|Publication number||US3820742 A|
|Publication date||Jun 28, 1974|
|Filing date||Feb 8, 1965|
|Priority date||Feb 8, 1965|
|Publication number||US 3820742 A, US 3820742A, US-A-3820742, US3820742 A, US3820742A|
|Original Assignee||Watkins R|
|Export Citation||BiBTeX, EndNote, RefMan|
|Referenced by (22), Classifications (7)|
|External Links: USPTO, USPTO Assignment, Espacenet|
United States Patent Watkins MISSILE GUIDANCE AND CONTROL SYSTEM  Inventor: Robert A. Watkins, 142 La Vista Grande, Santa Barbara, Calif.
Filed: Feb. 8, 1965 Appl. No.: 432,076
 3,820,742 June 28, 1974 Primary ExaminerVerlin R. Pendegrass Attorney, Agent, or FirmEdward J. Kelly; Herbert Berl [5 7 ABSTRACT A missile guidance and control system comprising: a
source of infrared radiation carried by a missile to be controlled, means carried by said missile for modulating said radiation, an infrared tracker responsive to said modulated radiation for generating tracking error signals when said missile deviates from a desired path, shimmer cancellation means connected to said tracker for eliminating shimmer noise from the tracking error signals, a computer connected to said tracker and responsive to said error signals for providing missile con-  References C'ted trol signals, an infrared transmitter responsive to said UNITED STATES PATENTS computer for transmitting said control signals to said 2,070,178 2/1937 Pottenger,Jr. et a1. 244/14 H UX missile, a missile carried rece ver for receiving said 2,513,367 7/1950 Scott 250/203 control signals, and means carried by said missile and 9 o n 4/1 D X responsive to the output of said receiver for control- 3,l8l,8l4 5 I965 Pittman 4 Claims, 6 Drawing Figures I AU 1 \1C1'1 SI1EQU1 EM E NT l TRACKER I TRANSMITTER I TRACKER MOTOR TIMER F1R|NG I I OPTICS T emu SIGNALI I "-I02 l I i I no Q I l 104 W F I PROJECTION PROJECTION I LQQ OPTICS OPTICS oerscron I20 I I PITCH YAW I I t I \II I [1/ \\I III! I I MODULAHW l I i i I RETICLE REFERENCE 308 304 \\I, 304' 308 I I l1 I I14 22 I l I I GALVO 21 OPTlCS OPTICS GALVO PbS BIAS MIRROR s MIRROR I J I l PROGRAM I24 I I I I l DETECTOR I 302 i// 302' 1 I i I REFERENCE 1 W I I AGC'd J 126 AMPUFIER 1 SIGNAL AGC LAMP LAMP l AMPLIFIER l I SHIMMER l I a REFERENCE I 3I0 I I znnon I GALVO GALVO I I DRIVER omven I SIGNAL-0R I AMPLIFIER AMPLIFIER I l---- -----I COMPUTER s 4 s 4 I YAW PITCH 20o FREQUENCY IERROR ERROR CLOCK W SHIFT IANGLE ANGLE 205 FREQUENCY osclttAronl' PITCH COMMAND I [202\ r H ERROR 205 i m PI c I CANT (DISTANCE OFF L05) PULSE 208 1, 1 1 m W ERROR W "2315? W I CORRECT (DISTANCE OFF L05) MODULATG" 2 9 YAW COMMAND LCQJTPANGILEIPEUL 2 J PATENTEDJUN28 I974 SHEET 3 BF 3 FIG. 3
FIELD OF VIEW FIG. 5 G
Robert Awoiki INVE R.
MISSILE GUIDANCE AND CONTROL SYSTEM This invention relates in general to missile guidance and control systems and more particularly to a system which utilizes a tracker that follows a modulated source of infrared radiation carried by the missile to be tracked.
In conventional and prior art infrared guidance and control systems, the radiance of the infrared source is the only detected quantity. These systems have the disadvantage that severe background problems arise unless some form of space filtering is used. The present invention overcomes this disadvantage by utilizing a high frequency modulation of the infrared source on the object to be tracked. Since the modulation of radiation from "natural objects is generally at a low frequency, it is possible to choose a sufficiently high modulation frequency so that background effects are negligible.
Accordingly, it is an object of this invention to provide an infrared guidance and control system that is substantially free of the effects of background radiation.
It is a further object of this invention to provide a guidance and control system wherein an optical system is programmed to the known range and position of an object to be tracked and controlled.
A still further object ofthis invention is to provide an infrared guidance and control system wherein the spurious effects of atmospheric modulation are reduced or eliminated.
The above objects and advantages of the present invention are obtained by providing a missile with a source of modulated infrared radiation and providing a means for measuring any deviation of this radiation from an established line of sight. A computer is provided for converting this measured deviation to pitch and yaw error signals which are in turn transmitted to the missile with an infrared transmitter. The missile is also equipped with an infrared receiver and flight control electronics to provide corrections to maintain the missile on the line of sight path. This system automatically guides the missile to its target utilizing infrared tracking and command links capable of performing under the deleterious effects of background noise, dust, haze. and the rocket motor plume.
A preferred embodiment of the invention will now be described in detail in conjunction with the accompanying drawings wherein:
FIG. I is a block diagram of the ground based equipment of a preferred embodiment of a guidance and control system according to the present invention;
FIG. 2 is a block diagram illustration of the missile carried equipment of the preferred embodiment of the invention;
FIG. 3 is a pictorial representation of a missile showing the location of the modulated source of infrared radiation and the infrared receiver of the preferred embodiment of the invention;
FIG. 4 is a cross sectional view of a modulated light source according to the invention; and
FIGS. 5a and 5b illustrate the chopper reticle located in the tracker portion of the system.
A missile guidance and control system according to the invention consists of launch site equipment shown in FIG. 1 and missile carried equipment shown in FIG. 2. The launch site equipment includes a tracker 100 which receives modulated infrared radiation from a missile to be tracked and provides a measure of any deviation of the missile from an established line of sight. This information is fed to a computer 200 which computes pitch and yaw correction commands to be transmitted to the missile by an infrared transmitter 300. The missile carries an infrared receiver 410, flight control electronics 450 and a modulated source of infrared radiation 490.
Tracker includes an optical device 102; which in a preferred embodiment is a variable focal length objective lens, programmed in time according to the known range vs. time characteristics of the missile being tracked. This programming is accomplished with motor 104 and timer circuit 106. Optical device 102 focuses the received radiation which passes through a beam splitting device 108, which can be for example a silicon mirror, which separates the radiation into two portions at the 1.0 micron point. The energy below 1.0 micron is fed to a PbS detector 110. This channel is used as a shimmer modulation cancellation loop. The 1.0 to 2.5 micron energy passes through a rotating half plane A.M. reticle I12 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 ampli-. fier 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 I10 and 114, an automatic gain controlled reference amplifier 124, and an AGC amplifier shimmer reference 126.
Computer 200 includes a 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 modulators 204 and 205.
Another piece of equipment located at the launch site is an infrared transmitter 300. The transmitter comprises two identical channels for transmitting pitch and yaw commands to the missile. Lamp 302, in the pitch channel, provides the infrared 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 identifying the parts as the pitch channel except that primes have been added.
The missile carried infrared receiver 410 utilizes an optical system 412 for focussing 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 and 420 which are in turn connected to PWM discriminators 422 and 424, respectively.
Another missile carried unit is the flight control electronics 450. This unit 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. The flight control unit will be discussed with more detail in the description of the operation of the system.
As shown in FIG. 3, missile 400 carries the missile LR. receiver 410 and the modulated LR. source 490 at the rear of the missile adjacent the missiles nozzle 402.
In operation, a tank gunner or other operator, 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 gunner thus establishes a line of sight to the target. 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 operators sight. Transmitter 300 and the telescope share part of a common aperture, in one embodiment this aperture is in a tank turret. This is accomplished by an interference filter-mirror (not shown) which separates the visual energy (0.4-0.8 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 0.4-1.0 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 mea sured by tracker 100 as an angle off the line of sight. This is accomplished by passing the radiation received from source 490 through lens 102 and mirror 108 which passes energy above 1.0 microns through reticle 112. Reticle 112 imparts from zero to 100 percent amplitude modulation on the signal from the modulated 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 percent modulation for the remainder of the total field. The A.M. reticle thus encodes position error into a polar coordinate 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. Gain control of this signal is required to prevent saturation and loss of error signal information. The shimmer modulation and signal detectors 110 and 114 receive 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 gunners firing button. The signal channel output is carrier modulated by reticle 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 effectively 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 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 then 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, and f for the yaw command and fig and f, for the pitch command.
The pitch and yaw commands are then amplified in power amplifiers 310 and 310 to drive galvanometers 308 and 308' in the transmitter. The pitch and yaw optical devices 304 and 304' are boresighted with each other and with the operators telescope. The pitch and yaw commands to the missile are then transmitted through projection optics 306 and 306, which are located in the integrated sight, to the missile receiver 412.
The missile receiver optical system collects radiated infrared 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 and f; are separated, rectified in opposite polarity, and summed to reproduce the yaw PWM command. The pitch command is also 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 dc. 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.
An embodiment of the modulated light source is illustrated in FIG. 4. Four tungsten sources 491 operating at a filament temperature of 3,200 K are disposed in a housing 492 and are positioned to the rear of a like number of optical projection systems 493. A common modulating reticle 494 is mounted on the shaft of and driven by a governed d.c. motor 495.
The missile carried source of modulated infrared radiation is imaged as an out-of-focus blur in the plane of tracker reticle 112, which is essentially a straight edge rotating about an axis located on that edge. Thus, the amount of low frequency modulation which the chopping reticle imposes on the missile source signal is a function of the radial displacement of the blur image from the axis of reticle rotation. For example, as illustrated in HQ. 5a, if the blur spot 1121; is located on the axis of rotation of reticle 112, no modulation is introduced, but if the blur spot is completely clear of the axis, the reticle modulation is percent. The phase of the reticle modulation indicates the direction in which the blur spot is off the reticle axis (i.e., up, down, left or right).
While this invention has been defined with reference to specific embodiments thereof, it will be appreciated that many modifications and changes may be made by those skilled in the art without departing from the spirit of the invention, as defined in the appended claims.
1. A missile guidance and control system comprising: a source of infrared radiation carried by a missile to be controlled, means carried by said missile for modulating said radiation, an infrared tracker responsive to said modulated radiation for generating tracking error signals when said missile deviates from a desired path, shimmer cancellation means connected to said tracker for eliminating shimmer noise from the tracking error signals, a computer connected to said tracker and responsive to said error signals for providing missile control signals, an infrared transmitter responsive to said computer for transmitting said control signals to said missile, a missile carried receiver for receiving said control signals, and means carried by said missile and responsive to the output of said receiver for controlling the flight of said missile.
2. A missile guidance and control system comprising: a self propelled missile, a modulated source of infrared radiation carried by said missile, an optical device positioned at a missile tracking site for focussing radiations received from said missile, a beam splitting device adjacent said optical device for separating the received radiation into two portions, a reticle disposed adjacent said beam splitting device for amplitude modulating the received radiation, a detector disposed adjacent said reticle for providing a signal related to position error of said missile, an error signal demodulator connected to said detector for providing pitch and yaw error signals, a reference generator connected to said demodulator, first and second pulse width modulators connected to said demodulator, first and second frequency shift oscillators connected to and driven by said pulse width modulators for generating pitch and yaw command signals, an infrared transmitter connected to said oscillators for transmitting said pitch and yaw commands to said missile, an infrared receiver carried by said missile for receiving said pitch and yaw commands, means responsive to the output of said receiver for separating said pitch and yaw commands, and flight control means carried by said missile for controlling the path of said missile in response to said pitch and yaw signals.
3. A missile guidance and control system as set forth in claim 2 wherein said flight control means comprises, a two axis gyro for sensing roll angle and for providing a yaw axis reference signal, a pitch accelerometer for providing a pitch damping reference signal, means for combining said pitch damping reference signal and said pitch command, means responsive to said combined pitch damping reference and pitch command signal for controlling the pitch of said missile, means connected to said gyro for combining said yaw axis reference signal and said yaw command signal, means responsive to said combined yaw reference and yaw command signal for controlling the yaw of said missile, and means responsive to the roll angle output of said gyro for controlling the roll of said missile.
4. A missile system comprising: a self propelled missile; a source of modulated infrared radiation carried by said missile; an infrared receiver positioned at a missile tracking site and including an optical device for focusing radiation received from said source, a rotating reticle disposed in the path of said focused radiation for amplitude modulating said radiation, and a signal detector disposed adjacent said reticle for receiving radiation from said optical device and generating an electrical signal in response thereto; means for programming said optical device to known range and time characteristics of said missile; means connected to and responsive to said receiver for generating a tracking error signal when the missile deviates from a desired path; and a beam splitting device disposed adjacent said optical device for separating said focused radiation into two portions; a second signal detector disposed adjacent said beam splitting device for receiving a portion of said radiation for generating a shimmer cancellation signal; and means applying said shimmer cancellation signal to said error signal generating means for eliminating shimmer noise from the error signal.
- UNITED STATES PATENT OFFICE CERTIFICAT F RR I N Patent No. 3, 20 1 Dated June28, 1974 Robert A; Watkins i i I Inventor(s) v I It is certified that error appears in the above-identified patent and that said Letters Patent are hereby corrected es shown below:-
On the cover sheet in s er t + Assigne e; United States' of America as represented bythe Secretary of the Arm)" Signed and sealed this 3rd day of December 1974.
1 (SEAL) AtteSt: I
McCOY GIBSON JR. 0. MARSHALL DANN Attesting Officer Commissioner of Patents i UNITED S" 1ATES PATENT OFFICE C RT E OF CORRECTION 2 Dated J n 2118, 7
7 Patent No.
Robert A. Watkins T Inventor(s) r It is certified that error appears in the above-identified patent and that said Letters Patent are hereby corrected as shown below:-
On the cover sheet insert  Assignee; United Statesief America as represented by the Secretary of the Arl ny' Signed and seeled this 3rd day of' December 1914;.
(SEAL) Arrest; v
- MCCOY My GIBSON JR. c. MARSHALL N Arresting Officer Qonunissioner of Patents
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|U.S. Classification||244/3.11, 244/3.16, 244/3.14|
|International Classification||F41G7/30, F41G7/20|