CN112158361A - Post-incident high-precision attitude determination method - Google Patents
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Abstract
A post high-precision attitude determination method belongs to the field of spacecraft attitude control and is suitable for a multi-load spacecraft with the requirement of post high-precision attitude determination. The method comprises the steps that firstly, at least two very high-precision star sensors are respectively installed on a plurality of loads; secondly, carrying out planet sensitive reference calibration according to the measurement data of the star sensor on each load; and finally, performing Kalman filtering correction according to the measurement data of the star sensor and the measurement data of the gyroscope on each load, and determining the post-event high-precision attitude of the load.
Description
Technical Field
The invention relates to a post high-precision attitude determination method, in particular to a multi-load multi-mode post high-precision attitude determination method, belongs to the field of spacecraft attitude control, and is suitable for a spacecraft with multiple loads and with the requirement of post high-precision attitude determination.
Background
In the existing post-incident high-precision attitude determination system for the satellite, a plurality of loads are generally subjected to post-incident attitude determination by taking a star sensor as a reference, but due to the influence factors such as ground installation, accurate measurement, on-orbit stress release, temperature and the like; after entering the space, the star sensor or the gyroscope has certain installation deviation and installation deformation, so that the installation precision and the measurement precision of the star sensor or the gyroscope have deviation, and further the subsequent attitude determination precision of a plurality of loads has errors.
Disclosure of Invention
The technical problem to be solved by the invention is as follows: the method comprises the steps of firstly, respectively installing at least two very high-precision star sensors on a plurality of loads; secondly, carrying out planet sensitive reference calibration according to the measurement data of the star sensor on each load; and finally, performing Kalman filtering correction according to the measurement data of the star sensor and the measurement data of the gyroscope on each load, and determining the post-event high-precision attitude of the load.
The purpose of the invention is realized by the following technical scheme:
a post-incident high-precision attitude determination method is used for a spacecraft provided with a plurality of loads and a plurality of gyros, wherein each load is provided with a plurality of star sensors, and the method comprises the following steps:
s1, selecting a reference star sensor for each load, and then calibrating other star sensors on the load; then, carrying out low-frequency error compensation on the reference star sensor;
s2, for any load, selecting a star sensor with effective measurement data on the load; obtaining an optical axis vector and a transverse axis vector of the effective star sensor;
s3, obtaining a filtering initial value by a TRIAD algorithm according to the effective star sensor selected in the S2;
s4, randomly selecting three orthogonal gyroscopes for resolving the three-axis angular velocity of the spacecraft, projecting the obtained three-axis angular velocity of the spacecraft onto the axes of other gyroscopes, and taking the three orthogonal gyroscopes as selected gyroscopes when the projected values of the three-axis angular velocity are equal to the measured values of any other gyroscopes;
and S5, carrying out filtering correction on the post attitude of the spacecraft according to the initial filtering value, the optical axis vector and the horizontal axis vector of the effective star sensor and the measurement data output by the selected gyroscope.
In the post-event high-precision attitude determination method, preferably, all the star sensors on each load are integrally mounted with the load.
Preferably, the post-event high-precision attitude determination method adopts the calibration result of the ground landmark points to perform low-frequency error compensation on the reference star sensor.
In the post-event high-precision attitude determination method, preferably, the attitude determination is performed by using the optical axis vectors of two star sensors, or the attitude determination is performed by using the optical axis vector and the horizontal axis vector of one star sensor, or the attitude determination is performed by using the optical axis vector and the horizontal axis vector after a plurality of star sensors are fused.
Preferably, the effective judgment method for the star sensor measurement data meets one of the following two conditions:
under the condition that the optical axis and the transverse axis data of any two or more star sensor measurement data are consistent, the two or more star sensor measurement data are effective;
and secondly, deviation of an optical axis included angle and a transverse axis included angle of the measurement data of any star sensor in the current period and the previous period does not exceed a theoretical value.
Preferably, the post-event high-precision attitude determination method performs filtering correction on the post-event attitude of the spacecraft by using a kalman filtering method.
Preferably, the post-event high-precision attitude determination method uses an extended kalman filtering method or an unscented kalman filtering method to perform filtering correction on the post-event attitude of the spacecraft.
Preferably, in the post-event high-precision attitude determination method, the star sensor is a very high-precision star sensor.
Preferably, the method for determining the post-event high-precision attitude determines the three-axis inertial angular velocity of the spacecraft by using the measurement data output by the selected gyroscope, and performs filtering correction on the post-event attitude of the spacecraft according to the initial filtering value, the optical axis vector and the horizontal axis vector of the effective star sensor and the three-axis inertial angular velocity of the spacecraft.
Compared with the prior art, the invention has the following beneficial effects:
(1) the method of the invention provides a post-incident high-precision attitude determination method based on the measurement data of a very-high-precision star sensor and a high-precision gyroscope which are integrally installed with a plurality of loads, solves the error caused by attitude determination of a plurality of loads by taking one star sensor as a reference, and improves the post-incident attitude determination precision of the plurality of loads;
(2) and each load selects a reference star sensor integrally installed with the load, and low-frequency error compensation is performed on the reference star sensor of each load according to the calibration result of the ground landmark points, so that the low-frequency influence of the star sensor is reduced, and the post-attitude determination precision of the load is greatly improved.
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FIG. 1 is a flow chart of the steps of the method of the present invention.
Detailed Description
In order to make the objects, technical solutions and advantages of the present invention more apparent, embodiments of the present invention will be described in detail with reference to the accompanying drawings.
A method for determining a post-event high-precision attitude is used for a spacecraft provided with a plurality of loads and a plurality of gyros, all star sensors on each load are integrally installed with the load, as shown in figure 1, a plurality of star sensors are installed on each load, and the star sensors are very high-precision star sensors, and the method comprises the following steps:
s1, selecting a reference star sensor for each load, and then calibrating other star sensors on the load; carrying out low-frequency error compensation on the reference star sensor by adopting a calibration result of the ground landmark point;
s2, for any load, selecting a star sensor with effective measurement data on the load; obtaining an optical axis vector and a transverse axis vector of the effective star sensor; determining the attitude by using the optical axis vectors of two star sensors, or determining the attitude by using the optical axis vector and the transverse axis vector of one star sensor, or determining the attitude by using the optical axis vector and the transverse axis vector after a plurality of star sensors are fused;
s3, obtaining a filtering initial value by a TRIAD algorithm according to the effective star sensor selected in the S2;
s4, randomly selecting three orthogonal gyroscopes for resolving the three-axis angular velocity of the spacecraft, projecting the obtained three-axis angular velocity of the spacecraft onto the axes of other gyroscopes, and taking the three orthogonal gyroscopes as selected gyroscopes when the projected values of the three-axis angular velocity are equal to the measured values of any other gyroscopes;
and S5, determining the three-axis inertial angular velocity of the spacecraft by using the measurement data output by the selected gyroscope, and performing filtering correction on the post attitude of the spacecraft by using an extended Kalman filtering method or an unscented Kalman filtering method according to the initial filtering value, the optical axis vector and the transverse axis vector of the effective star sensor and the three-axis inertial angular velocity of the spacecraft.
The effective judgment method for the star sensor measurement data meets one of the following two conditions:
under the condition that the optical axis and the transverse axis data of any two or more star sensor measurement data are consistent, the two or more star sensor measurement data are effective;
and secondly, deviation of an optical axis included angle and a transverse axis included angle of the measurement data of any star sensor in the current period and the previous period does not exceed a theoretical value.
Example 1:
a post-incident high-precision attitude determination method, namely a multi-load multi-mode post-incident high-precision attitude determination method, comprises the following specific implementation steps:
1) and two very high-precision star sensors are respectively arranged on the plurality of loads.
In order to realize high-precision attitude determination of the front-view camera and the rear-view camera and avoid the influences of ground installation and accurate measurement, on-orbit stress release, temperature, deformation of a load table body and the like, two very-high-precision star sensors are respectively installed on the main bodies of the front-view camera and the rear-view camera. The two star sensors (star sensors are called star sensors for short) integrally installed with the front-view camera are star sensor 1 and star sensor 2, and the two star sensors integrally installed with the rear-view camera are star sensor 3 and star sensor 4.
2) And selecting the reference star sensor of each load, and performing reference calibration of other star sensors according to the reference star sensor measurement data on each load.
The foresight camera takes the star sensor 2 as a reference star sensor and performs reference calibration of the star sensor 1 as an example, and similarly performs reference calibration of the star sensor of the rearview camera. The extrapolated output of the star sensor 1 at the time tst2 can be obtained from the measured output of the star sensor 1 at the time tst1 and the angular velocity of the satellite body system relative to the equatorial inertial system
Based on the extrapolated output
And the mounting matrix C of the star sensor 1SB1The relative inertial system attitude matrix of the star at time tst2 can be obtained
Similarly, the star sensor 2 is arranged according to the measured output at the tst2 moment of the star sensor 2 and the theoretical installation matrix of the star sensor 2
The relative inertial system attitude matrix of the star at the time tst2 can be obtained
Since the relative inertial system attitude matrices of the stars at the same time are equal, i.e. CBI1＝CBI2Then, the error matrix C between the theoretical installation matrix of the star sensor 1 and the calibrated installation matrix can be obtained according to the equationΔWherein
3) Performing low-frequency compensation on the reference star sensor: and compensating the low-frequency error of the reference star sensor by adopting the calibration result of the opposite landmark points. The compensated low frequency error employs a fitting algorithm as follows:
q1＝φkθLFEx；q2＝φkθLFEy；q3＝φkθLFEz；
wherein phik＝[cτ sτ c2τ s2τ c3τ s3τ c4τ s4τ c5τ s5τ]；
C is cos, s is sin; τ ═ ωLFE·t；
Wherein theta isLFExThe specific form of the low-frequency fitting coefficient of the N-order X-axis is as follows:
θLFEx＝[αLFEx1 βLFEx1......αLFExN βLFExN]T，
αLFEx1is a low frequency fitting coefficient, beta, of the order 1 cosine X-axisLFEx1Is a 1-order sine X-axis low-frequency fitting coefficient;
αLFExNis a low frequency fitting coefficient, beta, of the order N cosine X-axisLFExNIs a low-frequency fitting coefficient of an N-order sine X axis;
θLFEythe specific form of the low-frequency fitting coefficient is as follows:
θLFEy＝[αLFEy1 βLFEy1......αLFEyN βLFEyN]T，
αLFEy1is a low frequency fitting coefficient, beta, of the order 1 cosine Y-axisLFEy1Is a 1-order sine Y-axis low-frequency fitting coefficient;
αLFEyNis a low frequency fitting coefficient, beta, of the cosine Y-axis of the order NLFEyNIs a low-frequency fitting coefficient of an N-order sine Y axis;
θLFEzthe specific form of the low-frequency fitting coefficient of the Z axis of the Nth order is as follows:
θLFEz＝[αLFEz1 βLFEz1......αLFEzN βLFEzN]T
αLFEz1is a 1 st order cosine Z-axis low frequency fitting coefficient, betaLFEz1Is a 1-order sine Z-axis low-frequency fitting coefficient;
αLFEzNis a low frequency fitting coefficient, beta, of the cosine Z axis of the order NLFEzNIs a low-frequency fitting coefficient of an N-order sine Z axis; wherein ω isLFEDetermining the Fourier series fundamental frequency of the star sensitive low-frequency error by a ground calibration result, wherein t is the star time; using q1、q2、q3、q4And carrying out low-frequency compensation on the attitude quaternion output by the reference star sensor.
4) And carrying out filtering attitude correction according to the respective star sensor measurement data and gyro measurement data on each load, and carrying out post high-precision attitude determination in a multi-load multi-mode.
Step 1: and judging the validity of the star sensor data for any load, wherein one of the following conditions is met: under the condition I, the optical axis data and the transverse axis data of the two randomly selected star sensor data are consistent, and the two selected star sensor data are effective; and secondly, under the condition that the deviation of the optical axis included angle and the horizontal axis included angle of the measurement data of a certain star sensor and the measurement data of the period on the measurement data does not exceed the theoretical value (in a steady-state mode, the theoretical value is generally 5 degrees), the star sensor is effective.
Step 2: for any attitude determination star sensor of the load camera i: for the load camera 1, the optical axis vectors of the star sensor 1 and the star sensor 2 can be selected for attitude determination, the optical axis and the horizontal axis after the star sensor 1 and the star sensor 2 are fused can also be selected for attitude determination, or the optical axis and the horizontal axis of the star sensor 1 or the star sensor 2 are selected for attitude determination.
Step 3: and (3) determining an initial filtering value: and (3) determining the initial value of the attitude quaternion of the selected camera i by the star sensor selected by the camera by adopting a TRIAD algorithm. The TRIAD algorithm uses the two observation vectors to determine the three-axis attitude of the spacecraft.
Step 4: and (3) carrying out gyro data validity judgment: and selecting three orthogonal gyroscopes to be used for resolving the satellite triaxial angular velocity, projecting the acquired spacecraft triaxial angular velocity onto the axes of other gyroscopes, and when the projection value of the triaxial angular velocity is equal to the measurement value of any other gyroscope, enabling the data of the group of three orthogonal gyroscopes to be effective and participating in determining the satellite triaxial angular velocity.
Step 5: and according to the initial filtering value, the optical axis and the horizontal axis vector of each star sensor and the three-axis inertial angular velocity information of the satellite, selecting an extended Kalman filtering method (EKF) or an unscented Kalman filtering method (UKF) to determine the satellite inertial attitude quaternion and the three-axis constant drift of the gyroscope at the attitude determination time. Firstly, establishing an EKF (extended Kalman Filter) or UKF (unscented Kalman Filter) filtering state equation according to the measurement relation of a satellite sensor and a gyroscope, and taking the inertia attitude quaternion of the satellite and the three-axis constant value drift of the gyroscope as state quantities; secondly, components of the measurement errors of the optical axes measured by the two star sensors in the x and y directions of the star sensors are used as measurement quantities, or components of the measurement errors of the optical axes measured by a single star sensor and the measurement errors of the transverse axes in other two axes of the star sensors are used as measurement quantities; and finally, determining the satellite inertial attitude quaternion and the three-axis constant value drift of the gyroscope in the period according to the inertial attitude quaternion and the three-axis constant value drift of the gyroscope in the previous period, and the optical axis, the horizontal axis vector and the three-axis inertial angular velocity information of each star sensor.
Step 6: if the relative attitude of the load camera i in the specified time period needs to be determined, carrying out quaternion attitude extrapolation by using high-precision gyro data according to the inertia attitude quaternion and gyro constant drift of the high-precision absolute attitude determination in the starting time Step5 of the time period to obtain a relative high-precision attitude determination result of the load. The initial value of the inertial attitude quaternion determined by the relative attitude can also be derived from the calibration attitude of the landmark point.
Example 2:
a multi-load multi-mode post high-precision attitude determination method is characterized in that a satellite payload comprises a front-view camera 1 and a rear-view camera 2, two high-precision star sensors and 6 gyroscopes are respectively installed, and the method is specifically implemented as follows:
1) and two very high-precision star sensors are respectively arranged on the plurality of loads. Two very high precision star sensors are respectively arranged on the main bodies of the front-view camera 1 and the rear-view camera 2. The two star sensors integrally installed with the front-view camera 1 are star sensor 1 and star sensor 2, and the two star sensors integrally installed with the rear-view camera 2 are star sensor 3 and star sensor 4;
2) the front-view camera 1 is calibrated by taking the star sensor 2 as a reference and the rear-view camera 2 is calibrated by taking the star sensor 4 as a reference.
3) Performing low-frequency compensation on the reference star sensor: and compensating the low-frequency error of the reference star sensor by adopting the calibration result of the opposite landmark points.
4) And carrying out filtering attitude correction according to the respective star sensor measurement data and gyro measurement data on each load, and carrying out post high-precision attitude determination in a multi-load multi-mode.
Step 1: and (3) judging the validity of the data of the planetary sensor: the optical axis and the horizontal axis data of the two star sensors of the front-view camera and the rear-view camera are consistent.
Step 2: selecting a attitude determination star sensor mark of a load camera i: the load camera 1 selects the optical axis vectors of the star sensors 1 and 2 for attitude determination, and the load camera 2 selects the optical axis vectors of the star sensors 3 and 4 for attitude determination.
Step 3: and (3) determining an initial filtering value: and (3) determining the initial value of the attitude quaternion of the camera i by the star sensor selected by the camera by adopting a TRIAD algorithm.
The initial quaternion value of the camera 1 is [ -0.053300183008496330, 0.70885674550170863, 0.18658831394084169, 0.67813420939272817 ].
The initial quaternion value of the camera 2 is [ -0.053583398621007070, 0.70880721199223917, 0.18651569274507548, 0.67818364178136592 ].
Step 4: selecting a group of orthogonal gyroscopes 1, 3 and 5 to determine the three-axis angular velocity of the satellite, and judging the validity of the group of gyroscope data: the angular velocity of the satellite triaxial angular velocity calculated by the set of gyros projected onto the 2, 4 or 6 gyroscope axes is consistent with the actual measured angular velocity, so that the set of gyroscope data is effective and participates in the determination of the satellite triaxial angular velocity.
Step 5: and selecting an extended Kalman filtering method (EKF) to perform high-precision attitude filtering correction calculation.
Step 6: and determining the relative attitude of the load camera 1 within 500-520 seconds of a specified time period, and carrying out quaternion attitude extrapolation by using high-precision gyroscope data according to the inertia attitude quaternion determined by the load 1 at the starting time of the time period and the estimated value of gyroscope drift to obtain a relatively high-precision attitude determination result of the load.
Those skilled in the art will appreciate that those matters not described in detail in the present specification are well known in the art.
Although the present invention has been described with reference to the preferred embodiments, it is not intended to limit the present invention, and those skilled in the art can make variations and modifications of the present invention without departing from the spirit and scope of the present invention by using the methods and technical contents disclosed above.
Claims (9)
1. A post high-precision attitude determination method is used for a spacecraft provided with multiple loads and multiple gyros, and is characterized in that each load is provided with multiple star sensors, and the method comprises the following steps:
s1, selecting a reference star sensor for each load, and then calibrating other star sensors on the load; then, carrying out low-frequency error compensation on the reference star sensor;
s2, for any load, selecting a star sensor with effective measurement data on the load; obtaining an optical axis vector and a transverse axis vector of the effective star sensor;
s3, obtaining a filtering initial value by a TRIAD algorithm according to the effective star sensor selected in the S2;
s4, randomly selecting three orthogonal gyroscopes for resolving the three-axis angular velocity of the spacecraft, projecting the obtained three-axis angular velocity of the spacecraft onto the axes of other gyroscopes, and taking the three orthogonal gyroscopes as selected gyroscopes when the projected values of the three-axis angular velocity are equal to the measured values of any other gyroscopes;
and S5, carrying out filtering correction on the post attitude of the spacecraft according to the initial filtering value, the optical axis vector and the horizontal axis vector of the effective star sensor and the measurement data output by the selected gyroscope.
2. A method of post-incident high-accuracy attitude determination according to claim 1, wherein all the star sensors on each load are integrally mounted with the load.
3. The method of claim 1, wherein the calibration of the ground landmark points is used to compensate for low frequency errors of the reference star sensor.
4. The method of claim 1, wherein the attitude determination is performed by using optical axis vectors of two star sensors, or by using an optical axis vector and a horizontal axis vector of one star sensor, or by using an optical axis vector and a horizontal axis vector after the star sensors are fused.
5. The method for determining the post-event high-precision attitude according to any one of claims 1 to 4, wherein the effective judgment method of the star sensor measurement data is to satisfy one of the following two conditions:
under the condition that the optical axis and the transverse axis data of any two or more star sensor measurement data are consistent, the two or more star sensor measurement data are effective;
and secondly, deviation of an optical axis included angle and a transverse axis included angle of the measurement data of any star sensor in the current period and the previous period does not exceed a theoretical value.
6. The method for determining the post-event high-precision attitude according to any one of claims 1 to 4, wherein the post-event attitude of the spacecraft is corrected by filtering using a Kalman filtering method.
7. The method for determining the post-event high-precision attitude according to any one of claims 1 to 4, wherein the post-event attitude of the spacecraft is corrected by filtering using an extended Kalman filtering method or an unscented Kalman filtering method.
8. A post-incident high-precision attitude determination method according to any one of claims 1 to 4, wherein the star sensor is a very high-precision star sensor.
9. A post-event high-precision attitude determination method according to any one of claims 1 to 4, characterized in that the three-axis inertial angular velocity of the spacecraft is determined by using the measurement data output by the selected gyro, and the post-event attitude of the spacecraft is subjected to filtering correction according to the initial filtering value, the optical axis vector and the horizontal axis vector of the effective star sensor, and the three-axis inertial angular velocity of the spacecraft.
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