阅读说明：本技术 一种对星敏感器进行热变形补偿的方法 (method for carrying out thermal deformation compensation on star sensor ) 是由 陈桦 完备 杜耀珂 王嘉轶 刘美师 王文妍 贾艳胜 王禹 万亚斌 于 2019-09-11 设计创作，主要内容包括：本发明提供一种对星敏感器进行热变形补偿的方法,包括步骤：S1、在未发生热变形的星敏感器安装面,建立第一直角坐标系；S2、对星敏感器安装面进行温度控制,确定星敏感器安装面不发生热变形的基准温度；S3、以基准温度为起点,对星敏感器安装面进行升温操作,在星敏感器安装面选择一个测量点,测量该测量点在不同温度下相对于第一直角坐标系的变形度,得到若干组变形度测量向量；S4、通过多项式拟合所述若干组变形度测量向量,得到变形度测量向量的温度-形变拟合公式；S5、根据温度-形变拟合公式,计算星敏感器的热变形修正四元数；S6、根据热变形修正四元数修正星敏感器测量的惯性系四元数。本发明简单可靠,提高了卫星在轨姿态测量精度。(the invention provides a method for carrying out thermal deformation compensation on a star sensor, which comprises the following steps: s1, establishing a first rectangular coordinate system on the star sensor mounting surface without thermal deformation; s2, controlling the temperature of the star sensor mounting surface, and determining the reference temperature at which the star sensor mounting surface is not subjected to thermal deformation; s3, taking the reference temperature as a starting point, carrying out temperature rise operation on the star sensor mounting surface, selecting a measuring point on the star sensor mounting surface, and measuring the deformation degree of the measuring point relative to the first rectangular coordinate system at different temperatures to obtain a plurality of groups of deformation degree measuring vectors; s4, fitting the multiple groups of deformation measurement vectors through a polynomial to obtain a temperature-deformation fitting formula of the deformation measurement vectors; s5, calculating a thermal deformation correction quaternion of the star sensor according to a temperature-deformation fitting formula; and S6, correcting the quaternion of the inertial coefficient measured by the star sensor according to the thermal deformation correction quaternion. The method is simple and reliable, and improves the measurement precision of the on-orbit attitude of the satellite.)

1. a method for carrying out thermal deformation compensation on a star sensor is used for correcting an inertia system quaternion measured by the star sensor and is characterized by comprising the following steps:

S1, establishing a first rectangular coordinate system on the star sensor mounting surface without thermal deformation by taking the star sensor theoretical mounting coordinate system as a reference;

S2, performing a thermal test on the ground, performing temperature control on the star sensor mounting surface, and determining a reference temperature at which the star sensor mounting surface does not generate thermal deformation, wherein the reference temperature T ₀ at which the star sensor mounting surface does not generate thermal deformation is determined by measuring whether the star sensor mounting surface generates deformation;

s3, taking the reference temperature as a starting point, heating up the star sensor mounting surface, selecting a measuring point on the star sensor mounting surface, monitoring the deformation degree of the measuring point relative to the first rectangular coordinate system under different heating temperatures to obtain a plurality of groups of deformation degree measuring vectors, wherein the heating temperature is higher than the reference temperature T ₀;

S4, fitting the multiple groups of deformation measurement vectors through a polynomial to obtain a temperature-deformation fitting formula of the deformation measurement vectors;

S5, obtaining an attitude transformation matrix from the actual installation coordinate system of the thermally deformed star sensor to the theoretical installation coordinate system of the star sensor according to the temperature-deformation fitting formula; calculating and generating a thermal deformation correction quaternion according to the attitude transformation matrix;

And S6, monitoring the temperature information of the measuring point when the satellite operates in orbit, and correcting the quaternion of the inertial system measured by the star sensor according to the thermal deformation correction quaternion when the temperature information exceeds the reference temperature T ₀.

2. The method for compensating for the thermal deformation of the star sensor as claimed in claim 1, wherein the first orthogonal coordinate system is an orthogonal coordinate system in step S1, and includes an X _r axis, a Y _r axis and a Z _r axis which are perpendicular to each other, and the X _r axis, the Y _r axis and the Z _r axis intersect at a point O _r, wherein the X _r axis and the Y _r axis are located on the star sensor mounting surface which is not thermally deformed, and the Z _r axis is perpendicular to the star sensor mounting surface which is not thermally deformed.

3. a method for compensating a thermal deformation of a star sensor as claimed in claim 1, wherein the deformation degree measurement vector in step S3 is (T _i, α _i, β _i), where i is the number of measurements, T _i is the temperature of the i-th measurement, α _i is the deformation angle of the measurement point with respect to the Y _r axis at the temperature T _i, and β _i is the deformation angle of the measurement point with respect to the X _r axis at the temperature T _i.

4. the method for heat distortion compensation of a star sensor as claimed in claim 2, wherein in step S4, the sets of distortion measurement vectors are fitted by a first order polynomial, and the temperature-distortion fitting formula is:

Wherein T is the current temperature, alpha and beta are respectively the deformation angles of the measuring point relative to the Y _r axis and the X _r axis under the current temperature T, and K _Tα and K _Tβ are respectively the temperature coefficients of the measuring point relative to the Y _r axis and the X _r axis.

5. the method for thermally deforming and compensating the star sensor as claimed in claim 4, wherein the attitude transformation matrix in step S5 is:

according to the formula of converting the directional cosine matrix into quaternion, converting A _s→r into a thermal deformation correction quaternion q _s→r.

6. The method for compensating for thermal deformation of a star sensor as claimed in claim 5, wherein the quaternion q _i→a of the inertial coefficient measured by the star sensor after the correction in the step S6 is:

Wherein q _ic is an inertia system quaternion obtained by star sensor measurement, q _err is star sensor installation deviation, and q _s→b is a theoretical installation matrix of the star sensor from a theoretical installation coordinate system to a main body system.

Technical Field

the invention relates to the field of spacecraft control, in particular to a method for compensating thermal deformation of a star sensor, which is used for compensating attitude measurement errors of the star sensor caused by thermal deformation of a star sensor mounting surface.

background

With the improvement of satellite attitude control technology, the realization of high-precision attitude determination becomes a basic requirement for a satellite attitude determination system, and the accuracy of a star sensor which is taken as the most main measurement mechanism on the current satellite directly influences the accuracy of the whole attitude determination system.

The attitude determination system using the star sensor as a main measuring mechanism mainly has the following error sources: 1) the sensor has measurement errors such as noise, constant drift, various high and low frequency errors and the like; 2) the correction of the fusion algorithm may cause satellite attitude measurement errors, for example, the attitude determination accuracy finally obtained by selecting different filtering methods or fusion methods is different; 3) the star sensor and the star body have installation deviation, thermal deformation deviation and the like.

Disclosure of Invention

the invention aims to provide a method for carrying out thermal deformation compensation on a star sensor, which is used for calibrating an inertia coefficient quaternion measured by the star sensor. According to the invention, a measuring point is selected on the star sensor mounting surface, and a temperature-deformation fitting formula of the measuring point is obtained through a thermal test. And calculating to obtain a thermal deformation correction quaternion according to the temperature-deformation fitting formula and an attitude conversion matrix from an actual installation coordinate system to a theoretical installation coordinate system of the star sensor after the star sensor is thermally deformed, monitoring the temperature of the measuring point in an on-orbit manner, and performing thermal deformation compensation on the quaternion of the inertial system measured by the star sensor through the thermal deformation correction quaternion to improve the satellite attitude determination precision.

In order to achieve the above object, the present invention provides a method for compensating thermal deformation of a star sensor, which is used for correcting an inertia system quaternion measured by the star sensor, and comprises the following steps:

S1, establishing a first rectangular coordinate system on the star sensor mounting surface without thermal deformation by taking the star sensor theoretical mounting coordinate system as a reference;

s2, performing a thermal test on the ground, performing temperature control on the star sensor mounting surface, and determining a reference temperature at which the star sensor mounting surface does not generate thermal deformation, wherein the reference temperature T ₀ at which the star sensor mounting surface does not generate thermal deformation is determined by measuring whether the star sensor mounting surface generates deformation;

S3, taking the reference temperature as a starting point, heating up the star sensor mounting surface, selecting a measuring point on the star sensor mounting surface, monitoring the deformation degree of the measuring point relative to the first rectangular coordinate system under different heating temperatures to obtain a plurality of groups of deformation degree measuring vectors, wherein the heating temperature is higher than the reference temperature T ₀;

S4, fitting the multiple groups of deformation measurement vectors through a polynomial to obtain a temperature-deformation fitting formula of the deformation measurement vectors;

s5, obtaining an attitude transformation matrix from the actual installation coordinate system of the thermally deformed star sensor to the theoretical installation coordinate system of the star sensor according to the temperature-deformation fitting formula; calculating and generating a thermal deformation correction quaternion according to the attitude transformation matrix;

And S6, monitoring the temperature information of the measuring point when the satellite operates in orbit, and correcting the quaternion of the inertial system measured by the star sensor according to the thermal deformation correction quaternion when the temperature information exceeds the reference temperature T ₀.

In step S1, the first orthogonal coordinate system is an orthogonal coordinate system, and includes an X _r axis, a Y _r axis, and a Z _r axis that are perpendicular to each other, where the X _r axis, the Y _r axis, and the Z _r axis intersect at a point O _r, where the X _r axis and the Y _r axis are located on the star sensor mounting surface that is not thermally deformed, and the Z _r axis is perpendicular to the star sensor mounting surface that is not thermally deformed.

In step S3, the distortion measurement vector is (T _i, α _i, β _i), where i is the measurement frequency, T _i is the temperature of the ith measurement, α _i is the distortion angle of the measurement point relative to the Y _r axis at the temperature T _i, and β _i is the distortion angle of the measurement point relative to the X _r axis at the temperature T _i.

in step S4, the sets of deformation measurement vectors are specifically fitted by a first-order polynomial, and the temperature-deformation fitting formula is:

Wherein T is the current temperature, alpha and beta are respectively the deformation angles of the measuring point relative to the Y _r axis and the X _r axis under the current temperature T, and K _Tα and K _Tβ are respectively the temperature coefficients of the measuring point relative to the Y _r axis and the X _r axis.

In step S5, the posture conversion matrix is:

According to the formula of converting the directional cosine matrix into quaternion, converting A _s→r into a thermal deformation correction quaternion q _s→r.

in step S6, the modified inertial system quaternion q _i→a measured by the star sensor is:

Wherein q _ic is an inertia system quaternion obtained by star sensor measurement, q _err is star sensor installation deviation, and q _s→b is a theoretical installation matrix of the star sensor from a theoretical installation coordinate system to a main body system.

Compared with the prior art, the method carries out thermal deformation compensation on the star sensor by a method of ground test calibration and on-orbit temperature compensation. The deformation quantity of a measuring point of the star sensor mounting surface is measured for multiple times at different temperatures through a ground thermal test, so that the reference temperature of the star sensor mounting surface without deformation and a plurality of groups of deformation measurement vectors during deformation are obtained, and a degree-deformation formula of the deformation measurement vector temperature is obtained by utilizing a polynomial fitting method. When the satellite runs in orbit, the thermal deformation degree of the mounting surface of the star sensor can be calculated only according to the monitored temperature information of the measuring point, and then the quaternion of the inertia coefficient measured by the star sensor is compensated, so that a more accurate satellite attitude is obtained, and the satellite attitude determination precision is improved. According to the invention, the attitude measurement error caused by the thermal deformation of the mounting surface of the on-orbit can be effectively compensated by only adding one temperature measurement point to the mounting point of the star sensor, the compensation method is simple and reliable, and the determination precision of the on-orbit attitude can be effectively improved.

Drawings

In order to more clearly illustrate the technical solution of the present invention, the drawings used in the description will be briefly introduced, and it is obvious that the drawings in the following description are an embodiment of the present invention, and other drawings can be obtained by those skilled in the art without creative efforts according to the drawings:

FIG. 1 is a flow chart of a method for compensating for thermal deformation of a star sensor according to the present invention;

FIG. 2 is a schematic diagram of a first rectangular coordinate system when the star sensor mounting surface is not thermally deformed according to an embodiment of the present invention;

FIG. 3 is a schematic diagram illustrating deformation angles of measurement points when thermal deformation occurs on a star sensor mounting surface according to an embodiment of the present invention;

Detailed Description

The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.

as shown in fig. 1, the present invention provides a method for compensating thermal deformation of a star sensor, which is used for correcting an inertia system quaternion measured by the star sensor, and comprises the following steps:

S1, establishing a first rectangular coordinate system on a star sensor mounting surface which is not subjected to thermal deformation by taking a star sensor theoretical mounting coordinate system as a reference, wherein the first rectangular coordinate system comprises an X _r axis, a Y _r axis and a Z _r axis which are mutually perpendicular, and the X _r axis, the Y _r axis and the Z _r axis intersect at a point O _r, wherein the X _r axis and the Y _r axis are located on the star sensor mounting surface which is not subjected to thermal deformation, and the Z _r axis is perpendicular to the star sensor mounting surface which is not subjected to thermal deformation;

S2, performing a thermal test on the ground, heating the mounting surface of the star sensor, and determining a reference temperature T ₀ at which the mounting surface of the star sensor is not subjected to thermal deformation by measuring whether the mounting surface of the star sensor is deformed or not;

S3, taking the reference temperature as a starting point, heating up the star sensor mounting surface, selecting a measuring point on the star sensor mounting surface, measuring the deformation degree of the measuring point relative to the first rectangular coordinate system under different heating temperatures to obtain a plurality of groups of deformation degree measuring vectors, wherein the deformation degree measuring vectors are (T _i, alpha _i and beta _i), i is the measuring frequency, T _i is the temperature measured at the ith time, alpha _i is the deformation angle of the measuring point relative to the Y _r axis under the temperature T _i, and beta _i is the deformation angle of the measuring point relative to the X _r axis under the temperature T _i, and the heating temperature is higher than the reference temperature T ₀.

in the embodiment of the present application, as shown in fig. 2, a first rectangular coordinate system is established with the measurement point as the intersection point of the first rectangular coordinate system, when the heating temperature of the star sensor mounting surface is deformed beyond a reference temperature T ₀, and the section of the measurement point after deformation at a temperature T _i is denoted as S, as shown in fig. 3, the first rectangular coordinate system is rotated around an axis X _r by an angle α _i, and then rotated around an axis Y _r by an angle β _i, to form a second rectangular coordinate system O _r X _s Y _s Z _s, the second rectangular coordinate system is also a rectangular coordinate system, and includes an X _s axis, an axis Y _s axis, an axis Z _s axis, an axis X _s axis and an axis Y _s axis, which respectively correspond to the axis X _r, the axis Y _r axis, and the axis Z _r, and a plane O _r X _r Y _r falls on the section S of the measurement point, the angle α _r is the deformation angle of the measurement point at a temperature T _r relative to the axis Y _r.

S4, fitting the multiple groups of deformation measurement vectors through a polynomial to obtain a temperature-deformation fitting formula of the deformation measurement vectors; in an application embodiment of the present invention, the sets of deformation measurement vectors are fitted by a first-order polynomial, and the temperature-deformation fitting formula is:

Wherein T is the current temperature, alpha and beta are respectively the deformation angles of the measuring point relative to the Y _r axis and the X _r axis under the current temperature T, and K _Tα and K _Tβ are respectively the temperature coefficients of the measuring point relative to the Y _r axis and the X _r axis.

S5, calculating an attitude transformation matrix A _s→r from the actual installation coordinate system of the star sensor to the theoretical installation coordinate system of the star sensor after thermal deformation according to the temperature-deformation fitting formula, which is the prior art;

according to the formula of converting the directional cosine matrix into quaternion, a thermal deformation modified quaternion q _s→r is generated by calculation according to A _s→r, which is the prior art.

and S6, monitoring the temperature information of the measuring point when the satellite operates in an orbit, and correcting the quaternion of the inertial system measured by the star sensor according to the thermal deformation correction quaternion when the temperature information exceeds the reference temperature T ₀.

In step S6, the modified inertial system quaternion q _i→a measured by the star sensor is:

the method comprises the following steps of obtaining a theoretical installation matrix q _s→b of the star sensor from a theoretical installation coordinate system to a body system, wherein q _ic is an inertia system quaternion obtained by measurement of the star sensor, q _err is installation deviation of the star sensor, and q _s→b is the theoretical installation matrix from the theoretical installation coordinate system to the body system of the star sensor, and the body system is a fixed coordinate system of a spacecraft, so that the theoretical installation matrix q _s→b from the theoretical installation coordinate system to the body system of the star sensor belongs to the prior art in the field.

Compared with the prior art, the method carries out thermal deformation compensation on the star sensor by a method of ground test calibration and on-orbit temperature compensation. The deformation quantity of a measuring point of the star sensor mounting surface is measured for multiple times at different temperatures through a ground thermal test, so that the reference temperature of the star sensor mounting surface without deformation and a plurality of groups of deformation measurement vectors during deformation are obtained, and a degree-deformation formula of the deformation measurement vector temperature is obtained by utilizing a polynomial fitting method. When the satellite runs in orbit, the thermal deformation degree of the mounting surface of the star sensor can be calculated only according to the monitored temperature information of the measuring point, and then the quaternion of the inertia coefficient measured by the star sensor is compensated, so that a more accurate satellite attitude is obtained, and the satellite attitude determination precision is improved. According to the invention, the attitude measurement error caused by the thermal deformation of the mounting surface of the on-orbit can be effectively compensated by only adding one temperature measurement point to the mounting point of the star sensor, the compensation method is simple and reliable, and the determination precision of the on-orbit attitude can be effectively improved.

while the invention has been described with reference to specific embodiments, the invention is not limited thereto, and various equivalent modifications and substitutions can be easily made by those skilled in the art within the technical scope of the invention. Therefore, the protection scope of the present invention shall be subject to the protection scope of the claims.