ATA 27
Flight Controls
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FLIGHT CONTROL
SURFACES GENERAL
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A300/A310 FLIGHT CONTROL SURFACES
The control of the aircraft is achieved by:
• the primary flight controls
• the secondary flight controls
The secondary flight controls are the:
The primary flight controls ensure:
•
ROLL CONTROL achieved on each wing by:
one aileron
five roll spoilers, upper wing surfaces No. 3 through No. 7.
•
PITCH CONTROL achieved by two elevators hinged on the
trimmable horizontal stabilizer.
•
PITCH TRIM CONTROL achieved by the trimmable horizontal
stabilizer hinged on the aircraft structure.
•
YAW CONTROL achieved by one rudder.
MTT M540000 R3.3 01AUG01
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•
FLAPS
- three single slotted flaps on each wing
•
LIFT AUGMENTATION devices on each wing
- three slats
- one Krueger flap
- one notch flap—not applicable to A310
•
SPEEDBRAKES No. 1 through No. 5 on the upper surface of each
wing
•
GROUND SPOILERS No. 1 through No. 7 on the upper surface of
each wing
ATA 27
A300/A310
A300 Flight Control Surfaces
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Flight Compartment Controls and Indications
This illustration depicts all the controls and indications for the flight
surfaces located in the cockpit.
G. Speed Brake Control Panel
H. ECAM Display Control Panel
A. Servo Control Panel
I.
Aileron and Rudder Trim Switches
J.
Left ECAM Display Unit
B. Slats and Flaps Position Indicator
C. Pitch Trim and Yaw Damper Switch Panel
K. Master Warning and Caution Lights L/H
D. Flight Control Maintenance Test Panel
L.
Master Warning and Caution Lights R/H
E. Right ECAM Display Unit
F. Pitch Trim Wheel
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Flight Compartment Controls and Indicating
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FLIGHT CONTROLS HYDRAULIC POWER SUPPLY
The flight controls are powered by the three independent hydraulic
systems; redundancy is such that with two hydraulic systems failed, the
MTT M540000 R3.3 01AUG01
For Training Purposes Only
remaining system can operate the aircraft within an acceptable range of
the flight envelope.
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Flight Controls Hydraulic Power Supply
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FLIGHT CONTROLS - GREEN
HYDRAULIC POWER SUPPLY
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Flight Controls - Green Hydraulic Power Supply
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FLIGHT CONTROLS - BLUE
HYDRAULIC POWER SUPPLY
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Flight Controls - Blue Hydraulic Power Supply
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FLIGHT CONTROLS - YELLOW
HYDRAULIC POWER SUPPLY
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Flight Controls - Yellow Hydraulic Power Supply
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SERVO CONTROL P/B SWITCHES
1.
SERVO CTL PUSHBUTTON SWITCHES
2.
All these P/B switches are guarded. These P/B switches control
the servo shut-off valves for the individual hydraulic circuits Blue,
Green and Yellow.
•
A light comes on Amber when the flight control supply pressure in
the corres ponding hydraulic system has dropped (below 1450 PSI)
downstream of the servo control valve, or when the hydraulic supply
has been shut off. Illumination of an Amber LO PR light is
accompanied by ECAM activation.
NORMAL (P/B SWITCH PRESSED-IN)
Hydraulic power is supplied to the corresponding users as
soon as pressure is available in the corresponding hydraulic
system.
•
OFF (P/B SWITCH RELEASED-OUT)
The OFF light comes on White and the hydraulic power
supply to the corresponding users is shut off. The associated
JAM warning is inhibited and LO PR Amber illumination
confirms the OFF selection.
•
JAM
When a P/B switch is pressed-in, the associated JAM light
comes on Amber when a jamming is detected in the related
hydraulic control valves of rudder, elevator, ailerons or
trimmable horizontal stabilizer. Illumination of a JAM light is
accompanied by ECAM activation. The jammed control is
identified on the Warning Display (left CRT).
MTT M540000 R3.3 01AUG01
For Training Purposes Only
B, G, Y LO PR LIGHTS
NOTE: The SERVO CTL P/B switch positions and associated warnings
are repeated on the ECAM hydraulic system page.
-
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To prevent inadvertent complete deactivation of servo controls, only
two systems can be deactivated at a time by selection of SERVO
CTL P/B switches to OFF. When the third P/B switch is selected to
OFF all three systems are reactivated regardless of P/B switch
setting.
ATA 27
A300/A310
SERVO CTL (Servo Control) Panel
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ROLL CONTROL
SECTION
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ROLL CONTROL
improve the aerodynamic characteristics, a droop signal coming from the
slats control system moves the ailerons down 9.2° maximum when the
slats are extended. During cruise, the operational limits for aileron trim
are ±2°. The roll spoilers and speedbrakes are electrically signaled by
two identical computers (EFCU-Electrical Flight Control Units) that
elaborate the roll orders by processing the signals coming from the
control wheel position transducers units.
The roll control surfaces on each wing are:
• One (1) aileron powered by 3 servo controls
• 5 roll spoilers, each one powered by one (1) servo control.
The spoiler system is supplied from two normal bus bars (28 V DC and
26 V AC). If the normal buses have been cut off before landing, power is
supplied again to three spoiler groups by pressing the LAND RECOVERY
P/B switch on the overhead panel.
Each computer is composed of two control units and two monitoring
units. Each unit controls or monitors one group of surfaces. Each group is
made of one or two pairs of servo controls: spoilers 2-3, spoilers 4-1,
spoilers 5-7, spoilers 6. Thus, for a group of servo controls, the
corresponding control unit is in one computer and the monitoring unit is in
the other one. For the roll spoilers the control laws are such that they are
not usually used unless the control wheel is moved enough. An autopilot
servo actuator is mounted adjacent to the RH wing rear cable quadrant. It
drives the complete control via a detent lever which can be overridden by
the pilots.
From the two interconnected control wheels, the roll inputs are
transmitted to the ailerons by dual cables providing fail safe operation. In
each wing the inputs are transmitted to a differential unit receiving
additional inputs from:
• artificial feel unit
• aileron droop unit
• trim screw jack
In case of jamming in one control run, the interconnected spring strut can
be compressed to permit operation of the other control run to the other
wing. The pilot effort required on the wheel is between 34 lbs. and 90 lbs.
Spoiler control is still available but downgraded. Each servo control
linkage on the aileron includes a spring rod to protect it against a
runaway if an input lever on one jack remains in the open position.
INTERFACE WITH AUTOPILOT SYSTEM
An autopilot actuator is mounted adjacent to the right wing rear cable
quadrant; it drives the complete control via a detent lever which can be
overridden by the pilots. Dynamometric rods are installed upstream of the
cable tension regulators, they provide control signals to the control wheel
steering system.
The artificial feel is provided by a spring loaded rod. The, trim actuator is
electrically signaled by a control on the center pedestal. In order to
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Roll Control - Mechanical Aileron Control
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AILERON SYSTEM - COMPONENTS DESCRIPTION
1.
CABLE TENSION REGULATORS
Two tension regulators maintain a constant tension on the cables of
28.13 ± 5.30 lbf. They are identical apart from the input lever
position. They incorporate provision for installation of a special tool
used for installing the regulator on the aircraft.
2.
SERVO CONTROL ACTUATING SPRING ROD
The three ASA servo control actuating spring rods prevent runaway
of the control system if an input lever jams on its servo control body.
3.
CONTROL WHEEL INTERCONNECTING SPRING ROD
The two control wheels are interconnected by a spring rod in order to
allow one of the crew members to control half the surfaces in the
event of any single item jamming in the mechanical control system.
4.
RODS
Push-pull rods are adjustable or nonadjustable length, fitted with
replaceable ends.
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For Training Purposes Only
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5.
CABLES
The flexible cables (Dia. 3.2 mm/0.126 in.) are made of zinc-coated
carbon steel. The cable end fittings are equipped with barrels for
quick installation and fool proofing; turnbuckles are cliplocked.
Fairleads are of the roller type, for low friction purposes. The fairlead
supports allow passage of the cable end fittings. At bulkheads,
cables are fed through pressure seals.
6.
DYNAMOMETRIC RODS
The Flight Control Computer uses signals from the dynamometric
rods to detect the Captain's and First Officer's loads on the control
wheels. There are two rods for the pitch axis and two for the roll axis.
There is no rod in the yaw axis. The rods are placed in series in the
Flight Control linkages.
ATA 27
A300/A310
Aileron System - Components
4
1
5
1
3
2
6
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All Speed Aileron (ASA) - Mechanical Control
Each all speed aileron (ASA) is operated by three mechanically
controlled servo controls. The two interconnected control wheels drive
two symmetrical control systems composed of levers, rods, cables and
tension regulators routed along each side of the fuselage up to the input
levers of the servo controls. A differential and droop unit is installed in the
control linkage upstream of the servo controls. The unit receives two
inputs. One is from the control wheels (pilots input), the other is a droop
signal from the slat control system which droops the all speed ailerons
9.2° when the slats are extended in order to optimize aerodynamic
efficiency of the wing.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
When the droop signal is applied, all speed aileron deflection is not
simply modified by 9.2° throughout the travel range. Instead, response of
the all speed ailerons to control wheel motion is modified so that the
maximum up and down deflections remain close to those without droop
input. The droop signal also drives a differential mechanism between the
trim screwjack and the artificial feel unit. The mechanism pivots the
artificial feel unit, thus allowing the spring rod to remain at neutral.
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All Speed Aileron - Mechanical Control
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Aileron Trim
Trim control is electrically signaled. An electrical actuator installed in the
main gear W/W (center fuselage) drives two trim screwjacks via
sprockets, chains and cables. The actuator is controlled from panel
408VU located at the rear part of the center pedestal. Two switches on
this panel allow the crew to select constant speed displacement in the
appropriate direction. Trim position is indicated on scales at the top of the
control columns when the wheels are released.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
In each wing root, displacement of the trim screwjack drives the all speed
aileron servo control input linkage through the artificial feel unit, whose
spring rod remains at neutral. When the ailerons are drooped, the droop
signal drives a differential mechanism between the trim screwjack and
the artificial feel unit. The mechanism pivots the artificial feel unit, thus
allowing the spring rod to remain at neutral and the unit is held in this
position by the trim screwjack.
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Aileron Trim System
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Aileron Trim Components
A. AILERON TRIM ACTUATOR
B. ELECTRIC MOTOR CONTROL
The actuator is driven by a 28VDC electric motor through a reduction
gear and a torque limiter. The motor is a permanent magnet motor with
on-off control. A strong dynamic braking effect is obtained by shorting the
motor windings as soon as they are no longer energized (no static
braking on the actuator itself: trim irreversibility is provided by the
screwjacks downstream of the actuators).
•
•
•
Rotary stops limit output shaft rotation within the range allowed by
the screws.
When the electric motor is energized, it is protected by a torque
limiter when the stop limits are reached.
A rigging pin is used to set the output shaft at mid angular travel
(zero trim position and also zero reference for synchro transmitter
settings).
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For Training Purposes Only
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The electric motor windings of aileron trim actuator 9CG are
energized through contacts of two adjacent three position switches
(5CG) on control panel 408VU. The switches are spring loaded to the
center position and must both be moved simultaneously in the same
direction for the windings to be energized. The switch tabs are not
mechanically connected, to prevent run-away in the event of
mechanical jamming of one tab. The windings are shorted when the
two switches are in the center position.
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A300/A310
Aileron Trim System - Components
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All Speed Aileron - Artificial Feel Unit
There are two identical artificial feel units, each installed immediately
downstream of the all speed aileron servo control actuating spring rods.
The units each include a spring rod and are held in position by the trim
screwjacks.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
Their function is:
• To maintain servo control input linkage in trim position in the event of
disconnection of the control linkage upstream of the servo controls.
• To provide artificial feel loads proportional to control wheel
deflection.
• To provide accurate return of the surfaces to neutral.
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All Speed Aileron - Artificial Feel Unit
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Aileron Trim/spoiler and Speed Brake Switches
1.
AIL TRIM SWITCHES
Ailerons trim control is electrically powered. For safety purposes,
both switches must be moved and held in the same direction (L
WING or R WING) to energize the system. This action selects a
constant speed displacement in the corresponding direction. Full
travel of about 7° of aileron in each direction is achieved at a speed
of 0.4° per second.
2.
ON (P/B switch pressed-in):
Corresponding control system is activated. Each time a system is
activated, or corresponding hydraulic system on, or the aircraft
electrical network is energized, a 2 second safety BITE test is
triggered for the corresponding EFCU units (control and monitor).
•
OFF/R (P/B switch released-out):
The OFF/R light comes on White and the corresponding control
system is deactivated. If hydraulic pressure is available, the
actuators are automatically held in the retracted position. The
monitoring circuits are reset by ECAM activation. This action is
accompanied by ECAM activation.
•
FAULT:
When a P/B switch is pressed-in, the associated FAULT light
comes on Amber if a failure is detected by the monitoring
circuits, which then deactivate the control system. Illumination of
the FAULT light is accompanied by ECAM activation.
AILERON TRIM SCALES
A scale representing 14° of aileron movement (7° in each direction)
is engraved and painted on the top of each control column opposite
a pointer painted on the control wheel. With the control wheels
released, the crew can thus read the actual aileron trim value.
3.
•
SPLR & SPD BRK PUSHBUTTON SWITCHES
Each P/B switch is associated with one or two pairs of symmetrical
upper wing surfaces.
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Aileron Trim/spoiler and Speed Brake Switches
A
B
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RUDDER CONTROL
SECTION
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A300/A310
Rudder System - Yaw Control
The rudder, operated by 3 mechanically controlled servo controls,
receives pilot's inputs by a single cable run to a spring loaded artificial
feel unit connected to the trim screwjack. From this point up to the servo
controls, the commands are transmitted by dual rigid linkage, receiving
additional inputs from a rudder travel limiter, yaw damper and autopilot
servoactuators. The artificial feel is provided by a spring loaded rod. The
trim actuator is electrically signaled. It is driven by an electrical motor.
During cruise, the operational limits for rudder trim are ±1.5°.
INTERFACE WITH AUTOPILOT SYSTEM
An autopilot actuator is mounted adjacent to the artificial feel and trim
unit upstream of the variable stop lever; it drives the complete control via
a detent lever which can be overridden by the pilot. A yaw damper
actuator, mounted between the artificial feel and trim unit and the
variable stop lever, drives the rear control via a differential linkage. The
yaw damper actuator signals are added to those of the pilots, up to the
maximum travel allowed by the variable stop lever. The yaw damper
actuator is fail-safe, so that disconnection of the control is extremely
improbable.
The rudder travel limiter reduces the pedal and rudder deflection from
±30° at speed below 165 kt to ±5° at 308 kt and above. The orders are
delivered by two independent RUDDER TRAVEL channels, each one
included in a digital computer (Feel and Limitation Computer) receiving
inputs from the DADCs (Digital Air Data Computers) and the SFCCs
(Slats Flaps Control Computers). Each computer controls an electrical
motor driving a common electromechanical actuator coupled to variable
stop lever. Only one channel is normally active. The other is in standby.
A spring loaded rod positions the variable stop lever in the low speed
position in case of dual failure.
INTERFACE WITH MAIN WHEEL BRAKING SYSTEM
Levers are attached to each pedal, to provide braking inputs when the
pedals rotate about their axis.
INTERFACE WITH NOSE WHEEL STEERING
The nose wheel steering control is connected to the rudder control
through a hydraulic steering control coupler (engaged when the landing
gear is extended) and a spring rod, the threshold of which is lower than
the threshold of the rudder artificial feel and trim unit spring rod. The
spring rod prevents the nose wheel steering control from transmitting
inputs to the rudder control.
An. autopilot servo actuator is mounted adjacent to the artificial feel unit
upstream of the variable stop lever. It drives the complete control via a
detent lever which: car. be overridden by the pilot. Yaw damper
commands are transmitted via a differential unit canceling a feedback to
the pedals. A spring loaded rod on each servo control input avoids a
runaway of the rudder in case of jamming of one input lever in the open
position. Levers are attached to each pedal, to provide brake inputs
when the pedals rotate around their pivots.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
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A300/A310
Rudder System - Yaw Control
MTT M540000 R3.3 01AUG01
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A300/A310
Rudder System - Rudder Trim Actuator
One 28VDC electric motor is fitted in the actuator, directly coupled to the
reduction gear. It is energized when rudder travel is selected.
•
When the electric motor is energized, it is protected by a torque
limiter when the stops are reached (motor rotation is not stopped).
•
The motor is a permanent magnet motor with on-off control. A strong
dynamic braking effect is obtained by shorting the windings of the
motor when it is de-energized (no static braking on the actuator itself:
trim irreversibility is provided by the screwjack downstream of the
actuator).
•
•
The actuator includes a position transducer which delivers rudder
trim position signals to associated electrical circuits. The transducer
is a special RVDT, of the same type as those installed in the two
transducer units used for electrical roll control. The electrical
characteristics of the RVDT are monitored by associated computer
circuits.
Rotary stops limit output shaft rotation within the range allowed by
the screw.
•
A rigging pin is used to set the output shaft at mid angular travel
(zero trim position and also zero reference for transducer setting).
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For Training Purposes Only
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Rudder System - Rudder Control Input Components
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Rudder System Yaw Control - Rudder Trim Switches
1.
2.
RUD TRIM ROTARY SELECTOR
Rudder trim control is electrically powered. The rotary selector is
springloaded to the neutral (center) position. The direction of rudder
trim travel depends on the direction of rotary selector (NOSE L or
NOSE R). Full authority of rudder trim is about 21° in each direction.
RESET PUSHBUTTON SWITCH
It allows initiation of an automatic sequence controlled by the EFCUs
to position the rudder trim at 0° ±0.2°.
•
3.
•
Normal (P/B switch released-out)
Automatically or manually, the reset action is stopped and the
ON light goes off.
•
FAULT
The light comes on Amber if a failure of the reset function is
detected or if the actuator position transducer fails.
RUD TRIM POSITION INDICATOR
A digital indicator displays rudder trim direction (L or R) and value (0°
to 21°).
ON (P/B Switch pressed-in)
The ON light comes on White. The switch is latched during the
reset action and will release out automatically when reset is
achieved.
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For Training Purposes Only
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Rudder System Yaw Control - Rudder Mechanical/Hydraulic
3
1
2
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A300/A310
Rudder System - Rudder Artificial Feel
A spring assembly located in the artificial feel and trim unit restores a
resistance to pedal depression which is proportional to rudder
movement. A variable stop lever installed downstream of the servo
controls on the control linkage serves to reduce rudder deflection with
respect to pedal movement as the airspeed increases.
Spring function is:
•
•
•
•
RUDDER ARTIFICIAL FEEL
An artificial feel and trim unit is installed adjacent to the rear cable
quadrant. It consists of a trim screwjack and a fail-safe constant resisting
load spring rod, held in neutral position by the trim screwjack.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
•
27-44
2744
To maintain the downstream linkage and the input lever of the servo
controls at neutral in the event of disconnection of the control linkage
upstream of the artificial feel and trim unit
To provide artificial feel loads
proportional to rudder deflection
To provide accurate centering of the surface at neutral in the
absence of a control input
To maintain the upstream controls at neutral, when signals are
provided to the servo controls by the yaw damper actuator.
ATA 27
A300/A310
Rudder System - Rudder Artificial Feel
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Servo Controls
1.
SAFETY VALVES
2.
To preserve the Green system, safety valves are installed upstream of
the following components:
• Krueger selector solenoid valve (in case of engine failure)
• Rudder servo control (in case of inflight collision)
SERVO CONTROLS JAMMING DETECTION
There is one jamming detection circuit for each hydraulic system. If
jamming occurs the electronic circuitry inside the jamming detection
control unit receives 28V directly from the jamming detection microswitch
if a servo control is involved and from an intermediate logic if a THS
actuator hydraulic motor is involved.
Jamming detection is associated with the mechanically driven control
valves of the left and right all speed aileron, and left and right elevator
and rudder servo controls. It is also associated with the THS actuator
hydraulic motor control valves (for the Green and Yellow systems only).
NOTE: When a hydraulic system is selected OFF, the + 28V sent to the
corresponding jamming detection microswitches is cut off.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
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Rudder System - Servo Controls - Components
2
1
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Rudder System Travel - System 1 and 2 Pushbutton Switches
1.
RUD TRAVEL CONTROL PANEL
1. The P/B switches control channels 1 and 2 of the Feel and
Limitation Computers (FLC) for rudder travel limiting.
•
2.
ON (P/B switch pressed-in):
The corresponding system is engaged. Both systems may be
engaged simultaneously, but only system 1 is effectively
active. If system 1 fails, it is automatically deactivated and
system 2 becomes active.
•
OFF/R (P/B switch released-out):
The OFF/R light comes on White and the system involved is
disengaged. The monitoring circuits are reset by this action.
This indication is accompanied by ECAM activation.
•
FAULT:
When a P/B switch is pressed-in, its FAULT light comes on
Amber if a failure is detected in the respective system.
Illumination of the Amber FAULT light is accompanied by
ECAM activation. Both FAULT lights remain illuminated
when the switches are released-out and the OFF/R lights are
illuminated White. This constitutes a rudder disagree warning
(The variable stop lever is not in low speed position with
flaps extended 20° or more). Illumination of both FAULT
lights is accompanied by ECAM activation.
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For Training Purposes Only
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YAW DAMPER LEVERS
•
1 (or 2): The lever is magnetically latched in active position and
the yaw damper 1 (or 2) is engaged. If a failure is detected, the
YAW DAMPER 1 (or 2) lever trips to OFF.
•
OFF: The respective yaw damper is disengaged. When one
YAW DAMPER lever trips to OFF, the associated yaw damper
system disengages and the ECAM is activated. When both YAW
DAMPER levers trip to OFF the yaw damper function is lost and
the SCAM is activated.
ATA 27
A300/A310
Rudder System Travel and Yaw Damper Systems - Control Switches
B
A
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-49
2749
ATA 27
A300/A310
Rudder System - Rudder Travel Limiting
•
•
•
The rudder travel limiting system modifies control inputs to the servo
controls to vary rudder travel in relation to airspeed (Vc). Limitation is
such that the maximum deflection which can be achieved by the rudder
remains lower than the deflection which would induce limit loads on the
structure, throughout the flight envelope.
2.
TRANSDUCER UNIT
The actuator is servo controlled and is monitored through a
transducer unit driven by variable stop lever movement. The
transducer unit, comprising two inductive transducers, is identical to
the one used in the spoiler control system.
3.
SPRING - RETENTION ROD
In the event of a rupture or disconnection of an actuator attachment,
a retention rod limits actuator movement to prevent it from jamming
the variable stop lever. A spring returns the lever to the "low speed"
position where full control deflection (+30) is possible.
The system is composed of:
•
•
•
1.
A variable stop unit consisting of an articulated lever operated by an
electromechanical actuator and a transducer unit detecting lever
position. These items are all mounted on a frame assembly located
downstream of the differential between the AP and yaw damper
actuators.
Two control and monitoring computers designated FLC (Feel and
Limitation Computer).
One RUD TRAVEL control panel, one PITCH FEEL & RUD
TRAVEL maintenance panel and five electrical power supply circuit
breakers.
A nut/screw system, driven by means of a torque limiter
Mechanical end-of-travel stops
A torque limiter provided to protect the reduction system from
any abrupt jamming of the output shaft, particularly when it
reaches the mechanical stop.
VARIABLE STOP ACTUATOR DESCRIPTION
• Two AC motors, supplied with 26V-400 Hz
• A single reduction gear actuated by both motors, which are
rigidly connected
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-50
2750
ATA 27
A300/A310
Rudder System - Rudder Travel Limiting
2
1
3
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-51
2751
ATA 27
A300/A310
Rudder Travel and Pitch Feel Systems - Feel and Limitation Computer (FLC) - General
This computer contains the circuitry required for two functions: rudder
travel limiting and pitch feel. The FLC is a digital computer comprising
two different computation channels:
• Rudder travel limiting/pitch feel control channel
• Rudder travel limiting/pitch feel monitor channel
Safety of the systems is ensured by:
• control and monitor channel programs which are intentionally
different
• monitoring of digital computations which are performed by control
and monitor channels with the same input data, achieved by
comparison between the results of both channels, by means of
analog comparators
• power loop monitoring achieved by software means in each digital
channel.
If any indicator is on, the test of either RUDDER TRAVEL LIMITING
system or PITCH ARTIFICIAL FEEL system will not operate.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-52
2752
ATA 27
A300/A310
Rudder System and Pitch Feel - Feel and Limitation Computer (FLC 1/2)
FLC1/FIN 302CY1
FLC2/FIN 302CY2
MTT M540000 R3.3 01AUG01
For Training Purposes Only
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A300/A310
THIS PAGE INTENTIONALLY
LEFT BLANK
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-54
2754
ATA 27
A300/A310
ELEVATOR SYSTEM
PITCH CONTROL
SECTION
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-55
2755
ATA 27
A300/A310
Elevator System - Pitch Control
Pitch control is achieved by two elevators hinged on the horizontal
stabilizer, each actuated by three servo controls controlled by a dual
mechanical linkage through dynamometric rods, cable runs, an artificial
feel system linked to the cable run of the LH control column, and load
limiting rods. In normal operation the two elevators are controlled
together. In case of jamming in one control linkage during flight (take off
excluded), pitch control is provided by THS (Trimmable Horizontal
Stabilizer). If jamming occurs at take off, two uncoupling bellcranks
enable the elevator on the other side to be controlled by one or both
pilots.
downstream of the artificial feel system, a load limiting spring rod limits
the efforts in the elevators control linkage. A spring loaded rod on each
servo control input avoids a runaway of the elevator in case of jamming
of one input lever in the open position. An autopilot actuator is mounted
adjacent to the LH elevator. It drives the control via a detent lever which
can be overridden by the pilots.
Pitch trim is provided by adjustment of the horizontal stabilizer from +3°
(nose down) to -14° (nose up). It is actuated by a fail safe ball screw jack
driven by two independent hydraulic motors supplied respectively by
Green and Yellow systems and coupled by a differential gear through
pressure-off brakes. Horizontal stabilizer adjustment may be initiated:
• manually (AP disengaged) by trim wheels operation (mechanical
mode) or by action of the control wheel rocking levers (electrical
mode).
• automatically by AP trim, mach trim or alpha (angle of attack) trim
function.
A pitch uncoupling unit (locking rod plus solenoid) prevents accidental
asymmetrical deflection of the elevators during flight and allows
uncoupling of the RH and LH control systems during take off (locked at
speeds lower than 30 kt or higher than 195 kt). Artificial feel is provided
by the associated action of:
• a double action spring loaded rod
• a torsion bar driven by a variable gain mechanism which generates a
variable stiffness in the control. The variable gain mechanism is
actuated by either of two electrohydraulic actuators. Each actuator is
controlled by an independent PITCH FEEL channel, each one
included in a FLC (Feel and Limitation Computer).
Electrical and automatic trim signals are processed in two FAC (Flight
Augmentation Computers) and control two electrical motors. Trim speed
and trim authority depend on trim mode and aircraft configuration. The
motors drive the control linkage to the hydraulic valves which control the
hydraulic motors. The manual trim wheel run is connected to the same
linkage. Stall warning is provided by a stick shaker (electrical motor)
which is installed on each control column, and controlled by the FWC
(Flight Warning Computer)
PITCH FEEL systems are operative above 165 kt. Inputs are a function
of stabilizer position, airspeed and Mach number. In case of failure of two
systems, the mechanism returns to the sow speed position In each run,
MTT M540000 R3.3 01AUG01
For Training Purposes Only
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A300/A310
Elevator System - Pitch Control - Diagram
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-57
2757
ATA 27
A300/A310
Elevator System - Elevator Mechanical Control System - General and Components
A. GENERAL
4.
Rods: Identical to those used in the aileron control system
Each elevator is operated by three mechanically controlled servo
controls. The inputs from the control columns are transmitted to the
elevators by dual control systems. Each system is routed along one side
of the fuselage. The left and right systems are interconnected at two
points by detent bellcranks, one beneath the flight compartment floor, the
other between the two elevators.
5.
Cables: Identical to those used in the aileron control system
6.
Control column stops: Control column travel is limited in both
directions by non adjustable stops.
Elevator operational stops: Maximum input to the servo controls is
limited by adjustable stops located at a lever, close to each elevator.
Elevator travel stops: These are the stroke end stops
(non-adjustable) of the servo controls, never reached in normal
operation.
Elevator structural stops, when the servo controls are not installed:
The elevators rest on structural down stops, designed for that
purpose, which are not able to withstand any load other than the
weight of the elevators.
Adjustable levers: The length of a lever close to each elevator is
adjustable in order to maintain maximum travel of the elevators
within the design limits.
B. COMPONENT DESCRIPTION
1.
Cable tension regulators maintain a constant tension on the cables
(49.50 ±9.23 lbf).
2.
The servo control actuating spring rods:
• Provide flexibility in the control for any asymmetrical deflection of
the elevators in ground gusts
• Prevent runaway of the control system if an input lever jams on
its servo control body.
3.
NOTE: Rigging pin holes are provided at convenient places to facilitate
rigging.
A load limiting spring rod in each system, downstream of the artificial
feel unit, limits the design loads.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
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ATA 27
A300/A310
Elevator System - Elevator Mechanical Control
MTT M540000 R3.3 01AUG01
For Training Purposes Only
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2759
ATA 27
A300/A310
ARTIFICIAL FEEL (ELEVATOR)
The pitch artificial feel system creates load feel at the control column
which is variable with flight conditions, in order to reduce the variation of
force per g throughout the whole flight envelope. At high angle of attack,
the system causes an increase in the load feel at the control column
resulting in aircraft return to permissible angle of attack configuration.
1.
2.
Pitch Artificial Feel Actuator
Each actuator includes:
• A biased servovalve which modulates pressure in the actuator
large chamber, the small chamber being permanently supplied
with high pressure. In the event of an electrical failure, servovalve current is nulled and its control valve is displaced so that
the actuator is retracted.
• A solenoid valve, energized in normal operation
• A bypass which connects the large chamber to return in order to
retract the actuator when the solenoid valve is de-energized. It is
therefore redundant with respect to the servovalve bias.
• A position pickoff potentiometer.
Pitch Artificial Feel Unit
The artificial feel unit is composed of:
• spring box providing a force threshold
• A torsion bar driven by a variable gain mechanism which
generates variable load feel.
• Two electrohydraulic actuators, displacement of which produces
the kinematic gain variation.
• One return spring box used to retract the two actuators to the
position corresponding to "low speed" load feel, in the event of
double hydraulic failure.
3.
Pitch Upcoupling Unit
Provides connection of LH and RH elevators from 0-30 knots
airspeed during takeoff roll.
Above 30-195 knots, the LH and RH elevators are disconnected by
ADC 1/2 to allow either pilot to control the elevator (pitch) function. In
case of a jam in the elevator control system on the captain’s or first
officer’s control panels
Above 192 knots, both elevators will reconnect for full control of the
LH and RH elevator system runs.
During landing conditions, this process is repeated in the same
airspeed conditions.
Pitch Artificial Feel Unit - Operation
The actuators act on the gain variation mechanism by means of
levers. Gain is imposed by the actuator having extended the furthest.
In the event of jamming of the mechanism, a microswitch transmits a
pitch disagree warning signal. The artificial feel unit includes a "fail
safe" part to avoid loss of the force threshold and feel load at the
same time.
4.
Aft Detent Bellcrank
In the event of an elevator jam on the LH or RH elevators, the aft
detent bellcrack will release on the jammed side to prevent lockout of
the elevator system.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
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ATA 27
A300/A310
Elevator System - Pitch Artificial Feel - Components/Location
3
2
4
1
Artificial Feel Unit
MTT M540000 R3.3 01AUG01
For Training Purposes Only
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ATA 27
A300/A310
Elevator System - Pitch Uncoupling System - General and Components
The two elevator control channels can be uncoupled during the takeoff
phase in the event of jamming at any point on the control systems, by
means of two detent bellcranks; one installed between the control
columns, the other between the two elevators. A pitch uncoupling unit,
comprising a solenoid and rod, prevents any inadvertent uncoupling of
the two elevators after the takeoff phase in order to prevent asymmetrical
loads being applied to the structural attachments of the trimmable
horizontal stabilizer. The uncoupling unit solenoid is energized if airspeed
Vc is higher than 30 kts and lower than 195 kts.
COMPONENT DESCRIPTION
1.
Solenoid
The solenoid includes the following components:
• A low resistance draw coil, allowing high intensity current to
provide a high draw force when the coil is energized.
• A high resistance holding coil allowing low intensity current to
provide permanent operation capability of the solenoid.
• Two end of stroke switches, one for direct draw coil energization,
one for test purposes.
• A return spring, to lock the rod when the solenoid is
de-energized.
NOTE: The lower limit of 30 kts (minimum speed for which a Vc value
can be obtained from ADCs) has been introduced to prevent permanent
energization of the solenoid and power contactor coil when the aircraft
electrical network is energized on the ground.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
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2762
ATA 27
A300/A310
Elevator System - Pitch Uncoupling System - Schematic
MTT M540000 R3.3 01AUG01
For Training Purposes Only
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ATA 27
A300/A310
Trimmable Horizontal Stabilizer System - Pitch Controls
Both pitch trim wheels provide mechanical control of the Trimmable
Horizontal Stabilizer (THS). When a pitch trim control wheel is used to
override the electrical command, it disengages the electric actuators and
the PITCH TRIM levers trip to OFF. The trim range is from 14° nose up
to 3° nose down. Trim position is indicated in degrees on a scale
adjacent to each trim wheel which is painted Green over the normal take
off range (2° DN. 2.5° UP).
On each control wheel a rocking lever for pitch trim control is installed.
Up or down movement of the rocking levers activates the two electric
actuators which control the hydraulic motors for horizontal stabilizer
adjustment providing that at least one PITCH TRIM system is engaged
and AP is OFF or in CWS mode. The rocking levers are spring loaded to
neutral position. If both rocking levers are operated simultaneously, but in
opposite position, trimming action stops. If trimming by means of the
rocking levers lasts for more than 1 sec., an aural warning is activated.
NOTE: The pitch trim rate is:
• 0 . 9 ° /s when the speed is below 200 kts.
• 0.17°/s when the speed is above 240 kts. It varies
linearly from 0.9°/s to 0.17°/s when the speed is between
200 and 240 kts.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
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2764
ATA 27
A300/A310
Trimmable Horizontal Stabilizer System - Pitch Controls
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-65
2765
ATA 27
A300/A310
Elevator System - Elevator Surface Position Indicating
Position of the right elevator is indicated on the right SCAM display unit,
with the hydraulic systems available for the servo controls. There is no
special reference mark painted on the elevators, but on each side of the
APU tailcone, there is:
•
an engraved reference plate which indicates the neutral position of
the corresponding elevator.
•
an engraved placard with the following inscription:
VALID STABILIZER IN NEUTRAL POSITION.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
1.
ELEV AND STAB POSITION INDICATION
A White scale covering the full travel range is provided for elevator
and trimmable horizontal stabilizer position. An index indicating the
actual position of the surfaces moves along each scale. In addition,
each available hydraulic system on the THS is indicated by a Green
symbol (G,Y). In case of servo control low pressure detection, the
corresponding symbol becomes Amber.
27-66
2766
ATA 27
A300/A310
Flight Controls System - RH ECAM Page - System Display
Elevator and Horizontal Stabilizer Position Indication
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-67
2767
ATA 27
A300/A310
Pitch Feel and Trim System - Control Switches and Levers
1.
PITCH FEEL SYS 1 AND 2 PUSHBUTTON SWITCHES
2.
The P/B switches control channels 1 and 2 of the Feel and Limitation
Computers (FLC) for elevator control.
•
•
•
ON (P/B switch pressed-in): The corresponding system is
engaged. Both systems may be engaged simultaneously but
only one is effectively operating. If one system fails, it is
automatically deactivated and the other one continues to
operate.
OFF/R (P/B switch released-out): The OFF/R light comes on
White and the system involved is disengaged. The monitoring
circuits are reset by this action. This indication is accompanied
by SCAM activation.
PITCH TRIM 1 AND 2 LEVERS
•
1 (or 2): The lever is magnetically latched in the active position
and the pitch trim 1 (or 2) is engaged. If a failure is detected, the
corresponding PITCH TRIM lever trips to OFF.
•
OFF: The respective pitch trim is disengaged.
-
When one PITCH TRIM lever trips to OFF, all electrical
control modes of the THS are lost and the ECAM is
activated.
-
When both PITCH TRIM levers trip to OFF, all electrical
control modes of the THS are lost and the SCAM is
activated.
NOTE: Pitch trim disengages and the levers drop to OFF when trim
reaches full nose up or full nose down position (mechanical stops).
FAULT: When a P/B switch is pressed-in, the associated FAULT
light comes on Amber if a failure is detected in the corresponding
system. Illumination of the Amber FAULT light is accompanied
by ECAM activation.
Both FAULT lights remaining illuminated when the P/B switches
are released-out and the OFF/R lights are illuminated White,
constitutes a pitch disagree warning (The artificial feel unit
operates in high speed configuration when flaps are extended
20° or more ) . Illumination of both FAULT lights is accompanied
by ECAM activation.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-68
2768
ATA 27
A300/A310
Flight Deck Pitch Feel and Trim System - Control Switches and Levers
B
A
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-69
2769
ATA 27
A300/A310
THIS PAGE INTENTIONALLY
LEFT BLANK
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-70
2770
ATA 27
A300/A310
TRIMMABLE HORIZONTAL
STABILIZER (THS) SECTION
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-71
2771
ATA 27
A300/A310
Trimmable Horizontal Stabilizer System - General
Pitch trim control is achieved by a trimmable horizontal stabilizer (THS)
hinged on the rear part of the fuselage. The two elevators are hinged on
the THS. Their control systems are installed so that the elevators are in
line with the THS when the control columns are released. The THS is
driven by an actuator including a fail-safe ball screwjack, the structural
attachments of which are also fail-safe. Normal control of the actuator is
electrical, via the automatic pitch trim system.
INTERFACE WITH AUTOMATIC PITCH TRIM SYSTEM
Electrical control is achieved by means of two electric pitch trim
actuators, installed on the THS actuator. They drive the actuator
mechanical input. The two pitch trim actuators are controlled by two flight
augmentation computers (FAC) which deliver manual electric trim,
automatic trim, Mach trim and alpha trim signals. Manual electric trim
signals are provided by rocker switches mounted on the Captain's and
First Officer's control wheel horns. Electric limit switches detect THS end
of travel in aircraft nose up direction. The signals are used in the
automatic trim system to avoid automatic disconnection of this system
during automatic landings.
Stand-by controls are mechanical. The pilots can override electrical
control by the mechanical control system by applying sufficient force to
the control wheels. A torque limiter is mounted in each electric pitch trim
actuator for that purpose. The torque limiters remain automatically
released after their operation.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
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ATA 27
A300/A310
Trimmable Horizontal Stabilizer System - Diagram
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-73
2773
ATA 27
A300/A310
Trimmable Horizontal Stabilizer System - Hydraulic Actuation - Components/Location
The trimmable horizontal stabilizer is driven by an actuator which
includes two hydraulic motors, each powered by a different hydraulic
system.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-74
2774
ATA 27
A300/A310
Trimmable Horizontal Stabilizer System - Hydraulic Components/Location
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-75
2775
ATA 27
A300/A310
Trimmable Horizontal Stabilizer System - Actuator - General
The THS actuator consists of a fail-safe ball screwjack actuated by two
hydraulic motors coupled by a differential gear.
NORMAL OPERATION
Pressure-off brakes (9) are released. The ball screw is held by the
no-back brake formed by items (14) (15) (16) (17). Rotation of input shaft
(1), driven either by one of the electrical pitch trim actuators or by
MTT M540000 R3.3 01AUG01
For Training Purposes Only
the mechanical input, controls rotation of the ball screw through two
identical control loops, including input and feedback gear trains,
feedback differentials (4), control valves (5), hydraulic motors (8) and
actuate a power gear train through power differential (10). The control
stroke is limited by the actuator input shaft stop (2). The structural
components (ball screw and nut assembly, attachments to THS and
fuselage) and the power gear train are duplicated, the secondary load
path being normally unloaded.
27-76
2776
ATA 27
A300/A310
Trimmable Horizontal Stabilizer System - Actuator - Schematic
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-77
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ATA 27
A300/A310
WARNING LOGIC PITCH CONTROL SYSTEM
Depicted below are the various warnings displayed in the Flight Compartment in the
event the fault shown occurs in the Pitch Control System of the aircraft. Also
graphically shown are the Flight Phases at which the warnings will or will not be
displayed.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
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ATA 27
A300/A310
Warning Logic - Pitch Control System
MTT M540000 R3.3 01AUG01
For Training Purposes Only
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ATA 27
A300/A310
Alpha Probes System - Stall Warning - General
•
The dual stall warning system provides audio (cricket) and vibrating
(stick shaker) warning in case of impending stall.
•
On each control column, a stick shaker is installed and controlled by
the stall warning generator included in FWC 1 or 2.
•
The angle of attack is the governing parameter for stall warning,
together with slat extension.
•
Stall warnings are activated when angle of attack exceeds a
predetermined value.
•
The angle of attack is given by two alpha probes (one on each side
of the forward fuselage) which are electrically heated. Slat position is
transmitted by two synchro-transmitters, one for each FWC.
•
slat retraction is inhibited.
•
turn coordination of yaw damper is inhibited.
Stall warning is inhibited on the ground except during ALPHA PROBES
test.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-80
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ATA 27
A300/A310
Alpha Probes System - Stall Warning - Block Diagram
Captain
Captain
First Officer
First Officer
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-81
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ATA 27
A300/A310
Flight Control System - Stick Shaker - General
One stick shaker is installed on each control column, and is controlled by
the Flight Warning Computers 1/2.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-82
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A300/A310
Flight Control System - Stick Shaker
MTT M540000 R3.3 01AUG01
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27-83
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THIS PAGE INTENTIONALLY
LEFT BLANK
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-84
2784
ATA 27
A300/A310
LEADING EDGE LIFT DEVICES
SECTION
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-85
2785
ATA 27
A300/A310
Wing Leading Edge - Slat System - General
SLAT SYSTEM
KRUEGER FLAP AND NOTCH FLAP
There are three slat surfaces in each wing, the inboard, center and
outboard slats. They are guided on curved support tracks. The inboard
slat has three tracks and the center and outboard have four each. A
folding nose on each inner slat folds to clear the engine pylon when the
slats extend. The slats are actuated by ballscrew jacks, two for each
surface. Two friction brakes, one at each end of the wing transmission
system provide system irreversibility. Attached to each friction brake is a
position pick-off unit (PPU) for asymmetry and system monitoring.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
The Krueger flap and notch flap are provided to complete the wing
leading edge profile when the slats are extended. The Krueger flap and
notch flap are operated by individual hydraulic actuators. Both are
controlled by the SFCC and move to the extend and retract position
when the SFCC commands slat extension or retraction.
Slats and spoilers are numbered from inboard to outboard, each side
separately.
27-86
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ATA 27
A300/A310
A300 Wing Leading Edge - Slat System Components - Location
A300
SHOWN/A310 DO
NOT HAVE
NOTCH (OR
SLOT) FLAPS
MTT M540000 R3.3 01AUG01
For Training Purposes Only
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2787
ATA 27
A300/A310
Wing Leading Edge - Slat System - Hydraulic Operation - Diagram
The Power Control Unit (PCU) hydraulic supply is provided by the aircraft
hydraulic systems. The slat system No. 1 is supplied by the Blue system,
the slat system No. 2 is supplied by the Green system. If there is a single
system failure the system will still operate but at half speed.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-88
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ATA 27
A300/A310
Wing Leading Edge - Slat Hydraulic System - Schematic
* NOTE: A300 ONLY/A310 DO NOT
HAVE NOTCH OR SLOT FLAPS
*
MTT M540000 R3.3 01AUG01
For Training Purposes Only
*
27-89
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ATA 27
A300/A310
Wing Leading Edge - Slat Control System - General
The power is supplied to the ball screwjacks by a torque shaft driven by a
power control unit and protected by a system torque limiter for each
wing. Each ball screwjack also has its own torque limiter. All these torque
limiters include a latched lockout indicator and in case of overload of a
jack in the torque shaft system, they will freeze the system until a reverse
selection is attempted.
As soon as an order is given, the corresponding computer of each motor
sends signals to deliver the pressure to the motor, releases the involved
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-90
2790
pressure-off brake and controls the sense and speed of movement.
When the selected position is reached, the systems are de-energized,
applying the pressure-off brakes and stopping the movement.
In case of hydraulic failure, the corresponding motor remains locked by
its brake and the operating speed of the slats is reduced by half due to
the differential mechanism of the power unit gearbox. However, full
torque is still available. Three slat positions can be selected (0°, 15°, 30°)
by the five position control lever.
ATA 27
A300/A310
Wing Leading Edge - Slat System Control and Indicating - Schematic
ADC 1
ADC 2
* NOTE: A300 ONLY/A310 DO NOT
*
MTT M540000 R3.3 01AUG01
For Training Purposes Only
HAVE NOTCH OR SLOT FLAPS
27-91
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ATA 27
A300/A310
Wing Leading Edge - Slat - Hydraulic Power Drive System - General
The slat drive system comprises the power control unit (PCU), a
transverse torque shaft system and the screwjacks. In the PCU two
independent hydraulic motors, one controlled by the Blue valve block and
one controlled by the Green valve block, drive a summing gear. The
output is passed through torque shafts to a tee-gearbox which rotates the
motor drive direction by 90°. A pressure-off brake is provided between
each motor and the summing gear to lock the transmission system when
the slat system is static.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-92
2792
The two transverse outputs drive the ballscrew jacks through torque
limiters, a series of torque shafts, steady bearings and gearboxes. One
unidirectional friction brake is installed at each wing tip to provide system
irreversibility under compressive screwjack loads.
ATA 27
A300/A310
Wing Leading Edge - Slats Power Control Unit (PCU) Valve Block - Schematic
D
FW
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-93
2793
ATA 27
A300/A310
Wing Slats and Flaps Systems - Indicating and Control - Description
A single control lever located on the center pedestal permits slat and flap
control. The lever has five gated positions. It is not possible to select an
intermediate position (if the lever is held in between gates the system
drives to the last demanded position and after 10 sec. all the slat and flap
FAULT warnings illuminate). The slats and flaps are electrically signaled
by two identical digital computers (Slats Flaps Control Computers). Each
one is composed of one slat control channel and one flap control
channel.
1.
SLAT/FLAP POSITION INDICATOR STRIPS
Slat and flap positions are shown by White strips moving up and
down associated scales. The corresponding VFE (speed limit) is
placarded opposite each normal position (indicated by a round
number).
2.
SLAT AND FLAP LIGHTS
Come on Amber when the associated system is blocked.
SLATS (OR FLAPS) SYS 1 AND 2 FAULT LIGHTS
3.
Each light comes on Amber when the associated hydraulic motor is
inoperable. Both slats (or flaps) stop due to a system jam. In both cases,
a reverse selection is possible. If system jam is released, the system will
move to the commanded position.
KRUEGER LIGHT
Comes on Amber if either KRUEGER flap is not in correct position
10 sec. after a movement command. Illumination of KRUEGER light
is accompanied by ECAM activation.
4.
SPD BRK LIGHT
The light comes on Blue when the speed brake control lever is not in
RET position.
5.
Flashes Blue when the slat lock function is activated (inhibition of
complete slat and KRUEGER flap retraction at high angle of attack).
Both SLAT (or FLAP) FAULT lights and the associated Amber SLAT (or
FLAP) lights on the Slat/Flap Position Indicator will come on
simultaneously if a mechanical failure is detected. In this case, the
system is locked by the pressure-off brakes and there is no possibility of
recovery in flight.
NOTE: If a SFCC is not installed, the two associated FAULT light (one
SLAT FAULT light and one FLAP FAULT light) will come on.
Illumination of these lights is associated with ECAM activation.
Before selection of any position, the slat/flap control lever must be pulled
up. A block is provided for positions 2 and 4 to prevent the lever moving
straight through.
NOTE: All slat and flap FAULT lights will illuminate if the control lever
remains between two gated positions (after 10 sec.).
NOTE: All Slat and Flap FAULT lights will remain illuminated if Slat/Flap
Control Computer (SFCC) number 1 and 2 are removed from Electronic
Rack 90VU. The 28 Volt DC Interlock Relay in the system prevents
accidental dispatch of the aircraft if both SFCC 1/2 are removed from the
electronic rack or aircraft.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-94
2794
ATA 27
A300/A310
Wing Slats and Flaps Systems - Indication and Controls
B
A
C
0
0
15
0
SLATS
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-95
2795
FLAPS
15
15
15
20
30
40
ATA 27
A300/A310
A300/A310 Wing Leading Edge - Slat - Asymmetry Monitoring
Slat position is indicated to the flight crew by the SLATS vertical bar
display on the slat/flap position indicator. The display is driven by inputs
from the instrumentation Position Pick-off Unit (PPU) which also provides
independent slat position information for other systems.
ASYMMETRY AND POWER TRANSMISSION MONITORING
( SLATS)
The Asymmetry PPUs enable the SFCCs to monitor the transmission for
asymmetry and runaway conditions.
Slat position discrete and digital data are provided for other systems by
the SFCCs.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
If an asymmetry or runaway condition is detected, the PCU operation is
inhibited, preventing further movement of the transmission system.
27-96
2796
ATA 27
A300/A310
A300/A310 Slat Asymmetry Position Pick-Off Unit and Adapter - Component Location
B
B
B
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-97
2797
ATA 27
A300/A310
Wing Leading Edge - Krueger and Notch Flap Actuation
When the slats are extended, SFCC1 and SFCC2 send an extend
discrete signal to a Krueger selector solenoid valve, which has an extend
and retract solenoid. When the extend solenoid is energized, hydraulic
pressure is passed to the Krueger and notch flap actuators and to the all
speed aileron system. The Krueger and notch flaps extend. The solenoid
remains energized.
KRUEGER AND NOTCH FLAP ACTUATION
The Krueger selector solenoid valve is located in the hydraulic bay at
FR47. It is a three-position, four-port solenoid-operated shuttle valve.
The shuttle valve is springloaded to center. It moves to the center
position when both solenoids are deenergized. In this position, the
hydraulic pressure input in A is shut off and ports B, C and D are
interconnected.
When the slats are retracting and have passed the 15° position, the
retract solenoid is energized, the extend solenoid is deenergized and the
Krueger and notch flaps are retracted.
INTERFACE WITH THE AILERON SYSTEM
When the Krueger and notch flaps are supplied with pressure from the
Krueger selector solenoid valve, so also is the droop actuator in the all
speed aileron system. When the Krueger and notch flaps are extended,
the ailerons droop 9.2°. On retraction, the ailerons return to their normal
positions.
NOTE: The A310 Slat System does not have Notch or Slot Flap devices.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-98
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ATA 27
A300/A310
Wing Leading Edge - Krueger and Notch Flap Hydraulic Actuation - Schematic
A300 ONLY
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-99
2799
ATA 27
A300/A310
Wing Leading Edge - Krueger and Notch Flap - Components
KRUEGER FLAP ACTUATOR
The Krueger flap actuator is a double-acting actuator with mechanical
locking of the piston assembly in the extended position (flap retracted)
and hydraulic locking in the retracted position (flap extended). The piston
end is connected to the Krueger flap by a reverse link. The cylinder head
is mounted to the structure by two mounting blocks which allow the
actuator to pivot on the mounting during the operating cycle. The
actuator consists of a cylinder:. f head a cylinder and a valve block.
*NOTCH FLAP ACTUATOR / A300 ONLY
The notch flap actuator is a double-acting actuator and is hydraulically
locked in the retracted position. The piston rod is connected by an
eye-end to the notch flap and the cylinder is attached to the structure by
a shaft hinge to allow some pivoting during the retraction and extension
cycle.
SAFETY VALVE
The safety valve is located in the pressure line from the Green hydraulic
system to the solenoid selector valve. Its function is to prevent loss of
hydraulic fluid from the Green system should there be a major rupture in
the Krueger and notch flap actuating system.
*NOTE: The A310 Slat System does not have Notch or Slot Flap
devices.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-100
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ATA 27
A300/A310
Wing Leading Edge - Krueger and Notch Flap - Components
*
* NOTE: A300 ONLY/A310 DO NOT
HAVE NOTCH OR SLOT FLAPS
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-101
27101
ATA 27
A300/A310
Wing Leading Edge - Krueger and Notch Flap Control and Monitoring - General
In addition, to obtain better aerodynamic characteristics, a KRUEGER
flap and a NOTCH flap are provided on each wing and are located
between the inner slat and the fuselage.
Each KRUEGER flap and AIL droop actuator and each NOTCH actuator
are supplied from a KRUEGER selector solenoid valve supplied by the
Green circuit and controlled by the slats control system.
The KRUEGER and NOTCH flaps are extended when the slat/flap
control lever is moved from position 1 to 2 and remain extended for all
other selected positions. When slats 0° position is selected, the
KRUEGER flaps fold up under the leading edge and the NOTCH flap
retracts into the fuselage.
The KRUEGER jacks are mechanically locked in retracted position and
hydraulically locked in extended position.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-102
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ATA 27
A300/A310
Krueger/Notch Flap Control and Monitoring - Schematic
*
MTT M540000 R3.3 01AUG01
For Training Purposes Only
*
NOTE: A300 ONLY/A310
DO NOT HAVE NOTCH OR
SLOT FLAPS
27-103
27103
ATA 27
A300/A310
Wing Leading Edge - Slats/Krueger Flap Monitoring and Fault Warning System
The SFCC provides continuous monitoring of the slat system and the
Krueger and notch position. Fault warnings are generated for those faults
requiring pilot action or flight crew awareness. The faults are stored in
the SFCC including those which are purely maintenance data. The
warnings are displayed on one or more of the following:
•
•
•
•
SELF TEST
Self test facilities are provided to:
•
•
SLATS SYS 1 FAULT or SLATS SYS 2 FAULT annunciator (19CV)
on the overhead panel
BITE DISPLAY/SFCC1 or SFCC2 annunciators (52CV and 53CV) on
the FLIGHT CONTROL section panel 471VU
Left electronic centralized aircraft monitor (ECAM)
Fault indicator on the SFCC front panel.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
detect and indicate failure in redundant and dormant circuits
identify a failed line replaceable unit (LRU)
The self test can be initiated from the flight deck maintenance panel or by
using the BITE pushbutton switch on the SFCC front panel.
27-104
27104
ATA 27
A300/A310
Fault Indicators and Flight Controls Test Panel - Krueger System
MAINT PANEL 471VU
A300/SOME A310
1/FIN 21CV
2/FIN 22CV
A300 Indicator Shown
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-105
27105
ATA 27
A300/A310
Warning Logic Slat System
Depicted below are the various warnings displayed in the Flight Compartment in the event the fault
shown occurs in the Slat System of the aircraft. So shown graphically are the flight phases at which
the warnings will or will not be displayed.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-106
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ATA 27
A300/A310
A300/A310 Slats System - Warning Logic
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-107
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ATA 27
A300/A310
THIS PAGE INTENTIONALLY
LEFT BLANK
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-108
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ATA 27
A300/A310
WING FLAPS
SECTION
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-109
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ATA 27
A300/A310
A300 Wing Flap System - Description
Each wing has three flap sections. The three flaps are single slotted
fowler type and are guided by two tracks fitted with ball screwjacks. The
power is supplied to the ball screwjacks by a torque shaft driven by a
power control unit and protected by a system torque limiter for each
wing. Each ball screwjack has its own torque limiter.
All these torque limiters include a latched lock-out indicator and, in case
of overload of a jack in the torque shaft system, they will freeze the
system until a reverse selection is attempted. The Power Control Unit
(PCU) consists of two independent hydraulic motors coupled to a
differential mechanical system through pressure-off brakes that insure
the system irreversibility. The two motors are supplied by different
hydraulic circuits (Green and Yellow).
As soon as an order is given, the corresponding computer of each motor
sends signals to deliver the pressure to the motor, release the involved
pressure-off brakes and control the direction and speed of movement.
When the selected position is reached, the systems are de-energized,
applying the pressure-off brakes and stopping the movement.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
In case of hydraulic leakage, the corresponding motor remains locked
and the operating speed of the slats is reduced by half due to the
differential mechanism of the power unit gear box. However, full torque is
still available. Four flap positions can be selected (0°,15°,20°,40°) by
moving the slat flap control lever from position 2 to 5. Furthermore,
between the inboard and center flaps, there is an aileron droop signal
unit which commands the aileron to droop 9.2° maximum with slats
extension to 15°.
A load relief system is provided to minimize the design loads on the flap
support structure and the flap jacks. Load relief function can only engage
when the slat flap control lever is in gate 5. Load relief is activated within
the flap channels of the two SFCCs by using Calibrated Air Speed (CAS)
received from the two ADCs. Load relief logic is the following:
•
if CAS >178 kt, Flaps retract from 33.5° to 24°
•
if CAS <173 kt, Flaps extend from 24° to 33.5°
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ATA 27
A300/A310
A300 Wing Flap System - Components/Location
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-111
27111
ATA 27
A300/A310
Wing Flap System Hydraulics
The PCU hydraulic supply is provided by the aircraft hydraulic systems.
One motor is supplied from the Green system and one is supplied from
the Yellow system. If there is a single-system failure, one of the motors
will drive the system at half speed.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-112
27112
ATA 27
A300/A310
Wing Flaps Hydraulic System - Schematic
POWER CONTROL UNIT (PCU)
INSTALLATION LOCATION FUSELAGE FRAME (FR) 54
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-113
27113
ATA 27
A300/A310
Wing Flap System - Power Control Unit (PCU)
The flap drive system consists of two identical hydromechanical systems,
comprising two hydraulic motors, a differential gear and pressure-off
brakes, contained within a power control unit, (PCU) and a transverse
torque shaft system driving the ballscrew jacks. Static and dynamic
system irreversibility is provided by no-back friction brakes within
screwjacks 2, 4, 5 and 6. Screwjack 3 incorporates only one no-back
which operates only during flap extension under tensile loads. Screwjack
1 is not provided with no-back friction brakes.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-114
27114
ATA 27
A300/A310
Wing Flap System PCU - Schematic
FWD
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-115
27115
ATA 27
A300/A310
Wing Flap PCU Components
The PCU is located in the hydraulic compartment and provides the
driving power to the flap system.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
It consists of the following subassemblies:
•
Acceleration control valve
•
Valve blocks
• Transfer tubes
•
Hydraulic motors
•
Pressure-off brakes
•
Differential gearbox
•
Intermediate gearbox
•
Position pickoff units (PPU)
27-116
27116
ATA 27
A300/A310
Wing Flap PCU - Components and Valve Block
FW
D
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-117
27117
ATA 27
A300/A310
Wing Flap Controls
The flap system is controlled by two identical slat flap control computers
(SFCC1 and SFCC2). The SFCCs provide, in addition, monitoring and
test facilities.
Included in the monitoring is a flap relief function in which the flaps are
automatically retracted from a fully extended position should the aircraft
MTT M540000 R3.3 01AUG01
For Training Purposes Only
exceed the maximum flap extended speed (VFE) limitation for the flap
configuration. If the control lever setting has not been altered, the flaps
will automatically extend when the aircraft speed has been reduced
sufficiently. The air speed data is provided by the air data computer
(ADC).
27-118
27118
ATA 27
A300/A310
A300 Wing Flaps - Control System - Schematic
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-119
27119
ATA 27
A300/A310
Wing Flap Controls
The flap system is controlled by two identical slat flap control computers
(SFCC1 and SFCC2). The SFCCs provide, in addition, monitoring and
test facilities.
Included in the monitoring is a flap relief function in which the flaps are
automatically retracted from a fully extended position should the aircraft
MTT M540000 R3.3 01AUG01
For Training Purposes Only
exceed the maximum flap extended speed (VFE) limitation for the flap
configuration. If the control lever setting has not been altered, the flaps
will automatically extend when the aircraft speed has been reduced
sufficiently. The air speed data is provided by the air data computer
(ADC).
27-120
27120
ATA 27
A300/A310
A310 Wing Flaps - Control System - Schematic
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-121
27121
ATA 27
A300/A310
Warning Logic Flap System
Depicted below are the various warnings displayed in the Flight Compartment in the event the fault
shown occurs in the Flap Control System of the aircraft. Also graphically shown are the Flight Phases
at which the warnings will or will not be displayed.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-122
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ATA 27
A300/A310
A300 Flap System - Warning Logic
MTT M540000 R3.3 01AUG01
For Training Purposes Only
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ATA 27
A300/A310
THIS PAGE INTENTIONALLY
LEFT BLANK
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-124
27124
ATA 27
A300/A310
SPEEDBRAKES AND SPOILERS
SECTION
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-125
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ATA 27
A300/A310
Speed Brakes and Ground Spoilers
SPEEDBRAKES
There are two inner and three outer speedbrakes located on the upper
surface of each wing. The outer speedbrakes are also used as roll
spoilers. They are selected by a lever situated on the center pedestal.
Eleven positions can be selected from RET (retracted) to FULL (fully
extended) position.
Each speedbrake element is powered by one servo control which
receives pressure from an electrohydraulic valve group controlled by the
corresponding units of the EFCUs (Electrical Flight Control Units) and
also used for roll spoilers control.
"SPEED BRAKES EXTENDED" indication is given on the ECAM MEMO
page.
The ground spoilers are selected when the two following conditions are
fulfilled:
• SPEED BRAKE control lever pulled upwards (when it is in RET
position) or thrust reverser selected on one engine.
• both engine throttle levers in idle position.
The "aircraft on ground signal" is sent:
• during takeoff or landing when the two main landing gear aft wheel
speed is higher than 85 kt
• at landing, only if the ground spoiler preselection has been made by
speedbrake control lever selection, when:
main gear bogie beam is in the ground position.
radio altitude is lower that 5 ft.
These signals are inhibited 3 seconds after first shock absorber
compression.
GROUND SPOILERS
Speedbrakes and roll spoiler surfaces are used on the ground as ground
spoilers. Deflection angles become 50° for all surfaces. They are
automatically extended when:
•
they are selected.
•
the aircraft is on the ground.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
Automatic extension is also achieved for an aborted take off when these
conditions are fulfilled. Ground spoilers will remain extended during
bounces as long as both throttles are in the idle position and the SPEED
BRAKE control lever is pulled up.
Ground spoiler retraction after landing is achieved:
• Either by pressing the SPEED BRAKE control lever down
(preselection cancellation)
• or by moving one throttle lever out of the idle position.
27-126
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ATA 27
A300/A310
A300/A310 Speed Brakes and Ground Spoilers - Schematic
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-127
27127
ATA 27
A300/A310
Speed Brake Control Lever - Description
A control lever assembly located in the center pedestal enables the crew
to select the position of the speed brake surfaces arid to preselect the
ground spoiler function.
The control assembly is composed of:
• A control lever which drives the speed brake control transducer unit
via mechanical linkage.
• A fixed quadrant with eleven notches which lock the SPEED BRAKE
control lever in the selected position.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
Lever motion is guided by a slot. True slot prevents ground spoiler
preselection if the control lever is not in the RET position.
•
•
•
27-128
27128
A spring device which holds the control lever extended when the
ground spoiler function is preselected.
A cam which drives the ground spoiler preselection microswitches.
A spring which holds the cam in the "ground spoiler not preselected"
position when the control lever is not pulled upwards.
ATA 27
A300/A310
Speedbrake Control Lever - Selection Positions
CONTROL
LEVER
Speedbrake
Spoilers
Retracted
Speedbrake
Spoilers
Extended
FIXED
QUADRANT
CAM
MIROSWITCHES
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-129
27129
GUIDE
SLOT
ATA 27
A300/A310
Speedbrake Controls
•
A. SPLR & SPD BRK PUSHBUTTON SWITCHES
1.
Each P/B switch is associated with one or two pairs of
symmetrical upper wing surfaces.
•
•
ON (P/B switch pressed-in): The corresponding control
system is activated. Each time a system is activated, or
corresponding hydraulic system on, or the aircraft electrical
network is energized, a 2 sec. safety BITE test is triggered
for the corresponding EFCU units (control and monitor).
B. SPEEDBRAKE CONTROL LEVER
OFF/R (P/B switch released-out): The OFF/R light comes on
White and the corresponding control system is deactivated. If
hydraulic pressure is available, the actuators are
automatically held in the retracted position. The monitoring
circuits are reset by this action. This indication is
accompanied by ECAM activation.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
FAULT: When a P/B switch is pressed-in, the associated
FAULT light comes on Amber if a failure is detected by the
monitoring circuits which then deactivate the control system.
Illumination of the FAULT light is accompanied by SCAM
activation. The lever controls:
the position of the speedbrake eleven surface positions
from retracted (RET) to fully extended (FULL).
manual preselection of the ground spoiler function.
27-130
27130
2.
SPEEDBRAKE SELECTION
To select the speedbrake surfaces to the required position, the
pilot must press on the top of the lever and move it to the
corresponding notch. The control lever cannot be moved as long
as the ground spoiler function is preselected.
3.
GROUND SPOILER OPERATION
To arm the ground spoiler function, the control lever must be
lifted when in retracted position (RET).
ATA 27
A300/A310
Speedbrake Controls - Control Switches and Handle
B
A
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-131
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ATA 27
A300/A310
THIS PAGE INTENTIONALLY
LEFT BLANK
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-132
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ATA 27
A300/A310
MISCELLANEOUS FLIGHT
CONTROL AREAS
SECTION
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-133
27133
ATA 27
A300/A310
Flight Control Surface Position Indicating - FLT CTL Page
The flight control page of the system display (right SCAM display unit) is
called either by manual selection, or automatically when a warning
occurs on a system covered by this page.
1.
2.
3.
When the surface is completely retracted, it is symbolized by a small
dash.
When the surface is deflected more than 2°, a small arrow appears
above the dash.
B,G,Y SYMBOLS
Each available system on the flight controls and trimmable horizontal
stabilizer is indicated by a Green symbol. In case of servo control low
pressure detection, the corresponding symbols become Amber.
In normal conditions, these indications are Green. in case of failure
detected for a group of surface in the EFCUs, the numbers
identifying the surfaces involved appear below the dash.
PRIMARY FLIGHT CONTROLS SURFACE POSITION DISPLAY
A white scale covering the full travel range is provided for each
surface (elevators, rudder, all speed ailerons and trimmable
horizontal stabilizer). An index indicating the actual position of the
surfaces moves along each scale.
In case of hydraulic failure when a surface is extended, the color of
the corresponding symbol changes from Green to Amber and the
number appears below the symbol.
On the ground, after landing, the arrow of any surface extended
flashes as long as speedbrake surfaces No. 1 and 2 are not fully
retracted.
ROLL SPOILER AND SPEEDBRAKE SURFACE POSITION
DISPLAY
Each roll spoiler and speedbrake surface is represented as follows:
MTT M540000 R3.3 01AUG01
For Training Purposes Only
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ATA 27
A300/A310
Flight Control Surface Position Indicating - FLT CTL Page
Aircraft banking to right
MTT M540000 R3.3 01AUG01
For Training Purposes Only
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ATA 27
A300/A310
Flight Control Surface Position Indicating - WHEEL Page
The RH ECAM WHEEL Page called up automatically at landing
providing that the Spoiler Preselection Conditions are valid on aircraft
touchdown on the runway (Spoiler display on bottom of WHEEL Page).
Each spoiler is represented in the same manner as on the FTL CTL
(Fight Control) Page.
On the ground, the arrow corresponding to any spoiler extended after
landing and the control system has not failed, flashes as long as Spoilers
1 and 2 are not fully retracted.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
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ATA 27
A300/A310
Flight Control Surface Position Indicating - WHEEL Page
MTT M540000 R3.3 01AUG01
For Training Purposes Only
27-137
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ATA 27
A300/A310
Description Of Control Surface Position Indicating For The Three Axes
On all three axes synchro transmitters send control surface position
signals to the right ECAM display unit:
• Right all speed aileron: 6CT
• Left all speed aileron : 7CT
• Rudder: 10CT
• Trimmable horizontal stabilizer: synchro in position sensor: 11CT
• Elevator (right): 12CT
MTT M540000 R3.3 01AUG01
For Training Purposes Only
Travel of the above surfaces is indicated by displacement of an index,
symmetrically in the case of the ailerons. For the spoilers, pulses from
the EFCU are displayed on the right ECAM display unit by illumination of
the relevant spoiler arrows.
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ATA 27
A300/A310
Flight Control System - Control Surface Position Indicating Sensors - Location
• ASA (LH/RH Side) (2)
• Rudder
(1)
• Elevators
(1)
MTT M540000 R3.3 01AUG01
For Training Purposes Only
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A300/A310 Aircraft Takeoff (T/O) Configuration Test
When T/0 CONFIG TEST pushbutton is pressed and held, T/O power
application is simulated. This test will activate the appropriate warnings if
the aircraft is not at take-off configuration.
•
SLATS OR FLAPS are not in Take Of f Configuration. The Red T/O
CONFIG light comes on WLDP with associated CRC and ECAM
activation.
The warnings are canceled when the pushbutton is released. Flight
Controls conditions for warning and resultant warnings are:
•
SPEEDBRAKES OR GROUND SPOILERS are extended. In this
case, illumination of the Blue SPD BRK light on SFPI is accompanied
by the Red T/O CONFIG light on WLDP with associated CRC and
ECAM activation.
•
HORIZONTAL STABILIZER is not in Take Off Configuration (>3°UP
± 0.4 or > 2.3°DN ± 0.4). The Red T.O. CONFIG light comes on
WLDP with associated CRC and ECAM activation.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
Pressing T/O CONFIG TEST pushbutton also monitors the following
systems:
• DOORS (when not closed)
• LANDING GEAR (parking brake, brake temperature)
• PROBE HEAT (Standby or CAPT or F/O probes heat off).
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A300/A310 Aircraft Takeoff Configuration Test
MTT M540000 R3.3 01AUG01
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Maintenance Panel - Flight Controls Test Panel Pushbuttons
Only used on the ground to test the jamming detection microswitches,
with pressure for the respective circuit shut off.
1.
2.
B,G,Y TEST PUSHBUTTON SWITCHES
After selecting the related SERVO CTL P/B switch to OFF (on
overhead panel), the TEST P/B switch for the respective circuit is
magnetically latched when pressed-in and the TEST light comes on
White.
Left pushbutton tests pitch feel and rudder travel system 1. Right
pushbutton tests pitch feel and rudder travel system 2. When a TEST
pushbutton is pressed and held, the associated system must
disengage and its FAULT light comes on Amber.
For jamming detection test, the controls involved must be moved
rapidly. Successful test is indicated by flashing of the JAM light in the
related SERVO CTL P/B switch. If not successful, the fault isolation
procedure must be done on the face of the jamming detection control
box.
After selecting the SERVO CTL P/B switch to normal, the TEST P/B
switch is automatically released-out and the TEST light goes off.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
TEST PUSHBUTTONS
The pushbuttons control the test of PITCH FEEL and RUD TRAVEL
electrical systems and warning systems continuity. The test is
possible only if PITCH FEEL and RUD TRAVEL systems are
engaged on the control panel (overhead panel).
Successful test is indicated by White OK lights illumination.
• upper lights for PITCH FEEL SYS 1 and 2
• lower lights for RUD TRAVEL SYS 1 and 2.
3.
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OK LIGHTS
These lights illuminate White as long as the TEST pushbutton is
pressed and held, to indicate a successful test.
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Maintenance Panel - Flight Controls Test Panel Pushbuttons
2
3
1
MTT M540000 R3.3 01AUG01
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Maintenance Panel - Flight Controls Test Panel - Controls and Indications
1.
2.
TEST SELECTOR
• NORM FLT: Normal operating position, test circuits
disconnected, warnings canceled.
• GND SPLR: Checks that no undue condition is permanently
achieved in the ground spoilers logic for the EFCU involved when
the TEST P/B switch is pressed-in.
SEL LANE TEST 1 and 2 positions check the integrity of
selection lanes aircraft wirings.
EFCU TEST 3 and 4 positions check the integrity of each
EFCU logic. The corresponding FAULT lights on the SPLR &
SPD BRK panel will go off. This test requires all hydraulic
power to be cut off to all flight controls to have the FAULT
lights illuminated before test.
• PITCH CTL UNCOUPLING: Tests periodically, on ground, the
electrical circuits of the pitch uncoupling unit.
TEST 1 checks that the uncoupling unit rod is in the locked
position.
TEST 2 checks that the uncoupling unit moves to the
unlocking position, when the control solenoid is energized.
• SLATS/FLAPS: Commands a BITE sequence for the relevant
SFCC (SYS 1 or SYS 2) when the TEST P/B switch is pressed
in.
GND SPLR SEL LANE FAULT LIGHT
This light comes on White when a fault has been detected in TEST 1
or TEST 2 positions of the test selector.
MTT M540000 R3.3 01AUG01
For Training Purposes Only
3.
EFCU BITE DISPLAY LIGHT
This light comes on White when a fault has been detected by the
continuous monitoring of each EFCU. More details of the failure are
displayed of the face of the EFCUs.
4.
PTT PUSHBUTTON SWITCH
This P/B switch activates the test of the system selected by the test
selector. A TEST indication is integrated into the P/B switch.
•
PTT: When pressed-in and held, the selected system is tested.
•
TEST: The light comes on White when the test selector is set to a
system test position. It is extinguished when the test selector is in
NORM FLT position.
5.
SFCC 1 AND 2 BITE DISPLAY LIGHTS
These lights come on White when a fault has been detected by the
continuous monitoring of the SFCCs even if the failure does not
require crew action (no FAULT indication on the overhead panel).
More details of the failure are displayed on the face of the SFCC's.
6.
TEST RESULT OK LIGHT
This light comes on White when the test is successful.
NOTE: These lights (OK, FAULT, BITE, DISPLAY) will illuminate
providing that the ANN LTS switch is in READ position during the test.
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Maintenance Panel - Flight Controls Test Panel - Controls and Indications
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THIS PAGE INTENTIONALLY
LEFT BLANK
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COURSE CODE - M540000
MTT M540000 R3.3 01AUG01
For Training Purposes Only
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