US3295794A - Launcher and capture computer - Google Patents

Description

Jan. 3, 1967 I. H. SCHROADER ET AL 3,295,794

LAUNCHER AND CAPTURE COMPUTER 5 Sheets-Sheet 1 Filed July 24, 1963 MN 2004mm 39:: wozwmwta 85m 53m Irvin H. Schroeder Melvm E Hosea mm zocomw Nabisco III.) ll W N 1 m 85% mm 85mm mmEm zoo fizsixwomm EEMES 5&8 5%: Ill A 2925 5 EESEEE 33 15554 M02498 fitimzfifi 82356 MM "29E 1 85350 b 9 Isaac R. Hunter, Jr.

INVENTORS 535 to 2925 5 92 $23 15;? 50m BY Z i g 6 ATTORNEY Jan. 3, 1967 l. H. SCHROADER ET AL 3,295,794

LAUNCHER AND CAPTURE COMPUTER 5Sheets-Sheet 2 Filed July 24, 1963 I l l AEGG III IIIIL FIG. 5.

Irvin H. chroader Melvin E. Hosea Isaac R. Hunter, Jr.

INVENTORS BY MW ATTORNEY Jan. 3, 1967 l. H. SCHROADER ET AL 3,295,794

LAUNCHER AND CAPTURE COMPUTER 5 Sheets-Sheet 5 Filed July 24, 1963 0 \LAUNCHER FIG.'4.

Irvin H. Schroader Melvm E Hosea Isaac R. Hunter, Jr.

INVENTORS ATTORNEY Jan. 3, 1967 1. H. SCHROADER ET AL 3,295,794

LAUNCHER AND CAPTURE COMPUTER Filed July 24, 1963 5 sheets-sheet 4 AT SERVO I l l 80 I l MAG. LAUNCHER i BRAKE LOCK l .1 a2

00 MUG) TAGH as t F' G. 6 a.

Irvin H. Schroader Melvin E. Hosea Isaac R. Hunter, Jr. INVENTORS ATTORNEY Jan. 3, 1967 H. SCHROADER ET AL 3,295,794

LAUNCHER AND CAPTURE COMPUTER 5 Sheets-Shem 5 Filed July 24, 1965 CT A\ (R) BY a i 4 l V ATTORNEY I I I l j A (C) United States Patent 3,295,794 LAUNCHER AND CAPTURE COMPUTER Irvin H. Schroader, Simpsonville, Melvin E. Hosea, Silver Spring, and Isaac R. Hunter, Jr., Edgewater, Md., as-

signors to the United States of America as represented by the Secretary of the Navy Filed July 24, 1963, Ser. No. 297,469 9 Claims. (Cl. 244-14) The present invention relates in general to missile guidance systems and the like and more particularly to a new and improved launcher and guidance transmitter direction system for effecting accurate and dependable beam capture for beam-riding guided missiles.

US. application Serial No. 38,408, filed June 23, 1960, and now US. Patent No. 3,169,727, for Multiple Flight Course Second Order Programmer by Irvin H. Schroader, Melvin E. Hosea and Leo C. Mill-er, discloses a missile guidance system for beam-riding missiles which includes multiple fllight course programming. The present application presents a launcher and guidance transmitter control system for use with the missile guidance system of the referenced copending application and is designed to provide proper positioning of the launcher and guidance vtransmitter during missile launching so as to effect capture of the missile by the guidance beam.

In some respects the problems associated with missile launcher direction and with gun direction are similar in that the end result is to cause a missile or a projectile to intercept a designated point in space. The termination of the problem in a gun direction system occurs upon impact of the projectile with a target. As it concerns the launcher, the target is a designated point, at or near beam center of the guidance transmitter, from which the mid-course trajectory toward the ultimate target of the missile is begun. The problem of defining this point, referred to as the missile capture point, is the task of a launcher direction system. This system must establish the conditions which will enable a missile to arrive at the capture point in a specified length of time by appropriately positioning both the guidance transmitter beam and the launcher. The command position of the guidance transmitter beam should be determined from target position data in order that the capture point be in an autimum position for the missile with respect to the targe intercep point.

The ideal conditions for the direction of a launcher and a guidance transmiter would be to launch the missile from the site of the guidance transmitter and at the corresponding angular position. Assuming a condition of Zero wind velocity and neglecting the effects due to gravity, the missile would under these circumstances be in the capture position from the time of launching. However, because of obvious necessity the launcher and guidance transmitter must be widely separated. The result is a serious problem of parallax which arises to complicate any computation directed toward defining the capture point. On top of this problem the problems of wind velocity, wind direction, and gravity drop must also be considered.

A computer capable of commanding the launcher and guidance transmitter of a beam-rider guidance system so as to effect capture of the missile by the guidance beam must have the ability to compute the correct angular position for both based upon a specified set of input conditions which includes predicted target position intercept, launcher position, slant range to the required point of capture, ballistic wind velocity and direction, parallax angle, and gravity drop. To attain the latter, the instant invention contemplates, among other things, provision of a launcher and guidance transmitter direc- Patented Jan. 3, 1967 tion system which takes into account all of the aforementioned considerations in guiding the missile to accurate beam capture. Means are also provided for adapting the invention to the missile guidance system set forth in copending application Serial No. 38,408 previously mentioned and now US. Patent No. 3,169,727, such that the parameters used in control of that system may be fully applied in control of the direction system of the instant invention.

Accordingly, one object of the present invention is the provision of a new and improsed launcher and guidance transmitter direction system for use in a beam-rider guidance system.

Another object of the present invention is the provision of a launcher and guidance transmitter control system for directing a guided missile into a guidance beam with accuracy and reliability.

A further object of the present invention is the provision of a guidance computer for positioning the guidance transmitter and missile launcher in a beam-rider guidance system in compensation of parallax, wind and gravity errors thereby effecting accurate and dependable beam capture of the missile subsequent to launch.

Still another object of the present invention is the provision in a beam-rider guidance system of a guidance computer for positioning the guidance transmitter and missile launcher so as to effect beam capture of the missile, which computer operates on standard target position information obtainable with standard radars.

Other objects and many of the attendant advantages of this invention will be readily appreciated as the same becomes better understood by reference to the following detailed description when considered in connection with the accompanying drawings, wherein:

FIG. 1 is a schematic representation by block diagram of a complete missile guidance and control system of the type into which the invention may be incorporated;

FIG. 2 is a geometrical representation of the capture problem in the elevation plane;

FIG. 3 is a geometrical representation of the capture problem in three dimensions;

FIG. 4 is a geometrical representation of the parallax correction in the azimuth plane;

FIG. 5 is a simplified schematic representation by block diagram of the launcher and capture computer constructed in accordance with the instant invention;

FIG. 6a is a specific block diagram of the input data converter system to be used with the instant invention; and

FIG. 6b is a specific block diagram of the launcher and capture computer constructed in accordance with the instant invention.

Second order missile programmers of the type disclosed in the aforementioned application Serial No. 38,408 and now US. Patent No. 3,169,727 contemplate the manipulation of various input data to the computer in accordance with generalized systems equations, the constants of which are chosen in accordance with missile characteristics and target considerations to bring the missile to its target in the prescribed manner.

Referring now to the drawings, which illustrate one embodiment of the invention, there is shown in FIG. 1 a schematic representation by block diagram of a complete missile guidance system into which the instant invention may be incorporated.

FIG. 1 shows a search radar 20, incorporating a separate height finder 21, which sends out a radar search beam 13 to a target 12 and in turn receives a reflected signal 14 from the target. Many types of search radar instrumentation are suitable for utilization in carrying out the latter functions, including search radars having built in height finders or those which incorporate such units separately. The output of the search radar 20 and height finder 21 consists of information relating to rough values of target range and an even rougher value of target evaluation. The latter output data is then conveyed as input information to a tracking radar 22, which may take a great variety of forms but is here illustrated as incorporating a three-coordinate system utilizing train, traverse and elevation axes. Target traverse T target elevation E and target train T information is obtained from the search radar 20 and height finder 21 and this information enables the operation system of the tracking radar 22 to either actually lock on the target 12 or to throw the tracking radar into search operation.

The tracking radar 22 transmits its own radar beam 15 to the target 12 and receives a reflected signal 16 therefrom. The output data from the tracking radar 22 consists of target elevation E target train T target traverse T and target-missile slant range difference D D The latter output information from tracking radar 22 is derived for further use through synchro devices physically located at the tracking radar unit itself. Shaft rotations about the various axes of the tracking radar 22 are converted by means of such synchro devices to synchro signals, three wire lines being conventionally embodied for each synchro, to provide outputs from each synchro which are directly proportional to the number of degrees of radar shaft rotation from an established zero reference position.

The synchro output data from the tracking radar 22 is in turn fed to a computer section 23. The latter tracking data is first directed to the input converter servo section 24 of the computer 23 wherein the synchro input information is converted by means of servo-mechanism devices to electrical signals in the form of DC. voltages proportional to the original radar shaft rotation from which the input synchro signals are derived. Such DC. voltage signal form is required for use by the programmer section 25 of the computer 23. The programmer section 25 utilizes the latter DC. voltage outputs of the input converter servos as electrical voltage inputs to a computing section which manipulates such input voltages on the basis of prescribed generalized trajectory equations, and produces DC. output voltages which are then fed to an output converter servo section 26.

The output converter servo section 26 of computer 23 reconverts the DC. output voltage signals from the programmer section 25 to synchro signal form for transmission to and utilization by the guidance radar transmitters denoted generally as 27 in FIG. 1 of the drawings. The latter guidance radar transmitters 27 are physically located near the search and tracking radars and 22, respectively, and the launcher (not shown), or at least within a few hundred feet of the latter to avoid the introduction of severe parallax errors.

The guidance radar 27 transmits a radar signal 17 to the missile 19, but only in a single direction, that is, guidance radar 27 receives absolutely no reflected signal from the missile in the guidance system embodiment illustrated. The major distinction, therefore, between the guidance radar transmiter 27 and the tracking radar 22 is that the guidance transmitter is commanded into position by the output elevation and azimuth signals from the ground computer 23 as opposed to the complete lack of such control over the tracking radar. For some types of program trajectories, however, the guidance radar transmitter 27 may track in range. In such instances the tracking accomplished by the guidance transmitter 27 is distinguished from the operation of the tracking radar 22 in that tracking is not accomplished by the conventional method of receiving a signal from the missile constituting the reflection of an original sign-a1 generated in the guidance transmitter itself. On the contrary, missiles utilizing those types of program trajectories which require tracking in range by the guidance transmitter 27 carry a beacon 28, incorporated into the missile itself,

and which is triggered by the guidance beam 17 from the guidance transmitter 27 to generate a beacon signal 18 of its own. The latter beacon signal 18 is directed from the missile 19 to the guidance transmitter 27 so that the missile 19 is caused to follow the guidance beam as the beam moves in accordance with the program flight trajectory equation. Servomechanism devices within the missile 19 itself steer the missile in accordance with the transmitted guidance signals so that the missile is caused to always remain within the guidance beam after capture, provided the guidance beam moves in such a manner that the missile capabilities of linear velocity and lateral acceleration are not exceeded.

Before control of the missile can be effected, adequate capture of the missile by the guidance beam 17 emanating from the guidance transmitter 27 must be ensured. This feat can only be accomplished through proper positioning of the launcher and the guidance transmitter 27 prior to launch. Based upon target data and the physical relationship between the launcher and the guidance transmitter, a launcher and guidance transmitter direction system, which is incorporated into the ground computer 23, positions the launcher and guidance transmitter in accordance with prescribed generalized flight pattern equations so as to insure subsequent intercept by the missile of the capture point in the beam. These flight pattern equations will be explained by reference to FIGS. 2, 3 and 4.

Looking to FIG. 2 which represents a schematic diagram of the physical relationships in the elevation plane between the giudance transmitter located at G the launcher located at L and the point of intended capture C, the elevation of the missile at capture C with respect to the launcher L after gravity drop and wind corrections are added is ML= L-( (1) Where B is the elevation angle of the launcher, and 8 and w are the gravity drop and wind correction angles, respectively. With respect to the guidance transmitter G the elevation of the missile at capture is GG= L( where e is the parallax angle in the elevation plane between the guidance transmitter and the launcher. Re-

placing the correction terms (5+w-j-e) with a term AE Equation 2 becomes EGG:EL AEGG The instantaneous launcher elevation command derived from basic ballistics is L= T+ r Pr+ GG where E is the instantaneous target elevation, E is the rate of change of E and T is the predicted time from present to intercept of the target. Combining Equations 3 and 4 gives GG== T+ r Pr A similar set of equations can be developed in a like manner for the azimuth plane using a and a for the azimuth parallax and wind correction terms, respectively. A gravity drop term will be absent in this instance for obvious reasons. For the azimuth plane and to doctrine determined by target range and altitude. The angle thus inserted would replace the terms ET+ETTPI in Equation 5.

Considering the condition where the launcher is aimed at a point P in space and referring to FIG. 3, the position of the point P is defined by coordinates D E and A The line L will be the missile path provided there are no outside deflections. If all effects are neglected except that due to cross wind, the missile will be deflected in a slant plane due to the fact that the missile will try to align itself with the wind. As a result, the elevation plane containing the missile and the launcher will rotate in azimuth in a direction which will result in zero angular difference between the missile heading angle and the wind. If it is assumed that the deflection due to a cross Wind is some constant K multiplied by the component of the wind normal to the elevation plane containing L and P, the missile will have been deflected by an amount DCKVW sin (AWAL) s where D is the capture range in feet, V is the wind velocity in knots, A is the wind azimuth and A is the launcher azimuth. As a result of this wind deflection a missile aimed at P and launched from L will arrive at point C. The coordinates of C are D E and A and the length of the arc PC is represented by Equation 8. Projected into the ground plane the coordinates of C are R and A where R =D cos E The angle 0- which represents the change in azimuth due to a cross wind may now be approximated by small angles of o' as KV sin (A A (I cos E L (10) Referring to FIG. 4, it is seen that the parallax angle in the ground plane results from locating the guidance transmitter G at some distance X and Y from the launcher site L. From the figure it is seen that m=\ X +Y sin (AA (11) where A is the azimuth of the launcher with respect to the guidance transmitter, and the parallax angle is l Sm D cos E (12) From Equations 11 and 12 and for small values of a an approximation of the parallax angle a can be shown to be X COS Agq-Y Sin AGG D, cos EL 13 The gravity drop angle in the elevation plane is approximately by with respect to the launcher line of sight, where M has a value of 0.142. The change in elevation angle due to a wind component normal to the slant range D in the elevation plane is computed neglecting the gravity drop angle. The effect of a wind component is assumed to be an angular change given byw=KV cos (A -A' sin E (15) Once the correction terms 6 and w are included in the computation of the problem in the elevation plane, a new point C replaces point C and is defined with respect to the launcher by coordinates D E and A If no difference in height exists between the guidance transmitter and the launcher, and the guidance transmitter is to be aligned with point C, it must be depressed by an amount 6 with respect to angle E as shown in FIG. 3. From FIG. 4 it is seen that Y COS -A 6:M cos E radians and from 'FIG. 3 it is seen that h== /X l-Y cos (AA sin E (17) Since 6 is approximately equal to h/D for small values of c this parallax angle can be approximated as e g tx sin A -I-Y cos A (18) The manner in which the above generalized equations are used to direct the positioning of the launcher and guidance transmitter will now be described in connection with FIG. 5 which shows the basic block diagram of the launcher and guidance transmitter direction system according to the invention. Servo amplifier 30 combines signal values repersenting AE E T and E to produce a signal in accordance with Equation 4 which represents the required launcher elevation E In a like manner, servo amplifier 31 combines signal values A A T and AA to produce a signal in accordance with Equations 6 and 7 which represents the required launcher azimuth A These control signals E and A are applied to launcher 32 in control thereof.

A second pair of signals E and A represent actual repeated values of launcher elevation and azimuth, respectively, are derived from the launcher 32 itself and applied to servo amplifiers 33 and 34, respectively, for use in control of servo resolver unit 35. The resolver unit 35 through reception and processing of required servo data in accordance with Equations 14, 15 and 18 produces one set of control outputs to servo unit 36 representing values of 6, w and 5, respectively, which are combined in servo unit 36 to produce control proportional to AE In a like manner resolver unit 35 operating in accordance with Equations 10 and 13 produces a second set of control outputs which are applied to servo unit 37 and represent values of v and a, respectively. These signals are combined in servo unit 37 to produce control proportional to AA Servo units 36 and 37 control potentiometers 38 and 39 in such a way as to produce signals proportional to A'E and AA which signals are fed back to the input of servo amplifiers 3t) and 31, respectively, thereby completing the servo control loop. The servos 33 and 36 in combination control differential 40 which positions potentiometer 4 1 to produce an output proportional to E for control of the guidance transmitter elevation. In a like manner servos 34 and 37 in combination control differential 4-2 which positions potentiometer 43 to produce an output proportional to A for control of the guidance transmitter azimuth.

The system for effecting beam capture of the missile consists of an input data converter section shown in detail in FIG. 6a and the launcher order and capture computer shown in detail in FIG. 6b. The input data converter section is a dual purpose system providing servo data to both the launcher order and capture com puter, which represents the instant invention, and the multiple flight course programmer, which forms the subject matter of the aforementioned copending application Serial No. 38,408 and now US. Patent No. 3,169,727.

Looking to FIG. 6a, synchro data representative of the target traverse coordinate is fed to the T servo to position the servomotor 51 which operates a potentiometer 52, the voltage obtained from the slider arm 53 of the potentiometer 52 being directly proportional to T The preceding is accomplished by feeding the T synchro signal from tracking radar 22 in FIG. 1 to control transformer 54, which is designed specifically to turn out a sizable voltage to a high impedance input such as that of an amplifier, the output of the control transformer 54 being fed to the servo amplifier 55 which provides sufficient output power to drive the servomotor 51. The servomotor 51 in turn acts as a feedback device to position the rotor of the control transformer 54 in such a direction that the output of the control transformer 54 goes to zero, that is, the system performs as a nullin g type of device wherein an error signal is fed back to the control transformer to reduce its output to zero. At this point the shaft position of the servomotor 51 is in the equilibrium state and is representative of the shaft position of the transmitting synchro at the tracking radar 22. The latter operation succeeds in getting the radar data into the traverse section 50 in the form of a shaft position which is then readily converted to a DC. voltage by the servomotor driven slider arm 53 of the potentiometer 52. The latter potentiometer 52 is center tapped to ground and biased at each of its ends by positive and negative D.C. voltages respectively, typical values of which may be :100 volts DC. In the latter case, therefore, 100 volts D.C. would represent the desired number of degrees of traverse angle to be used as a standard, the voltage output from potentiometer 52 being directly proportional to some fractional portion of the latter standard.

The elevation servo 60 operates in substantially the same manner as the traverse servo 50 in converting E synchro data to DO. voltage form. The E synchro data is fed to a control transformer 64, the output of which is in turn directed to a servo amplifier 65. The output of the servo amplifier 65 is utilized to drive the servomotor 61 which in turn positions the rotor of control transformer 64 and simultaneously positions the slider arm 63 of a DC. biased potentiometer 62. The elevation potentiometer 62 is shown grounded at only one end and, therefore, in the illustrated embodiment, provides only positive values of elevation. Motor 61 is also used to position a differentiating device '66 is explained in greater detail in aforementioned c-o pending application Ser. No. 38,408 and now US. Patent No. 3,169,727.

In addition to driving slider arm 63 and differentiating device 66, the servomotor 61 drives a second calibrated potentiometer 56. The DC. voltage representing T is directed from potentiometer 52 and applied across the secant function potentiometer 56. Since the position of the slider arm of potentiometer 56 is controlled by the E servomotor 61, this type of arrangement will produce at the slider arm of secant function potentiometer 56 a voltage proportional to T sec E The output voltage from the secant potentiometer 56 is fed as an input to servo amplifier 70 which positions servomotor 71. The latter motor 71 in turn controls the magnitude of a feedback voltage from the potentiometer 72 for input to servo amplifier 70 to produce a null. The output shaft position of the motor 71, for a null condition in servo amplifier 70, is proportional to T sec E The T sec E servomotor 71 also drives the rotor of a diiferential generator 73, while synchro data representing target train T is fed from the tracking radar 22 to the stator Windings of the differential generator 73. The output from the differential generator 73 is therefore T +T sec E which, in accordance with Equation 1 of aforementioned copending application Serial No. 38,408, is the true value of target azimuth A;- in synchro form (see column 9, line 41 of US. Patent No. 3,169,727).

The latter synchro output from the differential generator 73 is then fed to the A servo section 80 wherein the servomotor 81 drives a generator 82 producing an output therefrom proportional to the taget azimuth A In addition to driving generator 82, servomotor 81 is also used to position a differentiating device 83, the shaft position of which is used to position the output slider arm 84 on potentiometer 85 producing an output proportional to :A the rate of change of target azimuth, in much the same manner as differentiating device 66 in E servo section 60 produces an output from potentiometer 67 proportional to iE/ Connected to motor-generator combination 8182 is a DC. tachometer 86 which generates a DC. voltage designated as A which serves as one of the launcher azimuth control signals.

The D D servo 90 and the T servo in the range servo section are employed to compute the time of flight of the missile to intercept. Information in synchro form relating to D D is applied via operational amplifier 91 to servo amplifier 92 which positions servomotor 93. The latter motor 93 in turn controls the magnitude of a feedback voltage from the potentiometer 94 for input to servo amplifier 92 to produce a null. Since the distance to the missile D is zero prior to firing, the output of servo 90 will be D In addition to positioning potentiometer 94, servomotor 93 also is used to position a differentiating device 95, the shaft position of which is used to position the output slider arm 96 of potentiometer 97 producing an output proportional to the rate of change of distance to the target D The quantity D D is supplied to servo amplifier 101 in T servo section 100 as in an input from the operational amplifier 91, which serves as the input to the D -D servo amplifier 92. The other input to the servo ampli fier 101 is obtained by summing D from potentiometer 97 and a voltage proportional to D obtained as a fixed voltage from potentiometer 98, which is part of the computer section 23 shown in FIG. 1. This summing proceduse is carried on by amplifier 99 which applies the result via computing potentiometer 102 to the input to T servo amplifier 101 thereby positioning servomotor 103. The servo amplifier 101 compares its inputs producing the solution The voltage D 3 is proportional to the nominal missile velocity in feet per second. However, since D =0 When the solution is required, the value of T is the time required for intercept if the missile were launched at that instant.

The servomotor 103 positions a pair of potentiometers 104 and 107. With the voltage proportional to E' derived from potentiometer 67 in E servo section 60 applied via amplifier 106 to potentiometer 104 whose arm is positioned by servomotor 103, a voltage proportional to the ballistic lead E T in the elevation plane is obtained. In a like manner, with the voltage proportional to A derived from servo section 80 applied via amplifier 109 to potentiometer 107 whose arm 108 is posi tioned by servomotor 103, a voltage proportional to the ballistic lead A T in the azimuth plane is obtained.

The remainder of the system comprising the launcher order and capture computer, shown in FIG. 6b, serves to perform the actual positioning of the launcher and the guidance transmitter through use of the synchro data obtained from the input data converter section shown in FIG. 6a.

The elevation control servo 110 accepts input voltages proportional to E E T and AE and produces a voltage output in accordance with Equation 4, which output serves to position servomotor 111. The latter motor 111 in turn controls the magnitude of a feedback voltage from the potentiometer 112 for input to servo amplifier 113 to produce a null. The servomotor 111 also drives a 60 c.p.s. generator 114 and a DC. tachometer 115 which provide voltages proportional to the launcher elevation command E and the rate of change of launcher elevation 13;, to launcher input lines 116 and 117, respectively, thereby effecting correct elevational positioning of the launcher 125.

The azimuth control servo 120 accepts input voltages proportional to A T and AA and produces a voltage 9. which serves to position servomotor 122. The magnitude of the target azimuth A derived from generator 82 in target azimuth servo. section 80 is then added to the output AA in dilferential generator 123. Through manipulation of Equation 6 and combination thereof with the above relationship for AA it is seen that the resultant voltage output from dilferential generator is proportional to the launcher azimuth command A This output along with the voltage proportional to the rate of change of launcher azimuth A derived from DC. tachometer 86 in target azimuth servo section 80 is applied to the launcher 125 effecting correct azimuth positioning thereof.

The repeated elevation angle of the launcher E is obtained from servo amplifier 130 operating from a control transformer 131 driven by generators at the launcher 125. The servo amplifier 130 positions the shaft on servomotor 132 proportional to the actual instantaneous launcher elevation and the latter motor 132 in turn serves to position resolvers 137, 138 and 139 in the AE servo section 135. In a like manner, the repeated azimuth angle of the launcher A is obtained from servo amplifier 150 operating from a control transformer 151 driven by generators at the launcher 125. The servo amplifier 150 positions the shaft on servomotor 152 proportional to the actual instantaneous launcher azimuth and the latter motor 152 in turn serves to control a pair of differentials 153 and 154 in the AA servo section 155. The opposite side of differential 153 is driven by a hand controlled dial 156 for representing A the wind azimuth angle. The output of differential 153 \=A A is used to position resolver 140, which receives an input KV on line 157 from manual control system 158 proportional to the predicted missile deflection due to cross wind velocity. The resulting outputs from resolver 140 are KV sin A, the former of which is applied as an input to resolver 137 and the latter of which is applied as an input to servo amplifier 160. Resolver 159 receives a pair of inputs from manual control system 158 which are equal to Y/Dc and X/Dc the parallax value ratios in the X and Y directions. With the shaft of resolver 159 positioned according to the derived value of A the resolver produces a pair of outputs 1/DC(X sin A -t-Y cos A which is applied to the input of resolver 136, and

sin A 'Y COS AG which is applied to the input of servo amplifier 160.

Servo amplifier 160 accepts input voltages equal to 1/DC(X cos A Y sin A which is equal to on cos E according to Equation 13, and KV sin which is equal to 1 cos E according to Equation 10. Since by definition a+oc is equal to AA servo amplifier 160 will initially provide a voltage output to servomotor 161 which is proportional to AA cos E However, resolvers 138 and 139 which have their shafts positioned by servomotor 132 in accordance with the launcher elevation E provide voltages proportional to +cos E and cos E to either side of center tapped potentiometer 156. Servomotor 161 in turn drives potentiometer 156 which applies a voltage equal to icos E to the input to servo amplifier 160 producing a resultant output equal 110 AAGG.

In addition to controlling potentiometer 156, servomotor 161 also controls differential 154 which combines A from servornotor 152 and AA from servomotor 161 according to Equation 6 producing an output proportional to the desired transmitter azimuth A This output from differential 154 positions the shaft on resolver 159 and also positions potentiometer 162, which applies the transmitter control signal A to the transmitter thereby effecting proper azimuth control thereof. Servomotor 161 also positions potentiometer 163 which applies the necessary feedback voltage proportional to AA to the input of launcher azimuth servo system 120.

Resolver 136 in AE servo section 136 receives an input voltage from resolver 159 proportional to sin A +Y COS A and with its rotor positioned according to the instantaneous value of E produces an output equal tothe elevation parallax angle e as defined in Equation 18. The resolver 137 receives an input equal to KV cos A from resolver and has its shaft positioned by servomotor 132 such that the output of resolver 137 is equal to the wind deflection in the elevation plane to as defined in Equation 15. As already described, resolver 138 provides an output voltage proportional to cos E This output is applied to a voltage divider consisting of resistors 144 and whose values are proportioned such that the voltage output at the center tap between the two resistors is equal to .142 cos E which is the gravity drop angle 5 as defined in Equation 14.

The voltages proportion to e, w and 6 derived from resolvers 136, 137 and 138, respectively, are applied as inputs to servo amplifier 141 in AE servo section 135. Since by definition AE =e+w+, an output from servo amplifier 141 proportional to AE positions servomotor 142. Servo motor 142 in turn drives a differential 146 which is also driven from the other end by servomotor 132 in proportion to E The output of the differential 146 according to Equation 3 is equal to the required transmitter elevation E and is used to control the shaft position of resolver 136 and the position of potentiometer 147, which controls the transmitter elevation angle. Servomotor 142 also controls the position of potentiometer 143 in the feedback circuit of servo amplifier 141 and controls the position of potentiometer 148, which provides a feedback voltage proportional to AE to launcher elevation servo 110.

The manual control system 158 provides, through a well-known system of manually cont-rolled potentiometers, voltage values proportional to wind velocity V wind direction A X and Y, the parallax values, and D the capture range.

The novel launcher and capture computer of the instant invention therefore provides complete automatic contol of guidance beam and launcher azimuth and elevation parameters so as to ensure beam capture of the missile at the proper time and position after launch. Although the invention has been described in connection with a specific missile programmed system, it is to be understood that the invention has general application to all beam-rider guidance systems given the proper radar input data.

Obviously, many modifications and variations of the present invention are possible in the light of the above teachings. It is therefore to be understood, that within the scope of the appended claims, the invention may be practiced otherwise than as specifically described.

What is claimed is:

1. Apparatus for effecting initial beam capture of a beam-riding guided missile comprising in combination,

tracking radar means for tracking a target and obtaining target position information,

guidance transmitter means for providing a beam to direct the missile towards said target after capture in accordance with a prescribed trajectory,

a missile launcher located remotely from said guidance transmitter at a known location relative to said guidance transmitter for effecting initial directing of said missile in addition to launching,

means for initially positioning said launcher and said guidance beam transmitter prior to launching the missile in accordance with derived control signals,

capture point computer means, and

means for applying input signals to said capture point computer means representative respectively of said target position information, the actual position of said launcher in elevation and azimuth, the location of said launcher relative to said guidance beam transmitter and various predicted external forces to which said missile will be subjected after launch,

said capture point computer means being programmed to manipulate said input signals in accordance with prescribed control and ballistic equations to derive control signals defining the pre-launch positioning of each of said guidance beam transmitter and launcher necessary to direct said guidance beam and said missile respectively towards a desired capture point on said prescribed trajectory, said capture point computer means being operably connected to apply said derived control signals to said initial positioning means to thereby insure capture of the missile after launch within the guidance beam.

2. The combination specified in claim 1 further including,

first means responsive to the existing positions of said launcher in elevation and azimuth for producing output signals proportional to the actual elevation and azimuth positions of said launcher,

second means connected to said tracking radar means for producing output signals defining the position of said target,

third means controlled in accordance with the predicted external forces to which said missile will be subjected after launch for producing output signals proportional to said external forces, and

fourth means controlled in accordance with the known relative locations of said launcher and said guidance transmitter for producing output signals defining the relative locations, and wherein,

said capture point computer means comprises combining means connected to receive as inputs the respective output signals from said first, second, third and fourth means and being programmed to combine said inputs according to said prescribed equations.

3. The combination specified in claim 2 wherein the predicted external forces to which said missile will be subjected after launch includes the force due to wind and the force due to gravity drop.

4. The combination specified in claim 2 wherein the output signal produced by said fourth means is indicative of the parallax angle that exists between said guidance transmitter and said launcher relative to said desired capture point.

5. The combination specified in claim 1 wherein said launcher positioning means comprises,

means for initially controlling the positioning of said launcher in accordance with said target position information, and

feedback means responsive to the control signals derived by said capture point computer means for correcting the .initial positioning control of said launcher in accordance with said derived control signals.

6. In a beam-riding missile guidance system including a launcher, a guidance transmitter positioned remotely from said launcher, and a target tracking radar producing output synchro information signals relating totarget traverse, target elevation, target train and slant range difference between the target and the missile, a launcher and capture computer for effecting initial positioning of the missile launcher and guidance transmitter so as to insure beam capture of the missile at a desired capture point after launch comprising,

an input data converter section connected to receive as input the output synchro information signals from said tracking radar and adapted to convert each input synchro signal to a proportional output voltage form,

input control circuit means for said launcher connected to receive the output voltages from said input data converter section, and capable of manipulating its inputs in accordance with the relations L T+ T PJ+ GGa and AL=AT+ATTPI+AAGG where s the launcher elevation command,

A is the launcher azimuth command,

E is the elevation of the target,

A is the azimuth of the target,

E5 is the first derivative of E A is the first derivative of A T is the time remaining from present until intercept of the target by the missile,

AE is the ballistic correction term with respect to the guidance transmitter in elevation plane, and

AA is the ballistic correction term with respect to the guidance transmitter in the azimuth plane,

first means responsive to the existing positions of said launcher in elevation and azimuth for deriving voltages proportional to the existing launcher elevation and azimuth positions, second means for deriving output voltages proportional to the predicted exterior ballistics of the missile in both the elevation and azimuth planes, third means responsive to the relative locations of said guidance transmitter and said launcher for deriving output voltages indicative of said relative locations, combining means connected to said first, second and third means for producing the output functions 6 is the ballistic gravity factor,

w is the ballistic wind correction factor in the elevation plane,

6 is the parallax angle between launcher and guidance transmitter in the elevation plane,

0' is the ballistic wind correction factor in the azimuth plane, and

a is the parallax angle between launcher and guidance transmitter in the azimuth plane,

means connected to said combining means, said input control circuit means for said launcher and said guidance transmitter for producing and supplying to said guidance transmitter the positioning control functions EGG=EL-AEGG, and AGG=AL AAGG: where E is the guidance transmitter elevation command, and v A is the guidance transmitter azimuth command,

whereby the guidance beam is transmitted towards said capture point, and

feedback means connected to said combining means and said input control circuit means for. said launcher for feeding back the AE and AA outputs of said combining means to said input control circuit means, hereby positioning said launcher relative to said transmitted guidance beam, so as to insure beam capture of the missile after launch at said capture 7 point.

7. In a beam-riding missile guidance system of the type described, a launcher and capture computer as defined in claim 6 wherein the ballistic gravity factor is determined by the relation 6=-M cos E radians where E is the launcher elevation command, and M is a constant.

8. In a beam-riding missile guidance system of the type described, a launcher and capture computer as defined in claim 6 wherein the ballistic wind correction factors are determined by the relations w=KV cos (A -A sin E and r=KV,,, sin (A -A cos E where V, is the instantaneous wind velocity,

A, is the instantaneous wind azimuth,

A is the launcher azimuth command,

E is the launcher elevation command, and K is a constant.

=SII1 EGGIIXP Sin AG -i-Y COS AGG] /D and 0z=X cos A Y sin A /D cos E where E -is the elevation command for the guidance transmitter,

A is the azimuth command for the guidance transmitter,

X is the distance from the launcher to the guidance transmitter along the X axis of a fixed coordinate system,

Y is the distance from the launcher to the guidance transmitter along the Y axis of a fixed coordinate system,

D is the beam capture range in feet, and

B is the launcher elevation command.

. References Cited by the Examiner UNITED STATES PATENTS 2,671,613 3/1954 Hansen 89-41 3,144,644 8/1964 Getting 8941 3,169,727 2/1965 Schroader et al 244-14 BENJAMIN A. BORCHELT, Primary Examiner.

0 M. F. HUBLER, Assistant Examiner.

Claims (1)

1. APPARATUS FOR EFFECTING INITIAL BEAM CAPTURE OF A BEAM-RIDING GUIDED MISSILE COMPRISING IN COMBINATION, TRACKING RADAR MEANS FOR TRACKING A TARGET AND OBTAINING TARGET POSITION INFORMATION, GUIDANCE TRANSMITTER MEANS FOR PROVIDING A BEAM TO DIRECT THE MISSILE TOWARDS SAID TARGET AFTER CAPTURE IN ACCORDANCE WITH A PRESCRIBED TRAJECTORY, A MISSILE LAUNCHER LOCATED REMOTELY FROM SAID GUIDANCE TRANSMITTER AT A KNOWN LOCATION RELATIVE TO SAID GUIDANCE TRANSMITTER FOR EFFECTING INITIAL DIRECTING OF SAID MISSILE IN ADDITION TO LAUNCHING, MEANS FOR INITIALLY POSITIONING SAID LAUNCHER AND SAID GUIDANCE BEAM TRANSMITTER PRIOR TO LAUNCHING THE MISSILE IN ACCORDANCE WITH DERIVED CONTROL SIGNALS, CAPTURE POINT COMPUTER MEANS, AND MEANS FOR APPLYING INPUT SIGNALS TO SAID CAPTURE POINT COMPUTER MEANS REPRESENTATIVE RESPECTIVELY OF SAID TARGET POSITION INFORMATION, THE ACTUAL POSITION OF SAID LAUNCHER IN ELEVATION AND AZIMUTH, THE LOCATION OF SAID LAUNCHER RELATIVE TO SAID GUIDANCE BEAM TRANSMITTER AND VARIOUS PREDICTED EXTERNAL FORCES TO WHICH SAID MISSILE WILL BE SUBJECTED AFTER LAUNCH, SAID CAPTURE POINT COMPUTER MEANS BEING PROGRAMMED TO MANIPULATE SAID INPUT SIGNALS IN ACCORDANCE WITH PRESCRIBED CONTROL AND BALLISTIC EQUATIONS TO DERIVE CONTROL SIGNALS DEFINING THE PRE-LAUNCH POSITIONING OF EACH OF SAID GUIDANCE BEAM TRANSMITTER AND LAUNCHER NECESSARY TO DIRECT SAID GUIDANCE BEAM AND SAID MISSILE RESPECTIVELY TOWARDS A DESIRED CAPTURE POINT ON SAID PRESCRIBED TRAJECTORY, SAID CAPTURE POINT COMPUTER MEANS BEING OPERABLY CONNECTED TO APPLY SAID DERIVED CONTROL SIGNALS TO SAID INITIAL POSITIONING MEANS TO THEREBY INSURE CAPTURE OF THE MISSILE AFTER LAUNCH WITHIN THE GUIDANCE BEAM.

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