US3905563A

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

Info

Publication number: US3905563A
Application number: US354386A
Authority: US
Inventors: Saburo Nagoshi
Original assignee: Fuji Heavy Industries Ltd
Current assignee: Subaru Corp
Priority date: 1972-09-28
Filing date: 1973-04-25
Publication date: 1975-09-16
Anticipated expiration: 1992-09-16
Also published as: DE2325355A1; DE2325355B2; JPS4970499A; FR2201773A5; JPS552555B2

Abstract

A system for controlling a missile motion in the homing mode comprises means for designating by an error signal a direction in which a missile is guided in the homing mode, and means for compounding with the error signal a predetermined biasing signal imparting acceleration in one or two dimensions normal to a direction of the missile derivation. The biasing signal causes shells from anti-aircraft guns and anti-missile missiles to increase the miss-distance to the missile, within the allowable maneuverability of the missile, whereby the missile evades defensive action and finally zeroes in on a target.

Description

Unite States Nagoshi 51 Sept. 16, 1975 SYSTEM FOR CONTROLLING A MISSILE MOTION IN THE HOMING MODE [75] Inventor: Saburo Nagoshi, Tokyo, Japan [73] Assignee: Fuji Heavy Industries Ltd., Tokyo,

Japan [22] Filed: Apr. 25, 1973 [21] Appl. No.: 354,386

F. Ross; Herbert Dubno [5 7] ABSTRACT A system for controlling a missile motion in the homing mode comprises means for designating by an error GENERATOR V [30] Foreign Application Priority Data signal a direction in which a missile is guided in'the Sept. 28, 1972 Japan 47-97345 homing mode and 1 for P with F error signal a predetermined biasing signal imparting [52] US. Cl. 244/3.15 acceleration in two dinfensions f f to 51 Int. Cl. F42b 15/02; F41 g 7/00 rectlon of the "1155116 The blame f f 58 Field of Search 244/3.15 causes from guns and -"1 missiles to increase the miss-distance to the missile, [56] References Cited within the allovvable maneuverability of the missile,

UNITED STATES PATENTS whereby the missile evades defensive action and finally zeroes in on a target. 2,992,423 7/1961 Floyd et a1. 244/3. 15 i 3,140,482 7/1964 Duncan et a1. 244/3.15 8 Cl ims, 16 Drawing Flgures a DEVIATION SERVO SERVO ST E G DETECTOR ACTUATOR E AMPLIFIER VALVE UNIT ACCELEROMETER x6 MOTION LATERAL- ACCELERATION sum 2 OF 9 PATENTEU 3591 1975 a MEZwWEEmE 5622 mgmgnsmsms 3305553 sun-:1 3 OF 9 MlSSlLE MOTIUN (ONE-DIMENSIONAL NAVIGATION) f l 1 0 2 4 6 8 I t (spa) \IITT 01214161320 PATEFHEU 38055563 sum u [If 9 MlSSllE rusm PATH 30%) (TWO-DIMENSIONAL munsanum m Bo v 1 a '9 E "Bo 2-5596 5 nsec 20-0sec IZ-Ssec IO-Usec MISSILE ACCELERATION (TWO-DIMENSIUNAL NAVIGATION) t (sec) MISSILE FLIGHT PATH flg) (TNREE'NMENSIONAL NMHOATION) Bo v 2 o Bo MISSILE AOOELERATION (THREE DIMENSIONAL NAVIGATION) Lsl 3,905,563 SHEET E OF 9 PUEFHED 51975 z 5 52% wzw MINEWDM TH)! PAIEIIIED I SW5 3. 905,563

SHEET 7 OF 9 MISSILE FLIGHT PATH (TWO-DIMENSIONAL NAVIGATION) I LL /g2 5|O 0 -5|O "Bo 2.5sec ZO-Usec 15-0 125m se -50- I'I-Ssec 50sec Ill-Uses 1.5 sec MISSILE ACCELERATION (TWO DIMENSIONAL NAVIGATION) i(sec) fjfT" 0 UJ wSELIBSYS 3,905,563

SHEET 8 OF 9 MISSILE FLIGHT PATH (THREE-DIMENSIUNAL IIMIGRIIUII) sec 15.0

-1uo -"sec 7-53 5 MISSILE ACCELERATION (THREE DIMENSIONAL NAVIGATION) SYSTEM FOR CONTROLLING A MISSILE MOTION IN THE HOMING MODE The present invention relates to a control system for guiding a missile in homing mode to a target, with the missile evading defensive action by anti-aircraft guns and anti-missile missiles.

A missile is generally controlled in the homing mode in a manner tending to reduce the error angle of the missile flight path to the target at the initial stage of derivation, thereby nearly maintaining a constantbearing course. Consequently, the missile flight path straightens gradually (unless the motion of a target necessitates considerable maneuvering of the missle) so that an intercept computer makes an exact forecast of the future missile position enabling effective defense with anti-aircraft guns or anti-missile missiles.

Conventional missile-control systems impart additional acceleration and deceleration along the missiles course with the intention of reducing a kill probability of the anti-aircraft guns or anti-missile missiles. Nevertheless, such control is hardly available to evade defensive action under actual conditions where the enemys weapons are often situated near the target. Since the effective intercepting time of shots of the anti-aircraft guns corresponds to the final period of the missile derivation in which the error angle to the target should be quickly eliminated, it is very difficult to evade such defensive action except by replacing the missile homing mode in the final period of derivation by a conventional programming mode that results in deflecting the missile from the target on account of the unavoidable accumulation of the missile divation error.

As object of the present invention is to provide means in such a system for so improving the control of missile navigation as to enable its effective evasion of defensive action all along the flight path.

l achieve this goal by adding a lateral acceleration bias in one or two dimensions to a error signal due to the missile deviation, with continued guidance of the missile flight toward the target for eventual collision therewith.

Thus, my present system includes aboard an in-flight homing missile such conventional components as sighting means for continuously determining the location of a line of sight from the missile to the target, detector means responsive to the sighting means for generating the aforementioned error signal as a function of a deviation of the missile flight path from the line of sight, steering means for changing the course of the missile, and control means therefor connected to the detector means for delivering to the steering means a corrective signal derived from the error signal. In accordance with the present improvement, I further provide biasing means for superimposing upon this corrective signal a lateral-acceleration signal effective in a direction perpendicular to an invariable reference direction generally parallel to the flight path, such as the axis (X) of a set of three orthogonal coordinate axes X, Y, Z. As described in another expression hereinafter, a differential signal (in which this lateral-acceleration signal is subtracted from the output of the feedback path involved in steering diminishes the error signal in the input of the control means so that its effect upon the steering means tends to change the existing flight path.

With two-dimensional navigation, two lateralacceleration signals are generated which are effective in directions parallel to axes Y and Z. In this case, to achieve the desired flight-path changing effect, these two signals must of course be superimposed upon respective Y and Z components of the error signal.

The various features and advantages of the present invention will become more clearly apparent from the following description given with reference to the accompanying drawing in which:

FIG. I is a diagram illustrating in principle a control system embodying the present invention;

FIG. 2 is a diagram illustrating details of the control system;

FIG. 3 is a geometrically exaggerated diagram of the missile deviation;

FIG. 4 is a plot of angular parameters in a tracking system;

FIG. 5 is a set of graphs of the missile motion determined by one-dimensional bias in the direction of deviation;

FIG. 6 is a vector diagram of the acceleration bias for the missile maneuvering;

FIG. 7 is an orthogonal coordinate system indicating the missile deviation with two-dimensional bias;

FIG. 8 is a graph of the missile acceleration required for a bias according to FIG. 7;

FIG. 9 is an orthogonal coordinate system indicating the missile deviation with three-dimensional bias;

FIG. 10 is a graph of the missile acceleration required for a bias according to FIG. 9;

FIG. 11 is a graph of total kill probabilities in terms of lateral acceleration;

FIG. 12 is an orthogonal coordinate system indicating the missile deviation with two-dimension bias;

FIG. 13 is a graph of the missile acceleration required for a bias according to FIG. 12;

FIG. 14 is an orthogonal coordinate system indicating the missile deviation with three-dimensional bias;

FIG. 15 is a graph of the missile acceleration required for a bias according to FIG. 14; and

FIG. 16 is a graph of the missile maneuvering acceleration necessary to keep a definite miss-distance from a surface-to-air missile.

In a control system according to my present invention there is illustrated in block form in FIG. 1 and error signal supplied from a missile-deviation detector a, which signal is not applied immediately to a control device b but is compounded with a lateral-acceleration bias from a generator c so that the compound signal is applied to the control device b as a control signal.

FIG. 2 shows details of such a control system in which the control device b includes a servo amplifier 1, a servo valve 2 and an actuator 3 for a steering unit 4, serially arranged in a forward path, as well as a feedback path from steering unit 4 to the input of amplifier 1 comprising aerodynamic motion transfer function unit 5 working into an accelerometer 6. Thus, a signal fed back from the output of the accelerometer 6 diminishes the bias signal from the output of generator 0 to produce a differential signal. A control signal is obtained by deducting the differential signal from the error signal supplied by the deviation detector a. Instead of the feedback signal from the output of the accelerometer, as shown, the feedback signal could come from the output of a rate gyro, or the deflection angle of a control surface could be utilized for the same purpose.

To point out the advantage of the missile navigation obtained by the superimposition of an acceleration bias in accordance with my invention, some test results will be reported hereinafter in cases with oneand twodimensional acceleration bias as well as, in some instances, an acceleration bias in a third dimension, i.e. in the direction of the missile bearing course, added to the error signal in a control system designed to guide the missile to a stationary target by proportional navigation in the homing mode.

1. The Missile Motion The X-axis of the missile coordinate system extends along a theoretical missile-collision course as shown in FIG. 3, wherein the homing trajectory of the missile leads from a starting position to a pinpoint S of the target. Furthermore, the missile positio n T at a time t is defined by coordinates x Y Proportional navigation requires that the change of the missile flight direction always be directly proportional to the angular rate of change of a line of sight which leads from the missile to the target (with a proportionality factor N where n missile lateral acceleration, i.e. the acceleration directed along the Y-axis of the missile coordinate system but with a negative sign g gravitational acceleration V missile velocity Now substitute d from Eq. (1) into Eq. (I).

NT'VT [a c d! (2) The line-of sight rate (dA- /dt) is measured by a tracking device which may be a radar, an optical system, or an infrared system. The tracking device points its antenna in the direction of the line of sight to the target, as accurately and as quickly as possible.

The principle of the measurement of (dk ldt) by means of the tracking device will be explained in connection with FIG. 4.

A line-of-sight angle relative to the fixed reference line A boresight angle of a tracking antenna relative to the same fixed reference line (if the tracking were perfect, the antenna boresight would coincide with the line of sight.)

e error angle, i.e. the angle corresponding to the difference between A and M For a perfect tracking system having no time lag, e O.

In the actual physical case. the antenna lags behind the instantaneous line of sight by an amount which is determined by the tracking time constant T so that we have Since the error signal 6 multiplied by a constant gain factor k equals the lateral acceleration n it may be written as The right-hand side of Eq. (2) is equal to that of Eq. (3) with the exception of the factor 1/1 r p), so that we may write:

Substitute this into Eq. (3),

(where y is a distance normal to the fixed reference line, i.e. measured on the Y-axis coordinate) Substitute this equation into Eq. (5):

Transpose the denominator of the right-hand side of Eq. (5') to the left-hand side:

The instantaneous line-of-sight angle M is given by:

(referring to FIG. 3)

where R remaining distance of the missile flight path Differentiate that with respect to time:

111- l dy V d1 R, d! R V from the illustration of FIG. 3)

Consequently, Eq. (6) is rewritten as (where z is a distance normal to the fixed reference line, i.e., measured on the Z-axis coordinate) 2. The Missile Acceleration The change of the missile velocity is defined with oneand two-dimensional acceleration as follows:

for 1= 0 V V, for l 1,, V1 V2 (1,, being the total homing flight time) I. In case of one-dimensional acceleration:

The missile velocity at a time t is determined by Therefore, the remaining distance of the missile flight path is 0 VTdt i t) 2 l 7 v 1;. Nu

c (I10 I9) Substitute t t in Eq. (9): (t is the hitting time of a shot from a defensive weapon.)

The hitting time is obtained from Eq. (10) which includes the hitting range (R from the target. 2. In case of two-dimensional acceleration:

The missile velocity at a time t is determined by:

Therefore, the remaining distance of the missile flight path is to R =3 V dt Vi(t ,"t)

Substitute t=t in Eq. (ll):

The hitting time is obtained from Eq. (13) which again includes the hitting range (R from the target. 3. Determination of Position and Motion Data of the Missile From information obtained by a defending FCS (Fire-Control System), the missile position and time indicated in FIG. 3 by reference characters F and are determined as follows: 1. Anti-aircraft (Guns The predicted time t coinciding with the instant when a shell fired at S has covered a distance (R )=1 may be written as (V, average velocity of the shell) but a delay from the calculation of the missile position and motion data to the firing of the shell amounts to 0.5 second in practice. 2. SAM (Surface-to-air missile) (ima but, supposing that the SAM takes an acceleration of 20.4g until it reaches a velocity V 2 mach (680m/s), a time required for acceleration is 680/20.4g 3.4 (sec). The SAM velocity may be assumed as 2 mach by subtracting a dead time from the time required for acceleration, this dead time being 3. 4 sec X k 1.7 sec. Eq. (15) employs a time of 2.2 secas mentioned above, which is given by the delay occurring between the calculation of the missile data and the SAM firing, added to the dead time, i.e.

0.5 +1.7 2.2(sec) 3. Determination of the Future Position of the Missile with Two-Dimensional Acceleration The FCS predicts the future position of the missile pointed at a location C in FIG. 3. Since the calculation of the FCS is based on position, velocity and acceleration data of the missile, the calculated time t,; does not coincide with the hitting time i (though the former co- 2 (11 I: ).ean'n incides with the latter with one-dimensional acceleration).

The remaining distance for homing is calculated by FCS as On the other hand, substitute t=t into Eqs. (12) and (11).

and substitute t=t,- into the Eq. (11) differentiated with respect to t:

These equations together give us the following equation:

l. Anti-aircraft Guns A predicated hitting time of the shell is defined as in Eq. (14):

2. Sam

A predicted hitting time of the SAM is defined as in Eq. (15).

The calculated time I is determined by solving Eqs. (17) and (18) for anti-aircraft guns'and by solving Eqs. (17) and (19) for the SAM, in using the technique of successive approximation.

5. The Miss-Distance of a Missile-lntercepting Shell On the assumption that the impact of the shell on the missile occurs when the shell and the missile simultaneously arrive at a point H whose projection on the X- axis represents the hitting time t the miss-distance m may be written as where m component in the Y-axis direction m 32 component in the Z-axis direction I. One-dimensional Acceleration- The Y-axis component is Similarly, the Z-axis component is Supposing the flight path of the shell to be straight, the displacement of the shell along the Y-axis at the hitting time t,, is

whence the miss-distance along the Y-axis may be written as the missile displacement along the Z-axis is dz (Z'r)1=r (Z'r)1=1 X (M r) The miss-distance along the Z-axis maybe written as The miss-distance will be obtained by substituting Eqs. (24) and (26) in Eq. (20).

6. The SAM Navigation The FCS predicts the factors of the missile in order to count the future position C of the missile aimed at by the SAM when the missile reaches a point F on the missile bearing course. The SAM starts along a collision course to the left of time t of 2.2 (sec) after the PCS prediction, this collision course being illustrated by the straight line SO in FIG. 3.

On the assumption that the SAM is derived by proportional navigation (with a proportionality factor NM),

the lateral acceleration n of the SAM in the Y-axis di- The illustration of FIG 3 yields rection is obtained as in Eq (4):

. v "Mu g 1 TMP AM 5 t where where R remaining distance of the SAM homing V average velocity of the SAM flight path. m time constant of the SAM Substitute this into Eq. (29): M, aiming angle of the SAM relative to the reference line (the lateral acceleration HM having a negative sign) 112M! 1M1 NMVM M M (FY11 7." 2 (I, RM )11 R "r '11 I ,2 =1

zly 1, NJIVM l/ll) t (30) d t with n g wherein, for the firm condition of the SAM,

(Y,., displacement of the SAM in the Y-axis direction) a 1m, Substitute the foregoing equation into Eq. (27): 111 =1 d! F! !T)I=I A!u)1=1 (R (F M i=1 [it P a N,\IV.\I lp .\l|l

' (where y angle of the SAM in the XY-plane, Transpose the denominator of the right-hand side of when almmg at the f Posmon C of the mlssfle this equation to the left hand Side fr)om the point S, relative to the fixed reference line Substitute these values into Eq. (30):

(an, (m, at 7 din d1! NMVM (28) y." yn NMVM NMVJI TM YM YT il'You Integrate both sides of Eq. (28): RM

(R.\I)I=ID t FIG. 3 indicates that n, =-N \,VMAM,, c 29) 1F) 11 F X (IF-IF) (1127 )t=tr- X z 2 y'w .\1( 'I)) (32) where C is the constant of integration. Equations for the SAM are also derived in conneclnsert the initial limiting conditions for 2 into Eq. tion with the component in the direction of the Z-axis,

(29) for defining the constant: ie:

1." N." N." V." V (1T) T." I R" Z." R." ZT M7": JI M RM)I=ID 21 FIT l I-) C W .\l !I)I= displacement of the SAM m the Z-axis direction,

y 02 angle of the SAM in the XZ-plane, when aim- (with 2,, t, 2.2 sec) ing at the future position C of the missile, from the point S relative to the fixed reference line. The integration constant C is defined by The next equations are derived from the illustration of FIG. 3.

5 Lil Q (R1'):=1,, d1 d1 T" 0 V (35) (RHH (36) going to zero in the limiting case t=0, with IO 1. One-dimensional Accelerating Substitute Eq. (9) into Eq. (36):

A .T )T V-V Tu M= (ml Munro (MW) (37) RT 2. Two-dimensional Acceleration Substitute (12) into (36); Substitute this equation into Eq. (41):

R v,+V 375 (38) I Simultaneously, the component in the Z-axis direc- I tron IS given by:

7. The Miss-Distance of the SAM With Reference AS Will be pp from 3, the hitting time is to the Missile The miss-distance of the SAM to the missle is based on the unavoidable fact that as the calculated maneuvering acceleration of the SAM attains and surpasses the g-limit, the actual maneuvering acceleration reach- 1)|=1,, es its saturation value. V1 (44) Accordingly, the evaluation of the components of the SAM miss-distance in the Y-axis and Z-axis directions must be executed with the addition of factors of gsaturation and desaturation. As the final period of the SAM derivation involves the fixed g-saturation, the following equations are established:

In order to determined the predicted time r,-, Eq. (14) can be used with anti-aircraft guns, and Eq. (15) with the SAM.

d I -l 2 z (ZT)l=! {(131)14 i I: |l u)+ 1: X 2 l The miss-distance will be obtained by substituting m The miss-distance of the shell is derived from Eqs. and m of Eqs. (39) and (40) into Eq. (20). (20), (21 and (22), and the SAM motion from the ap- According to the foregoing, the SAM miss-distance li rio of Eqs, (31 (32), (33), (34) and (35) as well is vectorially defined by Y-axis and Z-axis coordinates. as FIG, 3 i di t th t Equations (31 and (33) are available to calculated the miss-distance in the event that the bearing courses of R (V +V the missile and SAM are based on the same coordinate (45) system; otherwise, the value of Y and Z must be replaced by others obtained by coordinate conversion.

The preceding description explains the way of deterl nflss'dlstance of the SAM W reference to the mining the miss-distance with three-dimensional deviamlsslle derlved from the apphcatlon of q tion (39) and (40).

MISSILE NAVIGATION WITH MISSILE NAVIGATION WITH TWO-DIMENSIONAL MANEUVERING ONE-DIMENSIONAL MANEUVERING Since the velocity V is constant, Eq. (6) for the mis- When the missile is navigated in sea-skimming flight, sile motion may be rewritten as follows by integrating its interception by the SAM is difficult so that it is only its two sides: necessary to consider the use of anti-aircraft guns.

2 d y dy t 1 d1} N v r gjon y dt c +1 In this case, the missile motion is determined from Eq. (42), the hitting time from Eq. (44), the predicted time from Eq. (14) and the miss-distance of the shell from Eq. (21).

8. Example of Test Results Concerning Missile Deviation With One-Dimensional Maneuvering (to Evade Anti-Aircraft Gun) The equations to be applied are based on the following numerical values.

V 0.9 mach 1 0.5 sec 1,, sec

1 O-I 1.0 sec n "m, (constant) r= ll.0-l6.0 n,,,,=n,,,,

for:

time of flight to the effective range of gun of 3Km 3.8 sec firing rate 120 rounds/min lethal radius of shell 5m (with proximity fuse) time from FCS predicting the missile position to firing of gun 0.5 sec gun dispersion 0.00358 (10) where R distance measured from a firing position to a hitting position a standard deviation The equation of the predicted position of the missle In this example, the shells are successively launched at a firing rate within the hitting range (R) 500 4,000m.

Let us assume that the missile is deviated in its seaskimming flight and, though the explosion of shells concentrated in a region circumscribed by their lethal radius centered on the missile is theoretically effective to kill the missile, the explosion of some shells concentrated in the lower half of that region near the sea surface is too early on account of premature operation of the proximity fuse caused by the sea clutter so as to have no destructive effect on the missile.

Consequently, the cumulative kill probability per gun is shown in FIG. 11. This diagnosis indicates the need for increasing the evading capability of the missile, because the kill probability of the shell with reference to a missile subjected to lateral acceleration pursuant to this invention is limited to a value of 41% (n B 4.0), according to a curve I, in comparison with the case of a homing missile of conventional navigation where the limit is 99.5% (n 0).

9. Example of Test Results Concerning Missile Deviation With Two-Dimensional Maneuvering to Evade Anti-Aircraft GUN) If we consider the acceleration bias n as swung into the YZ-plane of the orthogonal coordinate system shown in FIG. 6, the missile motion may be plotted as illustrated in FIGS. 7 and 8 with the following numerical values: for:

t= 0 20.4 sec, d6/dt 5.45 o/sec.

ln l n (constant) where 0 angle due to lateralaccelera tion bias (-m) at a time t, equaling 0 for t 0 for:

t=20.4 22sec n 0 V 0.9 mach N1 4 T 0.5 sec 1 =22 sec (with n 11 n The cumulative kill probability of the anti-aircraft gun is represented by a curve II in FIG. 11 on the assumption of conditions indentical with those of the first example, i.e. with the missile not deviated in a seaskimming flight. Curve II indicates a reduction of the kill probability to a value of 6% (n 2 3.0), i.e. the missile deviation with two-dimensional maneuvering acceleration produces an evading efficiency higher than that given by the missile deviation with onedimensional maneuvering acceleration.

[0. Example of Test Results Concerning Missile Deviation With Three-Dimensional Maneuvering (to Evade Anti-Aircraft Gun) The variation of the missile speed conforms to the two-dimensional acceleration and is numerically defined as follows:

for: I O 15 sec V1 0.9 mach The other values are the same as in the second example so that the missile motionis apparent from FIGS. 9 and 10.

The cumulative kill probability of the anti-aircraft gun is determined by another curve III of FIG. 1 1 under conditions identical with those of the second example. The kill probability of the gun is further reduced to a value of 4% (n'flnmr) 3.0), i.e. the evading efficiency of the missile in this case is still higher than that given by the missile deviation withtwo-dimensional maneuvering acceleration.

l 1. Example of Test Results Concerning Missile Deviation With Two-DimensionalManeuvering (to Evade Anti-Missile Missile MISSILE) If we again consider the acceleration bias n as swung into the YZ-plane, the missile motion may be plotted as, illustrated in FIGS. 12 and 13 with the following numerical values: for:

t; 0 204 sec dO/dt 30 o/sec.

I n} H (constant) where 0 angle due to lateral-acceleration bias (-M) at a time t, again equaling O for t 0 for:

t=20.4 22 Sec n O V 0.9 mach N 'r 0.5 sec t 22 sec The SAM parameters are determined as follows:

V 2.0 mach TM 0.5 sec maneuvering-acceleration limit of SAM 14g (relative to both the Y-axis and the Z-axis) hitting distance of the SAM 1,000 4,000m

Under the above-mentioned conditions, the missile maneuvering accelerations along the Y-axis or the Z- axis are plotted in full lines in FIG. 16 for a missdistance of 20m. Cruve IV represents the case where the missile and the SAM share the same set of coordinates 6 0), whereas curve V applies where their coordinates differ (6 45) with the Y-axis and the Z-axis of the missile include respective angles of 45 with the corresponding axes of the SAM.

Consequently, the required maximum maneuvering accleration of the missle is about 6.7g in the case of 0 0 6.0g in the case of 6 45 12. Example of Test Results Concerning Missile Deviation With Three-Dimensional Maneuvering (to Evade Anti-Missile Missile The variation of the missile speed conforms to the one-dimensional acceleration and is numerically defined as follows:

for:

t O-22 sec V 0.9 0.6 X (1/22) mach The other values are the same as in the fourth example so that the missile motion is apparent from FIGS. 14 and 15.

In this case, the SAM parameters are determined under conditions similar to those of fourth example, the missile maneuvering accelerations along the Y-axis or the Z-axis being given by dotted lines in FIG. 16 and the required maximum maneuvering acceleration of the missile is about 5.5g in the case of 0 0 (Curve V1) 5.2 g in the case of 6 45 (Curve VII).

This result shows the advantage of three-dimensional acceleration over two-dimensional acceleration.

Though the numerical examples given above deal specifically with the navigation of a missle to a station-' ary target, the principles of my invention can be readily extended to situations where the target is in motion.

What is claimed is: 1. A system for guiding an in-flight homing missile to a target, comprising:

sighting means for tracking the location of a line of sight from the missile to the target; detector means responsive to said sighting means for generating an error signal depending upon any deviation of the missile flight path from said line of sight;

steering means for changing the course of the missile;

control means for said steering means connected to said detector means for delivering a corrective signal derived from said error signal to said steering means; and

biasing means connected to said control means for superimposing upon said corrective signal to a lateral-acceleration signal effective in a direction perpendicular to an invariable reference direction generally parallel to said flight path.

2. A system as defined in claim 1, further comprising circuit means for generating a feedback signal subtracted by said lateral-acceleration signal to obtain a differential signal which is from said error signal so as to obtain a control signal.

3. A system for guiding an in-flight homing missile to a target, comprising:

sighting means for tracking the location of a line of sight from the missile to the target;

detector means responsive to said sighting means for generating an error signal depending upon any deviation of the missile flight path from said line of sight;

steering means for changing the course of the missile;

biasing means connected to said control means for superimposing upon said corrective signal a lateralacceleration signal effective in a direction perpendicular to an invariable reference direction generally parallel to said flight path and further adding an ecceleration bias to the missile derivation direction.

4. A system as defined in claim 3, further comprising circuit means for generating a feedback signal subtracted by said lateral-acceleration signal to obtain a differential signal which is deducted from said error signal so as to obtain a control signal.

5. A system for guiding an in-flight homing missile to a target, comprising:

detector means responsive to said sighting means for generating an error signal depending upon any deviation of the missile flight path from said line of sight; I

steering means for changing the course of the missile;

control means for said steering means connected to said detector means for delivering a corrective signal derived from said error signal to said steering means, said control means establishing a set of three coordinate axes including a first axis generally parallel to said flight path and two other axes orthogonal thereto; and

biasing means connected to said control means for superimposing upon said corrective signal two lateral-acceleration signals effective in directions parallel to said other axes.

6. A system as defined in claim 5, further comprising circuit -means for generating feedback signals substracted by said lateral-acceleration signals to obtain a differential signal which is deducted from said error signal so as to obtain a control signal.

7. A system for guiding an in-flight homing missile to a target, comprising:

steering means for changing the course of the missile;

control means for said steering means connected to said detector means for delivering a corrective signal derived from said error signal to said steering means, said control means establishing a set of three coordinate axes including a first axis generally parallel to said flight path and two other axes nal so as to obtain a control signal.

l-' l= =l

Claims

1. A system for guiding an in-flight homing missile to a target, comprising: sighting means for tracking the location of a line of sight from the missile to the target; detector means responsive to said sighting means for generating an error signal depending upon any deviation of the missile flight path from said line of sight; steering means for changing the course of the missile; control means for said steering means connected to said detector means for delivering a corrective signal deriveD from said error signal to said steering means; and biasing means connected to said control means for superimposing upon said corrective signal to a lateral-acceleration signal effective in a direction perpendicular to an invariable reference direction generally parallel to said flight path.

3. A system for guiding an in-flight homing missile to a target, comprising: sighting means for tracking the location of a line of sight from the missile to the target; detector means responsive to said sighting means for generating an error signal depending upon any deviation of the missile flight path from said line of sight; steering means for changing the course of the missile; control means for said steering means connected to said detector means for delivering a corrective signal derived from said error signal to said steering means; and biasing means connected to said control means for superimposing upon said corrective signal a lateral-acceleration signal effective in a direction perpendicular to an invariable reference direction generally parallel to said flight path and further adding an ecceleration bias to the missile derivation direction.

5. A system for guiding an in-flight homing missile to a target, comprising: sighting means for tracking the location of a line of sight from the missile to the target; detector means responsive to said sighting means for generating an error signal depending upon any deviation of the missile flight path from said line of sight; steering means for changing the course of the missile; control means for said steering means connected to said detector means for delivering a corrective signal derived from said error signal to said steering means, said control means establishing a set of three coordinate axes including a first axis generally parallel to said flight path and two other axes orthogonal thereto; and biasing means connected to said control means for superimposing upon said corrective signal two lateral-acceleration signals effective in directions parallel to said other axes.

6. A system as defined in claim 5, further comprising circuit means for generating feedback signals substracted by said lateral-acceleration signals to obtain a differential signal which is deducted from said error signal so as to obtain a control signal.

7. A system for guiding an in-flight homing missile to a target, comprising: sighting means for tracking the location of a line of sight from the missile to the target; detector means responsive to said sighting means for generating an error signal depending upon any deviation of the missile flight path from said line of sight; steering means for changing the course of the missile; control means for said steering means connected to said detector means for delivering a corrective signal derived from said error signal to said steering means, said control means establishing a set of three coordinate axes including a first axis generally parallel to said flight path and two other axes orthogonal thereto; and biasing means connected to said control means for superimposing upon said corrective signal two lateral-acceleration signals effective in directions parallel to said other axes and further adding in acceleration bias to the missile derivation direction.

8. A system as defined in claim 7, further comprising circuit means for generating a feedback signal subtracted by said lateral-acceleration signal to obtain a differential signal which is deducted from said error signal so as to obtain a control signal.