A method for measuring micro angular vibration of a spacecraft

Abstract

The present invention belongs to the field of precision measurement technology and relates to a method for measuring micro angular vibration for spacecraft. The method includes: first step, calibrating the relationship matrix S between the gyro assembly and its coordinate system and the reference mirror coordinate system g ; second step, calibrating the relationship matrix S between the reference mirror coordinate system and the coordinate system of the spacecraft to be measured b ; third step, sampling the output of the gyro assembly; fourth step, calculating the rotation angle of the spacecraft to be measured; fifth step, obtaining the time-domain information of the micro angular vibration of the spacecraft to be measured; sixth step, obtaining the frequency-domain information of the micro angular vibration of the spacecraft to be measured. The high-frequency and high-precision micro angular vibration measurement method using a spatial four-frequency laser gyro as the core provided by the present invention can meet the urgent needs in aspects such as spacecraft micro vibration measurement, and ultimately significantly improve the earth observation ability and level of the spacecraft.

Description

Technical Field

[0001] The present invention belongs to the technical field of precision measurement, and relates to a method for measuring micro angular vibration for a spacecraft. Background Art

[0002] The main task of high-resolution optical imaging on a spacecraft is high-resolution earth observation, that is, using an optical payload (space imaging camera) carried on the spacecraft to image the ground, and through stabilization platform technology, optical imaging technology, and image processing technology, to obtain an image with as high a resolution as possible, which is mainly applied to fields such as national land census, urban planning, land right confirmation, road network design, crop yield estimation, and disaster prevention and reduction.

[0003] Currently, with the increasing requirement for the imaging resolution of earth observation, the limitation of spacecraft micro-vibration on the imaging resolution of its carried optical payload has become increasingly significant. The micro-vibration effect will seriously affect the pointing accuracy and attitude stability of the optical payload, causing problems such as image blurring and geometric distortion of the optical payload, and has become one of the important factors affecting the image quality of the optical payload. Therefore, it is necessary to use new sensors or new methods to measure the micro-vibration information of the spacecraft with high precision.

[0004] In view of the above problems, the present invention proposes a high-frequency and high-precision micro angular vibration measurement method with a four-frequency differential laser gyro as the core, which can meet the urgent needs in aspects such as spacecraft micro-vibration measurement, and ultimately significantly improve the earth observation ability and level of the spacecraft. Summary of the Invention

[0005] The technical problem to be solved by the present invention is: to provide a method for measuring micro angular vibration for a spacecraft, to solve the problem that there is no available high-frequency and high-precision micro-vibration measurement information for the spacecraft, and to provide an important technical support for the spacecraft to achieve high-precision earth observation and geometric positioning.

[0006] The technical solution adopted by the present invention is: a method for measuring micro angular vibration for a spacecraft, comprising the following steps:

[0007] The first step is to calibrate the relationship matrix S between the gyro assembly and its coordinate system and the coordinate system of the reference mirror g : Establish a gyro assembly coordinate system and a reference mirror coordinate system. The gyro assembly includes three mutually orthogonal four-frequency differential laser gyros. The gyro assembly and the reference mirror are called an angular vibration measurement system, and the two are fixedly connected. Calibrate the zero-bias means d 10 、d 20 、d 30 of the three mutually orthogonal four-frequency differential laser gyros, and calibrate the relationship matrix S between the gyro assembly coordinate system and the reference mirror coordinate system g ;

[0008] In the second step, calibrate the relationship matrix S between the reference mirror coordinate system and the coordinate system of the spacecraft under test b : Establish the x-axis of the coordinate system of the spacecraft under test b y-axis b z-axis b , fixedly install the calibrated angular vibration measurement system on the spacecraft under test, and calibrate the relationship matrix S between the reference mirror coordinate system and the coordinate system of the spacecraft under test b ;

[0009] In the third step, sample the output of the gyro assembly: Set the sampling time T, and continuously sample the output pulse numbers N1, N2, and N3 of the three mutually orthogonal four-frequency differential laser gyros within the sampling time T

[0010] In the fourth step, calculate the rotation angle of the spacecraft under test: The rotation angles of the x-axis b y-axis b z-axis b of the coordinate system of the spacecraft under test relative to the inertial space in three directions within the sampling time T

[0011]

[0012] In the fifth step, obtain the micro-angular vibration time-domain information of the spacecraft under test: Subtract the average rotation angle within the nT time from the rotation angles within n sampling times T to obtain the micro-angular vibration time-domain information of the spacecraft under test

[0013] In the sixth step, obtain the micro-angular vibration frequency-domain information of the spacecraft under test: Perform frequency-domain analysis on the micro-angular vibration time-domain information of the spacecraft under test obtained in the fifth step to obtain the frequency and amplitude information of the micro-angular vibration of the spacecraft under test

[0014] Furthermore, the three mutually orthogonal four-frequency differential laser gyros in the first step adopt space four-frequency differential laser gyros

[0015] Furthermore, the sampling of the output of the gyro assembly in the third step adopts pulse subdivision technology

[0016] The advantages of the present invention are as follows: The present invention provides a method for measuring micro-angular vibration of a spacecraft, which solves the problem that there is no available high-frequency and high-precision angular vibration measurement information for the spacecraft. At the same time, this method has the advantages of wide measurement bandwidth, high angular resolution, and wide application range Description of the Drawings

[0017] Figure 1 is the basic idea block diagram of the micro-angular vibration measurement method of the present invention

[0018] Figure 2 is the schematic diagram of the principle of the pulse subdivision technology of the present invention Detailed Embodiments ​​​

[0019] 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.

[0020] As Figure 1 shown, a method for measuring micro angular vibration of a spacecraft includes the following steps:

[0021] In the first step, calibrate the relationship matrix S between the gyro assembly and its coordinate system and the coordinate system of the reference mirror g : Establish the coordinate system of the gyro assembly and the coordinate system of the reference mirror. The gyro assembly includes three mutually orthogonal four-frequency differential laser gyros. The gyro assembly and the reference mirror are called the angular vibration measurement system, and the two are fixedly connected. Calibrate the zero-bias means d 10 、d 20 、d 30 of the three mutually orthogonal four-frequency differential laser gyros, and calibrate the relationship matrix between the coordinate system of the gyro assembly and the coordinate system of the reference mirror

[0022]

[0023] In the second step, calibrate the relationship matrix S between the coordinate system of the reference mirror and the coordinate system of the spacecraft to be measured b : Establish the coordinate system x b y b z b of the spacecraft to be measured, fixedly install the calibrated angular vibration measurement system on the spacecraft to be measured, and calibrate the relationship matrix between the coordinate system of the reference mirror and the coordinate system of the spacecraft to be measured

[0024]

[0025] In the third step, sample the output of the gyro assembly: Set the sampling time T, and continuously sample the output pulse numbers N1, N2, and N3 of the three mutually orthogonal four-frequency differential laser gyros within the sampling time T;

[0026] In the fourth step, calculate the rotation angle of the spacecraft to be measured: The coordinate system x b y b z b of the spacecraft to be measured rotates relative to the inertial space by rotation angles

[0027]

[0028] wherein, are respectively the rotation angles of the three mutually orthogonal four-frequency differential laser gyros within the coordinate system of the reference mirror, and θ 1 、θ 2 、θ 3They are the rotation angles of the three mutually orthogonal four-frequency differential laser gyros within the coordinate system of the gyro assembly, respectively.

[0029] Step 5: Obtain the time-domain information of the micro angular vibration of the spacecraft to be measured: Subtract the average value of the rotation angles within the n sampling times T from the rotation angles within the time of nT, and the time-domain information of the micro angular vibration of the spacecraft to be measured can be obtained.

[0030] Step 6: Obtain the frequency-domain information of the micro angular vibration of the spacecraft to be measured: Perform frequency-domain analysis on the time-domain information of the micro angular vibration of the spacecraft to be measured obtained in Step 5 to obtain the frequency and amplitude information of the micro angular vibration of the spacecraft to be measured.

[0031] Preferably, in Step 1, the four-frequency differential laser gyro uses a space four-frequency differential laser gyro, and in Step 3, pulse subdivision technology is used for sampling the output of the gyro assembly.

[0032] As Figure 2 shown, the method of using pulse subdivision technology is as follows: Assume that there are N pulse rising edges within the sampling time T (the sampling frequency is f s ). To measure the times of t 00 , t 01 , t 10 and t 11 , use a high-frequency signal with a frequency of f0. The number of pulses corresponding to the time within this high-frequency signal is measured as n 00 , n 01 , n 10 and n 11 respectively. Then the count within the sampling time T is

[0033] (N0 + N 00 - N 11 ) = N0 + n 00 / n 01 - n 10 / n 11 (4)

[0034] A method for measuring micro angular vibration of a spacecraft according to the present invention has the advantages of wide measurement bandwidth, high angular resolution, and wide application range, and can provide high-frequency and high-precision angular vibration measurement information for the spacecraft.

Claims

1. A method for measuring micro angular vibration of a spacecraft, characterized in that, It includes the following steps: The first step is to calibrate the relationship matrix S between the gyro assembly and its coordinate system and the reference mirror coordinate system g : Establish the gyro assembly coordinate system and the reference mirror coordinate system. The gyro assembly includes three mutually orthogonal four-frequency differential laser gyros. The gyro assembly and the reference mirror are called the angular vibration measurement system, and the two are fixedly connected. Calibrate the zero-bias means d 10 , d 20 , d 30 of the three mutually orthogonal four-frequency differential laser gyros, and calibrate the relationship matrix S between the gyro assembly coordinate system and the reference mirror coordinate system g ; Step 2: Calibrate the relationship matrix S between the reference mirror coordinate system and the coordinate system of the spacecraft to be measured b : Establish the x-axis b y-axis b z-axis b of the coordinate system of the spacecraft to be measured, fixedly install the calibrated angular vibration measurement system on the spacecraft to be measured, and calibrate the relationship matrix S between the reference mirror coordinate system and the coordinate system of the spacecraft to be measured b ; In the third step, sample the output of the gyro assembly: Set the sampling time T, and continuously sample the output pulse numbers N1, N2, and N3 of the three mutually orthogonal four-frequency differential laser gyros within the sampling time T. Step 4, calculate the rotation angle of the spacecraft under test: The x-axis of the coordinate system of the spacecraft under test b y b z b Rotation angles in three directions relative to the inertial space within the sampling time T In the fifth step, obtain the micro angular vibration time-domain information of the spacecraft to be measured: Subtract the average rotation angle within the nT time from the rotation angle within n sampling times T, and the micro angular vibration time-domain information of the spacecraft to be measured can be obtained. In the sixth step, obtain the micro angular vibration frequency-domain information of the spacecraft to be measured: Perform frequency-domain analysis on the micro angular vibration time-domain information of the spacecraft to be measured obtained in the fifth step to obtain the frequency and amplitude information of the micro angular vibration of the spacecraft to be measured.

2. The micro angular vibration measurement method for a spacecraft according to claim 1, wherein, The three mutually orthogonal four-frequency differential laser gyros described in the first step adopt space four-frequency differential laser gyros.

3. A micro angular vibration measurement method for a spacecraft according to claim 1 or 2, characterized in that, The sampling of the output of the gyro assembly described in the third step adopts pulse subdivision technology.

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