Micro-nano satellite angular momentum autonomous management method

Abstract

The invention discloses a micro-nano satellite angular momentum autonomous management method, and belongs to the technical field of satellite attitude control. The problem that an open-loop method is adopted in existing satellite angular momentum management, and the autonomous correction capacity is poor is solved. Comprising the steps that a satellite autonomously completes environment torque parameter identification through in-orbit flywheel rotating speed data, and autonomously updates parameters with high change speed along with task time extension; two task targets of angular momentum cumulative suppression and angular momentum unloading can be realized by operating the environment moment only through attitude bias relative to the sun orientation attitude and rotation of a sailboard, and intervention of active control momentum such as a magnetic torquer and jet unloading is not needed; manual intervention is not needed in the whole process, and angular momentum autonomous management is achieved. The method can autonomously carry out when the satellite has no task, and is good in reliability.

Description

Technical Field

[0001] This invention relates to a method for autonomous management of angular momentum of micro and nano satellites, belonging to the field of satellite attitude control technology. Background Technology

[0002] Existing satellite missions in orbit tend to use relatively simple methods such as spin, fixed attitude bias, and angular momentum bias to manage angular momentum, in order to delay flywheel saturation and unload it to some extent. The existing inertial angular momentum bias based on the Fengyun-4 remote sensing satellite and the zero-space angular momentum bias of the six-axis redundant flywheel system effectively avoids flywheel zero-crossing and saturation. By manipulating Chang'e-2 to actively rotate its solar panels and spin, generating turbine torque, the accumulation of angular momentum on the X-axis of the satellite system was effectively reduced. Based on the theoretical analysis of the gravity gradient torque of Chang'e-5, and combined with telemetry data, a satellite pitch axis attitude bias scheme was designed to extend the angular momentum unloading period. The autonomous angular momentum management scheme of the Queqiao-2 lunar communication relay satellite, by switching between three attitude biases and using gravity gradient torque to manage satellite angular momentum, does not effectively utilize the solar radiation pressure torque. These angular momentum management methods can only slow down the accumulation of satellite angular momentum. Even without the use of a magnetic torque converter, the flywheel will still slowly saturate. Depending on the mission scenario, the angular momentum still needs to be unloaded through the propulsion system every few days to several weeks.

[0003] Regarding angular momentum unloading, existing phased angular momentum management schemes for space stations include: one is to use a torque balance attitude tracking mode to unload the angular momentum accumulated by the space station's CMGs (Control Torque Gyroscopes) during the attitude stabilization mode using gravity gradient torque; the second is to use aerodynamic torque to manage the angular momentum of LEO (Low Earth Orbit) satellites by deflecting the satellite's solar panels; the third is to use an angular momentum unloading signal generator based on a strategy theme method that considers mission constraints to determine the timing of unloading, and to use gravity gradient torque and solar panel aerodynamic torque for angular momentum management; and the fourth is to use unscented Kalman filtering to predict the interplanetary magnetic field on-board, and to achieve magnetic torque unloading during orbit transfer based on this prediction.

[0004] Most of the studies on angular momentum management using environmental torque described above are open-loop methods. These rely heavily on ground control capabilities and require precise modeling and parameter prediction of the environmental torque. They lack on-orbit correction capabilities and exhibit poor autonomy. Some methods still rely on active management methods such as magnetic torque converters, with environmental torque playing only an auxiliary role. Most methods do not address angular momentum unloading, only extending the unloading interval by reducing angular momentum accumulation. Studies involving angular momentum unloading mostly employ traditional control methods, resulting in complex controller structures, high computational requirements, and a lack of practical mission validation. Summary of the Invention

[0005] To address the problem that existing satellite angular momentum management methods employ open-loop approaches and have poor autonomous correction capabilities, this invention provides a method for autonomous angular momentum management of micro / nano satellites.

[0006] The present invention provides a method for autonomous management of the angular momentum of micro / nano satellites, comprising obtaining an estimated value of the environmental moment in the satellite body coordinate system based on a satellite environmental moment model. ; Estimated values ​​of environmental torque By integrating and transforming, we obtain the cumulative estimate of the satellite's angular momentum. Extract independent parameters from the initial parameters of the satellite environmental moment model as the parameters to be solved; based on the collected estimated environmental moment values... The comparison results between the number of values ​​and the set data volume threshold are used to estimate the environmental torque values. and the cumulative estimate of satellite angular momentum The values ​​of the model parameters to be solved are calculated as model parameter values; the satellite enters the angular momentum management mode, switching between unloading mode and cruise mode to achieve autonomous angular momentum management: at the end of an orbital cycle, if the difference between the measured solar incidence angle and the calculated solar incidence angle in the current model parameter values ​​exceeds the incidence angle threshold, the measured solar incidence angle replaces the calculated solar incidence angle to update the current model parameter values; otherwise, the current model parameter values ​​remain unchanged; the current satellite flywheel speed is read, and the current cumulative true value of the satellite's angular momentum is calculated; if the current cumulative true value of the satellite's angular momentum exceeds the set angular momentum threshold, the satellite enters the unloading mode: by changing the control variables, the estimated environmental moment value for the next orbital cycle is changed. The magnitude and direction of the satellite's angular momentum accumulation estimate for the next orbital period are based on the current model parameter values. The current true value of the satellite's accumulated angular momentum is offset; if the current true value of the satellite's accumulated angular momentum does not exceed the set angular momentum threshold, then cruise mode is entered: by changing the control variables, the estimated value of the satellite's accumulated angular momentum for the next orbital period is based on the current model parameter values. The module length is minimized; then it enters the next orbital cycle, and decides again whether to enter cruise mode or unloading mode, until the end.

[0007] The beneficial effects of this invention are as follows: In this method, the satellite autonomously identifies environmental torque parameters using on-orbit flywheel rotation speed data and updates rapidly changing parameters as the mission duration increases. By controlling the environmental torque through attitude offset relative to the sun-oriented attitude and solar panel rotation, it can achieve two mission objectives: angular momentum accumulation suppression and angular momentum unloading, without the need for active control torque intervention such as magnetic torque generators or jet unloading. The entire process requires no manual intervention, achieving autonomous angular momentum management and significantly reducing ground support resource requirements. Mature reliance and reflection methods, along with convex optimization, are used to complete parameter identification and optimal control quantity calculation, resulting in a simple program structure and high reliability.

[0008] The method of this invention achieves angular momentum management by fully utilizing environmental torque: no jet unloading, no use of magnetic torquers, and no addition of additional actuators are required. The satellite only needs to manipulate its attitude deflection and solar panel rotation angle through the flywheel system and SADA (Solar Actuation and Control) mechanism, and can achieve efficient angular momentum management by utilizing the most common gravity gradient and solar pressure torque.

[0009] Optimal angular momentum management based on angular momentum accumulation function: without designing complex filters and control laws, it only needs to identify a finite number of parameters in the simple environmental torque function to efficiently predict the angular momentum accumulation in the next orbital period, and use a reliable convex optimization method to calculate the attitude deflection angle and sail rotation angle with the highest angular momentum management efficiency.

[0010] High degree of autonomy and reliability: The angular momentum management program requires no intervention from ground flight control personnel and can run autonomously when the satellite is not on a mission. It does not use immature mathematical methods with poor robustness, nor does it manipulate satellite equipment such as thrusters that could cause serious damage, thus ensuring high reliability. Attached Figure Description

[0011] Figure 1 This is a flowchart of the micro / nano satellite angular momentum autonomous management method described in this invention; Figure 2 This is a schematic diagram of orbital period identification based on the angular momentum accumulation curve; Figure 3 This is a schematic diagram illustrating the division of the pericentric and distal arc segments of an elliptical orbit. Figure 4 This is a schematic diagram illustrating the accumulation of angular momentum of a lunar orbiting satellite within 180 days during a lunar low orbit exploration mission. Figure 5 This is a schematic diagram of the accumulation of three-axis angular momentum in the inertial frame of a lunar satellite within 180 days; Figure 6 This is a comparative diagram showing the accumulation of angular momentum of satellites orbiting the moon over 30 days; Figure 7 This is a schematic diagram of the satellite attitude offset angle output by the lunar orbiting satellite angular momentum management working mode within 180 days; Figure 8 This is a schematic diagram of the solar panel rotation angle output by the lunar orbiting satellite angular momentum management working mode within 180 days; Figure 9 This is a schematic diagram illustrating the charging efficiency of a lunar satellite within 180 days. Figure 10 This is a schematic diagram illustrating the accumulation of angular momentum of the Mars communication relay satellite over 250 days during the Mars high-orbit communication relay mission. Figure 11 This is a schematic diagram of the accumulation of three-axis angular momentum in the inertial frame of the Mars communication relay satellite within 250 days; Figure 12 This is a diagram showing the cumulative angular momentum of Mars' high-orbit satellites over 6 days. Figure 13 This is a schematic diagram of the satellite attitude offset angle output by the Mars satellite angular momentum management working mode within 250 days; Figure 14 This is a schematic diagram of the solar panel rotation angle output by the Mars satellite angular momentum management working mode within 250 days; Figure 15 This is a diagram illustrating the charging efficiency of Mars' satellites over 250 days. Detailed Implementation

[0012] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0013] Specific Implementation Method 1: Combination Figure 1 As shown, this invention provides a method for autonomous management of the angular momentum of micro / nano satellites, including obtaining an estimated value of the environmental moment in the satellite's body coordinate system based on the satellite's environmental moment model. ; Estimated values ​​of environmental torque By integrating and transforming, we obtain the cumulative estimate of the satellite's angular momentum. Extract independent parameters from the initial parameters of the satellite environmental moment model as the parameters to be solved; based on the collected estimated environmental moment values... The comparison results between the number of values ​​and the set data volume threshold are used to estimate the environmental torque values. and the cumulative estimate of satellite angular momentum The values ​​of the model parameters to be solved are calculated as model parameter values; the satellite enters the angular momentum management mode, switching between unloading mode and cruise mode to achieve autonomous angular momentum management: at the end of an orbital cycle, if the difference between the measured solar incidence angle and the calculated solar incidence angle in the current model parameter values ​​exceeds the incidence angle threshold, the measured solar incidence angle replaces the calculated solar incidence angle to update the current model parameter values; otherwise, the current model parameter values ​​remain unchanged; the current satellite flywheel speed is read, and the current cumulative true value of the satellite's angular momentum is calculated; if the current cumulative true value of the satellite's angular momentum exceeds the set angular momentum threshold, the satellite enters the unloading mode: by changing the control variables, the estimated environmental moment value for the next orbital cycle is changed. The magnitude and direction of the satellite's angular momentum accumulation estimate for the next orbital period are based on the current model parameter values. The current true value of the satellite's accumulated angular momentum is offset; if the current true value of the satellite's accumulated angular momentum does not exceed the set angular momentum threshold, then cruise mode is entered: by changing the control variables, the estimated value of the satellite's accumulated angular momentum for the next orbital period is based on the current model parameter values. The module length is minimized; then it enters the next orbital cycle, and decides again whether to enter cruise mode or unloading mode, until the end.

[0014] In this embodiment, the satellite autonomously identifies environmental torque parameters on orbit based on environmental torque modeling, and predicts the cumulative integral of angular momentum based on the identified parameters, generating commands such as optimal attitude offset and solar panel rotation, so that the solar radiation pressure and gravity gradient torque cancel each other out in cruise mode, approaching balance, and significantly slowing down the angular momentum accumulation rate of the flywheel; when the flywheel is close to saturation, it automatically unloads through environmental torque.

[0015] Furthermore, environmental moment modeling and angular momentum integral prediction for micro / nano satellites: Angular momentum cumulative integral prediction: Based on the satellite environmental torque model, the solar radiation pressure torque of each plane and the overall satellite gravity gradient torque are added together to obtain the estimated environmental torque value. The expression is: , In the formula For roll angle, The pitch angle, Yaw angle This is the corner of the left solar panel. This is the corner of the right solar panel. For true near point angle, These are the initial parameters for the model. The solar radiation pressure torque acting on the satellite's central enclosure. This represents the solar radiation pressure torque experienced by the left solar panel. This represents the solar radiation pressure torque exerted on the right-side solar panel. This is the gravitational gradient torque of the entire star.

[0016] Environmental torque estimates The function's independent variable is only the attitude angle (including roll angle). Pitch angle and yaw angle The windshield angle and true perimeter angle are known quantities under normal operating conditions of the attitude determination system and navigation system.

[0017] The cumulative estimate of satellite angular momentum was obtained. The method is to express the control quantity as : , Integrating the environmental torque over time to obtain the accumulated angular momentum generated within one orbital period, and then predicting the accumulated angular momentum of the satellite in the next orbital period, is the basis for manipulating the environmental torque.

[0018] Satellite solar pressure torque and control quantity Related to calculating satellite solar radiation pressure torque Accumulated angular momentum generated over one orbital period : , In the formula For orbital period, For time; The optical pressure torque is independent of the orbital elements and is not affected by the satellite's current true perihelion angle. Therefore, it is independent of time, and integrating it over time is equivalent to integrating a constant over time.

[0019] Integrating the gravitational gradient torque over the orbital period over time can eliminate the periodic variation in torque caused by the true anomaly angle. For a circular orbit satellite, the orbital radius is a constant, and according to the periodic relationship of a circular orbit: In the formula Let be the orbital angular velocity; then the integral relationship is: , In the formula The gravitational gradient torque of the entire satellite over one orbital period The cumulative value of angular momentum generated, These are the gravitational parameters of the celestial body at the center of the satellite's orbit. The distance of the satellite from the center of mass of the central celestial body. Let be the unit vector pointing from the satellite's center of mass to the center of mass of the central celestial body. The moment of inertia matrix in the satellite's body coordinate system; , , , , , As an intermediate variable: , In the formula This represents the product of the satellite's YZ axes inertia. This represents the product of the satellite's X and Z axes inertia. This represents the product of the satellite's X and Y axes inertia. This represents the principal moment of inertia along the X-axis of the satellite in its body coordinate system. This represents the principal moment of inertia along the Y-axis of the satellite in its body coordinate system. This represents the principal moment of inertia of the satellite along the Z-axis in its body coordinate system. , , , , , As an intermediate variable; , In the formula , , , , , As an intermediate variable, , In the formula This is the calculated value for the angle of incidence of the sun. The inclination angle of the track; , , , , , , , , The nine elements of the coordinate rotation matrix for the satellite from the inertial frame to the body coordinate system are: , In the formula Let X be the coordinate rotation matrix from the inertial frame to the body coordinate system, defined by the rotation order along the XYZ axes. Yaw angle The coordinate transformation matrix, pitch angle The coordinate transformation matrix, For roll angle The coordinate transformation matrix.

[0020] Combination Figure 2 As shown, All orbital angular velocities in the expression Eliminate, become After accumulating points, It is only related to system parameters, attitude offset angle, and orbital period. This leads to an estimate of the satellite's cumulative angular momentum over one orbital period. : .

[0021] For satellites in elliptical orbits The orbital periodicity relationship does not exist, and In the expression It is a variable related to orbital time that cannot be described by analytical mathematical formulas, hence the estimated value of the cumulative angular momentum of the satellite during the orbital period. Theoretical integration cannot be completed. Most elliptical orbits have long orbital periods. Directly performing numerical integration of angular momentum for the entire orbital period using fixed time steps would require too many calculation steps, and each step would necessitate solving a Keplerian problem (calculating r from a known flight time), making on-board computation difficult. This implementation divides an orbital period into an average of 2000-5000 solution steps, based on the period length. The true anomaly angle λ and geocentric radius vector (distance from the satellite to the center of mass of the central celestial body) r are pre-calculated with high precision at the end of each step and stored in a file for lookup and interpolation during angular momentum integration.

[0022] Combination Figure 3 As shown, according to Kepler's theorem, the velocity of an elliptical orbiting satellite is fastest near its pericentric point. The time required for the pericentric angle to move from -45° to +45° from the pericentric point (0°) is only 1 / 10 or less of the orbital period. During this time, the rate of change of r and the gravitational gradient torque are the fastest, so a shorter step size should be used for angular momentum integration. Because the transit time for this arc is short, even with a very short integration step size, the total number of integration steps is still within an acceptable range. For the telecentric arc where the pericentric angle moves from +45° to -45° from the apocentric point (180°), the satellite transit time is long, the average angular velocity is slow, and r changes slowly. Therefore, a longer integration step size can be used to reduce the number of integration steps, speed up the numerical calculation, and the final angular momentum integration accuracy is still high.

[0023] Furthermore, methods for identifying environmental torque parameters: The method for extracting independent parameters from the initial parameters of the satellite environmental moment model is as follows: In the satellite environmental moment system, the parameters that cannot be changed by active maneuvering strategies such as attitude bias and solar panel rotation and are independent of time are the model parameters.

[0024] Environmental torque estimates In the expression, the solar radiation pressure torque on the satellite's central enclosure The mathematical model is as follows: , In the formula The pressure of sunlight. The effective sun-exposed area of ​​the satellite's central enclosure. The absorption coefficient of the surface material of the satellite's central enclosure. The diffuse reflectance coefficient of the surface material of the satellite's central enclosure. This represents the Z-axis component of the solar radiation pressure center position vector of the satellite's central enclosure in the satellite's body coordinate system. This represents the Y-axis component of the solar radiation pressure center position vector of the satellite's central enclosure in the satellite's body coordinate system. The X-axis component of the solar radiation pressure center position vector of the satellite's central enclosure in the satellite's body coordinate system. The specular reflectance coefficient of the surface material of the satellite's central enclosure; The angle of incidence of sunlight on the sunlit surface of the satellite's central enclosure: ; Extracting the solar radiation pressure torque on the satellite's central enclosure The mathematical model contains quantities that are only related to objective conditions and are represented as intermediate variables. , , , , : Based on the numerical relationships: ,get: Thus determine , , , These are independent parameters; These are derived parameters that can be obtained from other parameters.

[0025] Similarly, determine the initial parameters of the model that correspond to the gravitational gradient torque of the entire satellite. The eight independent parameters are related to the solar pressure torque on the left solar panel. and the solar radiation pressure torque on the right solar panel There are 6 related independent parameters, and a total of 18 independent parameters, which are used as the parameters of the model to be solved.

[0026] When the collected environmental torque estimate The number of values ​​is less than the set data volume threshold, so that the environmental torque estimate... The goal is to minimize the sum of squared errors between the model and the true values ​​of environmental torques, and to calculate the values ​​of the model parameters to be solved.

[0027] Least squares parameter identification method: In order to make full use of the flywheel speed data within one orbital cycle, accelerate the response speed, and reduce the number of control cycles, a parameter identification method based on environmental torque differential estimation curve fitting is proposed.

[0028] After the satellite completes its orbital maneuvers and enters a stable operating orbit, it continuously changes the control variable x at predetermined times to maintain the normal operation of the attitude stability closed-loop controller, absorbing the angular momentum generated by environmental torque into the flywheel system. Based on the precisely measured flywheel installation geometry and flywheel moment of inertia before launch, the flywheel speed is converted into an estimated cumulative value of the satellite's angular momentum within this system. The cumulative curve is obtained. By performing time difference analysis on this curve, the estimated value of the satellite environmental moment can be obtained.

[0029] After sampling environmental moment estimates at N time points, the trust region reflection algorithm is used to obtain the environmental moment estimates at N time points. If the number of elements is less than the set data volume threshold, the method for calculating the model parameter values ​​is as follows: , In the formula For the i-th environmental torque estimate The corresponding control quantity, For the i-th environmental torque estimate The corresponding true anterior angle, These are the model parameter values. For the i-th environmental torque estimate The corresponding true value of the environmental torque, where N is the estimated value of the environmental torque. The number of [elements]. The model parameter values ​​are calculated. The optimal value.

[0030] When the collected environmental torque estimate The number of values ​​should not be less than the set data volume threshold, so that the cumulative estimate of satellite angular momentum is obtained. The goal is to minimize the sum of squared errors between the actual cumulative values ​​of satellite angular momentum and the calculated values ​​of the model parameters.

[0031] Parameter identification methods based on environmental moment difference estimation rely on the true anomaly angle data of the satellite provided by the navigation system, making decoupling from the navigation system difficult and resulting in restart challenges after reference loss. This problem can be solved by performing parameter identification based on the accumulated angular momentum of one orbital period obtained through time integration of the environmental moment. Each time the satellite completes an orbital period, the current total angular momentum is sampled using the flywheel rotation speed. The difference between this sampled angular momentum and the value at the end of the previous period is used to obtain the accumulated angular momentum for that period. Using a trust-region reflection algorithm with a least-squares index, the collected environmental moment estimate is... The number of elements is not less than the set data volume threshold. The method for calculating the model parameter values ​​is as follows: , The cumulative estimate of the angular momentum of the i-th satellite The corresponding true value of the cumulative angular momentum of the satellite.

[0032] Calculate the model parameter values The optimal value. Combined with... Figure 2 It can be proven that the peak period of the satellite's cumulative angular momentum is... By measuring the interval between the two peaks of the flywheel speed curve, it is possible to detach from the navigation system, obtain the satellite orbit period, and then completely detach from the navigation system to complete the identification of environmental torque parameters.

[0033] In this embodiment, the solar incidence angle measurement is obtained by reading data from the solar sensor.

[0034] As an example, in unloading mode or cruise mode, the control quantity is solved using the interior point method of the obstacle function.

[0035] The optimal angular momentum management method in this embodiment is as follows: After completing orbital maneuvers, the satellite automatically enters angular momentum management mode. First, it enters the autonomous parameter identification process. The satellite continuously rotates its attitude offset angle and solar panel rotation angle according to the preset attitude trajectory, accumulating flywheel speed data under different control states. Through parameter identification methods based on environmental torque curve fitting or single-orbit periodic angular momentum accumulation, the unknown parameters of the environmental torque system are determined.

[0036] After parameter identification is completed, the satellite reads the current flywheel speed to obtain the total angular momentum accumulation. First, it decides whether to enter unloading mode based on the allowable angular momentum range. If the accumulation is low, it enters cruise mode and optimizes the satellite attitude and solar panel angle within the allowable control range using the fully identified angular momentum accumulation function. If the accumulation is high, unloading is required, and the satellite control state with the highest efficiency in unloading the current total angular momentum is optimized. After maintaining the above attitude offset and solar panel angle for one orbital cycle, the satellite reads the sun sensor data once to calculate the current solar incidence angle in the inertial frame. If the difference between this angle and the incidence angle at the time of parameter identification exceeds the allowable range, the relevant values ​​in the environmental moment parameters are recalculated using the current incidence angle, and the system parameters are updated. Other system parameters are slowly varying and do not require continuous re-identification in a short period. If the difference is small, the operating mode is determined to decide the operating mode for the next orbital cycle. After activating the angular momentum management mode, the satellite continuously switches between cruise mode and unloading mode, achieving long-term autonomous angular momentum management completely independent of ground control, which significantly reduces fuel consumption in the propulsion system.

[0037] Application scenarios of this invention: 1. Angular momentum management of high Earth orbit satellites: For probes in geosynchronous orbits, rapid revisit orbits, and other orbits with extremely high altitudes, where satellites are far from Earth and it is difficult to unload magnetic and aerodynamic torques, it is necessary to effectively utilize the two main environmental torques, gravity gradient and solar radiation pressure, to manage angular momentum, minimize the number of jet unloading operations, and extend the satellite's lifespan.

[0038] 2. Long-term autonomous survival of deep space probes: Most celestial bodies in the solar system lack magnetic fields, have thin or no atmospheres, making it impossible to unload magnetic and aerodynamic torques. Furthermore, their vast distances from Earth result in high communication delays and scarce ground control resources. The angular momentum autonomous management method of this invention can significantly reduce the number of angular momentum jet unloading operations, save fuel for the attitude and orbit control system, effectively extend mission duration, and reduce reliance on ground-based telemetry and control.

[0039] 3. Ultra-high precision formation flight orbit maintenance: Most multi-satellite serial formation flight missions rely on weak J2 gravitational perturbations of the central celestial body to maintain the formation orbit. Jet angular momentum unloading will generate unplanned orbital control forces, reducing formation maintenance accuracy and even causing the formation to diverge. The angular momentum management method of this invention does not require jet propulsion, does not generate orbital control forces, and can maintain the orbital accuracy of multi-satellite formations.

[0040] This invention proposes a method for predicting the cumulative integral of angular momentum: the theoretical integral of the cumulative angular momentum of environmental torque within one orbital period of a circular orbital satellite is derived, and a variable step-size numerical integration method is proposed for the cumulative angular momentum of an elliptical orbital satellite within one orbital period. Short step-size integration is adopted in the pericentric arc segment where the torque is large, changes drastically, but the time is short, while long step-size integration is adopted in the distal arc segment where the torque is small, changes slowly, and the time is long.

[0041] This invention also proposes a self-balancing angular momentum management method for highly elliptical orbits within an orbital period: For satellites in highly elliptical orbits, control quantity A is applied in the pericentric arc segment where the gravitational gradient torque is large, and control quantity B is applied in the apocentric arc segment where the gravitational gradient torque is almost non-existent, so that the angular momentum accumulated in the pericentric and apocentric arc segments within one orbital period is balanced with each other, and no long-term angular momentum accumulation occurs.

[0042] Autonomous angular momentum management workflow: The satellite, based on a set of predetermined program logic and the current angular momentum accumulation state, autonomously makes decisions to perform angular momentum management actions such as parameter identification, accumulation suppression, and momentum unloading, and optimizes the calculation of control quantities for the angular momentum accumulation function, thus achieving autonomous health management without ground control.

[0043] Experimental verification: The effectiveness of the method of this invention is verified by two typical deep space micro-nano satellite mission scenarios: lunar low-orbit orbit exploration and Mars high-orbit communication relay.

[0044] I. Lunar low-orbit orbit exploration mission: Lunar low-Earth orbit exploration is a key objective of microsatellite deep space exploration missions. The satellite simulation conditions are as follows: Table 1 Simulation parameters of lunar satellite orbit .

[0045] Table 2 Parameters for Modeling the Surface of Lunar Satellites .

[0046] Matrix of rotational inertia in satellite body coordinate system for: ; Environmental torque parameters are identified using a parameter identification method based on environmental torque difference estimation. The identification results are input into the controller. The lunar satellite angular momentum management results within 180 days are as follows: Figure 4 and Figure 5 As shown.

[0047] The autonomous angular momentum management controller uses only environmental torque, without employing any active control torque, to maintain the satellite's overall angular momentum below the unloading threshold over a long period. This ensures the normal operation of the attitude control system and avoids fuel consumption. The comparison of the satellite's accumulated angular momentum before and after activating the angular momentum management mode is as follows: Figure 6 As shown.

[0048] If the angular momentum management mode is not used, and the system remains fully oriented toward the sun while continuously unloading angular momentum using the jet propulsion system, 11 jet unloading events will be triggered within 30 days, unloading 22 Nms of angular momentum. Once the management mode is enabled, there is no need to initiate jet unloading within 30 days, and the angular momentum can be unloaded entirely by relying on the ambient torque.

[0049] To utilize environmental torque, the angular momentum management mode controls the satellite's attitude offset and solar panel rotation angle, as shown in the control curve. Figure 7 and Figure 8 As shown. The satellite solar panel has a large rotation angle control range, a small attitude deflection angle control range, and a long change period, resulting in a lower load on the attitude maneuvering system. The satellite solar panel charging efficiency is as follows: Figure 9 As shown, the charging power is maintained at around 50% of the maximum power for a long period, only decreasing to 30% during a few periods of angular momentum unloading, which meets the requirements for normal on-orbit operation.

[0050] II. Mars High-Orbit Communication Relay Mission: Communication relay for deep space probes is one of the main challenges of deep space exploration missions. Relay satellites require high data transmission power, high orbital altitude, and long mission cycles, necessitating fuel conservation and extended design lifespan. The simulation parameters for the Mars communication relay small satellite are as follows: Table 3. Mars satellite orbit simulation parameters .

[0051] Table 4 Modeling parameters for the surface of Mars' moons .

[0052] Matrix of rotational inertia in satellite body coordinate system for: , Due to the high precision requirements for angular momentum management of Mars communication relay satellites and the ample mission time, environmental parameters can be accurately identified using a parameter identification method based on the accumulation of angular momentum over a single orbital period. The results of the satellite's autonomous angular momentum management are as follows: Figure 10 and Figure 11 As shown. The high-orbit environment of Mars has a significant impact on torque; without effective management of angular momentum, the attitude control system's fuel consumption is rapid. When the Mars satellite is kept completely oriented towards the sun and the angular momentum management mode is not activated, the satellite's angular momentum accumulation is as follows. Figure 12 As shown, when the Mars satellite is fully oriented towards the Sun, it initiates 12 angular momentum jet unloading maneuvers within 6 days, unloading a total of 36 Nms of angular momentum. The fuel consumption is 200 times that under angular momentum management mode, which seriously affects the satellite's on-orbit lifespan.

[0053] The communication relay satellite is in an inclined high orbit 5000km above the Earth's surface. The gravitational gradient torque is too small compared to the solar pressure torque, preventing the generation of an unloading torque along the Z-axis in the inertial frame within the controllable range. It can only suppress accumulated velocity. Therefore, in angular momentum management mode, whenever the accumulated Z-axis angular momentum reaches a threshold, a jet propulsion unloading is initiated regardless of whether the satellite is in unloading or cruise mode, clearing the Z-axis angular momentum to zero. Jet propulsion was initiated three times within 250 days, unloading 7.5 NmS of angular momentum. The X and Y axis angular momentum can be unloaded solely by environmental torque, and the overall satellite angular momentum remains below the unloading threshold for an extended period. Satellite attitude, solar panel control status, and charging efficiency are as follows: Figures 13 to 15 As shown, the accumulated angular momentum of the solar pressure torque on the satellite's solar panels is relatively large. The controller continuously deflects the solar panels at a large angle to reduce the solar pressure generated by the wings. The satellite's solar charging efficiency has been hovering around 20% for a long time, limiting its power output.

[0054] While the invention has been described herein with reference to specific embodiments, it should be understood that these embodiments are merely examples of the principles and applications of the invention. Therefore, it should be understood that many modifications can be made to the exemplary embodiments, and other arrangements can be designed without departing from the spirit and scope of the invention as defined by the appended claims. It should be understood that different dependent claims and features described herein can be combined in ways different from those described in the original claims. It is also understood that features described in conjunction with individual embodiments can be used in other described embodiments.

Claims

1. A method for autonomous management of angular momentum of micro-nano satellites, characterized in that Comprising, According to a satellite environment torque model, an environment torque estimation value in a satellite body coordinate system is obtained ; Estimation of environmental torques Integration and deformation are performed to obtain an accumulated estimation of the satellite angular momentum ; Extracting the independent parameters in the model initial parameters of the satellite environment torque model as the model parameters to be solved; According to the estimated value of the environmental moment The number of the comparison result of the set data amount threshold value, respectively using the environmental moment estimation value And the satellite angular momentum cumulative estimation value Calculate the value of the model parameter to be solved as the model parameter value; Entering the angular momentum management mode to switch the satellite between the unloading mode and the cruising mode, and realizing the autonomous management of the angular momentum: At the end of an orbit period, if the difference between the measured value of the solar incidence angle and the calculated value of the solar incidence angle in the current model parameter value exceeds the incidence angle threshold, the measured value of the solar incidence angle is used to replace the calculated value of the solar incidence angle to update the current model parameter value, and the updated current model parameter value is obtained; Otherwise, the current model parameter value remains unchanged; Reading the current satellite flywheel speed, and calculating the real value of the current satellite angular momentum accumulation; If the current cumulative true value of the satellite's angular momentum exceeds the set angular momentum threshold, it enters unloading mode: by changing the control variables, the estimated environmental moment value for the next orbital period is altered. The magnitude and direction of the satellite's angular momentum accumulation estimate for the next orbital period are based on the current model parameter values. The current true value of the satellite's accumulated angular momentum is offset; if the current true value of the satellite's accumulated angular momentum does not exceed the set angular momentum threshold, then cruise mode is entered: by changing the control variables, the estimated value of the satellite's accumulated angular momentum for the next orbital period is based on the current model parameter values. The modulus is the smallest; Then entering the next orbit period, and again deciding to enter the cruising mode or the unloading mode until the end.

2. The micro-nano satellite angular momentum autonomous management method according to claim 1, characterized in that, Environmental moment estimate The expression for the environmental moment estimate is: ， wherein is the roll angle, is the pitch angle, is the yaw angle, is the left solar array rotation angle, is the right solar array rotation angle, is the true anomaly, is the initial model parameter, is the solar radiation pressure torque on the satellite central body, is the solar radiation pressure torque on the left solar array, is the solar radiation pressure torque on the right solar array, is the total gravity gradient torque on the satellite.

3. The micro-nano satellite angular momentum autonomous management method according to claim 2, characterized in that, obtaining a satellite angular momentum cumulative estimate the method is The control quantity is expressed as : ， Computing satellite solar pressure moments Angular momentum accumulation over an orbital period : ， In the formula is the orbital period, is the time; wherein is the orbital angular velocity; ， wherein is an orbital period integral of the satellite's gravity gradient torque is the accumulated value of the generated angular momentum, is the gravitational parameter of the central body around which the satellite orbits, is the distance of the satellite's center of mass from the central body's center of mass, is the unit vector pointing from the satellite's center of mass to the central body's center of mass, is the matrix of the satellite's moment of inertia in the satellite's body coordinate system; , , , , , is an intermediate variable: ， wherein is the product of inertia of the satellite along the YZ axis, is the product of inertia of the satellite along the XZ axis, is the product of inertia of the satellite along the XY axis, is the principal moment of inertia of the satellite along the X axis in the body coordinate system, is the principal moment of inertia of the satellite along the Y axis in the body coordinate system, is the principal moment of inertia of the satellite along the Z axis in the body coordinate system. , , , , , is an intermediate variable; ， wherein , , , , , are intermediate variables, ， wherein is the solar incidence angle calculated value, is the orbital inclination; , , , , , , , , are the nine elements of the coordinate rotation matrix from the inertial frame to the body frame for the satellite: ， wherein is a coordinate rotation matrix from the body frame to the inertial frame defined in the X-Y-Z axis rotation sequence, is a coordinate transformation matrix for the yaw angle is a coordinate transformation matrix for the roll angle is a coordinate transformation matrix for the pitch angle is a coordinate transformation matrix for the roll angle is a coordinate transformation matrix for the roll angle is a coordinate transformation matrix for the roll angle The all orbital angular velocities in the expression eliminate, to get a satellite angular momentum cumulative estimate within an orbit period : 。 4. The micro-nano satellite angular momentum autonomous management method according to claim 3, characterized in that, The method for extracting the independent parameters in the model initial parameters of the satellite environment torque model is: environmental moment estimate In the expression of the satellite central tank solar pressure moment The mathematical model is: ， In the formula P is the solar pressure intensity, A is the effective area of the satellite central box, a is the absorption coefficient of the surface material of the satellite central box, r is the diffuse reflection coefficient of the surface material of the satellite central box, Z is the Z-axis component of the solar pressure pressure center position vector of the satellite central box in the satellite body coordinate system, Y is the Y-axis component of the solar pressure pressure center position vector of the satellite central box in the satellite body coordinate system, X is the X-axis component of the solar pressure pressure center position vector of the satellite central box in the satellite body coordinate system, m is the mirror reflection coefficient of the surface material of the satellite central box; θ is the incident angle of the sunlight on the satellite central box. ； Extracting solar pressure torques on a satellite central tank the quantities related only to the objective conditions in the mathematical model and are expressed as intermediate variables , , , , : ， According to the digital relationship: ， Obtained: ， Thus determining , , , are independent parameters; Similarly, 8 independent parameters related to the torque of the whole satellite caused by the gravity gradient force 6 independent parameters related to the torque of the left solar array caused by the solar radiation pressure 6 independent parameters related to the torque of the right solar array caused by the solar radiation pressure 18 independent parameters in total, as the model parameters to be solved.

5. The micro-nano satellite angular momentum autonomous management method according to claim 4, characterized in that, When the number of collected environment moment estimation values is less than a set data amount threshold value, an environment moment estimation value is calculated, with the objective of minimizing the error square sum between the environment moment estimation value and the real environment moment value.

6. The micro-nano satellite angular momentum autonomous management method according to claim 5, characterized in that, When the number of collected environment moment estimation values is less than a set data amount threshold value, the method for calculating the model parameter value is: ， wherein is the i-th environment moment estimate is the corresponding control quantity, is the i-th environment moment estimate is the corresponding true anomaly, is the model parameter value, is the i-th environment moment estimate is the corresponding environment moment true value, N is the number of environment moment estimates of the environment moment estimates.

7. The micro-nano satellite angular momentum autonomous management method according to claim 6, characterized in that, When the number of collected environment moment estimation values is not less than a set data amount threshold, so that the satellite angular momentum cumulative estimation value is closest to the satellite angular momentum cumulative true value, the value of the model parameter to be solved is calculated.

8. The micro-nano satellite angular momentum autonomous management method according to claim 7, characterized in that, When the number of collected environment moment estimation values is not less than a set data amount threshold value, the method for calculating the model parameter value is: ， Cumulative estimate of the angular momentum of the i-th satellite Corresponding cumulative true value of the angular momentum of the satellite.

9. The micro-nano satellite angular momentum autonomous management method according to claim 1, characterized in that, The measured value of the solar incidence angle is obtained by reading the data of the solar sensor.

10. The micro-nano satellite angular momentum autonomous management method according to claim 1, characterized in that, In the unloading mode or the cruising mode, the control quantity is solved by the barrier function interior point method.

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