Patent GB-902547-A (GB)
Title: An instrument landing system
Inventor: MONCRIEFF-YEATES ALEXANDER JOH
IPC: G01S1/02
Filed: 1960-05-23
Abstract: 902,547. Instrument landing systems. NORTH AMERICAN AVIATION Inc. May 23, 1960, No. 18155/60. Class 40 (7). In an aircraft landing system the craft is flown from a first predetermined altitude, e.g. 800 ft., down a linear glide slope defined by known means, e.g. a radio ILS beacon, until a second predetermined altitude, e.g. 150 ft. is reached and during this initial calibration mode a signal is generated representative of the average value of the actual rate of descent hp. When the craft has descended to the second predetermined altitude the I.L.S. control signal is automatically disconnected and in the ensuing control mode the stored rate signal hp is continuously integrated in a programme servo to give a "programmed" altitude signal hp which varies in time in accordance with the linear extension of the initial glide slope down to the ground, said programmed altitude hp being continuously compared with the actual altitude h A  as determined by a radar altimeter to give an altitude error signal he from which is derived an aircraft altitude control signal which is applied to an auto pilot or to a centre-zero pilot indicator. As described, the final control signal is also a function of the pitch rate #, the joystick position #st and the acceleration α at an angle to the longitudinal axis of the aircraft in a plane containing the longitudinal axis and perpendicular to the pitch axis. When a tnird predetermined altitude, e.g. 60 ft. is reached the altitude control signal is automatically modified to ensure an exponential flare out. In the block diagram of the apparatus on the aircraft as shown, switches 31, 32 and 42 are shown in the control position; switch 13 is in the position shown in both the calibration and control modes and switch 55 is in the position shown during the flare-out portion of the control mode. As shown, the output hp from the programme servo 19 is combined with the output h A  from the radar altimeter 12 in a comparator 21 to give an error signal he. During the calibration mode the aircraft is kept on the glide slope by the signal from the I.L.S. receiver 15 so that the value of h A  corresponds to the desired glide slope and the programme servo 19 is controlled to reduce the error signal he to zero but during the control mode the varying values of hp define the desired glide slope and the aircraft is controlled to reduce he to zero. The programme servo 19 comprises a conventional electromechanical velocity servo with tachometer feedback, Fig. 3 (not shown), such that the output signal is the integral of the resultant input signal. In order to reduce the magnitude of the variable input signals to the servo 19, a preset signal representing the nominal rate of descent hnom suited to the aircraft is provided by a generator 17. In the calibration mode the error signal he from the comparator 21 is combined in a circuit 18 with the nominal descent rate signal hnom and is also integrated and stored in a " memory " servo 27 comprising another electromechanical velocity servo whose stored output #he dt is also applied to the combining circuit 18. The resultant output from the combining circuit 18 is applied to the input of the programme servo 19 so that when the error signal he reduces to zero the input to the programme servo 19 is equal to hp, i.e. hp = hnom + #he dt. Control mode. When the aircraft reaches the second predetermined altitude the switches 31, 32, 42 are automatically changed over to the positions shown so that the error signal he is disconnected from the combining circuit 18 and memory servo 27 but the stored memory signal #he dt and the nominal descent rate signal hnom are still applied to the combining circuit 18 so that the output altitude signal hp from the programme servo 19 continues to vary at the desired " programmed " rate hp. The error signal he is now combined in a circuit 34 with signals #, #st and I from the pitch rate gyro 40, stick position signal generator 38 and integrating rate gyro 36 respectively to produce the aircraft control signal E in accordance with the equation. E = K 1  he - K 2  #st + K 3 # + K 4 I where K 1 -K 4  are constants determined by the scaling resistors 33, 41 and other circuit constants. The steady state components of the stick position signal #st are removed in a " washout " circuit 39 which functions in the manner of a high-pass network of time constant 2 sees. The integration time constant of the gyro 36 is 10 sees. and the output signal I is given by where #I is the pitch rate signal generated by the gyro 36 and the term K 8  (#hp)<SP>2</SP>/#hnom is included by operation of the switch 55 when the aircraft descends to the third predetermined altitude to produce the flare out control. Steady state components of the acceleration signal α are removed in a " washout " circuit of time constant 10 sees. similar to the stick position " washout " circuit 39 and the flare out term (hp)<SP>2</SP>/#hnom is produced by applying a signal #hp from the input to the programme servo 19 to a suitably sealed potentiometer 52 whose wiper is controlled by the output shaft of the memory servo 27. Some circuit details of the apparatus are given in Fig. 3 (not shown). In general, the electrical signals are in the form of A.C. voltages and any D.C. input from the transducer is converted to A.C. by a modulator. The " washout " circuits 39 and 48 comprises CR networks which are effective only with D.C. signals and accordingly the outputs therefrom are converted back to A.C. Switching circuits, Fig. 4 (not shown). The switch 31, 32, 42 and 55 are actuated by relays controlled by cams which are driven by the programme servo 19 and the calibration mode is initiated by a switch operated by a cam driven by the output from the radar altimeter 12. The control system is automatically disabled (1) on landing by a switch mechanically connected to the shock absorbers, and (2) if the altimeter signal falls below a predetermined value for longer than a predetermined time, e.g. 1¢ sees., a manual reset switch is provided. Signal failure and the state of the altitude switches is also indicated visually by lamps.