| 摘要 |
1,190,330. Automatic control of aircraft. HONEYWELL Inc. Sept. 5, 1968 [Sept. 8, 1967], No.42309/68. Heading G3R. To maintain a constant altitude during rapid airspeed changes by controlling the pitch attitude, the aircraft control apparatus includes known pitch attitude control apparatus, (arranged above the broken line 10), and modifying control apparatus which introduces additional control signals for causing operation of the control surface actuator 74 to keep the aircraft at a predetermined altitude in the long term and also to oppose short term altitude changes which would result from indicated airspeed (I.A.S.) changes, a signal summing device 14 receiving outputs from a pitch attitude sensor 11, an altitude sensor 117, and computing means 111, 112, 129, responsive to the output of I.A.S. sensor 110 and having regard to both magnitude and rate of change of airspeed. The output of differentiating filter 111 is modified by a variable gain device 112, the gain of which is controlled by an I.A.S. signal from sensor 110 applied on line 114, and device 112 is arranged so that its gain is proportional to the reciprocal of the cube of the sensed airspeed, the output from device 112 passing via an integrator 129 to the summing device 14. The altitude sensor 117 gives an altitude error signal on line 130 which is integrated by integrator 129 or a separate integrator (not shown) before being applied to device 14. The output signal from gain device 112 is supplied through integrator 129 to provide a "quickened" corrective effect on the altitude during I.A.S. changes, whereas in the long term, the altitude is held constant by control signals derived from the altitude error signal on line 120, one output of gain device 122 going to integrator 129 and another output passing via network 124 and conductor 125 to summing device 23 to provide a conventional proportional control signal to the pitch attitude control apparatus. In an alternative, the gain device 112 may be linearly scheduled (see curve C of Fig. 2, (not shown), ), with a break point at a desired airspeed. In the pitch attitude control apparatus, sensor 11 is a displacement gyro, roll attitude sensor 25 is a horizon gyro for giving a corrective signal preventing loss of altitude during a banked turn, pitch rate sensor 40 is a conventional rate gyro, accelerometer 53 responds to yaw axis accelerations, sensor 60 may be the same as sensor 11 and the output of summing device 47 passes through an aircraft body bending filter 65 to a servo amplifier 70 controlling a summing valve 70 controlling the application of hydraulic oil to control surface actuator 74 controlling elevons or elevators. A second output from filter 65 passes to a summing device 78 where it is summed with a signal from a pitch rate sensor 80, which may be the same as sensor 40. Sensors 40, 80, could be replaced by means for differentiating the output of sensor 11. The output from device 78 is supplied via summing device 84 to a servo-amplifier 87 controlling an electrical servomotor 90, 91, controlling a parallel servo 94 via mechanical gearing 92, a rate feedback signal passing to device 84. A mechanical output from servo 94 via connection 100 moves the manual control column, and a force input from the column movement passes to valve 72, which also receives an electrical input which generates a mechanical torque in the valve. Operation of actuator 74 has proportional and integral terms, the former from conductor 71 and the latter from motor 90. |
