTITUDE 1000 FEET

ISA -10

TEMPERATURE C ISA ISA +10 ISA +15

ISA +20

251 355 1231 254 332 1219 252 305 1201 252 283 1212 253 264 1232

230 334 1180 251 335 1254 252 311 1252 253 290 1271 254 270 1290

236 323 1224 251 316 1300 253 295 1321 254 275 1344

210 291 1153 252 320 1329 253 297 1344 254 278 1370

238 306 1273 253 301 1374 254 280 1397

Figure 8-16 (Sheet 5 of 5) 8-100

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DESCENT AND HOLDING PERFORMANCE

The descent and holding performance on the following pages is based on flight test data and represents the average delivered aircraft. DESCENT PERFORMANCE SCHEDULE Figures 8-17 and 8-18 show times, distance and fuel used, for normal and high speed descents respectively, from a given altitude to sea level. An average descent weight of 16,000 pounds is assumed in the tables. Subtraction of performance values for two altitudes results in the time, distance and fuel required for descent between the two altitudes. The descent speed schedule is presented with each table. The power setting for descent is IDLE thrust. Data are shown without the use of spoilers. Descent performance is improved if spoilers are deployed. HOLDING OPERATIONS Figure 8-19 shows fuel flows and holding speed for various weights and altitude conditions. The holding speeds presented are sufficient to ensure a comfortable margin above shaker operation or low-speed buffet while maneuvering in a holding pattern.

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DESCENT PERFORMANCE SCHEDULE

NORMAL DESCENT

ALTITUDE 1000 Ft.

TIME Min.

DISTANCE N.M.

FUEL Lb.

51 49 47 45 43 41 39 37 35 33 31 29 27 25 23 21 19 17 15 13 11 9 7 5

17.6 16.6 15.4 14.1 12.9 11.9 11.0 10.2 9.6 9.1 8.6 8.3 7.9 7.5 7.1 6.6 6.2 5.8 5.3 4.9 4.4 3.7 2.9 2.1

114 106 97 88 80 72 66 60 55 52 48 46 43 40 37 34 31 28 25 23 20 16 13 9

167 157 144 131 118 107 98 90 83 78 74 70 67 63 59 55 51 48 44 41 37 31 25 19

DESCENT SPEED: 51,000 to 28,000 feet ..............................0.76 MI 28,000 to 10,000 feet .......................... 300 KIAS 10,000 feet and below ......................... 250 KIAS Figure 8-17 8-102
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DESCENT PERFORMANCE SCHEDULE

HIGH SPEED DESCENT
ALTITUDE 1000 Ft. TIME Min. DISTANCE N.M. FUEL Lb.

51 49 47 45 43 41 39 37 35 33 31 29 27 25 23 21 19 17 15 13 11 9 7 5

16.3 15.2 14.0 12.8 11.6 10.6 9.9 9.3 8.8 8.4 8.0 7.7 7.5 7.2 6.8 6.5 6.1 5.8 5.4 5.0 4.5 3.7 2.9 2.1

106 98 89 80 72 65 59 55 51 48 45 43 41 38 36 33 31 28 26 23 20 16 13 9

154 144 131 118 105 95 87 81 76 72 69 66 63 60 57 54 51 48 45 42 38 31 25 19

NOTE: The speed schedule portrayed below occurs when highspeed descent feature has been selected in the LVL CHG (Level Change) mode of the autopilot
DESCENT SPEED: 51,000 to 26,800 feet ............................. 0.76 MI 43,000 to 37,000 feet ................. 0.76 to 0.79 MI 37,000 to 27,000 feet ............................. 0.79 MI 27,000 to 14,500 feet .......................... 320 KIAS 14,500 to 15,000 feet ............... 330 to 250 KIAS 10,500 feet and below 250 KIAS Figure 8-18 PM-133 8-103

Pilots Manual

HOLDING OPERATIONS

WEIGHT 1000 LB 15 Mach Ind 41 Fuel - Lb/Hr Mach Ind 39 Fuel - Lb/Hr Mach Ind 37 Fuel - Lb/Hr Mach Ind 35 Fuel - Lb/Hr 784 .566 787 .541 789 .515 790 170 740 170 788 170 837 170 875 170 903 828 .581 837 .559 845 .537 853 175 787 175 831 175 877 175 915 175 949 871 .594 885 .575 898 .555 910 180 835 180 875 180 918 180 956 180 995 929 .612 943 .594 957 .575 970 185 882 185 919 185 960 185 999 185 1043 984 .627 999 .609 1011 .590 1023 190 930 190 965 190 1001 190 1042 190 1091 1027 .629 1036 .612 1048 .595 1060 195 978 195 1010 195 1044 195 1087 195 1143 1077 .634 1082 .615 1087 .595 1092 200 1025 200 1055 200 1087 200 1134 200 1194 1131 .641 1132 .621 1134 .600 1136 205 1073 205 1101 205 1132 205 1181 205 1247 1208 .658 1201 .636 1199 .613 1196 210 1123 210 1148 210 1178 210 1230 210 1301 765 .594 819 .604 876 .611 933 .628 972 .643 1024 .646 1100 .654 1158 .662 1214 .683 774 .604 832 .619 881 .636 932 .652 1002 .654 1061 .662 1118 .684 1174 .695 1231 .703 787 .626 835 .643 891 .654 962 .664 1021 .683 1080 .695 1136 .703 1214 .710 1323 .713 .650 16 .659 17 .672 18 .693 19 .704 20 .712 21 .715 22 .723 23 .741

ALTITUDE 1000 FEET

Mach Ind 33 Fuel - Lb/Hr Mach Ind 31 Fuel - Lb/Hr Mach Ind 29 Fuel - Lb/Hr KIAS 25 Fuel - Lb/Hr KIAS 20 Fuel - Lb/Hr KIAS 15 Fuel - Lb/Hr KIAS 10 Fuel - Lb/Hr KIAS 5 Fuel - Lb/Hr

Figure 8-19

8-104

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