Flight controls Flashcards
How are the Flight Controls operated?
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.
Primary flight controls include?
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).
The secondary flight controls consist of:
horizontal stabilizer trim
slats/flaps
multi-function spoilers and
ground spoilers along with the associated aileron and rudder trim systems
Primary Flight controls cont’d
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.
Ailerons, elevator and rudder are controlled by what?
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.
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).
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.
Flutter Dampers
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.
Aileron Disconnect
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.
How many PCU are installed on each Aileron?
2
How many PCUs are installed on the rudder
3
Roll disconnect handle
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.
Glareshield Roll Select switchlights
Left and right ROLL SEL illuminates when disconnect is pulled.
The green PLT ROLL lights indicate?
Both MFS on the wings are available.
Rudder
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.
Yaw damper
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.
Elevators
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.
Pitch disconnect
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.
Horizontal Stabilizer Trim
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.
Horizontal stabilizer trim priority
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.
Stabilizer trim control in order of priority is
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.
Mach Trim
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.
Slats and flaps
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.
Flap position indications are displayed on the EICAS primary page whenever one or more of the following conditions exists.
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.
