Attitude control system for short-wave satellite

Abstract

The invention provides an attitude control system for a short-wave satellite, and relates to the technical field of data processing, and the attitude control system comprises the steps: inputting a three-axis synthesis control torque instruction into a reaction flywheel controller, converting the instruction into a flywheel motor rotating speed adjustment instruction, and driving a reaction flywheel to execute; adjusting and analyzing the continuously updated attitude deviation angle based on the actual attitude change after the execution of the driving reaction flywheel, obtaining the updated gain parameter of the sliding mode controller, and feeding back the parameter of the updated sliding mode controller; performing control based on the updated parameters of the sliding mode controller, performing direction consistency verification on a newly output sun direction vector measurement value, and verifying an actual included angle; if the actual included angle is smaller than the preset precision threshold value, a new round of data acquisition cycle is triggered, and closed-loop control is achieved. According to the invention, the timeliness and effectiveness of attitude control are ensured, and the control efficiency and reliability are improved.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of data processing, in particular to a short-wave satellite attitude control system. BACKGROUND

[0002] As an important communication method, short-wave communication plays an irreplaceable role in the fields of aerospace, emergency communication, etc. As a key platform for realizing long-distance short-wave communication, the stability and accurate control of the attitude of the short-wave satellite directly determines the communication quality and task completion efficiency. With the continuous expansion of short-wave communication business and the increasing requirement for communication quality, the traditional satellite attitude control system cannot meet the precision attitude control requirements of the short-wave satellite in complex space environment.

[0003] When processing the measurement data such as the sun direction vector, the traditional attitude solving method often uses a simple vector comparison method, which fails to fully consider the nonlinear characteristics of the spherical coordinates, resulting in a large error in the calculation of the attitude deviation angle, and it is difficult to meet the requirements of the short-wave satellite for the attitude accuracy. SUMMARY

[0004] The technical problem to be solved by the present application is to provide a short-wave satellite attitude control system, which ensures the timeliness and effectiveness of attitude control and improves the efficiency and reliability of control.

[0005] To solve the above technical problems, the technical scheme of the present application is as follows: In a first aspect, a short-wave satellite attitude control system comprises: An acquisition module is configured to obtain angular velocity measurement values and sun direction vector measurement values in a satellite body coordinate system; A calculation module is configured to perform vector difference operation on the sun direction vector measurement values through an orbit attitude solver and the sun reference direction predicted by ephemeris, and convert the vector difference into spherical azimuth deviation and elevation angle deviation through a spherical coordinate mapping algorithm to obtain an attitude deviation angle; A control module is configured to input the attitude deviation angle into a sliding mode controller and generate a three-axis attitude correction torque command through a variable structure control algorithm; A compensation module is configured to calculate a gravity gradient interference torque and a solar pressure interference torque based on a preset orbit perturbation model, generate a compensation torque after compensation processing, and superimpose the compensation torque on the three-axis attitude correction torque command to generate a three-axis combined control torque command; An execution module is configured to input the three-axis combined control torque command into a reaction flywheel controller, convert it into a flywheel motor speed adjustment command and drive the reaction flywheel to execute; and The analysis module is configured to adjust and analyze the continuously updated attitude deviation angle based on an actual attitude change after the driving reaction flywheel is executed, to obtain an updated sliding mode controller gain parameter, and to feed back the parameter of the updated sliding mode controller; The verification module is configured to perform direction consistency verification on a newly output sun direction vector measurement value based on the parameter of the updated sliding mode controller, to verify an actual included angle, and to trigger a new round of data acquisition period to realize closed-loop control if the actual included angle is less than a preset precision threshold.

[0006] Further, the sun direction vector measurement value is subjected to vector difference operation with the sun reference direction predicted by ephemeris through an orbit attitude solver, and the vector difference is converted into a spherical azimuth angle deviation and an elevation angle deviation through a spherical coordinate mapping algorithm to obtain an attitude deviation angle, including: In the orbit attitude solver, the sun direction vector measurement value under the satellite body coordinate system and the sun reference direction vector predicted by ephemeris are obtained, component values on X, Y and Z axes are extracted respectively, subtraction operation on corresponding coordinate axes is performed based on the component values to generate axis component difference values, the component difference values are reorganized according to the orthogonal directions of the satellite body coordinate system to generate a sun direction vector difference; based on the satellite body coordinate system, the orbit normal axis is selected as an elevation angle reference axis, and an orbit plane reference surface is established; The sun direction vector difference is projected onto the orbit plane reference surface to obtain a plane projection vector; an azimuth angle of the plane projection vector relative to a preset reference direction axis in the orbit plane reference surface is calculated to obtain a spherical azimuth angle deviation; an included angle between the sun direction vector difference and the elevation angle reference axis is calculated to obtain a spherical elevation angle deviation; The spherical azimuth angle deviation and the spherical elevation angle deviation are fused to finally obtain the attitude deviation angle.

[0007] Further, the attitude deviation angle is input into the sliding mode controller, and a three-axis attitude correction torque instruction is generated through a variable structure control algorithm, including: Based on the attitude deviation angle and the historically stored attitude deviation angle, the sliding mode controller calculates a change rate of the attitude deviation angle with time to generate an attitude deviation change rate; the attitude deviation angle and the attitude deviation change rate are weighted and combined according to a preset proportional coefficient inside the sliding mode controller to generate a sliding mode surface input value; the sliding mode surface input value is input into the sliding mode controller to obtain a sliding mode surface state value; Based on the sliding mode surface input value and a preset equivalent control gain parameter, an equivalent control component is calculated; and based on the sign of the sliding mode surface state value and a preset switching control gain parameter, a switching control component is calculated; the equivalent control component and the switching control component are subjected to scalar addition operation and fusion to generate a sliding mode control signal; Based on the sliding mode control signal and the control distribution matrix pre-stored in the sliding mode controller, a three-axis control torque reference vector is calculated and generated, and the reference vector is distributed to each control axis according to the three orthogonal directions of the roll axis, the pitch axis and the yaw axis of the satellite body coordinate system, so as to generate a three-axis attitude correction torque instruction.

[0008] Further, based on the preset orbit perturbation model, the gravity gradient interference torque and the solar pressure interference torque are calculated, the compensation torque is generated after compensation processing, and the compensation torque is superimposed on the three-axis attitude correction torque instruction to generate a three-axis combined control torque instruction, including: Based on the real-time orbit data of the satellite and the preset orbit perturbation model, the orbit position parameters of the satellite are obtained; based on the orbit position parameters, the gravity gradient interference torque component in the orbit coordinate system is calculated and converted to the satellite body coordinate system; based on the solar direction vector measurement value, the solar pressure interference torque component in the satellite body coordinate system is calculated; The gravity gradient interference torque component and the solar pressure interference torque component are subjected to vector addition operation in the satellite body coordinate system to generate a total interference torque vector; The total interference torque vector is subjected to a negative operation to generate a compensation torque vector; the three-axis attitude correction torque instruction and the compensation torque vector are subjected to vector addition operation in the satellite body coordinate system to generate a three-axis combined control torque instruction.

[0009] Further, the three-axis combined control torque instruction is input into the reaction flywheel controller to be converted into a flywheel motor speed regulation instruction and drive the reaction flywheel to execute, including: The three-axis combined control torque instruction is decomposed to generate target control torque components of the roll, pitch and yaw axes; based on the target control torque components of each axis, the target angular acceleration of each axis is calculated in combination with the moment of inertia parameters; and based on the preset flywheel configuration information, the axis responsible for each flywheel and the distribution ratio are determined; Based on the target angular acceleration of each control axis, the flywheel configuration and the distribution ratio, the target torque required by each reaction flywheel is calculated; based on the target torque and the moment of inertia parameters of the flywheel, the target speed change amount required by the flywheel motor is calculated; Based on the target speed change amount, in combination with the current real-time speed feedback value of each flywheel motor obtained from the flywheel motor in real time, the flywheel motor speed regulation instruction value is calculated and generated; the flywheel motor speed regulation instruction value is executed through the driving reaction flywheel to change the speed.

[0010] Further, based on the actual attitude change after driving the reaction flywheel, the continuously updated attitude deviation angle is adjusted and analyzed to obtain the updated sliding mode controller gain parameter, and the parameter of the updated sliding mode controller is fed back, including: Based on the actual attitude change of the driven reaction flywheel executing the satellite, the updated angular velocity measurement value and the sun direction vector measurement value in the satellite body coordinate system are continuously obtained; the updated sun direction vector measurement value is input into the orbit attitude solver to generate an updated attitude deviation angle; Based on the updated attitude deviation angle, combined with the historically stored attitude deviation angle data, time series data is established; based on the time series data, the statistical characteristics of the attitude deviation angle in the preset time window are calculated; Based on the statistical characteristics of the attitude deviation angle, combined with the preset sliding mode controller gain adjustment strategy rule base for analysis; according to the analysis result, the gain adjustment factor for the equivalent control gain parameter and the switching control gain parameter of the sliding mode controller is generated; The gain adjustment factor is multiplied by the equivalent control gain parameter and the switching control gain parameter stored in the sliding mode controller respectively, and the updated equivalent control gain parameter value and the updated switching control gain parameter value are calculated; The equivalent control gain parameter value and the switching control gain parameter value are fed back to the internal parameter storage of the sliding mode controller to complete the online update of the sliding mode controller parameters.

[0011] Further, based on the updated parameters of the sliding mode controller, the direction consistency of the newly output sun direction vector measurement value is verified, and the actual included angle is verified; if the actual included angle is less than the preset precision threshold, a new round of data acquisition period is triggered, and closed-loop control is realized, including: Based on the updated parameters of the sliding mode controller, a complete control chain is executed to drive the satellite attitude adjustment, and the newly output sun direction vector measurement value under the new parameter control is obtained; The newly output sun direction vector measurement value is dot multiplied with the ephemeris predicted sun reference direction vector to calculate the actual direction included angle therebetween; the actual direction included angle is compared with the preset threshold to determine whether the actual direction included angle is less than the threshold; According to the judgment result, if the direction included angle is less than the threshold, a new round of data acquisition period is triggered; if it is not less than the threshold, the control flow is maintained and the control and verification are continued.

[0012] Further, based on the satellite real-time orbit data and the preset orbit perturbation model, the orbit position parameters of the satellite are obtained; based on the orbit position parameters, the gravity gradient disturbance torque component in the orbit coordinate system is calculated and converted to the satellite body coordinate system; based on the sun direction vector measurement value, the sun pressure disturbance torque component in the satellite body coordinate system is calculated, including: The position coordinate value of the orbit coordinate system is extracted based on the real-time orbit data of the satellite, a geocentric modulus scalar value is obtained through calculation, a gravity field strength scalar value is generated by combining the pre-stored gravity constant of the earth, and the main diagonal element value of the satellite rotational inertia is called to generate three-axis gravity gradient torque components; According to the three-axis gravity gradient torque components, a rotation transformation matrix is constructed in combination with the real-time attitude quaternion of the satellite, and a vector coordinate system conversion operation is performed through the matrix to obtain three-axis gravity gradient interference torque component values; Based on the unit vector of the sun incident direction, the effective projection area and the reflection coefficient are output from the satellite surface optical parameter database, the light pressure scalar value is calculated in combination with the solar light pressure intensity constant, and the solar light pressure interference torque component is generated through the torque fusion operation of the light pressure scalar value and the pre-calibrated position deviation vector.

[0013] In a second aspect, a computing device includes: One or more processors; Storage for storing one or more programs that, when executed by the one or more processors, cause the one or more processors to implement the system.

[0014] In a third aspect, a computer-readable storage medium stores a program that, when executed by a processor, implements the system.

[0015] The above scheme of the present application at least has the following beneficial effects: The vector difference operation is performed on the sun direction vector measurement value and the ephemeris prediction reference value by the orbit attitude solver, and the attitude deviation angle is obtained in combination with the spherical coordinate mapping algorithm, which realizes accurate quantization of the attitude deviation, quickly responds to the attitude deviation and generates a correction torque, effectively suppresses system disturbance, and improves the dynamic stability of attitude control.

[0016] The synthesized control torque is directly converted into a motor speed adjustment instruction by the reaction flywheel controller, which reduces the delay of the intermediate conversion link and improves the execution speed of the control instruction. The overall process is simple and efficient in link design from measurement, calculation, control to feedback, reducing the overall complexity. BRIEF DESCRIPTION OF DRAWINGS

[0017] Figure 1 is a short-wave satellite attitude control system schematic diagram provided by an embodiment of the present application.

[0018] Figure 2 is a three-axis synthesized control torque instruction flow schematic diagram of an embodiment of the present application. DETAILED DESCRIPTION

[0019] Exemplary embodiments of the present disclosure will be described in greater detail below with reference to the accompanying drawings. While exemplary embodiments of the present disclosure are shown in the drawings, it is understood that the present disclosure can be embodied in various forms and should not be limited by the embodiments set forth herein. Rather, these embodiments are provided so that the present disclosure can be thoroughly understood and fully conveyed to those skilled in the art.

[0020] As shown in Figure 1 An embodiment of the present application proposes a short-wave satellite attitude control system, comprising: An acquisition module is configured to obtain angular velocity measurement and sun direction vector measurement in a satellite body coordinate system; A calculation module is configured to perform vector difference operation on the sun direction vector measurement and the sun reference direction predicted by ephemeris through an orbit attitude solver, and convert the vector difference into a spherical azimuth deviation and an elevation angle deviation through a spherical coordinate mapping algorithm to obtain an attitude deviation angle; A control module is configured to input the attitude deviation angle into a sliding mode controller and generate a three-axis attitude correction torque command through a variable structure control algorithm; A compensation module is configured to calculate a gravity gradient interference torque and a solar pressure interference torque based on a preset orbit perturbation model, generate a compensation torque after compensation processing, and superimpose the compensation torque on the three-axis attitude correction torque command to generate a three-axis combined control torque command; An execution module is configured to input the three-axis combined control torque command into a reaction flywheel controller, convert it into a flywheel motor speed regulation command, and drive the reaction flywheel to execute; An analysis module is configured to adjust and analyze the continuously updated attitude deviation angle based on the actual attitude change after driving the reaction flywheel to execute, obtain updated sliding mode controller gain parameters, and feedback the parameters of the updated sliding mode controller; A verification module is configured to control based on the parameters of the updated sliding mode controller, verify the direction consistency of the newly output sun direction vector measurement, and verify the actual included angle; if the actual included angle is less than a preset precision threshold, a new round of data acquisition cycle is triggered to realize closed-loop control.

[0021] In the embodiment of the present application, the vector difference operation is performed on the sun direction vector measurement and the ephemeris predicted reference value through the orbit attitude solver, and the attitude deviation angle is obtained in combination with the spherical coordinate mapping algorithm, which realizes the accurate quantification of the attitude deviation, quickly responds to the attitude deviation and generates a correction torque, effectively suppresses the system disturbance, and improves the dynamic stability of the attitude control.

[0022] The synthetic control torque is directly converted into a motor speed adjustment instruction by the reaction flywheel controller, so that the delay of intermediate conversion links is reduced, and the execution speed of the control instruction is improved.

[0023] In a preferred embodiment of the present application, the sun direction vector measurement value is subjected to vector difference operation with the sun reference direction predicted by ephemeris through an orbit attitude solver, and the vector difference is converted into a spherical azimuth angle deviation and an elevation angle deviation through a spherical coordinate mapping algorithm to obtain an attitude deviation angle, including: In the orbit attitude solver, the sun direction vector measurement value in the satellite body coordinate system and the sun reference direction vector predicted by ephemeris are obtained, the component values of the two on the X, Y and Z axes are extracted respectively, subtraction operation on the corresponding coordinate axes is performed based on the component values to generate the component difference values of each axis, the component difference values are reorganized according to the orthogonal directions of the satellite body coordinate system to generate the sun direction vector difference; based on the satellite body coordinate system, the orbit normal axis is selected as the elevation angle reference axis, and an orbit plane reference surface is established; The sun direction vector difference is projected onto the orbit plane reference surface to obtain a plane projection vector; the azimuth angle of the plane projection vector relative to a preset reference direction axis in the orbit plane reference surface is calculated to obtain a spherical azimuth angle deviation; the included angle between the sun direction vector difference and the elevation angle reference axis is calculated to obtain a spherical elevation angle deviation; The spherical azimuth angle deviation and the spherical elevation angle deviation are fused to finally obtain the attitude deviation angle.

[0024] In the embodiment of the present application, from the satellite body coordinate system (right-handed coordinate system, the X axis is along the satellite longitudinal axis forward, the Y axis points to the right side of the satellite, and the Z axis is perpendicular to the orbit plane), three components (denoted as A x , A y , A_z) of the sun direction vector measurement value and the components (denoted as B x , B y , B_z) of the sun reference direction vector predicted by ephemeris are read; the value range of the component values is [-1, 1].

[0025] Subtraction operation is performed on the components of the same coordinate axis: X axis difference: Δ x =A x -B x ; Y axis difference: Δ y =A y -B y ; Z axis difference: Δ_z=A_z-B_z; the value range of each difference value is [-2, 2] (obtained by subtracting two [-1, 1] components).

[0026] Δ x , Δ y, Δ_z are combined according to the orthogonal relationship (X, Y, Z axes are perpendicular to each other) of the satellite body coordinate system, to form the sun direction vector difference Δ (Δ x , Δ y , Δ_z).

[0027] High angle reference axis: the selected Z axis is the orbit normal axis (perpendicular to the orbit plane), which is the reference for the high angle calculation; orbit plane reference plane: composed of the X axis and the Y axis (perpendicular to the Z axis), which is the reference plane for the azimuth angle calculation.

[0028] Spherical deviation angle calculation: project the sun direction vector difference Δ onto the orbit plane reference plane (X-Y plane) to obtain the plane projection vector Δ p (Δ x , Δ y , 0); during the projection process, the Z axis component is eliminated, and only the X and Y axis components are retained.

[0029] Spherical azimuth angle deviation calculation: select the X axis as the reference direction in the orbit plane (such as the forward direction of the satellite); by measuring the angle between the plane projection vector Δ p and the X axis, the azimuth angle deviation α is obtained.

[0030] Value range: 0°-360°, where 0° indicates that Δ p is in the same direction as the X axis, 90° indicates that it is in the same direction as the Y axis, 180° indicates that it is in the opposite direction of the X axis, and so on.

[0031] Spherical high angle deviation calculation: measure the angle between the sun direction vector difference Δ and the high angle reference axis (Z axis) to obtain the high angle deviation β; value range: -90°-90°, where 0° indicates that Δ is in the orbit plane (perpendicular to the Z axis), positive value indicates that Δ is deviated to the positive direction of the Z axis (above the orbit plane), and negative value indicates that it is deviated to the negative direction of the Z axis (below the orbit plane).

[0032] Integrate the azimuth angle deviation α (0°-360°) and the high angle deviation β (-90°-90°) through spatial geometric relationship: if α is in 0°-180° and β is positive, the attitude deviation is represented as "positive X axis deflection + upward tilt"; if α is in 180°-360° and β is negative, the attitude deviation is represented as "negative X axis deflection + downward tilt".

[0033] Finally output the comprehensive attitude deviation angle, the value range covers the possible attitude deviation interval of the satellite (usually -180°-180°, depending on the satellite attitude control capability).

[0034] Model construction: Construct the "solar vector difference-attitude deviation solving model", realize the conversion from vector difference to spherical angle deviation, and output the deviation that can be directly used for attitude control; Input layer: receive 6 parameters of solar direction vector measurement value (A x , A y , A_z) and reference vector (B x , B y , B_z).

[0035] Vector difference calculation module: execute component subtraction and vector reorganization; Projection conversion module: project the vector difference to the orbit plane; Angle solving module: calculate α and β based on geometric relationship, and fuse them into the final attitude deviation; Output attitude deviation angle (single value or α, β combination).

[0036] In the model, the satellite body coordinate system definition, the orbit plane equation (based on the orbit parameters), the reference direction axis pointing (such as the angle between the X axis and the orbit tangent) and the like are preset as the operation reference.

[0037] Model training: Through the satellite ground simulation system, 100,000 sets of solar vector measurement values (containing simulated noise) under different attitudes, corresponding ephemeris reference vectors, and real attitude deviations measured by high-precision turntable (as labels) are generated; Add extreme working condition data (such as measurement values when the sun is close to the horizon), expand the coverage range of the data set.

[0038] Training process: Substitute the input data into the model to obtain the initial attitude deviation output; Compare the model output with the real label, and calculate the azimuth angle deviation error (Δα) and the elevation angle deviation error (Δβ); adopt the mean absolute error (MAE) as the evaluation index.

[0039] Adjust the correction coefficient of the projection conversion module in the model (such as the correction value for compensating the slight tilt of the orbit plane), so that Δα and Δβ are less than 0.1° (satisfy the satellite attitude control accuracy requirement).

[0040] Iterative training: repeat the above process 500 times until the error of the model on the validation set (20% of the total data) is stable within the target range.

[0041] Introduce measurement noise (such as ±0.05° random error) and ephemeris error (such as ±0.1° orbit parameter deviation), test the output stability of the model.

[0042] For special scenes such as earth shadow area and strong sunlight interference, separately verify the adaptability of the model to ensure that the error increment does not exceed 0.2°.

[0043] Model implementation: The model is compiled into binary code executable by the onboard computer and stored in the FPGA chip of the satellite attitude control system to ensure that the time consumption of a single operation is less than or equal to 10 ms (to meet the real-time control requirement).

[0044] By component difference reorganization and spherical coordinate mapping, the three-dimensional vector difference is converted into an intuitive azimuth angle and elevation angle deviation, the multi-value problem of complex attitude angle calculation is avoided, the unique definition of the value range ensures the uniqueness of the angle description; the azimuth angle deviation corresponds to the attitude deviation in the orbital plane, and the elevation angle deviation corresponds to the deviation in the orbital normal direction, which is consistent with the actual motion law of the satellite, and the subsequent controller can generate a correction torque accordingly. The model is constructed based on geometric principles and does not require complex training, and the calculation process is simple. The accurate attitude deviation angle provides reliable input for the sliding mode controller, and indirectly improves the accuracy of the entire control system.

[0045] In a preferred embodiment of the present application, the attitude deviation angle is input into the sliding mode controller, and a three-axis attitude correction torque command is generated through a variable structure control algorithm, including: Based on the attitude deviation angle and the historically stored attitude deviation angle, the sliding mode controller calculates the rate of change of the attitude deviation angle with time to generate an attitude deviation rate; the attitude deviation angle and the attitude deviation rate are weighted and combined according to the preset proportional coefficient in the sliding mode controller to generate a sliding mode surface input value; and the sliding mode surface input value is input into the sliding mode controller to calculate the sliding mode surface state value; Based on the sliding mode surface input value and the preset equivalent control gain parameter, an equivalent control component is calculated; and based on the sign of the sliding mode surface state value and the preset switching control gain parameter, a switching control component is calculated; the equivalent control component and the switching control component are subjected to scalar addition operation and are fused to generate a sliding mode control signal; Based on the sliding mode control signal and the control distribution matrix pre-stored in the sliding mode controller, a three-axis control torque reference vector is calculated and generated, and the reference vector is distributed to each control axis according to the three orthogonal directions of the roll axis, the pitch axis and the yaw axis of the satellite body coordinate system to generate a three-axis attitude correction torque command.

[0046] In the embodiment of the present application, the sliding mode controller reads the current attitude deviation angle (denoted as θ t , the value range is -180°-180°), and retrieves the historically stored attitude deviation angle at the previous time (θ t-1 ) and the time interval (Δt, usually 0.1-1 seconds, depending on the satellite control period); the rate of change is calculated by the difference value and the time interval: attitude deviation rate = (θ t -θ t-1 ) ÷ Δt, the value range is determined by the satellite attitude dynamic characteristics, usually -10° / s-10° / s (if it exceeds this range, a fault protection is triggered).

[0047] The controller internally presets two proportional coefficients: attitude deviation weight coefficient (k1, value 0.5~2.0, adjusted according to control accuracy requirement) and change rate weight coefficient (k2, value 0.1~1.0, affecting response speed); the attitude deviation angle and change rate are multiplied by the corresponding coefficients respectively and summed to obtain the sliding mode surface input value: sliding mode surface input value = k1 x θ t + k2 x attitude deviation change rate, value range usually -50~50 (unitless, determined by coefficient and angle dimension combination).

[0048] The sliding mode controller substitutes the input value into the internally preset sliding mode surface function (such as linear function or saturation function) to output the sliding mode surface state value (s); the value range thereof is associated with the input value, usually -100~100, used to represent the deviation degree of the current state of the system from the ideal sliding mode.

[0049] The controller calls the preset equivalent control gain parameter (k e , value 0.1~5.0, set according to the inertia characteristics of the satellite) to multiply the sliding mode surface input value by the gain to obtain the equivalent control component: equivalent control component = k e x sliding mode surface input value, value range -250~250 (corresponding to the input value and gain range); the component is used to maintain the movement of the system on the ideal sliding mode.

[0050] The controller first judges the sign of the sliding mode surface state value (s): if s>0, the sign is +1; if s<0, the sign is -1; if s=0, the sign is 0; then the preset switching control gain parameter (k s , value 1.0~10.0, needs to be greater than the upper limit of the system disturbance) is called to multiply the sign by the gain to obtain the switching control component: switching control component = k s x sign, value range -10~10. The component is used to force the system state to converge to the sliding mode surface.

[0051] The equivalent control component and the switching control component are subjected to scalar addition operation to obtain the sliding mode control signal (u): u = equivalent control component + switching control component, value range -260~260; the signal integrates the steady-state control and dynamic adjustment capability to ensure that the system quickly converges to the target state.

[0052] The sliding mode controller calls the internally pre-stored control distribution matrix (3xn matrix, n being the number of actuators, the element value being determined by the installation position and efficiency of the actuator, value -1~1) to multiply the sliding mode control signal (u) by the matrix to obtain the three-axis control torque reference vector (T x , T y, T_z), the value range of each component is determined by the maximum torque of the satellite actuator, usually -5 ~ 5 N•m (Newton•meter).

[0053] According to the three orthogonal axes (roll axis X, pitch axis Y and yaw axis Z) of the satellite body coordinate system, the three components of the reference vector are directly assigned to the corresponding axes: the X axis corresponds to the roll axis correction torque (T x ), the Y axis corresponds to the pitch axis correction torque (T y ), and the Z axis corresponds to the yaw axis correction torque (T_z). The final generated three-axis attitude correction torque command needs to meet the physical limitations of each axis actuator (such as the minimum output torque ±0.1 N•m), and when exceeding the range, it is automatically truncated to the limit value.

[0054] The sliding mode control is not sensitive to model errors and spatial environmental disturbances, and even with parameter deviations, it can still guarantee control accuracy and stability better than traditional PID control. The fusion design of equivalent control component and switching control component can achieve small steady-state error of attitude deviation angle, meeting the attitude stability requirements of high-resolution remote sensing satellites. The control allocation matrix can be flexibly adapted to different satellite actuator configurations (such as flywheel and thruster combination), and through parameter adjustment, it can be applied to low-orbit, high-orbit and other types of satellites, with strong universality.

[0055] As shown in Figure 2 , based on the preset orbit perturbation model, the gravity gradient disturbance torque and the solar radiation pressure disturbance torque are calculated, the compensation torque is generated after compensation processing, and the compensation torque is superimposed on the three-axis attitude correction torque command to generate a three-axis combined control torque command, including: Based on the real-time orbit data of the satellite and the preset orbit perturbation model, the orbit position parameters of the satellite are obtained; based on the orbit position parameters, the gravity gradient disturbance torque component in the orbit coordinate system is calculated and converted to the satellite body coordinate system; based on the solar direction vector measurement value, the solar radiation pressure disturbance torque component in the satellite body coordinate system is calculated; The gravity gradient disturbance torque component and the solar radiation pressure disturbance torque component are subjected to vector addition operation in the satellite body coordinate system to generate a total disturbance torque vector; The total disturbance torque vector is subjected to a negative operation to generate a compensation torque vector; the three-axis attitude correction torque command and the compensation torque vector are subjected to vector addition operation in the satellite body coordinate system to generate a three-axis combined control torque command.

[0056] In the embodiment of the present application, based on the orbit data (such as semi-major axis, eccentricity, orbit inclination, etc.) transmitted back in real time by the satellite and the preset orbit perturbation model (including non-spherical terms of the earth's gravitational field, the sun and moon gravity, etc.), the position parameters of the satellite in the orbit coordinate system are calculated, including the geocentric distance (the distance from the satellite to the earth's center, the value range is about 400-40000 km, depending on the orbit type), the direction angle of the orbit radius vector (the value range is 0°-360°), etc.

[0057] According to the orbit position parameters and the inertia tensor parameters of the satellite (reflecting the mass distribution characteristics of the satellite, preset in the system), the three components (X_g, Y_g, Z_g) of the gravity gradient disturbance torque in the orbit coordinate system are calculated, the value range is usually -0.01-0.01 N•m (small satellite) to -1-1 N•m (large satellite), which is positively correlated with the mass and size of the satellite. or b、Y or b、Z or b), the value range is usually -0.01-0.01 N•m (small satellite) to -1-1 N•m (large satellite), which is positively correlated with the mass and size of the satellite.

[0058] Through the conversion matrix (calculated from the satellite attitude angle) between the orbit coordinate system and the satellite body coordinate system, the components in the orbit coordinate system are converted into the gravity gradient disturbance torque components (X_g, Y_g, Z_g) in the satellite body coordinate system, the value range remains unchanged.

[0059] Based on the sun direction vector measurement value, the satellite surface area parameter (preset, such as the sailboard area, the star body projection area) and the surface optical characteristics (such as the reflection coefficient, the value range is 0-1), the solar pressure disturbance torque components (X_s, Y_s, Z_s) in the satellite body coordinate system are calculated; the value range is -0.001-0.001 N•m (small satellite) to -0.1-0.1 N•m (satellite with large sailboard), which is significantly affected by the distance to the sun and the satellite attitude (such as the maximum value when the sailboard is directly opposite to the sun).

[0060] In the satellite body coordinate system, the gravity gradient disturbance torque components and the solar pressure disturbance torque components are subjected to vector addition operation: X-axis total disturbance torque: T_dx=X_g+X_s; Y-axis total disturbance torque: T_dy=Y_g+Y_s; Z-axis total disturbance torque: T_dz=Z_g+Z_s.

[0061] The total disturbance torque vector (T_dx, T_dy, T_dz) is generated, the value range is the superposition range of the sum of the components of each axis, usually -1.1-1.1 N•m (the maximum possible value of the two kinds of disturbances); the vector reflects the total disturbance effect of the space environment on the satellite attitude.

[0062] The total interference torque vector is subjected to a negative operation to obtain a compensation torque vector (T_cx, T_cy, T_cz), wherein: T_cx=-T_dx; T_cy=-T_dy; T_cz=-T_dz; the value range thereof is symmetrical to the total interference torque vector, and is -1.1-1.1 N·m, and is used for offsetting the interference of the space environment.

[0063] In the satellite body coordinate system, the generated three-axis attitude correction torque instruction (T_rx, T_ry, T_rz, the value range is -5-5 N·m) is subjected to a vector addition operation with the compensation torque vector: X-axis combined torque: T_total_x=T_rx+T_cx; Y-axis combined torque: T_total_y=T_ry+T_cy; Z-axis combined torque: T_total_z=T_rz+T_cz; The finally generated three-axis combined control torque instruction needs to meet the physical limit of the actuator (such as the maximum output torque ±5 N·m), and is automatically truncated to the limit value when exceeding the range, so as to ensure that the instruction is executable.

[0064] By compensating the main space interference torques such as gravity gradient and solar pressure, the satellite attitude stability can be improved to meet the high-precision task requirements such as high-resolution imaging and laser communication. The environmental interference torque is offset in advance, so that the attitude correction torque does not need to resist the interference, which can reduce the output power of the actuator (such as the reaction flywheel) and prolong its on-orbit life. The attitude drift caused by the interference torque is eliminated, the system is prevented from shaking due to continuous resistance to interference, and the attitude adjustment process is smoother. The compensation amount is automatically adjusted according to different orbit positions (such as perigee or apogee) and solar illumination conditions, so as to ensure that the satellite maintains stable control performance in the whole orbit period, and is especially suitable for elliptical orbit or sun-synchronous orbit satellites.

[0065] In a preferred embodiment of the present application, the three-axis combined control torque instruction is input into a reaction flywheel controller, is converted into a flywheel motor speed regulation instruction and drives the reaction flywheel to execute, comprising: The three-axis combined control torque instruction is subjected to decomposition processing to generate target control torque components of the rolling, pitching and yawing axes respectively; based on the target control torque components of the axes, the target angular accelerations of the axes are calculated; and based on the preset flywheel configuration information, the axes responsible for each flywheel and the distribution proportion thereof are determined; Based on the target angular accelerations of the control axes, the flywheel configuration and the distribution proportion, the target torque required by each reaction flywheel is calculated; based on the target torque and the moment of inertia parameter of the flywheel, the target speed change amount required by the flywheel motor is calculated; Based on the target speed variation, combined with the real-time speed feedback value of each flywheel motor, the flywheel motor speed adjustment instruction value is calculated and generated; the flywheel motor speed adjustment instruction value is changed through the driving reaction flywheel execution speed.

[0066] In the embodiment of the application, the three-axis synthetic control torque instruction (T_total_x, T_total_y, T_total_z) is directly decomposed according to the three orthogonal axes of the satellite body coordinate system to obtain the target control torque components of each axis: The target control torque of the rolling axis (X axis): T_x (the value range is -5~5 N•m, consistent with the synthetic torque range); the target control torque of the pitch axis (Y axis): T_y (the value range is -5~5 N•m); the target control torque of the yaw axis (Z axis): T_z (the value range is -5~5 N•m).

[0067] The moment of inertia parameters of each axis of the satellite (preset in the system, the moment of inertia of the X axis J_x, the moment of inertia of the Y axis J_y, and the moment of inertia of the Z axis J_z, the unit is kg•m 2 , the value range is usually 10~1000 kg•m 2 , which depends on the mass distribution of the satellite) are called, and the target angular acceleration is calculated by the ratio of the target control torque to the moment of inertia: The target angular acceleration of the rolling axis: α_x=T_x / J_x (the value range is -0.5~0.5 rad / s 2 , that is, -28.6° / s 2 ~28.6° / s 2 ); the target angular acceleration of the pitch axis: α_y=T_y / J_y (the value range is the same as above); the target angular acceleration of the yaw axis: α_z=T_z / J_z (the value range is the same as above).

[0068] Based on the preset flywheel configuration information (such as the common 4-flywheel orthogonal + diagonal layout), the physical installation angle and the control axis responsible for each flywheel are determined: For example, flywheel 1 is mainly responsible for the rolling axis (allocation proportion 70%) and the yaw axis (30%), flywheel 2 is mainly responsible for the pitch axis (80%) and the rolling axis (20%), and so on, and the sum of the allocation proportions of each axis is 100% (the value is 0~1).

[0069] The allocation proportion is set according to the moment output efficiency of the flywheel installation position, and the diagonal flywheel can participate in multi-axis control at the same time through vector decomposition.

[0070] According to the target angular acceleration of each control axis, the flywheel configuration and the allocation proportion, the shaft torque is distributed to each flywheel: For example, the target torque of flywheel 1 = alpha_x * J_x * 70% + alpha_z * J_z * 30% (the torque is reduced by the product of angular acceleration and moment of inertia); the target torque of a single flywheel ranges from -3 to 3 N•m (determined by total torque distribution and maximum output capacity of the flywheel).

[0071] The moment of inertia parameter (J_f, ranging from 0.1 to 10 kg•m 2 , which is positively correlated with the size of the flywheel) of each flywheel is called, and the angular acceleration change rate is calculated by the ratio of the target torque to the moment of inertia of the flywheel, and then the target speed change amount is obtained by combining the control period (Delta t = 0.1-1s): target speed change amount = (flywheel target torque / J_f) * Delta t (ranging from -300 to 300 r / min, i.e. the speed change amount per minute, limited by motor power).

[0072] The real-time speed feedback value (N_current, ranging from -6000 to 6000 r / min, negative sign indicating reverse rotation) of each flywheel is obtained from the speed sensor of the flywheel motor, and the speed instruction value to be adjusted is calculated by combining the target speed change amount: Speed adjustment instruction value = N_current + target speed change amount; the instruction value needs to be limited within the rated speed range of the flywheel (such as -6000-6000 r / min), and when it exceeds, it is automatically truncated to the maximum or minimum value.

[0073] The speed adjustment instruction value is converted into a motor driving signal (such as a pulse width modulation signal PWM), and the motor of the reaction flywheel is driven to rotate through a power amplification circuit, so that the actual speed of the flywheel gradually approaches the instruction value. During execution, the speed sensor feedbacks the adjusted speed in real time, forming a closed-loop control, until the speed error is less than or equal to 5 r / min (meeting the control accuracy requirement).

[0074] Through moment of inertia matching and proportional distribution, the combined torque is accurately converted into flywheel action to ensure the accuracy of attitude control. The oblique flywheel layout and multi-axis distribution strategy are adopted, and a single flywheel can participate in the torque output of multiple axes at the same time, which reduces the number of actuators, the weight and power consumption of the system compared with single-axis independent flywheels. The response time from the control torque instruction to the flywheel speed adjustment is shortened, which can quickly offset the space disturbance torque, so that the satellite attitude can recover to stable in a short time. Through speed range limitation and closed-loop feedback, the flywheel speed is prevented from exceeding, the load of each flywheel is balanced through torque distribution, and the service life is prolonged.

[0075] In a preferred embodiment of the present application, based on the actual attitude change after driving the reaction flywheel, the continuously updated attitude deviation angle is adjusted and analyzed to obtain updated sliding mode controller gain parameters, and the parameters of the updated sliding mode controller are fed back, including: Based on the actual attitude change of the rear satellite executed by the driving reaction flywheel, continuously obtain updated angular velocity measurement values and sun direction vector measurement values in the satellite body coordinate system; input the updated sun direction vector measurement values into the orbit attitude solver to generate updated attitude deviation angles; Based on the updated attitude deviation angles, combine the historically stored attitude deviation angle data to establish time series data; based on the time series data, calculate the statistical characteristics of the attitude deviation angles within a preset time window; Based on the statistical characteristics of the attitude deviation angles, analyze the preset sliding mode controller gain adjustment strategy rule base; according to the analysis result, generate gain adjustment factors for the equivalent control gain parameters and the switching control gain parameters of the sliding mode controller; Perform multiplication operation on the gain adjustment factors and the equivalent control gain parameters and the switching control gain parameters stored in the sliding mode controller respectively to calculate updated equivalent control gain parameter values and updated switching control gain parameter values; Feedback the equivalent control gain parameter values and the switching control gain parameter values to the internal parameter storage of the sliding mode controller to complete online update of the parameters of the sliding mode controller.

[0076] After driving the reaction flywheel, continuously obtain updated angular velocity measurement values (ω x , ω y , ω_z, value range -0.1~0.1 rad / s, i.e. -5.7° / s~5.7° / s) in the satellite body coordinate system through the angular velocity sensor (such as a gyroscope) of the satellite, and simultaneously obtain updated sun direction vector measurement values (A' x , A' y , A'_z, normalized value range [-1, 1]) through the sun sensor.

[0077] Input the updated sun direction vector measurement values into the orbit attitude solver to repeatedly calculate the logic (vector difference operation, spherical coordinate mapping, etc.) to generate updated attitude deviation angles (θ' t , value range -180°~180°) reflecting the actual attitude correction effect after the flywheel execution.

[0078] Retrieve the historically stored attitude deviation angle data (θ t-1 , θ t-2 ,..., θ t-100 , each period 0.1~1 second) of the past 100 control periods, and combine the updated θ' t to form a time series (length 101 points) containing the latest data.

[0079] Set the time window to the last 20 control cycles (about 2-20 seconds, adjust according to control accuracy requirements), calculate the statistical characteristics of the attitude deviation angle in the window: deviation mean: reflects the system steady-state deviation, value range -1°-1°; deviation maximum: reflects the dynamic overshoot, value range 0-5°; standard deviation of deviation rate of change: reflects the degree of attitude fluctuation, value range 0-0.5° / s; convergence time: from the peak value of the deviation to ≤0.1°, value range 0-30 seconds.

[0080] Call the preset sliding mode controller gain adjustment strategy rule base, which contains multiple sets of "statistical characteristics-adjustment strategy" mapping relationships, for example: If the deviation mean is >0.5°, it means that the equivalent control is insufficient, and the equivalent control gain needs to be increased; if the deviation maximum is >2° and the convergence time is >15 seconds, it means that the switching control is too weak, and the switching control gain needs to be increased; if the standard deviation of the deviation rate of change is >0.3° / s, it means that the system oscillation is too strong, and the switching control gain needs to be reduced.

[0081] According to the matching result of the statistical characteristics and the rule base, the gain adjustment factor is generated: The equivalent control gain adjustment factor (k ea , value range 0.9-1.1): 1.1 means an increase of 10%, 0.9 means a decrease of 10%; the switching control gain adjustment factor (k sa , value range 0.8-1.2): the adjustment amplitude is slightly larger than the equivalent control, in order to quickly suppress the oscillation or accelerate the convergence.

[0082] Retrieve the current storage of the equivalent control gain parameter of the sliding mode controller (k e _current, value 0.1-5.0), and multiply it by the equivalent control gain adjustment factor: Update the equivalent control gain k e _new=k e _current×k ea ; the result needs to be limited within the preset range (0.1-5.0), and the boundary value is taken when it is exceeded.

[0083] Retrieve the current switching control gain parameter (k s _current, value 1.0-10.0), and multiply it by the switching control gain adjustment factor: update the switching control gain k s _new=k s _current×k sa ; the result needs to be limited within the preset range (1.0-10.0), to ensure that the gain will not be too small to cause the anti-interference ability to decrease, or too large to cause the system to oscillate.

[0084] The calculated k e _new and ks The internal parameter memory (such as non-volatile memory) of the new writing sliding mode controller covers the original parameter value; the updating process adopts a "double buffering" mechanism: the new parameters take effect after the current control cycle ends, avoiding affecting the ongoing control operation during the updating process; the parameter updating frequency is consistent with the control cycle (0.1-1 seconds / time), realizing real-time adaptive adjustment of the sliding mode controller.

[0085] By continuously adjusting the gain parameters through feedback, the steady-state error of the attitude deviation can be further reduced from ±0.05° to ±0.02°, adapting to the performance drift caused by changes in mass distribution (such as fuel consumption) or equipment aging during satellite on-orbit operation. In different orbit stages (such as light or shadow area) or task modes (such as imaging / maneuvering), the gain is automatically adjusted to balance the response speed and stability, for example, reducing the switching gain to reduce oscillation during imaging, and increasing the equivalent gain to accelerate convergence during maneuvering. When the space disturbance torque suddenly changes (such as the sailboard receiving a sudden change in light angle), the switching control gain can be quickly increased to suppress the disturbance effect within 3-5 control cycles, avoiding attitude instability. By optimizing the gain parameters to reduce unnecessary torque output, the mechanical loss and energy consumption of the reaction flywheel are reduced, indirectly prolonging the on-orbit life of the actuator.

[0086] In a preferred embodiment of the present application, based on the updated parameters of the sliding mode controller, the newly output solar direction vector measurement is verified for direction consistency, and the actual included angle is verified. If the actual included angle is less than the preset precision threshold, a new data acquisition cycle is triggered, realizing closed-loop control, including: Based on the updated parameters of the sliding mode controller, a complete control chain is executed to drive the satellite attitude adjustment, and the newly output solar direction vector measurement under the control of the new parameters is obtained; The newly output solar direction vector measurement is dot multiplied with the ephemeris predicted solar reference direction vector to calculate the actual direction included angle therebetween; the actual direction included angle is compared with the preset threshold to determine whether the actual direction included angle is less than the threshold; According to the determination result, if the direction included angle is less than the threshold, a new data acquisition cycle is triggered; if it is not less than the threshold, the control flow is maintained and the control and verification are continued.

[0087] The updated sliding mode controller parameters (equivalent control gain k e _new, switching control gain k s _new) are used to re-execute the complete control flow: from solar direction vector measurement, attitude deviation calculation, sliding mode control torque generation, to disturbance torque compensation, flywheel driving and parameter updating, finally completing the satellite attitude adjustment.

[0088] Under the new parameter control, the adjusted solar direction vector measurement (denoted as A'' x 、A'' y 、A''_z) is collected by the sun sensor, which has been normalized and the range of each component is [-1, 1], reflecting the satellite's adjusted solar pointing state.

[0089] The ephemeris predicted solar reference direction vector (B x 、B y 、B_z) is obtained, with the range of each component being [-1, 1]; the dot product operation is performed on the newly output solar direction vector measurement (A'' x 、A'' y 、A''_z) and the reference vector, obtaining a scalar value (with the range [-1, 1], 1 indicating complete consistency in direction, and -1 indicating complete opposite direction) reflecting the consistency of the direction; based on the dot product result, the actual direction angle (θ_actual) between the two is calculated, with the range being 0°-180°: the closer the dot product is to 1, the smaller the angle (e.g., dot product = 0.9998 corresponds to an angle of approximately 1°); dot product = 0 corresponds to an angle of 90°.

[0090] The preset precision threshold (θ_threshold) is set according to the satellite mission requirements, usually 0.1°-0.5° (e.g., 0.1° for high-resolution remote sensing satellites, and 0.5° for general communication satellites); if the actual direction angle θ_actual<θ_threshold, it means that the current attitude accuracy meets the requirements; if θ_actual≥θ_threshold, it means that the attitude still needs to be further adjusted.

[0091] When θ_actual<θ_threshold, the system automatically triggers a new round of data collection period: the sun sensor re-collects the solar direction vector (as the initial measurement value for the next round of control), the ephemeris module updates the reference vector, and the attitude solver, sliding mode controller, etc. enter the next round of operation in turn, forming a closed-loop control cycle; the data collection period is consistent with the control period (0.1-1 seconds, ensuring real-time performance).

[0092] When θ_actual≥θ_threshold, the system does not trigger a new cycle and continues to maintain the current control flow: the sliding mode controller continuously outputs control torque based on the existing parameters, the flywheel maintains speed adjustment, and the sun sensor still collects data at the original frequency (updated every 0.1-1 seconds), until the next angle verification meets θ_actual<θ_threshold and enters a new cycle.

[0093] Verification of the azimuth angle ensures that attitude control always converges within preset accuracy thresholds, meeting core mission requirements such as the satellite's energy system (e.g., solar alignment of the sailboard) and attitude reference establishment. A new cycle is triggered only when the attitude meets accuracy requirements, avoiding unnecessary, high-frequency adjustments and reducing the onboard computer's computational load and flywheel energy consumption. A "measure-verify-trigger" mechanism forms negative feedback, effectively suppressing cumulative errors (e.g., deviations caused by sensor drift and control delays), ensuring stable satellite attitude during long-term operation. Preset thresholds can be dynamically adjusted according to the mission phase, balancing accuracy and efficiency requirements in different scenarios.

[0094] An embodiment of the present invention further provides a computing device comprising: a processor and a memory storing a computer program, wherein when the computer program is executed by the processor, the system described above is executed. All implementations in the above system embodiments are applicable to this embodiment and can achieve the same technical effects.

[0095] The embodiment of the present invention further provides a computer-readable storage medium storing instructions, which, when executed on a computer, cause the computer to execute the system described above. All implementations in the above system embodiments are applicable to this embodiment and can achieve the same technical effects.

[0096] The above is a preferred embodiment of the present invention. It should be pointed out that for ordinary technicians in this technical field, several improvements and modifications can be made without departing from the principles of the present invention. These improvements and modifications should also be regarded as within the scope of protection of the present invention.

Claims

1. A shortwave satellite attitude control system, characterized in that: include: The acquisition module is used to obtain the angular velocity measurement value and the sun direction vector measurement value in the satellite body coordinate system; The calculation module is used to calculate the vector difference between the measured value of the sun direction vector and the sun reference direction predicted by the ephemeris through the orbit attitude solver, and convert the vector difference into spherical azimuth deviation and altitude deviation through the spherical coordinate mapping algorithm to obtain the attitude deviation angle; The control module is used to input the attitude deviation angle into the sliding mode controller and generate the three-axis attitude correction torque command through the variable structure control algorithm; The compensation module is used to calculate the gravity gradient interference torque and the solar pressure interference torque based on the preset orbital perturbation model, generate the compensation torque after compensation processing, and synchronously superimpose the compensation torque on the three-axis attitude correction torque instruction to generate the three-axis synthetic control torque instruction; An execution module is used to input the three-axis synthetic control torque command into the reaction flywheel controller, convert it into a flywheel motor speed adjustment command and drive the reaction flywheel to execute; An analysis module is used to adjust and analyze the continuously updated attitude deviation angle based on the actual attitude change after the driving reaction flywheel is executed, obtain updated sliding mode controller gain parameters, and feedback to update the parameters of the sliding mode controller; A verification module is used to perform control based on the updated parameters of the sliding mode controller, verify the direction consistency of the newly output sun direction vector measurement value, and verify the actual angle; If the actual angle is less than the preset accuracy threshold, a new round of data collection cycle is triggered to achieve closed-loop control.

2. The shortwave satellite attitude control system according to claim 1, characterized in that: The measured solar direction vector value is passed through the orbit attitude solver to perform a vector difference operation with the solar reference direction predicted by the ephemeris. The vector difference is converted into spherical azimuth deviation and altitude deviation through the spherical coordinate mapping algorithm to obtain the attitude deviation angle, including: In the orbit attitude solver, the measured value of the solar direction vector in the satellite's coordinate system and the solar reference direction vector predicted by the ephemeris are obtained. The component values ​​of the two on the X, Y, and Z axes are extracted respectively. Based on the component values, subtraction operations are performed on the corresponding coordinate axes to generate the component differences of each axis. The component differences are reorganized according to the orthogonal directions of the satellite's coordinate system to generate the solar direction vector difference. Based on the satellite's coordinate system, the orbit normal axis is selected as the altitude angle reference axis, and the orbit plane reference plane is established. Project the solar direction vector difference onto the orbital plane reference plane to obtain a plane projection vector; calculate the azimuth of the plane projection vector relative to a preset reference direction axis in the orbital plane reference plane to obtain the spherical azimuth deviation; calculate the angle between the solar direction vector difference and the altitude angle reference axis to obtain the spherical altitude deviation; The spherical azimuth deviation and the spherical altitude deviation are combined to finally obtain the attitude deviation angle.

3. The shortwave satellite attitude control system according to claim 2, characterized in that: The attitude deviation angle is input into the sliding mode controller, and the three-axis attitude correction torque command is generated through the variable structure control algorithm, including: Based on the attitude deviation angle and the historically stored attitude deviation angle, the sliding mode controller calculates the rate of change of the attitude deviation angle over time to generate the attitude deviation change rate; the attitude deviation angle and the attitude deviation change rate are weighted and combined according to the proportional coefficient preset in the sliding mode controller to generate the sliding surface input value; the sliding surface input value is input into the sliding mode controller to obtain the sliding surface state value; An equivalent control component is calculated based on the sliding surface input value and a preset equivalent control gain parameter; a switching control component is calculated based on the sign of the sliding surface state value and a preset switching control gain parameter; a scalar addition operation is performed on the equivalent control component and the switching control component, and the resultant components are combined to generate a sliding mode control signal; Based on the sliding mode control signal and the control allocation matrix pre-stored in the sliding mode controller, the three-axis control torque reference vector is calculated and generated. Then, according to the three orthogonal directions of the roll axis, pitch axis and yaw axis of the satellite body coordinate system, the reference vector is allocated to each control axis to generate a three-axis attitude correction torque instruction.

4. The shortwave satellite attitude control system according to claim 3, characterized in that: Based on the preset orbital perturbation model, the gravity gradient interference torque and the solar pressure interference torque are calculated, and compensation processing is performed to generate the compensation torque. The compensation torque is then synchronously superimposed on the three-axis attitude correction torque command to generate the three-axis synthetic control torque command, including: Based on the satellite's real-time orbit data and the preset orbital perturbation model, the satellite's orbital position parameters are obtained; based on the orbital position parameters, the gravity gradient interference torque component in the orbital coordinate system is calculated and converted to the satellite's body coordinate system; based on the solar direction vector measurement value, the solar pressure interference torque component in the satellite's body coordinate system is calculated; Perform vector addition operation on the gravity gradient interference torque component and the solar pressure interference torque component in the satellite body coordinate system to generate the total interference torque vector; The total interference torque vector is negated to generate a compensation torque vector; the three-axis attitude correction torque instruction and the compensation torque vector are vector-added in the satellite body coordinate system to generate a three-axis synthetic control torque instruction.

5. The shortwave satellite attitude control system according to claim 4, characterized in that: The three-axis composite control torque command is input into the reaction flywheel controller, converted into a flywheel motor speed adjustment command and drives the reaction flywheel to execute, including: The three-axis composite control torque command is decomposed to generate the target control torque components for each of the three axes: roll, pitch, and yaw. Based on the target control torque components of each axis and the moment of inertia parameters, the target angular acceleration of each axis is calculated. The axis that each flywheel is responsible for and the distribution ratio are determined based on the preset flywheel configuration information. Based on the target angular acceleration of each control axis, the flywheel configuration and distribution ratio, the target torque required by each reaction flywheel is calculated; based on the target torque and the flywheel's moment of inertia parameter, the target speed change required by the flywheel motor is calculated; Based on the target speed change, combined with the current real-time speed feedback value of each flywheel motor obtained from the flywheel motor in real time, the speed adjustment command value of each flywheel motor is calculated and generated; the speed adjustment command value of each flywheel motor is used to execute the speed change by driving the reaction flywheel.

6. The attitude control system for shortwave satellites according to claim 5, characterized in that: Based on the actual attitude change after the drive reaction flywheel is executed, the continuously updated attitude deviation angle is adjusted and analyzed to obtain the updated sliding mode controller gain parameters, which are then fed back to update the sliding mode controller parameters, including: Based on the actual attitude change of the satellite after the reaction flywheel is driven, the angular velocity measurement value and the solar direction vector measurement value in the satellite body coordinate system are continuously obtained; the updated solar direction vector measurement value is input into the orbit attitude solver to generate the updated attitude deviation angle; Based on the updated attitude deviation angle, combined with the historically stored attitude deviation angle data, time series data is established; based on the time series data, the statistical characteristics of the attitude deviation angle within a preset time window are calculated; Based on the statistical characteristics of the attitude deviation angle, combined with the preset sliding mode controller gain adjustment strategy rule base, analysis is performed; based on the analysis results, gain adjustment factors for the equivalent control gain parameters and switching control gain parameters of the sliding mode controller are generated; Multiplying the gain adjustment factor by the equivalent control gain parameter and the switching control gain parameter stored in the sliding mode controller, respectively, to calculate an updated equivalent control gain parameter value and an updated switching control gain parameter value; The equivalent control gain parameter value and the switching control gain parameter value are fed back to the internal parameter memory of the sliding mode controller to complete the online update of the sliding mode controller parameters.

7. The shortwave satellite attitude control system according to claim 6, characterized in that: Based on the updated parameters of the sliding mode controller, the newly output sun direction vector measurement value is verified for direction consistency and actual angle; If the actual angle is less than the preset accuracy threshold, a new round of data collection cycle is triggered to achieve closed-loop control, including: Based on the updated sliding mode controller parameters, the complete control chain is executed to drive the satellite attitude adjustment and obtain the new output sun direction vector measurement value under the control of the new parameters; Perform a dot product operation on the newly output solar direction vector measurement value and the solar reference direction vector predicted by the ephemeris to calculate the actual direction angle between the two; compare the actual direction angle with a preset threshold to determine whether the actual direction angle is less than the threshold; According to the judgment result, if the direction angle is less than the threshold, a new round of data collection cycle is triggered; if it is not less than the threshold, the control process is maintained and control and verification continue.

8. The shortwave satellite attitude control system according to claim 7, characterized in that: Based on the satellite's real-time orbit data and the preset orbital perturbation model, the satellite's orbital position parameters are obtained; based on the orbital position parameters, the gravity gradient interference torque component in the orbital coordinate system is calculated and converted to the satellite's body coordinate system; The solar pressure interference torque component in the satellite coordinate system is calculated based on the measured value of the solar direction vector, including: Based on the orbital coordinate system position coordinate values ​​extracted from the satellite's real-time orbit data, the geocentric modulus scalar value is calculated, and the gravity field intensity scalar value is generated by combining it with the pre-stored earth gravitational constant. The main diagonal element value of the satellite's moment of inertia is then called to generate the three-axis gravity gradient torque component. According to the three-axis gravity gradient torque component, the rotation transformation matrix is ​​constructed in combination with the satellite real-time attitude quaternion, and the vector coordinate system conversion operation is performed through the matrix to obtain the three-axis gravity gradient interference torque component value; Based on the unit vector of the solar incidence direction, the satellite surface optical parameter database is retrieved to output the effective projected area and reflection coefficient. The light pressure scalar value is calculated in combination with the solar light pressure intensity constant. The solar light pressure interference torque component is generated through the torque fusion operation of the light pressure scalar value and the pre-calibrated position deviation vector.

9. A computing device, characterized in that include: one or more processors; A storage device for storing one or more programs, wherein when the one or more programs are executed by the one or more processors, the one or more processors implement the system according to any one of claims 1 to 8.

10. A computer-readable storage medium, characterized in that The computer-readable storage medium stores a program, which, when executed by a processor, implements the system according to any one of claims 1 to 8.

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