1350375 Photo-electric guidance systems USA ARMY SECRETARY OF 8 Feb 1966 [8 Feb 1965] 5561/66 Heading H4D [Also in Division F3] In a missile guidance and control system the missile carries an infra-red radiation source 490 (Figs. 3, 4, not shown) modulated at a high frequency, this modulated signal being received by a ground tracker 100 which generates error signals when the missile deviates from a desired path. These error signals are fed to a computer 200 which generates pitch and yaw signals transmitted via an infra-red transmitter 300 to a receiver 410 on the missile so as to control its flight. The modulation of the missile radiation source at a high frequency enables the tracker to be made insensitive to low frequency background radiation. More particularly, tracker 100 includes an optical device 102; having a variable focal length objective lens, programmed in time according to the known range vs. time characteristics of the missile being tracked by a motor 104 and timer circuit 106. Optical device 102 focuses the received radiation which passes through a beam splitting device 108, e.g. a silicon mirror, which separates the radiation into two portions at the 1À0 micron point. The energy below 1À0 micron is fed to a PbS detector 110. This channel is used as a shimmer modulation cancellation loop. The 1À0 to 2À5 micron energy passes through a rotating half plane A.M. reticle 112 and then to a PbS detector 114. Connected to the detector 114 is an amplifier 116 which is automatically gain controlled. An error signal demodulator 118 is connected to the output of amplifier 116 and a modulation reference generator 120 which is driven in synchronism with reticle 112. The tracker also includes a bias program 122 connected to detectors 110 and 114, an automatic gain controlled reference amplifier 124, and an AGC amplifier shimmer reference 126. Computer 200 includes cant angle corrector 202 having inputs connected to the error demodulator and to a cant angle input. A pair of pulse width modulators 204 and 205 are connected to the outputs of the cant angle corrector and are controlled by clock frequency 206. The computer also includes frequency shift oscillators 208 and 210 connected to pulse width modulator 204. The infra-red transmitter 300 comprises two identical channels for transmitting pitch and yaw commands to the missile. Lamp 302, in the pitch channel, provides the infra-red radiation which is controlled by optical device 304 and projected by means of projection optics 306. Optical device 304 is controlled by a galvanometer mirror 308 that is driven by galvanometer driver amplifier 310. The yaw channel of the transmitter is identical to the pitch channel and has the same reference numerals but with primes added. The missile carried infra-red receiver 410 utilizes an optical system 412 for focusing the received radiation on a PbS detector 414 which is connected to a signal amplifier 416. The signal amplifier has its output connected to band pass filters 418 which are in turn connected to PWM discriminators 422 and 424, respectively. The flight control electronics unit 450 carried by the missile has PWM demodulators 452 and 454 connected to the receiver discriminators and to pulse width modulators 456 and 458. The outputs of the modulators are fed to power amplifiers 460 and 462 which control hot gas valves 464 and 466 as described later. The missile 400 carries the missile I.R. receiver 410 and the modulated I.R. source 490 at its rear adjacent the missile nozzle 402 (Fig. 3, not shown). In operation, a tank gunner, utilizing an " Integrated Sight " visually acquires his target, fires the missile, and maintains the cross-hairs of his sight on the target until missile impact. The integrated sight may consist of a telescope having a common mounting with tracker 100 and being aligned therewith. The optical axes of tracker 100 and transmitter 300 are accurately boresighted to the line of sight established by the operator's sight. Transmitter 300 and the telescope share part of a common aperture in a tank turret. This is accomplished by an interference filter-mirror (not shown) which separates the visual energy (0À4-0À8 microns). Tracker 100 with a much smaller lens 102, uses the remainder of the common aperture but does not share its portion of the aperture with either the transmitter 300 or the telescope. This separation is required since the tracker utilizes energy from the missile modulated source 490 in the 0À4-1À0 micron region for shimmer cancellation. Upon leaving the launch site missile 400 enters the field of view of the tracker and transmitter where its displacement from the line of sight is measured by tracker 100 as an angle by passing the radiation received from source 490 through lens 102 and mirror 108 which passes energy above 1À0 microns through reticle 112. Reticle 112 imparts from zero to 100% amplitude modulation on the signal from the modulator source 490. The per cent modulation varies linearly with distance off the optical axis over a portion of the total field and then remains relatively constant at 100% modulation for the remainder of the field. The A.M. reticle thus encodes position error into a polar co-ordinate signal at the output of signal detector 114, at the reticle rotation rate. The amplitude modulated signal at the output of detector 114 is processed through automatic gain controlled amplifier 116. The shimmer modulator and signal detectors 110 and 114 received programmed bias voltages from source 122 depending on time from missile launch to reduce the large dynamic ranges over which the AGC amplifiers 116 and 124 in each loop must work. The program control is driven from timer circuit 106 which is actuated from a gunner's firing button. The signal channel output is carrier modulated by reticles 112 and also by shimmer. The shimmer detector output is a constant carrier having only the unwanted shimmer modulation. The demodulated output of the shimmer detector can thus be inserted as an open loop AGC control of the signal AGC amplifier 116. This cancels the shimmer noise in the signal. Reference generator 120 is driven by motor 104 and rotated in synchronism with reticle 112 to provide two 90 degree displaced 30 cps. signals which are used in synchronous demodulator 118 to recover the rectangular coordinate error signals. The outputs of demodulator 118, the pitch and yaw error signals representing distance off the LOS to the target are then sent to cant angle corrector 202, where a lg. bias in the pitch axis input for gravity correction is resolved into two axis components. These biases are added to the error signals. The two D.C. error signals, the outputs of the cant angle corrector, are then sent to pulse width modulators 204 and 205 where they are transformed to command signals using four separate frequency tones in a frequency shift keying technique. From the output of the frequency shift oscillators 208 and 210 the " on " and " off " time of the PWM modulators constitutes the " on " and " off " time for two separate frequencies f 1 and f 2 for the yaw command and f 3 and f 4 for the pitch command. The pitch and yaw commands are then amplified in power amplifiers 310 and 310<SP>1</SP> to drive galvanometers 308 and 308<SP>1</SP> in the transmitter. The pitch and yaw optical devices 304 and 304<SP>1</SP> are boresighted with each other and with the operator's telescope. The pitch and yaw commands to the missile are then transmitted through projection optics 306 and 306<SP>1</SP>, which are located in the integrated sight, to the missile receiver 412. The missile receiver optical system collects radiated infra-red radiation from the transmitter and focuses it on PbS detector 414. The output signal from detector 414 is amplified in amplifier 416 and fed to tone separation filters 418 and 420. Frequencies f 1 and f 2 are separated, rectified in opposite polarity, and summed to reproduce the yaw command. The pitch command is reproduced in like manner. A two-axis gyro 470 is used for roll angle sensing and for providing a yaw axis inertial reference. An accelerometer 472 provides a pitch damping reference. The output of yaw PWM demodulator 452 is subtracted from the D.C. output of a yaw gyro equalization network 474 and used to drive pulse width modulator 456. The PWM signal is then used to drive yaw hot gas valve 464, which produces a force to eliminate the yaw displacement error from LOS. The error signals in the pitch and roll channels are used to produce missile force commands from hot gas valves 466 and 476 in the same manner as in the yaw channel. The two-axis gyro is uncaged when the missile is launched and provides roll stabilization and a yaw inertial reference throughout the flight. Control of the missile to the line of sight with the hot gas valve force causes the modulated source 490 to be displaced to the LOS with the tracker error going to zero, thereby closing the outer loop of the guidance system.