GB2433368A

GB2433368A - Radar system

Info

Publication number: GB2433368A
Application number: GB8802859A
Authority: GB
Inventors: Anthony John Benson; Frederick Lawrence Pew
Original assignee: Marconi Co Ltd; Alenia Marconi Systems Ltd
Current assignee: BAE Systems Integrated System Technologies Ltd
Priority date: 1987-02-09
Filing date: 1988-02-09
Publication date: 2007-06-20
Anticipated expiration: 2008-02-09
Also published as: FR2903193A1; IT8867958A0; GB8802859D0; GB2433368B

Abstract

A missile seeker system for use with air-launched missiles against stationary ground targets. Tracking is effected in response to the 'look angle' ( c L) between missile axis and sightline. Bias in the measurements can produce a false look angle and consequently a displaced flight path. A signal <EMI ID=2.1 HE=8 WI=8 LX=1436 LY=918 TI=UI> <PC>is derived which is responsive to path errors, e.g. a sightline rate, and a correction signal ( c c) derived from this for adding in to the measured look angle. The correction signal is devised from the deviation from a predetermined path i.e. in azimuth, flight along the sightline and in elevation a two-stage path comprising a glide phase of constant look angle and a terminal phase comprising a circular transition to a vertical path to the target.

Description

Radar System This invention relates to a radar system and particularly

to a homing missile having a radar seeker primarily for use against stationary targets. In the case of a radar installation target the seeker need only be a passive seeker responding to the transmissions of the target radar but the origin of the radar signals is not relevant to the invention.

A coninon problem in seekers is the presence of bias to which a number of factors may contribute, e.g. distortion of the line of sight caused by radome inqerfections. The effect of such bias is that the difference characteristic, produced by the subtraction of spaced or angled antenna element signals has a null point which is off the target sightline and thus the seeker effectively squints' at the target and produces a steering control signal which tends to steer the missile off the required course. It will thus follow a bent' course where it should properly be straight, always aiming a few degrees (according to the magnitude of the bias) off the target.

This bias problem is particularly significant in modern systems using improved guidance laws and seeker distortion correction algorithms since the improvements provided by such techniques can be significantly reduced in the presence of bias.

It is an object of the present invention to remove or at least lessen this restriction on the efficiency of guidance law improvements and seeker distortion correctors.

According to the present invention, In a radar seeker system for a missile in which a measure of the target angle off the missile roll axis is derived and the missile steered in response to this angle and in which bias in the system produces a discrepancy between the true and measured values of said angle, a correction signal is incorporated with the measured value of said angle, the correction signal being derived from an estimated value of a course-indicating signal.

In a system for use with a missile adapted to descend upon a stationary target, the course-indicating signal having a zero value when the missile is correctly on course in the azimuth plane, the system preferably includes means for deriving an azimuth angle correction component from the course-indicating signal for incorporation with the azimuth component of the target angle.

The course-indicating signal may be a target azimuth sightline rate signal.

The means for deriving an azimuth angle correction component may comprise a proportional/integral controller adapted to operate on the sightline rate signal.

The system preferably includes means for estimating the value of the target angle in dependence upon the course-indicating signal, means for deriving an error signal from a comparison of the measured and estimated target angle signals, means for quantising a residual error signal to predetermined values, means for integrating the quantised value to produce a slowly changing correction signal for incorporation with the target angle, and means for combining the correction signal from said error signal to produce the residual error signal. The quantising means may produce one predetermined positive value, one predetermined negative value and one zero value in dependence upon the residual error signal being excessively positive, excessively negative, and of magnitude within predetermined positive and negative limits, respectively.

The target angle off the missile roll axis is preferably the target look angle and the system preferably includes means for estimating the target look angle from the sightline rate, the time-to-go and the missile incidence angle between roll axis and flight path.

One embodiment of a radar system incorporating bias correction in accordance with the invention will now be described, by way of example, with reference to the accompanying drawings, of which:-Figure 1 is a diagramatic view in elevation of a missile M on course for a radar target 1; Figure 2 is a block diagram of a bias correction scheme for an early phase of a missile's flight; and Figure 3 is a block diagram of a bias correction scheme for a terminal phase of the missile flight.

Referring to Figure 1 of the drawings, this shows a view in elevation of a missile M, which may be launched from an aircraft, on a flight path F with a requirement to descend vertically on a stationary target 1. While a vertical descent is one known requirement, other steep angles of descent, typically down to 300 off vertical, may be specified.

The various parameters involved are as follows: = body elevation angle (i.e. between horizontal space reference and missile roll axis) 0< = incidence angle (between the missile roll axis and the flight path F). This angle is sometimes known as the angle of attack.

<p≥ flight path elevation angle = true sightline angle (relative to horizontal space reference) = lead angle, i.e. angle between the flight path F and the line of sight S: = -65 = missile velocity along the flight path 81.. (elevation), (azimuth) = look angle, i.e. the angle between the antenna boresight, here the roll axis, and the target sight line: = 8d + O( (in elevation).

With no transverse forces on the missile the incidence angle should be zero. However, in azimuth a persistent incidence angle may result from crosswinds and in elevation a controlled incidence angle will be necessary to counter the effect of gravity.

In the early part of the flight, the glide phase', the missile flies on an economical' cruise path so called. In elevation the lead angle is controlled to a fixed predetermined angle, say 30 , so that the missile climbs at this angle and then levels out. No attempt is made during this phase to correct for bias in elevation.

Periodic computations are made during the glide phase to determine the lead angle that would be necessary at the current missile position for a circular terminal descent ending on a vertical path. Initially this lead angle is found to exceed the 30 to which the missile has been controlled hitherto and no action is taken. As the flight progresses however, the calculated lead angle will decrease and at 30 (at the point P say,) a smooth transition is permitted to the terminal phase having a circular path. The lead angle may, of course, be calculated in respect of other terminal path forms and other final descent angles.

During the glide phase the missile is controlled to null the azimuth sightline rate, which is taken as a suitable signal indicating adherence to the course, so that the missile flies directly along the azimuth line of sight. As will be explained, bias correction is employed in the azimuth plane in this phase.

Once into the terminal phase the azimuth bias correction is continued, as in Figure 2, and an elevation bias correction scheme as in Figure 3 is initiated.

Clearly, bias in the determination of the target sight line will produce faulty control of the flight in both its stages to the extent that the final descent angle may be significantly off the vertical and in addition the miss distance may be increased unacceptably.

The correction scheme for the glide phase will now be described with reference to Figure 2. In this phase, as mentioned previously, control is exercised only in azimuth, the requirement being to maintain the flight path in a vertical plane through the missile and the target.

Figure 2 represents both equipment and physical conditions or constraints. The homing guidance loop 2 operates in response to the measured azimuth look angle /S corrected by a signal Y" produced in accordance with the invention. This loop 2 produces missile steering signals which control the fin deflection. The result of fin deflection is dependent upon the seeker and missile dynamic characteristics 4. The missile body attitude in space, tVm' responds very rapidly to fin deflection, this feedback path being referenced 6.

The response of the missile body position in space is represented by the slow feedback path 8 and the kinematics 10 of the system. The sightline azimuth angle 1p5 will respond with the body position and is thus derived by way of the kinematics feedback path. This sightline angle is the true sightline angle by definition.

The body angle signal Y1m on feedback path 6 is subtracted from the sightline angle at 10 and it may be seen from the analogous situation in elevation (Figure 1) that the result is the azimuth look angle, which is therefore the true azimuth look angle This true look angle is then corrupted by various factors before it appears as the measured look angle 1fr. A scaling error 12 is imposed. A bias error is added at 13. A cross coupling error is added at 14. All of these contribute to the measured look angle 1/f applied to suming circuit 16 to be incremented by a controlled offset signal i.e. a correction signafl7. The output is then the A1 corrected, or modified look angle %y.

In order to make this correction a signal is employed which responds to deviation of the missile from the collision course. In this example it is the sightline rate signal which provides the necessary indication of deviation. Since, in the azimuth plane the missile should be flying along the sightline, any deviation will produce an angular velocity of the sightline, i.e. a finite sightline rate. An estimate of this signal is easily available in known manner from the seeker electronics and missile instruments 18 and is applied to a controller 20 for the purposes of the invention. The controller comprises a proportional/integral network and has the characteristic = fKi +

--J

where P is the output correction signal, K1 is the proportional constant, K2 is the integral constant, S is the Laplace operator ( d/dt) and is the sightline rate estimate.

In the present case the antenna is assumed to be strapped down, the look angle being determined from the antenna signal outputs.

In an alternative system where the antenna is gimbal mounted and stabilised the gimbal angles output from angle pick-offs may be corrected in response to the sightline rate.

The above bias correction system is employed for azimuth control throughout the flight.

The correction scheme employed in elevation for the terminal phase of the missile flight will be described with reference to Figure 3.

In the elevation plane, i.e. as shown in Figure 1, the measured look angle 6, with its in-built bias, is derived as before.

This is applied to a sunining circuit 16 for the addition of a correction signal.

Before deriving the correction signal it should be noted that, whereas in the glide phase the requirement was to null the sightline rate (in azimuth), in this terminal phase it should not be nulled in elevation ininediately since clearly the vertical descent angle would be sacrificed and in addition instability would be likely to arise on suddenly demanding a course down the line of sight. It is therefore required that the sightline rate should be constant (in the circular path case), and the lead angle should reduce at a limited and gradual rate. This is achieved by means of a very loosely coupled closed loop which almost gives open loop control,as will be explained.

Reverting to Figure 3, the measured look angle eL is compared, in a difference circuit 22, with an estimate of the look angle 6 which may be derived in various ways. For example, the sightline rate may be derived from the achieved flight-path. Its value may be seen from the geometry of the situation. Thus the component of the flight path velocity transverse to the sightline.

is approximately Vm.ed. This gives the sightline rate as VmOd/R, where R is the range.

Thus the estimated sightline rate = = and 6d = TF5 where T is the time to go', i.e. until impact.

Now, from Figure 1 the look angle is equal to the lead angle plus the incidence angle, and thus where o.is the elevation incidence angle.

This processing is performed by block 24 on the basis of the estimated factors 1, and. The time to go is calculated from the estimated range and known missile velocity.

An error signal E. is thus produced, this being limited to 6 degrees, or comparable value, by limiter 26. A filter 28 removes noise and rapid variations to produce a smoothed error signal which initiates the correction signal.

A residual error signal ee is formed from the sum (30) of the filtered error signal and the correction signal 8c which is to be added to the measured look angle. The residual error signal ee is subjected to a quantising process 32 which consists of producing the following output values of = -0.4 degrees/sec when ee)o.s degrees = 0 when 8e changes sign 8c = +0.4 degrees/sec when 8e< -0.5 degrees.

An integator 34 accumulates this output value and produces 6c the required correction signal, which is added to the measured look angle at 36 to produce the modified look angle 8. Thus the process 32 limits the rate of correction of the look angle, as required. This corrected signal is then employed as in Figure 2.

The output signal from process 32 is of course a quantised or step signal, the step values of which (0.4 degrees/sec) are chosen to produce a suitable rate of change of the correction signal 6c and corresponding slow reduction of the sightline rate.

The terminal system as described above is directed to the elevation plane but the same scheme is applied to the azimuth plane.

Again at this phase of the flight instability would be likely to arise by attempting to null the sightline rate directly and hence the open ended loop technique is again employed.

Claims

CLAIMS

1. A radar seeker system for a missile in which a measure of the target angle of f the missile roll axis is derived and the missile steered in response to this angle and in which bias in the system produces a discrepancy between the true and measured values of said angle, wherein a correction signal is incorporated with the measured value of said angle, said correction signal being derived from an estimated value of a course-indicating signal.

2. A system according to Claim 1, for use with a missile adapted to descend upon a stationary target and wherein said course-indicating signal has a zero value when the missile is correctly on course in the azimuth plane, the system including means for deriving an azimuth angle correction component from said course-indicating signal for incorporation with the azimuth component of said target angle.

3. A system according to Claim 2, wherein said course-indicating signal is a target azimuth sightline rate signal.

4. A system according to Claim 3, wherein said means for deriving an azimuth angle correction component comprises a proportional/integral controller adapted to operate on said sightline rate signal.

5. A system according to Claim 1, including means for estimating the value of said target angle in dependence upon said course-indicating signal, means for deriving an error signal from a comparison of the measured and estimated target angle signals, means for quantising a residual error signal to predetermined values, means for integrating the quantised values to produce a slowly changing correction signal for incorporation with said target angle, and means for combining the correction signal with said error signal to produce said residual error signal.

6. A system according to Claim 5 wherein said quantising means produces one predetermined positive value, one predetermined negative value and one zero value in dependence upon said residual error signal being excessively positive, excessively negative and of magnitude within predetermined positive and negative limits, respectively.

7. A system according to Claim 5 or Claim 6, wherein said course-indicating signal is a target sightilne rate signal.

8. A system according to Claim 7, wherein said target angle of f the missile roll axis is the target look angle and including means for estimating the target look angle from the sightline rate, the time-to-go and the missile incidence angle between roll axis and flight path.

9. A system according to any of Claims 5 to 8, for use in elevation control in the terminal phase of a missile flight, said terminal phase having a flight path of predetermined form incident upon a stationary target at a predetermined angle and means for initiating said terminal phase when the lead angle between flight path and sightline becomes equal to a predetermined value maintained for a glide phase preceding said terminal phase.

10. A system according to Claim 9, wherein azimuth control in said glide phase and in said terminal phase employs a said course-indicating signal having a zero value when the missile is correctly on course in the azinvth plane, the system including means for deriving an azimuth angle correction ocmponent from said course-indicating signal for incorporation with the azinuth component of said target angle.

11. A system according to any of Claims 1 to 10 wherein said measure of target angle off the missile roll axis is the look angle derived from the antenna characteristics of a strapped down antenna.

12. A system according to any of Claims 1 to 10, wherein said measure of target angle off the missile roll axis is determined from the antenna boresight of a stabilised gimbal-mounted antenna.

13. A radar seeker system substantially as hereinbefore described with reference to Figures 1 and 2 of the accompanying drawings.

14. A radar seeker system substantially as hereinbefore described with reference to Figures 1 and 3 of the accompanying drawings. I'

Amendments to the claims have been filed as follows 1. A radar seeker system for a missile in which a measure of the target angle off the missile roll axis is derived using radar signals and the missile steered in response to this angle and in which bias in the system produces a discrepancy between the true and measured values of said angle, wherein a correction signal is incorporated with the measured value of said angle, said correction signal being derived from an estimated value of a course-indicating signal representative of deviation of the missile from a predetermined flight path.

2. A system according to Claim 1, for use with a missile adapted to descend upon a stationary target and wherein said course-indicating signal has a zero value when the missile is correctly on course in the azimuth plane, the system including means for deriving an azimuth angle correction component from said course-indicating signal for incorporation with the azimuth component of said target angle.

3. A system according to Claim 2, wherein said course-indicating signal is a target azimuth sightline rate signal.

4. A system according to Claim 3, wherein said means for deriving an azimuth angle correction component comprises a proportional/integral controller adapted to operate on said sightline rate signal.

5. A system according to Claim 1, including means for estimating the value of said target angle in dependence upon said course-indicating signal, means for deriving an error signal from a comparison of the measured and estimated target angle signals, means for quantising a residual error signal to predetermined values, means for integrating the quantised values to produce a slowly changing correction signal for incorporation with said target angle, and means for combining the correction signal with said error signal to produce said residual error signal.

6. A system according to Claim 5 wherein said quantising means produces one predetermined positive value, one predetermined negative value and one zero value in dependence upon said residual error signal being excessively positive, excessively negative and of magnitude within predetermined positive and negative limits, respectively. 12.

7. A system according to Claim 5 or Claim 6, wherein said course-indicating signal Is a target sightline rate signal.

8. A system according to Claim 7, wherein said target angle of f the missile roll axis Is the target look angle and including means for estimating the target look angle from the sightilne rate, the time-to-go and the missile incidence angle between roll axis and flight path.

9. A system according to any of Claims 5 to 8, for use In elevation control in the terminal phase of a missile flight, said terminal phase having a flight path of predetermined form incident upon a stationary target at a predetermined angle and means for initiating said terminal phase when the lead angle between flight path and sightline becomes equal to a predetermined value maintained for a glide phase preceding said terminal phase.

10. A system according to Claim 9, wherein azimuth control in said glide phase and In said terminal phase employs a said course-indicating signal having a zero value when the missile Is correctly on course in the azimuth plane, the system including means for deriving an azimuth angle correction ocmponent from said course-indicating signal for incorporation with the azimuth component of said target angle.

11. A system according to any of Claims 1 to 10 wherein said measure of target angle off the missile roll axis is the look angle derived from the antenna characteristics of a strapped down antenna.

12. A system according to any of Claims 1 to 10, wherein said measure of target angle off the missile roll axis is determined from the antenna boresight of a stabihised gimbal-mounted antenna.

13. A radar seeker system substantially as hereinbefore described with reference to Figures 1 and 2 of the accompanying drawings.

14. A radar seeker system substantially as hereinbefore described with reference to Figures 1 and 3 of the accompanying drawings.