Patent GB-766146-A (GB)
Title: Simulated flight computing system having control response characteristics
IPC: G09B9/20
Filed: 1954-11-29
Abstract: 766,146. Aircraft trainers. DEHMEL, R. C. Nov. 29, 1954 [May 18, 1954], No. 34575/54. Class 4. Aircraft flight simulating apparatus comprises a plurality of interacting flight condition computing means responsive to simulated flight controls, including a control of one attitude condition (e.g. angle of pitch #) of the simulated aircraft, for deviring quantities representing attitude and translation conditions (e.g. vertical air speed) of said simulated aircraft in space, and modifying means for supplying a quantity representing rate of change of said attitude to one of the translation condition computing means, whereby the transient response of the aircraft to movement of the attitude control is simulated. The flight computer in Fig. 1 comprises systems computing airspeed (v), angle of attack (α), pitch (#), rate of pitch (#,y), rate of climb (dh/dt), and altitude (h). The airspeed computer comprises a summing amplifier 100 driving a motor 101 at a speed proportional to the magnitude of the output of the amplifier, and in a sense depending on the phase. The motor 101 drives a generator 101a providing a feed back signal Efb for amplifier 100, and also drives, through gearing 101b a shaft 101c positioning the slides of potentiometers 103, 104, 105 and 160, and an air speed indicator 24. In a similar way, the α, #, dh/dt and h computers comprise a summing amplifier, motor, feed back signal generator, and associated potentiometers. The motor 146 of the rate of climb system positions a vertical air speed indicator 151, and the motor 166 of the height computer positions an altitude indicator 153. Air speed (v) computer. A potentiometer 119 has its slider 122 adjusted by the motor 121 of the pitch computer, and is contoured so that the slider derives a signal -E sin #. This is fed to amplifier 100 to represent the component of gravity Gv along the flight path. A potentiometer 105 has its slider 108 positioned by motor 101, and the slider and one end of the potentiometer are connected to earth through a resistor R, the other end being connected to a voltage E, so that the slider picks off a voltage proportional to 1/v, which is fed to a potentiometer 109, the slider 110 of which is positioned by the pilot to represent throttle setting and thus engine output H. The slider 110 thus derives a signal H/v, which is the thrust T, fed to amplifier 100. A square law potentiometer 104 has its slider 107 positioned by motor 101, and a signal proportional to -v<SP>2</SP> is fed to amplifier 100 to represent a drag factor which is proportional to the square of the airspeed only. The potentiometer 116 is earthed through a resistor at its centre point, and each end is connected to a signal -v<SP>2</SP> derived from potentiometer 104. The slider 113 is positioned by the motor 111 of the angle of attack system, and the potentiometer 116 is contoured so that the slider 113 picks off a signal -CD v<SP>2</SP>, CD being the drag coefficient varying with α. This signal is fed to amplifier 100, which sums all the forces acting in the direction of flight, the resultant being proportional to the acceleration which is integrated by the motor-generator 101, 101a to rotate shaft 101 c to a position corresponding to the airspeed. Angle of attack (α) computer. The potentiometer 117 has a v<SP>2</SP> potential applied across its ends, derived from potentiometer 104, and is contoured so that its slider, positioned by motor 111, derives a signal CL v<SP>2</SP> where CL is the lift coefficient, which signal is fed to amplifier 112. The potentiometer 119 has a second slider 123 adapted to derive a signal E cos #, fed to amplifier 112 to represent the component of gravity Gα normal to the flight path. A signal #y from the rate of pitch computer is fed to a potentiometer 160, the slider of which is positioned by motor 101, thus deriving a signal V #y representing the centrifugal force Fc due to pitching movement fed to amplifier 112. The resultant of these forces is integrated by the motorgenerator 111, 111a to position shaft 111c in accordance with angle of attack. Rate of pitch and pitch computers. A potentiometer 118 is fed with a voltage derived from potentiometer 104 and contoured so that the slider 115 positioned by the angle of attack motor 111 derives a signal CM v<SP>2</SP> where CM is the moment coefficient, this signal being fed to a summing amplifier 125. A voltage proportional to airspeed derived from potentiometer 103 is applied to a potentiometer 124, the slider 124a of which is adjusted by the pilot as elevator control, driving a signal Mp representing the pitching moment resulting from elevator deflection. In addition, a signal proportional to d2h/dt2 is derived from the potentiometer 149 of the rate of climb system, and represents a stabilizing effect. The output of amplifier 125 is fed to a transformer 130, the secondary of which provides a feed back signal, and a signal for potentiometer 160 referred to above, and an input signal for the pitch summing amplifier 120 driving an integrating motor-generator 121, 121a, positioning shaft 121c. Rate of climb and altitude computers. The potentiometer 142 has voltages IV from potentiometer 103 applied to it, and is contoured so that the slider 143 positioned by the pitch system motor 121 derives a signal v sin # fed to amplifier 145. Further, the two sliders 142<SP>1</SP>, 144 derive voltages ŒV cos # fed to a potentiometer 140, contoured so that the slider 141, positioned by the angle of attack motor 111, derives a voltage v cos # sin α, applied to amplifier 145, giving a vertical air speed voltage v sin 0 - v cos # sin α = v sin (# - α) nearly, if α is small. A rate of pitch voltage # is derived from the feed back signal of generator 121a and is fed to amplifier 145, to serve as a stabilizing factor. A dh/d "answer " feed back signal is derived from a potentiometer 150, the slider of which is positioned by the motor 146. The feed back voltage from generator 146a is applied to a potentiometer 149 the slider 147(of which derives a signal led to the rate of pitch amplifier 125 as described above. The dh/dt " answer " signal is also fed to a summing amplifier 165 and integrating motor generator system 166, 166a to derive the simulated altitude. In the modification of Fig. 2 (not shown), the rate of pitch change is computed as #y cos ° - #z sin °, where ° is the angle of roll, #y the rate of rotation about the y or lateral axis, and #z the rate of rotation about the vertical axis, or rate of yaw. The term #2 sin ° is obtained by potentiometric multiplication of the rudder deflection and airspeed and applying the resulting signal #z to a sinusoidal characteristic potentiometer the slider of which is positioned by the angle of yaw (°) computing system. Similarly the term #y cos ° is derived. The resultant rate of pitch change signal is used in place of that employed in Fig. 1. At the same time, the # and ° computing systems are used to position a simulated attitude gyro.