CN101214860A - Method for self-determination choosing attitude determination mode during rail controlling course - Google Patents

Abstract

A method for autonomously selecting an attitude-determination mode during the orbital control process includes: (1) predicting the satellite inertia attitude according to gyro measurement data; (2) judging that whether the star sensor attitude correction is required to be induced: judging that whether the attitude angular velocity of the satellite exceeds the measurement scope of a gyro according to the measured information of the gyro; if not, continuously using the gyro to predict the inertia attitude of the satellite, transferring to the step (1) and predicting the inertia attitude of the satellite in the next period; if the attitude angular velocity of the satellite exceeds the threshold which is set according to the gyro measurement scope, a gyro transfinite mark is arranged; then judging that whether the attitude angular velocity of three axils of the satellite satisfies the working condition requirements of the star sensor, and if yes, transferring to the step (3) and if no, transferring to the step (1); (3)introducing the star sensor for the attitude correction. The method of the invention reduces the attitude error during the orbital control process, improves the orbital control precision and effectively ensures that the orbital control is accurately finished in time.

Description

The method of self-determination choosing attitude determination mode in the rail control process

Technical field

The present invention relates to the method that attitude is determined in a kind of spacecraft rail control process.Particularly bigger at rail control engine disturbance torque, cause measuring satellite angular velocities bigger, exceed under the gyro to measure scope situation, satellite can independently be introduced the method that the star sensor metrical information is accurately revised satellite attitude.

Background technology

Gyro attitude prediction method is comparatively commonly used in satellite attitude control process, but because the influence of its gyro wander makes it have only application within a short period of time could guarantee accuracy of attitude determination.And the scope of testing the speed of general high accuracy gyroscope is more limited, change greatly in measuring satellite angular velocities, exceed under the gyro to measure scope situation, only estimate that by gyro data satellite attitude can make the true attitude of estimation pose deviation satellite, influences the satellite control accuracy.The method of using gyro to estimate in the Satellite Orbit Maneuver process has been avoided comparatively complicated filtering algorithm, simplified on the star and realized, help improving the reliability of rail control, but the attitude misalignment that brings that is limited in scope owing to test the speed will make the firing attitude of satellite depart from needed nominal firing attitude, directly influence the rail control effect of satellite.For the satellites that control is had relatively high expectations concerning rail such as moon exploration and even survey of deep space, rail control error may cause the failure of whole task.

The method of gyro and star sensor integrated attitude determination is also comparatively commonly used in the three axis stabilization control of high precision satellite.But it is mainly used in the Steady-State Control process of satellite.When the attitude angular velocity that has when satellite is big, can produce adverse influence to the importance in star map recognition ability of star sensor, star sensor is output data normally, estimates attitude thereby can not revise satellite in time, still can make satellite estimate the actual attitude of pose deviation.Can influence rail control effect equally if in rail control process, use, cause rail control mission failure when serious.

When most geostationary orbits or middle low-orbit satellite transfer orbital control, attitude is determined only to utilize the metrical information of gyro to carry out.When the orbit maneuver engine disturbance torque is big, might cause satellite bigger attitude angular velocity to occur, when this cireular frequency surpasses the measurement range of gyro, only carrying out attitude prediction by gyro can cause satellite to determine that attitude and the actual attitude of satellite are not inconsistent, make the actual attitude of satellite constantly depart from firing attitude, influence rail control precision, may cause rail control failure when serious.

Summary of the invention

Technology of the present invention is dealt with problems: overcome the deficiencies in the prior art, attitude angular velocity is big during at the rail control that may occur, the situation that surpasses the gyro to measure scope, when having proposed that attitude angular velocity is in the gyro measurable range during a kind of rail control, utilize gyro data to estimate satellite attitude, occur exceeding the gyro to measure scope than great-attitude angle speed the time, the autonomous method of introducing the correction of star sensor attitude, promptly propose during a kind of rail control according to the measuring satellite angular velocities observed reading, the method of self-determination choosing attitude determination mode, reduce the attitude error in the rail control process, improve rail control precision, guarantee that effectively the rail control is timely, finish exactly.

The method of self-determination choosing attitude determination mode is meant in the rail control process celestial body cireular frequency hour in the rail control process, only selects gyro to estimate to carry out attitude to determine; When measuring satellite angular velocities is bigger, when surpassing the gyro to measure scope, independently select gyro to estimate to carry out attitude to determine with the method for star sensor attitude correction.

Technical solution of the present invention: the method for self-determination choosing attitude determination mode in the rail control process is characterized in that comprising:

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

At first calculate the attitude angular velocity of satellite, calculate three absolute angle speed then,, estimate the inertia attitude of satellite at last according to absolute angle speed according to the metrical information of gyro;

(2) judge whether to need to introduce the correction of star sensor attitude:

Metrical information according to gyro, judge whether the attitude angular velocity of satellite has surpassed the measurement range of gyro, if do not exceed measurement range, then continue to utilize gyro to estimate satellite inertia attitude, change step (1) over to, carry out the satellite inertia attitude prediction of following one-period, if surpass the threshold value of setting according to the gyro to measure scope, then set the sign that gyro transfinites, judge then whether the attitude angular velocity of three on satellite satisfies the service conditions requirement of star sensor, when the service conditions that satisfies star sensor requires, then change step (3) over to, otherwise continue to adopt the gyro to measure data to estimate satellite inertia attitude, change (1) over to.

(3) introduce star sensor and carry out the attitude correction.

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

(1) the present invention can independently select different method for determining posture for use according to actual conditions in rail control process, does not end the rail control easily, can guarantee that the rail control carries out smoothly.

(2) the present invention has guaranteed effectively that by the introducing of star sensor information high thrust rail control engine becomes the precision that attitude is determined in the rail process.Thereby in rail control process, carry out attitude control accurately and timely.

(3) this method is applicable to the high thrust transfer orbital control of spacecraft.Be particularly suitable in follow-up survey of deep space series satellite, using.Have good, the portable good characteristics of inheritance.

Description of drawings

Fig. 1 is the diagram of circuit of the inventive method;

Fig. 2 is rail control process attitude determination mode flow path switch figure;

Fig. 3 is the simulation curve of the inventive method;

Fig. 4 is the simulation curve of conventional approach;

Wherein, in the simulation curve of Fig. 3, Fig. 4, 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

The measurement range that is installed in the gyro that tests the speed on the satellite is limited.And in the Satellite Orbit Maneuver process, because rocking the interference that brings etc., interference that rail control engine start brings and liquid fuel all may cause bigger measuring satellite angular velocities, when this cireular frequency surpasses the gyro to measure scope, gyro can't provide correct angular velocity information, thereby satellite can't obtain attitude information accurately, can not guarantee attitude control accuracy, thereby influence rail control precision.

The present invention takes to estimate in the attitude that rail control process at first adopts the take off data of gyro to carry out satellite, the measurement of real-time judge gyro output simultaneously, near saturation value the time, provide the attitude angular velocity sign that transfinites, after treating that measuring satellite angular velocities steadily (satisfies star sensor and measures requirement), introduce the method that the star sensor data are revised the attitude of gyro estimation, to reduce the saturated evaluated error of bringing of gyro.

As shown in Figure 1, concrete implementation step of the present invention is as follows:

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

A, gyro data switching rate integration

Utilize the metrical information of three gyros can obtain satellite three-axis attitude angular velocity information.With the rate is example, and 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 △ g of three gyros _i, △ g _j, △ g _k, can obtain satellite three-axis attitude angle increment information △ g _x, △ g _y, △ g _zCalculate by following formula:

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

B, three absolute angle speed calculation

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] ^T, unit: radian/hour.The demarcation of gyroscope constant value drift can take ground demarcate in advance or star on method by other sensor information real-time calibrations such as star sensors obtain.

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

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

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

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 $\widehat{\stackrel{\&OverBar;}{q}}=-\widehat{\stackrel{\&OverBar;}{q}}$

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

Δ t is the sampling period in the above-mentioned formula, and function Eq utilizes measuring satellite angular velocities information to estimate the computing formula of satellite attitude quaternion.

(2) judge whether to need to introduce the correction of star sensor attitude

Metrical information according to gyro, judge whether the attitude angular velocity of satellite has surpassed the measurement range of gyro, if do not exceed measurement range, then continue to utilize gyro to estimate satellite inertia attitude, change step (1) over to, carry out the satellite inertia attitude prediction of following one-period, if surpass the threshold value of setting according to the gyro to measure scope, then set the sign that gyro transfinites, judge then whether the attitude angular velocity of three on satellite satisfies the service conditions requirement of star sensor, when the service conditions that satisfies star sensor requires, then change step (3) over to, otherwise continue to adopt the gyro to measure data to estimate satellite inertia attitude, change (1) over to.

(3) star sensor attitude modification method

After introducing the star sensor correction, 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.

Provide the concrete application of this method in lunar exploration satellite below:

As shown in Figure 2, lunar spacecraft adopts 6 three floating gyros, and wherein any three gyros all can be determined satellite inertia attitude angular velocity.The measurement range of each gyro is-0.9 °/s～0.9 °/s, considers certain design margin in engineering is used, and choosing the threshold value that attitude angular velocity surpasses the gyro to measure scope is 0.8 °/s.

Using 3 medium accuracy star sensors to carry out the satellite three-axis attitude in the lunar spacecraft determines.Star sensor has under the 0.5 °/s cireular frequency situation and can normally discern star chart at celestial body, provides attitude data.In engineering is used, consider certain design margin, choose and introduce the star sensor data to carry out the threshold values of attitude correction be 0.35 °/s.

According to above-mentioned implementation step, carried out l-G simulation test at certain lunar exploration satellite, simulation curve is as shown in Figure 3.Simulation process is: the rail control begins, measuring satellite angular velocities big (1.5 °/s), gyro output is saturated, the actual attitude of Satellite Attitude Estimation value and satellite engenders deviation, causes attitude deviation target ignition attitude.Behind the simulation time 20s, introduce the quick correction of star, as can be seen, very rapid convergence is in true attitude in the satellite attitude valuation from curve, and attitude error is less than 0.001 (rad) behind the 100s.

In order to compare the effect that the inventive method reaches, Fig. 4 has provided the simulation curve of not introducing the star sensor correction in rail control process.The simulation process condition is consistent with said process, and satellite only carries out attitude prediction by gyro data.From simulation curve, owing to do not introduce star sensor filtering correction, the satellite attitude valuation departs from the true attitude of satellite all the time, and attitude error is about 0.08 (rad) behind the 100s, and this error will influence rail control precision.

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 (4)

1. the method for self-determination choosing attitude determination mode in the rail control process is characterized in that comprising:

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

At first calculate the attitude angular velocity of satellite, calculate three absolute angle speed then,, estimate the inertia attitude of satellite at last according to absolute angle speed according to the metrical information of gyro;

(2) judge whether to need to introduce the correction of star sensor attitude:

Metrical information according to gyro, judge whether the attitude angular velocity of satellite has surpassed the measurement range of gyro, if do not exceed measurement range, then continue to utilize gyro to estimate satellite inertia attitude, change step (1) over to, carry out the satellite inertia attitude prediction of following one-period, if surpass the threshold value of setting according to the gyro to measure scope, then set the sign that gyro transfinites, judge then whether the attitude angular velocity of three on satellite satisfies the service conditions requirement of star sensor, when the service conditions that satisfies star sensor requires, then change step (3) over to, otherwise continue to adopt the gyro to measure data to estimate satellite inertia attitude, change step (1) over to.

(3) introduce star sensor and carry out the attitude correction.

2. the method for self-determination choosing attitude determination mode in the rail control process according to claim 1, it is characterized in that: the gyro in the described step (1) is selected rate for use.

3. the method for self-determination choosing attitude determination mode in the rail control process according to claim 1, it is characterized in that: the inertia attitude in the described step (1) provides with the quaternion form.

4. the method for self-determination choosing attitude determination mode in the rail control process according to claim 1 is characterized in that: the star sensor in the described step (2) is a CCD visual star sensor, and it is that attitude angular velocity is smaller or equal to 0.35 degree/second that its service conditions requires.

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