US2923496A - Single control missile guidance - Google Patents

Description

Feb. 2, 1960 J. F. GORDON S'INGLE CONTROL MISSILE GUIDANCE Filed July 25. 1952 5, Sheets-Sheet 1 JAMES F. GORDON INVENTOR.

2 L ATTORNEYS Filed July 25. 1952 GORDON 5 Sheets-Sheet 2 PULSE XTAL V PULSE ME M'SSILE DIVIDER GEN AuTo. PARAToR P'LOT 35 I? 36 ls DECODER 7 T ENCODER n GATE I MISSILE l T /34 /46 IB /|9 RADAR ENCODER PULSE RADAR RcvR m MOD. TRANs.

33 2| ANT. ANT.

CODE m SPACE D LINK Co S 11 cool; I

' 23 I T 47 3| F45 AUTO ,4a I RADAR DECODER RANG DECODER I RADAR TRANS. 1: TRACKER m RCVR 53 9 5e 48 l r W 1140- =-4 ANGLE RANGE 52 COMPUTER DATA CONTROL 29 STATION PULSE 1: DECODER MOD. I

7| 7 42 vlsuAL MANUAL IND. COMPONENT 2s /28 27 CONTROL [26 ENCODER u L PULSE w DELAY AMP H 30 CIRCUIT JAMES F. GORDON INVENTOR.

ATTO

Feb. 2, 1960 J. F. GORDON SINGLE CONTROL MISSILE GUIDANCE 5 Sheets-Sheet 3 Filed July 25, 1952 JAMES F. GORDON INVENTOR.

ATTORN Feb. 2, 1960 J. F. GORDON SINGLE CONTROL MISSILE GUIDANCE 5 Sheets-Sheet 4 Filed July 25, 1952 T L u I I h I I mwzofit mp u 323. a: m n I u n T l I l fi l H 300 fimmaou zov mozwa oz oo F L ll I] E. H 38 u L 2 50 R 3 165200 N A J I Ill! JAMES F. GORDON INVENTOR- ATTOR 9 Y5.

Feb. 2, 1960 Filed July 25. 1952 5 Sheets-Sheet 5 FIG. 5

as I COMPARATOR DEACTUATOR AUTO. PILOT I MISSILE I 34 9| I DECODER RCVR. m l

SPACE LINK TRANS CONTROL COMPARATOR ENCJgIDER STATION J JAMES F. GORDON INVENTOR.

2,923,496 SINGLE coNrnoi. MISSILE GUIDANCE James F. Gordon, Baltimore, Md., assignor to Bendix Aviation Corporation, Towson, Md., a corporation of Delaware Application July 25, 1952, Serial No. 300,868

'5 Claims. (Cl. 244-"-14) This invention relates generally to a method and means for guiding a moving vehicle and more particularly to the problem of mid-course guidance from a control station for a long range vehicle such as a missile or the like.

The problem of long range guidance of the flight of a missile from the point at which it is launched to its ultimate destination at a target is usually separated into three distinct phases when the distance to be traveled by the missile is of such great magntiude that no single location adjacent the course line can be in control of the missile throughout the complete flight. The phases for the flight of such a missile are usually designated as the launching phase, the mid-course guidance phase and the terminal phase. With this general plan there is provided in the launching phase some means for getting the missile into the air and started on a flight path in the general direction of the target. At some advance position intermediate the point of launching and the ultimate target there is setup a control station, the purpose of which is to intercept control of the missile as soon as it appears on the radio horizon from the launching site and control its flight past the controlstation to an ultimate release point which is generally on an opposite horizon. The function of the control station is completed once it has maneuvered the missile to the predetermined release point, whereupon the control of the flight of the missile is released to a homing guidance system in the missile itself or such other control as may be provided for terminal guidance in the final phase of the missile flight. While many arrangements have been provided heretofore for mid-course guidance, they have been relatively complex and have been limited in their application by a lack of flexibility for the various situations which are encountered. For example, many of the prior art systems have required that certain information with regard to the course to be flown be set in the missile before launching and, therefore, these settings may not be altered during the flight of the missile to meet any changing conditions or other unpredictable events which may occur.

It is the object of the present invention to provide method and means for guiding a missile in azimuth which are simple and reliable and readily adaptable to changing conditions encountered in the field.

A further object is the provision of missile guidance arrangements in which pre-set data, which is not subject to alteration during flight to compensate for unexpected irregularities, is not required to be supplied to the missile.

Another object is to provide mid-course guidance in which the guidance path is placed to agree With the position of the missile at the start of the guidance and the flight path from that point to the target is established as a straight line.

A still further object is the provision of such a system in which the flight path may be reestablished for a new missile position in the event that guidance control is lost momentarily.

Another object is the provision of a simple arrangement for comparing the actual missile position with the 2,923,495 Patented Feb. 2, 1960 desired position therefor and manual means for supplementing the guidance information, if desired.

These and other objects of the invention will be'rnore readily understood from the'following detailed description, taken in conjunction with the accompanying drawings, wherein:

Fig. 1 is a plan view of an azimuth control arrange ment;

Fig. 2 is a block diagram of a complete system for accomplishing azimuth guidance in accordance with the invention;

Fig. 3 is a block diagram functionally indicating the operation of the computer of the present invention;

Fig. 4 is a pulse timing diagram useful in explaining the operation of the invention;

Fig. 5 is a view showing the guidancerelationships in three dimension; and,

Fig. 6 is a partial block diagram showing a modification.

Referring now to Fig. 1, the azimuth control situation is shown wherein a control station C is located a fixed distance r fromarelease point R, which is in the vicinity of the target. In any tactical situation the" distance CR and the direction thereof, insome reference systems, are

known. The control station C is capable of measuring the range and bearing of a missile M continuously. The distance from C to M, r and the angle 9 between r and r are determined at the control station C'for the position of the missile M at some point, for example, Where it is first detected with suitable signal strength after being launched. Upon the determination of the quanti ties 1- and 6 fora'given point in space, the control station C computes the path MR by solution of the triangle CMR. The path MR, as thuscomputed, is a straight line and coincides at onefend thereof with the actual position of the missile M and the other end of the path is the release point R. The information thereafter required from the' controlstation C is merely that which will maintain the missile at the range r of points P (r,0') on the line MR for all subsequent values of 0.

The required relationship of the range 1' to the line MR for any angle 0 may be obtained from the equation of the iine MR in polar coordinates with C as the origin and CR the baseline for angular measurement. This equation of the line MR is rr; sin 0r r sin 0 +rr sin (0 0)'=0 from which the range 1'- as a function of 0 is T17'g Sill 01 -n sin 6+r sin (0 -0) With the relation (1) future values of r may be computed for any 0 after the missile has reached a point M for which r and 0 can be established. If suflicient time is available during the mid-course flight the range may be computed manually for a number of points along the desired course MR and corrections applied to the missile as required upon comparing the computed and actual ranges for each angle. For example, if for the angle 0 the computed range is r and the measured range r' at this angle discloses a missile position M a command must be given causing the missile to turn toward the control station C and decrease the measured range. In general, manual computation and control will not provide sufliciently rapid and continuous data for a high speed missile. To overcome this deficiency continuous and automatic computation and control may be employed.

In Fig. 2 a system for automatic computation and control in accordance with the present invention is disclosed. In themissile any desired form of propelling and steering mechanisms are provided. For controlling and correcta a j 2,928,498

ing the course upon which the missile flies there is provided therein a precise time reference, such as a crystal oscillator '11. The. frequency of the oscillatorll is applied to a divider 12 for producing a submultiple frequency having a period sufiiciently along to accommodate the control cycle for a particular application, as herein 'explained. The output of the divider'12 is applied to a pulse generator 13 which produces accurately I spaced timing pulses 14 at a predetermined point inthe cycle of the waves from the divider 12. The pulses 14 are cncoded by an encoder 15 which characterizes the signals produced thereby in response to the pulses v14 in accord ance with a predetermined code I. It will be understood that the encoding and decoding operations described herein may be of a conventional type and may include such characteristics, for example, as pulse width, pulse number, pulse spacing or the like, and signals so encoded only operate the channels which are equipped with a corresponding decoder. The encoder 15 produces code'I signals'16 which are selectively passed by a gating circuit 17 to a pulse modulator 18 which characterizes the transmission of a radar transmitter 19 from an antenna 21 in accordance therewith.

At the control station C transmission from the missile antenna 21 is received by an antenna 22 and detected and amplified to a useable level by a radar receiver 23. Code I signals from the receiver 23 are detected in a decoder 24 and passed to a pulse amplifier 25 from which they are applied to a controlled delay circuit 26. The delay circuit 26 provides a selectively controlled delay time between the input and output pulses thereof in accordance with a control signal applied thereto, as will be presently explained. The delayed code I signal 27 from the delay circuit 26 is applied via a switch 30 to an en'- coder 28, wherein the signal is characterized by code II. The characteraof the code II signals from encoder 28 is transmitted to the missile by means of a pulse modulator 29, radar transmitter 31 and'radar transmitting antenna 32.

The code II signal from the control station C is received I at the missile by a receiving antenna 33 and a radar receiver 34. Signals from the receiver '34 are applied to a decoder 35 which passes code'II signals 36 to the output circuit thereof. Code II signals 36 from the decoder 35 are applied to a pulse time comparator 37 under the control of gating circuits 17. The pulse time comparator 37 also receives the timing pulse signals 14 from the generator 13 and provides an output control signal, the amplitude and polarity of which represent the relative time positions of the two input pulse signals 14 and 36. This control signal from the comparator 37 V is applied to an automatic pilot 38'-which steers the missile in azimuth in response thereto. As will be more fully explained with reference to Fig. 4, the automatic pilot 38 provides'left and right steering forces to the missile in response to opposite polarity control signals from the comparator 37 and maintains zero turning forces in response to zero magnitude control'signals. V

Inorder that the guidance commands supplied from the coincidence type comparator just described to'the automatic pilot 38 in the missile shall call 'for the desired course tothe release point R the delay circuit 26 must introduce a time delay having a value which,-when added to the round trip echo time for the instantaneous range r, is equal to the period between the timing pulses 14 in the missile. The delay introduced by the circuit 26 is, there- -fore, controlled by' a .computer'41 and selectively by' supplementary control signals from a manual component device 42." The computer 41 is supplied with range data from an automatic range tracker 43,'which compares the'tirne interval'between the transmission of code II r 64. The value of r from the range keeper 54-is also 7 signals'fr'orn the antenna 32, byrmeans' of a detector 44 and a decoder 45, with the return from the missile of a code 111 signal. Code 111 signals in the missile originate in an encoder 46 in response to the code 11 signals 36 provided an accurate measure of the missile range 1'.

To obtain the azimuth angle position of the missile, a" tracking antenna is employed and for this purpose one of the control station antennas 22, 32 may be of the type which automatically follows the angular position of a particular target. 'With such an antenna an angle data take-off device 48 coupled thereto supplies output data representing the angle 0. Angle data from the device 48 and range data from the tracker 43 are supplied to the" 'computerfiil. w a 7 a 'The operation of the computer as, employed in the present invention will now be described with' reference to Fig.- 3.- It will be understood that the range 1' and the azimuth angle 6 which are supplied to the computer 41 may be in the form of a mechanical motion or an electricalfquantity or both. Accordingly, the mathematical operations indicated in Fig. 3 may be performed as mechanical motions or transformations of electrical Range data are applied in the computer 41 to an input coupling 5l which supplies a visual rangeindicator 52 and, through a switch 53, goes to a range keeping device 54 which reproduces the range datainput as long as switch 53 remains closed. A manually operated range data device 55 is provided with a manual control 60. The device 55 is capable of receiving set values corresponding to the distance ri which include the distances CR 'normally encountered. The outputs of the range keepers 54,255 are applied to a multiplier '56 which produces as an output the product r r 'The azimuth angle data 6 are applied to an inputcoupling 57, from which they are supplied through a switch 58 to'a sine function generator 59. The sine generator 59 provides in its output the 'sine of theangle 0 icontinuously for closed positions of'the switch 58 and maintains the last established value'upon the opening of switch 58. The opening of switch 58' establishes the angular constant 6 Upon reclosing the switch 58 new values are established according to the present value'of '0 at the'coupling 57. The sine output of generator 59 and a the product r r from the multiplier 56 are applied to a multiplier'61 to provide the output product thereof r r Sln01.'- a I V- The angular data 0 are also applied to a subtraction device 62 as is the selected'value thereof 0 from the right hand side of the switch 58. The difference 0 -0 from the subtractor '62 is applied to a sinegenerator 63 and the sine of the difference angle is applied to 'a multiplier applied to the multiplier 64 and the output product thereof is r sin (O -49) Another sine function generator 65 operates on the angle data 0 and supplies a multiplier 66 which also is supplied with the constant value r from the range set device 55. The output of the multiplier 66,

therefore, is r sin 0.

The outputs of the multlpliers 64 and 66 are applied to an adder 67'to produce the sum of these quantities. The outputs of the multiplier 61 and the adder 67 are applied .to a divider 68 as numerator and denominator; respectively, and the quotient produced at output 69 thereof is the functional value for'the range r according to Equaance Control remains unestablished due to tion of switch 30.

tion 1. The computed value'of range r is indicated by a visual indicator 71 in Fig. 2.

The operation of the system of Figs. -2 and 3 in guiding a missile in accordance with the maneuver of Fig. 1 will now be described with reference to the timing diagram of Fig. 4. In the launching phase the flight of the missile is established to come within the range of the-control station t," and the various propelling, navigational and control devices are rendered operative. The crystal oscillator 11 accurately establishes a time reference in the inissile in the form of pulses 14. The pulses 14 interrogate over the code I signals 16 seeking to establish guidance from the control station. At the control station C the tracking radar is alert and Waiting for the missile to come into some suitable range r at the point M Where g'u'idance-is to be established. During this interval range and azimuth angle data are continuously supplied to computer 41 fora period of time before guidance is established. The'manu'al component of the delay signal supplied by the device 32 'is normally set to zero or such other constant value as the unavoidable circuit delays mayrequire. The

control delay circuit 26, therefore, is calling for'the correct delay in the code 1, code II round trip signal period for the instantaneous position of the missile but the guidthe open pdsi- When the missile has reached some suitable position,

such as point M, the operator in the control'station opens switches 53, 58 and closes switch to initiate fguidan'ce along the path MR. Since the missile is at point M, opening switch 53 establishes the value r in the range keeper 54 and the opening of switch 58 establishes the angle 0 to the sine generator 59 and these values remain unchanged for the duration of the maneuver. Switch 30 which is simultaneously closed permits the code II transmission to be completed in response to the delayed code I signal 27. The code II signal reaches the missile after a further delay due to the distance traveled over the space path and the decoded signal 36 will be in time coincidence with the next timing pulse 14 for the on course condition which obta ns, by definition, at the point M. For all subsequent times the position of the missile, such as the position M, will determine whether or not the pulses 14 are in time coincidence with the code II pulses 36 received at the missile. The total time for such coin cidence is determined by the transmission delays over the space path of the code I and code II signals and the delay introduced by the circuit 25. The total of these delays would produce coincidence if the missile range were r for the angle 0. If the missile range is other than r, say r, the total delay as sensed by the comparator 37 will be other than that for which coincidence obtains and a corresponding steering correction signal of the correct polarity will be applied to the automatic pilot 38 to turn the missile toward the path MR. 7

The code III signals 72 returned from the missile to the control station provide on the range indicator 52 a visual indication of the actual range 1' to the missile which may be used for compmison with the desired range r as read from the indicator 71. If required, manual over-ride of the automatic steering control may be introduced by means of the device 42. If at any point, "as a result of uncontrollable conditions, the missile position M is so far from the course 'MR that the reacquisitiono f the course would be difficult or impossible, a new course may be readily established by reclosing switches 53, S8 and opening switch 30 until the system stabilizes upon the missiles present position. A new course to the release point R may then be established from the present missile position by reinitiating guidance as herein described with a new value of 1- and 0 for the new missile position.

In order that the operation of the circuit may occur as above described, the gate circuit '17 in the missile supplies a gating wave 73 which prevents transmitter 19 from transmitting on alternate ones of the pulses 14 and likewise prevents the comparator 37 from utilizing any sig- 'nals which occur during the transmission periods for the transmitter 19. p The wave 73 will normallyembrace the period on each sideof a pulse 14 equal to the maximum change in controlled delay from the circuit 26 that is expected to occur. The characteristics of the comparator 37 preferably include'a memory for supplying continuous outputdata to the pilot 38.during theintervals between thegates 73 when no data are received.

The system-as hereinbefore described provides midcourse guidance control which is adequate for many applications. In some situations, however, it' may be desirable to practice the invention with compensation for the altitude at whichthe missile travels. The three dimensional situation-is depicted in Fig. 5,-wherein the azimuth projection in the XY plane corresponds to Fig. 1. For this arrangement it is apparent that the slant range CMs is greater than r and that-r is equal to the productCMs cos (p Likewise the azimuth range r for any point P (r, 6) is obtained by multiplying the slant range to the point P (r, 0, 11 by "cos 4;. The maneuverof Fig. 1, therefore may be performed irrespective of the initial altitude of the missile by correcting the slant range measured by the system by a factor of cos This operation is shown in Fig.3 wherein a multiplier 75 is supplied with range data from the rangeinput 51 and the value of cos from a cosine generator 76. The generator 76'is supplied with elevational angle data from the tracking antennas 22, 32. For this purpose these antennas may be of the type providing a thin pencil shaped beam with automatic trackingin both azimuth and elevation. The output of the multiplier 75 is available at a terminal 77 and may be selectivelysupplied to the subsequent portions of the computer by means of a switch 78. Such selection interrupts the transmission of slant range data from the terminal 51 to the units 52, 54. It will be understood that the data supplied to the units 52, 54 for either position "of the switch 78 is the range data from which the computation is made and for situations in wh ch the difference between slant-range and the azimuth projection thereof becomes large, the connection to the terminal 77 is preferred.

output of the generator 98 is applied to a multiplier 99.

The other factor for the multiplier 99 is the output ofthe divider 68, obtained selectively by switch means 101. The output of the multiplier 99 alternatively supplies the output terminal 69 via a switch 102 when the divider 68 is disconnected therefrom by the switch 101.- The output of the multiplier 99 is the product of the computed on course range r and the secant of the angle :73. Therefore, the range data from the multiplier 99 supplied to the unit 26 of Fig. 2 will always be greater than the value of r computed by (1) except when the angle is .zero. The inverse range input-delay output function of the unit 26 providesa corresponding decrease in delay introduced intothe code I, code II signal round trip period thereby placing the on course signal at the point P (r, 0, n5) for a computed point P (r, 0). The missile, therefore, is guided to fly thespace path MsRs upon establishing the course MR thereof and in accordance with any desired altitude requirements established .by well known means, not shown.

In Fig. 6 one arrangement suitable for releasing control'of the vehicle upon its arrival at the release point R is indicated. The units now first described are shown as they would be applied to the system of Figs. 2 and '3. At the control station, data of the set range r from the unit '55 and 'the range from the range tracker 43 is supplied to a comparator 93. When the measured range 'is equal tor the comparator actuates an enc'oder'94which provides a code IV signal transmission from the trans- 'disable the generation of a code IV signal for a period until the measured range is less than 13.

It will be understood that small time delays are in evitable ina system of this type and in the calibration of the system will be accounted for. As an example, the

range indicator 52 actually measures the timeiinterval 2r+d between the pulse 74 at the start of code II and the pulse 75 at the end of code III. a The value indicated, however, is merely the actual range with the value d and the factor of two suitably calibrated' out of the indication. c

The'pr'esent invention is well adapted to automatic operation at the control station as well as inthe-vehicle. For example, for wartirneuse automatic control stations may be planted in enemy territory on land or in water and arranged to operate after a predetermined period for guiding missiles or occupied aircraft to-a release point established at the time the station is planted. These stations would then operate for some useful period while an attack was carried out with the aid of guidance signals therefrom.

Many equivalents of the present system will now be apparent for practicing the invention as here des ri ed and other embodiments may be devised for practi ing the method thereof and are to be understood as within the scope hereof. I 7

What is claimed is: r

1. A control system for guiding a radio controlled vehicle from a first point in space to a second point. said system fun tioning on signa s passin between said vehicle and a third point, where said first point is the location of said veh cle when said method of uidin commen es and where said third point is not colinear with said first and second points, comprising: means for producing. in said vehicle, a reference si nal; means for transmitting said signabfrom said vehicle: means at said third point f r receiving said transmitted signal; means for delavin said received signal; means fortransmitting said delaved signal; means in said vehicle for receivingtsaid second transmitted signal; meansfor retransmitting said second sig- 7 point for receiving said retransmitted second signal; means atsaid third point for determining range of said vehicle by comparing-said retransmitted second signal with said second transmitted signal; means at said third point for measuring the successive angles between said vehicle and said second point; means at said third point functioning on data cons'isting of the location of said second point, the initial range of saidranges, the initial angle of said angles, and said successive angles of said angles to pro duce an output to vary said delay means; and'rneans comparing said second transmitted signal received at said vehicle with said reference signal; said comparing means producingan effect tending to maintain said vehicle on said course.

3. A control system for guiding a radio controlled vehicle over a'course from a first point to a second point, said system functioning on signals passing between said vehicle and a third point, where said first point is the location of said vehicle when said guiding commences and where 'said third point is not colinear with said first and second points, comprisingzvmeans in said vehicle for producing a periodic time reference signal; a round trip communicationdata link between said vehicle and said third point for signals synchronized with said time reference signal; range and direction finder means for determining the range and bearing of said vehicle from said third pointymeans responsive to the range and bearing of said first point and predetermineddata of said second point to determine said course; computer means utilizing data of said course and the successive bearings of said vehicle to provide outputs representative of the respective ranges of said course from said third point; delay means I responsive to said outputs to introduce delay in said communicationlink representative of said ranges; and means in said vehicle comparing said communication link transmission interval, including said delay, with said reference signal, said comparing means producing an effect tending to maintain said vehicle on said course.

4. A control system for guiding a radio controlled vehicle over a course from a first point to a second point, said system functioning on signals passing between said vehicle and a third point, where said first point is the location of said vehicle when said guiding commences and'where said thirdpoint is not colinear with said'first and second points, comprising: means for producing a periodic time reference signal; a communication data link between said vehicle and said third point for signals syn ,7 chronized with said timereference signal; range and direcnal; means at said third point for receiving said retransmitted second signal; means at said third point. for determining range of said vehicle bv comparing said retransmitted second signal with said second transmitted V signal; means atsaid third point for measuring the suc- V cessive angles between said vehi le and said second point; means at said third point function ng ondataconsisting of the location of said second point. the initial range of said ranges, the initial angle of said angles, and said successive angles of said angles to produce ,an' output to vary said delay means; 'and means comparing said I second transmitted signal received at said vehicle with said reference signal; said comparing means producing an effect tending to maintain said vehicle on said course.

2. A control system for guiding a radio controlled ve- V hicle from a'first point in space to a second point, said system functioning on signals passing between'said vehicle and a third point, where said third point'is not colinear with said first and second points, comprising: means for producing, in said vehicle, a reference signal; means for transmitting said signal from said vehicle; means at said third pointwfor receiving said transmitted signal; means for delaying said received signal; means for transmitting said delayed signal; means in said vehicle for receiving said'second transmitted signal;'me'ans for retransmitting said second signal; means at said third tion finder means for determining the range and bearing of said vehicle from said third point; "means responsive to the range and bearing of said first point and predetermined data of said second point to determine said course;

computer means utilizing data of said course and the suc-' cessive bearings of said vehicle to provide outputs representative of the respective ranges of said course from .said third point; delay means responsive to said outputs' to introduce delay in said communication link representative of said ranges; and means comparing said communication link transmission interval, including said delay, with said reference signal, said comparing means producing an effect tendingto maintain saidvehicle on said course. ,i V e g 5. A control system for guiding a radioicontrolled vehicle over a course from a first point to a second point,

said system functioning on signals passing between said vehicle and a third point, Where 'said first point is the location of said vehicle wliensaid guiding commences and 'where said third point is not colinear with said first and second points, comprising: means for producing a periodic time reference signal; a communication data link between said vehicle and said third point for signals synchronized with said time reference signal; direction finder means for determining the bearing of said' vehicle from said third point; means for setting said course relative to References Cited in the file of this patent UNITED STATES PATENTS Ewing Jan. 30, Herbst Jan. 30, Watts Feb. 6, Omberg et al. Feb. 13, Fennessy et al Jan. 15, Haller Apr. 29, Getting May 31, Sunstein Dec. 20,

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