US3219294A - Homing system for guided missiles - Google Patents

Homing system for guided missiles Download PDF

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Publication number
US3219294A
US3219294A US157506A US15750661A US3219294A US 3219294 A US3219294 A US 3219294A US 157506 A US157506 A US 157506A US 15750661 A US15750661 A US 15750661A US 3219294 A US3219294 A US 3219294A
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United States
Prior art keywords
missile
control
axis
target
lateral
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Expired - Lifetime
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US157506A
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English (en)
Inventor
Welti Arno
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Siemens Schweiz AG
Albiswerk Zuerich AG
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Siemens Albis AG
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F41—WEAPONS
    • F41G—WEAPON SIGHTS; AIMING
    • F41G7/00—Direction control systems for self-propelled missiles
    • F41G7/20—Direction control systems for self-propelled missiles based on continuous observation of target position
    • F41G7/22—Homing guidance systems
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F41—WEAPONS
    • F41G—WEAPON SIGHTS; AIMING
    • F41G7/00—Direction control systems for self-propelled missiles
    • F41G7/20—Direction control systems for self-propelled missiles based on continuous observation of target position
    • F41G7/22—Homing guidance systems
    • F41G7/2253—Passive homing systems, i.e. comprising a receiver and do not requiring an active illumination of the target
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F41—WEAPONS
    • F41G—WEAPON SIGHTS; AIMING
    • F41G7/00—Direction control systems for self-propelled missiles
    • F41G7/20—Direction control systems for self-propelled missiles based on continuous observation of target position
    • F41G7/22—Homing guidance systems
    • F41G7/2273—Homing guidance systems characterised by the type of waves
    • F41G7/2293—Homing guidance systems characterised by the type of waves using electromagnetic waves other than radio waves

Definitions

  • My invention relates to a method and system for homing guidance of missiles, according to which a missile, herein understood to denote any object travelling through space and provided with suitable directional control means such as control surfaces or control jets, guides itself toward a target, which may likewise be travelling in space, to bring about a collision or near-collision of missile and target.
  • a missile herein understood to denote any object travelling through space and provided with suitable directional control means such as control surfaces or control jets, guides itself toward a target, which may likewise be travelling in space, to bring about a collision or near-collision of missile and target.
  • my invention relates to a homing method and system of the passive type in which energy from the target itself, for example heat radiation, is sensed by suitable detectors, for example infrared sensors, in the missile, for thereby tracking the target and determining and ⁇ correcting any departures from the collision course.
  • suitable detectors for example infrared sensors
  • a missile moving in space and a likewise moving target will collide under all circumstances if the target image remains at standstill on the viewing or picture area of the detector in the missile, it being only presumed that the collision has not yet occurred and that target and missile do not travel parallel to each other.
  • the immobility of the target image is tantamount to the fact that initially the line of sight from missile ⁇ to target retains its direction in space constant during the interval of time under observation. Consequently, if the missile is controlled so that the direction-al invariance of the line of sight is preserved at any moment, the impact condition is also satisfied at any moment.
  • Such directional invariance of the line of sight can be controlled after measuring its direction with the aid of target tracking, and referring it to the direction of a spacially fixed axis relative to a polar coordinate system fixed with respect to the missile.
  • the impact condition is met, according to the basic concept of the invention,
  • the impact condition can be graphically represented in the so-called phase space, a diagram which represents the condition of two magnitudes in mutual correlation.
  • phase space a diagram which represents the condition of two magnitudes in mutual correlation.
  • the angle differences Aa and Ak at an image point P are correlated as the two Cartesian coordinates of that point.
  • the impact condition is satisfied if the image point P remains at standstill in the phase space.
  • the spherical distance d would be constant only as long as, aside from the angle difference Aa and AA, the angle A2 also remained constant.
  • One reason for such limitation is the fact that the determination of the direction C and K (line of sight, and spacial- 1y fixed axis respectively) are also subject to limitation, for example, due to the constructionally limited turning range of a radiation sensing head, or the frame or gimbal stops of a gyro system.
  • the missile is supposed to take the course of most favorable impact probability. This requires roll-motion stabilization of the missile.
  • the missile is to be so controlled as to minimize the duration of the transition interval during which the missile passes from one to another collision course upon occurrence of course disturbance. This means, relative to the phase space, that the return of the image point to a standstill position again satisfying the impact condition should always occur on a straight path.
  • a stability condition satisfying those requirements can be derived from the phase space shown in FIG. 3. Assume that the image point has migrated from the initial standstill point P to a point P1 due to lateral deflection of the missile.
  • the directional regulator is now called upon to actuate the control surfaces or other control means in such a manner that the error from P0 to P1, after termination of a transitional interval, is corrected by placing the target image onto a new standstill point P2.
  • al and M denote the angular coordinates of the sighting line C from the missile to the target
  • a2 and A2 denote the angular coordinates of the stable axis K, these angular coordinates being continuously measured relative to a polar coordinate system fixed with respect to the missile.
  • the apparatus according to the invention for performing the missile-guiding method explained above comprises a gyro to provide a stable reference axis and a radiation detector for tracking the target. These two components control respective angle transmitters or resolvers which furnish the instantaneous llateral and elevational angles of the line of sight and of the stable axis respectively with respect to a polar coordinate system fixedly related to the missile.
  • the apparatus further comprises two networks for producing two control signals which form a measure of the time derivation (rate of change) of the difference between the lateral angles on the one hand, and of the time derivation (rate of change) of the difference between the elevational angles on the other hand; and these two signals control respective motors correlated to two coordinate axes of the flight directional control means in the sense of reducing the rate of angular change.
  • the system further comprises a differential amplifier to which one of the above-mentioned two control signals is directly supplied and to which the other control signal is supplied through a regulating amplifier whose gain is regulated in dependence upon the output signal of the differential amplifier in the sense toward reducing this output signal, and the adjusting member for gain regulating simultaneously controls an angle transmitter (resolver system) whose output signal constitutes a measure of the instantaneous amplification factor of the regulating arnplifier and, upon differentiation, controls a motor correlated to a third axis of the iiight directional control means toward reducing the change of this output signal.
  • an angle transmitter resolveer system
  • the above-described homing system is provided with another radiation detector which cooperates with the gyro for tracking a given point of the stable reference axis to thereby control the one appertaining resolver.
  • FIGS. l, 2 and 3 are explanatory diagrams already described above;
  • FIG. 4 is a block diagram of a missile-borne homing control apparatus, according to the invention.
  • FIG. 5 shows schematically the basic design -of an infrared radiation detector employed in the embodiment according to FIG. 4;
  • FIG. 6 is a circuit diagram of electrical equipment forming part of the same apparatus.
  • the guidance apparatus is provided with a radiation detector 1 for tracking the target Z.
  • the detector 1 is horizontally and vertically rotatable.
  • a set of servomotors 2 is connected with the detector unit for directing the sensing axis onto the target Z.
  • the detector 1 furnishes an error voltage which constitutes a measure for the departure of the optical axis from the line of sight C. This error voltage is impressed upon a servo-control stage 3 where it is converted to control signals for the servos 2 in the sense required for reducing the above-mentioned departure.
  • Servo devices of this type for automatic tracking of a target are known as such.
  • the radiation detector comprises an object lens 4, an image-area scanner 5, a field lens 6, and a radiation-responsive cell 7.
  • the optical and scanning components are accommodated within a housing 4a mounted on a supporting structure which comprises two shafts 8 and 9 extending perpendicular to each other and permitting the optical axis of the detector to be turned about the respective shaft axes.
  • the shafts 8 and 9 are connected with respective servomotors in the set 2 (FIG. 4).
  • the scanner 5 is a rotating raster disc with a raster pattern so designed that the beam of light impinging upon the photocell 7 is modulated in dependence upon the locality of the target-image point on the image area. Hence the output voltage of the cell 7 is modulated in the same manner.
  • This cell voltage is compared with a reference voltage indicative of the direction of the optical axis of the radiation detector and derived from the turning motion of the radiation detector. As long as the image point of the target coincides with the optical axis, the result of the comparison is zero and no error voltage is produced. When the image point of the target moves away from the optical axis, the voltage comparison results in a finite error voltage which is effective in the servo-control unit 3 to cause actuation of the servomotors in the sense of eliminating the departure of the target image point from the optical axis. More cornplete details can be had from the above-mentioned references.
  • the angles a1, A1 represent the coordinates of the line of sight C relative to a missilefixed polar coordinate system.
  • the lateral angle al of the missile target sighting line, the lateral angle a2 of the stable reference axis relative to a missile-fixed polar coordinate system, the elevation angle A1 of the missile target sighting line and the elevation angle A2 of the stable reference axis relative to the missile-fixed polar coordinate system may be continuously measured in any suitable manner known. There are a number of known methods for continuously measuring the angular displacement of an axis.
  • the space-fixed axis K is constituted by the main axis of a gyro system 11.
  • the position of the gyro axis relative to the missile-fixed polar coordinate system can be determined by ascertaining a point of the gyro axis.
  • a light source 12 is used for this purpose.
  • the source is mounted at the end of the gyro shaft and acts upon a radiation detector 13.
  • the detector 13 may be of the same type as the detector 1 used for tracking the target, and is likewise equipped for automatic tracking of the light source 12.
  • the detector 13 is connected with a servo-control stage 14 and a servomotor set 15 which correspond to the respective devices 3 and 2.
  • Any directional change of the missile axis manifests itself in a departure of the image point produced by the light source 12, from the optical axis of the radiation detector 13.
  • the resulting corrective motion is transmitted to the resolver system 16 mechanically coupled with the radiation detector 13.
  • the resolver system 16 operating in the same manner as the resolver system 10, furnishes two output voltages proportional to the angular coordinates a2, A2 of the stable reference axis inthe missile-fixed polar coordinate system.
  • any other electromagnetically detectable reference point may be provided on the gyro shaft and a correspondingly sensitive cell in the radiation detector.
  • the portion enclosed in FIG. 4 by a dot-and-dash line and denoted by A.C. constitutes a regulating system interposed between the above-described two groups of selftracking sensing components and the flight-direction control motors properof the missile.
  • the intermediate system A.C. is essentially an analog computer which determines from the measured input magnitudes a1, a2, A1 and A2, the control magnitudes required for flight control.
  • the computer portion A.C. is equipped with two networks 17 and 18 for forming the angle difference AAzAl-Ag and trazar-a2 respectively.
  • the computer further comprises two networks 19 and 20 for forming the differential quotients d(AA)/dt and KAM/dt.
  • control magnitudes supplied from the networks 19 and 20 are supplied to respective control motors 21 and 22 for actuating the elevational control (H) and lateral control (L) of the missile. These two control magnitudes also pass into a computer stage 23 in which a control signal is generated proportional to the time change in the ratio of the mentioned two control magnitudes, and the latter control signal is applied toy a control motor 24 for actuating the roll control (Q) of the missile.
  • FIG. 6 illustrates the circuit diagram of the regulating computer portion A.C. according to FIG. 4.
  • the volt-ages UA1, UA2 Ual and Uaz supplied from the resolvers and 16 (FIG. 4) and impressed upon respective pairs of input terminals 25, 26, 27 and 28, are indicative of the continuously measured angular coordinates A1, A2, al
  • the voltages may be linearly proportional to these coordinates.
  • the voltages UA1 and UA2, impressed across respective resistors 29 and 30 are series-opposed to each other so that the ydifference voltage UAA: UA1-UA2 is obtained across the series connection of the two resistors 29 and 3f).
  • Connected to the series connection is a differentiating member consisting of a longitudinal capacitor 31 and a transverse resistor 32. Across the resistor 32 there appears a voltage UH proportional to the differential quotient d(AA)/dt.
  • the voltage UH is -impressed upon output terminals 33 where it is available as control magnitude for elevational control. That is, this voltage UH serves to control the operation of motor 21 (FIG. 4) in the event a change in elevational angle is necessary for returning the missile to a collision course.
  • the resistors 34 and 35 connected across respective terminal pairs 27 and 28 are impressed by the voltages Ual and Ua2 in mutally series-opposed relation so that the series connection of the two resistors 34 and 35 furnishes the difference voltage UAazUal-Uaz.
  • a differentiating member is connected to the series connection and consists of a longitudinal capacitance member 36 and a transverse resistance member 37.
  • the voltage Us appearing across the resistance member 37 is proportional to the differential quotient d(Az)/dt. This Voltage is impressed across terminals 38 for lateral control of the motor 22 (FIG. 4).
  • the regulating portion A.C. of the system further comprises a differential amplifier 39 which receives the control voltage UH directly, and which is supplied with the control voltage US through a variable regulating amplifier 40.
  • the output of the regulating amplifier 40, and one input to the amplifier 39 is equal to the input voltage Us to the amplifier 40 times its amplification V; ⁇ or the product USV.
  • Connected to the output leads of the differential amplifier 39 is an actuator or motor 41 having a shaft 42.
  • the angle of rotation of the shaft 42 is designated The shaft 42 is coupled with a control member for varying the gain in the regulating amplifier 40 so that the amplification V of the regulating amplifier 40 is inversely proportional to the rotational angle This accomplishes an inverse feedback from the amplifier 39 through the motor M and the amplifier 40 so as to reduce the output voltage at the differential amplifier 39. This equalizes the input voltages to the amplifier.
  • the amplification V is controlled to be inversely proportional to so that l VN- Substituting for V in the equation USV: UH we obtain l Us?- Un and Us NH Substituting for Us and UH in the equation US/ UH We obtain .
  • a direct-voltage source 43 furnishing a constant voltage E
  • a potentiometer rheostat 44 whose slider 45 is mounted on the shaft 4Z so that the slider 45 taps olf the voltage E proportional to the rotation angle of the shaft 42.
  • a differentiating member consisting of a longitudinal capacitor 46 and a transverse resistor 47. This differentiating member forms the time derivation (rate of change) of the voltage E, whereby the voltage across the output terminals 48 is proportional to the differential quotient:
  • the homing method and ight regulating system according to the invention affords a number of advantages.
  • the means for performing the method, involving a zero principle are relatively simple and require relatively few and rather compact components, in comparison with known homing methods and systems.
  • the trajectory or collision course of the missile is not predetermined.
  • the flight guidance system sets the missile to a new impact course, thus always providing a new definition of the path of light.
  • the high degree of precision required for remote control operation is not necessary, and the tolerance limits can be kept conveniently wide. All regulating, controlling and computing operations are performed with relative magnitudes only. As a result, the invention affords an adaptation of the missile flight control to the intended purpose to an extent far beyond that heretofore attainable.
  • a missile-borne system for homing a guided missile onto a target comprising directional control means having a lateral control motor for control about a lateral axis of the missile, a vertical control motor for control about a vertical axis of the missile and a longitudinal control motor for control about a longitudinal axis of the missile; gyroscopic reference -means having a stable axis relative to space and indicating means for indicating said stable axis; a radiation sensor responsive to radiation from the target directed to receive radiation from said target for detecting departure of the sensor axis from the missile-target sighting line; detecting means directed at the indicating means of said gyroscopic reference means for detecting departure of said stable axis from a missile-fixed reference position; two servomotor means of which one is connected with said sensor for causing it to track the target and the other is connected to said detecting means for causing it to track said stable axis; two angle-transmitting revolving means connected to said respective radiation sensor and detecting means to issue
  • said gain regulating means comprises an actuator electrically connected with said differential amplifier output circuit and mechanically connected with said regulating amplifier for varying its gain, said angle transmitter comprising a source of constant direct voltage, a potentiometer connected across said source and having a slide contact, said slide contact being mechanically connected with said actuator to be displaced thereby in accordance with changes in gain of said regulating amplifier, said differential-forming network being electrically connected to said slide contact and said source.
  • a missile-homing system as claimed in claim 1, wherein said stable axis is constituted by a shaft of said gyroscopic reference means, said means for detecting departure of said stable axis from a missile-fixed reference position comprising a radiation member on said shaft, and a sensor responsive to said member and connected to said other servomotor means for tracking said member.
  • a missile-borne system for homing a guided missile onto a target comprising directional control means having a lateral control motor for control about a lateral axis of the missile, a vertical control motor for control about a vertical axis of the missile and a longitudinal control motor for control about a longitudinal axis of the missile; gyroscopic reference Imeans having a stable axis relative to space and indicating means for indicating said stable axis; a radiation sensor responsive to radiation from the target directed to receive radiation from said target for detecting departure of the sensor axis from the missile-target sighting line; detecting means directed at the indicating means of said gyroscopic reference means for detecting departure of said stable axis from a missile-fixed reference position; two servomotor means of which one is connected with said sensor for causing it to track the target and the other is connected to said detecting means for causing it to track said stable axis; two angle-transmitting revolving means connected to said respective radiation sensor and detecting means to issue respective

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Electromagnetism (AREA)
  • Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)
US157506A 1960-12-07 1961-12-06 Homing system for guided missiles Expired - Lifetime US3219294A (en)

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CH1366560A CH385034A (de) 1960-12-07 1960-12-07 Einrichtung zur selbsttätigen Regelung der Bewegung eines Flugkörpers mit Zielsuchlenkung

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Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3494576A (en) * 1967-03-06 1970-02-10 Telecommunications Sa Self-contained and automatic guidance system for directing a missile towards a radiation-emitting target
US4219170A (en) * 1977-07-08 1980-08-26 Mcdonnell Douglas Corporation Missile roll position processor

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE1781233C1 (de) * 1967-09-11 1985-10-31 British Aerospace Plc, London Verfahren und Vorrichtung zur Steuerung von Flugkörpern

Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2557401A (en) * 1945-01-10 1951-06-19 Arma Corp Remote control apparatus
US2992423A (en) * 1954-05-03 1961-07-11 Hughes Aircraft Co Rocket launch control systems
US3005981A (en) * 1944-05-19 1961-10-24 Rca Corp Radar control system for glide path control of aircraft

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3005981A (en) * 1944-05-19 1961-10-24 Rca Corp Radar control system for glide path control of aircraft
US2557401A (en) * 1945-01-10 1951-06-19 Arma Corp Remote control apparatus
US2992423A (en) * 1954-05-03 1961-07-11 Hughes Aircraft Co Rocket launch control systems

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3494576A (en) * 1967-03-06 1970-02-10 Telecommunications Sa Self-contained and automatic guidance system for directing a missile towards a radiation-emitting target
US4219170A (en) * 1977-07-08 1980-08-26 Mcdonnell Douglas Corporation Missile roll position processor

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SE305373B (is") 1968-10-21
GB943310A (en) 1963-12-04
CH385034A (de) 1965-02-26

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