Calibration method and device for measurement reference between star sensor and optical load

Abstract

The application relates to a method and device for calibrating a measurement reference between a star sensor and an optical load. The method comprises: adjusting an orbit angular velocity of an optical remote sensing satellite relative to inertial space to zero based on a preset damping method, and orienting an optical axis of the optical load to the inertial space; simultaneously performing a maneuvering scan on the inertial space by using the optical load and the star sensor to obtain a first attitude measurement result of the optical load on the inertial space and a second attitude measurement result of the star sensor on the inertial space; performing filtering processing on the first attitude measurement result and the second attitude measurement result to obtain corresponding first inertial quaternions and second inertial quaternions; and determining an error calibration result of the measurement reference between the optical load and the star sensor based on the first inertial quaternions and the second inertial quaternions. The application can eliminate the measurement deviation between the optical load and the star sensor, and ensure that the optical load and the star sensor keep a consistent measurement reference.

Description

Technical Field

[0001] This invention relates to the field of satellite attitude measurement technology, and in particular to a calibration method and apparatus for a measurement reference between a star sensor and an optical payload. Background Technology

[0002] A satellite consists of a payload and a platform. The platform is equipped with attitude measurement sensors (such as star sensors). Before a flight mission, the positional relationship between the payload and the star sensor is fixed. However, due to the mechanical vibrations during rocket launch and the on-orbit deformation and stress release caused by space heat, optical remote sensing satellites experience significant relative errors between the optical payload and the star sensor. This results in the inability to transmit the high-precision attitude data measured by the star sensor to the payload with equal accuracy, making it difficult to complete spaceborne remote sensing missions with high attitude determination accuracy requirements, such as space geographic mapping.

[0003] Therefore, there is an urgent need for a calibration method and apparatus for the measurement reference between the star sensor and the optical payload to solve the above-mentioned technical problems. Summary of the Invention

[0004] This invention provides a calibration method and apparatus for a measurement reference between a star sensor and an optical payload, which can eliminate measurement deviations between the optical payload and the star sensor and ensure that the optical payload and the star sensor maintain a consistent measurement reference.

[0005] In a first aspect, embodiments of the present invention provide a calibration method for a measurement reference between a star sensor and an optical payload, comprising:

[0006] Based on a pre-defined damping method, the orbital angular velocity of the optical remote sensing satellite relative to inertial space is adjusted to zero, and the optical axis of the optical payload is oriented to inertial space.

[0007] The inertial space is simultaneously scanned by an optical payload and a star sensor to obtain the first attitude measurement result of the optical payload on the inertial space and the second attitude measurement result of the star sensor on the inertial space, respectively.

[0008] The first attitude measurement result and the second attitude measurement result are filtered respectively to obtain the corresponding first inertial quaternion and second inertial quaternion;

[0009] Based on the first inertial quaternion and the second inertial quaternion, the error calibration result of the measurement reference between the optical payload and the star sensor is determined.

[0010] Secondly, embodiments of the present invention also provide a calibration device for a measurement reference between a star sensor and an optical payload, comprising:

[0011] The damping module is used to adjust the orbital angular velocity of the optical remote sensing satellite relative to inertial space to zero based on a preset damping method, and to orient the optical axis of the optical payload to inertial space.

[0012] The measurement module is used to simultaneously perform a maneuver scan of the inertial space using an optical payload and a star sensor, so as to obtain the first attitude measurement result of the optical payload on the inertial space and the second attitude measurement result of the star sensor on the inertial space, respectively.

[0013] The filtering module is used to filter the first attitude measurement result and the second attitude measurement result respectively to obtain the corresponding first inertial quaternion and second inertial quaternion.

[0014] The determination module is used to determine the error calibration result of the measurement reference between the optical payload and the star sensor based on the first inertial quaternion and the second inertial quaternion.

[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 calibrating a measurement reference between a star sensor and an optical payload. First, the optical payload undergoes an orientation conversion, changing from ground-oriented to inertial orientation. Then, a maneuver scan of the inertial space is performed, and the attitude measurement results of the optical payload and the star sensor are filtered to obtain a filtered inertial quaternion. Finally, based on the obtained inertial quaternion, the error between the optical payload and the star sensor measurement reference is determined to ensure that the optical payload and the star sensor maintain a consistent measurement reference. 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 calibration method for a measurement reference between a star sensor and an optical payload provided in an embodiment of the present invention;

[0020] Figure 2 This is a damping curve of the attitude angular velocity of the entire satellite provided in an embodiment of the present invention;

[0021] Figure 3 This is a reference error calibration curve of a star sensor relative to an optical payload provided in an embodiment of the present invention;

[0022] Figure 4 This is a reference error calibration curve of another star sensor relative to the optical payload provided in one embodiment of the present invention;

[0023] Figure 5 This is a hardware architecture diagram of an electronic device provided in an embodiment of the present invention;

[0024] Figure 6 This is a structural diagram of a calibration device for a measurement reference between a star sensor and an optical payload, provided in 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] Please refer to Figure 1 This invention provides a calibration method for a measurement reference between a star sensor and an optical payload, the method comprising:

[0027] Step 100: Based on a preset damping method, adjust the orbital angular velocity of the optical remote sensing satellite relative to inertial space to zero, and orient the optical axis of the optical payload to inertial space;

[0028] Step 102: Simultaneously perform a maneuver scan of the inertial space using an optical payload and a star sensor to obtain the first attitude measurement result of the optical payload on the inertial space and the second attitude measurement result of the star sensor on the inertial space, respectively.

[0029] Step 104: Filter the first attitude measurement result and the second attitude measurement result respectively to obtain the corresponding first inertial quaternion and second inertial quaternion.

[0030] Step 106: Based on the first inertial quaternion and the second inertial quaternion, determine the error calibration result of the measurement reference between the optical payload and the star sensor.

[0031] In this embodiment, the orientation of the optical payload is first converted from ground-oriented to inertial-oriented. Then, a maneuver scan of the inertial space is performed, and the attitude measurements of the optical payload and the star sensor are filtered to obtain a filtered inertial quaternion. Finally, the error between the optical payload and the star sensor's measurement reference is determined based on the obtained inertial quaternion to ensure that the optical payload and the star sensor maintain a consistent measurement reference.

[0032] The following description Figure 1 The execution method for each step is shown.

[0033] First, for step 100, based on a preset damping method, the orbital angular velocity of the optical remote sensing satellite relative to inertial space is adjusted to zero, and the optical axis of the optical payload is oriented to inertial space.

[0034] In practical applications, optical remote sensing satellites perform Earth measurements, with the optical payload conducting Earth remote sensing. The satellite has a certain orbital angular velocity relative to inertial space. During error calibration, this application requires converting the optical payload's orientation from Earth-oriented to inertial space-oriented. Therefore, the orbital angular velocity of the optical remote sensing satellite relative to inertial space needs to be adjusted to zero to facilitate adjusting the optical payload's orientation.

[0035] In some implementations, the specific process of adjusting the orbital angular velocity of an optical remote sensing satellite relative to inertial space to zero based on a preset damping method is as follows:

[0036] Based on the preset damping time and preset damping curve, the orbital angular velocity of the optical remote sensing satellite relative to inertial space is iteratively calculated using the following first formula until the orbital angular velocity of the optical remote sensing satellite relative to inertial space converges to zero.

[0037] The first formula is:

[0038] ω ri =0.5*ω0*(1–cos(π*t) m / T K ));

[0039] In the formula, ω0=[ω x0 ω y0 ω z0 ] T ω0 is the initial angular velocity of the satellite along its three axes, ω x0 ω y0 ω z0 ω represents the initial angular velocity of the satellite along the X, Y, and Z axes, respectively; ri T is the orbital angular velocity of the satellite relative to inertial space. k The preset damping time; tm ∈[0,T k [This refers to the moment during the maneuver.]

[0040] It should be noted that a longer damping time and a flatter damping curve result in a smoother damping process, while a shorter damping time leads to greater fluctuations in the damping process. Users can select preset damping times and preset damping curves as needed; this application does not impose specific limitations.

[0041] Then, for step 102, the optical payload and the star sensor simultaneously perform a maneuver scan of the inertial space. For each scan moment, the optical payload and the star sensor simultaneously measure the inertial space, thus obtaining a matching measurement pair and ensuring temporal consistency.

[0042] Then, for step 104, filtering can remove noise generated during the measurement process, resulting in more stable measurement results.

[0043] Finally, for step 106, based on the first inertial quaternion and the second inertial quaternion, the error calibration result of the measurement reference between the optical payload and the star sensor is determined.

[0044] In some implementations, step 106 is specifically implemented as follows:

[0045] Step A1: Convert the first inertial quaternion into the corresponding first inertial attitude matrix;

[0046] Step A2: Convert the second inertial quaternion into the corresponding second inertial attitude matrix;

[0047] Step A3: Using the optical payload as a reference, and based on the first inertial attitude array, the second inertial attitude array, the mounting array of the optical payload in the whole-star coordinate system, and the mounting array of the star sensor in the whole-star coordinate system, determine the reference error matrix of the star sensor.

[0048] Step A4: Determine the initial error angle of the star sensor relative to the optical payload based on the reference error matrix;

[0049] Step A5: Filter the initial error angle to obtain the filtered error angle;

[0050] Step A6: Use the filtered error angle as the calibration result of the measurement reference between the optical payload and the star sensor.

[0051] Steps A1 and A2 are commonly used methods and will not be described in detail here.

[0052] For step A3, the reference error matrix is ​​determined by the following formula:

[0053]

[0054] In the formula, ΔC tm For t m The reference error matrix at time C; PLB C is the mounting matrix of the optical payload in the whole-satellite coordinate system; PLI,tm For t m The first inertial attitude array at time C; STI,tm For t m The second inertial attitude array at time C; STB This is the mounting matrix of the star sensor in the whole-star coordinate system; the superscript T indicates the transpose of the matrix.

[0055] For step A4, the initial error angle of the star sensor relative to the optical payload is determined by the following formula:

[0056]

[0057] In the formula, DCM2ANG(ΔC) tm ,123) is the matrix ΔC tm Find the corresponding Euler angle function according to the 123 transformation order; θ0 and ψ0 are the initial error angles in the X, Y, and Z directions, respectively; ΔC tm For t m The baseline error matrix at time t.

[0058] For step A5, the initial error angle is filtered, and the filtered error angle is determined by the following formula:

[0059]

[0060] θ1 = θ1 + m*(θ0 - θ1)

[0061] ψ1=ψ1+m*(ψ0-ψ1)

[0062] In the formula, θ0 and ψ0 are the initial error angles in the X, Y, and Z directions, respectively; θ1 and ψ1 are the error angles after filtering in the X, Y, and Z directions, respectively; m is the filtering coefficient.

[0063] Finally, for step A6, the filtered error angle is used as the calibration result of the measurement reference between the optical payload and the star sensor.

[0064] Based on this calibration result, during the actual satellite detection process, the attitude measurement results of the star sensor in inertial space are obtained. Only the filtered error angle needs to be added or subtracted according to the actual situation to obtain the attitude data for the optical payload. This means that the high-precision attitude data measured by the star sensor can be transmitted to the optical payload with equal accuracy, facilitating the remote sensing satellite to accurately complete its detection mission.

[0065] The effectiveness of the method of the present invention is verified by a specific embodiment below.

[0066] Assuming the initial pitch axis of the entire satellite has an orbital angular velocity of 0.0635° / s, damping is achieved using the method proposed in this invention, and the damping curve is as follows. Figure 2 As shown. Damping time T k =160 seconds, then the maneuver time t m After 160 seconds, the satellite velocity drops to zero, and the satellite enters the scanning period. The reference error of the two star sensors is calibrated based on the optical payload.

[0067] Let the reference error angle of star sensor 1 (STS1) after calibration be {5.0, 2.0, -3.0} arcseconds, and the reference error angle of star sensor 2 (STS2) after calibration be {3.0, -5.0, 2.0} arcseconds, with a filter coefficient m = 0.0003. Then the simulation results of the two star sensors are as follows: Figure 3 and Figure 4 As shown in the figure, the errors of the two star sensors relative to the optical payload both converge to the preset value.

[0068] Therefore, it is evident that the calibration method provided in this application is correct and effective.

[0069] Furthermore, the method of the present invention:

[0070] (1) The on-orbit calibration of the reference between the optical payload and the star sensor was realized. Compared with the ground calibration, it reflects the real satellite state after experiencing rocket launch, thermal structural deformation, stress release and other processes, and can obtain a more realistic reference error matrix.

[0071] (2) A method for on-orbit data processing of the reference error between the optical payload and the star sensor is given. It does not require ground data processing, has better real-time performance, and meets the needs of autonomous and intelligent satellite operation.

[0072] (3) Achieving a unified reference between optical payloads and star sensors is beneficial for realizing aerospace remote sensing tasks such as high-precision Earth mapping.

[0073] like Figure 5 , Figure 6As shown, this embodiment of the invention provides a calibration device for a measurement reference between a star sensor and an optical payload. The device embodiment can be implemented through software, hardware, or a combination of both. From a hardware perspective, such as... Figure 5 The diagram shown is a hardware architecture diagram of an electronic device containing a calibration device for a measurement reference between a star sensor and an optical payload, provided in an embodiment of the present invention. (Except for...) Figure 5 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 6 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.

[0074] This embodiment provides a calibration device for a measurement reference between a star sensor and an optical payload, comprising:

[0075] The damping module 600 is used to adjust the orbital angular velocity of the optical remote sensing satellite relative to inertial space to zero based on a preset damping method, and to orient the optical axis of the optical payload to inertial space.

[0076] The measurement module 602 is used to simultaneously perform a maneuver scan of the inertial space using an optical payload and a star sensor, so as to obtain the first attitude measurement result of the optical payload on the inertial space and the second attitude measurement result of the star sensor on the inertial space, respectively.

[0077] The filtering module 604 is used to filter the first attitude measurement result and the second attitude measurement result respectively to obtain the corresponding first inertial quaternion and second inertial quaternion.

[0078] The determination module 606 is used to determine the error calibration result of the measurement reference between the optical payload and the star sensor based on the first inertial quaternion and the second inertial quaternion.

[0079] In some implementations, the damping module 600 is used to perform the following operations:

[0080] Based on the preset damping time and preset damping curve, the orbital angular velocity of the optical remote sensing satellite relative to inertial space is iteratively calculated using the following first formula until the orbital angular velocity of the optical remote sensing satellite relative to inertial space converges to zero.

[0081] The first formula is:

[0082] ω ri =0.5*ω0*(1–cos(π*t) m / TK ));

[0083] In the formula, ω0=[ω x0 ω y0 ω z0 ] T ω0 is the initial angular velocity of the satellite along its three axes, ω x0 ω y0 ω z0 ω represents the initial angular velocity of the satellite along the X, Y, and Z axes, respectively; ri T is the orbital angular velocity of the satellite relative to inertial space. k The preset damping time; t m ∈[0,T k [This refers to the moment during the maneuver.]

[0084] In some implementations, the determining module 606 is used to perform the following operations:

[0085] Convert the first inertial quaternion into the corresponding first inertial attitude matrix;

[0086] The second inertial quaternion is converted into the corresponding second inertial attitude matrix;

[0087] Using the optical payload as a reference, and based on the first inertial attitude array, the second inertial attitude array, the mounting array of the optical payload in the whole-star coordinate system, and the mounting array of the star sensor in the whole-star coordinate system, the reference error matrix of the star sensor is determined.

[0088] The initial error angle of the star sensor relative to the optical payload is determined based on the reference error matrix;

[0089] The initial error angle is filtered to obtain the filtered error angle;

[0090] The filtered error angle is used as the calibration result of the measurement reference between the optical payload and the star sensor.

[0091] In some implementations, the reference error matrix is ​​determined by the following formula:

[0092]

[0093] In the formula, ΔC tm For t m The reference error matrix at time C; PLB C is the mounting matrix of the optical payload in the whole-satellite coordinate system; PLI,tm For t m The first inertial attitude array at time C; STI,tm For t m The second inertial attitude array at time C; STBThis is the mounting matrix of the star sensor in the whole-star coordinate system; the superscript T indicates the transpose of the matrix.

[0094] In some embodiments, the initial error angle of the star sensor relative to the optical payload is determined by the following formula:

[0095]

[0096] In the formula, DCM2ANG(ΔC) tm ,123) is the matrix ΔC tm Find the corresponding Euler angle function according to the 123 transformation order; θ0 and ψ0 are the initial error angles in the X, Y, and Z directions, respectively; ΔC tm For t m The baseline error matrix at time t.

[0097] In some implementations, the initial error angle is filtered to obtain the filtered error angle, which is determined by the following formula:

[0098]

[0099] θ1 = θ1 + m*(θ0 - θ1)

[0100] ψ1=ψ1+m*(ψ0-ψ1)

[0101] In the formula, θ0 and ψ0 are the initial error angles in the X, Y, and Z directions, respectively; θ1 and ψ1 are the error angles after filtering in the X, Y, and Z directions, respectively; m is the filtering coefficient.

[0102] It is understood that the structures illustrated in the embodiments of the present invention do not constitute a specific limitation on a calibration device for a measurement reference between a star sensor and an optical payload. In other embodiments of the present invention, a calibration device for a measurement reference between a star sensor and an optical payload may include more or fewer components than illustrated, or combine some components, or split some components, or have different component arrangements. The illustrated components may be implemented in hardware, software, or a combination of software and hardware.

[0103] 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 in the method embodiment of the present invention, and will not be repeated here.

[0104] 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 calibration method for a measurement reference between a star sensor and an optical payload according to any embodiment of this invention.

[0105] 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 calibration method for a measurement reference between a star sensor and an optical payload according to any embodiment of this invention.

[0106] 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.

[0107] 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.

[0108] 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.

[0109] 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.

[0110] 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.

[0111] 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.

[0112] 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 calibration method for a measurement reference between a star sensor and an optical payload, characterized in that, include: Based on a pre-defined damping method, the orbital angular velocity of the optical remote sensing satellite relative to inertial space is adjusted to zero, and the optical axis of the optical payload is oriented to inertial space. The inertial space is simultaneously scanned by an optical payload and a star sensor to obtain the first attitude measurement result of the optical payload on the inertial space and the second attitude measurement result of the star sensor on the inertial space, respectively. The first attitude measurement result and the second attitude measurement result are filtered respectively to obtain the corresponding first inertial quaternion and second inertial quaternion; Based on the first inertial quaternion and the second inertial quaternion, the error calibration result of the measurement reference between the optical payload and the star sensor is determined; The aforementioned damping method, which adjusts the orbital angular velocity of the optical remote sensing satellite relative to inertial space to zero, includes: Based on the preset damping time and preset damping curve, the orbital angular velocity of the optical remote sensing satellite relative to inertial space is iteratively calculated using the following first formula until the orbital angular velocity of the optical remote sensing satellite relative to inertial space converges to zero. The first formula is: ω ri = 0.5 * ω0 * (1 - cos(π * t m / T K )); In the formula, ω0=[ω x0 ω y0 ω z0 ] T ω0 is the initial angular velocity of the satellite along its three axes, ω x0 ω y0 ω z0 ω represents the initial angular velocity of the satellite along the X, Y, and Z axes, respectively; ri This refers to the satellite's orbital angular velocity relative to inertial space. T k The preset damping time; t m ∈[0,T k [This refers to the moment during the maneuver.] 2. The method according to claim 1, characterized in that, The step of determining the error calibration result of the measurement reference between the optical payload and the star sensor based on the first inertial quaternion and the second inertial quaternion includes: Convert the first inertial quaternion into the corresponding first inertial attitude matrix; The second inertial quaternion is converted into the corresponding second inertial attitude matrix; Using the optical payload as a reference, and based on the first inertial attitude array, the second inertial attitude array, the mounting array of the optical payload in the whole-star coordinate system, and the mounting array of the star sensor in the whole-star coordinate system, the reference error matrix of the star sensor is determined. The initial error angle of the star sensor relative to the optical payload is determined based on the reference error matrix; The initial error angle is filtered to obtain the filtered error angle; The filtered error angle is used as the calibration result of the measurement reference between the optical payload and the star sensor.

3. The method according to claim 2, characterized in that, The reference error matrix is ​​determined by the following formula: In the formula, For t m The reference error matrix at time step; This is the mounting matrix of the optical payload in the whole-satellite coordinate system; For t m The first inertial attitude array at any given moment; For t m The second inertial attitude array at time; This is the mounting matrix of the star sensor in the whole-star coordinate system; the superscript T indicates the transpose of the matrix.

4. The method according to claim 3, characterized in that, The initial error angle of the star sensor relative to the optical payload is determined by the following formula: In the formula, DCM2ANG(ΔC) tm ,123) is the matrix ΔC tm Find the corresponding Euler angle function according to the 123 transformation order; θ0 and ψ0 are the initial error angles in the X, Y, and Z directions, respectively; For t m The reference error matrix at time t.

5. The method according to claim 4, characterized in that, The initial error angle is filtered to obtain the filtered error angle, which is determined by the following formula: θ1=θ1+m (θ0-θ1) ψ1=ψ1+m (ψ0-ψ1) In the formula, θ0 and ψ0 are the initial error angles in the X, Y, and Z directions, respectively; θ1 and ψ1 are the error angles after filtering in the X, Y, and Z directions, respectively; m is the filtering coefficient.

6. A calibration device for a measurement reference between a star sensor and an optical payload, characterized in that, The apparatus for implementing the method as described in any one of claims 1-5 comprises: The damping module is used to adjust the orbital angular velocity of the optical remote sensing satellite relative to inertial space to zero based on a preset damping method, and to orient the optical axis of the optical payload to inertial space. The measurement module is used to simultaneously perform a maneuver scan of the inertial space using an optical payload and a star sensor, so as to obtain the first attitude measurement result of the optical payload on the inertial space and the second attitude measurement result of the star sensor on the inertial space, respectively. The filtering module is used to filter the first attitude measurement result and the second attitude measurement result respectively to obtain the corresponding first inertial quaternion and second inertial quaternion. The determination module is used to determine the error calibration result of the measurement reference between the optical payload and the star sensor based on the first inertial quaternion and the second inertial quaternion.

7. 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 1-5.

8. 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 1-5.

Images