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OPERATING MANUAL
PSP 601A-6
SECTION 10
FLIGHT CONTROLS
TABLE OF CONTENTS
GENERAL
Page
A. Control Disconnect Systems 1
B. Power Control Units 1
C. Artificial Feel Mechanisms 2
D. Trim Systems 2
E. Control Surface, Trim and Flap Position Indicators 2
F. Gust Locks 2
ROLL CONTROL SYSTEM 3
A. Aileron Trim 3
B. Aileron Control Wheels 3
C. Artificial Feel Mechanisms 3
YAW CONTROL SYSTEM 4
A. Rudder Trim 4
B. Rudder Pedal Assemblies 4
C. Anti-Jam Mechanisms 4
D. Artificial Feel Mechanisms 4
PITCH CONTROL SYSTEM 5
A. Pitch Trim 5
B. Control Columns 6
C. Gain Change Mechanisms 6
D. Artificial Feel Mechanisms 6
E. Anti-Jam Mechanisms 6
WING FLAP SYSTEM 7
A, Flap Control Unit 7
B- Power Drive Unit 7
C. Asymmetry/Overspeed Detector and Brake Assemblies 8
10-CONTENTS
Page 1
Apr 02/87
OPERATING HANUU.
PSP 601A-6
Page
6. SPOILER SYSTEM 8
A. Flight Spoilers 8
B. Ground Spoilers 9
7. STALL PROTECTION SYSTEM 9
A. Angle-of-Attack Transducers 10
B. Stall Protection Computer 10
C. Stall Protection System Monitoring 11
D. Stick Shakers 11
E. Stick Pusher Sub-system 12
LIST OF ILLUSTRATIONS
Figure
Number Title Page
1 Flight Controls 13
2 Control Disconnect T-Handles 14
3 Flight Controls - Hydraulics 15
4 Control Surface Position Indicator and Servo Monitor Lights 16
5 Trim Controls and Trim Position Indicators 17
6 Control Wheel 18
7 Wing Flap Controls and Indicators 19
8 Spoiler Controls and Indicators 20
9 Stall Protection System Controls and Indicators 21
10-
Page
Apr
OPERATING MAMUAL
PS? 601A-6
SECTION 10
FLIGHT CONTROLS
1. GENERAL (Figures 1 and 3)
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 totaViy 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 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. 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. 1 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 to drive hydraulic
system No. 3.
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 (Figure 2)
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 mechanisms are
operated by the PITCH DISC and ROLL DISC T-handles on the centre pedestal.
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.
B. 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.
SECTION 10
Page 1
Apr 02/87
cttaueneter
OPERATING MANUAL
PSP 60U-6
Each PCU consists mainly of a control-valve-operated piston moving in a
cylinder.
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 pilots1 controls or from the automatic flight control
system (AFCS). A flight control monitoring unit monitors the operation of
the PCUs. The flight control monitoring unit receives inputs from PCU
proximity sensors and transmits warning signals to the servo monitor panel
in the flight compartment.
Artificial Feel Mechanisms
Because the primary flight control surfaces are fully power-operated,
artificial feel mechanisms, consisting of spring devices, are incorporated
in the control systems to simulate aerodynamic forces and provide a means
of sensing control loads under various flight conditions.
Trim Systems (Figures 5 and 6)
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 is 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.
Control Surface, Trim and Flap Position Indicators (Figures 4, 5 and 7)
Flight control surface positions and trim angles are displayed on
indicators located on the centre instrument panel. 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.
Gust Locks
Gust locking of the ailerons, rudder and elevators is provided by trapping
hydraulic fluid within the PCUs whenever hydraulic pressure is removed from
the PCUs. This arrangement locks the control surface against the effect of
gusts but permits restricted movement of the surface, if a sufficiently
large external force is applied continously.
SECTION 10
Page 2
Apr 02/87
OPERATING MANUAL
PSP 601A-6
2. ROLL CONTROL SYSTEM
Roll (lateral) control is achieved by hydraulically powered ailerons which are
controlled primarily from conventional column-mounted horn-type 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 so that
there is simultaneous movement of both ailerons; but i t is possible to isolate
a jammed aileron control circuit by means of a disconnect mechanism, thereby
allowing limited control (one aileron only) through the unjamrned circuit (refer
to Figures 1 and 2).
Control wheel movement is transmitted by cables and pulleys which incorporate
an a r t i f i c i a l feel unit to the PCUs located outboard in the wing, forward of
the rear spar.
Each PCU actuator is capable of aileron operation should there be a failure
associated with the adjacent actuator.
Signal inputs from the AFCS are made through 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 ).
A. Aileron Trim
An electrically driven actuator applies a bias to the primary control
circuit, when required, by operation of the AIL TRIM switches located on
the centre pedestal. The amount of trim applied to the ailerons is shown
on the control surface trim position indicator.
B. Aileron Control Wheels (Figure 6)
The aileron control wheels are horn-type handwheels, spline-mounted on 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 a r t i f i c i a l feel mechanisms provide the pilots with positive feel of the
power-operated control system and act as centering devices.
SECTION 10
Page 3
Apr 02/87
ctianenQer
OPERATING KMftJAL
PSP 601A-6
3. YAW CONTROL SYSTEM
Yaw (directional) control is achieved by a hydraulically powered rudder,
controlled primarily from conventional dual, cross-coupled 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.
Pedal assembly movement is transmitted by cables and pulleys which include
artificial feel mechanisms, load limiters and a trim mixing system.
In addition to control inputs from the pedal assembly, inputs from the
stability augmentation system of the AFCS are applied to the system through two
yaw dampers in the trim mixing system (refer to Section 4).
A. Rudder Trim
An electrically driven actuator 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.
B. Rudder Pedal Assemblies
Conventional rudder pedal assemblies enable foot control of the aircraft
wheel brake system and the rudder control system.
Each set of pedals is provided with a hand-operated adjusting mechanism to
cater to the individual requirements of pilots.
C. Anti-Jam Mechanisms
The two forward anti-jam mechanisms operate to nullify the effects of a
jammed cable circuit and 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.
D. Artificial Feel Mechanisms
Two artificial feel mechanisms provide the pilots with positive feel of the
power-operated system and act as a centering device for the system.
SECTION 10
Page 4
Apr 02/87
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OPERATING MANUAL
PS? 601A-6
4. PITCH CONTROL SYSTEM
Pitch (longitudinal) control is achieved primarily by two independent,
hydraulically powered elevators. Elevator movement i s controlled from
conventional control columns. Primary control is supplemented by an
e l e c t r i c a l l y actuated t r im system which varies the angle of incidence of the
horizontal stabilizer.
The pitch control system incorporates two parallel-connected PCUs for each
elevator, and a dual control system. Normally, both control systems are
interconnected so that there i s simultaneous movement of both elevators, but it
is possible to isolate a jammed c i r c u i t by means of a disconnect mechanism,
thereby providing limited p i t c h control (one elevator only) through the
remaining c i r c u i t (refer t o Figures 1 and 2).
Control column movement i s transmitted by cables and pulleys, through an
a r t i f i c i a l feel unit, to the PCUs.
Operation of the elevator PCUs i s similar to that of the aileron PCUs.
Signal inputs from the AFCS are made through the rear quadrant of the l e ft
elevator control system only. Therefore, should jamming of the l e f t cable
circuit occur, the autopilot inputs would no longer be available to the
elevator system.
A. Pitch Trim
The aircraft is trimmed iti 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 stability augmentation system (SAS) are
processed by a trim control unit to operate the electrically driven
stabilizer actuator. 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. Both
control wheels have a red disconnect button, PITCH TRIM DISC, which can be
pressed to remove power from the system and brake the actuator. In order
to enhance the longitudinal trim movement, the movement of the horizontal
stabilizer is accompanied by a degree of elevator movement that alters the
stabilizer/elevator camber. An elevator servo input is generated by the
horizontal stabilizer movement 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 trim rate. Manual trim commands from
the control wheel pitch trim switches produce a steady rate of stabilizer
movement of 1/2 degree per second. Depending on flap position, the
autopilot commands variable high or low trim rates of 0.1 to 0.5 degree per
second and 0.01 to 0.1 degree per second respectively. Mach trim commands
a variable rate of stabilizer movement between 0.01 and 0.1 degree per
second. Each of the electric motors driving the trim actuator is protected
against overspeed by a dual coil brake.
SECTION 10
Page 5
Apr 02/87
OPERATING HMU4L
PSP 601A-6
The control unit 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 and back-up function. Should a failure occur
within a controller channel or its associated motor, the control unit
automatically transfers to the back-up channel. In the event of an
overspeed condition, the control unit removes power from the drive motor
and operates the brake in the actuator. Channel failure, overspeed
condition and channel change are indicated by switch/lights on the centre
pedestal.
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 supplies position signals to the control surface trim
position indicator.
Control Columns
The pilot's and copilot's control columns each consist of a conventional
tubular column mounted vertically in a housing. A control column shaker,
which is a component part of the stall protection system, is mounted on the
column.
Gain Change Mechanisms
Two independent gain change mechanisms 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. The system is designed to ensure a reduced feel
force when rapid control column movement is required.
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 4).
SECTION 10
Page 6
Anr H?/R7
OPERATING MANUAL
PSP 601A-6
5. WING FLAP SYSTEM (Figures 1 and 7)
The flap system consists of externally hinged inboard and outboard
double-slotted 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
asymmetry/overspeed 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.
The flaps are extended or retracted in response to command signals from the
FLAPS control lever located on the centre pedestal.
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. The FLAPS FAIL light on the copilot's instrument panel comes
on when a system fault is detected.
A. Flap Control Unit
The flap control unit (FCU) 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 aural warnings for airspeed/flap, take-off/flap and gear-up/flap
configuration incompatibilities.
B. Power Drive Unit
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 drives the flexible
shaft assemblies and actuators. When the selected flap position is
reached, the motors are de-energized and the motor brakes are re-applied.
If power to one of the PDU motors fails, the associated brake is
automatically applied 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 on
the copilot's instrument panel comes on. The thermal switches reset once
the overheat condition has passed.
SECTION 10
Page 7
Apr 02/87
OPERATING MANUAL
PSP 601A-6
C. Asymmetry/Overspeed Detector and Brake Assemblies
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 o*e-speed.
6. SPOILER SYSTEM (Figures 1 and 8)
Wing l i f t modulation is achieved by the operation of flight and ground
spoilers- The flight spoilers may be extended to any position, between 0 and
MAX (40 degrees), required for the intended flight path. The ground spoilers
have only two positions, fully retracted during flight or fully deployed (45
degrees) when activated with the aircraft on the ground, to assist other
braking systems by dumping l i f t and increasing drag.
A. Flight Spoilers
The flight spoilers are two hydraulically powered panels, one hinged to the
upper surface of each wing, forward 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.
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 cables and pulleys.
The spoilers are fully retracted when the lever is in the fully forward
position. Pulling the spoiler control lever rearward deploys the flight
spoilers, spoiler panel deployment being proportional with control lever
movement.
Lever positions, when selected, are held by a serrated plate and plunger
mechanism.
Spoiler panel position is transmitted to the control surface position
indicator, the LH FLT SPLR and RH FLT SPLR lights and the LEFT and RIGHT
FLIGHT SPOILERS lights.
A detent mechanism on both of the spoiler wing circuits prevents
unacceptable spoiler asymmetry. If an asymmetry occurs, the detent
mechanism closes the affected spoiler when the spoilers are less than
one-half extended or retracts i t to the one-half extended position when the
spoilers are more than one-half extended- Operation of the LEFT and RIGHT
FLIGHT SPOILERS lights indicate 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
Page 8
Apr 02/87
cacnhaadiiaeinr Qer
OPERATING MANUAL
PSP 601A-6
B. Ground Spoilers
The ground spoilers are two hydraul ically powered panels, one hinged to the
upper surface of each wing, forward of the inboard flaps, and are
controlled electrically. Each panel is powered by one actuator supplied
from a dual hydraulic selector valve.
The ground spoilers deploy automatically when armed, with a
weight-on-wheels or wheel spin-up signal present, and the spoiler control
lever and throttle lever selected to the proper positions (refer to Figure
8).
A spoiler control unit 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, 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
ground spoilers, if extended, close and lock. If both throttle levers are
not pulled back to IDLE simultaneously, the SPLRS INOP light will come on.
Ground spoiler operation is monitored via the LH and RH GND SPLR and SPLRS
INOP lights. The system test is initiated via the GROUND SPOILERS switch.
7. STALL PROTECTION SYSTEM (Figure 9)
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 principal system components consist
of two trailing vane type angle-of-attack transducers, a dual-channel stall
protection computer, two altitude transducers, two lateral accelerometers and
two flap position transmitters. The system controls and indicators are:
Two stick shakers
A stick pusher sub-system
Stall protection test indicators
System warning lights and test switches
An aural warning horn (warbler)
SECTION 10
Page 9
Jul 19/05
cacnhaadiiaeinr Qer
OPERATING MANUAL
PSP 601A-6
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
continues to increase, activates the stick shakers to generate a stall warning
in the form of a mechanical vibration of the control columns. If the aircraft
angle of attack still continues to increase to the stick pusher trip point, the
aural warning horn sounds and the stick pusher sub-system 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.
Red STALL/PUSH lights flash whenever the aural warning horn and stick pusher
are operating.
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 is attached to an externally mounted trailing
vane. The trailing vane is moved by the local airflow which varies in
proportion to the aircraft angle of attack. The angles of attack sensed by
the left and right transducers are transmitted to the left and right
channels respectively 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).
B. Stall Protection Computer
The stall protection computer 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 1 degree per second,
the computer lowers the angle-of-attack trip points for the various system
functions. This action prevents the aircraft momentum in the pitching
plane from carrying it through the stall warning/stick pusher sequence into
the stall.
The two altitude transducers 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).
SECTION 10
Page 10
Apr 02/87
OPERATING MANUAL
PSP 601A-6
As the altitude transducers signal an increase in altitude between 2,000
and 15,000 feet, the computer progressively lowers the angle-of-attack trip
points for the stick shaker and pusher. Below 2,000 feet and above 15,000
feet, the trip points are constant- If one or both of the altitude signals
i s lost or if the difference between signals exceeds 2,000 feet, the
computer automatically applies the trip points associated with the 15,000
foot altitude.
The two lateral accelerometers 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
test.
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.
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 for
faults in the electrical circuitry.
Stick Shakers
There are two stick shakers, one on the p i l o t s and one on the copilot's
control column. Each shaker is a dc electric motor driving an eccentric
weight. The shakers operate independently of each other and are powered by
their respective stall protection computer channels. The noise of the
stick shakers operating is sufficiently loud to constitute an aural warning
of shaker operation.
SECTION 10
Page 11
Apr 02/87
OPERATING MANUAL
PS? 601A-6
E. Stick Pusher Sub-system
The stick pusher consists of a rotary actuator driven by a dc electric
motor which operates on the right elevator control. The pusher logic
circuits are so arranged that pusher signals must be transmitted
simultaneously from both channels of the stall protection computer befcr,- a
stick push can be initiated. When in operation, the stick pusher exerts an
80-pound forward push on the control columns. 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 disconnect the
pusher drive i f the aircraft reaches 0.5 G during the pitching manoeuvre
induced by the stick push.
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. The stick pusher is capable of
operating immediately when the switch is released. The stick pusher can be
deactivated by either of two PUSHER toggle switches, located on the pilot's
and copilot's STALL PROTECTION panels, which would cause flashing STALL
PROTECT FAIL lights to come on.
SECTION 10
Page 12
Apr 02/87
chanehtyer
OPERATING MANUAL
PSP 601A-6
AUTOPILOT SERVO
ACTUATOR
PILOTS CONTROL COLUMN
Flight Controls
Figure 1
SECTION 10
Page 13
Apr 02/87
OPERATING MANUAL
PSP 601A-6
PITCH DISC AND ROLL DISC T-HANDLES
Provides a disconnecting mechanism for control columns and
control wheels if a jam occurs in respective cable runs.
Puffing either handte disengages associated mechanism.
Then, rotating handle left or right secures handle in
disconnected position. Releasing handle into stowed position,
reconnects associated controls and re-abgns control column
or wheels, as appropriate.
When PITCH DISC handle is pufled. pflot controls left
elevator and copilot controls right elevator.
When ROLL DISC handte is potted, pilot controls left aileron
and copilot controls right aileron.
CENTRE PEDESTAL
Control Disconnect T-Handles
Figure 2
SECTION 10
Page 14
Apr 02/87
OPERATING MANUAL
PSP 601A-6
NO. 1 SYSTEM
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AND
BRAKE SYSTEMS
NO. 1 HYDRAUUC SYSTEM
NO. 2 HYDRAUUC SYSTEM
NO. 3 HYDRAUUC SYSTEM
Flight Controls - Hydraulics
Figure 3
SECTION 10
Page 15
Apr 02/87
OPERATING MANUAL
PS? 601A-6
CONTROL SURFACE POSITION
INDICATOR
Provioes a continuous indication of
control surface posrticns over Tht->r
operating range.
L AND R FLT SPLR
Right spader up indications.
Max 40 degrees
2/4
la
V4
28 degrees
16 degrees
5 degrees
L AND R AILERON
21.3 degrees
21.3 degrees
L AND R ELEVATOR
Up
Down
23.6 degrees
18.4 degrees
iR L N R \
F ELEVATOR
|A • •
L X R \
RUDDER
RUDDER
Left/right indications
LEFT 25 degrees
RIGHT 25 degrees
CENTRE INSTRUMENT PANEL
PITCH LIGHT
Amber PITCH light comes on when
proximity sensors detect a jammed
control varve or input linkage at the
elevator power control units.
NOTE
Wrth hyotaufie power off. servo
monitor panel lights are as follows:
- ROLL fight is on
- Y AW light is on
- PITCH light is out
-MON SAFE fight is on.
ROLL AND YAW LIGHTS
Amber ROLL and YAW faghts come
on whenever proximity sensors
detect a jammed control valve or
hydraulic pressure deficiency at the
respective power control units.
CENTRE PEDESTAL
MON SAFE LIGHT
Green MON SAFE light comes on
when all aileron and rudder PCUs
are unpressurized (all hydraulic
systems off) and all elevator PCUs
are unjammed.
Control Surface Position Indicator and
Servo Monitor Lights
Figure 4
SECTION 10
Page 16
Apr 02/87
chauencjer
OPERATING MANUAL
PSP 601A-6
TRIM POSITION INDICATOR
Provides a continuous indication
to trim position over their
operating range.
ROD N l AND NR
Nose (left) NL/noseright <NR)
indications
Left
Right
8.5 degrees
8.5 degrees
AIL LWD AND RWD
Up
Down
7.5 degrees
7.5 degrees
LWD RWD/
T R I M ^ ^
STAB NUP
Nose up (NUP) indications.
StabBizer moves from 0 to -9
degrees incidence. Green band
indicates take-off (TO) trim range.
CENTRE INSTRUMENT PANEL
RUDDER TRIM CONTROL
Control switch sets rudder trim left
and right.
CHANNEL INOPERATIVE
SWITCH/UGHT
CHAN 1 INOP
CHAN 2 INOP
Amber fights indicate failure in
respective channel.
Pressing swxtch/bght in conjunction
with OVSP/CHANGE CHAN
switch/light activates pitch trim
system.
PITCH TRIM
r—PUSHCMANI
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.
AILERON TRIM CONTROLS
Control switches sets aileron trim up
and down.
OVERSPEED/CHANNEL CHANGE
SWITCH/LIGHT
Amber lights indicate pitch trim
overspeed or channel change. Can
be used to change from one channel
to other for test.
Pressing switch/tight in conjunction
with CHAN 1/CHAN 2 switch/Hght
activates pitch trim system.
CENTRE PEDESTAL
Trim Controls and Trim Position
Figure 5
Indicators SECTION 10
Page 17
Apr 02/87
OPERATING MANUAL
PSP 601A-6
AUTOPILOT/STICK PUSHB* DISCONNECT
vvvrrcH
^ed pushbutton which, when pressed, diserrgages
•topfiot and deactivates stick pusher. When
;*teased, stick pusher system is immediately
reactivated but autopilot remains disengaged.
PITCH TRIM SWTICH
Enables piot to vary pitch trim
according to flight requirement.
RADIO K>rY
Light grey button which, when
pressed., switches on radio
transmitter.
FRONT VIEW
AUTOPILOT TOUCH CONTROL
Black button which, when pressed,
enables pilot to manoeuvre aircraft
without disconnecting autopilot.
PITCH TRIM DISCONNECT ^AJTCh
Red button which, when pressed, removes power
from system and brakes actuator to cater to a
possfcle runaway trim actuator. System is
reactivated with CHAN 1 INOP/CHAN 2 INOP
and OVSP/CHANGE CHAN switch/lights
(refer to figure 5).
REARVIEW
Control Wheel SECTION 10
"Sure 6 page 18
Apr 02/87
cacnhaadiiaeinr Qer
OPERATING MANUAL
PSP 601A-6
FLAPS
FAIL
OVHT
M0T1
OVHT
MOT 2
FLAP FAIL LIGHT
Amber light comes on to indicate a
flap asymmetry or speed response
fault.
PDU MOTOR OVERHEAT LIGHTS
Amber light comes on to indicate an
overheat condition in the associated
PDU motor.
FLAP POSITION INDICATOR
Provides a continuous angular
indication of the flaps over their
operating range.
COPILOT'S INSTRUMENT PANEL
FLAP CONTROL LEVER
Controls operation of flap power
drive unit (PDU).
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.
CENTRE PEDESTAL
Wing Flap Controls and Indicators SECTION 10
Figure 7 Page 19
Apr 02/87
cacnhaadiiaeinr Qer
OPERATING MANUAL
PSP 601A-6
FLIGHT SPOILER DEPLOYED INDICATON
Amber lights come on steady when flight spoilers
are not fully retracted. Lights come on flashing
and take-off configuration aural warning sounds
when N1 rpm is increased beyond 75% and flight
spoilers are not retracted.
GROUND SPOILER DEPLOYED INDICATION
Amber lights come on when ground spoilers are
at any position other than fully retracted.
LH FLT
SPLR
LH GND
SPLR
RH FLT
SPLR
RH GND
SPLR
GLARESHIELD
SPOILER CONTROL LEVER
To deploy flight spoilers, lever may be moved
rearwards to any one of eight detented positions
according to flight path requirements until MAX
position stop is reached.
FLIGHT SPOILERS LEFT AND RIGHT
INDICATION
Green lights come on when flight spoilers are
extended beyond one-half position.
GROUND SPOILERS SWiTCH
ON - Arms ground spoilers for deployment.
Ground spoilers deploy automatically if a weighton-
wheels or wheel spin-up signal is present and
either of the following two sets of conditions has
been met:
- Spoiler control lever at or above 0 through
to 1/4 positions and both throttle levers
have been advanced above IDLE then
returned to IDLE or SHUTOFF positions.
-Spoiler control lever is between 1/4 and
MAX positions and both throttle levers are
at IDLE or SHUTOFF positions.
OFF - Ground spoilers are disarmed and cannot
be deployed.
TEST - LH and RH GND SPLR and SPLRS INOP
lights come on to indicate correct operation of
ground spoiler control system. Refer to Volume 1.
NORMAL PROCEDURES for test procedure.
GROUND SPOILER INOP LIGHT
Amber light comes on if spoiler control unit
detects fault in ground spoiler hydraulic selector
valves or if both throttle levers are not pulled back
to IDLE simultaneously.
CENTRE PEDESTAL
Spoiler Controls and Indications SECTION 10
Figure 8 Page 20
Jul 19/05
ehauenper
OPERATING MANUAL
PSP601A-6
ALT COMP FAIL LIGHTS (2)
Red lights come on if one or both attitude 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
tights are on.
GLARESHIELD
STALL/PUSH LIGHTS (2)
Red lights flash when angle of attack reaches
stick pusher trip point.
STALL PROTECT FAIL WARNING LIGHTS (2i
Red warning lights flash in the following cases:
- To indicate a system fault.
Whenever one of th AP/SP DISC buttons on
the control wheels is pressed.
- During system test.
Lights come on steady when power is removed
from system.
RED SECTOR
YELLOW SECTOR
NOTE
Stick pusher can only be tested on the ground;
all other tests can be conducted on the ground
or in-flight.
BLUE SECTOR
SPS TEST INDICATORS (2)
Coloured sectors on indicator provide references
for stall warning/stick pusher sequence during
system test (refer to Volume 1. NORMAL
PROCEDURES). Indicator is nor calibrated to
provide in-flight angle of attack indication oi
approach speed reference.
PILOT'S AND COPILOTS SIDE PANELS
S T A L L
PROTECTION
TEST PUSHER
.ON
PILOTS STALL PROTECTION TEST SWITCH
Spring-loaded toggle switch. Holding switch on
activates serf-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. Stick pusher can be checked only when
pilot's and copilot's TEST switches are held on
simultaneously.
STICK PUSHER SYSTEM SWITCHES £2)
Two-position toggle switches wired in series
between stick pusher actuator and battery bus.
When both switches are sei to ON. power is
available for stick pusher operation.
If one switch is OFF. stick pusher cannot operate
and both STALL PROTECT FAIL lights come on
steady.
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.
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.
PILOT'S FACIA PANEL COPILOTS FACIA PANEL
Stall Protection System Controls and Indicators SECTION 10
figure 9 Page 21
Apr 02/87

dul

MANUAL 8 MANUAL 8 MANUAL 8 MANUAL 8 MANUAL 8 MANUAL 8 MANUAL 8

qiushengsean

好好学习

thank you

bocome

Airbus A380 operations at alternate airports