GB1036721A - Improvements in or relating to automatic control systems for aircraft - Google Patents

Abstract

1,036,721. Automatic aircraft landing systems. SMITHS INDUSTRIES Ltd. Sept. 14, 1962 [June 16, 1961], No. 21786/61. Heading B7G. An automatic landing system in an aircraft in which lift is a function of engine speed, in addition to airspeed and angle of attack, e.g. due to slip stream effects, and in which a change in pitch attitude is required during a flare phase following a phase in which the attitude is constant. and leading to touchdown, comprises means operative during the flare phase to actuate the elevators to control the pitch as a function of altitude, and further means operative during the flare phase to control an engine to vary the lift as a function of attitude to achieve smooth touchdown. In the Figure, a pitch rate gyro 1 supplies a signal forming one input to a summing amplifier 6. A vertical gyro 2 provides a pitch signal 0, and a pitch datum generator provides a signal # 0 the difference derived from differential 4 being scaled by a constant factor 1/t 1 in circuit 5 and fed to amplifier 6. A radio altimeter 14 and height datum generator 15 produce signals h and h 0 respectively, the difference derived from differential 16 being scaled by a factor ##/h 0 t 1 , where ## is the change in pitch required during the flare phase. A signal representing rate of climb Dh from a device 18 is scaled by a factor ##/h 0 in circuit 19, and fed to amplifier 6. Amplifier 6 energizes a motor 7 driving a tachometer generator 8 providing a feedback signal to amplifier 6. Motor 7 also provides the input signal to an actuator 11 controlling the elevators 12 and providing a feedback signal added to the actuator input by differential 10. The system solves continuously the equation: where # is elevator deflection and D is the time differential operator. The elevator deflection is therefore controlled so that substantially. If # 0 is the pitch angle being held during the approach, and h 0 is the height at which the flare phase commences, the above equation gives #=# 0 at h=h 0 , and #=# 0 +## at h=O, so that # changes linearly with h from the value # 0 obtaining at the beginning of flare to # 0 +##, a value suitable for touchdown. In the throttle control system, the h signal is fed to a differentiating circuit 18, providing outputs D<SP>2</SP>h fed to summing amplifier 21, and Dh scaled by a factor (1/t 2 + 1/t 3 ) and fed to amplifier 21. A flare base generator 22 provides a signal C, representing the height at which nominal touchdown occurs. This is chosen as somewhat less than the runway altitude to ensure a small rate of descent at touchdown. A signal h-C available at differential 23 is scaled by a factor 1/t 2 t 3 and fed to amplifier 21. The amplifier energizes motor 25 driving tachometer generator 26 providing a feedback signal scaled by a factor 1/K. Motor 25 drives the throttle actuator 29 providing a feedback signal to differential 28. This system solves the equation where T is the throttle setting, whereby T is controlled so that Oh= - 1/t 2 (h-C) substantially, and the aircraft flies an asymptotic approach to the flare base, with the " rapidity " of the approach determined by t 2 .