B737 NG Flight controls

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Overview of the B737NG flight control system.

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B737 NG Flight controls

  1. 1. B 737 NG Ground School. See the aircraft study guide at www.theorycentre.com The information contained here is for training purposes only. It is of a general nature it is unamended and does not relate to any individual aircraft. The FCOM must be consulted for up to date information on any particular aircraft.
  2. 2. FLIGHT CONTROLS
  3. 3. Introduction The primary flight control system uses conventional control wheel, column and pedals linked mechanically to hydraulic power control units which command the primary flight control surfaces; ailerons, elevators and rudder. The flight controls are powered by redundant hydraulic sources; system A and system B. Either hydraulic system can operate all primary flight controls. The ailerons and elevators may be operated manually if required (Manual Reversion). The rudder may be operated by the standby hydraulic system if system A and system B pressure is not available. The secondary flight controls, high lift devices consisting of trailing edge (TE) flaps and leading edge (LE) flaps and slats (LE devices), are powered by hydraulic system B. In the event hydraulic system B fails, the TE flaps can be operated electrically. Under certain conditions the power transfer unit (PTU) automatically powers the LE devices. The LE devices can also be extended using standby hydraulic pressure.
  4. 4. Fully powered rudder Hydraulic systems A-B and Stby. No tab Power assisted Elevators Hydraulic systems A and B Manual reversion Aerodynamic balance tab. Power assisted Ailerons. Hydraulic systems A and B. Manual reversion Aerodynamic balance tab. Primary Flight Controls
  5. 5. Movable horizontal stabiliser powered by an electric motor for manual trim or by the Autopilot. Manual via cables from the trim wheels in the cockpit. Four flight spoilers on each wing symmetrical pairs powered by hydraulic systems A and B Secondary Flight Controls Inboard and outboard double slotted T.E Flaps Normally Hydraulic system B Alternate Electric motor Outboard and Inboard spoilers Ground only Hydraulic system A 4 slats on each wing . Hydraulic system B Normal Standby Hydraulic Alternate 2 Krueger flaps On each wing inboard of the engines move with the slats
  6. 6. Blended winglets provide enhanced performance, extended range and increased fuel efficiency. Blended winglets offer operational and economic benefits to 737-800 customers. Mission block fuel is improved approximately 4 %. Range capability is increased by as much as 130 nm on the 737-800. The reduction in takeoff flap drag during the second segment of climb allows increased payload capability at takeoff-limited airports. Environmental benefits include a 6.5% reduction in noise levels around airports on takeoff and a 4% reduction in nitrogen dioxide emissions on a 2,000-nm flight
  7. 7. FLIGHT CONTROL Switches STBY RUD - activates standby hydraulic system pump and opens standby rudder shutoff valve to pressurize standby rudder power control unit. OFF - closes flight control shutoff valve isolating ailerons, elevators and rudder from associated hydraulic system pressure. ON (guarded position) - normal operating position.
  8. 8. Flight Control LOW PRESSURE Lights Illuminated (amber) - • indicates low hydraulic system (A or B) pressure to ailerons, elevator and Rudder (< 1200 PSI) • deactivated when associated FLIGHT CONTROL switch is positioned to STBY RUD and standby rudder ON light illuminates.
  9. 9. Flight SPOILER Switches ON (guarded position) – normal operating position. OFF – closes the respective flight spoiler shutoff valve. Note: Used for maintenance purposes only.
  10. 10. Roll Control The roll control surfaces consist of hydraulically powered ailerons and flight spoilers, which are controlled by rotating either control wheel. Ailerons The ailerons provide roll control around the airplane’s longitudinal axis. The ailerons are positioned by the pilots' control wheels. The Captain’s control wheel is connected by cables to the aileron power control units (PCUs) through the aileron feel and centring unit. The First Officer’s control wheel is connected by cables to the spoiler PCUs through the spoiler mixer. The two control wheels are connected by a cable drive system which allows actuation of both ailerons and spoilers by either control wheel. With total hydraulic power failure the ailerons can be mechanically positioned by rotating the pilots' control wheels. This is called manual reversion. Control forces are higher due to friction and aerodynamic loads.
  11. 11. Aerodynamic Balance tab to assist during manual reversion.
  12. 12. Aileron Transfer Mechanism If the ailerons or spoilers are jammed, force applied to the Captain’s and the First Officer’s control wheels will identify which system, ailerons or spoilers, is usable and which control wheel, Captain’s or First Officer’s, can provide roll control. If the spoiler system is jammed, force applied to the Captain’s control wheel provides roll control from the ailerons. The spoilers and the First Officer’s control wheel are inoperative. If the aileron control system is jammed, force applied to the First Officer’s control wheel provides roll control from the spoilers. There is a 12 degree lost motion device in the spoiler controls which must be exceeded before any spoiler movement will occur. The ailerons and the Captain’s control wheel are inoperative in this case.
  13. 13. Torsion Spring To Spoiler Mixer Unit. Transfer Mechanism If the right control wheel cannot move, the captain can operate the left wheel and override the force of the torsion spring and the force of the feel and centering mechanism. This allows control of the ailerons only. If the left control wheel cannot move, the first officer can operate the right control wheel and override the force of the torsion spring and the spring cartridge. This allows control of the flight spoilers only. To Aileron Feel and Centring unit Captain Control F/O Control. Aileron Transfer mechanism 12⁰ lost motion device
  14. 14. Spoiler mixer Main Wheel well Right side forward FO’s Control wheel input
  15. 15. Left Main Gear Captains control wheel input System B Aileron PCU’s System A Aileron feel and centring unit and trim actuator
  16. 16. Aileron PCU’s Main Wheel well Aileron Transfer Mechanism Connects Both control wheels Aileron Trim Electric motor repositions Feel unit and control wheel neutral Aileron Feel and centering unit Simple spring feel that holds control wheels neutral Spoiler mixer controls spoiler deflection proportionally to control wheel and speed brake commands.
  17. 17. The A and B FLT CONTROL switches control hydraulic shutoff valves. These valves can be used to isolate each aileron, as well as the elevators and rudder, from related hydraulic system pressure.
  18. 18. Aileron Trim Dual Aileron trim switches, located on the aft electronic panel, must be pushed simultaneously to command trim changes. The trim electrically repositions the aileron feel and centring unit, which causes the control wheel to rotate and redefines the aileron neutral position. The amount of aileron trim is indicated on a scale on the top of each control column. If aileron trim is used with the autopilot engaged, the trim is not reflected in the control wheel position. The autopilot overpowers the trim and holds the control wheel where it is required for heading/track control. Any aileron trim applied when the autopilot is engaged can result in an out of trim condition and an abrupt rolling movement when the autopilot is disconnected.
  19. 19. LIMITATIONS Autopilot/Flight Director System # Use of aileron trim with the autopilot engaged is prohibited.
  20. 20. Flight Spoilers Four flight spoilers are located on the upper surface of each wing. Each hydraulic system, A and B, is dedicated to a different set of spoiler pairs to provide isolation and maintain symmetric operation in the event of hydraulic system failure. Hydraulic pressure shutoff valves are controlled by the two flight SPOILER switches. Flight spoiler panels are used as speed brakes to increase drag and reduce lift, both in flight and on the ground. The flight spoilers also supplement roll control in response to control wheel commands. A spoiler mixer, connected to the aileron cable-drive, controls the hydraulic power control units on each spoiler panel to provide spoiler movement proportional to aileron movement. The flight spoilers rise on the wing with up aileron and remain faired on the wing with down aileron. When the control wheel is displaced more than approximately 10 , spoiler deflection is initiated.
  21. 21. Flight SPOILER Switches ON (guarded position) – normal operating position. OFF – closes the respective flight spoiler shutoff valve. Note: Used for maintenance purposes only.
  22. 22. • Spoilers • 4 Flight spoilers on each wing. • Spoiler rise on down-going wing with 10° control wheel rotation (equates to 1.6 trim units of rotation) NOTE; Trim indication is in units not degrees on Boeing aircraft
  23. 23. Flight spoiler movement depends on Roll command and Speed Brake lever position. Ground Spoilers UP or DOWN Position controlled by the SPOILER MIXER. Controlled by Ground Spoiler Control Valve and Ground Spoiler Bypass Valve.
  24. 24. 2 3 4 5 8 9 10 11 FLIGHT SPOILER ROLL CONTROL 1.6 Units 10º Spoilers 4&5 or 8&9 begin to rise
  25. 25. 2 3 4 5 8 9 10 11 Control Wheel 2, 3. 10, 11 4, 5, 8, 9 30 Degrees 4.8 Units 2 or 11 UP 3.5 3 or 10 UP 5.5 UP 10.5 70 Degrees 11.2 Units UP 33 UP 38 FLIGHT SPOILER ROLL CONTROL 4.8 Units 30º 2 or 11 UP 3.5º 3 or 10 UP 5.5º 4 & 5 or 8 & 9 UP 10.5º
  26. 26. 2 3 4 5 8 9 10 11 Control Wheel 2, 3. 10, 11 4, 5, 8, 9 30 Degrees 4.8 Units 2 or 11 UP 3.5 3 or 10 UP 5.5 UP 10.5 70 Degrees 11.2 Units UP 33 UP 38 FLIGHT SPOILER ROLL CONTROL 11.2 Units 70º 2 & 3 or 10 & 11 UP 33º 4 & 5 or 8 & 9 UP 38º
  27. 27. PITCH CONTROL Pitch Control The pitch control surfaces consist of hydraulically powered elevators and an electrically powered stabilizer. The elevators are controlled by forward or aft movement of the control column. The stabilizer is controlled by autopilot trim or manual trim.
  28. 28. Elevators The elevators provide pitch control around the airplane’s lateral axis. The elevators are positioned by the pilots control columns. The A and B FLT CONTROL switches control hydraulic shutoff valves for the elevators. Cables connect the pilots’ control columns to elevator power control units (PCUs) located in the tail compartment and powered by hydraulic system A and B. The elevators are interconnected by a torque tube. With loss of hydraulic system A and B the elevators can be mechanically positioned by forward or aft movement of the pilots’ control columns. Control forces are higher due to friction and aerodynamic loads.
  29. 29. Elevator Neutral Shift The stabilizer controlled elevator neutral shift function changes the neutral position of the elevators during stabilizer movement. The stabilizer moves elevator neutral shift rods. The neutral shift rods provide an input to the elevators through the Mach trim actuator and the feel and centring unit, When the feel and centring unit moves it also back drives the control cables which move the control columns to their new neutral position. The elevator is drooped 4 degrees when the stabilizer is at the neutral trim position indicated as 4 units of trim. When the stabilizer moves in either direction from neutral position, The elevators move up this is true with flaps UP. FCC Controlled Neutral Shift Functional The primary function of the flight control computer (FCC) controlled neutral shift is to reduce the pilot column force necessary to trim the airplane during initial climb out after takeoff with a forward CG and both engines operating. The FCCs provide inputs to the Mach trim actuator. The incremental elevator input commanded by the FCC controlled neutral shift varies with flap and stabilizer position. Very simply with flaps not up the combined FCC and stabiliser neutral shift function moves the elevators down with the stabiliser set between 0 and about 7 units. From 7 units to 16.9 units or maximum Nose up trim the elevators move UP Note. Trim 0 is maximum stabiliser L.E. Up or Nose down trim for Aft C. G. Trim 16.9 is maximum L.E Down or Nose UP trim for forward CG. Note. Manual trim wheel inputs, autopilot, and speed trim will not command elevator movement by the FCC controlled neutral shift function.
  30. 30. Elevator Tab Control Mechanism The elevator tab mechanism changes the function of the elevator tab. When the TE flaps are up, the elevator tab operates in the balance mode. When the TE flaps are not up, the elevator tab operates in the anti-balance mode which aids nose up control with the flaps not up. The elevator tab control mechanism attaches to the aft edge of each elevator. Functional Description When the flaps are up, the elevator tab operates in the balance mode. As the elevator moves, the tab moves in the opposite direction to elevator movement. In the balance mode, the tab moves 0.75 degrees for each degree of elevator movement. When the flaps are not up and hydraulic power is on, the elevator tab operates in the anti-balance mode. The actuator extends, moves the crank and repositions the elevator balance tab rods. When the elevator moves, the elevator tab moves in the same direction that the elevator moves. In the anti-balance mode, the tab moves 0.50 degrees for each degree of elevator movement. When the TE flaps are not up, and the FCCs send signals to the left solenoid control valve. This connects system A pressure to the actuator for the left tab. When the TE flaps are not up, there is a 10-second delay before the right actuator moves. The purpose of the 10-second delay is to improve autopilot performance. When power goes to operate the control valve it connects system B pressure to the right actuator. NOTE: With Both hydraulic system A and B failed in Manual reversion mode there is no pressure and the tabs remain in the balance mode to assist the pilot.
  31. 31. Elevator Torque tube Elevator control rod Elevator Elevator Tab. Elevator Elevator tab control mechanism Flaps UP balance mode Single acting actuator extended in Anti-Balance mode Elevator Hinge point
  32. 32. Elevator Control Column Override Mechanism In the event of a control column jam, an override mechanism allows the control columns to be physically separated. Applying force against the jam will breakout either the Captain’s or First Officer’s control column. Whichever column moves freely after the breakout can provide adequate elevator control. Although total available elevator travel is significantly reduced, there is sufficient elevator travel available for landing flare. Column forces are higher and exceed those experienced during manual reversion. If the jam exists during the landing phase, higher forces are required to generate sufficient elevator control to flare for landing. Stabilizer trim is available to counteract the sustained control column force.
  33. 33. Elevator Feel System The elevator feel computer provides simulated aerodynamic forces using airspeed (from the elevator pitot system) and stabilizer position. Feel is transmitted to the control columns by the elevator feel and centring unit. To operate the feel system the elevator feel computer uses either hydraulic system A or B pressure, whichever is higher. When either hydraulic system or elevator feel pitot system fails, excessive differential hydraulic pressure is sensed in the elevator feel computer and the FEEL DIFF PRESS light illuminates. Elevator Feel Computer Inputs The elevator feel computer receives pressure from hydraulic systems A and B, pitot pressure from the pitot tubes mounted on the vertical stabiliser, and mechanical input from the stabilizer position. The elevator feel computer uses pitot pressure and stabilizer input to control hydraulic pressure to the dual feel actuator in the feel and centering unit. The feel force of the feel and centring unit increases as the airspeed increases. Elevator Feel pitot Left and Right
  34. 34. Stall Warning System Natural stall warning (buffet) usually occurs at a speed prior to stall. In some configurations the margin between stall and natural stall warning is less than desired. Therefore, an artificial stall warning device, a stick shaker, is used to provide the required warning. The stall warning “stick shaker” consists of two eccentric weight motors, one on each control column. They are designed to alert the pilots before a stall develops. The warning is given by vibrating both control columns. The system is armed in flight at all times. The system is deactivated on the ground. The SMYD computers provide outputs for all stall warning to include stick shaker and signals to the pitch limit indicator and airspeed displays and the GPWS windshear detection and alert. STALL WARNING TEST Switches Push – on ground with AC power available: each test switch tests its respective stall management yaw damper (SMYD) computer. No.1 SMYD computer shakes Captain’s control column, No.2 SMYD computer shakes First Officer’s control column. Vibrations can be felt on both columns • inhibited while airborne. AFT OVERHEAD PANEL Pitch limit indicator. Displayed Flaps Not up and variable based on current conditions
  35. 35. Stall Identification Stall identification and control is enhanced by the yaw damper, the Elevator Feel Shift (EFS) module and the speed trim system. These three systems work together to help the pilot identify and prevent further movement into a stall condition. During high AOA operations, the SMYD reduces yaw damper commanded rudder movement. The EFS module increases hydraulic system A pressure to the elevator feel and centering unit during a stall. This increases forward control column force to approximately four times normal feel pressure. The EFS module is armed whenever an inhibit condition is not present. Inhibit conditions are: on the ground, radio altitude less than 100 feet and autopilot engaged. However, if EFS is active when descending through 100 feet RA, it remains active until AOA is reduced below approximately stickshaker threshold. There are no flight deck indications that the system is properly armed or activated. As airspeed decreases towards stall speed, the speed trim system trims the stabilizer nose down and enables trim above stickshaker AOA. With this trim schedule the pilot must pull more aft column to stall the airplane. With the column aft, the amount of column force increase as the EFS module increases pressure to the elevator feel and centering unit.
  36. 36. These are the components of the stall warning system: * Stall management yaw dampers (SMYDs) * Control column shakers * Elevator feel shift module (EFSM) * Stall warning test panel. The stall warning system shakes the control column when the airplane gets close to a stall. During a stall, the FCCs command the stabilizer to trim the airplane nose down. The EFSM and column cut-out switch modules operate to make sure the pilot cannot easily stop this automatic stabilizer movement with the elevator control column nose up input During a stall, the elevator feel shift module (EFSM) provides 850 psi system A pressure to the elevator feel computer and the dual feel actuator. This causes the feel force of the control column and the feel and centering unit to increase to approximately four times greater than high speed feel forces.
  37. 37. Flight control computer Nose down trim and nose up trim inhibit Elevator feel pitot system Stabiliser trim position input Feel computer adjusts hydraulic pressure according to speed and stabiliser position. Actual feel is provided by the higher of A or B pressure. During a stall System A pressure 850 psi. With a large difference (>25% & flaps are up.) FEEL DIFF PRESS light illuminates. With no Hydraulic pressure simple spring feel holds the control column neutral.
  38. 38. Mach tuck is an aerodynamic effect, whereby the nose of an aircraft tends to pitch downwards as the airflow around the wing reaches supersonic speeds. As the aircraft's wing approaches its critical Mach number, typically around Mach 0.8 the aircraft is traveling below Mach 1.0. However, the accelerated airflow over the upper surface of the wing exceeds Mach 1.0 and a shock wave is created at the point on the wing where the accelerated airflow returns from supersonic to subsonic airflow. While the air ahead of the shock wave is in laminar flow, a boundary layer separation is created aft of the shock wave, and that section of the wing fails to produce lift. In most aircraft susceptible to Mach tuck, the camber at the wing root is more pronounced than that of the wing tip. This means that when an aerofoil exceeds its critical Mach number, the shock wave, and resulting stall condition, will begin to form at the root. A second design element that leads to Mach tuck is that many aircraft which will approach the speed of sound are designed with swept wings. When the wing root stalls, the centre of pressure is shifted towards the wing tip. With a swept wing, this also means that the centre of pressure travels aft compared to the unmoving centre of mass of the aircraft which will generate a nose down pitching moment this is “Mach tuck."
  39. 39. Mach Trim System A Mach trim system provides speed stability at the higher Mach numbers. Mach trim is automatically accomplished above Mach .615 by adjusting the elevators with respect to the stabilizer as speed increases. The flight control computers use Mach information from the ADIRU to compute a Mach trim actuator position. The Mach trim actuator repositions the elevator feel and centering unit which adjusts the control column neutral position. As the speed of the airplane increases, the nose starts to drop. This is called Mach tuck. When the airplane airspeed is more than Mach 0.615, the Mach trim function of the FCC gives an up elevator command to keep the nose of the airplane level. This function operates with or without the autopilot or flight director. Limit airspeed to 280 knots/.82 Mach.
  40. 40. Stabilizer The horizontal stabilizer is positioned by a single electric trim motor controlled through either the stab trim switches on the control wheel or autopilot trim. The stabilizer may also be positioned by manually rotating the stabilizer trim wheel. Trim motor used by both Autopilot and Trim switches. Gear box Cable Drum turned by moving trim wheels and back drives trim wheels when trim moves electrically Screw jack turned by the motor or trim wheels
  41. 41. Stabiliser Trim Screw Jack. Attachment arm to the L.E. of the stabiliser As the trim actuator turns the screw jack a nut in the attachment arm runs up or down the screw thread moving the stabiliser Leading Edge.
  42. 42. Stabilizer Trim Stabilizer trim switches on each control wheel actuate the electric trim motor through the main electric stabilizer trim circuit when the airplane is flown manually. With the autopilot engaged, stabilizer trim is accomplished through the autopilot stabilizer trim circuit. The main electric and autopilot stabilizer trim have two speed modes: high speed with flaps extended and low speed with flaps retracted. If the autopilot is engaged, actuating either pair of stabilizer trim switches automatically disengages the autopilot. The stabilizer trim wheels rotate whenever electric stabilizer trim is actuated. Trim switches. Captain Left hand, FO Right hand. 2 switches both must be moved for trim.
  43. 43. The STAB TRIM MAIN ELECT cut out switch and the STAB TRIM AUTOPILOT cut out switch, located on the control stand, are provided to allow the autopilot or main electric trim inputs to be disconnected from the stabilizer trim motor. Main Electric and Auto pilot cut out switches
  44. 44. Manual Trim Auto pilot Cut out
  45. 45. Control column actuated stabilizer trim cut out switches stop operation of the main electric and autopilot trim when the control column movement opposes trim direction. i.e. If the pilot is pulling nose up nose down trim is inhibited. When the STAB TRIM override switch is positioned to OVERRIDE, electric trim can be used regardless of control column position. The control column activated cut out switches are bypassed.
  46. 46. STAB TRIM OVERRIDE SWITCH
  47. 47. Column Cut out Switches The column cut out switches contain a set of cam-operated switches. When the control column moves in the opposite direction from the stabilizer trim direction, electric trim stops.
  48. 48. Manual stabilizer control is accomplished through cables which allow the pilot to position the stabilizer by rotating the stabilizer trim wheels. The stabilizer is held in position by two independent brake systems. Manual rotation of the trim wheels can be used to override autopilot or main electric trim. The effort required to manually rotate the stabilizer trim wheels may be higher under certain flight conditions. Grasping the stabilizer trim wheel will stop stabilizer motion. Captains Stabiliser Trim wheel Folding Handle
  49. 49. Stabilizer Trim Operation with Forward or Aft CG In the event the stabilizer is trimmed to the end of the electrical trim limits, additional trim is available through the use of the manual trim wheels. If manual trim is used to position the stabilizer beyond the electrical trim limits, the stabilizer trim switches may be used to return the stabilizer to electrical trim limits. Stabilizer Position Indication and Green Band Stabilizer position is displayed in units on two STAB TRIM indicators located inboard of each stabilizer trim wheel. The STAB TRIM indicators also display the TAKEOFF green band indication. The trim authority for each mode of trim is limited to: • Main Electric Trim • flaps extended 0.05 to 14.5 units • flaps retracted 3.95 or 4.30 to 14.5 units (Aircraft option) • Autopilot Trim 0.05 to 14.5 units • Manual Trim -0.20 to 16.9 units. The green band range of the STAB TRIM indicator shows the takeoff trim range. An intermittent horn sounds if takeoff is attempted with the stabilizer trim outside the takeoff trim range.
  50. 50. Speed Trim System The speed trim system (STS) is a speed stability augmentation system designed to improve flight characteristics during operations with a low gross weight, aft centre of gravity and high thrust when the autopilot is not engaged. The purpose of the STS is to return the airplane to a trimmed speed by commanding the stabilizer in a direction opposite the speed change. The STS monitors inputs of stabilizer position, thrust lever position, airspeed and vertical speed and then trims the stabilizer using the autopilot stabilizer trim. As the airplane speed increases or decreases from the trimmed speed, the stabilizer is commanded in the direction to return the airplane to the trimmed speed. This increases control column forces to force the airplane to return to the trimmed speed. As the airplane returns to the trimmed speed, the STS commanded stabilizer movement is removed. STS operates most frequently during takeoff, climb and go-around. Conditions for speed trim operation are listed below: • STS Mach gain is fully enabled between 100 KIAS and Mach 0.60 with a fadeout to zero by Mach 0.68 • 10 seconds after takeoff • 5 seconds following release of trim switches • Autopilot not engaged • Sensing of trim requirement
  51. 51. QRH Procedure to deal with a runaway trim condition
  52. 52. YAW CONTROL
  53. 53. Yaw Control Yaw control is accomplished by a hydraulically powered rudder and a digital yaw damper system. The rudder is controlled by displacing the rudder pedals. The yaw damping functions are controlled through the stall management/yaw damper (SMYD) computers.
  54. 54. Rudder (with Rudder System Enhancement Program (RSEP) installed) The rudder provides yaw control about the airplane’s vertical axis. The A and B FLT CONTROL switches control hydraulic shutoff valves for the rudder and the standby rudder. Each set of rudder pedals is mechanically connected by cables to the input levers of the main and standby rudder PCUs. The main PCU consists of two independent input rods, two individual control valves, and two separate actuators; one for Hydraulic system A and one for Hydraulic system B. Either system A or B will move the rudder with limited authority. The standby rudder PCU is controlled by a separate input rod and control valve and powered by the standby hydraulic system. All three input rods have individual jam override mechanisms that allows input commands to continue to be transferred to the remaining free input rods if an input rod or downstream hardware is hindered or jammed. Pre Rudder system enhancements only 2 lights. After Enhancement program addition of STBY RUD ON light Indicating Standby rudder shutoff valve open. All aircraft should have been modified by 2008.
  55. 55. System A and B PCU’s and input control rods. Standby system PCU with input control rod
  56. 56. Rudder Pressure Limiter. At speeds above approximately 135 knots, both hydraulic system A and B pressure are each reduced within the main PCU control valve to approximately 2,200 psi. This function reduces the PCU output force by approximately 25% this limits full rudder authority in flight after takeoff and before landing. If either system A or B is low pressure the other system operates at full pressure. The main rudder PCU contains a Force Fight Monitor (FFM) that detects opposing pressure (force fight) between A and B actuators. This may occur if either system A or B input is jammed or disconnected. The FFM output is used to automatically turn on the Standby Hydraulic pump, open the standby rudder shutoff valve to pressurize the standby rudder PCU, and illuminate the STBY RUD ON, Master Caution, and Flight Control (FLT CONT) lights.
  57. 57. The standby rudder PCU is powered by the standby hydraulic system. The standby hydraulic system is provided as a backup if system A and/or B pressure is lost. With the standby PCU powered the pilot retains adequate rudder control capability. The Standby rudder PCU can be powered manually through the FLT CONTROL switches or automatically. An amber STBY RUD ON light illuminates when the standby rudder hydraulic system is pressurized. The standby rudder system can be pressurized with either Flight Control switch, or automatically during takeoff , landing and automatically by the Force Fight Monitor. The STBY RUD ON light illumination activates Master Caution and Flight Control warning lights on the Systems Annunciation Panel.
  58. 58. Rudder Trim The RUDDER trim control, located on the aft electronic panel, electrically repositions the rudder feel and centering unit which adjusts the rudder neutral position. The rudder pedals are displaced proportionately. The RUDDER TRIM indicator displays the rudder trim position in units. There is no tab on the rudder. Rudder Trim control Rudder Trim Indicator
  59. 59. Rudder control cables Rudder input rod with force transducer Rudder trim actuator Repositions control neutral and pedals move Centring spring Main PCU control rods Standby PCU control rod
  60. 60. Yaw Damper The yaw damper keeps the airplane stable around the vertical axis. When engaged, the yaw damper system gives input to the main or the standby rudder PCUs. The yaw damper system consists of a main and standby yaw damper. Both yaw dampers are controlled through Stall Management/Yaw Damper (SMYD) computers. During normal operation, SMYD 1 controls the rudder through the main rudder PCU. The SMYD computers receive inputs from both ADIRUs, both control wheels and the YAW DAMPER switch. SMYDs provide yaw damper inputs to the main rudder power control unit (PCU) or standby rudder PCU, as appropriate. YAW DAMPER Light Illuminated (amber) – yaw damper is not engaged. YAW DAMPER Switch OFF – disengages yaw damper. ON – Electrically held. Releases to OFF if Yaw Damper light illuminates • engages main yaw damper to main rudder power control unit if the B FLT CONTROL switch is in the ON position • engages standby yaw damper to standby rudder power control unit if both the A and B FLT CONTROL switches are in the STBY RUD position.
  61. 61. When engaged, the yaw damper system gives input to the main PCU yaw damper solenoid valve. The solenoid valve moves and sends pressure from system B to the electrohydraulic servo valve (EHSV). When the EHSV moves, it sends pressure to move the yaw damper actuator. The yaw damper actuator input mechanically adds to the pilot rudder pedal input. Then the control valve moves and sends pressure to the tandem actuator. This moves the PCU piston and the rudder. Rudder input from the yaw damper does not back drive the rudder pedals. The Yaw damper indicator show Main Yaw damper only inputs to the rudder system.
  62. 62. During normal operation the main yaw damper uses hydraulic system B and the SMYD computers provide continuous system monitoring. The YAW DAMPER Switch is held ON by an electro magnet and automatically moves to OFF, the amber YAW DAMPER light illuminates and the YAW DAMPER switch cannot be reset to ON when any of the following conditions occur: • SMYD senses a yaw damper system fault, • SMYD senses that the yaw damper does not respond to a command, • B FLT CONTROL switch is positioned to OFF or STBY RUD. During manual reversion flight (loss of hydraulic system A and B pressure), both FLT CONTROL switches are positioned to STBY RUD. In this case, the YAW DAMPER switch can be reset to ON when on the standby hydraulic system powers the standby yaw damper. SMYD 2 commands rudder movement for the wheel to rudder interconnect system (WTRIS) Standby yaw damping and turn coordination. During Standby Yaw Damper operation, The WTRIS commands a small amount of rudder movement as a function of the captains control wheel aileron input to help assists turns during flight control manual reversion operations.
  63. 63. Flight control surface position display. Option on some aircraft. Rudder position indicator shows both Main and standby Yaw damper inputs and the main yaw damper indicator may not be fitted. The flight spoiler position sensors are on panels 4 and 9 and send spoiler position to the FCC’s
  64. 64. Speed Brakes The speed brakes consist of flight spoilers and ground spoilers. Hydraulic system A powers all four ground spoilers, two on the upper surface of each wing. The SPEED BRAKE lever controls the spoilers. When the SPEED BRAKE lever is actuated all the spoilers extend when the airplane is on the ground and only the flight spoilers extend when the airplane is in the air. The SPEEDBRAKES EXTENDED light provides an indication of spoiler operation in-flight and on the ground. In-flight, the light illuminates to warn the crew that the speed brakes are extended while flaps are more than 10 or below 800 feet AGL. On the ground, the light illuminates when hydraulic pressure is sensed in the ground spoiler Bypass valve with the speed brake lever in the DOWN position.
  65. 65. In-Flight Operation (All Aircraft) Operating the SPEED BRAKE lever in flight causes all flight spoiler panels to rise symmetrically to act as speed brakes. Caution should be exercised when deploying flight spoilers during a turn, as they greatly increase roll rate. When the speed brakes are in an intermediate position roll rates increase significantly. Moving the SPEED BRAKE lever beyond the FLIGHT DETENT causes buffeting and is prohibited in flight.
  66. 66. •Speed Brakes Flight Spoilers • Speed Brake Lever has 4 or 5 positions DOWN, ARMED, (50%), FLIGHT DETENT & UP Note: the 50% position is on later aircraft only.
  67. 67. In Flight Operation with optional Load alleviation. The speed brake load alleviation feature limits the deployment of the speed brakes under certain high gross weight/airspeed combinations. Under these conditions, if the speed brakes are deployed to the FLIGHT DETENT, they automatically retract to 50 percent of the FLIGHT DETENT. The SPEED BRAKE lever moves to reflect the position of the speed brakes. Manual override is available. Increased force is needed to move the SPEED BRAKE lever beyond the 50 percent position with load alleviation active. The SPEED BRAKE lever must be held in place when manual override is used between 50 percent and the UP position. The SPEED BRAKE lever will remain stationary if moved to UP with load alleviation active. When load alleviation deactivates, the speed brakes can be manually returned to the FLIGHT DETENT position. OPTION. A lever stop feature is incorporated into the SPEED BRAKE lever mechanism. The lever stop prevents the SPEED BRAKE lever from being moved beyond the FLIGHT DETENT when the airplane is in flight with the flaps up. In the event of the loss of electrical power the lever stop is removed and full speed brake lever movement is available.
  68. 68. Speed Brakes Flight Spoilers Spoiler Mixer controls spoiler PCUs & the Ground Spoiler Control Valve Moving the speed brake past the Flight Detent position opens the Ground Spoiler Control Valve, It also causes a lot of buffet and is prohibited in flight.
  69. 69. Speed Brake Lever 50% Speed Brake Lever 50%. And Full Right Aileron. Spoiler Mixer reduces Speed Brake input on up going wing and increases the input to the down going wing .
  70. 70. •Speed Brakes Flight Spoilers • Avoid using when flaps are extended • Do not use below 1,000 ft RA • Do not use with Flap 15 • During flight, do not extend lever beyond FLIGHT DETENT
  71. 71. •Speed Brakes Flight Spoilers • Avoid using when flaps are extended • Do not use below 1,000 ft RA • Do not use with Flap 15 • During flight, do not extend lever beyond FLIGHT DETENT • SPEED BRAKES EXTENDED light: • In flight with lever extended beyond armed position & 800’ RA or Flap 10 • On ground with lever down and hydraulics pressure > 750 psi at the bypass valve indicates a fault.
  72. 72. Ground Operation During landing, the auto speed brake system operates when these conditions occur: • SPEED BRAKE lever is in the ARMED position • SPEED BRAKE ARMED light is illuminated • radio altitude is less than 10 feet (Radio alt From FCC’s) • landing gear strut compresses on touchdown or main landing gear wheels spin up (more than 60 kts). • both thrust levers are retarded to IDLE • The SPEED BRAKE actuator moves the lever to the UP position and the spoilers deploy. If a wheel spin-up signal is not detected, when the air/ground system senses ground mode (any gear strut compresses) the SPEED BRAKE lever moves to the UP position and flight spoiler panels deploy automatically. When the right main landing gear strut compresses, a mechanical linkage opens the ground spoiler bypass valve and the ground spoilers deploy. Note: Compression of any landing gear strut enables the flight spoilers to deploy. Compression of the right main landing gear strut enables the ground spoilers to deploy.
  73. 73. Ground Operation If the SPEED BRAKE lever is in the DOWN position during landing or rejected takeoff, the auto speed brake system operates when these conditions occur: • main landing gear wheels spin up (more than 60 kts) • both thrust levers are retarded to IDLE • reverse thrust levers are positioned to revers idle. The SPEED BRAKE lever actuator automatically moves to the UP position and spoilers deploy. After an RTO or landing, if either thrust lever is advanced, the SPEED BRAKE lever automatically moves to the DOWN detent and all spoiler panels retract. The spoiler panels may also be retracted by manually moving the SPEED BRAKE lever to the DOWN detent.
  74. 74. • Auto Speed Brakes • When armed, lever moves to UP & flight spoiler panels extend with main wheel spin-up or a MLG strut compressed • The ground spoilers deploy with right MLG compression • With the lever DOWN (landing or RTO), all spoilers will deploy if thrust Revers is selected Note: If either thrust lever is advanced after landing or RTO all spoilers will retract Ground Spoilers • Extend when lever in UP position & right main gear is compressed • If Ground Spoiler Bypass Valve fails do not operate speed brake lever in flight
  75. 75. 1 and 6 Ground only 7 and 12
  76. 76. SPEED BRAKE LEVER 2, 3. 10, 11 4, 5, 8, 9 1 6 7 12 Flight Detent UP 19.5º UP 24.5º UP UP 33º UP 38º 1 &12 52º 6 & 7 60º SPOILER SPEED BRAKE CONTROL 1 2 & 3 4 & 5 6 7 8 & 9 10 & 11 12
  77. 77. FLAPS AND SLATS
  78. 78. High lift wing with Short field performance modification. Includes a Triple slotted flap.
  79. 79. Airplane Flight Manual has a limitation restricting the use of flaps above 20,000 feet. The reason for the limitation is simple; Boeing does not demonstrate or test (and therefore does not certify) airplanes for operations with flaps extended above 20,000 feet. There is no Boeing procedure that requires the use of flaps above 20,000 feet. Since flaps are intended to be used during the takeoff and approach/land phases of flight, and since Boeing is not aware of any airports where operation would require the use of flaps above 20,000 feet, there is no need to certify the airplane for this. Flap Manoeuvring Speeds Flap - Speed Schedule/Manoeuvring Speeds The following tables contain flap manoeuvring speeds for various flap settings. The flap manoeuvring speed is the recommended operating speed during takeoff or landing operations. These speeds guarantee at least full manoeuver capability or at least 40 of bank (25 of bank and 15 overshoot) to stick shaker within a few thousand feet of the airport altitude. While the flaps may be extended up to 20,000 feet, less manoeuver margin to stick shaker exists for a fixed speed as altitude increases. Note: The flap manoeuvring speeds should not be confused with the minimum manoeuver speed which is displayed as the top of the lower amber band on the airspeed display.
  80. 80. Flap Extension During flap extension, selection of the flaps to the next flap position should be made when approaching, and before decelerating below, the maneuver speed for the existing flap position. The flap extension speed schedule varies with airplane weight and provides full maneuver capability or at least 40 of bank (25 of bank and 15 overshoot) to stick shaker at all weights. Note VREF 40 depends on current gross weight. VREF 40 + -- therefore varies accordingly.
  81. 81. Begin flap retraction at V2 + 15 knots, except for a flaps 1 takeoff. For a flaps 1 takeoff, begin flap retraction when reaching the flaps 1 manoeuvre speed. With airspeed increasing, subsequent flap retractions should be initiated when airspeed reaches the manoeuvre speed for the existing flap position. The manoeuvre speed for the existing flap position is indicated by the numbered flap manoeuvre speed bugs on the airspeed display. For flaps up manoeuvring, maintain at least: • “UP” Takeoff Flap Retraction Speed Schedule MCP speed is normally set to V2 for take off. The white bug represents V2 or MCP +15. The flight director will command MCP + 20 during initial climb. This is a flap 5 TO When speed is above V2 +15 flap 1 can be selected. Flaps should not be selected to UP until the speed is at least -1 with airspeed increasing. Minimum Manoeuvre speed is the top of the amber about 128 knots.
  82. 82. Leading edge Kruger flaps Extended Any time T.E. Flaps are not up.
  83. 83. L.E. Kruger Flaps. Kruger flaps are either retracted into the wing or extended there is no intermediate position. There is a Go around gate at Flap 1 position. This is used for single engine Go around. It also ensures that the L.E. devices are not inadvertently retracted. Flaps Up Retracted. Flaps 1 or more Extended.
  84. 84. Flaps and Slats The flaps and slats are high lift devices that increase wing lift and decrease stall speed during takeoff, low speed manoeuvring and landing. LE devices consist of four flaps and eight slats: two flaps inboard and four slats outboard of each engine. Slats extend to form a sealed or slotted leading edge depending on the TE flap setting. The TE devices consist of double slotted flaps inboard and outboard of each engine. TE flap positions 1–15 provide increased lift; positions 15–40 provide increased lift and drag. Flaps 15, 30 and 40 are normal landing flap positions. Flaps 15 is normally limited to airports where approach climb performance is a factor. Runway length and conditions must be taken into account when selecting a landing flap position. To prevent excessive structural loads from increased Mach at higher altitude, flap extension above 20,000 feet should not be attempted.
  85. 85. Flap and Slat Sequencing (There are differences depending on aircraft) LE devices and TE flaps are normally extended and retracted by hydraulic power from system B. When the FLAP lever is in the UP detent, all flaps and LE devices are commanded to the retracted or up position. Moving the FLAP lever aft allows selection of flap detent positions 1, 2, 5, 10, 15, 25, 30 or 40. The LE devices deployment is sequenced as a function of TE flaps deployment. When the FLAP lever is moved from the UP position to the 1, 2, or 5 (option10,15,25) position, the TE flaps extend to the commanded position and the LE: • flaps extend to the full extended position and • slats extend to the extend position. When the FLAP lever is moved beyond the 5 (25) position the TE flaps extend to the commanded position and the LE: • flaps remain at the full extended position and • slats extend to the full extended position.
  86. 86. The LE devices sequence is reversed upon retraction. Mechanical gates hinder inadvertent FLAP lever movement beyond flaps 1 for one engine inoperative go–around and flaps 15 for normal go–around. Indicator lights on the centre instrument panel provide overall LE devices position status. The LE DEVICES annunciator panel on the aft overhead panel indicates the positions of the individual flaps and slats. Flap 1 and flap 15 gates.
  87. 87. Leading Edge Flaps Transit (LE FLAPS TRANSIT) Light Illuminated (amber) – • any LE device in transit • any LE device not in programmed position with respect to TE flaps • a LE uncommanded motion condition exists (two or more LE flaps or slats have moved away from their commanded position) • during alternate flap extension until LE devices are fully extended and TE flaps reach flaps 10. Note: Light is inhibited during autoslat operation in flight. Leading Edge Flaps Extended (LE FLAPS EXT) Light Illuminated (green) – • all LE flaps extended and all LE slats in extended (intermediate) position (TE flap positions 1, 2 and 5) • all LE devices fully extended (TE flap positions 10 through 40).
  88. 88. Flap Load Relief (Some aircraft differ) The flaps/slat electronics unit (FSEU) provides a TE flap load relief function which protects the flaps from excessive air loads. This function is operative at the flaps 30 and flaps 40 positions only. The FLAP lever does not move, but the flap position indicator displays flap retraction and re–extension. When the flaps are set at 40, the TE flaps: • retract to 30 if airspeed exceeds 163 knots ( 1 knot above placard limit) • re–extend when airspeed is reduced below 158 knots. When the flaps are set at 30, the TE flaps: • retract to 25 if the airspeed exceeds 176 knots (Placard plus 1 Knot) • re–extend when airspeed is reduced below 171 knots. On some aircraft this function is operative at the flaps 10, 15, 25, 30 and flaps 40 positions. The FLAP lever does not move, but the flap position indicator displays flap retraction and re–extension. On some aircraft the FLAP LOAD RELIEF light illuminates when the TE flap load relief function is activated. Forward Right Panel when fitted.
  89. 89. Flap position Indicator shows position of the Left and Right T.E flaps and can show flap skew and asymmetry. Asymmetry and Skew Detection, Protection and Indication The FSEU monitors the TE flaps for asymmetry and skew conditions. It also monitors the LE devices for improper position and skew conditions on slats 2 through 7. If a flap on one wing does not align with the symmetrical flap on the other wing, there is a flap asymmetry condition. A skew condition occurs when a TE flap or LE slat panel does not move correctly causing the panel to twist during extension or retraction.
  90. 90. Asymmetry and Skew Detection, Protection and Indication The FSEU monitors the TE flaps for asymmetry and skew conditions. It also monitors the LE devices for improper position and skew conditions on slats 2 through 7. If a flap on one wing does not align with the symmetrical flap on the other wing, there is a flap asymmetry condition. A skew condition occurs when a TE flap or LE slat panel does not move correctly causing the panel to twist during extension or retraction. Trailing Edge Flap Asymmetry and Skew When the FSEU detects a trailing edge asymmetry or skew condition the FSEU: • closes the TE flap bypass valve • displays a needle split on the flap position indicator.
  91. 91. Asymmetry and Skew Detection, Protection and Indication The FSEU monitors the TE flaps for asymmetry and skew conditions. It also monitors the LE devices for improper position and skew conditions on slats 2 through 7. If a flap on one wing does not align with the symmetrical flap on the other wing, there is a flap asymmetry condition. A skew condition occurs when a TE flap or LE slat panel does not move correctly causing the panel to twist during extension or retraction. Leading Edge Device Improper Position or Skew When the FSEU detects a LE device in an improper position or a LE slat skew condition, the LE FLAPS TRANSIT light remains illuminated and one of the following indications is displayed on the LE DEVICES annunciator panel: • amber TRANSIT light illuminated • incorrect green EXT or FULL EXT light illuminated • no light illuminated. There is no skew detection of the outboard slats, 1 and 8, or for the LE flaps. Slat skew detection is inhibited during autoslat operations.
  92. 92. Uncommanded Motion Detection, Protection and Indication The FSEU provides protection from uncommanded motion by the LE devices or TE flaps. Leading Edge Uncommanded Motion Uncommanded motion is detected when no TE flap position or autoslat command is present and: • two LE flaps move on one wing, or • two or more slats move on one wing. The FSEU shuts down the LE control and illuminates the amber LE FLAPS TRANSIT light. In addition, to prevent uncommanded motion from occurring on the LE devices during cruise, the FSEU maintains pressure on the retract lines and depressurizes the extend and full extend lines. ( L.E. Cruise depressurisation.) Trailing Edge Uncommanded Motion Uncommanded motion is detected when no FLAP lever or flap load relief command is present and the TE flaps: • move away from the commanded position • continue to move after reaching a commanded position, or • move in a direction opposite to that commanded. The FSEU shuts down the TE drive unit by closing the TE flap bypass valve. The TE flap shutdown cannot be reset by the flight crew and they must use the alternate flap system to control TE flaps. The shutdown is indicated by the flap position indicator disagreeing with the FLAP lever position. There is no flap needle split.
  93. 93. Flap Placard speeds are displayed below the Landing Gear lever. These are T.E flap load limits. Flap Load Relief. If the Flaps are at 30 and speed exceeds 176 Knots the flaps will move to Flap 25. If the flaps are at 40 and speed exceeds 163 Knots the flaps will move to Flap 30. It is possible to have the flap lever at position 40 and the flaps at 25.
  94. 94. L.E. Devices are indicated individually by LED’s on the Aft Overhead panel. Amber in transit or not in commanded position. Green for extended or Fully extended. The Test button allows a test of all LED’s on the panel.
  95. 95. Flap lever via cables controls to main wheel well moves the TE flap control Valve.
  96. 96. Hydraulic system B pressure drives a hydraulic motor on the T.E Flap drive unit.
  97. 97. As the T.E Flaps move a mechanical linkage sends position information to the T.E control valve and also the L.E. control valve which allows the L.E. devices to move to the correct position
  98. 98. F.S.E.U. Flap Slat Electronics Unit. T.E. Flap Position Information. T.E. Flap Load Relief. T.E. Flap Skew and Asymmetry detection. ( Closes Bypass Valve) T.E. Uncommanded Motion detection. (Closes Bypass Valve) L.E. Flap and Slat Position Information. L.E. Cruise Depressurisation. L.E. Flap and Slat Uncommanded motion detection.
  99. 99. F.S.E.U. Flap Slat Electronics Unit. Rudder PCU for load limiter function. PSEU for T/O Configuration Warning. Ground Proximity warning module for Ground Proximity warnings.
  100. 100. NOT HERE This is what you asked for! WHEN FLAPS ARE PART OF A CHECKLIST. The response is: FLAP 10 GREEN LIGHT. This is what the FSEU is indicating you have!
  101. 101. Autoslats Autoslat operation is normally powered by hydraulic system B. An alternate source of power is provided by system A through a power transfer unit (PTU) if a loss of pressure is sensed from the higher volume system B engine driven pump. The PTU uses system A pressure to power a hydraulic motorized pump, pressurizing system B fluid to provide power for the autoslat operation. Most Aircraft At flap positions 1, 2 and 5 slats extended an autoslat function is available that moves the LE slats to full extended if the airplane approaches a stall condition. NOTE: No auto slat with flaps UP. The autoslat system is designed to enhance airplane stall characteristics at high angles of attack during takeoff or approach to landing. When TE flaps 1 through 5 are selected, the LE slats are in the extend position. As the airplane approaches the stall angle (within 1⁰ of stick shaker), the slats automatically begin driving to the full extended position. The slats return to the extend position when the pitch angle is sufficiently reduced below the stall critical attitude. The L.E FLAPS TRANSIT light is inhibited during autoslat operation. Alternative on some aircraft. At flap positions 1, 2, 5, 10, 15, and 25 an autoslat function is available that moves the LE slats to full extended if the airplane approaches a stall condition.
  102. 102. L.E. Slat fully Retracted. Slat Extended. Sealed position Flap 1 to 5. Slat Fully Extended Forming Aerodynamic slot for improved boundary layer control at high AOA. Flap 10 to 40 or Auto slat. Also Standby Hydraulic Alternate flap operation
  103. 103. Alternate Extension In the event that hydraulic system B fails, an alternate method of extending the LE devices and extending and retracting the TE flaps is provided. The TE flaps can be operated electrically through the use of two alternate flap switches. The guarded ALTERNATE FLAPS master switch closes a flap bypass valve to prevent hydraulic lock of the flap drive unit and arms the alternate flaps position switch. The ALTERNATE FLAPS position switch controls an electric motor that extends or retracts the TE flaps. The switch must be held in the DOWN position until the flaps reach the desired position. No asymmetry or skew protection is provided through the alternate (electrical) flap drive system. When using alternate flap extension the LE flaps and slats are driven to the full extended position using power from the standby hydraulic system. In this case the ALTERNATE FLAPS master switch energizes the standby pump and the ALTERNATE FLAPS position switch, held in the down position momentarily, fully extends the LE devices. Note: The LE devices cannot be retracted by the standby hydraulic system.
  104. 104. Lifting the guard and moving the Alternate flap arming switch to ARM Closes the T.E. Flap Bypass valve Commands the Standby Hydraulic pump to run.
  105. 105. Lifting the guard and moving the Alternate flap arming switch to ARM Closes the T.E. Flap Bypass valve Commands the Standby Hydraulic pump to run. Arms the Alternate Flap control switch.
  106. 106. The Alternate Flap control switch. Commands the T.E. Flap electric motor to move the T.E. Flaps. Opens the L.E. Alternate Shutoff valve and allows Stby Hydraulic pressure to fully extend the L.E Flaps and slats.
  107. 107. Hydraulic motor. Normal flap extension. Hydraulic System B Electric Motor. Alternate flap extension and Retraction Trailing edge Flap power drive unit. Main Wheel well aft wall.
  108. 108. Black and Red torque tube transfers power drive unit force to the Wing mounted flap actuators. Inboard flap inboard drive is in each wheel well. Torque tube. Ball screw. Flap attachment. Transmission 4 each wing.
  109. 109. LIMITATIONS Flight Controls Max flap extension altitude is 20,000 ft. Holding in icing conditions with flaps extended is prohibited. Alternate flap duty cycle: • When extending or retracting flaps with the ALTERNATE FLAPS position switch, allow 15 seconds after releasing the ALTERNATE FLAPS position switch before moving the switch again to avoid damage to the alternate flap motor clutch • After a complete extend/retract cycle, i.e., 0 to 15 and back to 0, allow 5 minutes cooling before attempting another extension. QRH LOSS OF SYSTEM B Hydraulic system B pressure is low. Plan a flaps 15 landing. Set VREF 15 or VREF ICE. Note: VREF ICE = VREF 15 + 10 knots: The trailing edge flaps can be operated with the alternate electrical system. Alternate flap extension time to flaps 15 is approximately 2 minutes.
  110. 110. QRH LEADING EDGE FLAPS TRANSIT Light(s) for only one leading edge device is illuminated: Limit airspeed to 300 knots (280 knots for turbulent air penetration) or 0.65 Mach, whichever is lower. Light(s) for more than one leading edge device is illuminated: Limit airspeed to 230 knots maximum.
  111. 111. Flap Placard speeds are displayed below the Landing Gear lever. These are T.E flap load limits. Alternate Flap Extension speed is 230 Knots as all L.E. devices will extend to the full extend position. This is the structural load limit for the L.E devices. During Alternate extension the LE FLAPS TRANSIT light will be illuminated until the flaps go past 5 (25 on some aircraft) Because the LE Slats are not in their correct sequenced position.
  112. 112. Elevator feel force is provided by the elevator feel computer. The computer receives inputs from where?
  113. 113. Elevator feel force is provided by the elevator feel computer. The computer receives inputs from where? Elevator Feel System The elevator feel computer provides simulated aerodynamic forces using airspeed (from the elevator pitot system) and stabilizer position. Feel is transmitted to the control columns by the elevator feel and centering unit. To operate the feel system the elevator feel computer uses either hydraulic system A or B pressure, whichever is higher. When either hydraulic system or elevator feel pitot system fails, excessive differential hydraulic pressure is sensed in the elevator feel computer and the FEEL DIFF PRESS light illuminates.
  114. 114. The system that detects opposing hydraulic pressure in the main rudder PCU is called the?
  115. 115. The system that detects opposing hydraulic pressure in the main rudder PCU is called the? The main rudder PCU contains a Force Fight Monitor (FFM) that detects opposing pressure (force fight) between A and B actuators. This may occur if either system A or B input is jammed or disconnected. The FFM output is used to automatically turn on the Standby Hydraulic pump, open the standby rudder shutoff valve to pressurize the standby rudder PCU, and illuminate the STBY RUD ON, Master Caution, and Flight Control (FLT CONT) lights.
  116. 116. How many flight spoilers are on each wing?
  117. 117. How many flight spoilers are on each wing?
  118. 118. The flight spoilers begin extension with a control wheel deflection of approximately: a) 10° b) 15° c) 20° d) 25°
  119. 119. The flight spoilers begin extension with a control wheel deflection of approximately: a) 10° b) 15° c) 20° d) 25°
  120. 120. What happens when the ALTERNATE FLAPS Master switch is armed?
  121. 121. What happens when the ALTERNATE FLAPS Master switch is armed?
  122. 122. The amber LE FLAPS TRANSIT light is inhibited under what conditions?
  123. 123. The amber LE FLAPS TRANSIT light is inhibited under what conditions?
  124. 124. Max flap extension altitude is?
  125. 125. Max flap extension altitude is? 20,000 ft.
  126. 126. What is the maximum duty cycle for the alternate flap electric motor?
  127. 127. What is the maximum duty cycle for the alternate flap electric motor? LIMITATIONS. Alternate flap duty cycle: • When extending or retracting flaps with the ALTERNATE FLAPS position switch, allow 15 seconds after releasing the ALTERNATE FLAPS position switch before moving the switch again to avoid damage to the alternate flap motor clutch. • After a complete extend/retract cycle, i.e., 0 to 15 and back to 0, allow 5 minutes cooling before attempting another extension.
  128. 128. L.E. Indicator FLIGHT CONTROL PANEL MACH and STALL warning test switches The END of FLIGHT CONTROLS. Now take the test at www.theorycentre.com For more information info@theorycentre.com

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