Satellite attitude control method and system, electronic equipment and storage medium

Abstract

The invention discloses a satellite attitude control method and system, electronic equipment and a storage medium, and relates to the technical field of satellite attitude control, and the method comprises the steps: obtaining an inertia parameter combination of a satellite under a satellite centroid body coordinate system when all rotating mechanisms in a solar panel driving device of the satellite are combined at each angular position; calculating a PID parameter combination corresponding to each inertia parameter combination; fitting a corresponding relation between the angular position of each rotating mechanism and all PID parameters; and calculating an expected control moment for adjusting the satellite from the current attitude to the final expected attitude by using a PID control law associated with the corresponding relationship, and controlling an execution mechanism of the satellite to output the expected control moment. The method not only can meet the control index requirements of the whole satellite attitude stability on the amplitude-frequency characteristic and the phase-frequency characteristic, but also can adapt to the inertia time-varying characteristic of the satellite caused by the rotation of the sailboard, the closed-loop control robustness is relatively high, and the engineering implementation operability is strong.

Description

Background Art

[0002] Satellite attitude control includes attitude determination, attitude stabilization control and attitude maneuvering control. Attitude determination is to study the orientation or pointing of the satellite relative to a reference benchmark, and then obtain the attitude angle parameters. Its accuracy depends on the accuracy of the attitude sensor and the attitude determination algorithm. Attitude stabilization control is to keep the satellite's attitude in the expected specified direction and specified value, and attitude maneuvering control is the reorientation process of the satellite from one attitude to another.

[0003] After entering orbit, the communication satellite first undergoes attitude control processes such as rate damping, sailboard deployment, and attitude capture, and then operates stably in the whole-satellite attitude earth-orientation mode to carry out communication payload work and provide communication business services.

[0004] Low-orbit communication satellites often carry large-area dual-wing dual-axis solar panels and operate in inclined or polar orbits. During the satellite's in-orbit operation, in order to ensure the on-board energy supply, the solar panels are driven to rotate through a two-dimensional drive mechanism to achieve the normal direction of the panels facing the sun. The solar array drive assembly (SADA) is a complex electromechanical device that drives the solar array to rotate by receiving instructions and outputting control torque. The dual-degree-of-freedom solar array drive mechanism is equipped with detection devices and drive devices in the directions of the two degrees of freedom. The second degree of freedom is controlled by an independent motor, which can realize the function of maneuvering separately from the satellite body.

[0005] As a classic attitude control method, PID control is an accurate and advanced control law. Based on the analysis and design of automatic control principles, the PID controller can consider factors such as system dynamic characteristics and bandwidth when designing, and reflect the robust performance of the system through amplitude margin and phase margin. Therefore, it is still used by most three-axis stabilized spacecraft.

[0006] A communication satellite carrying a dual-wing dual-axis large-area solar sail is in the whole-satellite attitude earth orientation mode. The dual-axis SADA moves according to the predetermined motion law. The movement of the sailboard causes the dynamic characteristics of the whole satellite to exhibit complex attitude dynamic characteristics, time-varying inertia, aerodynamic torque, gravity gradient torque, residual magnetic torque and other environmental interference torques, which vary greatly within the envelope of the sailboard movement, and have a great influence on the stability control accuracy and control robustness of the whole satellite attitude.

[0007] When a communication satellite carrying a large-area dual-wing dual-axis solar sail uses a PID controller, in order to ensure attitude stability when the movement of the sail causes a large change in the satellite's inertia, interference torque, etc., the attitude stability control engineering design is quite difficult. There is currently little research and engineering implementation on how to design and select PID parameters. Summary of the invention

[0008] The technical problem to be solved by the present invention is to address the deficiencies of the prior art and specifically provide a satellite attitude control method, system, electronic device and storage medium, 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] Obtaining the inertia parameter combination of the satellite in the satellite center of mass body coordinate system when all rotating mechanisms of the satellite's solar sail driving device are combined at each angular position;

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

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

[0013] The PID control law of the associated correspondence is used to calculate the expected control torque for adjusting the satellite from the current attitude to the final expected attitude, and the actuator of the satellite is controlled to output the expected 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 characteristics and phase-frequency characteristics of the attitude stability of the whole satellite, but also adapt to the time-varying characteristics of the satellite's inertia caused by the rotation of the sailboard. It is suitable for attitude closed-loop stability control of low-orbit communication satellites carrying large-area solar panels and a small amount of propellant. The closed-loop control has high robustness and the engineering implementation is highly operational.

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

[0017] Furthermore, the PID parameter combination corresponding to each inertia parameter combination is calculated, including:

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

[0019] Furthermore, the solar panel driving device includes three rotating mechanisms arranged in sequence, the rotating axis of the rotating mechanism in the middle is perpendicular to the rotating 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] Furthermore, the satellite is a satellite with time-varying inertia.

[0021] 2) In the 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 center of mass body coordinate system when all the rotating mechanisms in the satellite's solar sail driving device are combined at each angular position;

[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 PID parameters according to each angular position combination of all rotating mechanisms in the solar sail panel driving device and all PID parameter combinations;

[0026] The satellite attitude control module is used to calculate the desired control torque for adjusting the satellite from the current attitude to the final desired attitude using the PID control law of the associated correspondence, and control the satellite's actuator 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] Furthermore, the PID parameter combination calculation module is used to:

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

[0030] Furthermore, the solar panel driving device includes three rotating mechanisms arranged in sequence, the rotating axis of the rotating mechanism in the middle is perpendicular to the rotating 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.

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

[0032] 3) In a third aspect, the present invention further provides an electronic device, comprising a processor, the processor being coupled to a memory, the memory storing at least one computer program, the at least one computer program being loaded and executed by the processor, so that the electronic device implements any one of the above-mentioned satellite attitude control methods.

[0033] 4) In a fourth aspect, the present invention further provides a computer-readable storage medium having a computer program stored thereon, and when the computer program is executed by a processor, any of the above-mentioned satellite attitude control methods is implemented.

[0034] It should be noted that the beneficial effects achieved by the technical solutions of the second to fourth aspects of the present invention and the corresponding possible implementation methods can be found in the above-mentioned technical effects of the first aspect and its corresponding possible implementation methods, and will not be repeated 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 briefly introduces the drawings required for describing the embodiments of the present invention:

[0036] Figure 1 A schematic diagram of a flow chart of a satellite attitude control method according to an embodiment of the present invention;

[0037] Figure 2 is a schematic diagram of the structure of a solar sail panel driving device;

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

[0039] Figure 4 The figure is a schematic diagram of the structure of an electronic device according to an embodiment of the present invention. DETAILED DESCRIPTION

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

[0041] The technical solution of the present invention and how the technical solution of the present invention solves the above-mentioned technical problems are described in detail below with specific embodiments. The following specific embodiments can be combined with each other, and the same or similar concepts or processes may not be repeated in some embodiments. The embodiments of the present invention will be described below in conjunction with the accompanying drawings.

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

[0043] S1. Obtaining the inertia parameter combination of the satellite in the satellite center of mass body coordinate system when all rotating mechanisms of the satellite's solar sail driving device are combined at each angular position;

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

[0045] According to the unfolding state of the satellite's solar sail panels, a rotation mechanism is configured for each solar sail panel. The satellite includes two solar sail panels, which are respectively denoted as the first solar sail panel and the second solar sail panel. The rotation mechanism configured for the first solar sail panel is denoted as B1, and the rotation mechanism configured for the second solar sail panel is denoted as B2. A rotation mechanism is set between B1 and B2, denoted as A. B1, A and B2 are arranged in sequence. Figure 2 The cube in is used to represent the satellite body. The satellite body is: the remaining part of the satellite except the solar panels and the structure connecting the solar panels. The satellite body coordinate system is established with the center of mass of the satellite body as the origin: OXYZ. When both solar panels are deployed, the normal of the solar panel battery is consistent with the negative direction of the Z axis of the satellite body coordinate system. The position of B1 at this time is defined as the zero position of B1, the position of B2 at this time is defined as the zero position of B2, and the position of A at this time is defined as the zero position of A. When B1, B2 and A are all in the corresponding zero positions, the positive direction of the rotation axis of B1 is consistent with the negative direction of the X axis of the satellite 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 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 body coordinate system.

[0046] B1, B2 and A all use stepper motors inside. B1 is used to drive the first solar sail panel to rotate around the rotation axis of B1 through the stepper motor. B2 is used to drive the second solar sail panel to rotate around the rotation axis of B2 through the stepper motor. A is used to drive the first solar sail panel and the second solar sail panel to rotate around the rotation axis of A at the same time through the stepper motor. In this process, the right-hand rule is used to determine the rotation direction of the first solar sail panel and the second solar sail panel.

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

[0048] B1, B2 and A all support speed priority motion mode, position priority motion mode and position hold mode. The speed priority motion mode is: set the motion mechanism's angular velocity to move according to the set angular velocity magnitude and direction; the position priority motion mode is: set the motion position and motion speed of the motion mechanism at the same time, 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 determined by the rotating mechanism according to the shortest path principle; the position hold mode is for the driving mechanism to keep its current position.

[0049] Among them, A can rotate 360° continuously, B1 and B2 can move within the set range of motion, and the range of motion parameter setting is determined comprehensively based on the Beta angle and the occlusion of the star. Mechanical limit mechanisms are designed for B1 and B2 outside the comprehensively determined range of motion.

[0050] A solar sensor is arranged on each side of the positive direction of the normal line of the first solar panel and the second solar panel. The solar sensor is in a pyramid configuration. The solar sensor arranged on the first solar panel is recorded as the first solar sensor, and the solar sensor arranged on the second solar panel is recorded as the second solar sensor. The layout positions are as follows: Figure 3 The field of view of the first solar sensor and the second solar sensor is -90° to +90°. The first solar sensor and the second solar sensor collect solar vector information in the solar sensor measurement coordinate system set by themselves. The solar sensor measurement coordinate system can also be reset, and the electric propulsion thruster is configured in the negative direction of the X-axis of the satellite mass center body coordinate system.

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

[0052] Among them, each angular position combination corresponds to an inertia parameter combination, one 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, one inertia parameter combination includes 6 inertia parameters, specifically: Jxx, Jyy, Jzz, Jxy, Jxz and Jyz, Jxx represents: the moment of inertia of the star body around the X-axis of the center of mass body coordinate system, Jyy represents the moment of inertia of the star body around the Y-axis of the center of mass body coordinate system, Jzz represents the moment of inertia of the star body around the Z-axis of the center of mass body coordinate system, Jxy represents the inertia product of the star body about the OXY plane of the center of mass body coordinate system, Jxz represents the inertia product of the star body about the OXZ plane of the center of mass body coordinate system, Jyz represents the inertia product of the star body about the OYZ plane of the center of mass body coordinate system, and 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 movement range of the angular position of A is set to: -180° to +180°. B1 and B2 both adopt a step-by-step control strategy, and the step size of the step-by-step control strategy is set to 0.5°. It can also be set to other values ​​according to actual conditions. The set of angular positions of B1 (unit: degree) 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°, and the angular position (unit: degree) combination 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°, the motion range of the angular position of A is traversed: -180° to +180°, and multiple angular position combinations are obtained, until all angular positions of B1 and B2 are traversed, and all angular position combinations mentioned in S1 are obtained, which are respectively recorded as the first angular position combination, the second angular position combination, until the 3601*nth angular position combination. For details, see the serial numbers of the angular position combinations in Table 1.

[0056] S12, calculating the inertia parameter combination corresponding to each angular position combination, which can be specifically implemented using three-dimensional drawing software, such as SolidWorks.

[0057] Among them, the inertia parameter group corresponding to the first angular position combination is recorded as the first inertia parameter combination, the inertia parameter group corresponding to the second angular position combination is recorded as the second inertia parameter combination, until the inertia parameter group corresponding to the 3601*n angular position combination is recorded as the 3601*n inertia parameter combination, see Table 1 and Table 2 for details.

[0058] Table 1:

[0059]

[0060]

[0061] Table 2:

[0062]

[0063]

[0064] S2. Calculate the PID parameter combination corresponding to each inertia parameter combination, specifically:

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

[0066] Among them, the satellite attitude control accuracy index refers to: the satellite mission's requirements for the amplitude-frequency characteristics and phase-frequency characteristics of attitude stability control, which can also be set according to actual conditions.

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

[0068] As shown in Table 3, select 6 inertia parameters in the mth 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, the PID parameter combination for PID closed-loop control is determined using the frequency domain analysis design method, and the PID parameter combination corresponding to the mth inertia parameter combination is recorded as the mth PID parameter combination, specifically:

[0069] The PID parameter combination corresponding to the first inertia parameter combination is recorded as the first PID parameter combination, the PID parameter combination corresponding to the second inertia parameter combination is recorded as the second PID parameter combination, until the PID parameter combination corresponding to the 3601*n inertia parameter combination is recorded as the 3601*n PID parameter combination.

[0070] The mth PID parameter combination includes: Kp m 、Ki m and Kd m , Kp m Represents the proportional coefficient corresponding to the mth inertia parameter combination, Ki m Indicates the integral coefficient corresponding to the mth inertia parameter combination, Kd m Represents the differential coefficient corresponding to the mth inertia parameter combination. All inertia parameter combinations are traversed 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, the corresponding relationship between the angular position combination and the PID parameter combination is obtained, as shown in Table 4.

[0075] Table 4:

[0076]

[0077]

[0078] S3. According to each angular position combination of all rotating mechanisms in the solar sail panel driving device and all PID parameter combinations, the corresponding relationship between the angular position of each rotating mechanism and all PID parameters is fitted. The corresponding relationship is characterized by the following expression:

[0079]

[0080] Where 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 mth 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 All are 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 ,in, Mat B =Mat A .

[0091] The least squares method is used for fitting, and 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 can be calculated to get b 4_m 、b 3_m 、b 2_m 、b 1_m and b 0_m .

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

[0093] 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:

[0094]

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

[0096] The least squares method is used for fitting, and 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, the differential coefficient fitting coefficient of the Y axis of the satellite center of mass coordinate system corresponding to the mth angular position combination is recorded as W cy_m , as shown in Table 6.

[0101] Table 6:

[0102]

[0103] Traverse each angular position combination, calculate the closed-loop control PID parameter fitting coefficient of the Z axis of the satellite center of mass body coordinate system, the closed-loop control PID parameter fitting coefficient of the Z axis includes: the proportional coefficient fitting coefficient of the Z axis of the satellite center of mass body coordinate system, the proportional coefficient fitting coefficient of the Z axis of the satellite center of mass body coordinate system and the proportional coefficient fitting coefficient of the Z 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 Z axis of the satellite center of mass body coordinate system, the proportional coefficient fitting coefficient of the Z axis of the satellite center of mass body coordinate system and the proportional coefficient fitting coefficient of the Z axis of the satellite center of mass body coordinate system. The proportional coefficient fitting coefficient of the Z axis of the satellite center of mass body coordinate system corresponding to the mth angular position combination is recorded as W az_m , the integral coefficient fitting coefficient of the Z axis of the satellite mass center body coordinate system corresponding to the mth angular position combination is recorded as W bz_m , the differential coefficient fitting coefficient of the Z axis of the satellite center of mass coordinate system corresponding to the mth angular position combination is recorded as W cz_m , as shown in Table 7.

[0104] Table 7:

[0105]

[0106] S4. Using the PID control law of the associated correspondence, calculate the desired control torque for adjusting the satellite from the current attitude to the final desired attitude, and control the satellite's actuator to output the desired control torque. Specifically:

[0107] S40, obtaining 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 sail panel driving device;

[0108] S41. According to the angular position Ang_Pos_B1 of B1 at the current moment, traverse and search 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, k is a positive integer, and the value range is: 1 to n.

[0109] S42, obtain the closed-loop control PID parameter fitting coefficient of the X-axis of the satellite mass center 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 mass center 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 coefficient of the Y axis of the satellite mass center 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 mass center 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 coefficient of the Z axis of the satellite mass center body coordinate system at the current moment from Table 7: W az_k , W bz_k and W cz_k The PID parameters of the Z axis of the satellite center of mass 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 PID control theory (PID control law), calculate the desired control torque M of the satellite from the current attitude to the final desired attitude C The desired control torque for adjusting the satellite from the current attitude to the final desired attitude includes: the desired control torque in the X-axis, Y-axis and Z-axis directions of the satellite center of mass coordinate system that needs to be generated by the actuator, M C =-k D ·Δω-k I *∫Δqdt-k P ·Δq, Δω is the error angular velocity, Δq is the attitude angle error, k P is the scale 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] The actuator may be a flywheel, or other actuators may be selected according to actual conditions.

[0131] Optionally, in the above technical solution, the satellite is a satellite with time-varying inertia, specifically, the satellite is a low-orbit communication satellite, a high-orbit communication satellite or other satellite carrying a dual-axis SADA-driven solar panel.

[0132] The present invention adopts the classical PID control theory and the frequency domain design method to perform parameter design when the inertia is fixed, and fits the parameters when the inertia is different. It can not only meet the control index requirements of the amplitude-frequency characteristic and the phase-frequency characteristic of the attitude stability of the whole satellite, but also adapt to the time-varying characteristics of the rotational inertia of the sailboard of the whole satellite. It is suitable for the attitude closed-loop stability control of low-orbit communication satellites carrying large-area solar sailboards and a small amount of propellant. The closed-loop control has high robustness and strong operability in engineering implementation.

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

[0134] like Figure 3 As 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 mass center body coordinate system when all the rotating mechanisms in the satellite's solar sail panel driving device are combined at each angular position;

[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 PID parameters according to each angular position combination of all rotating mechanisms in the solar sail panel driving device and all PID parameter combinations;

[0138] The satellite attitude control module 204 is used to calculate the desired control torque for adjusting the satellite from the current attitude to the final desired attitude by using the PID control law of the associated correspondence, and control the satellite's actuator 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 satellite attitude control accuracy, the PID closed-loop control method is used, and the frequency domain design method is adopted to traverse each angular position combination and calculate the PID parameter combination corresponding to each inertia parameter combination.

[0141] Optionally, in the above technical solution, the solar panel driving device includes three rotating mechanisms arranged in sequence, the rotating axis of the rotating mechanism in the middle is perpendicular to the rotating axes of the rotating mechanisms on both sides, the rotating mechanisms on both sides are respectively connected to a solar panel, and a solar 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 a satellite attitude control system 200 provided in the above embodiment are the same as the beneficial effects of the above satellite attitude control method, which will not be described in detail here. In addition, when the system provided in the above embodiment realizes its functions, only the division of the above functional modules is used as an example. In practical applications, the above functions can be assigned to different functional modules as needed, that is, the system can be divided into different functional modules according to actual conditions to complete all or part of the functions described above. In addition, the system and method embodiments provided in the above embodiment belong to the same concept, and the specific implementation process is detailed in the method embodiment, which will not be described in detail here.

[0144] Among them, the satellite attitude control system of the present invention can be a computer program (including program code) running in a computer device. For example, the satellite attitude control system of the present invention is an application software that 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 by 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 adopt 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] The modules involved in the embodiments of the present invention may be implemented in software or hardware. The name of a module does not limit 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 in the memory and executable on the processor. When the processor executes the computer program, any one of the above-mentioned satellite attitude control methods is implemented. 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, such as Figure 4 As shown, Figure 4 The electronic device 4000 shown includes: a processor 4001 and a memory 4003. The processor 4001 and the memory 4003 are connected, such as through a bus 4002. Optionally, the electronic device 4000 may also include a transceiver 4004, which may be used for data interaction between the electronic device and other electronic devices, such as data transmission and / or data reception. It should be noted that in actual applications, the transceiver 4004 is not limited to one, and the structure of the electronic device 4000 does not constitute a limitation on the embodiments of the present invention.

[0149] 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 may implement or execute various exemplary logic blocks, modules and circuits described in conjunction with the disclosure of the present invention. 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 to transmit information between the above components. The bus 4002 may be a PCI (Peripheral Component Interconnect) bus or an EISA (Extended Industry Standard Architecture) bus. The bus 4002 may be divided into an address bus, a data bus, a control bus, etc. For ease of representation, Figure 4 In the figure, only one thick line is used to represent the bus 4002, but this does not mean that there is only one bus or one type of bus.

[0151] The memory 4003 can be a ROM (Read Only Memory) or other types of static storage devices that can store static information and instructions, a RAM (Random Access Memory) or other types of dynamic storage devices that can store information and instructions, or an EEPROM (Electrically Erasable Programmable Read Only Memory), a CD-ROM (Compact Disc Read Only Memory) or other optical disk storage, optical disk storage (including compressed optical disk, laser disk, optical disk, digital versatile disk, Blu-ray disk, etc.), a magnetic disk storage medium or other magnetic storage device, 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 to this.

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

[0153] Among them, the electronic device can also be a terminal device, and the terminal device can be any terminal device that can install applications and access web pages through applications, including at least one of a smart phone, a tablet computer, a laptop computer, a desktop computer, a smart speaker, a smart watch, a smart TV, and a smart car device.

[0154] It should be noted that Figure 4 The electronic device shown is only an example and should not bring any limitation to the functions and scope of use 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, and when the computer program is executed by a processor, any one of the above-mentioned satellite attitude control methods is implemented.

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

[0157] In an exemplary embodiment, a computer program product or a computer program is also provided, the computer program product or the computer program comprising computer instructions, the computer instructions being stored in a computer-readable storage medium. A processor of an electronic device reads the computer instructions from the computer-readable storage medium, and the processor executes the computer instructions, so that the electronic device performs any of the above-mentioned satellite attitude control methods.

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

[0159] It should be understood that the flow charts and block diagrams in the accompanying drawings illustrate the possible implementation architecture, functions and operations of the method and computer program product according to various embodiments of the present invention. In this regard, each box in the flow chart or block diagram can represent a module, a program segment, or a part of a code, and the module, the program segment, or a part of the code contains one or more executable instructions for realizing the specified logical function. It should also be noted that in some alternative implementations, the functions marked in the box can also occur in a sequence different from that marked in the accompanying drawings. For example, two boxes represented in succession can actually be executed substantially in parallel, and they can sometimes be executed in the opposite order, depending on the functions involved. It should also be noted that each box in the block diagram and / or flow chart, and the combination of the boxes in the block diagram and / or flow chart can be implemented with a dedicated hardware-based system that performs a specified function or operation, or can be implemented with a combination of dedicated hardware and computer instructions.

[0160] The computer-readable storage medium provided by the embodiment of the present invention may be, but is not limited to, an electrical, magnetic, optical, electromagnetic, infrared, or semiconductor system, device or component, or any combination thereof. More specific examples of computer-readable storage media may include, but are not limited to: an electrical connection with 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 thereof. In the present invention, a computer-readable storage medium may be any tangible medium containing or storing a program that may be used by or in conjunction with an instruction execution system, device, or component.

[0161] The computer-readable storage medium carries one or more programs. When the one or more programs are executed by the electronic device, the electronic device executes the method shown in the above embodiment.

[0162] The above description is only a preferred embodiment of the present invention and an explanation of the technical principles used. Those skilled in the art should understand that the disclosure scope involved in the present invention is not limited to the technical solution formed by a 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, the above features are replaced with the technical features with similar functions disclosed in the present invention (but not limited to) by each other.

[0163] It should be noted that the terms "first", "second", etc. in the specification and claims of the present application are used to distinguish similar objects and represent the definition of a specific order or sequence. The order of use of similar objects can be interchanged where appropriate, so that the embodiments of the present application described herein can be implemented in an order other than the order shown or described.

[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, namely: it can be complete hardware, it can be complete software (including firmware, resident software, microcode, etc.), or it can be a combination of hardware and software, which is generally referred to as "circuit", "module" or "system" herein. 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, and the computer-readable medium contains computer-readable program code.

[0165] Although the embodiments of the present invention have been shown and described above, it is to be understood that the above embodiments are exemplary and are not to be construed as limitations of the present invention. A person skilled in the art may change, modify, replace and vary the above embodiments within the scope of the present invention.

Claims

1. A satellite attitude control method, characterized in that: include: Obtaining the inertia parameter combination of the satellite in the satellite center of mass body coordinate system when all rotating mechanisms of the satellite's solar sail driving device are combined at each angular position; Calculate the PID parameter combination corresponding to each inertia parameter combination; According to each angular position combination of all rotating mechanisms in the solar sail panel driving device and all PID parameter combinations, the corresponding relationship between the angular position of each rotating mechanism and all PID parameters is fitted; The desired control torque for adjusting the satellite from the current attitude to the final desired attitude is calculated by using the PID control law associated with the corresponding relationship, and the actuator of the satellite is controlled to output the desired control torque.

2. A satellite attitude control method according to claim 1, characterized in that: Calculate the PID parameter combination corresponding to each inertia parameter combination, including: Under the constraint of satellite attitude control accuracy, the PID closed-loop control method is used, and the frequency domain design method is adopted to traverse each angular position combination and calculate the PID parameter combination corresponding to each inertia parameter combination.

3. A satellite attitude control method according to claim 1 or 2, characterized in that: The solar sail panel driving device includes three rotating mechanisms arranged in sequence, the rotating axis of the rotating mechanism in the middle is perpendicular to the rotating axes of the rotating mechanisms on both sides, the rotating mechanisms on both sides are respectively connected to a solar sail panel, and a sun sensor is arranged on each solar sail 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: It includes 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: acquire the inertia parameter combination of the satellite in the satellite center of mass body coordinate system when all the rotating mechanisms of the satellite's solar sail driving device are combined at each angular position; The PID parameter combination calculation module is used to: calculate the PID parameter combination corresponding to each inertia parameter combination; The fitting module is used to fit the corresponding relationship between the angular position of each rotating mechanism and all PID parameters according to each angular position combination of all rotating mechanisms in the solar sail panel driving device and all PID parameter combinations; The satellite attitude control module is used to calculate the desired control torque for adjusting the satellite from the current attitude to the final desired attitude by using the PID control law associated with the corresponding relationship, and control the satellite's actuator 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 for: Under the constraint of satellite attitude control accuracy, the PID closed-loop control method is used, and the frequency domain design method is adopted to traverse each angular position combination and calculate the PID parameter combination corresponding to each inertia parameter combination.

7. A satellite attitude control system according to claim 5 or 6, characterized in that: The solar sail panel driving device includes three rotating mechanisms arranged in sequence, the rotating axis of the rotating mechanism in the middle is perpendicular to the rotating axes of the rotating mechanisms on both sides, the rotating mechanisms on both sides are respectively connected to a solar sail panel, and a sun sensor is arranged on each solar sail 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: The invention comprises a memory, a processor and a computer program stored in the memory and executable on the processor, wherein when the processor executes the computer program, a satellite attitude control method as claimed in any one of claims 1 to 4 is implemented.

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

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