CN-118857309-B - Satellite attitude determination method, system and computer readable medium based on star sensor

CN118857309BCN 118857309 BCN118857309 BCN 118857309BCN-118857309-B

Abstract

The application relates to a satellite attitude determination method, a system and a computer readable medium based on a star sensor, wherein the satellite attitude determination method comprises the steps of establishing a sun-facing orientation coordinate system OX i2 Y i2 Z i2 according to a solar vector in an inertial coordinate system and a position vector in the inertial coordinate system, wherein O represents the mass center of a satellite, an X i2 axis is obtained by cross multiplication of the solar vector and a directional earth center vector, a Z i2 axis points to the negative direction of the solar vector, a Y i2 axis is obtained by an X i2 axis and a Z i2 axis according to a right-hand rule, constructing an installation matrix of the star sensor according to the sun-facing orientation coordinate system, calculating a first included angle between the star sensor and the solar vector and a second included angle between the star sensor and the earth center vector in the sun-facing orientation coordinate system according to the installation matrix, and controlling the attitude of the satellite to enable the first included angle to be kept within a first preset range and the second included angle to be kept within a second preset range, so that the star sensor is effective in the whole orbit period. The application can avoid the influence of light on the field of view of the star sensor.

Inventors

LI JINSONG
WANG LEI
YANG YINGQUAN
ZHOU HENG
Bi Xingzi
LIU BANG
ZHANG YONGHE
ZHANG XIAOFENG
LI DONG
LIU SHUANG

Assignees

中国科学院微小卫星创新研究院
上海微小卫星工程中心

Dates

Publication Date: 20260512
Application Date: 20230428

Claims (10)

1. The satellite attitude determination method based on the star sensor is characterized by comprising the following steps of: Establishing a sun-facing orientation coordinate system OX i2 Y i2 Z i2 according to a sun vector in an inertial coordinate system and a position vector in the inertial coordinate system, wherein O represents a satellite mass center, an X i2 axis is obtained by cross multiplication of the sun vector and a directional earth center vector, a Z i2 axis points to the negative direction of the sun vector, and a Y i2 axis is obtained by the X i2 axis and the Z i2 axis according to a right-hand rule; constructing an installation matrix of the star sensor according to the pair-day directional coordinate system OX i2 Y i2 Z i2 ; Calculating a first included angle between the star sensor and the solar vector and a second included angle between the star sensor and the geocentric vector in the pair of sun-oriented coordinate systems OX i2 Y i2 Z i2 according to the installation matrix; And controlling the attitude of the satellite, so that the first included angle is kept within a first preset range, and the second included angle is kept within a second preset range, thereby enabling the star sensor to be effective in the whole orbit period.
2. The satellite attitude determination method according to claim 1, wherein the step of establishing a pair of daily orientation coordinate systems OX i2 Y i2 Z i2 from the solar vector in the inertial coordinate system and the position vector in the inertial coordinate system further comprises: The solar vector S i under the inertial coordinate system is normalized using the following formula to obtain a normalized solar vector S z : According to the formula Z i2 ＝-S z , the negative direction of the normalized solar vector S z is taken as the Z i2 axis of the pair of solar orientation coordinate systems OX i2 Y i2 Z i2 .
3. The satellite attitude determination method according to claim 2, wherein the step of establishing a pair of daily directional coordinate systems OX i2 Y i2 Z i2 from the sun vector in the inertial coordinate system and the position vector in the inertial coordinate system further comprises using the formula according to the position vector R i in the inertial coordinate system: A unit vector R z pointing to the earth center is established.
4. The satellite attitude determination method according to claim 3, wherein the step of establishing a pair of day orientation coordinate system OX i2 Y i2 Z i2 from the sun vector in the inertial coordinate system and the position vector in the inertial coordinate system further comprises establishing an X i2 axis of the pair of day orientation coordinate system OX i2 Y i2 Z i2 using the following formula: X i ＝S z ×R z wherein S z represents the normalized solar vector, R z represents the unit vector pointing to the earth center, and x represents a vector cross-product operation.
5. The satellite attitude determination method according to claim 4, wherein the step of establishing a pair of day orientation coordinate system OX i2 Y i2 Z i2 from the sun vector in the inertial coordinate system and the position vector in the inertial coordinate system further comprises establishing the Y i2 axis of the pair of day orientation coordinate system OX i2 Y i2 Z i2 using the following formula: Y i ＝Z i2 ×X i2 Wherein Z i2 represents the Z i2 axis of the pair of day-oriented coordinate systems OX i2 Y i2 Z i2 and X i2 represents the X i2 axis of the pair of day-oriented coordinate systems OX i2 Y i2 Z i2 .
6. The method for satellite attitude determination according to claim 5, further comprising constructing an attitude matrix of said satellite according to said pair of daily directional coordinate systems OX i2 Y i2 Z i2 The step of controlling the attitude of the satellite includes controlling the attitude angle of the satellite.
7. The satellite attitude determination method of claim 1, wherein the inertial coordinate system comprises a J2000 equatorial geocentric coordinate system.
8. The satellite attitude determination method of claim 1, wherein the star sensor comprises a first star sensor and a second star sensor, the mounting matrix a bmASTR of the first star sensor comprising: and the second star sensor mounting matrix AbmBSTR includes: Wherein cosd (·) represents an angle cosine value, the numbers in cosd (·) represent angle values.
9. The utility model provides a satellite attitude determination system based on star sensor, set up in the star computer of satellite, its characterized in that, satellite attitude determination system includes: a memory for storing instructions executable by the processor; A processor for executing the instructions to implement the satellite attitude determination method according to any one of claims 1 to 8.
10. A computer readable medium storing computer program code, characterized in that the computer program code, when executed by a processor, implements the satellite attitude determination method according to any one of claims 1-8.

Description

Satellite attitude determination method, system and computer readable medium based on star sensor Technical Field The invention mainly relates to the technical field of aerospace, in particular to a satellite attitude determination method, system and computer readable medium based on a star sensor. Background With the development of aerospace industry, the task modes of the existing satellites are more and more, and are generally divided into earth-oriented satellites, inertial-oriented satellites and sun-oriented satellites. Satellites performing different tasks will use different coordinate systems, for example, earth-directed satellites will use an orbital coordinate system, inertial-directed satellites will use an inertial coordinate system, and earth-directed satellites will use a J2000 coordinate system or a custom coordinate system depending on the needs of the particular task. Satellites often carry a variety of mission loads, such as ground loads including cameras for remotely sensing imaging the ground, sky loads including oversized telescopes for long-term exposure to the sky to view inertial space, and daily loads including imagers for long-term observation of solar activity. For daily load of a main observation two-dimensional solar plane, the main observation two-dimensional solar plane is not sensitive to rotation of a load visual axis, and in the process of observing the sun by the load, the other two coordinate axes perpendicular to the load visual axis are required to be ensured to be relatively stable, so that a fixed attitude determination mode of a satellite is usually maintained for a long time. For conventional satellites, a star sensor is usually required to be installed, the star sensor is a high-precision space attitude measurement device, accurate space azimuth and reference can be provided for aerospace aircrafts such as satellites, and measurement errors among different star sensors can be reduced in an in-orbit calibration, precise temperature control and multi-star sensor information fusion mode. The star sensor is also an optical sensor, and in the process of long-term observation of a space target by satellite load, the field of view of the star sensor is easily influenced by sunlight, earth air light, stray light and the like, and after the field of view of the star sensor is influenced, the star sensor can possibly fail, so that the attitude determination precision and the effectiveness of the star sensor are influenced. In order to improve the reliability of satellite attitude determination, a standby satellite attitude determination scheme is generally set, for example, when all the satellite sensors fail, the satellite can switch to use other attitude sensors to perform attitude determination, and double-vector attitude determination or gyro integral attitude determination can be performed according to the magnetometer and the sun sensor, but the attitude determination modes not based on the satellite sensors may reduce the attitude determination precision of the satellite, so that the on-orbit task of satellite load is affected. In order to ensure the precision of satellite attitude determination, more than two star sensors can be configured on the satellite, and when the field of view of one star sensor is influenced by light to fail, the satellite can select to switch to use the other star sensor. However, in the process of switching different satellite sensors by the satellite, jitter of satellite attitude angles and satellite angular speeds may be generated, and although the jitter may be relieved by on-orbit calibration and the like, the jitter cannot be completely eliminated. Therefore, during satellite operation, it is desirable to keep the star sensor active for the full time, and to not switch the star sensor as much as possible. The coordinate system adopted by the current day directional satellite has the problem that the field of view of the star sensor is possibly influenced by light, so that the working of the star sensor is invalid. Disclosure of Invention The technical problem to be solved by the application is to provide a satellite attitude determination method, a satellite attitude determination system and a computer readable medium based on a satellite sensor, which can avoid that the field of view of the satellite sensor is not influenced by sunlight, earth-atmosphere light, stray light and the like, so that the satellite sensor on the satellite is effective in the whole orbit period. The technical scheme adopted by the application for solving the technical problems is that the satellite attitude determination method based on the star sensor comprises the steps of establishing a sun-facing orientation coordinate system OX i2Yi2Zi2 according to a sun vector in an inertia coordinate system and a position vector in the inertia coordinate system, wherein O represents a satellite centroid, an X i2 axis is obtained by cross multiplication of a sun vector