CN102878872A - Guidance information processing method aiming at seeker loss-of-lock conditions - Google Patents

Abstract

The invention discloses a guidance information processing method aiming at seeker loss-of-lock conditions. The guidance information processing method comprises the following steps of: calculating a theoretical line-of-sight rate according to the position and speed information of a guided missile, and employing different switching methods between the actually measured line-of-sight rate and the theoretical line-of-sight rate under the condition that the seeker loses a target according to different distances between the guided missile and the target; and calculating a theoretical frame angle according to the position and attitude angle information of the guided missile, and presetting the frame angle of the seeker under the condition that the seeker loses the target.

Description

A kind of guidance information processing method for target seeker losing lock situation

Technical field

Low-cost tactical missile zero-miss guidance Control System Design.

Background technology

At present, the high-performance of guided missile is that high-performance by target seeker, inertial navigation equipment and rudder system realizes basically.Wherein, target seeker has determined the performance of guided missile to a great extent, but for guided missile cheaply, can only remedy to a certain extent by advanced person's guidance control algolithm the deficiency of seeker performance.These deficiencies be mainly reflected in that line of sight rate measures that noise is large, the probability of target seeker lose objects is large and body between the aspects such as isolation is low.If the guidance control algolithm can not in addition to a certain degree the improvement of above-mentioned unfavorable factor, then can finally affect the accuracy at target of guided missile.

Advanced guidance control algolithm should be able to compensate the deficiency of target seeker system hardware performance to a certain extent, such as the noise that reduces line of sight rate, improve recapture target after the target seeker losing lock probability, improve the accuracy at target of guided missile in the target seeker lose objects situation etc.At present, the method that reduces the line of sight rate noise normally designs low pass filter, and the measured value of line of sight rate is carried out filtering, and in the situation of the omnidistance lock onto target of target seeker, conventional filtering method is ripe.But for the guided missile that adopts the modes such as laser guidance, when guided missile passed through cloud layer or target release smoke screen, the phenomenon of target seeker lose objects was inevitable, in this class situation, how to utilize existing information to guarantee the miss distance of guided missile, there is no effective method can be for reference.

Summary of the invention

Technology of the present invention is dealt with problems: in order to overcome the problem that existing low-cost tactical missile can't hit in the situation of target seeker lose objects, a kind of guidance information processing method for target seeker losing lock situation is provided, and the method can reduce miss distance in the situation of target seeker lose objects.

Technical solution of the present invention: a kind of guidance information processing method for target seeker losing lock situation may further comprise the steps:

(1) utilizes angular speed and the overload of rate gyroscope and Jia Biao measuring missile;

(2) utilize the angular speed of actual measurement and overload to carry out strapdown and resolve, obtain attitude angle, speed and the positional information of guided missile;

(3) according to above-mentioned missile velocity and the theoretical line of sight rate of positional information calculation; According to position and the attitude angle information of guided missile, theory of computation frame corners;

(4) when the target seeker losing lock, the processing of guidance information comprises line of sight rate and frame corners information processing; Line of sight rate is processed needs at first to judge that whether the guided missile residue flight time is greater than default time T, if greater than, then target seeker is in remote out-of-lock condition, the theoretical line of sight rate that calculates with step (3) replaces the actual measurement line of sight rate to carry out filtering, otherwise, target seeker is in closely out-of-lock condition, with the input as filtering of the actual measurement line of sight rate filtering output value of target seeker losing lock previous moment; In the two states processing procedure, when relocking target, with the input of actual measurement line of sight rate as filtering; The frame corners information processing is to adopt theoretical frame angle that step (3) calculates that the frame corners of target seeker is preset the enterprising line search in frame corners basis that is presetting.

The present invention's beneficial effect compared with prior art is:

(1) the present invention can for Guidance Law provides available guidance information in the situation of target seeker losing lock, overcome routine techniques for the helpless defective of target seeker losing lock phenomenon effectively.

(2) HWIL simulation data show that this programme can in the situation of target seeker long distance losing lock and long distance recapture, guarantee missile-target impact; Can in the situation of target seeker low coverage losing lock, guarantee miss distance is decreased in acceptable scope.

Description of drawings

Fig. 1 is flow chart of the present invention;

Fig. 2 is the trajectory correlation curve in the remote losing lock situation;

Fig. 3 is the trajectory correlation curve in the losing lock situation closely;

Fig. 4 is the partial enlarged drawing of Fig. 3 Trajectory-terminal.

The specific embodiment

Below in conjunction with accompanying drawing the present invention is elaborated, as shown in Figure 1, step of the present invention is as follows:

(1) utilizes angular speed and the overload of rate gyroscope and Jia Biao measuring missile;

(2) utilize the angular speed of actual measurement and overload to carry out strapdown and resolve, obtain attitude angle, speed and the positional information of guided missile; Specific algorithm is as follows:

$\left[\begin{array}{c}{n}_{\mathrm{xg}}\\ n_{\mathrm{yg}}\\ n_{\mathrm{zg}}\end{array}\right]=\left[\begin{array}{ccc}1& 0& 0\\ 0& \cos \mathrm{\γ}& \sin \mathrm{\γ}\\ 0& -\sin \mathrm{\γ}& \cos \mathrm{\γ}\end{array}\right]\left[\begin{array}{ccc}\cos \mathrm{\θ}& \sin \mathrm{\θ}& 0\\ -\sin \mathrm{\θ}& \cos \mathrm{\θ}& 0\\ 0& 0& 1\end{array}\right]\left[\begin{array}{ccc}\cos \mathrm{\ψ}& 0& -\sin \mathrm{\ψ}\\ 0& 1& 0\\ \sin \mathrm{\ψ}& 0& \cos \mathrm{\ψ}\end{array}\right]\left[\begin{array}{c}{n}_{\mathrm{xb}}\\ n_{\mathrm{yb}}\\ n_{\mathrm{zb}}\end{array}\right]$

Wherein, n _Xg, n _Yg, n _ZgBe respectively the guided missile three axles overload under the earth axes; ψ, θ, γ are respectively guided missile yaw angle, the angle of pitch and roll angle; n _Xb, n _Yb, n _ZbBe respectively the guided missile three axles overload under the missile coordinate system.

$\left[\begin{array}{c}{a}_{\mathrm{xg}}\\ b_{\mathrm{yg}}\\ a_{\mathrm{zg}}\end{array}\right]=\left[\begin{array}{ccc}\mathrm{<figure-callout\; id="0"\; label="g"\; filenames="HSA00000760101000031.png"\; state="\{\{state\}\}">g</figure-callout>}& 0& 0\\ 0& \mathrm{<figure-callout\; id="0"\; label="g"\; filenames="HSA00000760101000031.png"\; state="\{\{state\}\}">g</figure-callout>}& 0\\ 0& 0& g\end{array}\right](\left[\begin{array}{c}{n}_{\mathrm{xg}}\\ n_{\mathrm{yg}}\\ n_{\mathrm{zg}}\end{array}\right]+\left[\begin{array}{c}0\\ 1\\ 0\end{array}\right])$

Wherein, a _Xg, a _Yg, a _ZgBe respectively the guided missile 3-axis acceleration under the earth axes; G is acceleration of gravity.

$\left\{\begin{array}{c}{V}_{\mathrm{xg}}=V_{\mathrm{<figure-callout\; id="0"\; label="xg"\; filenames="HSA00000760101000031.png"\; state="\{\{state\}\}">xg</figure-callout>}0}+{\&Integral;}_0^t{a}_{\mathrm{xg}}\mathrm{dt}\\ V_{\mathrm{yg}}=V_{\mathrm{yg}0}+{\&Integral;}_0^t{a}_{\mathrm{yg}}\mathrm{dt}\\ V_{\mathrm{zg}}=V_{\mathrm{<figure-callout\; id="0"\; label="zg"\; filenames="HSA00000760101000031.png"\; state="\{\{state\}\}">zg</figure-callout>}0}+{\&Integral;}_0^t{a}_{\mathrm{zg}}\mathrm{dt}\end{array}\right.$

Wherein, V _Xg, V _Yg, V _ZgBe respectively the guided missile three axle speed under the earth axes; V _Xg0, V _Yg0, V _Zg0Be respectively the guided missile three axle initial velocities under the earth axes; T is missile flight time.

$\left\{\begin{array}{c}{x}_g=x_{\mathrm{<figure-callout\; id="0"\; label="g"\; filenames="HSA00000760101000031.png"\; state="\{\{state\}\}">g</figure-callout>}0}+{\&Integral;}_0^t{V}_{\mathrm{xg}}\mathrm{dt}\\ y_g={\mathrm{<figure-callout\; id="0"\; label="y\; g"\; filenames="HSA00000760101000031.png"\; state="\{\{state\}\}">y</figure-callout>}}_g+{\&Integral;}_0^t{V}_{\mathrm{yg}}\mathrm{dt}\\ z_g={\mathrm{<figure-callout\; id="0"\; label="z\; g"\; filenames="HSA00000760101000031.png"\; state="\{\{state\}\}">z</figure-callout>}}_g+{\&Integral;}_0^t{V}_{\mathrm{zg}}\mathrm{dt}\end{array}\right.$

Wherein, x _g, y _g, z _gBe respectively guided missile three shaft positions under the earth axes; x _G0, y _G0, z _G0Be respectively the guided missile three axle initial positions under the earth axes.

$\left\{\begin{array}{c}\mathrm{\&psi;}={\mathrm{\&psi;}}_0+{\&Integral;}_0^t\frac{{\mathrm{\&omega;}}_y\cos \mathrm{\&gamma;}-{\mathrm{\&omega;}}_z\sin \mathrm{\&gamma;}}{\cos \mathrm{\&theta;}}\mathrm{dt}\\ \mathrm{\&theta;}={\mathrm{\&theta;}}_0+{\&Integral;}_0^t({\mathrm{\&omega;}}_y\sin \mathrm{\&gamma;}+{\mathrm{\&omega;}}_z\cos \mathrm{\&gamma;})\mathrm{dt}\\ \mathrm{\&gamma;}={\mathrm{\&gamma;}}_0+{\&Integral;}_0^t({\mathrm{\&omega;}}_x-\tan \mathrm{\&theta;}({\mathrm{\&omega;}}_y\cos \mathrm{\&gamma;}-{\mathrm{\&omega;}}_z\sin \mathrm{\&gamma;}))\mathrm{dt}\end{array}\right.$

Wherein, ψ ₀, θ ₀, γ ₀Be respectively the initial yaw angle of guided missile, the angle of pitch and roll angle; ω _x, ω _y, ω _zBe respectively body three-axis attitude angular speed.

(3) according to above-mentioned speed and the theoretical line of sight rate of positional information calculation; According to position and the attitude angle information of guided missile, theory of computation frame corners;

$\left\{\begin{array}{c}{q}_{\mathrm{\&epsiv;}}=\arctan \left(\frac{y_g}{x_g}\right)\\ q_{\mathrm{\&beta;}}=\arcsin \left(\frac{z_g}{\sqrt{x_g^2+y_g^2+z_g^2}}\right)\end{array}\right.$

Wherein, q _ε, q _βBe respectively the theoretical pitching angle of sight and the theoretical driftage angle of sight.

$\left\{\begin{array}{c}{\stackrel{\&CenterDot;}{q}}_{\mathrm{\&epsiv;}}=\frac{V_{\mathrm{xg}}{y}_g-x_g{V}_{\mathrm{yg}}}{x_g^2+y_g^2}\\ {\stackrel{\&CenterDot;}{q}}_{\mathrm{\&beta;}}=\frac{1}{\sqrt{x_g^2+y_g^2}}\frac{(x_g^2+y_g^2)V_{\mathrm{zg}}-z_g(x_g{V}_{\mathrm{xg}}+y_g{V}_{\mathrm{yg}})}{x_g^2+y_g^2+z_g^2}\end{array}\right.$

Wherein,

Be respectively theoretical pitching line of sight rate and theoretical driftage line of sight rate.

$G=\left[\begin{array}{ccc}{\cos q}_{\mathrm{\&beta;}}& 0& {-\sin q}_{\mathrm{\&beta;}}\\ 0& 1& 0\\ {\sin q}_{\mathrm{\&beta;}}& 0& {\cos q}_{\mathrm{\&beta;}}\end{array}\right]\left[\begin{array}{ccc}{\cos q}_{\mathrm{\&epsiv;}}& {\sin q}_{\mathrm{\&epsiv;}}& 0\\ -{\sin q}_{\mathrm{\&epsiv;}}& {\cos q}_{\mathrm{\&epsiv;}}& 0\\ 0& 0& 1\end{array}\right]\left[\begin{array}{ccc}\cos \mathrm{\&psi;}& 0& \sin \mathrm{\&psi;}\\ 0& 1& 0\\ -\sin \mathrm{\&psi;}& 0& \cos \mathrm{\&psi;}\end{array}\right]\left[\begin{array}{ccc}\cos \mathrm{\&theta;}& -\sin \mathrm{\&theta;}& 0\\ \sin \mathrm{\&theta;}& \cos \mathrm{\&theta;}& 0\\ 0& 0& 1\end{array}\right]\left[\begin{array}{ccc}1& 0& 0\\ 0& \cos \mathrm{\&gamma;}& -\sin \mathrm{\&gamma;}\\ 0& \sin \mathrm{\&gamma;}& \cos \mathrm{\&gamma;}\end{array}\right]$

Wherein, G is for being tied to the coordinate conversion matrix of LOS coordinate system by missile body coordinate.

$\left\{\begin{array}{c}{\mathrm{\&theta;}}_g=57.3\&times;\arctan \left(\frac{G_{12}}{G_{11}}\right)\\ {\mathrm{\&psi;}}_g=57.3\&times;\arcsin (-G_{13})\end{array}\right.$

Wherein, θ _g, ψ _gBe respectively theoretical pitching frame angle and theoretical driftage frame corners; G ₁₁Be the 1st row the 1st column element among the matrix G; G ₁₂Be the 1st row the 2nd column element among the matrix G; G ₁₃Be the 1st row the 3rd column element among the matrix G.

(4) when the target seeker losing lock, the processing of guidance information comprises line of sight rate and frame corners information processing; Line of sight rate is processed needs at first to judge that whether the guided missile residue flight time is greater than default time T, if greater than, then target seeker is in remote out-of-lock condition, the theoretical line of sight rate that calculates with step (3) replaces the actual measurement line of sight rate to carry out filtering, otherwise, target seeker is in closely out-of-lock condition, with the input as filtering of the actual measurement line of sight rate filtering output value of target seeker losing lock previous moment; In the two states processing procedure, when relocking target, with the input of actual measurement line of sight rate as filtering; The frame corners information processing is to adopt theoretical frame angle that step (3) calculates that the frame corners of target seeker is preset, the enterprising line search in the frame corners that presets basis methods such as (can adopt carry out) rectangular searches, with the probability that improves the recapture target or the time that shortens the recapture target, thereby reduce miss distance.

The computing formula of above-mentioned default time T is

T＝K·T _b

Wherein, K is minimum control rigidity, T _bBe the body time constant.

Application example is as follows:

In the situation of target seeker long distance losing lock (9s after the emission～12s losing lock, the residue flight time is greater than 8s), trajectory contrasts as shown in Figure 2, wherein solid line is the trajectory that uses this programme to draw, miss distance is 0.29m, the trajectory of dotted line for not using this programme to draw, and miss distance is 91m.As seen, in the situation of target seeker long distance losing lock, use this programme obviously to reduce miss distance, can direct hit on the target.

(18s after the emission～21s losing lock in the situation of target seeker low coverage losing lock, the residue flight time is less than 8s), the trajectory contrast as shown in Figure 3, amplify as shown in Figure 4 the part of Trajectory-terminal, wherein solid line is the trajectory that uses this programme to draw, miss distance is 0.63m, the trajectory of dotted line for not using this programme to draw, and miss distance is 8.7m.As seen, in the situation of target seeker low coverage losing lock, use this programme obviously to reduce miss distance, can direct hit on the target.

The unspecified part of the present invention belongs to general knowledge as well known to those skilled in the art.

Claims (1)

1. guidance information processing method for target seeker losing lock situation is characterized in that may further comprise the steps:

(1) utilizes angular speed and the overload of rate gyroscope and Jia Biao measuring missile;

(2) utilize the angular speed of actual measurement and overload to carry out strapdown and resolve, obtain attitude angle, speed and the positional information of guided missile;

(3) according to above-mentioned missile velocity and the theoretical line of sight rate of positional information calculation; According to position and the attitude angle information of guided missile, theory of computation frame corners;

(4) when the target seeker losing lock, the processing of guidance information comprises line of sight rate and frame corners information processing; Line of sight rate is processed needs at first to judge that whether the guided missile residue flight time is greater than default time T, if greater than, then target seeker is in remote out-of-lock condition, the theoretical line of sight rate that calculates with step (3) replaces the actual measurement line of sight rate to carry out filtering, otherwise, target seeker is in closely out-of-lock condition, with the input as filtering of the actual measurement line of sight rate filtering output value of target seeker losing lock previous moment; In the two states processing procedure, when relocking target, with the input of actual measurement line of sight rate as filtering; The frame corners information processing is to adopt theoretical frame angle that step (3) calculates that the frame corners of target seeker is preset the enterprising line search in frame corners basis that is presetting.

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