DE3831587B3 - Radar seeker system for use with air launched missile, derives correction signal from estimated value of target azimuth sightline rate signal representing deviation of missile from preset flight path - Google Patents

Abstract

A controller derives a measure of target angle off the missile roll axis using radar signals such that the missile is steered based on the target angle. The bias in the system, produces a discrepancy between the true and measured values of the angle. A correction signal is incorporated with the measured value of the angle such that the correction signal is derived from an estimated value of target azimuth sightline rate signal representing deviation of the missile from predetermined flight path.

Description

The This invention relates to a radar system, and more particularly a missile serving a target flight with a radar seeker, and a missile, the first opposite stationary Goals is used. If the target is a radar installation, needs the seeker just to be a passive seeker who is on the broadcast of the target radar responds. However, the origin of the radar signals is not relevant for The invention.

A common one Problem in viewfinders is the presence of a one-sided error or a transfer to which a number of circumstances may contribute, for example one by imperfections Distortion of the sight or sight caused in the radar hood line. The effect of such displacement is that the difference characteristic, the spaced by subtraction of signals or angled Antenna elements is caused to have a zero point next to the one leading to the goal Visor or line of sight, so that the viewfinder effectively to the target "switches" and a steering control signal generated, which has the tendency of the missile of the required To divert course. The missile therefore follows a "curved" course where this Course should be appropriately straightforward in itself, so that always (dependent on of the size of the transfer) a few degrees next to the target is targeted.

This Dislocation problem is particularly important in modern system which improved steering laws and search distortion correction algorithms use as the improvements provided by such techniques be significantly reduced in the presence of displacement can.

It One object of the invention is this limitation in steering law improvements and eliminate or at least minimize search distortion corrections.

According to the invention is used in a radar search system for a missile, in which system once for that of the missile roll axis derived deviating target angle and the missile as a function of this angle and dislocation existing in the system is a discrepancy between the true and measured value of this angle, a correction signal included in the measured value of this angle, this correction signal is derived from an estimated Value for a course-indicating signal.

In a system for use with a missile that is prone to sinking a stationary destination is suitable, the course indicating signal has a zero value when the missile in the azimuth plane is right on course, the system preferably Means for deriving an azimuth angle correction component from the course indicating signal for inclusion in the azimuth component of the target angle.

The the course-indicating signal may be a destination azimuth line speed signal be.

The Means for deriving an azimuth angle correction component may be Proportional / integral control unit included for processing the line of sight velocity signal suitable is.

The System contains preferably means for estimating the value of the target angle as a function of the course indicating Signal, means for deriving an error signal from a comparison between the measured and estimated target angle signal, mean for quantizing a residual error signal to predetermined values, Means for integrating the quantized value to produce a slowly changing Correction signal for inclusion in the target angle and mean for uniting the correction signal with the error signal for obtaining the residual error signal. The quantizing means can be one predetermined positive value, a predetermined negative value and generate a zero value, depending on whether the residual error signal is excessively positive, excessively negative or of an amount that is within a predetermined positive and negative limits.

Of the from the missile roll axis deviating or on the missile roll axis Preferred target angles are preferably the viewing angle, and the system contains preferably means for estimating the viewpoint from the line of sight speed, the remaining one Flight time and the missile angle between the roll axis and the trajectory.

One embodiment a radar system with a dislocation correction according to the invention in the following by way of example with reference to the accompanying drawings in which:

1 is a schematic view in the elevation plane of a missile M, which is on course to a radar target T,

2 is a block diagram of an offset correction scheme for an early phase of a flight of the missile, and

3 FIG. 4 is a block diagram of an offset correction scheme for a final phase of missile flight. FIG.

1 of the drawings shows in the elevation or elevation plane a missile M which can be launched or released from an aircraft and which is on a trajectory F, with the requirement to descend vertically to a stationary target T. Although vertical descent is a known requirement, other steep descent angles, typically up to 30 ° from vertical, may also be prescribed.

• θ_m = Flugkörperelevationswinkel (d.h. der Winkel zwischen der horizontalen Raumreferenz und der Flugkörperrollachse).
• α = Anstellwinkel (Winkel zwischen der Flugkörperrollachse und der Flugbahn F). Dieser Winkel ist auch als Anströmwinkel bekannt.
• θ_f = Flugbahnelevationswinkel.
• θ_s = wahrer Visier- oder Sichtlinienwinkel (bezogen auf die horizontale Raumreferenz).
• θ_d = Vorhaltwinkel, d.h. Winkel zwischen der Flugbahn F und der Visier- oder Sichtlinie S: = θ_f – θ_s.
• V_m = Flugkörpergeschwindigkeit längs der Flugbahn.
• θ_L (Elevation), Ψ_L (Azimut) = Blickwinkel, d.h. der Winkel zwischen der Antennenziellinie, entspricht hier der Rollachse, und der zum Ziel führenden Sichtlinie: = θ_d + α (in der Elevationsebene).

The various parameters to consider are the following:

• θ _m = missile elevation angle (ie the angle between the horizontal space reference and the missile roll axis).
• α = angle of attack (angle between the missile roll axis and the trajectory F). This angle is also known as the angle of attack.
• θ _f = trajectory elevation angle.
• θ _s = true sight or sight line angle (relative to the horizontal space reference).
• θ _d = lead angle, ie angle between the trajectory F and the sight or sight line S: = θ _f - θ _s .
• V _m = missile velocity along the trajectory.
• θ _L (elevation), Ψ _L (azimuth) = angle of view, ie the angle between the antenna finish line, here corresponds to the roll axis, and the sight line leading to the target: = θ _d + α (in the elevation plane).

To grab on the missile no lateral forces on, would the Angle of attack be zero. In the azimuth direction, however, can a persistent angle of attack of cross winds, and in the elevation direction is then a controlled angle of attack required to counteract the effect of gravity.

in the early Part of the flight, the "gliding phase", the missile flies up an "economic", so-called march track. In the elevation direction of the lead angle to a fixed controlled predetermined angle, for example 30 °, so that the missile at this angle rises and then leveled out. During this phase will be no Attempted to correct in the elevation direction across from To carry out a transfer. While In this sliding phase periodic calculations are carried out to determine the lead angle at the current missile position is required to end up in a circular sinking, the ends in a vertical path. Initially, this lead angle exceeds on the missile previously controlled, a value of 30 °, and it is influenced by one Refrained. However, with the progress of the flight is increasing the calculated lead angle, and at 30 ° (for example, at the point P), becomes a smooth transition in the final phase with circular Railway allowed. The lead angle can of course with respect to other Endbahnformen and other final Absinkwinkel be calculated.

During the Sliding phase is the missile in such a way controlled that the Azimuth line speed is adjusted to zero, the used as a suitable signal to indicate compliance with the course so that the missile directly along the Azimuth Line of Sight flies. As will be explained, finds in this Phase in the azimuth plane offset correction.

Upon reaching the final phase, the azimuth offset correction continues as it is 2 and an elevation offset correction scheme will follow 3 initiated.

at The determination of the target sightline produces a shift in clearer Way a faulty flight control during both flight levels in to an extent that the final Absorbing angle considerably can deviate from the vertical and beyond the missing distance increased in an unacceptable manner can be.

The correction scheme for the sliding phase will be described below 2 described. At this stage, as noted above, control is exercised only in the azimuth direction, under the condition of keeping the trajectory in a vertical plane passing through the missile and the target.

2 represents both the equipment and the physical conditions or limits. The target flight steering loop 2 operates in response to the measured azimuth viewing angle Ψ ^ _L which is corrected by a signal Ψ _C generated according to the invention. This loop 2 generates missile steering signals that control the fins deflection. The result of the fin deflection depends on the viewfinder and missile dynamics 4 or the dynamic characteristics of the viewfinder and missile. The position of the missile in space, ie Ψ _m , responds very quickly to the fins deflection, with the associated return path with the reference number 6 is provided. The response of the missile position in space is through a slow return path 8th represented, and with 10 is called the kinematics of the system. The sight line azimuth angle Ψ _S responds to the missile position and is thus derived by means of the kinematic feedback path. This line of sight angle Ψ _S is by definition the true line of sight angle.

The missile angle signal Ψ _m on the return path 6 is the line of sight angle at the summation or subtraction point 10 subtracted, and in analogy to the situation in the elevation plane ( 1 ) one can recognize that the result of the Azimuth viewing angle, which is therefore the true azimuth viewing angle Ψ _L.

.This true point of view is distorted by various circumstances before appearing as the measured point of view Ψ ^ _L. A scaling error 12 is imposed. An offset error is added 13 added. A crosstalk error is added 14 added. All these circumstances contribute to the measured angle of view Ψ ^ _L , that of a summing circuit 16 is supplied to be incremented by a controlled offset or shift signal, ie from a correction signal Ψ _C. The output of 16 is then the corrected or modified angle

,

worin

Ψ_C

das Ausgangskorrektursignal,

K₁

die Proportionalkonstante,

K₂

die Integralkonstante,

S

der Laplaceoperator (= ^ d/dt) und

r ~_s

die Sichtliniengeschwindigkeitsschätzung ist.

To perform this correction, a signal is used that responds to the deviation of the missile from the collision course. In this example, it is the line-of-sight speed signal that provides the necessary indication of the deviation. Since in the azimuth plane the missile should fly along the line of sight, any deviation will cause an angular velocity of the line of sight, ie a finite line of sight velocity. An estimate of this signal is readily available in a manner known per se, from finder electronics and missile instrumentation 18 , and is for the purposes of the invention a controller 20 fed. The control unit 20 contains a proportional / integral network and has the characteristic

wherein

Ψ _C

the output correction signal,

K ₁

the proportional constant,

K ₂

the integral constant,

S

the Laplace operator (= ^ d / dt) and

r ~ _s

the line-of-sight speed estimation is.

in the In the case considered, it is assumed that the antenna is missile-proof is and that the Viewing angle is determined from the antenna signal outputs. at an alternative system in which the antenna is gimbaled and is stabilized the gimbal angle outputs from the angle sensors dependent on be corrected by the line of sight speed.

The The above offset correction system is applied to the azimuth control while of the entire flight.

The correction scheme used in the elevation plane during the final phase of missile flight will be described with reference to FIG 3 described.

In the elevation plane, ie in the in 1 The measured angle θ ^ _{L is derived} together with the inherent displacement as before. He then arrives at a summing circuit 36 in which a correction signal is added.

In front It should be noted in the derivation of the correction signal that, although in the glide phase the requirement was the line of sight speed (in the azimuth plane) to zero, in this final phase in the elevation level does not immediately make a null match should be because the vertical Absinkwinkel would be sacrificed and in addition the sudden Request of a longitudinal line-of-sight rate likely to cause instability would. It is therefore required that the Line of sight speed should be constant (in the case of the circular path) and that yourself the lead angle should change at a limited and gradual rate. This is achieved with the help of an extremely loose coupled closed loop, which is almost an open loop control results as explained becomes.

How out 3 can be seen, the measured angle of view θ ^ _L in a differential circuit 22 compared with an estimate of the viewing angle θ ~ _L , which estimated viewing angles can be derived in a variety of ways. The line of sight speed can be derived, for example, from the achieved trajectory. The relevant value can be found in the geometry of the situation. Thus, the component of the trajectory velocity transverse to the line of sight is about V _m · θ _d . Thus, for line of sight velocity, V _m θ _d / R, where R is the distance.

The estimated line-of-sight rate or line-of-sight speed is thus r ~ _s = V _m θ _d / R = θ _d / T
and θ ~ _d = T × r ~ _{s ,}
where T is the remaining flight time, ie until the impact.

Out 1 shows that the viewing angle is equal to the lead angle plus the angle of attack, so that: θ ~ L = T · r ~ s + α, where α is the elevation angle of attack.

This processing is thus from the block 24 based on the estimated factors T ~, r ~ _s and α ~. The remaining flight time is calculated from the estimated distance and the known missile speed.

Thus, an error signal ε generated by means of a limiter 26 is limited to a value of ± 6 ° or a comparable value. A filter 28 Removes noise and rapid changes, giving a smoothed error signal ε is obtained, which initiates the Korrek tursignal.

A residual error signal θ _e is formed from the sum (at 30 ) of the filtered error signal ε and the correction signal θ _C , which is added to the measured viewing angle. The residual error signal θ _e is at 32 a quantization process consisting of generating the following output values of θ ∙ _C :
θ ∙ _C = -0.4 degrees / second when θ _e > 0.5 degrees
θ ∙ _C = 0, if θ _e changes the sign
θ ∙ _C = +0.4 degrees / second when θ _e <-0.5 degrees.

An integrator 34 accumulates the output value θ ∙ _C and provides the required correction signal θ _C , which at 36 is added to the measured angle, to produce the modified viewing angle θ ^ _L. The process 32 thus limits the speed of correction of the viewing angle, as required. This corrected signal is then applied as in 2 shown.

The output signal θ ∙ _{C of} the process 32 is, of course, a quantized or step signal with its increments (0.4 degrees / second) chosen to obtain a suitable rate of change of the correction signal θ _C and a corresponding slow reduction in line-of-sight speed.

The end system described above is applied to the elevation plane, however, it is also applied to the azimuth plane in the same way. In this phase of the flight is in turn the probability of Occurrence of instability given when trying to direct the line of sight speed directly to zero. Consequently, in turn, the open-ended looping technique applied.

Claims (12)

A radar search system for a missile, in which system a measure of the target roll deviating from the missile roll axis is derived using radar signals and the missile is controlled in dependence on said angle, and a bias present in the system is a discrepancy between the true and measured value of that angle characterized in that a correction signal is derived from an estimated value of a heading-indicating signal representing the deviation of the missile from a predetermined trajectory, and this correction signal is included in the measured value of this angle. The system of claim 1, for use with a Missiles the sinking to a stationary one Target is suitable, wherein the course-indicating signal has a zero value has, if the missile in the azimuth plane right on course is which system means to Deriving an azimuth angle correction component from the cursor indicating Includes signal for inclusion in the azimuth component of the target angle. A system according to claim 2, wherein the course indicating Signal is a target azimuth line speed signal. A system according to claim 3, in which the means for Deriving an azimuth angle correction component a proportional / integral control included for acting on the line of sight velocity signal suitable is. A system according to claim 1, including means for estimating the Value of the target angle depending on from the course indicating signal, means for deriving an error signal from a comparison between the measured and estimated target angle signal, Means for quantizing a residual error signal to a predetermined one Values, means for integrating the quantized values for generating a slowly changing one Correction signal for inclusion in the target angle, and means for combining the correction signal with the error signal for generation a residual error signal. A system according to claim 5, wherein the quantizing means a predetermined positive value, a predetermined negative value Value and a zero value depending on of which generate whether the residual error signal in a surpassing Measure positive, in a surpassing Measure negative or of a size, which is within predetermined positive and negative limits. A system according to claim 5 or claim 6, in which the course indicating signal is a destination sighting speed signal is. A system according to claim 7, in which the of the missile roll axis deviating target angle is the sighting angle and which means to appreciate the Sighting angle from the line of sight speed, the remaining one Flight time and the missile angle between the roll axis and the trajectory. A system according to any one of claims 5 to 8, for use in elevation control in the final phase of a missile flight, wherein the final phase has a trajectory of predetermined shape which strikes a stationary target at a predetermined angle, including means for introducing the tail phase, when the lead angle between the trajectory and the line of sight becomes equal to a predetermined value, which is maintained during a preceding phase preceding the final phase. A system according to claim 9, wherein the azimuth control while the sliding phase and during the Final phase a named course-indicating signal with a zero value applies when the missile in the azimuth plane right on course is which system means to Deriving an azimuth angle correction component from that indicia Signal for inclusion in the azimuth component of the target angle contains. A system according to any one of claims 1 to 10, in which the measure of from the missile roll axis deviating target angle is the angle of view of the antenna characteristics a missile-resistant Antenna is derived. A system according to any one of claims 1 to 10, in which the measure of from the missile roll axis deviating target angle from the antenna aiming line of a stabilized, gimbaled antenna is derived.