Bombardier-Challenger_00-Flight_Controls庞巴迪挑战者飞行操纵

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庞巴迪挑战者飞行操纵

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canadair
chanencjer
OPERATING MANUAL
PSP 606
SECTION 70
FLIGHT CONTROLS
TABLE OF CONTENTS
Subject Page
GENERAL 1
Control Disconnect Systems
Power Control Units 2
Artificial Feel Mechanisms
Trim Systems 7
Control Surface and Trim Position Indicators
Gust Locks
Bypass Valves
Damping Valves
Relief Valves 9
ROLL CONTROL SYSTEM S
Ai leron Trim V
Aileron Control Wheels
Artificial Feel Mechanisms
Aileron Position Transmitters
Aileron Control Cable Tension Regulator 13
Aileron Flutter Dampers
YAW CONTROL SYSTEM 13
Rudder Trim 14
Rudder Pedal Assemblies
Ant i-Jam Mechanisms
I Artificial Feel Mechanisms
Rudder Position Transmitter 15
PITCH CONTROL SYSTEM 15
Horizontal Stabilizer 16
Control Columns 17
Gain Change Mechanisms
Artificial Feel Mechanisms
Anti-Jam Mechanisms 18
Elevator Position Transmitter
Elevator Flutter Dampers
WING FLAP SYSTEM 18
Flap Control Unit 21
Power Drive Unit
Flap Actuators 23
Flexible Shaft Drive Assemblies
Asymmetry/Overspeed Detector and Brake Assemblies
Flap Position Transmitter
10-
Page
Apr
cacntiaaauaeirn cjer
OPERATING MANUAL
PSP 606
Subject
SPOILER SYSTEM
Flight Spoilers
Ground Spoilers
STALL PROTECTION SYSTEM
Angle-of-Attack Transducers
Stall Protection Computer
Stall Protection System Monitoring
Stick Shakers
Stick Pusher Subsystem
Stall Protection System Test Indicators
Aural Warning Horn
Failure Warning Lights
System Test Switches
Systems Without Altitude Compensation
LIST OF ILLUSTRATIONS
Figure
Number Title
1 Flight Controls and Associated Instruments
2 Flight Controls Hydraulics
3 Servo Monitor Panel
4 Surface Trim Control Panel
5 Control Surface Position and Trim Position Indicators
6 Control Wheel
7 Wing Flap System Components
8 Wing Flap Controls and Indication
9 Spoiler Controls and Indication
10 Stall Protection System Controls and Indication
11 Stall Protection System Indicators
12 Stall Protection System Panels (2 Sheets)
13 Stall Margin Indicators
canadair
chaiienQer
OPERATING MANUAL
SECTION 10
FLIGHT CONTROLS
GENERAL (Figures 1 and 2)
The primary flight controls, consisting of roll control, yaw control, pitch
control, flight spoilers and ground spoilers, are fully powered from all three
hydraulic systems. Mechanical inputs from the pilots1 controls in the flight
compartment are conveyed via push-pull rods, quadrants and cables to power
control units (PCU). There is no interconnection between hydraulic systems,
and all PCUs are totally independent of each other. The secondary controls
consist of the wing trailing edge flaps and control surface trim systems, and
are electrically controlled and actuated.
The ailerons, elevators and the flight spoilers are each powered by two of the
three independent hydraulic systems. The rudder is powered by all three
systems and the ground spoilers are powered by No. 1 system only (refer to
Figure 2). The primary flight control systems are capable of continued safe
operation if jamming or disconnection of a component, loss of normal electrical
power and, with the exception of the spoilers, loss of hydraulic systems No. 7
and/or No. 2 occur.
Jamming or disconnection of a component is nullified by incorporation of dual
control circuits with anti-jam and/or disconnect mechanisms.
Loss of normal electrical power is overcome by an air-driven generator (ADG)
which is capable of supplying emergency electrical power and deploys
automatically if normal electrical power is lost.
Loss of hydraulic systems No. 1 and/or No. 2 is catered for by hydraulic system
No. 3 which supplies a PCU for each of the primary controls except spoilers.
A. Control Disconnect Systems
Control disconnect mechanisms are provided for disconnecting the control
columns (pitch control) and the control wheels (roll control), if a jam
occurs in their respective cable runs. The disconnect oechanisms are
located under the flight compartment floor and are operated by the PITCH
DISC and ROLL DISC T-handles on the centre pedestal (refer to Figure 1).
The handles are normally stowed in small recesses in the centre pedestal
when the disconnect mechanisms are engaged. When either handle is pulled
up, its associated mechanism is disengaged. The handle can be secured in
the disconnect position by rotating it left or right to engage a detent in
the centre pedestal. When the PITCH DISC handle is pulled, the pilot has
control of the left elevator and the copilot controls the right elevator.
When the ROLL DISC handle is pulled, the pilot controls the left aileron
and the copilot controls the right aileron. The controls can be
reconnected by releasing the handle to the stowed position and aligning the
control columns or the wheels as appropriate.
SECTION 10
Page 1
Apr 4/83
canaaair
cftanencjer
OPERATING MANUAL
If a jam occurs in the rudder control circuits, break-out bungees and an
anti-jam mechanism isolate the jammed circuit. Yaw control is retained by
both pilots.
Power Control Units
The primary flight control surfaces are fully power-operated by hydraulic
actuators known as power control units. To provide for failsafe operation
and eliminate fluid interflow between the three aircraft hydraulic systems,
each aileron is powered by a dual PCU consisting of two independent
actuators; each elevator is powered by two independent PCUs; and the rudder
is powered by three independent PCUs.
Although the PCUs of the ailerons, rudder and elevators differ in
appearance, the principle of design and operation is similar. Each PCU
consists mainly of a control-valve-operated piston moving in a cylinder.
The control valve is operated by a mechanical linkage and directs hydraulic
fluid under pressure to one side of the cylinder for actuation, or blocks
off pressure from both sides when actuation is completed.
The PCUs are connected to the control surfaces by rod-end attachments and
operate to move the control surfaces in the desired direction upon receipt
of a signal from the pilots' controls or from the automatic flight control
system (AFCS). A flight control monitoring unit, located in the underfloor
avionics bay, together with proximity sensors associated with each PCU,
monitors the operation of the PCUs. The flight control monitoring unit
receives inputs from the proximity sensors and transmits warning signals,
via the master caution and warning system, to the servo monitor panel in
the flight compartment (refer to Figure 3). Each proximity sensor forms
part of a solid state electrical circuit that produces voltage variations
when a metal target moves within a predetermined distance of the sensor.
The sensors perform the function of conventional microswitches but do not
require electrical contacts.
The proximity sensors of the roll and yaw control systems are integral
parts of their associated PCUs and are capable of detecting a PCU
malfunction caused by a jammed PCU control valve or a hydraulic supply
deficiency. The pitch control proximity sensors are mounted on the input
linkages of their associated PCUs and detect only PCU malfunctions caused
by a jammed PCU control valve.
Artificial Feel Mechanisms
Because the primary flight control surfaces are fully power-operated,
artificial feel mechanisms, consisting of cam-foilower-spring devices, are
incorporated in the control systems to simulate aeroctynamic forces and
provide a means of sensing control loads under various flight conditions.
SECTION 10
Page 2
May 28/82
cacntiaadilmenirQ sr
AILERON
OPERATING MANUAL
PSP 606
MASTER CAUTION AND INDICATOR LIGHTS
ELEVATOR
SPOILER CONTROL SWITCHES
AND SPLRS INOP WARNING LIGHT
j T ] SPOILER CONTROL
^ LEVER
CONTROL SURFACE TRIM PANEL
Items associated with stall protection system
are not shown (refer to Figure 10).
PILOT'S RUDDER PEDALS
EFFECTIVITY
71 Aircraft incorporating
SB 600-0452. For spoiler
control lever on other A/C,
refer to Figure 9.
Flight Controls and Associated Instruments
Figure 1 SECTION 10
Page 3/4
Feb 12/88
canadair
ctiaiiencjer
OPERATING MANUAL
NO. 1 SYSTEM NO. 3 SYSTEM NO. 2 SYSTEM
I ACC 0
TO
LANDING GEAR
AND
BRAKE SYSTEMS
LEGEND
NO. 1 HYDRAULIC SYSTEM
I NO. 2 HYDRAULIC SYSTEM
I NO. 3 HYDRAULIC SYSTEM
Flight Controls Hydraulics
Figure 2
SECTION 10
Page 5
May 28/82
OPERATING MANUAL
PSP 606
NOTE
Refer to section 4 for stabagmtn panel.
PITCH LIGHT
Amber PITCH light comes on when proximity sensors
detect a jammed control valve or input linkage at the
elevator power control units.
I SERVOMON
PITCH ROLL
MON SAFE LIGHT
Green MON SAFE fight comes on when ail
aileron and rudder PCUs are unpressurized
(all hydraulic systems off) and all elevator
PCUs are unjammed.
ITUryQSTAB AG MTN
MACH TRIM Y/D
I TEST
ON
QOFF DISENGA
NOTE
With hydraulic power off, servo monitor
panel lights are as follows:
— ROLL light is on
—YAW light is on
— PITCH light is out
-MON SAFE light is on.
ROLL AND YAW LIGHTS
Amber ROLL and YAW lights come on whenever
proximity sensors detect a jammed control valve or
hydraulic pressure deficiency at the respective power
control units.
Servo Monitor Panel SECTION 10
Figure 3 Page 6
Jun 12/86
canadair
chaiienQer
OPERATING MANUAL
PSP 606
D. Trim Systems (Figure 4)
Trim inputs are introduced into the roll and yaw control systems by
electrically driven actuators controlled by the AIL TRIM and RUD TRIM
switches on the centre pedestal. Pitch trim i s obtained by varying the
angle of incidence of the horizontal stabilizer. Signals from the pitch
trim switches on the control wheels, from the AFCS and from the stability
augmentation system (SAS) are processed by a control unit to operate an
electrically driven actuator which applies the required amount of
stabilizer deflection. The pitch trim disconnect switch on each control
wheel disconnects and brakes the pitch trim actuator in an emergency (refer
to Figure 6).
E. Control Surface and Trim Position Indicators (Figure 5)
Flight control surface positions and trim angles are displayed on
indicators located on the centre instrument panel in the flight
compartment. A flap position indicator on the copilot's instrument panel
displays flap position angles. Inputs to the position indicators are
provided by transmitters and trim actuators.
F* Gust Locks
To provide gust locking for the ailerons and the rudder, a bypass valve is
included in each of the rudder PCUs, and a damping valve is included in
each of the aileron PCUs. The valves are similar in construction and
actuation but operation of the damping valves differs slightly from that of
the bypass valves. Gust locking for the elevators is provided by relief
valves on the elevator PCUs.
(1) Bypass Valves
When hydraulic pressure is removed from the PCU, spring force moves
the valve spool to the bypass position at which the extend and retract
ports of the PCU are interconnected through a restrictor orifice in
the bypass valve. At the same time, the retract port of the PCU
control valve is blocked. These two actions lock the control surface
against the effect of gusts but permit restricted movement of the
surface, if a sufficiently large external force i s applied continously.
When hydraulic pressure is restored, the valve spool moves back to the
non-bypass position. In this position, the retract port of the PCU is
connected to the PCU cylinder and normal operation of the control
surface i s possible.
(2) Damping Valves
When hydraulic pressure is removed, spring force moves the valve to
the damping position. With the valve in this position, the extend
port of the PCU cylinder is connected through a restrictor port in the
damping valve to the pressure port of the PCU control valve which is
SECTION 10
Page 7
May 28/82
canadair
chaiienqer
OPERATING MANUAL
PSP 606
RUDDER TRIM CONTROL
Control switch sets rudder trim left and right.
RUD TRIM
OFF (§)
OVERSPEED/CHANNEL CHANGE SWfTCH/LIGHT
Amber lights indicate pitch trim oveopeed or channel
change. Can be used to change from one channel to
other for test.
Pressing switch/light in conjunction with CHAN 1/
CHAN 2 switch/light actrvsret pitch mm system.
PITCH TRIM/® AIL TRIM
r—PUSH—i / V^ — ^S
OVSP
OFF
CHANNEL INOPERATIVE SWITCH/LIGHT
CHAN 1 INOP
CHAN 2 INOP
Amber lights indicate failure in respective channel.
Pressing switch/light in conjunction with OVSP/
CHANGE CHAN switch/light activates pitch trim
system.
AILERON TRIM CONTROLS
Control switches sets aileron trim up and
down.
Surface Trim Control Panel
Figure 4
SECTION 10
Page 8
Jun 12/86
cacnhaadnaeirn qer
OPERATING MANUAL
open to the retract port of the PCU cylinder. When the PCU is in this
configuration, the aileron is gust locked but is capable of restricted
movement when a steady external force is applied to it.
When hydraulic pressure is applied, the pressure overcomes the spring
force and moves the valve spool to the non-damping position. In the
non-damping position, the damping valve connects the extend port of
the PCU cylinder to the return port of the PCU control valve and
connects the pressure line to the pressure port of the PCU control
valve, rendering the PCU operative.
(3) Relief Valves
The pressure relief valve on each elevator PCU connects the
PCU cylinder extend and retract pressure lines. When hydraulic
pressure is removed, a spring forces the valve closed. This action
gust locks the elevator but allows restricted movement of the surface
if a steady external force is applied to it.
When hydraulic pressure is applied, the relief valve opens and
connects the PCU cylinder extend and retract pressure lines to restore
normal operation of the elevator.
2. ROLL CONTROL SYSTEM
Roll (lateral) control is achieved by hydraulically powered ailerons which are
controlled primarily from conventional column-mounted, horn-type wheels through
a system of pulleys, cables, quadrants, push-pull rods, levers and bellcranks.
Aileron movement is limited by the operating range of the hydraulically
actuated PCUs to which the ailerons are connected, and by mechanical stops on
the pilot's and copilot's control wheels. Primary control is supplemented by
an electrically actuated trim system.
The roll control system incorporates a dual PCU for each aileron, and a dual
control system. Normally, both control systems are interconnected by a
cross-coupling shaft so that there is simultaneous movement of both ailerons;
but it is possible to isolate a jammed aileron control circuit by means of a
disconnect mechanism, thereby allowing limited control (one aileron only)
through the unjammed circuit (refer to paragraph I.A.).
Control wheel movement is transmitted by cable to a pulley located at the base
of each control column, then horizontally rearward under the flight compartment
floor where each of the twin cable circuits drives one of the two forward
quadrants which form a transverse cross-coupling shaft. From the forward
quadrants, control cables are routed under the cabin floor to rear quadrants,
located in the main landing gear bay, each of which incorporates an artificial
feel unit. Output from each rear quadrant is transmitted outboard, by cable,
to a PCU input quadrant located outboard in the wing, forward of the rear spar.
Each input quadrant has a cable tension regulator and is connected to the
input/feedback linkages of the dual PCU through a common linkage.
SECTION 10
Page 9
May 28/82
LANDRFLTSPLR
Right spoiler up indications.
Max 40 degrees
3/4 28 degrees
V2 16 degrees
V* 5 degrees
L AND R ELEVATOR
Up/down indications
Up 23.6 degrees
Down 18.4 degrees
canadair
ctianenQer
OPERATING MANUAL
PSP 606
L AND R AILERON
Wing up/down indications
Up 21.3 degrees
Down 21.3 degrees
RUDDER
Left/right indications
Left 20 degrees
Right 20 degrees
CONTROL SURFACE POSITION INDICATOR n
RUDNLANDNR
Nose left (NL)/nose right (NR) indications
Left 8.5 degrees
Right 8.5 degrees
AIL LWD AND RWD /»-
Wing up/down indications
Up 7.5 degrees
Down 7.5 degrees
CONTROL SURFACE TRIM POSITION INDICATOR
©
STAB NUP
Nose up (NUP) indications
Stabilizer moves from 0 to -9 degrees incidence.
Green band indicates take-off (TO) trim range.
NOTE
If input signals to trim indicator are lost, aileron
and rudder pointers move off scale 90 degrees
from zero index. Stabilizer pointer moves off scale
to a point between scale end points.
Control Surface Position and
Trim Position Indicators
Figure 5
SECTION 10
Page 10
Mar 01/85
canadair
chaiienQer
OPERATING MANUAL
The actuator pistons are connected to the ailerons by jointed toggles, and each
actuator is capable of aileron operation should there be a failure associated
with the adjacent actuator.
Movement of the PCU input/feedback linkage causes movement of the PCU control
valve which, in turn, actuates the PCU piston to move the aileron according to
control command. Piston movement also repositions the input/feedback linkage
to return the PCU control valve to a neutral position thus preventing further
movement of the ailerons until a subsequent control signal moves the PCU input/
feedback linkage. Signal inputs from the AFCS are made through the rear
quadrant of the right aileron system only. Therefore, should jamming of the
right control system occur, the autopilot inputs would not be transmitted to
the left aileron system, (refer to SECTION 4, AUTOMATIC FLIGHT CONTROL SYSTEM).
A. Aileron Trim
An electrically driven actuator, located in the main landing gear bay
between the rear quadrants, applies a bias to the primary control circuit,
when required, by operation of the AIL TRIM switches located on the centre
pedestal. The trim actuator is connected to the rear quadrants via
push-pull rods and bell cranks. The amount of trim applied to the ailerons
is shown on the control surface trim position indicator, located on the
left of the centre instrument panel.
B. Aileron Control Wheels (Figure 6)
The aileron control wheels are horn-type handwheels spline-mounted on the
control columns. The wheels are connected by cables to the aileron forward
quadrants via pulleys located near the base of the control columns. Each
control wheel mounts a pitch trim switch, a pitch trim disconnect switch,
an autopilot/stick pusher disconnect switch, an autopilot touch control
switch and a radio key.
C. Artificial Feel Mechanisms
Two artificial feel mechanisms, one at the rear of each rear quadrant,
provide the pilots with positive feel of the power-operated control system
and act as centering devices.
D. Aileron Position Transmitters
A position transmitter connected to each aileron transmits aileron position
signals to the control surface position indicator on the centre instrument
panel (refer to Figure 5). An additional aileron position transmitter,
connected to the left aileron, signals aileron position to the stability
augmentation computer of the automatic flight control system. The aileron
position signals are used to enhance yaw damping during rolling manoeuvres.
SECTION 10
Page 11
May 28/82
canatiair
ctianencjer
OPERATING MANUAL
AUTOPILOT/STICK PUSHER DISCONNECT SWITCH
Red pushbutton which, when pressed, disengages the
autopilot and deactivates the stick pusher. When
released, the stick pusher system is immediately
reactivated but the autopilot remains disengaged.
PITCH TRIM SWITCH
Enables the pilot to vary pitch trim
according to flight requirement.
RADIO KEY
Light grey button which, when
pressed, switches on the radio
transmitter.
FRONT VIEW
AUTOPILOT TOUCH CONTROL
Black button which, when pressed, enables
the pilot to manoeuvre the aircraft without
disconnecting the autopilot.
PITCH TRIM DISCONNECT SWITCH
Red button which, when pressed, disconnects the
pitch trim system to halt a possible trim
malfunction, for example, a runaway trim
actuator.
REAR VIEW
Control Wheel
Figure 6
SECTION 10
Page 12
May 28/82
canadair
chaiienper
OPERATING MANUAL
E. Aileron Control Cable Tension Regulator
The aileron PCU input quadrant incorporates a cable tension regulator. The
tension regulator maintains optimum control cable tension by compensating
for changes in tension caused by temperature variations, stretching of the
control cables and deflection of the control system components.
F. Aileron Flutter Dampers
Each aileron i s protected by a single hydraulic flutter damper assembly.
The flutter damper consists of a pre-charged hydraulic cylinder containing
a double-acting piston. The piston rod i s connected via a shear link to
the aileron at the outboard aileron hinge assembly. Flutter damping occurs
when hydraulic fluid i s forced from one side of the piston to the other
through small-diameter passages. The hydraulic fluid level in the damper
can be checked through an integral sight gauge.
3. YAW CONTROL SYSTEM
Yaw (directional) control i s achieved by a hydraulically powered rudder,
controlled primarily from conventional dual, cross-coupled pedals through a
system of push-pull rods, levers, quadrants, cables, pulleys and bellcranks.
Rudder movement i s limited by the operating range of the hydraulically actuated
rudder PCUs, and by mechanical stops which limit the movement of the control
pedals. Primary control is supplemented by an electrically actuated trim
system.
The yaw control system incorporates three independent, parallel-connected PCUs
and a dual control system which includes two anti-jamming mechanisms for
isolating or overriding the effects of a jammed circuit, enabling control to be
maintained via the intact circuit. The system i s also protected by anti-jam
mechanisms built into the PCU input levers, which act to isolate a jammed PCU.
Movement of the pedal assembly i s transmitted to forward quadrants by a
push-pull rod and lever system which includes a primary feel mechanism and
forward anti-jam mechanism. From each forward quadrant, the control signal is
transmitted, by cables, along each side of the fuselage, under the cabin floor,
to a corresponding rear quadrant in the rear fuselage. The two rear quadrants,
located centrally side by side, convey the dual control signal onwards as a
single signal, via a secondary feel mechanism, a rear anti-jam mechanism, two
load limiters and a trim mixing system, to the input torque tube of the PCUs.
The PCU input torque tube incorporates one input and three identical output
levers, and each output lever i s connected to the input/feedback linkage, which
transmits the control signal to the control valve of the PCU. Operation of the
rudder PCUs i s similar to that of the aileron PCUs (refer to paragraph 2 . ).
In addition to control inputs from the pedal assembly, inputs from the
stability augmentation system of the AFCS are applied to the system throuah two
yaw dampers in the trim mixing system (refer to SECTION 4, AUTOMATIC FLIGHT
CONTROL SYSTEM).
SECTION 10
Page 13
May 28/82
canadair
ctiaitenQer
OPERATING MANUAL
Rudder Trim
An electrically driven actuator, connected to the rudder PCUs via the trim
mixing system, applies a bias to the primary control circuit, when
required, by operation of the RUD TRIM control located on the centre
pedestal. The amount of trim applied to the rudder is shown on the control
surface trim indicator located on the left of the centre instrument panel*
Rudder Pedal Assemblies
Each rudder pedal is pivot-mounted on its own tubular pedestal which, in
turn, is pivot-mounted to lugs on a cross-mounted tube secured to the
structure below the flight compartment floor. The pedals are pivot-mounted
to enable foot control of the aircraft wheel brake system via control rods,
levers and cables. The tubular pedestals are pivot-mounted to convey
pilots1 foot movement to the rudder control system.
Each set of pedals is provided with a hand-operated adjusting mechanism to
cater for the individual requirements of pilots. The right-hand set of
pedals incorporates the primary artificial feel mechanism which consists of
a simple cam-follower-spring device.
Anti-Jam Mechanisms
The two forward anti-jam mechanisms, one located adjacent to each forward
quadrant under the flight compartment floor, and one rear anti-jam
mechanism, located adjacent to the rear quadrants in the rear fuselage,
operate to nullify the effects of a jammed cable circuit. Normally, with
both cable circuits unrestricted, the rear anti-jam mechanism acts as a
summing device so that movement of the rear quadrants, though in opposite
directions, is summed to produce twice the output movement of one
quadrant. If one rear quadrant cannot move because of a jammed condition
in the cable circuit, the forward anti-jam mechanisms alter the pivot
points of the forward quadrants to produce twice the normal movement of one
rear quadrant and thereby maintain normal pedal/rudder movement ratio.
The anti-jam mechanism on each rudder PCU acts as a push/pull rod for the
PCU input linkage during normal operation. If the input linkage cannot
move because of a jam in the PCU, the anti-jam mechanism breaks out to
isolate the defective PCU from the system. The remaining PCUs continue to
operate the rudder.
Artificial Feel Mechanisms
Two artificial feel mechanisms are included in the yaw control system. A
primary mechanism, incorporated into the copilot's pedal assembly, provides
both pilots with positive feel of the power-operated system and acts as a
centering device for the system. A secondary mechanism, included in the
rear linkage, caters for control system backlash in addition to providing
feel and acting as a centering device.
SECTION 10
Page 14
May 28/82
cacnhaadiiaeinr cjer
OPERATING MANUAL
E. Rudder Position Transmitter
A position transmitter, located at the bottom rudder hinge, transmits
rudder position signals continuously, over the full range of travel, to the
control surface position indicator on the left of the centre instrument
panel.
4. PITCH CONTROL SYSTEM
Pitch (longitudinal) control is achieved primarily by two independent,
hydraulically powered elevators which are hinge-mounted to the trailing edge of
the horizontal stabilizer. Elevator movement is controlled from conventional
control columns through a dual system of pulleys, cables, quadrants, push-pull
rods, levers and bellcranks. Elevator movement is limited by the operating
range of the hydraulically actuated PCUs and by mechanical stops which limit
the movement of the control columns. Primary control is supplemented by an
electrically actuated trim system which varies the angle of incidence of the
horizontal stabilizer, and is operated via a trim control unit, from switches
mounted on the pilot's and copilotfs control wheels.
The pitch control system incorporates two parallel-connected PCUs for each
elevator, and a dual control system. Normally, both control systems are
interconnected via the control column transverse coupling shaft so that there
is simultaneous movement of both elevators, but i t is possible to isolate a
jammed circuit by means of a disconnect mechanism, thereby providing limited
pitch control (one elevator only) through the remaining circuit (refer to
paragraph I.A.). Anti-jam mechanisms are included in each of the PCU input rod
linkages.
Control column movement is transmitted to forward quadrants by push-pull rods.
From each forward quadrant, control signals are conveyed by cables along each
side of the fuselage below the cabin floor to the respective rear quadrant
mounted on the rear face of the vertical stabilizer forward spar. Ouput from
each rear quadrant is transmitted, by push-pull rods, to a gain change
mechanism and to an artificial feel unit.
From the gain change mechanism, the control signal is transmitted to the PCU
input tube via a load limiter, bellcranks and push-pull rods. The PCU input
tube incorporates one input and two identical output levers, and each of the
output levers is connected to the input/feedback linkage of the PCU,
transmitting the control signal to the control valve of the PCU. Operation of
the elevator PCUs is similar to that of the aileron PCUs (refer to
paragraph 2.).
Signal inputs from the AFCS are made through the rear quadrant of the left
elevator control system only. Therefore, should jamming of the left cable
circuit occur, the autopilot inputs would no longer be available to the
elevator system.
SECTION 10
Page 15
May 28/82
canadair
chaiiencjer
OPERATING MANUAL
A. Horizontal Stabilizer
The aircraft is trimmed in pitch by varying the horizontal stabilizer angle
of incidence. Trim commands from the pilot's or copilot's control wheel
switches, the AFCS and the SAS, processed by a trim control unit, operate
the electrically driven stabilizer actuator. In order to enhance the
longitudinal trim movement, the movement of the horizontal stabilizer is
accompanied throughout its range of operation by a degree of elevator
movement that alters the stabilizer/elevator camber. The geometry is such
that an elevator servo input is generated as the horizontal stabilizer is
moved, the servo input being sufficient to produce the required elevator
deflection.
The electrically driven screw actuator, located at the top of the vertical
stabilizer, varies the horizontal stabilizer angle of incidence. The
actuator is driven by two electric motors directly connected to the drive
train each containing a high and low speed winding. Manual trim commands
from the pitch trim switches on the control wheels produce a steady rate of
stabilizer movement of one-half degree per second. Depending on flap
position, the autopilot commands variable high or low trim rates of 0.1 to
0.5 degrees per second and 0.01 to 0.1 degrees per second respectively.
Mach trim commands produce a variable rate of stabilizer movement between
0.01 and 0.1 degrees per second. Each of the electric motors driving the
trim actuator is protected against overspeed by a dual coil brake.
The control unit, located in the avionics bay, controls the rate and
direction of movement of the actuator. The unit consists of two
independent channels and operates from two power busses so that electrical
failure on one bus does not preclude operation of the stabilizer trim. A
pilot reset capability allows channel transfer at the pilot's option.
The system normally operates on channel No. 1, with channel No. 2
performing only a monitoring function. Should a failure occur within a
controller channel or its associated motor, the control unit automatically
signals that the channel is inoperative and transfers to the backup
channel. In the event of an overspeed condition, the control unit removes
power from the drive motor, operates the brake in the actuator and provides
a shutdown signal to the pilot.
Two trim position sensors on the actuator send signals to the control
unit. One sensor supplies the AFCS with stabilizer angle data and the
second is connected to the flight recorder. Both position sensors provide
travel limit signals for the control unit. Stabilizer trim position is
also an input to the take-off configuration warning system. A third
position sensor, located on the stabilizer rear spar, supplies position
signals to the control surface trim position indicator on the centre
instrument panel.
SECTION 10
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A panel mounted on the centre pedestal has two ganged amber switch/lights,
CHAN 1 INOP and CHAN 2 INOP, that indicate failures in their respective
pitch trim control unit channel (refer to Figure 4). Normally, channel
No. 1 is engaged and both switch/lights are out. A combined
overspeed/channel change switch/light, OVSP CHANGE CHAN, is also located on
the panel. For test purposes, this switch/light can be used tc change the
system from one channel to the other.
Pitch trim is activated by first pressing the CHAN 1 INOP/CHAN 2 INOP
switch/lights and then the OVSP CHANGE CHAN switch/light on the centre
pedestal control panel. Commands from the pilot's trim switch override
those from the copilot's trim switch, the AFCS and the SAS. Commands from
the copilot's trim switch override only those from the AFCS and the SAS.
The pilots' control wheels each have a red disconnect button, PITCH TRIM
DISC, which can be pressed to remove power from the system and brake the
actuator (refer to Figure 6). Re-engagement of the trim system is
accomplished by again pressing the CHAN 1 INOP/CHAN 2 INOP switch/light and
then the OVSP CHANGE CHAN switch.
Control Columns
The pilot's and copilot's control columns each consist of a conventional
tubular column mounted vertically in a housing. A push-pull rod connected
at the rear of the column base transmits column movement to the pitch
control system. A control column shaker, which is a component part of the
stall protection system, is mounted on the column.
Gain Change Mechanisms
The two independent gain change mechanisms consist of bell cranks and
push-pull rods. The mechanisms are located side by side in the vertical
stabilizer between the rear quadrants and the PCU input linkages, and are
identical in form and function.
The purpose of the gain change mechanisms is to convert the rotational
input from the rear quadrants into linear output in such a way as to ensure
that the rate of elevator movement increases as the control column is moved
from neutral to provide the required control response.
Artificial Feel Mechanisms
Two artificial feel mechanisms, one for each elevator, provide the pilots
with positive feel of the power-operated systems and act as centering
devices for the systems. Each unit consists of a main feel cam and
follower, a primary spring box and a secondary spring box.
The primary and secondary spring boxes load the feel mechanism cam
follower, and feel rate is achieved by movement of the cam follower along
the cam profile. The load exerted by the primary spring box depends on the
position (angle of incidence) of the horizontal stabilizer which varies
according to manual or automatic trim commands. The secondary spring box
SECTION 10
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is designed so that its contribution to feel force is released when the
control column input exceeds a predetermined load, ensuring a reduced feel
force when rapid control column movement is required.
E. Anti-Jam Mechanisms
The elevator anti-jam mechanisms act normally as push/pull rods for the PCU
input rod linkages* If a PCU input linkage cannot move because of a jam in
the PCU, the mechanism breaks out to isolate the defective PCU from the
system. The other PCU continues to operate the affected elevator.
When the mechanism breaks out, a proximity sensor is deactivated and the
amber PITCH light on the SERVO MONITOR panel comes on (refer to Figure 3).
F. Elevator Position Transmitter
A position transmitter, located on the rear spar of the horizontal
stabilizer, transmits elevator position signals continuously, over the full
range of travel, to the control surface position indicator on the l e f t of
the centre instrument panel.
G. Elevator Flutter Dampers
Each elevator is protected against aerodynamic flutter by two flutter
damper assemblies. Each damper consists of a pre-charged hydraulic
cylinder containing a double-acting piston. Flutter damping occurs when
loads are placed on the piston rod forcing hydraulic fluid from one side of
the piston to the other through small-diameter passages. The dampers
connect with the elevator immediately outboard and inboard of the elevator
centre hinge. The hydraulic fluid level in each damper can be checked
through an integral sight gauge.
5. WING FLAP SYSTEM (Figures 7 and 8)
The flap system consists of externally hinged inboard and outboard doubleslotted
flap panels mounted on the trailing edge of each wing. The panels are
electrically driven by a power drive unit (PDU) located in the main landing
gear bay* The motor action of the PDU is translated to eight actuators, two to
each flap panel, by flexible shaft assemblies. An a symmetry/over speed detector
and brake unit is incorporated in each flap drive system.
The outboard flaps have fixed leading edge vanes and the inboard flaps have
movable leading edge vanes which automatically extend or retract as the flaps
are lowered or raised. Each of the three hinges of the outboard flaps
incorporates a spring actuator which is connected by a rod to a bent up
trailing edge (BUTE) door hinged on the lower surface of the wing. Rollers on
the BUTE doors are kept in contact with cam-shaped fittings attached to the
vane and, as the flaps are lowered or raised, the movement of the BUTE doors is
governed by the cam profile. When the flaps are fully down, the BUTE doors
take up a raised position to direct airflow over the vane and flap assemblies.
SECTION 10
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OPERATING MANUAL
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ROTARY
SWITCH
FLAP LEVER QUADRANT n
FLAP LEADING EDGE \
VANE (INBOARD)
Wing Flap System Components
Figure 7
SECTION 10
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The flaps are extended or retracted in response to command signals from by the
FLAPS control lever located on the centre pedestal. Flap position is set by
feel detents on the flaps control lever quadrant. Four positions are provided
corresponding to the following operating modes:
Flight/taxiing 0 degrees
Take-off 20 degrees
Approach 30 degrees
Landing 45 degrees
The signals are fed to the PDU via the flap control unit. If the control unit
logic detects an anomaly such as flap asymmetry or overspeed, power is removed,
causing the PDU motor brakes and the asymmetry/overspeed detector brakes to
stop the system. When a system fault is detected, a signal is transmitted to
the warning system and a flap fail amber light, FLAPS FAIL, comes on above the
flap position indicator located on the copilot's instrument panel.
A. Flap Control Unit
The flap control unit (FCU) is located in the underfloor avionics bay and
is powered from dc bus No. 1 and dc bus No. 2. Although two power supplies
are provided, only one is necessary to operate the unit. The function of
the unit is to assess the flap extend/retract commands received from the
FLAPS control lever and provide the correct activating signal to the PDU.
Once a selected flap angle is reached, the flaps are locked in position by
the PDU motor brakes and the asymmetry/overspeed detector brake units.
The FCU also signals the aural warning unit (refer to Section 3) to
initiate the following aural warnings:
A wailer, when the airspeed is too great for the flap position
selected.
An intermittent horn, when the take-off configuration is incorrect
(aircraft on ground, either throttle lever set beyond HIGH IDLE and
flaps set to any position other than 20 degrees).
A horn, when flaps are set to more than 30 degrees with the landing
gear up.
B. Power Drive Unit
The PDU, located in the main landing gear bay at the aircraft centreline,
has a dual output shaft which is coupled to the left and right side flap
flexible drives. The two PDU motors are coupled to a mechanical
differential which drives the output shaft through a clutch and an output
gear train. With power applied to the PDU, the motor brakes are released
and the motor shafts rotate to drive the internal gear train to provide
SECTION 10
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PDU MOTOR OVERHEAT LIGHTS
Amber light comes on to indicate »n overheat
condition in the associated POU motor
FLAP FAIL LIGHT
Amber light comes on to indicate a flap
asymmetry or speed response fault.
FLAP CONTROL LEVEL
Lever is guarded to its full height to obviate
inadvertent operation.
Lever quadrant is marked with the four flight
modes:
Flight/Taxiing 0 degrees
Take-off 20 degrees
Approach 30 degrees
Landing 45 degrees
each mode corresponding with a detented
position of the lever.
FLAPS
FAIL
1 OVHT I
MOT 1
OVHT j
1 MOT 2 |i
FLAP POSITION INDICATOR
Provides a continuous angular indication of
the flaps over their operating range.
Wing Flap Controls and Indication SECTION 10
Figure 8 Page 22
May 28/82
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OPERATING MANUAL
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driving torque for the flexible shaft assemblies and actuators. When the
selected flap position is reached, the FCU responds to a PDU potentiometer
signal and opens a relay to de-energize the motors and apply the motor
brakes.
If power to one of the PDU motors fails, the associated brake is
automatically applied, locking PDU input to the differential and the second
motor continues to operate the system at half speed. In the event of
overheating of a PDU motor, thermal switches de-energize the applicable
motor and an amber overheat light, OVHT MOT 1 or OVHT MOT 2, comes on above
the flap position indicator. The thermal switches reset once the overheat
condition has passed.
Flap Actuators
Eight flap actuators are located on the flap hinge attachment fittings, two
actuators to each flap. The actuators are of the linear mechanical type
and consist of a housing assembly, worm and helical gears and a ball screw
assembly with a ball nut extension tube. Adapters for attachment of the
flexible shafts are provided. The flap is connected to a gimbdl block
attached to the ball nut. Two different gear reductions are used to
achieve a uniform movement of both flaps with respect to the swept wing
configuration.
Flexible Shaft Drive Assemblies
The flap drives, one to each wing, are located along the rear and auxiliary
spars in the wing trailing edge. The drives are in the form of two
flexible shafts, each made up of five segments connected from the PDU to
the four actuators and to the asymmetry/overspeed detector and brake
assembly in each wing.
Asymmetry/Overspeed Detector and Brake Assemblies
These components are located adjacent to the rear spar between the outboard
flap and aileron of each wing and are coupled by a segment of the flexible
shafts to the outboard actuators. The function of these assemblies is to
transmit signals to the FCU to provide positive braking action to the flaps
in the event of asymmetric movement of the left and right flaps, or
overspeed.
Flap Position Transmitters
Two flap position transmitters are provided, one each on the left and right
inboard flap assemblies at the respective inboard hinge boxes. Both
transmitters send flap position signals to the stall protection system
computer (refer to paragraph 7.B.).
A flap position potentiometer is contained in the right flap position
transmitter only. The potentiometer transmits flap position signals to the
flap position indicator on the copilot's instrument panel (refer to
Figure 8).
SECTION 10
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GROUND SPOILER DEPLOYED INDICATION
Amber lights come on when ground spoilers are at any
position other than fully retracted. On aircraft that do not
incorporate Canadair Service Bulletin 600-0368, lights do not
function if GROUND SPOILER spoiler switch is OFF.
FLIGHT SPOILER DEPLOYED INDICATION
Amber lights come on steady when flight spoilers are at any
position other than fully retracted. Lights come on flashing
and aural warning horn sounds if flight spoilers are deployed
and either engine is operating at an N1 rpm above 79%.
On aircraft not incorporating Canadair Service Bulletin
600-0385, lights come on steady only, and only when flight
spoilers are at any position other than fully retracted.
SPOILER CONTROL LEVER
To deploy flight spoilers lever may be moved rearwards to
any one of eight detented positions according to flight path
requirement until MAX position stop gate is reached. On aircraft
that do not incorporate Canadair Service Bulletin 600-0452,
ground spoilers are selected by pressing the button on top
of lever then lifting lever over stop gate to EXTEND position.
FLIGHT SPOILERS LEFT AND RIGHT INDICATION
Green lights come on when flight spoilers are extended
beyond one-half position.
GROUND SPOILERS SWITCH
Three position toggle switch.
ON - Ground spoilers are armed and deploy if deploy
conditions are met (refer to paragraph 6.B.).
OFF - Ground spoilers are disarmed and cannot be deployed.
TEST - LH and RH GND SPLR and SPLR INOP lights come
on to indicate correct operation of ground spoiler control
system.
GROUND SPOILER INOP LIGHT
Amber light comes on if spoiler control unit detects fault
in ground spoiler hydraulic selector valves.
SB 600-0452
NOT INCORPORATED
©
Spoiler Controls and
Figure 9
Indication SECTION 10
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OPERATING MANUAL
PSP 606
SPOILER SYSTEM (Figure 9)
Wing lift modulation is achieved by the operation of flight and ground
spoilers. The flight spoilers may be extended to any position, between 0 and
MAX FULL, as required for the intended flight path. The ground spoilers have
only two positions, fully retracted during flight or fully deployed when the
aircraft touches down, to assist other braking systems by dumping lift and
increasing drag.
A. Flight Spoilers
The flight spoilers are two hydraulically powered panels, one hinged to
each wing trailing edge upper surface, in the area of the outboard flaps,
and are controlled mechanically through pilot movement of a lever on the
centre pedestal. Each panel is powered by two hydraulically independent
PCUs secured to the wing auxiliary spar. Each PCU is independently
connected to its spoiler and is capable of spoiler operation should the
adjacent PCU fail either mechanically or hydraulically.
The spoiler control lever is connected to the PCUs via a push-pull rod,
pulley, cable and lever system. The spoilers are fully retracted when the
lever is in the fully forward position; this provides natural control,
similar to throttle lever movement, in that rearward movement of the
spoiler control lever deploys the flight spoilers and retards the aircraft.
Lever positions, when selected, are held by a serrated plate and plunger
mechanism. The lever is gated at the fully deployed position to prevent
inadvertent movement into the ground spoiler arming area.
A position transmitter, located inboard of the PCUs of each flight spoiler,
transmits position signals to the control surface position indicator on the
centre instrument panel. A proximity sensor switch, located between the
PCUs at each spoiler, senses spoiler position (retracted or extended) and
transmits a signal to amber LH FLT SPLR and RH FLT SPLR lights on the
glareshield which come on when the spoilers are not in the fully retracted
position. On airplanes which incorporate Canadair Service Bulletin
600-0385, the LH FLT SPLR and RH FLT SPLR lights come on flashing and the
take-off configuration aural warning horn sounds if the flight spoilers are
deployed and either engine is operating at an Nl rpm above 79%.
A detent mechanism on both of the spoiler wing circuits prevents
unacceptable spoiler asymmetry if a controlex cable disconnects. If a
cable disconnect occurs, the detent mechanism closes the affected spoiler
when the spoilers are less than one-half extended or retracts it to the
one-half extended position when the spoilers are more than one-half
extended.
Microswitches on each detent mechanism cause the LEFT and RIGHT FLIGHT
SPOILERS lights on the centre pedestal to come on when the flight spoilers
are more than one-half extended. Operation of the lights indicates that
the flight spoiler detent mechanism is serviceable and that blowback
protection in an asymmetrical spoiler condition has been reset to the
one-half extended position.
SECTION 10
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OPERATING MANUAL
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NOTE
^ 1 \ On airplanes 1051 and subsequent
which have this panel, the power
switch is placarded PUSHER.
PILOT'S STALL PROTECTION TEST PANEL
I STALL I
11 PUSH I
»1 STICK
^ INDICA
(IF INS"
PUSHER
TORS
TALLED)
C
ALT
COMP
FAIL
• ALTIT
w SYSTI
LIGHT
UDE COMPENSATION
EM FAILURE WARNING
S (IF INSTALLED)
<
STALL
PROTECT
FAIL
*M STALl
^ FAILU
LIGHT
STALL
PROTECTION
PUSHER 0N
V—''OFF
G SWITCH
TEST TEST
( S ) OFF (©)
OR
STALL
PROTECTION
G SWITCH
TEST TEST
(H) OFF (S) OFF
SPS TEST INDICATORS COPILOT'S STALL PROTECTION TEST PANEL
Stall Protection System Controls and Indication
Figure 10
SECTION 10
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B. Ground Spoilers
The ground spoilers are two hydraulically powered panels, one hinged to the
wing trailing edge upper surface each side, in the area of the inboard
flaps, and are controlled electrically. Each panel is powered by one
actuator secured to the wing auxiliary spar.
On aircraft incorporating Canadair SB 600-0452,
the ground spoilers deploy automatically if a weight-on-wheels or wheel
spin-up signal is present, the GROUND SPOILERS switch is in the ON position
and either of the following two conditions have been met:
The spoiler control lever is at the 0 position or between the 0 and
1/4 positions and both throttle levers have been advanced above HI
IDLE and then returned to the HI IDLE position or lower.
The spoiler control lever is between the 1/4 and MAX positions and
both throttle levers are at the HI IDLE position or lower.
A spoiler control unit in the underfloor avionic bay monitors
weight-on-wheels and wheel spin-up signals, throttle lever position, GROUND
SPOILERS switch position and the position of the two valves in the dual
hydraulic selector valve. When all of the conditions for ground spoiler
deployment have been met, the control unit energizes solenoids on the
hydraulic selector valves. The valves then open, hydraulic pressure is
applied at the ground spoiler actuators, the actuators unlock and the
spoilers are powered to the extended position. If the spoiler control unit
detects a difference in the positions of the hydraulic selector valves, the
SPLRS INOP light comes on, electrical power is removed from the selector
valve solenoids and the ground spoilers, if extended, close and lock. The
SPLRS INOP light also comes on and the ground spoilers retract if the
throttle levers are set to difference positions after the ground spoilers
deploy.
LH and RH GND SPLR lights on the glareshield come on when a proximity
switch near the associated spoiler centre hinge senses that the spoiler is
at any position other than fully closed.
The ground spoiler system is tested by setting the spoiler control lever to
0 and the GROUND SPOILERS switch to TEST. After a 2 second delay, the
following indications verify that the system is operating correctly:
The LH and RH GND SPLR lights come on for 4 seconds.
The SPLRS INOP light comes on immediately the LH and RH GND SPLR
lights go out.
All lights go out when the GROUND SPOILERS switch is moved from the
TEST position.
SECTION 10
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STALL
PUSH
O
STALL
PROTECT
FAIL
©
COMP
FAIL
RED
SECTOR YELLOW
SECTOR
STALL/PUSH LIGHTS (IF INSTALLED)
Red lights flash when angle of attack reaches stick pusher trip point.
STALL PROTECT FAIL WARNING LIGHTS
Red warning lights flash in the following cases:
—To indicate a system fault (refer to paragraph 7JLfor fault conditions).
—Whenever one of the AP/SP DISC buttons on the control wheels is
pressed.
—During system test.
Lights come on steady when power is removed from system.
ALT COMP FAIL LIGHTS
BLUE
SECTOR
Red lights come on if one or both altitude
signals to SPS computer are lost or if 2000
foot difference between them is detected.
15,000 foot angle of attack trip points are
applicable when lights are on.
SPS TEST INDICATORS
Colored sectors on indicator provide references
for stall warning/stick pusher sequence during
system test (refer to paragraph I.). Indicator is not
calibrated to provide in-flight angle of attack
indication or approach speed reference.
Stall Protection System Indicators
Figure 11
SECTION 10
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OPERATING MANUAL
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I On aircraft that do not incorporate Canadair SB 600-0452,
ground spoiler operation occurs only when all the following conditions
exist:
The GROUND SPOILERS switch is in the ON position.
The spoiler control lever is moved up and rearward through the stop
gate to the EXTEND position.
The left throttle lever is in the HIGH IDLE position or lower.
A weight-on-wheels (WOW) or spin-up signal is present.
A spoiler control unit located in the underfloor avionics bay receives
signals from the control lever, throttle lever and landing gear switches
and, when a signal is received concurrently from all three sources, the
control unit transmits a signal to the ground spoiler manifold solenoid
valves causing actuator operation for spoiler deployment.
A proximity sensor switch, located near the centre hinge, senses spciler
position (retracted or extended) and transmits a signal to the spoiler
control unit which, in turn, transmits a signal to amber LH GND SPLR and
RH GND SPLR lights on the glareshield which come on when the ground
spoilers are deployed. A SPLRS INOP amber light on the centre pedestal,
adjacent to the control lever, comes on if the ground spoilers fail to
deploy (refer to Figure 9).
The ground spoiler control system is tested by setting the GROUND SPOILERS
switch to TEST. After a 2 second delay the following indications verify
that the system is operating correctly:
The LH and RH GND SPLR lights come on for 3 seconds.
The SPLRS INOP light comes on immediately the LH and RH GND SPLR
lights go out.
All lights go out when the GROUND SPOILERS switch is moved from the
TEST position.
7. STALL PROTECTION SYSTEM (Figures 10, 11 and 12)
The stall protection system senses the aircraft angle of attack, provides the
flight crew with a visual and tactile warning of an impending stall and, if no
corrective action is taken, prevents flight into the stalled condition by
activating a stick pusher mechanism. The system consists of the following
principal components:
Two trailing vane type angle-of-attack transducers
A dual-channel stall protection computer
SECTION 10
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Two altitude transducers
Two lateral accelerometers
Two flap position transmitters
Two stick shakers
A stick pusher subsystem
Stall protection system test indicators
System warning lights and test switches
An aural warning horn (warbler)
NOTE: Some aircraft are fitted with a version of the stall protection system
that does not have the altitude compensation feature. These aircraft do
not have altitude transducers or altitude compensation failure warning
lights and have stall margin indicators fitted in place of stall
protection system test indicators (refer to paragraph J.).
When a dangerously high angle of attack is approached, the stall protection
computer applies continuous ignition to the engines and, if the angle of attack
is increased, activates the stick shakers to generate a stall warning in the
form of a mechanical vibration of the control columns. If the stall warning
occurs when the flaps are at the 45-degree position, an additional stall
warning is provided by the FAST/SLOW pointers on the pilot's and copilot's
attitude director indicators (refer to SECTION 11, FLIGHT INSTRUMENTS). If the
aircraft angle of attack continues to increase to the stick pusher trip point,
the aural warning horn sounds and the stick pusher subsystem forces the control
columns forward to effect recovery from the impending stall. When the aircraft
angle of attack has decreased to a preset point below the pusher trip point,
the aural warning horn stops and the stick pusher is deactivated. The stick
shakers and continuous ignition switch off automatically when the aircraft
angle of attack decreases through their respective trip points.
If installed, the red STALL/PUSH lights flash whenever the aural warning horn
and stick pusher are operating (refer to Figure 11).
If the autopilot is engaged when the aircraft approaches the stall, it is
automatically disengaged on a signal from the stall protection computer when
the aircraft angle of attack reaches the stick shaker trip point.
A. Angle-of-Attack Transducers
There are two angle-of-attack transducers, one on each side of the forward
fuselage. Each transducer consists of an externally mounted trailing vane
assembly connected by a shaft to an internally mounted potentiometer. The
trailing vane is calibrated in terms of slipstream angle of attack around a
fuselage datum and, as it is moved around this datum by the local airflow,
the transducer potentiometer produces a dc electrical signal, the voltage
of which varies in proportion to the aircraft angle of attack. The signals
SECTION 10
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from the left and right angle-of-attack transducers are transmitted,
respectively, to the left and right channels of the stall protection
computer.
The transducer trailing vanes are protected against ice by built-in heater
elements controlled from the ADS heater control panel (refer to SECTION 14,
ICE/RAIN PROTECTION).
Stall Protection Computer
The stall protection computer, located on the left side of the underfloor
avionics bay, is divided into two identical and independent (left and right)
channels. Each channel uses inputs from its associated angle of attack
transducer, altitude transducer, lateral accelerometer and flap position
transmitter to compute angle-of-attack trip points for auto-ignition, stick
shaker operation, aural warning and stick push. If the angle of attack
increases at a rate greater than one degree per second, the computer lowers
the angle-of-attack trip points for the various system functions. This
action prevents the aircraft's momentum in the pitching plane from carrying
it through the stall warning/stick pusher sequence into the stall.
The two altitude transducers are located in the avionics bay under the
flight compartment and provide altitude signals to the associated left and
right sides of the stall protection computer. The transducers are
connected to the left and right static systems via static source selectors
on the pilot's and copilot's side panels (refer to SECTION 11, FLIGHT
INSTRUMENTS for details of the pi tot/static system).
As the altitude transducers signal an increase in altitude between 2000 and
15,000 feet, the computer progressively lowers the angle-of-attack trip
points for the stick shaker and pusher. Below 2000 feet and above
15,000 feet, the trip points are constant. If one or both of the altitude
signals is lost or if the difference between signals exceeds 2000 feet, the
computer automatically applies the trip points associated with the 15,000
foot altitude and the ALT COMP FAIL lights on the glareshield come on.
The two lateral accelerometers in the underfloor avionic bay monitor skid
or sideslip and signal the corresponding channel of the computer. Each of
the computer channels uses the signals to generate compensated
angle-of-attack values produced by manoeuvres involving skid or sideslip.
The compensated angles insure that adequate stall protection is provided
during uncoordinated flight. The trip points are also lowered
progressively, on signals from the two flap position transmitters, as the
flaps move through the 0-, 20-, 30- and 45-degree positions. If one or
both of the flap position signals are lost, the computer automatically
applies the stick shaker, continuous ignition and stick pusher trip points
associated with the next higher flap setting.
The weight-on-wheels inputs from the landing gear control unit enable the
computer to disable the stick shakers and pusher and the system failure
warning lights while the aircraft is on the ground, except during system
SECTION 10
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OPERATING MANUAL
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test. If there is a failure in the weight-on-wheels signal to one of the
computer channels, the flashing STALL PROTECT FAIL light associated with
the channel comes on.
To prevent inadvertent operation of the stick pusher due to a failure in
one of the computer channels, the computer does not command a stick push
unless both of the computer channels signal a stick push simultaneously.
Each of the computer channels transmits an angle-of-attack signal to the
associated attitude director indicator to drive the instrument's speed
command display when the flaps are at 45 degrees. In-flight gust filtering
is provided by the system.
Stall Protection System Monitoring
The stall protection computer monitors the operation of the system for
possible mechanical defects in the angle-of-attack transducers and faults
in the electrical circuitry between the transducers, the lateral
accelerometers and the computer.
The computer compares the sideslip compensated signals from its left and
right channels and, if they are found to differ in magnitude by a preset
amount (within preset lateral acceleration limits), both of the flashing
STALL PROTECT FAIL lights come on. The computer also causes the lights to
come on if it detects a difference in the signals from the lateral
accelerometers.
Stick Shakers
There are two stick shakers, one each on the pilot's and the copilot's
control columns. Each shaker is a dc electric motor driving an eccentric
weight. The shakers operate independently of each other and are powered by
signals from the stick shaker circuits of their respective stall protection
computer channels. When the aircraft angle of attack reaches the shaker
trip point, the shaker vibration starts and, if the angle of attack
continues to increase, becomes a continuous vibration at the stick pusher
trip point. The noise of the stick shakers as they are operating is
sufficiently loud to constitute an aural warning of shaker operation.
Stick Pusher Subsystem
The stick pusher consists of a rotary actuator driven by a dc electric
motor which operates a capstan connected by cables to the right elevator
control quadrant. The pusher has an electronic control box with logic
circuits so arranged that pusher signals must be transmitted simultaneously
from both channels of the stall protection computer before a stick push can
be initiated. If this condition is met, the stick push signals are
amplified and sent via the pusher main power switch to the motor in the
rotary actuator. At the same time, the signal from the right channel of
the stall protection computer energizes the solenoid on an electromagnetic
clutch in the motor drive. The clutch allows the motor to drive the
capstan through a torque limiter so that an 80-pound forward push is
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exerted on the control columns. If installed, the red STALL/PUSH lights
flash whenever the stall protection system computer commands a stick push.
In order to prevent the aircraft from flying into a low or negative-g
condition during the stick push, two accelerometer switches in series with
the clutch of the rotary actuator motor disconnect the pusher drive, if the
aircraft reaches 0-5 g during the pitching manoeuvre induced by the stick
push. One of these switches can be tested using the G SWITCH TEST switch
on the copilot's facia panel (refer to paragraph 7.1.).
At any time, the pilot or copilot can stop the stick pusher and disconnect
the autopilot by pressing and holding the AP/SP DISC switch installed on
the left horn of each control wheel (refer to Figure 6). The stick pusher
is capable of operating immediately when the switch is released. The stick
pusher can be deactivated by opening the system circuit breakers on the
battery bus and dc essential bus circuit breaker panels or by setting
either the pilot's or copilot's PUSHER switch to OFF.
The stick pusher electronic control box contains monitoring circuits
capable of detecting a failure in the pusher circuits, failure of the
pusher power supplies or power amplifiers and loss of either of the signals
from the two channels of the stall protection computer. If any of these
failures occurs, the monitoring circuits cause the flashing STALL PROTECT
FAIL lights to come on. To prevent spurious warnings, the warning signals
are subject to a 3-second delay before they can generate a flight
compartment warning.
On FAA certified aircraft, POWER and PUSHER toggle switches are located,
respectively, on the pilot1s and copilot's STALL PROTECTION panels. Power
is supplied to the stick pusher system (from the battery bus) only when
both of the switches are set to ON (refer to Figure 14).
Stall Protection System Test Indicators
Two stall protection test indicators, located on the pilot's and copilot's
side panels, are driven by angle of attack signals from the left and right
channels of the stall protection computer. Each instrument provides a
reference indication during testing of the stall protection system but is
not calibrated for secondary use as an angle-of-attack indicator or
approach speed reference (refer to paragraph I. for details of system
testing). The indicators have a 5-volt ac lighting system which is part of
the integral lighting system.
Aural Warning Horn
The stall protection system aural warning horn sounds whenever one of the
two channels of the stall protection computer signals that the aircraft
angle of attack has reached the stick pusher trip point. Normally, the
sounding of the horn warns the flight crew of the stick pusher operation
but, if there is a loss of the signal from one of the channels of the stall
protection computer, the sounding of the horn indicates that flight crew
action is required to avoid the stall.
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OPERATING MANUAL
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PILOTS STALL PROTECTION TEST SWITCH
Spring-loaded toggle switch. Holding switch on activates
self-testing of stall protection system. During test, simulated
approach to stall is observed as pointer of left SPS TEST
INDICATOR moves from the blue to the red sector. If flaps
are set at 45 degrees, speed command pointer on associated
ADI moves from fast to slow, reaching full scale deflection at
stick shaker trip point. Stick pusher can be checked only whe
pilot's and copilot's TEST switches are held on
simultaneously.
NOTE
Stick pusher can only be tested on the ground; all other tests can
be conducted on the ground or in flight.
COPILOTS STALL PROTECTION TEST SWITCH
Spring-loaded toggle switch. Holding switch on activates
test of right side of system. Test is identical to test of left side
of system activated by pilot's TEST switch except that right
stick shaker operates and ALT COMP FAIL lights do not
come on.
G SWITCH TEST SWITCH
Spring-loaded toggle switch tests operation of one of the
accelerometer switches on stick pusher actuator. During
stick pusher test, correct operation of accelerometer switch
is indicated if stick pusher is immediately de-energized when
G SWITCH TEST switch is set to TEST.
STALL
^ PROTECTION ^
G SWITCH
TEST TEST
^
Stall Protection System Panels SECTION 10
Figure 12 (Sheet 1) Page 34
Mar 01/85
OPERATING MANUAL
PSP 606
STICK PUSHER SYSTEM SWITCHES
Two position toggle switches wired in series between stick
pusher actuator and battery bus. When both switches are set
to ON, power is available for stick pusher operation.
If one switch is OFF, stick pusher cannot operate and both
STALL PROTECTION FAIL lights come on steady.
NOTE
On some airplanes fitted with pusher system switches, the
pilot's switch is placarded PUSHER.
»"^riSr*r*u
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Failure Warning Lights
Two red STALL PROTECT FAIL lights are located on the left and right sides
of the glareshield in the flight compartment. The lights come on flashing
to warn the flight crew of any of the following system faults:
Loss of power to the stall protection computer. If only one channel
is affected, only the light associated with that channel comes on.
Failure of the stick shaker or stick pusher circuits in the stall
protection computer
Failure of one or both of the angle-of-attack transducers
A difference in the skid and sideslip compensated signals from the
angle-of-attack transducers (refer to paragraph 7.C.)
A difference in the signals from the lateral accelerometers
(refer to paragraph 7.C.)
Loss of the weight-on-wheels signal to the stall protection computer
in the flight mode
A failure in the electrical circuits of the stick pusher subsystem.
The lights also come on whenever one of the AP/SP DISC buttons on the
control wheels is pressed, and during the system test (refer to paragraph
7.1.). The lights come on steady whenever the system circuit breakers on
the battery bus and dc essential bus circuit breaker panels are opened.
System Test Switches (Figure 12)
Two spring-loaded toggle switches placarded STALL PROTECTION TEST and
located, respectively, on the pilot's and copilot's facia panels are used
to activate the self-test feature of the stall protection system. An
additional switch on the copilot's facia panel, G SWITCH TEST, is used to
test one of the acceleration switches in the stick pusher subsystem.
If the pilot's STALL PROTECTION TEST switch is held to TEST, the correct
operation of the system is indicated by the following sequence:
The two ALT COMP FAIL lights come on steady and remain on for the
entire test sequence.
The pointer of the pilot's SPS TEST INDICATOR first moves clockwise
then counterclockwise into the blue sector.
The two STALL PROTECT FAIL lights come on flashing when the pointer of
the SPS TEST INDICATOR starts moving clockwise. During the test
sequence the lights go out briefly then come on again flashing.
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STALL MARGIN INDICATORS
Each indicator presents visual display of margin available between aircraft's
actual speed and stall speed. Left and right indicators are driven by signals
from left and right channels of stall protection computer respectively.
—Cruise sector (green): Indicates stall margin available in cruising flight.
— Green and yellow sector: Indicates small margin, compensated for flap
angle, available when aircraft is maintaining 1.3 V3.
—Slow sector (yellow): Indicates that aircraft has assumed high angle of
attack and that dangerously low stall marain is available for continued flight.
Continuous ignition starts when pointer reaches upper half of sector.
—Stall warning sector (red and black): Stick shaker activated when pointer
reaches right edge of sector.
—Stall sector (red): Stick pusher activated when pointer reaches this
sector.
Stall Margin Indicators SECTION 10
Figure 13 Page 37
Mar 01/85
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OPERATING MANUAL
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Continuous ignition starts when the pointer of the SPS TEST INDICATOR
is within the blue sector*
The left stick shaker operates and the autopilot disconnect (AFCS)
lights come on steady when the pointer of the SPS TEST INDICATOR is
within the yellow sector.
The aural warning horn sounds and the STALL/PUSH lights (if installed)
come on flashing when the pointer of the SPS TEST INDICATOR is within
the red sector.
The aural warning horn, the stick shaker and continuous ignition stop
operating when the STALL PROTECTION TEST switch is released.
When the copilot's STALL PROTECTION TEST switch is held to test, the test
sequence is the same for the right side of the system except that the right
stick shaker operates and the ALT COMP FAIL lights remain out.
NOTE: If the flaps are set at the 45-degree position during the test, the
correct functioning of the FAST/SLOW stall warning on the pilot's
and copilot's attitude director indicators can be tested.
Testing of the stick pusher and the acceleration switch on the stick pusher
subsystem is carried out by holding both of the STALL PROTECTION TEST
switches on simultaneously to operate both channels of the stall protection
computer. The stick pusher operates when the pointers of the SPS TEST
indicators reach the red sector. Once the stick push has occurred, the
operation of the accelerometer switch in the stick pusher subsystem can be
checked by using the G SWITCH TEST switch. Correct operation of the
accelerometer switch is indicated if the stick pusher is immediately
de-energized and the control columns return to the neutral position when
the switch is set to the ON position.
The stick pusher can be tested only when the aircraft is on the ground; all
of the other tests described above can be carried out on the ground or in
flight.
Systems Without Altitude Compensation (Figure 13)
Some aircraft are equipped with a stall protection system that does not
have the altitude compensation feature. In these systems, the altitude
transducers, associated static source selectors and ALT COMP FAIL lights
are not installed. In addition, stall margin indicators are installed in
place of the SPS TEST indicators.
The operation of the system is identical to that described in the preceding
paragraphs except that altitude compensated angle-of-attack trip points are
not computed. The system test sequence using the stall margin indicators
is as follows:
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If the pi lot1s STALL PROTECTION TEST switch is held to TEST, the correct
operation of the system is indicated by the following sequence:
The two STALL PROTECT FAIL lights come on flashing momentarily, go
out, then come on flashing again.
The pointer of the left stall margin indicator moves from the CRUISE
to the STALL sector.
Continuous ignition starts on both engines when the pointer of the
pilotfs stall margin indicator passes through the upper half of the
yellow (SLOW) sector of the indicator.
The left stick shaker is activated when the pointer of the stall
margin indicator reaches the edge of the red and black stall warning
sector.
The stall protection aural warning horn sounds when the pointer of
stall margin indicator reaches the STALL sector and the STALL/PUSH
lights (if installed) on the glareshield come on.
the
When the switch is released, the stick shaker, aural warning, continuous
ignition and STALL/PUSH lights stop operating immediately and the pointer
of the pilot's stall margin indicator returns to the CRUISE sector. When
the copilot's STALL PROTECTION TEST switch is held on, the results are
identical for the right side of the stall protection system, except that
the STALL PROTECT FAIL lights come on only when the pointer of the
copilot's stall margin indicator is at the clockwise limit of its travel in
the green CRUISE sector, and when it reaches the red STALL sector.
NOTE: If the flaps are set at the 45-degree position during the test, the
correct functioning of the FAST/SLOW stall warning on the pilot's
and copilot's attitude director indicators can be checked when the
test switches are used.
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