Gyroscope on-orbit calibration method based on attitude maneuver

Abstract

The application provides a gyro on-orbit calibration method based on attitude maneuver. The attitude angle measured by a star sensor is used as the attitude angle reference before and after the satellite platform attitude rotation maneuver. The attitude angle of the platform attitude rotation maneuver is calculated by gyro integration. The scale factor of the gyro is calculated by the deviation of the platform attitude rotation maneuver attitude angle calculated by the star sensor and the gyro. The application provides sensitive data for the gyro through the satellite platform attitude rotation maneuver, and uses the star sensor with higher measurement precision than the gyro as the measurement reference to calibrate and correct the scale factor of the gyro, effectively solves the problem of the long-term on-orbit performance parameter decline of the gyro, and provides high-precision measurement guarantee for the high-precision and high-stability control of the long-life satellite platform and the rapid stable attitude maneuver control.

Description

Technical Field

[0001] This invention relates to the field of on-orbit performance calibration technology for satellite platform-supporting sensors, specifically to an on-orbit calibration method for gyroscopes based on attitude rotation maneuvers. Background Technology

[0002] As spacecraft mission requirements continue to increase, the demands on satellite platform performance and lifespan also rise. High-precision pointing, high-stability attitude control, and rapid, stable attitude maneuver control (hereinafter referred to as high-precision, high-stability control and rapid attitude maneuver control) over a long lifespan (8-15 years) are fundamental functions of the platform. Star sensors and gyroscopes are essential measurement sensors for satellite platforms to achieve these functions, providing the platform with the attitude angle and attitude angular velocity measurement information required for attitude control. Gyroscopes are inertial sensors, and their performance parameters (zero-point offset, scaling factor, etc.) change over time. To ensure the satellite platform's attitude control performance, the gyroscope's performance parameters need to be calibrated and corrected in orbit to ensure that its performance meets requirements throughout its entire lifespan, providing support and assurance for the platform to complete its missions. Summary of the Invention

[0003] The purpose of this invention is to provide an on-orbit calibration method for gyroscopes based on attitude maneuvers. This method can solve the problem of decreased measurement accuracy caused by changes in the scaling factor during long-term on-orbit flight of gyroscopes. By calibrating and correcting the scaling factor of the gyroscope in orbit, the measurement accuracy of the gyroscope can meet the performance index requirements throughout its entire life cycle. This provides the platform with measurement information for high-precision and high-stability control and rapid and stable attitude maneuver control, thus ensuring that the spacecraft can complete its mission.

[0004] To achieve the above objectives, the present invention is implemented through the following technical solution:

[0005] A method for on-orbit calibration of a gyroscope based on attitude maneuvering, wherein the attitude maneuvering refers to attitude maneuvering control that rotates a full revolution at a constant rotational angular velocity within a certain time period, the method comprising:

[0006] S1, the satellite platform maintains static stability control, collects the measurement data of the star sensor to calculate the platform's attitude angle, and uses it as the attitude angle P0 before the platform's attitude rotation maneuver;

[0007] S2, the satellite platform maintains the target attitude angular velocity ω t A positive attitude rotation maneuver is performed around the calibration axis. During the rotation, the output value of the gyroscope is collected. Using the attitude angle P0 before the platform's attitude rotation maneuver as the initial value, the attitude angle change P of the platform's attitude rotation maneuver is calculated through gyroscope integration. gyro ;

[0008] S3, the satellite platform maintains static stability control, collects measurement data from the star sensor to calculate the platform's attitude angle, which is used as the attitude angle P1 after the platform's attitude rotation maneuver; the attitude angle P after the attitude rotation maneuver is calculated by subtracting P0 from P1. ST As the reference attitude angle for platform attitude rotation maneuvering;

[0009] S4, Calculate the attitude angle P determined by the star sensor. ST The attitude angle P determined by the gyroscope gyro The deviation attitude angle ΔP between the two is calculated, and the deviation of the positive scaling factor Δk of the calibration axis is calculated. + =△P / ω t ;

[0010] S5, Repeat steps S1 to S4 to obtain the calibration axis negative scaling factor deviation Δk. - In step S2, the target attitude angular velocity ω is used to revolve around the calibration axis. t Perform a negative rotation;

[0011] S6, take the calibration axis positive scale factor deviation Δk + and negative scaling factor deviation Δk - The average value is used as the calibration axis scale factor correction amount Δk. 轴 ;

[0012] S7, the calibration axis scale factor correction amount Δk calculated in step S6. 轴 The gyroscope calibration factor is corrected to obtain a new scaling factor for the calibration axis.

[0013] Optionally, the attitude angular velocity used for the gyroscope integration is a measured value of the gyroscope output after zero-position offset correction.

[0014] Optionally, the satellite platform maintains static stability control, using the attitude angle information output by the acquired star sensor as the platform attitude angle reference, and performs on-orbit calibration and correction of the gyroscope's zero-position offset, specifically including:

[0015] Step a: The satellite platform maintains static stability control, using the attitude angle information output by the acquired star sensor as the platform attitude angle reference;

[0016] Step b: Calculate the quaternion q using the angular velocity output by the gyroscope. gyro ;

[0017] Step c, calculate the star sensor output quaternion q st Quaternion q calculated with a gyroscope gyro The deviation quaternion Δq between them is calculated, and the deviation quaternion Δq is normalized.

[0018] Step d: Filter the vector part of the deviation quaternion to obtain the zero offset ω0 of the gyroscope;

[0019] Step e involves correcting the gyroscope's measurement output using the calibrated gyroscope zero-position offset.

[0020] Optionally, the calibration axes are the X-axis, Y-axis, and Z-axis.

[0021] Optionally, the satellite platform maintains static and stable control of its orientation towards the Earth, wherein the platform stability is not less than 0.05° / s, the star sensor output attitude angle measurement accuracy is not less than 60″, the sampling period T is not less than 200ms, and the duration is not less than 300s.

[0022] Optionally, the attitude rotation maneuver control refers to the platform rotating 3 times at the target angular velocity, with the angular velocity control error during the rotation process not exceeding 0.05° / s.

[0023] Optionally, the star sensor has a measurement accuracy at least one order of magnitude higher than that of the calibrated gyroscope.

[0024] Optionally, the target attitude angular velocity ω t Not less than 10% of the gyroscope's measurement range.

[0025] Compared with the prior art, the present invention has the following advantages:

[0026] The method employed in this invention solves the problem of performance degradation caused by long-term on-orbit operation of gyroscopes. By calibrating and correcting the gyroscope's zero-position offset and scaling factor, high-precision and stable control, as well as rapid and stable attitude maneuver control, can be guaranteed throughout the spacecraft platform's entire life cycle, providing assurance for mission implementation. Attached Figure Description

[0027] To more clearly illustrate the technical solution of the present invention, the accompanying drawings used in the description will be briefly introduced below. Obviously, the drawings described below are one embodiment of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort:

[0028] Figure 1 A flowchart of an on-orbit calibration method for a gyroscope based on attitude maneuvering provided by the present invention;

[0029] Figure 2 This is a diagram illustrating the single-axis scale factor calibration process.

[0030] Figure 3 Diagram showing the calibration sequence for the three axes of a gyroscope;

[0031] Figure 4 This is a diagram of a single-axis attitude rotation maneuver. Detailed Implementation

[0032] The present invention will be further described in detail below with reference to the accompanying drawings and specific embodiments. The advantages and features of the present invention will become clearer from the following description. It should be noted that the drawings are in a very simplified form and use non-precise proportions, and are only used to facilitate and clarify the illustration of the embodiments of the present invention.

[0033] The scaling factor calibration of this invention uses the attitude rotation maneuver of the satellite platform (hereinafter referred to as the platform) as the reference angular velocity for gyroscope measurement, and the measurement value of the star sensor as the reference for the change in the platform's attitude angle before and after the attitude maneuver. To improve the accuracy of the scaling factor calibration, the platform should rotate to the target attitude angular velocity as quickly as possible, complete the full number of attitude rotations, and then de-rotate to a stable state as quickly as possible after the rotation ends. To obtain sufficient sampling data, the sampling period should be no less than 200ms, and the sampling time no less than 400s. Combined with... Figures 1-4 As shown, the method of the present invention includes the following steps:

[0034] S1, the satellite platform maintains static stability control by collecting measurement data from the star sensor to calculate the platform's attitude angle, which is then used as the attitude angle P0 before the platform's attitude rotation maneuver. This static stability control can be achieved by maintaining a ground-oriented attitude, where the platform stability is not less than 0.05° / s, the star sensor output attitude angle measurement accuracy is not less than 60″, the sampling period T is not less than 200ms, and the duration is not less than 300s. The star sensor measurement accuracy is at least one order of magnitude higher than the measurement accuracy of the calibrated gyroscope. Alternatively, the satellite platform can maintain static stability control by maintaining inertial orientation or solar orientation.

[0035] S2, the satellite platform maintains the target attitude angular velocity ω t (The target angular velocity ω is defined as an angular velocity that is not less than 10% of the gyroscope's measurement range.) t A positive attitude rotation maneuver is performed around the calibration axis (limited to 10° / s). During the rotation, the output value of the gyroscope is collected. Taking the attitude angle P0 before the platform's attitude rotation maneuver as the initial value, the attitude angle change P of the platform's attitude rotation maneuver is calculated by gyroscope integration. gyro ;

[0036] The aforementioned gyroscope calibration axis refers to any one of the X, Y, and Z axes of the gyroscope, with no specific order required;

[0037] The aforementioned attitude rotation maneuver control, such as Figure 4 As shown, this means that the platform rotates 3 times at the target angular velocity, and the angular velocity control error during the rotation process is no greater than 0.05° / s;

[0038] The attitude angular velocity used in the gyro integration is a measured value of the gyro output after zero-position offset correction. The satellite platform maintains stable attitude control, using the attitude angle information acquired from the star sensor output as the platform's attitude angle reference, and performs on-orbit calibration and correction of the gyro's zero-position offset. Specifically, this includes:

[0039] Step a: The satellite platform maintains static stability control, using the attitude angle information output by the acquired star sensor as the platform attitude angle reference;

[0040] Step b: Calculate the quaternion q using the angular velocity output by the gyroscope. gyro ;

[0041] Step c, calculate the star sensor output quaternion q st Quaternion q calculated with a gyroscope gyro The deviation quaternion Δq between them is calculated, and the deviation quaternion Δq is normalized.

[0042] Step d: Filter the vector part of the deviation quaternion to obtain the zero offset ω0 of the gyroscope;

[0043] Step e involves correcting the gyroscope's measurement output using the calibrated gyroscope zero-position offset.

[0044] S3, the satellite platform maintains stationary stability control relative to the Earth, collects measurement data from the star sensor to calculate the platform's attitude angles, which are then used as the attitude angle P1 after the platform's attitude rotation maneuver; the attitude angle P after the attitude rotation maneuver is calculated using P1-P0. ST The attitude angle serves as the reference angle for the platform's attitude rotation maneuver; in steps S1 and S3, the satellite platform is under attitude stationary control, serving as the attitude angle reference before and after the attitude rotation maneuver; the deviation P of the attitude angle change before and after the attitude rotation maneuver... ST The attitude angle change before and after the attitude rotation maneuver is measured by the star sensor as the benchmark.

[0045] S4, Calculate the attitude angle P determined by the star sensor. ST The attitude angle P determined by the gyroscope gyro The deviation attitude angle ΔP between the two is calculated, and the deviation of the positive scaling factor Δk of the calibration axis is calculated. + =△P / ω t ;

[0046] S5, Repeat steps S1 to S4 to obtain the calibration axis negative scaling factor deviation Δk. - In step S2, the target attitude angular velocity ω is used to revolve around the calibration axis. t Perform negative rotation; that is, when the positive and negative scaling factor deviations of the calibration axis are obtained, the satellite platform rotates around the calibration axis in the positive and negative directions with the target attitude angular velocity, respectively.

[0047] S6, take the calibration axis positive scale factor deviation Δk + and negative scaling factor deviation Δk - The average value is used as the calibration axis scale factor correction amount Δk. 轴 That is, the calibration axis scale factor correction is the average of the positive scale factor deviation and the negative scale factor.

[0048] S7, the calibration axis scale factor correction amount Δk calculated in step S6. 轴 The gyroscope calibration factor is corrected to obtain a new scaling factor for the calibration axis.

[0049] The present invention will be illustrated below using the calibration of the X-axis as a specific example.

[0050] Step 1: Estimate the X-axis zero offset

[0051] 1. The satellite platform maintains geostationary stability control. Geostationary stability is defined as a stability of not less than 0.05° / s. The attitude angle information output by the acquired star sensor is used as the platform attitude angle reference. The attitude angle measurement accuracy of the star sensor output is not less than 60″, the sampling period T is not less than 200ms, and the sampling time is not less than 1300s.

[0052] 2. Calculate the quaternion q using the angular velocity output by the gyroscope. gyro ;

[0053] 3. Calculate the star sensor output quaternion q. st Quaternion q calculated with a gyroscope gyro The deviation quaternion Δq between them is calculated, and the deviation quaternion Δq is normalized.

[0054] 4. Filter the vector part of the deviation quaternion to obtain the zero-position offset ω0 of the gyroscope;

[0055] 5. Correct the gyroscope's measurement output using the calibrated gyroscope zero-position offset.

[0056] Step 2: +X Scale Factor Calibration

[0057] 1. The platform maintains three-axis zero-attitude stability control relative to the ground, and uses a star sensor to measure the platform's attitude angle information; the platform stability is not less than 0.05° / s, the attitude measurement accuracy is not less than 60″, and the duration is not less than 300s; the platform attitude angle measured by the star sensor is used as the attitude angle P0 before the platform's attitude rotation maneuver.

[0058] 2. The platform revolves around the +X axis at the target attitude angular velocity ω tThe gyroscope rotates 3 times at 4° / s for no more than 400s, during which measurement information is collected from the gyroscope with a sampling period of 200ms. Using P0 as the initial attitude angle, the attitude angle P of the attitude rotation maneuver is calculated by gyroscope integration (the gyroscope output value is corrected for zero-position offset and then used as the gyroscope measurement value). gyro ;

[0059] 3. The platform maintains three-axis zero-attitude stability control relative to the ground, using a star sensor to measure the platform's attitude angle information; the platform stability is no less than 0.05° / s, the attitude measurement accuracy is no less than 60″, and the duration is no less than 300s; the platform attitude angle measured by the star sensor is used as the attitude angle P1 after the platform's attitude rotation maneuver; the attitude angle P after the attitude rotation maneuver is calculated by P1-P0. ST As the reference attitude angle for platform attitude rotation maneuvering;

[0060] 4. Calculate the attitude angle P determined by the star sensor. ST The attitude angle P determined by the gyroscope gyro The deviation angle ΔP between them;

[0061] 5. Calculate the scaling factor correction Δk in the +X direction. + =△P / ω t .

[0062] Step 3: X-axis scaling factor calibration

[0063] 1. The platform maintains three-axis zero-attitude stability control relative to the ground, and uses a star sensor to measure the platform's attitude angle information; the platform stability is not less than 0.05° / s, the attitude measurement accuracy is not less than 60″, and the duration is not less than 300s; the platform attitude angle measured by the star sensor is used as the attitude angle P0 before the platform's attitude rotation maneuver.

[0064] 2. The platform revolves around the -X axis at the target attitude angular velocity ω t The gyroscope rotates 3 times at 4° / s for no more than 400s, during which measurement information is collected from the gyroscope with a sampling period of 200ms. Using P0 as the initial attitude angle, the attitude angle P of the attitude rotation maneuver is calculated by gyroscope integration (the gyroscope output value is corrected for zero-position offset and then used as the gyroscope measurement value). gyro ;

[0065] 3. The platform maintains three-axis zero-attitude stability control relative to the ground, using a star sensor to measure the platform's attitude angle information; the platform stability is no less than 0.05° / s, the attitude measurement accuracy is no less than 60″, and the duration is no less than 300s; the platform attitude angle measured by the star sensor is used as the attitude angle P1 after the platform's attitude rotation maneuver; the attitude angle P after the attitude rotation maneuver is calculated by P1-P0. ST As the reference attitude angle for platform attitude rotation maneuvering;

[0066] 4. Calculate the attitude angle P determined by the star sensor. ST The attitude angle P determined by the gyroscope gyro The deviation angle ΔP between them;

[0067] 5. Calculate the X-axis scale factor correction Δk - =△P / ω t .

[0068] Step 4: Calculate the X-axis scale factor correction.

[0069] X-axis scale factor correction Δk x =(△k) + +△k - ) / 2

[0070] This completes the calculation of the X-axis scale factor correction.

[0071] Then, follow steps one through four to calculate the Y and Z axis scale factor corrections in sequence.

[0072] In summary, this invention uses the attitude angles measured by a star sensor as the attitude angle reference before and after the satellite platform's attitude rotation maneuver; it calculates the attitude angles of the platform's attitude rotation maneuver by integrating the gyroscope (using the gyroscope output value after zero-point correction as the gyroscope measurement value); and it calculates the gyroscope's scaling factor by using the attitude angle deviation of the platform's attitude rotation maneuver calculated by the star sensor and the gyroscope. This invention provides sensitive data for the gyroscope through the satellite platform's attitude rotation maneuver, and then uses a star sensor with higher measurement accuracy than the gyroscope as the measurement reference to calibrate and correct the gyroscope's scaling factor. This effectively solves the problem of long-term on-orbit performance parameter degradation of the gyroscope, providing high-precision measurement assurance for high-precision and high-stability control and rapid and stable attitude maneuver control of long-life satellite platforms.

[0073] Although the present invention has been described in detail through the preferred embodiments above, it should be understood that the above description should not be considered as a limitation of the present invention. Various modifications and substitutions to the present invention will be apparent to those skilled in the art after reading the above description. Therefore, the scope of protection of the present invention should be defined by the appended claims.

Claims

1. A method for on-orbit calibration of a gyroscope based on attitude maneuvering, characterized in that, The aforementioned attitude maneuver refers to attitude maneuver control that involves rotating a full circle at a constant rotational angular velocity within a certain time period. The method includes: S1, the satellite platform maintains static stability control, collects the measurement data of the star sensor to calculate the platform's attitude angle, and uses it as the attitude angle P0 before the platform's attitude rotation maneuver; S2, the satellite platform maintains the target attitude angular velocity ω t A positive attitude rotation maneuver is performed around the calibration axis. During the rotation, the output value of the gyroscope is collected. Using the attitude angle P0 before the platform's attitude rotation maneuver as the initial value, the attitude angle change P of the platform's attitude rotation maneuver is calculated through gyroscope integration. gyro ; S3, the satellite platform maintains static stability control, collects measurement data from the star sensor to calculate the platform's attitude angle, which is used as the attitude angle P1 after the platform's attitude rotation maneuver; the attitude angle P after the attitude rotation maneuver is calculated by subtracting P0 from P1. ST As the reference attitude angle for platform attitude rotation maneuvering; S4, Calculate the attitude angle P determined by the star sensor. ST The attitude angle P determined by the gyroscope gyro The deviation attitude angle ΔP between the two is calculated, and the deviation of the positive scaling factor Δk of the calibration axis is calculated. + =△P / ω t ; S5, Repeat steps S1 to S4 to obtain the calibration axis negative scaling factor deviation Δk. - In step S2, the target attitude angular velocity ω is used to revolve around the calibration axis. t Perform a negative rotation; S6, take the calibration axis positive scale factor deviation Δk + and negative scaling factor deviation Δk - The average value is used as the calibration axis scale factor correction amount Δk. 轴 ; S7, the calibration axis scale factor correction amount Δk calculated in step S6. 轴 The gyroscope calibration factor is corrected to obtain a new scaling factor for the calibration axis.

2. The on-orbit calibration method for a gyroscope based on attitude maneuvering according to claim 1, characterized in that, The attitude angular velocity used in the gyroscope integration is a measured value of the gyroscope output after zero-position offset correction.

3. The on-orbit calibration method for a gyroscope based on attitude maneuvering according to claim 2, characterized in that, The aforementioned satellite platform maintains static stability control, using the attitude angle information output by the acquired star sensor as the platform's attitude angle reference, and performs on-orbit calibration and correction of the gyroscope's zero-position offset, specifically including: Step a: The satellite platform maintains static stability control, using the attitude angle information output by the acquired star sensor as the platform attitude angle reference; Step b: Calculate the quaternion q using the angular velocity output by the gyroscope. gyro ; Step c, calculate the star sensor output quaternion q st Quaternion q calculated with a gyroscope gyro The deviation quaternion Δq between them is calculated, and the deviation quaternion Δq is normalized. Step d: Filter the vector part of the deviation quaternion to obtain the zero offset ω0 of the gyroscope; Step e involves correcting the gyroscope's measurement output using the calibrated gyroscope zero-position offset.

4. The on-orbit calibration method for a gyroscope based on attitude maneuvering according to claim 1, characterized in that, The calibration axes are the X-axis, Y-axis, and Z-axis.

5. The on-orbit calibration method for a gyroscope based on attitude maneuvering according to claim 1, characterized in that, The satellite platform maintains a stable attitude control for Earth orientation, wherein the platform stability is not less than 0.05° / s, the star sensor output attitude angle measurement accuracy is not less than 60″, the sampling period T is not less than 200ms, and the duration is not less than 300s.

6. The on-orbit calibration method for a gyroscope based on attitude maneuvering according to claim 1, characterized in that, The aforementioned attitude rotation maneuver control refers to the platform rotating 3 times at the target angular velocity, with the angular velocity control error during the rotation process not exceeding 0.05° / s.

7. The on-orbit calibration method for a gyroscope based on attitude maneuvering according to claim 1, characterized in that, The star sensor has a measurement accuracy that is at least an order of magnitude higher than that of the calibrated gyroscope.

8. The on-orbit calibration method for a gyroscope based on attitude maneuvering according to claim 1, characterized in that, The target attitude angular velocity ω t Not less than 10% of the gyroscope's measurement range.

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