Flight controls

Question 1

How are the Flight Controls operated?

Answer 1

The flight controls are operated conventionally with control wheels, control columns and rudder pedals for the captain and first officer. The control surfaces are either actuated hydraulically or electrically. The flight control systems include major control surfaces, components and subsystems that control the attitude of the aircraft during flight. The flight controls are divided into primary and secondary flight controls.

Question 2

Primary flight controls include?

Answer 2

elevators
rudder
ailerons

The ailerons, elevators and rudder are controlled by a network of cables, pulleys, push/pull rods and levers that transmit control inputs to the related system hydraulic power control units (PCUs)
The aileron and elevator controls are equipped with control disconnects which permit the captain or first officer to maintain sufficient lateral and longitudinal control in the event of a control jam. The rudder control is equipped with an anti-jam mechanism that permit both pilots to maintain sufficient directional control, however, additional force is required to obtain surface travel.
In the event of a total electrical power failure, the primary flight controls remain hydraulically powered by AC Motor Pump (ACMP) 3B. In an emergency, ACMP 3B will be electrically energized directly from the Air Driven Generator (ADG).

Question 3

The secondary flight controls consist of:

Answer 3

horizontal stabilizer trim
slats/flaps
multi-function spoilers and
ground spoilers along with the associated aileron and rudder trim systems

Question 4

Primary Flight controls cont’d

Answer 4

Ailerons and Multifunction Spoilers
Two separate lateral control systems are provided. The captain operates the left aileron and the first officer operates the right. Normally the aileron controls are interconnected and there is simultaneous and coordinated movement of all lateral control surfaces from either pilot station.
Turning either control wheel sends a mechanical signal (via cables and pulleys) to the aileron hydraulic control units. Two PCUs are used for each aileron.
Moving the control wheels also generates an electrical signal that is sent to the Spoiler Stabilizer Control Unit (SSCU 1 and SSCU 2). Dual redundant modules within each SSCU control the extension/retraction of the multifunction spoilers (MFS). The SSCUs combine the control wheel signals with other information to determine the required multifunction spoiler panel deflection for any given airplane configuration. A single PCU is used on each multifunction spoiler. The MFS operate on the down-going wing only to assist the ailerons in roll control at lower speeds.

Question 5

Ailerons, elevator and rudder are controlled by what?

Answer 5

by a network of cables, pulleys, push/pull rods and levers that transmit control inputs to the related system hydraulic power control units (PCUs)
The aileron and elevator controls are equipped with control disconnects which permit the captain or first officer to maintain sufficient lateral and longitudinal control in the event of a control jam. The rudder control is equipped with an anti-jam mechanism that permit both pilots to maintain sufficient directional control, however, additional force is required to obtain surface travel.

The primary flight controls are arranged conventionally with rudder pedals and a control wheel and column for both the captain and first officer. Movement of the cockpit controls is transmitted mechanically via cable and/or push rods to the aileron, elevator and rudder power control units (PCU) that hydraulically move the control surfaces.

Question 6

In the event of a total electrical power failure

Answer 6

the primary flight controls remain hydraulically powered by AC Motor Pump (ACMP) 3B. In an emergency, ACMP 3B will be electrically energized directly from the Air Driven Generator (ADG).

Question 7

Ailerons and Multifunction spoilers.

Answer 7

Two separate lateral control systems are provided. The captain operates the left aileron and the first officer operates the right. Normally the aileron controls are interconnected and there is simultaneous and coordinated movement of all lateral control surfaces from either pilot station.
Turning either control wheel sends a mechanical signal (via cables and pulleys) to the aileron hydraulic control units. Two PCUs are used for each aileron.
Moving the control wheels also generates an electrical signal that is sent to the Spoiler Stabilizer Control Unit (SSCU 1 and SSCU 2). Dual redundant modules within each SSCU control the extension/retraction of the multifunction spoilers (MFS). The SSCUs combine the control wheel signals with other information to determine the required multifunction spoiler panel deflection for any given airplane configuration. A single PCU is used on each multifunction spoiler. The MFS operate on the down-going wing only to assist the ailerons in roll control at lower speeds.

Question 8

Flutter Dampers

Answer 8

A flutter damper is installed on each of the ailerons. These double-acting shock absorbers prevent aileron control surface flutter when all hydraulic fluid is lost at the PCU during flight. On the ground, the flutter dampers provide a gust lock function. The aileron PCUs provide additional gust lock protection.

Question 9

Aileron Disconnect

Answer 9

Roll disconnect allows the flight crew to isolate the left control wheel and associated cable system from the right. Pulling the ROLL DISC handle separates the control wheel interconnect (torque tube) and advises the SSCU that the interconnect torque tube has been disconnected. Single- side roll control is then available. When the handle is pulled, a cross-side aileron/ MFS relationship is established. The captain moves the left aileron and right MFS; the first officer moves the right aileron and left MFS.
Pulling the ROLL DISC handle can isolate a jammed aileron control system. Pulling the handle isolates the faulted aileron system and provides the pilot with reduced lateral control (one aileron and opposite side MFS only) through the operable aileron system. Twenty seconds after pulling the handle, the SSCU commands two amber ROLL SEL lights on the glareshield to illuminate, and the EICAS caution SPOILERONS ROLL message to appear on the primary page.
Selecting the ROLL SEL switch on the side with the unjammed aileron provides the flying pilot with the use of the second spoileron. Pressing the ROLL SEL switch also removed the amber glareshield lights and caution message and replaces them with a green ROLL SEL glareshield light and the advisory message PLT ROLL CMD or CPLT ROLL CMD.

If uncommanded displacement of an aileron PCU occurs, a bungee breakout switch associated with the runaway aileron system sends a signal to the SSCU. The SSCU interprets the signal then commands both MFS to respond to control wheel inputs. The SSCU presents the advisory message PLT ROLL CMD or CPLT ROLL CMD and illuminates the green ROLL SEL glareshield light in front of the captain that should take control prior to ordering the ROLL DISC handle to be pulled.

Question 10

How many PCU are installed on each Aileron?

Answer 10

2

Question 11

How many PCUs are installed on the rudder

Answer 11

3

Question 12

Roll disconnect handle

Answer 12

Removes the interconnect feature of the control wheels.

To disconnect:
Pull and lock Roll Disconnect handle into the disconnect position.
When the Roll Disconnect is pulled, both pilots cable roll cable runs are separated.

Question 13

Glareshield Roll Select switchlights

Answer 13

Left and right ROLL SEL illuminates when disconnect is pulled.

Question 14

The green PLT ROLL lights indicate?

Answer 14

Both MFS on the wings are available.

Question 15

Rudder

Answer 15

The rudder provides directional control about the vertical axis. The rudder is hydraulically powered by 3 PCUs and controlled via cable runs and quadrants through displacement of either pilot’s rudder pedals. Displacement of either set of pedals mechanically sends a signal to the three hydraulic PCUs to move the rudder.
Two separate cable run systems with anti-jam/breakout protection (spring tension breakout) are provided. In the event of a jammed rudder control, both the captain’s and first officer’s rudder pedals remain operable, however additional pedal force will be required to move the rudder.
On the ground, trapped hydraulic fluid provides rudder control surface gust lock damping when the hydraulic systems are depressurized.
Rudder position is displayed on the EICAS FLIGHT CONTROLS synoptic page. A full-scale deflection of the rudder position indicator corresponds to maximum rudder travel.

Question 16

Yaw damper

Answer 16

Two independent yaw damper systems operate continuously in flight to improve the airplane’s directional stability and turn coordination.
Yaw damping improves the aircraft’s stability by damping out oscillations in yaw. These oscillations if not corrected could lead to the phenomenon called Dutch Roll.
The turn coordination function helps the aircraft into and out of turns. The yaw dampers operate independently of the autopilot system. The Flight Control Computers (FCCs) provide signals to operate the yaw damper actuators. Each computer commands its own yaw damper linear actuator. FCC 1 controls YD 1 and FCC 2 controls YD 2.
The yaw dampers are engaged by pushing the YD 1 and YD 2 ENGAGE switchlights on the YAW DAMPER panel. Each switchlight engages its associated yaw damper channel. The DISC push button when selected disengages the yaw damper.
When a yaw damper channel fails or disconnects, the YD 1(2) INOP status message is displayed. Should both yaw damper channels fail, the autopilot if engaged disconnects and a YAW DAMPER caution message is displayed.

Question 17

Elevators

Answer 17

Pitch control is provided by the elevators and supplemented by a moveable horizontal stabilizer. Each elevator is hydraulically powered by three PCUs and mechanically controlled via cable runs and quadrants through fore and aft displacement of either control column. Two pitch control systems are provided. The captain operates the left elevator and the first officer operates the right elevator. Normally, the control systems are interconnected and there is simultaneous movement of both elevators.
Elevator position is displayed on the EICAS FLIGHT CONTROLS synoptic page. A full-scale deflection of the elevators position indicator corresponds to maximum travel.

Question 18

Pitch disconnect

Answer 18

The PITCH DISC handle when pulled is used to isolate a jammed elevator control system. Pulling the handle provides the pilots reduced pitch control (one elevator only) through the operable control surface.

The PITCH DISC removes the interconnect feature of the control wheels.

When the PITCH DISC is pulled the Pilot controls the left elevator and the copilot controls the right elevator.

Question 19

Horizontal Stabilizer Trim

Answer 19

The stabilizer trim control system provides pitch trim by varying the horizontal stabilizer’s angle of incidence from +2 degrees (nose down) to -13 degrees (nose up). The stabilizer is positioned by a screw jack driven by two electric trim motors and controlled by the Spoiler Stabilizer Control Unit (SSCU). The control unit has two channels that are engaged by the CH1 or CH2 switchlights on the STAB TRIM control panel located on the center pedestal. Each trim motor has a brake to prevent trim runaway.
The SSCU receives inputs from the control wheel trim switches, the autopilot (AFCS), and the Mach Trim System. Emergency trim disconnect switches (PITCH TRIM DISC) are provided on each control wheel.

Operation of the horizontal stabilizer is continuously monitored and any fault detected is displayed on the appropriate EICAS screen. Should a single stab trim channel fail, a status message is displayed. When both stab trim channels fail, a caution message is displayed.

Question 20

Horizontal stabilizer trim priority

Answer 20

The stabilizer trim can be moved manually by control wheel trim switches or automatically via computer-controlled inputs from the AFCS or the Mach trim system. During AFCS operation, trim rate is influenced by flap movement.

If the horizontal stabilizer is in motion at the high or slow
rate for more than 3 seconds a clacker is activated to alert the pilots of a possible horizontal stabilizer trim runaway condition.

Question 21

Stabilizer trim control in order of priority is

Answer 21

Captain Manual Trim – The captain stab trim switches are highest in the order of priority. Selection of the captain’s trim switches moves the stabilizer at the highest rate of movement.

First officer Manual Trim – The first officer’s switches are next in the order of priority. The stabilizer moves at the highest rate.

Autopilot – The autopilot can trim the stabilizer at two rates. The high rate of movement occurs when the flaps are extending or retracting. The low rate of movement occurs when the flaps are stationary.

Auto Trim – As the flaps are extended or retracted, trim commands are provided by the SSCU’s to compensate for pitch changes in response to flap configuration change. The slat/flap electronic control unit (SFECU) provides flap information to the SSCU’s. During the extend selection from slats /flaps up, the slats will extend before flap movement. Trimming of the horizontal stabilizer will not occur until the flaps move (after slat extension).

Mach Trim – The automatic Mach trim is last in the order of priority. With the autopilot off and no pilot stab trim commands inputted, the Mach trim automatically moves the stabilizer to compensate for changes in airspeed. Mach trim moves the stabilizer at the slowest rate.

Question 22

Mach Trim

Answer 22

The Mach trim function makes allowances for the rearward shift of the aerodynamic center of pressure as the Mach number increases. Without correction, this shift in the center of pressure causes a negative stick force gradient and decreases longitudinal stability (Mach tuck) above Mach 0.4 when hand-flying the airplane.
The SSCU automatically adjusts the stabilizer as a function of the Mach number when the autopilot is not engaged. Both FCCs must signal the need for trim before any trim activity occurs.
The Mach trim system, using Mach speed information from the Air Data Computers (ADC), varies the angle of incidence of the stabilizer by commanding movement of the horizontal stabilizer actuator. Both channels of the SSCU must be operable and at least one STAB channel engaged for the Mach trim to function. Mach trim is selected by engaging the MACH TRIM\ switchlight located on the center pedestal.
The Mach trim system is continuously monitored and any fault detected is displayed on the primary EICAS page.

Question 23

Slats and flaps

Answer 23

The double-slotted flaps move rearwards and down when extending. The outboard flaps have fixed leading edge vanes and cams to operate the Bent Up Trailing Edge (BUTE) doors. The BUTE doors are used to direct airflow over the leading edge vanes during flap extension. The inboard flaps have spring-loaded leading edge vanes that automatically extend when the flaps are deployed. The slat/flap system receives power from the aircraft AC busses during normal operation and from the ADG, the system automatically configures to half speed to reduce power drain.
The Slat Flap Electronic Control Unit (SFECU) receives electrical commands from the slat flap control lever to initiate flap movement. When the SFECU commands a change in flap position, the flap brakes are released and the Power Drive Units (PDUs) mounted on a flap gearbox are energized. The gearbox rotates flex shafts to move the flap ballscrew actuators, which extend and/or retract the slats/flaps. When the desired setting is reached, the PDUs are de-energized and flap brakes are applied. The flaps are mechanically connected for simultaneous movement of the inboard and outboard flap sections.
An emergency flap switch is installed on the center pedestal to allow limited flap selection in the event of a mechanical failure of the slat/flap selector lever. When actuated, the EMER FLAP switch input to the SFECU overrides the flap control lever input and causes the slats to fully extend and the flaps to move to 20 degrees. The EMER FLAP switch input to the SFECUs is inhibited at speeds greater than Vfe flaps 20.
If the switch is returned to normal during flight, the SFECU will command the slats and flaps to the flap control lever selected position, if different than the slat extended with flaps at 20 degree position. If the switch is returned to normal while on ground, the slats and flaps will remain in the slat extended with flaps at 20 position regardless of the selector position. The on-ground scenario requires that the slat/flap selector be moved out of detent to restore flap control lever operation.

Question 24

Flap position indications are displayed on the EICAS primary page whenever one or more of the following conditions exists.

Answer 24

flaps are greater than zero degrees
• landing gear is not up and locked
• Brake Temperature Monitoring System (BTMS) is in the red range of operation.

Question 25

Spoiler stabilizer control unit (SSCU)

Answer 25

The control wheels and flight spoiler control lever sends signals to the Spoiler Stabilizer Control Units (SSCU 1 and SSCU 2). Dual redundant modules within each SSCU control the extension/retraction of the multifunction spoilers and ground lift dumping system.
In addition to the MFS and ground spoilers, the SSCU provides control signals to the horizontal stabilizer trim unit (HSTU) and the rudder travel limiter (RTL). The SSCU provides automatic power up self-test when the aircraft is on the ground and continuous system monitoring during all phases of operation.

Question 26

CPOST Computer On Self Test

Answer 26

When AC electrical power is initially applied to the aircraft the CPOST is conducted to verify the integrity of selected internal components of the SSCUs and other electronic hardware. The test is automatic and requires no pilot action.

Question 27

SPOST System Power on Self Test 1

Answer 27

SPOST 1 is performed immediately after the CPOST. It tests the integrity of the circuitry within each SSCU module and other interfaces that are external to the SSCU. The test is automatic and requires no pilot action.

Question 28

SPOST 2 System Power on Self Test

Answer 28

After a successful SPOST 1, following every 50th flight cycle, SPOST 2 is initiated when all three hydraulic systems are pressurized and both stabilizer trim channels (CH1 and CH2) are engaged. SPOST 2 checks the following flight subsystems:

Question 29

Roll Assist

Answer 29

Roll assist is provided through two pairs of multi-function spoilers (MFS) that operate separately to assist the ailerons to provide roll control. Each panel is electronically controlled, by the SSCU, and operated by a single hydraulically actuated power control unit (PCU). The pilots inputs to the roll assist system are through the spoiler control lever, the control wheels, and the two roll priority switch/lights located on the glare shield. Asymmetric deployment of the MFS is a function of captain and first officer control wheel deflection as sensed by two roll control input modules (RCIMs).

Question 30

Ground Lift Dumping System

Answer 30

Description
After touchdown or during a rejected takeoff, the inboard and outboard ground spoilers in conjunction with the multi-function spoilers are extended to spoil lift and increase drag to assist in aircraft braking. The ground lift dumping system (GLD) is normally automatic but can be activated manually by the pilot. The ground spoilers have no in-flight function.

Components and Operation
In the automatic and manual modes, arming, deployment and retraction is controlled by the two SSCUs. The SSCU will command operation of all spoiler panels to deploy if the GLD switch is armed, and two out of the following three conditions are satisfied:
• proximity sensor electronic unit (PSEU) (weight on wheels)
• anti-skid control unit (wheel speed greater than 16 knots)
• radio altimeter (7 feet)
In the event of malfunction, the GLD system can be manually disarmed through the SPOILER panel located on the center pedestal.
The position of all panels is shown on the EICAS F/CTL synoptic page. White triangles and position scales indicate relative deflection of the surfaces.
Inboard and outboard ground lift dumping devices are limited to two positions, either up or down. The spoilers are continuously monitored and any fault detected is displayed on the appropriate EICAS screen.

Question 31

GLD Arming Logic

Answer 31

The GLD circuit must be armed before deployment can take place. Arming can be accomplished automatically or manually.
GLD deployment during the landing or rejected takeoff is automatic. Should automatic deployment fail, the GLD devices can be manually deployed.
After landing or a rejected takeoff, the GLD devices automatically retract in accordance with SECU logic. A manual retract function is also provided.
During a touch and go sequence, the GLD devices will deploy when all extension parameters are met. Advancing the thrust lever for takeoff retracts the GLD devices and re-arms the system.

Question 32

Takeoff Configuration warning system

Answer 32

Description
The Takeoff Configuration Warning System monitors the position of the flaps, flight spoilers, parking brake, autopilot, aileron, rudder and horizontal stabilizer trims, to ensure that they are in a safe configuration for takeoff.

Components and Operation
The warning system is armed when the airplane is on the ground. When the airplane is in a safe takeoff configuration, a T/O CONFIG O/K advisory message is displayed. The T/O CONFIG O/K advisory message is removed from the status page upon airplane rotation.
If one of more of the monitored systems is in an unsafe takeoff configuration and both engines are accelerated (N1 greater than 70%), the master warning lights flash, aural alerts sound and warning messages are presented. The configuration warning indications are cancelled by correctly positioning the applicable control or retarding the thrust levers.

Question 33

Proximity Sensing Electronic Unit PSEU

Answer 33

The proximity sensing electronic unit (PSEU) processes information received from the ground spoilers, thrust levers and flap position and provides instructions to other aircraft systems.
Proximity switches and sensors measure the physical relationship between two aircraft components. If the components are close together, the sensors and/or switches provide a “near” signal. If the components are separated, a “far” signal is generated. Microswitches communicate information with regards to control selection. The PSS uses this information to generate appropriate signals for use by other aircraft subsystems.
The PSEU receives inputs from the sensors and controls the takeoff
configuration warning system and GLD deployment (through the SECU).
The PSEU also provides instruction to the following systems:
• SPS - Stall Protection System - Stick shaker and pusher are disabled on the ground and at rotation.
• HSTCU - Horizontal stabilizer trim control unit Built In Test (BITE), disabled in flight.
• FECU - Flap electronic control unit prevents reset of the flap asymmetry during flight and enables preflight test on ground.
• SECU - Spoiler electronic control unit can initiate automatic GLD deployment at touchdown with wheel spin-up and radio altitude less than 5 feet.
Failure of the proximity sensing system is indicated on EICAS.

Question 34

Stall Protection System SPS

Answer 34

Description
The stall protection system (SPS) provides the flight crew with aural, visual, and tactile (stick shaker) indications of an impending stall. If the captain does not take corrective action the system activates the stick pusher mechanism to prevent the airplane from entering the stall.

Components and Operation
A dual-channel stall protection system (SPS) computer monitors the following inputs:
• angle of attack - LH, RH Angle Of Attack (AOA) transducers (vanes)
• lateral acceleration - Attitude Heading Reference Systems (AHRS) or optional Inertial Reference Systems (IRS)
• flap position - LH, RH Flap position transmitters • weight on wheels - #1, #2 WOW
• Mach speed – ADC 1, ADC 2, and ISI

The Air Data Computers supply primary Mach data to the SPS computer for Mach compensation of the aircraft’s stall margin.
The SPS uses the above inputs to calculate the angle of attack trip points. When a high angle of attack is approached, the continuous (CONT) ignition is activated. If the angle of attack continues to increase, the stick shaker is activated and the autopilot is disengaged. If the angle of attack still continues to increase, the stick pusher is activated, STALL switchlights on the glareshield panel flash red and a warbler sounds.

In the event of an AOA increase rate greater than 1 degree per second, the stall protection computer lowers the activation trip point. This prevents the aircraft’s pitching momentum from carrying it through the stall warning/stick pusher sequence into the stall.

Question 35

Stick pusher disconnect features

Answer 35

An acceleration switch will disconnect the stick pusher if less than 0.5G is sensed during the stick pusher activation.
The stick pusher can be stopped by pressing and holding either the captain’s or first officer’s control wheel autopilot/stick pusher disconnect switch (AP/SP DISC). The stick pusher is capable of operating immediately when the AP/ SP DISC switch is released.
Should the SPS incorrectly activate the stick pusher, the stick pusher may be disabled by selecting either STALL PTCT PUSHER switch to OFF at the captain’s or first officer’s stall protection panel. Both switches must be ON for stick pusher activation.

Question 36

Controls and Indications

Captain’s reference manual Chapter 9-38

Answer 36

Conventional flight controls consist of the control column and wheel and rudder pedals at each pilot position. In the event of roll or pitch malfunction the associated faulted control system can be isolated by pulling the appropriate disconnect handle.
Aileron and rudder trim is provided.
The yaw damper system provides turn coordination and improves the aircraft’s stability by damping out yaw. It is engaged and disconnected at the center pedestal YAW DAMPER control panel.
The flight spoiler handle allows the pilots to selectively deploy or retract the flight spoilers. The flight spoilers are also operated automatically by the Ground Lift Dumping system. The Ground Lift Dumping system can be armed for either automatic or manual deployment or disarmed by the control switch on the center pedestal.
The horizontal stabilizer controls include:
• Pitch trim switches, (one set/control wheel) • STAB TRIM engage switches
• STAB disconnect switch (one/control wheel) • MACH TRIM engage/disengage switch
The flap lever on the center pedestal selects the electrically operated flaps. The Stall Protection System provides the flight crew with aural, visual, and tactile (stick shaker and pusher) indications of an impending stall. Stall warning switchlights illuminate upon activation of the stick pusher. When the stick pusher is activated, holding the AP/SP DISC switch on either control wheel will interrupt the pusher. To disarm the stick pusher system, select either STALL PTCT PUSHER switch to OFF.