Methods, computing devices, and storage media for solar transit prediction and solar transit avoidance

Abstract

Embodiments of the present invention relate to a method, computing device, and medium for solar outage prediction and avoidance. The solar outage prediction method includes obtaining the sun's position vector based on satellite broadcasts from the local satellite; obtaining multiple transformation matrices based on the satellite's position, velocity, attitude, and laser payload installation location information, including at least: a first transformation matrix from the inertial frame to the center-of-mass orbital coordinate system, a second transformation matrix from the center-of-mass orbital coordinate system to the satellite's coordinate system, and a third transformation matrix from the satellite's coordinate system to the laser payload's coordinate system, wherein the third transformation matrix is ​​calculated at least based on the laser payload's installation orientation and installation error angle along the three axes of the satellite; and calculating the field-of-view angle between the sun's position vector and the laser payload's line of sight from the local satellite based on the multiple transformation matrices; to determine whether the local satellite has entered the solar outage region. This improves the accuracy of solar outage prediction and enhances the efficiency of inter-satellite communication link recovery after a solar outage.

Description

Technical Field

[0001] The embodiments of the present invention generally relate to the field of satellite laser communication, and more specifically to a method, computing device and storage medium for solar outage prediction and avoidance. Background Technology

[0002] Solar interference is a common physical phenomenon in celestial motion, and it can significantly impact inter-satellite laser communication. For example, during solar interference, direct sunlight strikes the photosensitive surface of the tracking detector, severely degrading the optical tracking performance of the laser payload; it also causes the laser payload to overheat, potentially burning out the detector; and it affects the optical communication performance of the laser payload, increasing the overall bit error rate. Therefore, satellites need to take avoidance measures when encountering solar interference.

[0003] Traditional methods for predicting and avoiding solar outages typically involve only one or two steps of conversion between the coordinate system of sunlight and the coordinate system of the satellite laser payload. This traditional method does not take into account the deviations caused by installation errors of the satellite payload. Furthermore, traditional solar outage avoidance methods usually involve deflecting the two-dimensional turntable of the laser payload on the satellite directly exposed to the sun at the location of the outage, thus causing the laser payload to deviate from direct sunlight. This deflection of the laser payload's two-dimensional turntable reduces the efficiency of re-acquiring the satellite after the outage ends, resulting in long inter-satellite link re-establishment time. Moreover, inter-satellite communication links are frequently lost due to solar outages, requiring ground control stations to perform pointing operations to re-establish inter-satellite links, leading to high time and labor costs.

[0004] In summary, the shortcomings of traditional methods for predicting and avoiding solar outages are: low efficiency and inaccuracy in predicting solar outages, and low efficiency and high cost in restoring inter-satellite communication links. Summary of the Invention

[0005] To address the aforementioned problems, this invention provides a method, computing device, and storage medium for solar outage prediction and avoidance, which can improve the efficiency and accuracy of solar outage prediction and enhance the recovery efficiency of inter-satellite communication links after leaving the solar outage region.

[0006] According to a first aspect of the present invention, a method for predicting solar outage is provided, comprising: acquiring the position vector of the sun based on satellite broadcasts from a local satellite; acquiring multiple transformation matrices based on the position information, velocity information, attitude information, and laser payload installation position information of the local satellite body; the multiple transformation matrices including at least: a first transformation matrix from the inertial frame to the center-of-mass orbital coordinate system, a second transformation matrix from the center-of-mass orbital coordinate system to the satellite body coordinate system, and a third transformation matrix from the satellite body coordinate system to the laser payload coordinate system, the third transformation matrix being calculated at least based on the installation direction and installation error angle of the laser payload on the three axes of the satellite body; calculating the field of view angle between the sun's position vector and the laser payload's line of sight of the local satellite based on the acquired multiple transformation matrices; and determining whether the local satellite has entered a solar outage region based on the field of view angle.

[0007] According to a second aspect of the present invention, a method for avoiding solar interference is provided, comprising: determining whether a local satellite has entered a solar interference region according to the method provided in the first aspect of the present invention; and performing a solar interference avoidance operation in response to determining that the local satellite has entered a solar interference region, wherein the solar interference avoidance operation comprises at least: adjusting the angle of a precision-tracking fast-reflecting mirror of the local satellite's laser payload by an amount exceeding a predetermined threshold, such that sunlight deflects out of the camera field of view of the local satellite's laser payload.

[0008] According to a third aspect of the invention, a computing device is provided, the computing device comprising: at least one processing unit; at least one memory coupled to the at least one processing unit and storing instructions for execution by the at least one processing unit, the instructions, when executed by the at least one processing unit, causing the device to perform the steps of the method according to the first aspect.

[0009] According to a fourth aspect of the present invention, a computer-readable storage medium is provided, on which a computer program is stored, which, when executed by a machine, implements the method according to the first aspect.

[0010] According to a fifth aspect of the present invention, a computer program product is also provided, comprising a computer program that, when executed by a machine, performs the method of the first aspect of the present invention.

[0011] In some embodiments, based on the satellite's position information, velocity information, attitude information, and laser payload installation position information, multiple transformation matrices are obtained, including: calculating a first transformation matrix from the inertial frame to the center-of-mass orbital coordinate system based on the satellite's position information and velocity information; calculating a second transformation matrix from the center-of-mass orbital coordinate system to the satellite's body coordinate system based on the satellite's attitude information; and calculating a third transformation matrix from the satellite's body coordinate system to the laser payload coordinate system based on the installation direction and installation error angle of the laser payload on the three axes of the satellite body, in the rotation order of the x-axis, y-axis, and z-axis, wherein the installation error angle includes the x-axis error angle, y-axis error angle, and z-axis error angle.

[0012] In some embodiments, calculating the second transformation matrix from the centroid orbit coordinate system to the satellite body coordinate system based on the satellite attitude information of the local satellite includes: obtaining the attitude angles of the local satellite body based on the satellite attitude information of the local satellite, the attitude angles including roll angle, pitch angle and yaw angle; and calculating the second transformation matrix from the centroid orbit coordinate system to the satellite body coordinate system in the attitude adjustment order of roll angle, pitch angle and yaw angle for the attitude information of the satellite body.

[0013] In some embodiments, calculating the field-of-view angle between the Sun's position vector and the line-of-view of the local satellite's laser payload based on the acquired multiple transformation matrices includes: transforming the Sun's position vector to the laser payload coordinate system based on the acquired first, second, and third transformation matrices; calculating the pointing angle of the local satellite's laser payload turntable; and calculating the field-of-view angle between the Sun's position vector and the line-of-view of the local satellite's laser payload based on the pointing angle of the local satellite's laser payload turntable and the Sun's position vector.

[0014] In some embodiments, determining whether the local satellite has entered the solar interference region based on the field of view angle includes: calculating a first field of view angle between the local satellite's laser payload line of sight and the sun's position vector in the y-axis direction, and a second field of view angle between the local satellite's laser payload line of sight and the sun's position vector in the z-axis direction; and determining that the local satellite has entered the solar interference region in response to determining that the first field of view angle and / or the second field of view angle is less than or equal to a predetermined angle.

[0015] In some embodiments, the laser payload included in the spaceborne laser communication equipment is further configured with a light trapping device, which is disposed at a predetermined position inside the laser payload, and the inner wall of the light trapping device is provided with a light-absorbing material.

[0016] In some embodiments, adjusting the angle of the precision reflector of the local satellite's laser payload by an amount exceeding a predetermined threshold, so that sunlight deflects out of the camera's field of view of the local satellite's laser payload, includes: deflecting the precision reflector of the laser payload by a predetermined angle, so that sunlight entering the laser payload is deflected and then enters the light trapping device; and causing the sunlight entering the light trapping device to undergo multiple reflections in the light trapping device, thereby being gradually absorbed by the light-absorbing material in the light trapping device until the reflector is completely attenuated.

[0017] The above scheme fully considers various parameters such as the sun, the satellite body, and the payload installation, thereby accurately and efficiently calculating the field of view angle between the sun's position vector and the laser payload's line of sight of the satellite. This improves the efficiency and accuracy of solar outage prediction, allowing for timely detection of whether the satellite has entered the solar outage region and enabling precise solar outage avoidance operations. If the satellite is confirmed to have entered the solar outage region, the angle of the precision tracking mirror of the satellite's laser payload is adjusted by an amount exceeding a predetermined threshold, causing sunlight to deflect out of the camera's field of view of the satellite's laser payload, thus improving the efficiency of inter-satellite communication link recovery after leaving the solar outage region.

[0018] It should be understood that the description in this section is not intended to identify key or essential features of the embodiments of the present invention, nor is it intended to limit the scope of the invention. Other features of the invention will become readily apparent from the following description. Attached Figure Description

[0019] The above and other features, advantages, and aspects of the various embodiments of the present invention will become more apparent from the accompanying drawings and the following detailed description. In the drawings, the same or similar reference numerals denote the same or similar elements.

[0020] Figure 1 A schematic diagram is shown for implementing a solar interference avoidance environment for a spaceborne laser communication device according to an embodiment of the present invention.

[0021] Figure 2 A schematic diagram of a satellite structure for avoiding solar interference in a spaceborne laser communication device according to an embodiment of the present invention is shown.

[0022] Figure 3 A flowchart of a method for solar outage prediction according to an embodiment of the present invention is shown.

[0023] Figure 4 A flowchart of a method for obtaining multiple transformation matrices according to an embodiment of the present invention is shown.

[0024] Figure 5 A schematic diagram of the angle between the solar vector and the laser payload according to an embodiment of the present invention is shown.

[0025] Figure 6A schematic diagram of the structure of a satellite laser payload optical system according to an embodiment of the present invention is shown.

[0026] Figure 7 A schematic diagram of the optical path of the satellite laser payload optical system structure according to the present invention is shown.

[0027] Figure 8 A block diagram of an electronic device according to an embodiment of the present invention is shown.

[0028] Figure 9 A schematic diagram of the deflection of the precision reflector according to an embodiment of the present invention is shown.

[0029] Figure 10 A flowchart of a method for avoiding solar overpasses according to an embodiment of the present invention is shown. Detailed Implementation

[0030] The following description, in conjunction with the accompanying drawings, illustrates exemplary embodiments of the present invention, including various details to aid understanding. These details should be considered merely exemplary. Therefore, those skilled in the art will recognize that various changes and modifications can be made to the embodiments described herein without departing from the scope and spirit of the invention. Similarly, for clarity and brevity, descriptions of well-known functions and structures are omitted in the following description.

[0031] The term "comprising" and its variations as used herein signify open inclusion, i.e., "including but not limited to". Unless otherwise stated, the term "or" means "and / or". The term "based on" means "at least partially based on". The terms "one example embodiment" and "one embodiment" mean "at least one example embodiment". The term "another embodiment" means "at least one additional embodiment". The terms "first", "second", etc., may refer to different or the same objects. Other explicit and implicit definitions may also be included below.

[0032] As described above, the shortcomings of traditional methods for predicting and avoiding solar outages are that solar outage prediction is inefficient and inaccurate, and the recovery of inter-satellite communication links is inefficient and costly.

[0033] To at least partially address one or more of the aforementioned problems and other potential issues, an exemplary embodiment of the present invention proposes a scheme for solar interference avoidance in spaceborne laser communication equipment. In this scheme, the position vector of the sun is obtained based on the satellite broadcast of the local satellite; multiple transformation matrices are obtained based on the satellite's position information, velocity information, attitude information, and laser payload installation location information; the field of view angle between the sun's position vector and the laser payload's line of sight is calculated based on the obtained multiple transformation matrices; based on the field of view angle, it is determined whether the local satellite has entered the solar interference region; and in response to determining that the local satellite has entered the solar interference region, a solar interference avoidance operation is performed; furthermore, multiple transformation matrices... The invention includes at least: a first transformation matrix from the inertial frame to the center-of-mass orbital coordinate system, a second transformation matrix from the center-of-mass orbital coordinate system to the satellite body coordinate system, and a third transformation matrix from the satellite body coordinate system to the laser payload coordinate system. The third transformation matrix is ​​calculated based at least on the installation orientation and installation error angle of the laser payload along the three axes of the satellite body. Therefore, this invention fully considers various parameters such as the sun, the satellite body, and the payload installation, especially taking into account the installation deviation of the laser payload. This allows for accurate and efficient calculation of the field of view angle between the sun's position vector and the laser payload's line of sight, improving the accuracy of solar interference prediction. This enables the determination of whether the satellite has entered the solar interference region, allowing for timely and precise solar interference avoidance operations. Therefore, this invention can promptly avoid solar interference when the satellite enters the region and efficiently restore inter-satellite communication links after the satellite leaves the region.

[0034] Figure 1 A schematic diagram is shown for an operating environment 100 for implementing solar interference avoidance for a spaceborne laser communication device according to an embodiment of the present invention. Figure 1 As shown, the operating environment 100 includes a binary orbit 50 for Earth and the Sun (Sun 10 and Sun 20 illustrate two different positions of the Sun), and Satellite 1 30 and Satellite 2 40. Satellite 1 30 and Satellite 2 40 operate along orbit 50, and an inter-satellite laser communication link can be established between Satellite 1 30 and Satellite 2 40.

[0035] like Figure 1 As shown, the laser payload of Satellite 2-40 is facing away from the sun, which is the back-viewing end of the solar interference; the laser payload of Satellite 1-30 is facing the sun, which is the looking end of the solar interference. When the looking end of the laser payload faces the sun, and the angle between the line of sight of the laser payload and the sunlight is less than a certain range, a solar interference phenomenon occurs. For example, Figure 1 When the satellite is at position 10, the angle between the line of sight of the laser payload of satellite-30 and the sunlight is θ1; when the satellite is at position 20, the angle between the line of sight of the laser payload of satellite-30 and the sunlight is θ2. Figure 1The arc 60 indicates that the direction of solar interference is from Sun 10 to Sun 20. When θ1≤3°, Satellite 30 enters the solar interference region; when θ2<-3°, Satellite 30 leaves the solar interference region. In this embodiment, the angle between the laser payload line of the satellite and the sunlight is θ. When θ∈(-3°,+3°), the satellite is in the solar interference region.

[0036] It should be understood that Satellite 1 30 and Satellite 2 40 operate along orbit 50, and both satellites may enter the solar interference region (be at the solar interference observation end). Satellites located at the solar interference observation end need to perform solar interference avoidance operations to avoid equipment damage or communication interference caused by direct sunlight. The solution provided in this application can improve the execution efficiency and accuracy of solar interference avoidance operations, and improve the efficiency of restoring inter-satellite communication links after the solar interference ends.

[0037] Please refer to Figure 2 It shows a schematic diagram of the satellite's structure. Figure 2 The structure of satellite 200 shown can be that of either satellite 30 or satellite 40. For example... Figure 2 As shown, satellite 200 includes a control unit 212, a broadcast communication unit 214, and a laser payload 216. Control unit 212 is communicatively linked to broadcast communication unit 214, and control unit 212 is also communicatively linked to laser payload 216. For example, satellite 1 30 and satellite 2 40 can respectively perform broadcast communication through their respective broadcast communication units, and can respectively perform inter-satellite laser communication through their respective laser payloads. It should be understood that the communication interfaces in the above examples include, but are not limited to, CAN communication interfaces, RS422 communication interfaces, and LVDS communication interfaces. The above communication interfaces are only illustrative of data interaction via multiple communication interfaces; other communication interfaces may also be used in this invention.

[0038] In some embodiments, the satellite's control unit is also configured to interact with broadcast communication units, laser payloads, and control units of other satellites. In some embodiments, the control unit 212 may have one or more processing units, including dedicated processing units such as graphics processing units (GPUs), field programmable gate arrays (FPGAs), application-specific integrated circuits (ASICs), or general-purpose computing on graphics processing units (GPGPUs), as well as general-purpose processing units such as CPUs.

[0039] Regarding the broadcast communication unit 214, it is used for satellite broadcasting based on the local satellite to obtain the position vector of the sun.

[0040] Regarding the control unit 212, it is used to acquire multiple transformation matrices based on the satellite's position information, velocity information, attitude information, and laser payload installation position information. The multiple transformation matrices include at least: a first transformation matrix from the inertial frame to the center-of-mass orbital coordinate system, a second transformation matrix from the center-of-mass orbital coordinate system to the satellite body coordinate system, and a third transformation matrix from the satellite body coordinate system to the laser payload coordinate system. The third transformation matrix is ​​calculated based at least on the installation direction and installation error angle of the laser payload on the three axes of the satellite body.

[0041] Regarding the control unit 212, it is also used to calculate the field of view angle between the position vector of the sun and the line of sight of the local star's laser payload, based on the acquired multiple transformation matrices.

[0042] Regarding the control unit 212, it is also used to determine whether the local satellite has entered the solar interference region based on the field of view angle.

[0043] Regarding laser payload 216, it is used to perform solar interference avoidance operations in response to determining that the local satellite has entered the solar interference region.

[0044] Regarding the control unit 212, it is also configured to perform a solar interference avoidance operation in response to determining that the local satellite has entered the solar interference region, wherein the solar interference avoidance operation includes at least: adjusting the angle of the precision-tracking fast mirror of the local satellite's laser payload by an amount exceeding a predetermined threshold, so that sunlight deflects out of the camera field of view of the local satellite's laser payload.

[0045] Figure 3 A flowchart of a method 300 for solar outage prediction according to an embodiment of the present invention is shown. Method 300 may be derived from, for example... Figure 1 The satellites shown (Satellite 1-30, Satellite 2-40) can also be used. Figure 8 The method is performed at the illustrated electronic device 800. It should be understood that method 300 may also include additional steps not shown and / or the steps shown may be omitted, and the scope of the invention is not limited in this respect.

[0046] In step 302, the satellites (Satellite 1 30, Satellite 2 40) acquire the position vector of the sun based on their own satellite broadcasts.

[0047] Regarding this satellite, in the embodiments of this application specification, the satellite at the observation point during the solar outage is used as the "this satellite," for example, please refer to... Figure 1 , Figure 1The following is a detailed description of the scheme using Satellite-30 as the local satellite. It should be understood that, as the satellites operate, Satellite-40 can also be used as the local satellite when it is at the point of solar out-of-focus observation.

[0048] In step 304, the satellites (Satellite 1 30 and Satellite 2 40) acquire multiple transformation matrices based on the satellite's position information, velocity information, attitude information, and laser payload installation position information. These multiple transformation matrices include at least: a first transformation matrix from the inertial frame to the center-of-mass orbital coordinate system, a second transformation matrix from the center-of-mass orbital coordinate system to the satellite body coordinate system, and a third transformation matrix from the satellite body coordinate system to the laser payload coordinate system. The third transformation matrix is ​​calculated based at least on the installation direction and installation error angle of the laser payload on the three axes of the satellite body.

[0049] Regarding multiple transformation matrices, for example, the position vector A1 of the sun obtained by the local satellite from satellite broadcasts is usually the position vector of the sun in an inertial frame (e.g., the J2000 coordinate system). However, to determine whether the local satellite has entered the solar interference region, it is necessary to determine the angle between the sun's position vector and the local satellite's laser payload line of sight. The initial value of the local satellite's laser payload line of sight is usually in the satellite's laser payload coordinate system. In order to make the angle between the initial values ​​of the sun's position vector and the local satellite's laser payload line of sight calculated in the same coordinate system, this invention adopts the method of calculating the "angle between the sun's position vector and the local satellite's laser payload line of sight" in the satellite's laser payload coordinate system. For example, the transformation process of the sun's position vector A1 from "inertial frame → laser payload coordinate system" is completed through multiple transformation matrices.

[0050] Regarding multiple transformation matrices, for example, these include: a first transformation matrix for transforming the inertial frame to the center-of-mass orbit coordinate system; a second transformation matrix for transforming the center-of-mass orbit coordinate system to the center-of-mass orbit coordinate system; and a third transformation matrix for transforming the center-of-mass orbit coordinate system to the satellite body coordinate system and then to the laser payload coordinate system. The following will combine... Figure 4 The methods for obtaining multiple transformation matrices are explained in detail here, but will not be repeated.

[0051] In step 306, the satellites (Satellite 1 30, Satellite 2 40) calculate the field of view angle between the sun's position vector and the laser payload line of view of the local satellite based on the acquired multiple transformation matrices.

[0052] In some embodiments, calculating the angle between the Sun's position vector and the line-of-sight of the satellite's laser payload includes: calculating the Sun's position vector in the laser payload's field of view and the pointing vector of the laser payload's two-dimensional turntable based on multiple transformation matrices, so as to calculate the angle between the Sun's position vector and the line-of-sight of the satellite's laser payload. It should be understood that the pointing vector of the satellite's laser payload's two-dimensional turntable (i.e., the line-of-sight of the laser payload, i.e., the direction of the satellite's communication laser emission); and the adjustment of the laser payload's line-of-sight direction, are also achieved by controlling the two-dimensional turntable through the satellite's control unit.

[0053] In the above scheme, the angle between the sun's position vector and the satellite's laser payload line of sight is calculated by transforming the sun's position vector from the inertial frame to the satellite's laser payload coordinate system using multiple transformation matrices. For example, this achieves a sequential transformation from "inertial frame → center of mass orbit coordinate system → satellite body coordinate system → laser payload coordinate system." This transformation method is more direct and efficient in computation, as it involves continuous matrix multiplication of the initial value of the sun's position vector (in the inertial frame) with multiple transformation matrices. Compared to transforming the initial value of the laser payload line of sight from "laser payload coordinate system → inertial frame," which requires multiple matrix transpose operations, consuming more computation time and power, and compared to the matrix multiplication method used in this invention (which directly uses the original parameters of the laser payload, potentially reducing error sources), transpose calculations are also prone to introducing error risks. Therefore, for the satellite's operating system, the above-mentioned "inertial frame → laser payload coordinate system" transformation provided by this invention has a simpler computational flow, faster response speed under high-frequency computational demands, and lower power consumption.

[0054] Secondly, the aforementioned method offers advantages in engineering implementation and reliability. The transformation process from "inertial frame → center-of-mass orbital coordinate system → satellite body coordinate system → laser payload coordinate system" is more consistent with physical perception. Therefore, the relevant code of the satellite laser payload turntable operating system is more readable and easier to maintain, facilitating satellite system verification and debugging. Furthermore, centering the system on the laser payload increases its modularity, making it easier to maintain and expand to support multiple lines of sight or multiple payloads.

[0055] In step 308, the satellites (Satellite 1 30, Satellite 2 40) determine whether the local satellite has entered the solar interference region based on the field of view angle.

[0056] In some embodiments, determining whether the local satellite has entered the solar interference region based on the field of view angle includes: calculating a first field of view angle between the local satellite's laser payload line of sight and the sun's position vector in the y-axis direction, and a second field of view angle between the local satellite's laser payload line of sight and the sun's position vector in the z-axis direction; and determining that the local satellite has entered the solar interference region in response to determining that the first field of view angle and / or the second field of view angle is less than or equal to a predetermined angle.

[0057] For example, please refer to Figure 5 The diagram illustrates the first field of view angle θ between the laser payload's line of sight and the Sun's position vector along the y-axis. y And the second field of view θ between the laser payload line of this satellite and the position vector of the sun in the z-axis direction. z .

[0058] In some embodiments, for example, the first field of view angle is θ y The first field of view is θ z The predetermined critical value for solar eclipse is β; for example, in response to determining that -β ≤ the first field of view angle is θ. y When ≤β, it is determined that the local star has entered the solar transit region; also, for example, in response to determining that -β≤ the second field of view angle is θ z When ≤β, the local star is determined to have entered the solar transit region; for example, β=3°.

[0059] Therefore, the above-mentioned solution provided by the present invention can ensure that when the angle between the satellite's laser payload and the sun's position vector in any of the y-axis and z-axis meets a predetermined angle, the satellite is confirmed to have entered the solar interference region.

[0060] Figure 4 A flowchart illustrating a method 400 for obtaining multiple transformation matrices according to an embodiment of the present invention is shown. Method 400 may be performed by, for example... Figure 1 The satellites shown (Satellite 1-30, Satellite 2-40) can also be used. Figure 8 The method is performed at the illustrated electronic device 800. It should be understood that method 400 may also include additional steps not shown and / or the steps shown may be omitted, and the scope of the invention is not limited in this respect.

[0061] In step 402, the satellites (Satellite 1 30, Satellite 2 40) calculate the first transformation matrix from the inertial frame to the center-of-mass orbital coordinate system based on the position and velocity information of the satellite body itself.

[0062] Regarding the calculation of the first transformation matrix, for example, the position information of the local satellite in the inertial frame (J2000 coordinate system). Speed ​​information After normalizing the position and velocity information, we obtain formula (1):

[0063] (1).

[0064] The Z-axis of the centroid orbital coordinate system points from the satellite to the Earth's center, and the corresponding Z-direction vector in the J2000 coordinate system is given by formula (2):

[0065] (2).

[0066] The Y-axis direction of the center-of-mass orbital coordinate system is opposite to the normal direction of the orbital plane. In the J2000 coordinate system, the satellite's velocity vector and position vector are in the orbital plane. Therefore, the Y-direction vector satisfies formula (3):

[0067] (3).

[0068] It should be understood that the X, Y, and Z directions satisfy the right-hand coordinate relationship, and therefore conform to formula (4):

[0069] (4).

[0070] Therefore, the transformation matrix from the inertial frame to the center-of-mass orbital coordinate system, i.e., the first transformation matrix, is formula (5):

[0071] (5).

[0072] In step 404, the satellites (Satellite 1 30, Satellite 2 40) calculate the second transformation matrix from the center-of-mass orbit coordinate system to the satellite body coordinate system based on the satellite attitude information of the satellite itself.

[0073] In some embodiments, calculating the second transformation matrix from the centroid orbit coordinate system to the satellite body coordinate system based on the satellite's attitude information includes: obtaining the attitude angles of the satellite body, including roll angle, pitch angle, and yaw angle, based on the satellite's attitude information; and calculating the second transformation matrix from the centroid orbit coordinate system to the satellite body coordinate system in the order of attitude adjustment of roll angle, pitch angle, and yaw angle, based on the attitude information of the satellite body. Thus, the second transformation matrix is ​​constructed by fully considering the attitude of the satellite body at multiple angles, resulting in a more accurate transformation result.

[0074] Regarding the calculation of the second transformation matrix, for example, according to the definition of the satellite body coordinate system, the satellite body coordinate system and the centroid orbit coordinate system have the same orientation. The attitude matrix is ​​calculated from the attitude angles, which include the roll angle, pitch angle, and yaw angle. The attitude angles are expressed as... The values ​​correspond to the roll angle, pitch angle, and yaw angle respectively. The order of attitude angle adjustment is roll angle → pitch angle → yaw angle. The transformation matrix from the centroid orbit coordinate system to the satellite body coordinate system, i.e., the second transformation matrix, is calculated as shown in formula (6):

[0075] (6).

[0076] Therefore, the transformation matrix from the inertial frame to the satellite body coordinate system (fourth transformation matrix) is as shown in formula (7):

[0077] (7).

[0078] In step 406, the satellites (Satellite 1 30, Satellite 2 40) calculate the third transformation matrix from the satellite body coordinate system to the laser payload coordinate system based on the installation direction and installation error angle of the laser payload on the three axes of the satellite body, in the rotation order of the x-axis, y-axis, and z-axis. The installation error angle includes the x-axis error angle, the y-axis error angle, and the z-axis error angle.

[0079] In some embodiments, calculating the third transformation matrix from the satellite body coordinate system to the laser payload coordinate system based on the laser payload installation location information of the satellite body includes: obtaining the installation direction and installation error angle of the laser payload on the three axes of the satellite body based on the laser payload installation location information of the satellite body, wherein the installation error angle includes the x-axis error angle, y-axis error angle, and z-axis error angle; and calculating the third transformation matrix from the satellite body coordinate system to the laser payload coordinate system based on the installation direction and installation error angle of the laser payload, in the rotation order of the x-axis, y-axis, and z-axis. This fully considers the installation error of the laser payload in calculating the third transformation matrix, further improving the accuracy of the transformation result.

[0080] Regarding the calculation of the third transformation matrix, for example, if laser payload 216 is installed on the satellite body of satellite-30, the transformation relationship between the coordinate system of laser payload 216 and the coordinate system of satellite-30 is as follows: According to the installation method of laser payload 216, it can be seen that: laser payload 216 is installed on the -Z plane of satellite-30, the Z-axis direction of laser payload is the same as the Z-axis direction of satellite-30, the initial zero position of laser payload 216 points to the X direction of satellite-30, the X direction of laser payload 216 is the same as the X direction of satellite-30, the Y direction of laser payload 216 is consistent with the Y direction of satellite-30, and the installation matrix of laser payload 216 is as follows (8):

[0081] (8).

[0082] It is worth noting that this invention fully considers the possible installation errors of the laser payload, and here the same three-axis Euler angles as those used for satellites are adopted. Let the rotation angle of the laser payload around the X-axis be θ. The rotation angle around the Y-axis is The rotation angle around the Z-axis is If the rotation sequence is X→Y→Z, then the laser load mounting matrix is ​​calculated as shown in formula (9):

[0083] (9).

[0084] In some embodiments, calculating the field-of-view angle between the Sun's position vector and the line-of-view of the local satellite's laser payload based on the acquired multiple transformation matrices includes: transforming the Sun's position vector to the laser payload coordinate system based on the acquired first, second, and third transformation matrices; calculating the pointing angle of the local satellite's laser payload turntable; and calculating the field-of-view angle between the Sun's position vector and the line-of-view of the local satellite's laser payload based on the pointing angle of the local satellite's laser payload turntable and the Sun's position vector.

[0085] For example, the pointing angle of the two-dimensional turntable of laser payload 216 is calculated based on the pointing vector of laser payload 216, and the solar vector is transformed into the coordinate system of laser payload 216. The third transformation matrix from the satellite body coordinate system to the laser payload coordinate system is calculated. The transformation relationship is as follows: formula (10).

[0086] (10)

[0087] in, This represents the current solar vector (i.e., the position vector of the sun in the inertial frame). The solar vector represents the position vector of the sun in the coordinate system of laser payload 216 (i.e., the position vector of the sun in the coordinate system of laser payload 216).

[0088] For example, based on the solar vector broadcast in the satellite broadcast data, the current position vector of the sun can be obtained as shown in formula (11):

[0089] (11),

[0090] Where a, b, and c are the calculated values ​​of the solar position vector. Based on the solar position vector (11), the solar position vector in the coordinate system of the laser payload 216 is calculated as shown in formula (12):

[0091] (12)

[0092] Based on formula (11), we get d, e, and f are the corresponding position vector values, which allows us to calculate the pointing angles α1 and β1 of the two-dimensional turntable of the laser load 216 on the y-axis and z-axis, respectively.

[0093] For example, calculate the field of view angle between the sun's position vector and the line of sight of the laser payload 216 on satellite-30, determine the current pointing angles α2 and β2 of the two-dimensional turntable of the laser payload 216, and calculate the difference between the angle between the sun's position vector and the two-dimensional turntable of the laser payload 216 at the predetermined solar outburst critical value and the actual angle between the current sun's position vector and the two-dimensional turntable of the laser payload 216: and .

[0094] In some embodiments, the predetermined solar eclipse threshold value is 3°, and the minimum avoidance range is 5°, then the solar eclipse range is (-3°, +3°), for example, when and When any of the satellites enters the (-5°, +5°) range, laser payload 216 initiates solar interference avoidance maneuvers. It should be understood that the minimum avoidance range must be greater than the interference threshold, requiring sufficient time for the satellite to determine whether it has entered the interference zone and for the solar interference avoidance operation to be performed.

[0095] During the solar outage avoidance process, the laser payload 216 simultaneously calculates the angle between the current solar vector and the field of view of the laser payload 216. When the angle between the solar vector and the field of view of the laser payload 216 reaches the boundary of (-5°, +5°) again, the laser payload 216 begins to restore communication after the solar outage ends. The solar outage viewing end (e.g., satellite 1 30) performs operations such as searching, capturing, and scanning. For example, it enables satellite 1 30 to recapture satellite 2 40 after the solar outage ends. Satellite 2 40, which is at the back-viewing end of the solar outage, can remain stationary and staring during the solar outage process.

[0096] Figure 10 A flowchart of a method 1000 for solar overpass avoidance according to an embodiment of the present invention is shown. Method 1000 may be performed by, for example... Figure 1 The satellites shown (Satellite 1-30, Satellite 2-40) can also be used. Figure 8 The method is performed at the illustrated electronic device 800. It should be understood that method 1000 may also include additional steps not shown and / or the steps shown may be omitted, and the scope of the invention is not limited in this respect.

[0097] In step 1002, the satellites (Satellite 1 30 and Satellite 2 40) determine whether they have entered the solar interference region according to the solar interference prediction method provided by the present invention.

[0098] In step 1004, if the satellite (Satellite 1 30, Satellite 2 40) determines that it has entered the solar interference region, it performs a solar interference avoidance operation, wherein the solar interference avoidance operation includes at least: adjusting the angle of the precision tracking mirror of the satellite's laser payload by an amount exceeding a predetermined threshold, so that the sunlight deflects out of the camera field of view of the satellite's laser payload.

[0099] In some embodiments, the laser payload of the onboard laser communication equipment of the satellite (Satellite 1 30, Satellite 2 40) is further configured with a light trapping device, which is disposed at a predetermined position inside the laser payload, and the inner wall of the light trapping device is provided with a light-absorbing material.

[0100] In some embodiments, performing a solar outage avoidance operation includes: adjusting the angle of the precision tracking mirror of the local satellite's laser payload by an amount exceeding a predetermined threshold, so that sunlight deflects out of the camera's field of view of the local satellite's laser payload; and powering down the camera of the local satellite's laser payload.

[0101] In some embodiments, adjusting the angle of the precision-tracking fast reflector of the local satellite's laser payload by an amount exceeding a predetermined threshold, so that sunlight deflects out of the camera's field of view of the local satellite's laser payload, includes: deflecting the precision-tracking reflector of the laser payload by a predetermined angle, so that sunlight entering the laser payload is deflected and then enters the light trapping device; and causing the sunlight entering the light trapping device to undergo multiple reflections in the light trapping device, thereby being gradually absorbed by the light-absorbing material in the light trapping device until it is completely attenuated.

[0102] Regarding power-down control, for example, the satellite master control system issues a power-off command to the optical camera. Upon receiving the power-off command, the optical camera on the satellite's laser payload switches the power-on link switch of the optical camera on the laser payload from "on" to "off", thereby physically disconnecting the power-on link of the optical camera.

[0103] In some embodiments, please refer to Figure 6 The diagram illustrates the structure of a satellite laser payload optical system 600, which includes at least an optical antenna 602, a Decoupage mirror group 604, a precision tracking mirror 606, a PBS beam-splitting isolation optical path 608, a first beam splitter 610, a second beam splitter 642, a first reflector 644, a communication optical receiver 620, a pre-reflector 634, a transmitting optical system 632, a signal transmitting end 630, and an optical camera 640. Please refer to further details. Figure 7 This further illustrates Figure 6 The optical system structure shown in the diagram has the following optical paths: the red line indicates the transmitting optical path, the yellow line indicates the tracking optical path, and the green line indicates the receiving optical path.

[0104] In some embodiments, Figure 6 and Figure 7The satellite laser payload optical system structure shown features an integrated transmit / receive antenna design with parallel light input and output for easy beam tracking and control. The transmitting optical path is connected via a single-mode fiber with a mode field core diameter of approximately 10.58 μm (actual core diameter 9 μm) and a numerical aperture (NA) of 0.14. After passing through the optical transmission system, it is emitted by a 10x optical telescope with a divergence angle of approximately 67.2 rad. The receiving optical path is also connected via a single-mode fiber with a mode field core diameter of approximately 10.58 μm (actual core diameter 9 μm), a numerical aperture (NA) of 0.14, and an aperture of 18 mm. The PBS beam splitter and isolation optical path uses a PBS for beam combining and isolation. The transmitting light is S-beam as the emission source, and the receiving light is P-beam. The transmitted and received light beams are separated by the PBS and finally combined in the main optical path. The tracking optical path uses an integrated coarse / fine tracking camera. The instantaneous field of view of a single pixel of the tracking camera detector is 12 μrad, and the effective aperture of the coarse tracking camera lens is approximately Φ21 mm. In addition, it includes an optical head state switching unit (not shown) configured to complete rotation and filter switching; a fine tracking mirror 606 configured to achieve rapid scanning of the fine tracking field of view; and a drive advance mirror to achieve advance pointing of the emitted light path.

[0105] For information on adjusting the angle of the precision-tracking mirror of the satellite's laser payload by an amount exceeding a predetermined threshold, so that sunlight is deflected out of the camera's field of view, please refer to [reference needed]. Figure 6 and Figure 7 The optical camera 640 includes, for example, a camera group consisting of a main camera and a backup camera. Please refer to the yellow line to illustrate the tracking optical path. In the solar interference area, sunlight can pass through this optical path and through the various optical components of the optical system 600, and enter the optical camera 640, thereby causing damage to the optical camera with the laser payload.

[0106] Traditional solar interference avoidance operations, such as adjusting the azimuth and elevation angles of the laser payload's two-dimensional turntable to prevent sunlight from directly entering the aperture of the optical antenna, require the two-dimensional turntable to be adjusted back to its original position after leaving the solar interference area, and then the search for the target star must be restarted to recapture it. This solar interference recovery operation method is prone to increasing the difficulty of search and capture due to inaccurate positioning of the original position of the satellite itself.

[0107] In the above scheme, there is no need to adjust the pitch or azimuth of the two-dimensional turntable of the laser payload. Therefore, the angle of the azimuth motor can remain unchanged. Instead, the angle of the precision tracking mirror in the optical system is adjusted to deflect sunlight directly out of the camera's field of view. At the same time, the optical camera is powered down, further reducing the risk of camera damage. After the solar interference ends, the precision tracking mirror deflector is returned to its original position, the optical camera is powered on again, and the laser payload can directly search in its original position without adjusting its angle or pitch. The difficulty of recapturing the target star is reduced, and efficiency is guaranteed. In addition, even if some scattered sunlight still enters the photosensitive surface of the optical camera, because the optical lenses are usually coated with narrow-band filters, only some stray sunlight enters and will not cause damage to the optical camera.

[0108] It should be understood that in the traditional scheme described above, due to the different real-time states of the laser payload, its line-of-sight angle varies during operation. Therefore, for its two-dimensional turntable, the initial position during each solar outage avoidance operation may be different. However, within the optical system of the laser payload, ensuring the normal operation of each optical path and maintaining the constant angle of the fine-tuning mirror means that the position (angle) of the fine-tuning mirror is the same before each avoidance operation, and the original position is the same. For example, each avoidance operation deflects a predetermined deflection angle γ, and each recovery angle only requires deflecting back to γ. Compared to the traditional scheme, this scheme is more stable and does not require complex calculations. The deflection angle for each avoidance and recovery operation can be the same value. Therefore, the solar outage avoidance and recovery operations performed by the above scheme will not introduce errors caused by different original positions, and have more accurate and efficient operability.

[0109] In some embodiments, the precision heel reflector 606 includes a main precision heel reflector and a backup precision heel reflector, the main precision heel reflector and the backup precision heel reflector forming a precision heel reflector group.

[0110] For example, the ability to allow only a portion of sunlight to enter the camera's photosensitive surface (optical lenses are coated with narrow-band filters, allowing only a portion of stray sunlight to enter without damaging the optical camera), and to adjust the angle of the local satellite's laser payload by an amount exceeding a predetermined threshold, includes: adjusting the precision tracking mirror by a predetermined deflection angle, such as adjusting at least one of the primary precision tracking mirror and the backup precision tracking mirror.

[0111] When a solar interference occurs, direct sunlight can damage the optical camera. In the above method, by adjusting the angle of the fine-tuning reflector, the incident sunlight cannot directly hit the photosensitive surface of the optical camera when a solar interference occurs, thus avoiding camera damage.

[0112] Regarding adjusting the angle of the precision reflector of the local satellite's laser payload by an amount exceeding a predetermined threshold, for example, adjusting the angle of the precision reflector by a predetermined deflection angle. For example, the predetermined deflection angle is... .

[0113] Please refer to Figure 9 This illustrates the deflection of the fine-tuning mirror. For example, please refer to... Figure 9 As shown, the red dashed line 91 indicates the reflecting plane of the primary precision mirror 6061 in the precision mirror 606 under normal operating conditions. The blue dashed line 92 indicates the deflection angle, showing the position of the primary mirror and the reflector 6061 at the predetermined deflection angle. The reflecting plane after the mirror. For example, the precision mirror is deflected from its working position at a predetermined angle. After deflection, it reaches the predetermined deflection position, causing the sunlight to be deflected by 2. The angle. For example, the predetermined deflection angle. Determined based on the actual layout of the optical system, 2 This angle is typically greater than the operating angle range of the precision reflector (e.g., the operating angle range of a precision reflector is typically ±5 mrad). This deflects sunlight from the optical system of the satellite's laser payload, preventing it from reaching the optical camera and thus protecting the camera.

[0114] In some embodiments, the optical system 600 further includes a light trapping device 900. The light trapping device is configured with a light-absorbing material on its inner wall, such that light entering the light trapping device undergoes multiple diffuse reflections within the light trapping device, preventing the incoming light from escaping the light trapping device until the incoming light gradually attenuates.

[0115] In some embodiments, adjusting the angle of the precision reflector of the local satellite's laser payload by an amount exceeding a predetermined threshold, so that sunlight deflects out of the camera's field of view of the local satellite's laser payload, includes: deflecting the precision reflector of the laser payload by a predetermined angle, so that sunlight entering the laser payload is deflected and then enters a light trapping device, the light trapping device being disposed at a predetermined position inside the laser payload; and causing the sunlight entering the light trapping device to undergo multiple reflections in the light trapping device and be gradually absorbed by the light-absorbing material in the light trapping device until the sunlight is completely attenuated.

[0116] Regarding the location of the optical trap device, it can be positioned, for example, at a relatively forward position within the optical path of the laser load, such as a predetermined position between the precision reflector and the components following the precision reflector. Please refer to... Figure 9The light trap device 900 is positioned between the precision reflector 606 and the PBS beam-splitting isolation optical path. This allows sunlight entering the laser payload to be reflected by the precision reflector and immediately enter the light trap device, thus avoiding passing through other components and preventing damage to them.

[0117] In the above scheme, the sunlight entering the optical system of the satellite's laser payload is deflected and directed to a light trap device added within the optical system. This causes the sunlight to be captured by the light trap device and unable to escape. Furthermore, the sunlight gradually attenuates after entering the light trap device, thereby preventing other equipment in the optical system from being exposed to strong sunlight during solar interference. This ensures that the satellite's laser payload's optical system is not damaged by sunlight in the solar interference region.

[0118] In the above scheme, since the solar outage avoidance operation does not change the line-of-sight angle of the laser payload, the solar outage avoidance is achieved by adjusting the angle of the reflector in the optical system of the satellite laser payload. Therefore, after the satellite leaves the solar outage area, it only needs to maintain the current line-of-sight angle of the laser payload in its original position to search and scan, and it can recapture the target star. Compared with the traditional scheme of adjusting the turntable angle of the laser payload, the solar outage avoidance operation is more precise and efficient, and the communication recovery efficiency after the solar outage is also higher.

[0119] Figure 8 A schematic step diagram of an example electronic device 800 that can be used to implement embodiments of the contents of this specification is shown. For example, as Figure 1 The satellites shown (Satellite 1-30, Satellite 2-40) can be electronically generated. Figure 8 A schematic step diagram of an example electronic device 800 that can be used to implement embodiments of the contents of this specification is shown. For example, as Figure 1 The satellites shown (Satellite 1 30, Satellite 2 40) can be implemented by electronic device 800. As shown, electronic device 800 includes a central processing unit (CPU) 801, which can perform various appropriate actions and processes according to computer program instructions stored in read-only memory (ROM) 802 or loaded from storage unit 808 into random access memory (RAM) 803. The random access memory 803 can also store various programs and data required for the operation of electronic device 800. The CPU 801, ROM 802, and RAM 803 are interconnected via bus 804. Input / output (I / O) interface 805 is also connected to bus 804.

[0120] Multiple components in electronic device 800 are connected to input / output interface 805, including: input unit 806, such as keyboard, mouse, microphone, etc.; output unit 807, such as various types of monitors, speakers, etc.; storage unit 808, such as disk, optical disk, etc.; and communication unit 809, such as network card, modem, wireless transceiver, etc. Communication unit 809 allows electronic device 800 to exchange information / data with other devices through computer networks such as the Internet and / or various telecommunications networks.

[0121] The various processes and procedures described above, such as methods 300, 400, and 1000, can be executed by the central processing unit 801. For example, in some embodiments, methods 300, 400, and 1000 can be implemented as computer software programs tangibly contained in a machine-readable medium, such as storage unit 808. In some embodiments, part or all of the computer program can be loaded and / or installed on the electronic device 800 via read-only memory 802 and / or communication unit 809. When the computer program is loaded into random access memory 803 and executed by the central processing unit 801, one or more actions of methods 300, 400, and 1000 described above can be performed.

[0122] This invention relates to methods, apparatus, systems, electronic devices, computer-readable storage media, and / or computer program products. The computer program product may include computer-readable program instructions for performing various aspects of the invention.

[0123] Computer-readable storage media can be tangible devices capable of holding and storing instructions for use by an instruction execution device. Computer-readable storage media can be, for example—but not limited to—electrical storage devices, magnetic storage devices, optical storage devices, electromagnetic storage devices, semiconductor storage devices, or any suitable combination of the foregoing. More specific examples (a non-exhaustive list) of computer-readable storage media include: portable computer disks, hard disks, random access memory (RAM), read-only memory (ROM), erasable programmable read-only memory (EPROM or flash memory), static random access memory (SRAM), portable compact disc read-only memory (CD-ROM), digital multifunction disc (DVD), memory sticks, floppy disks, mechanical encoding devices, such as punch cards or recessed protrusions storing instructions thereon, and any suitable combination of the foregoing. The computer-readable storage media used herein are not to be construed as transient signals themselves, such as radio waves or other freely propagating electromagnetic waves, electromagnetic waves propagating through waveguides or other transmission media (e.g., light pulses through fiber optic cables), or electrical signals transmitted through wires.

[0124] The computer-readable program instructions described herein can be downloaded from computer-readable storage media to various computing / processing devices, or downloaded via a network, such as the Internet, local area network, wide area network, and / or wireless network, to an external computer or external storage device. The network may include copper transmission cables, fiber optic transmission, wireless transmission, routers, firewalls, switches, gateway computers, and / or edge computing devices. A network adapter card or network interface in each computing / processing device receives the computer-readable program instructions from the network and forwards them to computer-readable storage media within the respective computing / processing device.

[0125] The computer program instructions used to perform the operations of this invention may be assembly instructions, instruction set architecture (ISA) instructions, machine instructions, machine-dependent instructions, microcode, firmware instructions, state setting data, or source code or object code written in any combination of one or more programming languages, including object-oriented programming languages ​​such as Smalltalk, C++, etc., and conventional procedural programming languages ​​such as the "C" language or similar programming languages. The computer-readable program instructions may be executed entirely on the user's computer, partially on the user's computer, as a standalone software package, partially on the user's computer and partially on a remote computer, or entirely on a remote computer or server. In cases involving a remote computer, the remote computer may be connected to the user's computer via any type of network—including a local area network (LAN) or a wide area network (WAN)—or may be connected to an external computer (e.g., via the Internet using an Internet service provider). In some embodiments, electronic circuitry, such as programmable logic circuitry, field-programmable gate arrays (FPGAs), or programmable logic arrays (PLAs), is personalized by utilizing state information from the computer-readable program instructions. This electronic circuitry can execute the computer-readable program instructions to implement various aspects of the invention.

[0126] Various aspects of the present invention are described herein with reference to flowchart illustrations and / or step diagrams of methods, apparatus (systems), and computer program products according to embodiments of the invention. It should be understood that each step in the flowchart illustrations and / or step diagrams, as well as combinations of steps in the flowchart illustrations and / or step diagrams, can be implemented by computer-readable program instructions.

[0127] These computer-readable program instructions can be provided to a processing unit of a general-purpose computer, a special-purpose computer, or other programmable data processing apparatus to produce a machine such that, when executed by the processing unit of the computer or other programmable data processing apparatus, they create means for implementing the functions / actions specified in one or more steps of the flowchart and / or diagram of steps. These computer-readable program instructions can also be stored in a computer-readable storage medium that causes a computer, programmable data processing apparatus, and / or other device to operate in a particular manner. Thus, the computer-readable medium storing the instructions comprises an article of manufacture that includes instructions for implementing aspects of the functions / actions specified in one or more steps of the flowchart and / or diagram of steps.

[0128] Computer-readable program instructions may also be loaded onto a computer, other programmable data processing apparatus, or other device to cause a series of operational steps to be performed on the computer, other programmable data processing apparatus, or other device to produce a computer-implemented process, thereby causing the instructions executed on the computer, other programmable data processing apparatus, or other device to perform the functions / actions specified in one or more steps of a flowchart and / or a diagram of steps.

[0129] The flowcharts and step diagrams in the accompanying drawings illustrate the architecture, functionality, and operation of possible implementations of systems, methods, and computer program products according to various embodiments of the present invention. In this regard, each step in the flowchart or step diagram may represent a module, segment, or part of an instruction, which contains one or more executable instructions for implementing the specified logical function. In some alternative implementations, the functions indicated in the step may occur in a different order than those indicated in the drawings. For example, two consecutive step steps may actually be executed substantially in parallel, and they may sometimes be executed in reverse order, depending on the functions involved. It should also be noted that each step in the step diagram and / or flowchart, and combinations of steps in the step diagram and / or flowchart, can be implemented using a dedicated hardware-based system that performs the specified function or action, or using a combination of dedicated hardware and computer instructions.

[0130] The various embodiments of the present invention have been described above. These descriptions are exemplary and not exhaustive, nor are they limited to the disclosed embodiments. Many modifications and variations will be apparent to those skilled in the art without departing from the scope and spirit of the described embodiments. The terminology used herein is chosen to best explain the principles, practical application, or technical improvements to the embodiments in the market, or to enable others skilled in the art to understand the embodiments disclosed herein.

Claims

1. A method for avoiding solar overpasses, characterized in that, The method is applied to spaceborne laser communication equipment, and the method includes: The position vector of the sun is obtained based on the satellite broadcast of this satellite; Based on the satellite's position, velocity, attitude, and laser payload installation location information, multiple transformation matrices are obtained. These multiple transformation matrices include at least: a first transformation matrix from the inertial frame to the center-of-mass orbital coordinate system, a second transformation matrix from the center-of-mass orbital coordinate system to the satellite's coordinate system, and a third transformation matrix from the satellite's coordinate system to the laser payload's coordinate system. The third transformation matrix is ​​calculated based at least on the installation orientation and installation error angle of the laser payload along the three axes of the satellite's body. Based on the acquired multiple transformation matrices, the field of view angle between the Sun's position vector and the local star's laser payload line of view is calculated; Based on the field of view angle, determine whether the local star has entered the solar interference region; and Determine if the local satellite has entered the solar interference zone; In response to determining that the satellite has entered the solar interference region, a solar interference avoidance operation is performed, wherein the solar interference avoidance operation includes at least: adjusting the angle of the precision mirror of the satellite's laser payload by an amount exceeding a predetermined threshold, so that sunlight deflects out of the camera field of view of the satellite's laser payload.

2. The method according to claim 1, characterized in that, Based on the satellite's position, velocity, attitude, and laser payload installation location information, multiple transformation matrices are obtained, including: Based on the satellite's position and velocity information, calculate the first transformation matrix from the inertial frame to the center-of-mass orbital coordinate system; Based on the satellite's attitude information, calculate the second transformation matrix from the center-of-mass orbital coordinate system to the satellite's body coordinate system; and Based on the installation orientation and installation error angles of the laser payload on the three axes of the satellite body, the third transformation matrix from the satellite body coordinate system to the laser payload coordinate system is calculated in the rotation sequence of the x-axis, y-axis, and z-axis. The installation error angles include the x-axis error angle, y-axis error angle, and z-axis error angle.

3. The method according to claim 2, characterized in that, Based on the satellite's attitude information, the calculation of the second transformation matrix from the center-of-mass orbital coordinate system to the satellite's body coordinate system includes: Based on the satellite's attitude information, the attitude angles of the satellite body are obtained, including roll angle, pitch angle, and yaw angle; and Based on the attitude information of the satellite body, the second transformation matrix from the center-of-mass orbit coordinate system to the satellite body coordinate system is calculated in the order of attitude adjustment of roll angle, pitch angle and yaw angle.

4. The method according to claim 2, characterized in that, Based on the acquired transformation matrices, the field of view angle between the Sun's position vector and the local star's laser payload line of sight is calculated, including: Based on the obtained first, second, and third transformation matrices, the position vector of the sun is transformed into the laser payload coordinate system; Calculate the pointing angle of the turntable for the local satellite's laser payload; and Based on the pointing angle of the turntable of the local satellite's laser payload and the position vector of the sun, calculate the field of view angle between the position vector of the sun and the line of sight of the local satellite's laser payload.

5. The method according to claim 4, characterized in that, Determining whether the local satellite has entered the solar interference region based on the field of view angle includes: Calculate the first field-of-view angle between the laser payload's line of sight and the Sun's position vector along the y-axis, and the second field-of-view angle between the laser payload's line of sight and the Sun's position vector along the z-axis; and In response to determining that the first field of view angle and / or the second field of view angle are less than or equal to a predetermined angle, it is determined that the local satellite has entered the solar interference region.

6. The method according to any one of claims 1-5, characterized in that, The laser payload included in the spaceborne laser communication equipment is also equipped with a light trapping device, which is located at a predetermined position inside the laser payload, and the inner wall of the light trapping device is provided with a light-absorbing material.

7. The method according to claim 6, characterized in that, Adjusting the angle of the precision mirror of the satellite's laser payload by an amount exceeding a predetermined threshold, so that sunlight deflects out of the camera field of view of the satellite's laser payload, includes: The precision-guided mirror deflects the laser load at a predetermined angle, causing sunlight entering the laser load to be deflected and then incident on the light trapping device; and The sunlight entering the light trap device is reflected multiple times within the device, and is gradually absorbed by the light-absorbing material until it is completely attenuated.

8. A computing device, characterized in that, include: At least one processing unit; At least one memory coupled to the at least one processing unit and storing instructions for execution by the at least one processing unit, the instructions, when executed by the at least one processing unit, causing the device to perform the steps of the method according to any one of claims 1 to 7.

9. A computer-readable storage medium, characterized in that, It stores a computer program that, when executed by a machine, implements the method according to any one of claims 1 to 7.

10. A computer program product, characterized in that, Includes a computer program, which, when executed by a machine, performs the method according to any one of claims 1-7.

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