US3164339A - Missile programmer coast mode provision - Google Patents

Description

Jan. 5, 1965 l. H. scHRoADl-:R ETAL 3,164,339

MISSILE PROGRAMMER COAST MODE PROVISION Filed Jan. 31, 1961 4 sheets-sheet 1 l my,

ESE .E QE EE gaat SESNY Bt f M m Rham r/f I. om S Jan. 5, 1965 l. H. scHRoADER ETAL MssILE PROGRAMMER ooAsT MODE PROVISION Filed Jan. 31, 1961 4 Sheets-Sheet 2 Jan. 5, 1965 |..H. scHRoADr-:R ETAL 3,164,339

MIssILE PROGRAMMER coAsT Mom: PROVISION 4 Sheets-Shea?l 3 Filed Jan. 5l, 1961 INVENTORS /RV//V H. SCHROADER E. HosEA FREDER/CK E H/LTZ C7V. (O. BY M M4414 JMW ATTORNEYS Jan. 5, 1965 Filed Jan. 51, 1961 Type B Mgcourse Proqram Rega/'n track l. H. SCHRQADER ETAL 3,164,339

MISSILE PROGRAMMER COAST MODE PROVISION 4 Sheets-Sheet 4 Type A M/'dcourse Program C/Lass of rack Rega/'n frac/r fea-5,;

LEO C. M/LLE'R FREDERICK EHI/ TZ ATTORNEYS l United States Patent() 3,164,339 MISSILE PROG yl R COAST MDE PRVESIQN v lirvin H. Schroeder, Simpsonville, Melvin E. Hosea and Leo C. Miller, Silver Spring, and Frederick RHiltz, Kensington, Md., assignors to the-United Statesoi America as represented by thevdeeretaryjo tiieNavy Fiied Jan. 3l, 1961, Ser. No. 86,267

1d Claims. (Cl. 264-14) The present invention relates generally toimprovements in missile guidance systems and the like'and more particularly to a new and improved missile guidance system for beam-riding missiles which enables a programmed missile to coast in anapproximate ballistic trajectory to adesignated target during periods of temporary'loss of target track and to subsequently return to a standard program trajectory when target track is reacquired. l

This application presents an improvement in the missile j targets. Although such devices have,` generally served their purpose, they have not proved entirely satisfactory under all conditions of servicel and operation. In kthis regard, considerable diiculty has been experienced in minimizing the deleterious eliects of noisy radar input data upon both the computer and resulting programmed missile trajectory, with consequent eects upon missile `fuel consumption, missile range, and probability of successful target kill. A further and perhaps more critical problem confronting designers of'missile programmershas been the loss of control or Wild running of beam-riding missiles during even momentary losses of target track by ground radar and associated guidance transmitter equipment. In such instances, even after reacquisition of target track, beam-riders so lost cannot readily be recaptured by the guidance beam.v Furthermore, even in those few instances where` recapture of the stray is accomplished, the missile generally cannot be caused to return to a standard program trajectory t'o the target since the latter would usualiyrequire a hard maneuver: on the missile far beyond its design capabilities.A f j f The general purpose of this invention, therefore, is to provide a guided missiie programmer which embraces substantially ail ofi the advantages of previously employed missile guidance systems andyetl possesses none of vthe aforcdescribed disadvantages. To attain, the latter, the instant invention contemplates provision of'ak missile programming system capableof 'guiding beamLriding missiles along prescribed trajectories most suited to considerations oi missile performance characteristics and target' threat. The latter programming system performs in a manner insuring maximum eiiiciency through the smoothing of input radar data in the computing stages of the programmer itself; and simultaneously minimizesv the'. probability of loss-of a-programmed missile, with consequent failure of target kill, due to intermittent loss of target track during guidance system of the instant invention contemplates `the provision of a novel second order programmer computing section of the type disclosed kin "cope'nding application 'Serial No. 38,408, filed lune 23, 12o-(Liter Multiple Flight Course Second Order Programmer, by IrvinH. Schwader, Melvin E.^Hosea and Leo C. Miller. Inr'the instant in- F t ice vention, the second order smoothing Vcircuitry set forth inthe latter application is modified to include a novel coast. mode provision, the latter*l provision enabling a linear extrapolation for the programtrajectory existing at the instant the lcoastrnode is initiated. Means are also contemplated within the scope of the invention to enable the programmer to revert from coast programming mode to its normal programming mode, upon reacquisition of radarl target track, and toV correct any trajectory error developed during Vthe coast period withina time interval determined` by the amount of computer damping then in effect. e

Accordingly, one object of the present invention is the provision ofra new andfimproved missile guidance system.

Another object is to vprovide a new and improvedA missile guidance systemcapable of preventing loss of a programmed missile during periods of target track'loss by guidance equipment.

A further object of the instant invention is the 'provision of a new and improved missile guidance system enabling programmed missiles toE coast during periods of target track loss by guidance equipment and thereafter to return to a normal program dignt mode.

Still another object resides in the provision of a new and improved missile programmer` which minimizes missile t fuel' consumption, increases kmissile range, and insures maximum probability of target ln'll through themaintev FlG. 2*'illustrate's a; second order missile'programmer. computing circuit without 'the coast ,mode provisionl of the instant invention; Y 1 j `FIG.13 illustrates a'novel second order'missile programmer computing circuit embodying coastV mode provisions in accordance withthe instant invention;

FIGS. 4 and 5 are illustrative .ofthe mannerin which i the coast mode lprovision of; the instanttinvention operates in tWo instances of standard programs obtainable with the second order missile programmers setfforth in FIGS. 2 Varid3; Y r f FiG. 6 is a graph illustrating theimanner in which computer guidance .sensitivity may befvaried vin the second order missile programmers disc1osed.p y Y T he second order missile programmer shown in FIG.Y 2 of the drawingsv is of l'thetype shown andrdes/crib'edin copending application SerialENo. 38,408,; iledlune 23,

1960,'for Multiple FlightCourse Second Order Programmer, by'lrvin'H. SchroaderetV al'.v -`Inthelatter applicas ti'on,;a guidance computertwas disclosedwhich incorporated asecond orderV smoothingcircuit havingvariable gain sensitivity tothe end result that edects ofnoisyj input -second order smoother. Y. Second order missile programmers of the typeA disclosed Vin they aforementioned application Serial No, 38,408' con'-- Y, templatethe manipulation ofvarious input datato they Y vcomputer `inaccordance with generalized'systemsjequa--i tions, theconstantslof which are chosen-inacordance with vmissile cl'iaracteristicsY and target considerations-to bringlthe missile` tOtstarget in the prescribed manner.

The second `order computing system accomplishes the latter by the same functional manipulations performed upon any and all input data irrespective of whatever input parameters such data may represent. Therefore, such programmers enable a great variety of liight programs to be obtained for any given missile merely by varying the nature of the input data and constants for the generalized system equation manipulations in the programming computer section.

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 adaptable to second order smoothing and coast techniques.

FIGURE 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 reected 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 infomation relating to rough values of target range and azimuth and an even rougher value of target elevation. 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.

The target azimuth, target range and target elevation information from the search radar 2t) and height finder 21 enables the operator of the tracking radar 22 to either actually lock on the target 12 or to throw the tracking radar into searc operation, which is essentially a scan raster of plus or minus a discrete number yof degrees about a position which is believed by the operator of the tracking radar 22 to be the approximate target position as indicated by the rough information received from the search radar system. Other modes of radar systems for feeding :input data to the tracking radar 22 might include such systems as the Air Force SAGE type concentrated network of defense radars.

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, target train, target traverse, and slant range from the tracking radar 22 to the target 12. 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 nformation is converted by -means of servomechanism devices to electrical signals in the form of D.C. voltages proportional to the original radar shaft rotation from which the input synchro signals were derived. Such D.C. voltage signal form is required for use by the programmer section 25 of the computer 23. The program- 'mer section 25 utilizes the latter D.C. 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, andproduces D C. output voltages which are then fed to an output converter servo section 26.

l The output converter servo section 26 of computer 23 recouverts the D.C. output voltage signals from the programmer section 2S 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 tracking and search radars 22 and 20, respectively, or at least Within a few hundred feet of the latter radars 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 transmitter 27 and traclcing 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 a reflection yof an original signal 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 b-y the guidance beam 17 from the guidance transmitter 27 to generate a beacon signal 18 of its own. The latter beacon signal 1S is directed from the missile 19 to the guidance transmitter 27, the time of delay yof arrival of the beacon signal 18 at the guidance transmitter 27 being a measure of the slant range to the missile.

Missiles utilizing multiple ight course programmers of the type shown in FIG. 1 must, of necessity, be beamriders. Such a missile is captured by the radar beam 17 emanating from 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 ight 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 missiles capabilities of linear velocity and lateral acceleration are not exceeded.

In actual operation utilizing multiple iiight course programmers of the type disclosed, prior to launching, the tactical operator elects in accordance with the nature of the missile to be fired and the target to be intercepted, one of the standard program trajectories within vthe scope of the programmer. The operator then proceeds to manipulate various switches in the guidance transmitter, e.g., for tracking in missile range or not, depending upon the nature of the program trajectory chosen, and simultaneously operate switches in the programmer section 25 of the computer 23 to select proper constants and input data channels for the desired program. All of the latter operations for multiple iiight course second order programmers are set forth in the aforementioned cop/ending application, Serial No. 38,408.

After the missile is initially red, a discrete period of time elapses before capture of the missile 19 by the guidance beam 13 is accomplished. Thereafter, EG, the elevation of the guidance transmitter, must approach ET, the elevation of the target, in a manner determined'by the prescribed programming equations set up in the programmer section 25 of computer 23. FIGS. 4 and 5 of the drawings illustrate the form of ECT-ET found desirable .in instances where twostandard programs, arbitrarily designated type A and type B, are utilized. It will be noted from FIGS. 4 and 5 that, whereas the type A program illustrated causes the missile to dive on the target in a nearly linear collision course, the type B program is of a nature which causes the missile to approach asymtotically to the line of sight from the tracking radar to the target prior to collision. j

The type A program is the more accurate of the two types of flight programs iillustrated as being obtainable with the Multiple Flight Course Second OrderProgrammer. Although the generalized system equations for both the type A and type B programs are similar, the type A program trajectory is varied from that of the type B program by varying the nature Yof the parameters which are fed into the computer section as input data.

The type A program with its direct intercept collision course is generally chosen for multiple target and wherever the potential threat is very great, as where the target may be carrying a thermonuclear device. On the other hand, the type B programtrajectory is usually f eo t: G eodt l 1) where G and P denote the times of guidance initiation and the present, respectively, and eo represents the generalized system equation, the variation of the constants and'input data of which enables the wide variety of programs obtainable with such second order programmers. The'second order system equation is expressed in terms of 61` the first derivative of e0, rather than conventional eo formv for sake of simplicity as will be subsequently made evident. Thus, the behavior of the second order Vequation Vcomputing circuit set forth' in the aforementioned cog pending application Serial No. 38,408 is uniquely described in terms of the following generalized system equation: e' Y t=P Y Q e0= 2w(e-ei) 0]; 6 w2(e" ei)dt+ constant',

where eo is the output data from the computer, ei is the input data to the computer, e'0 is the first derivative of e0, Y

j circuitry unmodified and as modified by the coast mode wn is the gain sensitivity factor `of the computer, and I a is a constant.

An important feature of lthe circuitry utilized inthe multiple flight second order programmer for computing the various program trajectories in accordance with the v above equations, is the manner in Whichfthe value of wn, the gain sensitivity factor,vis varied'. A low value or wn allows eo to change only yslowly and to be' relatively insensitive to changes in ei, the input data. On the other hand, a high value of wn allows e0 to change rapidlyand to be more sensitive to fluctuations in ep Therefore, the manner in which wn, the gain sensitivity factor, is programmed is a major aspect'of second order programming computer operation. t

It has been determined empirically that adequate -smoothing of noisy input data and desired flight trajectory accuracy will be realized vif 'the gain sensitivity factor,

'wn is programmed in accordance with the equation wn TPH-0 j where TF1 is the time remaining until interceptV oftlie target by the missile and vc isla constant. pirical programming equation for wn insures values-"of Such an er'ni gain sensitivitylying between zero and unity and eliminates the situation where a gain sensitivity of infinity i would be required, the later case being one which would obtain if the gain sensitivity constant c were omitted` from Equation 3. The constant c in Equation 3 is alsov related to the constant oin Equation 2 by the relation any desired form for presently existing or subsequently developed future missiles and, therefore, the variability of cvalues provides for a wide range of adaptability; By way of example, a value of c=4 is chosen to illustrate the principles of operation of ythe programming computer in conjunction with the type A and type B'pro` grammed Aflight trajectories depicted in FIGS. 4 and 5 of the drawings. Setting 6:4, the value of wn'in' Equation 3 is expressed by 4 1 'wp- TPr-i lY and according to Equation 4, l

Thus, Equation 2 for the generalized system equation ecomes f n .1t=P,v-f, H

e,= aant-WG asada .y (t) where laref-ei. H i l The value of the constant c is varied in Vaccordance withthetype of trajectory to `be programmedrand practical embodiments of the' Multiple FlightfCourse Second Order Programmer inc quantity as desired.

Referring now to FIG, 2 of the drawings, there is illustrated a novel secondv order computing circuit for a missile programmer substantially as disclosed, in copending application Serial No.A 38,408 filed I une 23, 1960 by lrvin H. 1 Schwader, Melvin E. Hosea and Leo C. Miller. rlatter circuitry will be described in detail here in order to make more readily apparent the distinctions between such improvement of Vthe instant invention.

FIG. 2 shows a second order computingor-srnoothing circuit comprising afseries of four amplifiersindicated as Y291, 2,02,V 2li?, andZtM, respectivelyg Amplifiers Zill'and 202V aref conventional DC. Vanalog type summing'amplifiersjwcllkno'wn 'inthe art and commercially.-available; v

Amplifiers 2l3land 264 areconventional integrating Varnplifiers. f The input to amplifier 2% ispo, thenfirst derivative-of en, vthe output of "amplifierZtM being -eorThe i latter quantity is Ydirected to the output converter r'servo section 256 prior. to transmission to the guidance transm-iter 27 and is'ralso simultaneously'fed back as ari-"addi-V Vtionalinput to amplifier Ztifwhich receives the'inputda'ta eijthe nature of which `is determined by the desired flight Y `trajectory program. `'The input to amplifier Zttlis-thu's ei-,eo andfthev output, in View of the '180 phase shift j which takes place'in .the amplifier Zfilpwithresultant Areversal insign, isI therefore ye' ,-efor 5.1 For purposes A of solving Equationlthe quantities' aneand w' must, y be obtained. The ,latter is 'accroinplishedV asxd'e'scribed f below.

The output of TheV specific value of the constant c in Equation 3 for orporate means for varying? this The amplifier 2631, Whic'hwis ais"A directed to .v

Y 'lapotentiometer 2h55, the shaftiposition of which is equal j l inf valueto wn, the gain Asensitivity factor. Hence,-sin'ce the. input to potentiometer 205 is a voltage corresponding r`toytlie'value ofv e, ,the outputtherefrom is equal jto`- wnmultiplied by or 'whe The'latter' quantity is fed asan input to amplifier 202 and is also simultaneously directed to a second potentiometer 206 whose shaft position is likewise adjusted to the value of wn, the output from potentiometer 206, therefore, being onze. The latter quantity is, in turn, directed as an input to integrating amplifier 203. The manner in which the output voltage proportional to wn is obtained to control the shaft positions of potentiometers 205 and 206, in accordance with Equation 7, forms no basis of the instant invention and, hence, is referred to only generally in FIG. 2 as the wn circuitry section 47. A suitable embodiment of the latter wn circuitry section 47 is adequately set forth in the aforementioned copending application Serial No. 38,408. For purposes of the instant discussion, it will sufiice to say that wn is limited to values between zero and unity, in accordance with Equation 3, and since the shaft positions of both potentiometers 205 and 206 are at all times set to the value of wn, therefore the respective outputs from the latter potentiometers are functions of the input voltages across the potentiometers and directly proportional to wn. It is to be noted, of course, that the use of the second potentiometer 206 connected in the manner shown to obtain @n2 is merely a close approximation method since the effect of the second potentiometer 206 is to ioad the first potentiometer 205. Hence, to increase the accuracy of such an approximation approach, the second potentiometer 206 is supplied with a somewhat higher value of total resistance than that of the first potentiometer 205.

It has been established that the input to amplifier 202 is wne while the input to integrating amplifier 203 is wn2e. The gain of amplifier 202 is adjusted to provide the multiplication factor of 2, called for in Equations 2 and 7, as well as the usual reversal in sign through inherent phase shift. Similarly, the gain of amplifier 203 is adjusted to provide multiplication by the factor oin Equation 2, which in the illustrative example is set equal to one-half. Thus, the output of amplifier 202 is equal to l Zame whereas the output of integrating amplifier 203 is By varying the nature of the input data ej to amplifier 201 and the value of the constant initially fed to the integrating amplifier 203 prior to commencement of integration at the compute signal, the nature of the output fiight trajectories represented by eo can be varied to suit Vvarious missile and target conditions. Input data is varied by means of selective exploitation of output data available from an input converter servo section 24 and also through suitable utilization of a variety of function generators designated generally as 207 in FIG..2 of the drawings and which may be selectively placed either in or out of the input circuit to amplifier 201 by means of a single switch 20S or a plurality of such switches controlling a plurality of possible input function generators.

It is to be noted that the second order computing circuit described above allows eo to gradually approach ei, until actual intercept with the target, in a manner and degree dependent upon the rapidity with which eo can change, the latter being controlled by the magnitude of the gain sensitivity factor wn which also controls the shaft position settings of the potentiometers 20S and 206.

:If wn is allowed to reach its maximum value of unity, then ;eo can approach e, very rapidly since the feedback` loop gain is very high, whereas if the wn settings of potentiom- , eters 205 and 206 are near their lower limits that is with wn considerably less than unity, then the feedback loop gain is very low and eo will follow input data e, rather slowly, and with a substantial lag. A primary purpose, therefore, of the secondorder smoothing programmer circuitry is to cause e0 to approach e1 from an initial value of eo and proceed through a transient midcourse phase whose nature and time duration is controlled by the manner in which the gain sensitivity factor, wn, is programmed. A large value of wn increases the tightness of following of eo and reduces the transient error due to rapid variation in the input voltages. On the other hand, a low value of wn provides heavy smoothing of noisy radar data. In practice, therefore, the manner of programming wn is a compromise between the suppression of noise on the input data ei, to prevent vibration of the missile Wing fiaps and consequent increased drag and possible damage to missile wing servos, and yet maintain maximum probability of target kill. Thus, wn is programmed in time in such a manner as to enable eo to follow e1 to an eX- tremely close tolerance as actual collision with the target approaches, that is, a large value of on is utilized to increased the tightness of following during the latter portion of the programmed fiight trajectory, when the missile is closed to the target, whereas a small value of wn is utilized for the initial portion of the programmed flight trajectory to provide heavy smoothing of noisy radar input data. The latter serves to minimize missile wing vibration, servo damage and air drag, and enhance fuel economy. FIG. 6 illustrates the manner in which the wn is varied during a typical flight program in accordance with Equation 3 with the value of c arbitrarily set equal to four. y

Referring now to FIG. 3 of the drawings, which illustrates a second order computing circuit for a missile programmer incorporating the novel coast provisions of the instant invention, there is shown a guidance computer circuit comprising a series of four ampliiers 201, 202', 203' and 204' in an arrangement which allows for ready switching from the normal programming mode to the coast programming mode whenever required. The major distinctions between the circuitry illustrated in FIG. 3 of the drawings as opposed to that' of FIG. 2 of the dnawings is in the modification to amplifiers 202, 203 and 204 of FIG. 2 which have been replaced by new amplifiers 202', 203' and 204', respectively, in FIG. 3. Y

In essence, the summing amplifier 202 in FIG. 2 has been changed to the integrating amplifier 202' in FIG. 3 while amplifier 203 has been modified in that the new integrating amplifier 203' has been provided with a switching arrangement 209 for selectively grounding the input to the integrating amplifier 203'. It will be noted that the input circuitry to amplifier 202' in FIG. 3 is shortcircuited to ground and the elle voltage obtained from potentiometer 205 is directed to the initial Lcondition circuit of amplifier 202'. This is the normal programming Vmode -for integrating amplifier 202'. The latter arrangement insures that the output of amplifier 202' will follow the input une but with the usual phaseV inversion previously provided by amplifier 202 in FIG. 2 to provide reversal in sign. Amplifier 204 has been modified infFIG. 3 by substituting amplifier 204! which receives the'output signals from both amplifiers 202' and 203. However, the input channel lof amplifier 204', which receives output from amplifier 202', has been modied to raise thefgain of that channel to two in order to compensate for the reduction in gain in new amplifier 202' to unity, as op- 'posed to the previous gain of two in the unmodified am- ,under the control of a relay system in the vground tracking radar 22 ofAFIG. l. Iuthe event target trackis lost at any time after. a programmed missile has been launched, relays in the tracking radar 22arc actuated to place the radar 22 in the search condition. The same relays in tracking radar 22 which ei'iect the change from track to Search condition simultaneously switch integrating amplier 202 from the initial condition to the compute state and also operate switch 209 to short-circuit to ground the input to integrating amplifier 203. This is in contrast to operation during a normal programming mode when these relays are in the initial conditionV state for amplifier 202 and the ungrounded input state for integrating ampiier 2%. When target track is lost and the coast mode is in eiect, both 202 and 2% are in the integrate state with their inputs grounded. Therefore, Whatever charge has accumulated on the integrating capacitors of amplifiers Z692 and 203 during the normal programming mode is held at a iixed value equal to the value it had at coast mode initiation. In essenceptherefore, do, the rate of change of the program, is held constant at the value it had at coast initiation. Thus, the program continues at the same rate of change it had at loss of target track since do is stored continuously for use during the interval when radar target track is interrupted. Since the input en to integrating amplier 204 isheld constant, the etect, therefore, is to continue the rate of change of missile position during intervals of loss of track at the same rate at which it was changing when loss of target track rst occurred. The latter is an attempt to extrapolate during tracking loss intervals between the program trajectory in eeet at loss of track and that which is to be expected when track is regained.

Accuracy of extrapolation by means of a coast mode provision is limited primarilyv by three factors. These are: (l) lack of linearity of target motion during coast intervals; (2) inherent non-linearity of the standard program trajectory prior to loss; (3) integrator instability or drift. All three of these factors are functions of time in the sense that increased coast intervals will result in increased trajectory deviations. However, it has been determined empirically that the apparatus of the instant invention, when used even with the most non-linear of standard programs, limits deviations in program elevation and azimuth angles to no more than one half a degree, even for coast intervals of as long as two minutes duration.

Referring now to FIGS. 4 and 5 of the drawings, plots of EG-ET against time are vshown fortwo types of standard second order programs previously designated as type A and type l These graphs illustrate the manner in which the missile trajectory is varied afterinitiation of the coast mode. It will be observed that for the typek A midcourse prograrnshown in FIG-4, whichfis of the linear variety, extrapolation'error during the coast tory and would remain so during any coast interval. Y For the type B midcourse program, on the other hanihdeA picted in FIG. 5 of the drawings, it is noted that -the' tary losses of target track, frequently encountered with distant targets.

Obviously, many modiiications 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 specically described. l

What is claimed is:

l. Apparatus for controlling the night course of -a guided missile toV a target comprising a guidancetcomi puter, a source of D.C. voltage signals relating toy target and missile position data for use by said guidance cornputer, said computer including input amplifiers for receiving said DC. voltage signals, amplifier and integration circuit means responsive to the respective output sig-A. i

nals from said'input ampliiier for acting upon said output signals in accordance with the relations where C' y Tri-FC' Y where Tp, is the time remaining until intercept of the VVtarget bythe missile and C is a constant,Y

means tofeed back theDCJOutput signals--frorn said Y 'computer tothe inputs of said input amplifiers, A'normally extrapolation error is somewhat greater due to inherent non-linearity of 'the manner in which EG-ET varies with time. However, even in this instance, it will be noted that the return to the normal Vprogramming modef is accom-` plished with minimial error. These situations are in contrast to that which would prevail in'A the event'such a second order programming computer were not provided' with coast mode provision not only provides vthe versatility of a wide variety of programmed iiight trajectories, the inherent advantage of the second orderV programmer, but also adds profoundly to the economy and'probability yandV Y 7 0. Hence, the improved second order missile programmer inactivated means to hold the value ofco constant ,inthe 'active state during periodsV of loss` or target trackbyv a `tracking radar. f 4 Y I' v- 'Y 2.' Apparatus fiorcontrolling the ightfffcourrse ofl a` l guided missile toa target comprising aguidance computer, Ea' source of voltage'v-signals relating to targetfarid-missile vposition ,data for use by said guidance computer, .said

. computer includinginput ampliers for'freceiving y voltage signals, amplifierA and integration ycircuitmeans l responsive to 'the respective output'signals from sa .put `amplifiers r=P Y 'arias at? e9.: 2mn (e0 fei) ficr @Keg-ef) dijiconstant where e, is kthe input data tofpsaid computerf.v Y l f .on is .the gain sensitivityfactor of thefc'omputer, l: Y aisafconstannV Y vP isthe presen-t time, and 75,.

G islthe time of guidanceinitiation i l 'i k `for acting upon 'saidoutput signals',in"'ac5 `y GO cordance with the relations Y j to produce output signals proportional to desired guidance transmitter positions, means to vary the magnitude of the computer gain sensitivity factor wn as a function of predicted time to go to target intercept to smooth noisy target and missile input data until maximum maneuverability of the missile is required, means to feed back the output signals from said computer to the inputs of said input amplifiers, normally inactivated means in said guidance computer to hold the value of o constant in the active state during periods of loss of target track by a tracking radar.

3. Apparaus for controlling the flight trajectory of guided missile to its target comprising a guidance computer, means to supply input voltages to said computer proportional to target and missile position parameters, means Within said guidance computer to manipulate said parameters in accordance with the relations Where and eo is the output data from said computer,

ei is the input data to said computer,

o is the first derivatives of eo,

wn is the gain sensitivity factor of the computer, o' is a constant,

P is the present time, and

G is the time of guidance initiation means to feed back the output voltages from said guidance computer to its input, normally inactivated means in said guidance computer to hold the value of e'o constant in the active state during periods of loss of target track by a tracking radar, and means delaying maximum gain sensitivity of said computing circuit until the missile is close to the target, whereby noisy radar input data is smoothed until maximum maneuverability of the missile is required and loss of missile control due to loss of target track is avoided.

4. In an apparatus for controlling the flight trajectory of a guided missile, a guidance computer comprising elevation and azimuth computing sections, each of said computing sections including input summing amplifier means for receiving a plurality of input functions relating to missile and target position data and whose output consists of a voltage proportional to the sum -of said input functions, a first potentiometer, means to direct the output voltage `from said, summing amplifier means acrosslsaid first potentiometer, a first integrating amplifier Whose input is grounded, a second potentiometer, means to derive a voltage from said first potentiometer proportional to the gain sensitivity factor of the guidance computer and to direct said voltage to the initial condition circuit of said ,first integrating amplifier and simultaneously across said first integrating amplifier from the initial condition state t to the compute state and by grounding the input to said second integrating amplier.

5. In an apparatus for controlling the fiight trajectory of a guided missile, a guidance computer comprising elevation and azimuth computing sections, each fof said com- A puting sections including input summing amplifier means for receiving a plurality of input functions relating to missile and target position data and Whose output consists of a voltage proportional to the sum of said input functions, a first integrating amplifier Whose input is grounded, a second integrating amplifier, means to derive a first voltage from the output of said summing amplifier means proportional to the gain sensitivi-ty factor of the guidance computer and to direct said voltage to the initial condition circuit of said first integrating amplifier, means to derive a second voltage from the output of said input summing amplier means proportional to the square of the gain sensitivity factor of the guidance computer and to direct said second Voltage as input to said second integrating amplifier, third integrating amplifier means for receiving as inputs the output voltages from said first and second integrating amplifiers, normally inactivated means for maintaining said output voltages from said first and second integrating amplifiers constant during periods of loss of target track by a tracking radar by switching said first integrating amplifier from the initial condition state to the compute state and by grounding the input to said Second integrating amplifier.

6. In an apparatus for controlling the flight trajectory of a guided missile, a guidance computer comprising elevation and azimuth computing sections, each of said computing sections including input summing amplifier means for receiving a plurality of input functions relating to missile and target position data and whose output consists of a Voltage proportional to the sum of said input functions, a first integrating amplifier whose input is grounded, a second integrating amplifier, means to derive a first voltage function from the output of said input summing amplifier means and to direct said first voltage function to the initial condition circuit of said first integrating amplifier, means to derive a second voltage function from the output of said input summing amplifier means and to direct said second voltage function as input to said second integrating amplifier, third integrating amplifier means for receiving as inputs the output voltages from said first and second integrating amplifiers, normally inactivated means for maintaining said output voltages from said first and second integrating amplifiers constant during periods of loss of target track by a tracking radar byswitching said first integrating amplifier from the initial condition state to the compute state and by short-circuiting t0 ground-the input to said second integrating amplifier.

7. In an apparatus for controlling the flight trajectory of a guided missile, a guidance computer comprising input summing amplifier means for receiving a plurality of input functions relating to missile and target position data and Whose output consists of a voltage proportional to the sum of said input functions, a first integrating amplifier Whoseinput is grounded, said' first integrating amplifier being provided with an initial condition circuit, a second integrating amplifier, means to derive a first voltage function from the output of said input summing amplifier means and to direct said first voltage function to the initial condition circuit of said first integrating amplifier, means to derive a second voltage function from the output of said input summing amplifier means and to direct said second voltage function as'input to said second integrating amplifier, third integrating amplier means for receiving as inputs the output voltages from said first and second integrating amplifiers, normally inactivated means for maintaining the inputs toV said third integrating amplifier means constant during periods of loss of target track by a tracking radar by switching said first integrating amplifier from the initial condition state to the compute state and by simultaneously short-circuiting to ground the input to said second integrating amplifier.

8. In 'an apparatus for controlling the flight trajectory of "a guided missile, a guidance computer comprising a first integrating amplifier Whose input is grounded and having an initial condition circuit, a second integrating amplifier, means to provide a first voltage function relating to missile and target position data as input to the initial condition circuit of said first integrating amplifier, means to provide a second voltage function relating to missile and target position data as input to said second integrating amplifier, third integrating amplifier means for receiving as inputs the output voltages from said first and second integrating amplifiers, normally inactivated means for maintaining the inputs to said third integrating amplifier means constant during periods of loss of target track by a tracking radar by switching said first integrating amplier from the initial condition state to the compute state and by simultaneously short-circuiting to ground the input to said second integrating amplifier.

9. In an apparatus for controlling the fiiglit trajectory of a guided missile, a guidance computer comprising input summing amplifier means for receiving ya plurality of input functions relating to missile and target position data and whose output consists of a voltage proportional to the sum of said input functions, first and second integrating amplifiers, means to derive a first voltage function from the output of said input summing amplifier leans and to direct said first voltage function as input to said first integrating amplier, means to derive a second voltage function from the output of said input summing amplifier means and todirect said second voltage function as input to said second integrating amplifier, third integrating amplifier means for'receiving as inputs the output voltages from said first and second integrating amplifiers, normally inactivated means to maintain the voltage outputs from said first and second integrating amplifiers constant during loss of target track.

l0. ln an apparatus for controlling the flight trajectory of a guided missile, a guidance computer comprising input summing amplifier means for receiving a plurality of input functions relating to missile and target position data and Whose output consists of a voltage proportional to the sum of said input functions, first and second integrating amplifiers, means to derive a first voltage function from the output of said input summing amplifier means and to direct said first voltage function as input to said first integrating amplifier, means to derive a second voltage function from the output of said input summing amplifier means and to direct said second voltage function as input to said second integrating amplifier, third integrating amplifier means for receiving as inputs the output voltages fromsaid first and second integrating amplifiers, and Whose output is an electrical signal relating to programmed missile position, and normally inactivated means to maintain the rate of change of output electrical signals from said third integrating amplifier means constant during periods of loss of target track.

having an initial condition circuit, a second integrating amplier, a third integrating amplifier for receiving as inputs the output signals from said first and second integrating ampliers and -Whose output is an electrical signal relating to programmed missile position, normally inactivated means to switch said first integrating amplifier from the initial condition ystate to the compute state and to short-circuit to ground the input circuit to said second integrating amplifier during periods of loss of target track.

13. In an apparatus for controllingthe flight trajectory of a. guided missile, a guidance computer comprising a first integrating amplifier, a second integrating amplifier, a third integrating amplifier for receiving as inputs the output voltages from said first and second integrating arnplifiers and whose output is an electrical signal relating to programmed missile position, normally inactivated means to maintain the output voltages from said first and second integrating amplifiers constant during periods of loss of target track.

14. A missile guidance computer which acts upon signal outputs from a target tracking means comprising first and second integrating amplifiers coupled tol receiveV outputs from the tracking means; a third integrating amplifier for receiving as inputs the output electrical signals from said first and second 4integrating amplifiers; and condition responsive means activated upon loss of target signal to maintain the rate of change of output electrical signals from said third integrating amplifier constant during periods of loss of target signal.

References Cited bythe Examiner UNITED STATES PATENTS 2,401,779 6/46 Swartzel 330--147X 2,709,804 5/55 Chance et al. 343-7.4

2,956,236 10/60 Klutadt 330--123 SAMUEL FEINBERG, Primary Examiner.

CHESTER L. JUSTUS, Examiner.

Claims (1)

11. IN AN APPARATUS FOR CONTROLLING THE FLIGHT TRAJECTORY OF A GUIDED MISSILE, A GUIDANCE COMPUTER COMPRISING FIRST AND SECOND INTEGRATING AMPLIFIERS, A THIRD INTEGRATING AMPLIFIER FOR RECEIVING AS INPUTS THE OUTPUT VOLTAGES FROM SAID FIRST AND SECOND INTEGRATING AMPLIFIERS, NORMALLY INACTIVATED MEANS TO MAINTAIN THE RATE OF CHANGE OF OUTPUT SIGNALS FROM SAID THIRD INTEGRATING AMPLIFIER CONSTANT DURING PERIODS OF LOSS OF TARGET TRACK.

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