A spacecraft orbit tracking guidance method for earth-moon space autonomous flight
By employing an autonomous spacecraft orbit tracking and guidance method, utilizing onboard computers for model predictive control, and autonomously deploying mid-course correction maneuvers, the problem of reliance on ground-based telemetry and control in existing technologies has been solved, achieving autonomous and low-cost guidance and control.
Patent Information
- Application Number
- CN202411947189.5
- Authority / Receiving Office
- CN · China
- Patent Type
- Patents(China)
- Current Assignee / Owner
- Filing Date
- 2024-12-27
- Publication Date
- 2025-10-17
- Estimated Expiration
- 2044-12-27
AI Technical Summary
Current spacecraft orbit tracking and guidance relies on ground-based telemetry and control and human intervention in the orbit, which increases resource investment costs and communication latency, and the lack of autonomy limits guidance and control performance.
The spacecraft employs an autonomous flight orbit tracking and guidance method, utilizes onboard computers for model predictive control, optimizes multi-pulse orbit tracking through rolling optimization, and autonomously deploys mid-course correction maneuvers to reduce reliance on ground-based telemetry and control.
It achieves autonomous guidance and control without ground-based telemetry and control or human intervention, reducing resource investment costs, avoiding the impact of communication delays and transmission errors, and improving the autonomy and efficiency of guidance and control.
Smart Images
Figure CN119759063B_ABST
Abstract
Description
TECHNICAL FIELD
[0001] The present application relates to the technical field of spacecraft guidance control technology, and particularly relates to a spacecraft orbit tracking guidance method for autonomous flight in the earth-moon space. BACKGROUND
[0002] The spacecraft orbit tracking guidance refers to making the spacecraft fly according to a pre-planned orbit or a reference orbit, comparing the actual flight state parameters (for example, position, speed, attitude, etc.) of the spacecraft with the predetermined orbit parameters by continuously measuring the actual flight state parameters of the spacecraft, and generating control instructions based on the deviation between the two to correct the flight path of the spacecraft, so that the spacecraft can be as close as possible to the predetermined orbit.
[0003] At present, the spacecraft orbit tracking guidance mainly uses the ground-based measurement and control system for telemetry and remote control, which needs the staff to be in the loop. This orbit tracking guidance mode is non-autonomous, which increases the ground measurement and control burden and resource investment cost. SUMMARY
[0004] The present application provides a spacecraft orbit tracking guidance method for autonomous flight in the earth-moon space, which can realize the calculation of the spacecraft guidance control algorithm without relying on ground measurement and control and human-in-the-loop to make the spacecraft enter the target orbit, thereby reducing the resource investment cost of the spacecraft orbit tracking guidance.
[0005] The technical solution of the present application to solve the above technical problems is as follows:
[0006] In the first aspect, the present application provides a spacecraft orbit tracking guidance method for autonomous flight in the earth-moon space, in which the current time is determined as an initial time, and the state parameters corresponding to the target spacecraft at the initial time are obtained. The state parameters include position parameters and speed parameters. A first nominal maneuver time is determined. The first nominal maneuver time is the closest nominal maneuver time to the initial time among the nominal maneuver times corresponding to the target spacecraft, and the first nominal maneuver time is later than the initial time. Based on the initial time, the first nominal maneuver time and a preset maximum optimization time interval, an end time is determined. The end time is the time when the target spacecraft enters a nominal orbit predicted. Based on the initial time, the state parameters of the target spacecraft at the initial time, the end time and the nominal orbit state parameters corresponding to the target spacecraft, a correction maneuver strategy is determined. The nominal orbit state parameters include preset position parameters and preset speed parameters of the target spacecraft flying in the nominal orbit. The correction maneuver strategy is used to apply multiple pulse maneuvers to the target spacecraft to make the target spacecraft enter the nominal orbit at the end time. The correction 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. The correction maneuver strategy is executed based on the fact that the correction maneuver strategy satisfies a preset condition.
[0007] Based on the technical solution, the application can be further improved as follows.
[0008] Further, the preset conditions include a first preset condition and a second preset condition. The correction maneuver strategy corresponding correction maneuver total speed increment and the target correction maneuver time are obtained. The correction maneuver strategy corresponding correction maneuver total speed increment is determined based on each pulse value included in the correction maneuver strategy, and the target correction maneuver time corresponding to the correction maneuver strategy is determined based on the time of the earliest pulse maneuver included in the correction maneuver strategy. Based on the target correction maneuver time corresponding to the correction maneuver strategy satisfying the first preset condition, and the correction maneuver total speed increment corresponding to the correction maneuver strategy satisfying the second preset condition, the correction maneuver strategy is executed.
[0009] Further, based on the initial time, the state parameter of the target spacecraft at the initial time, the terminal time, and the nominal orbit state parameter corresponding to the target spacecraft, a plurality of groups of candidate correction maneuver strategies are determined. For any one group of candidate correction maneuver strategies in the plurality of groups of candidate correction maneuver strategies, the correction maneuver total speed increment corresponding to any one group of candidate correction maneuver strategies is calculated. The candidate correction maneuver strategy with the smallest correction maneuver total speed increment in the plurality of groups of candidate correction maneuver strategies is determined as the correction maneuver strategy.
[0010] Further, based on the first preset relationship and any one group of candidate correction maneuver strategies, the correction maneuver total speed increment corresponding to any one group of candidate correction maneuver strategies is obtained. The first preset relationship includes:
[0011]
[0012] wherein, represents the minimum value of the correction maneuver total speed increment corresponding to any one group of candidate correction maneuver strategies. represents the initial time, represents the state parameter of the target spacecraft at the initial time, represents the terminal time, and x represents the nominal orbit state parameter corresponding to the target spacecraft, represents the pulse value corresponding to the i-th pulse maneuver included in any one group of candidate correction maneuver strategies. is the derivative of x, and x represents the state parameter of the target spacecraft. is the derivative of r, and r represents the position parameter of the target spacecraft. is the derivative of v, and v represents the velocity parameter of the target spacecraft. a represents the acceleration parameter of the target spacecraft, which is determined based on the state parameter of the target spacecraft. represents the time of applying the pulse maneuver to the target spacecraft for the i-th time based on any one group of candidate correction maneuver strategies. represents the state parameters of the target spacecraft at the initial time, represents the state parameters of the target spacecraft at the initial time, represents the state parameters of the target spacecraft at the initial time, represents the state parameters of the target spacecraft at the initial time. represents the time of the i+1th pulse maneuver to the target spacecraft based on any one of the sets of candidate correction maneuver strategies. nav represents the preset minimum time interval of the correction maneuver. represents the state parameters of the target spacecraft at the initial time, represents the state parameters of the target spacecraft at the initial time, represents the state parameters of the target spacecraft at the initial time.
[0013] Further, the first preset condition includes that the difference between the target correction maneuver time and the initial time is less than or equal to the preset rolling optimization time interval. The second preset condition includes that the total velocity increment of the correction maneuver is greater than the preset minimum execution velocity increment.
[0014] Further, based on the second preset relationship, the initial time, the first nominal maneuver time, and the preset maximum optimization time interval, the terminal time is determined. The second preset relationship includes:
[0015]
[0016] wherein, represents the terminal time, represents the first nominal maneuver time, represents the initial time, Δt opt represents the preset maximum optimization time interval.
[0017] Further, after the preset minimum time interval of the correction maneuver, the correction maneuver strategy is updated.
[0018] Further, based on the correction maneuver strategy not satisfying the preset condition, after the preset rolling optimization time interval, the correction maneuver strategy is updated.
[0019] Further, the initial correction maneuver parameters corresponding to the target spacecraft are set. The initial correction maneuver parameters include the preset maximum optimization time interval, the preset rolling optimization time interval, the preset minimum execution velocity increment, and the preset minimum time interval of the correction maneuver.
[0020] The beneficial effects of the present application are: reducing the measurement and control pressure, avoiding the influence of communication delay and transmission error, without relying on ground measurement and control and human-in-the-loop, the spacecraft can be guided to enter the target orbit by solving the control algorithm, and the resource investment cost of the spacecraft orbit tracking guidance is reduced.
[0021] In the second aspect, the present application provides a spacecraft orbit tracking guidance system for autonomous flight in the Earth-Moon space, which can execute the spacecraft orbit tracking guidance method for autonomous flight in the Earth-Moon space according to any one of the first aspect.
[0022] In the third aspect, the present application 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, and the computer program code comprises computer instructions, when the computer instructions are executed by the processor, the electronic device executes the spacecraft orbit tracking guidance method for autonomous flight in the Earth-Moon space according to any one of the first aspect.
[0023] In the fourth aspect, the present application provides a computer readable storage medium, comprising computer instructions, when the computer instructions are executed on the electronic device, the electronic device executes the spacecraft orbit tracking guidance method for autonomous flight in the Earth-Moon space according to any one of the first aspect.
[0024] In the fifth aspect, the present application provides a computer program product, when the computer program product is executed on the computer, the computer executes the spacecraft orbit tracking guidance method for autonomous flight in the Earth-Moon space according to any one of the first aspect.
[0025] It can be understood that the beneficial effects of the system of the second aspect, the electronic device of the third aspect, the computer readable storage medium of the fourth aspect, and the computer program product of the fifth aspect provided above can refer to the beneficial effects of the first aspect and any possible design of the first aspect, which will not be repeated here. BRIEF DESCRIPTION OF DRAWINGS
[0026] Figure 1 The flowchart of the spacecraft orbit tracking guidance method for autonomous flight in the Earth-Moon space provided by the present application. DETAILED DESCRIPTION
[0027] The technical solutions in the embodiments of the present application will be described below with reference to the drawings in the embodiments of the present application. In the description of the present application, unless otherwise specified, " / " represents an "or" relationship between the objects before and after the " / ", for example, A / B can represent A or B; in the present application, "and / or" is only a description of the relationship between the associated objects, which means that there can be three relationships, for example, A and / or B, which means that there are three cases of A alone, A and B together, and B alone, where A and B can be singular or plural. In the description of the present application, unless otherwise specified, "multiple" means two or more than two. "At least one of the following" or the like means any combination of the items, including any combination of single or multiple items. For example, at least one of a, b, or c can mean a, b, c, a-b, a-c, b-c, or a-b-c, where a, b, and c can be single or multiple. In addition, in order to clearly describe the technical solutions of the embodiments of the present application, in the embodiments of the present application, "first", "second", and the like are used to distinguish the same items or similar items with basically the same function and effect. Those skilled in the art can understand that "first", "second", and the like do not limit the quantity and execution order, and "first", "second", and the like do not necessarily mean different. At the same time, in the embodiments of the present application, "exemplary" or "for example" means to serve as an example, illustration or description. Any embodiment or design scheme described as "exemplary" or "for example" in the embodiments of the present application should not be interpreted as more preferred or more advantageous than other embodiments or design schemes.
[0028] Due to the existence of errors such as navigation, propulsion system errors, spacecraft cannot fly along the pre-designed nominal orbit, and need to be guided and corrected. The existing spacecraft midcourse correction guidance control method mainly uses ground-based measurement and control system for telemetry and remote control, which needs human-in-the-loop. This guidance control mode is non-autonomous and has limitations: 1. Increases the ground measurement and control burden and resource investment cost; 2. Communication delay and transmission error exist, which is not conducive to improving the guidance control performance. With the increase in the number of future lunar space spacecraft, space flight gradually realizes "flight", and this limitation will become more prominent.
[0029] In view of the above problems, the present application provides a spacecraft orbit tracking guidance method for autonomous flight in lunar space, which can reduce the measurement and control pressure, avoid the influence of communication delay and transmission error, and realize the calculation of the spacecraft guidance control algorithm without relying on ground measurement and control and human-in-the-loop, so as to make the spacecraft enter the target orbit and reduce the resource investment cost of spacecraft orbit tracking guidance.
[0030] It should be noted that the method provided by the present application requires that the spacecraft has autonomous orbit determination capability (determining the position, velocity and other state information of the spacecraft in space autonomously on the spacecraft), and the nominal orbit is determined in advance and loaded into the on-board computer. The basic principle of the method provided by the present application is that the model predictive control method is used to solve the multi-pulse orbit tracking optimal control problem by using the on-board computer to roll, and the intermediate correction maneuver is autonomously deployed according to the solving result.
[0031] Referring to Figure 1 The spacecraft orbit tracking guidance method for autonomous flight in the Earth-Moon space provided by the present application comprises the following steps S100-S105:
[0032] S100: setting initial correction maneuver parameters corresponding to the target spacecraft.
[0033] The initial correction maneuver parameters comprise a preset maximum optimization time interval, a preset rolling optimization time interval, a preset minimum execution speed increment and a preset minimum correction maneuver time interval.
[0034] In some embodiments, the initial correction maneuver parameters can further comprise a correction maneuver number.
[0035] S101: determining the current time as an initial time, and acquiring state parameters corresponding to the target spacecraft at the initial time.
[0036] The state parameters can comprise position parameters and velocity parameters.
[0037] S102: determining a first nominal maneuver time.
[0038] The first nominal maneuver time is the closest nominal maneuver time to the initial time among the nominal maneuver times corresponding to the target spacecraft, and the first nominal maneuver time is later than the initial time.
[0039] S103: determining an end time based on the initial time, the first nominal maneuver time and the preset maximum optimization time interval.
[0040] The end time is the time when the target spacecraft enters the nominal orbit.
[0041] In some embodiments, the end time can be determined based on a second preset relationship, the initial time, the first nominal maneuver time and the preset maximum optimization time interval.
[0042] In some embodiments, the second preset relationship can comprise:
[0043]
[0044] Wherein, The end time is represented by t end, denotes a first nominal maneuver time, denotes an initial time, Δt opt denotes a preset maximum optimization time interval.
[0045] S104: Determine a correction maneuver strategy based on the initial time, the state parameter of the target spacecraft at the initial time, the terminal time, and the nominal orbit state parameter corresponding to the target spacecraft.
[0046] The nominal orbit state parameter includes preset position parameters and preset velocity parameters of the target spacecraft flying in the nominal orbit. The correction maneuver strategy is used to apply multiple impulse maneuvers to the target spacecraft so that the target spacecraft enters the nominal orbit at the terminal time. The correction maneuver strategy includes a time and an impulse value of each impulse maneuver applied to the target spacecraft. The impulse value includes a size and a direction of the corresponding impulse maneuver.
[0047] In some embodiments, a plurality of groups of alternative correction maneuver strategies can be determined based on the initial time, the state parameter of the target spacecraft at the initial time, the terminal time, and the nominal orbit state parameter corresponding to the target spacecraft. For any one of the plurality of groups of alternative correction maneuver strategies, a correction maneuver total velocity increment corresponding to the any one of the plurality of groups of alternative correction maneuver strategies can be calculated. The alternative correction maneuver strategy with the minimum correction maneuver total velocity increment among the plurality of groups of alternative correction maneuver strategies is determined as the correction maneuver strategy.
[0048] In some embodiments, the correction maneuver total velocity increment corresponding to any one of the plurality of groups of alternative correction maneuver strategies can be obtained based on a first preset relationship and the any one of the plurality of groups of alternative correction maneuver strategies. The first preset relationship can include:
[0049]
[0050]
[0051] wherein, denotes a minimum value of the correction maneuver total velocity increment corresponding to any one of the plurality of groups of alternative correction maneuver strategies. denotes an initial time, Δt denotes a state parameter corresponding to the target spacecraft at the initial time, denotes a terminal time, denotes a nominal orbit state parameter corresponding to the target spacecraft, denotes an impulse value corresponding to an i-th impulse maneuver included in any one of the plurality of groups of alternative correction maneuver strategies. is a derivative of x, and x denotes a state parameter corresponding to the target spacecraft. is a derivative of r, and r denotes a position parameter of the target spacecraft. is a derivative of v, v represents a velocity parameter of the target spacecraft. a represents an 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 a time at which a pulse maneuver is applied to the target spacecraft based on any one of the sets of alternative correction maneuver strategies. represents a time at which a pulse maneuver is applied to the target spacecraft based on any one of the sets of alternative correction maneuver strategies. represents a state parameter of the target spacecraft corresponding to the time at which a pulse maneuver is applied to the target spacecraft based on any one of the sets of alternative correction maneuver strategies. represents a state parameter of the target spacecraft corresponding to the time at which a pulse maneuver is applied to the target spacecraft based on any one of the sets of alternative correction maneuver strategies. represents a state parameter of the target spacecraft corresponding to the time at which a pulse maneuver is applied to the target spacecraft based on any one of the sets of alternative correction maneuver strategies. represents a time at which a pulse maneuver is applied to the target spacecraft based on any one of the sets of alternative correction maneuver strategies. nav represents a preset minimum time interval of the correction maneuver. represents a state parameter of the target spacecraft corresponding to the initial time, represents a state parameter of the target spacecraft corresponding to the terminal time, represents a nominal orbit state parameter of the target spacecraft corresponding to the terminal time.
[0052] In some embodiments, the first relationship can be solved based on a sequence convex optimization algorithm to obtain the optimal pulse maneuver and the corresponding time and the minimum total velocity increment required
[0053] S105: Based on the correction maneuver strategy satisfying the preset condition, the correction maneuver strategy is executed.
[0054] In some embodiments, in order to avoid the situation that the correction maneuver is too small and is frequently executed in the middle, a maneuver execution judgment condition can be introduced, that is, a preset condition is introduced. The preset condition includes a first preset condition and a second preset condition. When the determined correction maneuver strategy satisfies the above preset condition, the orbit correction maneuver can be executed based on the correction maneuver strategy, otherwise the orbit correction maneuver is not executed.
[0055] Specifically, the correction maneuver total velocity increment corresponding to the correction maneuver strategy and the target correction maneuver time can be obtained. Based on the target correction maneuver time corresponding to the correction maneuver strategy satisfying the first preset condition, and the correction maneuver total velocity increment corresponding to the correction maneuver strategy satisfying the second preset condition, the correction maneuver strategy is executed.
[0056] wherein the correction maneuver total velocity increment corresponding to the correction maneuver strategy is determined based on each pulse value included in the correction maneuver strategy, and the target correction maneuver time corresponding to the correction maneuver strategy is determined based on a time at which the earliest pulse maneuver is applied according to the correction maneuver strategy.
[0057] In some embodiments, the first preset condition comprises that a difference between the target correction maneuver moment and the initial moment is less than or equal to a preset rolling optimization time interval. The first preset condition can ensure that the optimal midcourse correction maneuver moment to be executed is close enough to the current moment.
[0058] In some embodiments, the second preset condition comprises that the total velocity increment of the correction maneuver is greater than a preset minimum execution velocity increment. The second preset condition can avoid too small and frequent execution of the midcourse correction maneuver, and keep the actual flight trajectory in the vicinity of the nominal orbit.
[0059] For example, when the following conditions are met and a first pulse maneuver is executed at the moment based on the corresponding correction maneuver strategy Otherwise, the corresponding correction maneuver strategy is not executed.
[0060] In some embodiments, the correction maneuver strategy can be updated after a preset correction maneuver minimum time interval based on the correction maneuver strategy meeting the preset condition. That is, steps S101-S104 can be executed again to re-determine the correction maneuver strategy after the preset correction maneuver minimum time interval based on the correction maneuver strategy meeting the preset condition, and the re-determined correction maneuver strategy can be executed based on the re-determined correction maneuver strategy meeting the preset condition.
[0061] In some embodiments, the correction maneuver strategy can be updated after a preset rolling optimization time interval based on the correction maneuver strategy not meeting the preset condition. That is, steps S101-S104 can be executed again to re-determine the correction maneuver strategy after the preset rolling optimization time interval based on the correction maneuver strategy not meeting the preset condition, and it can be determined whether to execute the re-determined correction maneuver strategy based on whether the re-determined correction maneuver strategy meets the preset condition.
[0062] It can be seen that the method provided by the present application can overcome the influence of various error factors in the actual flight process by rolling optimization, deploy the orbit correction maneuver in time according to the actual flight state, and evaluate the size of the spacecraft deviating from the nominal orbit in real time by taking the total velocity increment required for orbit correction as an index, and combine the maneuver execution conditions to realize autonomous deployment of midcourse correction maneuver to realize autonomous flight into orbit.
[0063] In some aspects, multiple embodiments of the present application can be combined, and the combined aspects can be implemented. Optionally, some operations in the flow of various method embodiments are combined, and / or the order of some operations is changed. Also, the order of execution or performance of the operations in the flows of various embodiments is immaterial since the flow is merely exemplary. Moreover, other operations in addition to those listed can be performed, and / or the order of performance can be varied from that described. One of ordinary skill in the art will recognize that the operations of the disclosed aspects can be implemented in electronic hardware, computer software, or any combination thereof. Unless explicitly stated otherwise, embodiments of the present application are not constrained to any single aspect of implementation.
[0064] In addition, some of the steps in the method embodiments can be optional, and some of the steps in the method embodiments can be replaced by other possible steps. Or, some of the steps in the method embodiments can be optional, and can be deleted in some use scenarios. Or, other possible steps can be added in the method embodiments. Also, the method embodiments can be implemented individually, or in combination.
[0065] From the above description of the embodiments, those skilled in the art can clearly understand that, for the convenience and brevity of description, only the division of the above functional modules is taken as an example for description, and in actual application, the above functions can be completed by different functional modules according to needs, that is, the internal structure of the system is 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 system and method can be implemented in other ways. For example, the division of the above-described system embodiments is merely a logical function division, and other division manners can be adopted in actual implementation, for example, a plurality of units or components can be combined or integrated into another system, or some features can be ignored or not executed. In addition, the coupling or direct coupling or communication connection between the displayed or discussed mutual units can be indirect coupling or communication connection through some interfaces, systems or units, and can be electrical, mechanical or other forms.
[0067] In addition, each functional unit in each embodiment of the present application can be integrated in one processing unit, or each unit can exist physically, or two or more units can be integrated in one unit. The integrated unit can be realized in the form of hardware, or in the form of software functional unit.
[0068] The integrated unit, if implemented in the form of a software function unit and sold or used as an independent product, can be stored in a readable storage medium. Based on such understanding, the technical solutions of the embodiments of the present application can be embodied in the form of a software product in essence or in the part that contributes to the present application, or the whole or part of the technical solutions can be embodied in the form of a software product stored in a storage medium, including a plurality of instructions for causing an apparatus (which can be a single-chip microcomputer, a chip, etc.) or a processor to execute all or part of the steps of the method described in the embodiments of the present application. The foregoing storage medium includes a U disk, a mobile hard disk, a read-only memory (ROM), a random access memory (RAM), a magnetic disk or an optical disk, and various media that can store program codes.
[0069] The above is only a specific implementation of the present application, but the protection scope of the present application is not limited thereto, and any change or replacement within the technical scope disclosed in the present application should be covered in the protection scope of the present application. Therefore, the protection scope of the present application should 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 moment as the initial moment, and obtain the state parameters corresponding to the target spacecraft at the initial moment; The state parameters include position parameters and speed parameters; determining a first nominal maneuvering time; wherein 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 an end time based on the initial time, the first nominal maneuvering time, and a preset maximum optimization time interval; the end time being a predicted time when the target spacecraft enters the nominal orbit; Determining a corrected 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 velocity parameters of the target spacecraft flying in the nominal orbit; the correction 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 timing 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 maneuvering strategy meets a preset condition, the modified maneuvering strategy is executed.
2. The method according to claim 1, characterized in that The preset conditions include a first preset condition and a second preset condition; and executing the modified maneuvering strategy based on the modified maneuvering strategy satisfying the preset conditions includes: Obtaining a corrected maneuver total speed increment and a target corrected maneuver time corresponding to the corrected maneuver strategy; the corrected maneuver total speed increment corresponding to the corrected maneuver strategy is determined based on each pulse value included in the corrected maneuver strategy; and the target corrected maneuver time corresponding to the corrected maneuver strategy is determined based on the earliest pulse maneuver time included in the corrected maneuver 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 corrected maneuvering total speed increment corresponding to the corrected maneuvering strategy satisfying the second preset condition.
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, the terminal moment, and the nominal orbital state parameters corresponding to the target spacecraft includes: Determining a plurality of sets of alternative corrective 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 corrective maneuvering strategies, calculating a corrective maneuvering total speed increment corresponding to the one of the plurality of alternative corrective 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 Calculating the corrected maneuvering total speed increment corresponding to any one set of alternative corrected maneuvering strategies includes: Based on a first preset relationship and any one of the set of alternative corrected maneuvering strategies, obtaining a corrected maneuvering total speed increment corresponding to any one of the set of alternative corrected maneuvering strategies; the first preset relationship includes: in, 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 corresponding to the initial moment; represents the 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 target spacecraft is subjected to the pulse maneuver 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 state parameters corresponding to the target spacecraft before applying a pulse maneuver 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 a pulse maneuver is applied to the target spacecraft at any given moment; represents the time when the target spacecraft is subjected to the pulse maneuver for the i+1th time based on any of the 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 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 of 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 the 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 maneuver strategy based on the initial time, the state parameters of the target spacecraft at the initial time, and the nominal orbital state parameters corresponding to the target spacecraft at the terminal time, the method further includes: Based on the fact that the modified maneuvering strategy does not meet 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 moment as the initial moment and obtaining the state parameters corresponding to the target spacecraft at the initial moment, the method further includes: Initial correction maneuver parameters corresponding to the target spacecraft are set; the initial correction maneuver parameters include the preset maximum optimization time interval, the preset rolling optimization time interval, the preset minimum execution speed increment, and the preset correction maneuver minimum time interval.
10. An electronic device, characterized in that: include: A memory, 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 the spacecraft orbit tracking and guidance method for autonomous flight in the cis-lunar space as described in any one of claims 1 to 9.