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

Abstract

The present application relates to a satellite attitude management method, a satellite attitude management device, and a computer-readable medium. The present application provides a satellite attitude management method, including: obtaining a target local system solar vector, which is the local system solar vector when the sailboard current value is maximum; and adjusting the current satellite attitude according to the current constraint, the sensor constraint, and the target local system solar vector to obtain the target satellite attitude, including: establishing an intermediate coordinate system according to the local system's satellite Z-axis vector and the target local system solar vector; obtaining a first conversion matrix, which is the conversion matrix from the intermediate coordinate system to the target inertial system; obtaining a second conversion matrix, which is the conversion matrix from the local system to the intermediate coordinate system; obtaining an attitude matrix according to the first conversion matrix and the second conversion matrix, which is the projection of the local system in the target inertial system; and obtaining the target satellite attitude according to the attitude matrix.

Description

Technical Field

[0001] The present application mainly relates to the field of satellite attitude control technology, and specifically 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 which collisions with space debris are one of the most important problems. Even the impact of tiny debris can cause deformation of the solar panel structure or short circuit of the circuit, making it impossible for the ground to accurately determine the direction of the solar panel based on the real-time data sent back by the satellite, thus posing a huge risk to the stable operation of the satellite. At present, the conventional satellite attitude management method is to control the closed-loop attitude according to the preset guidance law to the target satellite attitude based on the mission requirements, while rotating the solar panel to maintain the maximum angle of incidence of sunlight on the solar cell array to obtain more energy. However, this method can perform attitude control when the solar panel is pointing according to the normal operating mode, but it cannot guarantee that sufficient energy can be obtained when the solar panel is pointing unknownly. Summary of the Invention

[0003] In response to the above technical problems, the present application provides a satellite attitude management method for solving the technical problem that the satellite cannot obtain sufficient energy supply when the direction of the solar panels is unknown.

[0004] The technical solution adopted by the present application to solve the above-mentioned technical problems is a satellite attitude management method, including: obtaining a target local system solar vector, the target local system solar vector is the local system solar vector when the panel current value is maximum; and adjusting the current satellite attitude according to the current constraint, the sensor constraint and the target local system solar vector to obtain the target satellite attitude, including: establishing an intermediate coordinate system according to the satellite Z-axis vector of the local system and the target local system solar vector; obtaining a first conversion matrix, the first conversion matrix is the conversion matrix from the intermediate coordinate system to the target inertial system; obtaining a second conversion matrix, the second conversion matrix is the conversion matrix from the local system to the intermediate coordinate system; obtaining an attitude matrix according to the first conversion matrix and the second conversion matrix, the attitude matrix is the projection of the local 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 local system solar vector, it also includes: judging whether the satellite meets the light pressure unloading condition, including: in response to the whole satellite angular momentum under the current satellite attitude being greater than the whole satellite angular momentum under the target attitude and the target satellite attitude, or, in response to the first angle between the whole satellite angular momentum and the angular momentum increment under the target satellite attitude belonging to the first range, or in response to the second angle between the target local system solar vector and the whole satellite angular momentum under the current satellite attitude belonging to the second range, judging that the satellite meets the light pressure unloading condition.

[0006] In an embodiment of the present application, the current constraint is configured such that the sailboard current under the target satellite attitude is greater than the minimum current required for satellite operation.

[0007] In one embodiment of the present application, the sensor constraint is configured as follows: a third angle between the optical axis of the sensor of the system and the celestial body vector of the system 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 system solar vector to obtain the target satellite attitude also 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-axis, Y-axis, and Z-axis components of the whole satellite angular momentum under the target satellite attitude are respectively less than a first threshold.

[0009] In one embodiment of the present application, the adjusting of the current satellite attitude according to the current constraint, the sensor constraint and the target system solar vector to obtain the target satellite attitude also includes: measurement and control constraints, and the measurement and control constraints are configured as: the fourth angle between the system geocentric vector and the system satellite Z-axis vector belongs to the third range.

[0010] In one embodiment of the present application, the first conversion matrix is obtained using the following formula:

[0011] ,

[0012] in, and are the first transformation matrices, For the corner.

[0013] In one embodiment of the present application, the second conversion matrix is obtained using the following formula:

[0014] ,

[0015] in, is the second transformation matrix, is the x-axis component of the solar vector of the target system, is the y-axis component of the solar vector of the target system, is the z-axis component of the solar vector of the target system.

[0016] In one embodiment of the present application, the rotation angle is calculated using the following formula:

[0017] ,

[0018] in, is the turning 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.

[0019] 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 falling within the fourth range, the rotation angle is recalculated using the following formula: .

[0020] In one embodiment of the present application, obtaining the posture matrix according to the first transformation matrix and the second transformation matrix includes: obtaining the posture matrix using the following formula:

[0021] ,

[0022] in, and are the posture matrices, is the first transformation matrix, is the second transformation matrix.

[0023] In one embodiment of the present application, the target satellite attitude is obtained according to the attitude matrix using the following formula:

[0024] ,

[0025] ,

[0026] in, is the target satellite attitude, The rotation angle of the satellite around the target system's sun vector is required for the satellite to adjust from its current attitude to the target satellite attitude. is the unit matrix, e is the unit vector of the target system's solar vector, and E is the matrix composed of the unit vector e of the target system's solar vector.

[0027] In one embodiment of the present application, the following formula is used to calculate E:

[0028] ,

[0029] in, is the x-axis component of the unit vector of the target system's solar vector, is the y-axis component of the unit vector of the target system's solar vector, is the z-axis component of the unit vector of the target system's solar vector.

[0030] In one embodiment of the present application, the following formula is used to calculate the angular momentum of the target satellite attitude:

[0031] ,

[0032] Among them, H1 represents the angular momentum of the target satellite attitude, is the target satellite attitude, is the posture matrix, represents the transposed matrix of the attitude matrix, and H0 is the angular momentum of the current satellite attitude.

[0033] In one embodiment of the present application, obtaining the target local system solar vector includes: using a neural network model to obtain the target local system solar vector, wherein the input data of the neural network model includes any local system solar vector, and the output data includes the windsurfing panel current value.

[0034] In one embodiment of the present application, after obtaining the target satellite attitude, it also includes: taking the target satellite attitude as the current satellite attitude, optimizing the current satellite attitude according to a 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.

[0035] In one embodiment of the present application, the target satellite attitude is used as the current satellite attitude, and the current satellite attitude is optimized according to the constraint function to obtain the optimized target satellite attitude, which includes: obtaining the target system solar vector; judging whether the satellite meets the light pressure unloading condition; when the satellite meets the light pressure unloading condition, judging in turn 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, wherein the first constraint function includes current constraint, sensor constraint, measurement and control constraint and angular momentum constraint; the second constraint function includes the current constraint, the sensor constraint and the measurement and control constraint; the third constraint function includes the current constraint and the sensor constraint; and using the constraint function corresponding to the first solution to optimize the current satellite attitude to obtain the optimized target satellite attitude.

[0036] In one embodiment of the present application, in response to the absence 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 posture adjustment is less than the first interval; in response to the interval between the current time and the end time of the most recent posture adjustment being less than the first interval, the interval between the start time of the next posture adjustment and the end time of the most recent posture adjustment is adjusted to the second interval, and whether the remaining constraint functions have feasible solutions is determined in turn, 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 posture adjustment being greater than the first interval, the interval between the start time of the next posture adjustment and the end time of the most recent posture adjustment is adjusted to the third interval, and whether the remaining constraint functions have feasible solutions is determined in turn, wherein the third interval is less than the first interval.

[0037] In one embodiment of the present application, optimizing the current satellite attitude according to the constraint function includes: optimizing the current satellite attitude using a sequential quadratic programming algorithm to obtain an optimized target satellite attitude.

[0038] The present application also provides a satellite attitude management device, comprising: a memory for storing instructions executable by a processor; and a processor for executing the instructions to implement the above method.

[0039] The present application also provides a computer-readable medium storing computer program code, which implements the above method when executed by a processor.

[0040] This application obtains the target satellite attitude by obtaining the target system solar vector and adjusting the current satellite attitude according to the current constraint, sensor constraint and target system solar vector, so that the satellite can obtain sufficient energy supply when the solar wing pointing is unknown. BRIEF DESCRIPTION OF THE DRAWINGS

[0041] In order to make the above-mentioned objects, features and advantages of the present application more clearly understood, the specific embodiments of the present application are described in detail below with reference to the accompanying drawings, wherein:

[0042] Figure 1 is a flow chart of a satellite attitude management method according to an embodiment of the present application;

[0043] Figure 2 Schematic diagram of a neural network model of a satellite attitude management method according to an embodiment of the present application;

[0044] Figure 3 is a flow chart of a satellite attitude management method according to another embodiment of the present application;

[0045] Figure 41 is a schematic diagram of a second angle of a satellite attitude management method according to an embodiment of the present application;

[0046] Figure 5 The angle distribution diagram between the target system sun vector and the angular momentum increment of the satellite attitude management method according to one embodiment of the present application is shown;

[0047] Figure 6 2 is a 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;

[0048] Figure 7 A flowchart of a satellite attitude management method according to another embodiment of the present application is shown;

[0049] Figure 8 A flowchart of a satellite attitude management method according to an embodiment of the present application is shown;

[0050] Figure 9 A schematic diagram showing changes in satellite panel input current and satellite total load current in a satellite attitude management method according to an embodiment of the present application is shown;

[0051] Figure 10 shows an angular momentum change curve of a satellite attitude management method according to an embodiment of the present application;

[0052] Figure 11 This is a system block diagram of a satellite attitude management device according to an embodiment of the present application. DETAILED DESCRIPTION

[0053] In order to make the above-mentioned objectives, features and advantages of the present application more obvious and easy to understand, the specific implementation methods of the present application are described in detail below with reference to the accompanying drawings.

[0054] In the following description, many specific details are set forth to facilitate a full understanding of the present application. However, the present application may also be implemented in other ways different from those described herein. Therefore, the present application is not limited to the specific embodiments disclosed below.

[0055] As used in this application, unless the context clearly indicates otherwise, the words "a," "an," "an," and / or "the" are not intended to refer to the singular but may include the plural. Generally speaking, the terms "comprise" and "include" only indicate the inclusion of the steps and elements specifically identified, and these steps and elements do not constitute an exclusive list. A method or apparatus may also include other steps or elements.

[0056] The present application provides a satellite attitude management method, which uses the solar vector of the system corresponding to the maximum sailboard current value as the target solar vector of the system, and adjusts the current satellite attitude according to the constraint conditions and the target solar vector of the system to obtain the target satellite attitude, thereby ensuring that the satellite can obtain sufficient energy when the solar wing pointing is unknown. The satellite attitude management method of the present application also determines whether the satellite meets the light pressure unloading conditions. When the light pressure unloading conditions are met, the satellite unloads the angular momentum of the entire satellite under natural light pressure, thereby reducing the number of jet unloading times of the satellite in orbit, saving satellite propulsion fuel, and extending the service life of the satellite. In addition, the satellite attitude management method of the present application achieves that the satellite can obtain sufficient energy for a long time by obtaining the optimized target satellite attitude, thereby meeting the long-term operation of the satellite.

[0057] Next, the satellite attitude management method of the present application will be described in detail with reference to the accompanying drawings.

[0058] Figure 1 FIG. 1 is a flow chart of a satellite attitude management method 100 according to an embodiment of the present application. Figure 1 As shown, the present application provides a satellite attitude management method 100, comprising:

[0059] Step S10: obtaining a target solar vector of the system, where the target solar vector of the system is the solar vector of the system when the panel current value is the maximum; and

[0060] Step S20: Adjusting the current satellite attitude according to the current constraint, the sensor constraint, and the target local system solar vector to obtain the target satellite attitude, including: establishing an intermediate coordinate system according to the local system's satellite Z-axis vector and the target local system solar vector; obtaining a first conversion matrix, which is a conversion matrix from the intermediate coordinate system to the target inertial system; obtaining a second conversion matrix, which is a conversion matrix from the local system to the intermediate coordinate system; obtaining an attitude matrix according to the first conversion matrix and the second conversion matrix, which is a projection of the local system in the target inertial system; and obtaining the target satellite attitude according to the attitude matrix.

[0061] The "local system" in this application's satellite attitude management method refers to the satellite's own coordinate system. This method obtains the target local system solar vector and adjusts the current satellite attitude based on current constraints, sensor constraints, and the target local system solar vector to achieve the target satellite attitude. This allows the satellite to obtain sufficient energy even when the solar wing's orientation is unknown.

[0062] In some embodiments, in step S10, obtaining the target local system solar vector includes: using a neural network model to obtain the target local system solar vector, wherein the input data of the neural network model includes any local system solar vector, and the output data includes the panel current value.

[0063] Figure 2 FIG. 1 is a schematic diagram showing a neural network model 11 of a satellite attitude management method according to an embodiment of the present application. Figure 2 As shown, the neural network model 11 includes an input layer 111, a hidden layer 112, and an output layer 113.

[0064] In some embodiments, the present application optimizes and iterates neural network model 11 by inputting multiple local solar vectors as input data into input layer 111 of neural network model 11, causing output layer 113 to output the panel current values corresponding to the multiple local solar vectors. Subsequently, based on the optimized neural network model 11, the panel current value output when any local solar vector is used as input data is estimated, and the local solar vector corresponding to the maximum panel current value is used as the target local solar vector.

[0065] Figure 3 Shown Figure 2 The illustrated embodiment shows the distribution of the panel current value under any solar vector input of the system, wherein the horizontal axis is the azimuth angle of the solar vector of the system, in degrees (deg); the vertical axis is the pitch angle of the solar vector of the system, in degrees (deg); the legend on the right shows that different colors represent the corresponding panel current values, in amperes (A). Figure 3 As shown in the figure, when the azimuth angle of the solar vector of the system is in the range of [200, 220] and the pitch angle range is [-60, -40], the corresponding sailboard current value is large. To ensure the energy supply of the satellite, the solar vector of the system with an azimuth angle range of [200, 220] and a pitch angle range of [-60, -40] can be selected as the target solar vector of the system.

[0066] It is understood that due to differences in the satellite itself and the environment, the resulting panel current value distribution diagram for any local solar vector input condition will be different. In some embodiments, a local solar vector within an appropriate azimuth and elevation angle range can be selected as the target local solar vector based on the actual environmental conditions using a neural network model.

[0067] In some embodiments, after obtaining the target local system solar vector in step S10, it also includes: judging whether the satellite meets the light pressure unloading condition, including: in response to the whole satellite angular momentum under the current satellite attitude being greater than the whole satellite angular momentum under the target satellite attitude, or, in response to the first angle between the whole satellite angular momentum under the target satellite attitude and the direction of the angular momentum increment belonging to the first range, or in response to the second angle between the target local system solar vector and the whole satellite angular momentum under the current satellite attitude belonging to the second range, judging that the satellite meets the light pressure unloading condition.

[0068] The total satellite angular momentum refers to the sum of the angular momentum of the entire satellite system (including the satellite itself and any components such as momentum wheels). When a satellite meets the conditions for solar pressure unloading, an angular momentum constraint can be added in step S20. During attitude adjustments, the satellite can utilize natural solar pressure to unload the total satellite angular momentum. This reduces the number of jet unloading operations on the satellite in orbit, conserves propellant, and extends the satellite's service life.

[0069] It can be understood that the result of natural light pressure unloading the entire satellite's angular momentum is that the entire satellite's angular momentum in the target satellite's attitude after unloading is less than the entire satellite's angular momentum in the current satellite's attitude before unloading. In other words, when the entire satellite's angular momentum in the current satellite's attitude is greater than the entire satellite's angular momentum in the target satellite's attitude, the satellite meets the light pressure unloading condition.

[0070] In some embodiments, the whole-satellite angular momentum at the current satellite attitude is greater than the whole-satellite angular momentum at the target satellite attitude means that the modulus of the whole-satellite angular momentum at the current satellite attitude is greater than the modulus of the whole-satellite angular momentum at the target satellite attitude.

[0071] In some embodiments, the light pressure unloading condition is expressed by formula (1):

[0072] (1)

[0073] Among them, H1 is the whole satellite angular momentum under the target satellite attitude, H0 is the whole satellite angular momentum under the current satellite attitude, is the increment of angular momentum.

[0074] In some embodiments, Represents the increment of angular momentum in one day.

[0075] In some embodiments, when the first angle θ1 between the whole satellite angular momentum and the angular momentum increment direction in the target satellite attitude is an obtuse angle, that is, the first range is: When , the light pressure unloading condition is met.

[0076] In some embodiments, when the second range of the second angle θ2 between the target local system sun vector and the whole satellite angular momentum at the current satellite attitude is [45,135], the relationship expressed in formula (1) is satisfied, that is, the satellite satisfies the light pressure unloading condition, and the whole satellite angular momentum at the current satellite attitude is greater than the whole satellite angular momentum at the target satellite attitude.

[0077] In some embodiments, a distribution diagram of the angle θ' between the target local system solar vector and the angular momentum increment can be drawn based on the first range of the second angle θ2, the range of the target local system solar vector and the whole-star angular momentum increase telemetry. Figure 4 FIG. 1 is a schematic diagram showing the second angle of the satellite attitude management method according to an embodiment of the present application. Figure 4As shown, θ2 is the second angle, H0 is the angular momentum of the entire satellite under the current satellite attitude, is the angular momentum increment, is the target system solar vector, θ' is the angle between the target system solar vector and the angular momentum increment. Figure 5 This figure shows the distribution of the angle θ' between the target local solar vector and the angular momentum increment for one embodiment of the present application. The horizontal axis represents the azimuth angle of the local solar vector, in degrees (deg); the vertical axis represents the pitch angle of the local solar vector, in degrees (deg). The legend on the right shows different colors representing the corresponding angle θ' between the target local solar vector and the angular momentum increment, in degrees (deg).

[0078] like Figure 5 As shown in the figure, when the second angle θ2 is in the range of [45, 135] and the target local system solar vector range is selected to be [200, 220] in the azimuth angle range and [-60, -40] in the pitch angle range, the angle θ' between the target local system solar vector and the angular momentum increment is in the range of [74, 92]. In this case, the first angle θ1 between the whole satellite angular momentum in the target satellite attitude and the direction of the angular momentum increment is an obtuse angle, which satisfies that the whole satellite angular momentum in the current satellite attitude is greater than the whole satellite angular momentum in the target satellite attitude.

[0079] In some embodiments, after determining whether the satellite meets the optical pressure unloading condition, step S20 is executed.

[0080] In some embodiments, the current constraint in step S20 is configured such that the panel current under the target satellite attitude is greater than the minimum current required for satellite operation.

[0081] It should be understood that the satellite can only operate after attitude adjustment if the current of the sailboard in the target satellite attitude is greater than the minimum current required for satellite operation. Therefore, the current constraint is the energy condition that the satellite must meet for operation.

[0082] In some embodiments, the current constraint is expressed using formula (2):

[0083] (2)

[0084] in, is the minimum current required by the satellite, The sailboard current under the target satellite attitude, , A neural network model learned from in-orbit data.

[0085] In some embodiments, the sensor constraint in step S20 is configured as follows: the third angle θ3 between the optical axis of the sensor of the system and the celestial body vector of the system is greater than the celestial body suppression angle .

[0086] This application uses an inertial orientation attitude control mode. This mode measures the satellite's attitude using attitude sensors (such as gyroscopes and star sensors) and compares it with the inertial coordinate system. A controller generates control commands based on the deviation signal. Because the satellite relies on star sensors for attitude determination, at least one star sensor must be available after attitude adjustments. Therefore, sensor constraints are mandatory for satellite attitude adjustments.

[0087] In some embodiments, the sensor constraint is expressed using formula (3):

[0088] (3)

[0089] Among them, the optical axis of the sensor in the satellite body coordinate system is , represents the celestial vector, is the suppression angle of the corresponding celestial body. In some embodiments, the celestial body vector The solar vector of the target system , Moon Vector or geocentric vector .

[0090] In some embodiments, step S20 further includes: adjusting the current satellite attitude according to the angular momentum constraint, where the angular momentum constraint is configured as follows: the moduli of the X-, Y-, and Z-axis components of the whole satellite angular momentum under the target satellite attitude are respectively less than a first threshold.

[0091] In some embodiments, the angular momentum constraint is expressed using formula (4):

[0092] (4)

[0093] in, is the angular momentum of the entire satellite under the target satellite attitude The magnitude of the X-axis component, is the modulus of the Y-axis component of the whole satellite angular momentum H1 under the target satellite attitude, is the modulus of the Z-axis component of the whole satellite angular momentum H1 under the target satellite attitude, is the satellite angular momentum unloading threshold.

[0094] In some embodiments, the whole satellite angular momentum H1 under the target satellite attitude is obtained by formula (5):

[0095] (5)

[0096] Where T is the time from the completion of the last posture adjustment to the start of the next posture adjustment, in days.

[0097] In some embodiments, the uninstall threshold Determined according to the satellite's onboard hardware configuration.

[0098] When the satellite meets the aforementioned conditions for light pressure unloading, an angular momentum constraint can be added in step S20. By utilizing natural light pressure to unload the entire satellite's angular momentum, the number of on-orbit propulsion jets can be reduced, extending the service life of the cis-lunar satellite.

[0099] 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 a fourth angle θ4 between the system's geocentric vector and the system's satellite Z-axis vector falls within a third range. It should be understood that the TT&C antenna is mounted on the satellite's ±Z axes. To ensure TT&C visibility after satellite attitude adjustment, the fourth angle θ4 between the system's geocentric vector and the system's satellite Z-axis vector must fall within the third range so that the satellite meets TT&C requirements. The satellite Z-axis vector includes the satellite's ±Z-axis directions.

[0100] In some embodiments, the third range of the measurement and control constraint can be expressed using formula (6) or formula (7):

[0101] (6)

[0102] (7)

[0103] in, is the geocentric vector of this system, is the half-beam angle of the measurement and control antenna, is the satellite Z-axis vector of this system.

[0104] In some embodiments, the above-mentioned current constraints, sensor constraints, angular momentum constraints and measurement and control constraints can be pre-set, and then the target satellite attitude is calculated and obtained.

[0105] In some embodiments, the intermediate coordinate system of step S20 Axis is the target system solar vector ; The intermediate coordinate system The axis is determined by the right-hand rule. Axis is the target system solar vector , so that the satellite's sailboard can face the target system's sun vector after attitude adjustment , to obtain the maximum sailboard current value, so that the satellite meets the current constraint.

[0106] In some embodiments, the intermediate coordinate system is obtained using the following formula (8): axis:

[0107] (8).

[0108] When the satellite orbits the target, the solar vector During rotation, the geocentric vector falls into the satellite's body coordinate system Axis vector and the target system solar vector Composition of ZOS b The plane can ensure that the geocentric vector and the satellite Z axis vector of this system The fourth included angle θ4 between them is the smallest, satisfying the measurement and control constraints.

[0109] In some embodiments, the direction of the satellite Z-axis vector should be set according to time. Figure 6 Schematic diagram showing the direction of the satellite Z-axis vector according to an embodiment of the present application. Figure 6 As shown in the figure, the period of the Earth-Moon space libration point satellite orbiting the Earth is about 27 days. When the target system solar vector When the satellite's Z-axis projection direction is negative, the satellite uses the -Z-axis pointing to the earth between the 22nd and the 7th of the next month, and uses the +Z-axis pointing to the earth between the 7th and the 22nd. When the desired target is the sun vector of the system When the satellite's Z-axis projection direction is positive, the satellite uses the Z-axis pointing to the Earth between the 22nd and the 7th of the next month, and uses the -Z-axis pointing to the Earth between the 7th and the 22nd.

[0110] In some embodiments, in order to ensure that the geocentric vector and the satellite Z axis vector of the system The fourth angle θ4 between them is the smallest, which satisfies the measurement and control constraints. In step S20, according to the angle Calculate the transformation matrix from the intermediate coordinate system to the target inertial system, that is, the first transformation matrix.

[0111] In some embodiments, the first conversion matrix in step S20 is obtained using the following formula (9):

[0112] (9)

[0113] in, and are the first transformation matrices, For the corner.

[0114] In some embodiments, the following formula (10) is used to calculate the rotation angle: :

[0115] (10)

[0116] in, For corners, 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.

[0117] In some embodiments, in response to the geocentric vector of the target inertial system and the Z-axis vector of the intermediate coordinate system The fifth angle θ5 between them belongs to the fourth range, and the rotation angle is recalculated using the following formula (11):

[0118] (11).

[0119] In some embodiments, the fourth range is: the geocentric vector of the target inertial system The fifth angle θ5 between the Z-axis vector of the intermediate coordinate system is a right angle or an obtuse angle, that is, ,in, , is the geocentric vector in the intermediate coordinate system.

[0120] By putting the intermediate coordinate system Axis is the target system solar vector , the intermediate coordinate system The axis is set to the satellite Z axis vector of this system and the target system sun vector Composition of ZOS b Surface, and according to the corner Calculate the first transformation matrix so that the satellite satisfies the measurement and control constraints after attitude adjustment.

[0121] After obtaining the first transformation matrix in step S20, the second transformation matrix is obtained: the transformation matrix of the system to the intermediate coordinate system, and the attitude matrix: the projection of the system in the target inertial system.

[0122] In some embodiments, the second transformation matrix in step S20 is obtained using the following formula (12):

[0123] (12)

[0124] in, is the second transformation matrix, The solar vector of the target system The x-axis component of The solar vector of the target system The y-axis component of The solar vector of the target system The z-axis component of .

[0125] In some embodiments, obtaining the posture matrix according to the first conversion matrix and the second conversion matrix in step S20 includes: obtaining the posture matrix using the following formula (13):

[0126] (13)

[0127] in, is the posture matrix, is the first transformation matrix, is the second transformation matrix.

[0128] The actual process of satellite attitude adjustment is that the satellite orbits the target system sun vector Rotation angle θ. In some embodiments, in step S20, the attitude of the target satellite is obtained according to the attitude matrix using the following formula (14) and formula (15):

[0129] (14)

[0130] (15)

[0131] in, is the target satellite attitude, The rotation angle of the satellite around the target system's sun vector is required for the satellite to adjust from its current attitude to the target satellite attitude. is the unit matrix, e is the unit vector of the target system's solar vector, and E is the matrix composed of the unit vector e of the target system's solar vector.

[0132] In some embodiments, E is calculated using the following formula (16):

[0133] (16).

[0134] In some embodiments, the target satellite attitude The whole-satellite angular momentum of the target satellite attitude can be calculated, and the calculated whole-satellite angular momentum of the target satellite attitude satisfies the light pressure unloading condition and angular momentum constraint mentioned above. In some embodiments, the following formula (17) is used to calculate: the whole-satellite angular momentum of the target satellite attitude:

[0135] (17)

[0136] Among them, H1 represents the whole satellite angular momentum of the target satellite attitude, is the target satellite attitude, is the posture matrix, represents the transposed matrix of the attitude matrix, and H0 is the whole satellite angular momentum of the current satellite attitude.

[0137] In some embodiments, after obtaining the target satellite attitude in step S20, an inertial orientation attitude control mode may be used to adjust the current satellite attitude to the target satellite attitude.

[0138] In some embodiments, after obtaining the target satellite attitude, the method 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.

[0139] Next, we will combine the Figure 7 and Figure 8 , expand and explain how to obtain the optimized target satellite attitude.

[0140] Figure 7 A flowchart 700 of a satellite attitude management method according to another embodiment of the present application is shown.

[0141] In some embodiments, 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 includes:

[0142] Step S71: Obtain the target system solar vector;

[0143] Step S72: Determine whether the satellite meets the optical pressure unloading condition;

[0144] Step S73: When the satellite meets the optical pressure unloading condition, determine whether there are feasible solutions for the first constraint function, the second constraint function, and the third constraint function in sequence until a first solution is obtained, wherein the first constraint function includes current constraint, sensor constraint, measurement and control constraint, and angular momentum constraint; the second constraint function includes current constraint, sensor constraint, and measurement and control constraint; and the third constraint function includes current constraint and sensor constraint; and

[0145] Step S74: Optimize the current satellite attitude using the constraint function corresponding to the first solution to obtain the optimized target satellite attitude.

[0146] In the above embodiment, by successively determining whether there are feasible solutions to 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, the optimized target satellite attitude that can meet the constraint conditions to the greatest extent can be obtained.

[0147] Step S71 is equivalent to step S10, and similarly, the neural network model algorithm can be used to obtain the target system's solar vector. Step S72 can also determine the solar pressure unloading condition based on the solar pressure unloading condition described above. Step S74 is equivalent to step S20, and similarly, the optimized target satellite attitude can be obtained based on step S20.

[0148] In step S73 , since the current constraint and the sensor constraint are the conditions that must be met for satellite attitude adjustment, the first constraint function, the second constraint function and the third constraint function all include the current constraint and the sensor constraint.

[0149] In some embodiments, in response to the absence of a feasible solution for the first constraint function or the second constraint function in step S73, it is determined whether the interval between the current time and the end time of the most recent posture adjustment is less than the first interval; in response to the interval between the current time and the end time of the most recent posture adjustment being less than the first interval, the interval between the start time of the next posture adjustment and the end time of the most recent posture adjustment is adjusted to the second interval, and it is determined in turn 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 posture adjustment being greater than the first interval, the interval between the start time of the next posture adjustment and the end time of the most recent posture adjustment is adjusted to the third interval, and it is determined in turn whether the remaining constraint functions have feasible solutions, wherein the third interval is less than the first interval.

[0150] It is understandable that over time, the satellite's surrounding environment and the satellite's own conditions will change. When no feasible solution exists for the first or second constraint function, by determining the relationship between the interval between the current time and the end time of the most recent 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 most recent attitude adjustment, the optimized target satellite attitude corresponding to different time periods can be better obtained. The satellite attitude management method of the present application obtains the optimized target satellite attitude by adjusting the attitude adjustment interval, and can be used for long-term attitude management of in-orbit satellites.

[0151] In some embodiments, the first, second, and third intervals may be set based on the actual onboard environment and satellite conditions. In some embodiments, the second interval does not exceed: 2 * half beam angle / (360 / 28) days. Where 360 / 28 represents the angular velocity of the moon around the earth.

[0152] In some embodiments, optimizing the current satellite attitude based on the constraint function in step S73 includes optimizing the current satellite attitude using a sequential quadratic programming (SQP) algorithm to obtain an optimized target satellite attitude. In some embodiments, the SQP algorithm can yield multiple solutions, each corresponding to multiple panel current values, with a minimum panel current value among the multiple panel current values. In some embodiments, the panel current values corresponding to each solution can be calculated using different minimum time units as the calculation frequency, depending on practical needs, including using days or hours as the minimum calculation frequency. For example, if each solution is calculated using days as the minimum unit, assuming the interval between two attitude adjustments is 7 days, each solution has 7 corresponding panel current values, and the minimum panel current value is the minimum of the 7 panel current values. In some embodiments, the objective function can be configured to compare the minimum panel current values among the multiple solutions and select the solution corresponding to the largest minimum panel current value as the optimized target satellite attitude. If the first constraint function does not converge and no feasible solution is found, the second constraint function is evaluated. If convergence is not achieved, the third constraint function is evaluated. The decision variable can be a satellite attitude quaternion.

[0153] By setting the first, second, and third constraint functions and optimizing the target satellite attitude as input, the optimized target satellite attitude is obtained. Based on the optimization results, the satellite attitude is adjusted and controlled so that after the satellite attitude adjustment is completed, all the above constraints are met as much as possible and sufficient energy is obtained.

[0154] Next, we will combine the Figure 8 , further expand on how to optimize the current satellite attitude according to the constraint function and obtain the optimized target satellite attitude.

[0155] Figure 8 FIG8 shows a flow chart 800 of a satellite attitude management method according to an embodiment of the present application. Figure 8 As shown, the current satellite attitude is optimized according to the constraint function to obtain the optimized target satellite attitude, including:

[0156] Step S801: Satellite Panel Status Assessment. Step S801 is equivalent to steps S71 and S10 described above. In step S801, the satellite panel current value corresponding to any local solar vector is obtained using a neural network algorithm model. The local solar vector with the largest corresponding satellite panel current value is then selected as the target local solar vector.

[0157] Step S802: Determine whether the light pressure unloading condition is met. Step S802 is equivalent to step S72 described above. If the light pressure unloading condition is met, execute step S803.

[0158] Step S803: Optimize the target satellite attitude using the first constraint function. Set the first constraint function to a constraint function consisting of current constraints, sensor constraints, measurement and control constraints, and angular momentum constraints. Then, execute step S804.

[0159] Step S804: Determine whether the first constraint function has a feasible solution. In step S804, it is determined whether the satellite can simultaneously satisfy the current constraint, sensor constraint, measurement and control constraint, and angular momentum constraint. If a feasible solution exists, attitude optimization is completed. Then, step S74 is performed to obtain the optimized target satellite attitude, and the satellite attitude is adjusted based on the optimized target satellite attitude. If no feasible solution exists, step S805 is executed.

[0160] Step S805: Determine whether the posture adjustment interval T is less than 5 days. Specifically, determine whether the interval between the current time and the end time of the most recent posture adjustment is less than 5 days. If the interval between the current time and the end time of the most recent posture adjustment is greater than or equal to 5 days, proceed to step S806; if the interval is less than 5 days, proceed to step S807.

[0161] Step S806: Reduce the interval between the two posture adjustments. Reduce the interval between the two posture adjustments, that is, reduce the interval between the start time of the next posture adjustment and the end time of the most recent posture adjustment. Then, return to step S803.

[0162] Step S807: Set the posture adjustment interval T to 10 days. That is, set the interval between the start time of the next posture adjustment and the end time of the most recent posture adjustment to 10 days. Then, proceed to step S808.

[0163] Step S808: Optimize the target posture using the second constraint function. The second constraint function includes current constraints, sensor constraints, and measurement and control constraints. Then, execute step S809.

[0164] Step S809: Determine whether the second constraint function has a feasible solution. In step S809, determine whether the satellite can simultaneously satisfy the current constraint, sensor constraint, and TT&C constraint. If a feasible solution exists, attitude optimization is completed. Then, step S74 is performed to obtain the optimized target satellite attitude, and the satellite attitude is adjusted based on the optimized target satellite attitude. If no feasible solution exists, step S810 is executed.

[0165] Step S810: Determine whether the posture adjustment interval T is less than 5 days. Specifically, determine whether the interval between the current time and the end time of the most recent posture adjustment is less than 5 days. If the interval between the current time and the end time of the most recent posture adjustment is greater than or equal to 5 days, proceed to step S811; if the interval is less than 5 days, proceed to step S812.

[0166] Step S811: Reduce the interval between the two posture adjustments. Reduce the interval between the two posture adjustments, that is, reduce the interval between the start time of the next posture adjustment and the end time of the most recent posture adjustment. Then, return to step S808.

[0167] Step S812: Set the posture adjustment interval T to 10 days. That is, set the interval between the start time of the next posture adjustment and the end time of the most recent posture adjustment to 10 days. Then, proceed to step S813.

[0168] Step S813: Target attitude optimization is performed using the third constraint function. The third constraint function includes current constraints and sensor constraints. Since these constraints must be satisfied, attitude optimization can be directly completed using the third constraint function. The optimized target satellite attitude is then obtained in step S74, and the satellite attitude is adjusted based on the optimized target satellite attitude.

[0169] The satellite attitude management method described above can also be used as a long-term satellite attitude management method for in-orbit satellites. In some embodiments, the target satellite attitude is first obtained according to the above method. The target satellite attitude is then used as the current satellite attitude. The optimized target satellite attitude is obtained according to the constraint function and the satellite attitude is adjusted to the optimized target satellite attitude. The optimized target satellite attitude can then be used as the current satellite attitude. The constraint function can be determined to have a feasible solution at different attitude adjustment intervals, and the optimized target satellite attitude can be obtained again. Furthermore, through iterative optimization, long-term satellite attitude management can be achieved.

[0170] Figure 9 The following is a schematic diagram showing the change of satellite panel input current and satellite total load current in a satellite attitude management method according to an embodiment of the present application. Figure 9 As shown in Figure 3, during several attitude adjustments, the satellite panel input current is always greater than the load current, that is, the satellite always maintains energy balance and meets the current constraint.

[0171] Figure 10 The angular momentum change curve of the satellite attitude management method according to one embodiment of the present application is shown. The horizontal axis is time; the vertical axis is the angular momentum of the entire satellite, in Nms. Figure 10 As shown in the figure, during the three satellite attitude adjustments, the Y-axis component of the satellite's angular momentum can be significantly reduced without the satellite actively unloading, while the X-axis and Z-axis components of the satellite's angular momentum change slightly. Overall, the satellite's angular momentum can be naturally reduced by unloading it with the help of natural solar pressure.

[0172] The satellite attitude management method of the present application uses a neural network model to estimate the specific pointing of the solar wing based on the solar vector and sailboard 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 angular 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 number of jet unloadings of the in-orbit satellite is reduced by unloading the angular momentum through light pressure, saving satellite propulsion fuel and extending the service life of the satellite. By successively judging whether there are feasible solutions to 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 function, it is possible to obtain an optimized target satellite attitude that can meet the constraint conditions to the greatest extent. By adjusting the attitude adjustment interval and iterative optimization, the optimized target satellite attitude is obtained. The satellite attitude management method of the present application can be used for long-term attitude management of in-orbit satellites.

[0173] Compared with traditional satellite attitude management methods, the satellite attitude management method of the present application has the following significant advantages: 1) The target satellite attitude calculated by the satellite attitude management method of the present application satisfies current constraints, sensor constraints, measurement and control constraints, and sensor constraints while the satellite can obtain maximum energy, providing a reliable method for long-term on-orbit management of satellites; 2) The solar light pressure is used to naturally reduce the angular momentum of the star, so that the three-axis angular momentum can be maintained in a smaller range for a longer time, avoiding frequent unloading operations, reducing the consumption of propulsion fuel, and increasing the life of the satellite.

[0174] The present application also includes a satellite attitude management device, comprising a memory and a processor, wherein the memory is used to store instructions executable by the processor; and the processor is used to execute the instructions to implement the satellite attitude management method described above.

[0175] Figure 11 This is a system block diagram of a satellite attitude management device according to an embodiment of the present application. 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 used on a personal computer, the satellite attitude management device 1100 may also include a hard disk 1106. The internal communication bus 1101 enables data communication between components of the satellite attitude management device 1100. The processor 1102 can make decisions and issue prompts. In some embodiments, the processor 1102 may be composed of one or more processors. The communication port 1105 enables data communication between the satellite attitude management device 1100 and the external environment. In some embodiments, the satellite attitude management device 1100 can send and receive information and data from a network via the communication port 1105. The satellite attitude management device 1100 may also include various forms of program storage units and data storage units, such as a hard disk 1106, a read-only memory (ROM) 1103, and a random access memory (RAM) 1104. These 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 components of the method. The results of the processor processing are transmitted to the user equipment via a communication port and displayed on the user interface.

[0176] 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.

[0177] The present application also includes a computer-readable medium storing computer program code, which, when executed by a processor, implements the satellite attitude management method described above.

[0178] 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 a product. For example, a computer-readable storage medium may include, but is not limited to, magnetic storage devices (e.g., hard disks, floppy disks, magnetic strips), optical disks (e.g., compact disks (CDs), digital versatile disks (DVDs)), smart cards, and flash memory devices (e.g., 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.

[0179] It should be understood that the embodiments described above are merely illustrative. The embodiments described herein may be implemented in hardware, software, firmware, middleware, microcode, or any combination thereof. For hardware implementation, the processor may be implemented within 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.

[0180] Some aspects of this application may be implemented entirely in hardware, entirely in software (including firmware, resident software, microcode, etc.), or a combination of hardware and software. These hardware and software components may be referred to as "data blocks," "modules," "engines," "units," "components," or "systems." A processor may be one or more application-specific integrated circuits (ASICs), digital signal processors (DSPs), digital signal processing devices (DSPs), programmable logic devices (PLDs), field-programmable gate arrays (FPGAs), processors, controllers, microcontrollers, microprocessors, or combinations thereof. Furthermore, various aspects of this application may be embodied as computer products embodied in one or more computer-readable media, including computer-readable program code. For example, computer-readable media may include, but are not limited to, magnetic storage devices (e.g., hard disks, floppy disks, magnetic tapes), optical disks (e.g., compact disks, digital versatile disks, DVDs), smart cards, and flash memory devices (e.g., cards, sticks, key drives, etc.).

[0181] A computer-readable medium may include a propagated data signal embodying computer program code, for example, in baseband or as part of a carrier wave. The propagated signal may be in a variety of forms, including electromagnetic, optical, etc., or a suitable combination thereof. A computer-readable medium may be any computer-readable medium other than a computer-readable storage medium that can be connected to an instruction execution system, apparatus, or device to communicate, propagate, or transmit the program for use. The program code on the computer-readable medium may be transmitted via any suitable medium, including radio, cable, fiber optic cable, radio frequency signal, or similar medium, or any combination of the above.

[0182] The basic concepts have been described above. It will be apparent to those skilled in the art that the above disclosures are merely illustrative and do not constitute limitations on this application. Although not explicitly stated herein, those skilled in the art may make various modifications, improvements, and amendments to this application. Such modifications, improvements, and amendments are suggested in this application and remain within the spirit and scope of the exemplary embodiments of this application.

[0183] At the same time, this application uses specific terms to describe the embodiments of this application. For example, "one embodiment," "an embodiment," and / or "some embodiments" refer to a certain feature, structure, or characteristic related to at least one embodiment of this application. Therefore, it should be emphasized and noted that "one embodiment," "an embodiment," or "an alternative embodiment" mentioned twice or multiple times in different locations 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 may be appropriately combined.

[0184] In some embodiments, numbers describing the number of components and attributes are used. It should be understood that such numbers used in the description of the embodiments are modified by the modifiers "about", "approximately" or "substantially" in some examples. Unless otherwise stated, "about", "approximately" or "substantially" indicate that the numbers are allowed to vary by ±20%. Accordingly, in some embodiments, the numerical parameters used in the specification are approximate values, which may vary according to the characteristics required by the individual embodiments. In some embodiments, the numerical parameters should take into account the specified significant digits and adopt the general method of retaining the digits. Although the numerical domains and parameters used to confirm the breadth of their range in some embodiments of the present application are approximate values, in specific embodiments, the settings of such numerical values are as accurate as possible within the feasible range.

Claims

1. A satellite attitude management method, characterized in that: include: Obtaining a target solar vector of the system, where the target solar vector of the system is the solar vector of the system when the panel current value is maximum; as well as Adjusting the current satellite attitude according to the current constraint, the sensor constraint, and the target local system sun vector to obtain the target satellite attitude includes: Establishing an intermediate coordinate system based on the satellite Z-axis vector of the system and the target system sun vector; Obtaining a first transformation matrix, where the first transformation matrix is a transformation matrix from the intermediate coordinate system to the target inertial system; Obtaining a second transformation matrix, where the second transformation matrix is a transformation matrix from the local coordinate system to the intermediate coordinate system; Obtaining a posture matrix according to the first transformation matrix and the second transformation matrix, wherein the posture matrix is a projection of the local system in the target inertial system; and The target satellite attitude is obtained according to the attitude matrix.

2. The satellite attitude management method according to claim 1, wherein: After obtaining the target local system solar vector, the method also includes: judging whether the satellite meets the light pressure unloading condition, including: in response to the whole satellite angular momentum under the current satellite attitude being greater than the whole satellite angular momentum under the target satellite attitude, or in response to the first angle between the whole satellite angular momentum and the angular momentum increment under the target satellite attitude belonging to the first range, or in response to the second angle between the target local system solar vector and the whole satellite angular momentum under the current satellite attitude belonging to the second range, judging that the satellite meets the light pressure unloading condition.

3. The satellite attitude management method according to claim 1, wherein: The current constraint is configured such that the panel current under 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 follows: a third angle between the optical axis of the sensor of the system and the celestial body vector of the system is greater than the celestial body suppression angle.

5. The satellite attitude management method according to claim 1, wherein: The step of adjusting the current satellite attitude according to the current constraint, the sensor constraint, and the target system solar vector to obtain the target satellite attitude further includes: adjusting the current satellite attitude according to the angular momentum constraint, wherein the angular momentum constraint is configured such that the moduli of the X-axis components of the whole satellite angular momentum under the target satellite attitude are respectively less than a first threshold.

6. The satellite attitude management method according to claim 1, wherein: The method of adjusting the current satellite attitude according to the current constraint, the sensor constraint and the target system solar vector to obtain the target satellite attitude also includes: adjusting the current satellite attitude according to the measurement and control constraint, and the measurement and control constraint is configured as follows: the fourth angle between the system geocentric vector and the satellite Z-axis vector of the system belongs to the third range.

7. The satellite attitude management method according to claim 1, characterized in that: The first conversion matrix is obtained using the following formula: ， in, and are the first transformation matrices, For corners; The rotation angle is calculated using the following formula: ， in, 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.

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

9. The satellite attitude management method according to claim 7, wherein: In response to a fifth angle between the geocentric vector of the target inertial system and the Z-axis vector of the intermediate coordinate system falling within a fourth range, the rotation angle is recalculated using the following formula: .

10. The satellite attitude management method according to claim 1, wherein: Obtaining the posture matrix according to the first conversion matrix and the second conversion matrix includes: obtaining the posture matrix using the following formula: ， in, and are the posture matrices, is the first transformation matrix, is the second transformation matrix.

11. The satellite attitude management method according to claim 10, wherein: The target satellite attitude is obtained according to the attitude matrix using the following formula: ， ， in, is the target satellite attitude, The rotation angle of the satellite around the target system's sun vector is required for the satellite to adjust from its current attitude to the target satellite attitude. is the unit matrix, e is the unit vector of the target system's solar vector, and E is the matrix composed of the unit vector e of the target system's solar vector.

12. The satellite attitude management method according to claim 11, wherein: The E is calculated using the following formula: ， in, is the x-axis component of the unit vector of the target system's solar vector, is the y-axis component of the unit vector of the target system's solar vector, is the z-axis component of the unit vector of the target system's solar vector.

13. The satellite attitude management method according to claim 1, wherein: The following formula is used to calculate the whole satellite angular momentum of the target satellite attitude: ， Among them, H1 represents the whole satellite angular momentum of the target satellite attitude, is the target satellite attitude, is the posture matrix, represents the transposed matrix of the attitude matrix, and H0 is the whole satellite angular momentum of the current satellite attitude.

14. The satellite attitude management method according to claim 1, wherein: The obtaining of the target local solar vector includes: obtaining the target local solar vector using a neural network model, wherein the input data of the neural network model includes any local solar vector, and the output data includes the windsurfing panel current value.

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

16. The satellite attitude management method according to claim 15, wherein: 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 includes: Obtaining the target system's solar vector; Determine whether the satellite meets the optical pressure unloading conditions; When the satellite meets the optical pressure unloading condition, determining in sequence whether a first constraint function, a second constraint function, and a third constraint function have feasible solutions until a first solution is obtained, wherein the first constraint function includes a current constraint, a sensor constraint, a measurement and control constraint, and an angular momentum constraint; the second constraint function includes the current constraint, the sensor constraint, and the measurement and control constraint; and the third constraint function includes the current constraint and the sensor constraint; and The constraint function corresponding to the first solution is used to optimize the current satellite attitude to obtain an optimized target satellite attitude.

17. The satellite attitude management method according to claim 16, wherein: Also includes: In response to the first constraint function or the second constraint function not having a feasible solution, determining whether an interval between a current time and an end time of a most recent posture adjustment is less than a first interval; In response to the interval between the current time and the end time of the most recent posture adjustment being less than the first interval, adjusting the interval between the start time of the next posture adjustment to the end time of the most recent posture adjustment to a second interval, and sequentially determining 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 posture adjustment being greater than the first interval, the interval between the start time of the next posture adjustment and the end time of the most recent posture adjustment is adjusted to a third interval, and whether the remaining constraint functions have feasible solutions is determined in turn, wherein the third interval is smaller than the first interval.

18. The satellite attitude management method according to claim 17, wherein optimizing the current satellite attitude according to the constraint function comprises: The current satellite attitude is optimized using a sequential quadratic programming algorithm to obtain an optimized target satellite attitude.

19. A satellite attitude management device, characterized in that: include: a memory for storing instructions to be executed by the processor; A processor, configured to execute the instructions to implement the method according to any one of claims 1 to 18.

20. 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 to 18 when executed by a processor.

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