1. General
A. The SP300 Digital Flight Control System (DFCS) is part of the Flight Management System (FMS). The FMS is the integration of four major systems including:DFCS - Digital Flight Control System A/T - Autothrottle System, IRS - Inertial Reference System, FMCS - Flight Management Computer System
B. The SP300 Digital Flight Control System (DFCS) is an all digital system that includes the following functions:
Autopilot
Flight Director
Mach Trim
Speed Trim
Altitude Alert
C. Two independent DFCS channels are provided. Flight director indications for the captain are from flight control computer A; flight director indications for the first officer are from flight control computer B. No flight director switching between A and B system is available for the captain and first officer; flight director command signals are connected directly to the attitude director indicators (ADI's) on the P1 and P3 panels. The flight control computers use analog and digital input signals which are digitally processed to provide guidance commands for the autopilot and flight director systems. Either autopilot A and/or autopilot B can be engaged for autopilot control of the airplane to pilot selected guidance commands on the mode control panel. Both autopilot channels may be engaged simultaneously to accomplish automatic approaches. Control wheel steering (CWS) is used for manual control of the airplane through the autopilot system. Autopilot actuators for pitch and roll are connected to the appropriate control system linkage.
D. Independence of each system channel is assured by isolation of power supplies, sensors, computers, airplane wire bundles and shelf wiring harnesses. Separate hydraulic systems supply pressure for the autopilot actuators as well as the flight control power control units (PCU's).
F. Altitude Alert System
(1) The altitude alert system compares the airplane altitude with the altitude selected on the mode control panel. A visual and aural warning is provided when the airplane approaches (option) or departs from the selected altitude. No warnings are issued while the airplane is within perscribed limits from the selected altitude. The warning is inhibited during approach by flap position logic.
G. Mach Trim System
(1) The mach trim system provides automatic elevator quadrant displacement as a function of mach number in the mach tuck (nosedown) region. Mach data from the air data computer is used to generate a servo position command signal. This signal is applied to the mach trim actuator. The actuator provides elevator pitch-up position to maintain the correct attitude when airplane mach tuck speed is attained. The mach trim actuator changes the position of the column neutral point without altering feel/force characteristics.
H. Speed Trim System
(1) The speed trim system controls the stabilizer trim motor to provide positive speed stability during certain low speed, high thrust flight conditions. The system is automatically engaged when the autopilot is disengaged, flaps are not up, engine speed exceeds a prescribed amount and control column limit switches are not activated.
2.Flight Control Computer (FCC)
A. Two flight control computers (FCC) designated as A and B are mounted on the E1-3 shelf in the electrical and electronics equipment bay. The units are retained in the rack by cam-loc handles located on the front of the unit which engage adjustable rack-mounted swivel bolts. The flight control computers are standard 1 ATR-long boxes. Electrical connections to the units are provided by four DPX2 connectors located on the rear of the unit which mate with rack-mounted DPX2 connectors. Three capped test connectors and a cooling fan are located on the front panel of the flight control computers.
B. The FCC contains keyed plug-in printed circuit cards which are installed vertically within the unit. The cards have built in lever assist handles for ease of removal. The unit power supply consists of plug-in cards located across the rear of the unit. The electronic circuitry on the boards are latest state-of-the-art design with solid state micro-electronics used extensively.
C. The FCC interfaces with analog and digital data signals. The mode computation is performed by the digital computer. Analog to digital (A/D) and digital to analog (D/A) converters within the computer provide system compatability. The computer is designed to interface with ARINC 429 digital data.
D. Airplane compatibility is determined by the three option pins in the FCC. These pins shall be used to compare against the allowable combination of FCC and airplane. The failure to pass the compatibility check shall result in the computer failing the power up test.
E. Operation of the FCC is centered around two digital processors. The CPU 1 processor performs most mode logic and control while CPU 2 does parallel computations for dual-channel autoland functions and provides the pitch inner loop and limiting functions. Computations are split between the two processors to preclude one processor failure from causing a simultaneous roll and pitch Autopilot hardover. Control and logic calculations are shared between the processors for critical applications. All signals to and from the FCC and data transferred between the processors are controlled by a direct memory access controller.
F. The inputs, data valids, and existing logic states are combined in logic software to compute the current operating mode states. The selected software generates autopilot commands for output to the airplane control surfaces, and Flight Director commands for output to the pitch and roll F/D bars.
G. The CPU 1 software executes periodic self-tests to detect failures in software execution or in RAM or ROM access. During autoland and autopilot go-around modes, CPU 1 software monitors surface commands generated by CPU 2 for comparison with CPU 1 commands (and visa versa) to detect unacceptable discrepancies. Failure in the self-test or in cross-CPU redundancy monitoring are recorded by the continuous monitor.
H. System A interfaces with system A electrical power, system A hydraulic power and the captain's instruments, attitude, compass, air data and navigation systems. System B interfaces with system B electrical power, system B hydraulic power and the first officers instruments, attitude, compass, air data and navigation systems. The systems are isolated to provide two completely independent systems. Both systems operate simultaneously during dual channel approach mode.
3. Mode Control Panel (MCP)
A. The mode control panel (MCP) is mounted on the glare shield, accessible to both pilots. The unit is held by four screws located on the face of the unit and two screws located on the bottom of the unit. The MCP provides the primary interface between the pilots and the DFCS. The MCP contains the engagement control, mode selection control and control parameter selection associated with the Altitude Alert, A/P, F/D, Autothrottle, and, to a limited extend, the FMCS. Interconnection to airplane wiring is by 3 connectors located on the rear panel. The panel also provides control of preset heading, preset course, altitude select engage control and speed selection.
B. The A/P (autopilot) ENGAGE switches (A and B) are 3-position paddle switches that are solenoid-held in OFF, CWS, and CMD positions. If proper logic is made, the solenoid releases the switch allowing movement from the OFF position. These switches allow engagement of the desired autopilot channel, A or B in CWS (control wheel steering) mode or engagement of A or B or both during approach in CMD (command) mode. If engage logic is satisfied, then the switch will hold in CWS or CMD position (with override capability) and are spring-loaded to the OFF position when the system is disengaged. The switches control discrete signal lines.
C. The F/D (flight director) ON/OFF switches provide control of the flight director command output to the ADI. Discrete signals are provided to the respective ADI and digital data is provided to the FCCs. A master light (MA) is located above each F/D switch to indicate which system (during slaved operation) is providing mode control.
D. The A/T (autothrottle) ARM/OFF switch is a two-position, solenoid-held switch for arming the autothrottle. On the ground, the FMCS must be engaged before the autothrottle can be armed. The ARM position permits A/T control of thrust levers when compatible mode selection exists between the A/P and A/T.
4.Automatic Flight Controls (AFC) Accessory Unit
A. The AFC accessory unit is mounted on the E1-3 shelf in the electrical and electronic equipment bay. The unit is a 3/8 ATR long unit which is retained in the rack by a camlock handle which engages in a rack-mounted adjustable swivel bolt. Electrical connections are provided by a DPX connector on the rear of the unit which mates with a rack-mounted connector.
B. The unit contains several relays, time delay gates, a circuit interrupter and other components used for interconnecting the DFCS.
5. Flight Mode Annunciator (FMA)
A. One flight mode annunciator (FMA) is installed on the P1 panel and the other on the P3 panel. Each FMA is composed of three sections: one section provides AP/FD system (AFDS) mode annunciation with engage status and warning; one section provides A/T annunciation and warning; and the third section provides N1 LIMIT annunciation and FMC warning.
B. The AFDS mode annunciators on the left side are: L NAV-HDG SEL, VOR LOC, G/S, FLARE, VNAV PATH-VNAV SPD, ALT ACQ-ALT HOLD, V/S, MCP SPD-TO/GA and A/P STATUS annunciators: CWS ROLL-SINGLE CH, CWS PITCH-A/P OFF. The AT mode annunciators on the center are: FMC SPD-N1, MCP SPD-RETARD, ARM-A/T LIMIT, THR HLD-GA. The N1 annunciators on the right side are: CRZ, CLB, REDUCED-CON, TO-GA. Warning lights, A/P, A/T, and FMC, with push-to-reset feature, are installed in the lower right area of the FMA. A 3-position test switch is located adjacent to the FMC warning light.
C. Each of the mode annunciators is a prismatic electromechanical device which displays the annunciator legends. The device consists of a three-faced prism and two prism-driving solenoids. Each face of the prism is colored and engraved with an annunciator legend; however, only two faces are engraved. The third face is blank (black) and is in the blank display position when deactivated. The prism is electromechanically rotated in either direction about the longitudinal axis so that one of the other faces (upper and lower) can appear in the display position. A solenoid actuated by a remote mode signal (dc ground) drives the prism mechanism. The prism is driven 120 degrees in the required direction to display the corresponding mode legend. When the solenoid loses the signal, the prism reverts to the original position (display of the blank face). The second solenoid is used with another mode and drives the prism in the opposite direction. The prisms are illuminated constantly by background lights.
D. The warning lights are composed of colored lens caps, red for A/P and A/T and amber for FMC. The lights are illuminated by application of a ground. Pressing the warning light assembly operates a switch which completes 28 volts dc to the corresponding autopilot channel to extinguish the light and/or stop a warning horn.
E. The FMA contains self-test circuits to test the operation of the annunciator electromechanisms. A 3-position toggle switch is provided to perform the test. Test 1 is used to check the upper prism face electromechanisms. When the switch is held to 1, a dc ground is applied to all solenoids corresponding to the upper prism faces. All the upper face legends appear simultaneously in the display position if the related circuits and mechanisms are functioning properly. Similarly, the test 2 switch is used to test the lower prism face electromechanisms. The test switches also illuminate the warning lights. A power supply and logic circuits are within the FMA for operating the mode annunciators. The pilot's panel lighting circuits provide power for the FMA background lights.
6. Attitude Director Indicator (ADI)
A. The attitude director indicator (ADI) provides attitude information and flight director commands. An ADI is located on each of the pilots panels (P1 and P3). The flight director commands are described in this section. The attitude display is described in Inertial Reference System.
B. An inverted V command bar on ADI displays both pitch and roll commands. The command bars move up or down to represent pitch commands, or tilt to the left or right to represent roll commands. When the pitch and roll commands being generated by the flight control computer are followed, the command bars will rest on top of the fixed airplane symbol.
C.The input commands for pitch or roll are amplified then connected to a feedback summing point and to a motor. The motor drives the roll or pitch movement of the V bar command and also the feedback potentiometer. When the command bar has moved the proper amount the feedback potentiometer signal will cancel the input signal and the command bar will stop. If the valid signal is lost, the roll command bar drives to level and the pitch DC command causes the bar to be biased up and out of view.
D. A CMPTR flag in the ADI is controlled by DC valid voltage from the FCC.
7.Control Wheel Steering (CWS) Force Transducers
A.Three identical force transducers are installed in the captains and first officers control columns, two in the pitch axis and one in the roll axis. The pitch axis force transducers are located in the forward control quadrants below each control column. The roll axis force transducer is mounted between the control shaft and aileron drum below the captain's control column.
B.Each force transducer provides dual electrically isolated ac output signals that are proportional to the force applied to the control wheel or column. The signals are used by the autopilot computers when the system is engaged in CWS (manual) mode or in CMD mode with no flight mode selected. When the autopilot is engaged in a CMD mode the CWS signals are monitored. CWS signals, above a preset level (high detent), will cause the CMD mode to disconnect and the CWS mode to engage.
8.Aileron Force Limiter
A.The force limiter is splined to the aileron drum at the bottom of the captain's control column. The force limiter consists of an electromechanical clutch, switch, a spring-loaded arm and cam assembly, plus motor-driven limit switches. The electromechanical clutch and switch are activated when the roll axis is engaged. The force limiter mechanically limits the control wheel angle by progressively increasing the force within the system to cam out autopilot control at the A/P actuator. Motor-driven switches within the force limiter, which are controlled by trailing edge flap position, determine the control wheel angle limit. With trailing edge flaps up, the limit is 17 degrees. With trailing edge flaps down (1 unit position and beyond) the limit is 25 degrees. This limits the rate of roll which the autopilot can command; thereby, limiting autopilot hardover maneuver capability. The clutch switch ensures the force limiter is engaged. If the clutch disengages, the switch opens the A/P engage switch holding coil circuit.
B.The captain or first officer may override the force limiter by applying sufficient force on the control wheel to drive through the limiting mechanism. The autopilot continues to drive the A/P actuator throughout the control wheel steering regime after the control wheel is returned to less than 25 or 17 degrees.