Method and device for improving the inertial attitude accuracy of a satellite payload

Abstract

The application relates to the technical field of satellite load attitude measurement, in particular to a method and device for improving the accuracy of satellite load inertia attitude. The method comprises the following steps: based on the angle monitoring quaternion of the optical axis of a star sensor relative to a load reference, three-axis error angles of the star sensor relative to the load reference are determined; the three-axis error angles are filtered to obtain filtered three-axis error angles; based on the filtered three-axis error angles, a load installation matrix is corrected to obtain a corrected load installation matrix; and the load inertia attitude is calculated based on the corrected load installation matrix. According to the application, low-frequency errors can be compensated, and the accuracy of the satellite load inertia attitude is improved.

Description

Technical Field

[0001] This invention relates to the field of satellite payload attitude measurement technology, and in particular to a method and apparatus for improving the accuracy of satellite payload inertial attitude. Background Technology

[0002] Optical remote sensing satellites consist of an optical payload and a platform, on which star sensors and gyroscopes are mounted. During satellite operation, due to the alternating thermal environment of space, periodic structural deformations occur between the payload and the star sensor. These payload attitude measurement errors caused by structural deformation are commonly referred to as low-frequency errors. Because of these low-frequency errors, the accuracy of the satellite payload inertial attitude determined by related technologies is relatively low.

[0003] Therefore, there is an urgent need for a method and device to improve the accuracy of satellite payload inertial attitude to solve the above-mentioned technical problems. Summary of the Invention

[0004] This invention provides a method and apparatus for improving the accuracy of satellite payload inertial attitude, which can compensate for low-frequency errors and improve the accuracy of satellite payload inertial attitude.

[0005] In a first aspect, embodiments of the present invention provide a method for improving the accuracy of satellite payload inertial attitude, comprising:

[0006] Based on the quaternion of the angle between the optical axis of the star sensor and the load reference, the triaxial error angle of the star sensor relative to the load reference is determined.

[0007] The triaxial error angles are filtered to obtain the filtered triaxial error angles.

[0008] The load installation matrix is ​​corrected based on the filtered triaxial error angles to obtain the corrected load installation matrix.

[0009] The load inertial attitude is calculated based on the corrected load installation matrix.

[0010] Secondly, embodiments of the present invention also provide an apparatus for improving the accuracy of satellite payload inertial attitude, comprising:

[0011] The error angle determination module is used to monitor the quaternion of the angle between the star sensor's optical axis and the load reference, and determine the three-axis error angle of the star sensor relative to the load reference.

[0012] The filtering module is used to filter the triaxial error angles to obtain the filtered triaxial error angles;

[0013] The correction module is used to correct the load mounting matrix based on the filtered triaxial error angles to obtain the corrected load mounting matrix.

[0014] The calculation module is used to calculate the load inertial attitude based on the corrected load mounting matrix.

[0015] Thirdly, embodiments of the present invention also provide an electronic device, including a memory and a processor, wherein the memory stores a computer program, and when the processor executes the computer program, it implements the method described in any embodiment of this specification.

[0016] Fourthly, embodiments of the present invention also provide a computer-readable storage medium having a computer program stored thereon, which, when executed in a computer, causes the computer to perform the methods described in any embodiment of this specification.

[0017] This invention provides a method and apparatus for improving the accuracy of satellite payload inertial attitude. By monitoring the quaternion of the angle between the star sensor's optical axis and the payload reference, the triaxial error angle of the star sensor relative to the payload reference is determined, unifying the star sensor measurement results to the payload reference. By correcting the payload mounting matrix based on the filtered triaxial error angle, the impact of low-frequency errors in star sensor measurements on payload attitude determination can be effectively reduced. That is, the reference between the payload and the star sensor can be compensated in real-time on-orbit, fully reflecting the satellite's true state compared to ground calibration, thus obtaining a more accurate payload mounting matrix. Calculating the payload inertial attitude using the corrected payload mounting matrix improves the accuracy of the satellite payload's inertial attitude. Therefore, this application can compensate for low-frequency errors and improve the accuracy of the satellite payload's inertial attitude. Attached Figure Description

[0018] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0019] Figure 1 This is a schematic diagram of a method for improving the inertial attitude accuracy of satellite payloads according to an embodiment of the present invention;

[0020] Figure 2(a) is a simulation curve of the X-direction load camera attitude measurement error provided in an embodiment of the present invention;

[0021] Figure 2(b) is a simulation curve of the attitude measurement error of the Y-direction load camera provided in an embodiment of the present invention;

[0022] Figure 2(c) is a simulation curve of the attitude measurement error of the Z-direction load camera provided in an embodiment of the present invention;

[0023] Figure 3 This is a hardware architecture diagram of an electronic device provided in an embodiment of the present invention;

[0024] Figure 4 This is a structural diagram of a device for improving the accuracy of satellite payload inertial attitude according to an embodiment of the present invention. Detailed Implementation

[0025] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are some embodiments of the present invention, but not all embodiments. All other embodiments obtained by those skilled in the art based on the embodiments of the present invention without creative effort are within the scope of protection of the present invention.

[0026] As mentioned above, low-frequency error is one of the most important factors affecting the accuracy of load pointing measurement and pointing determination. Moreover, due to its real-time changing characteristics during load operation, it is difficult to calibrate through pre-measurement or post-processing, which seriously affects the accuracy of load inertial attitude.

[0027] Based on this, the inventors proposed that low-frequency errors can be compensated on-board in real time to eliminate their impact.

[0028] Please refer to Figure 1 This invention provides a method for improving the inertial attitude accuracy of satellite payloads, the method comprising:

[0029] Step 100: Based on the quaternion of the angle between the star sensor optical axis and the load reference, determine the triaxial error angle of the star sensor relative to the load reference;

[0030] Step 102: Filter the triaxial error angles to obtain the filtered triaxial error angles;

[0031] Step 104: Correct the load mounting matrix based on the filtered triaxial error angles to obtain the corrected load mounting matrix;

[0032] Step 106: Calculate the load inertial attitude based on the corrected load installation matrix.

[0033] In this embodiment, the triaxial error angle of the star sensor relative to the load reference is determined by monitoring the quaternion of the angle between the star sensor's optical axis and the load reference, thus unifying the star sensor measurement results under the load reference. By correcting the load mounting matrix based on the filtered triaxial error angle, the impact of low-frequency errors in star sensor measurements on load attitude determination can be effectively reduced. This means the reference between the load and the star sensor can be compensated in real-time on-orbit, fully reflecting the satellite's true state compared to ground calibration, thereby obtaining a more accurate load mounting matrix. Calculating the load inertial attitude using the corrected load mounting matrix improves the accuracy of the satellite load inertial attitude. Therefore, this application can compensate for low-frequency errors and improve the accuracy of the satellite load inertial attitude.

[0034] The following description Figure 1 The execution method for each step is shown.

[0035] First, for step 100, based on the quaternion of the angle between the star sensor optical axis and the load reference, the triaxial error angle of the star sensor relative to the load reference is determined.

[0036] The specific implementation process of step 100 is as follows:

[0037] Step A1, let the monitoring quaternion of the angle between the star sensor's optical axis and the load reference be [q1, q2, q3, q4]. T ;

[0038] Step A2: Normalize the included angle monitoring quaternion to obtain a normalized quaternion;

[0039] Step A3: Determine the pose matrix corresponding to the normalized quaternion;

[0040] Step A4: Determine the three-axis error angle of the star sensor relative to the load reference based on the attitude matrix.

[0041] In step A1, q1, q2, q3, and q4 are the elements of the quaternion, depending on the actual situation.

[0042] In one implementation, step A2 is performed as follows:

[0043] Determine if norm([q1, q2, q3]) is not greater than 1;

[0044] If so, then the quaternion monitoring the angle between the star sensor's optical axis and the load reference is normalized to q. CSr_ests =[q1, q2, q3, sqrt(1–norm([q1, q2, q3]))] T ;

[0045] If not, then normalize the quaternion of the angle between the star sensor's optical axis and the load reference to q. CSr_ests =[0, 0, 0, 1] T ;

[0046] In the formula, norm([q1, q2, q3]) is the 2-norm of [q1, q2, q3]; q CSr_ests It is a normalized quaternion.

[0047] In one implementation, in step A3, the attitude matrix corresponding to the normalized quaternion is determined by the following formula:

[0048] C CSr_ests =q2dcm(q CSr_ests );

[0049] In the formula, C CSr_ests q2dcm is the pose matrix corresponding to the normalized quaternion; q2dcm is the function for converting the normalized quaternion into pose.

[0050] In one implementation, in step A4, the triaxial error angle of the star sensor relative to the load reference is determined by the following formula:

[0051] [α CSr ,β CSr γ CSr ]=dcm2ang(C CSr_ests ,123)

[0052] In the formula, α CSr ,β CSr γ CSr These are the error angles along the X, Y, and Z axes, respectively; dcm2ang is the attitude matrix C based on the 123 rotation sequence. CSr_ests A function to convert to attitude angles.

[0053] Then, for step 104, the corrected load mounting matrix is ​​determined by the following formula:

[0054]

[0055] In the formula, C B_CS The corrected load mounting matrix; The load mounting matrix before correction; α CSr LPF β CSr LPF γ CSr LPF , respectively, are the error angles in the X, Y, and Z axes after filtering; dcm is a function that converts the attitude angles into an attitude matrix according to the 321 rotation sequence.

[0056] By correcting the payload installation matrix based on the filtered triaxial error angles, the impact of low-frequency errors in star sensor measurements on payload attitude determination can be effectively reduced. This means the reference between the payload and the star sensor can be compensated in real-time on-orbit, fully reflecting the satellite's true state compared to ground calibration, thus obtaining a more accurate payload installation matrix. Using the corrected payload installation matrix to calculate the payload inertial attitude improves the accuracy of the satellite's payload inertial attitude calculation.

[0057] Furthermore, the on-orbit data processing method of the star sensor relative to the payload reference of the present invention does not require ground data processing, has better real-time performance, and meets the needs of autonomous and intelligent satellite operation.

[0058] Finally, for step 106, the load inertial attitude includes the quaternion of the load inertial attitude.

[0059] The reliability of the method of the present invention will be demonstrated below with the following embodiments.

[0060] like Figures 2(a) to 2(c) As shown, a low-frequency error with a period of 5600 seconds exists between the star sensor and the two payload cameras, with an error amplitude of tens to hundreds of arcseconds. The method proposed in this invention is introduced at the 6000th second of the simulation. Based on the quaternion monitoring the angle between the star sensor's optical axis and the payload reference, the reference correction is completed after filtering, effectively eliminating the attitude error of the payload cameras.

[0061] Therefore, it can be seen that the present invention can effectively eliminate low-frequency errors in load attitude determination.

[0062] like Figure 3 , Figure 4 As shown, this invention provides a device for improving the accuracy of satellite payload inertial attitude. The device can be implemented in software, hardware, or a combination of both. From a hardware perspective, as... Figure 3 The diagram shown is a hardware architecture diagram of an electronic device containing a device for improving the inertial attitude accuracy of a satellite payload, provided in an embodiment of the present invention. Except for... Figure 3 In addition to the processor, memory, network interface, and non-volatile memory shown, the electronic device in the embodiment may also include other hardware, such as a forwarding chip responsible for processing packets. Taking software implementation as an example, such as... Figure 4 As shown, a device in a logical sense is formed by the CPU of the electronic device in which it is located reading the corresponding computer program from the non-volatile memory into the memory for execution.

[0063] This embodiment provides a device for improving the accuracy of satellite payload inertial attitude, comprising:

[0064] Error angle determination module 400 is used to determine the triaxial error angle of the star sensor relative to the load reference based on the quaternion of the angle between the optical axis of the star sensor and the load reference.

[0065] Filtering module 402 is used to filter the triaxial error angle to obtain the filtered triaxial error angle;

[0066] The correction module 404 is used to correct the load mounting matrix based on the filtered triaxial error angles to obtain the corrected load mounting matrix.

[0067] Calculation module 406 is used to calculate the load inertial attitude based on the corrected load mounting matrix.

[0068] In some implementations, the error angle determination module 400 is used to perform the following operations:

[0069] Let the quaternion for monitoring the angle between the optical axis of the star sensor and the load reference be [q1, q2, q3, q4]. T ;

[0070] The included angle monitoring quaternion is normalized to obtain a normalized quaternion;

[0071] Determine the attitude matrix corresponding to the normalized quaternion;

[0072] The three-axis error angles of the star sensor relative to the load reference are determined based on the attitude matrix.

[0073] In some implementations, when the error angle determination module 400 performs normalization processing on the included angle monitoring quaternion to obtain a normalized quaternion, it performs the following operations:

[0074] Determine if norm([q1, q2, q3]) is not greater than 1;

[0075] If so, then the quaternion monitoring the angle between the star sensor's optical axis and the load reference is normalized to q. CSr_ests =[q1, q2, q3, sqrt(1–norm([q1, q2, q3]))] T ;

[0076] If not, then normalize the quaternion of the angle between the star sensor's optical axis and the load reference to q. CSr_ests =[0, 0, 0, 1] T ;

[0077] In the formula, norm([q1, q2, q3]) is the 2-norm of [q1, q2, q3]; q CSr_ests It is a normalized quaternion.

[0078] In some implementations, the pose matrix corresponding to the normalized quaternion is determined by the following formula:

[0079] C CSr_ests =q2dcm(q CSr_ests );

[0080] In the formula, C CSr_ests q2dcm is the pose matrix corresponding to the normalized quaternion; q2dcm is the function for converting the normalized quaternion into pose.

[0081] In some implementations, the triaxial error angle of the star sensor relative to the load reference is determined by the following formula:

[0082] [α CSr ,β CSr γ CSr ]=dcm2ang(C CSr_ests ,123)

[0083] In the formula, α CSr ,β CSr γ CSr These are the error angles along the X, Y, and Z axes, respectively; dcm2ang is the attitude matrix C based on the 123 rotation sequence. CSr_ests A function to convert to attitude angles.

[0084] In some implementations, the corrected load mounting matrix is ​​determined by the following formula:

[0085]

[0086] In the formula, C B_CS The corrected load mounting matrix; The load mounting matrix before correction; α CSr LPF β CSr LPF γ CSr LPF , respectively, are the error angles in the X, Y, and Z axes after filtering; dcm is a function that converts the attitude angles into an attitude matrix according to the 321 rotation sequence.

[0087] It is understood that the structures illustrated in the embodiments of the present invention do not constitute a specific limitation on a device for improving the inertial attitude accuracy of a satellite payload. In other embodiments of the present invention, a device for improving the inertial attitude accuracy of a satellite payload may include more or fewer components than illustrated, or combine some components, or split some components, or arrange different components. The illustrated components may be implemented in hardware, software, or a combination of software and hardware.

[0088] The information interaction and execution process between the modules in the above-mentioned device are based on the same concept as the method embodiment of the present invention, and the specific details can be found in the description of the method embodiment of the present invention, and will not be repeated here.

[0089] This invention also provides an electronic device, including a memory and a processor. The memory stores a computer program, and when the processor executes the computer program, it implements a method for improving the inertial attitude accuracy of a satellite payload according to any embodiment of this invention.

[0090] This invention also provides a computer-readable storage medium storing a computer program, which, when executed by a processor, causes the processor to perform a method for improving the inertial attitude accuracy of a satellite payload according to any embodiment of this invention.

[0091] Specifically, a system or apparatus equipped with a storage medium may be provided, on which software program code implementing the functions of any of the embodiments described above is stored, and the computer (or CPU or MPU) of the system or apparatus may read and execute the program code stored in the storage medium.

[0092] In this case, the program code read from the storage medium can itself implement the function of any of the above embodiments, and therefore the program code and the storage medium storing the program code constitute part of the present invention.

[0093] Examples of storage media used to provide program code include floppy disks, hard disks, magneto-optical disks, optical disks (such as CD-ROM, CD-R, CD-RW, DVD-ROM, DVD-RAM, DVD-RW, DVD+RW), magnetic tapes, non-volatile memory cards, and ROMs. Alternatively, program code can be downloaded from a server computer via a communication network.

[0094] Furthermore, it should be clear that not only can the program code read by the computer be executed, but also the operating system or other components operating on the computer can be instructed based on the program code to perform some or all of the actual operations, thereby realizing the function of any of the embodiments described above.

[0095] Furthermore, it is understood that the program code read from the storage medium is written to the memory set in the expansion board inserted into the computer or to the memory set in the expansion module connected to the computer. Then, based on the instructions of the program code, the CPU or other components installed on the expansion board or expansion module execute some and all of the actual operations, thereby realizing the function of any of the above embodiments.

[0096] It should be noted that, in this document, relational terms such as "first" and "second" are used only to distinguish one entity or operation from another, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Furthermore, the terms "comprising," "including," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such a process, method, article, or apparatus. Without further limitations, an element defined by the phrase "comprising one..." does not exclude the presence of other identical elements in the process, method, article, or apparatus that includes said element.

[0097] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, and not to limit them; although the present invention has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that modifications can still be made to the technical solutions described in the foregoing embodiments, or equivalent substitutions can be made to some of the technical features; and these modifications or substitutions do not cause the essence of the corresponding technical solutions to deviate from the spirit and scope of the technical solutions of the embodiments of the present invention.

Claims

1. A method for improving the accuracy of satellite payload inertial attitude, characterized in that, include: Based on the quaternion of the angle between the optical axis of the star sensor and the load reference, the triaxial error angle of the star sensor relative to the load reference is determined. The triaxial error angles are filtered to obtain the filtered triaxial error angles. The load installation matrix is ​​corrected based on the filtered triaxial error angles to obtain the corrected load installation matrix. The load inertial attitude is calculated based on the corrected load installation matrix; The method for determining the triaxial error angle of the star sensor relative to the load reference based on the angle monitoring quaternion of the star sensor optical axis relative to the load reference includes: Let the quaternion for monitoring the angle between the optical axis of the star sensor and the load reference be [q1, q2, q3, q4]. T ; The included angle monitoring quaternion is normalized to obtain a normalized quaternion; Determine the attitude matrix corresponding to the normalized quaternion; The three-axis error angles of the star sensor relative to the load reference are determined based on the attitude matrix. The corrected load mounting matrix is ​​determined by the following formula: C B_CS =C0 B_CS*dcm(-α CSr LPF ,-β CSr LPF ,-γ CSr LPF ,321) In the formula, C B_CS C0 is the corrected load mounting matrix; C0 B_CS is the uncorrected load mounting matrix; α CSr LPF β CSr LPF γ CSr LPF , respectively, are the error angles in the X, Y, and Z axes after filtering; dcm is a function that converts the attitude angles into an attitude matrix according to the 321 rotation sequence.

2. The method according to claim 1, characterized in that, The normalization process for the included angle monitoring quaternion, resulting in a normalized quaternion, includes: Determine if norm([q1, q2, q3]) is not greater than 1; If so, then the quaternion monitoring the angle between the star sensor's optical axis and the load reference is normalized to q. CSr_ests =[q1, q2, q3, sqrt(1–norm([q1, q2, q3]))] T ; If not, then normalize the quaternion of the angle between the star sensor's optical axis and the load reference to q. CSr_ests =[0, 0, 0, 1] T ; In the formula, norm([q1, q2, q3]) is the 2-norm of [q1, q2, q3]; q CSr_ests It is a normalized quaternion.

3. The method according to claim 2, characterized in that, The attitude matrix corresponding to the normalized quaternion is determined by the following formula: C CSr_ests =q2dcm(q CSr_ests ); In the formula, C CSr_ests q2dcm is the pose matrix corresponding to the normalized quaternion; q2dcm is the function for converting the normalized quaternion into pose.

4. The method according to claim 3, characterized in that, The triaxial error angles of the star sensor relative to the load reference are determined by the following formula: [a CSr ,b CSr ,c CSr ]=dcm2ang(C CSr_ests ,123) In the formula, α CSr ,β CSr γ CSr These are the error angles along the X, Y, and Z axes, respectively; dcm2ang is the attitude matrix C based on the 123 rotation sequence. CSr_ests A function to convert to attitude angles.

5. A device for improving the accuracy of satellite payload inertial attitude, characterized in that, The apparatus for implementing the method as described in any one of claims 1-4 comprises: The error angle determination module is used to monitor the quaternion of the angle between the star sensor's optical axis and the load reference, and determine the three-axis error angle of the star sensor relative to the load reference. The filtering module is used to filter the triaxial error angles to obtain the filtered triaxial error angles; The correction module is used to correct the load mounting matrix based on the filtered triaxial error angles to obtain the corrected load mounting matrix. The calculation module is used to calculate the load inertial attitude based on the corrected load mounting matrix.

6. A computing device comprising a memory and a processor, wherein the memory stores a computer program, and the processor, when executing the computer program, implements the method as described in any one of claims 1-4.

7. A computer-readable storage medium having a computer program stored thereon, which, when executed in a computer, causes the computer to perform the method of any one of claims 1-4.

Images