GB795526A - Flight control system - Google Patents

Abstract

795,526. Automatic steering control systems. SPERRY RAND CORPORATION. Oct. 18, 1954, No. 29993/54. Class 38 (4). In an automatic control system for a guided missile, the radio control is inoperative during the propulsion period when stabilization is effected gyroscopically. The missile vertical and horizontal control surfaces are positioned by electrical signals from a free gyroscope situated at the missile centre of gravity two accelerometers at the missile nodal point and two rate gyroscopes. The spin axis of the free gyroscope is aligned with the missile longitudinal axis at the instant of launching and subsequent deviations of the missile axis in yaw or pitch, from the reference axis, are detected by the gimbol system. The accelerometers determine accelerations normal to the missile longitudinal axis. Damping of body oscillations is effected by the rate gyroscopes. In one embodiment, Fig. 7, the yaw wings 95, are positioned by the hydraulic system 187, 189, 190 under the control of electric error signals from the directional reference gyroscope 110, accelerometer 100 and rate gyroscope 166. Similarly, the pitch wings 90, are controlled from the gyroscope 110, accelerometer 139 and the rate gyroscope 165. Damping of the rate gyroscopes is effected by the oil-filled dampers 170 and prior to launching the rotor 111 of 110 is locked with its spin axis aligned with the missile longitudinal axis by the caging device 114. In a second embodiment, Fig. 3, the accelerations perpendicular and parallel to the missile longitudinal axis are measured at the accelerometers 20 and 10 respectively and then applied to the inverse tangent computer 40. The output signal of the latter, which is proportional to the angle between the resultant acceleration and the missile axis, is added to the output of the directional reference gyroscope 50, and applied to the servomotor of the control surfaces 90.