Spacecraft orbit tracking guidance method for earth-moon space autonomous flight

Abstract

The invention provides a spacecraft orbit tracking guidance method for earth-moon space autonomous flight, which can determine the current moment as the initial moment, and obtain the state parameter corresponding to the target spacecraft at the initial moment. And determining a first nominal maneuvering moment, wherein the first nominal maneuvering moment is the nominal maneuvering moment closest to the initial moment in the nominal maneuvering moments corresponding to the target spacecraft. An end time is determined based on the initial time, the first nominal maneuver time, and the maximum optimization time interval. And determining a modified maneuvering strategy based on the initial moment, the state parameter of the target spacecraft at the initial moment, the tail end moment and the nominal orbit state parameter corresponding to the target spacecraft. And the corrected maneuvering strategy is executed based on the condition that the corrected maneuvering strategy meets the preset condition, so that the spacecraft enters the target orbit, and the resource investment cost of spacecraft orbit tracking guidance is reduced.

Description

Technical Field

[0001] The present invention relates to the technical field of spacecraft guidance and control technology, and in particular to a spacecraft orbit tracking and guidance method for autonomous flight in cis-lunar space. Background Art

[0002] Spacecraft orbit tracking guidance refers to making the spacecraft fly according to a pre-planned orbit or reference orbit. It continuously measures the actual flight status parameters of the spacecraft (for example, position, speed, attitude, etc.), compares them with the predetermined orbital parameters, and generates control instructions based on the deviation between the two to correct the flight path of the spacecraft, so that the spacecraft can get as close to the predetermined orbit as possible.

[0003] At present, spacecraft orbit tracking and guidance mainly utilizes ground-based measurement and control systems for telemetry and remote control. This process requires staff to be in the loop. This orbit tracking and guidance mode is non-autonomous, which increases the ground measurement and control burden and resource investment costs. Summary of the invention

[0004] The present invention provides a spacecraft orbit tracking and guidance method for autonomous flight in the cis-lunar space. It can solve the guidance and control algorithm of the spacecraft so as to make the spacecraft enter the target orbit without relying on ground measurement and control and man-in-the-loop, thereby reducing the resource investment cost for the spacecraft orbit tracking and guidance.

[0005] The technical solution of the present invention to solve the above technical problems is as follows:

[0006] In a first aspect, the present invention provides a spacecraft orbit tracking and guidance method for autonomous flight in the cis-lunar space, in which the current moment is determined as the initial moment, and the state parameters corresponding to the target spacecraft at the initial moment are obtained. The state parameters include position parameters and speed parameters. The first nominal maneuvering moment is determined. The first nominal maneuvering moment is the nominal maneuvering moment closest to the initial moment among the nominal maneuvering moments corresponding to the target spacecraft, and the first nominal maneuvering moment is later than the initial moment. The terminal moment is determined based on the initial moment, the first nominal maneuvering moment, and the preset maximum optimization time interval. The terminal moment is the predicted moment when the target spacecraft enters the nominal orbit. Based on the initial moment, the state parameters of the target spacecraft at the initial moment, the terminal moment, and the nominal orbit state parameters corresponding to the target spacecraft, a modified maneuvering strategy is determined. The nominal orbit state parameters include the preset position parameters and the preset speed parameters of the target spacecraft flying in the nominal orbit. The modified maneuvering strategy is used to apply multiple pulse maneuvers to the target spacecraft so that the target spacecraft enters the nominal orbit at the terminal moment. The modified maneuvering strategy includes the moment and pulse value of each pulse maneuver applied to the target spacecraft. The pulse value includes the size and direction of the corresponding pulse maneuver. Based on the correction maneuver strategy meeting the preset conditions, the correction maneuver strategy is executed.

[0007] Based on the above technical solution, the present invention can also be improved as follows.

[0008] Furthermore, the preset condition includes a first preset condition and a second preset condition. Obtain the total speed increment of the corrected maneuver and the target corrected maneuvering time corresponding to the corrected maneuvering strategy. The total speed increment of the corrected maneuvering corresponding to the corrected maneuvering strategy is determined based on each pulse value included in the corrected maneuvering strategy, and the target corrected maneuvering time corresponding to the corrected maneuvering strategy is determined based on the earliest time of applying the pulse maneuver included in the corrected maneuvering strategy. The corrected maneuvering strategy is executed based on the target corrected maneuvering time corresponding to the corrected maneuvering strategy satisfying the first preset condition, and the total speed increment of the corrected maneuvering corresponding to the corrected maneuvering strategy satisfying the second preset condition.

[0009] Further, based on the initial moment, the state parameters of the target spacecraft at the initial moment, and the terminal moment, the nominal orbit state parameters corresponding to the target spacecraft, multiple sets of alternative correction maneuvering strategies are determined. For any set of alternative correction maneuvering strategies among the multiple sets of alternative correction maneuvering strategies, the correction maneuvering total speed increment corresponding to any set of alternative correction maneuvering strategies is calculated. The alternative correction maneuvering strategy with the smallest correction maneuvering total speed increment among the multiple sets of alternative correction maneuvering strategies is determined as the correction maneuvering strategy.

[0010] Further, based on the first preset relationship and any group of candidate corrected maneuvering strategies, the corrected maneuvering total speed increment corresponding to any group of candidate corrected maneuvering strategies is obtained. The first preset relationship includes:

[0011]

[0012] in, It represents the minimum value of the total speed increment of the corrected maneuver corresponding to any set of alternative corrected maneuver strategies. represents the initial time, represents the state parameters of the target spacecraft at the initial moment, represents the terminal moment, x represents the nominal orbital state parameter corresponding to the target spacecraft, It represents the pulse value corresponding to the i-th pulse maneuver included in any set of alternative correction maneuver strategies. is the derivative of x, where x represents the state parameter corresponding to the target spacecraft. is the derivative of r, where r represents the position parameter of the target spacecraft. is the derivative of v, v represents the velocity parameter of the target spacecraft. a represents the acceleration parameter of the target spacecraft, and the acceleration parameter of the target spacecraft is determined based on the state parameter of the target spacecraft. represents the time when the target spacecraft is subjected to the pulse maneuver for the i-th time based on any set of alternative corrected maneuver strategies. Represents a set of alternative modified maneuver strategies based on The state parameters corresponding to the target spacecraft before applying pulse maneuvers to the target spacecraft at all times, Represents a set of alternative modified maneuver strategies based on The state parameters corresponding to the target spacecraft after the pulse maneuver is applied to the target spacecraft at all times. represents the time when the target spacecraft is subjected to the pulse maneuver for the i+1th time based on any set of alternative modified maneuver strategies. Δt nav Indicates the preset minimum time interval for correction maneuvers. represents the state parameters corresponding to the target spacecraft at the initial moment, represents the state parameters corresponding to the target spacecraft at the end moment, Represents the nominal orbital state parameters corresponding to the target spacecraft at the terminal moment.

[0013] Further, the first preset condition includes that the difference between the target correction maneuvering time and the initial time is less than or equal to the preset rolling optimization time interval. The second preset condition includes that the correction maneuvering total speed increment is greater than the preset minimum execution speed increment.

[0014] Further, based on the second preset relationship, the initial moment, the first nominal maneuvering moment, and the preset maximum optimization time interval, the terminal moment is determined. The second preset relationship includes:

[0015]

[0016] in, Indicates the end moment, represents the first nominal maneuvering time, represents the initial time, Δt opt Indicates the preset maximum optimization time interval.

[0017] Further, after a preset minimum time interval of the correction maneuver, the correction maneuver strategy is updated.

[0018] Further, based on the fact that the modified maneuvering strategy does not satisfy the preset condition, the modified maneuvering strategy is updated after a preset rolling optimization time interval.

[0019] Furthermore, initial correction maneuvering parameters corresponding to the target spacecraft are set, wherein the initial correction maneuvering parameters include a preset maximum optimization time interval, a preset rolling optimization time interval, a preset minimum execution speed increment, and a preset correction maneuvering minimum time interval.

[0020] The beneficial effects of the present invention are: reducing measurement and control pressure, avoiding the influence of communication delays and transmission errors, and being able to solve the guidance and control algorithm for the spacecraft so that the spacecraft can enter the target orbit without relying on ground measurement and control and humans in the loop, thereby reducing the resource investment cost for spacecraft orbit tracking and guidance.

[0021] In a second aspect, the present invention provides a spacecraft orbit tracking and guidance system for autonomous flight in the Earth-Moon space. The system can execute the spacecraft orbit tracking and guidance method for autonomous flight in the Earth-Moon space as described in any one of the first aspects above.

[0022] In a third aspect, the present invention provides an electronic device comprising: a memory, one or more processors; the memory and the processor are coupled; wherein the memory stores computer program code, the computer program code comprises computer instructions, and when the computer instructions are executed by the processor, the electronic device executes the spacecraft orbit tracking and guidance method for autonomous flight in the Earth-Moon space as described in any one of the first aspects above.

[0023] In a fourth aspect, the present invention provides a computer-readable storage medium comprising computer instructions, which, when executed on an electronic device, enables the electronic device to execute the spacecraft orbit tracking and guidance method for autonomous flight in the Earth-Moon space as described in any one of the first aspects above.

[0024] In a fifth aspect, the present invention provides a computer program product, which, when executed on a computer, enables the computer to execute the method for spacecraft orbit tracking and guidance for autonomous flight in the Earth-Moon space as described in any one of the first aspects above.

[0025] It can be understood that the beneficial effects that can be achieved by the system of the second aspect, the electronic device described in the third aspect, the computer-readable storage medium described in the fourth aspect, and the computer program product described in the fifth aspect provided above can refer to the beneficial effects in the first aspect and any possible design method thereof, and will not be repeated here. BRIEF DESCRIPTION OF THE DRAWINGS

[0026] Figure 1 A schematic flow chart of a spacecraft orbit tracking and guidance method for autonomous flight in cis-lunar space provided by the present invention. DETAILED DESCRIPTION

[0027] The technical solutions in the embodiments of the present application will be described below in conjunction with the drawings in the embodiments of the present application. Among them, in the description of the present application, unless otherwise specified, " / " indicates that the objects associated before and after are in an "or" relationship, for example, A / B can represent A or B; "and / or" in the present application is only a description of the association relationship of the associated objects, indicating that there can be three relationships, for example, A and / or B, which can represent: A exists alone, A and B exist at the same time, and B exists alone, where A and B can be singular or plural. And, in the description of the present application, unless otherwise specified, "multiple" refers to two or more than two. "At least one of the following" or similar expressions refers to any combination of these items, including any combination of single items or plural items. For example, at least one of a, b, or c can represent: a, b, c, ab, ac, bc, or abc, where a, b, c can be single or multiple. In addition, in order to facilitate the clear description of the technical solutions of the embodiments of the present application, in the embodiments of the present application, the words "first", "second" and the like are used to distinguish between the same items or similar items with substantially the same functions and effects. Those skilled in the art will understand that the words "first", "second" and the like do not limit the quantity and execution order, and the words "first", "second" and the like do not necessarily limit the differences. At the same time, in the embodiments of the present application, the words "exemplary" or "for example" are used to indicate examples, illustrations or descriptions. Any embodiment or design described as "exemplary" or "for example" in the embodiments of the present application should not be interpreted as being more preferred or more advantageous than other embodiments or design.

[0028] Due to the existence of error factors such as navigation and propulsion system errors, the spacecraft cannot fly along the pre-designed nominal orbit and needs to perform mid-course correction guidance. The existing spacecraft mid-course correction guidance control method mainly uses ground-based measurement and control systems for telemetry and remote control. This process requires humans in the loop. This guidance and control mode is non-autonomous and has limitations: 1. It increases the burden of ground measurement and control and resource investment costs; 2. There are communication delays and transmission errors, which are not conducive to improving the guidance and control performance. With the increase in the number of spacecraft in the Earth-Moon space in the future, space flight will gradually become "flight-based", and this limitation will become more and more prominent.

[0029] In view of the above problems, the present invention provides a spacecraft orbit tracking and guidance method for autonomous flight in the Earth-Moon space, which can reduce the measurement and control pressure, avoid the influence of communication delays and transmission errors, and can solve the guidance and control algorithm of the spacecraft to enable the spacecraft to enter the target orbit without relying on ground measurement and control and humans in the loop, thereby reducing the resource investment cost for spacecraft orbit tracking and guidance.

[0030] It should be noted that the method provided by the present invention requires the spacecraft to have autonomous orbit determination capability (the spacecraft autonomously determines the position, speed and other status information of the spacecraft in space), and the nominal orbit is predetermined and loaded into the onboard computer. The basic principle of the method provided by the present invention is: using the model predictive control method, using the onboard computer to roll-solve the multi-pulse orbit tracking optimal control problem, and autonomously deploying mid-course correction maneuvers based on the solution results.

[0031] See also Figure 1 The spacecraft orbit tracking and guidance method for autonomous flight in the cis-lunar space provided by the present invention comprises the following steps S100-S105:

[0032] S100: Setting initial correction maneuvering parameters corresponding to the target spacecraft.

[0033] Among them, the initial correction maneuver parameters include a preset maximum optimization time interval, a preset rolling optimization time interval, a preset minimum execution speed increment, and a preset correction maneuver minimum time interval.

[0034] In some embodiments, the initial correction maneuver parameters may also include a correction maneuver amount.

[0035] S101: Determine the current time as the initial time, and obtain the state parameters corresponding to the target spacecraft at the initial time.

[0036] Among them, the state parameters may include position parameters and speed parameters.

[0037] S102: Determine a first nominal maneuvering time.

[0038] The first nominal maneuvering time is the nominal maneuvering time closest to the initial time among the nominal maneuvering times corresponding to the target spacecraft, and the first nominal maneuvering time is later than the initial time.

[0039] S103: Determine the end time based on the initial time, the first nominal maneuvering time, and the preset maximum optimization time interval.

[0040] Among them, the terminal moment is the predicted moment when the target spacecraft enters the nominal orbit.

[0041] In some embodiments, the terminal time may be determined based on the second preset relationship, the initial time, the first nominal maneuvering time, and the preset maximum optimization time interval.

[0042] In some embodiments, the second preset relationship may include:

[0043]

[0044] in, Indicates the end moment, represents the first nominal maneuvering time, represents the initial time, Δt opt Indicates the preset maximum optimization time interval.

[0045] S104: Determine a modified maneuvering strategy based on the initial moment, the state parameters of the target spacecraft at the initial moment, the terminal moment, and the nominal orbital state parameters corresponding to the target spacecraft.

[0046] The nominal orbit state parameters include the preset position parameters and preset speed parameters of the target spacecraft flying in the nominal orbit. The modified maneuver strategy is used to apply multiple pulse maneuvers to the target spacecraft so that the target spacecraft enters the nominal orbit at the end time. The modified maneuver strategy includes the time and pulse value of each pulse maneuver applied to the target spacecraft. The pulse value includes the size and direction of the corresponding pulse maneuver.

[0047] In some embodiments, multiple sets of alternative correction maneuvering strategies may be determined based on the initial moment, the state parameters of the target spacecraft at the initial moment, the terminal moment, and the nominal orbital state parameters corresponding to the target spacecraft. For any set of alternative correction maneuvering strategies among the multiple sets of alternative correction maneuvering strategies, the correction maneuvering total speed increment corresponding to any set of alternative correction maneuvering strategies may be calculated. The alternative correction maneuvering strategy with the smallest correction maneuvering total speed increment among the multiple sets of alternative correction maneuvering strategies is determined as the correction maneuvering strategy.

[0048] In some embodiments, based on the first preset relationship and any group of alternative corrective maneuvering strategies, the corrective maneuvering total speed increment corresponding to any group of alternative corrective maneuvering strategies can be obtained. The first preset relationship may include:

[0049]

[0050]

[0051] in, It represents the minimum value of the total speed increment of the corrected maneuver corresponding to any set of alternative corrected maneuver strategies. represents the initial time, represents the state parameters of the target spacecraft at the initial moment, Indicates the end moment, represents the nominal orbital state parameters corresponding to the target spacecraft, It represents the pulse value corresponding to the i-th pulse maneuver included in any set of alternative correction maneuver strategies. is the derivative of x, where x represents the state parameter corresponding to the target spacecraft. is the derivative of r, where r represents the position parameter of the target spacecraft. is the derivative of v, v represents the velocity parameter of the target spacecraft. a represents the acceleration parameter of the target spacecraft, and the acceleration parameter of the target spacecraft is determined based on the state parameter of the target spacecraft. represents the time when the target spacecraft is subjected to the pulse maneuver for the i-th time based on any set of alternative corrected maneuver strategies. Represents a set of alternative modified maneuver strategies based on The state parameters corresponding to the target spacecraft before applying pulse maneuvers to the target spacecraft at all times, Represents a set of alternative modified maneuver strategies based on The state parameters corresponding to the target spacecraft after the pulse maneuver is applied to the target spacecraft at all times. represents the time when the target spacecraft is subjected to the pulse maneuver for the i+1th time based on any set of alternative modified maneuver strategies. Δt nav Indicates the preset minimum time interval for correction maneuvers. represents the state parameters corresponding to the target spacecraft at the initial moment, represents the state parameters corresponding to the target spacecraft at the end moment, Represents the nominal orbital state parameters corresponding to the target spacecraft at the terminal moment.

[0052] In some embodiments, the first relationship can be solved based on a sequential convex optimization algorithm to obtain the optimal pulse maneuver And the corresponding time and the minimum total speed increment required

[0053] S105: Execute the modified maneuvering strategy based on the fact that the modified maneuvering strategy meets the preset conditions.

[0054] In some embodiments, in order to avoid the situation where the mid-course correction maneuvers are too small and performed frequently, a maneuver execution judgment condition can be introduced, that is, a preset condition can be introduced. The preset condition includes a first preset condition and a second preset condition. When the determined correction maneuver strategy meets the above preset conditions, the orbit correction maneuver can be performed based on the correction maneuver strategy, otherwise the orbit correction maneuver is not performed.

[0055] Specifically, the corrected maneuvering total speed increment and the target corrected maneuvering time corresponding to the corrected maneuvering strategy can be obtained. The corrected maneuvering strategy is executed based on the target corrected maneuvering time corresponding to the corrected maneuvering strategy satisfying the first preset condition and the corrected maneuvering total speed increment corresponding to the corrected maneuvering strategy satisfying the second preset condition.

[0056] The total speed increment of the corrected maneuver corresponding to the corrected maneuver strategy is determined based on the pulse values ​​included in the corrected maneuver strategy, and the target corrected maneuver time corresponding to the corrected maneuver strategy is determined based on the earliest time of applying the pulse maneuver included in the corrected maneuver strategy.

[0057] In some embodiments, the first preset condition includes that the difference between the target corrected maneuvering time and the initial time is less than or equal to a preset rolling optimization time interval. The first preset condition can ensure that the optimal mid-course corrected maneuvering time to be executed is close enough to the current time.

[0058] In some embodiments, the second preset condition includes that the total speed increment of the correction maneuver is greater than the preset minimum execution speed increment. The second preset condition can avoid excessively small and frequent execution of mid-course correction maneuvers and keep the actual flight trajectory in the vicinity of the nominal trajectory. nearby.

[0059] For example, it can be satisfied and At the time Perform the first impulse maneuver based on the corresponding modified maneuver strategy Otherwise, the corresponding corrective maneuver strategy is not executed.

[0060] In some embodiments, the modified maneuvering strategy may be updated based on the modified maneuvering strategy satisfying the preset condition and after a preset modified maneuvering minimum time interval has passed. That is, based on the modified maneuvering strategy satisfying the preset condition and after a preset modified maneuvering minimum time interval has passed, steps S101-S104 may be performed again to redetermine the modified maneuvering strategy, and based on the redetermined modified maneuvering strategy satisfying the preset condition, the redetermined modified maneuvering strategy is executed.

[0061] In some embodiments, based on the fact that the modified maneuvering strategy does not meet the preset conditions, the modified maneuvering strategy is updated after a preset rolling optimization time interval. That is, based on the fact that the modified maneuvering strategy does not meet the preset conditions, steps S101-S104 may be performed again after a preset rolling optimization time interval to redetermine the modified maneuvering strategy, and based on whether the redetermined modified maneuvering strategy meets the preset conditions, it is determined whether to execute the redetermined modified maneuvering strategy.

[0062] It can be seen that the method provided by the present invention can overcome the influence of various error factors in the actual flight process through rolling optimization, deploy orbit correction maneuvers in time according to the actual flight status, and use the total speed increment required for orbit correction as an indicator to evaluate the size of the spacecraft's deviation from the nominal orbit in real time, and combine the maneuver execution conditions to realize autonomous deployment of mid-course correction maneuvers to achieve autonomous flight into orbit.

[0063] In some schemes, multiple embodiments of the present application can be combined, and the combined scheme can be implemented. Optionally, some operations in the process of each method embodiment are optionally combined, and / or the order of some operations is optionally changed. In addition, the execution order between the steps of each process is only exemplary and does not constitute a restriction on the execution order between the steps. There can also be other execution orders between the steps. It is not intended to indicate that the execution order is the only order in which these operations can be performed. A person of ordinary skill in the art will think of a variety of ways to reorder the operations described herein. In addition, it should be noted that the process details involved in a certain embodiment of this article are also applicable to other embodiments in a similar manner, or different embodiments can be used in combination.

[0064] In addition, some steps in the method embodiment may be equivalently replaced by other possible steps. Alternatively, some steps in the method embodiment may be optional and may be deleted in certain usage scenarios. Alternatively, other possible steps may be added to the method embodiment. Furthermore, the method embodiments may be implemented separately or in combination.

[0065] Through the description of the above implementation methods, technical personnel in the relevant field can clearly understand that for the convenience and simplicity of description, only the division of the above-mentioned functional modules is used as an example. In actual applications, the above-mentioned functions can be assigned to different functional modules as needed, that is, the internal structure of the system can be divided into different functional modules to complete all or part of the functions described above.

[0066] In the several embodiments provided in the present application, it should be understood that the disclosed systems and methods can be implemented in other ways. For example, the system embodiments described above are only schematic. For example, the division of the modules or units is only a logical function division. There may be other division methods in actual implementation, such as multiple units or components can be combined or integrated into another system, or some features can be ignored or not executed. Another point is that the mutual coupling or direct coupling or communication connection shown or discussed can be through some interfaces, indirect coupling or communication connection of systems or units, which can be electrical, mechanical or other forms.

[0067] In addition, each functional unit in each embodiment of the present application may be integrated into one processing unit, or each unit may exist physically separately, or two or more units may be integrated into one unit. The above-mentioned integrated unit may be implemented in the form of hardware or in the form of software functional units.

[0068] If the integrated unit is implemented in the form of a software functional unit and sold or used as an independent product, it can be stored in a readable storage medium. Based on this understanding, the technical solution of the embodiment of the present application is essentially or the part that contributes or all or part of the technical solution can be embodied in the form of a software product, which is stored in a storage medium, including several instructions to enable a device (which can be a single-chip microcomputer, chip, etc.) or a processor (processor) to perform all or part of the steps of the method described in each embodiment of the present application. The aforementioned storage medium includes: U disk, mobile hard disk, read-only memory (ROM), random access memory (RAM), disk or optical disk and other media that can store program code.

[0069] The above contents are only specific implementation methods of the present application, but the protection scope of the present application is not limited thereto. Any changes or substitutions within the technical scope disclosed in the present application shall be included in the protection scope of the present application. Therefore, the protection scope of the present application shall be subject to the protection scope of the claims.

Claims

1. A spacecraft orbit tracking and guidance method for autonomous flight in cis-lunar space, characterized in that: include: Determine the current time as the initial time, and obtain the state parameters corresponding to the target spacecraft at the initial time; The state parameters include position parameters and speed parameters; Determining a first nominal maneuvering time; the first nominal maneuvering time is a nominal maneuvering time closest to the initial time among the nominal maneuvering times corresponding to the target spacecraft, and the first nominal maneuvering time is later than the initial time; Determining a terminal moment based on the initial moment, the first nominal maneuvering moment, and a preset maximum optimization time interval; the terminal moment is a predicted moment when the target spacecraft enters the nominal orbit; Determining a modified maneuvering strategy based on the initial moment, the state parameters of the target spacecraft at the initial moment, the terminal moment, and the nominal orbital state parameters corresponding to the target spacecraft; The nominal orbit state parameters include preset position parameters and preset speed parameters of the target spacecraft flying in the nominal orbit; the corrected maneuver strategy is used to apply multiple pulse maneuvers to the target spacecraft so that the target spacecraft enters the nominal orbit at the terminal time; The modified maneuver strategy includes the time and pulse value of each pulse maneuver applied to the target spacecraft; The pulse value includes the magnitude and direction of the corresponding pulse maneuver; Based on the fact that the modified maneuver strategy meets a preset condition, the modified maneuver strategy is executed.

2. The method according to claim 1, characterized in that: The preset condition includes a first preset condition and a second preset condition; and executing the modified maneuver strategy based on the modified maneuver strategy satisfying the preset condition includes: Obtaining a corrected maneuvering total speed increment and a target corrected maneuvering time corresponding to the corrected maneuvering strategy; the corrected maneuvering total speed increment corresponding to the corrected maneuvering strategy is determined based on each pulse value included in the corrected maneuvering strategy; the target corrected maneuvering time corresponding to the corrected maneuvering strategy is determined based on the earliest time of applying a pulse maneuver included in the corrected maneuvering strategy; Based on the target corrected maneuvering time corresponding to the corrected maneuvering strategy satisfying the first preset condition, and the corrected maneuvering total speed increment corresponding to the corrected maneuvering strategy satisfying the second preset condition, the corrected maneuvering strategy is executed.

3. The method according to claim 2, characterized in that The determining of a corrected maneuvering strategy based on the initial moment, the state parameters of the target spacecraft at the initial moment, and the terminal moment, the nominal orbital state parameters corresponding to the target spacecraft, comprises: Determining a plurality of sets of alternative correction maneuvering strategies based on the initial moment, the state parameters of the target spacecraft at the initial moment, the terminal moment, and the nominal orbital state parameters corresponding to the target spacecraft; For any one of the plurality of alternative corrected maneuvering strategies, calculating a corrected maneuvering total speed increment corresponding to the one of the plurality of alternative corrected maneuvering strategies; An alternative correction maneuver strategy having a minimum correction maneuver total speed increment among the plurality of groups of alternative correction maneuver strategies is determined as the correction maneuver strategy.

4. The method according to claim 3, characterized in that The calculating the corrected maneuvering total speed increment corresponding to any set of the alternative corrected maneuvering strategies comprises: Based on a first preset relationship and any one of the set of candidate corrected maneuvering strategies, a corrected maneuvering total speed increment corresponding to any one of the set of candidate corrected maneuvering strategies is obtained; the first preset relationship includes: in, represents the minimum value of the total speed increment of the corrected maneuver corresponding to any one of the set of alternative corrected maneuver strategies; represents the initial time; represents the state parameters of the target spacecraft corresponding to the initial moment; represents said terminal moment; represents the nominal orbital state parameters corresponding to the target spacecraft; represents the pulse value corresponding to the i-th pulse maneuver included in any set of alternative corrected maneuver strategies; is the derivative of x, where x represents the state parameter corresponding to the target spacecraft; is the derivative of r, where r represents the position parameter of the target spacecraft; is the derivative of v, v represents the velocity parameter of the target spacecraft; a represents the acceleration parameter of the target spacecraft, and the acceleration parameter of the target spacecraft is determined based on the state parameter of the target spacecraft; represents the time when the pulse maneuver is applied to the target spacecraft for the i-th time based on any one of the set of alternative modified maneuver strategies; Indicates that based on any set of alternative modified maneuver strategies The state parameters corresponding to the target spacecraft before the pulse maneuver is applied to the target spacecraft at any moment; Indicates that based on any set of alternative modified maneuver strategies state parameters corresponding to the target spacecraft after applying a pulse maneuver to the target spacecraft at any moment; represents the time when the target spacecraft is subjected to the pulse maneuver for the i+1th time based on any one of the set of alternative modified maneuver strategies; Δt nav Indicates the preset minimum time interval for correction maneuvers; Indicates the state parameters corresponding to the target spacecraft at the initial moment; represents the state parameter corresponding to the target spacecraft at the end moment, Represents the nominal orbital state parameters corresponding to the target spacecraft at the end moment.

5. The method according to claim 4, characterized in that The first preset condition includes that the difference between the target corrected maneuvering moment and the initial moment is less than or equal to a preset rolling optimization time interval; the second preset condition includes that the total speed increment of the corrected maneuvering is greater than a preset minimum execution speed increment.

6. The method according to claim 5, characterized in that The determining the terminal time based on the initial time, the first nominal maneuvering time, and a preset maximum optimization time interval includes: The end time is determined based on a second preset relationship, the initial time, the first nominal maneuvering time, and the preset maximum optimization time interval; the second preset relationship includes: in, represents said terminal moment; represents the first nominal maneuvering moment; represents the initial time; Δt opt Indicates the preset maximum optimization time interval.

7. The method according to claim 6, characterized in that After executing the modified maneuvering strategy based on the modified maneuvering strategy satisfying a preset condition, the method further includes: After the preset minimum time interval of the correction maneuver has passed, the correction maneuver strategy is updated.

8. The method according to claim 7, characterized in that After determining the modified maneuvering strategy based on the initial moment, the state parameters of the target spacecraft at the initial moment, the terminal moment, and the nominal orbital state parameters corresponding to the target spacecraft, the method further includes: Based on the fact that the modified maneuvering strategy does not satisfy a preset condition, the modified maneuvering strategy is updated after the preset rolling optimization time interval.

9. The method according to claim 8, characterized in that Before determining the current time as the initial time and obtaining the state parameters corresponding to the target spacecraft at the initial time, the method further includes: The initial correction maneuvering parameters corresponding to the target spacecraft are set; the initial correction maneuvering parameters include the preset maximum optimization time interval, the preset rolling optimization time interval, the preset minimum execution speed increment, and the preset correction maneuvering minimum time interval.

10. An electronic device, characterized in that: include: A memory and one or more processors; the memory is coupled to the processor; wherein the memory stores computer program code, the computer program code includes computer instructions, and when the computer instructions are executed by the processor, the electronic device executes a spacecraft orbit tracking and guidance method for autonomous flight in the cis-lunar space as described in any one of claims 1 to 9.

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