A method and system for evaluating stability margin of a long-range retro-orbital spacecraft

Abstract

The application provides a method and system for evaluating the stability margin of a long-distance retrograde orbit spacecraft. First, the current orbit state information of the target spacecraft is obtained. Then, based on the evaluation allowed time range, evaluation time interval, speed error size range and speed error size interval, an error parameter combination set is constructed. Next, based on the current orbit state information and the error parameter combination set, the stability evaluation parameters and the eclipse evaluation parameters corresponding to the error parameter combination are determined through a high-precision ephemeris model. Second, based on the stability evaluation parameters and the eclipse evaluation parameters, the target error parameter combination is determined. Finally, the stability margin is determined based on the target error parameter combination. This method combines the characteristics of long-distance retrograde orbits, engineering application requirements and the influence of eclipses, and selects parameters through the error parameter combination set to determine the stability margin. In this way, the efficiency and accuracy of determining the stability margin of the spacecraft can be effectively improved.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of spacecraft orbit control, in particular to a stable margin evaluation method and system for a distant retrograde orbit spacecraft. BACKGROUND

[0002] With the continuous development of space technology, the Earth-Moon space exploration has gradually become a research hotspot in the global space field. Among them, the distant retrograde orbit (DRO) as a unique three-body dynamics orbit in the Earth-Moon space has the characteristics of prograde around the Earth and retrograde around the Moon, and has the advantages of low energy entry, stable parking and global accessibility. Therefore, the DRO in the Earth-Moon space can provide a highly stable parking orbit, which is an extremely important test field and transfer station in the development of the Earth-Moon space and deep space exploration.

[0003] Due to the characteristics of the DRO, the spacecraft running on the DRO is sensitive to speed changes. When the tangential speed changes by a few meters, it will cause significant changes in the stability of the spacecraft orbit and the eclipse. However, in the long-term mission of the DRO spacecraft, a certain degree of speed error is inevitable. On the one hand, in the Earth-Moon space, the orbit state of the spacecraft running on the DRO for a long time contains a fixed orbit error. On the other hand, the spacecraft experiences disturbances or active momentum unloading, causing changes in the orbit state. Therefore, in view of the above situation, it is necessary to determine the allowed speed error margin (i.e. stable margin) of the spacecraft to determine whether an orbit maintenance maneuver is needed.

[0004] However, the existing stable margin evaluation method mainly focuses on low Earth orbits. However, the Earth-Moon space environment of the DRO is significantly different from the low Earth orbit. The spacecraft running on the DRO faces more severe eclipse threats, and long eclipse phenomena frequently occur, which may cause long-term energy supply interruption for spacecraft relying on solar power, seriously affecting its normal operation. Therefore, there is an urgent need for a stable margin evaluation method for a distant retrograde orbit spacecraft to effectively solve the problem of being unable to accurately determine the stable margin of the DRO spacecraft, thereby ensuring the stability of the spacecraft running on the DRO. SUMMARY

[0005] The technical problem to be solved by the present application is the problem of inefficient and inaccurate orbit maintenance maneuver for a distant retrograde orbit spacecraft.

[0006] To solve the above technical problems, the present application provides a stable margin evaluation method and system for a distant retrograde orbit spacecraft, which specifically adopts the following technical solutions:

[0007] In a first aspect, the present application provides a method for evaluating stability margin of a spacecraft in a distant retrograde orbit. The method can be applied to a target spacecraft in a distant retrograde orbit. The method comprises the following steps. First, current orbit state information of the target spacecraft is obtained. The current orbit state information is used to represent the position and velocity of the target spacecraft at the current time. Then, an error parameter combination set is constructed based on an evaluation allowed time range, an evaluation time interval, a velocity error size range, and a velocity error size interval. The error parameter combination set includes a plurality of error parameter combinations. Each error parameter combination includes an error occurrence time and a velocity error size. The error occurrence time is determined according to the evaluation allowed time range and the evaluation time interval. The velocity error size is determined according to the velocity error size range and the velocity error size interval. The evaluation allowed time range and the velocity error size range are both associated with the current orbit state information. Next, based on the current orbit state information and the error parameter combination set, a stability evaluation parameter and a solar eclipse evaluation parameter corresponding to each error parameter combination are determined by a high-precision ephemeris model. The stability evaluation parameter is used to represent the stability of the target spacecraft after a velocity error is generated based on the error parameter combination. The solar eclipse evaluation parameter is used to represent the solar eclipse condition of the target spacecraft after the velocity error is generated based on the error parameter combination. Then, a target error parameter combination is determined based on the stability evaluation parameter and the solar eclipse evaluation parameter. The target error parameter combination is the error parameter combination that satisfies a stability threshold for the stability evaluation parameter and satisfies a solar eclipse evaluation threshold for the solar eclipse evaluation parameter. Finally, a stability margin is determined based on the target error parameter combination.

[0008] The method first obtains the current orbit state information of the target spacecraft. Then, the error parameter combination set is constructed based on the evaluation allowed time range, the evaluation time interval, the velocity error size range, and the velocity error size interval. Next, the stability evaluation parameter and the solar eclipse evaluation parameter corresponding to each error parameter combination are determined based on the current orbit state information and the error parameter combination set by the high-precision ephemeris model. Then, the target error parameter combination is determined based on the stability evaluation parameter and the solar eclipse evaluation parameter. Finally, the stability margin is determined based on the target error parameter combination. The method combines the characteristics of the distant retrograde orbit, engineering application requirements, and solar eclipse effects, and constructs the error parameter combination set. The method can comprehensively cover the error parameter combinations while ensuring the screening efficiency, so as to screen more accurate target error parameter combinations, and then determine the stability margin based on the target error parameter combination. In this way, the efficiency and accuracy of determining the stability margin of the spacecraft in the distant retrograde orbit can be effectively improved, and the stability of the spacecraft operation is improved.

[0009] In conjunction with the first aspect, in one alternative implementation, determining the stability margin based on the combination of multiple target error parameters, when multiple combinations of target error parameters are known, includes: first, determining the maximum and minimum speed error values ​​among the multiple combinations of target error parameters; then, determining the range from the minimum speed error value to the maximum speed error value as the stability margin.

[0010] In conjunction with the first aspect, in one alternative implementation, the stability margin, given a given target error parameter combination, is 0 to the magnitude of the velocity error in the target error parameter combination.

[0011] In conjunction with the first aspect, in one alternative implementation, the aforementioned stability evaluation parameter is: the flight time during which the target spacecraft maintains bounded flight with the Moon within a first preset time period. Based on the current orbital state information and the set of error parameter combinations, a high-fidelity ephemeris model is used to determine the stability evaluation parameter and eclipse evaluation parameter corresponding to each error parameter combination. This includes: first, based on the current orbital state information, using a high-fidelity ephemeris model to determine the first predicted orbital state information of the target spacecraft within the first preset time period after a velocity error arises based on the error parameter combination. Then, based on the first predicted orbital state information and the Moon's position information, the flight time is determined.

[0012] In conjunction with the first aspect, in one alternative implementation, the aforementioned eclipse evaluation parameter is: the longest eclipse duration for the target spacecraft within a second preset time period. Based on the current orbital state information and the set of error parameter combinations, a high-fidelity ephemeris model is used to determine the stability evaluation parameter and eclipse evaluation parameter corresponding to each error parameter combination. This includes: first, based on the current orbital state information, using a high-fidelity ephemeris model to determine the second predicted orbital state information of the target spacecraft within a second preset time period after a velocity error arises based on the error parameter combinations. Then, based on the second predicted orbital state information and the sun's position information, the longest eclipse duration is determined.

[0013] In conjunction with the first aspect, in one alternative implementation, the aforementioned stability threshold includes: a flight time greater than or equal to a bounded flight time threshold; the bounded flight time threshold is 10 years.

[0014] In conjunction with the first aspect, in one alternative implementation, the aforementioned eclipse evaluation thresholds include: the longest eclipse duration being less than the eclipse duration threshold; and the eclipse duration threshold being 1 hour.

[0015] In conjunction with the first aspect, in one alternative implementation, the aforementioned set of error parameter combinations is represented by an error parameter grid; wherein the horizontal axis of the error parameter grid represents the time when the error occurs, the vertical axis of the error parameter grid represents the magnitude of the velocity error, and each grid point of the error parameter grid represents an error parameter combination.

[0016] In conjunction with the first aspect, in one alternative implementation, the above-mentioned speed error range is [-3,3] m / s, and the speed error interval is 0.01 m / s.

[0017] Secondly, this invention provides a stability margin assessment system for a long-distance retrograde orbit spacecraft, applicable to target spacecraft in long-distance retrograde orbits. The system includes: an acquisition module, an assembly construction module, an evaluation module, a screening module, and a stability margin determination module. The acquisition module acquires the current orbital state information of the target spacecraft, which characterizes the spacecraft's position and velocity at the current moment. The assembly construction module constructs an error parameter combination set based on an allowable evaluation time range, an evaluation time interval, a velocity error magnitude range, and a velocity error magnitude interval. This error parameter combination set includes multiple error parameter combinations, each including an error occurrence time and a velocity error magnitude. The error occurrence time is determined based on the allowable evaluation time range and the evaluation time interval, while the velocity error magnitude is determined based on the velocity error magnitude range and the velocity error magnitude interval. Both the allowable evaluation time range and the velocity error magnitude range are associated with the current orbital state information. The evaluation module can be used to determine the stability evaluation parameters and eclipse evaluation parameters corresponding to each error parameter combination based on the current orbital state information and a set of error parameter combinations, using a high-fidelity ephemeris model. The stability evaluation parameters characterize the stability of the target spacecraft after velocity errors are introduced based on the error parameter combinations, while the eclipse evaluation parameters characterize the eclipse event of the target spacecraft after velocity errors are introduced based on the error parameter combinations. The screening module can be used to determine the target error parameter combinations based on the stability evaluation parameters and eclipse evaluation parameters; the target error parameter combinations are those where both the stability evaluation parameters and eclipse evaluation parameters satisfy a stability threshold. The stability margin determination module can be used to determine the stability margin based on the target error parameter combinations.

[0018] Thirdly, the present invention provides an electronic device, comprising: a memory and one or more processors; the memory being coupled to the processors; wherein the memory stores computer program code, the computer program code including computer instructions, which, when executed by the processor, cause the electronic device to perform the method provided by the first aspect and any of its alternative implementations.

[0019] Fourthly, the present invention provides a computer-readable storage medium including computer instructions that, when executed on an electronic device, cause the electronic device to perform the method provided by the first aspect and any alternative implementation thereof.

[0020] Understandably, the beneficial effects of the stability margin assessment system for long-distance retrograde orbit spacecraft provided in the second aspect, the electronic equipment in the third aspect, and the computer-readable storage medium in the fourth aspect can be referenced to the beneficial effects of the first aspect and any of its possible design embodiments, which will not be elaborated here. Attached Figure Description

[0021] Figure 1 A flowchart illustrating the stability margin assessment method for a long-distance retrograde orbit spacecraft provided in this application embodiment;

[0022] Figure 2 A schematic diagram illustrating the stability margin results provided in the embodiments of this application;

[0023] Figure 3 A schematic diagram of the stability margin assessment system for a long-distance retrograde orbit spacecraft provided in this application embodiment. Detailed Implementation

[0024] The embodiments will now be described in detail, examples of which are illustrated in the accompanying drawings. When the following description relates to the drawings, unless otherwise indicated, the same numbers in different drawings represent the same or similar elements. The embodiments described below do not represent all embodiments consistent with this application. They are merely examples of systems and methods consistent with some aspects of this application as detailed in the claims.

[0025] With the continuous development of aerospace technology, lunar space exploration has gradually become a research hotspot in the global aerospace field. Among them, the long-distance retrograde orbit (DRO), as a unique three-body dynamic orbit in the lunar space, has the characteristics of orbiting the Earth progradely and orbiting the Moon retrogradely, and has significant advantages such as low-energy orbit insertion, stable parking, and accessibility across the entire lunar orbit. Therefore, the lunar space DRO can provide highly stable parking orbits and is an extremely important test field and transit station in the development of the lunar space and deep space exploration.

[0026] Due to the inherent characteristics of the DRO (Depth-of-Reference Spacecraft), spacecraft operating on the DRO are highly sensitive to velocity changes. Even meter-level changes in tangential velocity can significantly alter orbital stability and eclipse conditions. However, a certain degree of velocity error is unavoidable during long-term DRO missions. Firstly, in the Earth-Moon space, the orbital state of a spacecraft orbiting on the DRO for an extended period includes orbit determination errors. Secondly, disturbances or active momentum unloading can alter the orbital state. Therefore, considering these factors, it is necessary to determine the permissible velocity error margin (i.e., stability margin) for the spacecraft to assess whether orbital maintenance maneuvers are required.

[0027] However, existing stability margin assessment methods primarily focus on low Earth orbit (LEO). The Earth-Moon space environment in which the DRO operates differs significantly from that of LEO. Spacecraft operating on the DRO face a more severe threat of solar eclipses, with frequent prolonged eclipses. This could lead to prolonged power outages for solar-powered spacecraft, severely impacting their normal operation. Therefore, a stability margin determination method applicable to long-distance retrograde orbit spacecraft is urgently needed to effectively address the problem of inaccurately determining the stability margin of DRO spacecraft, thereby ensuring the stability of spacecraft operating on the DRO.

[0028] To address the aforementioned issues, this application provides a stability margin assessment method and system for spacecraft in long-distance retrograde orbits. This method can be applied to target spacecraft in long-distance retrograde orbits. Specifically, addressing the engineering requirements of DRO assessment, this method constructs a set of error parameter combinations based on the spacecraft's current orbital information and using discretized error generation times and velocity error magnitudes. Further, simulation calculations are performed on multiple error parameter combinations within this set. Under a high-fidelity ephemeris model, the spacecraft's orbital state after applying velocity errors is evaluated, including stability and eclipse conditions, for statistical analysis of the orbital state. Finally, the stability margin is determined based on the target error parameter combinations that meet threshold requirements. This method combines the characteristics of long-distance retrograde orbits, engineering application requirements, and the impact of eclipses by constructing a set of error parameter combinations. This ensures efficient screening while comprehensively covering all error parameter combinations, enabling the search for more accurate combinations and improving the efficiency and accuracy of determining the stability margin of long-distance retrograde orbit spacecraft, thereby enhancing the stability of spacecraft operation.

[0029] The solutions provided in the embodiments of this application will be described below with reference to the accompanying drawings.

[0030] Specifically, Figure 1 A flowchart illustrating the stability margin assessment method for a long-distance retrograde orbit spacecraft provided in this application embodiment is shown below. Figure 1As shown, the stability margin assessment method for long-distance retrograde orbit spacecraft provided in this application includes the following steps S101-S105:

[0031] S101. Obtain the current orbital status information of the target spacecraft.

[0032] In this embodiment, the first step is to obtain the current orbital state information of the target spacecraft. This facilitates the accurate construction of the error parameter combination set and the accurate evaluation of the target spacecraft's stability and the impact of the solar eclipse after velocity errors are generated based on the error parameter combination. For example, the current orbital state information can be obtained through a ground observation station.

[0033] The current orbital state information can be used to characterize the target spacecraft's position and velocity at the current moment. For example, the current orbital state information may include: position parameters along the x-axis, y-axis, and z-axis, velocity parameters along the x-axis, y-axis, and z-axis, and the current time parameter.

[0034] S102. Based on the allowable evaluation time range, evaluation time interval, speed error magnitude range, and speed error magnitude interval, construct a set of error parameter combinations.

[0035] Then, based on the allowable evaluation time range, evaluation time interval, speed error magnitude range, and speed error magnitude interval, the error time and speed error magnitude can be discretized to construct a set of error parameter combinations for determining the target error parameter combination.

[0036] The error parameter combination set includes multiple error parameter combinations, each containing an error occurrence time and a velocity error magnitude. The error occurrence time is determined based on the allowable evaluation time range and the evaluation time interval, while the velocity error magnitude is determined based on the velocity error magnitude range and the velocity error magnitude interval. Both the allowable evaluation time range and the velocity error magnitude range are associated with the current orbital state information.

[0037] In some embodiments, the aforementioned set of error parameter combinations can be represented by an error parameter grid; wherein the horizontal axis of the error parameter grid represents the time when the error occurs, the vertical axis represents the magnitude of the velocity error, and each grid point of the error parameter grid represents an error parameter combination. Thus, this error parameter grid facilitates rapid and comprehensive searching (filtering) of error parameter combinations, thereby improving the efficiency of filtering target error parameters that meet the threshold.

[0038] In one implementation, each grid point can also correspond to the predicted orbital state of the target spacecraft during the S103 processing, that is, the orbital state of the target spacecraft after the velocity error is generated based on the error parameter combination corresponding to the grid point, so as to facilitate the rapid screening of the target error parameter combination.

[0039] In some embodiments, the allowable evaluation time range can be 60 days, and the evaluation interval can be 1 day, that is, the error occurrence time is discretized with a one-day interval. The validity period of the allowable error range is 60 days, that is, within 60 days, the tangential velocity error generated by the target spacecraft is acceptable within a given range (i.e., meets the requirements of engineering applications), and the stability of the target spacecraft and its susceptibility to solar eclipses will not change. The velocity error magnitude range can be [-3, 3] m / s, and the velocity error magnitude interval can be 0.01 m / s. Thus, a velocity error magnitude within the range of [-3, 3] m / s can ensure the stability of the target spacecraft during DRO operation. A velocity error magnitude interval of 0.01 m / s can ensure the accuracy of determining the maneuver magnitude.

[0040] The allowable evaluation time range can be used to characterize the effective period of the allowable velocity error range, and the velocity error magnitude range can be used to characterize the allowable velocity error of the target spacecraft within the allowable evaluation time range. The aforementioned allowable evaluation time range, evaluation time interval, velocity error magnitude range, and velocity error magnitude interval can be preset based on current orbital state information, ensuring DRO stability, and incorporating prior knowledge and practical application requirements.

[0041] S103. Based on the current orbital state information and the set of error parameter combinations, determine the stability evaluation parameters and eclipse evaluation parameters corresponding to each error parameter combination through a high-fidelity ephemeris model.

[0042] Furthermore, based on the current orbital state information determined in S101 and the set of error parameter combinations constructed in S102, a high-fidelity ephemeris model is used for simulation to simulate and determine the target spacecraft's operation after velocity errors are generated based on the error parameter combinations. This involves determining the stability evaluation parameters and eclipse evaluation parameters corresponding to each error parameter combination. This facilitates accurate selection from the set of error parameter combinations to determine the target error parameter combination.

[0043] Among them, stability evaluation parameters can be used to characterize the stability of the target spacecraft after velocity errors are generated based on the combination of error parameters, and eclipse evaluation parameters can be used to characterize the eclipse situation of the target spacecraft after velocity errors are generated based on the combination of error parameters.

[0044] In some embodiments, the stability evaluation parameter can be the flight time during which the target spacecraft maintains bounded flight with the Moon within a first preset time period. Bounded flight refers to the flight distance between the target spacecraft and the Moon within a first preset flight range. For example, the first preset flight range can be 40,000 km to 140,000 km.

[0045] Based on the current orbital state information and the set of error parameter combinations, S103 determines the stability evaluation parameters and eclipse evaluation parameters corresponding to each error parameter combination using a high-fidelity ephemeris model. Specifically, these include:

[0046] First, based on the current orbital state information, the first predicted orbital state information of the target spacecraft within the first preset time period is determined by a high-fidelity ephemeris model after the velocity error is generated based on the combination of error parameters.

[0047] Specifically, based on the current orbital state information, the orbital state can be integrated using a high-fidelity ephemeris model to obtain the first predicted orbital state information of the target spacecraft within the first preset time period.

[0048] The first preset time period can be combined with a bounded flight time threshold and preset based on prior knowledge and actual application requirements. For example, if the bounded flight time threshold is 10 years, the first preset time period can be 15 years. This ensures the stability of the target spacecraft over a longer period of time.

[0049] Then, based on the first predicted orbital state information and the lunar position information, the flight time is determined.

[0050] Specifically, the position information of the target spacecraft can be determined based on the first predicted orbital state information. Then, combined with the position information of the moon, the flight time during which the target spacecraft remains within the first preset flight range from the moon can be determined.

[0051] In some embodiments, the aforementioned eclipse evaluation parameter can be: the longest eclipse duration for the target spacecraft within a second preset time period. This application does not specifically limit the type of eclipse; for example, the type of eclipse can be: an Earth shadow eclipse, a lunar shadow eclipse, etc.

[0052] Then, based on the current orbital state information and the set of error parameter combinations, S103 above determines the stability evaluation parameters and eclipse evaluation parameters corresponding to each error parameter combination through a high-fidelity ephemeris model. Specifically, it may also include:

[0053] First, based on the current orbital state information, the second predicted orbital state information of the target spacecraft within a second preset time period is determined by a high-fidelity ephemeris model after the velocity error is generated based on the combination of error parameters.

[0054] Specifically, the orbital state can first be integrated using a high-fidelity ephemeris model based on the current orbital state information to obtain the predicted orbital state information of the target spacecraft in the second preset time period.

[0055] The second preset time period can be preset based on prior knowledge and actual application needs; for example, the second preset time period can be 180 days. In this way, the impact of the solar eclipse on the target spacecraft can be accurately evaluated within this second preset time period.

[0056] Then, based on the second predicted orbital state information and the sun's position information, the longest duration of the eclipse is determined.

[0057] Specifically, based on the second predicted orbital state information and the sun's position information, the positional relationship between the target spacecraft and the sun can be determined, and the equivalent shadow duration of the solar eclipse can be used as the eclipse duration. Furthermore, the longest eclipse duration can be determined.

[0058] The equivalent shadow duration can be expressed as follows:

[0059] ;

[0060] in, Indicates the duration of the equivalent shadow. This represents the occlusion rate of a celestial body when the target spacecraft is in the total shadow area. =1; Indicates the start time of the penumbra. Indicates the end time of the penumbra.

[0061] S104. Based on the stability evaluation parameters and the eclipse evaluation parameters, determine the target error parameter combination.

[0062] Among them, the target error parameter combination is the error parameter combination in which the stability evaluation parameter satisfies the stability threshold and the eclipse evaluation parameter satisfies the eclipse evaluation threshold.

[0063] In some embodiments, when the stability evaluation parameter is the flight time during which the target spacecraft maintains bounded flight with the moon within a first preset time period, the stability threshold may include: the flight time being greater than or equal to the bounded flight time threshold.

[0064] To ensure the stability of the target spacecraft, the bounded flight time threshold can be 10 years, meaning the target spacecraft needs to maintain bounded flight time with the moon for at least 10 years within the first preset time period. In some implementations, the bounded flight time threshold can also be preset according to the needs of the actual application.

[0065] In some embodiments, when the eclipse evaluation parameter is the longest eclipse duration of the target spacecraft within a second preset time period, the eclipse evaluation threshold may include: the longest eclipse duration being less than the eclipse duration threshold. The eclipse duration threshold may be 1 hour, meaning the longest eclipse duration of the target spacecraft within the second preset time period (e.g., within 180 days) needs to be less than 1 hour.

[0066] S105. Determine the stability margin based on the combination of target error parameters.

[0067] Finally, based on the magnitude of the velocity error in the target error parameter combination determined in S104, the stability margin can be determined.

[0068] In some embodiments, when a target error parameter combination is determined in S104, the stability margin is: 0 to the magnitude of the velocity error in the target error parameter combination.

[0069] In some embodiments, when multiple combinations of target error parameters are determined in S104, determining the stability margin based on the combinations of target error parameters includes:

[0070] First, determine the maximum and minimum speed error values ​​among multiple combinations of target error parameters.

[0071] Then, the range from the minimum speed error value to the maximum speed error value is determined as the stability margin.

[0072] For example, Figure 2 A schematic diagram illustrating the stability margin results provided in the embodiments of this application, as shown below. Figure 2 As shown, the set of error parameter combinations is represented by error parameter grid 201, with the eclipse evaluation parameter being the longest eclipse duration of the target spacecraft within the second preset time period. In this example, the horizontal axis of error parameter grid 201 represents the time when the error occurs, the vertical axis represents the magnitude of the velocity error, and each grid point in error parameter grid 201 represents an error parameter combination. The color of the region corresponding to a grid point in error parameter grid 201 represents the magnitude of the longest eclipse duration of the error parameter combination corresponding to that grid point. Finally, based on the target error parameter combination, the stability margin can be determined to be [-0.38, 0.14] m / s.

[0073] It should be noted that, Figure 2 The colors used to represent the duration of the longest solar eclipse are for illustrative purposes only and are not intended to be limiting.

[0074] The stability margin assessment method for long-distance retrograde orbit spacecraft provided in the above embodiments of this application first obtains the current orbital state information of the target spacecraft. Then, based on the allowable assessment time range, assessment time interval, velocity error magnitude range, and velocity error magnitude interval, a set of error parameter combinations is constructed. Next, based on the current orbital state information and the set of error parameter combinations, a high-fidelity ephemeris model is used to determine the stability evaluation parameters and eclipse evaluation parameters corresponding to each error parameter combination. Then, based on the stability evaluation parameters and eclipse evaluation parameters, the target error parameter combination is determined. Finally, the stability margin is determined based on the target error parameter combination. This method combines the characteristics of long-distance retrograde orbits, engineering application requirements, and the impact of eclipses by constructing a set of error parameter combinations. This method can comprehensively cover error parameter combinations while ensuring screening efficiency, thereby selecting more accurate target error parameter combinations and determining the stability margin based on these combinations. This effectively improves the efficiency and accuracy of determining the stability margin of long-distance retrograde orbit spacecraft, thus improving the stability of spacecraft operation.

[0075] This application also provides a stability margin assessment system for spacecraft in long-distance retrograde orbits, which can be applied to target spacecraft in long-distance retrograde orbits. Specifically, Figure 3 A schematic diagram of the stability margin assessment system for a long-distance retrograde orbit spacecraft provided in this application embodiment is shown below. Figure 3 As shown, the stability margin assessment system 300 for the long-distance retrograde orbit spacecraft includes: an acquisition module 301, an assembly construction module 302, an evaluation module 303, a screening module 304, and a stability margin determination module 305.

[0076] The acquisition module 301 can be used to acquire the current orbital state information of the target spacecraft, which is used to characterize the position and velocity of the target spacecraft at the current moment.

[0077] The set construction module 302 can be used to construct a set of error parameter combinations based on the allowable evaluation time range, evaluation time interval, velocity error magnitude range, and velocity error magnitude interval. The set of error parameter combinations includes multiple error parameter combinations, each of which includes an error occurrence time and a velocity error magnitude. The error occurrence time is determined based on the allowable evaluation time range and evaluation time interval, and the velocity error magnitude is determined based on the velocity error magnitude range and velocity error magnitude interval. Both the allowable evaluation time range and the velocity error magnitude range are associated with the current orbital state information.

[0078] The evaluation module 303 can be used to determine the stability evaluation parameters and eclipse evaluation parameters corresponding to each error parameter combination based on the current orbital state information and the set of error parameter combinations, using a high-fidelity ephemeris model. The stability evaluation parameters are used to characterize the stability of the target spacecraft after velocity errors are generated based on the error parameter combinations, and the eclipse evaluation parameters are used to characterize the eclipse situation of the target spacecraft after velocity errors are generated based on the error parameter combinations.

[0079] The screening module 304 can be used to determine the target error parameter combination based on the stability evaluation parameters and the eclipse evaluation parameters; the target error parameter combination is the error parameter combination in which the stability evaluation parameters meet the stability threshold and the eclipse evaluation parameters meet the eclipse evaluation threshold.

[0080] The stability margin determination module 305 can be used to determine the stability margin based on the combination of target error parameters.

[0081] The stability margin assessment system for long-distance retrograde orbit spacecraft provided in the above embodiments of this application firstly acquires the current orbital state information of the target spacecraft through an acquisition module. Then, a set construction module constructs a set of error parameter combinations based on the allowable assessment time range, assessment time interval, velocity error magnitude range, and velocity error magnitude interval. Next, an evaluation module determines the stability evaluation parameters and eclipse evaluation parameters corresponding to each error parameter combination based on the current orbital state information and the set of error parameter combinations using a high-fidelity ephemeris model. Secondly, a screening module determines the target error parameter combination based on the stability evaluation parameters and the eclipse evaluation parameters. Finally, a stability margin determination module determines the stability margin based on the target error parameter combination. This method combines the characteristics of long-distance retrograde orbits, engineering application requirements, and the impact of eclipses by constructing a set of error parameter combinations. This method can comprehensively cover error parameter combinations while ensuring screening efficiency, thereby selecting more accurate target error parameter combinations and determining the stability margin based on these combinations. This effectively improves the efficiency and accuracy of determining the stability margin of long-distance retrograde orbit spacecraft, thus enhancing the stability of spacecraft operation.

[0082] This invention also provides an electronic device, which may include a display screen, a memory, and one or more processors. The display screen, memory, and processors are coupled. The memory stores computer program code, which includes computer instructions. When the processor executes the computer instructions, the electronic device can perform various methods or steps as described in the embodiments of the stability margin assessment method for long-distance retrograde orbit spacecraft. Of course, this electronic device includes, but is not limited to, the aforementioned display screen, memory, and one or more processors.

[0083] This invention also provides a computer-readable storage medium for storing computer instructions for operating the stability margin assessment method for the aforementioned long-distance retrograde orbit spacecraft.

[0084] Through the above description of the embodiments, those skilled in the art can clearly understand that, for the sake of convenience and brevity, only the division of the above functional modules is used as an example. In actual applications, the above functions can be assigned to different functional modules as needed, that is, the internal structure of the device can be divided into different functional modules to complete all or part of the functions described above.

[0085] In the description of this invention, it should be understood that the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of indicated technical features. Thus, a feature defined as "first" or "second" may explicitly or implicitly include at least one of that feature. In the description of this invention, "a plurality of" means at least two, such as two, three, etc., unless otherwise explicitly specified.

[0086] In the description of this specification, the references to terms such as "one embodiment," "some embodiments," "example," "specific example," or "some examples," etc., refer to specific features, structures, materials, or characteristics described in connection with that embodiment or example, which are included in at least one embodiment or example of the present invention. In this specification, the illustrative expressions of the above terms do not necessarily refer to the same embodiment or example. Furthermore, the specific features, structures, materials, or characteristics described may be combined in any suitable manner in one or more embodiments or examples. Moreover, without contradiction, those skilled in the art can combine and integrate the different embodiments or examples described in this specification, as well as the features of different embodiments or examples.

[0087] Similar parts between the embodiments provided in this application can be referred to mutually. The specific implementation methods provided above are only a few examples under the overall concept of this application and do not constitute a limitation on the scope of protection of this application. For those skilled in the art, any other implementation methods extended from the solution of this application without creative effort shall fall within the scope of protection of this application.

Claims

1. A method for evaluating the stability margin of a long-distance retrograde orbit spacecraft, characterized in that, The method, applied to a target spacecraft in a distant retrograde orbit, includes: The current orbital state information of the target spacecraft is obtained, and the current orbital state information is used to characterize the position and velocity of the target spacecraft at the current moment; Based on the allowable evaluation time range, evaluation time interval, velocity error magnitude range, and velocity error magnitude interval, an error parameter combination set is constructed. This set includes multiple error parameter combinations, each comprising an error occurrence time and a velocity error magnitude. The error occurrence time is determined based on the allowable evaluation time range and the evaluation time interval, while the velocity error magnitude is determined based on the velocity error magnitude range and the velocity error magnitude interval. Both the allowable evaluation time range and the velocity error magnitude range are associated with the current orbital state information. Based on the current orbital state information and the set of error parameter combinations, a stability evaluation parameter and an eclipse evaluation parameter corresponding to each error parameter combination are determined using a high-fidelity ephemeris model. The stability evaluation parameter is used to characterize the stability of the target spacecraft after a velocity error is generated based on the error parameter combination, and the eclipse evaluation parameter is used to characterize the eclipse situation of the target spacecraft after a velocity error is generated based on the error parameter combination. Based on the stability evaluation parameters and the eclipse evaluation parameters, a target error parameter combination is determined; the target error parameter combination is an error parameter combination in which the stability evaluation parameters satisfy a stability threshold and the eclipse evaluation parameters satisfy an eclipse evaluation threshold. The stability margin is determined based on the combination of the target error parameters. When multiple combinations of the target error parameters are determined, determining the stability margin based on the combinations of the target error parameters includes: Determine the maximum and minimum speed error values ​​among the multiple combinations of target error parameters; The range from the minimum speed error value to the maximum speed error value is defined as the stability margin. Given a given target error parameter combination, the stability margin is: 0 to the magnitude of the velocity error in the target error parameter combination.

2. The method according to claim 1, characterized in that, The stability evaluation parameter is: the flight time during which the target spacecraft maintains bounded flight with the moon within a first preset time period; The step of determining the stability evaluation parameters and eclipse evaluation parameters corresponding to each error parameter combination based on the current orbital state information and the set of error parameter combinations using a high-fidelity ephemeris model includes: Based on the current orbital state information, the first predicted orbital state information of the target spacecraft within the first preset time period is determined by the high-fidelity ephemeris model after a velocity error is generated based on the error parameter combination. The flight time is determined based on the first predicted orbital state information and the lunar position information.

3. The method according to claim 1, characterized in that, The solar eclipse evaluation parameter is: the longest solar eclipse duration of the target spacecraft within a second preset time period; The step of determining the stability evaluation parameters and eclipse evaluation parameters corresponding to each error parameter combination based on the current orbital state information and the set of error parameter combinations using a high-fidelity ephemeris model includes: Based on the current orbital state information, the second predicted orbital state information of the target spacecraft within the second preset time period is determined by the high-fidelity ephemeris model after a velocity error is generated based on the error parameter combination. The longest duration of the solar eclipse is determined based on the second predicted orbital state information and the position information of the sun.

4. The method according to claim 2, characterized in that, The stability threshold includes: the flight time is greater than or equal to a bounded flight time threshold; the bounded flight time threshold is 10 years.

5. The method according to claim 3, characterized in that, The eclipse evaluation thresholds include: the longest eclipse duration is less than the eclipse duration threshold; and the eclipse duration threshold is 1 hour.

6. The method according to claim 1, characterized in that, The set of error parameter combinations is represented by an error parameter grid; wherein, the horizontal axis of the error parameter grid represents the time when the error occurs, the vertical axis of the error parameter grid represents the magnitude of the velocity error, and each grid point of the error parameter grid represents one of the error parameter combinations.

7. The method according to claim 1, characterized in that, The speed error range is [-3, 3] m / s, and the speed error interval is 0.01 m / s.

8. A stability margin assessment system for a long-distance retrograde orbit spacecraft, employing the stability margin assessment method for a long-distance retrograde orbit spacecraft as described in claim 1, characterized in that, For target spacecraft in long-distance retrograde orbits, the system includes: an acquisition module, an ensemble construction module, an evaluation module, a screening module, and a stability margin determination module; wherein, The acquisition module is used to acquire the current orbital state information of the target spacecraft, and the current orbital state information is used to characterize the position and velocity of the target spacecraft at the current moment; The set construction module is used to construct an error parameter combination set based on the allowable evaluation time range, evaluation time interval, velocity error magnitude range, and velocity error magnitude interval. The error parameter combination set includes multiple error parameter combinations, each including an error occurrence time and a velocity error magnitude. The error occurrence time is determined based on the allowable evaluation time range and the evaluation time interval, and the velocity error magnitude is determined based on the velocity error magnitude range and the velocity error magnitude interval. Both the allowable evaluation time range and the velocity error magnitude range are associated with the current orbital state information. The evaluation module is used to determine the stability evaluation parameter and eclipse evaluation parameter corresponding to each error parameter combination based on the current orbital state information and the set of error parameter combinations, using a high-fidelity ephemeris model; wherein, the stability evaluation parameter is used to characterize the stability of the target spacecraft after a velocity error is generated based on the error parameter combination, and the eclipse evaluation parameter is used to characterize the eclipse situation of the target spacecraft after a velocity error is generated based on the error parameter combination; The filtering module is used to determine a target error parameter combination based on the stability evaluation parameters and the eclipse evaluation parameters; the target error parameter combination is an error parameter combination in which the stability evaluation parameters satisfy a stability threshold and the eclipse evaluation parameters satisfy an eclipse evaluation threshold. The stability margin determination module is used to determine the stability margin based on the target error parameter combination.

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