Method and device for determining lunar orbiting satellite orbit based on radio and laser ranging

Abstract

The present invention discloses a method and device for determining the orbit of a lunar-orbiting satellite based on radio and laser ranging, which relates to the technical field of lunar-orbiting satellite orbit calculation, and includes: acquiring the observation data of radio tracking stations and laser stations; establishing a high-precision dynamic model of the lunar-orbiting satellite; establishing a mathematical model for radio and laser ranging of the lunar-orbiting satellite, processing the observation data according to the observation equation, and obtaining the solution values of the dynamic parameters of the satellite at the initial orbit determination moment; substituting the solution values of the dynamic parameters into the satellite perturbed motion equation to calculate and obtain the orbit determination result of the lunar-orbiting satellite. The present invention uses combined orbit determination by radio ranging and laser ranging to achieve precise orbit determination of the lunar-orbiting satellite. Compared with the traditional orbit determination method, the accuracy and reliability are greatly improved, and the requirements for the number and geometric distribution of ground stations are reduced, thus reducing the measurement and control pressure on existing ground stations.

Description

Technical Field

[0001] The present invention relates to the technical field of lunar satellite orbit calculation, and particularly to a method and device for determining the orbit of a lunar satellite based on radio and laser ranging. Background Art

[0002] In lunar exploration missions, the precise orbit determination of lunar satellites is one of the key factors to ensure mission success. Currently, the traditional lunar satellite orbit determination technology is ground-based radio observation technology, including S / X ranging and velocity measurement and very long baseline interferometry (VLBI) measurement technology, which have been fully applied in lunar exploration projects at home and abroad. However, the radio orbit determination technology has limited accuracy and has limitations in lunar exploration missions with high-precision requirements.

[0003] Lunar laser ranging is currently the measurement method with the highest accuracy for the Earth-Moon distance. Its characteristics of high accuracy and long distance have attracted more and more attention in the determination of lunar satellite orbits. However, compared with near-Earth satellites and lunar surface corner reflectors, the initial orbit accuracy of lunar satellites is low, resulting in difficulty for lasers to align and search for satellites. Currently, this method has not been applied to the orbit determination of lunar satellites. And laser ranging is greatly affected by weather, and the amount of effective data obtained is small, and it cannot be used alone for the orbit determination of lunar satellites. A single radio or laser ranging technology may still be affected by various factors in a complex space environment. Therefore, the method of combining radio ranging and laser ranging can comprehensively utilize the advantages of both to improve the accuracy and reliability of lunar satellite orbit determination.

[0004] The Chinese patent application with the publication number CN103363994A proposes a satellite precise orbit determination technology based only on radio carrier phase observation, and designs a technical solution for precisely tracking multiple space targets while observing data of a ground monitoring device driven only by an atomic frequency standard carrier. However, it only uses a single measurement method, and the orbit determination accuracy is low.

[0005] The Chinese patent application with the publication number CN117848354A proposes a multi-modal information fusion optoelectronic detection, positioning and orbit determination device and method for space targets, and uses laser and optical imaging methods for space target positioning. However, it uses optical imaging to obtain the two-dimensional position information of the target, and it is difficult for the telescope to track when the satellite is far away. Summary of the Invention

[0006] To solve the above technical problems, the present invention provides a method and device for determining the orbit of a lunar satellite based on radio and laser ranging, so as to improve the accuracy and reliability of lunar satellite orbit determination.

[0007] To achieve the above object, the present invention adopts the following technical solutions:

[0008] A method for determining the orbit of a lunar-orbiting satellite based on radio and laser ranging, comprising the following steps:

[0009] Step 1: Obtain the observation data of the radio tracking station and the laser station;

[0010] Step 2: Establish a high-precision dynamic model of the lunar-orbiting satellite;

[0011] Step 3: Establish a mathematical model for radio and laser ranging of the lunar-orbiting satellite;

[0012] Step 4: Process the observation data according to the high-precision dynamic model of the lunar-orbiting satellite and the mathematical model for radio and laser ranging of the lunar-orbiting satellite to obtain the calculated values of the dynamic parameters of the satellite at the initial orbit determination moment;

[0013] Step 5: Substitute the calculated values of the dynamic parameters of the satellite at the initial orbit determination moment into the satellite's perturbed motion equation to solve and obtain the orbit determination result of the lunar-orbiting satellite.

[0014] Furthermore, obtain the positions and velocities of the main celestial bodies in the solar system according to the JPL ephemeris;

[0015] Furthermore, the observation equation is for the actual state parameters of the lunar-orbiting satellite and the state parameters of the corresponding preset orbit;

[0016] Furthermore, perform data preprocessing on the radio ranging distance information to generate a data file in the format required for orbit calculation;

[0017] Furthermore, the content of the data preprocessing includes: satellite transmission delay and reception delay, clock face time differences between the satellite and the ground and between the ground and the satellite, ionospheric delay correction, tropospheric delay correction, relativistic delay correction, Sagnac effect correction, data format conversion;

[0018] Furthermore, perform data preprocessing on the laser ranging distance information to generate a data file in the format required for orbit calculation;

[0019] Furthermore, the content of the data preprocessing includes: system delay correction, tropospheric refraction correction, relativistic delay correction, earth tide correction, centroid offset correction of the retroreflector, space-time correction, data format conversion;

[0020] Furthermore, obtain the calculated values of the dynamic parameters of the satellite at the initial orbit determination moment according to the processed observation data and the satellite dynamic equation;

[0021] Utilize the orbit measurement data, combine with the satellite dynamic equation, and perform initial orbit calculation of the satellite based on the Laplace method;

[0022] Within the orbit determination arc segment, utilize the observed values for orbit determination to solve and obtain the orbit parameters of each satellite.

[0023] Calculate the theoretical observation value of the said observation value according to the said solution parameters;

[0024] Calculate the residuals according to the said observation value and the said theoretical observation value, and eliminate the observation values whose residuals exceed the threshold.

[0025] Furthermore, within the orbit determination arc segment, use the observation values for orbit determination to obtain the solution parameters, specifically including:

[0026] Perform orbit integration using the initial orbit information to obtain a reference orbit and a state transition matrix sampled at a certain time interval;

[0027] Linearize the observation equation according to the said reference orbit and the said state transition matrix;

[0028] Solve the linearized observation equation using the linear optimal estimation method to obtain the solution parameters.

[0029] Furthermore, the said observation values include radio observation data and laser observation data; the said theoretical observation values include the theoretical observation values of radio observation data and laser observation data.

[0030] An embodiment of the present invention also provides a lunar orbiting satellite orbit determination device based on radio and laser ranging, including the following modules:

[0031] Observation data acquisition module, which acquires the observation data of the radio tracking station and the laser station;

[0032] Mechanical model establishment module, which establishes a high-precision dynamic model of the lunar orbiting satellite;

[0033] Mathematical model establishment module, which establishes a radio and laser ranging mathematical model of the lunar orbiting satellite;

[0034] Solution module, which processes the observation data according to the radio and laser ranging mathematical model of the lunar orbiting satellite to obtain the solution value of the dynamic parameters of the satellite at the initial orbit determination moment;

[0035] Result acquisition module, which substitutes the solution value of the dynamic parameters of the satellite at the initial orbit determination moment into the satellite's perturbed motion equation to calculate and obtain the orbit determination result of the lunar orbiting satellite.

[0036] An embodiment of the present invention also provides an electronic device, including a memory, a processor, and a computer program stored on the memory and executable on the processor. When the processor executes the program, it implements the steps of the above-mentioned lunar orbiting satellite orbit determination method based on radio and laser ranging.

[0037] An embodiment of the present invention further provides a non-transitory computer-readable storage medium, on which a computer program is stored. When the computer program is executed by a processor, the steps of the above-mentioned lunar satellite orbit determination method based on radio and laser ranging are implemented.

[0038] The beneficial effects of the present invention are as follows:

[0039] 1. The present invention uses combined orbit determination by radio ranging and laser ranging to achieve precise orbit determination of lunar satellites. Compared with traditional orbit determination methods, the accuracy and reliability are greatly improved.

[0040] 2. The present invention uses combined observations of radio ranging and laser ranging, reducing the requirements for the number and geometric distribution of ground stations, and alleviating the measurement and control pressure on existing ground stations. Description of the Drawings

[0041] Figure 1 is a flowchart of the lunar satellite orbit determination method based on radio and laser ranging according to an embodiment of the present invention;

[0042] Figure 2 is a flowchart of the data preprocessing of the present invention;

[0043] Figure 3 is a flowchart of the orbit improvement processing of the present invention. Detailed Embodiments

[0044] To make the objectives, technical solutions and advantages of the present invention clearer and more understandable, the present invention will be further described in detail below with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the present invention and are not used to limit the present invention. In addition, the technical features involved in the various embodiments of the present invention described below can be combined with each other as long as they do not conflict with each other.

[0045] The objective of the present invention is to provide a lunar satellite orbit determination method based on radio and laser ranging to solve the problems of low orbit determination accuracy of current lunar orbit satellites and the limitation of ground measurement and control station resources.

[0046] To make the above objectives, features and advantages of the present invention more obvious and understandable, the present invention will be further described in detail below with reference to the accompanying drawings and specific embodiments.

[0047] Precise orbit determination is a process of using various types of observation data containing observation errors to accurately fit the satellite orbit through dynamics. Obviously, the accuracy of orbit determination calculations is largely limited by the accuracy of measurement data, the accuracy of the observation model, and the accuracy of the dynamics model, in addition to the observation geometry. Precise orbit determination involves the processing of various types of observation data, including accurate observation modeling and the correction of various measurement errors. The dynamic effects on satellites of different orbit types are also different, so the dynamic models considered in orbit determination calculations are also different. In addition, there are certain errors in both the dynamic model and the observation model, especially in the dynamic model part. To sum up, in the process of orbit calculation, it is necessary to select an appropriate mathematical model and optimize the measurement data to improve the accuracy of the calculation results.

[0048] As Figure 1 shown, the method for determining the orbit of a lunar-orbiting satellite based on radio and laser ranging according to the embodiment of the present invention includes the following steps:

[0049] Step 1: Obtain the observation data of the radio tracking station and the laser station. Both the radio and laser observation data are used as observation data.

[0050] Step 2: Establish a high-precision dynamic model for the lunar-orbiting satellite:

[0051] For the lunar-orbiting satellite, the coordinate system selects the epoch (J2000.0) lunar-centered inertial coordinate system. Considering factors such as the gravitational forces of the sun, moon, earth, and other celestial bodies on the lunar satellite, lunar solid tide perturbation, lunar physical libration, solar radiation pressure, jet unloading, and high-precision ephemeris, the high-precision dynamic model of the lunar-orbiting satellite is shown as follows:

[0052] ;

[0053] In the above acceleration formula (i.e., the latter three formulas), the first term on the right side of the equal sign is the lunar center gravitational model, and the second and third terms are the third body gravitational models caused by the sun and the earth.

[0054] The above formula can be abbreviated as:

[0055] ;

[0056] In the formula, is the state vector of the satellite at time t; is the differential of ; is the state vector of the state model; are the positions and velocities of the lunar satellite in the X, Y, and Z directions respectively; is the system nonlinear continuous state transition function of the state model; represents the system noise; Denote the vector from the lunar center to the satellite; Denote the vector from the Earth center to the satellite; The vector from the solar center to the satellite; Are the gravitational constants of the sun, moon, and Earth respectively; Is the vector from the solar center to the satellite; Is the vector from the lunar center to the Earth center in the Earth-centered coordinate system; Is the vector from the lunar center to the solar center; Is the coordinate of the lunar position in the heliocentric coordinate system; Is the coordinate of the Earth position in the Earth-centered coordinate system; Is the coordinate of the solar position in the heliocentric coordinate system; Are the system noises respectively.

[0057] Step 3: Establish the mathematical models for lunar satellite radio ranging and laser ranging:

[0058] The observed value of lunar satellite laser ranging is the distance from the laser station to the satellite. Assume the original observed distance is , and the observation equation is:

[0059] ;

[0060] Among them, Is the actual distance from the laser station to the satellite, Is the ranging error caused by the change in the station position due to the Earth tide; Is the ranging error caused by the refraction effect of light in the atmosphere; Is the ranging error caused by the relativistic effect of light in the gravitational field; Is the deviation of the reflection point of the laser on the satellite surface from the centroid; Is the system delay error of the station.

[0061] Radio ranging is a two-way distance measurement in a time-division system. The ranging value contains the distance and clock difference information between the ground station and the satellite. The radio two-way ranging observation equation is:

[0062] ;

[0063] Among them, Are the observed data for the uplink and downlink respectively; Are the geometric distances from the ground station to the satellite and from the satellite to the ground station at the ranging moment respectively; Are the sums of the space signal delay corrections for the uplink and downlink, specifically including ionospheric delay correction, tropospheric delay correction, relativistic delay correction, and Sagnac effect correction, etc. Are the transmission and reception delays of the satellite and ground equipment respectively; Is the time difference between the two satellite-ground and ground-satellite clock faces; is the random error, and c is the speed of light.

[0064] Step Four: Preprocess the observation data (i.e., Figure 2 the laser ranging / radio ranging data):

[0065] As Figure 2 shown, the specific steps of the preprocessing include:

[0066] Step 4.1: Based on the initial state parameters of the satellite and the high-precision lunar-orbiting satellite dynamics model, use the numerical integration method to integrate the satellite orbit to generate theoretical orbit prediction data (i.e., Figure 2 the spatio-temporal conversion):

[0067] Step 4.2: Process the observation data according to the observation equation and the theoretical orbit prediction data, and deduct the error terms in the data to obtain the standard point data of the satellite observation;

[0068] In laser ranging, the systematic error is the ground station laser ranging system delay error; in radio ranging, the systematic error is the transmission and reception time delays between the ground and the satellite and the clock face time difference, including (i.e., perform Figure 2 the systematic delay correction):

[0069] Step 4.2.1: Tropospheric refraction correction; based on the temperature, humidity, and pressure data measured at the station, use the atmospheric refraction correction model to complete the tropospheric refraction correction for satellite ranging and velocity measurement; perform tropospheric delay correction.

[0070] Step 4.2.2: Relativistic delay correction: Calculate the bending of light caused by the gravitational field of celestial bodies according to the distances from the satellite to various celestial bodies to complete the relativistic time delay correction.

[0071] Step 4.2.3: Earth tide correction: Calculate the coordinate deformation of the ground station according to the solid tide, ocean tide, atmospheric load tide generated by the external gravitational force on the Earth, and the solid polar tide and ocean polar tide parameters caused by the centrifugal perturbation due to the Earth's rotation to complete the Earth tide correction.

[0072] Step 4.2.4: Retroreflector centroid offset correction: Complete the centroid offset correction according to the distance from the retroreflector on the satellite to the satellite centroid.

[0073] Step 4.2.5: Satellite antenna phase center correction: Correct the phase offset of the satellite antenna;

[0074] Step 4.2.6: Ionospheric delay correction, Sagnac effect correction.

[0075] Step 4.3 Select the initial satellite orbit determination time, and obtain the calculated values of the dynamic parameters of the satellite at the initial orbit determination time according to the standard point data observed by the satellite and the high-precision dynamic model of the lunar-orbiting satellite.

[0076] After that, data format conversion is performed, that is, format conversion is carried out after the correction processing of the orbit measurement data to generate a data file in the format required for orbit calculation.

[0077] Step Five: Solve the position coordinates of each satellite in the formation, as Figure 3 shown, including the following steps:

[0078] Step 5.1 Use the orbit measurement data to calculate and generate the initial orbit information of the satellite for use in orbit improvement (that is, Figure 3 the initial orbit calculation using the orbit measurement data combined with the satellite dynamic equation in

[0079] Based on the Laplace method, use the radio observation data to complete the initial orbit calculation of the satellite. The dynamic model used in the initial orbit calculation is the model in Step Two, which is applicable to lunar-orbiting satellites.

[0080] Step 5.2 Use the initial orbit information to perform orbit integration, and calculate the reference orbit and the state transition matrix;

[0081] Step 5.3 Linearize the observation equation according to the reference orbit and the state transition matrix:

[0082] Step 5.4 Perform linear optimal estimation on the linearized observation equation to obtain the calculated parameters:

[0083] The state differential equation of satellite motion can be expressed as:

[0084] ;

[0085] where, ; and represent the state quantities to be estimated; represents the dynamic parameters; represents the state quantity to be estimated at the initial time of the orbit determination arc; represents the initial time of the orbit determination arc; represents the state quantity at the initial time of the orbit determination arc; represents the satellite position; represents the satellite velocity; represents the satellite acceleration; represents other parameters to be estimated in the dynamic model.

[0086] Expand in the above formula at the reference state , and denote , after omitting the higher-order terms, it can be expressed as a linear equation:

[0087] ;

[0088] Among them, the intermediate function ;

[0089] Its solution can be expressed as ;

[0090] The solution is ;

[0091] Among them, represents the state transition matrix, represents the derivative of the state transition matrix, represents the initial value of the state transition matrix; represents the identity matrix; represents the reference state; represents the difference between the state to be estimated and the reference state; represents time; represents the derivative of the difference between the state to be estimated and the reference state; represents the difference between the state to be estimated and the reference state at time

[0092] The observable quantity of the satellite at time is , represents the measurement noise, represents the state vector of the satellite at time; represents the observation time; represents the observation data corresponding true value; i represents the i-th data.

[0093] Expanding the above formula at the reference state and only considering the first-order terms, we have:

[0094] ;

[0095] Among them, represents the actual value, represents the observed value, represents the observation partial derivative of the observable quantity with respect to the state quantity at the observation epoch, represents the observation partial derivative with respect to the improved epoch state quantity.

[0096] A linear equation can be obtained ;

[0097] Among them, represents the state quantity to be estimated; represents the random error.

[0098] Step 5.5 Calculate the theoretical observation values according to the solution parameters:

[0099] Solve for the best estimate value, and usually use the least squares method for parameter estimation. Denote the weight matrix of the observed quantities as and, according to the linear unbiased minimum variance estimation, the estimate value can be obtained as follows:

[0100] ;

[0101] wherein, represents transpose; k represents the amount of observed data.

[0102] Step 5.6 Calculate the residuals and eliminate the observed values whose residuals exceed the threshold:

[0103] Both the state equation and the observation equation are the results of linearized approximation, and the errors brought by the non-linear part are solved by continuous iteration, and the observed values whose residuals exceed the threshold are eliminated.

[0104] The embodiment of the present invention also provides a lunar orbiting satellite orbit determination device based on radio and laser ranging, including the following modules:

[0105] Observation data acquisition module, which acquires the observation data of the radio tracking station and the laser station;

[0106] Mechanical model establishment module, which establishes a high-precision dynamic model of the lunar orbiting satellite;

[0107] Mathematical model establishment module, which establishes a radio and laser ranging mathematical model of the lunar orbiting satellite;

[0108] Solution module, which processes the observation data according to the high-precision dynamic model of the lunar orbiting satellite and the radio and laser ranging mathematical model of the lunar orbiting satellite to obtain the solution value of the dynamic parameters of the satellite at the initial orbit determination moment;

[0109] Result acquisition module, which substitutes the solution value of the dynamic parameters of the satellite at the initial orbit determination moment into the satellite perturbed motion equation to calculate and obtain the orbit determination result of the lunar orbiting satellite.

[0110] The embodiment of the present invention also provides an electronic device, including a memory, a processor and a computer program stored on the memory and executable on the processor, and when the processor executes the program, it implements the steps of the above-mentioned lunar orbiting satellite orbit determination method based on radio and laser ranging.

[0111] The embodiment of the present invention also provides a non-transitory computer-readable storage medium, on which a computer program is stored, and when the computer program is executed by a processor, it implements the steps of the above-mentioned lunar orbiting satellite orbit determination method based on radio and laser ranging.

[0112] Those skilled in the art should understand that the embodiments of the present invention can be provided as a method, a system, or a computer program product. Therefore, the present invention can take the form of a completely hardware embodiment, a completely software embodiment, or an embodiment combining software and hardware aspects. Moreover, the present invention 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-ROM, optical storage, etc.) that contain computer-usable program code. The solutions in the embodiments of the present invention can be implemented in various computer languages. For example, object-oriented programming languages such as Java and interpreted scripting languages such as JavaScript, etc.

[0113] The present invention is described with reference to the flowcharts and / or block diagrams of methods, apparatuses (systems), and computer program products according to the embodiments of the present invention. It should be understood that each flow and / or block in the flowchart and / or block diagram, and the combination of flows and / or blocks in the flowchart and / or block diagram, can be implemented by computer program instructions. These computer program instructions can be provided to the processor of a general-purpose computer, a special-purpose computer, an embedded processor, or other programmable data processing devices to generate a machine, such that the instructions executed by the processor of the computer or other programmable data processing devices generate a device for implementing the functions specified in Figure 1 one flow or multiple flows and / or blocks Figure 1 one block or multiple blocks.

[0114] These computer program instructions can also be stored in a computer-readable memory that can direct a computer or other programmable data processing device to work in a specific manner, such that the instructions stored in the computer-readable memory generate a manufactured article including an instruction device, and the instruction device implements the functions specified in Figure 1 one flow or multiple flows and / or blocks Figure 1 one block or multiple blocks.

[0115] These computer program instructions can also be loaded onto a computer or other programmable data processing device, such that a series of operation steps are executed on the computer or other programmable device to generate a computer-implemented process, and thus the instructions executed on the computer or other programmable device provide steps for implementing the functions specified in Figure 1 one flow or multiple flows and / or blocks Figure 1 one block or multiple blocks.

[0116] Although the preferred embodiments of the present invention have been described, additional changes and modifications can be made by those skilled in the art once they learn of the basic creative concept. Therefore, the appended claims are intended to be interpreted to include the preferred embodiments as well as all changes and modifications that fall within the scope of the present invention.

[0117] Obviously, those skilled in the art can make various changes and modifications to the present invention without departing from the spirit and scope of the present invention. Thus, if these modifications and variations of the present invention fall within the scope of the claims of the present invention and their equivalent technologies, the present invention is also intended to include these modifications and variations.

Claims

1. A method for determining the orbit of a lunar satellite based on radio and laser ranging, characterized in that: The steps include: Step 1: Obtain observation data from radio tracking stations and laser stations; Step 2: Establish a high-precision dynamic model of the lunar satellite; Step 3: Establish mathematical models of radio and laser ranging of lunar satellites; Step 4: Process the observation data according to the high-precision dynamics model of the lunar satellite and the mathematical model of radio and laser ranging of the lunar satellite to obtain the dynamic parameter solution value of the satellite at the initial time of orbit determination, including: Step 4.1 Based on the initial state parameters of the satellite and the high-precision dynamic model of the lunar satellite, the satellite orbit is integrated using a numerical integration method to generate theoretical orbit prediction data; Step 4.2: Process the observation data according to the observation equation and theoretical orbit prediction data, deduct the error term in the observation data, and obtain the standard point data of satellite observation; Step 4.3: Select the initial time of satellite orbit determination, and obtain the solution value of the satellite's dynamic parameters at the initial time of orbit determination based on the standard point data of satellite observation and the high-precision dynamic model of the lunar satellite; Step 5: Substitute the calculated values ​​of the satellite's dynamic parameters at the initial moment of orbit determination into the satellite's motion equation to obtain the orbit determination result of the lunar satellite.

2. The method for determining the orbit of a lunar satellite based on radio and laser ranging according to claim 1, characterized in that: In the step 1, the observation values ​​of radio ranging and laser ranging are obtained from the ground station.

3. The method for determining the orbit of a lunar satellite based on radio and laser ranging according to claim 1, characterized in that: In step 2, the high-precision dynamic model of the lunar satellite takes into account the central celestial body gravity model and the third body gravity model caused by the sun and the earth.

4. The method for determining the orbit of a lunar satellite based on radio and laser ranging according to claim 1, characterized in that: The step three comprises: The lunar satellite laser ranging observation equation is constructed as follows: ； in, is the actual distance from the laser station to the satellite, The original observation distance is, The ranging error caused by the change of station position due to the earth's tides; The distance measurement error caused by the refraction effect of light in the atmosphere; The distance measurement error caused by the relativistic effect of light in the gravitational field; is the deviation of the laser reflection point on the satellite surface from the center of mass; is the system delay error of the station.

5. The method for determining the orbit of a lunar satellite based on radio and laser ranging according to claim 4, characterized in that: The step three comprises: The observation equation for radio ranging of lunar satellite is constructed as follows: ； in, They are the observation data of uplink and downlink respectively; are the geometric distances from the ground station to the satellite and from the satellite to the ground station at the time of ranging, respectively; are the sum of uplink space signal delay correction and the sum of downlink space signal delay correction respectively; They are satellite launch, satellite reception delay, ground equipment launch, and ground equipment reception delay; The time difference between the two clock faces is: star-earth, and earth-star; is the random error, and c is the speed of light.

6. The method for determining the orbit of a lunar satellite based on radio and laser ranging according to claim 1, characterized in that: The step five comprises: Step 5.1: Use the standard point data of the satellite observation in step 4 as the initial state parameters, and combine it with the high-precision dynamic model of the lunar satellite to calculate and generate the initial orbit information of the satellite for orbit improvement; Step 5.2: Based on the high-precision dynamics model of the lunar satellite, the initial orbit information of the satellite is used to perform orbit integration and calculate the reference orbit and state transfer matrix; Step 5.3: Linearize the observation equation according to the reference orbit and state transfer matrix; Step 5.4 performs linear optimal estimation on the linearized observation equation to obtain the solution parameters; Step 5.5 calculates theoretical observation values ​​according to the solution parameters; Step 5.6 calculates the residuals and removes observations whose residuals exceed the threshold.

7. A device for determining the orbit of a lunar satellite based on radio and laser ranging, characterized in that: Includes the following modules: Observation data acquisition module, which acquires observation data from radio tracking stations and laser stations; Mechanical model building module, to build a high-precision dynamic model of the lunar satellite; Mathematical model building module, which builds mathematical models of radio and laser ranging of lunar satellites; The solution module processes the observation data according to the high-precision dynamics model of the lunar satellite and the mathematical model of radio and laser ranging of the lunar satellite to obtain the solution value of the satellite's dynamic parameters at the initial moment of orbit determination; The result acquisition module substitutes the calculated values ​​of the satellite's dynamic parameters at the initial moment of orbit determination into the satellite's motion equation to obtain the orbit determination result of the lunar satellite.

8. 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 program, the steps of the method for determining the orbit of a lunar satellite based on radio and laser ranging as described in one of claims 1 to 6 are implemented.

9. A non-transitory computer-readable storage medium having a computer program stored thereon, characterized in that: When the computer program is executed by a processor, the steps of the method for determining the orbit of a lunar satellite based on radio and laser ranging as described in any one of claims 1 to 6 are implemented.

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