Micro-nano satellite angular momentum autonomous management method

Abstract

The application discloses a kind of micro-nano satellite angular momentum autonomous management method, belongs to satellite attitude control technical field.The application is in view of the problem that existing satellite angular momentum management adopts open loop method, and autonomous correction capability is poor.The application comprises: satellite is identified through on-orbit flywheel rotating speed data environment torque parameter autonomously, and parameter that changes speed is faster is updated autonomously along with task time extension;Only through the attitude bias of relative sun orientation and sail rotation, angular momentum accumulation inhibition and angular momentum unloading two task objectives can be realized by manipulating environment torque, without active control torque intervention such as magnetic torque device, jet unloading;The whole process does not need artificial intervention, and angular momentum autonomous management is realized.The application can be autonomously carried out when satellite has no task, and reliability is good.

Description

TECHNICAL FIELD

[0001] The present application relates to a micro-nano satellite angular momentum autonomous management method, belonging to the technical field of satellite attitude control. BACKGROUND

[0002] The existing satellites in on-orbit mission tend to use simple methods such as spinning, fixed attitude bias, angular momentum bias, etc. to manage angular momentum, in order to delay the flywheel saturation speed and unload to some extent. The existing inertia angular momentum bias based on Fengyun No. 4 remote sensing satellite and the zero-space angular momentum bias of the six-skew redundant flywheel system effectively avoid flywheel zero and saturation. By manipulating the Chang'e No. 2 active rotating solar panel and spinning, generating turbine torque, the accumulation of angular momentum on the X-axis of the satellite body is effectively slowed down; on the basis of the theoretical analysis of the gravity gradient torque of Chang'e No. 5, combined with the satellite pitch axis attitude bias scheme designed by the telemetry data, the angular momentum unloading period is prolonged. The autonomous angular momentum management scheme of Tianwen No. 2 lunar communication relay satellite uses the gravity gradient torque to manage the satellite angular momentum by switching between three attitude biases, without effectively utilizing the solar radiation pressure torque. These angular momentum management methods can only slow down the accumulation of satellite angular momentum, and the flywheel will still be saturated slowly when the magnetic torque device is not used; according to different task scenarios, the angular momentum needs to be unloaded once every few days to several weeks through the propulsion system.

[0003] In terms of angular momentum unloading, the existing space station phased angular momentum management scheme is one of the torque balance attitude tracking modes, which uses the gravity gradient torque to unload the angular momentum accumulated by the CMGs (control moment gyro) of the space station in the attitude stabilization mode; the second is to deflect the solar panel of the satellite to use the aerodynamic torque to manage the angular momentum of the LEO (low earth orbit) satellite; the third is to use the gravity gradient torque and the solar panel aerodynamic torque for angular momentum management based on the decision of the angular momentum unloading signal generator of the strategy theme method considering the task constraints; the fourth is to use the interplanetary magnetic field predicted on the satellite based on the unscented Kalman filter to realize the magnetic torque unloading during the orbit transfer.

[0004] The above research on angular momentum management using environmental torques is mostly an open-loop method, which relies heavily on ground control ability and needs accurate modeling and parameter estimation of environmental torques, cannot be corrected on-orbit, and has poor autonomy. Some methods do not deviate from active management methods such as magnetic torque devices, and environmental torques only play an auxiliary role. Most of the methods do not involve angular momentum unloading, but only slow down the accumulation of angular momentum to prolong the unloading interval. The research involving angular momentum unloading mostly uses traditional control methods, which have complex controller structures, high computational power requirements, and lack of practical task verification. SUMMARY

[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 the application realizes complete utilization of angular momentum management of environmental torque: without jet unloading, without using magnetic torque, without adding additional actuators, the satellite only needs to manipulate its attitude deflection angle and sailboard rotation angle through the flywheel system and SADA (solar sail drive mechanism), and to use the most common gravity gradient and solar light pressure torque, so as to realize efficient angular momentum management.

[0009] Optimal angular momentum management based on angular momentum accumulation function: without designing complex filters and control laws, only a limited number of parameters in the simple environmental torque function need to be identified, so as to efficiently predict the angular momentum accumulation in the next orbit period, and to calculate the most efficient attitude deflection angle and sailboard rotation angle for angular momentum management by using a reliable convex optimization method.

[0010] High autonomy and reliability: the angular momentum management program does not need ground flight control personnel intervention at all, can be autonomously performed when the satellite is not in task, and does not use immature mathematical methods with poor robustness, does not manipulate the thruster and other satellite equipment that can cause serious damage, and has good reliability. BRIEF DESCRIPTION OF DRAWINGS

[0011] Figure 1 is a flowchart of the micro-nano satellite autonomous angular momentum management method described in the application;

[0012] Figure 2 is an orbit period identification schematic diagram based on the angular momentum accumulation curve;

[0013] Figure 3 is a schematic diagram of the division of the near-periapsis arc segment and the far-periapsis arc segment of the elliptical orbit;

[0014] Figure 4 is an angular momentum accumulation schematic diagram of the 180-day lunar orbit satellite in the lunar low-orbit circumlunar exploration mission;

[0015] Figure 5 is a three-axis angular momentum accumulation schematic diagram of the 180-day lunar orbit satellite in the lunar low-orbit circumlunar exploration mission;

[0016] Figure 6 is a comparison schematic diagram of the angular momentum accumulation of the 30-day lunar orbit satellite;

[0017] Figure 7 is a satellite attitude deflection angle schematic diagram output by the 180-day lunar orbit satellite angular momentum management working mode;

[0018] Figure 8 is a solar sailboard rotation angle schematic diagram output by the 180-day lunar orbit satellite angular momentum management working mode;

[0019] Figure 9 is a 180-day lunar orbit satellite charging efficiency schematic diagram;

[0020] Figure 10 is a schematic diagram of the angular momentum accumulation of the Mars communication relay satellite in 250 days in the Mars high-orbit communication relay mission;

[0021] Figure 11 is a schematic diagram of the three-axis angular momentum accumulation of the Mars communication relay satellite in 250 days in the inertial system;

[0022] Figure 12 is a schematic diagram of the comparison of the angular momentum accumulation of the Mars high-orbit satellite in 6 days;

[0023] Figure 13 is a schematic diagram of the satellite attitude offset angle output by the angular momentum management working mode of the Mars satellite in 250 days;

[0024] Figure 14 is a schematic diagram of the solar panel rotation angle output by the angular momentum management working mode of the Mars satellite in 250 days;

[0025] Figure 15 is a schematic diagram of the charging efficiency of the Mars satellite in 250 days. DETAILED DESCRIPTION

[0026] The technical solutions in the embodiments of the present application will be clearly and completely described below with reference to the drawings in the embodiments of the present application. Obviously, the described embodiments are only part of the embodiments of the present application, rather than all the embodiments. Based on the embodiments in the present application, all other embodiments obtained by those of ordinary skill in the art without creative work fall within the scope of protection of the present application.

[0027] DETAILED DESCRIPTION Figure 1 , the present application provides a micro-nano satellite angular momentum autonomous management method, which comprises the following steps: obtaining an environment torque estimation value in a satellite body coordinate system according to a satellite environment torque model ; integrating and deforming the environment torque estimation value to obtain a satellite angular momentum accumulation estimation value ; extracting an independent parameter in a model initial parameter of the satellite environment torque model as a to-be-solved model parameter; comparing the number of the collected environment torque estimation values with a set data amount threshold value, and respectively using the environment torque estimation value and the satellite angular momentum accumulation estimation value 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.

[0028] 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.

[0029] Furthermore, environmental moment modeling and angular momentum integral prediction for micro / nano satellites:

[0030] 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:

[0031] ,

[0032] 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.

[0033] 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.

[0034] The cumulative estimate of satellite angular momentum was obtained. The method is to express the control quantity as :

[0035] ,

[0036] 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.

[0037] Satellite solar pressure torque and control quantity Related to calculating satellite solar pressure torque Accumulated angular momentum generated over one orbital period :

[0038] ,

[0039] In the formula For orbital period, For time;

[0040] 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.

[0041] 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:

[0042] In the formula Let be the orbital angular velocity; then the integral relationship is:

[0043] ,

[0044] In the formula The gravitational gradient torque of the entire satellite over one orbital period The cumulative value of angular momentum generated, G is the gravitational parameter of the central body around which the satellite orbits, R is the distance of the satellite from the center of mass of the central body, is the unit vector pointing from the center of mass of the satellite to the center of mass of the central body, is the matrix of the moments of inertia in the body coordinate system of the satellite; , , , , , are intermediate variables:

[0045] ,

[0046] where is the product of inertia about the YZ axis of the satellite, is the product of inertia about the XZ axis of the satellite, is the product of inertia about the XY axis of the satellite, is the principal moment of inertia about the X axis in the body coordinate system of the satellite, is the principal moment of inertia about the Y axis in the body coordinate system of the satellite, is the principal moment of inertia about the Z axis in the body coordinate system of the satellite;

[0047] , , , , , are intermediate variables;

[0048] ,

[0049] where , , , , , are intermediate variables,

[0050] ,

[0051] where is the calculated value of the solar incidence angle, is the orbit inclination; , , , , , , , , are the nine elements of the coordinate rotation matrix from the inertial system to the body coordinate system of the satellite:

[0052] ,

[0053] where is the coordinate rotation matrix from the inertial frame to the body frame defined in the X-Y-Z rotation sequence, is the yaw angle is the coordinate transformation matrix, is the pitch angle is the coordinate transformation matrix, is the roll angle is the coordinate transformation matrix.

[0054] Combining Figure 2 , all the orbital angular velocities in the expression of are eliminated and become . After the integration is completed, only relates to the system parameters, the attitude bias angle, and the orbital period. Further, an estimation of the satellite angular momentum accumulation in an orbital period is obtained:

[0055] .

[0056] For a satellite in an elliptical orbit, the orbital period relationship does not exist in , and in the expression of is a variable related to the orbital time that cannot be described by an analytical mathematical formula, so the estimation of the satellite angular momentum accumulation in an orbital period cannot be completed by theoretical integration. The orbital period of most elliptical orbits is long, and if the angular momentum numerical integration of the entire orbital period is directly completed with a fixed time step, the number of steps to be calculated is too large, and the Kepler problem of finding r for each step requires solving the flight time, which is difficult to calculate on the satellite. In this embodiment, one orbital period is divided into 2000-5000 orbital solving steps according to the length of the period, and the true anomaly angle λ and the geocentric radius (the distance from the satellite to the center of the central celestial body) r at the end of each step are calculated in advance with high precision and stored in a file, ready for table lookup and interpolation when performing angular momentum integration.

[0057] Figure 3As shown, according to Kepler's law, the elliptical orbit satellite passes through the vicinity of the pericenter at the fastest speed, and the time required for the pericenter angle to pass from -45° through the pericenter (0°) to +45° is only 1 / 10 of the orbital period or less, and the change speed of r is the fastest during this period, and the change speed of the gravity gradient torque is also the fastest, so a shorter step should be taken to perform the angular momentum accumulation integral. Since the arc segment passes through in a short time, even if a very short integral step is taken, the total integral step number is still within an acceptable range. For the apocenter arc segment from +45° through the apocenter (180°) to -45°, the satellite passes through for a long time, the average angular velocity is slow, r changes slowly, and therefore a longer integral step can be taken to reduce the integral step number and speed up the numerical calculation, and the final angular momentum integral accuracy is still high.

[0058] Further, the environmental torque parameter identification method:

[0059] The method for extracting the independent parameters in the model initial parameters of the satellite environmental torque model is as follows: in the satellite environmental torque system, the quantity that cannot be changed by active maneuvering strategies such as attitude bias and sail rotation and is independent of time is the model parameter.

[0060] The environmental torque estimation value The mathematical model of the solar radiation pressure torque experienced by the satellite central box is:

[0061] ,

[0062] In the formula, p is the solar radiation pressure, is the effective sunlit area of the satellite central box, is the absorption coefficient of the surface material of the satellite central box, is the diffuse reflection coefficient of the surface material of the satellite central box, is the Z-axis component of the solar radiation pressure pressure center position vector of the satellite central box in the satellite body coordinate system, is the Y-axis component of the solar radiation pressure pressure center position vector of the satellite central box in the satellite body coordinate system, is the X-axis component of the solar radiation pressure pressure center position vector of the satellite central box in the satellite body coordinate system, is the mirror reflection coefficient of the surface material of the satellite central box; is the incident angle of the solar radiation on the sunlit plane of the satellite central box:

[0063] ;

[0064] The mathematical model of the solar radiation pressure torque experienced by the satellite central box only contains quantities related to objective conditions, and is expressed as an intermediate variable ,​ , , , :

[0065] Based on numerical relationships:

[0066] ,get:

[0067] Thus determine , , , These are independent parameters; These are derived parameters that can be obtained from other parameters.

[0068] 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.

[0069] 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.

[0070] 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.

[0071] 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.

[0072] 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:

[0073] ,

[0074] 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.

[0075] 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.

[0076] 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:

[0077] ,

[0078] The cumulative estimate of the angular momentum of the i-th satellite The corresponding true value of the cumulative angular momentum of the satellite.

[0079] 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.

[0080] In this embodiment, the solar incidence angle measurement is obtained by reading data from the solar sensor.

[0081] As an example, in unloading mode or cruise mode, the control quantity is solved using the interior point method of the obstacle function.

[0082] 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.

[0083] 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.

[0084] Application scenarios of this invention:

[0085] 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.

[0086] 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.

[0087] 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.

[0088] 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.

[0089] 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.

[0090] 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.

[0091] 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.

[0092] I. Lunar low-orbit orbit exploration mission:

[0093] Lunar low-Earth orbit exploration is a key objective of microsatellite deep space exploration missions. The satellite simulation conditions are as follows:

[0094] Table 1 Simulation parameters of lunar satellite orbit

[0095] .

[0096] Table 2 Parameters for Modeling the Surface of Lunar Satellites

[0097] .

[0098] Matrix of rotational inertia in satellite body coordinate system for:

[0099] ;

[0100] 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.

[0101] 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.

[0102] 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.

[0103] 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.

[0104] 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:

[0105] Table 3. Mars satellite orbit simulation parameters

[0106] .

[0107] Table 4 Modeling parameters for the surface of Mars' moons

[0108] .

[0109] Matrix of rotational inertia in satellite body coordinate system for:

[0110] ,

[0111] 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 operations 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.

[0112] 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 solar pressure torque angular momentum accumulated 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.

[0113] 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... include, Based on the satellite environmental moment model, the estimated environmental moment values ​​in the satellite body coordinate system are obtained. ; Estimated values ​​of environmental torque By integrating and transforming, we obtain the cumulative estimate of the satellite's angular momentum. ; Independent parameters are extracted from the initial parameters of the satellite environmental moment model and used as the parameters of the model to be solved. Based on the collected environmental torque estimates 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 Calculate the values ​​of the parameters of the model to be solved and use them as model parameter values; Entering the angular momentum management operating mode enables the satellite to switch between unloading mode and cruise mode, achieving autonomous angular momentum management: At the end of an orbital period, 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 is used to replace the calculated solar incidence angle to update the current model parameter values, resulting in the updated current model parameter values. Otherwise, the current model parameter values ​​remain unchanged; Read the current satellite flywheel speed and calculate the current true value of the satellite's cumulative angular momentum; 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 it enters the next orbital cycle, and decides again whether to enter cruise mode or unload mode, until the end.

2. The method for autonomous management of angular momentum of micro / nano satellites according to claim 1, characterized in that, Environmental torque estimates 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.

3. The method for autonomous management of angular momentum of micro / nano satellites according to claim 2, characterized in that, The cumulative estimate of satellite angular momentum was obtained. The method is as follows: The control quantity is expressed as : ， Calculate the solar radiation pressure torque of the satellite Accumulated angular momentum generated over one orbital period : ， In the formula For orbital period, For time; In the formula It is the orbital angular velocity; ， 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 Y and Z 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; Will All orbital angular velocities in the expression Eliminating these values ​​yields an estimate of the cumulative angular momentum of the satellite over one orbital period. : 。 4. The method for autonomous management of angular momentum of micro / nano satellites according to claim 3, characterized in that, The method for extracting independent parameters from the initial parameters of the satellite environmental moment model is as follows: 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. This represents 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 numerical relationships: ， get: ， Thus determine , , , These are independent parameters; 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.

5. The method for autonomous management of angular momentum of micro / nano satellites according to claim 4, characterized in that, 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.

6. The method for autonomous management of angular momentum of micro / nano satellites according to claim 5, characterized in that, When the collected environmental torque estimate 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.

7. The method for autonomous management of angular momentum of micro / nano satellites according to claim 6, characterized in that, 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.

8. The method for autonomous management of angular momentum of micro / nano satellites according to claim 7, characterized in that, When the collected environmental torque estimate 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.

9. The method for autonomous management of angular momentum of micro / nano satellites according to claim 1, characterized in that, The solar incidence angle is measured by reading data from the solar sensor.

10. The method for autonomous management of angular momentum of micro / nano satellites according to claim 1, characterized in that, In unloading mode or cruise mode, the control quantity is solved using the interior point method of the obstacle function.

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