An on-orbit identification method for gyro nonlinearity and pointing system errors

Abstract

The present invention provides an on-orbit identification method for gyro nonlinearity and pointing system errors, comprising: calculating the platform angular velocity based on star sensor dynamic data; establishing an equation relationship between the gyro system error and the platform angular velocity; constructing observation equations for the nonlinear error and installation error angle of each gyro meter head; using the least squares method to perform on-orbit identification of the gyro meter head system error; and calculating the installation pointing angle of the gyro meter head in the body coordinate system. On-orbit identification of gyro system errors can effectively reduce the attitude angular velocity calculation error and gyro integrated attitude angle calculation error introduced by gyro nonlinearity and installation error in high-dynamic satellites. The present invention can simultaneously identify the nonlinear error and installation angle pointing error of the gyro meter head. The gyro meter head system error identification dimension is high, and the gyro system error estimation accuracy is high. This method can effectively improve the attitude angular velocity measurement accuracy and spatial pointing measurement accuracy of remote sensing satellites under high dynamic conditions.

Description

Technical Field

[0001] The present invention relates to accurate measurement and calculation of spacecraft high-dynamic attitude angular velocity, belongs to the field of spacecraft attitude determination, and specifically relates to an on-orbit identification method for gyroscope nonlinearity and pointing system errors. Background Art

[0002] Gyroscopes are essential components for measuring the attitude and angular velocity of spacecraft. The accuracy of gyro attitude measurement directly impacts the stability of the spacecraft and its dynamic attitude and pointing control capabilities. Gyro measurement errors include gyro constant drift, gyro angle random walk error, gyro nonlinearity, and gyro installation and pointing errors. At low angular velocities, gyro angle random walk noise and gyro constant drift are the primary contributors to errors in spacecraft attitude angular velocity measurement. When the spacecraft is in high-dynamic conditions with high angular velocities, the primary causes of gyro angular velocity errors are gyro nonlinearity and gyro installation errors. Furthermore, under high dynamic conditions, spacecraft typically use gyro angular velocity integration to obtain platform attitude information. While gyro angle random walk noise has little impact on the attitude integral, gyro nonlinearity and installation errors are systematic errors that significantly impact the gyro integrated attitude angle.

[0003] Typically, gyro installation error angles and gyro nonlinearity measurements obtained during ground-based calibration are used to calculate the gyro angular velocity on-orbit. In reality, due to the effects of launch vibrations, long-term thermal deformation during on-orbit operation, and changes in the spacecraft's parameters, the gyro's installation error and nonlinearity values ​​can differ significantly from those on-orbit. This requires re-identification and correction of the gyro's own nonlinear and installation error based on on-orbit measurement data.

[0004] The general gyro system error identification method considers the gyro combination as a whole, identifies the three installation error angles of the orthogonal system, and cannot identify the nonlinear error of the gyro at the same time. Summary of the Invention

[0005] In view of the defects of the prior art, the present invention provides an on-orbit identification method of gyro nonlinearity and pointing system error using dynamic data of star sensors, and an on-orbit identification method of gyro nonlinear error and installation error angle based on dynamic data of star sensors.

[0006] The present invention provides a method for identifying gyro nonlinear errors and pointing system errors, comprising the steps of:

[0007] S1. Calculate the platform angular velocity based on the star sensor dynamic data;

[0008] S2. Establish the equation relationship between the gyro system error and the platform angular velocity;

[0009] S3, constructing the observation equations of the nonlinear error and installation error angle of each gyro head;

[0010] S4. Use the least square method to perform on-orbit identification of the gyro meter system error;

[0011] S5. Calculate the corrected installation orientation of the gyro head in the sensor coordinates.

[0012] Compared with the prior art, the present invention has the following beneficial effects:

[0013] 1. The present invention provides an on-orbit identification method for gyro nonlinearity and pointing system errors using dynamic data from star sensors. The method can simultaneously identify gyro nonlinearity and installation pointing error angles, and the dimensionality of the gyro system error is high. The identification method is simple in structure and the identification results are accurate.

[0014] 2. The on-orbit identification method for gyro nonlinearity and pointing system errors using star sensor dynamic data provided by the present invention performs identification based on the star sensor dynamic data observed by the spacecraft on orbit. This method can accurately obtain the actual on-orbit system error data of the spacecraft's gyroscope without over-reliance on ground calibration information, and has strong engineering practicality. BRIEF DESCRIPTION OF THE DRAWINGS

[0015] Other characteristics, objectives and advantages of the present invention will become more apparent upon reading the following detailed description given by way of non-limiting example with reference to the accompanying drawings.

[0016] Figure 1 This is a flow chart of the on-orbit identification method of gyro nonlinearity and pointing system errors using star sensor dynamic data of the present invention;

[0017] Figure 2 is the installation error angle of the gyro head 1 in the specific embodiment The estimated process curve of

[0018] Figure 3 is the installation error angle of the gyro head 1 in the specific embodiment The estimated process curve of

[0019] Figure 4 is the nonlinear error of the gyro head 1 in the specific embodiment The estimated process curve of

[0020] Figure 5 is the installation error angle of the gyro meter head 2 in the specific embodiment The estimated process curve of

[0021] Figure 6 is the installation error angle of the gyro meter head 2 in the specific embodiment The estimated process curve of

[0022] Figure 7is the nonlinear error of the gyro meter 2 in the specific implementation mode The estimated process curve of

[0023] Figure 8 The gyro meter head 3 installation error angle in the specific embodiment is The estimated process curve of

[0024] Figure 9 The gyro meter head 3 installation error angle in the specific embodiment is The estimated process curve of

[0025] Figure 10 is the nonlinear error of the gyro meter 3 in the specific embodiment Estimated process curve. DETAILED DESCRIPTION

[0026] The present invention will be described in detail below with reference to specific embodiments. The following embodiments will help those skilled in the art further understand the present invention, but are not intended to limit the present invention in any way. It should be noted that a person skilled in the art may make various modifications and improvements without departing from the scope of the present invention. These modifications and improvements are all within the scope of protection of the present invention.

[0027] like Figure 1 As shown, the method for identifying gyro nonlinear error and pointing system error provided by the present invention includes the following steps:

[0028] S1. Calculate the platform angular velocity w based on the star sensor dynamic data b ;

[0029] S2. Establish the gyro system error and platform angular velocity w b The equation relationship of

[0030] S3, constructing the observation equations of the nonlinear error and installation error angle of each gyro head;

[0031] S4. Use the least square method to perform on-orbit identification of the gyro meter system error;

[0032] S5. Calculate the corrected installation orientation of the gyro head in the sensor coordinates.

[0033] In step S1, the deviation quaternion δq(k)=[δq0(k)δq1(k)δq2(k)δq3(k)] is obtained based on the front and back shot data of the star sensor under high dynamic conditions. T , δ symbol represents the meaning of deviation. Then the platform inertial angular velocity w is estimated based on the deviation quaternion b ,

[0034]

[0035] Where, θ(k)=2*arccos(q0(k)), w n (k) represents the scalar magnitude of angular velocity, e x (k) represents the X-direction component of the angular velocity vector, e y (k) represents the Y-direction component of the angular velocity vector, e z (k) represents the Z-direction component of the angular velocity vector, T m (k) represents the sampling period of the star sensor.

[0036] In step S2, the gyro system error and the platform angular velocity w b The equation relationship is as follows

[0037]

[0038] Among them, δ xz , δ xy , δ yz , δ yx , δ zy , δ zx is the gyro installation error angle, w′=(I+Λ)w, w′ represents the angular velocity including the gyro nonlinear error, w′ x , w′ y , w′ z are the components of w′ on the X, Y, and Z axes respectively. w = w gm -b g -η g , w gm is the actual measurement value of the gyroscope, b g is the gyro constant drift, η g is the gyro white noise, I is the identity matrix, Λ is the gyro diagonal scale error matrix, Δ is the gyro installation error matrix, A gb0 It is the transformation matrix of the gyro's initial coordinate system relative to this system.

[0039] In step 3, the satellite performs a high angular velocity attitude maneuver along the gyro axis to be identified. The angular velocity in the same direction as the gyro axis is relatively large, while the angular velocities of the other two axes are both small. The observation equations for the gyro head nonlinearity and the installation error angle are further obtained as follows:

[0040] When the platform rotates around the gyro head 1 (sensor X axis), the installation error angle δ of the gyro head 1 xz , δ xy and nonlinear error λ x The relationship is as follows:

[0041]

[0042]

[0043] where w x 、w y 、w z The X, Y, and Z axis components of w are the three axis components after the constant drift of the measuring gyro is deducted. x , Δw′ y , Δw′ z are the deviation angular velocities of the X, Y, and Z axes, respectively, and are calculated using formula (3).

[0044] Then construct the observation equation:

[0045] z k =h k X+v k (4)

[0046] Among them, the observation quantity z k =[y(1) y(2) y(3)] T , state quantity X=[λ x δ xz δ xy ] T , observation matrix h k =eye(3),v k represents the observation noise.

[0047] When the platform rotates around the gyro head 2 (sensor Y axis), the installation error angle δ of the gyro head 2 yz , δ yx and nonlinear error λ y The relationship is as follows:

[0048]

[0049]

[0050] Then the observation equation is constructed as follows:

[0051] z k =h k X+v k (6)

[0052] Among them, the observation quantity z k =[y(1) y(2) y(3)] T , state quantity X=[δ yz λ y δ yx ] T , observation matrix h k =eye(3).

[0053] When the platform rotates around the gyro head 3 (sensor Z axis), the installation error angle δ of the gyro head 3zy , δ zx and nonlinear error λ z The relationship is as follows:

[0054]

[0055]

[0056] Then the observation equation is constructed as follows:

[0057] z k =h k X+v k (8)

[0058] Among them, the observation quantity z k =[y(1)y(2)y(3)] T , state quantity X=[δ zy δ zx λ z ] T , observation matrix h k =eye(3).

[0059] In step 4, the recursive least squares estimation method is used to estimate the nonlinear error of the gyro installation error angle. The state estimation value obtained by the least squares recursion is

[0060]

[0061] in, K k is the innovation correction matrix,

[0062] P k It is an intermediate parameter and has no specific meaning.

[0063] In step 5, the vector directions of the installation error angle of table head i (i=1, 2, 3) in the sensor coordinate system are calculated as follows:

[0064]

[0065]

[0066]

[0067] Further calculate the vector direction of the gyro head in the body coordinate system, that is, the installation direction of the head i in the body coordinate system A bg0 Represents the conversion matrix from the gyro head to the gyro assembly body coordinate system.

[0068]

[0069] The following provides a specific embodiment of a practical application of the on-orbit identification method of gyro nonlinearity and pointing system errors based on star sensor dynamic data provided by the present invention.

[0070] First, the simulation input for gyro nonlinearity and pointing system error identification using a star sensor is given:

[0071] The three-axis measurement accuracy of the star sensor is 6″, and the gyro angle random walk coefficient is 0.001° / h 1 / 2 , the constant drift estimation residual is 0.01° / h, and the data acquisition period is 0.2s. The vectors of the three gyro heads in the body coordinate system are: Taking gyro heads 1, 2, and 3 as the X, Y, and Z axes of the rectangular coordinate system XmYmZm, the transformation matrix from the satellite system to the rectangular coordinate system is: Installation error angle δ of gyro head 1 xz =80", δ xy =80″, nonlinear error λ x 00.0003; Installation error angle δ of gyro head 2 yz =-100", δ yx =120″, nonlinear error λ x =-0.0003; Installation error angle δ of gyro head 3 zy =200", δ zx =-150″, nonlinear error λ z =0.0003.

[0072] 1. The spacecraft maneuvers at an angular velocity of [0,0,-1]° / s for 60 seconds. The platform attitude angular velocity is obtained based on the dynamic data of the star sensor. Based on the system error observation equation of gyro head 1, a least squares recursive filter is performed with a filtering time of 60 seconds. The estimated value of the installation error angle of gyro head 1 is obtained. Nonlinear error estimate The gyro meter 1 system error estimation process curve is shown in Figures 2 to 4 . We can further obtain the vector direction of header 1 in this system:

[0073]

[0074] 2. The spacecraft maneuvers at an angular velocity of [-1, 0, 0]° / s for 60 seconds. The platform attitude angular velocity is obtained based on the dynamic data of the star sensor. Based on the system error observation equation of gyro head 2, a least squares recursive filter is performed with a filtering time of 60 seconds. The estimated value of the installation error angle of gyro head 2 is obtained. Nonlinear error estimate The error estimation process curve of gyro meter 2 system is shown in Figures 5 to 7 . We can further obtain the vector direction of header 2 in this system:

[0075]

[0076] 3. The spacecraft maneuvers at an angular velocity of [0,1,0]° / s for 60 seconds. The platform attitude angular velocity is obtained based on the dynamic data of the star sensor. Based on the system error observation equation of gyro head 3, a least squares recursive filter is performed with a filtering time of 60 seconds. The estimated value of the installation error angle of gyro head 3 is obtained. Nonlinear error estimate The error estimation process curve of gyro meter 3 system is shown in Figures 8 to 10 . We can further obtain the vector direction of header 3 in this system:

[0077]

[0078] The present invention provides an on-orbit identification method for gyro nonlinearity and pointing errors based on star sensor dynamic data. By treating each gyro head in a system as a separate pointing unit, the spatial pointing error of each gyro head can be identified, increasing the spatial dimension of the gyro head installation error angle and simultaneously identifying the gyro head nonlinearity error.

[0079] This method identifies gyro installation error angles with a higher spatial dimension, can simultaneously identify gyro nonlinear errors, and more comprehensively identifies system errors. This method is simple in construction and provides accurate identification results, effectively improving the accuracy of attitude angular velocity measurements and spatial pointing measurements of remote sensing satellites under high dynamic conditions.

[0080] The above describes the specific implementation of the present invention. It should be understood that the present invention is not limited to the above specific implementation, and those skilled in the art may make various variations or modifications within the scope of the claims, which do not affect the essence of the present invention.

Claims

1. A method for on-orbit identification of gyro nonlinearity and pointing system errors, characterized in that: The steps include: S1. Calculate the platform angular velocity based on the star sensor dynamic data; S2. Establish the equation relationship between the gyro system error and the platform angular velocity; S3, constructing the observation equations of the nonlinear error and installation error angle of each gyro head; S4. Use the least square method to perform on-orbit identification of the gyro meter system error; S5. Calculate the corrected installation orientation of the gyro head in the body coordinate system; In step S5, the gyro head i is installed in the body coordinate system. Where i = 1, 2, 3; A is the vector direction of the gyro head i installation error angle in the sensor coordinate system, bg0 Represents the conversion matrix from the gyro head to the gyro assembly body coordinate system; The vector direction of the gyro head i installation error angle in the sensor coordinate system Among them, δ xz , δ xy , δ yz , δ yx , δ zy , δ zx Mount the gyro with the error angle.

2. The on-orbit identification method for gyro nonlinearity and pointing system errors according to claim 1, wherein: In step S1, the deviation quaternion δq(k)=[δq0(k)δq1(k)δq2(k)δq3(k)] is obtained based on the front and back shot data of the star sensor under high dynamic conditions. T , and then estimate the platform angular velocity w based on the deviation quaternion b , Where, θ(k)=2*arccos(q0(k)), w n (k) represents the scalar magnitude of angular velocity, e x (k) represents the X-direction component of the angular velocity vector, e y (k) represents the Y-direction component of the angular velocity vector, e z (k) represents the Z-direction component of the angular velocity vector, T m (k) represents the sampling period of the star sensor.

3. The on-orbit identification method for gyro nonlinearity and pointing system errors according to claim 1, wherein: In step S2, the gyro installation error angle δ xz , δ xy , δ yz , δ yx , δ zy , δ zx The platform angular velocity w estimated by the star sensor b , the transformation matrix of the gyro initial coordinate system relative to this system The relationship is as follows: Where w′=(I+Λ)w, w′ represents the angular velocity including gyro nonlinear error, w′ x , w′ y 、w z ′ are the components of w′ on the X, Y, and Z axes, respectively, w = w gm -b g -η g , w gm is the actual measurement value of the gyroscope, b g is the gyro constant drift, η g is the gyro white noise, I is the identity matrix, Λ is the gyro diagonal scale error matrix, Δ is the gyro installation error matrix, A gb0 It is the transformation matrix of the gyro's initial coordinate system relative to this system.

4. The on-orbit identification method for gyro nonlinearity and pointing system errors according to claim 3, wherein: In step S3, the satellite is allowed to perform a high angular velocity attitude maneuver along the gyro axis to be identified. The high angular velocity indicates the angular velocity in the same direction as the gyro axis, and the angular velocities of the other two axes are both low angular velocities. The gyro head nonlinearity and installation error angle observation equation is further obtained: z k =h k X+v k Among them, the observation quantity z k =[y(1)y(2)y(3)] T , observation matrix h k =eye(3),v k represents the observation noise; When the platform rotates around the sensor's X axis, the state quantity X=[λ x δ xz δ xy ] T , the installation error angle δ of gyro head 1 xz , δ xy and nonlinear error λ x The relationship is as follows: where w x 、w y 、w z are the X, Y, and Z axis components of w, which are the three axis components after the gyro is measured after deducting the constant drift, Δw′ x , Δw′ y , Δw z ′ are the deviation angular velocities of the X, Y, and Z axes, respectively, and are calculated using formula (3).

5. The on-orbit identification method for gyro nonlinearity and pointing system errors according to claim 4, characterized in that: When the platform rotates around the sensor's Y axis, the state quantity X=[δ yz λ y δ yx ] T , the installation error angle δ of gyro head 2 yz , δ yx and nonlinear error λ y The relationship is as follows:

6. The on-orbit identification method for gyro nonlinearity and pointing system errors according to claim 5, characterized in that: When the platform rotates around the sensor Z axis, the installation error angle δ of the gyro head 3 zy , δ zx and nonlinear error λ z The relationship is as follows:

7. The on-orbit identification method for gyro nonlinearity and pointing system errors according to claim 6, wherein: In step S4, the recursive least squares estimation method is used to estimate the nonlinear error of the gyro installation error angle; the state estimation value obtained by the least squares recursion is in, K k is the innovation correction matrix, P k It is an intermediate parameter with no specific meaning.

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