CN101214861A - Star sensor attitude determination method at self-determination retrieve rail controlling fault - Google Patents

Abstract

A star sensor attitude-determination method during the autonomous orbital control fault recovery includes: (1) predicting the satellite inertia attitude according to data measured by a gyro; (2) calculating the filtering modified innovation amount according to the inertia attitude of the satellite and the optical axis vector and the lateral axis vector under an inertial coordinate system which are measured and output by the star sensor, and calculating the error of the innovation amount between the front and the back periodicals, which is used for judging the consistency of star sensor data; (3) judging the consistency of the star sensor data; (4) fixing the attitude of the star sensor according to the two vectors; (5) introducing the star sensor which is combined with the gyro for the correction of the satellite attitude under the condition that the star sensor data arranges the initial value of the attitude estimation. The method can improves the reliability of the orbital control fault recovery, saves the time for the fault recovery and ensures that the orbital control is accurately recovered in time.

Description

Star sensor method for determining posture during a kind of autonomous recovery rail controlling fault

Technical field

The satellite attitude that the present invention relates in a kind of spacecraft rail controlling fault rejuvenation redefines method.Particularly after fault handling, when recovering the rail control, introduce the method that star sensor information is decided appearance again.

Background technology

During Satellite Orbit Maneuver control at present, attitude is determined generally all to utilize the metrical information of gyro to carry out attitude prediction.If break down in the rail control process, then stop this rail control, pending fault carries out the rail control after getting rid of once more, and the attitude valuation overcorrection of satellite during the rail control once more is consistent with the actual attitude of satellite.But this method need just can make satellite recover the rail control than the long time, influence the time of failure recovery, can not satisfy the requirement (rail control uniqueness requires to be meant that rail control enforcement will be in time, and rail control attitude is wanted accurately) of rail control window uniqueness.If satellite is independently finished fault handling and failure recovery on star, then in the satellite failure process, the Satellite Attitude Estimation value may depart from actual attitude.In failover procedure, if only carrying out attitude by gyro data estimates, satellite carries out attitude and estimates on a wrong initial value basis, can make the estimation attitude and the actual attitude of satellite differ bigger, thereby influence failure recovery and follow-up rail control effect, be difficult to satisfy the requirement of uniqueness (accurately implementing the rail control); If the attitude estimation is estimated by gyro and star sensor correction combination is finished, then with only compare with gyro estimation attitude, can reduce to estimate the error of attitude and actual attitude gradually, but because the attitude correction is limited, the time of eliminating evaluated error is longer, make the time of failure recovery longer, can influence the realization that rail control window uniqueness requires (rail control in time) equally.

Summary of the invention

Technology of the present invention is dealt with problems: overcome the deficiencies in the prior art, satellite estimates that attitude and the actual attitude of satellite differ big or need the long period to recover the situation of attitude during at failure recovery, and the attitude when having proposed independently to recover rail controlling fault on a kind of star is determined method.This method can improve the reliability that rail controlling fault recovers, and saves the time of failure recovery, guarantees to recover in time, exactly track control.

Technical solution of the present invention: the star sensor method for determining posture during a kind of autonomous recovery rail controlling fault is characterized in that comprising:

(1) estimates satellite inertia attitude according to the gyro to measure data;

(2) according to the required new breath amount of optical axis vectorial sum transverse axis vector calculation of filtered correction under the inertial coordinates system of satellite inertia attitude and star sensor measurement output, and calculate former and later two adjacent periods error of breath amount newly, be used to judge the conformability of star sensor data;

(3) the star sensor data consistency is differentiated: at first judge whether to put the initial value that attitude is estimated with the star sensor data, if put, change step (5) over to, if do not put, then whether the error of the new breath amount of calculating in the determining step (2) is within the range of permission double at least, if within the range of permission, then change step (4) over to, if not in allowed limits, then change step (1) over to, continue to utilize gyro data to estimate the inertia attitude of satellite;

(4) the two vectors of star sensor are decided appearance: utilize star sensor optical axis and the component of transverse axis vector under satellite body system of axes and inertial coordinates system, calculate satellite inertia attitude, and upgrade the initial value of Satellite Attitude Estimation with this result, the sign of initial value was put in setting, change step (1) then over to, carry out the attitude prediction of next cycle;

(5) put under the situation of the initial value that attitude estimates in the star sensor data, introduced star sensor, with gyro combination the carrying out correction of satellite attitude.

The present invention's advantage compared with prior art is:

(1) the present invention is used in combination the two vector method for determining posture and the gyro star sensor combined filter modification method of star sensor, can estimate satellite attitude timely and accurately, and the assurance satellite is accurately and timely realized the rail control.

(2) the present invention is when rail controlling fault recovers, at first utilize the star sensor take off data directly to carry out geometry and decide appearance, determine the satellite attitude initial value, then utilizing conventional gyro star sensor combined filter method to carry out attitude again estimates, shorten fault recovery process effectively, guaranteed the timely and validity that satellite failure recovers.

(3) the present invention carries out the differentiation of star sensor data validity by the comparison of star sensor data, has guaranteed the correctness of star sensor data.

(4) method of the present invention is reliable, is easy to realize on the star.Can be widely used in the track control of various satellites.Be specially adapted to key point or unique window and become the track control of the spacecraft of rail requirement, become rail and pericynthion braking for the first time as the perigee for the third time of lunar spacecraft.

Description of drawings

Star sensor method for determining posture diagram of circuit when Fig. 1 independently recovers rail controlling fault for the present invention;

Fig. 2 is for adopting the simulation curve figure of the inventive method;

Fig. 3 is for adopting the simulation curve figure of conventional approach;

Wherein, in the simulation curve of Fig. 2, Fig. 3, on behalf of qw (0), curve 2, curve 1 represent qw (1), curve 3 to represent qw (2), is respectively the vector part of the actual attitude quaternion of satellite; On behalf of q (0), curve 5, curve 4 represent q (1), curve 6 to represent q (2), is respectively the star upper estimate of satellite attitude quaternion vector part.

The specific embodiment

Star sensor method for determining posture during autonomous recovery rail controlling fault of the present invention is meant and becomes the rail et out of order, by on the star during autonomous the recovery, utilizes the metrical information of star sensor, carries out the attitude of satellite and determines.Utilize method of geometry to determine the attitude of satellite,, adopt gyro star sensor combined filter method afterwards, carry out the attitude of satellite and estimate with the attitude valuation on this posture renewal star.In order to guarantee the correct reliable of star sensor data, before the use star sensor carries out deciding appearance, at first carry out the conformability of star sensor measurement output and differentiate.Utilize the star sensor measuring amount that obtains in continuous two sampling periods--the error of-optical axis vector is judged, if double above comparison optical axis vector error thinks that then the star sensor data are actvies in allowed band.Utilize the optical axis vectorial sum transverse axis vector of this star sensor to carry out once two vectors and decide appearance calculating, obtain the inertia attitude in this moment of satellite.With this inertia posture renewal Satellite Attitude Estimation value, as the initial value of the estimation of the attitude in the rail control rejuvenation; Introduce star sensor at last, with gyro combination the carrying out correction of satellite attitude.

The concrete implementing procedure of method of the present invention is referring to Fig. 1, and concrete steps are as follows:

(1) gyro is estimated the inertia attitude: estimate satellite inertia attitude according to the gyro to measure data.

The conversion of a gyro data

Utilize the metrical information of three gyros can obtain satellite three-axis attitude angular velocity information.According to the principle of work of rate, each sampling can obtain the angle step of gyro in this sampling period, through calculating the attitude angular velocity of three on satellite after suitably changing.

(k) the installation matrix B in the satellite body system of axes is calculated gyro output transition matrix A=B for numbering i, j according to three gyros that participate in deciding appearance ^-1, in conjunction with the measurement output Δ g of three gyros _i, Δ g _j, Δ g _k, can obtain satellite three-axis attitude angle increment information Δ g _x, Δ g _y, Δ g _zCalculate by following formula:

$\left[\begin{array}{c}\mathrm{\Δ}{g}_x\\ \mathrm{\Δ}{g}_y\\ \mathrm{\Δ}{g}_z\end{array}\right]=A\left[\begin{array}{c}\mathrm{\Δ}{g}_i\\ \mathrm{\Δ}{g}_j\\ \mathrm{\Δ}{g}_k\end{array}\right]$

Three absolute angle speed calculation of b

According to the result of calculation of gyro data conversion, calculate the mean angular velocity in the sampling period: $\widehat{\mathrm{\ω}}={\left[{\widehat{\mathrm{\ω}}}_x{\widehat{\mathrm{\ω}}}_y{\widehat{\mathrm{\ω}}}_z\right]}^T.$ Note cutting gyroscope constant value drift amount =[ in the calculating through demarcating _x _y _z] ^TUnit: radian/hour.

${\widehat{\mathrm{\ω}}}_x=\mathrm{\Δ}{g}_x/\mathrm{\Δt}-{\hat{b}}_x/3600$

${\widehat{\mathrm{\ω}}}_y=\mathrm{\Δ}{g}_y/\mathrm{\Δt}-{\hat{b}}_y/3600$

${\widehat{\mathrm{\ω}}}_z=\mathrm{\Δ}{g}_z/\mathrm{\Δt}-{\hat{b}}_z/3600$

C inertia attitude prediction

According to three absolute angle speed that calculate, calculate the three axis angular rate increments of satellite in this cycle, adopt following formula to estimate the inertia attitude of satellite, with the quaternion form $\stackrel{\widehat{-}}{q}=\left[\begin{array}{cccc}{\hat{q}}_1& {\hat{q}}_2& {\hat{q}}_3& {\hat{q}}_4\end{array}\right]$ Provide.

$\mathrm{\Δ}\hat{g}=\widehat{\mathrm{\ω}}\·\mathrm{\Δt}$

$\stackrel{\widehat{-}}{q}=\stackrel{\widehat{-}}{q}+\frac{1}{2}\mathrm{Eq}\left(\stackrel{\widehat{-}}{q}\right)\mathrm{\Δ}\hat{g}$

If ${\hat{q}}_4<0,$ Then $\stackrel{\widehat{-}}{q}=-\stackrel{\widehat{-}}{q}$

$\stackrel{\widehat{-}}{q}=\stackrel{\widehat{-}}{q}/\mathrm{Norm}\left(\stackrel{\widehat{-}}{q}\right)$

Annotate: Δ t is the sampling period in the above-mentioned formula.Function Eq utilizes measuring satellite angular velocities information to estimate the computing formula of satellite attitude quaternion.Norm is the function of quaternion delivery.Function definition is as follows:

Function a=Norm (B)

Function name: Norm

Input: B=[b1, b2 ... bn] T, n represents the dimension of vectorial B

Output: a

The function content:

{

$a=\sqrt{b_1^2+b_2^2+\&CenterDot;\&CenterDot;\&CenterDot;+b_n^2}$

}

Function A=Eq (B)

Function name: Eq

Input: B=[b ₁, b ₂, b ₃, b ₄] ^T

Output: A=(a _Ij) _{4 * 3}

The function content:

{

a ₁₁＝b ₄

a ₁₂＝-b ₃

a ₁₃＝b ₂

a ₂₁＝b ₃

a ₂₂＝b ₄

a ₂₃＝-b ₁

a ₃₁＝-b ₂

a ₃₂＝b ₁

a ₃₃＝b ₄

a ₄₁＝-b ₁

a ₄₂＝-b ₂

a ₄₃＝-b ₃

}

(2) star sensor Measurement and Data Processing

According to the optical axis vector Z under the inertial coordinates system of used star sensor measurement output _IWith transverse axis vector X _IThe new breath amount δ Z of calculation of filtered correction.And calculate the error delta Z of new breath amount of former and later two cycles, be used for the conformability of follow-up judgement star sensor data.

Y _I＝Z _I×X _I

$\mathrm{\&delta;}{Z}_x=X_B\&times;(\mathrm{Aq}\left(\stackrel{\widehat{-}}{q}\right)\&CenterDot;X_I)$

$\mathrm{\&delta;}{Z}_y=Y_B\&times;(\mathrm{Aq}\left(\stackrel{\widehat{-}}{q}\right)\&CenterDot;Y_I)$

$\mathrm{\&delta;}{Z}_z=Z_B\&times;(\mathrm{Aq}\left(\stackrel{\widehat{-}}{q}\right)\&CenterDot;Z_I)$

$\mathrm{\&delta;Z}=\frac{1}{2}(\mathrm{\&delta;}{Z}_x+\mathrm{\&delta;}{Z}_y+\mathrm{\&delta;}{z}_z)$

ΔZ＝|δZ _pst-δZ|

δ Z _Pst=δ Z (quantity of information that this cycle of record calculates is used for the comparison of next computation of Period)

Wherein, Z _IThe expression of sensor optical axis vector under geocentric inertial coordinate system that measures for star sensor; X _IThe expression of sensor transverse axis vector under geocentric inertial coordinate system that measures for star sensor; X _B, Y _B, Z _BBe respectively three coordinate axle (transverse axis (X of star sensor system of axes _B), optical axis (Z _B), Y _BWith X _B, Z _BSatisfy right-hand rule, Y _B=Z _B* X _B) expression under the satellite body system of axes, determine according to the concrete installation site of star sensor on satellite.Aq () estimates the attitude transition matrix of the satellite body system of axes of Attitude Calculation with respect to inertial coordinates system according to satellite.δ Z _PstThe new breath amount in last cycle for record.

(3) the star sensor data consistency is differentiated

At first judge whether to utilize the star sensor take off data to put the initial value that attitude is estimated.If put, then change step (5) over to, carry out follow-up star sensor filtering correction.If do not put the initial value that attitude is estimated, then judge new breath magnitude of error Δ Z whether double more than all within the range of permission.If condition satisfies, then change step (4) over to, carry out two vectors and decide appearance, and upgrade the Satellite Attitude Estimation value with deciding the appearance result.

(4) the two vectors of star sensor are decided appearance

Utilize star sensor optical axis and the component of transverse axis vector under satellite body system of axes and inertial coordinates system, calculate the attitude transition matrix A:A=Avv (X of satellite body system of axes with respect to inertial coordinates system _I, Z _I, X _B, Z _B).This function is decided appearance for carrying out two vectors at the component of satellite body system of axes and inertial coordinates system respectively with two vectors, and calculating satellite body system of axes is with respect to the function of the attitude quaternion of inertial coordinates system, and function definition is as follows:

Function A=Avv (X _1I, X _2I, X _1b, X _2b)

Function name Avv

Input: X _1I, X _2I, X _1b, X _2b(being unit vector)

Output: A=(a _Ij) _{3 * 3}

The function content:

{

V _2I＝X _1I×X _2I；

V _2I＝V _2I/Norm(V _2I)；

V _3I＝X _1I×V _2I；

V _2b＝X _1b×X _2b；

V _2b＝V _2b/Norm(V _2b)；

V _3b＝X _1b×V _2b；

$A=X_{1b}{X}_{1I}^T+V_{2b}{V}_{2I}^T+V_{3b}{V}_{3I}^T$

}

Decide the attitude transition matrix A that appearance is determined according to two vectors, calculate satellite inertia attitude quaternion: $\stackrel{-}{q}=\mathrm{Qa}\left(A\right).$ This function is the general formula with attitude transform matrix calculations satellite quaternion, and function definition is as follows:

Function A=Qa (B)

Function name: Qa

Input: B=(b _Ij) _{3 * 3}

Output: A=[a ₁, a ₂, a ₃, a ₄] ^T

The function content:

if(b ₁₁+b ₂₂+b ₃₃+1≥0.004)

{

$a_4=\left(\sqrt{b_{11}+b_{22}+b_{33}+1}\right)/2$

a ₁＝(b ₂₃-b ₃₂)/(4·a ₄)

a ₂＝(b ₃₁-b ₁₃)/(4·a ₄)

a ₃＝(b ₁₂-b ₂₁)/(4·a ₄)

}

Else if (b ₁₁+ b ₂₂+ b ₃₃+ 1＜0.004 and 1-b ₁₁+ b ₂₂-b ₃₃〉=0.004)

{

$a_2=\left(\sqrt{1-b_{11}+b_{22}+b_{33}}\right)/2\&CenterDot;\mathrm{sgn}(b_{31}-b_{13})$

(if b ₃₁-b ₁₃=0, then get sgn (b ₃₁-b ₁₃)=1)

a ₁＝(b ₂₁+b ₁₂)/(4·a ₂)

a ₃＝(b ₃₂+b ₂₃)/(4·a ₂)

a ₄＝(b ₃₁-b ₁₃)/(4·a ₂)

}

Else if (b ₁₁+ b ₂₂+ b ₃₃+ 1＜0.004 and 1-b ₁₁+ b ₂₂-b ₃₃＜0.004

And 1+b ₁₁-b ₂₂-b ₃₃〉=0.004)

{

$a_1=\left(\sqrt{1+b_{11}-b_{22}-b_{33}}\right)/2\&CenterDot;\mathrm{sgn}(b_{23}-b_{32})$

(if b ₂₃-b ₃₂=0, then get sgn (b ₂₃-b ₃₂)=1)

a ₂＝(b ₂₁+b ₁₂)/(4·a ₁)

a ₃＝(b ₁₃+b ₃₁)/(4·a ₁)

a ₄＝(b ₂₃-b ₃₂)/(4·a ₁)

}

else

{

$a_3=\left(\sqrt{1-b_{11}-b_{22}+b_{33}}\right)/2\&CenterDot;\mathrm{sgn}(b_{12}-b_{21})$

(if b ₁₂-b ₂₁=0, then get sgn (b ₁₂-b ₂₁)=1)

a ₁＝(b ₁₃+b ₃₁)/(4·a ₃)

a ₂＝(b ₂₃+b ₃₂)/(4·a ₃)

a ₄＝(b ₁₂-b ₂₁)/(4·a ₃)

}

Upgrade quaternion estimated valve on satellite star with this result of calculation: $\stackrel{\widehat{-}}{q}=\stackrel{-}{q}$ (

Be the attitude quaternion of estimating in the satellite), and be provided with one and with star sensor the sign of Satellite Attitude Estimation initial value be set, be used to judge whether down-stream also needs to continue to judge the conformability of star sensor data.

(5) introduce star sensor attitude modification method

After satellite attitude is estimated initial value and is provided with, adopt conventional star sensor filtering algorithm to carry out real-time attitude correction.The filtering algorithm design gets final product according to the Kalman filtering principle of classics.The filtering algorithm that adopts in the lunar spacecraft is as follows:

$\left[\begin{array}{c}\mathrm{\&delta;}\hat{q}\\ \mathrm{\&Delta;}\hat{b}\end{array}\right]=K_s\mathrm{\&delta;Z}$

$\stackrel{\widehat{-}}{q}=\stackrel{\widehat{-}}{q}+\mathrm{Eq}\left(\stackrel{\widehat{-}}{q}\right)\&CenterDot;\mathrm{\&delta;}\hat{q}$

If ${\hat{q}}_4<0,$ Then $\stackrel{\widehat{-}}{q}=-\stackrel{\widehat{-}}{q}$

$\stackrel{\widehat{-}}{q}=\stackrel{\widehat{-}}{q}/\mathrm{Norm}\left(\stackrel{\widehat{-}}{q}\right)$

＝+Δ

Here K _sBe the filtering coefficient of correction,  is the scalar quantity of gyroscope constant value drift.

According to above-mentioned implementation step, carried out l-G simulation test at certain lunar exploration satellite, simulation curve is as shown in Figure 2.Simulation process is: in the rail control ignition process, satellite breaks down, and causes attitude deviation target ignition attitude, independently changes rate damping pattern by rail control directional pattern at simulation time in the time of 272 seconds.In the rate damping pattern, introducing star sensor data are carried out pair vectors and are decided appearance, utilize and decide the appearance result, are provided with the attitude estimated valve of satellite, after this continue to use star sensor and gyro to unite filtering method and carry out the attitude estimation.From simulation curve, the estimation attitude of satellite has departed from actual attitude in the failure process, because the star sensor data are decided appearance result's introducing, very rapid convergence is in true attitude to make the estimation attitude of satellite, and the satellite attitude error is less than 0.0005 (rad).

In order to compare the effect that the inventive method reaches, Fig. 3 has provided and has adopted conventional method to carry out the simulation curve that failure recovery is decided appearance.Simulation process is consistent with above-mentioned simulation process, does not just introduce star sensor and decides the attitude valuation that the appearance data are provided with satellite.From simulation curve, do not decide the appearance result owing to introduce the two vectors of star sensor, behind the failure recovery, the attitude quaternion of estimating on the satellite star still differs bigger with the true attitude of satellite, and the satellite attitude error is about 0.08 (rad).

System described above is a kind of situation of the present invention, and those skilled in the art can carry out under the situation of the present invention variously augmenting, improving and change not departing from according to different requirements and design parameters, and therefore, the present invention is widely.

Claims (3)

1. autonomous star sensor method for determining posture when recovering rail controlling fault is characterized in that comprising:

(1) estimates satellite inertia attitude according to the gyro to measure data;

(2) according to the required new breath amount of optical axis vectorial sum transverse axis vector calculation of filtered correction under the inertial coordinates system of satellite inertia attitude and star sensor measurement output, and calculate former and later two adjacent periods error of breath amount newly, be used to judge the conformability of star sensor data;

(3) the star sensor data consistency is differentiated: at first judge whether to put the initial value that attitude is estimated with the star sensor data, if put, change step (5) over to, if do not put, then whether the error of the new breath amount of calculating in the determining step (2) is within the range of permission double at least, if within the range of permission, then change step (4) over to, if not in allowed limits, then change step (1) over to, continue to utilize gyro data to estimate the inertia attitude of satellite;

(4) the two vectors of star sensor are decided appearance: utilize star sensor optical axis and the component of transverse axis vector under satellite body system of axes and inertial coordinates system, calculate satellite inertia attitude, and upgrade the initial value of Satellite Attitude Estimation with this result, the sign of initial value was put in setting, change step (1) then over to, carry out the attitude prediction of next cycle;

(5) put under the situation of the initial value that attitude estimates in the star sensor data, introduced star sensor, with gyro combination the carrying out correction of satellite attitude.

2. the star sensor method for determining posture during a kind of autonomous recovery rail controlling fault according to claim 1, it is characterized in that the method for estimating satellite inertia attitude in the described step (1) is: the attitude angular velocity that at first calculates satellite according to the metrical information of gyro, calculate three absolute angle speed then, according to absolute angle speed, estimate the inertia attitude of satellite at last;

3. the star sensor method for determining posture during a kind of autonomous recovery rail controlling fault according to claim 1, it is characterized in that: the method for the error delta Z of the new breath amount δ Z of calculation of filtered correction and new breath amount is in the described step (2):

Y _I＝Z _I×X _I

$\mathrm{\&delta;}{Z}_x=X_B\&times;(\mathrm{Aq}\left(\stackrel{\widehat{-}}{q}\right)\&CenterDot;X_I)$

$\mathrm{\&delta;}{Z}_y=Y_B\&times;(\mathrm{Aq}\left(\stackrel{\widehat{-}}{q}\right)\&CenterDot;Y_I)$

$\mathrm{\&delta;}{Z}_z=Z_B\&times;(\mathrm{Aq}\left(\stackrel{\widehat{-}}{q}\right)\&CenterDot;Z_I)$

$\mathrm{\&delta;Z}=\frac{1}{2}(\mathrm{\&delta;}{Z}_x+\mathrm{\&delta;}{Z}_y+\mathrm{\&delta;}{Z}_z)$

ΔZ＝|δZ _pst-δZ|

δZ _pst＝δZ

Wherein, Z _IThe expression of sensor optical axis vector under geocentric inertial coordinate system that measures for star sensor; X _IThe expression of sensor transverse axis vector under geocentric inertial coordinate system that measures for star sensor; X _B, Y _B, Z _BBe respectively three expressions of coordinate axle under the satellite body system of axes, i.e. transverse axis X of star sensor system of axes _B, optical axis Z _B, Y _BWith X _B, Z _BSatisfy right-hand rule, Y _B=Z _B* X _B, determine that according to the concrete installation site of star sensor on satellite Aq () estimates the attitude transition matrix of the satellite body system of axes of Attitude Calculation with respect to inertial coordinates system, δ Z according to satellite _PstFor writing down the new breath amount in last cycle.