(2)  Manual Mode
(a)  
The basic pitch channel operating mode is the manual (pitch attitude hold) mode. To engage the pitch channel in the manual mode (mode select switch in MAN position and pitch mode select switch in OFF position), the flight control system B switch must be on, the vertical gyro must be up to speed, the main electric stabilizer trim not in operation, the stabilizer trim cutout switch in the normal position, a valid pitch calibrator must be installed, the flaps must be synchronized and stabilizer trim servo speed change relay must be good (2-second time delay), 30-volt dc pitch channel power must be good, a hot short NOT signal must be available, a pitch control wheel steering out of detent NOT, stabilizer mistrimmed NOT signal must be available and air data computer airspeed interlock must be good. On airplanes with pitch channel part No. 2588810-903, -904, derived rate must be good (Fig. 16).

(b)  
The interlock signals are processed through a logic chain in the pitch channel, summed, and fed to one input of an AND gate in the pitch control panel. The 28 volts dc from the autopilot disengage switches provides a corresponding 1 input to the AND gate which triggers the AND thus providing a 1 output to one input of a second AND gate in the control panel. When integrity of the autopilot disconnect circuits are satisfied, 28 volts dc from the engage interlock circuit breaker triggers the second AND gate. This action allows the pitch engage switch to unlatch and energizes the pitch engage switch solenoid. The pitch engage switch solenoid is held to ground through the circuit interrupter. Except for the trim monitor circuit and the CWS circuit, failure of any of the prescribed interlocks or deliberate operation of the disconnect switches will disengage the pitch channel. After engagement, trim monitor signal will not cause the pitch channel to disengage. This is also true for control wheel steering inputs.

(c)  
Upon engagement, if no other mode is selected, the pitch channel automatically operates in the pitch attitude hold mode. The motor clamp circuit for the pitch computer is activated, locking the servo assembly in the position corresponding to the vertical gyro sensed pitch attitude existing upon engagement (Fig. 15). Changes in airplane pitch attitude are sensed in the vertical gyro. The CT, directly coupled to the vertical gyro, provides an output signal to summing point 6 proportional tot he attitude change. This signal is routed to the valve amplifier which commands elevator movement to restore original

582 
May 01/01  BOEING PROPRIETARY - Copyright . - Unpublished Work - See title page for details.  22-11-0 Page 55 

Roll Engage Logic Circuit  5C8 
22-11-0  Figure 16  Jun 20/85 
Page 56 
BOEING PROPRIETARY - Copyright . - Unpublished Work - See title page for details. 

pitch attitude. When the elevator moves as a result of the command signal, the elevator position transducer signal is fed to summing point 3 to null the command signal and stops the elevator movement. As the airplane returns to the original pitch attitude, the CT error signal diminishes to zero. The elevator position signal, at summing point 3, restores the elevator to neutral position. Damping of airplane movement is provided by the signal from pitch rate gyro which is shaped in the wipeout (bandpass) filter in the rate gyro filter. Output from the pitch rate gyro is routed through a filter and summing point 6, where it is added to the pitch attitude error signal. Gain of the variable gain amplifier is inversely proportional to the signal from the air data computer to compensate pitch command signals for changes in airplane control dynamics as a function of airspeed.
(3) 
 Lift Compensation

(a)  During turn maneuvers, a lift compensation (pitch up) signal, proportional to the versine (1-cosine) of the commanded bank angle is routed through a variable gain amplifier to summing point 6. At summing point 6, this signal is summed with the pitch computer CT output to provide an elevator pitchup command from the valve amplifier. Elevator motion stops when position transducer signal at summing point 3 cancels the pitch command signal. The amplifier gain is inversely proportional to a signal from the air data computer to compensate the lift compensation signals for changes in airplane control dynamics.

(4)  
Thrust Compensation (if installed)

(a)  
An input from the auto throttle system is fed to summing point 2 to give down elevator command for acceleration and to give up command for deceleration. This action complements the auto throttle system.

(b)  
Thrust compensation is not effective when the pitch computer is clamped (pitch attitude hold) or when the pitch CWS is out of detent. The thrust compensation input to summing point 2 is grounded by energized switch PS-12 when the CWS is out of detent.
　
中國航空網 m.k6050.com
航空翻譯 www.aviation.cn
本文鏈接地址：737 AMM 飛機維護手冊 AUTO FLIGHT 自動飛行 1(24)