|Publication number||US3737122 A|
|Publication date||Jun 5, 1973|
|Filing date||May 7, 1971|
|Priority date||May 7, 1971|
|Publication number||US 3737122 A, US 3737122A, US-A-3737122, US3737122 A, US3737122A|
|Inventors||B Papa, E Solov|
|Original Assignee||Singer General Precision|
|Export Citation||BiBTeX, EndNote, RefMan|
|Patent Citations (5), Non-Patent Citations (2), Referenced by (17), Classifications (5)|
|External Links: USPTO, USPTO Assignment, Espacenet|
United States Patent 91 Solov et a1.
[ 1 June 5, 1973 OTHER PUBLICATIONS Guided Missles Fundamentals, Air Force Manual 52-31, Superintendent of Documents, Government Printing Office, Wash. 25, DC. p. 428-435.
ABSTRACT A range control system is provided for a missile or projectile which utilizing its inflight prediction of the impact range error, provides means for correcting for this error. The system is carried in the vehicle, and it includes a comparator system which compares outputs from a vehicle-mounted accelerometer system with predetermined values. Any differences are fed to a function generator which produces a control signal. The control signal from the function generator operates a reaction control motor which, in turn, compensates for any variations of the vehicle from its course, so that the predetermined range of the vehicle 6 Claims, 3 Drawing Figures l e/ocz/y M2), 5;
I I leellw'aac/er I Irvine 01/ 69/1 Talley/afar pzrcix/f  TACTICAL MISSILE RANGE CONTROL SYSTEM Air Force Training Manual, page 462.  Inventors: Edwin G. Solov; Bernard F. Papa,
both of Wayne, NJ.
 Assignee: Singer-General Precision, Inc., Little Falls, NJ. i D Primary Exammer--Robert F. Stahl Filed: May 7, 1971 Assistant Examiner-James M. I-Ianley [211 App. 141,313 Attorney-S. A. Giarratana  11.5. C1 ..244/3.21, 235/150.26  Int. Cl. ..F41g 7/00, F42b 15/02 1  Field of Search ..244/3.15, 3.21, 3.22; 235/150.2,.150.26, 150.27
 References Cited UNITED STATES PATENTS 3,363,858 1/1968 Dobbins et a1. ..244/3.14 3,568,954 3/1971 3,184,182 5/1965 2,932,467 4/1960 3,412,961 11/1968 is maintained.
/Z4 8 A t/bre t ll f eyi ar I M... I 1 4/11 4r 15 a4 1 f lavfg/ K Flor/Aw! daze/earner I PATENTEU JUN 5 I973 SHEET 2 OF 2 INVENTO 9-5 /01 6. So/a V (9 er e/dk'papa mm M. Kauai;
TACTICAL MISSILE RANGE CONTROL SYSTEM BACKGROUND OF THE INVENTION As is well known, if a projectile departs from its course at any time during its flight, these departures will affect the preset range of the vehicle and cause it to depart from its predetermined impact point. Even in the case of guided missiles, if the vehicle leaves its predetermined course, there is a variation in gravity acceleration which produces errors in the data programmed into the guidance system. The vehicle-mounted range control system of the present invention provides a low cost means of achieving range control for battlefield and other tactical missiles and projectiles. The system provides for the vehicle to be held at all times on its predetermined course, so that its established range is accurately and precisely maintained.
BRIEF DESCRIPTION OF THE DRAWINGS FIG. 1 is a schematic representation of a projectile or missile, and showing in block form a control system incorporating the concepts of the invention, the control system being represented as mounted within the vehicle;
FIG. 2 is a more detailed block diagram of a range control system representative of one embodiment of the invention; and
FIG. 3 is a schematic circuit diagram of the system shown in block diagram of FIG. 2.
DETAILED DESCRIPTION OF THE ILLUSTRATED EMBODIMENT In FIG. 1, the missile or projectile is represented as 10, and the control system representative of one embodiment of the invention, and which is mounted within the vehicle is shown, as mentioned above, in block form. The system of FIG. 1 includes, for example, an accelerometer 12 which is designated an axial accelerometer, and which is mounted with its sensitive axis coincident with the vehicle longitudinal axis. The system also includes a comparator system designated by the block 14 and which, as will be described, and as shown in FIG. 2, includes three integrators 16, 18 and 20. The integrator 20 provides a track and hold function; and the integrators 18 and 16 serve for doubly integrating the output of the axial accelerometer 12.
The control system also includes a small reaction control rocket motor 22, or equivalent device, which is aligned with its thrust axis perpendicular to the center line of the vehicle 10 and in a vertical plane. A clock or timer circuit 24 is provided for controlling the sequence of the system. It is assumed that the vehicle incorporates the usual system for maintaining its roll orientation relative to the vertical, and such a system is not shown.
As shown in FIG. 1, the output from the comparator system 14 is fed to a function generator 26 whose output is applied to a zero detector system 28 which, in turn, controls the reaction control motor 22. As shown in FIG. 2, the zero detector system 28 includes a zero detector number 1, and a zero detector number 2, the outputs from the zero detectors number 1 and 2 being passed through an and gate 30 to the reaction control motor 22.
As will be described, the vehicle 10 is maintained on its predetermined course by controlling the length of time the reaction control motor 22 is operated. This control is obtained by comparing the actual readings derived from the accelerometer 12 with calculated readings, and by feeding the difference into the function generator 26. The function generator 26 then provides the control signal for the reaction motor 22.
As shown in more detail in FIG. 2, the clock circuit 24 controls three single-pole double-throw relay switches designated Y1, Y2 and Y3. When these switches are in their down position, the system is controlling. However, when the clock circuit moves the switches to their upper position, the output from the axial accelerometer 12 is integrated first in the integrator 18 so that the accelerations are transformed into velocities ('v), and the output from the velocity integrator 18 is integrated by the integrator 16 so that the velocities are transformed into arc lengths (5).
At the same time, the accelerometer track and hold integrator 20 is connected to the output of a comparator circuit 50, the inputs of which are connected to the output of the accelerometer 12 and to the output of the integrator 20. The output (a) from the integrator 20 is also applied to a comparator circuit 52 in which it is compared with a predetermined nominal acceleration value (n Likewise, the output (V) of the integrator 18 is applied to a comparator 54 in which it is compared with a predetermined nominal value (v,,), and the output of which is applied to a gain Ky. The output (s) of the integrator 16 is applied to a comparator 56 in which it is compared with a predetermined nominal value (s and the output is applied to a gain K The outputs of the gain K Ky and K, are all applied to a summing network 58, and the output 8R of the summing network 58 is applied to the function generator 26.
The output of the clock circuit (1) is applied to a comparator 60 in which it is compared with a predetermined time (t,,) and the output of the comparator 60 is applied directly to the zero detector No. 1 and to a summing network 62. The output (At) of the function generator 26 is applied to the summing network 62, as is a predetermined value (T The output of the network 62 is applied to the zero detector No. 2.
Although the various components designated above will be shown and described in somewhat more detail in FIG. 3, it should be pointed out at this time that the individual components such as the clock circuit 24, the axial accelerometer 12, the integrators, the function generator 26, the summing networks, the comparators, the zero detectors, the and" gate 30 and the reaction control motor 22 are aall well known to the art, so that a detailed description of these individual components is deemed to be unnecessary.
In the operation of the system of FIG. 2, any departures of the vehicle 10 from its flight path will produce error signals, and these rror signals will cause the time histories of the outputs from the accelerometer 12, and will cause the outputs of the integrators 18, 16 and 20 to differ from certain predetermined nominal values (v,,), (s,,) and (0,) respectively. At some time (2,.) during the flight the clock circuit 24 operates the relay switches Y1, Y2 and Y3 to establish the input of the three integrators 16, 18 and 20 at a reference level so that the outputs of the three integrators remain constant thereafter. At this point, the three constant values of the three integrators, namely s, v and a are compared in the respective comparators 56, 54 and 52 with their respective nominal values s,,, v, and a,,.
The three difference outputs from the comparators 56, 54 and 52 are linearly combined in the summing network 58 to provide an estimate of the error that would occur if no correction were made. This error is designated 6R. In order to effectuate the correction, the small reaction motor 22 is utilized. As mentioned above, the reaction motor 22 is oriented with its thrust axis in a vertical plane and normal to the center line of the vehicle 10. The direction of thrust of the reaction control motor 22 is generally upward.
Independent of the aforesaid in-flight estimate of range error, the motor 22 begins to fire at a predetermined time (t) during the flight if the range error is estimated at zero. That is, if the estimate shows that the vehicle is on its correct course towards the target, then the motor 22 burns for a duration of T, seconds. If the estimate is that the vehicle will fall short of the target, the duration of the burn of the motor 22 is extended for AT seconds, where AT is a function of the estimated range error as derived from the function generator 26. This function is determined in the function generator by simulation and in most cases will turn out to be a simple proportionality. If, on the other hand, the estimate is that the vehicle will overshoot the target, the duration of the burn of the motor 22 is shortened by a period of time (AT) which is, again, the same function pf estimated range error.
The implementation of the foregoing is achieved, for example, by the system of FIG. 2 in which the output of the accelerometer 12 is fed to the integrators 16, 18 and 20 through the aforesaid integrator for said relay switches Y1, Y2 and Y3. The output (a) from the integrator 20 is fed back to its input through the comparator 50 form a track and hold circuit. The output (v) of the velocity integrator 18 is fed through the relay switch Y1 to the integrator 16 to provide the arc length data (s). The three relay switches Y1, Y2 and Y3 are actuated by the clock 24 which serves as a timer to control the entire process. It is pointed out that the relay switches Y1, Y2 and Y3 may be solid state type switches, rather than electromagnetically operated relay switches, if so desired.
The respective outputs s, v and a are compared in the summing or comparator networks 56, 54 and 52 with the respective nominal values s v and a,,, and the coefficients of combination, that is the gains K Ky and K are determined using a so-called static Kalman-filter, or linear regression, technique, as designated by the blocks K Ky and K,,.
At a pre-selected observation time (I the clock 2d causes the three switches Y1, Y2 and Y3 to establish the three integrator inputs 16, 18 and at a reference level, as mentioned above, and this in turn causes their outputs to hold at their values at the time t,,. The three outputs from the integrators, summed as described plied to the zero detector No. 2 by the summing network 62, where it is subtracted from the sum T, AT. Normally, |AT| T,,, so that T +AT is positive. Thus, when t-t s 0,- the net input to the zero detector No. 2 is positive, and its output O is a l As t-t becomes more positive, a point is finally reached when t-t B 2 T,,+AT. At this time, the required burn time of the motor 22 has ended, and the input to the zero detector No. 2 goes negative, setting O to 0.
The signals Q, and Q, are fed to the and gate 30 whose output controls the reaction control motor 22.
When 0 Q, l, the motor 22 burns. Otherwise,
is inoperative. The operation of the reaction control motor may be summarized by the following table:
Reaction Control I I Q1 Q Motor 22 r 2,, 0 0 1 Off 0 s r-z, T AT 1 1 On 1-; 2 T,, A T 1 0 Off As illustrated in FIG. 3, the block diagram of FIG. 2
may be implemented, for example, by operational amplifiers Al-A9, connected as shown. The and" gate 30 is formed, for example, by a pair of diodes D1 and D2, and a power amplifier A8 is interposed between the and gate and the reaction control motor 22. The operational amplifiers A1, A3 and A4 form the integrators l6, l8 and 20 respectively.
The invention provides, therefore, a low cost control system capable of providing accurate in-flight range controls for tactical missiles. It will be appreciated that although a particular embodiment of the invention has been shown, alternate mechanizations are possible. For example, a differential digital analyzer (DDA), or other digital means may be used to implement the foregoing integrations and computations. Moreover, the rocket motor 22 may be replaced by fins, or other steering means for the vehicle. Also, for more inexpensive systems, one or more of the integrators 16, 18, 20 may be omitted, although this would result in a concomitant degradation of system performance. Alternately, additional sensors and/or integrators may be incorporated to enhance system performance.
Therefore, although a particular embodiment of the control system of the present invention has been shown and described modifications may be made. It is intended to cover all modifications which fall within the scope of the invention in the following claims.
APPENDIX Derivation of Estimator The following linear regression, or static Kalman filter technique, represents a method which may be used to determine the estimating gains K Ky and K First consider the three integrator outputs a, v and s as an observation vector y:
The vector y can be related to a vector of error sources, 6, and a measurement noise vector, v, by
y Hg v where: H is a Jacobian matrix (of partials) of the form H=[8Y,/8e,],i= l, 2, 3;j= l, n
where g is an n-dimensional vector of the form The error sources 6,,
6,, include the contributors to range error, such as It is desired to estimate, from the observation 1, the 20 on-target error, e If this estimate is to be a linear function of y, then we can write the estimate as,
R 5 where: 2,; is the estimate of e and [9' is the row vector ofgains [K Ky K The error in the estimate is:
. 30 The criterion for selecting the gain vector, E, 15 a least square error criterion. The expectation of the square of the error is given by To minimize this function with respect to 5, we take the gradient of E [E lll Then,
1 an; ut
w E [11'] E +1) +y)'] ruz E ("ii E [12' l gu'l tx 1'] Since 3 and g are independent,
w HE [2 1] E [z 2'] Let:
Since 68 and y are independent, E [6,; y] Q and Q E [68 s1= E [(92) 5] 55] Qeu ed Substituting these results into the expression for & gives:
A In mechanizing this system, we denote 2,; E 6R, and
81% l [a, v, s]
E (K Ky, K [HC H C,,9 HC
What is claimed is:
l. A vehicle-mounted range control system for a missile, projectile, or other vehicle comprising:
accelerometer means mounted on the vehicle with its sensitive axis coincident with the longitudinal axis of the vehicle and having an output at which electrical signals representative of accelerations of the vehicle in the direction of said longitudinal axis are produced;
a first and second integrator each having an input and an output;
a first switch having a first position for coupling said first integrator to the output of said accelerometer and a second switch having a first position for coupling the input of said second integrator to the output of said first integrator;
comparator means coupled to the output of said integrators for comparing the output therefrom with predetermined nominal values;
a third integrator arranged as a track and hold circuit having an input and an output, with said output being coupled to said comparator means;
a third switch having a first position for coupling the input of said third integrator to said accelerometer and a second position for coupling said third inte' grator input to a point of reference potential;
a course control unit for the vehicle; and
function generator means coupled to said comparator means and responsive to the outputs therefrom for controlling the operation of said course control unit.
2. The range control system defined in claim 1, in which each of said first and second switches includes a second position for coupling the inputs of said first and second integrators to a point of reference potential, and which includes a clock timing circuit for actuating said first, second, and third switches selectively to their first and second positions.
3. The range control system defined in claim 1 in which said first integrator produces output signals representative of the velocity of the vehicle along the longitudinal axis, said second integrator is selectively coupled to said first integrator for producing output signals representative of the arc length of the vehicle along its flight path, said third integrator produces track and hold output signals representative of the vehicle acceleration, and further including a summing circuit rebetween said function generator means and said course control unit, and further comparator means interposed between said clock timer circuit and said zero detector circuit means for establishing a predetermined time interval at which said course control unit is to be activated.
6. The range control system defined in claim 1, in which said course control unit comprises a reaction control rocket motor.
|Cited Patent||Filing date||Publication date||Applicant||Title|
|US2932467 *||Aug 11, 1955||Apr 12, 1960||English Electric Co Ltd||Ballistic missiles|
|US3184182 *||Jan 18, 1960||May 18, 1965||Chrysler Corp||Pulsed thrust velocity control of a projectile|
|US3363858 *||Sep 23, 1958||Jan 16, 1968||Navy Usa||Doppler homing system|
|US3412961 *||Jul 27, 1966||Nov 26, 1968||Nasa Usa||Means for suppressing or attenuating bending motion of elastic bodies|
|US3568954 *||Apr 4, 1966||Mar 9, 1971||Mccorkle William C Jr||Directional control-automatic meteorological compensation (d.c.-automet) inertial guidance system for artillery missiles|
|1||*||Air Force Training Manual, page 462.|
|2||*||Guided Missles Fundamentals, Air Force Manual 52 31, Superintendent of Documents, Government Printing Office, Wash. 25, D.C. p. 428 435.|
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|US4456862 *||Sep 22, 1982||Jun 26, 1984||General Dynamics, Pomona Division||Augmented proportional navigation in second order predictive scheme|
|US5248114 *||Jun 20, 1974||Sep 28, 1993||Ankeney Dewey P||Adaptive autopilot|
|US6345785 *||Nov 7, 2000||Feb 12, 2002||The United States Of America As Represented By The Secretary Of The Army||Drag-brake deployment method and apparatus for range error correction of spinning, gun-launched artillery projectiles|
|US6378435 *||Apr 3, 1995||Apr 30, 2002||General Dynamics Decision Systems, Inc.||Variable target transition detection capability and method therefor|
|US6672533||Aug 9, 2000||Jan 6, 2004||Saab Ab||Method and guidance system for guiding a missile|
|US7446291 *||Oct 3, 2005||Nov 4, 2008||Lockheed Martin Corporation||Augmented proportional navigation guidance law using angular acceleration measurements|
|US7795565 *||Jan 3, 2008||Sep 14, 2010||Lockheed Martin Corporation||Guidance system with varying error correction gain|
|US7834300||Feb 7, 2006||Nov 16, 2010||Bae Systems Information And Electronic Systems Integration Inc.||Ballistic guidance control for munitions|
|US20070241227 *||Feb 7, 2006||Oct 18, 2007||Zemany Paul D||Ballistic Guidance Control for Munitions|
|US20090173820 *||Jan 3, 2008||Jul 9, 2009||Lockheed Martin Corporation||Guidance system with varying error correction gain|
|US20150301521 *||Apr 17, 2015||Oct 22, 2015||Masitek Instruments Inc.||Systems, methods, devices and computer readable medium for real and near-real time sensor data capture and analysis|
|EP0231161A2 *||Jan 28, 1987||Aug 5, 1987||Aktiebolaget Bofors||Apparatus for reducing projectile spread|
|EP0231161A3 *||Jan 28, 1987||Aug 31, 1988||Aktiebolaget Bofors||Apparatus for reducing projectile spread|
|EP1225327A1 *||Oct 12, 2001||Jul 24, 2002||Saab Dynamics Aktiebolag||Range control of a rocket-propelled projectile|
|WO2001014820A1 *||Aug 9, 2000||Mar 1, 2001||Saab Bofors Dynamics Ab||Method and guidance system for guiding a missile|
|WO2006086528A2 *||Feb 7, 2006||Aug 17, 2006||Bae Systems Information And Electronic Systems Integration Inc.||Ballistic guidance control for munitions|
|WO2006086528A3 *||Feb 7, 2006||Sep 28, 2006||Mark A Carlson||Ballistic guidance control for munitions|
|U.S. Classification||244/3.21, 701/505|