CN105987652A - Attitude angular rate estimation system and ammunition using same - Google Patents

Abstract

The invention discloses an attitude angular rate estimation system. The attitude angular rate estimation system comprises a guidance module 01, an accelerometer module 02 and a calculation module 03. The guidance module 01 comprises a steering engine used for obtaining a steering deflection angle and transmitting the steering deflection angle signal to the calculation module 03. The accelerometer module 02 comprises an accelerometer used for measuring the overload and transmitting the measured overload signal to the calculation module 03. The calculation module 03 obtains the estimation value of the attitude angular rate according to the steering deflection angle signal and the overload signal. The attitude angular rate estimation system can replace an angular rate gyroscope, lowers the cost, stabilizes the missile attitude under the various interference conditions, guarantees that the flying attitude angular deviation of a missile is within the allowed range, adjusts the flying direction of the missile, corrects the flying path and enables the missile to hit a target accurately.

Description

Attitude angular rate estimating system and apply its ammunition

Technical field

The present invention relates to missile guidance technical field, particularly relate to one and be applied to three times Missile attitude angular speed estimating system in the automatic pilot of road and evaluation method.

Background technology

In tactical missile field, automatic pilot Successful utilization was more than 50 years.Automatically The main task of pilot is to increase body to damp, stablize pneumatic gain, holding system Stability, very fast instruction respond, provide high maneuverability and ensure at arbitrary height big The robustness etc. of scope aerial mission.

Three-loop autopilot is by gesture feedback loop, attitude angular velocity feedback circuit Constituting with overload feedback loop, wherein overload loop is main feedback loop, and its effect is Accelerate body response speed.In order to improve body damping, increase body stability, logical Often design inner looping in main feedback loop, be referred to as damping inner looping by this inner looping.Resistance The feedback signal of Buddhist nun's inner looping is missile attitude angular speed, is typically surveyed by angular rate gyroscope ?.

Angular rate gyroscope is a kind of fine measuring instrument, and its structure is the most complicated, and power consumption is also The biggest.Due to limited space on guided missile, not allowing it to account for much room, this allows for Its processing is highly difficult, and processing technique requires very strict, and cost is at a relatively high.

When being fitted without angular rate gyroscope on guided missile, it is impossible to return in directly obtaining damping The feedback signal on road, it is therefore desirable to indirectly construct feedback signal；When installing on guided missile There is an angular rate gyroscope, but for some reason, such as when terminal-guided shell is launched High overload causes angular speed to damage, when causing the output of angular rate gyroscope unavailable, also Need indirectly to construct feedback signal.

Owing to guided missile is single use weapon, needing situation about equipping our troops in a large number Under, if reducing cost, improving reliability, right and wrong while ensureing control accuracy The most significant.

For above-mentioned reasons, existing three-loop autopilot has been carried out deeply by the present inventor Enter research, in order to design a kind of attitude angular rate estimating system reliably, for guided missile Design provides actual reference, shortens the guided missile lead time, reduces guided missile development cost.

Summary of the invention

In order to overcome the problems referred to above, present inventor has performed and study with keen determination, design one Plant reliable attitude angular rate estimating system, under various interference conditions, stablize guided missile Attitude, it is ensured that missile flight attitude angle deviation, in allowed band, adjusts flying of guided missile Line direction, revises flight path, makes guided missile pinpointing.

In particular it is object of the present invention to offer following aspect:

First aspect, a kind of attitude angular rate estimating system, it is characterised in that this is System includes guide module 01 processed, accelerometer module 02 and computing module 03,

Wherein, described guide module 01 processed includes steering wheel, and it is used for obtaining angle of rudder reflection, and Described angle of rudder reflection signal is transferred to computing module 03,

Described accelerometer module 02 includes accelerometer, and it is used for measuring overload, and And by measurement to overload signal be transferred to computing module 03,

Described computing module 03, is used for receiving angle of rudder reflection signal and overload signal, and root According to described angle of rudder reflection signal and overload signal, it is thus achieved that the estimated value of attitude angular rate.

Second aspect, according to system described in first aspect, it is characterised in that calculate mould Block 03 obtains the estimated value of attitude angular rate according to following formula (1),

$\stackrel{\·\·}{\widehat{\mathrm{\θ}}}=(A-GC)\stackrel{\·}{\widehat{\mathrm{\θ}}}+{\mathrm{Ga}}_y+{\mathrm{B\δ}}_z---\left(1\right)$

Wherein,Represent the estimated value of attitude angular rate,Represent estimating of attitude angular rate Calculation valueDerivative, a_yRepresent overload, δ_zRepresenting angle of rudder reflection, G represents feedback matrix,

$A=\left[\begin{array}{cc}-b_{\mathrm{\α}}& 1\\ -a_{\mathrm{\α}}& -a_{\mathrm{\ω}}\end{array}\right],B=\left[\begin{array}{c}-b_{\mathrm{\δ}}\\ -a_{\mathrm{\δ}}\end{array}\right],C=\left[\begin{array}{cc}{\mathrm{Vb}}_{\mathrm{\α}}-{\mathrm{ca}}_{\mathrm{\α}}& -{\mathrm{ca}}_{\mathrm{\ω}}\end{array}\right],$

b_αRepresent normal force coefficient, a_αRepresent static-stability moment coefficient, a_ωRepresent damping Coefficient, b_δRepresent lift of rudder coefficient, a_δRepresenting steerage rate moment coefficient, V represents ammunition Speed, c represents the distance between accelerometer module 02 and ammunition barycenter.

The third aspect, according to system described in first or second aspect, it is characterised in that Feedback matrix G is obtained by the POLE PLACEMENT USING of this system,

Preferably, feedback matrix G is obtained by following formula (2)-(4),

$\left[\begin{array}{c}\stackrel{\·}{\mathrm{\α}}\\ \stackrel{\·\·}{\mathrm{\θ}}\end{array}\right]=\left[\begin{array}{cc}-b_{\mathrm{\α}}& 1\\ -a_{\mathrm{\α}}& -a_{\mathrm{\ω}}\end{array}\right]\left[\begin{array}{c}\mathrm{\α}\\ \stackrel{\·}{\mathrm{\θ}}\end{array}\right]+\left[\begin{array}{c}-b_{\mathrm{\δ}}\\ -a_{\mathrm{\δ}}\end{array}\right]\[{\mathrm{\δ}}_z\]---\left(2\right)$

$\[a_y\]=\left[\begin{array}{cc}{\mathrm{Vb}}_{\mathrm{\α}}-{\mathrm{ca}}_{\mathrm{\α}}& -{\mathrm{ca}}_{\mathrm{\ω}}\end{array}\right]\left[\begin{array}{c}\mathrm{\α}\\ \stackrel{\·}{\mathrm{\θ}}\end{array}\right]+\[{\mathrm{Vb}}_{\mathrm{\δ}}-{\mathrm{ca}}_{\mathrm{\δ}}\]\[{\mathrm{\δ}}_z\]---\left(3\right)$

$\left[\begin{array}{c}\stackrel{\·}{\widehat{\mathrm{\α}}}\\ \stackrel{\·\·}{\widehat{\mathrm{\θ}}}\end{array}\right]=(A-GC)\left[\begin{array}{c}\widehat{\mathrm{\α}}\\ \stackrel{\·}{\widehat{\mathrm{\θ}}}\end{array}\right]+{\mathrm{Ga}}_y+{\mathrm{B\δ}}_z---\left(4\right)$

$A=\left[\begin{array}{cc}-b_{\mathrm{\α}}& 1\\ -a_{\mathrm{\α}}& -a_{\mathrm{\ω}}\end{array}\right],B=\left[\begin{array}{c}-b_{\mathrm{\δ}}\\ -a_{\mathrm{\δ}}\end{array}\right],C=\left[\begin{array}{cc}{\mathrm{Vb}}_{\mathrm{\α}}-{\mathrm{ca}}_{\mathrm{\α}}& -{\mathrm{ca}}_{\mathrm{\ω}}\end{array}\right],$

Wherein, α represents the angle of attack,Represent the first derivative of α,Represent attitude angular rate Actual value,RepresentFirst derivative,Represent the state variable of this system,Generation Table angle of attack estimated value,Represent angle of attack estimated valueDerivative.

Fourth aspect, according to system described in any one of the first to the third aspect, its feature Being, this system also includes accelerometer noise processed module, and described accelerometer is made an uproar Sonication module is for processing the noise that accelerometer causes.

5th aspect, a kind of attitude angular rate evaluation method, it is characterised in that the party Method includes:

Step 1): obtain angle of rudder reflection by guide module 01 processed, by accelerometer module 02 measures overload；

Step 2): by angle of rudder reflection δ_z, angle of attack, overload a_y, attitude angular rate true ValueWith all other aerodynamic coefficients, it is thus achieved that feedback matrix G；

Step 3): obtain attitude angle by described feedback matrix G, angle of rudder reflection and overload Rate estimation value

Wherein, all other aerodynamic coefficients described include normal force coefficient b_α, quiet surely Determine moment coefficient a_α, damped coefficient a_ω, lift of rudder coefficient b_δ, steerage rate moment coefficient a_δ, V represents ammunition speed, distance c between accelerometer module 02 and ammunition barycenter.

6th aspect, according to method described in the 5th aspect, it is characterised in that described appearance State angular speed estimated valueObtained by formula (1),

$\stackrel{\·\·}{\widehat{\mathrm{\θ}}}=(A-GC)\stackrel{\·}{\widehat{\mathrm{\θ}}}+{\mathrm{Ga}}_y+{\mathrm{B\δ}}_z---\left(1\right)$

Wherein,Represent the estimated value of attitude angular rate,Represent estimating of attitude angular rate Calculation valueDerivative, a_yRepresent overload, δ_zRepresenting angle of rudder reflection, G represents feedback matrix,

Preferably, feedback matrix G is obtained by following formula (2)-(4),

$\left[\begin{array}{c}\stackrel{\·}{\mathrm{\α}}\\ \stackrel{\·\·}{\mathrm{\θ}}\end{array}\right]=\left[\begin{array}{cc}-b_{\mathrm{\α}}& 1\\ -a_{\mathrm{\α}}& -a_{\mathrm{\ω}}\end{array}\right]\left[\begin{array}{c}\mathrm{\α}\\ \stackrel{\·}{\mathrm{\θ}}\end{array}\right]+\left[\begin{array}{c}-b_{\mathrm{\δ}}\\ -a_{\mathrm{\δ}}\end{array}\right]\[{\mathrm{\δ}}_z\]---\left(2\right)$

$\[a_y\]=\left[\begin{array}{cc}{\mathrm{Vb}}_{\mathrm{\α}}-{\mathrm{ca}}_{\mathrm{\α}}& -{\mathrm{ca}}_{\mathrm{\ω}}\end{array}\right]\left[\begin{array}{c}\mathrm{\α}\\ \stackrel{\·}{\mathrm{\θ}}\end{array}\right]+\[{\mathrm{Vb}}_{\mathrm{\δ}}-{\mathrm{ca}}_{\mathrm{\δ}}\]\[{\mathrm{\δ}}_z\]---\left(3\right)$

$\left[\begin{array}{c}\stackrel{\·}{\widehat{\mathrm{\α}}}\\ \stackrel{\·\·}{\widehat{\mathrm{\θ}}}\end{array}\right]=(A-GC)\left[\begin{array}{c}\widehat{\mathrm{\α}}\\ \stackrel{\·}{\widehat{\mathrm{\θ}}}\end{array}\right]+{\mathrm{Ga}}_y+{\mathrm{B\δ}}_z---\left(4\right)$

$A=\left[\begin{array}{cc}-b_{\mathrm{\α}}& 1\\ -a_{\mathrm{\α}}& -a_{\mathrm{\ω}}\end{array}\right],B=\left[\begin{array}{c}-b_{\mathrm{\δ}}\\ -a_{\mathrm{\δ}}\end{array}\right],C=\left[\begin{array}{cc}{\mathrm{Vb}}_{\mathrm{\α}}-{\mathrm{ca}}_{\mathrm{\α}}& -{\mathrm{ca}}_{\mathrm{\ω}}\end{array}\right],$

Wherein, α represents the angle of attack,Represent the first derivative of α,Represent attitude angular rate Actual value,RepresentFirst derivative,Represent the state variable of this system,Generation Table angle of attack estimated value,Represent angle of attack estimated valueDerivative.

7th aspect, a kind of guided munition, it is characterised in that be provided with in this ammunition Attitude angular rate estimating system as described in terms of first to the 6th.

Eighth aspect, a kind of guided munition, it is characterised in that be provided with in this ammunition Attitude angular rate estimating system as described in terms of first to the 6th, on this guided munition It is additionally provided with angular rate gyroscope.

9th aspect, according to the guided munition described in eighth aspect, it is characterised in that This ammunition is arranged just like first to the attitude angular rate estimating system described in the 6th aspect, It is additionally provided with measurement apparatus on this guided munition,

Wherein, described measurement apparatus is for obtaining the missile attitude of angular rate gyroscope output Angular speed, and judge whether described attitude angular rate can be used, when measurement apparatus judges When described attitude angular rate is unavailable, enables described attitude angular rate estimating system and substitute Angular rate gyroscope works.

Tenth aspect, according to the guided munition described in the 9th aspect, it is characterised in that The disabled situation of described attitude angular rate includes: attitude angular rate is zero；Attitude angle Speed is constant value；Attitude angular rate vary more than normal range.

The present invention is had the advantage that to include:

(1) under various interference conditions, missile attitude is stablized, it is ensured that missile flight Attitude angle deviation, in allowed band, adjusts the heading of guided missile, revises flight road Line, makes guided missile pinpointing；

(2) angular rate gyroscope can be substituted, it is not necessary to increase any hardware configuration, Reduce manufacturing cost；

(3) it is effectively improved Body Damping Characteristic.

Accompanying drawing explanation

Fig. 1 illustrates the attitude angular rate estimation according to a kind of preferred implementation of the present invention The structural representation of system.

Fig. 2 illustrates the structure of the computing module according to a kind of preferred implementation of the present invention Simplify figure.

Fig. 3 illustrates and comprises attitude angular rate according to a kind of preferred implementation of the present invention The three-loop autopilot structure chart of estimating system.

Fig. 4 illustrates the attitude angular rate estimation according to a kind of preferred implementation of the present invention The flow chart of steps of method.

Fig. 5 illustrates and passes through attitude angular rate according to a kind of preferred implementation of the present invention The attitude angle that the attitude angular rate estimated value that estimating system obtains records with angular rate gyroscope Speed actual value comparison diagram.

Fig. 6 illustrates the attitude angular rate estimation according to a kind of preferred implementation of the present invention The valuation error of system.

Fig. 7 illustrates and passes through attitude angular rate according to a kind of preferred implementation of the present invention The step response of the attitude angular rate that estimating system obtains and actual attitude angular rate Step response comparison diagram.

Fig. 8 illustrates and passes through attitude angular rate according to a kind of preferred implementation of the present invention The attitude angle that the attitude angular rate estimated value that estimating system obtains records with angular rate gyroscope Speed actual value comparison diagram.

Fig. 9 illustrates and passes through attitude angular rate according to a kind of preferred implementation of the present invention The angle of pitch that estimating system obtains and actual angle of pitch comparison diagram.

Figure 10 illustrate according to a kind of preferred implementation of the present invention by attitude angle speed The angle of rudder reflection signal that rate estimating system obtains and actual angle of rudder reflection signal contrast figure.

Figure 11 illustrate according to a kind of preferred implementation of the present invention by attitude angle speed The overload that rate estimating system obtains and actual overload comparison diagram.

Figure 12 illustrate according to a kind of preferred implementation of the present invention by attitude angle speed The trajectory tilt angle that rate estimating system obtains and actual trajectory tilt angle comparison diagram.

Figure 13 illustrate according to a kind of preferred implementation of the present invention by attitude angle speed The ballistic curve that rate estimating system obtains and actual ballistic curve comparison diagram.

Drawing reference numeral illustrates:

01-guide module

02-accelerometer module

03-computing module

Detailed description of the invention

Below by drawings and Examples, the present invention is described in more detail.By this A little explanations, the features and advantages of the invention will become more apparent from clearly.

The most special word " exemplary " means " as example, embodiment or say Bright property ".The here as any embodiment illustrated by " exemplary " should not necessarily be construed as excellent In or be better than other embodiments.Although the various aspects of embodiment shown in the drawings, But unless otherwise indicated, it is not necessary to accompanying drawing drawn to scale.

In one preferred embodiment, as it is shown in figure 1, the present invention provide appearance State angular speed estimating system, including guide module 01 processed, accelerometer module 02 and calculating Module 03,

Wherein, described guide module 01 processed includes steering wheel, and it is used for obtaining angle of rudder reflection, and And described angle of rudder reflection signal is transferred to computing module 03,

Described accelerometer module 02 includes accelerometer, and it is used for measuring overload, and And by measurement to overload signal be transferred to computing module 03,

Described computing module 03, is used for receiving angle of rudder reflection signal and overload signal, according to Described angle of rudder reflection signal and overload signal, it is thus achieved that the estimated value of attitude angular rate.

In one preferred embodiment, computing module 03 structure simplifies figure such as Fig. 2 Shown in, wherein,Represent the estimated value of attitude angular rate, a_yRepresent overload, δ_zRepresent Angle of rudder reflection, G represents feedback matrix, and the physical significance of G is attitude angle speed in the present invention The speed that rate estimated value approaches to attitude angular rate actual value.

The equation being obtained this attitude angular rate estimating system by Fig. 2 is:

$\stackrel{\·\·}{\widehat{\mathrm{\θ}}}=(A-GC)\stackrel{\·}{\widehat{\mathrm{\θ}}}+{\mathrm{Ga}}_y+{\mathrm{B\δ}}_z---\left(1\right)$

Wherein,RepresentFirst derivative.

In one preferred embodiment, the state of this attitude angular rate estimating system Equation is expressed as formula (5),

$\{\begin{array}{c}\stackrel{\·}{x}=Ax+Bu\\ y=Cx+Du\end{array}---\left(5\right)$

Wherein, x represents the state vector of this system, and u represents the input of this system, y Represent the output of this system.

In one preferred embodiment, angle of rudder reflection δ is chosen_zFor the input of this system, I.e. u=[δ_z], choose angle of attack and attitude angular rate actual valueState for this system becomes Amount, i.e.Choose overload a_yFor the output of this system, i.e. y=[a_y], should The state equation (5) of system is converted into two-dimensional state space form, it is thus achieved that formula (2) With formula (3), obtain formula (4) according to formula (1),

$\left[\begin{array}{c}\stackrel{\·}{\mathrm{\α}}\\ \stackrel{\·\·}{\mathrm{\θ}}\end{array}\right]=\left[\begin{array}{cc}-b_{\mathrm{\α}}& 1\\ -a_{\mathrm{\α}}& -a_{\mathrm{\ω}}\end{array}\right]\left[\begin{array}{c}\mathrm{\α}\\ \stackrel{\·}{\mathrm{\θ}}\end{array}\right]+\left[\begin{array}{c}-b_{\mathrm{\δ}}\\ -a_{\mathrm{\δ}}\end{array}\right]\[{\mathrm{\δ}}_z\]---\left(2\right)$

$\[a_y\]=\left[\begin{array}{cc}{\mathrm{Vb}}_{\mathrm{\α}}-{\mathrm{ca}}_{\mathrm{\α}}& -{\mathrm{ca}}_{\mathrm{\ω}}\end{array}\right]\left[\begin{array}{c}\mathrm{\α}\\ \stackrel{\·}{\mathrm{\θ}}\end{array}\right]+\[{\mathrm{Vb}}_{\mathrm{\δ}}-{\mathrm{ca}}_{\mathrm{\δ}}\]\[{\mathrm{\δ}}_z\]---\left(3\right)$

$\left[\begin{array}{c}\stackrel{\·}{\widehat{\mathrm{\α}}}\\ \stackrel{\·\·}{\widehat{\mathrm{\θ}}}\end{array}\right]=(A-GC)\left[\begin{array}{c}\widehat{\mathrm{\α}}\\ \stackrel{\·}{\widehat{\mathrm{\θ}}}\end{array}\right]+{\mathrm{Ga}}_y+{\mathrm{B\δ}}_z---\left(4\right)$

$A=\left[\begin{array}{cc}-b_{\mathrm{\α}}& 1\\ -a_{\mathrm{\α}}& -a_{\mathrm{\ω}}\end{array}\right],B=\left[\begin{array}{c}-b_{\mathrm{\δ}}\\ -a_{\mathrm{\δ}}\end{array}\right],C=\left[\begin{array}{cc}{\mathrm{Vb}}_{\mathrm{\α}}-{\mathrm{ca}}_{\mathrm{\α}}& -{\mathrm{ca}}_{\mathrm{\ω}}\end{array}\right],$

Wherein,Represent angle of attack estimated value,Represent angle of attack estimated valueDerivative,WithIt is intermediate variable, can be eliminated during calculating.V represents ammunition speed, c Represent the distance between accelerometer module 02 and ammunition barycenter, before ammunition barycenter For just, being negative after ammunition barycenter, with missile coordinate system as reference, the body longitudinal axis Pointing to bullet direction is front, and deviating from bullet direction is rear, b_αRepresent normal force system Number, a_αRepresent static-stability moment coefficient, a_ωRepresent damped coefficient, b_δRepresent lift of rudder Coefficient, a_δRepresent steerage rate moment coefficient, when calculating the aerodynamic coefficient of guided missile, Conventional coordinate system includes missile coordinate system, velocity coordinate system, earth axes and bullet Road coordinate system.

In one preferred embodiment, a characteristic point of this system is chosen, i.e. When ammunition speed V value is 606 meter per second, every aerodynamic derivative at this speed It is as shown in table 1 below,

Table 1

Every aerodynamic derivative in table 2 is substituted in formula (2)-(4), it is thus achieved that under Formula (5),

$\begin{array}{c}A-GC=\left[\begin{array}{cc}-1.07& 1\\ -168.3& -5.93\end{array}\right]-\left[\begin{array}{c}{G}_1\\ G_2\end{array}\right]\left[\begin{array}{cc}537.42& -3.9138\end{array}\right]\\ =\left[\begin{array}{cc}-1.07-537.342G_1& 1+3.9138G_1\\ -168.3-537.342G_2& -5.93+3.91138G_2\end{array}\right]\end{array}---\left(5\right)$

Therefore, the proper polynomial of this attitude angular rate estimating system is:

$\begin{array}{c}f\left(\mathrm{\λ}\right)=\det \[\mathrm{\λ}I-(A-GC)\]\\ =\det \left[\begin{array}{cc}\mathrm{\λ}-(-1.07-537.342G_1)& -(1+3.9138G_1)\\ -(-168.3-537.342G_2)& \mathrm{\λ}-(-5.93+3.9138G_2)\end{array}\right]\end{array}---\left(6\right)$

Wherein, the eigenvalue of λ representative formula (5), I represents unit matrix.

In one preferred embodiment, the limit of this attitude angular rate estimating system For 1～5 times of three-loop autopilot dominant pole, preferably 2～5 times.

In all of closed-loop pole of three-loop autopilot system, the distance imaginary axis is Near and around without the limit of closed-loop zero, and remaining limit is away from the imaginary axis, then away from The response component corresponding to limit that the imaginary axis is nearest plays a leading role in system responds, Such closed-loop pole is referred to as three-loop autopilot dominant pole.

In one preferred embodiment, the limit of this attitude angular rate estimating system For [λ_p,λ_p], then the desired character multinomial of this system is:

$f_p\left(\mathrm{\λ}\right)=\det \left[\begin{array}{cc}\mathrm{\λ}-{\mathrm{\λ}}_p& 0\\ 0& \mathrm{\λ}-{\mathrm{\λ}}_p\end{array}\right]=(\mathrm{\λ}-{\mathrm{\λ}}_p)(\mathrm{\λ}-{\mathrm{\λ}}_p)---\left(7\right)$

In one preferred embodiment, the proper polynomial of this system equal to this is The desired character multinomial of system, i.e. formula (8), and then obtain feedback matrix G.

F (λ)=f_p(λ) (8)

In one preferred embodiment, the feedback matrix G obtained is substituted into formula (1) In, it is thus achieved that the attitude angular rate estimated value of this system estimation

In one preferred embodiment, when the pole of this attitude angular rate estimating system 1 times of the dominant pole that point is three-loop autopilot, i.e. [λ_p λ_p]= Time [-2.25-2.25], obtain feedback matrix by formula (8)

In one preferred embodiment, when the pole of this attitude angular rate estimating system 4 times of the dominant pole that point is three-loop autopilot, i.e. [λ_p λ_p]= Time [-10-10], obtain feedback matrix by formula (8)

In one preferred embodiment, the estimating system of this attitude angular rate is comprised The structure of three-loop autopilot as it is shown on figure 3, wherein, antihunt signal is estimated System is attitude angular rate estimating system of the present invention, a_ycRepresent guided missile acceleration Instruction, k_DCRepresent pilot closed loop gain regulation coefficient, k_ARepresent driver-automobile closed-loop system ginseng Number,Representative represents body transmission function, k_gRepresent angular rate gyroscope gain, c (s) Represent the differential of distance between accelerometer and barycenter, ω_IRepresent driver-automobile closed-loop system ginseng Number, a_yc、k_DC、k_A、 C (s) and ω_IIt is the intrinsic of three-loop autopilot Parameter, is all pre-stored on guided missile, and above-mentioned parameter can be according to the type of guided missile Modify with the task performed by guided missile and set.

In one preferred embodiment, this attitude angular rate estimating system also includes Seeker processing module and accelerometer noise module,

Wherein, seeker processing module is for making an uproar of processing that missile homer causes Sound, accelerometer noise processed module is for processing the noise that accelerometer causes, excellent Choosing, described seeker processing module and accelerometer noise processed module are Wave filter, seeker processing module is used for filtering the noise that missile homer causes, The noise that accelerometer noise processed module accelerometer causes, and then improve this system Precision.

The attitude angular rate estimating system provided based on the present invention, by as shown in Figure 4 Method obtain attitude angular rate estimated value.

A kind of guided munition provided according to the present invention, is provided with attitude angle in this ammunition Rate estimation system, this ammunition does not has angular rate gyroscope, therefore the volume of this ammunition Less, and this ammunition obtains ammunition in real time by described attitude angular rate estimating system Attitude angular rate, and then adjust guided missile heading, revise flight path, make Guided missile pinpointing, wherein, according to type and the task configuration attitude angle of guided missile The limit of rate estimation system, and then obtain feedback matrix G.

Arrange just like attitude angle according in this ammunition of a kind of guided munition that the present invention provides Rate estimation system, this guided munition is additionally provided with angular rate gyroscope and measurement apparatus,

Wherein, described measurement apparatus is for obtaining the missile attitude of angular rate gyroscope output Angular speed, and judge whether described attitude angular rate can be used, when measurement apparatus judges When described attitude angular rate is unavailable, enables described attitude angular rate estimating system and substitute Angular rate gyroscope works.

The disabled situation of described attitude angular rate includes: attitude angular rate is zero；Appearance State angular speed is constant value；Attitude angular rate vary more than normal range, wherein said Refer to that the attitude angular rate recorded is between-25deg/s to+25deg/s normal range.

In one preferred embodiment, can by this attitude angular rate estimating system To obtain the angle of pitch of guided missile, angle of rudder reflection, overload, trajectory tilt angle and ballistic curve.

Embodiment

When not considering seeker and accelerometer noise, by Matlab software Contrast attitude angular rate estimating system provided by the present invention estimation attitude angular rate with The impact that three-loop autopilot is produced by the attitude angular rate that angular rate gyroscope is measured, Analyze the effect of this attitude angular rate estimating system, comparing result such as Fig. 5, Fig. 6 and Tu Shown in 7, wherein, the valuation error in Fig. 6 is deducted by the estimated value of attitude angular rate The actual value of attitude angular rate obtains, and, from above-mentioned comparing result, attitude The limit of angular speed estimating system can be taken as three-loop autopilot dominant pole 4 times, and the attitude angular rate estimated by attitude angular rate estimating system and angle speed The effect of the attitude angular rate of rate gyro to measure is basically identical, and therefore this appearance is described State angular speed estimating system can be effectively improved Body Damping Characteristic, meets body work and wants Ask.

When considering seeker and accelerometer noise, imitated by Matlab software The attitude angular rate of true attitude angular rate estimating system estimation and angular rate gyroscope are measured Attitude angular rate, as shown in figures 8-13, the real system shown in figure is simulation result It is provided with the three-loop autopilot of angular rate gyroscope, and, above-mentioned emulation tie Fruit understands, and by the attitude angular rate curve of this attitude angular rate estimating system acquisition, bows Elevation angle curve, angle of rudder reflection curve, overload curves, trajectory tilt angle curve and relative trajectory Curve is bent with the attitude angular rate of the three-loop autopilot being provided with angular rate gyroscope respectively Line, angle of pitch curve, angle of rudder reflection curve, overload curves, trajectory tilt angle curve and phase Consistent to ballistic curve, therefore illustrate that this attitude angular rate estimating system can be effectively improved Body Damping Characteristic, meets body job requirement.

Above in association with preferred embodiment describing the present invention, but this A little embodiments are only exemplary, only play illustrative effect.On this basis, The present invention can be carried out multiple replacement and improvement, these each fall within the protection of the present invention In the range of.

Claims (10)

1. an attitude angular rate estimating system, it is characterised in that this system includes system Guide module (01), accelerometer module (02) and computing module (03),

Wherein, described guide module processed (01) includes steering wheel, and it is used for obtaining angle of rudder reflection, And described angle of rudder reflection signal is transferred to computing module (03),

Described accelerometer module (02) includes accelerometer, and it is used for measuring overload, And by measurement to overload signal be transferred to computing module (03),

Described computing module (03) is used for receiving angle of rudder reflection signal and overload signal, and The estimated value of attitude angular rate is obtained according to described angle of rudder reflection signal and overload signal.

System the most according to claim 1, it is characterised in that computing module (03) The estimated value of attitude angular rate is obtained according to following formula (1),

$\stackrel{\·\·}{\widehat{\mathrm{\θ}}}=(A-GC)\stackrel{\·}{\widehat{\mathrm{\θ}}}+{\mathrm{Ga}}_y+{\mathrm{B\δ}}_z---\left(1\right)$

Wherein,Represent the estimated value of attitude angular rate,Represent estimating of attitude angular rate Calculation valueDerivative, a_yRepresent overload, δ_zRepresenting angle of rudder reflection, G represents feedback matrix,C=[Vb_α-ca_α -ca_ω],

b_αRepresent normal force coefficient, a_αRepresent static-stability moment coefficient, a_ωRepresent damping Coefficient, b_δRepresent lift of rudder coefficient, a_δRepresenting steerage rate moment coefficient, V represents ammunition Speed, c represents the distance between accelerometer module (02) and ammunition barycenter.

System the most according to claim 1 or claim 2, it is characterised in that feedback matrix G Obtained by the POLE PLACEMENT USING of this system,

Preferably, feedback matrix G is obtained by following formula (2)-(4),

$\left[\begin{array}{c}\stackrel{\·}{\mathrm{\α}}\\ \stackrel{\·\·}{\mathrm{\θ}}\end{array}\right]=\left[\begin{array}{cc}-b_{\mathrm{\α}}& 1\\ -a_{\mathrm{\α}}& -a_{\mathrm{\ω}}\end{array}\right]\left[\begin{array}{c}\mathrm{\α}\\ \stackrel{\·}{\mathrm{\θ}}\end{array}\right]+\left[\begin{array}{c}-b_{\mathrm{\δ}}\\ -a_{\mathrm{\δ}}\end{array}\right]\[{\mathrm{\δ}}_z\]---\left(2\right)$

$\[a_y\]=\[{\mathrm{Vb}}_{\mathrm{\α}}-{\mathrm{ca}}_{\mathrm{\α}}-{\mathrm{ca}}_{\mathrm{\ω}}\]\left[\begin{array}{c}\mathrm{\α}\\ \stackrel{\·}{\mathrm{\θ}}\end{array}\right]+\[{\mathrm{Vb}}_{\mathrm{\δ}}-{\mathrm{ca}}_{\mathrm{\δ}}\]\[{\mathrm{\δ}}_z\]---\left(3\right)$

$\left[\begin{array}{c}\stackrel{\·}{\widehat{\mathrm{\α}}}\\ \stackrel{\·\·}{\widehat{\mathrm{\θ}}}\end{array}\right]=(A-GC)\left[\begin{array}{c}\widehat{\mathrm{\α}}\\ \stackrel{\·}{\widehat{\mathrm{\θ}}}\end{array}\right]+{\mathrm{Ga}}_y+{\mathrm{B\δ}}_z---\left(4\right)$

C=[Vb_α-ca_α -ca_ω],

Wherein, α represents the angle of attack,Represent the first derivative of α,Represent attitude angular rate Actual value,RepresentFirst derivative,Represent the state variable of this system,Generation Table angle of attack estimated value,Represent angle of attack estimated valueDerivative.

4. according to system described in any one of claim 1-3, it is characterised in that this is System also includes accelerometer noise processed module, described accelerometer noise processed module For processing the noise that accelerometer causes.

5. an attitude angular rate evaluation method, it is characterised in that the method includes:

Step 1): obtain angle of rudder reflection by guide module processed (01), pass through accelerometer Module (02) measures overload；

Step 2): by angle of rudder reflection signal δ_z, angle of attack, overload signal a_y, attitude angle Speed actual valueWith all other aerodynamic coefficients, it is thus achieved that feedback matrix G；

Step 3): obtain attitude angle by described feedback matrix G, angle of rudder reflection and overload Rate estimation value

Wherein, all other aerodynamic coefficients described include normal force coefficient b_α, quiet surely Determine moment coefficient a_α, damped coefficient a_ω, lift of rudder coefficient b_δ, steerage rate moment coefficient a_δ, Ammunition speed V, distance c between accelerometer module (02) and ammunition barycenter.

Method the most according to claim 5, it is characterised in that described attitude angle speed Rate estimated valueObtained by formula (1),

$\stackrel{\·\·}{\widehat{\mathrm{\θ}}}=(A-GC)\stackrel{\·}{\widehat{\mathrm{\θ}}}+{\mathrm{Ga}}_y+{\mathrm{B\δ}}_z---\left(1\right)$

Wherein,Represent the estimated value of attitude angular rate,Represent estimating of attitude angular rate Calculation valueDerivative, a_yRepresent overload, δ_zRepresenting angle of rudder reflection, G represents feedback matrix,

Preferably, feedback matrix G is obtained by following formula (2)-(4),

$\left[\begin{array}{c}\stackrel{\·}{\mathrm{\α}}\\ \stackrel{\·\·}{\mathrm{\θ}}\end{array}\right]=\left[\begin{array}{cc}-b_{\mathrm{\α}}& 1\\ -a_{\mathrm{\α}}& -a_{\mathrm{\ω}}\end{array}\right]\left[\begin{array}{c}\mathrm{\α}\\ \stackrel{\·}{\mathrm{\θ}}\end{array}\right]+\left[\begin{array}{c}-b_{\mathrm{\δ}}\\ -a_{\mathrm{\δ}}\end{array}\right]\[{\mathrm{\δ}}_z\]---\left(2\right)$

$\[a_y\]=\[\begin{array}{cc}{\mathrm{Vb}}_{\mathrm{\α}}-{\mathrm{ca}}_{\mathrm{\α}}& -{\mathrm{ca}}_{\mathrm{\ω}}\end{array}\]\left[\begin{array}{c}\mathrm{\α}\\ \stackrel{\·}{\mathrm{\θ}}\end{array}\right]+\[{\mathrm{Vb}}_{\mathrm{\δ}}-{\mathrm{ca}}_{\mathrm{\δ}}\]\[{\mathrm{\δ}}_z\]---\left(3\right)$

$\left[\begin{array}{c}\stackrel{\·}{\widehat{\mathrm{\α}}}\\ \stackrel{\·\·}{\widehat{\mathrm{\θ}}}\end{array}\right]=(A-GC)\left[\begin{array}{c}\widehat{\mathrm{\α}}\\ \stackrel{\·}{\widehat{\mathrm{\θ}}}\end{array}\right]+{\mathrm{Ga}}_y+{\mathrm{B\δ}}_z---\left(4\right)$

C=[Vb_α-ca_α -ca_ω],

Wherein, α represents the angle of attack,Represent the first derivative of α,Represent attitude angular rate Actual value,RepresentFirst derivative,Represent the state variable of this system,Generation Table angle of attack estimated value,Represent angle of attack estimated valueDerivative.

7. a guided munition, it is characterised in that arrange in this ammunition and want just like right Seek the attitude angular rate estimating system described in 1-6.

8. a guided munition, it is characterised in that arrange in this ammunition and want just like right Seek the attitude angular rate estimating system described in 1-6, this guided munition is additionally provided with angle speed Rate gyro.

Guided munition the most according to claim 8, it is characterised in that this guidance Measurement apparatus it is additionally provided with on ammunition,

Wherein, described measurement apparatus is for obtaining the missile attitude of angular rate gyroscope output Angular speed, and judge whether described attitude angular rate can be used, when measurement apparatus judges When described attitude angular rate is unavailable, enables described attitude angular rate estimating system and substitute Angular rate gyroscope works.

Guided munition the most according to claim 9, it is characterised in that described appearance The disabled situation of state angular speed includes: attitude angular rate is zero；Attitude angular rate is Constant value；Attitude angular rate vary more than normal range.