GB2433307A

GB2433307A - Missile control system

Info

Publication number: GB2433307A
Application number: GB8805553A
Authority: GB
Inventors: Anthony John Benson
Original assignee: Marconi Co Ltd; Alenia Marconi Systems Ltd
Current assignee: BAE Systems Integrated System Technologies Ltd
Priority date: 1987-03-09
Filing date: 1988-03-09
Publication date: 2007-06-20
Anticipated expiration: 2008-03-09
Also published as: SE8804394A; DE3831440B3; IT8867959A0; GB8805553D0; SE8804394D0; GB8705450D0; FR2928748A1; GB2433307B

Abstract

An anti-radar missile control system which is required to impose a glide phase and a terminal phase on the missile flight. The glide phase produces a constant angle of look and is required to pass smoothly into the terminal phase which has a circular or near circular path descending vertically or near vertically on the target. This basic form is subject to substantial error resulting from the large angle of look that obtains for a considerable part of the terminal phase. The terminal phase could be improved, sharpening up the curve and reducing the dwell at high look angle, by increasing the navigation gain. This, however, tends to increase the instability significantly. The invention effectively simulates a high gain trajectory by adding a bias to the look angle so inducing the glide phase to remain operative longer. This in effect determines a larger intercept angle on impact. By removing the bias during the terminal phase the intercept angle is brought back to the vertical (or other predetermined value) and a sharper curve results so gaining the benefit of a high gain system without lessening the stability.

Description

Missile Control System This invention relates to a missile flight path

control system particularly for a missile having a radar seeker for detecting and locking on to a stationary or near stationary target. Such missiles find application, for example, against ground or ship based radar installations.

It is a comon requirement for such missiles to fly in a shallow glide path for an initial phase which would take the missile directly over the target but at a considerable altitude, and for the missile to dive in a terminal phase when sufficiently close to the target so as to impact vertically or near vertically.

The transition between the glide phase and the terminal, or dive, phase is initiated in response to the angle-of-look, i.e. the angle between the antenna boresight and the missile-target sightline.

In a missile with a strapped down antenna, i.e. one whose boresight is fixed in alignment with the missile roll axis, the angle-of-look is thus the angle between the missile axis and the target sightline.

The terminal phase may conveniently employ a circular flight path which implies a navigation constant of 2 for the guidance system, such that the available rate of change of flight path direction is twice the rate of change of the line of sight angle in space. Such a circular path may be required to terminate vertically on the target and also to be tangential to the glide path at the transition.

Calculations can be made at successive points in the glide phase to determine the target line of sight, from the angle of look (between boresight and line of sight to target) and the missile body angle in space. The angle between the line of sight and the vertical gives the necessary lead angle (between flight path and sight line) to obtain a circular terminal phase. If this lead angle exceeds the actual lead angle (and, more conveniently, the actual look angle, assuming the boresight is close to the flight path) then a smooth transition to a circular path is not possible. This situation will obtain early in the glide phase when the complement of the line of sight angle is large and likely to exceed greatly the actual look angle. As the glide phase continues, a point is reached when the value of the necessary look angle has fallen to equal the actual look angle. A comparison is made and the equality (or near equality) is used to initiate the terminal phase.

In such a system the seeker measurements, and in particular the look angle, will generally contain errors which increase missile dispersion and degrade guidance accuracy. Such errors tend to increase with angle of look and are therefore particularly troublesome in missiles with fixed antennae where the target angle off boresight can be large.

An object of the present invention is therefore to try to reduce the initial period in the terminal phase for which the angle of look is large in a flight such as described above by way of example.

One way of achieving this object is to increase the navigation constant (gain) from the value 2 quoted above for a circular path. This would sharpen the initial curvature of the terminal phase and cause the missile to adopt a smaller angle of look more rapidly, but unfortunately the increased gain tends to cause instability, with consequent increase in dispersion and miss distance.

According to one aspect of the present invention, In a missile control system for a missile adapted to descend upon a stationary or near stationary target, the system including a radar seeker for controlling the flight-path in dependence, in elevation, upon the angle of look between the seeker boresight and the missile-target line of sight, the flight path comprising a glide phase and a terminal phase and means for initiating a transition between the glide phase and the terminal phase in dependence upon the line of sight angle in space in relation to a signal dependent upon the angle of look, a bias signal is incorporated with the angle of look signal so as to tend to reduce the effective angle of look in relation to the line of sight angle and thus delay the onset of the transition, the system further including means for reducing the bias signal to zero during the terminal phase.

The transition may be determined by a comparison between, on one hand, the angle between the line of sight and the terminal intercept path, and, on the other, the biased angle of look.

The bias signal may be maintained constant during the glide phase and reduced exponentially during the terminal phase.

Where the system is subject to inherent bias in measurement of the angle of look there may be included means for reducing the bias signal exponentially during a first part of the terminal phase and means for reducing net bias, i.e., including inherent bias, by a repeated quantisation and integration process for the remainder of the terminal phase.

According to another aspect of the invention, a missile control system for a missile having a trajectory comprising a glide phase and a terminal phase in which the missile is adapted to descend upon a stationary or near stationary target, the system including a radar seeker for controlling the flight path in elevation in dependence upon a determined angle of look between the seeker boresight and the missile-target line of sight, a method of reducing the average value of the angle of look in the terminal phase in which the predetermined terminal intercept angle is effectively increased temporarily to sharpen the trajectory during the terminal phase.

A system and method for controlling the flight path of a ground attack missile, which may be air or ground launched, will now be described, by way of example, with reference to the accompanying drawings, of which: Figure 1 is a diagram of alternative missile flight path trajectories towards the latter part of the flight; Figure 2 is a block diagram of a control signal path for the terminal phase of the flight; and Figure 3 is a flow chart for a bias reduction process effective in the terminal phase.

Referring to the drawings, Figure 1 shows basic and modified trajectories 1 and 3 which have identical glide phase (up to the point P) but different transitions to the terminal terminal, or dive' phase, i.e. at P and P' respectively. During the glide phase the missile is controlled to maintain, or at least to output, a constant angle of look A,of say, 300, between the missile seeker boresight BS and the missile-target line of sight LOS. In this system the seeker antenna is strapped down and the boresight coincides with the missile roll axis.

In the early stages of the flight the angle of look may exceed the preset figure of 30 . However, the angle output to the guidance system is limited to 30 .

The incidence angleo.,, between the missile axis (the boresight BS) and the flight path settles down to a small angle as the flight progresses, depending upon the curvature of the trajectory, gravity, atmospheric conditions, and the lag of the guidance control system. The basic law of the glide phase trajectory is: X + 30 = 0 so that A is controlled to -30 and any disparity produces an error signal to the autopilot. The autopilot controls the steering fin angle and thus the body angle em between the missile axis and the horizontal. A change of body angle produces a corresponding correction to the angle of look.

The terminal phase is controlled according to a different law, such as to produce basically a circular or near circular trajectory. The trajectory 1 is such a circular path to which a vertical line through the target T is tangential at the point of interception (i.e. at the target) and which is continuous with the glide path at the transition P. The circular terminal path 1 is not an ideal trajectory for several reasons. As mentioned above, angle measurements by the seeker are subject to errors which tend to increase with the look angle. In addition the missile is still accelerating (i.e. its path is curved) all the way to interception rather than being on target' for the final portion and merely requiring fine adjustment.

The control system law for the terminal phase is as follows: (X +8m+9O )k -(9o +em) = 0 where X is the measured angle of look, is the body angle, and k is the navigation constant, i.e. gain.

The first bracket is provided from the measured values of 8m and a preset 900 input. The gain of the control system, k, is chosen as 2 to produce the circular trajectory 1. The second bracket function is performed by the autopilot to produce a net error signal which is nulled to control the missile attitude and maintain the circular course. The above terminal law is a simplification in that if the missile were controlled according to it, the line of sight would bisect the angle between the vertical through the missile and the seeker boresight. Other factors such as inherent bias in the seeker, missile inertia, gravity, autopilot lag, cause variations from the above circular path law which is therefore only an approximation to the actual trajectory.

However, the more or less circular terminal path could be improved by increasing the value of k in the above guidance law. The transition to the terminal phase would be deferred and the trajectory would be sharper so increasing the incidence initially and reducing the angle of look. This would, however, only be achieved at the expense of considerable instability and is therefore not a practical solution. According to the present embodiment of the invention, a bias is introduced into the control signal path which in effect lowers the seeker boresight and reduces the look angle magnitude for ground based targets. A positive, i.e. upward bias signal is required for this purpose, equivalent to a look angle of, say, +5 , so that an actual target 5 below the physical boresight would give a zero measurement.

This bias may alternatively be seen as a variation on the pre-set 900 input determining the intercept angle. Thus the above bias of 5 may be seen as producing an intercept angle of 95 (i.e. overshooting and returning), if the bias were maintained. The above criterion for determining the onset of the terminal phase, i.e. the transition, means that in both biased and unbiased cases the actual angle of look at the transition is equal to the angle between the line of sight and the terminal intercept path. In the biased condition, however, this condition occurs later in the flight, i.e. at point P1 in Figure 1, and the curvature of the ensuing circular path is greater, i.e., with smaller radius.

The duration of the early part of the (biased) circular curve, during which the angle of look is greater and hence more troublesome, is therefore much shorter than for the unbiased case.

Removal of the bias progressively, sharpens the biased circular curve to tend to reduce overshoot'. This again tends to speed up the reduction of the angle of look both by virtue of the improved flight path and the improved body angle resulting from the bias correction.

Figure 2 shows the signal path to the autopilot Incorporating the above bias. Thus the measured angle of look A, the measured body angle 8m (derived from missile gyros in known manner), and the preset intercept angle (= 900) are sununed at 5 to provide what would be the normal input to the controller for a circular terminal path such as 1 in Figure 1. In this embodiment a controlled bias initially equivalent to 50 is added into the signal path at 7 from a bias control circuit shown in Figure 3.

The composite signal is then subject to the navigation gain k of the navigation controller following which the intercept angle S (= 90 ) is removed at 9 and the controlled bias at 15. The resulting signal 9d represents the body angle demand fed to the autopilot 11. In the autopilot the demanded body angle 8d may be compared with the measured body angle em at 13 but it is found better to subtract the flight path angle 6 which differs from the body angle by the incidence thus taking account of the incidence. The resulting difference signal ee is an error signal which is nulled by the autopilot.

It will be apparent from Figure 2 that the two 90 intercept angle inputs at 5 and 9 could be combined in a single preset input at 9 of value 90 (k -1). Similarly the controlled bias inputs at 7 and could be combined in a single input Bb(k -1) at 15. However, there may be a provision to vary k (within the limit imposed by instability) and in this case the input could not be pre-set.

Both the glide control path and the terminal control path of Figure 2 have outputs available to the autopilot. These outputs are compared, and while the body angle demanded by the terminal signal path exceeds that demanded by the glide signal path the latter takes precedence and determines the look angle at -30 . As the flight progresses, the terminal control path output signal to the autopilot decreases until, when it equals -30 , the terminal signal path of Figure 2 takes control and provides the autopilot input signal.

It will be apparent that as the terminal control signal takes effect at the glide/terminal transition, the bias signal 8b imediately takes effect. Figure 3 is a flow chart illustrating how It is removed progressively, so causing the intercept angle to revert to the design figure of 900 (or other intercept angle).

A comparator 17 which determines the dominant circuit of the glide signal path and the (Fig. 2) terminal signal path, indicates the phase 19,'glide'or'terminal In the glide phase the bias signal 8b is constant and 6b = 0 (21). Function 23 then sets the bias at the initial value, here 50, plus the integral of 8b* The bias thus remains at 50 throughout the glide phase and, although applied at 7 and 15 in Figure 2, remains ineffective because the glide phase control is dominant.

In the terminal phase a check (25) is made on the time lapsed in that phase. If the flight has more than 9 seconds to go before impact, i.e. early in the terminal phase, the function 27 is initiated and a rate of decline of 0b of 0*26b is specified. This (negative) value is integrated in function 23 and used to decrement the basic 50 value. Since 8b is specified as a function of 6b the decrement reduces with time and 8b falls exponentially. This reduction of eb continues up to the last nine seconds of the flight when the decision takes an alternative path providing reduction of the bias in accordance with an algorithm 29. After 9 seconds of the terminal phase the imposed bias will have been largely removed and any remaining bias will be, at least partially, inherent bias resulting from inaccurate angle of look measurement, etc. The algorithm 29 provides a reduction of the net bias.

According to this algorithm the remaining bias 8b (which is a combination of the 50 and the accumulated correction so far) is quantised very coarsely by the function 29 as follows. When 6b is greater than 0.5 , 8b takes the value -0.4 per second; when 8b is less than -0.5 , 8b takes the value +0.4 per second; and when eb falls back through zero from either direction, eb takes the value zero. The quantised value 8b is then integrated to give the correcting signal for application to Figure 2. It will be appreciated that the above quantisatlon figures are chosen merely by way of example and may be varied to suit the circumstances.

The bias decrementing system employed prior to the last 9 seconds of the terminal phase can be maintained If desired throughout the whole terminal phase so reducing the bias towards zero quite smoothly. In this case the decision 25 is deleted from Figure 3 and the terminal' output from 19 is applied direct to the function 27.

In either case it may be seen that the invention provides an improved shape of trajectory which sweeps quickly through the difficult region of high angle of look making use of incidence angle control at the same time. The effect of a high navigation constant is achieved, i.e. passing through a high acceleration region in good time before intercept so leaving the path straighter at the end, but without the basic increase in instability which is inherent in a high gain system.

Claims

CLAIMS

1. A missile control system for a missile adapted to descend upon a stationary or near stationary target, the system including a radar seeker for controlling the flight-path in dependence, in elevation, upon the angle of look between the seeker boresight and the missile-target line of sight, the flight path comprising a glide phase and a terminal phase and means for initiating a transition between the glide phase and the terminal phase in dependence upon the line of sight angle in space in relation to a signal dependent upon said angle of look, wherein a bias signal is incorporated with the angle of look signal so as to tend to reduce the effective angle of look in relation to said line of sight angle and thus delay the onset of said transition, the system further including means for reducing said bias signal to zero during said terminal phase.

2. A system according to Claim 1, wherein said transition is determined by a comparison between, on one hand, the angle between the line of sight and the terminal intercept path, and, on the other, the biased angle of look.

3. A system according to Claim 1 or Claim 2, wherein said bias signal is maintained constant during the glide phase and is reduced exponentially during the terminal phase.

4. A system according to any preceding claim, subject to inherent bias in measurement of said angle of look and including means for reducing said bias signal exponentially during a first part of said terminal phase and means for reducing net bias by a repeated quantisation and integration process for the remainder of the terminal phase.

5. In a missile control system for a missile having a trajectory comprising a glide phase and a terminal phase in which the missile is adapted to descend upon a stationary or near stationary target, the system including a radar seeker for controlling the flight path in elevation in dependence upon a determined angle of look between the seeker boresight and the missile-target line of sight, a method of reducing the average value of the angle of look in the terminal phase in which the predetermined terminal intercept angle is effectively increased temporarily to sharpen the trajectory during the terminal phase.

6. A missile control system substantially as hereinbefore described with reference to the accompanying drawings.

7. A method of sharpening the trajectory of the terminal phase of a ground attack missile substantially as hereinbefore described with reference to the accompanying drawings. 2-.

Amendments to the claims have been filed as follows 1. A missile control system for a missile adapted to descend upon a stationary or near stationary target, the system including a radar seeker for controlling the flight-path in dependence, in elevation, upon the angle of look between the seeker boresight and the missile-target line of sight, the flight path comprising a glide phase and a terminal phase and means for initiating a transition between the glide phase and the terminal phase in dependence upon the line of sight angle in space in relation to a signal dependent upon said angle of look, wherein a bias signal is incorporated with the angle of look signal so as to tend to reduce the effective angle of look in relation to said line of sight angle and thus delay the onset of said transition, the system further including means for reducing said bias signal to zero during said terminal phase.

2. A system according to Claim 1, wherein said bias signal is maintained constant during the glide phase and is reduced exponentially during the terminal phase.

3. A system according to Claim 1 or Claim 2, subject to inherent bias in measurement of said angle of look and including means for reducing said bias signal exponentially during a first part of said terminal phase and means for reducing net bias by a repeated quantisation and integration process for the remainder of the terminal phase.

4. In a missile control system for a missile having a trajectory comprising a glide phase and a terminal phase in which the missile is adapted to descend upon a stationary or near stationary target, the system including a radar seeker for controlling the flight path in elevation in dependence upon a determined angle of look between the seeker boresight and the missile-target line of sight, a method of reducing the average value of the angle of look in the terminal phase, by the incorporation of a bias signal, which is reduced to zero during the terminal phase, with an angle of look signal whereby the predetermined terminal intercept angle is effectively increased temporarily to sharpen the trajectory during the terminal phase.

5. A missile control system substantially as hereinbefore described with reference to the accompanying drawings.

6. A method of sharpening the trajectory of the terminal phase of a ground attack missile substantially as hereinbefore described with reference to the accompanying drawings.