GB910242A - Stable optical tracking fire control system - Google Patents

Abstract

910,242. Radiant-energy aircraft gun sights. NORTH AMERICAN AVIATION Inc. May 23, 1960, No. 18074/60. Class 40 (7). [Also in Groups XXI and XXXIII] Relates to a fire control system for a fighter aircraft in which the pilot maintains the fighter 5, Fig. 3, on a lead pursuit course towards a moving target 1 such that a missile fired at any time from the fighter in a fixed direction relative thereto will strike the target by maintaining a " disturbed " optical sight coincident with the visual image of the target. According to the present invention the optical sight is displaced from a datum position in accordance with the algebraic sum of (1) the angular displacement of the target from the longitudinal axis of the fighter as determined by an automatic tracking radar or an equivalent infra-red system and (2) a steering error signal computed as a function of the range r, range rate r, direction and angular velocity # of the target relative to the fighter as determined by the radar and the velocity vector Vo of the missile relative to the fighter as a function of the roll #, pitch # and angle of attack α of the fighter and the pressure ratio Ps/Pso, where Ps is the static air pressure at the fighter altitude and Pso is the static pressure at sea level. When the target is not visible the target and sight images may be displayed on an equivalent cathode-ray tube display. Assuming that the target is following a straight line path, the steering error, which is zero when the fighter is on a lead pursuit course, is given by where #M is the miss distance vector between the predicted positions C and D of the target and the missile explosion point at a time T f after firing the missile. The steering error is computed as components Mx/T # , My/T # , Mz/T # along x, y, z axes fixed relative to the fighter, as shown in Fig. 4, and the value of T # is given by equating Mx - to zero. The resultant value of T f thus T f gives the instant of nearest approach of the missile to the target and substitution of this value in the expressions for the z and y components of M/T # gives the pitch and yaw steering error signals. As shown in Fig. 2 and Fig. 5 (not shown), the output data from the tracking radar 7 is coupled to the computer 11 and the output yaw and pitch steering error signals are combined at 14 and 15 with signals from the radar aerial 6 representing the azimuth and elevation of the target to give resultant yaw and pitch signals which actuate servos 20, 21 controlling a sight head 22 so that an image 4, Fig. 1, of the sight reticle as projected on to the windscreen 2 of the fighter is displaced horizontally and vertically from the visual image 1 of the target as seen by the pilot in accordance with the yaw My Mz and pitch steering errors - and -. Thus T f T f the pilot controls the fighter to follow the correct course by maintaining the images 1 and 4 coincident. Computer.-If the orthogonal co-ordinate axes of the radar aerial are i, j and k with the beam along the i axis then when the beam is tracking the target the rate of change r in equation (7) above of the target range vector relative to the fighter is given by where #l i , # lj , and #l k are unit vectors along the aerial axes i, j, and k, and # k and # j are the rates of rotation of the radar beam about the k and j aerial axes as determined by rate gyros mounted on the aerial. Equation (7) may thus be expressed in i, j and k components which are then transformed into x, y and z components by means of resolvers set in accordance with the azimuth # and elevation # of the radar beam to give the x, y and z components of #M/T # and hence the time of flight T f and yaw and pitch steering error signals as described above. Certain components of the computer 11 may be as described in Specification 900,047 and the radar data may be smoothed before entering the computer by a vector filter as in Specification 824,177.