US3905563A - System for controlling a missile motion in the homing mode - Google Patents

System for controlling a missile motion in the homing mode Download PDF

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Publication number
US3905563A
US3905563A US354386A US35438673A US3905563A US 3905563 A US3905563 A US 3905563A US 354386 A US354386 A US 354386A US 35438673 A US35438673 A US 35438673A US 3905563 A US3905563 A US 3905563A
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Prior art keywords
missile
signal
acceleration
error signal
target
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Expired - Lifetime
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US354386A
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English (en)
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Saburo Nagoshi
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Subaru Corp
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Fuji Heavy Industries Ltd
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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
    • F41G7/224—Deceiving or protecting means
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F41—WEAPONS
    • F41G—WEAPON SIGHTS; AIMING
    • F41G7/00—Direction control systems for self-propelled missiles
    • F41G7/34—Direction control systems for self-propelled missiles based on predetermined target position data
    • G—PHYSICS
    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
    • G05D1/00—Control of position, course, altitude or attitude of land, water, air or space vehicles, e.g. using automatic pilots
    • G05D1/12—Target-seeking control

Definitions

  • ABSTRACT A system for controlling a missile motion in the homing mode comprises means for designating by an error GENERATOR V [30] Foreign Application Priority Data signal a direction in which a missile is guided in'the Sept. 28, 1972 Japan 47-97345 homing mode and 1 for P with F error signal a predetermined biasing signal imparting [52] US. Cl. 244/3.15 acceleration in two dinfensions f f to 51 Int. Cl.
  • the present invention relates to a control system for guiding a missile in homing mode to a target, with the missile evading defensive action by anti-aircraft guns and anti-missile missiles.
  • a missile is generally controlled in the homing mode in a manner tending to reduce the error angle of the missile flight path to the target at the initial stage of derivation, thereby nearly maintaining a constantbearing course. Consequently, the missile flight path straightens gradually (unless the motion of a target necessitates considerable maneuvering of the missle) so that an intercept computer makes an exact forecast of the future missile position enabling effective defense with anti-aircraft guns or anti-missile missiles.
  • object of the present invention is to provide means in such a system for so improving the control of missile navigation as to enable its effective evasion of defensive action all along the flight path.
  • my present system includes aboard an in-flight homing missile such conventional components as sighting means for continuously determining the location of a line of sight from the missile to the target, detector means responsive to the sighting means for generating the aforementioned error signal as a function of a deviation of the missile flight path from the line of sight, steering means for changing the course of the missile, and control means therefor connected to the detector means for delivering to the steering means a corrective signal derived from the error signal.
  • I further provide biasing means for superimposing upon this corrective signal a lateral-acceleration signal effective in a direction perpendicular to an invariable reference direction generally parallel to the flight path, such as the axis (X) of a set of three orthogonal coordinate axes X, Y, Z.
  • a differential signal in which this lateral-acceleration signal is subtracted from the output of the feedback path involved in steering diminishes the error signal in the input of the control means so that its effect upon the steering means tends to change the existing flight path.
  • FIG. I is a diagram illustrating in principle a control system embodying the present invention.
  • FIG. 2 is a diagram illustrating details of the control system
  • FIG. 3 is a geometrically exaggerated diagram of the missile deviation
  • FIG. 4 is a plot of angular parameters in a tracking system
  • FIG. 5 is a set of graphs of the missile motion determined by one-dimensional bias in the direction of deviation
  • FIG. 6 is a vector diagram of the acceleration bias for the missile maneuvering
  • FIG. 7 is an orthogonal coordinate system indicating the missile deviation with two-dimensional bias
  • FIG. 8 is a graph of the missile acceleration required for a bias according to FIG. 7;
  • FIG. 9 is an orthogonal coordinate system indicating the missile deviation with three-dimensional bias
  • FIG. 10 is a graph of the missile acceleration required for a bias according to FIG. 9;
  • FIG. 11 is a graph of total kill probabilities in terms of lateral acceleration
  • FIG. 12 is an orthogonal coordinate system indicating the missile deviation with two-dimension bias
  • FIG. 13 is a graph of the missile acceleration required for a bias according to FIG. 12;
  • FIG. 14 is an orthogonal coordinate system indicating the missile deviation with three-dimensional bias
  • FIG. 15 is a graph of the missile acceleration required for a bias according to FIG. 14.
  • FIG. 16 is a graph of the missile maneuvering acceleration necessary to keep a definite miss-distance from a surface-to-air missile.
  • FIG. 1 In a control system according to my present invention there is illustrated in block form in FIG. 1 and error signal supplied from a missile-deviation detector a, which signal is not applied immediately to a control device b but is compounded with a lateral-acceleration bias from a generator c so that the compound signal is applied to the control device b as a control signal.
  • FIG. 2 shows details of such a control system in which the control device b includes a servo amplifier 1, a servo valve 2 and an actuator 3 for a steering unit 4, serially arranged in a forward path, as well as a feedback path from steering unit 4 to the input of amplifier 1 comprising aerodynamic motion transfer function unit 5 working into an accelerometer 6.
  • a signal fed back from the output of the accelerometer 6 diminishes the bias signal from the output of generator 0 to produce a differential signal.
  • a control signal is obtained by deducting the differential signal from the error signal supplied by the deviation detector a.
  • the feedback signal could come from the output of a rate gyro, or the deflection angle of a control surface could be utilized for the same purpose.
  • the Missile Motion The X-axis of the missile coordinate system extends along a theoretical missile-collision course as shown in FIG. 3, wherein the homing trajectory of the missile leads from a starting position to a pinpoint S of the target. Furthermore, the missile positio n T at a time t is defined by coordinates x Y Proportional navigation requires that the change of the missile flight direction always be directly proportional to the angular rate of change of a line of sight which leads from the missile to the target (with a proportionality factor N where n missile lateral acceleration, i.e. the acceleration directed along the Y-axis of the missile coordinate system but with a negative sign g gravitational acceleration V missile velocity Now substitute d from Eq. (1) into Eq. (I).
  • the line-of sight rate (dA- /dt) is measured by a tracking device which may be a radar, an optical system, or an infrared system.
  • the tracking device points its antenna in the direction of the line of sight to the target, as accurately and as quickly as possible.
  • a line-of-sight angle relative to the fixed reference line A boresight angle of a tracking antenna relative to the same fixed reference line (if the tracking were perfect, the antenna boresight would coincide with the line of sight.)
  • e error angle i.e. the angle corresponding to the difference between A and M
  • e O For a perfect tracking system having no time lag, e O.
  • the instantaneous line-of-sight angle M is given by:
  • the missile velocity at a time t is determined by Therefore, the remaining distance of the missile flight path is 0 VTdt i t) 2 l 7 v 1;.
  • the hitting time is obtained from Eq. (10) which includes the hitting range (R from the target. 2.
  • Eq. (10) which includes the hitting range (R from the target. 2.
  • the missile velocity at a time t is determined by:
  • the hitting time is obtained from Eq. (13) which again includes the hitting range (R from the target.
  • SAM Surface-to-air missile
  • V 2 mach velocity
  • a time required for acceleration 680/20.4g 3.4 (sec).
  • the SAM velocity may be assumed as 2 mach by subtracting a dead time from the time required for acceleration, this dead time being 3. 4 sec X k 1.7 sec.
  • Eq. (15) employs a time of 2.2 secas mentioned above, which is given by the delay occurring between the calculation of the missile data and the SAM firing, added to the dead time, i.e.
  • FCS predicts the future position of the missile pointed at a location C in FIG. 3. Since the calculation of the FCS is based on position, velocity and acceleration data of the missile, the calculated time t,; does not coincide with the hitting time i (though the former co- 2 (11 I: ).ean'n incides with the latter with one-dimensional acceleration).
  • a predicted hitting time of the SAM is defined as in Eq. (15).
  • the calculated time I is determined by solving Eqs. (17) and (18) for anti-aircraft guns'and by solving Eqs. (17) and (19) for the SAM, in using the technique of successive approximation.
  • miss-distance m may be written as where m component in the Y-axis direction m 32 component in the Z-axis direction I.
  • m component in the Y-axis direction m 32 component in the Z-axis direction I One-dimensional Acceleration- The Y-axis component is Similarly, the Z-axis component is Supposing the flight path of the shell to be straight, the displacement of the shell along the Y-axis at the hitting time t, is
  • the miss-distance along the Z-axis maybe written as The miss-distance will be obtained by substituting Eqs. (24) and (26) in Eq. (20).
  • the SAM Navigation The FCS predicts the factors of the missile in order to count the future position C of the missile aimed at by the SAM when the missile reaches a point F on the missile bearing course.
  • the SAM starts along a collision course to the left of time t of 2.2 (sec) after the PCS prediction, this collision course being illustrated by the straight line SO in FIG. 3.
  • the miss-distance of the SAM to the missle is based on the unavoidable fact that as the calculated maneuvering acceleration of the SAM attains and surpasses the g-limit, the actual maneuvering acceleration reach- 1)
  • 1,, es its saturation value. V1 (44) Accordingly, the evaluation of the components of the SAM miss-distance in the Y-axis and Z-axis directions must be executed with the addition of factors of gsaturation and desaturation. As the final period of the SAM derivation involves the fixed g-saturation, the following equations are established:
  • Eq. (14) can be used with anti-aircraft guns, and Eq. (15) with the SAM.
  • the miss-distance will be obtained by substituting m
  • the miss-distance of the shell is derived from Eqs. and m of Eqs. (39) and (40) into Eq. (20). (20), (21 and (22), and the SAM motion from the ap- According to the foregoing, the SAM miss-distance li rio of Eqs, (31 (32), (33), (34) and (35) as well is vectorially defined by Y-axis and Z-axis coordinates.
  • FIG, 3 i di t th t Equations (31 and (33) are available to calculated the miss-distance in the event that the bearing courses of R (V +V the missile and SAM are based on the same coordinate (45) system; otherwise, the value of Y and Z must be replaced by others obtained by coordinate conversion.
  • the missile motion is determined from Eq. (42), the hitting time from Eq. (44), the predicted time from Eq. (14) and the miss-distance of the shell from Eq. (21).
  • V 0.9 mach 1 0.5 sec 1,, sec
  • FIG. 11 This diagnosis indicates the need for increasing the evading capability of the missile, because the kill probability of the shell with reference to a missile subjected to lateral acceleration pursuant to this invention is limited to a value of 41% (n B 4.0), according to a curve I, in comparison with the case of a homing missile of conventional navigation where the limit is 99.5% (n 0).
  • the cumulative kill probability of the anti-aircraft gun is represented by a curve II in FIG. 11 on the assumption of conditions indentical with those of the first example, i.e. with the missile not deviated in a seaskimming flight. Curve II indicates a reduction of the kill probability to a value of 6% (n 2 3.0), i.e. the missile deviation with two-dimensional maneuvering acceleration produces an evading efficiency higher than that given by the missile deviation with onedimensional maneuvering acceleration.
  • the cumulative kill probability of the anti-aircraft gun is determined by another curve III of FIG. 1 1 under conditions identical with those of the second example.
  • the kill probability of the gun is further reduced to a value of 4% (n'flnmr) 3.0), i.e. the evading efficiency of the missile in this case is still higher than that given by the missile deviation withtwo-dimensional maneuvering acceleration.
  • V 2.0 mach TM 0.5 sec maneuvering-acceleration limit of SAM 14g (relative to both the Y-axis and the Z-axis) hitting distance of the SAM 1,000 4,000m
  • the missile maneuvering accelerations along the Y-axis or the Z- axis are plotted in full lines in FIG. 16 for a missdistance of 20m.
  • Cruve IV represents the case where the missile and the SAM share the same set of coordinates 6 0), whereas curve V applies where their coordinates differ (6 45) with the Y-axis and the Z-axis of the missile include respective angles of 45 with the corresponding axes of the SAM.
  • the SAM parameters are determined under conditions similar to those of fourth example, the missile maneuvering accelerations along the Y-axis or the Z-axis being given by dotted lines in FIG. 16 and the required maximum maneuvering acceleration of the missile is about 5.5g in the case of 0 0 (Curve V1) 5.2 g in the case of 6 45 (Curve VII).
  • a system for guiding an in-flight homing missile to a target comprising:
  • sighting means for tracking the location of a line of sight from the missile to the target; detector means responsive to said sighting means for generating an error signal depending upon any deviation of the missile flight path from said line of sight;
  • control means for said steering means connected to said detector means for delivering a corrective signal derived from said error signal to said steering means;
  • biasing means connected to said control means for superimposing upon said corrective signal to a lateral-acceleration signal effective in a direction perpendicular to an invariable reference direction generally parallel to said flight path.
  • a system for guiding an in-flight homing missile to a target comprising:
  • sighting means for tracking the location of a line of sight from the missile to the target
  • detector means responsive to said sighting means for generating an error signal depending upon any deviation of the missile flight path from said line of sight;
  • control means for said steering means connected to said detector means for delivering a corrective signal derived from said error signal to said steering means;
  • biasing means connected to said control means for superimposing upon said corrective signal a lateralacceleration signal effective in a direction perpendicular to an invariable reference direction generally parallel to said flight path and further adding an ecceleration bias to the missile derivation direction.
  • a system for guiding an in-flight homing missile to a target comprising:
  • sighting means for tracking the location of a line of sight from the missile to the target
  • control means for said steering means connected to said detector means for delivering a corrective signal derived from said error signal to said steering means, said control means establishing a set of three coordinate axes including a first axis generally parallel to said flight path and two other axes orthogonal thereto; and
  • biasing means connected to said control means for superimposing upon said corrective signal two lateral-acceleration signals effective in directions parallel to said other axes.
  • a system for guiding an in-flight homing missile to a target comprising:
  • sighting means for tracking the location of a line of sight from the missile to the target
  • detector means responsive to said sighting means for generating an error signal depending upon any deviation of the missile flight path from said line of sight;
  • control means for said steering means connected to said detector means for delivering a corrective signal derived from said error signal to said steering means, said control means establishing a set of three coordinate axes including a first axis generally parallel to said flight path and two other axes nal so as to obtain a control signal.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • General Engineering & Computer Science (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Radar, Positioning & Navigation (AREA)
  • Remote Sensing (AREA)
  • Physics & Mathematics (AREA)
  • General Physics & Mathematics (AREA)
  • Automation & Control Theory (AREA)
  • Aiming, Guidance, Guns With A Light Source, Armor, Camouflage, And Targets (AREA)
US354386A 1972-09-28 1973-04-25 System for controlling a missile motion in the homing mode Expired - Lifetime US3905563A (en)

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JP9734572A JPS552555B2 (fr) 1972-09-28 1972-09-28

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Cited By (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2474686A1 (fr) * 1980-01-29 1981-07-31 Europ Propulsion Systeme d'auto-guidage simplifie pour engin du type obus ou roquette
US4347996A (en) * 1980-05-22 1982-09-07 Raytheon Company Spin-stabilized projectile and guidance system therefor
WO1983001298A1 (fr) * 1981-10-08 1983-04-14 Skarman, Bengt Procede et appareil de guidage d'un corps aerodynamique ayant un dispositif de radioralliement
US4456862A (en) * 1982-09-22 1984-06-26 General Dynamics, Pomona Division Augmented proportional navigation in second order predictive scheme
US4541591A (en) * 1983-04-01 1985-09-17 The United States Of America As Represented By The Secretary Of The Navy Guidance law to improve the accuracy of tactical missiles
EP0222571A3 (fr) * 1985-10-31 1988-05-04 British Aerospace Public Limited Company Guidage de missile sur ligne de visée
WO1994000731A1 (fr) * 1992-06-30 1994-01-06 Grushin Petr D Procede et dispositif de regulation de la pression d'admission d'un projectile
US5379968A (en) * 1993-12-29 1995-01-10 Raytheon Company Modular aerodynamic gyrodynamic intelligent controlled projectile and method of operating same
US5425514A (en) * 1993-12-29 1995-06-20 Raytheon Company Modular aerodynamic gyrodynamic intelligent controlled projectile and method of operating same
US7350744B1 (en) * 2006-02-22 2008-04-01 Nira Schwartz System for changing warhead's trajectory to avoid interception
US20090173820A1 (en) * 2008-01-03 2009-07-09 Lockheed Martin Corporation Guidance system with varying error correction gain

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPS57162028U (fr) * 1981-04-07 1982-10-12

Citations (3)

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Publication number Priority date Publication date Assignee Title
US2992423A (en) * 1954-05-03 1961-07-11 Hughes Aircraft Co Rocket launch control systems
US3140482A (en) * 1954-06-07 1964-07-07 North American Aviation Inc System providing error rate damping of an autonavigator
US3167276A (en) * 1961-09-29 1965-01-26 Honeywell Inc Control apparatus

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2992423A (en) * 1954-05-03 1961-07-11 Hughes Aircraft Co Rocket launch control systems
US3140482A (en) * 1954-06-07 1964-07-07 North American Aviation Inc System providing error rate damping of an autonavigator
US3167276A (en) * 1961-09-29 1965-01-26 Honeywell Inc Control apparatus

Cited By (15)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4397430A (en) * 1980-01-29 1983-08-09 Societe Europeenne De Propulsion Simplified homing system for a missile of the shell or rocket type
EP0033283A3 (en) * 1980-01-29 1981-12-02 Societe Europeenne De Propulsion Societe Anonyme Dite Simplified self-steering system for a missile such as a shell or rocket
FR2474686A1 (fr) * 1980-01-29 1981-07-31 Europ Propulsion Systeme d'auto-guidage simplifie pour engin du type obus ou roquette
US4347996A (en) * 1980-05-22 1982-09-07 Raytheon Company Spin-stabilized projectile and guidance system therefor
US4529151A (en) * 1981-10-08 1985-07-16 Saab-Scania Aktiebolag Method and an apparatus for steering an aerodynamic body having a homing device
WO1983001298A1 (fr) * 1981-10-08 1983-04-14 Skarman, Bengt Procede et appareil de guidage d'un corps aerodynamique ayant un dispositif de radioralliement
US4456862A (en) * 1982-09-22 1984-06-26 General Dynamics, Pomona Division Augmented proportional navigation in second order predictive scheme
US4541591A (en) * 1983-04-01 1985-09-17 The United States Of America As Represented By The Secretary Of The Navy Guidance law to improve the accuracy of tactical missiles
EP0222571A3 (fr) * 1985-10-31 1988-05-04 British Aerospace Public Limited Company Guidage de missile sur ligne de visée
WO1994000731A1 (fr) * 1992-06-30 1994-01-06 Grushin Petr D Procede et dispositif de regulation de la pression d'admission d'un projectile
US5379968A (en) * 1993-12-29 1995-01-10 Raytheon Company Modular aerodynamic gyrodynamic intelligent controlled projectile and method of operating same
US5425514A (en) * 1993-12-29 1995-06-20 Raytheon Company Modular aerodynamic gyrodynamic intelligent controlled projectile and method of operating same
US7350744B1 (en) * 2006-02-22 2008-04-01 Nira Schwartz System for changing warhead's trajectory to avoid interception
US20090173820A1 (en) * 2008-01-03 2009-07-09 Lockheed Martin Corporation Guidance system with varying error correction gain
US7795565B2 (en) * 2008-01-03 2010-09-14 Lockheed Martin Corporation Guidance system with varying error correction gain

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JPS4970499A (fr) 1974-07-08
FR2201773A5 (fr) 1974-04-26
DE2325355B2 (de) 1981-04-30
DE2325355A1 (de) 1974-04-04
JPS552555B2 (fr) 1980-01-21

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