Design method for transfer orbit from DRO orbit to lunar low orbit circular orbit

Abstract

The invention relates to a method for designing a transfer orbit from a DRO orbit to a lunar low-orbit circular orbit. The method comprises the following steps: establishing a mapping data table and an orbital transfer speed increment three-dimensional data pair according to a preset DRO orbit size parameter and a DRO aircraft initial moment phase angle; dividing the orbital transfer speed increment three-dimensional data pair into two groups according to positive and negative conditions of an orbital transfer speed increment elevation angle, acquiring ascending node longitude and transfer flight time corresponding to the data pair of the speed increment in the minimum transfer process of each group to calculate orbital transfer moments, and selecting a minimum value in the two orbital transfer moments as a selected orbital transfer moment; correcting the phase at the selected orbital transfer moment, and calculating an orbital transfer speed increment projection according to the corrected phase; and adopting a Newton iteration method to take the projection of the selected orbital transfer moment and orbital transfer speed increment as an iteration initial value to carry out iteration solution, and obtaining the transfer orbital transfer moment and the transfer orbital transfer speed increment. By adopting the method, the transfer track design precision can be improved.

Description

Technical Field

[0001] The present application relates to the field of data processing technology, and in particular to a method for designing a transfer orbit from a DRO orbit to a low-moon circular orbit. Background Art

[0002] DRO (Distant Retrograde Orbit) is a periodic orbit proposed under the assumption of a circular restricted three-body model. In actual problems, although the forces on the DRO orbiter are far more complex than the circular restricted three-body force model, its periodicity and stability can still be maintained to a certain extent. Therefore, it has unique application value in the exploration of the Earth-Moon space and deep space beyond the Moon. Transferring from the DRO orbit to a low-moon circular orbit is an important mode of DRO orbit application, which is of great value in the application of lunar exploration and lunar transfer.

[0003] The common Earth-Moon space transfer orbit design problem is based on a circular restrictive three-body model, an elliptical restrictive three-body model or a restrictive four-body model. Although this type of analysis and design can obtain relatively concise theoretical results and can also perform qualitative characteristic analysis, due to the large difference between the force model and the actual force conditions, and the fact that the spacecraft orbit itself is sensitive to the initial value in the multi-celestial body system, in actual engineering calculations, the error of the calculation result often cannot meet the iterative initial value accuracy requirements. The use of a numerical integration method based on a high-precision force model can effectively improve the calculation accuracy, and through intelligent algorithms to perform search iterations, high-precision calculation results can be obtained in a global range. However, due to the sensitivity of the spacecraft orbit itself to the initial value, this type of intelligent search algorithm often converges slowly, and the solution time and convergence reliability are difficult to guarantee. From the current public information, there is still a lack of a design method with high calculation accuracy, fast convergence and reliability for the orbit design problem of DRO orbit transfer to a low-moon circular orbit. Summary of the invention

[0004] Based on this, it is necessary to provide a transfer orbit design method from DRO orbit to low-lunar circular orbit with high precision and fast convergence speed to address the above technical problems.

[0005] A method for designing a transfer orbit from a DRO orbit to a low-moon orbit circular orbit, the method comprising: Obtain the DRO orbit-related parameters and target orbit-related parameters at the time of the spacecraft's orbit change; the DRO orbit-related parameters include the DRO orbit size parameters and the DRO spacecraft's initial phase angle; the target orbit-related parameters include the target orbit lunar altitude, the target orbit inclination, and the ascending node longitude at the initial time of the target orbit; A mapping data table is established according to the preset DRO orbit size parameters and the phase angle of the DRO spacecraft at the initial moment; a mapping data table corresponding to the value closest to the preset DRO orbit size parameters is selected according to the DRO orbit size parameters at the time of the spacecraft's orbit change; and a three-dimensional data pair of orbit change velocity increment whose perigee orbit height and orbit surface inclination meet the preset scattering range is selected according to the target orbit lunar surface height and the target orbit inclination; The three-dimensional data pairs of orbit change velocity increments within the preset scattering range are divided into two groups according to the positive and negative conditions of the orbit change velocity increment elevation angles, corresponding to the two transfer orbits of ascending and descending transfers, and the orbit change time is calculated by obtaining the data pairs of the minimum velocity increment in each transfer process corresponding to the ascending node longitude and the transfer flight time, and the minimum value of the two orbit change times is selected as the selected orbit change time; The phase of the selected orbit change moment is corrected to obtain the corrected phase; the orbit change velocity increment projection is calculated according to the corrected phase; the Newton iteration method is used to iteratively solve the selected orbit change moment and the orbit change velocity increment projection as the iteration initial values ​​to obtain the transfer orbit change moment and the transfer orbit change velocity increment.

[0006] In the above-mentioned transfer orbit design method from DRO orbit to low-moon orbit, this application comprehensively and systematically solves the accuracy and convergence problems faced in the design of DRO orbit transfer to low-moon orbit circular orbit by accurately obtaining orbit parameters, establishing efficient mapping data tables, carefully screening orbit data pairs, reasonably determining the orbit change time, and adopting scientific phase correction and iterative solution methods. This method is closely adapted to the high-precision force model, providing a reliable and efficient solution for orbit design in aerospace engineering, and has important practical application value. BRIEF DESCRIPTION OF THE DRAWINGS

[0007] Figure 1 A schematic diagram of a flow chart of a method for designing a transfer orbit from a DRO orbit to a low-moon circular orbit in one embodiment; Figure 2 FIG. 4 is a diagram showing the internal structure of a computer device in one embodiment. DETAILED DESCRIPTION

[0008] In order to make the purpose, technical solution and advantages of the present application more clearly understood, the present application is further described in detail below in conjunction with the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the present application and are not used to limit the present application.

[0009] In one embodiment, Figure 1 As shown, a transfer orbit design method from a DRO orbit to a low lunar orbit circular orbit is provided, comprising the following steps: Step 102, obtaining DRO orbit-related parameters and target orbit-related parameters at the time of the spacecraft orbit change; the DRO orbit-related parameters include DRO orbit size parameters and the DRO spacecraft initial phase angle; the target orbit-related parameters include the target orbit lunar altitude, the target orbit inclination and the ascending node longitude at the initial time of the target orbit.

[0010] The DRO orbit-related parameters include the DRO orbit size parameters and the DRO spacecraft initial phase angle, which accurately describe the state of the spacecraft on the initial orbit. The target orbit-related parameters, including the target orbit lunar altitude, target orbit inclination, and the longitude of the ascending node at the initial moment of the target orbit, clearly define the characteristics of the target orbit that the spacecraft will eventually reach. By accurately mastering these parameters, subsequent calculations can be carried out closely around the actual operating state and target state of the spacecraft, avoiding calculation deviations caused by inaccurate parameters, and providing reliable data support for accurate calculations based on high-precision force models.

[0011] Step 104, establishing a mapping data table according to the preset DRO orbit size parameters and the phase angle of the DRO spacecraft at the initial moment; selecting a mapping data table corresponding to the value closest to the preset DRO orbit size parameters according to the DRO orbit size parameters at the moment of the spacecraft's orbit change; and selecting a three-dimensional data pair of orbit change velocity increments whose perigee orbit height and orbital surface inclination meet a preset scatter range according to the target orbit lunar surface height and the target orbit inclination.

[0012] The mapping data table is established based on the preset DRO orbit size parameters and the phase angle of the DRO spacecraft at the initial moment, which greatly optimizes the calculation process. In practical applications, when faced with the DRO orbit size parameters at the moment of the spacecraft's orbit change, the mapping data table corresponding to the value closest to the preset DRO orbit size parameters can be quickly selected. This method avoids blind searching in a large amount of data and greatly improves the speed and accuracy of data matching. Through efficient data screening, unnecessary calculations are reduced, which provides a strong guarantee for the subsequent rapid and accurate determination of the orbit transfer plan and plays a positive role in improving the convergence speed.

[0013] According to the target orbit lunar surface altitude and target orbit inclination, select the three-dimensional data pairs of orbit change velocity increments whose perigee orbit altitude and orbital surface inclination meet the preset scatter range to ensure orbit accuracy. In a complex multi-celestial body system, the spacecraft orbit is extremely sensitive to the initial value, and a slight deviation may cause the orbit calculation results to be far from the actual requirements. By strictly limiting the screening of data pairs within the preset scatter range, orbital schemes that do not meet the accuracy requirements can be effectively excluded, so that subsequent calculations are based on orbital parameters that are closer to actual needs. This fine screening of orbital parameters ensures that the entire orbit design process is compatible with the high-precision force model, thereby significantly improving the calculation accuracy and overcoming the problem of low accuracy of the traditional simplified force model.

[0014] Step 106, from the three-dimensional data pairs of orbit change velocity increments that meet the preset scatter range, they are divided into two groups according to the positive and negative conditions of the altitude angles of the orbit change velocity increments, corresponding to the two transfer orbits of ascending orbit and descending orbit transfer, and the data pairs of the minimum velocity increment in each transfer process are obtained to calculate the orbit change time according to the corresponding ascending node longitude and transfer flight time, and the minimum value of the two orbit change times is selected as the selected orbit change time.

[0015] From the three-dimensional data of the orbit change speed increment within the preset dispersion range, it is divided into two groups according to the positive and negative conditions of the high and low angles of the orbit change speed increment, corresponding to the two transfer orbits of ascending and descending transfer, and then the data of the speed increment in each group of the smallest transfer process is obtained to calculate the orbit change time corresponding to the ascending node longitude and transfer flight time, and the minimum of the two orbit change times is selected as the selected orbit change time, which fully considers the various actual situations in the orbit transfer process, and finds the optimal orbit change time through detailed analysis of different transfer orbits. Selecting the minimum orbit change time, while meeting the orbit accuracy requirements, minimizes the energy consumption and time cost of the aircraft during the orbit transfer process, and also further improves the convergence stability of the entire orbit design scheme, avoiding orbit deviation and convergence problems caused by unreasonable orbit change time selection, and adopting this application has a fast convergence speed.

[0016] Step 108, correct the phase of the selected orbit change moment to obtain a corrected phase; calculate the orbit change velocity increment projection based on the corrected phase; use the Newton iteration method to iteratively solve the selected orbit change moment and the orbit change velocity increment projection as the iteration initial values ​​to obtain the transfer orbit change moment and the transfer orbit change velocity increment.

[0017] The phase of the selected orbit change moment is corrected to obtain the corrected phase, and the orbit change velocity increment projection is calculated based on the corrected phase. The Newton iteration method is used to iteratively solve the selected orbit change moment and orbit change velocity increment projection as the initial values ​​of the iteration, and finally the transfer orbit change moment and transfer orbit change velocity increment are obtained. The phase correction further optimizes the initial conditions of the orbit calculation, so that the iterative process can converge to the real orbit transfer parameters more accurately, with the characteristics of fast convergence speed and high precision. Combining the carefully determined initial values ​​of the iteration with the Newton iteration method, it is possible to quickly and accurately search for orbit transfer parameters that meet the high-precision force model in a global range, thereby obtaining high-precision calculation results, which effectively solves the problems of slow convergence speed, long solution time and difficult to ensure convergence reliability of traditional intelligent search algorithms.

[0018] In the above-mentioned transfer orbit design method from DRO orbit to low-moon orbit, this application comprehensively and systematically solves the accuracy and convergence problems faced in the design of DRO orbit transfer to low-moon orbit by accurately obtaining orbit parameters, establishing an efficient mapping data table, carefully screening orbit data pairs, reasonably determining the orbit change time, and adopting scientific phase correction and iterative solution methods. This method is closely adapted to the high-precision force model, providing a reliable and efficient solution for orbit design in aerospace engineering, and has important practical application value.

[0019] In one embodiment, a mapping data table is established according to the preset DRO orbit size parameters and the DRO vehicle initial phase angle, including: According to the preset DRO orbit size parameters and the phase angle of the DRO spacecraft at the initial moment, interval sampling is performed within the preset range of the orbit change speed increment. The lunar surface altitude, inclination, ascending node longitude and transfer flight time at the time of flight to the perigee are calculated through numerical integration of the high-precision force model. A mapping data table of the orbit change speed increment three-dimensional data pairs and the lunar surface altitude, inclination, ascending node longitude and transfer flight time at the time of perigee is established.

[0020] In a specific embodiment, the DRO orbit phase at the time of the aircraft change , and the DRO orbit size parameters are 60000km, 65000km, 70000km, 75000km, and 80000km respectively. The orbit change speed increment size is sampled at intervals of 1m / s in the range of [200m / s, 350m / s], the orbit change speed increment azimuth is sampled at intervals of 0.5° in the range of [100°, 130°], and the orbit change speed increment elevation angle is sampled at intervals of 0.5° in the range of [-15°, 15°]. Through numerical integration based on a high-precision force model, the lunar surface altitude, inclination, ascending node longitude, and transfer flight time at the time of flight to the perigee are calculated, thereby establishing a three-dimensional data pair of orbit change speed increment. , and the lunar surface height at perigee ,inclination , Longitude of ascending node and transfer flight time Mapping data table.

[0021] In one embodiment, the lunar surface altitude, inclination, ascending node longitude and transfer flight time at the time of flight to the perigee are calculated by numerical integration of a high-precision force model, and a mapping data table of the orbit change velocity increment three-dimensional data pairs and the lunar surface altitude, inclination, ascending node longitude and transfer flight time at the time of perigee is established, including: The lunar surface altitude, inclination, ascending node longitude and transfer flight time at the time of flight to the perigee are calculated by numerical integration of the high-precision force model, and the mapping data table of the three-dimensional data of the orbit change velocity increment and the lunar surface altitude, inclination, ascending node longitude and transfer flight time at the time of perigee is established as follows: ; ; ; ; in, Represents the three-dimensional data pair of track change speed increment, Indicates the lunar height at perigee, The inclination angle at the time of perigee, represents the longitude of the ascending node at the time of perigee, Indicates the transfer flight time at the perigee time, Indicates the orbit change speed increment parameter - lunar height, Indicates the track change speed increment parameter - inclination, Indicates the orbit change speed increment parameter - the longitude of the ascending node, Indicates the incremental parameter of the change-track speed - transfer flight time, Indicates the speed increment during the transfer process. Indicates the track change speed increment azimuth, Indicates the height angle of the track change speed increment.

[0022] In one embodiment, the preset DRO orbit size parameters include 60,000 km, 65,000 km, 70,000 km, 75,000 km, and 80,000 km.

[0023] In a specific embodiment, according to a given DRO track size parameter , among 60000km, 65000km, 70000km, 75000km, and 80000km, select the mapping data table corresponding to the value closest to the value.

[0024] In one embodiment, selecting a three-dimensional data pair of orbit change velocity increments whose perigee orbit altitude and orbital surface inclination meet a preset scattering range according to the target orbit lunar surface altitude and the target orbit inclination includes: According to the target orbit lunar surface height and target orbit inclination, the orbital height and orbital surface inclination of the perigee point are selected to meet the preset scattering range of the orbit change velocity increment three-dimensional data pairs. ; in, and , represents the lunar altitude of the target orbit, represents the target orbit inclination, represents the orbital height of the perigee, represents the orbital inclination, It indicates the allowable deviation range of the perigee orbit altitude relative to the target orbit lunar surface altitude. Indicates the allowable deviation range of the orbital inclination relative to the target orbital inclination.

[0025] In a specific embodiment, according to a given target orbit height and orbital inclination , filter out the perigee orbit height and orbital inclination Satisfy a certain distribution range Value, that is, in the mapping data table, find and At the same time, the following conditions are met Data pairs. In the obtained data pairs, according to The positive and negative cases are divided into two groups, corresponding to the two transfer orbits of ascending and descending transfer. The smallest data pair is denoted as ( , , ), and the corresponding ascending node longitude and transfer flight time can be obtained, recorded as , .

[0026] In one embodiment, obtaining data of each set of minimum velocity increments during the transfer process and calculating the orbit change time for the corresponding ascending node longitude and transfer flight time includes: Obtain the data of the minimum velocity increment in each group of transfer processes, and calculate the orbit change time for the corresponding ascending node longitude and transfer flight time: ; in, is the lunar revolution angular velocity, is the average orbital angular velocity of the DRO spacecraft in the L2 coordinate system, The data representing each set of the minimum velocity increment in the transfer process corresponds to the transfer flight time, represents the longitude of the ascending node, represents the longitude of the ascending node at the initial moment of the target orbit, Indicates the phase angle of the DRO aircraft at the initial moment, Indicates the initial time.

[0027] In a specific embodiment, is the average orbital angular velocity of the DRO spacecraft (in the L2 coordinate system), which can be obtained by interpolation according to Table 1 based on the given DRO size parameters.

[0028] Table 1

[0029] Positive and negative groups and Two orbit change times are calculated separately, and the smaller value is taken as the selected orbit change time.

[0030] In one embodiment, the phase at the selected track change time is corrected to obtain a corrected phase, including: The phase at the selected orbit change time is calculated as ; in, Indicates the phase angle of the DRO aircraft at the initial moment, is the average orbital angular velocity of the DRO spacecraft in the L2 coordinate system, represents the selected track change time, Indicates the initial time.

[0031] In one embodiment, the phase of the selected track change time includes the azimuth and height angle The corrected result is recorded as , ,have

[0032] Among them, the correction factor and Obtained according to the interpolation table obtained from the previous simulation experiment.

[0033] In a specific embodiment, The interpolation table is shown in Table 2. The interpolation table is shown in Table 3.

[0034] Table 2

[0035] Table 3

[0036] In one embodiment, calculating the track change velocity increment projection according to the corrected phase includes: The projection of the track change velocity increment is calculated based on the corrected phase: ; in, represents the corrected azimuth, represents the corrected elevation angle, represents the phase of the selected track change time, Indicates the minimum track change speed increment, Represents a coordinate transformation matrix.

[0037] It should be understood that although Figure 1 The steps in the flowchart are shown in sequence as indicated by the arrows, but these steps are not necessarily executed in the order indicated by the arrows. Unless otherwise specified in this document, there is no strict order restriction for the execution of these steps, and these steps can be executed in other orders. Moreover, Figure 1 At least part of the steps may include multiple sub-steps or multiple stages. These sub-steps or stages are not necessarily executed at the same time, but can be executed at different times. The execution order of these sub-steps or stages is not necessarily sequential, but can be executed in turn or alternately with other steps or at least part of the sub-steps or stages of other steps.

[0038] In one embodiment, a computer device is provided. The computer device may be a terminal, and its internal structure diagram may be as follows: Figure 2As shown. The computer device includes a processor, a memory, a network interface, a display screen and an input device connected through a system bus. Among them, the processor of the computer device is used to provide computing and control capabilities. The memory of the computer device includes a non-volatile storage medium and an internal memory. The non-volatile storage medium stores an operating system and a computer program. The internal memory provides an environment for the operation of the operating system and the computer program in the non-volatile storage medium. The network interface of the computer device is used to communicate with an external terminal through a network connection. When the computer program is executed by the processor, a transfer orbit design method from a DRO orbit to a low-moon circular orbit is implemented. The display screen of the computer device can be a liquid crystal display screen or an electronic ink display screen, and the input device of the computer device can be a touch layer covered on the display screen, or a key, trackball or touchpad set on the computer device housing, or an external keyboard, touchpad or mouse, etc.

[0039] Those skilled in the art will understand that Figure 2 The structure shown in the figure is only a block diagram of a part of the structure related to the solution of the present application, and does not constitute a limitation on the computer device to which the solution of the present application is applied. The specific computer device may include more or fewer components than those shown in the figure, or combine certain components, or have a different arrangement of components.

[0040] Those of ordinary skill in the art can understand that all or part of the processes in the above-mentioned embodiment methods can be implemented by instructing the relevant hardware through a computer program, and the computer program can be stored in a non-volatile computer-readable storage medium. When the computer program is executed, it can include the processes of the embodiments of the above-mentioned methods. Among them, any reference to memory, storage, database or other media used in the embodiments provided in this application may include non-volatile and / or volatile memory. Non-volatile memory may include read-only memory (ROM), programmable ROM (PROM), electrically programmable ROM (EPROM), electrically erasable programmable ROM (EEPROM) or flash memory. Volatile memory may include random access memory (RAM) or external cache memory. By way of illustration and not limitation, RAM is available in many forms, such as static RAM (SRAM), dynamic RAM (DRAM), synchronous DRAM (SDRAM), double data rate SDRAM (DDRSDRAM), enhanced SDRAM (ESDRAM), synchronous link (Synchlink) DRAM (SLDRAM), memory bus (Rambus) direct RAM (RDRAM), direct memory bus dynamic RAM (DRDRAM), and memory bus dynamic RAM (RDRAM), etc.

[0041] The technical features of the above embodiments may be combined arbitrarily. To make the description concise, not all possible combinations of the technical features in the above embodiments are described. However, as long as there is no contradiction in the combination of these technical features, they should be considered to be within the scope of this specification.

[0042] The above-described embodiments only express several implementation methods of the present application, and the descriptions thereof are relatively specific and detailed, but they cannot be understood as limiting the scope of the invention. It should be pointed out that, for a person of ordinary skill in the art, several modifications and improvements can be made without departing from the concept of the present application, and these all belong to the protection scope of the present application. Therefore, the protection scope of the present application shall be subject to the attached claims.

Claims

1. A transfer orbit design method from a DRO orbit to a low-moon orbit circular orbit, characterized in that: The method comprises: Acquire the DRO orbit-related parameters and target orbit-related parameters at the time of the spacecraft orbit change; the DRO orbit-related parameters include the DRO orbit size parameters and the DRO spacecraft initial phase angle; the target orbit-related parameters include the target orbit lunar surface altitude, the target orbit inclination and the ascending node longitude at the initial time of the target orbit; A mapping data table is established according to the preset DRO orbit size parameters and the phase angle of the DRO spacecraft at the initial moment; a mapping data table corresponding to the value closest to the preset DRO orbit size parameters is selected according to the DRO orbit size parameters at the time of the spacecraft's orbit change; and a three-dimensional data pair of orbit change velocity increment whose perigee orbit height and orbit surface inclination meet the preset scattering range is selected according to the target orbit lunar surface height and the target orbit inclination; The three-dimensional data pairs of orbit change velocity increments within the preset scattering range are divided into two groups according to the positive and negative conditions of the orbit change velocity increment elevation angles, corresponding to the two transfer orbits of ascending and descending transfers, and the orbit change time is calculated by obtaining the data pairs of the minimum velocity increment in each transfer process corresponding to the ascending node longitude and the transfer flight time, and the minimum value of the two orbit change times is selected as the selected orbit change time; The phase of the selected orbit change moment is corrected to obtain a corrected phase; the orbit change velocity increment projection is calculated according to the corrected phase; the Newton iteration method is used to iteratively solve the selected orbit change moment and the orbit change velocity increment projection as iteration initial values ​​to obtain the transfer orbit change moment and the transfer orbit change velocity increment.

2. The method according to claim 1, characterized in that A mapping data table is established based on the preset DRO orbit size parameters and the DRO spacecraft initial phase angle, including: According to the preset DRO orbit size parameters and the phase angle of the DRO spacecraft at the initial moment, interval sampling is performed within the preset range of the orbit change speed increment. The lunar surface altitude, inclination, ascending node longitude and transfer flight time at the time of flight to the perigee are calculated through numerical integration of the high-precision force model. A mapping data table of the orbit change speed increment three-dimensional data pairs and the lunar surface altitude, inclination, ascending node longitude and transfer flight time at the time of perigee is established.

3. The method according to claim 2, characterized in that The lunar surface altitude, inclination, ascending node longitude and transfer flight time at the time of flight to the perigee are calculated by numerical integration of the high-precision force model, and a mapping data table of the three-dimensional data of the orbit change velocity increment and the lunar surface altitude, inclination, ascending node longitude and transfer flight time at the time of perigee is established, including: The lunar surface altitude, inclination, ascending node longitude and transfer flight time at the time of flight to the perigee are calculated by numerical integration of the high-precision force model, and the mapping data table of the three-dimensional data of the orbit change velocity increment and the lunar surface altitude, inclination, ascending node longitude and transfer flight time at the time of perigee is established as follows: in, Represents the three-dimensional data pair of track change speed increment, Indicates the lunar height at perigee, The inclination angle at the time of perigee, represents the longitude of the ascending node at the time of perigee, Indicates the transfer flight time at the perigee time, Indicates the orbit change speed increment parameter - lunar height, Indicates the track change speed increment parameter - inclination, Indicates the orbit change speed increment parameter - the longitude of the ascending node, Indicates the incremental parameter of the change-track speed - transfer flight time, Indicates the speed increment during the transfer process. Indicates the incremental azimuth of the track change speed, Indicates the height angle of the track change speed increment.

4. The method according to claim 1, characterized in that: The preset DRO orbit size parameters include 60,000 km, 65,000 km, 70,000 km, 75,000 km, and 80,000 km.

5. The method according to any one of claims 1 to 4, characterized in that: According to the target orbit lunar surface altitude and the target orbit inclination, the orbit change velocity increment three-dimensional data pairs whose perigee orbit altitude and orbit surface inclination meet the preset scattering range are selected, including: According to the target orbit lunar surface height and target orbit inclination, the orbital height and orbital surface inclination of the perigee point are selected to meet the preset scattering range of the orbit change velocity increment three-dimensional data pairs. in, and , represents the lunar altitude of the target orbit, represents the target orbit inclination, represents the orbital altitude of the perigee, represents the orbital inclination, It indicates the allowable deviation range of the perigee orbit altitude relative to the target orbit lunar surface altitude. Indicates the allowable deviation range of the orbital inclination relative to the target orbital inclination.

6. The method according to claim 1, characterized in that Obtain the data of the speed increment in each group of the minimum transfer process and calculate the orbit change time for the corresponding ascending node longitude and transfer flight time, including: Obtain the data of the minimum velocity increment in each group of transfer processes, and calculate the orbit change time for the corresponding ascending node longitude and transfer flight time: in, is the lunar revolution angular velocity, is the average orbital angular velocity of the DRO spacecraft in the L2 coordinate system, The data representing each set of the minimum velocity increment in the transfer process corresponds to the transfer flight time, represents the longitude of the ascending node, represents the longitude of the ascending node at the initial moment of the target orbit, represents the phase angle of the DRO aircraft at the initial moment, Indicates the initial time.

7. The method according to claim 1, characterized in that Correcting the phase of the selected track change time to obtain a corrected phase includes: The phase at the selected orbit change time is calculated as in, Indicates the phase angle of the DRO aircraft at the initial moment, is the average orbital angular velocity of the DRO spacecraft in the L2 coordinate system, represents the selected track change time, Indicates the initial time.

8. The method according to claim 6, characterized in that The method further comprises: The phase of the selected orbit change time includes the azimuth and height angle The corrected result is recorded as , ,have Among them, the correction factor and Obtained according to the interpolation table obtained from the previous simulation experiment.

9. The method according to claim 1, characterized in that: The trajectory change velocity increment projection is calculated based on the corrected phase, including: The projection of the track change velocity increment is calculated based on the corrected phase: in, represents the corrected azimuth, represents the corrected elevation angle, represents the phase of the selected track change time, Indicates the minimum track change speed increment, Represents a coordinate transformation matrix.

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