CN112414398A - Method for calibrating measurement precision of star sensor by on-orbit satellite - Google Patents
Method for calibrating measurement precision of star sensor by on-orbit satellite Download PDFInfo
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- CN112414398A CN112414398A CN202011135311.0A CN202011135311A CN112414398A CN 112414398 A CN112414398 A CN 112414398A CN 202011135311 A CN202011135311 A CN 202011135311A CN 112414398 A CN112414398 A CN 112414398A
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- G—PHYSICS
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- G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
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- G01C21/24—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 specially adapted for cosmonautical navigation
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
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Abstract
The invention discloses a method for calibrating the measurement accuracy of a star sensor by an in-orbit satellite, which aims at the satellite provided with more than three star sensors and comprises the following steps under the condition that a plurality of star sensors are effective: (1) respectively determining attitude angle measurement values of the star sensors; (2) taking optical axis vectors measured by any two star sensors as input, combining a star sensor mounting matrix to carry out double-vector geometric attitude determination, and taking the average value of the attitude determination results of the two groups of closest star sensors at the same time as the attitude angle 'true value' of the satellite; (3) and determining the error between the attitude angle measurement value of each star sensor and the true value of the satellite attitude angle, and calibrating the measurement precision of each star sensor, so as to solve the problem that the star sensor cannot be directly calibrated due to high accuracy of the attitude determination result.
Description
Technical Field
The invention belongs to the field of aerospace measurement and control, and relates to a rapid determination method for calibrating the measurement precision of a single star sensor by using the output data of a plurality of star sensors of an on-orbit satellite to perform double-vector attitude determination. The method is suitable for the precision calibration of the single star sensor by the satellite provided with more than three star sensors.
Background
When the satellite performs on-orbit performance identification on the platform in the early stage of orbit operation or performs on-orbit health evaluation on the platform in the late stage of operation, the measurement data of the star sensor needs to be evaluated. Under the condition that a single star sensor is installed on a satellite, the star sensor has the highest measurement precision, and the technical index conformance verification can be only carried out on the measurement data of the star sensor.
With the development of technology, star sensors with small size and low price are in development trend, and a plurality of star sensors are widely used on a satellite platform. For a satellite with a plurality of star sensors mounted on the star, any two star sensors can be used for double-vector geometric attitude determination, the attitude determination result is used as the true value of the attitude angle of the satellite, the measured value of the attitude angle of each star sensor is compared with the true value of the attitude angle of the satellite, the precision of each star sensor can be calibrated, and the current situation that the coincidence verification is only carried out on the measured data of the star sensors is solved. When the on-orbit performance identification or health evaluation is carried out on the satellite platform, the measurement data of the star sensor needs to be evaluated. Under the condition that a single star sensor is installed on a satellite, the measurement data of the star sensor is the component with the highest precision of the attitude measurement data on the satellite, so that the technical index conformance verification can be only carried out on the measurement data of the star sensor, and the precision calibration cannot be carried out.
Disclosure of Invention
The invention aims to provide a method for calibrating the measurement precision of a star sensor by an on-orbit satellite, which can use a plurality of star sensors to carry out double-vector attitude determination aiming at the on-orbit satellite provided with more than three star sensors and carry out precision calibration on a single star sensor by utilizing the attitude determination result.
The method of the invention has the main ideas that: under the condition that a plurality of star sensors are effective, firstly, the attitude angle measurement value of each star sensor is determined respectively. And secondly, taking optical axis vectors measured by any two star sensors as input, combining a star sensor mounting matrix to carry out double-vector geometric attitude determination, and taking the average value of the attitude determination results of the two groups of closest star sensors at the same time as the true value of the attitude angle of the satellite. And finally, determining the error between the attitude angle measurement value of each star sensor and the true value of the satellite attitude angle, and calibrating the measurement precision of each star sensor.
The adopted technical scheme comprises the following steps:
step one, collecting satellite telemetering data of a platform during an on-orbit test;
processing satellite platform telemetry data to obtain four elements effective for the star time and the plurality of star sensors and the generation time of the four elements;
thirdly, judging the effectiveness of the star sensors according to the satellite platform telemetering data and the quaternary prime data of the star sensors;
under the condition that a plurality of star sensors are effective, optical axis vectors measured by any two star sensors are respectively used as input, and double-vector geometric attitude determination is carried out by combining star sensor installation matrixes;
step five, taking the average value of the combined attitude determination results of two groups of two star sensors which are closest to each other at the same moment as the true value of the attitude angle of the satellite;
respectively determining errors of the attitude angle measurement value of each star sensor and the true value of the satellite attitude angle, determining the average value (E) and the standard deviation (sigma) of the errors, and determining the measurement precision of each star sensor: i E +3 σ.
The invention has the beneficial effects that: the method provided by the invention can be used for carrying out double-vector attitude determination on the on-orbit satellite provided with more than one star sensor, and carrying out precision calibration on a single star sensor by utilizing the attitude determination result, thereby solving the problem that the prior art only carries out technical index conformity verification on the measurement data of the star sensor.
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FIG. 1: flow chart of the method of the invention.
Detailed Description
The present invention will be further described with reference to the accompanying drawings and examples, which include but are not limited to the following examples.
As shown in fig. 1, taking an orbiting satellite equipped with a star sensor 1, a star sensor 2 and a star sensor 3 as an example, the present invention is implemented as follows:
1) and determining a relative inertial system attitude matrix of the star sensor system according to the four elements of the star sensor.
Given that the star sensor output four elements at a certain moment is [ q1 q2 q3 q4], wherein q1 is a scalar quantity, the attitude matrix at the moment is:
an attitude matrix A1 of four-element conversion of the star sensor 1, an attitude matrix A2 of four-element conversion of the star sensor 2 and an attitude matrix A3 of four-element conversion of the star sensor 3 are respectively determined.
2) The roll, pitch, yaw angles of the satellite sensitive measurements are determined (213 order) from the a1, a2, A3 matrices and the satellite sensitive mounting matrix.
(1) Determining a relative inertial system attitude matrix of a satellite body system
Obtaining a relative inertial system attitude matrix A1i1b of the satellite system according to the attitude matrix A1 of the satellite system relative inertial system measured by the star sensor 1 and the installation matrix A1anzhuang of the satellite system by the star sensor 1: a1i1b ═ A1anzhuang ═ A1.
Obtaining a satellite system relative inertia system attitude matrix A2i2b according to a satellite system relative inertia system attitude matrix A2 measured by the satellite sensor 2 and a satellite sensor 2 to a system installation matrix A2 anzhuang: a2i2b ═ A2anzhuang ═ A2.
Obtaining a satellite system relative inertia system attitude matrix A3i3b according to a satellite system relative inertia system attitude matrix A3 measured by a satellite sensor 3 and a satellite sensor 3 tied to a system installation matrix A3 anzhuang: a3i3b ═ A3anzhuang ═ A3.
(2) The attitude matrix of the satellite system relative to the inertial system is converted into a roll angle phi, a pitch angle theta and a yaw angle psi (213 sequence)
Star sensitive 1 measurement angle: phi 1 is arcsin (-a32), theta 1 is arctan (a31/a33), and psi 1 is arctan (a12/a 22). Based on the attitude matrix A1i1b, a32 is the third row, second column element; a31 is the third row, first column element, a33 is the third row, third column element; a12 is the first row and second column element, a22 is the second row and second column element.
Star sensitivity 2 measurement angle: phi 2 is arcsin (-a32), theta 2 is arctan (a31/a33), and psi 2 is arctan (a12/a 22). Based on the attitude matrix A2i2b, a32 is the third row, second column element; a31 is the third row, first column element, a33 is the third row, third column element; a12 is the first row and second column element, a22 is the second row and second column element.
Star sensitivity 3 measurement angle: phi 3 is arcsin (-a32), theta 3 is arctan (a31/a33), and psi 3 is arctan (a12/a 22). Based on the attitude matrix A3i3b, a32 is the third row, second column element; a31 is the third row, first column element, a33 is the third row, third column element; a12 is the first row and second column element, a22 is the second row and second column element.
3) Determining roll, pitch, yaw angles of the system relative to the inertial system using star-sensitive dual vector attitude determination (213)
(1) Standard star sensor A, B double vector posture determination
The attitude matrix (relative inertial system of the system) determined by the star sensor A, B double-vector attitude determination is AAB: AAB MSAB MRABT.
Wherein MSAB [ U1, U1 × U2/| U1 × U2|, U1 × (U1 × U2/| U1 × U2|) ], U1 is the third column vector in A1anzhuang, U2 is the third column vector in A2anzhuang, and U3 is the third column vector in A3 anzhuang; MRAB ═ V1, V1 × V2/| V1 × V2|, V1 × (V1 × V2/| V1 × V2|) ], V1 is the third row vector in a1, V2 is the third row vector in a2, and V3 is the third row vector in A3.
The AAB is converted into roll angle phi, pitch angle theta, yaw angle psi (213 order) measured in dual vectors from star sensor A, B. Phi AB ═ arcsin (-a32), where a32 is the third row, second column element of the AAB matrix; θ AB ═ arctan (a31/a33), where a31 is the third row and first column element of the AAB matrix, and a33 is the third row and third column element of the AAB matrix; ψ AB ═ arctan (a12/a22), a12 is the first row and second column elements of the AAB array, and a22 is the second row and second column elements of the AAB array.
(2) Determining the rolling, pitching and yaw angles of the relative inertial system of the system by determining pairwise double-vector attitude determination of the star sensor 1, the star sensor 2 and the star sensor 3 (213)
According to the standard A, B dual-vector attitude determination step, determining rolling phi 12, pitching theta 12 and yaw angle phi 12 of the star sensor 1 and the star sensor 2 relative to the inertial system respectively; determining rolling phi 23, pitching theta 23 and yaw angle psi 23 of the system relative to an inertial system by using the star sensors 2 and 3; the star sensor 1 and 3 double-vector attitude determination determines the rolling phi 13, the pitching theta 13 and the yaw angle phi 13 of the system relative to the inertial system.
4) Determining true values of attitude angles "
Comparing the three groups of attitude angle results of the three groups of double-vector attitude determination, taking the two groups of results which are closest to each other, and solving the average value as a true value. The determination process is as follows:
(1) determining the two sets of results that are closest
|φ12-φ13|+|θ12-θ13|+|ψ12-ψ13|＝delta1
|φ12-φ23|+|θ12-θ23|+|ψ12-ψ23|＝delta2
|φ23-φ13|+|θ23-θ13|+|ψ23-ψ13|＝delta3
Taking the minimum value of delta1, delta2 and delta3, assuming that delta1 is the minimum, it indicates that star sensors 1 and 2 are the closest to the results of double vector attitude determination of star sensors 1 and 3.
(2) Determining true value of attitude angle measurement "
φREAL＝(φ12+φ13)/2
θREAL＝(θ12+θ13)/2
ψREAL＝(ψ12+ψ13)/2
5) Determining star sensitivity measurement accuracy
(1) Star sensitive 1 measurement error
d_φ1＝φ1-φREAL
d_θ1＝θ1-θREAL
d_ψ1＝ψ1-ψREAL
(2) Star sensitive 2 measurement error
d_φ2＝φ2-φREAL
d_θ2＝θ2-θREAL
d_ψ2＝ψ2-ψREAL
(3) Star sensitive 3 measurement error
d_φ3＝φ3-φREAL
d_θ3＝θ3-θREAL
d_ψ3＝ψ3-ψREAL
(4) Determining the measurement accuracy of each star sensor
Determining the average value (E) and the standard deviation (sigma) of the measurement errors of the star sensors according to the formula (1), and determining the measurement precision of the star sensors: i E +3 σ.
Where Xi is an error value.
Claims (2)
1. A method for calibrating measurement accuracy of star sensors of an in-orbit satellite is characterized by comprising the following steps of: (1) respectively determining attitude angle measurement values of the star sensors; (2) taking optical axis vectors measured by any two star sensors as input, combining a star sensor mounting matrix to carry out double-vector geometric attitude determination, and taking the average value of the attitude determination results of the two groups of closest star sensors at the same time as the attitude angle 'true value' of the satellite; (3) and determining the error between the attitude angle measurement value of each star sensor and the true value of the satellite attitude angle, and calibrating the measurement precision of each star sensor.
2. The method for calibrating the measurement accuracy of the star sensor by the orbiting satellite according to claim 1, wherein: the double-vector geometric pose determination of the step (2) further comprises: and performing double-vector attitude determination by utilizing output data of a plurality of star sensors of the on-orbit satellite, and calibrating the measurement precision of a single star sensor by using the result.
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Cited By (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113063444A (en) * | 2021-04-02 | 2021-07-02 | 北京控制工程研究所 | Method and system for calibrating optical axis measurement reference deviation of sub-arc-second precision star sensor |
| CN113447043A (en) * | 2021-05-21 | 2021-09-28 | 北京控制工程研究所 | GNSS-based satellite astronomical navigation system error autonomous calibration method and system |
| CN113932802A (en) * | 2021-10-12 | 2022-01-14 | 中国科学院微小卫星创新研究院 | Priority changing method and system for multiple star sensors |
| CN114088112A (en) * | 2021-10-27 | 2022-02-25 | 中国空间技术研究院 | Satellite attitude determination precision evaluation method and system |
| CN114140540A (en) * | 2021-12-06 | 2022-03-04 | 长光卫星技术有限公司 | Remote sensing satellite star sensor installation calibration method based on image control points |
| CN114234962A (en) * | 2021-11-10 | 2022-03-25 | 上海航天控制技术研究所 | Multi-star sensor on-orbit thermal deformation correction method, storage medium and electronic equipment |
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- 2020-10-22 CN CN202011135311.0A patent/CN112414398A/en active Pending
Cited By (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113063444A (en) * | 2021-04-02 | 2021-07-02 | 北京控制工程研究所 | Method and system for calibrating optical axis measurement reference deviation of sub-arc-second precision star sensor |
| CN113063444B (en) * | 2021-04-02 | 2024-03-15 | 北京控制工程研究所 | Sub-angle second precision star sensor optical axis measurement reference deviation calibration method and system |
| CN113447043A (en) * | 2021-05-21 | 2021-09-28 | 北京控制工程研究所 | GNSS-based satellite astronomical navigation system error autonomous calibration method and system |
| CN113932802A (en) * | 2021-10-12 | 2022-01-14 | 中国科学院微小卫星创新研究院 | Priority changing method and system for multiple star sensors |
| CN113932802B (en) * | 2021-10-12 | 2024-05-14 | 中国科学院微小卫星创新研究院 | Priority changing method and system for multiple star sensors |
| CN114088112A (en) * | 2021-10-27 | 2022-02-25 | 中国空间技术研究院 | Satellite attitude determination precision evaluation method and system |
| CN114234962A (en) * | 2021-11-10 | 2022-03-25 | 上海航天控制技术研究所 | Multi-star sensor on-orbit thermal deformation correction method, storage medium and electronic equipment |
| CN114234962B (en) * | 2021-11-10 | 2023-09-12 | 上海航天控制技术研究所 | Multi-star sensor on-orbit thermal deformation correction method, storage medium and electronic equipment |
| CN114140540A (en) * | 2021-12-06 | 2022-03-04 | 长光卫星技术有限公司 | Remote sensing satellite star sensor installation calibration method based on image control points |
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