In-orbit correction method for mounting reference error of imaging satellite gyroscope in motion

Abstract

The invention discloses an in-orbit correction method for gyro installation reference errors of an in-orbit imaging satellite, which comprises the following steps of: performing iterative filtering adjustment on a correction coefficient in real time according to in-orbit data by taking an estimation error and a gyro angular velocity generated in a calculation process of calculating gyro constant drift through Kalman filtering in a conventional satellite in orbit as input; and one-step correction coefficient calculation is combined in the process of calculating the angular velocity of the gyroscope, and the correction coefficient after filtering convergence is substituted into angular velocity calculation, so that the angular velocity error caused by the installation reference error of the gyroscope is corrected in real time.

Description

Technical Field

[0001] The present invention relates to the field of satellite imaging processing, and particularly to an on-orbit correction method for the gyro installation reference error of a satellite with imaging during motion. Background Art

[0002] For a satellite with the ability of agile maneuvering and imaging during motion, in order to achieve high precision and high stability of satellite imaging during motion, it is necessary to correct the gyro installation reference error of the satellite. For the angular velocity error caused by the gyro installation error and the deformation of the whole satellite of traditional steady-state imaging satellites, since the inertial angular velocity is small and stable during the imaging process, its angular velocity error is close to a constant value and can be uniformly calculated into the gyro constant drift. For the installation error, there is currently a method of estimating and correcting the real installation matrix through on-orbit data. The above two methods have certain limitations in the use of agile maneuvering and imaging-during-motion satellites, which are mainly reflected in the following aspects:

[0003] 1) The angular velocity error caused by the gyro installation reference error is related to the magnitude and direction of the satellite inertial angular velocity. The angular velocity error caused by the installation error included in the constant drift calculated under steady state is no longer accurate under large angular velocity maneuvers in any direction.

[0004] 2) For satellites with agile maneuvering requirements, during the maneuvering process, the satellite inertial angular velocity is large, which causes the star sensor data to be invalid. During some time periods of the maneuvering process, the gyro angular velocity integration needs to be used as the attitude reference. The angular velocity error will cause a deviation in the attitude maneuvering effect, resulting in a secondary maneuvering process and affecting the time efficiency of agile maneuvering.

[0005] 3) During the imaging-during-motion process, a high angular velocity tracking accuracy is required to ensure the imaging quality of the camera. Compared with the traditional mode of steady-state imaging after agile maneuvering, there is still a certain angular velocity during the imaging-during-motion process and this angular velocity changes in real time. The angular velocity error will cause a decrease in the angular velocity tracking accuracy, thereby affecting the imaging quality.

[0006] 4) For the currently commonly used method of estimating the installation matrix based on on-orbit data, there is a contradiction that a large amount of data is required for accurate calculation and it cannot be calculated in real time on orbit, while if the amount of data used is small, the calculation accuracy will decrease. Therefore, it is not suitable for on-board real-time calculation and correction. To obtain data, a special device needs to be designed and the ground correction loop is long, resulting in poor timeliness. Summary of the Invention

[0007] The technical problem solved by the present invention is: overcoming the deficiencies of the prior art, providing an on-orbit correction method for the gyro installation reference error of a satellite with imaging during motion, and correcting the angular velocity error caused by the gyro installation reference error of the satellite with imaging during motion.

[0008] The technical solution of the present invention is: an on-orbit correction method for the gyro installation reference error of a satellite with imaging in motion, including:

[0009] S1: Through the theoretical installation matrix of the gyro, the inertial attitude output by the star sensor is converted to the gyro theoretical installation system, and the star sensor attitude q of the gyro theoretical installation system relative to the inertial system is obtained i→s ;

[0010] S2: Calculate the constant drift of the gyro under steady-state conditions and subtract it from the gyro angular velocity. The angular velocity after subtraction is denoted as ω groy ;

[0011] S3: Take A s←s' (k)·ω groy (k) as the angular velocity input, and take the star sensor attitude q i→s (k) in the gyro theoretical installation system as the star sensor attitude input, and perform one iteration using Kalman filtering to obtain the updated estimated attitude and the estimated attitude error Q e (k);

[0012] S4: According to Q e (k) and ω groy (k), perform one correction on the correction coefficient matrix A s←s' (k) to obtain A s←s' (k + 1);

[0013] Take A s←s' (k + 1)·ω groy (k + 1), q i→s (k + 1) as the input, repeat the processes of S3 and S4 for further iteration; after meeting the iteration termination condition, substitute the current A s←s' ×ω groy into the star angular velocity calculation to achieve the correction of the angular velocity error caused by the gyro installation reference for the satellite with imaging in motion.

[0014] Furthermore, in step S1, according to the maneuvering mode of the agile maneuvering satellite, design the theoretical installation position of the gyro so that the gyro is not perpendicular to the angular velocity direction to obtain the theoretical installation matrix A s←b .

[0015] Furthermore, in step S1, based on the theoretical installation matrix A s←b , obtain q b→s by converting the rotation matrix to a quaternion. The inertial attitude output by the star sensor is q i→b , and then obtain the star sensor attitude q i→s of the gyro theoretical installation system relative to the inertial system = q i→b ·q b→s .

[0016] Further, in step S2, under inertial orientation conditions, star sensor-gyroscope combined Kalman filtering is performed to obtain the gyro constant drift under steady-state conditions.

[0017] Further, the specific implementation manner of step S3 is as follows:

[0018] Taking A s←s' (k)·ω groy (k) as the angular velocity estimate value, the quaternion reference value is calculated; according to the quaternion reference value and the star sensor attitude q i→s (k) of the gyro theoretical installation system relative to the inertial system, the deviation quaternion measurement value is obtained, and the vector part of the deviation quaternion measurement value is used as the estimated attitude error Q e (k), combined with the gain filtering coefficient, the estimated attitude is updated.

[0019] Further, in step S4, when correcting the correction coefficient matrix A s←s' (k), the correction amount ΔA s←s' (k) of the error correction coefficient is calculated according to the following formula:

[0020]

[0021] where q e1 , q e2 , q e3 are the vector parts of q e , and the deducted angular velocity ω groy = [ω x ω y ω z , and ω x ω y ω z are the three-axis components of ω groy .

[0022] Further, the iteration termination condition is: when the satellite allows access to this error correction and it is determined that A s←s' has converged to a stable state.

[0023] Further, it is applied to an agile maneuvering satellite with sun orientation in a steady state.

[0024] The present invention also provides an on-board computing terminal, including a memory, a processor, and a computer program stored on the memory and executable on the processor. When the processor executes the computer program, the steps of the method described above are implemented.

[0025] The present invention also provides a computer program product, which implements the steps of the method described above when executed by a processor.

[0026] The advantages of the present invention compared with the prior art are as follows:

[0027] (1) Existing gyro calibration technologies generally collect data based on specific working conditions and perform calibration and correction through off-line calculations on the ground after collection. However, this method directly uses the real-time data of the attitude angular velocity generated during slow rotation or maneuvering in orbit for filtering, and the filtered and converged data is directly used for in-orbit correction, with higher usage efficiency and a wider range of usage scenarios, effectively solving the angular velocity error problem of satellites imaging during motion.

[0028] (2) By using this real-time filtering method, for the in-orbit thermal deformation caused by the influence of the change of in-orbit external heat flux, the influence of the installation error on the imaging effect can be reduced through filtering during the imaging process during motion, while traditional off-line methods cannot calculate the new errors generated by the thermal deformation after off-line calibration.

[0029] (3) This method uses filtering to avoid the problem of limited on-board computing power caused by calculating a large amount of data at one time, and is more suitable for the on-board computing environment with limited computing power. BRIEF DESCRIPTION OF THE DRAWINGS

[0030] Figure 1 It is a schematic diagram of the main process of the method of the present invention. DETAILED DESCRIPTION OF THE INVENTION

[0031] To better understand the technical solution of the present invention, the specific implementation of the present invention will be described below. The main idea of the present invention is to combine the calculation process of calculating gyro drift through Kalman filtering during the in-orbit operation of a conventional satellite, calculate in combination with a one-step correction coefficient during the calculation of the gyro angular velocity, and filter and adjust the correction coefficient in real time according to the in-orbit data, and substitute the filtered and converged correction coefficient into the angular velocity operation. Refer to Figure 1 , which specifically includes the following steps:

[0032] Step 1: According to the common maneuvering modes of agile maneuvering satellites, and based on the principle of avoiding the gyro head being perpendicular to the angular velocity direction, design the theoretical installation position of the gyro head, and denote the theoretical installation matrix as A s←b , and convert the inertial attitude output by the star sensor to the theoretical installation system of the gyro, and denote this attitude value as q i→s .

[0033] Step 2: Use the conventional star sensor-gyro combined Kalman filtering method to calculate and deduct the constant drift of the gyro head under the steady-state working conditions of inertial maintenance or low angular velocity, and denote the deducted angular velocity as ω groy .

[0034] Step 3: Take A s←s' (k)·ω groy (k) as the angular velocity input value, and the star sensor attitude q i→s(k) is the star sensor attitude input. One iteration is performed using the Kalman filter to calculate the updated estimated attitude and the estimated attitude error Q obtained during the process e (k); the correction coefficient matrix A s←s' (k) has an initial value of the unit diagonal matrix;

[0035] Step 4: Based on the estimated error Q e (k) and ω groy (k), perform one correction on A s←s' (k) to obtain A s←s' (k + 1); when correcting the correction coefficient matrix A s←s' (k), calculate the correction amount ΔA of the error correction coefficient according to the following formula s←s' (k):

[0036]

[0037] where q e1 、q e2 、q e3 are the vector parts of q e , and the deducted angular velocity ω groy = [ω x ω y ω z , and ω x ω y ω z are the three-axis components of ω groy .

[0038] Step 5: Use A s←s' (k + 1)·ω groy (k + 1), q i→s (k + 1) as inputs for further iteration, and repeat the processes of Steps 3 and 4.

[0039] Finally, when on-board access to this error correction is allowed and it is determined that A s←s' converges to a relatively stable state, substitute the current A s←s' ×ω groy into the on-board angular velocity calculation to complete the correction of the angular velocity error caused by the gyro installation reference.

[0040] The following takes an agile maneuvering satellite with sun-pointing in the steady state as an example to describe the present invention in detail:

[0041] Step 1: For an agile maneuvering satellite with sun-pointing in the steady state, there is usually a fixed sun-pointing plane. During sun-pointing, the satellite can rotate around its sun-pointing axis. The rotation angular velocity is optimal within a certain range of variable angular velocities, which can be determined by uploading specific rules from the ground or by other constraints (such as specific field-of-view avoidance for cameras, sensors, etc.). Taking the rotation of the satellite around the +Y axis as an example, the head directions of the gyroscopes can be designed such that none of them are in the XOZ plane of the satellite, so that all three head sensors in the installation system have sufficient measured angular velocities during the rotation. If the star sensor can output effective attitude under the angular velocity conditions during in-motion imaging, the working conditions during in-motion imaging can also be used for correction. Similarly, the gyroscope installation should be avoided perpendicular to the angular velocity direction of the typical working conditions of imaging. The theoretical installation matrix is denoted as A s←b 。

[0042] From the theoretical installation matrix A s←b , the quaternion q can be obtained by converting the rotation matrix into a quaternion b→s , the inertial attitude q output by the star sensor i→b , and the star sensor attitude q of the gyroscope installation system relative to the inertial system can be obtained by multiplying the quaternions i→s =q i→b ·q b→s

[0043] Step 2: The constant drift of the gyroscope head can be obtained by performing joint Kalman filtering of the star sensor and gyroscope under inertial orientation conditions The angular velocity after deduction is denoted as ω groy =[ω x ω y ω z 。

[0044] Step 3: Taking as the angular velocity input value and the star sensor attitude q i→s (k) of the gyroscope theoretical installation system relative to the inertial system as the star sensor attitude input q ST (k)=q i→s (k), perform one iteration using the Kalman filtering method. In the full text, (k) represents the kth iteration. The specific calculation process is as follows:

[0045] a) Calculate the quaternion reference value q R

[0046]

[0047] where the estimated attitude initial value is the inertial attitude q i→s (k - 1) measured by the star sensor under the gyroscope installation system in the previous beat;

[0048]

[0049] The relevant parameters are calculated as follows:

[0050]

[0051] In the formula, is the estimated value of the angular velocity; is the three-axis components of; is the 2-norm of; Φ is the Euler axis rotation angle; T c is the system control period, taking a constant value.

[0052] b) Calculate the deviation quaternion measurement error Q e

[0053] The deviation quaternion measurement value q e is calculated as follows:

[0054]

[0055] If the scalar part of the deviation quaternion measurement value q e is negative, then this value is negated, i.e., q e =-q e .

[0056] The deviation quaternion measurement error Q e is calculated as follows:

[0057]

[0058] In the formula, q e1 , q e2 , q e3 are the vector parts of q e .

[0059] c) Estimate the attitude

[0060]

[0061] Step 4: According to the estimation error Q e (k) and ω groy (k), perform a correction on A s←s' (k) to obtain A s←s' (k + 1). The specific calculation process is as follows:

[0062] a) Adjust the filtering gain coefficient k according to Q e (k) A

[0063]

[0064] In the formula, k A0is the designed filter gain coefficient, which is a constant value.

[0065] b) Calculate the correction amount ΔA of the error correction coefficient s←s' (k)

[0066]

[0067] c) Update the error correction coefficient A s←s' (k + 1)

[0068] A' s←s' (k + 1) = A s←s' (k) + ΔA s←s' (k)

[0069] Divide A' s←s' (k + 1) into three column vectors, namely

[0070] A' s←s' (k + 1) = [A1 A2 A3]

[0071] Perform normalization operations on the three column vectors respectively to obtain A s←s' (k + 1)

[0072]

[0073] Iterate steps 3 and 4. When the error correction is allowed on the satellite and it is determined that A s←s' converges to a relatively stable state, substitute the current A s←s' ×ω groy into the satellite angular velocity calculation to complete the correction of the angular velocity error caused by the gyro installation reference.

[0074] It can be understood that the present invention is described by way of examples. Those skilled in the art know that without departing from the spirit and scope of the present invention, various changes or equivalent replacements can be made to these features and examples. Additionally, under the teaching of the present invention, these features and examples can be modified to adapt to specific situations without departing from the spirit and scope of the present invention. Therefore, the present invention is not limited by the specific examples disclosed herein, and the examples that can fall within the scope of the claims of this application all belong to the scope protected by the present invention.

[0075] The content not detailedly described in the specification of the present invention belongs to the well-known technology of those skilled in the art.

Claims

1. An in-orbit correction method for the gyro installation reference error of a satellite imaging in motion, characterized in that, Including: S1: Convert the inertial attitude output by the star sensor to the gyro theoretical installation system through the theoretical installation matrix of the gyroscope, and obtain the star sensor attitude q of the gyro theoretical installation system relative to the inertial system i→s ; S2: Calculate the gyro constant drift under steady-state conditions and subtract it from the gyro angular velocity. The angular velocity after subtraction is denoted as ω groy ; S3: Take A s←s' (k)·ω groy (k) as the angular velocity input, take the star sensor attitude q i→s (k) in the gyro theory installation system as the star sensor attitude input, perform one iteration using the Kalman filter to obtain the updated estimated attitude and the estimated attitude error Q e (k) obtained during the process; S4: According to Q e (k) and ω groy (k), perform one correction on the correction coefficient matrix A s←s' (k) to obtain A s←s' (k + 1); Take A s←s' (k + 1)·ω groy (k + 1), q i→s Take (k + 1) as the input, repeat the processes of S3 and S4 for further iteration; after meeting the iteration termination condition, substitute the current A s←s' ×ω groy into the star angular velocity calculation to achieve the correction of the angular velocity error brought by the gyro installation reference for the imaging satellite in motion.

2. The on-orbit correction method for the gyro installation reference error of the moving medium imaging satellite according to claim 1, wherein: In the step S1, according to the maneuvering mode of the agile maneuvering satellite, the theoretical installation position of the gyroscope is designed so that the gyroscope is not perpendicular to the angular velocity direction, and the theoretical installation matrix A is obtained s←b .

3. The on-orbit correction method for the gyro installation reference error of the moving imaging satellite according to claim 1, characterized in that: In the step S1, based on the theoretical installation matrix A s←b , q is obtained by the method of converting the rotation matrix into quaternion b→s . The inertial attitude output by the star sensor is q i→b . Furthermore, the star sensor attitude q of the gyro theoretical installation system relative to the inertial system is obtained i→s = q i→b · q b→s .

4. The on-orbit correction method for the gyro installation reference error of the moving medium imaging satellite according to claim 1, characterized in that: In the step S2, under the condition of inertial orientation, star sensor and gyroscope integrated Kalman filtering is performed to obtain the gyro constant drift under steady-state working conditions.

5. The on-orbit correction method for the gyro installation reference error of the satellite for moving object imaging according to claim 1, wherein: The specific implementation manner of the step S3 is as follows: With A s←s' (k)·ω groy Taking (k) as the estimated value of angular velocity, calculate the quaternion reference value; according to the star sensor attitude q i→s (k) of the gyro theory installation system relative to the inertial system, obtain the measured value of the deviation quaternion, and take the vector part of the measured value of the deviation quaternion as the estimated attitude error Q e (k), combined with the gain filtering coefficient, update the estimated attitude Update.

6. The on-orbit correction method for the gyro installation reference error of the moving imaging satellite according to claim 1, characterized in that: In the step S4, when correcting the correction coefficient matrix A s←s' (k), the correction amount ΔA s←s' (k) of the error correction coefficient is calculated according to the following formula: where q e1 , q e2 , q e3 is the vector part of q e , and the angular velocity ω groy = [ω x ω y ω z , ω x ω y ω z being the three-axis components of ω groy .

7. The in-orbit correction method for the gyro installation reference error of an imaging satellite in motion according to claim 1, wherein: The iteration termination condition is: the error correction is allowed to be accessed on the satellite and it is judged that A s←s' After converging to a stable state.

8. The on-orbit correction method for the gyro installation reference error of the moving medium imaging satellite according to claim 1, characterized in that: Applied to an agile maneuvering satellite with sun orientation in a steady state.

9. A on-board computing terminal, characterized in that: It includes a memory, a processor, and a computer program stored on the memory and executable on the processor. When the processor executes the computer program, the steps of the method according to any one of claims 1 to 8 are implemented.

10. A computer program product, characterized in that: When the computer program product is executed by the processor, the steps of the method according to any one of claims 1 to 8 are implemented.

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