A satellite attitude control method, system, electronic device and storage medium

Abstract

The present invention discloses a satellite attitude control method, system, electronic device and storage medium, which relates to the technical field of satellite attitude control. The method includes: obtaining the inertia parameter combinations of the satellite in the satellite centroid body coordinate system when all the rotating mechanisms in the satellite's solar panel drive device are at each angular position combination; calculating the PID parameter combinations corresponding to each inertia parameter combination; fitting out the corresponding relationship between the angular position of each rotating mechanism and all the PID parameters; using the PID control law associated with the corresponding relationship to calculate the expected control torque for the satellite to adjust from the current attitude to the final expected attitude, and controlling the actuator of the satellite to output the expected control torque. The present invention can not only meet the control index requirements of the overall satellite attitude stability for the amplitude-frequency characteristic and phase-frequency characteristic, but also adapt to the characteristics of the time-varying inertia caused by the rotation of the solar panels of the satellite, and has relatively high closed-loop control robustness and strong engineering implementation operability.

Description

Background Art

[0002] The attitude control of a satellite includes attitude determination, attitude stabilization control, and attitude maneuver control. Attitude determination is to study the orientation or pointing of the satellite relative to a certain reference datum, and then obtain attitude angle parameters. Its accuracy depends on the accuracy of the attitude sensors and the attitude determination algorithm. Attitude stabilization control is to keep the attitude of the satellite at the expected specified direction and specified value. Attitude maneuver control is the reorientation process of making the satellite transition from one attitude to another.

[0003] After the communication satellite is put into orbit, it first undergoes attitude control processes such as rate damping, solar panel deployment, and attitude capture, and then operates stably in the whole-satellite attitude earth-pointing mode to carry out communication payload work and provide communication service.

[0004] Low-Earth-orbit communication satellites often carry double-wing and double-axis large-area solar panels and operate in inclined or polar orbits. During the on-orbit operation of the satellite, to ensure the on-board energy supply, the solar panels are driven to rotate by a two-dimensional drive mechanism to make the normal of the solar panel face the sun. The solar array drive assembly (SADA) is a complex electromechanical device that outputs a control torque to drive the solar panel to rotate by receiving instructions. The two-degree-of-freedom solar panel drive mechanism is provided with detection devices and drive devices in the directions of two degrees of freedom. The second degree of freedom is controlled by an independent motor, and the function of separating and maneuvering from the satellite body can be realized.

[0005] As a classic attitude control method, PID control is an accurate and advanced control law. Based on the analysis and design of the automatic control principle, when designing a PID controller, factors such as the dynamic characteristics and bandwidth of the system can be considered, and the robustness performance of the system can be reflected by the amplitude margin and phase margin. Therefore, it is still adopted by the vast majority of three-axis stabilized spacecraft so far.

[0006] In the whole-satellite attitude earth-pointing mode of a communication satellite carrying double-wing and double-axis large-area solar panels, the two-axis SADA moves according to a predetermined motion law. The movement of the solar panel causes the dynamic characteristics of the whole satellite to show complex attitude dynamics characteristics, time-varying inertia, aerodynamic torque, gravity gradient torque, residual magnetic torque and other environmental disturbance torques, which vary greatly within the movement envelope of the solar panel and have a great impact on the attitude stabilization control accuracy and control robustness of the whole satellite.

[0007] When a communication satellite carrying double-wing and double-axis large-area solar panels adopts a PID controller, to ensure attitude stability when the inertia and disturbance torque of the whole satellite change greatly due to the movement of the solar panel, the engineering design difficulty of attitude stabilization control is relatively large. There is less research and engineering implementation on how to design and select PID parameters at present. Summary of the Invention

[0008] The technical problem to be solved by the present invention is to provide a satellite attitude control method, system, electronic device and storage medium in view of the deficiencies of the prior art, specifically as follows:

[0009] 1) In the first aspect, the present invention provides a satellite attitude control method, and the specific technical solution is as follows:

[0010] Obtain the inertia parameter combinations of the satellite in the satellite centroid body coordinate system when all the rotating mechanisms in the solar panel drive device of the satellite are at each angular position combination;

[0011] Calculate the PID parameter combinations corresponding to each inertia parameter combination;

[0012] According to each angular position combination of all the rotating mechanisms in the solar panel drive device and all the PID parameter combinations, fit the corresponding relationship between the angular position of each rotating mechanism and all the PID parameters;

[0013] Using the PID control law of the associated correspondence, calculate the desired control torque for the satellite to adjust from the current attitude to the final desired attitude, and control the actuator of the satellite to output the desired control torque.

[0014] The beneficial effects of a satellite attitude control method provided by the present invention are as follows:

[0015] It can not only meet the control index requirements of the amplitude-frequency characteristic and phase-frequency characteristic for the overall satellite attitude stability, but also adapt to the characteristics of the time-varying inertia caused by the rotation of the solar panels. It is applicable to the attitude closed-loop stable control of low-orbit communication satellites carrying large-area solar panels and a small amount of propellant, with relatively high closed-loop control robustness and strong engineering implementation operability.

[0016] On the basis of the above solution, a satellite attitude control method of the present invention can also be improved as follows.

[0017] Further, calculating the PID parameter combinations corresponding to each inertia parameter combination includes:

[0018] Under the constraint of the satellite attitude control accuracy index, use the PID closed-loop control method and adopt the frequency-domain design method to traverse each angular position combination, and calculate the PID parameter combinations corresponding to each inertia parameter combination.

[0019] Further, the solar panel drive device includes three rotating mechanisms arranged in sequence. The rotation axis of the rotating mechanism in the middle is perpendicular to the rotation axes of the rotating mechanisms on both sides. The rotating mechanisms on both sides are respectively connected to a solar panel, and a sun sensor is arranged on each solar panel.

[0020] Further, the satellite is a satellite with time-varying inertia.

[0021] 2) Second aspect, the present invention also provides a satellite attitude control system, and the specific technical solution is as follows:

[0022] It includes an inertia parameter combination acquisition module, a PID parameter combination calculation module, a fitting module, and a satellite attitude control module;

[0023] The inertia parameter combination acquisition module is used to: acquire the inertia parameter combination of the satellite in the satellite centroid body coordinate system when all the rotating mechanisms in the solar panel drive device of the satellite are at each angular position combination;

[0024] The PID parameter combination calculation module is used to: calculate the PID parameter combination corresponding to each inertia parameter combination;

[0025] The fitting module is used to: fit the corresponding relationship between the angular position of each rotating mechanism and all the PID parameters according to each angular position combination of all the rotating mechanisms in the solar panel drive device and all the PID parameter combinations;

[0026] The satellite attitude control module is used to: use the PID control law with associated correspondence to calculate the desired control torque for the satellite to adjust from the current attitude to the final desired attitude, and control the actuator of the satellite to output the desired control torque.

[0027] Based on the above solution, a satellite attitude control system of the present invention can also be improved as follows.

[0028] Further, the PID parameter combination calculation module is used to:

[0029] Under the constraint of the satellite attitude control accuracy index, use the PID closed-loop control method, and adopt the frequency domain design method to traverse each angular position combination to calculate the PID parameter combination corresponding to each inertia parameter combination.

[0030] Further, the solar panel drive device includes three rotating mechanisms arranged in sequence. The rotation axis of the middle rotating mechanism is perpendicular to the rotation axes of the two rotating mechanisms on both sides. Each of the two rotating mechanisms on both sides is connected to a solar panel, and a sun sensor is arranged on each solar panel.

[0031] Further, the satellite is a satellite with time-varying inertia.

[0032] 3) Third aspect, the present invention also provides an electronic device. The electronic device includes a processor, the processor is coupled with a memory, and at least one computer program is stored in the memory. The at least one computer program is loaded and executed by the processor so that the electronic device implements any one of the above satellite attitude control methods.

[0033] 4) Fourth aspect, the present invention further provides a computer-readable storage medium, on which a computer program is stored. When the computer program is executed by a processor, the above-mentioned satellite attitude control method is implemented.

[0034] It should be noted that for the beneficial effects obtained by the technical solutions and corresponding possible implementation manners of the second to fourth aspects of the present invention, reference may be made to the technical effects of the first aspect and its corresponding possible implementation manners described above, which will not be elaborated here. BRIEF DESCRIPTION OF THE DRAWINGS

[0035] In order to more clearly illustrate the technical solutions in the embodiments of the present invention, the following will briefly introduce the drawings required for description in the embodiments of the present invention:

[0036] Figure 1 It is a schematic flowchart of a satellite attitude control method according to an embodiment of the present invention;

[0037] Figure 2 It is a schematic structural diagram of a solar panel drive device;

[0038] Figure 3 It is a schematic structural diagram of a satellite attitude control system according to an embodiment of the present invention;

[0039] Figure 4 It is a schematic structural diagram of an electronic device according to an embodiment of the present invention. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0040] The principles and features of the present invention are described below. The examples given are only for explaining the present invention and are not intended to limit the scope of the present invention.

[0041] The technical solutions of the present invention and how the technical solutions of the present invention solve the above technical problems will be described in detail below with specific embodiments. These several specific embodiments can be combined with each other. For the same or similar concepts or processes, they may not be repeated in some embodiments. The embodiments of the present invention will be described below with reference to the drawings.

[0042] As Figure 1 shown, a satellite attitude control method according to an embodiment of the present invention includes the following steps:

[0043] S1. Obtain the inertia parameter combinations of the satellite in the satellite centroid body coordinate system when all the rotating mechanisms in the solar panel drive device of the satellite are at each angular position combination;

[0044] Among them, the solar panel drive device includes three rotating mechanisms arranged in sequence. The rotation axis of the rotating mechanism in the middle is perpendicular to the rotation axes of the rotating mechanisms on both sides. The rotating mechanisms on both sides are respectively connected to a solar panel, and a sun sensor is arranged on each solar panel. Specifically:

[0045] According to the deployment state of the satellite's solar panels, a rotating mechanism is configured for each solar panel. The satellite includes two solar panels, denoted as the first solar panel and the second solar panel respectively. The rotating mechanism configured for the first solar panel is denoted as B1, and the rotating mechanism configured for the second solar panel is denoted as B2. A rotating mechanism is arranged between B1 and B2, denoted as A. B1, A, and B2 are arranged in sequence. Figure 2 The cube in it is used to represent the satellite body. The satellite body is the remaining part of the satellite excluding the solar panels and the structures connecting the solar panels. Taking the centroid of the satellite body as the origin, a satellite centroid body coordinate system: O-X-Y-Z is established. When both solar panels are deployed, the normal line of the solar cell on the solar panel is consistent with the negative direction of the Z axis of the satellite centroid body coordinate system. Define the position of B1 at this time as the zero position of B1, define the position of B2 at this time as the zero position of B2, and define the position of A at this time as the zero position of A. When B1, B2, and A are all in their respective zero positions, the positive direction of the rotation axis of B1 is consistent with the negative direction of the X axis of the satellite centroid body coordinate system, the positive direction of the rotation axis of B2 is consistent with the positive direction of the X axis of the satellite centroid body coordinate system, and the positive direction of the rotation axis of A is consistent with the positive direction of the Y axis of the satellite centroid body coordinate system.

[0046] Stepper motors are used inside B1, B2, and A. B1 is used to drive the first solar panel to rotate around the rotation axis of B1 through the stepper motor. B2 is used to drive the second solar panel to rotate around the rotation axis of B2 through the stepper motor. A is used to drive the first solar panel and the second solar panel to rotate around the rotation axis of A simultaneously through the stepper motor. During this process, the right-hand rule is used to determine the rotation directions of the first solar panel and the second solar panel.

[0047] Among them, B1, B2, and A return status information through telemetry. The status information includes angle measurement information and whether the rotating mechanism is normal, etc. Moreover, B1, B2, and A can receive instructions to implement functions such as "move at a set speed", "move to a set angular position", and "remain in a holding state after moving to a set position".

[0048] Both B1, B2, and A support a speed - priority motion mode, a position - priority motion mode, and a position - holding mode. The speed - priority motion mode is: setting the angular velocity of the motion mechanism to move according to the set angular velocity magnitude and direction; the position - priority motion mode is: simultaneously setting the motion position and motion speed of the motion mechanism, and the rotating mechanism moves to the set position at the set speed, where the set motion speed only sets the speed magnitude, and the motion direction is judged by the rotating mechanism internally according to the principle of the shortest path; the position - holding mode is that the driving mechanism keeps the current position unchanged.

[0049] Among them, A can rotate continuously by 360°, and B1, B2 can move within the set motion range. The motion - range parameters are set comprehensively according to the Beta angle and star occlusion. Mechanical limit mechanisms are designed outside the comprehensively determined motion range of B1 and B2.

[0050] On one side of the normal positive direction of the solar cells of the first solar panel and the second solar panel, a sun sensor is respectively arranged. The sun sensor is in a pyramid configuration. The sun sensor arranged on the first solar panel is denoted as the first sun sensor, and the sun sensor arranged on the second solar panel is denoted as the second sun sensor. The layout positions are as Figure 3 shown. The field - of - view ranges of the first sun sensor and the second sun sensor are: - 90° to + 90°. The first sun sensor and the second sun sensor collect solar vector information in the sun - sensor measurement coordinate system set by themselves, and the sun - sensor measurement coordinate system can also be reset. An electric - propulsion thruster is configured in the negative X - axis direction of the satellite - centroid body coordinate system.

[0051] When the sun enters the field - of - view ranges of the first sun sensor and the second sun sensor, solar vector information can be obtained. If the sun does not enter the field - of - view ranges of the first sun sensor and the second sun sensor, the obtained solar vector information is a zero vector.

[0052] Among them, each angular - position combination corresponds to an inertia - parameter combination. An angular - position combination includes 3 angular positions, specifically including: the angular position of B1, the angular position of B2, and the angular position of A. An inertia - parameter combination includes 6 inertia parameters, specifically: Jxx, Jyy, Jzz, Jxy, Jxz, and Jyz. Jxx represents: the moment of inertia of the star around the X - axis of the centroid body coordinate system; Jyy represents the moment of inertia of the star around the Y - axis of the centroid body coordinate system; Jzz represents the moment of inertia of the star around the Z - axis of the centroid body coordinate system; Jxy represents the product of inertia of the star about the OXY plane of the centroid body coordinate system; Jxz represents the product of inertia of the star about the OXZ plane of the centroid body coordinate system; Jyz represents the product of inertia of the star about the OYZ plane of the centroid body coordinate system. The unit of the inertia parameter is Kg*m*m.

[0053] The specific implementation process of S1 is as follows:

[0054] S10. A adopts 360-degree continuous rotation control, and the motion range of the angular position of A is set as: -180° to +180°. Both B1 and B2 adopt the stepped stepping control strategy, and the step size of the stepped stepping control strategy is set as 0.5°. It can also be set as other values according to the actual situation. The set of angular positions (unit: degree) of B1 is: B1_1, B1_2, B1_3... B1_n, where n is the number of angular positions of B1, B1_2 - B1_1 = 0.5°, B1_3 - B1_2 = 0.5°... B1_n - B1_n-1 = 0.5°. The combination of angular positions (unit: degree) of B2 is: -B1_1, -B1_2, -B1_3... -B1_n.

[0055] S11. When the angular position of B1 is B1_1, the angular position of B2 is -B1_1. With a step size of 0.1°, traverse the motion range of the angular position of A: -180° to +180°, and obtain multiple angular position combinations until all the angular positions of B1 and B2 are traversed, and obtain all the angular position combinations mentioned in S1, which are respectively recorded as the 1st angular position combination, the 2nd angular position combination, until the 3601*nth angular position combination, as shown in the serial numbers of the angular position combinations in Table 1 specifically.

[0056] S12. Calculate the inertia parameter combinations corresponding to each angular position combination. Specifically, it can be implemented by using 3D drawing software, such as solidworks, etc.

[0057] Among them, the inertia parameter group corresponding to the 1st angular position combination is recorded as the 1st inertia parameter combination, the inertia parameter group corresponding to the 2nd angular position combination is recorded as the 2nd inertia parameter combination, until the inertia parameter group corresponding to the 3601*nth angular position combination is recorded as the 3601*nth inertia parameter combination, as shown in Table 1 and Table 2 specifically.

[0058] Table 1:

[0059]

[0060]

[0061] Table 2:

[0062]

[0063]

[0064] S2. Calculate the PID parameter combinations corresponding to each inertia parameter combination. Specifically:

[0065] Under the constraint of the satellite attitude control accuracy index, the PID closed-loop control method is used, and the frequency-domain design method is adopted to traverse each angular position combination, and the PID parameter combination corresponding to each inertia parameter combination is calculated.

[0066] Among them, the satellite attitude control accuracy index refers to the requirements of the satellite mission for the amplitude-frequency characteristic and phase-frequency characteristic of attitude stable control, and can also be set according to the actual situation.

[0067] The specific implementation process of S2 is as follows:

[0068] As shown in Table 3, select 6 inertia parameters in the m-th inertia parameter combination: Jxx m 、Jyy m 、Jzz m 、Jxy m 、Jxz m and Jyz m , m = 1, 2, 3, 4... 3601×n. According to the satellite attitude control accuracy index, use the frequency-domain analysis and design method to determine the PID parameter combination for PID closed-loop control, and record the PID parameter combination corresponding to the m-th inertia parameter combination as the m-th PID parameter combination. Specifically:

[0069] Record the PID parameter combination corresponding to the 1st inertia parameter combination as the 1st PID parameter combination, record the PID parameter combination corresponding to the 2nd inertia parameter combination as the 2nd PID parameter combination, and so on until the PID parameter combination corresponding to the 3601*n-th inertia parameter combination is recorded as the 3601*n-th PID parameter combination.

[0070] The m-th PID parameter combination includes: Kp m 、Ki m and Kd m , Kp m represents the proportional coefficient corresponding to the m-th inertia parameter combination, Ki m represents the integral coefficient corresponding to the m-th inertia parameter combination, and Kd m represents the differential coefficient corresponding to the m-th inertia parameter combination. Traverse all inertia parameter combinations to obtain the PID parameter combination corresponding to each inertia parameter combination, as shown in Table 3.

[0071] Table 3:

[0072]

[0073]

[0074] According to Table 1, Table 2 and Table 3, obtain the corresponding relationship between the angular position combination and the PID parameter combination, as shown in Table 4.

[0075] Table 4:

[0076]

[0077]

[0078] S3. Fit the corresponding relationship between the angular position of each rotating mechanism and all PID parameter combinations according to each angular position combination of all rotating mechanisms in the solar panel drive device. This corresponding relationship is characterized by the following expression:

[0079]

[0080] Among them, Ang_A represents the angular position of A; y_kp m represents the proportionality coefficient calculated according to the angular position Ang_A of A when the angular position of B1 is B1_m and the angular position of B2 is -B1_m, where B1_m and -B1_m are obtained from the m-th angular position combination; y_ki m is the integral coefficient calculated according to the angular position Ang_A of A when the angular position of B1 is B1_m and the angular position of B2 is -B1_m; y_kd m is the differential coefficient calculated according to the angular position Ang_A of A when the angular position of B1 is B1_m and the angular position of B2 is -B1_m; a 4_m 、a 3_m 、a 2_m 、a 1_m 、a 0_m 、b 4_m 、b 3_m 、b 2_m 、b 1_m 、b 0_m 、c 4_m 、c 3_m 、c 2_m 、c 1_m and c 0_m are all coefficients.

[0081] a 4_m 、a 3_m 、a 2_m 、a 1_m 、a 0_m 、b 4_m 、b 3_m 、b 2_m 、b 1_m 、b 0_m 、c 4_m 、c 3_m 、c 2_m 、c 1_m and c 0_m The specific acquisition process is as follows:

[0082] ①a 4_m 、a 3_m 、a 2_m 、a 1_m and a 0_m The acquisition process is as follows:

[0083] When the angular position of B1 is B1_m and the angular position of B2 is -B1_m, the different values of Ang_A within the motion range and the proportional coefficients in the PID parameter combinations corresponding to different angular position combinations are read from Table 4, and the following can be obtained:

[0084]

[0085] Denoted as: Y_Kp_m=Mat A *W a_m ,in,

[0086] The least squares method is used for fitting, and we get: W a_m =(Mat A T *Mat A ) -1 *Mat A T *Y_Kp_m, according to W a_m =(Mat A T *Mat A ) -1 *Mat A T *Y_Kp_m, calculate a 4_m 、a 3_m 、a 2_m 、a 1_m and a 0_m .

[0087] ②b 4_m 、b 3_m 、b 2_m 、b 1_m and b 0_m The acquisition process is as follows:

[0088] When the angular position of B1 is B1_m and the angular position of B2 is -B1_m, the different values of Ang_A within the motion range and the integral coefficients in the PID parameter combinations corresponding to different angular position combinations are read from Table 4, and the following can be obtained:

[0089]

[0090] Denoted as: Y_Ki_m=Mat B*W b_m , where Mat B = Mat A .

[0091] Using the least squares method for fitting, we get: W b_m = (Mat B T * Mat B ) -1 * Mat B T * Y_Ki_m. According to W b_m = (Mat B T * Mat B ) -1 * Mat B T * Y_Ki_m, b 4_m , b 3_m , b 2_m , b 1_m and b 0_m .

[0092] ③ The acquisition process of c 4_m , c 3_m , c 2_m , c 1_m and c 0_m is as follows:

[0093] When the angular position of B1 is B1_m and the angular position of B2 is -B1_m, read different values of Ang_A within the motion range from Table 4, and the integral coefficients in the PID parameter combinations corresponding to different angular position combinations, we can get:

[0094]

[0095] Denoted as: Y_Kd_m = Mat C * W c_m , where Mat C = Mat A .

[0096] Using the least squares method for fitting, we get: W c_m = (Mat C T * Mat C ) -1 * Mat C T * Y_Kd_m. According to W c_m = (Mat C T * MatC ) -1 *Mat C T *Y_Kd_m, can calculate c 4_m 、c 3_m 、c 2_m 、c 1_m and c 0_m .

[0097] Traverse each angular position combination, calculate the closed-loop control PID parameter fitting coefficient of the X-axis of the satellite center of mass body coordinate system, the closed-loop control PID parameter fitting coefficient of the X-axis includes: the proportional coefficient fitting coefficient of the X-axis of the satellite center of mass body coordinate system, the proportional coefficient fitting coefficient of the X-axis of the satellite center of mass body coordinate system and the proportional coefficient fitting coefficient of the X-axis of the satellite center of mass body coordinate system. Specifically calculate the corresponding relationship between the angular position of B1 and the angular position of B2 of each angular position combination and the proportional coefficient fitting coefficient of the X-axis of the satellite center of mass body coordinate system, the proportional coefficient fitting coefficient of the X-axis of the satellite center of mass body coordinate system and the proportional coefficient fitting coefficient of the X-axis of the satellite center of mass body coordinate system. The proportional coefficient fitting coefficient of the X-axis of the satellite center of mass body coordinate system corresponding to the mth angular position combination is recorded as W ax_m , the integral coefficient fitting coefficient of the X-axis of the satellite mass center coordinate system corresponding to the mth angular position combination is recorded as W bx_m , the differential coefficient fitting coefficient of the X-axis of the satellite center of mass coordinate system corresponding to the mth angular position combination is recorded as W cx_m , as shown in Table 5.

[0098] Table 5:

[0099]

[0100] Traverse each angular position combination, calculate the closed-loop control PID parameter fitting coefficient of the Y axis of the satellite center of mass body coordinate system, the closed-loop control PID parameter fitting coefficient of the Y axis includes: the proportional coefficient fitting coefficient of the Y axis of the satellite center of mass body coordinate system, the proportional coefficient fitting coefficient of the Y axis of the satellite center of mass body coordinate system and the proportional coefficient fitting coefficient of the Y axis of the satellite center of mass body coordinate system. Specifically, calculate the corresponding relationship between the angular position of B1 and the angular position of B2 of each angular position combination and the proportional coefficient fitting coefficient of the Y axis of the satellite center of mass body coordinate system, the proportional coefficient fitting coefficient of the Y axis of the satellite center of mass body coordinate system and the proportional coefficient fitting coefficient of the Y axis of the satellite center of mass body coordinate system. The proportional coefficient fitting coefficient of the Y axis of the satellite center of mass body coordinate system corresponding to the mth angular position combination is recorded as W ay_m , the integral coefficient fitting coefficient of the Y axis of the satellite mass center body coordinate system corresponding to the mth angular position combination is recorded as W by_m, denote the fitting coefficient of the differential coefficient of the Y-axis of the satellite centroid body coordinate system corresponding to the m-th angular position combination as W cy_m , as shown in Table 6

[0101] Table 6:

[0102]

[0103] Traverse each angular position combination, calculate the fitting coefficients of the closed-loop control PID parameters of the Z-axis of the satellite centroid body coordinate system. The fitting coefficients of the closed-loop control PID parameters of the Z-axis include: the proportional coefficient fitting coefficient of the Z-axis of the satellite centroid body coordinate system, the integral coefficient fitting coefficient of the Z-axis of the satellite centroid body coordinate system, and the differential coefficient fitting coefficient of the Z-axis of the satellite centroid body coordinate system. Specifically, calculate the corresponding relationship between the angular positions of B1 and B2 and the proportional coefficient fitting coefficient of the Z-axis of the satellite centroid body coordinate system, the integral coefficient fitting coefficient of the Z-axis of the satellite centroid body coordinate system, and the differential coefficient fitting coefficient of the Z-axis of the satellite centroid body coordinate system for each angular position combination. Denote the proportional coefficient fitting coefficient of the Z-axis of the satellite centroid body coordinate system corresponding to the m-th angular position combination as W az_m , denote the integral coefficient fitting coefficient of the Z-axis of the satellite centroid body coordinate system corresponding to the m-th angular position combination as W bz_m , denote the differential coefficient fitting coefficient of the Z-axis of the satellite centroid body coordinate system corresponding to the m-th angular position combination as W cz_m , as shown in Table 7

[0104] Table 7:

[0105]

[0106] S4. Use the PID control law of the correlation correspondence relationship to calculate the expected control torque for the satellite to adjust from the current attitude to the final desired attitude, and control the actuator of the satellite to output the expected control torque. Specifically:

[0107] S40. Obtain the angular position Ang_Pos_A of A at the current moment, the angular position Ang_Pos_B1 of B1 at the current moment, and the angular position Ang_Pos_B2 of B1 at the current moment from the telemetry data of the solar panel drive device

[0108] S41. According to the angular position Ang_Pos_B1 of B1 at the current moment, traverse and search for B1_1, B1_2, B1_3, B1_4…, B1_n-2, B1_n-1, and B1_n to determine the position interval of Ang_Pos_B1, that is, B1_k ≤ Ang_Pos_B1 < B1_k+1, where k is a positive integer, and the value range is: 1 to n

[0109] S42. Obtain the closed-loop control PID parameter fitting coefficients of the X-axis of the satellite centroid body coordinate system at the current moment from Table 5: W ax_k 、W bx_k 、W cx_k . The PID parameters of the X-axis of the satellite centroid body coordinate system at the current moment include k px 、k ix and k dx , as follows:

[0110] k px = W ax_k (1)*(Ang_Pos_A) 4 + W ax_k (2)*(Ang_Pos_A) 3 + W ax_k (3)*(Ang_Pos_A) 2 + W ax_k (4)*Ang_Pos_A + W ax_k (5)

[0111] k ix = W bx_k (1)*(Ang_Pos_A) 4 + W bx_k (2)*(Ang_Pos_A) 3 + W bx_k (3)*(Ang_Pos_A) 2 + W bx_k (4)*Ang_Pos_A + W bx_k (5)

[0112] k dx = W cx_k (1)*Ang_Pos_A 4 + W cx_k (2)*Ang_Pos_A 3 + W cx_k (3)*Ang_Pos_A 2

[0113] + W cx_k (4)*Ang_Pos_A + W cx_k (5)

[0114] S43. Obtain the closed-loop control PID parameter fitting coefficients of the Y-axis of the satellite centroid body coordinate system at the current moment from Table 6: W ay_k 、W by_k 、W cy_k . The PID parameters of the Y-axis of the satellite centroid body coordinate system at the current moment include k py 、k iy and kdy , as follows:

[0115] k py = W ay_k (1)*(Ang_Pos_A) 4 + W ay_k (2)*(Ang_Pos_A) 3 + W ay_k (3)*(Ang_Pos_A) 2

[0116] + W ay_k (4)*Ang_Pos_A + W ay_k (5)

[0117] k iy = W by_k (1)*(Ang_Pos_A) 4 + W by_k (2)*(Ang_Pos_A) 3 + W by_k (3)*(Ang_Pos_A) 2

[0118] + W by_k (4)*Ang_Pos_A + W by_k (5)

[0119] k dy = W cy_k (1)*(Ang_Pos_A) 4 + W cy_k (2)*(Ang_Pos_A) 3 + W cy_k (3)*(Ang_Pos_A) 2

[0120] + W cy_k (4)*Ang_Pos_A + W cy_k (5)

[0121] S44. Obtain the closed-loop control PID parameter fitting coefficients of the Z-axis of the satellite centroid body coordinate system at the current moment: W az_k , W bz_k and W cz_k , the PID parameters of the Z-axis of the satellite centroid body coordinate system at the current moment include k pz , k iz and k dz , as follows:

[0122] k pz = W az_k (1)*(Ang_Pos_A)4 +W az_k (2)*(Ang_Pos_A) 3 +W az_k (3)*(Ang_Pos_A) 2

[0123] +W az_k (4)*Ang_Pos_A+W az_k (5)

[0124] k iz =W bz_k (1)*(Ang_Pos_A) 4 +W bz_k (2)*(Ang_Pos_A) 3 +W bz_k (3)*(Ang_Pos_A) 2

[0125] +W bz_k (4)*Ang_Pos_A+W bz_k (5)

[0126] k dz =W cz_k (1)*(Ang_Pos_A) 4 +W cz_k (2)*(Ang_Pos_A) 3 +W cz_k (3)*(Ang_Pos_A) 2

[0127] +W cz_k (4)*Ang_Pos_A+W cz_k (5)

[0128] S45. Based on the PID control theory (PID control law), calculate the desired control torque M for the satellite to adjust from the current attitude to the final desired attitude C , the desired control torque for the satellite to adjust from the current attitude to the final desired attitude includes: the desired control torques in the X-axis, Y-axis, and Z-axis directions of the satellite's body coordinate system at the center of mass that need to be generated by the actuator, M C =-k D ·Δω - k I *∫Δqdt - k P ·Δq, where Δω is the error angular velocity, Δq is the attitude angle error, k P is the proportional coefficient matrix, k I is the integral coefficient matrix, k D is the differential coefficient matrix, k D 、k P and kI are all diagonal matrices, where:

[0129]

[0130] Among them, the actuator can be a flywheel, or other actuators can be selected according to the actual situation.

[0131] Optionally, in the above technical solution, the satellite is a satellite with time-varying inertia. Specifically, the satellite is a low-earth-orbit communication satellite, a geostationary communication satellite, or other satellites carrying a two-axis SADA-driven solar panel.

[0132] The present invention adopts the classical PID control theory, uses the frequency-domain design method to design the parameters for the fixed inertia, and fits the parameters for different variables. It can not only meet the control index requirements of the amplitude-frequency characteristic and phase-frequency characteristic for the overall satellite attitude stability, but also adapt to the characteristic of the time-varying inertia of the solar panel of the overall satellite. It is applicable to the attitude closed-loop stability control of low-earth-orbit communication satellites carrying large-area solar panels and a small amount of propellant, with high closed-loop control robustness and strong engineering implementation operability.

[0133] In the above embodiments, although the steps are numbered S1, S2, etc., they are only specific embodiments given by the present invention. Those skilled in the art can adjust the execution order of S1, S2, etc. according to the actual situation, and this is also within the protection scope of the present invention. It can be understood that in some embodiments, it may include some or all of the above embodiments.

[0134] As Figure 3 shown, a satellite attitude control system 200 according to an embodiment of the present invention includes an inertia parameter combination acquisition module 200, a PID parameter combination calculation module 201, a fitting module 203, and a satellite attitude control module 204;

[0135] The inertia parameter combination acquisition module 201 is used to: acquire the inertia parameter combination of the satellite in the satellite centroid body coordinate system when all the rotating mechanisms in the solar panel drive device of the satellite are at each angular position combination;

[0136] The PID parameter combination calculation module 202 is used to: calculate the PID parameter combination corresponding to each inertia parameter combination;

[0137] The fitting module 203 is used to: fit the corresponding relationship between the angular position of each rotating mechanism and all the PID parameters according to each angular position combination of all the rotating mechanisms in the solar panel drive device and all the PID parameter combinations;

[0138] The satellite attitude control module 204 is used to: calculate the desired control torque for the satellite to adjust from the current attitude to the final desired attitude by using the PID control law with associated correspondence relationships, and control the actuator of the satellite to output the desired control torque.

[0139] Optionally, in the above technical solution, the PID parameter combination calculation module 201 is used to:

[0140] Under the constraint of the satellite attitude control accuracy index, by using the PID closed-loop control method and traversing each angular position combination by the frequency domain design method, calculate the PID parameter combination corresponding to each inertia parameter combination.

[0141] Optionally, in the above technical solution, the solar panel drive device includes three rotating mechanisms arranged in sequence. The rotation axis of the middle rotating mechanism is perpendicular to the rotation axes of the rotating mechanisms on both sides. The rotating mechanisms on both sides are respectively connected to a solar panel, and a sun sensor is arranged on each solar panel.

[0142] Optionally, in the above technical solution, the satellite is a satellite with time-varying inertia.

[0143] It should be noted that the beneficial effects of the satellite attitude control system 200 provided in the above embodiment are the same as those of the above satellite attitude control method, and will not be elaborated here. In addition, when the system provided in the above embodiment realizes its functions, only the division of the above function modules is used for illustration. In actual applications, the above functions can be allocated to different function modules according to needs, that is, the system can be divided into different function modules according to actual situations to complete all or part of the functions described above. In addition, the system provided in the above embodiment and the method embodiment belong to the same concept, and the specific implementation process can be seen in the method embodiment, which will not be elaborated here.

[0144] Among them, the satellite attitude control system of the present invention can be a computer program (including program codes) running in a computer device. For example, the satellite attitude control system of the present invention is an application software and can be used to execute the corresponding steps in the satellite attitude control method of the present invention.

[0145] In some embodiments, the satellite attitude control system of the present invention can be implemented in a combination of software and hardware. As an example, the satellite attitude control system of the present invention can be a processor in the form of a hardware decoding processor, which is programmed to execute the satellite attitude control method of the present invention. For example, the processor in the form of a hardware decoding processor can employ one or more application specific integrated circuits (ASICs), DSPs, programmable logic devices (PLDs), complex programmable logic devices (CPLDs), field programmable gate arrays (FPGAs), or other electronic components.

[0146] Among them, the modules involved in the embodiments of the present invention can be implemented in software or in hardware. Among them, the name of the module does not constitute a limitation to the module itself in some cases.

[0147] An electronic device according to an embodiment of the present invention includes a memory, a processor, and a computer program stored on the memory and executable on the processor. When the processor executes the computer program, it implements any one of the above satellite attitude control methods. That is to say, an electronic device according to an embodiment of the present invention may include, but is not limited to: a processor and a memory; the memory is used to store the computer program; the processor is used to execute the satellite attitude control method shown in any embodiment of the present invention by calling the computer program.

[0148] In an alternative embodiment, an electronic device is provided, as Figure 4 shown, Figure 4 the electronic device 4000 shown includes: a processor 4001 and a memory 4003. Among them, the processor 4001 and the memory 4003 are connected, such as connected through a bus 4002. Optionally, the electronic device 4000 may further include a transceiver 4004, and the transceiver 4004 can be used for data interaction between this electronic device and other electronic devices, such as data sending and / or data receiving, etc. It should be noted that in practical applications, the transceiver 4004 is not limited to one, and the structure of the electronic device 4000 does not constitute a limitation to the embodiments of the present invention.

[0149] The processor 4001 may be a CPU (Central Processing Unit), a general-purpose processor, a DSP (Digital Signal Processor), an ASIC (Application Specific Integrated Circuit), an FPGA (Field Programmable Gate Array), or other programmable logic devices, transistor logic devices, hardware components, or any combination thereof. It can implement or execute various exemplary logical blocks, modules, and circuits described in connection with the disclosure of the present invention. The processor 4001 may also be a combination that implements computing functions, such as a combination of one or more microprocessors, a combination of a DSP and a microprocessor, etc.

[0150] The bus 4002 may include a path for transmitting information between the above components. The bus 4002 may be a PCI (Peripheral Component Interconnect) bus, an EISA (Extended Industry Standard Architecture) bus, or the like. The bus 4002 may be divided into an address bus, a data bus, a control bus, etc. For ease of representation, Figure 4 only a thick line is used to represent the bus 4002 in the figure, but it does not mean that there is only one bus or one type of bus.

[0151] The memory 4003 may be a ROM (Read Only Memory) or other type of static storage device that can store static information and instructions, a RAM (Random Access Memory), or other type of dynamic storage device that can store information and instructions. It may also be an EEPROM (Electrically Erasable Programmable Read Only Memory), a CD-ROM (Compact Disc Read Only Memory), or other optical disc storage, optical disc storage (including compact discs, laser discs, optical discs, digital versatile discs, Blu-ray discs, etc.), magnetic storage media, or other magnetic storage devices, or any other medium that can be used to carry or store the desired program code in the form of instructions or data structures and can be accessed by a computer, but is not limited thereto.

[0152] The memory 4003 is used to store the application program code (computer program) for executing the solution of the present invention, and is controlled by the processor 4001 for execution. The processor 4001 is used to execute the application program code stored in the memory 4003 to implement the content shown in the foregoing method embodiments.

[0153] Among them, the electronic device may also be a terminal device, and the terminal device may be any terminal device that can install an application and access a web page through the application, including at least one of a smart phone, a tablet computer, a notebook computer, a desktop computer, a smart speaker, a smart watch, a smart TV, and a smart vehicle-mounted device.

[0154] It should be noted that Figure 4 The electronic device shown is only an example and should not impose any restrictions on the functions and usage scope of the embodiments of the present invention.

[0155] A computer-readable storage medium according to an embodiment of the present invention stores a computer program thereon, and when the computer program is executed by a processor, the above-mentioned satellite attitude control method is implemented.

[0156] Optionally, the computer-readable storage medium may be a read-only memory (ROM), a random access memory (RAM), a compact disc read-only memory (CD-ROM), a magnetic tape, a floppy disk, an optical data storage device, etc.

[0157] In an exemplary embodiment, a computer program product or a computer program is further provided. The computer program product or the computer program includes computer instructions, and the computer instructions are stored in a computer-readable storage medium. The processor of the electronic device reads the computer instructions from the computer-readable storage medium, and the processor executes the computer instructions, so that the electronic device executes the above-mentioned satellite attitude control method.

[0158] Computer program code for performing the operations of the present invention may be written in one or more programming languages or combinations thereof. The above-mentioned programming languages include object-oriented programming languages such as Java, Smalltalk, C++, and also include conventional procedural programming languages such as the "C" language or similar programming languages. The program code may be executed entirely on the user's computer, partially on the user's computer, executed as a stand-alone software package, partially on the user's computer and partially on a remote computer, or entirely on a remote computer or server. In the case of a remote computer, the remote computer may be connected to the user's computer through any type of network, including a local area network (LAN) or a wide area network (WAN), or it may be connected to an external computer (e.g., through the Internet using an Internet service provider).

[0159] It should be understood that the flowcharts and block diagrams in the accompanying drawings illustrate the possible architectures, functions, and operations of the methods and computer program products according to various embodiments of the present invention. In this regard, each block in the flowchart or block diagram may represent a module, a program segment, or a part of code that contains one or more executable instructions for implementing the specified logical function. It should also be noted that in some alternative implementations, the functions marked in the blocks may occur in a different order than that marked in the accompanying drawings. For example, two consecutive blocks shown may actually be executed substantially in parallel, and they may sometimes be executed in the reverse order, depending on the functions involved. It should also be noted that each block in the block diagram and / or flowchart, and the combination of blocks in the block diagram and / or flowchart, may be implemented by a dedicated hardware-based system for performing the specified functions or operations, or may be implemented by a combination of dedicated hardware and computer instructions.

[0160] The computer-readable storage medium provided by the embodiments of the present invention may be, but is not limited to, an electrical, magnetic, optical, electromagnetic, infrared, or semiconductor system, apparatus, or device, or any combination of the above. More specific examples of the computer-readable storage medium may include, but are not limited to: an electrical connection having one or more wires, a portable computer disk, a hard disk, a random access memory (RAM), a read-only memory (ROM), an erasable programmable read-only memory (EPROM or flash memory), an optical fiber, a portable compact disk read-only memory (CD-ROM), an optical storage device, a magnetic storage device, or any suitable combination of the above. In the present invention, the computer-readable storage medium may be any tangible medium that contains or stores a program that can be used by or in conjunction with an instruction execution system, apparatus, or device.

[0161] The above computer-readable storage medium carries one or more programs, which, when executed by the electronic device, cause the electronic device to execute the method shown in the above embodiments.

[0162] The above description is only a preferred embodiment of the present invention and an explanation of the applied technical principles. Those skilled in the art should understand that the scope of disclosure involved in the present invention is not limited to the technical solution formed by the specific combination of the above technical features, but should also cover other technical solutions formed by any combination of the above technical features or their equivalent features without departing from the above disclosed concept. For example, a technical solution formed by mutually replacing the above features with technical features (but not limited to) having similar functions disclosed in the present invention.

[0163] It should be noted that the terms "first", "second", etc. in the specification and claims of this application are used to distinguish similar objects, and represent a limitation on a specific order or sequence. Under appropriate circumstances, the order of use of similar objects can be interchanged so that the embodiments of this application described here can be implemented in an order other than the illustrated or described order.

[0164] Those skilled in the art know that the present invention can be implemented as a system, method or computer program product. Therefore, the present invention can be specifically implemented in the following forms, that is, it can be completely hardware, can also be completely software (including firmware, resident software, microcode, etc.), or can also be in the form of a combination of hardware and software, which is generally referred to as "circuit", "module" or "system" in this article. In addition, in some embodiments, the present invention can also be implemented in the form of a computer program product in one or more computer-readable media, which contains computer-readable program code.

[0165] Although the embodiments of the present invention have been shown and described above, it can be understood that the above embodiments are exemplary and should not be construed as limiting the present invention. Those of ordinary skill in the art can make changes, modifications, substitutions and variations to the above embodiments within the scope of the present invention.

Claims

1. A satellite attitude control method, characterized in that, Including: Obtain the combination of inertia parameters of the satellite in the satellite body coordinate system with the satellite's center of mass when all the rotating mechanisms in the satellite's solar panel drive device are at each angular position combination; Calculate the corresponding PID parameter combination for each inertia parameter combination; According to each angular position combination of all the rotating mechanisms in the solar panel drive device and all the PID parameter combinations, fit the corresponding relationship between the angular position of each rotating mechanism and all the PID parameters; Using the PID control law associated with the corresponding relationship, calculate the desired control torque for the satellite to adjust from the current attitude to the final desired attitude, and control the actuator of the satellite to output the desired control torque.

2. The satellite attitude control method according to claim 1, characterized in that Calculating the corresponding PID parameter combination for each inertia parameter combination includes: Under the constraint of the satellite attitude control accuracy index, use the PID closed-loop control method and adopt the frequency-domain design method to traverse each angular position combination to calculate the corresponding PID parameter combination for each inertia parameter combination.

3. A satellite attitude control method according to claim 1 or 2, characterized in that, The solar panel drive device includes three rotating mechanisms arranged in sequence. The rotation axis of the rotating mechanism in the middle is perpendicular to the rotation axes of the rotating mechanisms on both sides. The rotating mechanisms on both sides are respectively connected to a solar panel, and a sun sensor is arranged on each solar panel.

4. A satellite attitude control method according to claim 1 or 2, characterized in that, The satellite is a satellite with time-varying inertia.

5. A satellite attitude control system, characterized in that, Including an inertia parameter combination acquisition module, a PID parameter combination calculation module, a fitting module, and a satellite attitude control module; The inertia parameter combination acquisition module is used to: obtain the combination of inertia parameters of the satellite in the satellite body coordinate system with the satellite's center of mass when all the rotating mechanisms in the satellite's solar panel drive device are at each angular position combination; The PID parameter combination calculation module is used to: calculate the corresponding PID parameter combination for each inertia parameter combination; The fitting module is used to: according to each angular position combination of all the rotating mechanisms in the solar panel drive device and all the PID parameter combinations, fit the corresponding relationship between the angular position of each rotating mechanism and all the PID parameters; The satellite attitude control module is used to: using the PID control law associated with the corresponding relationship, calculate the desired control torque for the satellite to adjust from the current attitude to the final desired attitude, and control the actuator of the satellite to output the desired control torque.

6. A satellite attitude control system according to claim 5, characterized in that, The PID parameter combination calculation module is used to: Under the constraint of the satellite attitude control accuracy index, use the PID closed-loop control method and adopt the frequency-domain design method to traverse each angular position combination to calculate the corresponding PID parameter combination for each inertia parameter combination.

7. A satellite attitude control system according to claim 5 or 6, characterized in that, The solar panel drive device includes three rotating mechanisms arranged in sequence. The rotation axis of the rotating mechanism in the middle is perpendicular to the rotation axes of the rotating mechanisms on both sides. The rotating mechanisms on both sides are respectively connected to a solar panel, and a sun sensor is arranged on each solar panel.

8. A satellite attitude control system according to claim 5 or 6, characterized in that, The satellite is a satellite with time-varying inertia.

9. An electronic device, characterized in that, Including a memory, a processor, and a computer program stored on the memory and executable on the processor. When the processor executes the computer program, it implements the satellite attitude control method according to any one of claims 1 to 4.

10. A computer-readable storage medium, characterized in that, A computer program is stored on the computer-readable storage medium, and when the computer program is executed by a processor, it implements a satellite attitude control method according to any one of claims 1 to 4.

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