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CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

CHAPTER 27
FLIGHT CONTROLS

CONTENTS
Page

INTRODUCTION......... 27-1

GENERAL .. 27-1

INDICATING AND WARNING SYSTEMS 27-4

AUTOPILOT INTERCONNECT 27-5

AILERON CONTROL SySTEM....... 27-7

Components Description and Operation 27-7

Aileron Operation........ 27-13

Aileron Trim Components.......................................................................................... 27-15

Aileron Trim Operation.............. 27-15

Autopilot Interface...................................................................................................... 27-15

RUDDER CONTROL SySTEM.......... 27-17

Components Descri ption and Operation.................................................................... 27-17

Rudder Operation.............. 27-19

Rudder Trim Operation 27-21

Rudder Bias Operation... 27-23

ELEVATOR CONTROL SYSTEM 27-27

Components Descri ption and Operation.................................................................... 27-27

Elevator Operation...................................................................................................... 27-29

CONTROL (GUST) LOCK SYSTEM 27-31

General....................................................................................................................... . 27-31
Control Lock Installation 27-31

FOR TRAINING PURPOSES ONLY 27-i

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CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

HORIZONTAL STABILIZER TRIM SYSTEM 27-33

General 27-33

Horizontal Stabilizer Trim System Components 27-35

Horizontal Stabilizer Trim System Power Sources 27-40

Trim System Monitors................................................................................................ 27-41

Primary Horizontal Stabilizer Trim System Operation 27-43

Secondary Horizontal Stabilizer Trim System Operation 27-44

Autopilot Trim System Operation.............................................................................. 27-45

FLAP SySTEM................................................................................................................. 27-47

General 27-47

Components Description and Operation.................................................................... 27-49

System Operation....................................................................................................... 27-55

Indicating And Warning Functions............................................................................. 27-56

Flap Altitude Speedbrake Monitoring (FASM) System 27-57

SPEEDBRAKE SYSTEM 27-59

General 27-59

Components Description and Operation.................................................................... 27-61

Hydraulic Fluid Circulation 27-63

Hydraulic Temperature Monitoring System... 27-63

Operation......... 27-65

SPOILER SYSTEM 27-67

General 27-67

Components Description and Operation.................................................................... 27-67

Ground Spoiler Operation.......................................................................................... 27-67

Roll Spoiler Operation 27-69

27-ii FOR TRAINING PURPOSES ONLY

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CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

Nonservoactuator 27-71

Emergency Descent and Aerodynamic Braking after Touchdown 27-71

Spoiler Holddown System 27-73

ANGLE-OF-ATTACK AND STALL WARNING SYSTEM 27-79

Components Description and Operation.................................................................... 27-79

Operation.................................................................................................................... 27-81

MAINTENANCE CONSIDERATIONS 27-83

General Maintenance Practices.................................................................................. 27-83

Inspections.................................................................................................................. 27-86

Functional Checks........ 27-87

FAULTANALySIS............................................................................................................ 27-89

LIMITATIONS.. 27-89

FOR TRAINING PURPOSES ONLY

27-iii
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IllUSTRATIONS
Figure Title

Page
27-1 Flight Control Surfaces .. 27-2

27 -2 Control Wheel 27-6

27 -3 Aileron Crossover Quadrant 27-6

27-4 Aileron Control System and Aileron/Spoiler Disconnect Mechanism 27-8

27-5 Aileron Cutoff Valve and Force Link Assembly 27-10

27-6 Aileron Power Boost Electrical Schematic 27-12

27-7 Aileron Trim and Feel Mechanism 27-14

27-8 Aileron Autopilot Servo Installation 27-14

27 -9 Rudder Pedal Installation.................................................................................. 27-16

27-10 Rudder and Bellcrank.................. 27-16

27 -11 Rudder Control System Diagram...................................................................... 27-18

27-12 Aileron and Rudder Trim Controls 27-20

27-13 Rudder Trim System Diagram 27-20

27-14 Rudder Bias Control and Indication 27-22

27 -15 Rudder Bias Installation.................................................................................... 27-22

27 -16 Rudder Bias System.......................................................................................... 27-24

27 -17 Rudder Bias Unequal Thrust............................................................................. 27-24

27 -18 RUD BIAS Switch OFF.................................................................................... 27-25

27 -19 Control Colulun.... 27-26

27 -20 Foward Elevator Control................................................................................... 27-26

27-21 Aft Elevator Quadrant and Idler Arm Assembly 27-28

27-22 Flight Control Surfaces and Throttle Lock Installation 27-30

FOR TRAINING PURPOSES ONLY

27-v
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27-23 Horizontal Stabilizer Trim System 27-32

27-24 Horizontal Stabilizer Trim Actuator 27-34
27-25 Controls and Advisory Units 27-36
27-26 Pitch Trim Controls and Indicators 27-36
27-27 Horizontal Stabilizer Trim System Components 27-38
27-28 Horizontal Stabilizer Trim Electrical Schematic 27-42
27-29 Wing Flaps 27-46
27-30 Flap System Components 27-48
27-31 Flap Carriages. 27-50
27-32 Flap System Block Diagram 27-52
27-33 BITE Indicators... 27-52
27-34 Flap Controls and Indicators............................................................................ 27-54
27-35 Spoiler and Speed brake Panels 27-58
27-36 Speedbrake Control System.......... 27-60
27-37 Fluid Temperature Monitoring System 27-62
27-38 Speedbrake Control and Indicator 27-64
27-39 Speed brake Actuator Installation..... 27-64
27-40 Spoiler System.................................................................................................. 27-66
27-41 Spoiler Mixer Assembly............. 27-68
27-42 Ground Spoiler Nonservoactuator Flow Pattern 27-70
27-43 Spoiler Holddown Components......... 27-72
27-44 Spoi ler Normal Operation 27-74
27-45 Spoiler Operation In HOLD DOWN Position 27-76
27-46 Stick Shaker Installation 27-78
27-47 Angle-of-Attack/Stall Warning System Components 27-78

27-vi FOR TRAINING PURPOSES ONLY

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27-48 Angle-of-Attack/Stall Warning System 27-80

27-49 Roll Spoiler Cable Tension Graph (Typical) 27-82

27-50 Cable Inspection 27-82

27 -51 CESCARD Example 27-84

27 -52 Cable Pressure Seal 27-84

27-53 Aileron Control System Troubleshooting Chart 27-88

FOR TRAINING PURPOSES ONLY 27-vii

 Flight~~~ty
CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

CHAPTER 27
FLIGHT CONTROLS

INTRODUCTION
The primary flight controls of the Model 650 Series consist of ailerons, rudders, and el-
evators. The ailerons are hydraulically boosted but can be operated manually; the rud-
der and elevators are manually actuated with no powered assist. Rudder and aileron trim
is mechanical; pitch trim is accomplished electrically.

Secondary flight controls consist of electrically operated flaps and hydraulically actu-
ated spoilers. The spoilers can be used as speedbrakes to assist in lateral control, for rapid
descent, and to supplement braking after touchdown.

The elevators are also manually actuated with

GENERAL no power assist. Control column movement is
mechanically transmitted through quadrants,
The rudder is manually actuated (with no cables, and an idler assembly to torque tubes
power boost) through conventional rudder attached to the elevators. The elevators can also
pedals. Rudder pedal deflection is transmitted be positioned by the autopilot elevator servo.
through cables to a bellcrank attached to a Deflection limit stops are conventional bolts
torque tube at the base of the rudder. In addi- that stop the lower idler arm at the extremes
tion to manual operation, the rudder can also of its travel. The stop bolts in the vertical fin
be actuated by the autopilot rudder servo. are considered to be the primary stops.

FOR TRAINING PURPOSES ONLY 27-1

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ROLL
SPOILER
SPEEDBRAKES
Figure 27-1. Flight Control Surfaces

77-7 FOR TRAINING PURPOSFS ONI Y

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CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

Secondary stops of the same type are part of NOTES

the forward sector installation.

Rotation of either control wheel mechanically

controls the power boosted ailerons. Cables
provide a control input to the aileron crossover
quadrant. If hydraulic pressure is lost, the
ailerons revert to manual operation.

Roll control is a combination of powered

ailerons and roll spoilers. Control wheel ro-
tation also provides control inputs to roll
spoiler servoactuators for hydraulic actuation
of the roll spoilers. A spring cartridge in the
system provides load feel.

The wing flaps are powered by a reversible 28-

VDC motor controlled by an electronic con-
troller and a flap selector. The six Fowler flaps
are positioned by mechanical actuators through
flexible drivers resembling large speedometer
cables. Split flap protection is provided by an
asymmetry protection system that shuts down
the flaps upon command from the controller if
a split flap condition is detected.

Speedbrakes and spoilers comprise eight hy-

draulically actuated panels on the top surface
of the wing. The two center panels on each
wing can be deployed as speedbrakes. All
eight panels can be deployed simultaneously
for rapid descent or to improve braking effi-
ciency. Control of all panels is from controls
on the pedestal.

There are eleven vortex generators on the top

surface of the wing forward of the outboard
spoiler panels.

All flight controls, including primary,

secondary, and trim tabs, are shown on Figure
27-1.

FOR TRAINING PURPOSES ONLY

27-3
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INDICATING AND A HYD TEMP LOW annunciator monitors

wing spoiler hydraulic fluid temperature when
WARNING SYSTEMS the aircraft is in cold soak conditions. A BITE
indicator box in the tail cone displays which
Three indicators (instruments with movable of eight thermal switches triggered the HYD
pointers) provide control surface position in- TEMP LOW annunciator.
formation: a horizontal pitch trim indicator,
a flap position indicator, and a speedbrake
position indicator. NOTES
A total of 12 annunciator lights advise the
crew of out-of-the-ordinary control surface
conditions:

• PRI TRIM FAIL

• SEC TRIM FAULT
• NO TAKEOFF
• SPEED BRAKE

• SPOILERS UP

• SPOILER HOLD DOWN

• RUDDER BIAS
• AIL BOOST OFF
• GUST LOCK
• STALL WARN
• FLAPS INOP

• FLAP OVERSPEED

Three audio warnings interface with the flight

control system: no takeoff warning horn, land-
ing gear warning horn (FLAPS position), and
aileron boost disengage.

The angle-of-attack system indicator, stick

shaker, indexer, and annunciators also inter-
face with the flight controls.

An ENG SYNC annunciator illuminates if the

flap selector is moved out of the UP position
while engine sync is engaged.

27-4 FOR TRAINING PURPOSES ONI Y

 CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

AUTOPILOT NOTES
INTERCONNECT
The Spe;rry autopilot system provides auxil-
iary control through the selected mechanical
servos.

The aileron servo is located under the cabin

floor and is cable-connected directly to the
aileron crossover quadrant.

The elevator servo is mounted on the canted

bulkhead in the aft tail cone and is cable-con-
nected to the elevator primary cables.

The rudder servo is located on the same canted

bulkhead and is cable-connected to the rudder
quadrant at the base of the rudder torque tube.

FOR TRAINING PURPOSES ONLY 27-5

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~."...•....

CONTROL
WHEEL

Figure 27-2. Control Wheel

FLOORBOARD
STRUCTURE
COPILOT
SECTOR

FORCE LINK
PILOT
DISCONNECT
SECTOR

OUTER TORQUE
TUBE (UPPER)

SEPARATION -.

OUTER TORQUE
TUBE (LOWER)

~ OUTER
III:l!!!!' OUTPUT
I TORQUE STRUCTURE
SUPPORT
TUBE

Figure 27-3. Aileron Crossover Quadrant

27-6 FOR TRAINING PURPOSES ONI Y

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AILERON CONTROL At the bottom end of the inner torque tube, the
crossover arm is attached. The crossover arm
SYSTEM is the means by which the spoiler mixer box
is connected to the aileron system. A crossover
bungee assembly connects the crossover arm
COMPONENTS DESCRIPTION to the spoiler mixer box.
AND OPERATION
Control Wheels NOTES
The control wheels (Figure 27-1) provide
aileron control inputs through cables to the
crossover control quadrant. Each control wheel
is attached by a shaft to a cable drum, a cast-
ing with provisions for connecting cables at
separate points. The control wheels are not me-
chanically connected within the control
columns.

Crossover Quadrant Assembly

The quadrant assembly (Figure 27-2) is built
around inner and outer vertical torque tubes.
One quadrant is attached to the inner torque
tube; two quadrants are attached to the outer
torque tube. The outer torque tube is a two-
part unit in which the upper half (the outer
input torque tube) mounts the pilot quadrant,
which receives the cables from the pilot con-
trol wheel. The upper half of the outer torque
tube also mounts the outer input torque tube
arm to which the force link attaches.

The lower half of the outer tube mounts the

lower quadrant from which the actuating cables
run outboard to the aileron quadrants. The lower
tube also mounts the outer output torque arm to
which the hydraulic actuator attaches. Thus,
any force applied to the pilot and co pilot con-
trol wheels must pass through the combined
force link and actuator assemblies before being
transmitted to the lower quadrant.

The inner torque tube mounts the co pilot

quadrant that receives the cables from the co
pilot control wheel. This quadrant has provi-
sion for two sets of cables-one set from the
co pilot control wheel and one set from the
aileron autopilot servo.

FOR TRAINING PURPOSES ONLY

27-7
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CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

RIGHT
AILERON
.'

RIGHT
AILERON
QUADRANT

COPILOT ,..

CO~~~ ~ ' :: .:. -" •,:' •••">'.:.. <~.::.~..........~;:':'::,,';:,'.';:::'::

''''''''''':::''':''''''''''
'
.......... \~~\~....... AILERON •......•.•.•......... .....
,.<.., ~ ..~.....' PILOT CROSSOVER ......•.•• ..••....•.• . •. ;;.
" ',.:..... CABLES QUADRANT .•••..••••
: .... ,:;;::.. PILOT
.... ,..:.... CONTROL
LEFT AILERON
WHEEL
QUADRANT

T.HANDLE ~,..j V
~~~

~. /~~
/y~ ...
/"Jr~
~
~ V
ij
/,---
/</ / / . ~\~ LEVER
L ",,?.::--:;.\\
I ASSEMBLY

<.: ~I PEDESTAL /
~ \)
~ ~ ~:. E:~~I .:~~~\

V""
~ / DISCO~NEf~:3

COLLAR

Figure 27-4. Aileron Control System and Aileron/Spoiler Disconnect Mechanism

27-8 FOR TRAINING PURPOSES ONLY

 CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

Aileron hydraulic power boost is applied to the Since there are two separate roll control sys-
ailerons through the crossover quadrant by a tems, certification regulations require that a
hydraulic actuator connected to the lower out- means be provided to separate them. In this
put torque tube. The actuator is controlled by manner, a jam in one system will not disable
a force link assembly connected to the upper the other system.
input torque tube. Movement of the control
wheel is boosted through the crossover quad- A red AILERON SPOILER DISCONNECT
rant assembly for movement of the ailerons. T-handle is provided to disconnect the two
systems. The handle is located on the pedestal
An autopilot servo and motor located on the and is covered with a clear plastic guard.
aircraft structure is cable-connected to a quad-
rant on the crossover quadrant assembly. When If a jam is encountered in lateral control,
the autopilot quadrant is moved by the servoac- pulling the T-handle separates the quadrants
tuator, the quadrant assembly moves, transmit- which are normally mechanically intercon-
ting the movement to the ailerons. nected. The pilot wheel then controls the
ailerons, which still have load feel and can be
The aileron boost actuator and force link at- either hydraulically powered or manually ac-
tachment to the crossover quadrant are shown tuated. The copilot wheel controls only the roll
in the following illustration. spoilers, which do not have any load feel.
Once the handle is pulled, the quadrants can-
not be reconnected in flight. They must be
Aileron/Spoiler Disconnect reset by maintenance personnel.
Mechanism
With the crossover quadrant assembly as de- WARNING I
scribed above, the pilot control wheel controls
the ailerons, both manually and hydraulically, This system is only to be used in the
and the copilot control wheel controls only the event of a jammed roll control con-
spoiler mixer box, since the autopilot cables dition. It is not used for any other
receive their input from the autopilot servo. purpose.
The spoiler mixer box will be discussed later
in this chapter.

An interlock of the two torque tubes has been

built into the crossover assembly. The inter-
lock, known as the aileron spoiler intercon-
nect/disconnect, is installed at the top of the
crossover assembly. It consists of a mechan-
ical interconnect in the form of two .375-inch
diameter dowel pins. The dowel pins extend
through the pilot and co pilot quadrants, caus-
ing them to move as one unit.

Provision has been made for disconnecting

the interlock in an emergency. Mechanical
components of the aileron/spoiler disconnect
mechanism are shown in Figure 27-8.

FOR TRAINING PURPOSES ONLY 27-9

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@@

/
FORCE LINK
ASSEMBLY

..... ' .
............•..
."
..
'

TO

AILERON CUTOFF VALVE

Figure 27-5. Aileron Cutoff Valve and Force Link Assembly

27-10 FOR TRAINING PURPOSES ONLY

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CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

Aileron Boost Actuator Aileron Quadrant

The piston end of the hydraulic boost actuator The aileron quadrants are located in the dry
is attached to the aircraft structure; the fixed end bay area of each wing and are cable-connected
is connected to the crossover quadrant outer to the crossover quadrant. The aileron quad-
output torque tube arm. The actuator incorpo- rant transmits motion via a pushrod to the
rates a bypass valve, servo control valve actu- aileron. Aileron travel is limited by stop bolts.
ated by the input lever, and a check valve.

Aileron Deflection
Force Link The required aileron deflection is:
The force link assembly (Figure 27-5) is a
• Up-12.5°
self-contained unit serving as a link between
the crossover quadrant and the boost actuator
• Down-12.5°
input lever. It incorporates an internal com-
pression spring sandwiched between two
slides, a cam, cam follower, and override mi- NOTES
croswitch. The force link also electrically
closes a hydraulic cutoff valve to stop boost
function when excessive force is applied. The
fixed end of the unit attaches to the outer input
torque tube arm of the crossover quadrant;
the clevis rod fixed end connects to the boost
actuator input lever.

Aileron Cutoff Valve

The solenoid-operated hydraulic cutoff valve
(Figure 27-5) is located under the cabin floor
near the crossover quadrant.

The three-way, two-position valve is normally

closed. With the solenoid energized, fluid
pressure is routed through a filter to the aileron
boost actuator. Return fluid from the boost ac-
tuator is routed back through the cutoff valve.
With the solenoid deenergized, pressure is
cut off and actuator fluid is routed to system
return through the valve. This prevents fluid
lock in the actuator.

The valve is deenergized closed by the over-

ride switch on the force link, cutting off pres-
sure to the aileron boost actuator. An additional
out of synchronization switch, located on the
crossover quadrant outer output torque tube
arm, will also deenergize the cutoff valve if
the input and output torque tubes do not move
together due to failure of the force link/actu-
ator hardware.

FOR TRAINING PURPOSES ONLY 27-11

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CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

AILERON
28 VDC
AURAL POWER
ALERT HORN
WARN 2

•t
PF15 PF201
AIL S225
BOOST B ABOVE
HYDRAULIC
OFF 500
PRESSURE
PSI
II A SWITCH

PF48
NORMALLY
K30
C
POWERED
AILERON
POWER
D

CLOSED

POWERED
AILERON
SHUTOFF
VALVE II

S26 K31
AILERON ON POWERED
BOOST OFF AILERON
POWER RESET RESET
SWITCH

II

S218
QUADRANT
t
OUT OF SYNC

•
DISCONNECT
K141
POSITION POWERED
MONITOR AILERON
SWITCH
~ LATCH

S27
FORCE LINK
DISCONNECT
AILERON SWITCH
BOOST

28 VDC

Figure 27-6. Aileron Power Boost Electrical Schematic

27-12 FOR TRAINING PURPOSES ONLY

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CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

AILERON OPERATION (tension or compression) is applied to the

force link, the internal slide-spring of the
Hydraulic pressure for the powered aileron force link compresses, allowing the cam to
boost is from the hydraulic system. The boost move away from the cam follower. The over-
system is controlled with the three-position ride switch deactivates, and electrical power
AIL BOOST switch with positions decaled to the cutoff valve through the aileron boost
RESET, OFF, and ON. power relay is removed.

The powered aileron system is actuated by The AIL BOOST OFF annunciator illumi-
placing the AIL BOOST switch to RESET, nates, and the aileron control system reverts
then to ON. Electrical power will be directed to manual operation. The fail-safe function of
to the hydraulic cutoff valve through the the aileron system enables the flight crew to
aileron boost power relay. manually operate the ailerons through the
same system without the assistance of the
The relay seeks 28 VDC through the aileron power boost.
boost override switch mounted on the force
link assembly and the out of synchronization A quadrant position monitor (out of synchro-
switch mounted on the crossover quadrant nization) switch, mounted between the upper
outer torque tube. With the electrical circuit and lower sections of the outer shaft, also re-
complete, hydraulic power will be applied to moves DC power from the shutoff valve in
the actuator. Rotation of either control wheel the event that the two sections of the shaft are
causes movement of the actuator input lever not in synchronization. This activates only if
through the force link and the outer input the force link is mechanically disconnected
torque tube of the crossover quadrant. The from the actuator or if the linkage and the ac-
actuator piston rod end extends/retracts to ro- tuator movement did not follow the linkage.
tate the outer output torque tube and lower If the force link is disconnected upon initial
quadrant of the crossover quadrant assembly. power application, the valve cannot be pow-
The lower quadrant is cable-connected to the ered open because the force link must trip to
left and right aileron quadrants for movement enable the circuit to be completed. An elec-
of the ailerons. trical schematic of the aileron boost system is
shown in Figure 27-6.
If the AIL BOOST switch is inadvertently
placed to OFF with the powered aileron sys-
tem activated, the hydraulic cutoff valve will
close, blocking pressure to the boost actuator.
The AIL BOOST OFF annunciator will illu-
minate and the warning horn will sound.
Provided DC power has not been removed
from the latching relay, powered aileron can
be restored by placing the AIL BOOST switch
to ON. If DC power has been interrupted, the
force link must be tripped to reset the latch-
ing relay (move the control wheel to the stop,
then release). The AIL BOOST switch must
then be placed to RESET, then to ON to open
the cutoff valve and restore pressure. The
force link assembly is designed to sense ex-
cessive force being applied by the control
wheel(s) due to a possible actuator malfunc-
tion. When an axial load of 15-30 pounds

FOR TRAINING PURPOSES ONLY 27-13

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CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

PilOT QUADRANT

TRIM ACTUATOR (JACKSCREW)

Figure 27-7. Aileron Trim and Feel Mechanism

\ ..•. \ ..

\ \.
'; ",\
\ "~7
UPPER QUADRANT
OF
CROSSOVER
QUADRANT

Figure 27-8. Aileron Autopilot Servo Installation

27-14 FOR TRAINING PURPOSES ONLY

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CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

AILERON TRIM COMPONENTS AUTOPILOT INTERFACE

The aileron trim components include the me- The ailerons can also be positioned by the au-
chanical trim actuator, a spring-loaded feel topilot aileron servo (Figure 27-8). The servo
cartridge, an idler link, and connecting link- incorporates an electrical clutch that engages
age (Figure 27-7). Roll trim inputs from the only when the autopilot roll axis is engaged.
cockpit are chain and cable transmitted to the It also incorporates a mechanical slip clutch
jackscrew trim actuator and to the pilot quad- set to a predetermined value. In the event of
rant of the crossover quadrant. a malfunction, the servo can be manually over-
ridden through the slip clutch or disconnected
by disengaging the autopilot.
AILERON TRIM OPERATION
The aileron trim system incorporates aileron
feel and centering functions. Trim is achieved NOTES
by rotating the aileron trim wheel on the aft
end of the pedestal.

The trim function of the trim feel system is to-

tally mechanical, using chains and cables be-
tween the trim control mechanism in the center
console and the trim actuator in the aft cabin.
A chain and sprocket mechanism in the cen-
ter console converts to cables beneath the
cockpit floor. From there, it is routed aft to the
trim actuator where the mechanism reverts to
chain and sprocket.

Extending or retracting the jackscrew of the

actuator results in movement of the pilot quad-
rant for aileron trim. This is accomplished
without compressing the cartridge since the
quadrant is free to move.

Conversely, if a control wheel load from the

cockpit is exerted on the quadrant, the car-
tridge will compress, producing feel. There is
no trim tab on either aileron.

There are no mechanical stops in the aileron

trim system; the stop function is built into the
trim wheel.

FOR TRAINING PURPOSES ONLY 27-15

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CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

TO PILOT RIGHT
PEDAL AND
COPILOT PEDALS

4
PEDAL ADJUST
LEVER

PILOT LEFT PEDAL SHOWN

Figure 27-9. Rudder Pedal Installation

.....•.:... ....
..:.... ;.: ..
/~:::
~'. ':~.:..
"

\\\ . . ;"" ......•....~.:::.,.:_~ ) .

.......: .

:....._,...~..-=:::;'t!:;;"'"
,,' :<C...

AUTOPILOT
CABLE

Figure 27-10. Rudder and Bellcrank

27-16 FOR TRAINING PURPOSES ONLY

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CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

RUDDER CONTROL A solenoid shutoff valve, mounted on the ac-

tuator, controls engine bleed-air flow between
SYSTEM the two chambers of the actuator. Bleed-air
lines from the left and fight engine bleed-air
system connect to the actuator.
COMPONENTS DESCRIPTION
AND OPERATION A thermostatically controlled heater blanket is
secured to the actuator and shutoff valve to pre-
Rudder Pedals vent freezing of moisture inside the actuator.
Both sets of rudder pedals are interconnected
with torque tubes (Figure 27-9). Lever arms NOTES
on the torque tubes attach to and actuate sys-
tem cables. A pedal adjust lever on each pedal
provides for adjustment to three positions.

Bellcrank
The rudder bellcrank is attached to the base of
the rudder torque tube (Figure 27 -10). Cables
from the pedals ride in the lower groove of the
bellcrank, while autopilot servo cables ride in
the upper groove.

Rudder Trim Actuator

The mechanical rudder trim actuator incorpo-
rates two jackscrews actuated by chain-driven
sprockets. The primary sprocket is driven by
a chain and cable arrangement deriving mo-
tion from a trim wheel in the cockpit. The pri-
mary sprocket is chain connected to the
secondary sprocket; thus, both sprockets si-
multaneously drive the jackscrews. One of
the connecting pushrods is adjustable; the
other is not.

The rudder trim tab is a small airfoil hinged

to the trailing edge of the rudder with piano
wire hinges. The tab is moved by the trim ac-
tuator or by servo linkage as the rudder moves.

Rudder Bias Components

The pneumatically actuated rudder bias actu-
ator is located in the tail cone at FS 571.00.
The piston rod end is attached to a torque arm
on the rudder torque tube, the butt end of the
actuator to the aircraft structure.

FOR TRAINING PURPOSES ONLY 27-17

 FlightSafety
International

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

LEVER
ARM

Figure 27-11. Rudder Control System Diagram

FOR TRAINING PURPOSES ONLY

27-18
 FlightSafety
InIematlonaJ

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

RUDDER OPERATION NOTES

The rudder control system is a manually ac-
tuated cable system with no power boost. Foot
pressure applied to the rudder pedals is trans-
mitted via cables to the rudder bell crank. A
simplified pictorial diagram of the rudder sys-
tem is shown in Figure 27-11.

The left and right lever arms, welded to the

rudder pedal torque tubes interconnect where
the primary cables and the loop cable join at
the forward pressure bulkhead.

From the lever arms, the rudder control cables

are routed under the cockpit floor and aft
along the left cabin floor through pressure
seals in the aft pressure bulkhead.

In the tai I cone, the cables run beneath the bag-

gage compartment, through the canted bulk-
heads, to the rudder torque tube bellcrank.

The rudder bellcrank is installed at the base

of the rudder torque tube and receives both the
primary and autopilot cables.

The rudder is of conventional construction

with a Kevlar skin and is the only control sur-
face with a trim tab. The ribs and leading edge
are made of aluminum alloy; the side skins are
made from Kevlar.

There are no rig pin holes in the rudder system.

Rigging of the rudder should follow the step-
by-step instructions in the maintenance manual.

Rudder travels are:

• Left-25° + 1/_0°

• Right-25°+I/-0°

The rudder can also be positioned by the au-

topilot rudder servo. The servo incorporates
an electrical clutch that engages only when the
autopilot yaw axis is engaged. It also incor-
porates a mechanical slip clutch set to a pre-
determined value. In the event of a
malfunction, the servo can be manually over-
ridden through the slip clutch or disconnected
by disengaging the autopilot.

FOR TRAINING PURPOSES ONLY 27-19

 F ightSafety
international

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

AILERON
TRIM
WHEEL

RUDDER
TRIM

-
WHEEL

Figure 27-12. Aileron and Rudder Trim Controls

TRIM WHEEL

SPROCKET
JACKSCREW
ACTUATOR
TURNBUCKLE

TURNBUCKLE

SEAL

Figure 27-13. Rudder Trim System Diagram

27-20 FOR TRAINING PURPOSES ONLY

 FlightSafety
IntematlanaI

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

RUDDER TRIM OPERATION NOTES

The rudder trim tab is hinged to the trailing
edge of the rudder and is actuated by a chain
and sprocket arrangement at the trim wheel and
at the trim actuator. Between these points,
movement is transmitted by cables.

Rudder trim is initiated with the rudder trim

wheel on the aft end of the pedestal (Figure
27-12).

Actuation of the trim wheel is instinctive; ro-

tate the wheel right to move the nose of the air-
craft right. Travel deflection stops are built into
the control uni t and are not adjustable. If cor-
rect tab travel cannot be obtained, check for
proper chain installation (that the chains are
centered on the sprockets) and the correct pre-
installation dimensions for the actuator
jackscrews.

The trim actuator is a dual jackscrew. Each

jackscrew has quadrathreads-four threads
on each stem with thread starts 90 apart. 0

With this configuration, it is possible to make

length adjustments of a quarter turn while
still maintaining the bearing rod end in a hor-
izontal plane (to mate with the clevis ends of
the tab pushrods). It is very important to have
each jackscrew extended exactly the same di-
mension out of the actuator body or binding
will result.

When removing the rudder, install a long 10

x 32 bolt through both jackscrews to prevent
loss of synchronization.

The rudder trim cable routing is shown; it is

a simple cable run with pulleys, fairleads, and
pressure seals (Figure 27-13).

FOR TRAINING PURPOSES ONLY 27-21

 FlightSafety
InlematlonaJ

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

Figure 27-14. Rudder Bias Control and Indication

SOLENOID
SHUTOFF
VALVE

,• .,#

,"
.,'""

~~~~l~ CA ~<.--'\C:)'f 0 ~o f
l~ ~\.JS\ua f l(2- /--0'(c. c> ~\ (( l\ (
({ r 'J
'WARNING o
RUDDER BIAS ARM IS INS'l"ALLED WITH FIVE .j
MS20470AD4 RIVETS. NO SUBSTITUTES OR OVERSIZE
RIVETS ARE AUTHORIZED.

Figure 27-15. Rudder Bias Installation

27-22 FOR TRAINING PURPOSES ONLY

 FlightSafety
InIematIonaI

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

RUDDER BIAS OPERATION NOTES

The rudder bias system (Figure 27-14) assists
in reducing the yaw forces resulting from en-
gine failure. Normal operating pressures at
the bias actuator are 50 psi at maximum power
settings. At high power settings, the actuator
applies up to 550 inch-pounds of torque to
the rudder torque tube. This translates to ap-
proximately 125-150 inch-pounds of force at
the rudder pedals to assist the pilot. When on,
the system is fully automatic in operation.

With the RUD BIAS switch in NORM, the so-

lenoid shutoff valve is deenergized and spring-
loaded closed. Rudder action is not affected.

If either engine loses power or fails, pressure

to one end of the actuator is reduced accord-
ingly, and the rudder is deflected against the
resultant asymmetrical thrust.

Loss of electrical power results in continued op-

eration of the system since the solenoid shut-
off valve is normally in the deenergized position.

Placing the RUD BIAS switch in OFF ener-

gizes the solenoid shutoff valve open, connect-
ing both sides of the actuator to disable the
system as indicated by illumination of the
amber RUDDER BIAS annunciator.

Figure 27 -15 shows the rudder bias installation.

FOR TRAINING PURPOSES ONLY 27-23

 F ightSafety
IntematIonaI

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

•
TEST

RUDDER
BIAS
SWITCH
OFF.

RUDDER
-
NORM
BIAS ACTUATOR

THRUST REVERSER
SWITCH (STOWED)

BOTH ENGINES OPERATING,

EQUAL THRUST

LEGEND
• LEFT ENGINE BLEED AIR

• RIGHT ENGINE BLEED AIR

Figure 27-16. Rudder Bias System

•
TEST

OFF.

NORM
-
•
-
THRUST REVERSER
SWITCH (STOWED)

LEGEND
• LEFT ENGINE BLEED AIR

o RIGHT ENGINE BLEED AIR (ZERO)

RUDDER DEFLECTED LEFT

Figure 27-17. Rudder Bias Unequal Thrust

27-24 FOR TRAINING PURPOSES ONLY

 FlightSafety
international

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

Raising either or both thrust reverser levers to NOTES

the deploy position also energizes the bypass
valve open and illuminates the RUDDER BIAS
annunciator. Illumination of the RUDDER
BIAS annunciator should be verified prior to
increasing reverse thrust (Figure 27-16,27-17,
and27-18).

The actuator is anti-iced by an electrical heat

blanket powered through the RH PITOT/
STATIC switch on the center panel. The power
is cycled on and off by a thermostat on the
blanket.

Placing the RUD BIAS switch in TEST applies J" '0fj ~~ ~ ?fU(JOC<
power to the actuator heating element, by-
e'
--~
passing the thermostat so that proper opera- J 0'(1 J
tion can be verified by a minimum reading of
7 amps on the RUDDER BIAS HTR ammeter.

Uncommanded rudder movement on the

ground with both engines operating constitutes
a system malfunction and must be corrected
prior to flight.

TEST
•
OFF

•
-
NORM

•
-
THRUST REVERSER
SWITCH (STOWED)

LEGEND
• LEFT ENGINE BLEED AIR

• RIGHT ENGINE BLEED AIR

Figure 27-18. RUD BIAS Switch OFF

FOR TRAINING PURPOSES ONLY 27-25

 F ightSafety
InlemalIonaI

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

TO
COPILOT ELEVATOR
COLUMN PUSHROD

PIVOT PIN

Figure 27-19. Control Column

ITOAFT
ELEVATOR
QUADRANT

SECONDARY STOPS
(NONADJUSTABLE)

Figure 27-20. Foward Elevator Control

27-26 FOR TRAINING PURPOSES ONLY

 FlightSafety
international

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

ELEVATOR CONTROL NOTES

SYSTEM
COMPONENTS DESCRIPTION
AND OPERATION
Control Columns
The control columns are attached to a two-
piece torque tube under the cockpit floor
(Figure 27-19). An arm attached to the torque
tube is pushrod-connected to the forward
quadrant.

Forward Elevator Quadrant

The forward quadrant at FS 130.00 is pushrod-
connected to the control column torque tube.
Control cables attached to the quadrant extend
rearward to the aft quadrant (Figure 27-20).

Aft Elevator Quadrant

The aft quadrant and idler arm assembly lo-
cated within the vertical fin receive cable move-
ment from the forward quadrant. The quadrant
is pushrod-connected to lower and upper idler
arms which, in turn, actuate pushrods for move-
ment of the elevators. Mechanical stops are
provided to limit elevator travel.

FOR TRAINING PURPOSES ONLY

27-27
 Flight~c[~ty
CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

.... ~;:-.................. ......

. ..
, ,I, ••.•.•.• ~~
" ,I,' ••••............
, ,,

. ,..
, "
,,'" ,','
,

.' ."
" .'-,
,
,. .-:..",
......
4........ .'
.'

..••.
Figure 27-21. Aft Elevator Quadrant and Idler Arm Assembly

27-28 FOR TRAINING PURPOSES ONLY

 FlightSafety
Intematlonal

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

ELEVATOR OPERATION NOTES

The elevator control system is a manually ac-

QJl i fL 5tot leA v- u/

tuated cable and mechanical linkage system
with no power boost. Fore and aft movement
of either control column is transmitted through
a pushrod to the forward quadrant. Cables
from the forward quadrant actuate the aft quad- uCf' I
V (
"-
rant and idler arm assemblies (Figure 27-2 I)
that move the elevator torque tubes for eleva- , I , ')

tor deflection. Surface travel is limited by stop G r,' ,..., v'L.

l.-
bolts contacting idler arm linkage.
e ,) v J
The elevator can also be positioned by the au- /I
topilot elevator servo. The servo incorporates
an electrical clutch that engages only when the
J I

,J
..-- J
autopilot pitch axis is engaged. It also incorpo-
rates a mechanical slip clutch set to a predeter-
mined value. In the event of a malfunction, the
servo can be manually overridden by disengag- L1 I

ing the autopilot.

Full travel for the elevators (from the stream-

lined position) is:
• Up-151J2 +0/- 112°
• Down-15 +1/_1°

FOR TRAINING PURPOSES ONLY 27-29

 F ightSafety
Inlematlonal

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

ELEVATOR
CABLES

LOCK TRAY

RECEPTACLE

AILERON CABLE

SWAGE BALL

SWAGE BALL

TRAY RUDDER CABLES

PULL
CABLE

Figure 27-22. Flight Control Surfaces and Throttle Lock Installation

27-30 FOR TRAINING PURPOSES ONLY

 FlightSafety
Internallonal

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

CONTROL (GUST) When the CONTROL

pulled to engage the system,
LOCK T-handle
rotating
is
the
LOCK SYSTEM handle 900 clockwise locks it in position.
Switches are actuated by lock tray move-
ment to illuminate the amber GUST LOCK
GENERAL and NO TAKEOFF annunciators. To disen-
gage the system, the handle is rotated coun-
The control lock system locks the primary
terclockwise 90 and pushed in, allowing
0

control cables in the neutral position and

springs to return the lock trays and throttle
blocks the power levers from being advanced
bellcranks to the released position. With the
beyond the IDLE Position. Citation VII con-
control locks engaged, it is possible to bring
trollocks permit the engine to be started with
the throttle levers out of cutoff and up to the
the control locks engaged in order to allow
IDLE position, to permit taxiing with con-
taxiing with the control locks engaged, a pre-
trol locks engaged.
ferred procedure in gusty weather. It is not pos-
sible to apply power beyond idle rpm with
the locks engaged. A GUST LOCK annunci-
ator is illuminated when the control locks are
NOTES
engaged. The control locks interface with the
no-takeoff warning system.

CONTROL lOCK
INSTAllATION
The control and throttle lock installation
(Figure 27-22) consists of left and right lock
trays, swaged balls on the primary control ca-
bles, and throttle lock bellcranks actuated by
a cable connected to a handle under the pilot
instrument panel.

To engage the flight control and throttle lock

system, both throttles must be positioned to
IDLE and the flight controls neutralized.
Pulling the CONTROL LOCK T-handle to its
limit of travel pivots a bell crank beneath the
cabin floor.

Cables connected to the bellcrank pull the

lock trays forward until the receptacles engage
the swaged balls on the control cables, immo-
bilizing the primary flight controls. At the
same time, throttle lock bell cranks are posi-
tioned to move a plate into position against the
striker on each throttle lever assembly in the
throttle quadrant. This prevents the throttles
from being advanced more than 1.25 inches
forward of IDLE.

FOR TRAINING PURPOSES ONLY

 FlightSafety International

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

"">:;:;?)
........ .../
SECONDARY /~~~ •.•. --- _
~
TRIM ACTUATOR ~ ."....pr - -
fI ~, •.•./, ,
r- ,_ "
I
I ,(I r
I

I
,,'

I ,I "
I , t) PRIMARY
..•. II 7
" J TRIM ACTUATOR
1/ " I , V ••.•
\ .....
, I J/ I
" I ..•. 1

,, ".>- -- ~- ~ -' -
\ , , -A /........ 1
\ '" ,.'/ ". /
\\ .... \)' /.? ~ // "'.... ~ /
\ ~;:~,
-'~/
/.... .",
)"'.....:-.... /
/
ACTUATOR
\ ' " /. j ....
/ ....
~, CONTROL UNIT
-(~ / •••.
// /;.J.;
\ ,~ /~-\ / ..../ ~ ,.-/j
\ >;%~ 11\ \\ //// ~ •••• ',//,
TRIM
\ \ , ~\..J'y/ ~ // :' /
, 'A ~-/ ;' .'" RELAY BOX
\\ \ ,// ./
COPILOT'S CONTROL \ \ " /'
•••.>.:"/
p,
\1\ \\
TRIM
ADVISORY
WHEEL TRIM ,,/ / / / \-) \\....J I •.•./ UNIT
SWITCHES \ / / / ,/ •.•.')---
(/ \) / /
/ -, / / MAIN POWER JUNCTION
/// \- \,) / / / ~.!....__ BOX RCCB MONITOR
// /, ,- ,) //............ - -"':::::--""':t... DISCONNECT RELAY
AUXILIARY / / \) / /.... /. ~-::... ----
.... \ \' // / .....
:...----7::--=:.~_~
RELAY ( \ / / / / - ~ -:..."'::z-=--=--,"">: __
n'ANEL I A \ / I /........ - - -:; <::"""_'-:::-::
V (.1. .--~

/><
r, , ) _ _
",,-,( II
',II .... t... •••.L.----------
-',11\ ' -----______ - -"

'X-:// .
,/~ I - - - - - _ _ _ _ /
I

I
,
.•.
\\ •.•. ~:...-
....-::./

I .•.••••.•••. SECONDARY TRIM

\ - .J •••. CONTROL MODULE
AND POSITION INDICATOR

PILOT'S CONTROL
WHEEL TRIM SWITCHES

DUAL REMOTE
AUDIO AMPLIFIER

Figure 27-23. Horizontal Stabilizer Trim System

27-32 FOR TRAINING PURPOSES ONLY

 Flight~j[~ty
CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

HORIZONTAL NOTES
STABILIZER TRIM
SYSTEM
GENERAL
The Citation 650 Series horizontal stabilizer
trim system (Figure 27-23) allows the flight
crew to alter the incidence of the horizontal
stabilizer to provide pitch trim control. The
stabilizer is positioned with an actuator which
is electrically driven by a primary or second-
ary trim system. Three modes of control are
available: primary, secondary, and autopilot
operation.

The primary system positions the stabilizer by

driving the actuator with an AC electric motor.
In the event of an electrical fault in the primary
system, the secondary system can drive the ac-
tuator through a separate DC electric motor.
To ensure redundancy and isolation, the pri-
mary and secondary systems are controlled by
completely separate electrical circuits. Also,
the primary system wiring is routed along the
left side of the aircraft while the secondary sys-
tem wiring is on the right side. When the au-
topilot is engaged, the primary system will be
controlled by the autopilot computer. The au-
topilot will automatically trim the stabilizer
to reduce the torque requirement necessary for
the elevator servo to position the elevator.

Pitch trim system monitoring is provided by

three annunciator lights: the PRI TRIM FAIL
and SEC TRIM FAULT lights on the main an-
nunciator panel inform the flight crew of a fault
in the respective system. The NO TAKEOFF
light on the main annunciator panel warns the
flight crew if the horizontal stabilizer is incor-
rectly positioned for takeoff.

A horizontal stabilizer trim position indica-

tor mounted in the center pedestal provides a
continuous indication of trim position. An in-
terface with the aircraft audio system pro-
duces a clacker output from the cockpit
speakers if the trim system remains in motion
for longer than approximately one second.

FOR TRAINING PURPOSES ONLY 27-33

 FlightSafety
Intematlonal

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

RAM
PRIMARY TRIM ACTUATOR

DC ELECTRIC BRAKE

AC MOTOR

P487

DC ELECTRIC CLUTCH SECONDARY

TRIM INPUT
SHAFT

SECONDARY TRIM ACTUATOR

INTERCONNECT
CHAIN

P497

Figure 27-24. Horizontal Stabilizer Trim Actuator

27-34 FOR TRAINING PURPOSES ONLY

 FlightSafety
IntemslIonaJ

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

HORIZONTAL STABILIZER Secondary Trim Actuator

TRIM SYSTEM COMPONENTS The secondary trim actuator (Figure 27-24) is
mounted just below the primary trim actuator
Horizontal Stabilizer Trim in the vertical stabilizer structure. The actua-
Actuators tors are interconnected with a chain. The actu-
ator consists of a single-speed 28-VDC
Primary Trim Actuator reversible electric motor, DC electric clutch
The primary trim actuator (Figure 27-24) is and brake, clutch monitor switch, and two re-
mounted within the vertical stabilizer and at- lays for motor/brake control.
taches to the forward portion of the horizontal
stabilizer. The actuator has a ram which is ex- Operation of the brake and clutch in the sec-
tended or retracted to move the stabilizer ondary trim actuator is identical to the primary
through a range of 1.50 nosedown to 12.50 trim actuator. When selecting secondary trim,
noseup. The actuator contains an AC electric power is removed from the clutch in the pri-
motor, DC electric clutch and brake, limit and mary trim actuator, disconnecting the AC
no-takeoff switches, position potentiometer, motor from the primary actuator geartrain.
DC tachometer, a mechanical input shaft, and At the same time, the clutch in the secondary
mechanical geartrain. actuator is electrically energized to connect
the DC motor to the primary actuator geartrain
The actuator motor is powered by two-phase through the external chain. The normally en-
26 VAC. Speed control is provided by vary- gaged brake in the secondary actuator locks
ing the frequency of the AC voltage applied the DC motor and energizes to release when
to the motor. A normally engaged DC brake trim movement is commanded by the second-
is mounted on one end of the motor shaft. ary system.
During trim movement, electrical power is
applied to release the brake and allow the
motor to rotate freely. After reaching the cor-
rect position, power is removed from the brake
and it reengages. The holding action of the
brake prevents the actuator from retracting
or extending as airloads are applied to the
horizontal stabilizer.

A normally disengaged DC clutch is installed

between the AC motor and the actuator geartrain.
The clutch is electrically energized to connect
the primary trim AC motor to the actuator
geartrain anytime electrical power is applied to
the aircraft.

Adjustable switches within the actuator are

used to control the limit of actuator retraction
and the operation of the NO TAKEOFF annun-
ciator. A potentiometer is used as the source for
stabilizer position information and a magnetic
pickup as the source of stabilizer rate of travel.

FOR TRAINING PURPOSES ONLY 27-35

 FlightSafety
InlematIonaI

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

........~
.....~

.:?T";:::'~:~"""'"
.!.f ACTUATOR
i.i CONTROL
.f.-") UNIT

'" \'
\
.....•:::.:.... TRIM
r\" ..... ADVISORY
UNIT

NOTE:
BOTH COMPONENTS ON TAil
CONE ELECTRICAL RACK

Figure 27-25. Controls and Advisory Units

Figure 27-26. Pitch Trim Controls and Indicators

27-36 FOR TRAINING PURPOSES ONLY

 FlightSafety Inlematlonal

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

Actuator Control Unit Remote-Controlled Circuit

The primary trim actuator control unit is Breaker and Monitor
mounted on the left equipment rack in the tail Disconnect Relay
cone (Figure 27-25). The control unit gener-
The remote-controlled circuit breaker
ates a two-phaseAC variable-frequency square
(RCCB) is located in the left halfofthe main
wave voltage to drive the primary trim actu-
power junction box. An RCCB combines the
ator motor whenever the proper command in-
functions of a circuit breaker and a relay in
puts are present. The AC voltage is also
one component. The trim monitor discon-
rectified and used to release the primary trim
nect relay is mounted in the left half of the
actuator brake. Additionally, the control unit
main power junction box. The monitor relay
provides the monitor functions for the pri-
is energized by the actuator control unit if
mary trim system.
the primary trim system is valid. The mon-
itor relay, when energized, completes the
control circui t for the RCCB.
Stabilizer Trim Relay Box
The stabilizer trim relay box is mounted ad-
jacent to the actuator control unit on the left Horizontal Stabilizer Trim
equipment rack. It houses the primary trim Relays
brake relay (K251). Arming the primary trim
system energizes the relay and electrically Relays to control the primary horizontal trim
connects the primary trim brake to the actua- system are housed in the main electrical junc-
tor control unit. tion box (Figure 27 -26). The relays and their
functions are listed below.
• K341 Trim valid relay
Pilot and Copilot Control
Wheel Switches • K233 Trim-up relay

Both the pilot and copilot control wheels are • K241 Trim-down relay
equipped with a split rocker switch for primary
trim control. One element of the split rocker • K249 Trim latch relay
is used to issue an up or down command while
the other issues a valid signal. Both elements • K239 Trim-up time-delay relay
of the split rocker switch must be operated to
• K237 Trim-down time-delay relay
achieve stabilizer movement.
• K235 Autopilot trim relay
Each control wheel is also equipped with a trim
disconnect switch labeled AP/TRIM/NWS • K34 Monitor disconnect relay
DISC. Pressing the disconnect switch removes
power from the split rocker switch and can- • K251 Actuator motor brake relay
cels the trim direction and valid commands.

Qk~ ~ 'r C ...-..,7 Q( \ ('\\C1 t \{) (fi}JO.

~6 TCc\J c.:M. r CO 1\ CJ n'lu h, tttet:r6 to ~

~ U I

~ 'J ~- t t!'" v7 f D~ q v
r (J
'x\\JJ~ G {c \1 q I:J
c)t:;?t ::.J (?r ~r

FOR TRAINING PURPOSES ONLY

 F ightSafety
International

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

HORIZONTAL
STABILIZER
P491
ADVISORY UNIT

SECTION A-A

DETAIL B

DETAIL A
HORIZONTAL STABILIZER
POSITION INDICATOR
EL PANEL

DETAIL C

Figure 27-27. Horizontal Stabilizer Trim System Components

27-38 FOR TRAINING PURPOSES ONLY

 FlightSafety
international

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

Secondary Trim Control Horizontal Stabilizer

Module Position Indicator
The secondary trim control module is mounted The horizontal stabilizer trim position indica-
in the center pedestal behind the power levers. tor is mounted on the center pedestal next to the
It consists of a guarded ON-OFF switch, split secondary trim control module. The indicator
rocker switch, and two relays. Positioning the provides a visual indication of the stabilizer
switch to ON removes power from the pri- position from +20 nosedown to -120 noseup
mary trim system and engages the clutch in the with a takeoff range from _40 to -100 noseup.
secondary trim actuator. Moving the split
rocker switch energizes the relays within the The position indicator is a 0- to 5- VDC volt-
control module to apply a trim direction and meter which responds to signals from the hor-
valid command to the secondary trim actua- izontal stabilizer trim advisory unit.
tor DC motor and brake.

The secondary trim system does not include Dual Remote Audio Amplifier
any switches to limit the actuator travel. The
The dual remote audio amplifier is located
DC motor drives the primary actuator until it
under the cockpit floor. The amplifier receives
reaches the internal mechanical stops.
the clacker signal from the horizontal stabi-
lizer advisory unit, amplifies it, and delivers
Horizontal Stabilizer Trim it to the pilot/copilot headphones or cockpit
speakers.
Advisory Unit
The horizontal stabilizer trim advisory unit is With the battery switch in the EMER position,
mounted alongside the actuator control unit the clacker signal will not be amplified and is
and stabilizer trim relay box on the left equip- audible only through the crew headphones.
ment rack in the tail cone (Figure 27-27). The
advisory unit supplies a reference voltage to
the position potentiometer within the primary
trim actuator. The returned voltage is
processed by the advisory unit and sent to the
position indicator.

The advisory unit includes two potentiometers

identified as OFFSET and GAIN. Adjusting the
potentiometers alters the signal sent to the po-
sition indicator to ensure that the display agrees
with the actual stabilizer position. The GAIN
potentiometer is used to correct the noseup
indication, and the OFFSET is used for the
nosedown.

The advisory unit also monitors the magnetic

pickup within the primary trim actuator. If
the advisory unit senses trim motion for longer
than one second, a clacker output generated
within the advisory unit is sent to the dual re-
mote audio amplifier.

FOR TRAINING PURPOSES ONLY 27-39

 CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

HORIZONTAL STABILIZER Secondary Trim System

TRIM SYSTEM POWER The secondary trim system is part of the DC
SOURCES emergency bus system. The circuit receives its
power from a 7 .5-amp circuit breaker labeled
Primary Trim System SEC PITCH on the pilot circuit-breaker panel.
The secondary trim system normally has power
The primary trim circuit is part of the DC hot
available when any of the following conditions
battery bus system. Power is available to ~he
exists:
primary system when any of the followIng
conditions exists: • Battery switch is in the BATT or EMER
position.
• Aircraft is in flight, regardless of the bat-
tery switch position.
• Battery switch is OFF if an engine-
driven starter-generator is on the line.
• Aircraft is on the ground, with the battery
switch in the BATT or EMER position.
Horizontal Stabilizer Trim
The primary trim system employs two sepa- Advisory System
rate circuits to supply 28 VDC to the actua-
The horizontal stabilizer trim advisory system
tor control unit. Both circuits have electrical
power available any time the aircraft batte~-
provides a visual indic~ti~n 0Tth~ stabili~er
position and an aural indIcatIOn If the trIm
ies are installed and connected. One source IS
system is in motion for more than one second.
used by the command/monitor circuitry while
The advisory circuit operates independent of
a separate source is provided for the power
whether the primary or secondary trim system
drive circuitry.
is in operation.
The command/monitor circuitry is powered by
The advisory circuit is supplied by a I-amp
the 2-amp PITCH CONTROL circui~ brea~er
circuit breaker labeled H TRIM ADVISE on
on the pilot circuit-breaker panel, in ~erI~s
the copilot circuit-breaker panel and is part of
with the 5-amp PITCH CONTROL cIrcuIt
the DC avionics emergency bus. Both the bat-
breaker in the main power junction box.
tery switch and the AVIONICS POWER switch
must be in the ON position for normal oper-
The power drive circuitry is supplied from a
ation of the visual and aural advisory circuit.
20-amp remote-controlled circuit break~r
The advisory circuit is energized with the bat-
(RCCB) installed in the left half of the main
tery switch in the EMER position, regardless
power junction box. A Yz-amp circuit breaker
of AVIONICS POWER switch position.
labeled PITCH POWER on the flight deck is
used to indicate if the circuit-breaker function
of the RCCB has tripped, removing power
from the trim system.

27-40 FOR TRAINING PURPOSES ONLY

 FlightSafety
Intemalk>nal

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

TRIM SYSTEM MONITORS NOTE

PRI TRIM FAIL Annunciator The SEC TRIM FAULT annunciator

The red PRI TRIM FAIL annunciator light is may illuminate when the secondary
controlled by the actuator control unit. trim switch is first placed in the ON
Illumination of the PRI TRIM FAIL annunci- position. Actuate the secondary trim
ator also triggers the MASTER WARNING split rocker switches, and verify that
lights. The actuator control unit provides a the light extinguishes.
ground to illuminate the PRI TRIM FAIL light
if one of the following faults is detected:
NO TAKEOFF Annunciator
• Loss of 28- VDC electrical power to the
actuator control unit The amber NO TAKEOFF annunciator is con-
trolled by a switch in the primary trim actua-
• Electrical power to disengage the brake tor. The switch provides a ground to illuminate
is not available with a valid trim com- the light if the trim is not in the takeoff range
mand. (-4 to -1 0°) while the aircraft is on the ground.
• Trim movement opposite of the flight
crew command
NOTES
• Trim movement without a flight crew
command
• Trim rate of travel too fast or slow
• Excessive voltage input to, or current
output from, the actuator control unit

SEC TRIM FAULT Annunciator

The amber SEC TRIM FAULT annunciator in-
dicates that the secondary trim actuator clutch
is improperly positioned with respect to the
position of the secondary trim ON-OFF switch:
• With the secondary trim switch posi-
tioned to OFF, the clutch should be dis-
engaged. Failure of the clutch to release
is indicated by illumination of the SEC
TRIM FAULT annunciator.
• With the secondary trim switch posi-
tioned to ON, the clutch should be en-
gaged. Failure of the clutch to properly
engage is indicated by illumination of
the SEC TRIM FAULT annunciator.

FOR TRAINING PURPOSES ONLY

27-41
 FlightSafety international

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

HORIZTRIM
ADVISORY UNIT
STABILIZER
P491 CB122

- -
POSITION INDICATOR
P489
28VDC IN L 28VDC
HORIZ TRIM ADVISE A POSITION HIGH
DC GROUND N ---1/ 1 r--- B POSITION LOW
POSITION DRIVE V ~
HIGH OUT STABILIZER
POSITION DRIVE S TRIM ACTUATOR
P488
LOW OUT
5VDC OUT G T POSITION
REFERENCE SIGNAL
POSITION COMMON J S POSITION SEND
LOW
POSITION H V POSITION SEND
SIGNAL IN HIGH
RATE SENSOR A D RATE SENSOR
HIGH IN SEND HIGH

----
RATE SENSOR B E RATE SENSOR
LOW IN SEND LOW

ACTUATOR
CONTROL UNIT
P486

- K SPEED
CONTROL 1 COMMON
y SPEED
CONTROL 1 HIGH
X RATE SENSOR HIGH

J RATE SENSOR LOW

ANNUNCIATOR
PANEL
R MON FAlLOUT
PRI
TRIM FAIL [~ -
---- A 28-VDC MON TEST

IND1
• OFF
28 VDC ---a--.:
S19
REMOTE AUDIO UNIT
TEST SWITCH P513

AUDIO OUT HIGH D 122 AUX A UNMUTE HIGH

AUDIO OUT LOW E 123 AUX A UNMUTE LOW

~

Figure 27-28. Horizontal Stabilizer Trim Electrical Schematic

27-4? FOR TRAINING PURPOSFS nNI v

 CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

PRIMARY HORIZONTAL NOTES

STABILIZER TRIM SYSTEM
OPERATION
The primary trim system is armed by depress-
ing both halves of either the pilot or copilot
split rocker switch. This action energizes the
K249 trim latch relay and supplies power to
the actuator control unit and primary trim ac-
tuator clutch. The actuator control unit ex-
tinguishes the PRI TRIM FAIL annunciator
and energizes the K34 monitor disconnect
relay if no faults are detected. The energized
monitor disconnect relay completes the con-
trol circuit for the RCCB to supply the actu-
ator control unit with motor power. Figure
27-28 is a detailed electrical schematic of the
horizontal stabilizer trim system.

Depressing both elements of the split rocker

switches on either the pilot or copilot control
wheel applies valid and up or down commands
to the actuator control unit which energizes the
primary trim brake to release and drive the AC
motor to extend or retract the actuator ram.
Limit switches in the actuator are in series with
the up/down command signal and open to limit
the travel of the actuator.

Rate of travel for the actuator ram is auto-

matically controlled and monitored by the ac-
tuator control unit. The rate of travel
progressively increases as the stabilizer is
moving from the nosedown to the noseup
position. Approximately 40 seconds is nec-
essary for the actuator to move from the
noseup electrical limit to the nosedown limit.
To prevent overheating, the primary trim ac-
tuator should not be operated more than 7
minutes in a 20-minute period.

The primary trim system can be disengaged by

depressing the pilot or copilot AP/TRIM/NWS
DISC switch on the control wheel, by arming
the secondary trim system, or by removing
power from the circuit.

FOR TRAINING PURPOSES ONLY 27-43

 FlightSafety
International

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

SECONDARY HORIZONTAL NOTES

STABILIZER TRIM SYSTEM
OPERATION
The secondary trim system is armed by placing
the guarded secondary trim ON-OFF switch
in the ON position. This disengages the pri-
mary trim actuator clutch, engages the second-
ary trim actuator clutch, removes power from
the control wheel trim switches, and disengages
the autopilot. The visual and aural position in-
dicating systems remain operational.

Moving the secondary trim split rocker switch

up or down energizes the respective up or down
relay within the secondary trim actuator. The en-
ergized relay applies power to the brake, caus-
ing it to release and control the direction in
which the secondary trim DC motor will rotate.
The secondary trim actuator drives the primary
trim actuator through the interconnect chain
and primary trim actuator mechanical input
shaft. The secondary actuator has no limit
switches and relies on a torque limiter and the
mechanical stops in the primary actuator to
stop movement of the stabilizer.

The secondary trim DC motor is a single-speed

type and drives the stabilizer at a constant rate
of travel. It requires approximately 74 seconds
to move the stabilizer from one mechanical
stop to the other. To prevent overheating, the sec-
ondary trim actuator should not be operated
more than 5 minutes in a I O-minute period.

FOR TRAINING PURPOSES ONLY

27-44
 FlightSafety
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CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

AUTOPILOT TRIM SYSTEM • Pressing the pilot or copilot TCS switch

on the control wheel
OPERATION
With the autopilot engaged, the primary trim • Pressing the GO AROUND switch on the
system is automatically controlled by the au- power lever
topilot computer. Autopilot trim is commanded
when a sustained elevator force is detected for
longer than 3 seconds. If this condition is NOTES
present, the autopilot computer delivers an
up or down command to the actuator control
unit, which trims the aircraft and reduces the
elevator servo force.

Before the autopilot can be engaged, the pri-

mary trim system must be armed by momen-
tarily commanding trim up or down with either
the pilot or copilot control wheel split rocker
switch. This energizes the trim latch relay
(K249) to supply power to the actuator con-
trol unit and autopilot computer. Pressing the
AP button on the autopilot controller then en-
ergizes the autopilot trim relay (K235) and
transfers control of the primary trim system
to the autopilot computer.

During autopilot trim, the stabilzier rate of

travel is approximately half the speed of nor-
mal primary trim operation, requiring approx-
imately 66 seconds for the stabilizer to move
from electrical stop to the other stop with au-
topilot trim. The computer also monitors the
position of the flap control handle. Moving the
flap handle in or out of the UP detent alters
the autopilot trim system delay to provide a
smooth pitch transition as the flaps are extend-
ing or retracting.

Autopilot trim control is disengaged by:

• Pressing the pilot or copilot AP/TRIM/
NWS DISC switch on the control wheel

• Arming the secondary trim by placing

the secondary trim ON-OFF switch in
the ON position

• Commanding primary trim movement

with the pilot or copilot split rocker
switch

FOR TRAINING PURPOSES ONLY 27-45

 FlightSafety
InIematlonal

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

Figure 27-29. Wing Flaps

FOR TRAINING PURPOSES ONLY

27-46
 FlightSafety
International

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

FLAP SYSTEM NOTES

GENERAL
There are a total of six Fowler-type flaps (three
per side) of composite construction embody-
i ng high strength and low weight (Figure 27-
29). The flaps move aft and downward within
flap islands which are, in effect, guide tracks
within which the flap rollers move.

The flaps are electrically controlled and driven

and are mechanically actuated.

The wing flap system provides lift to the air-

craft by the positioning of the six flap panels.
The mechanical positioning is achieved
through conversion ofrotary drive power into
linear force and motion. The contoured shape
of the flap tracks determines the angularity of
the flap panels to wing datum, and thus, their
effect on the aircraft in flight.

Flap carriages (which are assemblies to mount

the rollers that move within the flap tracks) are
attached to the inboard and outboard ends of
each flap panel. The carriages are connected into
the flap actuating mechanism by links (short
push-pull rods). All six flap panels are always
at the same position in relation to each other.

FOR TRAINING PURPOSES ONLY

27-47
 FlightSafety
Internatlonal

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

- '-.tP"
t
RIGHT FLEX~
DRIVESHAFT

<:»::>-...
...... .~:.:
.....
. ...~
. ~' ..::
..:: .

../..:.:fif.:...
... :.>. ....
....
-- ...~::~::::::~--.
•...•..•~.

~ BALLSCREW
NUT
ANGLE GEARBOX

~~
FLEX DRIVE
FROM PDU

NO.1
ACTUATOR
- . 'L!d (TYPICAL)

~~~

~ ~
~~~

Figure 27-30. Flap System Components

27-48 FOR TRAINING PURPOSES ONLY

 FlightSafety
international

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

COMPONENTS DESCRIPTION NOTES

AND OPERATION
Figure 27-30 depicts the various components
of the flap system.

Power Drive Unit

The power drive unit (POU) is installed at FS
355.00, BL 0.00, between the main gear wheel
wells. Access to the unit is gained by remov-
ing access plates 173AB and 174AB (refer to
Chapter 6 of the maintenance manual).

The drive unit is a 28- VOC bidirectional motor

capable of retracting or extending the flaps.
Left and right flex drives are engaged by square
drive fittings at each end of the drive unit.

Angle Gearboxes
The angle gearbox is a mechanical unit that
transmits rotary motion at a 90 angle. Input
0

to the unit is from the flex drive from the

POU. Output is through a square drive to the
flap actuator. The angle gearboxes are at the
inboard end of the left and right inboard flaps.

Flap Actuator
There are four flap actuators in each wing. The
actuators are mechanical units driven by flex
drives (except the No. I actuator which is
driven by the attached angle gearbox). As the
unit is actuated, a ballscrew nut is driven along
the threaded shaft. The flap, mechanically
connected to the ballscrew nut, is positioned
accordingly.

FOR TRAINING PURPOSES ONLY 27-49

 FlightSafety
Intematlonal

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

TYPICAL OUTBOARD END OF FLAP

CARRIAGE

TYPICAL INBOARD END OF FLAP

/
.'
• ENSURE THAT ALL SURFACES NOTED ARE
FREE OF PAINT

*
BETWEEN
SURFACES. INSIDE BOTH
ARMS

FLAP ACTUATING CARRIAGE COMPONENTS

*
Figure 27-31. Flap Carriages

27-50 FOR TRAINING PURPOSES ONLY

 FlightSafety
international

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

Flap Carriages NOTES

Flap carriages are located in each flap island
(Figure 27 -31). They carry the load of the flap
panels and transmit the travel of the flap ac-
tuator ballscrew nuts to the flap panels dur-
ing extension and retraction. Each flap carriage
has an inner and outer half. Bearings at the car-
riages travel in the flap island tracks to guide
and position the flap panels at the correct
aerodynamic angle for flight, takeoff, ap-
proach, and landing of the aircraft.

The inner and outer carriage halves sandwich

two Teflon-Dacron track bearings which travel
in the flap island tracks. A third bearing (at the
upper track) is positioned at the link(s) to the
actuator ballscrew nut; the link transfers mo-
tion from the actuator to the flaps via the
tracks. At the flap island support tracks, the
bearings roll and follow the tracks during flap
extension and retraction. In addition, each
outer carriage half and yoke incorporates an
elastomeric spherical bearing that permits
flexibility of the carriage-to-yoke center at-
tach bolt.

Correct assembly and alignment of the car-

riages is critical for freedom of movement
and satisfactory operation of the flap system.

FOR TRAINING PURPOSES ONLY 27-51

 F ightSafety
InlemalIonaI

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

TEST SWITCH SELECTED POSITION

INOP AND OVERHEAT LIGHTS

NO TAKEOFF
FLAP POSITION

ANGLE OF ATIACK FLAP RESET

POSITION

GEAR WARNING

MOTOR CONTROL

TO POWER
LEFT DRIVE
WING UNIT

FLEXIBLE
RIGHT WING SHOWN DRIVE

ANGLE
GEAR
NUMBER
1
NUMBER
2
NUMBER
3
l NUMBER
4
ASYMMETRY

BOX ACTUATOR ACTUATOR ACTUATOR ACTUATOR BRAKE

Figure 27-32. Flap System Block Diagram

1 2 3 4 5 6 7 8 9

• • • • • • • • •
1. FEEDBACK FAULT 6. NO COMMAND

2. MOTOR FAULT 7. • CONTROLLER

INTERNAL FAULT
3. COMMAND ERROR • REFERENCE OUT OF RANGE
• COMPARATOR FAULT
4. • POWER FAULT • SHUT DOWN FAULT
• POWER RELAY SHORT
• SOFT START/STOP MODULE 8. OVERSHOOT FAULT
OPEN
9. PREFLIGHT TEST
5. • GROUND FAULT FAIL
• HIGH-SPEED RELAY SHORT
• SOFT START/STOP MODULE
SHORT

Figure 27-33. BITE Indicators

FOR TRAINING PURPOSES ONLY

 FlightSafety
Inlemallonal

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

Controller CU 102 Soft-Start/Stop Module

The flap controller is located on the tail cone The soft-start (soft-stop) module, soft-starts
forward electrical rack at FS 470.00. Controller the drive motor, ramping it up to operating
logic interfaces with the flap position control speed.
signal (cockpit selector) and all other system
information monitored by the controller. If all As the flaps approach their selected position,
is well, the controller permits the POU to motor the soft-stop function slows the motor down
in the direction requested by the flap selector. to preven t an overshoot.
It also contains circuitry to protect the system
against asymmetry, loss of or improper com- A dynamic brake is applied at shutdown to pre-
mand and/or position signals, incorrect flap vent coasting of the drive motor.
power drive (POU) rotation, noncommanded
flap operation, and overcurrent of the power NOTES
control circuitry. Output signals to the no-
takeoff system, the dual angle-of-attack sys-
tems, the flap position indicator, and the
FLAPS INOP light are also directed through
the controller The flap system block diagram
is detailed in Figure 27-32.

The controller monitors flap position in mil-

livolt signal values. The controller's com-
parator circuitry continuously monitors the
di fferential between left and right sides of the
flap system. If the controller detects a differ-
ential that indicates an impending asymmetry
condition, it disables the system by removing
electrical power.

The face of the controller incorporates nine

BITE indicators, which tell the technician
why the FLAPS INOP light illuminated and
the flaps shut down.

The BITE Indicators are shown In Figure

27-33.

Flap Relay Box

The flap relay box, located under the baggage
compartment floor, contains the relays that
control the flap drive motor, as follows:
• Flap power relay
• Flap extend relay
• Flap retract relay
• High-speed relay
• Flap brake relay K272
• Soft-start/stop module CU 102

FOR TRAINING PURPOSES ONLY 27-53

 FlightSafety
international

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

Figure 27-34. Flap Controls and Indicators

~r-s
I tc,
~~Ic..', .: {
"\ J
\....( /
r
I

<1- No u'''" (

27-54 FOR TRAINING PURPOSES ONLY

 FlightSafety
International

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

SYSTEM OPERATION ings). The resolver "resolves": it reads and in-

terprets the field relationship of rotor to sta-
The flaps are controlled by the flap selector tor and transmits the electrical information
handle (Figure 27 -34) on the center console. signal to the flap controller. The preceding
The selector has four preselect positions: 0°, simplified explanation demonstrates that the
r, 20°, and FULL down (38 to 39°) (Figure resolver system is consistent in its feedback
27-34). signal quality.

When a flap posItIOn is selected, mi-

croswitches within the selector transmit that Resolvers
selected position information to the controller.
Resolvers are AC devices; therefore, a means
If the controller detects no faults or conflicts,
is required to change aircraft 28- VDC bus power
it commands the PDU to drive in the selected
to the AC power required by the resolvers.
direction. While the flaps are in motion, po-
sition resolvers at the outboard end of each sys-
The converters (one left/one right) are located
tem (left wing and right wing) feed back a
to the right side of the tailcone maintenance
progressive position information signal to the
door. Each converter receives 28 VDC and
controller. When the flap feedback signals to
converts it to the AC voltage required by the
the controller match the command signal input
resolver.
from the flap selector, the controller removes
power from the PDU (drive motor).
The converter receives the AC feedback from
the resolver, interprets and processes it, con-
The flap controller processes the position
verts it back to DC current, and transmits it
feedback signal from the left and right re-
to the flap controller.
solvers, averages them, and sends a linear DC
voltage signal to the flap position indicator.
Flap position information is utilized by the
Voltage values for the output signal to the in-
controller for the flap position indicator, the
dicator range from 2.0 volts with the flaps re-
split-flap protection system, the no-takeoff
tracted to approximately 9.4 volts with the
warning system, the landing gear warning
flaps fully extended.
horn system, the flap altitude speedbrake mon-
itoring system, and the angle-of-attack system.
Those same voltage values are also used for
inputs to the no-takeoff warning system, to C?e>'h(On~f},Lro ek l rd. 'CQ(!..JcN'1 d.. c
the landing gear warning horn system, to the ~1'CWr
flap overspeed warning system, and to the NOTE
angle-of-attack system, all of which monitor
flap position.
The converters are sealed compo-
nents; therefore, servicing is limited
to removal and replacement. It is an
Flap Position Sensing acceptable troubleshooting proce-
Resolver flap position transmitters are stan- dure to switch converters and see if
dard equipment on the Citation VII. the squawk follows the converter.

Resolvers were selected because of their ac-

curacy and consistency.

Resolvers are totally electrical in their func-

tion. While driven by the flap flexible drives
(mechanical) internally, the movable rotor
(winding) rotates adjacent to the stator (wind-

FOR TRAINING PURPOSES ONLY

27-55
 FlightSafety
international

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

INDICATING AND WARNING • Any error in ground, on-ground, or

FUNCTIONS ground-in-air logic as seen by PCB 5

• The rotary test switch in TRIM/FLAP

Flap Position Indicator position
The flap position indicator on the center con-
sole displays flap position in degrees.
NOTES
Flap Controller Fault Light
The microprocessor flap controller is self-
monitoring and has a fault condition annun-
ciator. The fault annunciator is located to the
left of the flap selector and is labeled FLAPS
INOP. During flight or normal operation, il-
lumination of the FLAPS INOP light indi-
cates that the flap controller has detected a
fault in the flap system or its own electronic
circuits and has shut down.

During the preflight cockpit checks, when the

rotary test switch is selected to the TRIM/FLAP
position, the light illuminates for approximately
seven seconds while the controller performs an
integrity check of the total system. If the light
fails to extinguish after seven seconds, the con-
troller has detected a fault.

The fault annunciator is also a reset button.

When the fault condition has been cleared,
pressing the annunciator resets the flap sys-
tem for operation.

Flap Overs peed Light

The flap overspeed annunciator lights are lo-
cated above the pilot airspeed indicator and
the copilot altimeter. Through an interface
with printed circuit boards 5 and 45, the flap
overpseed light illuminates under any of the
following conditions:
• Flaps extended 1.2 and airspeed
0
above
215 KIAS

• Loss of flap position signal

• Loss of airspeed signal

27-56 FOR TRAINING PURPOSES ONLY

 FlightSafety
international

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

FLAP ALTITUDE SPEEDBRAKE NOTES

MONITORING (FASM) SYSTEM
The FASM system is concerned with sink
rate and primarily sink rate during final L6~ I[0f~ 0J: (
approach.
'b' Ur ~~ .
When the flap selector is out of the 0° detent
and the speedbrakes are extended to any po- V
oJ
mcc~{ S(ICCfC"
sition, the no-takeoff warning horn sounds
continuously when the aircraft is within 500
feet of the ground (as determined by the radar
altimeter). The FASM monitors a configura-
tion that could be hazardous as the aircraft
approaches the runway threshold.

NO TAKEOFF
Annunciator Light
The NO TAKEOFF annunciator illuminates
when both of the following conditions are
present:
• Aircraft is on the wheels (squat switch
actuated).
• Control configuration on the aircraft is
unacceptable for takeoff.

The flap controller and the flap indicator in-

terface with the no-takeoff warning system.

Normal flap position for takeoff is 20°; flap po-

sition for takeoff from a high-altitude runway
is 7°. Therefore, a flap position of less than 7°
or greater than 20° is unacceptable for takeoff
and causes the NO TAKEOFF annunciator to
illuminate. The flap system is only one of many
inputs to the no-takeoff system.

FOR TRAINING PURPOSES ONLY

27-57
 FlightSafety
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CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

..--- RIGHT
AILERON

NUMBER 7 SPEEDBRAKE

NUMBER6SPEEDBRAKE

NUMBER 5
GROUND SPOILER

NUMBER 4
GROUND SPOILER

NUMBER 3 SPEEDBRAKE

NUMBER 1SPEEDBRAKE
NUMBER 1
ROLL SPOILER

Figure 27-35. Spoiler and Speedbrake Panels

FOR TRAINING PURPOSES ONLY

27-58
 FlightSafety
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CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

SPEEDBRAKE SYSTEM NOTES

GENERAL
The speed brake panels (2, 3,6, and 7) have a C1~ CtcL.J()( s~d
maximum open position of 47° +3, -0, but
variable deflection of any position from zero '010
to4r is attainable by selective positioning of
the control lever. They are controlled by a r ,,' Z-
system of cables connected to the control lever
to the left of the throttles. Spoiler and speed-
j Q-(\ _t(lfre C, t'r/, ,j,

brake panel locations are shown.

~~14wr ('ffl
The aircraft has a total of eight spoiler pan-
els on the upper surface of the wings. Some
of the panels have dual functions, depending
on the flight configuration of the aircraft at the
time they are deployed (Figure 27-35).

The functions are:

• Roll spoiler
• Speedbrake
• Emergency descent
• Aerodynamic braking after touchdown
• Spoiler speedbrake holddown

Hydraulic pressure for the spoilers is from

the main hydraulic power (3,000 psi) system.

FOR TRAINING PURPOSES ONLY 27-59

 Flight~jl!~ty
CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

FEEDBACK
ARM.LINKAGE
INPUT
LEVER

SPEEDBRAKE ACTUATOR (TYPICAL)

SPEEDBRAKE LEVER

THROTTLE QUADRANT

TORQUE SHAFT

SPEEDBRAKE
CROSSOVER
QUADRANT

NO.3 ACTUATOR

TO NO.6 AND NO.7

ACTUATORS

Figure 27-36. Speedbrake Control System

27-60 FOR TRAINING PURPOSES ONLY

 F ightSafety
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CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

COMPONENTS DESCRIPTION NOTES

AND OPERATION
Speedbrake Center Quadrant
The center quadrant is located beneath the aft
cabin center floorboards at FS 335.10. The
quadrant is actuated by a Controlex assembly
from the speed brake lever in the cockpit.
Conventional control cables from the quadrant
extend to and actuate the bellcranks located at
WS 98.50, left and right. Pushrods connect the
inboard bellcranks to the speed brake bellcranks.

Speed brake Bellcranks

The bellcranks installed at WS 132.04 and
WS 169.00 are similar. Each has an upper and
a lower bearing, lug-mounted clevis, and rig
pin provisions. Rotation of the bellcranks at
each location results in the rotation of the
input lever at each individual servoactuator,
and it is at this point that the mechanical
speed brake selection mechanism interfaces
with the hydraulic system.

Speedbrake Actuators
The speedbrake actuators are located and iden-
tified by the panels they actuate. They are ser-
voactuators, fully modulated positional
hydraulic units incorporating a mechanical
input lever, input control valve, mechanical
feedback, holddown valve, thermostat valve,
and a manual pressure release button (Figure
27-36).

FOR TRAINING PURPOSES ONLY 27-61

 FlightSafety
Inlemallonal

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

HYD TEMP LOW ANNUNCIATOR

'.~~
RESET
LH I2 3 4 5 6 7 8
rrt;~.:Jof.l~
,M ••• "

POWER MUST DE APPLIED FOR

• I.
RH

INDICATORS TO BE DISPLAYED

FLUID TEMPERATURE MONITOR UNIT

Figure 27-37. Fluid Temperature Monitoring System

27-62 FOR TRAINING PURPOSES ONLY

 F ightSafety
international

~ITATION 650 SERIES MAINTENANCE TRAINING MANUAL

HYDRAULIC FLUID Fluid Temperature Monitor Unit

CIRCULATION The monitor unit is mounted on the tail cone
electrical rack, left of the tail cone access door.
Because MIL-H-83282 B fluid becomes slug-
gish at very cold temperatures, provision is
Containing a logic circuitry on a printed cir-
made to circulate the fluid in the wing trail-
cuit board, it monitors fluid temperatures from
ing edge area.
sensors at each speedbrake/spoiler actuator in
the system. If abnormally low fluid tempera-
The speedbrake and roll spoiler actuators in-
tures are detected, the appropriate BITE (built-
corporate thermostat val ves which open to
in troubleshooting equipment) indicator on
circulate warm fluid from the main system
the monitor unit trips and a white ball ap-
into the actuators when the fluid temperature
pears. In addtion, the HYD TEMP LOW an-
in the actuators drops to -10°F. This prevents
nunciator illuminates. The indicator can be
fluid in the actuators from being cold soaked,
reset with the rest button and will reset the
ensuring that speed brake and/or ground spoil-
BITE to black.
ers do not deploy sluggishly, causing a non-
commanded roll.

The circulation at the actuators is dependent

NOTES
upon the actuators being properly rigged. The
mechanical input arm of the actuator must be
against its stop to allow ingress of warm fluid
when the panels are retracted.

NOTE
Refer to the Chapter/Section 27-60-
00 of the maintenance manual
(Spoiler/SpeedbrakeAdjustment and
Test) for specific rigging instructions.

In addition to the built-in thermal bleed at the

actuators, an orifice bleed plate is installed in the
rear spar at wing station 230.00. The orifice plate
contiuously bleeds fluid to return (to achieve
fluid movement) at the outboard end of the spoil-
er hydraulic tubing.

HYDRAULIC TEMPERATURE
MONITORING SYSTEM
The hydraulic temperature monitoring sys- ecc<rJ~ GlQ ilurn!nCt lCt lu~, ";.I'(1pl({trc-r ~
tem (Figure 27-37) monitors the fluid temper- CfJ<?tCr ~C5 '=>?e0d bfG I-ff0 u;J roco fX1rs.
ature at the speed brake and roll spoiler ~ue rE '1 U 'J i' (}J
I •
~ 5G
eoll Cl(l k dt2>rue ~ C«v1t. 0tr ~
actuators. It also monitors the temperature of
the fluid in the return line from the spoiler sys-
105 5 (J J
tem. If the temperature at anyone of the eight t <:'0 \() f1 Q
f 0 .)'0 I-oIJ 00vJ(l.
sensors drops to -30° :t 3°F, two HYD TEMP
LOW annunciator lights will be illuminated.
These annunciators are located near the top of
both the left and right instrument panels.

FOR TRAINING PURPOSES ONLY 27-63

 F ightSafety
international

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

SPOILER LEVER

Figure 27-38. Speedbrake Control and Indicator

LEFT NO.2 SPEEDBRAKE ACTUATOR

RIGHT NO. 7 SPEED BRAKE ACTUA'TOR

Figure 27-39. Speed brake Actuator Installation

27-64 FOR TRAINING PURPOSES ONLY

 FlightSafety
InlematIonaI

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

OPERATION NOTES
The speedbrakes consist ofthe two center spoiler
panels on each wing. They can be hydraulically
deployed to any position up to 50 by selective
0

positioning of the speedbrake lever.

As the speedbrake lever (Figure 27-38) is moved

aft, cables routed beneath the cabin floor and into
the wing actuate quadrants and bell cranks that
provide inputs to the speedbrake servoactuators.

All four speedbrakes are simultaneously po-

sitioned in proportion to lever movement.
Speedbrake position is shown in percent of ex-
tension on the SPEED BRAKES POSITION
indicator.

Inputs to the indicator are from a speedbrake

position transmitter located adjacent to, and ac-
tuated by, the left outboard speedbrake actuator
(Figure 27-39).

NOTE
A similar transmitter is installed at
the right outboard speed brake actu-
ator. The right transmitter sends its
information to the angle-of-attack
computer. Both transmitters direct
their signals to PCB 8.

On Canadian-certified aircraft, both signals are

fed into a spoiler asymmetry monitoring circuit.

Whenever the speed brakes are not fully

stowed, an amber SPEED BRAKE annuncia-
tor light illuminates. The annunciator is con-
trolled by a proximi ty swi tch and target
bracket. The proximity switch is mounted to
the wing structure; the target bracket is
mounted on the movable spoiler panel. If the
target bracket moves .050 to .070 inch from
the proximity switch, the SPEED BRAKE an-
nunciator illuminates.

FOR TRAINING PURPOSES ONLY 27-65

 FlightSafety
Inlemational

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

TO SPOILER MIXER
AND GROUND SPOILER
ACTUATORS

TO SPOILER
CONTROL
LEVER

SPOILER CONTROL VALVE

.-,-.-
~
GROUND SPOILER ACTUATOR
(NONSERVO)

Figure 27-40. Spoiler System

FOR TRAINING PURPOSES ONLY

27-66
 FlightSafety
irtematlonal

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

SPOILER SYSTEM GROUND SPOILER

OPERATION
GENERAL The inboard spoiler at panels 4 and 5 are uti-
lized as ground spoilers. Unlike the roll spoil-
The spoilers (Figure 27-40) are divided into two ers and speed brakes, the ground spoiler panels
distinct systems-roll spoilers and ground spoil- can be commanded to only two positions-full
ers. The roll spoilers at panels I and 8 (see Figure up (300) and stowed. Unlike the other com-
27-35), operate with the ailerons to increase roll mand signals, the signal to command full-up
rate. The ground spoilers at panels 4 and 5, and deployment is hydraulic. Main hydraulic sys-
the speedbrakes at 2, 3, 6, and 7, are used to tem pressure is directed to the servoactuator
improve aerodynamic braking after touchdown by a spoiler control valve that is mechani-
and for emergency descent. cally actuated by the spoiler lever on the speed-
brake handle.
COMPONENTS DESCRIPTION The ground spoilers are individually moni-
AND OPERATION tored by amber lights inside the speedbrake po-
sition indicator. These will illuminate when
Spoi ler Actuators the left and right panels are not fully stowed.
If either of these lights illuminate, an amber
The spoiler actuators are located and identi-
SPOILERS UP annunciator light illuminates
fied by the panels they actuate. The ground
as well.
spoiler actuators are nonservoactuators and are
either fully extended or fully retracted. The roll
spoiler actuators are servoactuators, fully
modulated positional hydraulic units.
NOTES

Servoactuators incorporate a mechanical input

lever, input control valve, mechanical feed-
<i?0knc~o ~ ~~d
back, holddown valve, thermostat valve, in d (C£<(!f()()
I
C(f-
~B 0n
and a manual pressure release button.
Nonservoactuators feature a control input Cst\..; nc. u...; ,,'
I 0 0 < -c owl' ..J 0
mechanism, selector valve, holddown valve,
and manual pressure release button.
es ~r~ tbrtcol cf-f {)aP.

Spoiler Control Valve

The spoiler control valve is mounted below the
pilot pedestal and is pushrod-connected to
the spoiler lever on the speedbrake control
lever. The valve is a two-position, three-way,
mechanically operated valve with pressure, re-
turn, and control ports. The valve input arm
compensates for the movement of the speed-
brake control lever with free travel of the input
arm; the spoiler lever moves the input arm to
its final actuation to open the valve.

FOR TRAINING PURPOSES ONLY

27-67
 FlightSafety International

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

'.....•...... ....
.. '
"
.
\ .'
'.
(.':: . :::....
.

FEEDBACK
ARM-LINKAGE ," ('f.' .'~
.~~
../" :~/"~ })l ::.!:::.... . ::..................................... .;:.r. ' .

,.::~:..:.....
• ' .:. ':"': ~'::'.':. ~.~ •••.• ; •• - •••••••••• h •••••

.
'~"
. ....
.' ::,.:;:::1:"
.......•:....

ROLL SPOILER
ACTUATOR SPOILER MIXER
ASSEMBLY

GROUND/EMERGENCY
DESCENT GUIDE ROD

ROLL SPOILER
CABLES

AILERON
CROSSOVER
QUADRANT
ASSEMBLY

Figure 27-41. Spoiler Mixer Assembly

27-6R FOR TRAINING PURPOSES ONLY

 FlightSafety
lnlemallonaJ

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

ROLL SPOILER OPERATION Secondarily, the roll spoilers interconnect hy-

draulically with the ground spoilers and will
The roll spoilers have dual functions. function as speed brakes (spoilers I and 8 both
Primarily, they interconnect with the aileron fully deployed) for emergency descent or land-
controls and assist in lateral control during ing roll aerodynamic braking. This function
turns. The roll spoilers (I or 8) will deflect up- will be discussed under Emergency Descent
ward in coordination with the upward deflec- and Aerodynamic Braking After Touchdown.
tion of the aileron adjacent to it. The degree
of spoiler deflection is proportional to the ex-
tent of aileron movement. The interconnection NOTES
of aileron and roll spoiler function is accom-
plished by the spoiler mixer assembly that is
mechanical in function (Figure 27-41).

The spoiler mixer assembly is located at FS

328.0 in the lower fuselage.

Located on the bottom of the box is the aileron

input arm which is bungee-connected to the
inner torque tube of the aileron crossover
quadrant assembly. The input arm rotates an
internal cam assembly for the roll function of
the mixer assembly. A hydraulic actuator is
mounted on top of the box and connects to and
actuates the mixer guide rod for the emer-
gency descent and landing function of the
mixer assembly. Two output quadrant assem-
blies under each end of the mixer assembly
provide for connecting the left and right roll
spoiler control cables.

The mixer box is designed to provide a 3.5 dead 0

band (no spoiler operation) until the aileron

has moved up 3.50• After the dead band is passed,
the spoiler will move up proportionally and
will be fully extended at full-up aileron.

Cables run from the mixer box to the bell-

crank for the outboard spoiler. The bel1crank
is connected to the hydraulic servoactuator;
movement applied to the mixer box will be pro-
portionally transferred to the hydraulic spoiler
actuating valve.

The roll spoiler actuators are mechanically

and hydraulically identical to the speed brake
actuators.

FOR TRAINING PURPOSES ONLY

27-69
 Flight~!fi!ty
CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

..... ,,~;:;..: : , .. . .::

'. ~ .;/'

....----~'i. •..
~.;'-_./.
...
/. ~ '.
.... //. .'-':':'~",,"_.
'
;;- .~...
.. ,

. . "
~
~.~.
", .- . '.. ,,::~
HOLDDOWN
ACCUMULATOR

SPOILER HOLDDOWNI
NO TAKEOFF
PRESSURE
SWITCH MANUAL
(1,200 PSI) BLEEDDOWN
VALVE

AUXILIARY
HYDRAULIC
PUMP DUMP
VALVE

Figure 27-43. Spoiler Holddown Components

27-7? FOR TRAINING PURPOSFS ONI V

 F ightSafety
InlemalIonal

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

SPOILER HOLDDOWN SYSTEM Spoiler Holddown, No-Takeoff,

Pressure Switch
General
The pressure switch in the holddown hydraulic
The spoiler holddown system (Figure 27-43) line illuminates the SPOILER HOLD DOWN
is designed to hold the spoilers retracted under annunciator and the NO TAKEOFF annunci-
pressure. The pressure can be from any of ator when subjected to holddown pressure.
three sources.

NOTES
Spoiler Holddown Components
The spoiler holddown valve is a three-way,
two-position, normally open valve. In the open
position, main hydraulic system pressure
passes through the valve for normal opera-
tion of all spoiler/speedbrake actuators. In
the closed position, main pressure is cut off
to all spoiler/speedbrake actuators and is di-
rected to the holddown ports. Return fluid
from both wings passes through the holddown
valve when it is open or closed. The valve re-
ceives 28- VDC power from the HYD
SPOILER circuit breaker through the
SPOILER holddown switch on the pilot
pedestal.

Spoiler Holddown Accumulator

The holddown accumulator has nitrogen
precharge and hydraulic fluid chambers sep-
arated by a floating piston. A nitrogen pres-
sure gage and service valve are located in the
accumulator panel in the tail cone baggage
compartment. The accumulator is hydrauli-
cally pressurized during main hydraulic sys-
tem operation and is the source of pressure for
spoiler/speedbrake holddown operation in the
event of main system pressure failure.

Spoiler Accumulator Manual

Bleeddown Valve
The manual bleeddown valve is a manually op-
erated, spring-loaded closed valve mounted in
the accumulator gage panel in the tail cone bag-
gage compartment. When held open, hydraulic
pressure in the spoiler holddown accumulator
is ported to the system return line.

FOR TRAINING PURPOSES ONLY

27-73
 F ightSafety
InlemaIIonaI

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

-- r-- ------------l
SPOILER HOLDDOWN MODULE

I
,-
I

I
I
I
L _

HOLDDOWN
ACCUMULATOR

PRESSURE
SWITCH
HOLDDOWN
PRESSURE
SWITCH I
I
I
I
I
I
I
I

----------------------~

LEGEND
• PRESSURE

• RETURN
MECHANICAL

ELECTRICAL

Figure 27-44. Spoiler Normal Operation

27-74 FOR TRAINING PURPOSES ONLY

 FlightSafety
international

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

Spoiler Holddown System

Operation
In normal operating conditions, the spoilers
are held in the stowed position by hydraulic
pressure commanded by the servo valves
(Figure 27-44).

However, if system pressure is lost, the spoil-

ers can float up, causing an increase in drag
with a resulting decrease in performance.
To guard against this, a holddown system is
incorporated.

It is important to understand that the only

conditions that cause spoiler holddown to ac-
tivate are loss of hydraulic pressure or the
pilot intentionally shutting off main system
pressure to the spoilers with a switch in the
cockpit.

The spoiler holddown module located under

the baggage compartment floor, contains two
valves: a solenoid valve controlled by a cock-
pit switch and a pressure-operated valve con-
trolled by hydraulic system pressure and a
spring. If hydraulic pressure is greater than
spring force, the valve is closed; if spring
force is greater than system pressure, the valve
is pushed open.

Main system pressure enters the spoiler mod-

ule and passes through the spoiler holddown
valve. The pressure normally takes two paths
as shown on the facing page.

Pressure to the spoiler actuators also oper-

ates a pressure-operated valve which is held
in a position that prevents pressure from the
spoiler holddown accumulator from entering
the holddown line. The other path taken is to
the spoiler holddown accumulator. It is
charged by main system pressure and has a ni-
trogen precharge of 1,500 psi.

FOR TRAINING PURPOSES ONLY

27-75
 FlightSafety
InIemationaI

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

-
SPOILER HOLDDOWN MODULE
r-- ------------, ,-

-
I I

-- I
I
I
L
--+

HOLD DOWN
ACCUMULATOR

PRESSURE
SWITCH
HOLDDOWN
PRESSURE
SWITCH I
I
I
I
I
I
I
I

----------------------- I

LEGEND
• PRESSURE
RETURN
MECHANICAL

ELECTRICAL

Figure 27-45. Spoiler Operation in HOLD DOWN Position

27-76 FOR TRAINING PURPOSES ONLY

 FlightSafety
International

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

If the SPOILER switch is placed in the HOLD The auxiliary pressure dump valve is under the
DOWN position, DC power is applied to en- baggage compartment floor, mounted on the
ergize the solenoid valve to the position shown underside of the tray that holds the spoiler
in Figure 27-45. holddown module.

Pressure to the servo valves on the actuators The name, auxiliary pressure dump valve,
and to the pressure-operated valve on the refers to its original purpose. The valve is
spoiler holddown line is shut off, allowing deenergized open at auxiliary pump termina-
the spring to push the valve open. tion and dumps residual auxiliary pump pres-
sure to return, allowing the four shuttle valves
The path from the holddown accumulator to to return to their normal positions.
the stow side of the spoiler actuators is opened
by the pressure-operated valve, and the spoil- If operation of the pump is not desired, the
ers will stow. However, since main hydraulic AUX HYD PWR switch should be placed in
pressure is still available to the accumulator, OFF. In the OFF switch position, the auxiliary
it is actually the main system pressure hold- pump will not respond to inputs from the 1,200
ing the spoilers down. psi switch or the SPOILER holddown switch.
As can be seen, the effect is the same if main
system pressure is lost the spoiler holddown CAUTION
valve will energize as main system pressure
drops below 1,200 psi. Ensure that the AUX HYD PWR
switch is OFF when the aircraft nor-
In addition to main system pressure charging mal DC power is removed after
the holddown accumulator, the electric aux- flight. The pump is powered from the
iliary pump can also charge the accumulator hot battery bus and will drain the
through a check valve. The AUX HYD PWR aircraft batteries.
switch on the center switch panel has three po-
sitions. With the switch positioned to NORM, In the event that both main and auxiliary hy-
the auxiliary hydraulic pump will operate if draulic pressure is lost, the holddown accumu-
the SPOILER switch is positioned to HOLD lator will still provide the necessary pressure
DOWN or if main hydraulic system pressure to keep the spoilers fully stowed.
drops below 1,200 psi.

The electrical ground that triggers pump op-

eration also energizes the auxiliary pressure
dump valve. As the energized auxiliary pres-
sure dump valve changes position, it directs
auxiliary pump pressure to:
• Roll holddown release valves (spring-
force pressure-operated valves that will
now block holddown pressure from the
roll spoilers).
• Roll shuttle valves to provide operating
pressure to reactivate the roll spoilers.
• Spoiler holddown accumulator to en-
sure that it remains charged while main
system pressure is lost

FOR TRAINING PURPOSES ONLY 27-77

 FlightSafetyinternational

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

.. ....
'

." ." ..
'
,"

NOTE:
SHAKERS INSTALLED
BOTH COLUMNS
ON (

Figure 27-46. Stick Shaker Installation

TRANSMITIER
PROBE

~
RIGHT TRANSMITIER SHOWN
LEFT OPPOSITE

"~':"
..•.::\::~ .
....... .. "'"!:::.,

Figure 27-47. Angle-of-AttackiStall Warning System Components

FOR TRAINING PURPOSES ONLY

27-78
 F ightSafety
InlematIonaI

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

ANGLE-OF-ATTACK NOTES
AND STALL WARNING
SYSTEM
COMPONENTS DESCRIPTION
AND OPERATION
Stick Shaker
A stick shaker assembly is mounted on each
control column (Figure 27-46). Each shaker
consists of a housing enclosing a 28-VDC
motor and rotating weights. The shakers are
/
. energized through the angle-of-attack com-
puter when the proper threshold is reached.

Computer
The angle-of-attack computer is located on the
right side of the aircraft at FS 157.00 and WL
118.34. It is the central component of the stall
warning and indicating system (Figure 27-
47). Using a composite flap angle (average of
positions of potentiometers on asymmetry
brakes) and the right speedbrake angle as in-
dependent variables, it synthesizes angle-of-
attack thresholds which are compensated for
wing configuration. Whenever the in-flight
angle reaches or exceeds the threshold, the
computer supplies shaker power.

The right speed brake position transmitter sup-

plies inputs to the computer.

Transmitter Probe
Angle-of-attack transmitter probes located
on both sides of the aiprIane at FS 165.00 and
WL 119.81 are utilized to provide aircraft at-
titude information to the system. The transmit-
ter is a null-sensing internal vane instrument
employing a conical pressure-sensing probe
which extends through the aircraft fuselage
perpendicular to the local airflow. Directly
coupled to the probe are a paddle (internal
vane) and the electrical output elements. The
paddle has two blades. There are two separa-
tors in the transmitter housing so that two
paddle chambers are formed.

FOR TRAINING PURPOSES ONLY

27-79
 F ightSafety
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CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

-
-

AOA PROBE
Figure 27-48. Angle-of-AttacklStall Warning System

27-80 FOR TRAINING PURPOSES ONLY

 FlightSafety
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CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

OPERATION The approach indexer provides a heads up

display of deviation from the approach refer-
Angle-of-attack readouts are provided by an ence. The display is in the form of three lighted
indicator, a speed command on the attitude di- symbols which are used to indicate relative
rector indicator (ADI), and an indexer; stall speed to the approach reference. The follow-
warning is provided by a stick shaker on each ing angle-of-attack indications occur:
control"column (Figure 27 -48). Inputs to this
system are from angle-of-attack transmitter • Angle-of-attack high-top (red) chevron
probes, flap position sensors, and spoiler and lighted
speedbrake position sensors. Electrical power
is supplied by the right branch DC bus. • Angle-of-attack on reference-(green)
circle lighted
The angle-of-attack transmitter probes, one on
each side of the forward fuselage, detect air- • Angle-of-attack low-bottom (amber)
flow angle and provide signals to the angle- chevron lighted
of-attack computer.
The top chevron points down, indicating that
The computer receives angle-of-attack infor- the pitch attitude must be decreased. The bot-
mation from the transmitters, flap position tom chevron points up to indicate that the
sensors, and spoiler and speedbrake position angle of attack can be increased.
sensors to drive the pointer on the face of the
indicator. The indicator has four colored A second indexer for the copilot is available
arcs-green, white, yellow, and red. as an optional installation.

The green arc (0 to 0.55) is the normal oper- Two amber lights are located above the ANGLE
ating range of the aircraft. The white arc (0.55 OF ATTACK indicator. Illumination oftheAOA
to 0.65) covers the area between the normal PROBE light indicates a fault in a probe or
operating range and the caution area. The mid- transmitter, but the system remains operative
dle of the white arc (0.6) represents the opti- since the computer selects the most conserva-
mum landing approach airspeed (V AAP)' The tive input. Faulty inputs from the flap spoiler
yellow range (0.65 to 0.75) is a caution area or speedbrake sensors will illuminate the FLAPS
where the aircraft is approaching a critical SPDBK/SP light. In this case, a stick shaker ac-
angle of attack. tuation range of 0.715 to 0.825 in the red arc is
selected. The OFF flag is visible and the index-
The red arc (0.75 to 1.0) is a warning area and ers and the FAST/SLOW indicator on the ADI
represents the beginning of low-speed buf- are inoperative.
fet to full stall. At an indication of 0.82 :t.02
in the warning range, the stick shakers are ac- If a fault occurs in the stall warning system that
tivated. renders it inoperative, the amber STALL WARN
annunciator illuminates. The ANGLE OF AT-
The indicator displays lift information with 0 TACK indicator OFF flag is visible, the
representing zero lift and 1.0 representing FAST/SLOW indicator on the ADI is inopera-
full stall. Lift is presented as a precentage, and, tive, and the stick shakers do not operate.
with speedbrake and flap position information,
the display is valid for all aircraft configura-
tions and weights. Therefore, at 1.0 where
full stall occurs, 100% of the available lift is
being produced.

FOR TRAINING PURPOSES ONLY 27-81

 FlightSafety
International

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

TENSION (POUNDS)
30

20

10

o
O' 10' 20' 30' 40' SO' 50' 70' 80' 90' 100' 110'

TEMPERATURE (oF)

Figure 27-49. Roll Spoiler Cable Tension Graph (Typical)

BROKEN WIRE DETECTED VISUALLY

WHEN CABLE WAS REMOVED
AND BENT

NORMAL TECHNIQUE FOR

BENDING CABLE AND
CHECKING FOR BROKEN WIRES

,
DO NOT BEND INTO LOOP SMALLER
THAN 50 CABLE DIAMETERS

Figure 27-50. Cable Inspection

27-82 FOR TRAINING PURPOSES ONLY

 Flight~jI!~ty
CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

MAINTENANCE NOTES
CONSIDERATIONS
GENERAL MAINTENANCE
PRACTICES
NOTE
The following maintenance prac-
tices are of a general and abbreviated
nature. Complete procedures per-
taining to specific maintenance tasks
are found in Chapter 27 of the main-
tenance manual.

Control Cables
Control cable tension must be set to predeter-
mined values for proper flight control system
operation. Cable tension varies with cable
size, length, and usage, and is predicated on
ambient temperature. Complete cable ten-
sioning information is found in Chapter 27 of
the maintenance manual. A typical cable ten-
sioning graph is shown in Figure 27-49.

Cable Inspections
Control cables are subjected to a variety of en-
vironmental conditions and forms of deteri-
oration that result in wire/strand breakage or
corrosion. Broken wires can be detected by
passing a cloth along the length of the cable
(Figure 27-50).

The absence of snags is not positive proof

that broken wires do not exist. Remove the
cable and bend it into a loop, as shown. Broken
wires will then be readily apparent.

Control cables should also be inspected for

corrosion damage as outlined in Chapter 51 of
the maintenance manual.

FOR TRAINING PURPOSES ONLY

27-83
 FlightSafety
Inlematlooal

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

C274000 TASK: LUBRICATE HORIZONTAL STABILIZER AFT HINGE BOLTS

Tools And Special Equipment

Grease Gun Commercially Available

Consumables
Grease, Silicone with Teflon MIL-G-83261
Lubricant. LPS3 Heavy Duty Rust Inhibitor

HORIZONTAL STABILIZER
1. Lubricate Horizontal Stabilizer Aft Hinge Bolts.
AFT HINGE BOLT
LUBRICATION POINT

Figure 27-51. CESCARD Example

SEAL
RETAINING RING
ADAPTER

PRESSURE SEAL (NONPRESSURIZED SIDE)

(PRESSURIZED SIDE)

PACK WITH
GREASE
PIN 5565450-28

- - /
~ RETAINING RING

Figure 27-52. Cable Pressure Seal

27-84 FOR TRAINING PURPOSES ONLY

 FlightSafety
Internatlonal

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

Example of Flight Control The CESCARD Maintenance work cards in-

Lubrication Instructions clude lubrication instructions and illustra-
tions on the appropriate task card. An example
The aft hinge bolt (pivot point) of the horizon- is given in (Figure 27-51).
tal stabilizer is lubricated with grease MIL-
G-83261 (silicone with Teflon), using the
following procedure: Cable Pressure Seals
I. Gain access to the grease fitting through Cable pressure seals of molded synthetic rub-
a hole/notch in the bullet fairing. ber are installed on the control cables where
they pass through a pressure bulkhead (Figure
2. Pump grease until grease can be seen 27-52).
extruding from between ears of the
pivot fitting. The seal is split longitudinally to the center
to permit removal without disturbing the con-
3. Wipe off excess grease. trol cable. A light silicone grease packed into
the seal lubricates the cable. Retainer rings
The scissor assembly at the horizontal stabi- maintain the seal in place. Specific instructions
lizer primary trim actuator is lubricated with for seal maintenance are provided in Chapter
LPS-3 spray at the upper attach fitting and the 20 of the maintenance manual.
lower mounting bracket.

The primary flight control (gust) lock cranks, Pulleys

pins, aileron push-pull rod, and bearing sur-
Cable pulleys should be periodically rotated
faces are lubricated with MIL-G-23827 or
to provide a new bearing surface for the cable
MIL-G-81322 grease. Refer to Scheduled
Lubricati on/Cleaning-Description and since it sometimes operates in a small arc.
Operation. Various cable system malfunctions may be
analyzed by observing pulley conditions.
Common wear patterns related to given mal-
Each flap actuator ballscrew is lubricated with
functions are shown.
a MS24204-1 grease gun nozzle adapter using
MIL-G-23827 grease. Remove button plug
from the lower pan of flap islands to gain ac-
cess to grease fitting on Nos. 2, 3, and 4 left
Control Surface Rigging
and right flap actuators. • Rig pins must be a slip fit in the rig pin
hole. Incorrect size of rig pin or a cable
Flap yokes, gimbal yokes, and flap end fittings tension load can damage structural
in the areas from which paint has been re- components.
moved to bare steel should be lubricated with
LPS 3 on all bare surfaces. After application, • Turnbuckles must be adjusted alter-
all excess lubricant must be wiped clean with nately to prevent cable tension load
a clean cloth. LPS 3 will prevent rusting of the against rig pins.
steel parts.

The special lubrication bolts securing the flap

carriage rollers are lubricated by applying
MIL-G-21164 grease at the fitting in the bolt
head until the grease extrudes from between
the roller and the yoke. Wipe off all excess
grease.

FOR TRAINING PURPOSES ONLY

27-85
 FlightSafety
International

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

INSPECTIONS • Aileron quadrants-Inspect for secu-

rity of installation and corrosion,
Flight control systems inspections are per- pushrods for corrosion and evidence of
formed in accordance with the schedule spec- damage.
i fied and as ou tI i ned in Chapter 5 of the
maintenance manual. The inspections include, • Aileron cables-Inspect for broken
but are not limited to, the following items per- wires, fraying, chafing, clad damage
formed at various time intervals specified in and cleanliness, pulleys for freedom of
the maintenance manual: rotation.
• Control column-Inspect torque tube,
• Aileron trim control and indication-
pushrods for evidence of damage,
Inspect for security of installation, free-
columns for freedom of movement.
dom of movement. Verify surface
displacement is an equal amount in each
• Ailerons-Check for travel and free-
direction for equal amount of trim input.
dom of movement.
• Rudder-Inspect hinges, hinge bolts,
• Rudder bias system-Inspect actuator,
torque tube for evidence of damage, wear,
tubing, and hoses for security of instal-
failed fasteners, security of installation.
lation, evidence of leakage. Check rud-
der torque arm for security of
• Rudder pedals and cables-Inspect ped-
attachment and cracks, solenoid shut-
als, torque tube, brake arm, and adjust-
off valve for operation.
ment levers for security of installation,
evidence of damage, corrosion, and
• Elevator front quadrant and forward ca-
cleanliness. Check cable seals at aft
bles-Inspect stop bolts, pushrods for
pressure bulkhead for deterioration and
corrosion, evidence of damage, secu-
lubrication.
rity of installation. Check cables for
fraying, chafing, broken strands.
• Rudder bellcrank assembly-Inspect
bel1crank, stops for corrosion, evidence
• Elevator rear quadrant-Inspect quad-
of damage, and security of installation.
rant, pushrods, stops, torque arms, idler
arms for corrosion and evidence of dam-
• Horizontal stabilizer, pilot and copilot
age, failed fasteners, security of instal-
pitch trim switches-Perform split trim
lation. Check aft cables for fraying, switches check.
chafing, broken strands.
• Flap power drive assembly-Inspect for
• Horizontal stabilizer primary trim actu- security of installation, evidence of
ator-Perform torque check on motor overheating, and obvious damage to
brake. electrical components.
• Flap drive actuators, gearboxes, flexi-
• Aileron assemblies-Inspect hinges for ble drives-Inspect for security of in-
damage and security, pivot bolts for cor- stallation, cleanliness, chafing, evidence
rosion and safety, counterweights for of damage.
damage and attachment. • Flap carriages-Inspect for cracks,
damage, cleanliness. Check roller bear-
ings for cleanliness, nicks, evidence of
damage.

27-86 FOR TRAINING PURPOSES ONLY

 FlightSafety
Inlemational

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

• Flap inboard and outboard yokes- FUNCTIONAL CHECKS

Inspect both yokes for damage and
attachment. Functional (operational) checks are some-
• Flaps-Inspect all linkages for attach- times performed after maintenance has been
ment, function, safety. accomplished on a system. They may also be
performed when a system is suspected of mal-
• Flap asymmetry brakes-Perform func- functioning. Functional checks of a system are
tional test with breakout box.
frequently a part of a maintenance procedure
• Spoiler and speedbrake linkage- and are not called out separately. They some-
Inspect for damage. Check pushrods, times include maintenance tasks such as final
bellcranks, pivot mounts for security adjustment, torquing, and safetying.
and damage.
• Speedbrakes and ground spoilers- Functional or operational checks of the follow-
Inspect all linkages, bellcranks, pivot ing systems are specified in Chapter 27 of the
bolts for attachment, damage, wear. maintenance manual:
• Spoiler/speedbrake hydraulic fluid tem- • Flight control lock system
peratures monitor unit-Inspect for se-
curity of installation, damage to • Aileron control system
electrical wiring. Observe BITE equip- • Rudder bias system
ment for indication of sensor failure, • Horizontal trim system (including pri-
sensors for security of installation and mary and secondary split trim switches)
leaks, wiring for damage.
• Flap system (including asymmetry
• Spoiler/aileron disengage mechanism- brake)
Inspect mechanism and all related parts.
Perform operational check and reset. • Aileron/spoiler interconnect/disconnect
system
In addition to scheduled inspections, the • Roll spoilers system
Components Time Limits schedule in Chapter
• Speedbrake system
5 of the maintenance manuall requires re-
placement of the following components at the • Spoiler/speedbrake holddown system
number of flying hours specified: • Emergency descent spoiler system
• Horizontal stabilizer primary actuator-
6,000 hours.

• Center flap assemblies-15,000 hours

FOR TRAINING PURPOSES ONLY 27-87

 FlightSafety
International

CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

LOST MOTION BETWEEN

CONTROL WHEELS AND
AILERONS.

CHECK CABLE TENSION.

IF -

I
OK, CHECK FOR NOT OK, ADJUST
PROPERLY ROUTED CABLES TO PROPER
CABLES. IF - TENSION.

I
OK, CHECK FOR NOT OK, PROPERLY
DEFECTIVE PULLEYS, ROUTE CABLES IN
GUARDS, AND IMPROPER PULLEYS AND FAIRLEADS.
PULLEY INSTALLATIONS.
IF -

I
NOT OK, REPLACE
OK, CHECK FOR
DEFECTIVE PULLEYS,
DEFECTIVE AILERON
GUARDS, AND/OR
QUADRANTS OR
PROPERLY INSTALL
ACTUATOR ASSEMBLY.
PULLEYS.

NOT OK, REPLACE

DEFECTIVE QUADRANT
OR ACTUATOR ASSEMBLY

Figure 27-53. Aileron Control System Troubleshooting Chart

FOR TRAINING PURPOSES ONLY

27-88
 CITATION 650 SERIES MAINTENANCE TRAINING MANUAL

FAULT ANALYSIS NOTES

Isolation of a fault or malfunction can be accom-
plished by a systematic analysis of the trouble. A
reproduction of an aileron control system trou-
blehsooting chart from Chapter 27 of the mainte-
nance manual is shown in (Figure 27-53).

LIMITATIONS
If any of the following systems do not oper-
ate satisfactorily during preflight testing,
maintenance action is required prior to flight:
• Primary trim system

• Secondary trim system

• Rudder bias system (including heater)

• Speedbrake/spoiler system

• Roll spoilers

FOR TRAINING PURPOSES ONLY 27-89