GB882334A - Improvements in or relating to apparatus for the control of an aircraft's speed - Google Patents

Abstract

882,334. Automatic speed control systems for aircraft. SMITH & SONS (ENGLAND) Ltd., S. Feb. 24, 1960 [Feb. 25, 1959], No. 6488/59. Class 38 (4). In an aircraft automatic speed control system, signals dependent on rate of pitch of the aircraft are supplied through a filter having special characteristics and combined with signals dependent on aircraft speed error to control propulsive thrust of the aircraft. As applied to the control of a four-engined aircraft, a computer 42, Fig. 1, associated with a pressure head 43, produces two identical signals proportional to deviation of airspeed from a value determined by the setting of a knob 16. One of these signals controls the outboard engines and the other the inboard engines. The outboard engine control signal on line 3 is applied to a summator 4 together with a signal derived from a rate of pitch gyro 14 through a filter 15 which is provided with rectifiers and decay networks to eliminate transient effects such as would be caused by turbulence. The resultant signal from summator 4 is applied through amplifiers 8, 8a to control servomotors 23, 23a adjusting outboard engine throttles 28, 28a to maintain constant airspeed. Each servomotor control system is provided with position and rate feedback. The inboard engine throttles 28b, 28c are controlled in a similar manner by the second airspeed error signal on line 3a and a signal derived from a second rate gyro 14a. The two airspeed error signals are also respectively applied to electro-mechanical integrators 12, 12a producing rotary outputs which are combined in a differential gear 19 so that the addition is effective to operate generators 22- 22c applying signals proportional to the integral of airspeed error to the servomotor amplifiers. A differential output from gear 19 operates a rate generator 29 which supplies a signal to a warning and cut-out device 30 in the event of inequality of the two inputs to the differential gear 19 under fault conditions. The signals from the two summators 4, 4a are compared by a unit 7 which also supplies a signal to warning device 30 under fault conditons. The signals from the pitch rate gyros may be added to the inputs of the integrators. Operation of ganged switches 6, 9, 6a, 9a to the other position shown disables automatic airspeed control and permits signal sources 44, 44a to progressively close the throttles through the integrators during landing. In a modification for controlling a two-engined aircraft A.C. control signals are employed and a single integrator of combined airspeed error signals and pitch rate signals applies its outputs to the stators of signal generators providing positional feedsback in the throttle servo-systems. Warning of faulty operation in this case arises from a difference between signal generators adjusted by the two servomotors.