Satellite attitude management method, satellite attitude management device and computer readable medium

Abstract

The invention relates to a satellite attitude management method, a satellite attitude management device and a computer readable medium. The satellite attitude management method provided by the invention comprises the following steps: acquiring a target system sun vector, wherein the target system sun vector is the system sun vector when the current value of a sailboard is maximum; and adjusting the current satellite attitude according to the current constraint, the sensor constraint and the sun vector of the target body system to obtain the target satellite attitude: establishing an intermediate coordinate system according to the satellite Z-axis vector of the body system and the sun vector of the target body system; obtaining a first conversion matrix, wherein the first conversion matrix is a conversion matrix from the intermediate coordinate system to the target inertial system; obtaining a second conversion matrix, wherein the second conversion matrix is a conversion matrix from the system to the intermediate coordinate system; obtaining an attitude matrix according to the first conversion matrix and the second conversion matrix, wherein the attitude matrix is a projection of the system in the target inertial system; and obtaining a target satellite attitude according to the attitude matrix.

Description

Technical Field

[0001] This application mainly relates to the field of satellite attitude control technology, and specifically relates to a satellite attitude management method, a satellite attitude management device, and a computer-readable medium. Background Art

[0002] Satellites operate in a complex space environment for a long time and face many challenges. Among them, space debris collision is one of the important problems. Even the impact of tiny debris may cause deformation of the solar panel structure or short circuit of the circuit, making it impossible for the ground to accurately judge the solar panel pointing based on the real-time data sent back by the satellite, thus posing a huge risk to the stable operation of the satellite. Currently, conventional satellite attitude management methods are to control the closed-loop attitude to the target satellite attitude according to the mission requirements according to the preset guidance law, and at the same time rotate the solar panel to keep the sunlight irradiating the solar array at the maximum incident angle, so as to obtain more energy. However, this method can perform attitude control when the solar panel points in the conventional working mode, but it cannot guarantee sufficient energy acquisition for the case where the solar panel pointing is unknown. Summary of the Invention

[0003] In view of the above technical problems, this application provides a satellite attitude management method for solving the technical problem that the satellite cannot obtain sufficient energy supply when the solar panel pointing is unknown.

[0004] The technical solution adopted by this application to solve the above technical problems is a satellite attitude management method, including: obtaining a target body-fixed solar vector, where the target body-fixed solar vector is the body-fixed solar vector when the sail current value is the largest; and adjusting the current satellite attitude according to current constraints, sensor constraints, and the target body-fixed solar vector to obtain a target satellite attitude, including: establishing an intermediate coordinate system according to the satellite Z-axis vector of the body-fixed system and the target body-fixed solar vector; obtaining a first transformation matrix, where the first transformation matrix is the transformation matrix from the intermediate coordinate system to the target inertial system; obtaining a second transformation matrix, where the second transformation matrix is the transformation matrix from the body-fixed system to the intermediate coordinate system; obtaining an attitude matrix according to the first transformation matrix and the second transformation matrix, where the attitude matrix is the projection of the body-fixed system in the target inertial system; and obtaining the target satellite attitude according to the attitude matrix.

[0005] In one embodiment of the present application, after obtaining the target heliospheric solar vector, it further includes: determining whether the satellite meets the light pressure unloading condition, including: in response to the total angular momentum of the entire satellite in the current satellite attitude being greater than the total angular momentum of the target satellite in the target attitude, or, in response to the first included angle between the total angular momentum of the target satellite in the target attitude and the angular momentum increment belonging to a first range, or in response to the second included angle between the target heliospheric solar vector and the total angular momentum of the entire satellite in the current satellite attitude belonging to a second range, determining that the satellite meets the light pressure unloading condition.

[0006] In one embodiment of the present application, the current constraint is configured as: the current of the solar panel in the target satellite attitude is greater than the minimum current required for the satellite to operate.

[0007] In one embodiment of the present application, the sensor constraint is configured as: the third included angle between the optical axis of the heliospheric sensor and the heliospheric celestial body vector is greater than the celestial body suppression angle.

[0008] In one embodiment of the present application, adjusting the current satellite attitude according to the current constraint, the sensor constraint, and the target heliospheric solar vector to obtain the target satellite attitude further includes: adjusting the current satellite attitude according to the angular momentum constraint, and the angular momentum constraint is further configured as: the moduli of the X, Y, and Z axis components of the total angular momentum of the entire satellite in the target satellite attitude are respectively less than a first threshold.

[0009] In one embodiment of the present application, adjusting the current satellite attitude according to the current constraint, the sensor constraint, and the target heliospheric solar vector to obtain the target satellite attitude further includes: the TT&C constraint, and the TT&C constraint is configured as: the fourth included angle between the heliospheric geocentric vector and the heliospheric satellite Z-axis vector belongs to a third range.

[0010] In one embodiment of the present application, the first transformation matrix is obtained by using the following formula: , wherein, and are both the first transformation matrix, is the rotation angle.

[0011] In one embodiment of the present application, the second transformation matrix is obtained by using the following formula: , wherein, is the second transformation matrix, is the x-axis component of the target heliospheric solar vector, is the y-axis component of the target heliospheric solar vector, is the z-axis component of the target heliospheric solar vector.

[0012] In one embodiment of the present application, the following formula is used to calculate the rotation angle: , where, is the rotation angle, is the y-axis component of the geocentric vector in the intermediate coordinate system, is the z-axis component of the geocentric vector in the intermediate coordinate system.

[0013] In one embodiment of the present application, in response to the fifth angle between the geocentric vector of the target inertial system and the Z-axis vector of the intermediate coordinate system belonging to the fourth range, the following formula is used to recalculate the rotation angle: .

[0014] In one embodiment of the present application, obtaining the attitude matrix according to the first transformation matrix and the second transformation matrix includes: obtaining the attitude matrix by using the following formula: , where, and are both the attitude matrix, is the second transformation matrix, is the first transformation matrix.

[0015] In one embodiment of the present application, the following formula is used to obtain the target satellite attitude according to the attitude matrix: , , where, is the target satellite attitude, is the rotation angle required for the satellite to adjust from the current attitude to the target satellite attitude around the target body-fixed solar vector, is the identity matrix, e is the unit vector of the target body-fixed solar vector, and E is the matrix composed of the unit vector e of the target body-fixed solar vector.

[0016] In one embodiment of the present application, the following formula is used to calculate the E: , where, is the x-axis component of the unit vector of the target body-fixed solar vector, is the y-axis component of the unit vector of the target body-fixed solar vector, is the z-axis component of the unit vector of the target body-fixed solar vector.

[0017] In one embodiment of the present application, the following formula is used to calculate the angular momentum of the target satellite attitude: , where H1 represents the angular momentum of the target satellite attitude, is the target satellite attitude, is the attitude matrix, represents the transpose matrix of the attitude matrix, and H0 is the angular momentum of the current satellite attitude.

[0018] In an embodiment of the present application, the obtaining of the target body-fixed solar vector includes: using a neural network model to obtain the target body-fixed solar vector, where the input data of the neural network model includes any of the body-fixed solar vectors, and the output data includes the sail current value.

[0019] In an embodiment of the present application, after obtaining the target satellite attitude, it further includes: taking the target satellite attitude as the current satellite attitude, and optimizing the current satellite attitude according to a constraint function to obtain an optimized target satellite attitude; where the constraint function includes a first constraint function, a second constraint function, and a third constraint function.

[0020] In an embodiment of the present application, the taking the target satellite attitude as the current satellite attitude and optimizing the current satellite attitude according to a constraint function to obtain an optimized target satellite attitude includes: obtaining the target body-fixed solar vector; determining whether the satellite meets the light pressure unloading condition; when the satellite meets the light pressure unloading condition, sequentially determining whether there is a feasible solution for the first constraint function, the second constraint function, and the third constraint function until the first solution is obtained, where the first constraint function includes current constraint, sensor constraint, TT&C constraint, and angular momentum constraint; the second constraint function includes the current constraint, the sensor constraint, and the TT&C constraint; the third constraint function includes the current constraint and the sensor constraint; and optimizing the current satellite attitude using the constraint function corresponding to the first solution to obtain an optimized target satellite attitude.

[0021] In an embodiment of the present application, in response to the non-existence of a feasible solution for the first constraint function or the second constraint function, it is determined whether the interval between the current time and the end time of the most recent attitude adjustment is less than a first interval; in response to the interval between the current time and the end time of the most recent attitude adjustment being less than the first interval, the interval between the start time of the next attitude adjustment and the end time of the most recent attitude adjustment is adjusted to a second interval, and it is sequentially determined whether the remaining constraint functions have feasible solutions, where the second interval is greater than the first interval; and in response to the interval between the current time and the end time of the most recent attitude adjustment being greater than the first interval, the interval between the start time of the next attitude adjustment and the end time of the most recent attitude adjustment is adjusted to a third interval, and it is sequentially determined whether the remaining constraint functions have feasible solutions, where the third interval is less than the first interval.

[0022] In an embodiment of the present application, optimizing the current satellite attitude according to the constraint function includes: using a sequential quadratic programming algorithm to optimize the current satellite attitude to obtain an optimized target satellite attitude.

[0023] The present application also provides a satellite attitude management device, including: a memory for storing instructions executable by a processor; and a processor for executing the instructions to implement the above method.

[0024] The present application also provides a computer-readable medium storing computer program code, where the computer program code implements the above method when executed by a processor.

[0025] The present application obtains the target body-fixed solar vector and adjusts the current satellite attitude according to the current constraint, the sensor constraint, and the target body-fixed solar vector to obtain the target satellite attitude, enabling the satellite to obtain sufficient energy supply when the solar array pointing is unknown. BRIEF DESCRIPTION OF THE DRAWINGS

[0026] To make the above objects, features, and advantages of the present application more obvious and understandable, the following detailed description of the specific embodiments of the present application is provided in conjunction with the accompanying drawings, where: Figure 1 is a flowchart of a satellite attitude management method according to an embodiment of the present application; Figure 2 is a schematic diagram of a neural network model of a satellite attitude management method according to an embodiment of the present application; Figure 3 is a flowchart of a satellite attitude management method according to another embodiment of the present application; Figure 4 is a schematic diagram of a second included angle of a satellite attitude management method according to an embodiment of the present application; Figure 5It is the distribution diagram of the included angle between the target heliospheric vector and the angular momentum increment of the satellite attitude management method according to an embodiment of the present application; Figure 6 It is the schematic diagram of the direction of the satellite Z-axis vector of the satellite attitude management method according to an embodiment of the present application; Figure 7 It shows the flowchart of the satellite attitude management method according to another embodiment of the present application; Figure 8 It shows the flowchart of the satellite attitude management method according to an embodiment of the present application; Figure 9 It shows the schematic diagram of the change of the input current of the satellite solar panel and the total load current of the satellite of the satellite attitude management method according to an embodiment of the present application; Figure 10 It shows the angular momentum change curve of the satellite attitude management method according to an embodiment of the present application; Figure 11 It is the system block diagram of the satellite attitude management device according to an embodiment of the present application. Detailed implementation manners

[0027] To make the above objects, features, and advantages of the present application more obvious and understandable, the following detailed description of the specific implementation manners of the present application is provided in conjunction with the accompanying drawings.

[0028] In the following description, many specific details are set forth in order to fully understand the present application. However, the present application may be implemented in other ways different from those described herein. Therefore, the present application is not limited by the specific embodiments disclosed below.

[0029] As shown in the present application, unless the context clearly indicates an exception, words such as "a", "an", "one", and / or "the" are not specifically singular and may also include the plural. Generally speaking, the terms "comprising" and "including" only indicate the inclusion of the clearly identified steps and elements, and these steps and elements do not constitute an exclusive list. The method or device may also include other steps or elements.

[0030] The present application provides a satellite attitude management method. The body-fixed solar vector corresponding to the maximum sail current value is used as the target body-fixed solar vector, and the current satellite attitude is adjusted according to the constraint conditions and the target body-fixed solar vector to obtain the target satellite attitude, ensuring that the satellite can obtain sufficient energy when the pointing direction of the solar panel is unknown. The satellite attitude management method of the present application also determines whether the satellite meets the light pressure unloading condition. When the light pressure unloading condition is met, the satellite unloads the total angular momentum under natural light pressure, reducing the number of jet unloads of the on-orbit satellite, saving satellite propulsion fuel, and extending the service life of the satellite. In addition, the satellite attitude management method of the present application realizes that the satellite can obtain sufficient energy for a long time and meets the long-term operation of the satellite by obtaining the optimized target satellite attitude.

[0031] Next, the satellite attitude management method of the present application will be described in detail with reference to the accompanying drawings.

[0032] Figure 1 It is a flowchart of a satellite attitude management method 100 according to an embodiment of the present application. As Figure 1 shown, the present application provides a satellite attitude management method 100, including: Step S10: Obtain a target body-fixed solar vector, where the target body-fixed solar vector is the body-fixed solar vector when the sail current value is the largest; and Step S20: Adjust the current satellite attitude according to the current constraint, the sensor constraint, and the target body-fixed solar vector to obtain the target satellite attitude, including: establishing an intermediate coordinate system according to the satellite Z-axis vector of the body-fixed system and the target body-fixed solar vector; obtaining a first transformation matrix, where the first transformation matrix is the transformation matrix from the intermediate coordinate system to the target inertial system; obtaining a second transformation matrix, where the second transformation matrix is the transformation matrix from the body-fixed system to the intermediate coordinate system; obtaining an attitude matrix according to the first transformation matrix and the second transformation matrix, where the attitude matrix is the projection of the body-fixed system in the target inertial system; and obtaining the target satellite attitude according to the attitude matrix.

[0033] In the satellite attitude management method of the present application, the body-fixed system refers to the satellite body coordinate system. By obtaining the target body-fixed solar vector and adjusting the current satellite attitude according to the current constraint, the sensor constraint, and the target body-fixed solar vector to obtain the target satellite attitude, the satellite can obtain sufficient energy supply when the pointing direction of the solar panel is unknown.

[0034] In some embodiments, in step S10, obtaining the target body-fixed solar vector includes: using a neural network model to obtain the target body-fixed solar vector, where the input data of the neural network model includes any body-fixed solar vector, and the output data includes the sail current value.

[0035] Figure 2The figure shows a schematic diagram of the neural network model 11 of the satellite attitude management method according to an embodiment of the present application. As Figure 2 shown, the neural network model 11 includes an input layer 111, a hidden layer 112, and an output layer 113.

[0036] In some embodiments, the present application optimizes and iterates the neural network model 11 by taking multiple body-fixed solar vectors as input data and inputting them into the input layer 111 of the neural network model 11, so that the output layer 113 outputs the sail current values corresponding to the multiple body-fixed solar vectors, and completes the modeling. Then, according to the optimized neural network model 11, the sail current value output when any body-fixed solar vector is used as input data is estimated, and the body-fixed solar vector corresponding to the maximum sail current value is used as the target body-fixed solar vector.

[0037] Figure 3 shows Figure 2 the distribution diagram of the sail current values in the case of any body-fixed solar vector input in the shown embodiment. Among them, the abscissa is the azimuth angle of the body-fixed solar vector, with the unit of degree (deg); the ordinate is the elevation angle of the body-fixed solar vector, with the unit of degree (deg); the legend on the right shows that different colors represent the corresponding sail current values, with the unit of ampere (A). As Figure 3 shown, when the azimuth angle range of the body-fixed solar vector belongs to [200, 220] and the elevation angle range belongs to [-60, -40], the corresponding sail current values are relatively large. To ensure the satellite energy supply, the body-fixed solar vector with the azimuth angle range of [200, 220] and the elevation angle range of [-60, -40] can be selected as the target body-fixed solar vector.

[0038] It can be understood that due to the differences in the satellite itself and the environment, the distribution diagrams of the sail current values in the case of any body-fixed solar vector input are different. In some embodiments, according to the actual environmental conditions, the body-fixed solar vector within the appropriate azimuth angle and elevation angle range can be selected as the target body-fixed solar vector by means of the neural network model.

[0039] In some embodiments, after obtaining the target body-fixed solar vector in step S10, it further includes: judging whether the satellite meets the light pressure unloading condition, including: in response to the whole-satellite angular momentum in the current satellite attitude being greater than the whole-satellite angular momentum in the target satellite attitude, or, in response to the first included angle between the whole-satellite angular momentum in the target satellite attitude and the angular momentum increment direction belonging to the first range, or in response to the second included angle between the target body-fixed solar vector and the whole-satellite angular momentum in the current satellite attitude belonging to the second range, judging that the satellite meets the light pressure unloading condition.

[0040] Among them, the total angular momentum of the entire satellite refers to the total angular momentum of the entire satellite system (including the satellite body and components such as momentum wheels that may be carried). When the satellite meets the light pressure unloading condition, the satellite can add angular momentum constraints in step S20 and use natural light pressure to unload the total angular momentum of the entire satellite during attitude adjustment, so as to reduce the number of jet unloads of the on-orbit satellite, save propulsion fuel, and extend the service life of the satellite.

[0041] It can be understood that from the result of unloading the total angular momentum of the entire satellite by natural light pressure: the total angular momentum of the target satellite attitude after unloading is less than the total angular momentum of the current satellite attitude before unloading. That is, when the total angular momentum of the current satellite attitude is greater than the total angular momentum of the target satellite attitude, the satellite meets the light pressure unloading condition.

[0042] In some embodiments, the fact that the total angular momentum of the current satellite attitude is greater than the total angular momentum of the target satellite attitude means that the magnitude of the total angular momentum of the current satellite attitude is greater than the magnitude of the total angular momentum of the target satellite attitude.

[0043] In some embodiments, the light pressure unloading condition is represented by formula (1): (1) Among them, H1 is the total angular momentum of the target satellite attitude, H0 is the total angular momentum of the current satellite attitude, is the angular momentum increment.

[0044] In some embodiments, represents the angular momentum increment in one day.

[0045] In some embodiments, when the first included angle θ1 between the total angular momentum of the target satellite attitude and the direction of the angular momentum increment is an obtuse angle, that is, the first range is: when, the light pressure unloading condition is met.

[0046] In some embodiments, when the second range of the second included angle θ2 between the target body system solar vector and the total angular momentum of the current satellite attitude is [45, 135], the relationship represented in formula (1) is met, that is, the satellite meets the light pressure unloading condition, and the total angular momentum of the current satellite attitude is greater than the total angular momentum of the target satellite attitude.

[0047] In some embodiments, according to the first range of the second included angle θ2, the range of the target body system solar vector, and the total angular momentum increase telemetry, a distribution diagram of the included angle θ' between the target body system solar vector and the angular momentum increment can be drawn. Figure 4 Shows a schematic diagram of the second included angle of the satellite attitude management method according to an embodiment of the present application. As Figure 4 shown, θ2 is the second included angle, H0 is the total angular momentum of the current satellite attitude, is the angular momentum increment, is the target satellite system solar vector, and θ' is the angle between the target satellite system solar vector and the angular momentum increment. Figure 5 The figure shows the distribution diagram of the angle θ' between the target satellite system solar vector and the angular momentum increment in an embodiment of the present application. Among them, the abscissa is the azimuth angle of the satellite system solar vector, with the unit of degree (deg); the ordinate is the elevation angle of the satellite system solar vector, with the unit of degree (deg); the legend on the right shows that different colors represent the corresponding angle θ' between the target satellite system solar vector and the angular momentum increment, with the unit of degree (deg).

[0048] As Figure 5 shown, when the range of the second angle θ2 is [45, 135] and the selected azimuth angle range is [200, 220] and the elevation angle range is [-60, -40] corresponding to the target satellite system solar vector range, the range of the angle θ' between the target satellite system solar vector and the angular momentum increment is [74, 92]. At this time, the following is satisfied: the first angle θ1 between the overall satellite angular momentum in the target satellite attitude and the direction of the angular momentum increment is an obtuse angle, and the overall satellite angular momentum in the current satellite attitude is greater than the overall satellite angular momentum in the target satellite attitude.

[0049] In some embodiments, after determining whether the satellite meets the light pressure unloading condition, step S20 is executed.

[0050] In some embodiments, the current constraint in step S20 is configured as: the current of the solar panel in the target satellite attitude is greater than the minimum current required for the satellite to operate.

[0051] It should be understood that only when the current of the solar panel in the target satellite attitude is greater than the minimum current required for the satellite to operate, the satellite can operate after attitude adjustment. Therefore, the current constraint is an energy condition that the satellite operation must meet.

[0052] In some embodiments, the current constraint is expressed by formula (2) as: (2) Wherein, is the minimum current required for the satellite, the current of the solar panel in the target satellite attitude, , is a neural network model learned through on-orbit data.

[0053] In some embodiments, the sensor constraint in step S20 is configured as: the third angle θ3 between the optical axis of the satellite system sensor and the satellite system celestial body vector is greater than the celestial body suppression angle .

[0054] This application adopts an inertial orientation and attitude control mode to measure the satellite attitude through attitude sensors (such as gyroscopes and star sensors), compare it with the inertial coordinate system, and the controller generates control commands according to the deviation signal. Since the satellite relies on star sensors for attitude determination, at least one star sensor of the satellite should be ensured to be available after attitude adjustment. Therefore, the sensor constraint is a constraint condition that must be satisfied when the satellite adjusts its attitude.

[0055] In some embodiments, the sensor constraint is expressed by formula (3) as: (3) Wherein, the optical axis of the sensor in the satellite body coordinate system is , represents the celestial vector, is the suppression angle corresponding to the celestial body. In some embodiments, the celestial vector is the solar vector in the target body coordinate system, the lunar vector or the geocentric vector .

[0056] In some embodiments, step S20 further includes: adjusting the current satellite attitude according to the angular momentum constraint, and the angular momentum constraint is configured as: the moduli of the X, Y, and Z axis components of the overall satellite angular momentum in the target satellite attitude are respectively less than the first threshold.

[0057] In some embodiments, the angular momentum constraint is expressed by formula (4) as: (4) Wherein, is the modulus of the X axis component of the overall satellite angular momentum in the target satellite attitude, is the modulus of the Y axis component of the overall satellite angular momentum H1 in the target satellite attitude, is the modulus of the Z axis component of the overall satellite angular momentum H1 in the target satellite attitude, is the satellite angular momentum unloading threshold.

[0058] In some embodiments, the overall satellite angular momentum H1 in the target satellite attitude is obtained through formula (5): (5) Wherein, T is: the time from the completion of the last attitude adjustment to the start of the next attitude adjustment, with the unit of day.

[0059] In some embodiments, the unloading threshold is determined according to the on-board hardware configuration of the satellite.

[0060] When the satellite meets the light pressure unloading conditions described above, angular momentum constraints can be added in step S20. Under the condition of meeting the angular momentum constraints, the natural light pressure is used to unload the angular momentum of the whole satellite, which can reduce the number of on-orbit propulsion jets of the satellite and extend the service life of the satellite in the Earth-Moon space.

[0061] In some embodiments, step S20 further includes: adjusting the current satellite attitude according to the TT&C constraints, where the TT&C constraints are configured such that the fourth angle θ4 between the geocentric vector of this system and the satellite Z-axis vector of this system belongs to a third range. It should be understood that the TT&C antennas are assembled on the ±Z axes of the satellite. To ensure TT&C visibility after the satellite attitude adjustment, the fourth angle θ4 between the geocentric vector of this system and the satellite Z-axis vector of this system needs to belong to the third range, so that the satellite meets the TT&C requirements. Among them, the satellite Z-axis vector includes the ±Z-axis directions of the satellite.

[0062] In some embodiments, the third range of the TT&C constraints can be expressed by formula (6) or formula (7): (6) (7) Where, is the geocentric vector of this system, is the half beam angle of the TT&C antenna, is the satellite Z-axis vector of this system.

[0063] In some embodiments, the above current constraints, sensor constraints, angular momentum constraints, and TT&C constraints can be preset, and then the target satellite attitude can be calculated and obtained.

[0064] In some embodiments, the axis of the intermediate coordinate system in step S20 is the target heliocentric vector of this system; the axis of the intermediate coordinate system is determined by the right-hand rule. By setting the axis of the intermediate coordinate system as the target heliocentric vector of this system, the solar panels after the satellite attitude adjustment can face the target heliocentric vector of this system, obtaining the maximum solar panel current value, and thus enabling the satellite to meet the current constraints.

[0065] In some embodiments, the following formula (8) is used to obtain the axis of the intermediate coordinate system: (8).

[0066] When the satellite rotates around the target heliocentric vector of this system, the geocentric vector falls into the axis vector of the satellite body coordinate system and the target system sun vector constituting the Z - O - S b within the plane, it can ensure that the geocentric vector and the satellite Z - axis vector of the system have the smallest fourth angle θ4 between them, meeting the TT&C constraints.

[0067] In some embodiments, the direction of the satellite Z - axis vector should be set according to time. Figure 6 shows a schematic diagram of the direction of the satellite Z - axis vector in an embodiment of the present application. As Figure 6 shown, the satellite around the Earth at the Earth - Moon libration point has an orbital period of about 27 days. When the required target system sun vector has a negative projection direction on the satellite Z - axis, the satellite uses the - Z - axis to point to the Earth between the 22nd and the 7th of the next month, and uses the + Z - axis to point to the Earth between the 7th and the 22nd; when the required target system sun vector has a positive projection direction on the satellite Z - axis, the satellite uses the Z - axis to point to the Earth between the 22nd and the 7th of the next month, and uses the - Z - axis to point to the Earth between the 7th and the 22nd.

[0068] In some embodiments, in order to ensure that the geocentric vector and the satellite Z - axis vector of the system have the smallest fourth angle θ4 between them, meeting the TT&C constraints, in step S20, according to the rotation angle calculate the transformation matrix from the intermediate coordinate system to the target inertial system, that is, the first transformation matrix.

[0069] In some embodiments, the first transformation matrix in step S20 is obtained by the following formula (9): (9) wherein, and are both the first transformation matrix, is the rotation angle.

[0070] In some embodiments, the rotation angle is calculated by the following formula (10): (10) wherein, is the rotation angle, is the y - axis component of the geocentric vector in the intermediate coordinate system, is the z - axis component of the geocentric vector in the intermediate coordinate system.

[0071] In some embodiments, in response to the fifth angle θ5 between the geocentric vector of the target inertial system and the Z - axis vector of the intermediate coordinate system belonging to the fourth range, the rotation angle is recalculated by the following formula (11): (11).

[0072] In some embodiments, the fourth range is: the geocentric vector of the target inertial system and the fifth angle θ5 between the Z-axis vector of the intermediate coordinate system is a right angle or an obtuse angle, that is, wherein, , is the geocentric vector in the intermediate coordinate system.

[0073] By making the axis of the intermediate coordinate system be the target body-fixed solar vector , and making the axis of the intermediate coordinate system be the satellite Z-axis vector of the body-fixed system and the target body-fixed solar vector to form the Z-O-S b plane, and calculating the first transformation matrix according to the rotation angle so that the satellite attitude adjustment satisfies the TT&C constraints.

[0074] After obtaining the first transformation matrix in step S20, obtain the second transformation matrix: the transformation matrix from the body-fixed system to the intermediate coordinate system, and the attitude matrix: the projection of the body-fixed system in the target inertial system.

[0075] In some embodiments, the second transformation matrix in step S20 is obtained by the following formula (12): (12) wherein, is the second transformation matrix, is the x-axis component of the target body-fixed solar vector , is the y-axis component of the target body-fixed solar vector , is the z-axis component of the target body-fixed solar vector .

[0076] In some embodiments, obtaining the attitude matrix according to the first transformation matrix and the second transformation matrix in step S20 includes: obtaining the attitude matrix by the following formula (13): (13) wherein, is the attitude matrix, is the second transformation matrix, is the first transformation matrix.

[0077] The actual process of satellite attitude adjustment is that the satellite rotates by an angle θ around the target body-fixed solar vector . In some embodiments, in step S20, the following formulas (14) and (15) are used to obtain the target satellite attitude according to the attitude matrix: (14) (15) Wherein, is the target satellite attitude, is the rotation angle of the satellite around the target body-fixed solar vector from the current attitude to the target satellite attitude, is the identity matrix, e is the unit vector of the target body-fixed solar vector, and E is the matrix composed of the unit vector e of the target body-fixed solar vector.

[0078] In some embodiments, the following formula (16) is used to calculate E: (16).

[0079] In some embodiments, through the target satellite attitude the overall satellite angular momentum of the target satellite attitude can be calculated, and the calculated overall satellite angular momentum of the target satellite attitude satisfies the light pressure unloading condition and angular momentum constraint described above. In some embodiments, the following formula (17) is used to calculate the overall satellite angular momentum of the target satellite attitude: (17) Wherein, H1 represents the overall satellite angular momentum of the target satellite attitude, is the target satellite attitude, is the attitude matrix, represents the transpose matrix of the attitude matrix, and H0 is the overall satellite angular momentum of the current satellite attitude.

[0080] In some embodiments, after obtaining the target satellite attitude in step S20, the inertial orientation attitude control mode can be adopted to adjust the current satellite attitude to the target satellite attitude.

[0081] In some embodiments, after obtaining the target satellite attitude, it further includes: taking the target satellite attitude as the current satellite attitude, optimizing the current satellite attitude according to the constraint function, and obtaining the optimized target satellite attitude; wherein, the constraint function includes a first constraint function, a second constraint function, and a third constraint function.

[0082] Next, it will be combined with the attached Figure 7 and Figure 8 to explain how to obtain the optimized target satellite attitude.

[0083] Figure 7 Fig. 700 shows the flowchart of the satellite attitude management method according to another embodiment of the present application.

[0084] In some embodiments, taking the target satellite attitude as the current satellite attitude and optimizing the current satellite attitude according to the constraint function to obtain the optimized target satellite attitude includes: Step S71: Obtain the target body-fixed solar vector; Step S72: Determine whether the satellite meets the light pressure unloading condition; Step S73: When the satellite meets the light pressure unloading condition, sequentially determine whether there is a feasible solution for the first constraint function, the second constraint function, and the third constraint function until the first solution is obtained. Among them, the first constraint function includes current constraint, sensor constraint, TT&C constraint, and angular momentum constraint; the second constraint function includes current constraint, sensor constraint, and TT&C constraint; the third constraint function includes current constraint and sensor constraint; and Step S74: Optimize the current satellite attitude using the constraint function corresponding to the first solution to obtain the optimized target satellite attitude.

[0085] In the above embodiments, by sequentially determining whether there is a feasible solution for the first constraint function, the second constraint function, and the third constraint function until the first solution is obtained, gradually relaxing the restrictions of the constraint function, an optimized target satellite attitude that can maximally meet the constraint conditions can be obtained.

[0086] Among them, step S71 is equivalent to the previous step S10, and the neural network model algorithm can also be used to obtain the target body-fixed solar vector. Step S72 can also judge the light pressure unloading condition according to the light pressure unloading condition described above. Step S74 is equivalent to step S20, and the optimized target satellite attitude can also be obtained according to step S20 described above.

[0087] In step S73, since the two constraints of current constraint and sensor constraint are the conditions that must be met for satellite attitude adjustment, therefore, the first constraint function, the second constraint function, and the third constraint function all include current constraint and sensor constraint.

[0088] In some embodiments, in response to the non-existence of a feasible solution for the first constraint function or the second constraint function in step S73, determine whether the interval between the current time and the end time of the most recent attitude adjustment is less than the first interval; in response to the interval between the current time and the end time of the most recent attitude adjustment being less than the first interval, adjust the interval between the start time of the next attitude adjustment and the end time of the most recent attitude adjustment to the second interval, and sequentially determine whether there is a feasible solution for the remaining constraint functions, where the second interval is greater than the first interval; and in response to the interval between the current time and the end time of the most recent attitude adjustment being greater than the first interval, adjust the interval between the start time of the next attitude adjustment and the end time of the most recent attitude adjustment to the third interval, and sequentially determine whether there is a feasible solution for the remaining constraint functions, where the third interval is less than the first interval.

[0089] It can be understood that as time changes, both the environment around the satellite and the satellite itself will change. When there is no feasible solution for the first constraint function or the second constraint function, by judging the relationship between the interval between the current time and the end time of the nearest attitude adjustment and the first interval and adjusting the interval between the start time of the next attitude adjustment and the end time of the nearest attitude adjustment, an optimized target satellite attitude corresponding to different time periods can be better obtained. The satellite attitude management method of the present application can obtain an optimized target satellite attitude by adjusting the attitude adjustment interval, and can be used for long-term attitude management of on-orbit satellites.

[0090] In some embodiments, the first interval, the second interval, and the third interval can be set according to the actual on-board environment and satellite conditions. In some embodiments, the second interval does not exceed: 2 * half beam angle / (360 / 28) days. Among them, 360 / 28 represents the angular velocity of the moon orbiting the earth.

[0091] In some embodiments, optimizing the current satellite attitude according to the constraint function in step S73 includes: using the sequential quadratic programming algorithm (SQP) to optimize the current satellite attitude to obtain an optimized target satellite attitude. In some embodiments, multiple solutions can be obtained using the SQP algorithm, each solution corresponding to multiple sail current values, and there is a minimum sail current value among the multiple sail current values. In some embodiments, according to actual needs, different minimum time units can be used as the calculation frequency to calculate the sail current values corresponding to each solution, including: using days or hours as the minimum unit as the calculation frequency. For example, when calculating each solution with days as the minimum unit, assuming the interval between two consecutive attitude adjustment times is 7 days, then each solution has 7 corresponding sail current values, and the minimum sail current value is the smallest value among the 7 sail current values. In some embodiments, the objective function can be set as: comparing the minimum sail current values among multiple solutions, and taking the solution corresponding to the largest value among the multiple minimum sail current values as the optimized target satellite attitude. When the first constraint function does not converge and no feasible solution appears, then judge the second constraint function. If it does not converge, then continue to judge the third constraint function. The decision variable can be set as the satellite attitude quaternion.

[0092] By setting the first constraint function, the second constraint function, and the third constraint function, and optimizing with the target satellite attitude as the input, an optimized target satellite attitude is obtained. According to the optimization result, the satellite attitude is adjusted and controlled so that after the satellite attitude adjustment is completed, all the above constraints are satisfied as much as possible, and sufficient energy can be obtained.

[0093] Next, it will be combined with the attached Figure 8 , and further expand how to optimize the current satellite attitude according to the constraint function to obtain an optimized target satellite attitude.

[0094] Figure 8The flowchart 800 of the satellite attitude management method according to an embodiment of the present application is shown. As Figure 8 shown, optimizing the current satellite attitude according to the constraint function to obtain the optimized target satellite attitude includes: Step S801: Satellite solar panel state evaluation. Step S801 is equivalent to steps S71 and S10 described above. In step S801, the satellite solar panel current value corresponding to any solar vector in the local system is obtained through the neural network algorithm model, and then the solar vector in the local system with the largest corresponding satellite solar panel current value is selected as the target solar vector in the local system.

[0095] Step S802: Determine whether the light pressure unloading condition is satisfied. Step S802 is equivalent to step S72 described above. When the light pressure unloading condition is satisfied, step S803 is executed.

[0096] Step S803: Optimize the target satellite attitude using the first constraint function. The first constraint function is set as the constraint function composed of current constraint, sensor constraint, TT&C constraint, and angular momentum constraint. Then step S804 is executed.

[0097] Step S804: Determine whether the first constraint function has a feasible solution. In step S804, it is judged whether the satellite can simultaneously satisfy the current constraint, sensor constraint, TT&C constraint, and angular momentum constraint. When there is a feasible solution, the attitude optimization is completed, and then the optimized target satellite attitude is obtained through step S74, and the satellite attitude is adjusted according to the optimized target satellite attitude. When there is no feasible solution, step S805 is executed.

[0098] Step S805: Determine whether the attitude adjustment time interval T is less than 5 days. That is, determine whether the interval between the current time and the end time of the most recent attitude adjustment is less than 5 days. When the interval between the current time and the end time of the most recent attitude adjustment is greater than or equal to 5 days, step S806 is executed; when it is less than 5 days, step S807 is executed.

[0099] Step S806: Narrow the interval between two attitude adjustments. Among them, narrowing the interval between two attitude adjustments means narrowing the interval between the start time of the next attitude adjustment and the end time of the most recent attitude adjustment. Then, return to step S803.

[0100] Step S807: Set the attitude adjustment interval T = 10 days. That is, set the interval between the start time of the next attitude adjustment and the end time of the most recent attitude adjustment to 10 days. Then, continue to execute step S808.

[0101] Step S808: Optimize the target attitude using the second constraint function. The second constraint function includes current constraint, sensor constraint, and TT&C constraint. Then, step S809 is executed.

[0102] Step S809: Determine whether the second constraint function has a feasible solution. In step S809, it is determined whether the satellite can simultaneously satisfy the current constraint, the sensor constraint, and the TT&C constraint. When there is a feasible solution, the attitude optimization is completed. Then, the optimized target satellite attitude is obtained through step S74, and the satellite attitude is adjusted according to the optimized target satellite attitude. When there is no feasible solution, step S810 is executed.

[0103] Step S810: Determine whether the attitude adjustment time interval T is less than 5 days. That is, determine whether the interval between the current time and the end time of the most recent attitude adjustment is less than 5 days. When the interval between the current time and the end time of the most recent attitude adjustment is greater than or equal to 5 days, step S811 is executed; when it is less than 5 days, step S812 is executed.

[0104] Step S811: Narrow the interval between two attitude adjustments. Among them, narrowing the interval between two attitude adjustments means narrowing the interval between the start time of the next attitude adjustment and the end time of the most recent attitude adjustment. Then, return to step S808.

[0105] Step S812: Set the attitude adjustment interval T = 10 days. That is, set the interval between the start time of the next attitude adjustment and the end time of the most recent attitude adjustment to 10 days. Then, continue to execute step S813.

[0106] Step S813: Perform target attitude optimization using the third constraint function. The third constraint function includes the current constraint and the sensor constraint. Since the current constraint and the sensor constraint are constraint conditions that must be satisfied, when using the third constraint function, the attitude optimization can be directly completed. Then, the optimized target satellite attitude is obtained through step S74, and the satellite attitude is adjusted according to the optimized target satellite attitude.

[0107] The above satellite attitude management method of the present application can also be used as a long-term satellite attitude management method for on-orbit satellites. In some embodiments, first, the target satellite attitude is obtained according to the above method. Then, the target satellite attitude is used as the current satellite attitude, and the optimized target satellite attitude is obtained according to the constraint function and the satellite attitude is adjusted to the optimized target satellite attitude. Then, the optimized target satellite attitude can be used as the current satellite attitude again, and it is judged whether the constraint function has a feasible solution at different attitude adjustment intervals, and the optimized target satellite attitude is continuously obtained. Furthermore, through iterative optimization, long-term management of the satellite attitude is realized.

[0108] Figure 9 The schematic diagram of the input current of the satellite solar panel and the change of the total load current of the satellite shows a satellite attitude management method according to an embodiment of the present application. As Figure 9As shown, during several attitude adjustments, the input current of the satellite solar panel is always greater than the load current, that is, the satellite always maintains energy balance and meets the current constraint.

[0109] Figure 10 The angular momentum change curve of the satellite attitude management method according to an embodiment of the present application is shown. Among them, the abscissa is: time; the ordinate is: the total satellite angular momentum, and the unit is Nms. As Figure 10 shown, during the three attitude adjustments of the satellite, the Y-axis component of the total satellite angular momentum can be significantly reduced without the satellite actively unloading, and the X-axis component and Z-axis component of the total satellite angular momentum change less. Overall, the total satellite angular momentum can be naturally reduced by relying on natural light pressure unloading.

[0110] The satellite attitude management method of the present application estimates the specific pointing of the solar wing by using a neural network model based on the solar vector and solar panel current data of the system. The solar vector of the system corresponding to the maximum current is used as the target solar vector of the system, ensuring that sufficient energy can be obtained when the pointing of the solar wing is unknown. According to the included angle relationship between the target solar vector of the system corresponding to the maximum current and the angular momentum before and after attitude adjustment, it is judged whether the light pressure unloading condition is met. When the light pressure unloading condition is met, the angular momentum is reduced by light pressure unloading, which reduces the number of jet unloadings of the on-orbit satellite, saves satellite propulsion fuel, and extends the service life of the satellite. By sequentially judging whether there is a feasible solution for the first constraint function, the second constraint function, and the third constraint function until the first solution is obtained, and gradually relaxing the restrictions of the constraint functions, an optimized target satellite attitude that can most meet the constraint conditions can be obtained. By adjusting the attitude adjustment interval and iterative optimization to obtain the optimized target satellite attitude, the satellite attitude management method of the present application can be used for the long-term attitude management of on-orbit satellites.

[0111] Compared with the traditional satellite attitude management method, the significant advantages of the satellite attitude management method of the present application are: 1) The target satellite attitude calculated by the satellite attitude management method of the present application can obtain the maximum energy while meeting the current constraint, the sensor constraint, the TT&C constraint, and the sensor constraint, providing a reliable method for the long-term on-orbit management of the satellite; 2) Using the sunlight pressure to naturally reduce the angular momentum of the satellite body can keep the three-axis angular momentum within a smaller range for a longer time, avoiding frequent unloading operations, reducing the consumption of propulsion fuel, and increasing the satellite life.

[0112] The present application also includes a satellite attitude management device, including a memory and a processor. Among them, the memory is used to store instructions executable by the processor; the processor is used to execute the instructions to implement the satellite attitude management method described above.

[0113] Figure 11 is the system block diagram of the satellite attitude management device according to an embodiment of the present application. Refer to Figure 11As shown, the satellite attitude management device 1100 may include an internal communication bus 1101, a processor 1102, a read-only memory (ROM) 1103, a random access memory (RAM) 1104, and a communication port 1105. When applied to a personal computer, the satellite attitude management device 1100 may further include a hard disk 1106. The internal communication bus 1101 can enable data communication among the components of the satellite attitude management device 1100. The processor 1102 can make judgments and issue prompts. In some embodiments, the processor 1102 may be composed of one or more processors. The communication port 1105 can enable data communication between the satellite attitude management device 1100 and the outside. In some embodiments, the satellite attitude management device 1100 can send and receive information and data from a network through the communication port 1105. The satellite attitude management device 1100 may further include different forms of program storage units and data storage units, such as the hard disk 1106, the read-only memory (ROM) 1103, and the random access memory (RAM) 1104, which can store various data files used for computer processing and / or communication, as well as possible program instructions executed by the processor 1102. The processor executes these instructions to implement the main part of the method. The result processed by the processor is transmitted to the user device through the communication port and displayed on the user interface.

[0114] The above-mentioned satellite attitude management method can be implemented as a computer program, stored in the hard disk 1106, and loaded into the processor 1102 for execution to implement the satellite attitude management method of the present application.

[0115] The present application further includes a computer-readable medium storing computer program code, and the computer program code implements the satellite attitude management method described above when executed by a processor.

[0116] When the satellite attitude management method of the present application is implemented as a computer program, it can also be stored in a computer-readable storage medium as an article of manufacture. For example, the computer-readable storage medium may include, but is not limited to, magnetic storage devices (such as hard disks, floppy disks, magnetic strips), optical disks (such as compact discs (CDs), digital versatile discs (DVDs)), smart cards, and flash memory devices (such as electrically erasable programmable read-only memories (EPROMs), cards, sticks, key drives). In addition, the various storage media described herein can represent one or more devices and / or other machine-readable media for storing information. The term "machine-readable medium" may include, but is not limited to, wireless channels and various other media (and / or storage media) that can store, contain, and / or carry code and / or instructions and / or data.

[0117] It should be understood that the embodiments described above are merely illustrative. The embodiments described herein can be implemented in hardware, software, firmware, middleware, microcode, or any combination thereof. For a hardware implementation, the processor can be implemented in one or more application specific integrated circuits (ASICs), digital signal processors (DSPs), digital signal processing devices (DSPDs), programmable logic devices (PLDs), field programmable gate arrays (FPGAs), processors, controllers, microcontrollers, microprocessors, and / or other electronic units designed to perform the functions described herein, or a combination thereof.

[0118] Some aspects of the present application can be executed entirely by hardware, entirely by software (including firmware, resident software, microcode, etc.), or by a combination of hardware and software. The above hardware or software can all be referred to as "data blocks", "modules", "engines", "units", "components", or "systems". The processor can be one or more application specific integrated circuits (ASICs), digital signal processors (DSPs), digital signal processing devices (DAPDs), programmable logic devices (PLDs), field programmable gate arrays (FPGAs), processors, controllers, microcontrollers, microprocessors, or a combination thereof. In addition, aspects of the present application may be embodied as a computer product located in one or more computer-readable media, the product including computer-readable program code. For example, the computer-readable media may include, but is not limited to, magnetic storage devices (such as hard disks, floppy disks, magnetic tapes...), optical discs (such as compact discs CD, digital versatile discs DVD...), smart cards, and flash memory devices (such as cards, sticks, key drives...).

[0119] The computer-readable media may contain a propagated data signal having computer program code embodied therein, for example, on a baseband or as part of a carrier wave. The propagated signal may have various forms of representation, including electromagnetic form, optical form, etc., or a suitable combination thereof. The computer-readable media can be any computer-readable media other than a computer-readable storage medium, which can be connected to an instruction execution system, apparatus, or device to effect communication, propagation, or transmission for use of the program. The program code located on the computer-readable media can be propagated through any suitable medium, including radio, cable, fiber optic cable, radio frequency signal, or similar media, or any combination of the above media.

[0120] The basic concepts have been described above. Obviously, for those skilled in the art, the above invention disclosure is merely an example and does not constitute a limitation to the present application. Although not explicitly stated herein, those skilled in the art may make various modifications, improvements, and corrections to the present application. Such modifications, improvements, and corrections are proposed in the present application, so such modifications, improvements, and corrections still fall within the spirit and scope of the exemplary embodiments of the present application.

[0121] Meanwhile, specific terms are used in this application to describe the embodiments of this application. For example, "an embodiment", "one embodiment", and / or "some embodiments" mean a certain feature, structure, or characteristic related to at least one embodiment of this application. Therefore, it should be emphasized and noted that the "one embodiment" or "an embodiment" or "an alternative embodiment" mentioned twice or more at different positions in this specification does not necessarily refer to the same embodiment. In addition, certain features, structures, or characteristics in one or more embodiments of this application can be appropriately combined.

[0122] In some embodiments, numbers are used to describe components and the quantity of attributes. It should be understood that such numbers used to describe embodiments are, in some examples, modified by the modifiers "about", "approximately", or "substantially". Unless otherwise specified, "about", "approximately", or "substantially" indicate that the said numbers allow a ±20% variation. Accordingly, in some embodiments, the numerical parameters used in the specification are approximate values, and such approximate values can change according to the characteristics required by individual embodiments. In some embodiments, the numerical parameters should consider the specified significant digits and adopt the method of retaining general digits. Although the numerical ranges and parameters used to confirm the breadth of the scope in some embodiments of this application are approximate values, in specific embodiments, the setting of such numerical values is as precise as possible within the feasible range.

Claims

1. A satellite attitude management method, characterized in that, Including: Obtaining a target body-frame solar vector, where the target body-frame solar vector is the body-frame solar vector when the sail current value is the largest; And Adjusting the current satellite attitude according to current constraints, sensor constraints, and the target body-frame solar vector to obtain a target satellite attitude, including: Establishing an intermediate coordinate system based on the satellite Z-axis vector of the body-frame and the target body-frame solar vector; Obtaining a first transformation matrix, where the first transformation matrix is the transformation matrix from the intermediate coordinate system to the target inertial system; Obtaining a second transformation matrix, where the second transformation matrix is the transformation matrix from the body-frame to the intermediate coordinate system; Obtaining an attitude matrix according to the first transformation matrix and the second transformation matrix, where the attitude matrix is the projection of the body-frame in the target inertial system; and Obtaining the target satellite attitude according to the attitude matrix.

2. The satellite attitude management method according to claim 1, wherein, After obtaining the target body-frame solar vector, it further includes: judging whether the satellite meets the light pressure unloading condition, including: in response to the overall satellite angular momentum in the current satellite attitude being greater than the overall satellite angular momentum in the target satellite attitude, or, in response to the first included angle between the overall satellite angular momentum in the target satellite attitude and the angular momentum increment belonging to a first range, or in response to the second included angle between the target body-frame solar vector and the overall satellite angular momentum in the current satellite attitude belonging to a second range, judging that the satellite meets the light pressure unloading condition.

3. The satellite attitude management method according to claim 1, characterized in that The current constraint is configured as: the sail current in the target satellite attitude is greater than the minimum current required for satellite operation.

4. The satellite attitude management method according to claim 1, wherein, The sensor constraint is configured as: the third included angle between the optical axis of the body-frame sensor and the celestial vector of the body-frame is greater than the celestial suppression angle.

5. The satellite attitude management method according to claim 1, characterized in that Adjusting the current satellite attitude according to current constraints, sensor constraints, and the target body-frame solar vector to obtain a target satellite attitude further includes: adjusting the current satellite attitude according to angular momentum constraints, where the angular momentum constraints are configured as: the magnitudes of the X, Y, and Z axis components of the overall satellite angular momentum in the target satellite attitude are respectively less than a first threshold.

6. The satellite attitude management method according to claim 1, characterized in that, Adjusting the current satellite attitude according to current constraints, sensor constraints, and the target body-frame solar vector to obtain a target satellite attitude further includes: adjusting the current satellite attitude according to TT&C constraints, where the TT&C constraints are configured as: the fourth included angle between the geocentric vector of the body-frame and the satellite Z-axis vector of the body-frame belongs to a third range.

7. The satellite attitude management method according to claim 1, wherein, The first transformation matrix is obtained by using the following formula: ， Among them, and are both the first conversion matrix, is the rotation angle.

8. The satellite attitude management method according to claim 1, characterized in that, The second transformation matrix is obtained by using the following formula: ， wherein, is the second conversion matrix, is the x-axis component of the target body system solar vector, is the y-axis component of the target body system solar vector, is the z-axis component of the target body system solar vector.

9. The satellite attitude management method according to claim 7, characterized in that, The rotation angle is calculated by using the following formula: ， Among them, is the said corner, is the y-axis component of the geocentric vector in the intermediate coordinate system, is the z-axis component of the geocentric vector in the intermediate coordinate system.

10. The satellite attitude management method according to claim 9, wherein In response to the fifth included angle between the geocentric vector of the target inertial system and the Z-axis vector of the intermediate coordinate system belonging to the fourth range, the following formula is used to recalculate the rotation angle: .

11. The satellite attitude management method according to claim 1, characterized in that, Obtaining the attitude matrix according to the first transformation matrix and the second transformation matrix includes: obtaining the attitude matrix by using the following formula: ， Among them, and are both the attitude matrices, is the second conversion matrix, is the first conversion matrix.

12. The satellite attitude management method according to claim 11, wherein The target satellite attitude is obtained according to the attitude matrix by using the following formula: ， ， Among them, is the attitude of the target satellite, is the rotation angle required for the satellite to adjust from the current attitude to the attitude of the target satellite around the target body-fixed solar vector, is the identity matrix, e is the unit vector of the target body-fixed solar vector, and E is the matrix composed of the unit vector e of the target body-fixed solar vector.

13. The satellite attitude management method according to claim 12, characterized in that, The E is calculated by using the following formula: ， Among them, is the x-axis component of the unit vector of the target body system solar vector, is the y-axis component of the unit vector of the target body system solar vector, is the z-axis component of the unit vector of the target body system solar vector.

14. The satellite attitude management method according to claim 1, wherein, The overall satellite angular momentum of the target satellite attitude is calculated by using the following formula: ， where H1 represents the overall angular momentum of the target satellite attitude, is the target satellite attitude, is the attitude matrix, represents the transpose matrix of the attitude matrix, and H0 is the overall angular momentum of the current satellite attitude.

15. The satellite attitude management method according to claim 1, characterized in that, Obtaining the target body-frame solar vector includes: obtaining the target body-frame solar vector by using a neural network model, where the input data of the neural network model includes any body-frame solar vector, and the output data includes the sail current value.

16. The satellite attitude management method according to claim 1, wherein, After obtaining the target satellite attitude, the following steps are further included: taking the target satellite attitude as the current satellite attitude, and optimizing the current satellite attitude according to a constraint function to obtain an optimized target satellite attitude; wherein, the constraint function includes a first constraint function, a second constraint function, and a third constraint function.

17. The satellite attitude management method according to any one of claims 1-16, taking the target satellite attitude as the current satellite attitude, and optimizing the current satellite attitude according to a constraint function to obtain an optimized target satellite attitude includes: Obtaining the target body-fixed solar vector; Judging whether the satellite meets the condition of light pressure unloading; When the satellite meets the condition of light pressure unloading, sequentially judging whether the first constraint function, the second constraint function, and the third constraint function have feasible solutions until the first solution is obtained, wherein the first constraint function includes current constraint, sensor constraint, TT&C constraint, and angular momentum constraint; the second constraint function includes the current constraint, the sensor constraint, and the TT&C constraint; the third constraint function includes the current constraint and the sensor constraint; And Optimizing the current satellite attitude by using the constraint function corresponding to the first solution to obtain an optimized target satellite attitude.

18. The satellite attitude management method according to claim 17, wherein It further includes: In response to the non-existence of a feasible solution for the first constraint function or the second constraint function, judging whether the interval between the current time and the end time of the most recent attitude adjustment is less than a first interval; In response to the interval between the current time and the end time of the most recent attitude adjustment being less than the first interval, adjusting the interval between the start time of the next attitude adjustment and the end time of the most recent attitude adjustment to a second interval, and sequentially judging whether the remaining constraint functions have feasible solutions, wherein the second interval is greater than the first interval; and In response to the interval between the current time and the end time of the most recent attitude adjustment being greater than the first interval, adjusting the interval between the start time of the next attitude adjustment and the end time of the most recent attitude adjustment to a third interval, and sequentially judging whether the remaining constraint functions have feasible solutions, wherein the third interval is less than the first interval.

19. The satellite attitude management method according to claim 18, wherein optimizing the current satellite attitude according to the constraint function includes: Using the sequential quadratic programming algorithm to optimize the current satellite attitude to obtain an optimized target satellite attitude.

20. A satellite attitude management device, characterized in that, It includes: A memory for storing instructions executed by a processor; A processor for executing the instructions to implement the method according to any one of claims 1-19.

21. A computer-readable medium storing computer program code, characterized in that, The computer program code implements the method according to any one of claims 1-19 when executed by a processor.

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