High-precision attitude determination method and device for high-frequency estimation and timing correction of rotating load

Abstract

The present invention discloses a high-precision attitude determination method and device for high-frequency estimation and timing correction of rotating payloads, belonging to the field of satellite attitude control. The method comprises: when the satellite is in a star-sensitive invisible sky area, the short-period attitude estimation of the gyroscope is used for attitude determination; when the satellite enters a star-sensitive visible sky area, the initial filter estimated attitude of the gyroscope is reset using the star-sensitive attitude quaternion, and a short-period attitude estimation is performed based on the reset gyro initial filter estimated attitude, so as to perform fusion attitude determination using the short-period attitude estimation of the gyroscope and the star-sensitive correction; wherein the attitude estimation period of the gyroscope is less than the star-sensitive correction period. The present invention reduces the attitude quaternion estimation period of the gyroscope, increases the attitude estimation frequency of the gyroscope, and uses a long-short period fusion attitude determination strategy in the star-sensitive visible sky area to effectively reduce the deviation in the attitude determination process caused by the motion characteristics of the high angular velocity rotating payload, thereby achieving high-precision attitude determination of high-speed rotating payloads.

Description

Technical Field

[0001] The present invention relates to the field of satellite attitude control technology, and in particular to a high-precision attitude determination method and device for high-frequency estimation and timing correction of a rotating load. Background Art

[0002] The three-axis attitude determination of the payload platform for high-dynamic imaging missions directly depends on the accuracy of the high-dynamic star sensor and the gyro joint attitude determination of the payload platform. In the traditional attitude determination method, since the rotation speed of the payload platform around the roll axis is not high, the gyro attitude estimation period can be consistent with the star sensor attitude correction period. However, when the payload platform rotates at high speed around the roll axis, the gyro measurement accuracy decreases significantly. In order to complete the high-precision and high-stability fusion attitude determination of the high-speed rotating payload, it is necessary to reduce the gyro attitude quaternion estimation period.

[0003] Therefore, there is an urgent need to provide a high-precision attitude determination method and device for high-frequency estimation and timing correction of rotating loads. Summary of the invention

[0004] In order to solve the problem that traditional attitude determination methods are difficult to achieve the expected accuracy when directly applied to high-speed rotating payload satellites, an embodiment of the present invention provides a high-precision attitude determination method and device for high-frequency estimation and timing correction of rotating payloads.

[0005] On the one hand, a high-precision attitude determination method for high-frequency estimation and timing correction of rotating loads is provided, which is applied to an onboard control system, and the method comprises:

[0006] When the satellite is in a sky area where the star sensor is not visible, the attitude is determined by using the short-period attitude estimation of the gyroscope.

[0007] When the satellite enters the visible sky area of ​​the star-sensitive system, the star-sensitive attitude quaternion is used to reset the initial filter estimated attitude of the gyro, and a short-period attitude estimation is performed based on the reset initial filter estimated attitude of the gyro, so as to use the short-period attitude estimation of the gyro and the star-sensitive correction to perform fusion attitude determination; wherein the attitude estimation period of the gyro is less than the star-sensitive correction period, and the star-sensitive correction is used to periodically correct the initial filter estimated attitude of the gyro in each control period.

[0008] On the other hand, a high-precision attitude determination device for high-frequency estimation and timing correction of rotating loads is provided, which is used to implement the steps described in any method embodiment of the specification, and is arranged in an onboard control system, and the device includes:

[0009] The first attitude determination unit is used for determining the attitude of the satellite using the short-period attitude estimation of the gyroscope in the sky area where the satellite is not visible to the star sensor;

[0010] The second attitude determination unit is used to reset the initial filter estimated attitude of the gyroscope using the star-sensitive attitude quaternion when the satellite enters the star-sensitive visible sky area, and perform short-period attitude estimation based on the reset gyroscope initial filter estimated attitude, so as to perform fusion attitude determination using the gyroscope's short-period attitude estimation and star-sensitive correction; wherein the gyroscope's attitude estimation period is smaller than the star-sensitive correction period, and the star-sensitive correction is used to periodically correct the initial filter estimated attitude of each control period of the gyroscope.

[0011] On the other hand, a computer device is provided, which includes a memory and a processor, wherein the memory is used to store a computer program, and the processor is used to execute the computer program stored in the memory to implement the steps of the above-mentioned method.

[0012] On the other hand, a computer-readable storage medium is provided, wherein a computer program is stored in the storage medium, and when the computer program is executed by a processor, the steps of the above-mentioned method are implemented.

[0013] On the other hand, a computer program product is provided, comprising a computer program, wherein the computer program implements the steps of the above method when executed by a processor.

[0014] The technical solution provided by the present invention can at least bring the following beneficial effects:

[0015] By reducing the gyro's attitude quaternion estimation period, increasing the gyro's attitude estimation frequency, and using a long-short cycle fusion attitude determination strategy in the star-sensitive visible sky area, the deviation in the attitude determination process caused by the motion characteristics of the large angular velocity rotating payload can be effectively reduced, and high-precision attitude determination of high-speed rotating payloads can be achieved. In addition, when the satellite enters the star-sensitive visible sky area from the star-sensitive invisible sky area, the star-sensitive attitude quaternion is used to reset the gyro's initial filter estimation attitude, which can ensure the accuracy and convergence time of the gyro constant drift estimation and avoid introducing the gyro extrapolation model deviation in the non-visible area. Therefore, this solution can meet the high-precision attitude determination requirements of the optical satellite on-orbit payload platform under the high-speed motion characteristics. BRIEF DESCRIPTION OF THE DRAWINGS

[0016] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings required for use in the embodiments or the description of the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0017] Figure 1 This is a flow chart of a high-precision attitude determination method for high-frequency estimation and timing correction of a rotating load provided by an embodiment of the present invention;

[0018] Figure 2 This is a structural diagram of a high-precision attitude determination device for high-frequency estimation and timing correction of a rotating load provided by an embodiment of the present invention;

[0019] Figure 3 It is a hardware architecture diagram of a computer device provided by an embodiment of the present invention. DETAILED DESCRIPTION

[0020] In order to make the purpose, technical solutions and advantages of the embodiments of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are part of the embodiments of the present invention, rather than all the embodiments. Based on the embodiments in the present invention, all other embodiments obtained by ordinary technicians in this field without making creative work are within the scope of protection of the present invention.

[0021] The specific implementation of the above concept is described below.

[0022] Please refer to Figure 1 , a high-precision attitude determination method for high-frequency estimation and timing correction of a rotating load provided by an embodiment of the present invention is applied to an onboard control system, the method comprising:

[0023] Step 100: The satellite is in a sky area where the star sensor is not visible, and the attitude is determined by using the short-period attitude estimation of the gyroscope;

[0024] Step 102: When the satellite enters the visible sky area of ​​the star-sensing system, the initial filter estimated attitude of the gyro is reset using the star-sensing attitude quaternion, and a short-period attitude estimation is performed based on the reset initial filter estimated attitude of the gyro, so as to use the short-period attitude estimation of the gyro and the star-sensing correction to perform fusion attitude determination; wherein the attitude estimation period of the gyro is smaller than the star-sensing correction period.

[0025] In the embodiment of the present invention, by reducing the gyro's attitude quaternion estimation period, increasing the gyro's attitude estimation frequency, and using a long-short cycle fusion attitude determination strategy in the star-sensitive visible sky area, the deviation in the attitude determination process caused by the motion characteristics of the large angular velocity rotating payload is effectively reduced, and high-precision attitude determination of the high-speed rotating payload is achieved. In addition, when the satellite enters the star-sensitive visible sky area from the star-sensitive invisible sky area, the star-sensitive attitude quaternion is used to reset the gyro's initial filter estimation attitude, which can ensure the accuracy and convergence time of the gyro constant drift estimation and avoid introducing the gyro extrapolation model deviation in the non-visible area. Therefore, this solution can meet the high-precision attitude determination requirements of the optical satellite on-orbit payload platform under the high-speed motion characteristics.

[0026] Described below Figure 1 How the various steps are performed.

[0027] For step 100:

[0028] In the traditional low-speed payload attitude determination method, since the rotation speed of the payload platform around the roll axis is not high, the gyro attitude estimation cycle can be consistent with the star-sensing attitude correction cycle, that is, the gyro attitude estimation cycle is the same as the star-sensing attitude correction cycle, and the cycle time is relatively long. However, for the optical satellite with high dynamic imaging mission payload of the present invention, the payload platform rotates at high speed around the roll axis during the on-orbit operation, and the traditional gyro low-frequency attitude estimation method cannot meet the high-precision and high-stability attitude determination requirements of the high-speed rotating payload on orbit.

[0029] In this step, as the payload platform rotates around the rotation axis, the star-sensing combination enters the visible sky area and the invisible sky area in turn. The star-sensing combination uses the gyro information to perform attitude extrapolation in the invisible sky area, and only uses the gyro for satellite attitude determination. The traditional gyro attitude estimation cycle is consistent with the star-sensing attitude correction cycle, both of which are 200ms, while the gyro attitude estimation cycle of this step is 10ms, so it is a short-cycle high-frequency attitude estimation.

[0030] In the embodiment of the present invention, a plurality of short-cycle task slices are provided in each control cycle of the gyroscope. The first short-cycle task slice in each control cycle is used to perform gyroscope attitude estimation calculation based on the initial filtered estimated attitude of the current control cycle to obtain the extrapolated attitude; each of the remaining short-cycle task slices performs one-shot gyroscope attitude extrapolation estimation calculation on the extrapolated attitude to update the extrapolated attitude. For example, the control cycle can be 200ms, the short-cycle task slice is 10ms, and one control cycle contains 20 consecutive short-cycle task slices.

[0031] In the embodiment of the present invention, since the star sensor has no measurement data in the sky area where the star sensor is not visible, the "star sensor correction reset flag" is always 0. Therefore, the short-period attitude estimation method of the gyroscope in the sky area where the star sensor is not visible is:

[0032] For each control cycle, the following are executed:

[0033] Set the counter Cnt to record the cumulative calculation times of the short-cycle task slice in each control cycle. Cnt is cleared at the beginning of each control cycle, and Cnt+1 is added each time the short-cycle task slice is called.

[0034] When the counter Cnt is 0, the extrapolated posture obtained in the previous control cycle is assigned to the starting filtered estimated posture of the current control cycle;

[0035] Execute each short-cycle task slice in turn to perform attitude extrapolation;

[0036] The orbit information at the start time of the current control cycle is recorded to obtain the conversion matrix for converting the inertial system to the orbital system at the start time. The conversion matrix and the satellite angular velocity information are further used to perform orbit extrapolation in each short-cycle mission slice, and the conversion matrix is ​​updated to convert the extrapolated attitude from the inertial system to the orbital system using the conversion matrix to obtain the orbital attitude used to control the satellite.

[0037] It can be seen that in the invisible sky area, by reducing the gyroscope's attitude quaternion estimation period and increasing the gyroscope's attitude estimation frequency, the deviation in the attitude determination process caused by the motion characteristics of the large angular velocity rotating payload can be effectively reduced, and high-precision attitude determination of the high-speed rotating payload in the invisible sky area can be achieved.

[0038] Regarding step 102:

[0039] In some implementations, the star-sensitive correction period is equal in length to the control period of the gyro, and has the same starting time; the star-sensitive correction is used to regularly correct the starting filtered estimated attitude of each control period of the gyro.

[0040] In the embodiment of the present invention, the short-period task slice further includes validity judgment of gyro measurement data and gyro data processing before gyro attitude estimation calculation.

[0041] The steps of star-sensitivity correction are star-sensitivity validity judgment, star-sensitivity measurement data preprocessing, timing inversion attitude filter correction based on star-sensitivity measurement, and resetting the initial filter estimation attitude of the gyro based on the corrected star-sensitivity attitude quaternion.

[0042] In this embodiment, since the star-sensing correction cycle is equal to the control cycle of the gyro and has the same starting time, after the star-sensing obtains the star-sensing attitude quaternion of the current control cycle, and the dual star-sensing is valid, the "star-sensing correction reset flag" is set to 1. It can be understood that when it is set to 1, the gyro is executing a short-cycle task slice in the middle of the current control cycle, and the starting filtered estimated attitude of the gyro in the current control cycle is reset to update and correct the gyro extrapolated attitude of each beat of the current control cycle.

[0043] Specifically, the short-period attitude estimation method of the gyroscope is:

[0044] For each control cycle, the following are executed:

[0045] Set the counter Cnt to record the cumulative calculation times of the short-cycle task slice in each control cycle. Cnt is cleared at the beginning of each control cycle, and Cnt+1 is added each time the short-cycle task slice is called.

[0046] When the counter Cnt is 0, the posture estimated by the short cycle is assigned to the starting filtered estimated posture of the current control cycle;

[0047] Execute each short-cycle task slice in turn, and monitor in real time whether the "star sensitivity correction reset flag" is 1;

[0048] When the "star sensitivity correction reset flag" is 0, each short-cycle task slice is executed normally in sequence;

[0049] When the "star sensitivity correction reset flag" is 1, the star sensitivity attitude quaternion obtained by star sensitivity correction is used to reset the starting filtered estimated attitude of the gyro attitude estimation of the current control cycle;

[0050] Based on the initial filtered estimated attitude of the current control cycle after reset, re-update the extrapolated attitude of each short-cycle task slice that has been executed, and continue to execute the short-cycle task slice corresponding to the current counter based on the new extrapolated attitude, and set the "star sensitivity correction reset flag" to 0;

[0051] The orbit information at the start time of the current control cycle is recorded to obtain the conversion matrix for converting the inertial system to the orbital system at the start time. The conversion matrix and the satellite angular velocity information are further used to perform orbit extrapolation in each short-cycle mission slice, and the conversion matrix is ​​updated to convert the extrapolated attitude from the inertial system to the orbital system using the conversion matrix to obtain the orbital attitude used to control the satellite.

[0052] In this embodiment, it is monitored whether the "star sensitivity correction reset flag" is set to 1. When it is set to 1, the gyroscope is executing a short-cycle task slice in the middle position of this control cycle. By resetting the starting filter estimated attitude of the gyroscope in this control cycle, the extrapolated attitude of each short-cycle task slice in this control cycle is updated and corrected, thereby realizing a long- and short-cycle multi-star sensitivity and multi-gyroscope fusion attitude determination strategy in the star-sensitive visible sky area, which can effectively reduce the deviation in the attitude determination process caused by the motion characteristics of the large angular velocity rotating load, and realize high-precision attitude determination of high-speed rotating loads.

[0053] In some implementations, the short-period attitude estimation process of the gyroscope further includes:

[0054] From 150 ms before the start time of each control cycle to the start time, the cumulative value of the gyro measured angular velocity in this time period is accumulated and calculated.

[0055] Timing inversion attitude filter correction based on star-sensing measurement, including:

[0056] Obtain the preprocessed star-sensitivity measurement data; wherein the sampling time of the star-sensitivity measurement data is 150ms before the start time of the current star-sensitivity correction cycle;

[0057] The accumulated value of the angular velocity measured by the gyro is compensated to the preprocessed star-sensing measurement data to correct the star-sensing attitude quaternion at the start time of the current star-sensing correction cycle.

[0058] In this embodiment, the motion characteristics of the high-speed rotation of the payload platform make the influence of the time compensation factor very prominent, so it is necessary to perform time compensation calculations on the delayed star-sensing data of the payload platform and the time mismatch problem caused by the calculation process.

[0059] Since the aberration is processed on a single machine, the star-sensitivity measurement data has a 150ms delay. If the starting time of the star-sensitivity correction cycle, that is, the star-sensitivity measurement data at the starting time of the control cycle, is directly obtained in each star-sensitivity correction cycle, then the star-sensitivity measurement data that can be obtained by the onboard control system 150ms after the starting time cannot be completed in the remaining 50ms because the control cycle is 200ms. Therefore, in each star-sensitivity correction cycle, the star-sensitivity measurement data sampled 150ms before the starting time is obtained. By accumulating the cumulative value of the gyro measurement angular velocity in the time period from 150ms before the starting time of each control cycle to the starting time in the short-cycle attitude estimation process of the gyro, the star-sensitivity measurement data in the first 150ms after preprocessing is time-compensated to obtain the star-sensitivity attitude quaternion at the starting time of the current star-sensitivity correction cycle, and the timing inversion attitude filter correction of the star-sensitivity measurement is completed. The corrected star-sensitivity attitude quaternion is further used to reset the starting filter estimation attitude of the gyro, and the high-precision and high-stability fusion attitude determination of the high-speed payload is completed.

[0060] Please refer to Figure 2 The embodiment of the present invention provides a high-precision attitude determination device for high-frequency estimation and timing correction of rotating loads, which is used to implement the steps of any method embodiment in the specification and is arranged in an onboard control system. The device includes:

[0061] The first attitude determination unit 201 is used for determining the attitude of the satellite by using the short-period attitude estimation of the gyroscope when the satellite is in a sky area where the star sensor is not visible;

[0062] The second attitude determination unit 202 is used to reset the initial filter estimated attitude of the gyroscope using the star-sensitive attitude quaternion when the satellite enters the star-sensitive visible sky area, and perform short-period attitude estimation based on the reset gyroscope initial filter estimated attitude, so as to perform fusion attitude determination using the gyroscope's short-period attitude estimation and star-sensitive correction; wherein the gyroscope's attitude estimation period is smaller than the star-sensitive correction period.

[0063] In one embodiment of the present invention, a plurality of short-cycle task slices are provided in each control cycle of the gyro in the first attitude determination unit 201 and the second attitude determination unit 202. The first short-cycle task slice in each control cycle is used to perform gyro attitude estimation calculation based on the initial filtered estimated attitude of the current control cycle to obtain an extrapolated attitude; each of the remaining short-cycle task slices performs a one-shot gyro attitude extrapolation estimation calculation for the extrapolated attitude to update the extrapolated attitude;

[0064] The second attitude determination unit 202 sets the star-sensitivity correction period to be equal in length to the gyro control period and to have the same starting time; the star-sensitivity correction is used to regularly correct the starting filtered estimated attitude of each gyro control period.

[0065] In one embodiment of the present invention, the short-period task slices in the first attitude determination unit 201 and the second attitude determination unit 202 further include validity determination of gyro measurement data and gyro data processing before gyro attitude estimation calculation;

[0066] The steps of star sensitivity correction in the second attitude determination unit 202 are star sensitivity validity judgment, star sensitivity measurement data preprocessing, timing inversion attitude filtering correction based on star sensitivity measurement, and resetting the initial filtering estimated attitude of the gyro based on the corrected star sensitivity attitude quaternion.

[0067] In one embodiment of the present invention, the short-period attitude estimation method of the gyroscope in the first attitude determination unit 201 and the second attitude determination unit 202 is:

[0068] For each control cycle, the following are executed:

[0069] Set the counter Cnt to record the cumulative calculation times of the short-cycle task slice in each control cycle. Cnt is cleared at the beginning of each control cycle, and Cnt+1 is added each time the short-cycle task slice is called.

[0070] When the counter Cnt is 0, the posture estimated by the short cycle is assigned to the starting filtered estimated posture of the current control cycle;

[0071] Execute each short-cycle task slice in turn, and monitor in real time whether the "star sensitivity correction reset flag" is 1;

[0072] When the "star sensitivity correction reset flag" is 0, each short-cycle task slice is executed normally in sequence;

[0073] When the "star sensitivity correction reset flag" is 1, the star sensitivity attitude quaternion obtained by star sensitivity correction is used to reset the starting filtered estimated attitude of the gyro attitude estimation of the current control cycle;

[0074] Based on the initial filtered estimated attitude of the current control cycle after reset, re-update the extrapolated attitude of each short-cycle task slice that has been executed, and continue to execute the short-cycle task slice corresponding to the current counter based on the new extrapolated attitude, and set the "star sensitivity correction reset flag" to 0;

[0075] The orbit information at the start time of the current control cycle is recorded to obtain the conversion matrix for converting the inertial system to the orbital system at the start time. The conversion matrix and the satellite angular velocity information are further used to perform orbit extrapolation in each short-cycle mission slice, and the conversion matrix is ​​updated to convert the extrapolated attitude from the inertial system to the orbital system using the conversion matrix to obtain the orbital attitude used to control the satellite.

[0076] In one embodiment of the present invention, the short-period attitude estimation process of the gyroscope in the second attitude determination unit 202 further includes:

[0077] From 150 ms before the start time of each control cycle to the start time, the cumulative value of the gyro measured angular velocity in this time period is accumulated and calculated.

[0078] In one embodiment of the present invention, when the second attitude determination unit 202 performs the timing inversion attitude filtering correction based on the star-sensing measurement, it is used to:

[0079] Obtain the preprocessed star-sensitivity measurement data; wherein the sampling time of the star-sensitivity measurement data is 150ms before the start time of the current star-sensitivity correction cycle;

[0080] The accumulated value of the angular velocity measured by the gyro is compensated to the preprocessed star-sensing measurement data to correct the star-sensing attitude quaternion at the start time of the current star-sensing correction cycle.

[0081] It should be noted that the high-precision attitude determination device for high-frequency estimation and timing correction of rotating loads provided in the above embodiment is only illustrated by the division of the above functional units. In practical applications, the above functions can be assigned to different functional units as needed, that is, the internal structure of the device can be divided into different functional units to complete all or part of the functions described above. In addition, the above device embodiment and the method embodiment belong to the same concept, and the specific implementation process is detailed in the method embodiment, which will not be repeated here.

[0082] The embodiment of the present application also provides a computer device, please refer to Figure 3 The computer device includes a processor and a memory, wherein the memory stores at least one instruction, at least one program, code set or instruction set, and the at least one instruction, at least one program, code set or instruction set is loaded and executed by the processor to implement the high-precision attitude determination method for high-frequency estimation and timing correction of rotating loads provided in the above-mentioned method embodiments.

[0083] An embodiment of the present application also provides a computer-readable storage medium, on which is stored at least one instruction, at least one program, code set or instruction set, and the at least one instruction, at least one program, code set or instruction set is loaded and executed by a processor to implement the high-precision attitude determination method for high-frequency estimation and timing correction of rotational loads provided in the above-mentioned method embodiments.

[0084] An embodiment of the present application also provides a computer program product, which includes a computer program. A processor of a computer device reads the computer program from a computer-readable storage medium, and the processor executes the computer program, so that the computer device executes the high-precision attitude determination method for high-frequency estimation and timing correction of rotational loads in any of the above-mentioned embodiments.

[0085] For the convenience of description, the above system or device is described by dividing it into various modules or units according to its functions. Of course, when implementing the present application, the functions of each unit can be implemented in the same or multiple software and / or hardware.

[0086] It can be known from the description of the above implementation methods that those skilled in the art can clearly understand that the present application can be implemented by means of software plus a necessary general hardware platform. Based on such an understanding, the technical solution of the present application can be essentially or partly embodied in the form of a software product that contributes to the prior art. The computer software product can be stored in a storage medium such as ROM / RAM, a magnetic disk, an optical disk, etc., and includes several instructions for enabling a computer device (which can be a personal computer, a server, or a network device, etc.) to execute the methods of each embodiment of the present application or some parts of the embodiments.

[0087] Finally, it should be noted that, in this article, relational terms such as first, second, third and fourth are only used to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Moreover, the terms "include", "comprise" or any other variants thereof are intended to cover non-exclusive inclusion, so that a process, method, article or device including a series of elements includes not only those elements, but also other elements not explicitly listed, or also includes elements inherent to such process, method, article or device. In the absence of further restrictions, the elements defined by the sentence "comprise a ..." do not exclude the presence of other identical elements in the process, method, article or device including the elements.

[0088] The above are only preferred implementations of the present application. It should be pointed out that for ordinary technicians in this technical field, several improvements and modifications can be made without departing from the principles of the present application. These improvements and modifications should also be regarded as the scope of protection of the present application.

Claims

1. A high-precision attitude determination method for high-frequency estimation and timing correction of rotating loads, characterized in that: Applied to an onboard control system, the method comprises: When the satellite is in a sky area where the star sensor is not visible, the attitude is determined by using the short-period attitude estimation of the gyroscope. When the satellite enters the visible sky area of ​​the star-sensing system, the gyro's initial filter estimated attitude is reset using the star-sensing attitude quaternion, and a short-period attitude estimation is performed based on the reset gyro's initial filter estimated attitude, so as to use the gyro's short-period attitude estimation and star-sensing correction to perform fusion attitude determination; wherein, the gyro's attitude estimation period is shorter than the star-sensing correction period; Several short-cycle task slices are set in each control cycle of the gyro. The first short-cycle task slice in each control cycle is used to perform gyro attitude estimation calculation based on the starting filtered estimated attitude of the current control cycle to obtain the extrapolated attitude; each of the remaining short-cycle task slices performs a one-shot gyro attitude extrapolation estimation calculation on the extrapolated attitude to update the extrapolated attitude; The star-sensitivity correction period is equal to the control period of the gyro, and the starting time is the same; the star-sensitivity correction is used to regularly correct the starting filter estimation attitude of each control period of the gyro; The short-term attitude estimation method of the gyroscope is: For each control cycle, the following are executed: Set the counter Cnt to record the cumulative calculation times of the short-cycle task slice in each control cycle. Cnt is cleared at the beginning of each control cycle, and Cnt+1 is added each time the short-cycle task slice is called. When the counter Cnt is 0, the posture estimated by the short cycle is assigned to the starting filtered estimated posture of the current control cycle; Execute each short-cycle task slice in turn, and monitor in real time whether the star-sensing correction reset flag is 1; When the star-sensitivity correction reset flag is 0, each short-period task slice is executed in sequence normally; When the star-sensitivity correction reset flag is 1, the starting filtered estimated attitude of the gyro attitude estimation of the current control cycle is reset with the star-sensitivity attitude quaternion obtained by the star-sensitivity correction; Taking the initial filtered estimated attitude of the current control cycle after reset as the benchmark, re-update the extrapolated attitude of each short-cycle task slice that has been executed, and continue to execute the short-cycle task slice corresponding to the current counter based on the new extrapolated attitude, and set the star-sensitivity correction reset flag to 0; The orbit information at the start time of the current control cycle is recorded to obtain the conversion matrix for converting the inertial system to the orbital system at the start time. The conversion matrix and the satellite angular velocity information are further used to perform orbit extrapolation in each short-cycle mission slice, and the conversion matrix is ​​updated to convert the extrapolated attitude from the inertial system to the orbital system using the conversion matrix to obtain the orbital attitude used to control the satellite.

2. The method according to claim 1, characterized in that The short-cycle task slice also includes validity judgment of gyro measurement data and gyro data processing before gyro attitude estimation calculation; The steps of star-sensitivity correction are star-sensitivity validity judgment, star-sensitivity measurement data preprocessing, timing inversion attitude filter correction based on star-sensitivity measurement, and resetting the initial filter estimation attitude of the gyro based on the corrected star-sensitivity attitude quaternion.

3. The method according to claim 2, characterized in that The short-term attitude estimation process of the gyro also includes: From 150 ms before the start time of each control cycle to the start time, the cumulative value of the gyro measured angular velocity in this time period is accumulated and calculated.

4. The method according to claim 3, characterized in that The timing inversion attitude filtering correction based on star-sensitive measurement includes: Acquire preprocessed star sensitivity measurement data; wherein the sampling time of the star sensitivity measurement data is 150ms before the start time of the current star sensitivity correction cycle; The accumulated value of the gyroscopic angular velocity is compensated to the preprocessed star-sensing measurement data to correct and obtain the star-sensing attitude quaternion at the start time of the current star-sensing correction cycle.

5. A high-precision attitude determination device for high-frequency estimation and timing correction of rotating loads, used to implement the steps of any of the methods described in claims 1-4 above, characterized in that: Set up in the on-board control system, the device includes: The first attitude determination unit is used for determining the attitude of the satellite using the short-period attitude estimation of the gyroscope in the sky area where the satellite is not visible to the star sensor; The second attitude determination unit is used to reset the initial filter estimated attitude of the gyroscope using the star-sensitive attitude quaternion when the satellite enters the star-sensitive visible sky area, and perform short-period attitude estimation based on the reset gyroscope initial filter estimated attitude, so as to perform fusion attitude determination using the gyroscope's short-period attitude estimation and star-sensitive correction; wherein the gyroscope's attitude estimation period is smaller than the star-sensitive correction period, and the star-sensitive correction is used to periodically correct the initial filter estimated attitude of each control period of the gyroscope.

6. A computer device, characterized in that: The computer device includes a memory and a processor, the memory is used to store a computer program, and the processor is used to execute the computer program stored in the memory to implement the steps of any one of the methods described in claims 1-4.

7. A computer-readable storage medium, characterized in that: The storage medium stores a computer program, and when the computer program is executed by a processor, the steps of the method described in any one of claims 1 to 4 are implemented.

8. A computer program product, characterized in that The method comprises a computer program, wherein when the computer program is executed by a processor, the steps of the method according to any one of claims 1 to 4 are implemented.

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