Patent GB-868199-A (GB)
Title: Aircraft control system
IPC: B64C27/32
Filed: 1957-06-27
Abstract: 868,199. Fluid-pressure servomotor control systems. SPERRY RAND CORPORATION. June 27, 1957 [June 28, 1956], No. 20359/57. Class 135. [Also in Groups XXXIII and XXXV] Control apparatus for an aircraft having means adjustable to control the attitude of the aircraft about an axis, comprises a mechanical differential having its output connected to operate a device controlling a main actuator connected to adjust the attitude-controlling means, a manual controller mechanically connected to one input of the differential, automatic control means providing an output in dependence of angular movement of the aircraft about the axis, which output controls the control device of a secondary actuator connected to the second input of the differential, means operable to produce an ordered rate of change of attitude signal, and means including the automatic control means and responsive to the order signal to produce an output to actuate the control means of the secondary actuator and thus cause the aircraft to change its attitude at a rate corresponding to the ordered rate. In the Figure, a pitch-responsive gyro 11 in a helicopter is coupled to the rotor 31 of a pick-off 30, the stator 32 of which is connected to the stator of a pickoff 36 whose rotor 38 is on the shaft of an integrating servomotor generator combination 80 having an input on line 72 from a switch arm 90b of switch 90. The voltage induced in rotor 38 thus represents the difference between the angle of pitch and a reference pitch determined by the time integral of the signal on line 72. This difference, or pitch error signal, is fed through a demodulator 41, switch arm 15a and resistor 14, to amplifier 16. A pitch rate error signal is also provided as an amplifier input by capacitor 13. Amplifier 16 energizes a motor 122 controlling a valve 123 which receives pressure fluid from a line 131 if a valve 130 is opened by a solenoid 128 on closure of a switch arm 90c of a switch 90, and distributes the pressure fluid to a rotary actuator 124 connected by a pin 25 to one end of a differential lever 18. A manual control column 20 is connected to the other end of lever 18 by rod 29 and pivot 26. An intermediate point on the lever is pivoted at 27 to the rod 56 of a valve 59 on a fluid pressure servo jack 60, the valve receiving pressure fluid from line 129 and distributing it to the cylinder. The cylinder piston is tied to aircraft structure so that cylinder 60 moves to re-seat valve 59 in known manner. The output shaft 28 of the servo is connected to the swash plate of the helicopter. Motor 122 also drives the wiper 46 of a potentiometer 47 energized by a battery 48 with a grounded centre tap. The derived voltage is fed over line 115 and resistor 66 to amplifier 16, over line 68, switch arm 15b and line 92 to the ON contact of switch arm 90a of switch 90, and over line 67 to the OFF contact of switch arm 90b. The control column 20 is connected by pivot 104 and rod 105 to a casing 100 containing springs 106, 107 tending to centre a piston 108. The piston rod 109 is movable by a lead screw 86 driven by a motor 22 energized from switch arm 90a over line 91 and amplifier 23. Attached to piston 108 is switch gear 103 which deenergizes a relay 15 controlling switch arms 15a, 15b to open the contacts whenever the control column is moved from its trimmed position. The pilot can depress a button 96 on the control column, close switch 96d and short-circuit switch gear 103, to maintain coil 15 energized. The button also closes switch 96b to energize a relay 24, the arm 24a of which connects either with line 116 or a line 118 connected to the ON contact of switch arm 90a and the OFF contact of switch arm 90b. Piston 108 drives the wiper 102 of a potentiometer 101 on casing 100 energized by a battery with a grounded centre tap, the derived voltage being fed over line 99 to relay arm 24a. During manual control with the automatic control system disengaged (switch 90 OFF), switch arm 90c is open, and valve 130 closed. Pressure fluid is cut off from valve 123, and actuator 124 is inoperative, and locked in a neutral position by a finger 52 engaged in a recess 53 in the actuator by a spring 57 acting on a piston 50 on the finger. (When switch 90c is closed, the fluid pressure in line 131 acts on the piston and withdraws the finger from the recess.) The control column 20 operates valve 59 and jack 60 through rod 29 and lever 18 in conventional manner, to control the cyclic pitch of the rotor and control the pitching attitude of the helicopter. During manual control, departure of the pitch angle from the reference determined by the angular position of rotor 38 results in a signal on line 42 which unbalances the inputs of amplifier 16. Motor 122 is operated, and wiper 46 derives a feedback signal fed over lines 67 and 72 to integrator 80, which thus rotates rotor 38 until it picks up a zero signal. The pitch reference is thus continuously adjusted to agree with the actual pitch, and the automatic system may be engaged smoothly at any time, and will operate with a reference pitch equal to the pitch at the time of engagement. To trim the control column feel, the pilot presses button 96. Relay 24 is then energized, and the signal on wiper 102, representing the feel force, is fed over line 99, arm 24a, arm 90a and line 91, to energize the trim motor 22 to trim out the feel. When the automatic system is engaged (switch 90 ON) and the helicopter flown " hands off," valve 130 is opened by solenoid 128, and pressure fluid is supplied to valve 123 and actuator 124, and finger 52 is withdrawn from recess 53. Any departure of the helicopter from the reference pitch attitude generates a signal on line 42 which unbalances amplifier 122 and operates valve 123, and actuator 124 operates lever 18, valve 59, and jack 60 to apply a compensating cyclic rotor pitch change. Wiper 46 derives a feedback signal fed over line 115 to amplifier 16 to prevent large control effects developing from small pitch departures. The feedback signal is also fed over line 68, arm 15b, line 92, and arm 90a, to energize the trim motor 22. Should the pitch departure be persistent, e.g. caused by a movement of the centre of gravity of the helicopter, the trim motor will move piston 108, casing 100, control column 20, rod 29, and the differential lever 18, to control the cyclic rotor pitch. This relieves the automatic system from producing long-term control effects. During automatic control, the human pilot can effect short-term manoeuvres by displacing the control column from its trimmed position, e.g. to avoid an obstacle. Such displacement results in a pitching rate by manual operation of valve 59 and jack 60, and operates switch gear 103 to de-energize relay 15, and open arm 15a, so that amplifier 16 receives a pitch rate signal from line 42 through capacitor 13, without a pitch error signal. This pitch rate signal is compared in amplifier 16 with the control column force signal fed from wiper 102 over line 99, arm 24a and line 116. If the pitch rate does not correspond to the control column force, motor 122 and valve 123 are operated to produce a correcting rotor cyclic pitch change. Thus the human pilot can demand a rate of pitch change by exerting a corresponding force on the control column. On release of the control column, the automatic system returns the helicopter to the reference pitch attitude. To change the reference pitch attitude during automatic control, the human pilot displaces the control column and depresses button 96. Relay 24 is energized, and a control column force signal on wiper 102 is fed over line 99, arm 24a, line 118, arm 90b and line 72 to integrator 80, which changes the reference pitch at a rate proportional to the control column force. Amplifier 16 is unbalanced and operates motor 122 to produce cyclic rotor pitch change until the helicopter attains the new reference pitch. Switch arm 96a being closed, relay 15 remains energized despite operation of switch gear 103. The feedback signal on wiper 46 is thus fed over line 68, arm 15b, line 92 and arm 90a to the trim motor 22, to trim the neutral position of the control column to correspond to the new reference pitch. This process is continuous, and affects the signal on wiper 102, so that the human pilot must apply a constant force, rather than a constant displacement, to control column 20 to produce a constant rate of change of reference pitch. In a modification, Fig. 3 (not shown), switch 96, when operated to effect trimming of the control column during manual control, or alteration of the reference pitch during automatic control, supplies a constant voltage of appropriate sign to the trim motor or the integrator as appropriate, in place of the signal on wiper 102. Specifications 638,971 and 830,481 are referred to.