A method and device for predicting angular momentum of a lunar orbit satellite and an electronic device

Abstract

A lunar orbit satellite angular momentum prediction method, device and electronic equipment. The method comprises: acquiring satellite historical angular momentum and attitude quaternion telemetry data, dividing the angular momentum into multiple periods with the rapid change process as a node, and performing polynomial fitting on the angular momentum and attitude quaternion in each period to complete and smooth the missing data due to the control constraint. Based on the completed historical angular momentum and attitude quaternion, the historical angular momentum increment of the satellite in the system is calculated. Based on the satellite ephemeris, the historical angle between the body axis with the largest solar pressure area of the satellite and the sun position vector in the system is calculated. Using a preset order sine and cosine model, the historical angle and its corresponding historical angular momentum increment are fitted to obtain the fitting coefficients of the three-axis directions in the system, and an angular momentum increment fitting model is constructed. The model is used to predict the angular momentum prediction value of the satellite in the system, and the accurate prediction of the angular momentum is realized.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of aerospace technology, and particularly relates to a method and device for predicting angular momentum of a lunar-orbit satellite and an electronic device. BACKGROUND

[0002] The attitude control of a satellite is mainly achieved by rotating a momentum wheel driven by a motor, and then exchanging angular momentum with the satellite. However, the rotation speed of the momentum wheel is not infinite, and once the critical value is reached, the initial rotation speed needs to be restored through unloading operation. Although the speed increment generated by the unloading of the momentum wheel is relatively limited, in some subtle and sensitive three-body orbit environment in the earth-moon space, such a small change can also cause an unignorable divergence effect of the orbit. Therefore, it is particularly important to improve the accuracy of the satellite angular momentum prediction. SUMMARY

[0003] Embodiments of the present application provide a method and device for predicting angular momentum of a lunar-orbit satellite and an electronic device, which are used to improve the accuracy of the satellite angular momentum prediction.

[0004] In a first aspect, embodiments of the present application provide a method for predicting angular momentum of a lunar-orbit satellite, which comprises:

[0005] Obtaining a telemetry value of angular momentum of a satellite three-axis in a body system at a prediction start time and a satellite attitude quaternion, calculating a conversion matrix of the satellite from an inertial system to the body system based on the satellite attitude quaternion, the satellite attitude quaternion being used to describe the satellite attitude, and elements in the conversion matrix representing projection components of unit vectors in the inertial system in the body system;

[0006] Obtaining an angular momentum vector of the satellite in the body system based on the telemetry value of angular momentum and a reference angular momentum, determining an angular momentum vector of the satellite in the inertial system through coordinate transformation based on the conversion matrix and the angular momentum vector in the body system, and the reference angular momentum representing a reference value of the angular momentum three-axis direction when the angular momentum of the satellite is zero after unloading;

[0007] Inputting a solar direction angle into an angular momentum increment fitting model to obtain an angular momentum increment of the satellite in the body system, the solar direction angle representing an angle between the axis with the largest solar pressure area of the satellite in the body system and a solar position vector, and the angular momentum increment fitting model being obtained by fitting a corresponding relationship between historical solar direction angles and historical angular momentum increments, and the axis with the largest solar pressure area representing a coordinate axis with the largest solar pressure area of the satellite in the body system;

[0008] Determining a next-step angular momentum prediction value of the satellite in the body system based on the angular momentum increment in the body system, the conversion matrix, the telemetry value of angular momentum and the reference angular momentum, and the next-step angular momentum prediction value being positively correlated with the angular momentum increment in the body system if the conversion matrix and the reference angular momentum are unchanged.

[0009] In the above method, the angular momentum increment of the satellite can be more accurately predicted by fitting the corresponding relationship between the historical sun direction angle and the historical angular momentum increment, the prediction accuracy of the predicted value of the angular momentum of the next step length of the satellite in the body system is improved, and more accurate data support is provided for subsequent complex task execution of the satellite. By obtaining the angular momentum telemetry value of the satellite in the body system at the prediction start time, the current angular momentum state of the satellite can be understood in real time. By the reference angular momentum, it can be ensured that when the satellite is unloaded and the angular momentum of the satellite is zero, the indicated value of the angular momentum three-axis direction is accurate. By calculating the conversion matrix of the satellite converted from the inertial system to the body system, the unit vector in the inertial system can be projected into the body system to realize coordinate transformation. By fitting the corresponding relationship between the historical sun direction angle and the historical angular momentum increment, the future angular momentum increment of the satellite can be more accurately predicted. The angular momentum increment fitting model and the real-time sun direction angle data are used to calculate the angular momentum increment, so that the angular momentum increment fitting model can be predicted in real time, and the accuracy of the predicted value of the angular momentum of the next step length of the satellite in the body system is improved.

[0010] Optionally, the sun direction angle is obtained in the following manner:

[0011] The sun position vector of the sun in the inertial system at the prediction start time is calculated by searching the planet ephemeris.

[0012] The satellite ephemeris data of the satellite is obtained from the database to obtain the satellite position vector of the satellite in the inertial system at the prediction start time.

[0013] The relative position vector of the sun with respect to the satellite in the inertial system is determined based on the vector difference between the sun position vector and the satellite position vector.

[0014] The relative position vector of the sun with respect to the satellite in the body system is determined by coordinate transformation based on the conversion matrix and the relative position vector in the inertial system.

[0015] The ratio between the component of the relative position vector in the body system in the axis with the largest solar pressure area and the module of the relative position vector in the body system is subjected to inverse cosine function operation to obtain the sun direction angle.

[0016] In the above method, high-precision sun position information can be obtained by searching the planet ephemeris. The high accuracy of the satellite position vector is ensured by obtaining the satellite ephemeris data, thereby ensuring the accuracy of the relative position vector. The accurate sun direction angle is determined by the relative position vector, which can improve the accuracy of the predicted value of the angular momentum of the next step length of the satellite in the body system in subsequent prediction.

[0017] Optionally, the angular momentum increment fitting model is fitted in the following manner:

[0018] In the current system, for each historical sun direction angle in a historical sun direction angle set, a preset sine model of different order is used to fit the historical angular momentum increment corresponding to the historical sun direction angle in the current system, to obtain fitting coefficients of the satellite in three-axis directions, and the historical sun direction angle set includes the historical angle between the sun position vector and the axis with the largest solar pressure area of the satellite at a time node corresponding to each step in a historical time period in the current system;

[0019] An angular momentum increment fitting model is constructed based on the fitting coefficients.

[0020] In the above method, by fitting each historical sun direction angle in a historical time period with the historical angular momentum increment corresponding to the historical sun direction angle in the current system, the model can better learn the rules between the historical sun direction angle and the historical angular momentum increment, thereby enhancing the generalization ability of the model.

[0021] Optionally, the historical angular momentum increment corresponding to the historical sun direction angle in the current system is obtained in the following manner:

[0022] A historical telemetry data set is obtained, and the historical telemetry data set is standardized, the historical telemetry data set including a plurality of satellite time in the current system of the satellite in a historical time period, historical angular momentum telemetry values corresponding to the satellite time, and historical satellite attitude quaternions corresponding to the satellite time;

[0023] For each time in the standardized historical telemetry data set, a vector difference between the historical angular momentum telemetry value corresponding to the time and a reference angular momentum is calculated to obtain an instantaneous angular momentum of the satellite in the current system at the time;

[0024] Based on the conversion matrix of the satellite from the inertial system to the current system at each time and the instantaneous angular momentum, the instantaneous angular momentum of the satellite in the inertial system is determined through coordinate transformation, and the conversion matrix of the satellite from the inertial system to the current system at each time is obtained by calculating the historical satellite attitude quaternion at each time;

[0025] A vector difference between the instantaneous angular momentum of the satellite in the inertial system at the next time and the instantaneous angular momentum of the satellite in the inertial system at each time is calculated to obtain the historical angular momentum increment of the satellite in the inertial system;

[0026] Based on the conversion matrix of the satellite from the inertial system to the current system at each time and the historical angular momentum increment of the satellite in the inertial system, the historical angular momentum increment of the satellite in the current system is determined through coordinate transformation.

[0027] In the above method, by standardizing the historical telemetry data set, the data can be more unified, which helps to improve the efficiency and accuracy of subsequent data processing. By calculating the vector difference between the historical angular momentum telemetry value at each time and the reference angular momentum, the instantaneous angular momentum of the satellite in the body system at each time can be obtained, which helps to more accurately understand the angular momentum state of the satellite. Using the conversion matrix, flexible coordinate transformation can be performed between the inertial system and the body system, which can analyze the angular momentum state of the satellite from multiple angles and improve the accuracy of the predicted angular momentum value of the next step of the satellite in the body system.

[0028] Optionally, the historical telemetry data set includes a historical angular momentum telemetry data set and a historical attitude quaternion telemetry data set, and the standardization processing of the historical telemetry data set specifically includes:

[0029] The related data of the angular momentum rapid change process in the historical angular momentum telemetry data set and the historical attitude quaternion telemetry data set are deleted, and the angular momentum rapid change process at least includes satellite rapid attitude adjustment, momentum wheel unloading, and orbit control.

[0030] The historical angular momentum telemetry data set and the historical attitude quaternion telemetry data set are divided into multiple angular momentum periods with the angular momentum rapid change process as a node, and each angular momentum period includes a historical angular momentum telemetry data sub-set and a historical attitude telemetry data sub-set in different time periods.

[0031] The historical angular momentum telemetry data sub-set and the historical attitude telemetry data sub-set in each angular momentum period are fitted and processed to complete and smooth the missing data points of the historical angular momentum telemetry data sub-set and the historical attitude telemetry data sub-set.

[0032] In the above method, by deleting the related data of the angular momentum rapid change process, such as the data in the period of satellite rapid attitude adjustment, momentum wheel unloading, and orbit control, the interference of special states on data analysis can be avoided, thereby improving the stability of the data. By dividing the processed data into multiple angular momentum periods, the periodic variation rule of the satellite angular momentum can be identified and analyzed, thereby improving the accuracy of subsequent data analysis. By fitting and processing, the missing data points can be completed and smoothed, making the historical telemetry data set more complete and continuous, and reducing the analysis error caused by missing or abnormal data.

[0033] Optionally, the fitting and processing of the historical angular momentum telemetry data sub-set and the historical attitude telemetry data sub-set in each angular momentum period to complete and smooth the missing data points of the historical angular momentum telemetry data sub-set and the historical attitude telemetry data sub-set specifically includes:

[0034] For each angular momentum period, a preset first-order polynomial is used to fit the components of each historical angular momentum telemetry value in the three-axis directions of the satellite in the body system to obtain first fitting coefficients, and a first fitting formula is determined based on the first fitting coefficients;

[0035] Each component of each historical satellite attitude quaternion is fitted using a preset second-order polynomial to obtain second fitting coefficients, and a second fitting formula is determined based on the second fitting coefficients;

[0036] Using the first fitting formula, based on the known data points of the historical angular momentum telemetry data subset and the preset step size, the missing data points of the historical angular momentum telemetry data subset are determined, and the completed historical angular momentum telemetry data subset is smoothed;

[0037] Using the second fitting formula, based on the known data points of the historical attitude quaternion telemetry data subset and the preset step size, the missing data points of the historical attitude quaternion telemetry data subset are determined, and the completed historical attitude quaternion telemetry data subset is smoothed.

[0038] In the above method, through polynomial fitting, the missing values in the historical angular momentum telemetry data subset and the historical attitude telemetry data subset can be completed respectively, so that the historical angular momentum telemetry data subset and the historical attitude telemetry data subset are more complete. At the same time, since the fitting process itself has the characteristics of smoothing data, it can reduce noise and outliers in the data, and improve the continuity and smoothness of the data.

[0039] Optionally, the above method for determining the next step angular momentum prediction value of the satellite in the body system based on the angular momentum increment in the body system, the conversion matrix, the angular momentum telemetry value and the reference angular momentum comprises:

[0040] The transpose of the conversion matrix is multiplied by the transpose of the angular momentum increment in the body system to obtain the angular momentum increment of the satellite in the inertial system;

[0041] Based on the angular momentum increment in the inertial system and the vector sum of the angular momentum vector in the inertial system, the next step angular momentum vector of the satellite in the inertial system is determined;

[0042] The conversion matrix is multiplied by the next step angular momentum vector in the inertial system to determine the next step angular momentum vector of the satellite in the body system through coordinate transformation;

[0043] Based on the vector sum of the next step angular momentum vector in the body system and the reference angular momentum, the next step angular momentum prediction value of the satellite in the body system is determined.

[0044] In the method, the angular momentum increment in the body system is accurately converted into the inertial system through multiplication operation of the transpose of the conversion matrix and the transpose of the angular momentum increment, consistency of the angular momentum vector between different coordinate systems is ensured, and accuracy of subsequent data analysis is improved. The angular momentum vector in the inertial system represents absolute motion state of the satellite in space, and the angular momentum vector in the body system represents motion state of the satellite relative to the body coordinate system. By knowing the angular momentum prediction values of the satellite at different time instants and in different coordinate systems, the satellite attitude and orbit can be more accurately controlled subsequently, and stable operation of the satellite is ensured.

[0045] In a second aspect, an embodiment of the present application provides an angular momentum prediction device for a lunar orbit satellite, the device comprising:

[0046] A transceiving module is configured to acquire angular momentum telemetry values of three axes of the satellite in the body system at a prediction start moment and satellite attitude quaternions, the satellite attitude quaternions being used to describe the satellite attitude.

[0047] A processing module is configured to calculate a conversion matrix of the satellite from the inertial system to the body system based on the satellite attitude quaternions, elements in the conversion matrix representing projection components of unit vectors in the inertial system in the body system.

[0048] The processing module is further configured to obtain an angular momentum vector of the satellite in the body system based on the angular momentum telemetry values and a reference angular momentum, determine an angular momentum vector of the satellite in the inertial system through coordinate transformation based on the conversion matrix and the angular momentum vector in the body system, and the reference angular momentum represents a reference value of the angular momentum three axes when the angular momentum of the satellite is zero after unloading.

[0049] The processing module is further configured to input a solar direction angle into an angular momentum increment fitting model to obtain an angular momentum increment of the satellite in the body system, the solar direction angle representing an angle between the axis with the largest solar pressure area of the satellite in the body system and a position vector of the sun, the angular momentum increment fitting model being obtained by fitting a corresponding relationship between historical solar direction angles and historical angular momentum increments, and the axis with the largest solar pressure area represents a coordinate axis with the largest solar pressure area of the satellite in the body system.

[0050] The processing module is further configured to determine a next-step angular momentum prediction value of the satellite in the body system based on the angular momentum increment in the body system, the conversion matrix, the angular momentum telemetry values and the reference angular momentum, and the next-step angular momentum prediction value is positively correlated with the angular momentum increment in the body system if the conversion matrix and the reference angular momentum are constant.

[0051] In a third aspect, an embodiment of the present application provides an electronic device, comprising a memory, a processor and a computer program stored in the memory and executable by the processor, when the computer program is executed by the processor, the processor implements any of the methods in the first aspect.

[0052] In a fourth aspect, the embodiments of the present application further provide a computer readable storage medium, wherein the computer readable storage medium stores a computer program, and the computer program is executed by a processor to implement any of the methods in the first aspect.

[0053] In a fifth aspect, the embodiments of the present application further provide a computer program product, comprising a computer program, and the computer program is executed by a processor to implement any of the methods in the first aspect.

[0054] The technical effects brought by any of the implementation manners of the second aspect to the fifth aspect can refer to the technical effects brought by the corresponding implementation manners of the first aspect, which will not be described here. BRIEF DESCRIPTION OF DRAWINGS

[0055] Figure 1 A flow chart of a method for predicting angular momentum of a lunar-orbit satellite according to an embodiment of the present application;

[0056] Figure 2 A schematic diagram of a relationship between an angular momentum increment of a satellite in an X-axis of a local system and an angle with a sun direction according to an embodiment of the present application;

[0057] Figure 3 A schematic diagram of a relationship between an angular momentum increment of a satellite in a Y-axis of a local system and an angle with a sun direction according to an embodiment of the present application;

[0058] Figure 4 A schematic diagram of a relationship between an angular momentum increment of a satellite in a Z-axis of a local system and an angle with a sun direction according to an embodiment of the present application;

[0059] Figure 5 A structural schematic diagram of an angular momentum prediction device for a lunar-orbit satellite according to an embodiment of the present application;

[0060] Figure 6 A structural schematic diagram of a control device according to an embodiment of the present application. DETAILED DESCRIPTION

[0061] In the following, some terms in the embodiments of the present application are explained to facilitate understanding by those skilled in the art.

[0062] 1. “and / or”, describing the association relationship of the associated objects, indicating that there can be three relationships, for example, A and / or B, which can represent the following three cases: A exists alone, A and B exist simultaneously, and B exists alone. The character “ / ” generally represents an “or” relationship between the associated objects before and after it.

[0063] 2、The terms "first", "second", etc. in the specification and claims of the present application and in the above-described drawings are used to distinguish similar objects, and do not necessarily have to be used to describe a specific order or sequence. It should be understood that the data thus used can be interchanged under appropriate circumstances, so that the embodiments described herein can be implemented in an order other than that illustrated or described herein.

[0064] 3、Inertial frame, also known as Inertial Frame of Reference, is a crucial concept in physics. It refers to a reference frame in which an object will remain at rest or in uniform linear motion when it is not subject to external forces. Inertial frame is also known as inertial coordinate system or simply inertial system. In an inertial frame, time flows uniformly, and space is uniform and isotropic.

[0065] 4、This system, also known as body coordinate system or satellite body coordinate system, is a reference frame established relative to the satellite itself. In this reference frame, the coordinate origin is usually chosen at the center of mass or a specific point of the satellite, and the coordinate axes are determined according to the structure, function and design requirements of the satellite. This system is mainly used to describe the attitude, orbit and motion state of the satellite itself.

[0066] 5、Satellite ephemeris data, also known as Two-Line Orbital Element (TLE), is a kind of expression invented by Celestrak in the United States to describe the position and velocity of space flying bodies (such as satellites, spacecraft, etc.), i.e. two-line orbital data system. It is based on the mathematical relationship between the six orbital parameters of Kepler's law to determine the time, coordinate, orientation, velocity and other parameters of the flying body.

[0067] 6、Satellite time, usually refers to the time record on the satellite, which may be synchronized or calibrated based on a certain time standard (such as UTC time, GPS time, etc.). In satellite systems, satellite time is used to record the running state of the satellite, data collection time and other key information.

[0068] In order to make the purpose, technical scheme and advantages of the present application clearer, the present application will be described in further detail below in conjunction with the drawings. Obviously, the described embodiments are only a part of the embodiments of the present application, not all embodiments. Based on the embodiments in the present application, all other embodiments obtained by those skilled in the art without creative labor fall within the scope of protection of the present application.

[0069] The application scenarios described in this application are for the purpose of more clearly illustrating the technical solutions of this application, and do not constitute a limitation on the technical solutions provided in this application. Those skilled in the art will understand that with the emergence of new application scenarios, the technical solutions provided in this application are also applicable to similar technical problems. In the description of this application, unless otherwise stated, "multiple" means two or more.

[0070] It is understood that this application applies to situations where the solar radiation pressure area of ​​a satellite along one axis is significantly greater than that along the other two axes within the satellite's own system. The axis with the largest solar radiation pressure area represents the coordinate axis with the largest solar radiation pressure area within the satellite's own system, and can be the X-axis, Y-axis, or Z-axis of the system. For example, this application applies to situations where the solar radiation pressure area of ​​a lunar orbiting satellite along the Z-axis is significantly greater than that along the X-axis and Y-axis within the satellite's own system. In this case, the axis with the largest solar radiation pressure area is the Z-axis.

[0071] The letters A, B, C, etc. in the formula of this application, whether appearing in upright or italic form, all have the same meaning.

[0072] It should be noted that, in the satellite's body coordinate system, the axis with the largest solar radiation pressure area can be any one of the X-axis, Y-axis, or Z-axis in this system. The following example uses the Z-axis as the axis with the largest solar radiation pressure area in this system to illustrate the embodiments of this application.

[0073] like Figure 1 The flowchart shown is a method for predicting the angular momentum of a lunar orbiting satellite according to an embodiment of this application. Specifically, it may include the following steps.

[0074] Step S101: Obtain the telemetry values ​​of the satellite's angular momentum along its three axes and the satellite's attitude quaternion at the predicted start time. Based on the satellite's attitude quaternion, calculate the transformation matrix of the satellite from the inertial frame to the local frame.

[0075] The satellite attitude quaternion is used to describe the satellite attitude. The elements in the transformation matrix represent the projected components of the unit vector in the inertial frame into this frame.

[0076] For example, suppose the satellite attitude quaternion at the start of the prediction is Q = [Q c [Q1, Q2, Q3]. Where, Q c Q1, Q2, and Q3 are components of the satellite attitude quaternion. Based on the satellite attitude quaternion, the transformation matrix M of the satellite from the inertial frame to its home frame is calculated. bi The values ​​of each element in the matrix satisfy the following formula:

[0077]

[0078] wherein M bi (0,0) represents an element of the first row and the first column in the conversion matrix. M bi (0,1) represents an element of the first row and the second column in the conversion matrix. M bi (0,2) represents an element of the first row and the third column in the conversion matrix. M bi (1,0) represents an element of the second row and the first column in the conversion matrix. M bi (1,1) represents an element of the second row and the second column in the conversion matrix. M bi (1,2) represents an element of the second row and the third column in the conversion matrix. M bi (2,0) represents an element of the third row and the first column in the conversion matrix. M bi (2,1) represents an element of the third row and the second column in the conversion matrix. M bi (2,2) represents an element of the third row and the third column in the conversion matrix.

[0079] In step S102, the angular momentum vector of the satellite in the body system is obtained based on the angular momentum remote sensing value and the reference angular momentum, and the angular momentum vector of the satellite in the inertial system is determined through coordinate transformation based on the conversion matrix and the angular momentum vector in the body system.

[0080] wherein the reference angular momentum represents the marked value of the three-axis direction of the angular momentum of the satellite when the angular momentum of the satellite after unloading is zero.

[0081] Since the satellite usually adopts a redundant design of four momentum wheels, when the overall angular momentum of the satellite returns to zero, the X, Y and Z three-axis angular momentum remote sensing values are not zero. Assuming that the marked values of the X, Y and Z three-axis angular momentum when the angular momentum of the satellite returns to zero, i.e. the reference angular momentum, are H inib =[h inibx h iniby h inibz ], the angular momentum remote sensing value at the prediction start time is H yb0 =[h ybx0 h yby0 h ybz0 ], the actual angular momentum vector of the satellite in the body system H b0 satisfies the following formula:

[0082] H b0 =H yb0 -H inib

[0083] For example, the actual angular momentum vector is (0, 0, 0). The satellite displays an angular momentum telemetry value of (0, 0, 0). But since the satellite usually adopts a redundant design of four momentum wheels, the satellite displays a value other than zero when the overall angular momentum of the satellite returns to zero, such as (1, 1, 1). The reference angular momentum is (1, 1, 1). For another example, assuming that the reference angular momentum is (2, 2, 2) and the actual angular momentum vector of the satellite is (1, 1, 1), the satellite displays an angular momentum telemetry value of (3, 3, 3).

[0084] It can be understood that if the satellite does not adopt a redundant design, the angular momentum telemetry value displayed by the satellite is the same as the actual angular momentum vector of the satellite. At this time, the reference angular momentum is (0, 0, 0).

[0085] In an alternative embodiment, after obtaining the angular momentum vector of the satellite in the body system, the transpose of the conversion matrix converted from the inertial system to the body system can be multiplied by the angular momentum vector of the satellite in the body system to determine the angular momentum vector of the satellite in the inertial system through coordinate transformation.

[0086] For example, assuming that the conversion matrix converted from the inertial system to the body system is M bi , the angular momentum vector of the satellite in the body system is H b0 , and the angular momentum vector of the satellite in the inertial system is H i0 , the following formula is satisfied:

[0087]

[0088] Step S103: inputting the sun direction angle into the angular momentum increment fitting model to obtain the angular momentum increment of the satellite in the body system.

[0089] The sun direction angle represents the angle between the axis with the largest solar pressure area of the satellite in the body system and the sun position vector. The angular momentum increment fitting model is obtained by fitting the corresponding relationship between the historical sun direction angle and the historical angular momentum increment. The axis with the largest solar pressure area represents the coordinate axis with the largest solar pressure area of the satellite in the body system.

[0090] The following describes how to obtain the sun direction angle:

[0091] In an alternative embodiment, the sun position vector in the inertial frame at the beginning of the prediction can be calculated by looking up the ephemeris. The satellite position vector in the inertial frame at the beginning of the prediction can be obtained from the satellite ephemeris data in the database. The relative position vector of the sun with respect to the satellite in the inertial frame can be determined based on the vector difference between the sun position vector and the satellite position vector. The relative position vector of the sun with respect to the satellite in the body frame can be determined by coordinate transformation based on the transformation matrix and the relative position vector in the inertial frame. The sun direction angle can be obtained by taking the inverse cosine function of the ratio between the component of the relative position vector in the body frame along the axis of the largest solar pressure area and the modulus of the relative position vector in the body frame.

[0092] For example, the sun position vector in the inertial frame at the beginning of the prediction can be calculated by looking up the JPL ephemeris. sun = [r sunx r suny r sunz ]. The satellite position vector in the inertial frame at the beginning of the prediction can be obtained from the satellite ephemeris data in the database. sat = [r satx r saty r satz ].

[0093] The relative position vector of the sun with respect to the satellite in the inertial frame can be determined based on the vector difference between the sun position vector and the satellite position vector. i , satisfying the following equation:

[0094] Δr i = r sun - r sat

[0095] The transformation matrix M bi from the inertial frame to the body frame can be multiplied by the relative position vector in the inertial frame to determine the relative position vector of the sun with respect to the satellite in the body frame by coordinate transformation. b , satisfying the following equation:

[0096]

[0097] The sun direction angle can be obtained by taking the inverse cosine function of the ratio between the component of the relative position vector in the body frame along the axis of the largest solar pressure area and the modulus of the relative position vector in the body frame. satisfying the following equation:

[0098]

[0099] The following describes how to obtain the fitting model of the angular momentum increment:

[0100] In an alternative embodiment, a preset different order of sine and cosine model can be used to fit each historical solar direction angle and the corresponding historical angular momentum increment in the inertial system for each historical solar direction angle in the historical solar direction angle set, which includes the historical angle between the axis with the largest solar pressure area of the satellite at each time node in each step in the historical time period and the sun position vector in the inertial system. The fitting coefficients of the satellite in each direction of the three axes are obtained based on the fitting coefficients. The angular momentum increment fitting model is constructed based on the fitting coefficients.

[0101] The historical angular momentum increment corresponding to the historical solar direction angle in the inertial system is obtained in the following manner:

[0102] A historical telemetry data set is obtained, and the historical telemetry data set is standardized. The historical telemetry data set includes a plurality of satellite star times of the satellite in the inertial system, historical angular momentum telemetry values corresponding to the satellite star times, and historical satellite attitude quaternions corresponding to the satellite star times in the historical time period. For each time in the standardized historical telemetry data set, the vector difference between the historical angular momentum telemetry value corresponding to each time and the reference angular momentum is calculated to obtain the instantaneous angular momentum of the satellite in the inertial system. Based on the conversion matrix of the satellite from the inertial system to the inertial system at each time and the instantaneous angular momentum, the instantaneous angular momentum of the satellite in the inertial system is determined through coordinate transformation. The conversion matrix of the satellite from the inertial system to the inertial system at each time is obtained by calculating the historical satellite attitude quaternion at each time. The vector difference between the instantaneous angular momentum of the next time in the inertial system at each time and the instantaneous angular momentum of the satellite in the inertial system at each time is calculated to obtain the historical angular momentum increment of the satellite in the inertial system. Based on the conversion matrix of the satellite from the inertial system to the inertial system at each time and the historical angular momentum increment of the satellite in the inertial system, the historical angular momentum increment of the satellite in the inertial system is determined through coordinate transformation. The time difference between any two adjacent times in the time sequence is a preset step.

[0103] For example, the historical telemetry data set can include a historical angular momentum telemetry data set and a historical attitude quaternion telemetry data set. After obtaining the historical telemetry data set, the historical telemetry data set can be standardized in the following manner:

[0104] The related data of the angular momentum rapid change process in the historical angular momentum telemetry data set and the historical attitude quaternion telemetry data set is deleted. The angular momentum rapid change process at least includes satellite rapid attitude adjustment, momentum wheel unloading, and orbit control. The processed historical angular momentum telemetry data set and the historical attitude quaternion telemetry data set are divided into a plurality of angular momentum periods with the angular momentum rapid change process as a node. Each angular momentum period includes a historical angular momentum telemetry data subset and a historical attitude telemetry data subset in different time periods. The historical angular momentum telemetry data subset and the historical attitude telemetry data subset in each angular momentum period are fitted and processed to complete and smooth the missing data points of the historical angular momentum telemetry data subset and the historical attitude telemetry data subset.

[0105] For example, the historical telemetry data set after the satellite enters the lunar mission orbit is obtained from the database. The historical telemetry data set can include a historical angular momentum telemetry data set S H0 and a historical attitude quaternion telemetry data set S Q0 .

[0106] wherein S H0 contains satellite star time (i.e., time marker of satellite operation) and historical angular momentum telemetry value corresponding to satellite star time in the system. S Q0 contains satellite star time and historical satellite attitude quaternion.

[0107] The historical angular momentum telemetry data set S H0 and the historical attitude quaternion telemetry data set S Q0 respectively satisfy the following formula:

[0108]

[0109] wherein i is an integer. T i is satellite star time. H bxi , H byi , H bzi are respectively components of historical angular momentum telemetry value in X, Y, and Z directions of satellite three-axis in the system. Q ci , Q 1i , Q 2i , Q 3i are components of historical attitude quaternion telemetry data.

[0110] In the historical angular momentum telemetry data set S H0 and the historical attitude quaternion telemetry data set S Q0 , the related data of the angular momentum rapid change process is identified. For example, the angular momentum rapid change process can include: rapid attitude adjustment process T Ai (i=1,...,k), orbit control process T Ci(i = 1,..., m), momentum wheel unloading T Wi (i = 1,..., n). The relevant data of the angular momentum rapid change process is deleted from the original angular momentum telemetry data set S H0 and the original attitude telemetry data set S Q0 to obtain the processed historical angular momentum telemetry data set S H and the historical attitude quaternion telemetry data set S Q .

[0111] The processed historical angular momentum telemetry data set S H and the historical attitude quaternion telemetry data set S Q are sorted in chronological order to obtain the time node sequence T i (i = 1,..., t, t = k + m + n). According to the time period divided according to the different processes of the angular momentum rapid change, the processed historical angular momentum telemetry data set and the historical attitude quaternion telemetry data set are divided into multiple angular momentum periods.

[0112] Each angular momentum period satisfies the following formula:

[0113]

[0114] wherein t is an integer. In the T1 to T2 time period, the first angular momentum period includes the historical angular momentum telemetry data sub-set S Q1 and the historical attitude telemetry data sub-set S H1 . In the T2 to T3 time period, the second angular momentum period includes the historical angular momentum telemetry data sub-set S Q2 and the historical attitude telemetry data sub-set S H2 . The tthangular momentum period includes the historical angular momentum telemetry data sub-set S Qt and the second historical attitude telemetry data sub-set S Ht .

[0115] For example, assuming that the time sequence in the historical telemetry data set is from 0:00 to 24:00 on a certain day. In the event sequence, the angular momentum rapid change process includes satellite rapid attitude adjustment, momentum wheel unloading, and orbit control. Assuming that the satellite rapid attitude adjustment is performed from 5:00 to 5:02, the momentum wheel unloading is performed from 12:00 to 12:02, and the orbit control is performed from 18:00 to 18:05. Then the time sequence from 0:00 to 24:00 can be divided into 4 angular momentum periods. That is, 0:00 to 5:00 is the time sequence corresponding to the first angular momentum period. 5:02 to 12:00 is the time sequence corresponding to the second angular momentum period. 12:02 to 18:00 is the time sequence corresponding to the third angular momentum period. 18:05 to 24:00 is the time sequence corresponding to the fourth angular momentum period.

[0116] Due to the influence of the tracking arc segment limit and the telemetry minimum graduation value, the historical angular momentum telemetry data subset and the historical attitude telemetry data subset in the angular momentum period will have discontinuity and data bounce problems. In order to obtain more accurate, continuous and smooth data, it is necessary to fit the historical angular momentum telemetry data subset and the historical attitude telemetry data subset in each angular momentum period, and complete and smooth the missing data points of the historical angular momentum telemetry data subset and the historical attitude telemetry data subset.

[0117] For example, the missing data points of the historical angular momentum telemetry data subset and the historical attitude telemetry data subset can be completed in the following way:

[0118] For each angular momentum period, a preset first-order polynomial is used to fit the components of each historical angular momentum telemetry value in the three-axis directions of the satellite under the body system to obtain first fitting coefficients, and a first fitting formula is determined based on the first fitting coefficients. Each component of the historical satellite attitude quaternion is fitted by a preset second-order polynomial to obtain second fitting coefficients, and a second fitting formula is determined based on the second fitting coefficients. The missing data points of the historical angular momentum telemetry data subset are determined based on the known data points of the historical angular momentum telemetry data subset and a preset step size using the first fitting formula, and the completed historical angular momentum telemetry data subset is smoothed. The missing data points of the historical attitude quaternion telemetry data subset are determined based on the known data points of the historical attitude quaternion telemetry data subset and a preset step size using the second fitting formula, and the completed historical attitude quaternion telemetry data subset is smoothed.

[0119] For example, a 5th order polynomial is used to fit the components of the historical angular momentum telemetry value [H bx H by H bz ] in the three-axis directions (x, y, z) of the satellite under the body system, and the interpolation coefficients [a1 b1 c1 d1 e1 f1], [a2 b2 c2 d2 e2 f2] and [a3 b3 c3 d3 e3 f3] in the x, y and z directions are fitted.

[0120] Wherein, the 5th order polynomial satisfies the following formula:

[0121]

[0122] A 3rd order polynomial is used to fit each component of the historical satellite attitude quaternion [Q c Q1 Q2 Q3], and the interpolation coefficients [a1 b1 c1 d1 e1 f1], [a2 b2 c2 d2 e2 f2] and [a3 b3 c3 d3 e3 f3] in the x, y and z directions are fitted. cThe interpolation coefficients [a0 b0 c0 d0], [a1 b1 c1 d1], [a2 b2 c2 d2] and [a3 b3 c3 d3] of the four elements Q1, Q2, Q3 and Q4.

[0123] wherein the fourth-order polynomial satisfies the following formula:

[0124]

[0125] After the interpolation coefficient fitting is completed, the first fitting formula can be used to determine the missing data points of the historical angular momentum telemetry data subset based on the known data points of the historical angular momentum telemetry data subset and the preset step length, and the completed historical angular momentum telemetry data subset is smoothed. The second fitting formula is used to determine the missing data points of the historical attitude quaternion telemetry data subset based on the known data points of the historical attitude quaternion telemetry data subset and the preset step length, and the completed historical attitude quaternion telemetry data subset is smoothed.

[0126] It can be understood that if the ground has attitude quaternion prediction data. If there is prediction data, the data is directly used for subsequent calculation. If there is no prediction data, the telemetry attitude quaternion can be interpolated and fitted.

[0127] In an optional embodiment, after the historical telemetry data set is standardized, for each time in the standardized historical telemetry data set, the vector difference between the historical angular momentum telemetry value corresponding to each time and the reference angular momentum is calculated to obtain the instantaneous angular momentum of the satellite in the body system at each time. The transpose of the conversion matrix of the satellite from the inertial system to the body system is multiplied by the instantaneous angular momentum to determine the instantaneous angular momentum of the satellite in the inertial system through coordinate transformation.

[0128] For example, assuming that the historical angular momentum telemetry value corresponding to a historical time is H ybq =[h ybxq h ybyq h ybzq ], and the reference angular momentum is H inib =[h inibx h iniby h inibz ]. Wherein q is an integer. The instantaneous angular momentum H bq of the satellite in the body system at the historical time satisfies the following formula:

[0129] H bq =H ybq -H inib

[0130] The transpose of the conversion matrix of the satellite from the inertial system to the body system at the historical time is multiplied by the instantaneous angular momentum to determine the instantaneous angular momentum of the satellite in the inertial system through coordinate transformation. H bq is multiplied by the conversion matrix M iq of the satellite from the inertial system to the body system, the historical angular momentum increment ΔH i1 of the satellite in the body system is determined through coordinate transformation, and satisfies the following formula:

[0131]

[0132] After obtaining the instantaneous angular momentum H i2 of the satellite in the inertial system at each time, a vector difference between the instantaneous angular momentum H i2 of the satellite in the inertial system at each time and the instantaneous angular momentum H i1 of the satellite in the inertial system at the next time of each time is calculated to obtain the historical angular momentum increment ΔH i of the satellite in the inertial system.

[0133] For example, it is assumed that the instantaneous angular momentum H i1 of the satellite in the inertial system at a certain time in the history is H i2 . The instantaneous angular momentum H i2 of the satellite in the inertial system at the next time of this time is H i1 . The vector difference between H i2 and H i1 is calculated to obtain the historical angular momentum increment ΔH i of the satellite in the inertial system at this time, and satisfies the following formula:

[0134] ΔH i = H i2 - H i1

[0135] Based on the conversion matrix M bi of the satellite from the inertial system to the body system at each time and the historical angular momentum increment ΔH i of the satellite in the inertial system, the historical angular momentum increment ΔH b of the satellite in the body system is determined through coordinate transformation, and satisfies the following formula:

[0136] ΔH b = M bi ΔH i = [ΔH bx ΔH by ΔH bz ]

[0137] In an optional embodiment, after obtaining each historical solar direction angle and the historical angular momentum increment corresponding to the historical solar direction angle in the body system, a preset sine model of different orders is respectively adopted to fit each historical solar direction angle in the historical solar direction angle set and the historical angular momentum increment corresponding to the historical solar direction angle in the body system, to obtain fitting coefficients of the satellite in each direction of the three-axis, and a fitting model of the angular momentum increment is constructed based on the fitting coefficients.

[0138] For example, it is assumed that the angular momentum increment ΔH b of the satellite in the body system is ΔH bx . The fitting model of the angular momentum increment is ΔH byΔH bz For the component of angular momentum increment in X-axis ΔH bx , a five-order sinusoidal model can be used to fit ΔH bx , and the values of a xi , b xi , c xi (i = 1,..., 5) are obtained. Wherein, i is a positive integer.

[0139] The five-order sinusoidal model satisfies the following formula:

[0140]

[0141] For the component of angular momentum increment in Y-axis ΔH by , a seven-order sinusoidal model can be used to fit ΔH by , and the values of a yi , b yi , c yi (i = 1,..., 7) are obtained.

[0142] The seven-order sinusoidal model satisfies the following formula:

[0143]

[0144] For the component of angular momentum increment in Z-axis ΔH bz , a three-order sinusoidal model can be used to fit ΔH bz , and the values of a zi , b zi , c zi (i = 1,..., 3) are obtained.

[0145] The three-order sinusoidal model satisfies the following formula:

[0146]

[0147] After obtaining the fitting coefficients in three-axis directions, the above angular momentum increment fitting model is constructed.

[0148] In an optional embodiment, it is assumed that the preset step length is Δt. The solar direction angle at the current time is The solar direction angle is input into the angular momentum increment fitting model, and the angular momentum increment ΔH b0 of the satellite in the body system is obtained, that is, ΔH bx0 ΔH by0 ΔH bz0 , and each component of the angular momentum increment satisfies the following formula:

[0149]

[0150] Step S104, determining the angular momentum prediction value of the satellite in the current system for the next step length based on the angular momentum increment in the current system, the conversion matrix, the angular momentum telemetry value, and the reference angular momentum.

[0151] Wherein, if the conversion matrix and the reference angular momentum are constant, the angular momentum prediction value for the next step length is positively correlated with the angular momentum increment in the current system.

[0152] In an alternative embodiment, the transpose of the conversion matrix can be multiplied by the transpose of the angular momentum increment in the current system to obtain the angular momentum increment of the satellite in the inertial system. Based on the angular momentum increment in the inertial system and the vector sum of the angular momentum vector in the inertial system, the angular momentum vector of the satellite in the inertial system for the next step length is determined. The conversion matrix is multiplied by the angular momentum vector of the satellite in the inertial system for the next step length to determine the angular momentum vector of the satellite in the current system for the next step length through coordinate transformation. Based on the vector sum of the angular momentum vector of the satellite in the current system for the next step length and the reference angular momentum, the angular momentum prediction value of the satellite in the current system for the next step length is determined.

[0153] For example, it is assumed that the angular momentum increment of the satellite in the current system is ΔH b0 = [ΔH bx0 ΔH by0 ΔH bz0 ]. The conversion matrix is M bi . Then the angular momentum increment of the satellite in the inertial system is ΔH i0 . ΔH i0 satisfies the following formula:

[0154]

[0155] It is assumed that the angular momentum vector in the inertial system is H i0 . Based on the angular momentum increment in the inertial system ΔH i0 and the vector sum of the angular momentum vector in the inertial system H i0 , the angular momentum vector of the satellite in the inertial system for the next step length H i1 is determined. H i1 satisfies the following formula;

[0156] H i1 = H i0 + ΔH i0

[0157] The conversion matrix M bi is multiplied by the angular momentum vector of the satellite in the inertial system for the next step length H i1 to determine the angular momentum vector of the satellite in the current system for the next step length H b1 through coordinate transformation. H b1 satisfies the following formula:

[0158] H b1 =M bi ×H i1

[0159] Assume the reference angular momentum is H inib =[h inibx h iniby h inibz Based on the next-long angular momentum vector H in this system. b1 With reference angular momentum H inib =[h inibx h iniby h inibz The vector sum of ] determines the predicted angular momentum H of the satellite in this system for the next long distance. yb1 H yb1 Satisfy the following formula:

[0160] H yb1 =H b1 +H inib

[0161] After calculating the predicted angular momentum value of a satellite over a future period, it can be used to predict the saturation time of the satellite's momentum wheel, plan its unloading timing, and avoid unexpected unloading of the momentum wheel. Because the orbital stability of some Earth-Moon systems is poor, unexpected unloading of the momentum wheel may cause significant deviations in the orbit, leading to the failure of some on-orbit experiments with high orbital accuracy requirements.

[0162] In another possible scenario, angular momentum prediction methods can be used for satellite angular momentum management. For example, a Halo orbiting satellite at the Earth-Moon L2 point can maintain its orbital configuration by unloading momentum wheels. Based on the obtained angular momentum increment fitting model, considering the satellite's demand for angular momentum changes, the satellite's on-orbit attitude can be planned to achieve controlled growth of satellite angular momentum. Unloading to maintain the orbital configuration at appropriate times can greatly reduce the orbital maintenance frequency of Halo orbiting satellites.

[0163] In another possible scenario, satellite fault diagnosis can be based on angular momentum predictions. For example, by comparing the predicted angular momentum values ​​with the actual telemetry values, it can be determined whether a potential fault exists. If there is a significant deviation between the predicted and actual angular momentum telemetry values, then the satellite's attitude control system or related sensors are faulty.

[0164] Understandably, when the axis with the largest solar radiation pressure area of ​​the satellite in this system is not the Z-axis, the coordinate axis related terms in the formula also need to be replaced accordingly.

[0165] If the X-axis is the axis of maximum solar radiation pressure area, the variables involving the Z-axis in the original formula should be replaced with X, while the Y-axis remains unchanged.

[0166] If the Y-axis is the maximum solar pressure area axis, the variable related to the Z-axis in the original formula should be replaced by Y, and the X-axis remains unchanged.

[0167] The angular momentum prediction method when the satellite's maximum solar pressure area axis is the X-axis or Y-axis under the current system is consistent with the angular momentum prediction method when the satellite's maximum solar pressure area axis is the Z-axis under the current system. Only the coordinate axes in the formula need to be replaced accordingly, which will not be repeated here.

[0168] The following examples of Figure 1 are provided:

[0169] Taking Chang'e-4 relay satellite as an example, the satellite has been running in the Halo orbit of the Earth-Moon L2 point for a long time. The Z-axis of the satellite is a large umbrella-shaped parabolic antenna, and the Y-axis is a solar sail. The solar pressure area of the satellite in the Z-axis direction is much larger than that in the X-axis and Y-axis directions, so it is suitable for this application.

[0170] The historical angular momentum telemetry data set is obtained by searching the database for the satellite's angular momentum telemetry data from January to December 2022. The historical attitude quaternion telemetry data set is obtained through the historical attitude prediction file. Due to the large amount of data, the telemetry data of the first hour of the first two arc segments of the historical angular momentum telemetry data set is displayed at 5-minute intervals as shown in Table 1. The angular momentum unit is Nms, and the same applies below.

[0171] Table 1: Angular momentum original telemetry data

[0172]

[0173]

[0174] In the entire data set, find the fast attitude adjustment process TAi(i=1,...,k), the orbit control process TCi(i=1,...,m), the momentum wheel unloading TWi(i=1,...,n), etc. Taking January 2022 as an example, the results are shown in Table 2:

[0175] Table 2: Angular momentum fast change events (January 2022)

[0176] Beijing time Key events 2022-01-03T09:00 Momentum wheel offload 2022-01-05T19:00 Orbit maintenance 2022-01-08T18:00 Momentum wheel offload 2022-01-11T21:00 Momentum wheel offload 2022-01-15T14:30 Momentum wheel offload 2022-01-19T08:00 Momentum wheel offload 2022-01-21T22:30 Orbit maintenance 2022-01-23T23:00 Momentum wheel offload 2022-01-26T09:00 Momentum wheel offload 2022-01-27T17:10 Momentum wheel offload 2022-01-30T09:20 Fast attitude slews

[0177] Delete the relevant data of the angular momentum fast change process in the historical angular momentum telemetry data set and the historical attitude quaternion telemetry data set. Divide the historical angular momentum telemetry data set and the historical attitude quaternion telemetry data set into multiple angular momentum periods with the angular momentum fast change process as the node. According to the angular momentum fast change events in Table 2, the angular momentum data in January 2022 can be divided into 10 angular momentum periods as shown in Table 3.

[0178] Table 3: Rapid change of angular momentum events (January 2022)

[0179]

[0180]

[0181] For each subset of historical angular momentum telemetry data and each subset of historical attitude telemetry data within an angular momentum period, the angular momentum data [Hx, Hy, Hz] in each direction of the satellite's three axes under the body system is fitted using a 5th order polynomial to obtain the fitting coefficients corresponding to each direction of the satellite's three axes. The fitting coefficients are as follows: bx H by H bz ] are fitted using a 5th order polynomial to obtain the fitting coefficients corresponding to each direction of the satellite's three axes. The fitting coefficients are as follows:

[0182]

[0183] The fitting formula is constructed using the fitting coefficients corresponding to each direction of the satellite's three axes. Based on the known data points of the subset of historical angular momentum telemetry data and a preset step size, the missing data points of the subset of historical angular momentum telemetry data are determined using the fitting formula, and the completed subset of historical angular momentum telemetry data is smoothed. Taking the third angular momentum period as an example, the fitted data is shown in Table 4.

[0184] Table 4: Interpolated and fitted angular momentum data

[0185]

[0186]

[0187] Since the ground has satellite attitude quaternion prediction data, the telemetry attitude quaternion does not need to be interpolated and fitted.

[0188] In each angular momentum period, the angle between the satellite's Z-axis in the body system and the sun vector and the corresponding angular momentum increment in the body system are calculated. When the satellite's angular momentum is zero, the X, Y, Z three-axis angular momentum reference value is (1.5, 1.5, -1.5). Taking the fourth angular momentum period as an example, the calculated Z-axis and sun vector angle and the X, Y, Z three-axis angular momentum increment in the body system are shown in Table 5, with the angle unit in degrees.

[0189] Table 5: Z-axis and sun vector angle and angular momentum increment

[0190]

[0191]

[0192] The angle between the Z-axis and the sun vector is taken as the independent variable As the independent variable, the sine and cosine model is used to fit the three-direction angular momentum increment ΔH bx , ΔH by , ΔH bz , and the fitting coefficient values are as follows, only the Y-axis coefficient a yi , b yi , and c yi (i = 1,...,7) are shown, and the rest is omitted.

[0193]

[0194] The relationship between the angular momentum increment of the satellite in the X-axis of the body system and the angle between the sun direction is shown in FIG. 6. Figure 2

[0195] The relationship between the angular momentum increment of the satellite in the Y-axis of the body system and the angle between the sun direction is shown in FIG. 7. Figure 3

[0196] The relationship between the angular momentum increment of the satellite in the Z-axis of the body system and the angle between the sun direction is shown in FIG. 8. Figure 4

[0197] Based on the angular momentum increment in the body system, the conversion matrix, the angular momentum telemetry value, and the reference angular momentum, the angular momentum prediction value of the next step of the satellite in the body system is determined. Let the prediction time period be the start time T s = 2023-02-01T10:00:00.000, the end time T e = 2023-02-01T23:00:00.000, the start time T s The angular momentum telemetry value of the satellite in the body system H yb0 = [1.5 1.5 -1.5], and the satellite attitude in the prediction time period is taken as an example under the following two commonly used attitude working conditions.

[0198] Example 1:

[0199] The satellite attitude in the prediction period is positive Z-axis deviated by 37 degrees to the positive X-axis for sun orientation, and the earth is constrained in the XZ plane. The satellite angular momentum prediction result is shown in Table 6.

[0200] Table 6: Angular momentum prediction result of example 1

[0201]

[0202] Example 2:

[0203] The satellite attitude in the prediction period is positive Z-axis for moon orientation, and the sun is constrained in the XZ plane. The satellite angular momentum prediction result is shown in Table 7.

[0204] ​​​Table 7: Angular momentum prediction results of example two

[0205]

[0206]

[0207] The above method is used to predict the angular momentum of the satellite in the lunar relay tracking attitude mode and the offset sun inertial flight attitude mode, and the correctness of the method is verified.

[0208] Verification one: angular momentum prediction in lunar relay tracking attitude mode

[0209] The satellite was adjusted and unloaded on November 23, 2024, and then entered the lunar relay tracking mode. After unloading, the satellite telemetry angular momentum data was collected, the time was 18:00:00.457 on November 23, 2024, and the angular momentum telemetry value was [1.435 1.435 -1.435] Nms. With this set of data as the initial value, the subsequent angular momentum is predicted and compared with the telemetry result, and the results are shown in Tables 8, 9 and 10.

[0210] Table 8: 0.69-day prediction deviation (2024-11-24T10:27:52)

[0211] Type X-axis angular momentum Y-axis angular momentum Z-axis angular momentum Telemetry results 1.435 0.665 -1.505 Forecast results 1.469 0.558 -1.520 Forecast bias 0.034 -0.107 -0.015

[0212] Table 9: 1.96-day prediction deviation (2024-11-25T16:59:53)

[0213] Type X-axis angular momentum Y-axis angular momentum Z-axis angular momentum Telemetry results 1.680 -1.120 -1.680 Forecast results 1.734 -1.276 -1.662 Forecast bias 0.054 -0.156 0.018

[0214] Table 10: 2.96-day prediction deviation (2024-11-26T17:00:00)

[0215] Type X-axis angular momentum Y-axis angular momentum Z-axis angular momentum Telemetry results 2.345 -2.310 -1.820 Forecast results 2.368 -2.089 -1.776 Forecast bias 0.023 0.221 0.044

[0216] From the above table, it can be seen that the angular momentum deviation is mainly in the Y-axis direction, the 0.7-day prediction deviation is about -0.11 Nms, the 2.0-day prediction deviation is about -0.16 Nms, and the 3.0-day prediction deviation is about 0.22 Nms.

[0217] Verification two: angular momentum prediction in sun offset inertial flight mode

[0218] The satellite was adjusted and unloaded on December 17, 2024, and then entered an inertial flight attitude offset to the sun, with the +Z axis offset to the sun by 32 degrees, corresponding to the attitude quaternion [0.159148512 -0.188430480.82269712 -0.51218664]. After unloading, the satellite telemetry angular momentum data was collected, and the time was 11:40:03.510 on December 17, 2024, and the angular momentum telemetry value was [1.470 1.505 -1.575] Nms. With this set of data as the initial value, the subsequent angular momentum is predicted, and the results are compared with the telemetry results, as shown in Tables 11, 12 and 13.

[0219] Table 11 0.94-day prediction deviation (2024-12-18T10:09:55)

[0220] Type X-axis angular momentum Y-axis angular momentum Z-axis angular momentum Telemetry results 1.365 0.245 -1.610 Forecast results 1.357 0.231 -1.635 Forecast bias -0.008 -0.014 -0.025

[0221] Table 12 1.98-day prediction deviation (2024-12-19T11:10:04)

[0222] Type X-axis angular momentum Y-axis angular momentum Z-axis angular momentum Telemetry results 1.260 -1.085 -1.680 Forecast results 1.240 -1.038 -1.695 Forecast bias -0.020 0.047 -0.015

[0223] Table 13 3.12-day prediction deviation (2024-12-20T14:31:08)

[0224]

[0225]

[0226] From the above table, under the inertial flight attitude offset to the sun, the prediction error is relatively small. The 0.94-day prediction deviation is about 0.03 Nms, the 2.0-day prediction deviation is about 0.05 Nms, and the 3.1-day prediction deviation is about 0.10 Nms.

[0227] Figure 5 A structure diagram of an angular momentum prediction device for a lunar orbit satellite provided by an embodiment of the present application is shown in FIG. 1. As shown in FIG. 1, the device includes a transceiver module 501 and a processing module 502. Figure 5

[0228] The transceiver module 501 is configured to obtain the angular momentum telemetry value of the satellite three-axis in the body system at the prediction start time and the satellite attitude quaternion, and the satellite attitude quaternion is used to describe the satellite attitude.

[0229] The processing module 502 is configured to calculate the conversion matrix of the satellite converted from the inertial system to the body system based on the satellite attitude quaternion, and the elements in the conversion matrix represent the projection components of the unit vectors in the inertial system in the body system.

[0230] ​The processing module 502 is further configured to obtain an angular momentum vector of the satellite in the body system based on the angular momentum telemetry value and the reference angular momentum, and determine an angular momentum vector of the satellite in the inertial system by coordinate transformation based on the conversion matrix and the angular momentum vector in the body system, wherein the reference angular momentum represents a marked value of three-axis direction of the angular momentum of the satellite when the angular momentum of the satellite is zero after unloading.

[0231] The processing module 502 is further configured to input a solar direction angle into the angular momentum increment fitting model to obtain an angular momentum increment of the satellite in the body system, wherein the solar direction angle represents an angle between an axis with the largest solar pressure area of the satellite in the body system and a solar vector, and the angular momentum increment fitting model is obtained by fitting a corresponding relationship between historical solar direction angles and historical angular momentum increments, and the axis with the largest solar pressure area represents a coordinate axis with the largest solar pressure area of the satellite in the body system.

[0232] The processing module 502 is further configured to determine a next-step angular momentum prediction value of the satellite in the body system based on the angular momentum increment in the body system, the conversion matrix, the angular momentum telemetry value and the reference angular momentum, wherein if the conversion matrix and the reference angular momentum are constant, the next-step angular momentum prediction value is positively correlated with the angular momentum increment in the body system.

[0233] Optionally, the solar direction angle is obtained in the following manner, and the processing module 502 is further configured to:

[0234] The solar position vector of the sun in the inertial system at the start time of the prediction is calculated by searching the JPL planetary ephemeris.

[0235] The satellite position vector of the satellite in the inertial system at the start time of the prediction is obtained from the satellite ephemeris data in the database.

[0236] The relative position vector of the sun with respect to the satellite in the inertial system is determined based on the vector difference between the solar position vector and the satellite position vector.

[0237] The relative position vector of the sun with respect to the satellite in the body system is determined by coordinate transformation based on the conversion matrix and the relative position vector in the inertial system.

[0238] The solar direction angle is obtained by performing an inverse cosine function operation on the ratio between the component of the relative position vector in the body system in the axis with the largest solar pressure area and the module of the relative position vector in the body system.

[0239] Optionally, the angular momentum increment fitting model is fitted in the following manner, and the processing module 502 is further configured to:

[0240] The satellite three-axis direction under the system respectively adopts preset different order sine and cosine model to fit each historical solar direction angle and the historical solar direction angle corresponding historical angular momentum increment in the system, to obtain the fitting coefficient of the satellite in three-axis direction, and the historical solar direction angle set includes the historical angle between the time node corresponding to each step in the historical time period and the maximum solar pressure area axis of the satellite under the system and the sun position vector;

[0241] The fitting coefficient is used to construct an angular momentum increment fitting model.

[0242] Optionally, the historical solar direction angle corresponding to the historical angular momentum increment in the system is obtained by the following method:

[0243] The transceiver module 501 is further configured to obtain a historical telemetry data set;

[0244] The processing module 502 is further configured to standardize the historical telemetry data set, and the historical telemetry data set includes a plurality of satellite star times of the satellite under the system in a historical time period, historical angular momentum telemetry values corresponding to the satellite star times, and historical satellite attitude quaternions corresponding to the satellite star times;

[0245] For each time in the standardized historical telemetry data set, the vector difference between the historical angular momentum telemetry value corresponding to each time and the reference angular momentum is calculated to obtain the instantaneous angular momentum of the satellite under the system at each time;

[0246] Based on the conversion matrix of the satellite from the inertial system to the system at each time and the instantaneous angular momentum, the instantaneous angular momentum of the satellite under the inertial system is determined through coordinate transformation, and the conversion matrix of the satellite from the inertial system to the system at each time is obtained by calculating the historical satellite attitude quaternion at each time;

[0247] The vector difference between the instantaneous angular momentum of the satellite under the inertial system at the next time of each time and the instantaneous angular momentum of the satellite under the inertial system at each time is calculated to obtain the historical angular momentum increment of the satellite under the inertial system;

[0248] Based on the conversion matrix of the satellite from the inertial system to the system at each time and the historical angular momentum increment of the satellite under the inertial system, the historical angular momentum increment of the satellite under the system is determined through coordinate transformation.

[0249] Optionally, the historical telemetry data set includes a historical angular momentum telemetry data set and a historical attitude quaternion telemetry data set, and the processing module 502 is specifically configured to:

[0250] delete the data related to the angular momentum rapid change process in the historical angular momentum telemetry data set and the historical attitude quaternion telemetry data set, the angular momentum rapid change process at least including satellite rapid attitude adjustment, momentum wheel unloading, orbit control;

[0251] divide the historical angular momentum telemetry data set and the historical attitude quaternion telemetry data set into a plurality of angular momentum periods with the angular momentum rapid change process as a node, each angular momentum period including a historical angular momentum telemetry data sub-set and a historical attitude telemetry data sub-set in different time periods;

[0252] fit the historical angular momentum telemetry data sub-set and the historical attitude telemetry data sub-set in each angular momentum period, complete and smooth the missing data points of the historical angular momentum telemetry data sub-set and the historical attitude telemetry data sub-set.

[0253] Optionally, the historical angular momentum telemetry data sub-set and the historical attitude telemetry data sub-set in each angular momentum period are fitted to complete and smooth the missing data points of the historical angular momentum telemetry data sub-set and the historical attitude telemetry data sub-set, and the processing module 502 is specifically configured to:

[0254] for each angular momentum period, a preset first-order polynomial is used to fit the components of each historical angular momentum telemetry value in the three-axis directions of the satellite in the body system to obtain first fitting coefficients, and a first fitting formula is determined based on the first fitting coefficients;

[0255] a preset second-order polynomial is used to fit each component of each historical satellite attitude quaternion to obtain second fitting coefficients, and a second fitting formula is determined based on the second fitting coefficients;

[0256] the first fitting formula is used to determine the missing data points of the historical angular momentum telemetry data sub-set based on the known data points of the historical angular momentum telemetry data sub-set and a preset step size, and the completed historical angular momentum telemetry data sub-set is smoothed;

[0257] the second fitting formula is used to determine the missing data points of the historical attitude quaternion telemetry data sub-set based on the known data points of the historical attitude quaternion telemetry data sub-set and a preset step size, and the completed historical attitude quaternion telemetry data sub-set is smoothed.

[0258] Optionally, the next step angular momentum prediction value of the satellite in the body system is determined based on the angular momentum increment in the body system, the conversion matrix, the angular momentum telemetry value, and the reference angular momentum, and the processing module 502 is specifically configured to:

[0259] the transpose of the conversion matrix is multiplied by the transpose of the angular momentum increment in the body system to obtain the angular momentum increment of the satellite in the inertial system;

[0260] Based on the increment of angular momentum in the inertial frame and the vector sum of the angular momentum vectors in the inertial frame, determine the angular momentum vector of the satellite in the next step length in the inertial frame;

[0261] The transformation matrix is ​​multiplied by the angular momentum vector of the next length in the inertial frame, and the angular momentum vector of the next length in the satellite's own frame is determined by coordinate transformation.

[0262] Based on the vector sum of the angular momentum vector of the next length under this system and the reference angular momentum, the predicted value of the angular momentum of the satellite under this system for the next length is determined.

[0263] Based on the same technical concept, this application also provides an electronic device that can realize the function of the aforementioned lunar orbit satellite angular momentum prediction device.

[0264] Figure 6 This is a schematic diagram of the structure of an electronic device provided in an embodiment of this application.

[0265] At least one processor 601 and a memory 602 connected to at least one processor 601. In this embodiment, the specific connection medium between the processor 601 and the memory 602 is not limited. Figure 6 The example shown is the connection between processor 601 and memory 602 via bus 600. Bus 600 is... Figure 6 The connections between other components are indicated by thick lines and are for illustrative purposes only, not as limiting information. The 600 bus can be divided into address bus, data bus, control bus, etc., for ease of representation. Figure 6 The term is represented by a single thick line, but this does not imply that there is only one bus or one type of bus. Alternatively, the processor 601 can also be called a controller; there is no restriction on the name.

[0266] In this embodiment, memory 602 stores instructions executable by at least one processor 601. By executing the instructions stored in memory 602, at least one processor 601 can execute the angular momentum prediction method for a lunar orbiting satellite discussed above. Processor 601 can implement... Figure 6 The functions of each module in the device shown.

[0267] The processor 601 is the control center of the device. It can connect to various parts of the control device through various interfaces and lines. By running or executing instructions stored in memory 602 and calling data stored in memory 602, the processor can perform various functions and process data, thereby monitoring the device as a whole.

[0268] In one possible design, the processor 601 can include one or more processing units, and the processor 601 can integrate an application processor and a modem processor, where the application processor mainly processes operating systems, driver interfaces, and application programs, and the modem processor mainly processes wireless communication. It can be understood that the modem processor can also not be integrated into the processor 601. In some embodiments, the processor 601 and the memory 602 can be implemented on the same chip, and in some embodiments, they can also be implemented on separate chips respectively.

[0269] The processor 601 can be a general-purpose processor, such as a central processing unit (CPU), a digital signal processor, an application-specific integrated circuit, a field programmable gate array, or other programmable logic device, a discrete gate or transistor logic device, a discrete hardware component, and can implement or execute the methods, steps, and logic block diagrams disclosed in the embodiments of the present application. The general-purpose processor can be a microprocessor or any conventional processor. The steps of the method for predicting the angular momentum of a lunar-orbit satellite disclosed in the embodiments of the present application can be directly embodied as execution completed by a hardware processor, or executed by a combination of hardware and software modules in the processor.

[0270] The memory 602 is a non-volatile computer-readable storage medium, which can be used to store non-volatile software programs, non-volatile computer-executable programs, and modules. The memory 602 can include at least one type of storage medium, such as flash memory, a hard disk, a multimedia card, a card-type memory, random access memory (RAM), static random access memory (SRAM), programmable read-only memory (PROM), read-only memory (ROM), electrically erasable programmable read-only memory (EEPROM), a magnetic memory, a magnetic disk, an optical disk, and the like. The memory 602 is any other medium capable of carrying or storing desired program codes in the form of instructions or data structures and capable of being accessed by a computer, but is not limited thereto. The memory 602 in the embodiments of the present application can also be a circuit or any other device capable of realizing a storage function, for storing program instructions and / or data.

[0271] By designing and programming the processor 601, the code corresponding to the method for predicting the angular momentum of a lunar-orbit satellite introduced in the foregoing embodiments can be fixed into the chip, so that the chip can execute the method for predicting the angular momentum of a lunar-orbit satellite when running. Figure 1A method for predicting angular momentum of a lunar-orbit satellite is shown in the embodiment. How to design and program the processor 601 is known to those skilled in the art, and thus is not described herein.

[0272] It should be noted that the above electronic device provided by the embodiments of the present application can realize all the method steps achieved by the method embodiments and achieve the same technical effects. Therefore, the same parts and beneficial effects of the method embodiments are not described in detail herein.

[0273] The embodiments of the present application also provide a computer readable storage medium, which stores computer executable instructions. The computer executable instructions are used for causing a computer to execute the method for predicting angular momentum of a lunar-orbit satellite in the above embodiments.

[0274] Those skilled in the art should understand that the embodiments of the present application can be provided as a method, a system, or a computer program product. Therefore, the present application can take the form of an entirely hardware embodiment, an entirely software embodiment, or an embodiment combining software and hardware aspects. Moreover, the present application can take the form of a computer program product implemented on one or more computer-usable storage media (including, but not limited to, disk storage, CD-ROMs, optical storage media, etc.) containing computer usable program code.

[0275] The present application is described with reference to the flowcharts and / or block diagrams of the method, device (system), and computer program product according to the present application. It should be understood that each flow and / or block in the flowcharts and / or block diagrams, and the combination of flows and / or blocks in the flowcharts and / or block diagrams can be implemented by computer program instructions. These computer program instructions can be provided to a processor of a general-purpose computer, a special-purpose computer, an embedded processor, or other programmable data processing apparatus to produce a machine, so that the instructions executed by the processor of the computer or other programmable data processing apparatus produce a device that implements the functions specified in the flowcharts and / or block diagrams. Figure 1 The functions specified in one or more flows and / or blocks in the flowcharts and / or block diagrams. Figure 1 The functions specified in one or more flows and / or blocks in the flowcharts and / or block diagrams.

[0276] These computer program instructions can also be stored in a computer readable storage medium that can direct the computer or other programmable data processing apparatus to work in a specific manner, so that the instructions stored in the computer readable storage medium produce a manufactured product including instruction devices that implement the functions specified in the flowcharts and / or block diagrams. Figure 1 The functions specified in one or more flows and / or blocks in the flowcharts and / or block diagrams. Figure 1 The functions specified in one or more flows and / or blocks in the flowcharts and / or block diagrams.

[0277] These computer program instructions can also be loaded into a computer or other programmable data processing devices, so that a series of operational steps are performed on the computer or other programmable data processing devices to generate computer-implemented processes, thus the instructions executed on the computer or other programmable data processing devices provide processes for implementing the flow Figure 1 The flowchart and / or block diagram in the drawings show the architecture, functionality, and operation of possible implementations of systems, methods and computer program products according to various embodiments of the present application. In this regard, each block in the flowchart and / or block diagrams can represent a module, segment, or portion of code, which comprises one or more executable ​ The flowchart and / or block diagram in the drawings show the architecture, functionality, and operation of possible implementations of systems, methods and computer program products according to various embodiments of the present application. In this regard, each block in the flowchart and / or block diagrams can represent a module, segment, or portion of code, which comprises one or more executable

[0278] It will be apparent that those skilled in the art will be able to make various modifications and variations without departing from the spirit and scope of the application. Accordingly, it should be understood that the application is intended to be illustrative only and not limiting of the scope of the application, as claimed.

Claims

1. A method for predicting the angular momentum of a lunar orbiting satellite, characterized in that, The method includes: Obtain the telemetry values ​​of the satellite's three axes and the satellite's attitude quaternion at the predicted start time in this system. Based on the satellite attitude quaternion, calculate the transformation matrix of the satellite from the inertial frame to this system. The satellite attitude quaternion is used to describe the satellite's attitude, and the elements in the transformation matrix represent the projection components of the unit vector in the inertial frame into this system. Based on the telemetry value of angular momentum and the reference angular momentum, the angular momentum vector of the satellite in the local system is obtained. Based on the transformation matrix and the angular momentum vector in the local system, the angular momentum vector of the satellite in the inertial frame is determined by coordinate transformation. The reference angular momentum represents the indicated value of the three-axis direction of angular momentum when the satellite's angular momentum is zero after unloading. Input the solar direction angle into the angular momentum increment fitting model to obtain the satellite's angular momentum increment under the specified system. The solar direction angle represents the angle between the axis of the satellite with the largest solar radiation area and the solar position vector under the specified system. The angular momentum increment fitting model is obtained by fitting the correspondence between historical solar direction angles and historical angular momentum increments. The axis with the largest solar radiation area represents the coordinate axis of the satellite with the largest solar radiation area under the specified system. Based on the angular momentum increment under the system, the transformation matrix, the angular momentum telemetry value, and the reference angular momentum, the predicted angular momentum value of the satellite for the next length under the system is determined. If the transformation matrix and the reference angular momentum remain unchanged, the predicted angular momentum value for the next length is positively correlated with the angular momentum increment under the system.

2. The method according to claim 1, characterized in that, The solar direction angle is obtained in the following way: By consulting the planetary ephemeris, the Sun's position vector in the inertial frame at the predicted start time is calculated; The satellite ephemeris data is obtained from the database to obtain the satellite position vector in the inertial frame at the start time of the prediction; Based on the vector difference between the solar position vector and the satellite position vector, the relative position vector of the sun with respect to the satellite in the inertial frame is determined; Based on the transformation matrix and the relative position vector in the inertial frame, the relative position vector of the sun with respect to the satellite in the local frame is determined by coordinate transformation; The solar direction angle is obtained by performing an inverse cosine function operation on the ratio between the component of the relative position vector in the system with the axis of maximum solar radiation pressure and the magnitude of the relative position vector in the system.

3. The method according to claim 1, characterized in that, The angular momentum increment fitting model was obtained using the following method: In the aforementioned system, for each of the three axes of the satellite, sine and model of different preset orders are used to fit each historical solar direction angle in the historical solar direction angle set with the historical angular momentum increment corresponding to the historical solar direction angle in the aforementioned system, to obtain the fitting coefficients of the satellite in each of the three axes. The historical solar direction angle set includes the historical angle between the axis with the largest solar radiation pressure area of ​​the satellite at each step point in the historical time period and the solar position vector in the aforementioned system. The angular momentum increment fitting model is constructed based on the fitting coefficients.

4. The method according to claim 3, characterized in that, The historical solar orientation angle is obtained from the historical angular momentum increment corresponding to this system in the following manner: Acquire a historical telemetry data set and standardize the historical telemetry data set. The historical telemetry data set includes multiple satellite times of the satellite under the system within the historical time period, the historical angular momentum telemetry values ​​corresponding to the satellite times, and the historical satellite attitude quaternions corresponding to the satellite times. For each moment in the standardized historical telemetry data set, the vector difference between the historical angular momentum telemetry value corresponding to each moment and the reference angular momentum is calculated to obtain the instantaneous angular momentum of the satellite in the system at each moment; Based on the transformation matrix of the satellite from the inertial frame to the local frame at each time point and the instantaneous angular momentum, the instantaneous angular momentum of the satellite in the inertial frame is determined by coordinate transformation. The transformation matrix of the satellite from the inertial frame to the local frame at each time point is obtained by calculating the historical satellite attitude quaternions at each time point. Calculate the vector difference between the instantaneous angular momentum of the satellite in the inertial frame at the next moment after each moment and the instantaneous angular momentum at each moment, to obtain the historical angular momentum increment of the satellite in the inertial frame; Based on the transformation matrix of the satellite from the inertial frame to the local frame at each time point and the historical angular momentum increment of the satellite in the inertial frame, the historical angular momentum increment of the satellite in the local frame is determined by coordinate transformation.

5. The method according to claim 4, characterized in that, The historical telemetry data set includes a historical angular momentum telemetry data set and a historical attitude quaternion telemetry data set. The standardization process for the historical telemetry data set specifically includes: Delete the relevant data on the rapid change process of angular momentum in the historical angular momentum telemetry data set and the historical attitude quaternion telemetry data set. The rapid change process of angular momentum includes at least the satellite's rapid attitude adjustment, momentum wheel unloading, and orbit control. Using the rapid change process of angular momentum as a node, the historical angular momentum telemetry data set and the historical attitude quaternion telemetry data set are divided into multiple angular momentum periods. Each angular momentum period includes a subset of historical angular momentum telemetry data and a subset of historical attitude telemetry data within different time periods. The historical angular momentum telemetry data subset and the historical attitude telemetry data subset within each angular momentum cycle are fitted to complete and smooth the missing data points of the historical angular momentum telemetry data subset and the historical attitude telemetry data subset.

6. The method according to claim 5, characterized in that, The fitting process for the historical angular momentum telemetry data subset and the historical attitude telemetry data subset within each angular momentum cycle, to complete and smooth missing data points in the historical angular momentum telemetry data subset and the historical attitude telemetry data subset, specifically includes: For each angular momentum period, a first fitting coefficient is obtained by fitting the components of each historical angular momentum telemetry value in each direction of the satellite's three axes under this system using a preset first-order polynomial, and a first fitting formula is determined based on the first fitting coefficient. A second fitting coefficient is obtained by fitting each component of the attitude quaternion of each historical satellite with a preset second-order polynomial, and a second fitting formula is determined based on the second fitting coefficient. Using the first fitting formula, based on the known data points of the historical angular momentum telemetry data subset and the preset step size, the missing data points of the historical angular momentum telemetry data subset are determined, and the completed historical angular momentum telemetry data subset is smoothed. Using the second fitting formula, based on the known data points of the historical attitude quaternion telemetry data subset and the preset step size, the missing data points of the historical attitude quaternion telemetry data subset are determined, and the completed historical attitude quaternion telemetry data subset is smoothed.

7. The method according to claim 1, characterized in that, The determination of the predicted angular momentum of the satellite for the next length under the current system based on the angular momentum increment under the current system, the transformation matrix, the angular momentum telemetry value, and the reference angular momentum specifically includes: The angular momentum increment of the satellite in the inertial frame is obtained by multiplying the transpose of the transformation matrix with the transpose of the angular momentum increment in the system. Based on the increment of angular momentum in the inertial frame and the vector sum of the angular momentum vectors in the inertial frame, the angular momentum vector of the satellite in the next step length in the inertial frame is determined; The transformation matrix is ​​multiplied by the angular momentum vector of the next step length in the inertial frame, and the angular momentum vector of the satellite of the next step length in the local frame is determined by coordinate transformation. Based on the vector sum of the angular momentum vector of the next length under the current system and the reference angular momentum, the predicted value of the angular momentum of the satellite under the current system is determined.

8. An angular momentum prediction device for a lunar orbiting satellite, characterized in that, The device includes: The transceiver module is used to acquire the angular momentum telemetry values ​​of the satellite's three axes and the satellite attitude quaternion at the predicted start time in this system. The satellite attitude quaternion is used to describe the satellite attitude. The processing module is used to calculate the transformation matrix of the satellite from the inertial frame to the local frame based on the satellite attitude quaternion, wherein the elements in the transformation matrix represent the projection components of the unit vector in the inertial frame into the local frame; The processing module is further configured to obtain the angular momentum vector of the satellite in the local system based on the angular momentum telemetry value and the reference angular momentum, and determine the angular momentum vector of the satellite in the inertial frame through coordinate transformation based on the transformation matrix and the angular momentum vector in the local system. The reference angular momentum represents the indicated value of the three-axis direction of the angular momentum when the satellite's angular momentum is zero after unloading. The processing module is also used to input the solar direction angle into the angular momentum increment fitting model to obtain the angular momentum increment of the satellite in the local system. The solar direction angle represents the angle between the axis of the satellite with the largest solar radiation area and the solar position vector in the local system. The angular momentum increment fitting model is obtained by fitting the correspondence between historical solar direction angles and historical angular momentum increments. The axis with the largest solar radiation area represents the coordinate axis of the satellite with the largest solar radiation area in the local system. The processing module is further configured to determine the predicted angular momentum value of the satellite for the next length under the current system based on the angular momentum increment under the current system, the transformation matrix, the angular momentum telemetry value, and the reference angular momentum. If the transformation matrix and the reference angular momentum remain unchanged, the predicted angular momentum value for the next length is positively correlated with the angular momentum increment under the current system.

9. An electronic device, comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, characterized in that, When the processor executes the computer program, it implements the method as described in any one of claims 1-7.

10. A computer-readable storage medium, characterized in that, The computer-readable storage medium stores a computer program for causing the computer to perform the method of any one of claims 1-7.

11. A computer program product, characterized in that, When the computer program product is invoked by a computer, it causes the computer to perform the method as described in any one of claims 1-7.

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