CN110553667B - Method for carrying out thermal deformation compensation on star sensor - Google Patents
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Abstract
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.
Description
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.
The prior art provides different solutions for the error sources, for example, the accuracy of fusion of the star sensor and other sensors is improved by optimizing a filtering algorithm, the measurement error is compensated by identifying the noise of the star sensor, the structural thermal deformation is reduced by a new material and new technology, and the like. The method for solving the measurement error of the star sensor in the prior art needs a large amount of theoretical research and engineering practice to obtain a certain effect, and has high calculation complexity and complex process.
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 thermal deformation does not occur on the star sensor mounting surface; determining the reference temperature T of the star sensor mounting surface without thermal deformation by measuring whether the star sensor mounting surface deforms or not0；
S3, taking the reference temperature as a starting point, and carrying out temperature rise operation on the star sensor mounting surface; selecting a measuring point on a star sensor mounting surface, and 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; the heating temperature is higher than the reference temperature T0；
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;
s6, monitoring the temperature information of the measuring point when the satellite runs in orbit; when the temperature information exceeds the reference temperature T0And correcting the quaternion of the inertial system measured by the star sensor according to the thermal deformation correction quaternion.
In step S1, the first rectangular coordinate system is an orthogonal rectangular coordinate system and includes mutually perpendicular XrAxis, YrAxis and ZrAxis, said XrAxis, YrAxis, ZrThe axes intersecting at a point Or: wherein XrAxis, YrAxis falling on star sensor mounting surface without thermal deformation, ZrThe axis is perpendicular to the star sensor mounting surface without thermal deformation.
The distortion measurement vector in step S3 is (T)i,αi,βi) (ii) a Where i is the number of measurements, TiFor the temperature measured i, alphaiIs a temperature TiThe measurement points are relative to YrAngle of deformation of the shaft, betaiIs a temperature TiThe measurement points are relative to XrThe deformation angle of the shaft.
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, and alpha and beta are the relative position of the measuring point to Y at the current temperature TrAxis, XrThe deformation angle of the shaft; kTα、KTβRespectively, measured point relative to YrAxis, XrTemperature coefficient of the shaft.
In step S5, the posture conversion matrix is:
converting A into quaternion formula according to direction cosine matrixs→rConverted to a thermal distortion corrected quaternion qs→r。
Inertial system quaternion q measured by the modified star sensor in step S6i→aComprises the following steps:
wherein q isicThe quaternion of the inertial system is obtained by the measurement of the star sensor; q. q.serrInstalling deviation for the star sensor; q. q.ss→bThe star sensor is a theoretical installation matrix from a theoretical installation coordinate system to the main 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.
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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 the star sensor mounting surface without thermal deformation by taking the star sensor theoretical mounting coordinate system as a reference; as shown in FIG. 1, the first rectangular coordinate system comprises X's perpendicular to each otherrAxis, YrAxis and ZrAxis, said XrAxis, YrAxis, ZrThe axes intersecting at a point Or: wherein XrAxis, YrAxis falling on star sensor mounting surface without thermal deformation, ZrThe axis is perpendicular to the star sensor mounting surface without thermal deformation. Recording the first rectangular coordinate system as OrXrYrZr；
S2, performing a thermal test on the ground, heating the mounting surface of the star sensor, and determining the reference temperature T 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 not0；
S3, taking the reference temperature as a starting point, and carrying out temperature rise operation on the star sensor mounting surface; selecting a measuring point on a star sensor mounting surface, and 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; the deformation degree measurement vector is (T)i,αi,βi) (ii) a Where i is the number of measurements, TiFor the temperature measured i, alphaiIs a temperature TiThe measurement points are relative to YrAngle of deformation of the shaft, betaiIs a temperature TiThe measurement points are relative to XrThe deformation angle of the shaft. The heating temperature is higher than the reference temperature T0。
As shown in fig. 2, in the embodiment of the present application, the 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 exceeds the reference temperature T0Deformation occurs, temperature TiLower deformationThe section of the latter measurement point is denoted S. As shown in FIG. 3, the first rectangular coordinate system is around XrAxis of rotation alphaiAfter angling, rewinding YrShaft rotation betaiThe angle forms a second rectangular coordinate system OrXsYsZs. The second rectangular coordinate system is also an orthogonal rectangular coordinate system, including corresponding to X respectivelyrAxis, YrAxis, ZrX of axissAxis, YsAxis, ZsA shaft. XsAxis, YsPlane O formed by the shaftrXsYsFalls on the tangent plane S of the measuring point. Angle alphaiI.e. the temperature TiThe measurement points are relative to YrAngle of deformation of the shaft, betaiIs a temperature TiThe measurement points are relative to XrThe deformation angle of the shaft.
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, and alpha and beta are the relative position of the measuring point to Y at the current temperature TrAxis, XrThe deformation angle of the shaft; kTα、KTβRespectively, measured point relative to YrAxis, XrTemperature coefficient of the shaft.
S5, calculating an attitude transformation matrix A 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 formulas→rThis is prior art;
according to the direction cosine matrix to the quaternionAccording to formula As→rCalculating and generating thermal deformation correction quaternion qs→rThis is the prior art.
S6, monitoring the temperature information of the measuring point when the satellite operates in orbit, and when the temperature information exceeds the reference temperature T0And correcting the quaternion of the inertial system measured by the star sensor according to the thermal deformation correction quaternion.
Inertial system quaternion q measured by the modified star sensor in step S6i→aComprises the following steps:
wherein q isicThe quaternion of the inertial system is obtained by the measurement of the star sensor; q. q.serrInstalling deviation for the star sensor; q. q.ss→bThe star sensor is a theoretical installation matrix from a theoretical installation coordinate system to the main system. The system is a spacecraft self-fixed coordinate system, and a theoretical installation matrix q from a theoretical installation coordinate system to the system of the star sensor is obtaineds→bAre known in the art.
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.
Claims (5)
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 thermal deformation does not occur on the star sensor mounting surface; determining the reference temperature T of the star sensor mounting surface without thermal deformation by measuring whether the star sensor mounting surface deforms or not0；
S3, taking the reference temperature as a starting point, and carrying out temperature rise operation on the star sensor mounting surface; selecting a measuring point on a star sensor mounting surface, and 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; the heating temperature is higher than the reference temperature T0；
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 step S4, 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, and alpha and beta are respectivelyThe measurement point is relative to Y at the current temperature TrAxis, XrThe deformation angle of the shaft; kTα、KTβRespectively, measured point relative to YrAxis, XrTemperature coefficient of the shaft;
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;
s6, monitoring the temperature information of the measuring point when the satellite runs in orbit; when the temperature information exceeds the reference temperature T0And correcting the quaternion of the inertial system measured by the star sensor according to the thermal deformation correction quaternion.
2. The method for compensating for thermal deformation of a star sensor of claim 1, wherein the first rectangular coordinate system is an orthogonal rectangular coordinate system in step S1, and comprises mutually perpendicular X' SrAxis, YrAxis and ZrAxis, said XrAxis, YrAxis, ZrThe axes intersecting at a point Or: wherein XrAxis, YrAxis falling on star sensor mounting surface without thermal deformation, ZrThe axis is perpendicular to the star sensor mounting surface without thermal deformation.
3. The method for compensating for thermal deformation of a star sensor as claimed in claim 1, wherein the distortion measure vector in step S3 is (T)i,αi,βi) (ii) a Where i is the number of measurements, TiFor the temperature measured i, alphaiIs a temperature TiThe measurement points are relative to YrAngle of deformation of the shaft, betaiIs a temperature TiThe measurement points are relative to XrThe deformation angle of the shaft.
4. The method for thermally deforming and compensating the star sensor as claimed in claim 1, wherein the attitude transformation matrix in step S5 is:
converting A into quaternion formula according to direction cosine matrixs→rConverted to a thermal distortion corrected quaternion qs→r。
5. The method for compensating for thermal deformation of a star sensor as claimed in claim 4, wherein the modified quaternion q of the inertia coefficient measured by the star sensor in step S6i→aComprises the following steps:
wherein q isicThe quaternion of the inertial system is obtained by the measurement of the star sensor; q. q.serrInstalling deviation for the star sensor; q. q.ss→bThe star sensor is a theoretical installation matrix from a theoretical installation coordinate system to the main system.
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