An attitude measurement method for a spacecraft

Abstract

The present invention belongs to the technical field of attitude measurement, and relates to a combined attitude measurement method of inertia and star sensor, in particular to an attitude measurement method for spacecraft. The method includes: First step, establish a coordinate system and calibrate the relationship matrix between coordinate systems; Second step, calibrate the zero-bias means of the three ring laser gyros; Third step, sample the outputs of the gyro assembly and the star sensor assembly respectively; Fourth step, calculate the rotation angle of the spacecraft to be measured within the sampling time T; Fifth step, obtain the first attitude matrix of the spacecraft to be measured; Sixth step, obtain the second attitude matrix of the spacecraft to be measured; Seventh step, estimate and correct the errors of the gyro assembly; Eighth step, obtain the real-time attitude data of the spacecraft to be measured. The present invention can provide attitude measurement information with high precision, high bandwidth and high real-time performance for spacecraft, and provide technical support for the spacecraft to achieve high-precision earth observation and geometric positioning.

Description

Technical Field

[0001] The present invention belongs to the technical field of attitude measurement, and relates to a combined attitude measurement method of an inertial sensor and a star sensor, and particularly to an attitude measurement method for a spacecraft. Background Art

[0002] The real-time high-precision measurement of the spacecraft attitude is of great significance for its attitude control, navigation, positioning, remote sensing observation, etc. Especially for high-resolution optical imaging satellites, there are higher requirements for their attitude measurement. The main task of high-resolution optical imaging on the spacecraft is high-resolution earth observation, that is, using the optical payload (space imaging camera) carried on the spacecraft to image the ground, and through the stable platform technology, optical imaging technology and image processing technology, to obtain images with as high a resolution as possible, which are mainly applied to fields such as urban planning, road network design, crop yield estimation and disaster prevention and mitigation, providing information guarantee for the implementation of the plan.

[0003] In view of the above problems, the present invention proposes a high-precision, high-bandwidth and high-real-time attitude measurement method with a space four-frequency differential laser gyro as the core and assisted by a star sensor, which can meet the urgent needs in aspects such as the spacecraft earth imaging resolution and spacecraft attitude control, and ultimately significantly improve the attitude control accuracy of China's spacecraft. Summary of the Invention

[0004] The technical problem to be solved by the present invention is: to propose an attitude measurement method for a spacecraft, to solve the problem that there is no high-precision, high-bandwidth and high-real-time attitude measurement information available in the fields of high-resolution earth imaging of the spacecraft and spacecraft attitude control, and to provide technical support for the spacecraft to achieve high-precision earth observation and geometric positioning.

[0005] The technical solution adopted by the present invention is: an attitude measurement method for a spacecraft, including the following steps:

[0006] First step, establish a coordinate system and calibrate the relationship matrix between coordinate systems: fixedly install the combined attitude measurement system on the spacecraft to be measured, the combined attitude measurement system includes a gyro assembly and a star sensor assembly, the gyro assembly includes three four-frequency differential laser gyros, and after installation, the sensitive axes of the three laser gyros are not parallel to each other, the star sensor assembly includes two star sensors, and after installation, the sensitive axes between the two star sensors are not parallel to each other; respectively establish the coordinate system x b y b z b of the spacecraft to be measured, the coordinate system x g y g z g of the gyro assembly, and the coordinate system x i y i z iand the x-axis of the star sensor assembly coordinate system m y m z m , calibrate the x-axis of the gyro assembly coordinate system g y g z g and the relationship matrix between the x-axis of the spacecraft to be measured coordinate system b y b z b Calibrate the x-axis of the star sensor assembly coordinate system m y m z m and the relationship matrix between the x-axis of the spacecraft to be measured coordinate system b y b z b

[0007] Step 2: Calibrate the zero-bias means of the three ring lasers: Calibrate the zero-bias means d 10 、d 20 、d 30 ;

[0008] Step 3: Sample the outputs of the gyro assembly and the star sensor assembly respectively: Sample the output of the star sensor assembly at the sampling frequency f1, sample the output of the gyro assembly at the sampling frequency f2, and satisfy f2 >> f1. Assume the output pulse numbers N1, N2, N3 of the three ring lasers within the sampling time T, and T = 1 / f2;

[0009] Step 4: Calculate the rotation angles of the spacecraft to be measured within the sampling time T: Use the output of the gyro assembly to calculate the rotation angles of the spacecraft to be measured in three directions in the inertial space within the sampling time T

[0010]

[0011] where are the rotation angles of the spacecraft to be measured in the x i y i z i directions within the inertial coordinate system;

[0012] Step 5: Obtain the first attitude matrix of the spacecraft to be measured According to the rotation angles θ k of the spacecraft to be measured from time t k+1 to time t k , use the rotation vector algorithm to calculate the rotation quaternion of the spacecraft to be measured coordinate system from time t k to time t k+1 ​​​ Then, based on the rotation quaternion calculate the attitude quaternion so as to calculate the first attitude matrix of the spacecraft to be measured relative to the inertial space

[0013] Step 6: Obtain the second attitude matrix of the spacecraft to be measured Using the output of the star sensor assembly and the relationship matrix calculate the second attitude matrix of the spacecraft to be measured

[0014]

[0015] wherein is the output of the star sensor assembly;

[0016] Step 7: Estimate and correct the errors of the gyro assembly: Establish the Kalman filter error model of the combined attitude measurement system, compare the second attitude matrix with the first attitude matrix and use the Kalman filter error model to estimate the drift error and attitude angle error of the gyro assembly in real time, and correct the output of the gyro assembly;

[0017] Step 8: Obtain the real-time attitude data of the spacecraft to be measured: Use the drift error estimated in real time to correct the drift of the gyro assembly, and use the attitude angle error estimated in real time to correct the first attitude matrix and finally give the real-time attitude data of the spacecraft to be measured according to the corrected first attitude matrix of the spacecraft to be measured.

[0018] Furthermore, after the combined attitude measurement system is fixedly installed, one sensitive axis of each of the two star sensors is respectively parallel to the sensitive axes of any two of the three ring laser gyros.

[0019] Furthermore, after the combined attitude measurement system is fixedly installed, the three ring laser gyros are perpendicular to each other or form any three faces of a regular tetrahedron.

[0020] Furthermore, the three ring laser gyros adopt space four-frequency differential laser gyros.

[0021] The present invention can provide high-precision, high-bandwidth and high-real-time attitude measurement information for a spacecraft, and provide technical support for the spacecraft to achieve high-precision earth observation and geometric positioning. Description of the Drawings

[0022] Figure 1This is the basic idea block diagram of the attitude measurement method of the present invention. Specific embodiments

[0023] The present invention will be further described below in conjunction with the accompanying drawings, but the protection scope of the present invention should not be limited thereby.

[0024] As Figure 1 shown, an attitude measurement method for a spacecraft includes the following steps:

[0025] In the first step, a coordinate system is established, and the relationship matrix between coordinate systems is calibrated: The combined attitude measurement system is fixedly installed on the spacecraft to be measured. The combined attitude measurement system includes a gyro component and a star sensor component. The gyro component includes three four-frequency differential laser gyros, and after installation, the sensitive axes of the three laser gyros are not parallel to each other. The star sensor component includes two star sensors, and after installation, the sensitive axes of the two star sensors are not parallel to each other; The coordinate system x b y b z b of the spacecraft to be measured, the coordinate system x g y g z g of the gyro component, the inertial coordinate system x i y i z i and the coordinate system x m y m z m of the star sensor component are respectively established, and the relationship matrix g y g z g between the coordinate system x b y b z b of the gyro component and the coordinate system x of the spacecraft to be measured is calibrated, and the relationship matrix m y m z m between the coordinate system x b y b z b of the star sensor component and the coordinate system x

[0026] In the second step, the zero-bias means of the three laser gyros are calibrated: The zero-bias means d 10 、d 20 、d 30 of the three laser gyros are respectively calibrated;

[0027] Step 3: Sample the outputs of the gyro assembly and the star sensor assembly respectively: Sample the output of the star sensor assembly with a sampling frequency f1, and sample the output of the gyro assembly with a sampling frequency f2, and satisfy f2 >> f1. Assume the output pulse numbers N1, N2, and N3 of the three ring lasers during the sampling time T, and T = 1 / f2;

[0028] Step 4: Calculate the rotation angle of the spacecraft under test during the sampling time T: Use the output of the gyro assembly to calculate the rotation angles of the spacecraft under test in three directions in the inertial space during the sampling time T

[0029]

[0030] where, are the rotation angles of the spacecraft under test in the x i y i z i directions in the inertial coordinate system in the x, y, and z directions respectively;

[0031] Step 5: Obtain the first attitude matrix of the spacecraft under test According to the rotation angle θ k of the spacecraft under test from time t k+1 to time t k , use the rotation vector algorithm to calculate the rotation quaternion of the spacecraft under test from time t k to time t k+1 Then, according to the rotation quaternion calculate the attitude quaternion Thus, calculate the first attitude matrix of the spacecraft under test relative to the inertial space

[0032] Use the output of the star sensor assembly and the relationship matrix to calculate the second attitude matrix of the spacecraft under test

[0033]

[0034] where, is the output of the star sensor assembly;

[0035] Step 7: Estimate and correct the error of the gyro assembly: Establish the Kalman filter error model of the combined attitude measurement system, and compare the second attitude matrix with the first attitude matrix Compare, and use the Kalman filter error model to estimate the drift error and attitude angle error of the gyro assembly in real time, and correct the output of the gyro assembly;

[0036] The eighth step is to obtain the real-time attitude data of the spacecraft to be measured: correct the drift of the gyro assembly by using the drift error obtained by real-time estimation, and correct the first attitude matrix by using the attitude angle error obtained by real-time estimation and finally give the real-time attitude data of the spacecraft to be measured according to the corrected first attitude matrix of the spacecraft to be measured. Give the real-time attitude data of the spacecraft to be measured.

[0037] Preferably, the three ring laser gyros are space four-frequency differential ring laser gyros. After the combined attitude measurement system is fixedly installed, the three ring laser gyros are perpendicular to each other or form any three faces of a regular tetrahedron, and one sensitive axis of each of the two star sensors is parallel to the sensitive axes of any two of the three ring laser gyros.

[0038] An attitude measurement method for a spacecraft proposed by the present invention is a high-precision, high-bandwidth, and high-real-time attitude measurement method with a space four-frequency differential ring laser gyro as the core and assisted by star sensors, which can meet the urgent needs in aspects such as the ground imaging resolution of the spacecraft and the attitude control of the spacecraft, and finally significantly improve the attitude control accuracy of our country's spacecraft.

Claims

1. A method for attitude measurement of a spacecraft, characterized in that Including the following steps: Step 1: Establish a coordinate system and calibrate the relationship matrix between coordinate systems. Fix the combined attitude measurement system on the spacecraft to be measured. The combined attitude measurement system includes a gyro assembly and a star sensor assembly. The gyro assembly includes three four-frequency differential laser gyros, and after installation, the sensitive axes of the three laser gyros are not parallel to each other. The star sensor assembly includes two star sensors, and after installation, the sensitive axes of the two star sensors are not parallel to each other. Respectively establish the coordinate system x b y b z b of the spacecraft to be measured, the coordinate system x g y g z g of the gyro assembly, the coordinate system x m y m z m of the star sensor assembly, and the inertial coordinate system x i y i z i . Calibrate the relationship matrix between the coordinate system x g y g z g of the gyro assembly and the coordinate system x b y b z b of the spacecraft to be measured . Calibrate the relationship matrix between the coordinate system x m y m z m of the star sensor assembly and the coordinate system x b y b z b of the spacecraft to be measured Step 2, calibrate the zero-bias means of the three ring laser gyros: Calibrate the zero-bias means d 10 , d 20 , d 30 ; In the third step, sampling is respectively performed on the output of the gyro assembly and the output of the star sensor assembly: the output of the star sensor assembly is sampled at a sampling frequency f1, the output of the gyro assembly is sampled at a sampling frequency f2, and f2 >> f1 is satisfied. Assuming the number of output pulses N1, N2, and N3 of the three ring lasers during the sampling time T, and T = 1 / f2; In the fourth step, calculate the rotation angle of the spacecraft to be measured during the sampling time T: use the output of the gyro assembly to calculate the rotation angles of the spacecraft to be measured in three directions in the inertial space during the sampling time T Among them, are respectively the rotation angles of the spacecraft to be measured in the x i y i z i directions of the inertial coordinate system in the x, y, and z directions; Step 5: Obtain the first attitude matrix of the spacecraft to be measured According to t k moment to t k+1 moment, the rotation angle θ of the spacecraft to be measured k , use the rotation vector algorithm to calculate the rotation quaternion of the coordinate system of the spacecraft to be measured from t k moment to t k+1 moment Then, according to the rotation quaternion calculate the attitude quaternion Thus, calculate the first attitude matrix of the spacecraft to be measured relative to the inertial space Step 6: Obtain the second attitude matrix of the spacecraft to be measured Using the output of the star sensor assembly and the relationship matrix Calculate the second attitude matrix of the spacecraft to be measured Among them, is the output of the star sensor assembly; Step 7, estimate and correct the errors of the gyro assembly: establish the Kalman filter error model of the combined attitude measurement system, and use the second attitude matrix to compare with the first attitude matrix and use the Kalman filter error model to estimate the drift error and attitude angle error of the gyro assembly in real time, and correct the output of the gyro assembly; Step 8: Obtain the real-time attitude data of the spacecraft to be measured: Correct the drift of the gyro assembly by using the drift error estimated in real time, and correct the first attitude matrix by using the attitude angle error estimated in real time. Finally, give the real-time attitude data of the spacecraft to be measured according to the corrected first attitude matrix of the spacecraft to be measured. for correction, and finally, based on the corrected first attitude matrix of the spacecraft to be measured give the real-time attitude data of the spacecraft to be measured.

2. The attitude measurement method for a spacecraft according to claim 1, characterized in that, After the combined attitude measurement system is fixedly installed, one sensitive axis of each of the two star sensors is respectively parallel to the sensitive axes of any two of the three ring lasers.

3. A method for attitude measurement of a spacecraft according to claim 1, characterized in that, After the combined attitude measurement system is fixedly installed, the three ring lasers are perpendicular to each other or form any three faces of a regular tetrahedron.

4. A method for attitude measurement of a spacecraft according to any one of claims 1-3, characterized in that, The three ring lasers adopt a spatial four-frequency differential ring laser gyro.

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