CN105371852A - Optimal gyroscope combination selection method based on residual error generators - Google Patents

Abstract

The invention discloses an optimal gyroscope combination selection method based on residual error generators. The method comprises steps as follows: (1), four gyroscopic instruments of four standby gyroscopic assemblies including three upright gyroscopic assemblies and one oblique gyroscopic assembly are divided into four groups of gyroscope combinations; (2), the three-axis angular speed of each satellite is calculated according to measurement values of the four groups of gyroscopic combinations according to mounting matrixes of the standby gyroscopic assemblies, and three-axis angular speeds of the four satellites are obtained; (3), the three-axis attitude provided by a star sensor and the three-axis angular speeds, calculated by the four groups of gyroscopic combinations, of the satellites are used for forming the residual error generators, and four attitude residual errors are obtained totally; (4), an evaluation function of each attitude residual error is solved, the four evaluation functions are sequenced according to numerical magnitudes, and the smallest group is the optimal gyroscope combination. The gyroscope combination with the optimal performance is dynamically selected in orbit in real time, redundancy of the gyroscopic assemblies on the satellite is effectively used, angular speed information with higher accuracy can be provided for an attitude control system, and the method has certain fault-tolerant capacity for gyroscopic faults.

Description

A kind of optimum Gyro choosing method based on Residual Generation device

Technical field

The invention belongs to attitude of satellite determination technical field, relate to a kind of online preferred method of gyroscope instrument, especially relate to a kind of method choosing three best gyros participation angular velocity calculation of performance when alternative gyrounit exists hardware redundancy.

Background technology

Gyro is used for the inertia angular velocity of responsive satellite, is inertial measurement unit main in Satellite Attitude Determination System.In view of its importance, satellite can adopt redundant configuration usually.At present conventional high accuracy gyroscope is optic fiber gyroscope component and hemispherical reso nance gyroscope assembly, and this two classes gyrounit all have employed three positive oblique redundancy configurations, namely uses the responsive three axle inertia angular velocity informations of four gauge outfits.There is Analysis design between four gyroscope instruments due to three positive one oblique gyrounits, therefore select wherein three can determine satellite three axis angular rate.Ideally, select any three in four gyros can calculate identical satellite three axis angular rate, but because gyroscope instrument exists constant value drift, random walk equal error, therefore all there is deviation in various degree in the measured value of each gyroscope instrument and true value.In order to reach the object that high-precision attitude is determined, three gyros selecting deviation relatively little are needed to resolve for satellite three axis angular rate.

According to the pertinent literature delivered both at home and abroad, resolving satellite three axis angular rate method for the gyrounit of three positive one oblique configurations at present has: use least-squares algorithm makes four gyroscope instruments all participate in resolving; Choose three fixing gyroscope instruments according to ground test result to participate in resolving.All there is certain limitation in these methods, such as four gyroscope instruments all participate in the method resolved, although take full advantage of all metrical informations, because four gyroscope instruments also exist performance difference, so the program easily makes the satellite three axis angular rate precision that calculates on the low side; In another kind of scheme, although selected three gyros of best performance according to the test result on ground, but space environment can make gyro performance change from the different of ground and in factors such as the vibrations of launching stage satellite of entering the orbit, actual performance when therefore the test result on ground can not represent in-orbit.

Summary of the invention

The problem of satellite three axis angular rate is participated in resolving in order to solve the gyroscope instrument cannot choosing best performance when existing gyrounit exists redundancy in real time, the invention provides a kind of optimum Gyro choosing method based on Residual Generation device, the Gyro of Dynamic Selection performance optimum in real time is in-orbit achieved by the method constructing one group of Residual Generation device, efficiently utilize the redundancy of gyrounit on star, more high-precision angular velocity information can be provided for attitude control system, and for gyro failure, there is certain fault-tolerant ability.

The object of the invention is to be achieved through the following technical solutions:

Based on an optimum Gyro choosing method for Residual Generation device, comprise the steps:

One, four gyroscope instruments in alternative three positive one oblique gyrounits are divided into four groups of Gyros, often organize containing three different gyros;

Two, according to the installation matrix of alternative gyrounit, utilize the measured value of four groups of Gyros to calculate satellite three axis angular rate respectively, obtain four groups of satellite three axis angular rates;

Three, satellite three axis angular rate that the three-axis attitude provided by star sensor calculates with four groups of Gyros respectively constructs Residual Generation device, obtains four groups of attitude residual errors altogether;

Four, ask for the evaluation function often organizing attitude residual error, sorted according to numerical values recited by four groups of evaluation functions afterwards, that minimum group is optimum Gyro.

The present invention compared with prior art, its advantage and beneficial effect as follows:

(1) real-time Dynamic Selection in-orbit, excludes the gyro of poor-performing or fault, for attitude and heading reference system provides high precision angular velocity information;

(2) star sensor is the measurement component of Satellite attitude and orbit control system standard configuration, and measuring accuracy is usually above at least one order of magnitude of gyro, and therefore this method engineering can realize and not increase extra cost;

(3) algorithm of star sensor and gyroscope structure Residual Generation device is simple, and On-board software easily realizes, and reliability is high.

Accompanying drawing explanation

Fig. 1 is four gyro layouts of three positive one oblique configuration gyrounits.

Embodiment

Below in conjunction with accompanying drawing, technical scheme of the present invention is further described; but be not limited thereto; everyly technical solution of the present invention modified or equivalent to replace, and not departing from the spirit and scope of technical solution of the present invention, all should be encompassed in protection scope of the present invention.

The invention provides a kind of optimum Gyro choosing method based on Residual Generation device, concrete implementation step is as follows:

One, four gyrounits are divided into four groups of alternative Gyros:

As shown in Figure 1, O _nx _ny _nz _nbe the instrument coordinate system of three positive one oblique configuration gyrounits, this gyrounit has four gyros, is labeled as gyro 1, gyro 2, gyro 3, gyro 4 respectively, wherein the direction of measurement of gyro 1, gyro 2, gyro 3 respectively with the O of instrument coordinate system _nx _naxle, O _ny _naxle, O _nz _naxle is parallel, the direction of measurement of gyro 4 and the O of instrument coordinate system _nx _naxle, O _ny _naxle, O _nz _nthe angle of axle is respectively α ₁, α ₂, α ₃, three angles are not all 0.

Two, four groups of satellite three axis angular rates are calculated:

Under equipment selects gyrounit instrument coordinate system, three axis angular rates are respectively ω _xn, ω _yn, ω _zn, four gyro to measure of alternative gyrounit export and are respectively ω ₁, ω ₂, ω ₃, ω ₄, then three axis angular rate ω under the instrument coordinate system that calculates of these four groups of Gyros _n1, ω _n2, ω _n3, ω _n4be respectively:

${\mathrm{\ω}}_{n1}=\left[\begin{array}{c}{\mathrm{\ω}}_{xn1}\\ {\mathrm{\ω}}_{yn1}\\ {\mathrm{\ω}}_{zn1}\end{array}\right]=\left[\begin{array}{ccc}1& 0& 0\\ 0& 1& 0\\ 0& 0& 1\end{array}\right]\left[\begin{array}{c}{\mathrm{\ω}}_1\\ {\mathrm{\ω}}_2\\ {\mathrm{\ω}}_3\end{array}\right];$

${\mathrm{\ω}}_{n2}=\left[\begin{array}{c}{\mathrm{\ω}}_{xn2}\\ {\mathrm{\ω}}_{yn2}\\ {\mathrm{\ω}}_{zn2}\end{array}\right]={\left[\begin{array}{ccc}1& 0& 0\\ 0& 1& 0\\ {\mathrm{cos\α}}_1& {\mathrm{cos\α}}_2& {\mathrm{cos\α}}_3\end{array}\right]}^{-1}\left[\begin{array}{c}{\mathrm{\ω}}_1\\ {\mathrm{\ω}}_2\\ {\mathrm{\ω}}_4\end{array}\right];$

${\mathrm{\ω}}_{n3}=\left[\begin{array}{c}{\mathrm{\ω}}_{xn3}\\ {\mathrm{\ω}}_{yn3}\\ {\mathrm{\ω}}_{zn3}\end{array}\right]={\left[\begin{array}{ccc}1& 0& 0\\ 0& 0& 1\\ {\mathrm{cos\α}}_1& {\mathrm{cos\α}}_2& {\mathrm{cos\α}}_3\end{array}\right]}^{-1}\left[\begin{array}{c}{\mathrm{\ω}}_1\\ {\mathrm{\ω}}_3\\ {\mathrm{\ω}}_4\end{array}\right];$

${\mathrm{\ω}}_{n4}=\left[\begin{array}{c}{\mathrm{\ω}}_{xn4}\\ {\mathrm{\ω}}_{yn4}\\ {\mathrm{\ω}}_{zn4}\end{array}\right]={\left[\begin{array}{ccc}0& 1& 0\\ 0& 0& 1\\ {\mathrm{cos\α}}_1& {\mathrm{cos\α}}_2& {\mathrm{cos\α}}_3\end{array}\right]}^{-1}\left[\begin{array}{c}{\mathrm{\ω}}_2\\ {\mathrm{\ω}}_3\\ {\mathrm{\ω}}_4\end{array}\right].$

The transition matrix that equipment selects gyrounit instrument coordinate to be tied to satellite body coordinate system is A, then ω _n1, ω _n2, ω _n3, ω _n4be projected to three axis angular rate ω under satellite body coordinate system _b1, ω _b2, ω _b3, ω _b4be respectively:

ω _bi＝A·ω _ni(i＝1,2,3,4)。

Four groups of satellite three axis angular rates are obtained according to above step.

Three, construct Residual Generation device, obtain four groups of attitude residual errors:

The particular type of Residual Generation device is unrestricted, can choose suitable wave filter/observer according to actual conditions.Here, we choose conventional EKF method design Residual Generation device.

According to the satellite three axis angular rate ω that above-mentioned Gyro calculates _bcan be expressed as follows:

ω _b＝ω+b+v _g。

In above formula, b, v _gbe respectively drift and the white noise of gyro, ω is real satellite three axis angular rate.

Definition error quaternion is star sensor measured value q and estimation hypercomplex number between difference DELTA q, then have:

$q=\hat{q}\⊗\mathrm{\Δ}q$

Because Δ q is in a small amount, approximate [1,0,0,0] ^t, after formula conversion, can obtain:

$\mathrm{\Δ}{\stackrel{\·}{Q}}_q=-\[{\widehat{\mathrm{\ω}}}_b\×\]{\mathrm{\ΔQ}}_q-\frac{1}{2}(\mathrm{\Δ}b+v_g).$

$\mathrm{\Δ}\stackrel{\·}{b}=\stackrel{\·}{b}=v_b$

In above formula, Δ Q _qfor the vector section of Δ q, Δ b is the evaluated error of gyroscopic drift, ${\widehat{\mathrm{\ω}}}_b={\mathrm{\ω}}_b-b.$

Getting state variable is x=[Δ Q _q ^t, Δ b ^t] ^t, then state equation and the observation equation that can obtain extended Kalman filter are:

$\stackrel{\·}{x}=\left[\begin{array}{cc}-\[{\widehat{\mathrm{\ω}}}_b\×\]& -\frac{1}{2}{I}_{3\×3}\\ 0_{3\×3}& 0_{3\×3}\end{array}\right]\left[\begin{array}{c}\mathrm{\Δ}{Q}_q\\ \mathrm{\Δ}b\end{array}\right]+\left[\begin{array}{cc}-\frac{1}{2}{I}_{3\×3}& 0_{3\×3}\\ 0_{3\×3}& I_{3\×3}\end{array}\right]\left[\begin{array}{c}{v}_g\\ v_b\end{array}\right].$

Δz＝[I _3×30 _3×3]x

In above formula, I _{3 × 3}it is the unit matrix of 3 dimensions.Above-mentioned extended Kalman filter namely can as Residual Generation device, Δ z=Δ Q _qbe attitude residual error.

Respectively the measured value of four groups of satellite three axis angular rates and star sensor is constructed Residual Generation device, four attitude residual errors can be obtained, be respectively Δ z ₁, Δ z ₂, Δ z ₃, Δ z ₄.

Four, determine the evaluation function often organizing attitude residual error, select optimum Gyro:

Evaluation function r adopts the form of 2-norm, then:

r _i＝||Δz _i||(i＝1,2,3,4)。

By r ₁, r ₂, r ₃, r ₄sort according to size.If r ₁minimum, then Gyro 1 is optimum Gyro; If r ₂minimum, then Gyro 2 is optimum Gyro; If r ₃minimum, then Gyro 3 is optimum Gyro; If r ₄minimum, then Gyro 4 is optimum Gyro.

Claims (3)

1., based on an optimum Gyro choosing method for Residual Generation device, it is characterized in that described method step is as follows:

One, four gyroscope instruments in alternative three positive one oblique gyrounits are divided into four groups of Gyros, often organize containing three different gyros;

Two, according to the installation matrix of alternative gyrounit, utilize the measured value of four groups of Gyros to calculate satellite three axis angular rate respectively, obtain four groups of satellite three axis angular rates;

Three, satellite three axis angular rate that the three-axis attitude provided by star sensor calculates with four groups of Gyros respectively constructs Residual Generation device, obtains four groups of attitude residual errors altogether;

Four, ask for the evaluation function often organizing attitude residual error, sorted according to numerical values recited by four groups of evaluation functions afterwards, that minimum group is optimum Gyro.

2. the optimum Gyro choosing method based on Residual Generation device according to claim 1, is characterized in that the method for described structure Residual Generation device is EKF method.

3. the optimum Gyro choosing method based on Residual Generation device according to claim 1, is characterized in that described evaluation function adopts the form of 2-norm.