Bombardier-Challenger_01-Flight_Controls飞行操纵

航空发表于 2010-5-9 09:01:41

Bombardier-Challenger_01-Flight_Controls
 
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航空发表于 2010-5-9 09:01:51

OPERATING MANUAL PSP 601A-6 SECTION 10 FLIGHT CONTROLS TABLE OF CONTENTS GENERAL Page A. Control Disconnect Systems 1 B. Power Control Units 1 C. Artificial Feel Mechanisms 2 D. Trim Systems 2 E. Control Surface, Trim and Flap Position Indicators 2 F. Gust Locks 2 ROLL CONTROL SYSTEM 3 A. Aileron Trim 3 B. Aileron Control Wheels 3 C. Artificial Feel Mechanisms 3 YAW CONTROL SYSTEM 4 A. Rudder Trim 4 B. Rudder Pedal Assemblies 4 C. Anti-Jam Mechanisms 4 D. Artificial Feel Mechanisms 4 PITCH CONTROL SYSTEM 5 A. Pitch Trim 5 B. Control Columns 6 C. Gain Change Mechanisms 6 D. Artificial Feel Mechanisms 6 E. Anti-Jam Mechanisms 6 WING FLAP SYSTEM 7 A, Flap Control Unit 7 B- Power Drive Unit 7 C. Asymmetry/Overspeed Detector and Brake Assemblies 8 10-CONTENTS Page 1 Apr 02/87 OPERATING HANUU. PSP 601A-6 Page 6. SPOILER SYSTEM 8 A. Flight Spoilers 8 B. Ground Spoilers 9 7. STALL PROTECTION SYSTEM 9 A. Angle-of-Attack Transducers 10 B. Stall Protection Computer 10 C. Stall Protection System Monitoring 11 D. Stick Shakers 11 E. Stick Pusher Sub-system 12 LIST OF ILLUSTRATIONS Figure Number Title Page 1 Flight Controls 13 2 Control Disconnect T-Handles 14 3 Flight Controls - Hydraulics 15 4 Control Surface Position Indicator and Servo Monitor Lights 16 5 Trim Controls and Trim Position Indicators 17 6 Control Wheel 18 7 Wing Flap Controls and Indicators 19 8 Spoiler Controls and Indicators 20 9 Stall Protection System Controls and Indicators 21 10- Page Apr OPERATING MAMUAL PS? 601A-6 SECTION 10 FLIGHT CONTROLS 1. GENERAL (Figures 1 and 3) The primary flight controls, consisting of roll control, yaw control, pitch control, flight spoilers and ground spoilers, are fully powered from all three hydraulic systems. Mechanical inputs from the pilots1 controls in the flight compartment are conveyed via push/pull rods, quadrants and cables to power control units (PCU). There is no interconnection between hydraulic systems, and all PCUs are totaViy independent of each other. The secondary controls consist of the wing trailing edge flaps and control surface trim systems, and are electrically controlled and actuated. The ailerons, elevators and flight spoilers are each powered by two of the three independent hydraulic systems. The rudder is powered by all three systems and the ground spoilers are powered by No. 1 system only. The primary flight control systems are capable of continued safe operation if jamming or disconnection of a component, loss of normal electrical power and, with the exception of the spoilers, loss of hydraulic systems No. 1 and/or No. 2 occur. Jamming or disconnection of a component is nullified by incorporation of dual control circuits with anti-jam and/or disconnect mechanisms. Loss of normal electrical power is overcome by an air-driven generator (ADG) which is capable of supplying, emergency electrical power to drive hydraulic system No. 3. Loss of hydraulic systems No. 1 and/or No. 2 is catered for by hydraulic system No. 3 which supplies a PCU for each of the primary controls except spoilers. A. Control Disconnect Systems (Figure 2) Control disconnect mechanisms are provided for disconnecting the control columns (pitch control) and the control wheels (roll control), if a jam occurs in their respective cable runs. The disconnect mechanisms are operated by the PITCH DISC and ROLL DISC T-handles on the centre pedestal. If a jam occurs in the rudder control circuits, break-out bungees and an anti-jam mechanism isolate the jammed circuit. Yaw control is retained by both pilots. B. Power Control Units The primary flight control surfaces are fully power-operated by hydraulic actuators known as power control units. To provide for failsafe operation and eliminate fluid interflow between the three aircraft hydraulic systems, each aileron is powered by a dual PCU consisting of two independent actuators; each elevator is powered by two independent PCUs; and the rudder is powered by three independent PCUs. SECTION 10 Page 1 Apr 02/87 cttaueneter OPERATING MANUAL PSP 60U-6 Each PCU consists mainly of a control-valve-operated piston moving in a cylinder. The PCUs are connected to the control surfaces by rod-end attachments and operate to move the control surfaces in the desired direction upon receipt of a signal from the pilots1 controls or from the automatic flight control system (AFCS). A flight control monitoring unit monitors the operation of the PCUs. The flight control monitoring unit receives inputs from PCU proximity sensors and transmits warning signals to the servo monitor panel in the flight compartment. Artificial Feel Mechanisms Because the primary flight control surfaces are fully power-operated, artificial feel mechanisms, consisting of spring devices, are incorporated in the control systems to simulate aerodynamic forces and provide a means of sensing control loads under various flight conditions. Trim Systems (Figures 5 and 6) Trim inputs are introduced into the roll and yaw control systems by electrically driven actuators controlled by the AIL TRIM and RUD TRIM switches on the centre pedestal. Pitch trim is obtained by varying the angle of incidence of the horizontal stabilizer. Signals from the pitch trim switches on the control wheels, from the AFCS and from the stability augmentation system (SAS) are processed by a control unit to operate an electrically driven actuator which applies the required amount of stabilizer deflection. The pitch trim disconnect switch on each control wheel disconnects and brakes the pitch trim actuator in an emergency. Control Surface, Trim and Flap Position Indicators (Figures 4, 5 and 7) Flight control surface positions and trim angles are displayed on indicators located on the centre instrument panel. A flap position indicator on the copilot's instrument panel displays flap position angles. Inputs to the position indicators are provided by transmitters and trim actuators. Gust Locks Gust locking of the ailerons, rudder and elevators is provided by trapping hydraulic fluid within the PCUs whenever hydraulic pressure is removed from the PCUs. This arrangement locks the control surface against the effect of gusts but permits restricted movement of the surface, if a sufficiently large external force is applied continously. SECTION 10 Page 2 Apr 02/87 OPERATING MANUAL PSP 601A-6 2. ROLL CONTROL SYSTEM Roll (lateral) control is achieved by hydraulically powered ailerons which are controlled primarily from conventional column-mounted horn-type wheels. Primary control is supplemented by an electrically actuated trim system. The roll control system incorporates a dual PCU for each aileron, and a dual control system. Normally, both control systems are interconnected so that there is simultaneous movement of both ailerons; but i t is possible to isolate a jammed aileron control circuit by means of a disconnect mechanism, thereby allowing limited control (one aileron only) through the unjamrned circuit (refer to Figures 1 and 2). Control wheel movement is transmitted by cables and pulleys which incorporate an a r t i f i c i a l feel unit to the PCUs located outboard in the wing, forward of the rear spar. Each PCU actuator is capable of aileron operation should there be a failure associated with the adjacent actuator. Signal inputs from the AFCS are made through the right aileron system only. Therefore, should jamming of the right control system occur, the autopilot inputs would not be transmitted to the left aileron system (refer to Section 4 ). A. Aileron Trim An electrically driven actuator applies a bias to the primary control circuit, when required, by operation of the AIL TRIM switches located on the centre pedestal. The amount of trim applied to the ailerons is shown on the control surface trim position indicator. B. Aileron Control Wheels (Figure 6) The aileron control wheels are horn-type handwheels, spline-mounted on the control columns. Each control wheel mounts a pitch trim switch, a pitch trim disconnect switch, an autopilot/stick pusher disconnect switch, an autopilot touch control switch and a radio key. C. Artificial Feel Mechanisms Two a r t i f i c i a l feel mechanisms provide the pilots with positive feel of the power-operated control system and act as centering devices. SECTION 10 Page 3 Apr 02/87 ctianenQer OPERATING KMftJAL PSP 601A-6 3. YAW CONTROL SYSTEM Yaw (directional) control is achieved by a hydraulically powered rudder, controlled primarily from conventional dual, cross-coupled pedals- Primary control is supplemented by an electrically actuated trim system. The yaw control system incorporates three independent, parallel-connected PCUs and a dual control system which includes two anti-jamming mechanisms for isolating or overriding the effects of a jammed circuit, enabling control to be maintained via the intact circuit. The system i s also protected by anti-jam mechanisms built into the PCU input levers, which act to isolate a jammed PCU. Pedal assembly movement is transmitted by cables and pulleys which include artificial feel mechanisms, load limiters and a trim mixing system. In addition to control inputs from the pedal assembly, inputs from the stability augmentation system of the AFCS are applied to the system through two yaw dampers in the trim mixing system (refer to Section 4). A. Rudder Trim An electrically driven actuator applies a bias to the primary control circuit, when required, by operation of the RUD TRIM control located on the centre pedestal. The amount of trim applied to the rudder is shown on the control surface trim indicator. B. Rudder Pedal Assemblies Conventional rudder pedal assemblies enable foot control of the aircraft wheel brake system and the rudder control system. Each set of pedals is provided with a hand-operated adjusting mechanism to cater to the individual requirements of pilots. C. Anti-Jam Mechanisms The two forward anti-jam mechanisms operate to nullify the effects of a jammed cable circuit and maintain normal pedal/rudder movement ratio. The anti-jam mechanism on each rudder PCU acts as a push/pull rod for the PCU input linkage during normal operation. If the input linkage cannot move because of a jam in the PCU, the anti-jam mechanism breaks out to isolate the defective PCU from the system. The remaining PCUs continue to operate the rudder. D. Artificial Feel Mechanisms Two artificial feel mechanisms provide the pilots with positive feel of the power-operated system and act as a centering device for the system. SECTION 10 Page 4 Apr 02/87 canaaair ctiauencjer OPERATING MANUAL PS? 601A-6 4. PITCH CONTROL SYSTEM Pitch (longitudinal) control is achieved primarily by two independent, hydraulically powered elevators. Elevator movement i s controlled from conventional control columns. Primary control is supplemented by an e l e c t r i c a l l y actuated t r im system which varies the angle of incidence of the horizontal stabilizer. The pitch control system incorporates two parallel-connected PCUs for each elevator, and a dual control system. Normally, both control systems are interconnected so that there i s simultaneous movement of both elevators, but it is possible to isolate a jammed c i r c u i t by means of a disconnect mechanism, thereby providing limited p i t c h control (one elevator only) through the remaining c i r c u i t (refer t o Figures 1 and 2). Control column movement i s transmitted by cables and pulleys, through an a r t i f i c i a l feel unit, to the PCUs. Operation of the elevator PCUs i s similar to that of the aileron PCUs. Signal inputs from the AFCS are made through the rear quadrant of the l e ft elevator control system only. Therefore, should jamming of the l e f t cable circuit occur, the autopilot inputs would no longer be available to the elevator system. A. Pitch Trim The aircraft is trimmed iti pitch by varying the horizontal stabilizer angle of incidence. Trim commands from the pilot's or copilot's control wheel switches, the AFCS and the stability augmentation system (SAS) are processed by a trim control unit to operate the electrically driven stabilizer actuator. Commands from the pilot's trim switch override those from the copilot's trim switch, the AFCS and the SAS. Commands from the copilot's trim switch override only those from the AFCS and the SAS. Both control wheels have a red disconnect button, PITCH TRIM DISC, which can be pressed to remove power from the system and brake the actuator. In order to enhance the longitudinal trim movement, the movement of the horizontal stabilizer is accompanied by a degree of elevator movement that alters the stabilizer/elevator camber. An elevator servo input is generated by the horizontal stabilizer movement to produce the required elevator deflection. The electrically driven screw actuator, located at the top of the vertical stabilizer, varies the horizontal stabilizer angle of incidence. The actuator is driven by two electric motors directly connected to the drive train each containing a high and low trim rate. Manual trim commands from the control wheel pitch trim switches produce a steady rate of stabilizer movement of 1/2 degree per second. Depending on flap position, the autopilot commands variable high or low trim rates of 0.1 to 0.5 degree per second and 0.01 to 0.1 degree per second respectively. Mach trim commands a variable rate of stabilizer movement between 0.01 and 0.1 degree per second. Each of the electric motors driving the trim actuator is protected against overspeed by a dual coil brake. SECTION 10 Page 5 Apr 02/87 OPERATING HMU4L PSP 601A-6 The control unit controls the rate and direction of movement of the actuator. The unit consists of two independent channels and operates from two power busses so that electrical failure on one bus does not preclude operation of the stabilizer trim. A pilot reset capability allows channel transfer at the pilot's option, The system normally operates on channel No. 1, with channel No. 2 performing only a monitoring and back-up function. Should a failure occur within a controller channel or its associated motor, the control unit automatically transfers to the back-up channel. In the event of an overspeed condition, the control unit removes power from the drive motor and operates the brake in the actuator. Channel failure, overspeed condition and channel change are indicated by switch/lights on the centre pedestal. Two trim position sensors on the actuator send signals to the control unit. One sensor supplies the AFCS with stabilizer angle data and the second is connected to the flight recorder. Both position sensors provide travel limit signals for the control unit. Stabilizer trim position is also an input to the take-off configuration warning system. A third position sensor supplies position signals to the control surface trim position indicator. Control Columns The pilot's and copilot's control columns each consist of a conventional tubular column mounted vertically in a housing. A control column shaker, which is a component part of the stall protection system, is mounted on the column. Gain Change Mechanisms Two independent gain change mechanisms ensure that the rate of elevator movement increases as the control column is moved from neutral to provide the required control response. Artificial Feel Mechanisms Two artificial feel mechanisms, one for each elevator, provide the pilots with positive feel of the power-operated systems and act as centering devices for the systems. The system is designed to ensure a reduced feel force when rapid control column movement is required. Anti-Jam Mechanisms The elevator anti-jam mechanisms act normally as push/pull rods for the PCU input rod linkages. If a PCU input linkage cannot move because of a jam in the PCU, the mechanism breaks out to isolate the defective PCU from the system. The other PCU continues to operate the affected elevator. When the mechanism breaks out, a proximity sensor is deactivated and the amber PITCH light on the SERVO MONITOR panel comes on (refer to Figure 4). SECTION 10 Page 6 Anr H?/R7 OPERATING MANUAL PSP 601A-6 5. WING FLAP SYSTEM (Figures 1 and 7) The flap system consists of externally hinged inboard and outboard double-slotted flap panels mounted on the trailing edge of each wing. The panels are electrically driven by a power drive unit (PDU) located in the main landing gear bay. The motor action of the PDU is translated to eight actuators, two to each flap panel, by flexible shaft assemblies. An asymmetry/overspeed detector and brake unit is incorporated in each flap drive system. The outboard flaps have fixed leading edge vanes and the inboard flaps have movable leading edge vanes which automatically extend or retract as the flaps are lowered or raised. The flaps are extended or retracted in response to command signals from the FLAPS control lever located on the centre pedestal. The signals are fed to the PDU via the flap control unit. If the control unit logic detects an anomaly such as flap asymmetry or overspeed, power is removed, causing the PDU motor brakes and the asymmetry/overspeed detector brakes to stop the system. The FLAPS FAIL light on the copilot's instrument panel comes on when a system fault is detected. A. Flap Control Unit The flap control unit (FCU) is powered from dc bus No. 1 and dc bus No. 2. Although two power supplies are provided, only one is necessary to operate the unit. The function of the unit is to assess the flap extend/retract commands received from the FLAPS control lever and provide the correct activating signal to the PDU. Once a selected flap angle is reached, the flaps are locked in position by the PDU motor brakes and the asymmetry/overspeed detector brake units. The FCU also signals the aural warning unit (refer to Section 3) to initiate aural warnings for airspeed/flap, take-off/flap and gear-up/flap configuration incompatibilities. B. Power Drive Unit Two PDU motors are coupled to a mechanical differential which drives the output shaft through a clutch and an output gear train. With power applied to the PDU, the motor brakes are released and the motor drives the flexible shaft assemblies and actuators. When the selected flap position is reached, the motors are de-energized and the motor brakes are re-applied. If power to one of the PDU motors fails, the associated brake is automatically applied and the second motor continues to operate the system at half speed. In the event of overheating of a PDU motor, thermal switches de-energize the applicable motor and an amber overheat light on the copilot's instrument panel comes on. The thermal switches reset once the overheat condition has passed. SECTION 10 Page 7 Apr 02/87 OPERATING MANUAL PSP 601A-6 C. Asymmetry/Overspeed Detector and Brake Assemblies The function of these assemblies is to transmit signals to the FCU to provide positive braking action to the flaps in the event of asymmetric movement of the left and right flaps, or o*e-speed. 6. SPOILER SYSTEM (Figures 1 and 8) Wing l i f t modulation is achieved by the operation of flight and ground spoilers- The flight spoilers may be extended to any position, between 0 and MAX (40 degrees), required for the intended flight path. The ground spoilers have only two positions, fully retracted during flight or fully deployed (45 degrees) when activated with the aircraft on the ground, to assist other braking systems by dumping l i f t and increasing drag. A. Flight Spoilers The flight spoilers are two hydraulically powered panels, one hinged to the upper surface of each wing, forward of the outboard flaps, and are controlled mechanically through pilot movement of a lever on the centre pedestal. Each panel is powered by two hydraulically independent PCUs. Each PCU is independently connected to its spoiler and is capable of spoiler operation should the adjacent PCU fail either mechanically or hydraulically. The spoiler control lever is connected to the PCUs via cables and pulleys. The spoilers are fully retracted when the lever is in the fully forward position. Pulling the spoiler control lever rearward deploys the flight spoilers, spoiler panel deployment being proportional with control lever movement. Lever positions, when selected, are held by a serrated plate and plunger mechanism. Spoiler panel position is transmitted to the control surface position indicator, the LH FLT SPLR and RH FLT SPLR lights and the LEFT and RIGHT FLIGHT SPOILERS lights. A detent mechanism on both of the spoiler wing circuits prevents unacceptable spoiler asymmetry. If an asymmetry occurs, the detent mechanism closes the affected spoiler when the spoilers are less than one-half extended or retracts i t to the one-half extended position when the spoilers are more than one-half extended- Operation of the LEFT and RIGHT FLIGHT SPOILERS lights indicate that the flight spoiler detent mechanism is serviceable and that blowback protection in an asymmetrical spoiler condition has been reset to the one-half extended position. SECTION 10 Page 8 Apr 02/87 cacnhaadiiaeinr Qer OPERATING MANUAL PSP 601A-6 B. Ground Spoilers The ground spoilers are two hydraul ically powered panels, one hinged to the upper surface of each wing, forward of the inboard flaps, and are controlled electrically. Each panel is powered by one actuator supplied from a dual hydraulic selector valve. The ground spoilers deploy automatically when armed, with a weight-on-wheels or wheel spin-up signal present, and the spoiler control lever and throttle lever selected to the proper positions (refer to Figure 8). A spoiler control unit monitors weight-on-wheels and wheel spin-up signals, throttle lever position, GROUND SPOILERS switch position and the position ' of the two valves in the dual hydraulic selector valve. When all of the conditions for ground spoiler deployment have been met, hydraulic pressure is applied at the ground spoiler actuators, the actuators unlock and the spoilers are powered to the extended position. If the spoiler control unit detects a difference in the positions of the hydraulic selector valves, the ground spoilers, if extended, close and lock. If both throttle levers are not pulled back to IDLE simultaneously, the SPLRS INOP light will come on. Ground spoiler operation is monitored via the LH and RH GND SPLR and SPLRS INOP lights. The system test is initiated via the GROUND SPOILERS switch. 7. STALL PROTECTION SYSTEM (Figure 9) The stall protection system senses the aircraft angle of attack, provides the flight crew with a visual and tactile warning of an impending stall and, if no corrective action is taken, prevents flight into the stalled condition by activating a stick pusher mechanism. The principal system components consist of two trailing vane type angle-of-attack transducers, a dual-channel stall protection computer, two altitude transducers, two lateral accelerometers and two flap position transmitters. The system controls and indicators are: Two stick shakers A stick pusher sub-system Stall protection test indicators System warning lights and test switches An aural warning horn (warbler) SECTION 10 Page 9 Jul 19/05 cacnhaadiiaeinr Qer OPERATING MANUAL PSP 601A-6 When a dangerously high angle of attack is approached, the stall protection computer applies continuous ignition to the engines and, if the angle of attack continues to increase, activates the stick shakers to generate a stall warning in the form of a mechanical vibration of the control columns. If the aircraft angle of attack still continues to increase to the stick pusher trip point, the aural warning horn sounds and the stick pusher sub-system forces the control columns forward to effect recovery from the impending stall. When the aircraft angle of attack has decreased to a preset point below the pusher trip point, the aural warning horn stops and the stick pusher is deactivated. The stick shakers and continuous ignition switch off automatically when the aircraft angle of attack decreases through their respective trip points. Red STALL/PUSH lights flash whenever the aural warning horn and stick pusher are operating. If the autopilot is engaged when the aircraft approaches the stall, it is automatically disengaged on a signal from the stall protection computer when the aircraft angle of attack reaches the stick shaker trip point. A. Angle-of-Attack Transducers There are two angle-of-attack transducers, one on each side of the forward fuselage. Each transducer is attached to an externally mounted trailing vane. The trailing vane is moved by the local airflow which varies in proportion to the aircraft angle of attack. The angles of attack sensed by the left and right transducers are transmitted to the left and right channels respectively of the stall protection computer. The transducer trailing vanes are protected against ice by built-in heater elements controlled from the ADS heater control panel (refer to Section 14). B. Stall Protection Computer The stall protection computer is divided into two identical and independent (left and right) channels. Each channel uses inputs from its associated angle-of-attack transducer, altitude transducer, lateral accelerometer and flap position transmitter to compute angle-of-attack trip points for auto-ignition, stick shaker operation, aural warning and stick push. If the angle of attack increases at a rate greater than 1 degree per second, the computer lowers the angle-of-attack trip points for the various system functions. This action prevents the aircraft momentum in the pitching plane from carrying it through the stall warning/stick pusher sequence into the stall. The two altitude transducers provide altitude signals to the associated left and right sides of the stall protection computer. The transducers are connected to the left and right static systems via static source selectors on the pilot's and copilot's side panels (refer to Section 11). SECTION 10 Page 10 Apr 02/87 OPERATING MANUAL PSP 601A-6 As the altitude transducers signal an increase in altitude between 2,000 and 15,000 feet, the computer progressively lowers the angle-of-attack trip points for the stick shaker and pusher. Below 2,000 feet and above 15,000 feet, the trip points are constant- If one or both of the altitude signals i s lost or if the difference between signals exceeds 2,000 feet, the computer automatically applies the trip points associated with the 15,000 foot altitude. The two lateral accelerometers monitor skid or sideslip and signal the corresponding channel of the computer. Each of the computer channels uses the signals to generate compensated angle-of-attack values produced by manoeuvres involving skid or sideslip. The compensated angles insure that adequate stall protection is provided during uncoordinated flight. The trip points are also lowered progressively, on signals from the two flap position transmitters, as the flaps move, through the 0-, 20-, 30- and 45-degree positions- If one or both of the flap position signals are lost, the computer automatically applies the stick shaker, continuous ignition and stick pusher trip points associated with the next higher flap setting. The weight-on-wheels inputs from the landing gear control unit enable the computer to disable the stick shakers and pusher and the system failure warning lights while the aircraft is on the ground, except during system test. To prevent inadvertent operation of the stick pusher due to a failure in one of the computer channels, the computer does not command a stick push unless both of the computer channels signal a stick push simultaneously. Stall Protection System Monitoring The stall protection computer monitors the operation of the system for possible mechanical defects in the angle-of-attack transducers and for faults in the electrical circuitry. Stick Shakers There are two stick shakers, one on the p i l o t s and one on the copilot's control column. Each shaker is a dc electric motor driving an eccentric weight. The shakers operate independently of each other and are powered by their respective stall protection computer channels. The noise of the stick shakers operating is sufficiently loud to constitute an aural warning of shaker operation. SECTION 10 Page 11 Apr 02/87 OPERATING MANUAL PS? 601A-6 E. Stick Pusher Sub-system The stick pusher consists of a rotary actuator driven by a dc electric motor which operates on the right elevator control. The pusher logic circuits are so arranged that pusher signals must be transmitted simultaneously from both channels of the stall protection computer befcr,- a stick push can be initiated. When in operation, the stick pusher exerts an 80-pound forward push on the control columns. Red STALL/PUSH lights flash whenever the stall protection system computer commands a stick push. In order to prevent the aircraft from flying into a low or negative G condition during the stick push, two accelerometer switches disconnect the pusher drive i f the aircraft reaches 0.5 G during the pitching manoeuvre induced by the stick push. At any time, the pilot or copilot can stop the stick pusher and disconnect the autopilot by pressing and holding the AP/SP DISC switch installed on the left horn of each control wheel. The stick pusher is capable of operating immediately when the switch is released. The stick pusher can be deactivated by either of two PUSHER toggle switches, located on the pilot's and copilot's STALL PROTECTION panels, which would cause flashing STALL PROTECT FAIL lights to come on. SECTION 10 Page 12 Apr 02/87 chanehtyer OPERATING MANUAL PSP 601A-6 AUTOPILOT SERVO ACTUATOR PILOTS CONTROL COLUMN Flight Controls Figure 1 SECTION 10 Page 13 Apr 02/87 OPERATING MANUAL PSP 601A-6 PITCH DISC AND ROLL DISC T-HANDLES Provides a disconnecting mechanism for control columns and control wheels if a jam occurs in respective cable runs. Puffing either handte disengages associated mechanism. Then, rotating handle left or right secures handle in disconnected position. Releasing handle into stowed position, reconnects associated controls and re-abgns control column or wheels, as appropriate. When PITCH DISC handle is pufled. pflot controls left elevator and copilot controls right elevator. When ROLL DISC handte is potted, pilot controls left aileron and copilot controls right aileron. CENTRE PEDESTAL Control Disconnect T-Handles Figure 2 SECTION 10 Page 14 Apr 02/87 OPERATING MANUAL PSP 601A-6 NO. 1 SYSTEM RESERVOIR NO. 3 SYSTEM RESERVOIR NO. 2 SYSTEM RESERVOIR LEFT* ENGINE PUMP ELECT PUMP 2 mm ACCUMULATOR ^ 4 L-v RIGHT ENGINE PUMP £ i i n i t i i i i i f l i i t iH z z •» iis i Z 5 ACCUMULATOR f l l l l l t l l l f l l l l l t l l l l l l l l l l l l l t U I I I l f l l l l l l £ a i f t f i i i i i i i i i i i i i t i t ix • j i t i i i i i i i i m i i i i i i i i i i i i i s i i« LEGEND TO LANDING GEAR AND BRAKE SYSTEMS NO. 1 HYDRAUUC SYSTEM NO. 2 HYDRAUUC SYSTEM NO. 3 HYDRAUUC SYSTEM Flight Controls - Hydraulics Figure 3 SECTION 10 Page 15 Apr 02/87 OPERATING MANUAL PS? 601A-6 CONTROL SURFACE POSITION INDICATOR Provioes a continuous indication of control surface posrticns over Tht->r operating range. L AND R FLT SPLR Right spader up indications. Max 40 degrees 2/4 la V4 28 degrees 16 degrees 5 degrees L AND R AILERON 21.3 degrees 21.3 degrees L AND R ELEVATOR Up Down 23.6 degrees 18.4 degrees iR L N R \ F ELEVATOR |A • • L X R \ RUDDER RUDDER Left/right indications LEFT 25 degrees RIGHT 25 degrees CENTRE INSTRUMENT PANEL PITCH LIGHT Amber PITCH light comes on when proximity sensors detect a jammed control varve or input linkage at the elevator power control units. NOTE Wrth hyotaufie power off. servo monitor panel lights are as follows: - ROLL fight is on - Y AW light is on - PITCH light is out -MON SAFE fight is on. ROLL AND YAW LIGHTS Amber ROLL and YAW faghts come on whenever proximity sensors detect a jammed control valve or hydraulic pressure deficiency at the respective power control units. CENTRE PEDESTAL MON SAFE LIGHT Green MON SAFE light comes on when all aileron and rudder PCUs are unpressurized (all hydraulic systems off) and all elevator PCUs are unjammed. Control Surface Position Indicator and Servo Monitor Lights Figure 4 SECTION 10 Page 16 Apr 02/87 chauencjer OPERATING MANUAL PSP 601A-6 TRIM POSITION INDICATOR Provides a continuous indication to trim position over their operating range. ROD N l AND NR Nose (left) NL/noseright <NR) indications Left Right 8.5 degrees 8.5 degrees AIL LWD AND RWD Up Down 7.5 degrees 7.5 degrees LWD RWD/ T R I M ^ ^ STAB NUP Nose up (NUP) indications. StabBizer moves from 0 to -9 degrees incidence. Green band indicates take-off (TO) trim range. CENTRE INSTRUMENT PANEL RUDDER TRIM CONTROL Control switch sets rudder trim left and right. CHANNEL INOPERATIVE SWITCH/UGHT CHAN 1 INOP CHAN 2 INOP Amber fights indicate failure in respective channel. Pressing swxtch/bght in conjunction with OVSP/CHANGE CHAN switch/light activates pitch trim system. PITCH TRIM r—PUSHCMANI NOTE If input signals to trim indicator are lost, aileron and rudder pointers move off scale 90 degrees from zero index. Stabilizer pointer moves off scale to a point between scale end points. AILERON TRIM CONTROLS Control switches sets aileron trim up and down. OVERSPEED/CHANNEL CHANGE SWITCH/LIGHT Amber lights indicate pitch trim overspeed or channel change. Can be used to change from one channel to other for test. Pressing switch/tight in conjunction with CHAN 1/CHAN 2 switch/Hght activates pitch trim system. CENTRE PEDESTAL Trim Controls and Trim Position Figure 5 Indicators SECTION 10 Page 17 Apr 02/87 OPERATING MANUAL PSP 601A-6 AUTOPILOT/STICK PUSHB* DISCONNECT vvvrrcH ^ed pushbutton which, when pressed, diserrgages •topfiot and deactivates stick pusher. When ;*teased, stick pusher system is immediately reactivated but autopilot remains disengaged. PITCH TRIM SWTICH Enables piot to vary pitch trim according to flight requirement. RADIO K>rY Light grey button which, when pressed., switches on radio transmitter. FRONT VIEW AUTOPILOT TOUCH CONTROL Black button which, when pressed, enables pilot to manoeuvre aircraft without disconnecting autopilot. PITCH TRIM DISCONNECT ^AJTCh Red button which, when pressed, removes power from system and brakes actuator to cater to a possfcle runaway trim actuator. System is reactivated with CHAN 1 INOP/CHAN 2 INOP and OVSP/CHANGE CHAN switch/lights (refer to figure 5). REARVIEW Control Wheel SECTION 10 "Sure 6 page 18 Apr 02/87 cacnhaadiiaeinr Qer OPERATING MANUAL PSP 601A-6 FLAPS FAIL OVHT M0T1 OVHT MOT 2 FLAP FAIL LIGHT Amber light comes on to indicate a flap asymmetry or speed response fault. PDU MOTOR OVERHEAT LIGHTS Amber light comes on to indicate an overheat condition in the associated PDU motor. FLAP POSITION INDICATOR Provides a continuous angular indication of the flaps over their operating range. COPILOT'S INSTRUMENT PANEL FLAP CONTROL LEVER Controls operation of flap power drive unit (PDU). Lever quadrant is marked with the four flight modes: Flight/Taxiing 0 degrees Take-off 20 degrees Approach 30 degrees Landing 45 degrees Each mode corresponding with a detented position of the lever. CENTRE PEDESTAL Wing Flap Controls and Indicators SECTION 10 Figure 7 Page 19 Apr 02/87 cacnhaadiiaeinr Qer OPERATING MANUAL PSP 601A-6 FLIGHT SPOILER DEPLOYED INDICATON Amber lights come on steady when flight spoilers are not fully retracted. Lights come on flashing and take-off configuration aural warning sounds when N1 rpm is increased beyond 75% and flight spoilers are not retracted. GROUND SPOILER DEPLOYED INDICATION Amber lights come on when ground spoilers are at any position other than fully retracted. LH FLT SPLR LH GND SPLR RH FLT SPLR RH GND SPLR GLARESHIELD SPOILER CONTROL LEVER To deploy flight spoilers, lever may be moved rearwards to any one of eight detented positions according to flight path requirements until MAX position stop is reached. FLIGHT SPOILERS LEFT AND RIGHT INDICATION Green lights come on when flight spoilers are extended beyond one-half position. GROUND SPOILERS SWiTCH ON - Arms ground spoilers for deployment. Ground spoilers deploy automatically if a weighton- wheels or wheel spin-up signal is present and either of the following two sets of conditions has been met: - Spoiler control lever at or above 0 through to 1/4 positions and both throttle levers have been advanced above IDLE then returned to IDLE or SHUTOFF positions. -Spoiler control lever is between 1/4 and MAX positions and both throttle levers are at IDLE or SHUTOFF positions. OFF - Ground spoilers are disarmed and cannot be deployed. TEST - LH and RH GND SPLR and SPLRS INOP lights come on to indicate correct operation of ground spoiler control system. Refer to Volume 1. NORMAL PROCEDURES for test procedure. GROUND SPOILER INOP LIGHT Amber light comes on if spoiler control unit detects fault in ground spoiler hydraulic selector valves or if both throttle levers are not pulled back to IDLE simultaneously. CENTRE PEDESTAL Spoiler Controls and Indications SECTION 10 Figure 8 Page 20 Jul 19/05 ehauenper OPERATING MANUAL PSP601A-6 ALT COMP FAIL LIGHTS (2) Red lights come on if one or both attitude signals to SPS computer are lost or if 2000 foot difference between them is detected. 15,000 foot angle of attack trip points are applicable when tights are on. GLARESHIELD STALL/PUSH LIGHTS (2) Red lights flash when angle of attack reaches stick pusher trip point. STALL PROTECT FAIL WARNING LIGHTS (2i Red warning lights flash in the following cases: - To indicate a system fault. Whenever one of th AP/SP DISC buttons on the control wheels is pressed. - During system test. Lights come on steady when power is removed from system. RED SECTOR YELLOW SECTOR NOTE Stick pusher can only be tested on the ground; all other tests can be conducted on the ground or in-flight. BLUE SECTOR SPS TEST INDICATORS (2) Coloured sectors on indicator provide references for stall warning/stick pusher sequence during system test (refer to Volume 1. NORMAL PROCEDURES). Indicator is nor calibrated to provide in-flight angle of attack indication oi approach speed reference. PILOT'S AND COPILOTS SIDE PANELS S T A L L PROTECTION TEST PUSHER .ON PILOTS STALL PROTECTION TEST SWITCH Spring-loaded toggle switch. Holding switch on activates serf-testing of stall protection system. During test, simulated approach to stall is observed as pointer of left SPS TEST INDICATOR moves from the blue to the red sector. Stick pusher can be checked only when pilot's and copilot's TEST switches are held on simultaneously. STICK PUSHER SYSTEM SWITCHES £2) Two-position toggle switches wired in series between stick pusher actuator and battery bus. When both switches are sei to ON. power is available for stick pusher operation. If one switch is OFF. stick pusher cannot operate and both STALL PROTECT FAIL lights come on steady. G SWITCH TEST SWITCH Spring-loaded toggle switch tests operation of one of the accelerometer switches on stick pusher actuator- During stick pusher test, correct operation of accelerometer switch is indicated if stick pusher is immediately de-energized when G SWITCH TEST switch is set to TEST. COPILOTS STALL PROTECTION TEST SWITCH Spring-loaded toggle switch. Holding switch on activates test of right side of system. Test is identical to test of left side of system activated by pilot s TEST switch except that right stick shaker operates and ALT COMP FAIL lights do not come on. PILOT'S FACIA PANEL COPILOTS FACIA PANEL Stall Protection System Controls and Indicators SECTION 10 figure 9 Page 21 Apr 02/87

dul 发表于 2011-2-11 15:20:17

MANUAL 8

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

qiushengsean 发表于 2011-3-1 03:59:51

好好学习
thank you

bocome 发表于 2011-7-31 10:25:57

Airbus A380 operations at alternate airports

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航空论坛_航空翻译_民航英语翻译_飞行翻译's Archiver

Bombardier-Challenger_01-Flight_Controls飞行操纵

MANUAL 8