A launch time and orbit transfer impulse joint planning method for a moon-earth transfer mission

Abstract

The application discloses a launch time and orbit transfer impulse joint planning method for a moon-earth transfer mission, and the method comprises the following steps: obtaining initial orbit, launch time window, propulsion and reentry corridor and other task constraints, adopting a coarse and fine layered strategy to generate candidate launch time, parameterizing a phase of a periodic orbit and quickly searching for available phase points; selecting a single impulse, two impulses or multiple small impulse templates, constructing candidate combinations, quickly estimating candidate trajectories according to a simplified dynamic model, screening feasible schemes through constraint checking, obtaining optimal initial values through multi-index scoring and local encryption, carrying out trajectory propagation and multi-target iterative correction under a full perturbation high-precision model, and after completing convergence and engineering constraint review, outputting a final trajectory and control command. The application realizes joint optimization of launch time and impulse parameters, considers calculation efficiency and accuracy, is suitable for on-board real-time planning, and can output a moon-earth transfer trajectory meeting reentry corridor and multiple engineering constraints.

Description

Technical Field

[0001] This invention relates to the field of spacecraft orbit design and guidance control technology, and in particular to a method for jointly planning the departure time and orbit change pulse for a lunar-Earth transfer mission. Background Technology

[0002] With the continuous advancement of deep space exploration and manned / lunar exploration missions, the long-term operating orbits of spacecraft in the Earth-Moon space are becoming increasingly diversified, typically including near-linear halo orbits (NRHO), distant retrograde orbits (DRO), and other periodic orbits near translation points. These orbits are characterized by long-term residence, low energy, and excellent communication / observation geometry, and are therefore widely used for relay, space station, and rover exploration missions.

[0003] In practical missions, spacecraft returning to Earth from the aforementioned orbits often involves more than just reaching the vicinity of Earth. They must meet a set of critical engineering constraints: if the mission ultimately re-enters the atmosphere and returns to Earth (or the return capsule re-enters), certain entry corridor conditions must be met at the designated re-entry interface. These conditions include, for example, the velocity range at the re-entry interface altitude, the flight path angle range, the arrival time window, and the latitude and longitude window of the entry point. If the mission only needs to return to near-Earth for subsequent processing, the terminal conditions are often described as a type of entry corridor or a set of constraints, rather than being strictly equivalent to a specific reference orbit state. Furthermore, in-orbit operation involves initial state deviations, orbital perturbations, execution errors, and constraint changes (such as changes in communication windows, attitude limitations, and obstruction). This makes it difficult for offline, one-time orbit design to cover all in-orbit scenarios, necessitating online or quasi-online replanning.

[0004] However, traditional lunar-Earth transfer or return designs rely heavily on high-precision ground dynamic models and large-scale global optimization iterations, which involve large computational loads, numerous parameter adjustments, and are difficult to deploy directly to onboard computers for real-time operation. In addition, many processes adopt a serial approach of first selecting the departure time and then optimizing the pulse, making it difficult to find the comprehensive optimal combination of departure time and pulse sequence within a given time window, and also difficult to uniformly balance the conflicts between multiple constraints. Summary of the Invention

[0005] To address the shortcomings of existing technologies, this invention provides a joint planning method for the departure time and orbit change pulses of a lunar-Earth transfer mission. This invention can achieve joint optimization of departure time and pulse sequence, unified modeling of multiple constraints, and multi-fidelity hierarchical solution. It has the capability of real-time or near-real-time deployment on satellite and outputs an engineering-executable lunar-Earth return trajectory.

[0006] The technical solution of this invention is: a method for jointly planning the departure time and orbit change pulse of a lunar-Earth transfer mission, comprising: Obtain the mission input set for the lunar-Earth transfer mission. The mission input set includes the spacecraft's initial orbit, departure time window, propulsion constraints, flight time constraints, geometric and mission constraints, and terminal reentry or return corridor constraints. Based on the task input set, a coarse-to-fine layered candidate departure time sequence is generated within the departure time window. The periodic orbit is phase parameterized, and the set of available departure phase points for each candidate departure time is quickly retrieved based on the phase parameters. Select one or more pulse templates according to the task requirements, and specify the feasible range of pulse time for each pulse template. Combine the set of available start phase points at each candidate start time with the selected pulse template as a candidate combination. Each candidate combination is estimated based on a simplified dynamic model, resulting in a set of corresponding candidate lunar-Earth transfer trajectory schemes; Perform constraint checks on all candidate lunar-Earth transfer trajectory schemes, eliminate those that do not meet the constraints, and obtain a set of feasible candidate lunar-Earth transfer trajectory schemes; By using multiple indicators to comprehensively score and screen the set of feasible candidate lunar-Earth transfer trajectory schemes, a set of preferred candidate transfer trajectories is obtained. Then, a local encrypted search is performed on the set of preferred candidate transfer trajectories to obtain an initial set of preferred candidate transfer trajectories. Under the full perturbation high-precision dynamic model, each preferred candidate initial value in the preferred candidate transfer trajectory initial value set is used as the initial solution, and trajectory accurate propagation and multiple target iterative correction are carried out respectively to obtain a high-precision corrected trajectory set; Convergence checks and engineering constraint verifications are performed on the high-precision corrected trajectory set. Unconverged corrected trajectories are removed to obtain the high-precision feasible trajectory set. The high-precision feasible trajectory set is comprehensively evaluated based on the total orbit change, flight time, terminal corridor margin, control execution margin, and mission preference. The high-precision feasible trajectory with the best comprehensive evaluation is selected as the final lunar-Earth transfer trajectory, and the corresponding control commands are output.

[0007] Secondly, the present invention provides a joint planning system for the departure time and orbit change pulse of a lunar-Earth transfer mission, comprising: The data acquisition module is used to acquire the mission input set for the lunar-Earth transfer mission. The mission input set includes the spacecraft's initial orbit, departure time window, propulsion constraints, flight time constraints, geometric and mission constraints, and terminal reentry or return corridor constraints. The departure phase point retrieval module is used to generate a coarse-to-fine tiered candidate departure time sequence within the departure time window based on the task input set, perform phase parameterization on the periodic orbit, and quickly retrieve the set of available departure phase points for each candidate departure time based on the phase parameters. The template selection module is used to select one or more pulse templates according to task requirements, and to specify the feasible range of pulse time for each pulse template. It combines the set of available start phase points at each candidate start time with the selected pulse template as a candidate combination. The initial trajectory estimation module is used to estimate each candidate combination based on a simplified dynamic model, thereby obtaining a set of corresponding candidate lunar-Earth transfer trajectory schemes; The feasible trajectory generation module is used to perform constraint checks on all candidate lunar-Earth transfer trajectory schemes, eliminate candidate lunar-Earth transfer trajectory schemes that do not meet the constraints, and obtain a set of feasible candidate lunar-Earth transfer trajectory schemes. The trajectory initial value generation module uses multiple indicators to comprehensively score and filter the set of feasible candidate lunar-Earth transfer trajectory schemes to obtain the preferred candidate transfer orbit set, and performs local encrypted search on the preferred candidate transfer orbit set to obtain the preferred candidate transfer trajectory initial value set. The trajectory correction module is used to perform precise trajectory propagation and multiple target shooting iteration correction under the full perturbation high-precision dynamic model, using each preferred candidate initial value in the preferred candidate transfer trajectory initial value set as the initial solution, to obtain a high-precision corrected trajectory set. The control command output module performs convergence checks and engineering constraint verifications on the high-precision corrected trajectory set, eliminates non-converged corrected trajectories, and obtains a high-precision feasible trajectory set. The high-precision feasible trajectory set is comprehensively evaluated based on the total orbit change, flight time, terminal corridor margin, control execution margin, and mission preference. The high-precision feasible trajectory with the best comprehensive evaluation is selected as the final lunar-Earth transfer trajectory, and the corresponding control commands are output.

[0008] Thirdly, the present invention also provides an electronic device, comprising: At least one processor; and A memory that is communicatively connected to the at least one processor; The memory stores a computer program that can be executed by the at least one processor. When the computer program is executed by the at least one processor, it implements the method for jointly planning the departure time and orbit change pulse of a lunar transfer mission.

[0009] The beneficial effects of this invention are as follows: 1. This invention enables the representation of periodic orbits of translational points such as NRHO / DRO using phase parameters through periodic orbit phase parameterization and fast phase retrieval within a window, and rapidly generates a set of available phase points within the departure time window, significantly reducing the cost of candidate generation; 2. This invention transforms the terminal target from a strict reference orbital state error into a set of corridor constraints at EI, making the lunar-Earth return background logic self-consistent and the engineering reasonable. 3. This invention uses candidate departure time series, departure phase point set, pulse template, and terminal corridor parameterization to jointly plan the departure time and pulse parameters as a whole, and quickly form convergent initial trajectory values ​​for candidate combinations. 4. This invention achieves large-scale rapid screening by simplifying the dynamic model, and improves the speed and accuracy of screening by performing differential correction and verification on a small number of preferred schemes through a fully perturbated high-precision dynamic model. 5. Based on the state transition matrix / sensitivity information, this invention iteratively corrects the pulse vector and piecewise time to ensure that the terminal strictly enters the corridor and satisfies multiple constraints. Attached Figure Description

[0010] Figure 1 This is a flowchart illustrating the method of Embodiment 1 of the present invention; Figure 2 This is a schematic diagram of the lunar-Earth transfer trajectory in Embodiment 1 of the present invention. Detailed Implementation

[0011] The specific embodiments of the present invention will be further described below with reference to the accompanying drawings: like Figure 1 and 2 As shown, this embodiment provides a method for jointly planning the departure time and orbit change pulse for a lunar-Earth transfer mission, including: S1. Obtain the mission input set for the lunar-Earth transfer mission. The mission input set includes the spacecraft's initial orbit, departure time window, propulsion constraints, flight time constraints, geometric and mission constraints, and terminal reentry or return corridor constraints. S2. Based on the task input set, generate a coarse-to-fine layered candidate departure time sequence within the departure time window, perform phase parameterization on the periodic orbit, and quickly retrieve the set of available departure phase points for each candidate departure time based on the phase parameters. S3. Select one or more pulse templates according to the task requirements, and specify the feasible range of pulse time for each pulse template. Combine the set of available start phase points and the selected pulse template at each candidate start time as a candidate combination. S4. Estimate each candidate combination based on a simplified dynamic model to obtain the corresponding set of candidate lunar-Earth transfer trajectory schemes; S5. Perform constraint checks on all candidate lunar-Earth transfer trajectory schemes, eliminate candidate lunar-Earth transfer trajectory schemes that do not meet the constraint conditions, and obtain a feasible set of candidate lunar-Earth transfer trajectory schemes. S6. Using multiple indicators, a comprehensive evaluation and screening of the set of feasible candidate lunar-Earth transfer trajectory schemes is performed to obtain a set of preferred candidate transfer trajectories. A local encrypted search is then performed on the set of preferred candidate transfer trajectories to obtain a set of initial values ​​for preferred candidate transfer trajectories. S7. Under the full perturbation high-precision dynamic model, each preferred candidate initial value in the preferred candidate transfer trajectory initial value set is used as the initial solution, and trajectory accurate propagation and multiple target iterative correction are carried out respectively to obtain a high-precision corrected trajectory set. S8. Perform convergence checks and engineering constraint verifications on the high-precision corrected trajectory set, remove non-converged corrected trajectories, and obtain the high-precision feasible trajectory set. Perform a comprehensive evaluation of the high-precision feasible trajectory set based on total orbit change, flight time, terminal corridor margin, control execution margin, and mission preference. Select the high-precision feasible trajectory with the best comprehensive evaluation as the final lunar-Earth transfer trajectory and output the corresponding control commands. In this embodiment, in step S1, the initial orbit of the spacecraft includes the type of periodic orbit and the current initial state. or current phase ,in, These are the initial position vector and the initial velocity vector, respectively.

[0012] In this embodiment, in step S1, the spacecraft's departure time window is the return window given by the mission plan, which can be represented by an absolute epoch or relative to the current time. Within the return window, telemetry and control visibility and attitude pointing constraints can be further superimposed to form a usable sub-window.

[0013] In this embodiment, the propulsion constraints include the modified pulse limit, single pulse limit, and pulse count limit.

[0014] The flight time constraints are given by factors such as mission timing, thermal control and power supply life, and include minimum and maximum flight time requirements.

[0015] In this embodiment, in step S2, based on the task input set, a coarse-to-fine layered candidate departure time sequence is generated within the departure time window. The periodic orbit is then phase-parameterized, and the available departure phase point set for each candidate departure time is quickly retrieved based on the phase parameters. This includes: S21. A two-layer time sampling strategy of first performing a global coarse search and then a local fine search is adopted to improve the search accuracy under limited on-board computing power, generating a coarse-fine layer candidate departure time sequence, as follows: Input the absolute departure time window given by the task, overlay the telemetry and control visibility, Earth-Moon occlusion, and attitude pointing taboo intervals, remove unusable time periods, and obtain the available departure sub-window; Within the available sub-windows, set a fixed coarse step size. A uniform candidate start time sequence is generated to obtain a coarse-grained candidate time set, wherein the coarse step size is... Selected based on window length and onboard computing power; Fine step size is applied to the coarse-grained candidate time set. Encrypted sampling yields a fine-grained set of candidate time points; The fine-grained candidate time set and the coarse-grained candidate time set are merged and deduplicated to form a coarse-fine layered candidate departure time sequence.

[0016] S22. Perform phase parameterization on the periodic orbit and quickly retrieve the set of available phase points for each candidate departure time based on the phase parameters, as detailed below: Phase parameter representation of the periodic orbits of the NRHO and DRO lunar translation points is used to achieve rapid mapping of orbital states at any departure time. For each candidate departure time in the coarse-to-fine layer candidate departure time sequence Based on the orbital period and the current phase The available departure phases for the corresponding candidate departure times are obtained through phase propagation calculations. ; S23. Output the initial orbital state of the spacecraft at the corresponding candidate departure time using the phase-state index table. .

[0017] In this embodiment, in step S3, the pulse template includes: The single-pulse template T1 is defined as applying a main orbit change pulse near the candidate departure time. The main orbit change pulse completes the energy adjustment, orbital plane adjustment and geometric shaping of the lunar-Earth transfer orbit, and completes the direct cut from the periodic orbit to the lunar-Earth transfer orbit. The two-pulse template T2 refers to applying two orbital change pulses on the transfer trajectory, which are divided into a main departure pulse and a mid-course correction pulse. The main departure pulse is used to provide the main energy required for the lunar-Earth transfer and to determine the basic shape of the transfer trajectory. The mid-course correction pulse is used to eliminate orbital propagation errors, adjust flight geometry, and correct the reentry point state to ensure that the spacecraft accurately enters the reentry corridor. The multi-pulse template T3 is defined as applying three or more small correction pulses on multiple segments of the transfer trajectory.

[0018] In this embodiment, a time description method is established for each type of pulse template, and a legal ignition time interval is defined; Based on the candidate start times of the coarse-fine layer candidate start time sequence, upper and lower bound constraints are set for the time offset of each pulse; the minimum time interval between adjacent pulses is limited for the two-pulse template T2 and the multi-small-pulse template T3.

[0019] In this embodiment, a fixed initial pulse direction guessing strategy is preset for each type of pulse template, including: A pulse is applied along the direction of the spacecraft's orbital velocity to adjust the orbital energy and flight speed; A pulse is applied radially along the Earth-Moon line to adjust the perigee altitude and reentry point position. A pulse is applied along the orbital plane normal to adjust the orbital plane and reentry latitude and longitude. A multi-directional weighted combination is adopted, and the proportion of each directional component is automatically allocated according to the task type to form a comprehensive optimal initial direction.

[0020] In this embodiment, in step S4, each candidate combination is estimated based on a simplified dynamic model to obtain a corresponding set of candidate lunar-Earth transfer trajectory schemes; specifically as follows: S41. Transform the continuous re-entry or return corridor constraints into a finite number of discrete terminal sample points. For each terminal sample point, construct a standard terminal state that satisfies the corridor constraints, as follows: The re-entry interface intersection point is described by several configurable parameters, including the re-entry arrival time, the geographical location of the entry point, the speed at the interface, and the flight path angle. A set of discrete terminal states is formed by uniformly or boundary sampling of several configurable parameters according to the allowable range of the task. For each terminal sample point, a standard terminal state that satisfies the corridor constraint is constructed as the target terminal condition for the propagation of the lunar-Earth transfer trajectory.

[0021] S42. Under the simplified dynamic model, construct a starting state... To the terminal location For boundary value problems, either the Lambert solution or the boundary value solution is used as the initial transition velocity; specifically as follows: The two-body model, the conical splicing model, or the simplified CRTBP model are selected as the simplified dynamic model; For each group of starting states To the terminal location The Lambert algorithm is used to solve for the lunar-Earth transfer orbit, and the initial transfer velocity that satisfies the position boundary conditions is obtained; The difference between the starting trajectory velocity and the initial transfer velocity obtained from the solution is used to obtain the starting pulse. .

[0022] S43. Determine the starting pulse and intermediate correction point settings based on the selected pulse template to obtain the candidate lunar-Earth transfer trajectory with initial pulse values; details are as follows: For a single-pulse template T1, the trigger pulse will be... As a single main pulse near the departure time, it is used to form the nominal lunar-Earth transfer orbit; For the two-pulse template T2, a starting pulse is applied at the starting time. During the transfer process, a mid-course correction point is set. When the spacecraft propagates to this correction point, based on the deviation between the actual or predicted state at the correction point and the nominal orbit, terminal reentry, or return corridor constraints, a boundary value problem or target-hitting correction problem is constructed from the state at the correction point to the target in the terminal corridor, and the first correction pulse is solved. When the deviation or constraint default exceeds a preset threshold, the first correction pulse is applied. ; For the multi-pulse template T3, a starting pulse is applied at the starting time. And set multiple mid-course correction points during the transfer; when the spacecraft propagates to the... When there are one correction point, according to the correction point Given the state deviation and terminal corridor constraints at point 1, resolve the solution for point 2. Correction pulse When the corresponding deviation or constraint default exceeds a preset threshold, the first... One correction pulse.

[0023] In this embodiment, the total orbit change of the candidate lunar-Earth transfer trajectory scheme It is the sum of the starting pulse and all the correction pulses that have been applied.

[0024] S44. Propagate the lunar-Earth transfer orbit with initial pulse values ​​in the simplified dynamic model to obtain the intersection state of the lunar-Earth transfer orbit and the reentry interface EI.

[0025] S45. Extract the reentry interface (EI) index from the propagation results. The reentry interface (EI) index includes the velocity at the reentry interface. Flight path angle latitude and longitude of the entry point Arrival time And the distance to the corridor boundary and the flight time T.

[0026] In this embodiment, in step S5, a constraint check is performed on all candidate lunar-Earth transfer trajectory schemes, and candidate lunar-Earth transfer trajectory schemes that do not meet the constraint conditions are eliminated to obtain a feasible set of candidate lunar-Earth transfer trajectory schemes, including: The constraint checks performed on all candidate lunar-Earth transfer trajectory schemes include propulsion constraints, flight time constraints, geometry and mission constraints, and terminal reentry or return corridor constraints. For all feasible candidate lunar-Earth transfer trajectory schemes with corridor feasibility, corridor margin and corridor violation degree are calculated respectively. The corridor margin is expressed as: ; The aforementioned corridor default degree The expression is as follows: In the formula, For the first Normalization margin of each corridor constraint; For the first One corridor constraint parameter; , These are the maximum and minimum corridor constraint parameters at the re-entry interface, respectively. This represents the number of corridor constraint parameters.

[0027] In this embodiment, step S6 involves using multiple indicators to comprehensively score and filter the set of feasible candidate lunar-Earth transfer trajectory schemes to obtain a preferred set of candidate transfer trajectories. A local encrypted search is then performed on the preferred set of candidate transfer trajectories to obtain an initial set of preferred candidate transfer trajectories; including: S61. All feasible candidate lunar-Earth transfer trajectory schemes are comprehensively scored according to propulsion consumption, flight time, corridor margin and geometric and mission constraints. The comprehensive scores are sorted from smallest to largest, and the top N feasible candidate lunar-Earth transfer trajectory schemes are selected as the preferred candidate transfer trajectory set. When there are multiple conflicting objectives, a set of preferred candidate transfer trajectories is obtained by using hierarchical priority sorting or Pareto screening. ; In the formula, This indicates the overall score; Indicates the corrected pulse, Indicates the total flight time; For the degree of default in the corridor; For corridor margin; These are the weighting coefficients.

[0028] S62. Taking each preferred candidate transfer trajectory scheme as the center, local densification is performed in the neighborhood of its corresponding departure time, phase point or related discrete parameters to generate a new neighborhood candidate transfer trajectory combination and perform trajectory propagation and constraint check to obtain the initial value set of preferred candidate transfer trajectories.

[0029] In this embodiment, each preferred candidate transfer trajectory scheme is re-discretized with a local encryption step size smaller than the coarse search step size to generate a new neighborhood candidate transfer trajectory combination.

[0030] In this embodiment, the optimal candidate transfer trajectory set includes preferred departure time, trajectory phase and initial state, pulse template, pulse times, and terminal corridor parameters.

[0031] In this embodiment, in step S7, under the fully perturbation high-precision dynamic model, each preferred candidate initial value in the preferred candidate transfer trajectory initial value set is used as the initial solution, and trajectory accurate propagation and multiple target shooting iterative correction are carried out respectively to obtain a high-precision corrected trajectory set; specifically as follows: S71. Establish a high-precision dynamic model consistent with the real flight environment, and simultaneously incorporate all key perturbation terms during trajectory integration; The high-precision dynamic model includes gravitational perturbations of the Earth, Moon, Sun and corresponding celestial bodies in the solar system, higher-order non-spherical gravitational terms of the Earth and Moon, and solar radiation pressure perturbations; Simultaneously integrate the spacecraft's position and velocity status, and integrate the state transition matrix.

[0032] S72. Using the pulse application time as the segment node, the entire lunar-Earth transfer trajectory is divided into multiple uncontrolled free-flight arc segments. The pulse is applied only at the nodes between the arc segments. The initial state of each arc segment is obtained by superimposing the end state of the previous arc segment with the pulse velocity increment of the current node.

[0033] S73. Using pulse magnitude, pulse direction, and pulse application time as optimization variables to be corrected, and with reentry corridor constraints, flight time constraints, propulsion constraints, and geometric and mission constraints as constraints, iterative corrections are carried out; specifically as follows: Substitute the current initial pulse value into the high-precision dynamic model and integrate it to the reentry interface to calculate the residual between the actual state of the terminal and the target state of the reentry corridor. A linear relationship between the change of pulse variable and the change of terminal residual is established by using the state transition matrix. Under the premise of satisfying the upper limit of single pulse, the upper limit of corrected pulse, and the pulse time window constraints, the least squares or constraint optimization method is used to solve the optimal pulse correction amount. Substitute the corrected pulse parameters back into the high-precision dynamic model integral to update the terminal residual; Repeat the iterative process until the terminal residual is less than the preset convergence threshold, the terminal state is completely within the re-entry corridor, and all constraints are satisfied.

[0034] An embodiment of this application also provides a joint planning system for the departure time and orbit change pulse of a lunar-Earth transfer mission, comprising: The data acquisition module is used to acquire the mission input set for the lunar-Earth transfer mission. The mission input set includes the spacecraft's initial orbit, departure time window, propulsion constraints, flight time constraints, geometric and mission constraints, and terminal reentry or return corridor constraints. The departure phase point retrieval module is used to generate a coarse-to-fine tiered candidate departure time sequence within the departure time window based on the task input set, perform phase parameterization on the periodic orbit, and quickly retrieve the set of available departure phase points for each candidate departure time based on the phase parameters. The template selection module is used to select one or more pulse templates according to task requirements, and to specify the feasible range of pulse time for each pulse template. It combines the set of available start phase points at each candidate start time with the selected pulse template as a candidate combination. The initial trajectory estimation module is used to estimate each candidate combination based on a simplified dynamic model, thereby obtaining a set of corresponding candidate lunar-Earth transfer trajectory schemes; The feasible trajectory generation module is used to perform constraint checks on all candidate lunar-Earth transfer trajectory schemes, eliminate candidate lunar-Earth transfer trajectory schemes that do not meet the constraints, and obtain a set of feasible candidate lunar-Earth transfer trajectory schemes. The trajectory initial value generation module uses multiple indicators to comprehensively score and filter the set of feasible candidate lunar-Earth transfer trajectory schemes to obtain the preferred candidate transfer orbit set, and performs local encrypted search on the preferred candidate transfer orbit set to obtain the preferred candidate transfer trajectory initial value set. The trajectory correction module is used to perform precise trajectory propagation and multiple target shooting iteration correction under the full perturbation high-precision dynamic model, using each preferred candidate initial value in the preferred candidate transfer trajectory initial value set as the initial solution, to obtain a high-precision corrected trajectory set. The control command output module performs convergence checks and engineering constraint verifications on the high-precision corrected trajectory set, eliminates non-converged corrected trajectories, and obtains a high-precision feasible trajectory set. The high-precision feasible trajectory set is comprehensively evaluated based on the total orbit change, flight time, terminal corridor margin, control execution margin, and mission preference. The high-precision feasible trajectory with the best comprehensive evaluation is selected as the final lunar-Earth transfer trajectory, and the corresponding control commands are output.

[0035] In this embodiment, the initial orbit of the spacecraft acquired by the data acquisition module includes the orbital type of the current period and the current initial state. or current phase ,in, These are the initial position vector and the initial velocity vector, respectively.

[0036] In this embodiment, the spacecraft's departure time window acquired by the data acquisition module is the return window given by the mission plan, which can be represented by absolute epoch or relative to the current time. Within the return window, telemetry and control visibility and attitude pointing constraints can be further superimposed to form usable sub-windows.

[0037] In this embodiment, the propulsion constraints acquired by the data acquisition module include the corrected pulse upper limit, single pulse upper limit, and pulse count upper limit.

[0038] The flight time constraints are given by factors such as mission timing, thermal control and power supply life, and include minimum and maximum flight time requirements.

[0039] In this embodiment, the departure phase point retrieval module generates a coarse-to-fine layered candidate departure time sequence within the departure time window based on the task input set, performs phase parameterization on the periodic orbit, and quickly retrieves the set of available departure phase points for each candidate departure time according to the phase parameters, including: A two-layer time sampling strategy, consisting of a global coarse search followed by a local fine search, is employed to improve search accuracy under limited on-board computing power, generating a coarse-fine layer candidate departure time sequence, as follows: Input the absolute departure time window given by the task, overlay the telemetry and control visibility, Earth-Moon occlusion, and attitude pointing taboo intervals, remove unusable time periods, and obtain the available departure sub-window; Within the available sub-windows, set a fixed coarse step size. A uniform candidate start time sequence is generated to obtain a coarse-grained candidate time set, wherein the coarse step size is... Selected based on window length and onboard computing power; Fine step size is applied to the coarse-grained candidate time set. Encrypted sampling yields a fine-grained set of candidate time points; The fine-grained candidate time set and the coarse-grained candidate time set are merged and deduplicated to form a coarse-fine layered candidate departure time sequence.

[0040] The periodic orbit is phase parameterized, and the set of available phase points for each candidate departure time is quickly retrieved based on the phase parameters, as detailed below: Phase parameter representation of the periodic orbits of the NRHO and DRO lunar translation points is used to achieve rapid mapping of orbital states at any departure time. For each candidate departure time in the coarse-to-fine layer candidate departure time sequence Based on the orbital period and the current phase The available departure phases for the corresponding candidate departure times are obtained through phase propagation calculations. ; Output the initial orbital state of the spacecraft at the corresponding candidate departure time using the phase-state index table. .

[0041] In this embodiment, the pulse template of the template selection module includes: The single-pulse template T1 is defined as applying a main orbit change pulse near the candidate departure time. The main orbit change pulse completes the energy adjustment, orbital plane adjustment and geometric shaping of the lunar-Earth transfer orbit, and completes the direct cut from the periodic orbit to the lunar-Earth transfer orbit. The two-pulse template T2 refers to applying two orbital change pulses on the transfer trajectory, which are divided into a main departure pulse and a mid-course correction pulse. The main departure pulse is used to provide the main energy required for the lunar-Earth transfer and to determine the basic shape of the transfer trajectory. The mid-course correction pulse is used to eliminate orbital propagation errors, adjust flight geometry, and correct the reentry point state to ensure that the spacecraft accurately enters the reentry corridor. The multi-pulse template T3 is defined as applying three or more small correction pulses on multiple segments of the transfer trajectory.

[0042] In this embodiment, a time description method is established for each type of pulse template, and a legal ignition time interval is defined; Based on the candidate start times of the coarse-fine layer candidate start time sequence, upper and lower bound constraints are set for the time offset of each pulse; the minimum time interval between adjacent pulses is limited for the two-pulse template T2 and the multi-small-pulse template T3.

[0043] In this embodiment, a fixed initial pulse direction guessing strategy is preset for each type of pulse template, including: A pulse is applied along the direction of the spacecraft's orbital velocity to adjust the orbital energy and flight speed; A pulse is applied radially along the Earth-Moon line to adjust the perigee altitude and reentry point position. A pulse is applied along the orbital plane normal to adjust the orbital plane and reentry latitude and longitude. A multi-directional weighted combination is adopted, and the proportion of each directional component is automatically allocated according to the task type to form a comprehensive optimal initial direction.

[0044] In this embodiment, the initial trajectory estimation module estimates each candidate combination based on a simplified dynamic model, obtaining a corresponding set of candidate lunar-Earth transfer trajectory schemes; specifically as follows: The continuous re-entry or return corridor constraints are transformed into a finite number of discrete terminal sample points. For each terminal sample point, a standard terminal state satisfying the corridor constraints is constructed, as follows: The re-entry interface intersection point is described by several configurable parameters, including the re-entry arrival time, the geographical location of the entry point, the speed at the interface, and the flight path angle. A set of discrete terminal states is formed by uniformly or boundary sampling of several configurable parameters according to the allowable range of the task. For each terminal sample point, a standard terminal state that satisfies the corridor constraint is constructed as the target terminal condition for the propagation of the lunar-Earth transfer trajectory.

[0045] Under the simplified dynamic model, construct the starting state To the terminal location For boundary value problems, either the Lambert solution or the boundary value solution is used as the initial transition velocity; specifically as follows: The two-body model, the conical splicing model, or the simplified CRTBP model are selected as the simplified dynamic model; For each group of starting states To the terminal location The Lambert algorithm is used to solve for the lunar-Earth transfer orbit, and the initial transfer velocity that satisfies the position boundary conditions is obtained; The difference between the starting trajectory velocity and the initial transfer velocity obtained from the solution is used to obtain the starting pulse. .

[0046] Based on the selected pulse template, the starting pulse and intermediate correction point settings are determined, resulting in a candidate lunar-Earth transfer trajectory with initial pulse values; details are as follows: For a single-pulse template T1, the trigger pulse will be... As a single main pulse near the departure time, it is used to form the nominal lunar-Earth transfer orbit; For the two-pulse template T2, a starting pulse is applied at the starting time. During the transfer process, a mid-course correction point is set. When the spacecraft propagates to this correction point, based on the deviation between the actual or predicted state at the correction point and the nominal orbit, terminal reentry, or return corridor constraints, a boundary value problem or target-hitting correction problem is constructed from the state at the correction point to the target in the terminal corridor, and the first correction pulse is solved. When the deviation or constraint default exceeds a preset threshold, the first correction pulse is applied. ; For the multi-pulse template T3, a starting pulse is applied at the starting time. And set multiple mid-course correction points during the transfer; when the spacecraft propagates to the... When there are one correction point, according to the correction point Given the state deviation and terminal corridor constraints at point 1, resolve the solution for point 2. Correction pulse When the corresponding deviation or constraint default exceeds a preset threshold, the first... One correction pulse.

[0047] In this embodiment, the total orbit change of the candidate lunar-Earth transfer trajectory scheme It is the sum of the starting pulse and all the correction pulses that have been applied.

[0048] In a simplified dynamic model, a lunar-Earth transfer orbit with initial pulse values ​​is propagated to obtain the intersection state of the lunar-Earth transfer orbit and the reentry interface EI.

[0049] The reentry interface (EI) metrics are extracted from the propagation results. These EI metrics include the velocity at the reentry interface. Flight path angle latitude and longitude of the entry point Arrival time And the distance to the corridor boundary and the flight time T.

[0050] In this embodiment, the feasible trajectory generation module performs constraint checks on all candidate lunar-Earth transfer trajectory schemes, eliminates candidate lunar-Earth transfer trajectory schemes that do not meet the constraints, and obtains a set of feasible candidate lunar-Earth transfer trajectory schemes, including: The constraint checks performed on all candidate lunar-Earth transfer trajectory schemes include propulsion constraints, flight time constraints, geometry and mission constraints, and terminal reentry or return corridor constraints. For all feasible candidate lunar-Earth transfer trajectory schemes with corridor feasibility, corridor margin and corridor violation degree are calculated respectively. The corridor margin is expressed as: ; The aforementioned corridor default degree The expression is as follows: In the formula, For the first Normalization margin of each corridor constraint; For the first One corridor constraint parameter; , These are the maximum and minimum corridor constraint parameters at the re-entry interface, respectively. This represents the number of corridor constraint parameters.

[0051] In this embodiment, the trajectory initial value generation module uses multiple indicators to comprehensively score and filter the set of feasible candidate lunar-Earth transfer trajectory schemes to obtain a preferred candidate transfer trajectory set, and performs a local encrypted search on the preferred candidate transfer trajectory set to obtain a preferred candidate transfer trajectory initial value set; including: All feasible candidate lunar-Earth transfer trajectory schemes are comprehensively scored according to propulsion consumption, flight time, corridor margin, and degree of satisfaction of geometric and mission constraints. The comprehensive scores are sorted from smallest to largest, and the top N feasible candidate lunar-Earth transfer trajectory schemes are selected as the preferred candidate transfer trajectory set. When there are multiple conflicting objectives, a set of preferred candidate transfer trajectories is obtained by using hierarchical priority sorting or Pareto screening. ; In the formula, This indicates the overall score; Indicates the corrected pulse, Indicates the total flight time; For the degree of default in the corridor; For corridor margin; These are the weighting coefficients.

[0052] Taking each preferred candidate transfer trajectory scheme as the center, local densification is performed in the neighborhood of its corresponding departure time, phase point or related discrete parameters to generate a new neighborhood candidate transfer trajectory combination and perform trajectory propagation and constraint checks to obtain the initial value set of preferred candidate transfer trajectories.

[0053] In this embodiment, each preferred candidate transfer trajectory scheme is re-discretized with a local encryption step size smaller than the coarse search step size to generate a new neighborhood candidate transfer trajectory combination.

[0054] In this embodiment, the optimal candidate transfer trajectory set includes preferred departure time, trajectory phase and initial state, pulse template, pulse times, and terminal corridor parameters.

[0055] In this embodiment, under the fully perturbation high-precision dynamic model, the trajectory correction module uses each preferred candidate initial value in the preferred candidate transfer trajectory initial value set as the initial solution, and performs trajectory accurate propagation and multiple target-hitting iterative correction to obtain a high-precision corrected trajectory set; specifically as follows: Establish a high-precision dynamic model consistent with the real flight environment, and simultaneously incorporate all key perturbation terms during trajectory integration; The high-precision dynamic model includes gravitational perturbations of the Earth, Moon, Sun and corresponding celestial bodies in the solar system, higher-order non-spherical gravitational terms of the Earth and Moon, and solar radiation pressure perturbations; Simultaneously integrate the spacecraft's position and velocity status, and integrate the state transition matrix.

[0056] Using the pulse application time as the segment node, the entire lunar-Earth transfer trajectory is divided into multiple uncontrolled free-flight arc segments. The pulse is applied only at the nodes between the arc segments. The initial state of each arc segment is obtained by superimposing the end state of the previous arc segment with the pulse velocity increment of the current node.

[0057] The pulse magnitude, pulse direction, and pulse application time are used as optimization variables to be corrected, and reentry corridor constraints, flight time constraints, propulsion constraints, and geometric and mission constraints are used as constraints to carry out iterative correction; as detailed below: Substitute the current initial pulse value into the high-precision dynamic model and integrate it to the reentry interface to calculate the residual between the actual state of the terminal and the target state of the reentry corridor. A linear relationship between the change of pulse variable and the change of terminal residual is established by using the state transition matrix. Under the premise of satisfying the upper limit of single pulse, the upper limit of corrected pulse, and the pulse time window constraints, the least squares or constraint optimization method is used to solve the optimal pulse correction amount. Substitute the corrected pulse parameters back into the high-precision dynamic model integral to update the terminal residual; Repeat the iterative process until the terminal residual is less than the preset convergence threshold, the terminal state is completely within the re-entry corridor, and all constraints are satisfied.

[0058] Embodiments of this application also provide an electronic device, including: At least one processor; and A memory that is communicatively connected to the at least one processor; The memory stores a computer program that can be executed by the at least one processor. When the computer program is executed by the at least one processor, it implements the method for jointly planning the departure time and orbit change pulse of a lunar transfer mission.

[0059] In this embodiment, the memory can be implemented by any type of volatile or non-volatile storage device or a combination thereof, such as static random access memory (SRAM), electrically erasable programmable read-only memory (EEPROM), erasable programmable read-only memory (EPROM), programmable read-only memory (PROM), read-only memory (ROM), magnetic storage, flash memory, magnetic disk, or optical disk. A processor, coupled to the memory, is used to execute computer programs stored in the memory.

[0060] Computer programs include computer program code, which can be in the form of source code, object code, executable files, or certain intermediate forms.

[0061] The embodiments and descriptions above are merely illustrative of the principles and preferred embodiments of the present invention. Various changes and modifications may be made to the present invention without departing from its spirit and scope, and all such changes and modifications fall within the scope of the present invention as claimed.

Claims

1. A method for jointly planning the departure time and orbit change pulse for a lunar-Earth transfer mission, characterized in that, include: Obtain the mission input set for the lunar-Earth transfer mission. The mission input set includes the spacecraft's initial orbit, departure time window, propulsion constraints, flight time constraints, geometric and mission constraints, and terminal reentry or return corridor constraints. Based on the task input set, a coarse-to-fine layered candidate departure time sequence is generated within the departure time window, the periodic orbit is phase parameterized, and the set of available departure phase points for each candidate departure time is quickly retrieved based on the phase parameters. Select one or more pulse templates according to the task requirements, and specify the feasible range of pulse time for each pulse template. Combine the set of available start phase points at each candidate start time with the selected pulse template as a candidate combination. Each candidate combination is estimated based on a simplified dynamic model, resulting in a set of corresponding candidate lunar-Earth transfer trajectory schemes; Perform constraint checks on all candidate lunar-Earth transfer trajectory schemes, eliminate those that do not meet the constraints, and obtain a set of feasible candidate lunar-Earth transfer trajectory schemes; By using multiple indicators to comprehensively score and screen the set of feasible candidate lunar-Earth transfer trajectory schemes, a set of preferred candidate transfer trajectories is obtained. Then, a local encrypted search is performed on the set of preferred candidate transfer trajectories to obtain an initial set of preferred candidate transfer trajectories. Under the full perturbation high-precision dynamic model, each preferred candidate initial value in the preferred candidate transfer trajectory initial value set is used as the initial solution, and trajectory accurate propagation and multiple target iterative correction are carried out respectively to obtain a high-precision corrected trajectory set; Convergence checks and engineering constraint verifications are performed on the high-precision corrected trajectory set. Unconverged corrected trajectories are removed to obtain the high-precision feasible trajectory set. The high-precision feasible trajectory set is comprehensively evaluated based on the total orbit change, flight time, terminal corridor margin, control execution margin, and mission preference. The high-precision feasible trajectory with the best comprehensive evaluation is selected as the final lunar-Earth transfer trajectory, and the corresponding control commands are output.

2. The method for jointly planning the departure time and orbit change pulse of a lunar-Earth transfer mission according to claim 1, characterized in that: Based on the task input set, a coarse-to-fine tiered candidate departure time sequence is generated within the departure time window. The periodic orbit is phase-parameterized, and the available departure phase points for each candidate departure time are quickly retrieved based on the phase parameters, including: A two-layer time sampling strategy of first global coarse search and then local fine search is adopted to improve the search accuracy under limited on-board computing power and generate a coarse-fine layer candidate departure time sequence. The periodic orbit is phase parameterized, and the set of available phase points for each candidate departure time is quickly retrieved based on the phase parameters. The initial orbital state of the spacecraft at the corresponding candidate departure time is output through the phase-state index table.

3. The method for jointly planning the departure time and orbit change pulse of a lunar-Earth transfer mission according to claim 1, characterized in that: The pulse templates include a single-pulse template T1, a two-pulse template T2, and a multi-pulse template T3.

4. The method for jointly planning the departure time and orbit change pulse of a lunar-Earth transfer mission according to claim 1, characterized in that: Based on a simplified dynamic model, each candidate combination is estimated to obtain a corresponding set of candidate lunar-Earth transfer trajectory schemes, including: The continuous re-entry or return corridor constraints are transformed into a finite number of discrete terminal sample points. For each terminal sample point, a standard terminal state that satisfies the corridor constraints is constructed. Under the simplified dynamic model, a boundary value problem is constructed from the starting state to the terminal position, and the Lambert solution or the boundary value solution is used as the initial transition velocity. The departure trajectory velocity is subtracted from the initial transfer velocity obtained by the solution to obtain the departure pulse; the selection of the pulse template determines whether to set an intermediate correction point and the number of correction points. For pulse templates with intermediate correction points, when the spacecraft propagates to the corresponding correction point, the correction pulse is recalculated based on the correction point state deviation and the terminal re-entry or return corridor constraints. By propagating a lunar-Earth transfer trajectory with initial pulse values ​​in a simplified dynamic model, a set of corresponding candidate lunar-Earth transfer trajectory schemes is obtained. Obtain the intersection status of the lunar-Earth transfer orbit and the reentry interface (EI), and extract the indices at the reentry interface (EI) from the propagation results.

5. The method for jointly planning the departure time and orbit change pulse for a lunar-Earth transfer mission according to claim 4, characterized in that: Under the simplified dynamic model, a boundary value problem is constructed from the starting state to the terminal position, using the Lambert solution or the boundary value solution as the initial transition velocity, including: The two-body model, the conical splicing model, or the simplified CRTBP model are selected as the simplified dynamic model; For each set of starting states to terminal positions, the Lambert algorithm is used to solve the lunar-Earth transfer trajectory, and the initial transfer velocity that satisfies the position boundary conditions is obtained.

6. The method for jointly planning the departure time and orbit change pulse for a lunar-Earth transfer mission according to claim 1, characterized in that: The constraint checks performed on all candidate lunar-Earth transfer trajectory schemes include propulsion constraints, flight time constraints, geometry and mission constraints, and terminal reentry or return corridor constraints. Corridor margin and corridor default are calculated for all feasible candidate lunar-Earth transfer trajectory schemes with corridor feasibility.

7. The method for jointly planning the departure time and orbit change pulse of a lunar-Earth transfer mission according to claim 1, characterized in that: A comprehensive evaluation and screening of feasible candidate lunar-Earth transfer trajectory schemes is conducted using multiple indicators to obtain a preferred set of candidate transfer trajectories. A localized intensive search is then performed on this preferred set of candidate transfer trajectories to obtain an initial set of preferred candidate transfer trajectories, including: All feasible candidate lunar-Earth transfer trajectory schemes are comprehensively scored according to propulsion consumption, flight time, corridor margin, and degree of satisfaction of geometric and mission constraints. The comprehensive scores are sorted from smallest to largest, and the top N feasible candidate lunar-Earth transfer trajectory schemes are selected as the preferred candidate transfer trajectory set. Taking each preferred candidate transfer trajectory scheme as the center, local densification is performed in the neighborhood of its corresponding departure time, phase point or related discrete parameters to generate a new neighborhood candidate transfer trajectory combination and perform trajectory propagation and constraint checks to obtain the initial value set of preferred candidate transfer trajectories.

8. The method for jointly planning the departure time and orbit change pulse of a lunar-Earth transfer mission according to claim 1, characterized in that: Under the fully perturbation high-precision dynamic model, each preferred candidate initial value in the preferred candidate transfer trajectory initial value set is used as the initial solution. Trajectory accurate propagation and multiple target-hitting iterative correction are then performed to obtain a high-precision corrected trajectory set, including: Establish a high-precision dynamic model consistent with the real flight environment, and simultaneously incorporate all key perturbation terms during trajectory integration; Using the pulse application time as the segment node, the entire lunar-Earth transfer trajectory is divided into multiple uncontrolled free-flight arc segments. The pulse is applied only at the nodes between the arc segments. The initial state of each arc segment is obtained by superimposing the end state of the previous arc segment with the pulse velocity increment of the current node. The pulse magnitude, pulse direction, and pulse application time are used as optimization variables to be corrected. Reentry corridor constraints, flight time constraints, propulsion constraints, and geometric and mission constraints are used as constraints to conduct iterative corrections, including: Substitute the current initial pulse value into the high-precision dynamic model and integrate it to the reentry interface to calculate the residual between the actual state of the terminal and the target state of the reentry corridor. A linear relationship between the changes in pulse variables and the changes in terminal residuals is established using the state transition matrix. Under the premise of satisfying the upper limit of a single pulse, the upper limit of the total orbit change, and the pulse time window constraints, the optimal pulse correction amount is solved by least squares or constraint optimization methods. Substitute the corrected pulse parameters back into the high-precision dynamic model integral to update the terminal residual; Repeat the iterative process until the terminal residual is less than the preset convergence threshold, the terminal state is completely within the re-entry corridor, and all constraints are satisfied.

9. A joint planning system for the departure time and orbit change pulse of a lunar-Earth transfer mission, characterized in that, include: The data acquisition module is used to acquire the mission input set for the lunar-Earth transfer mission. The mission input set includes the spacecraft's initial orbit, departure time window, propulsion constraints, flight time constraints, geometric and mission constraints, and terminal reentry or return corridor constraints. The departure phase point retrieval module is used to generate a coarse-to-fine tiered candidate departure time sequence within the departure time window based on the task input set, perform phase parameterization on the periodic orbit, and quickly retrieve the set of available departure phase points for each candidate departure time based on the phase parameters. The template selection module is used to select one or more pulse templates according to task requirements, and to specify the feasible range of pulse time for each pulse template. It combines the set of available start phase points at each candidate start time with the selected pulse template as a candidate combination. The initial trajectory estimation module is used to estimate each candidate combination based on a simplified dynamic model, thereby obtaining a set of corresponding candidate lunar-Earth transfer trajectory schemes; The feasible trajectory generation module is used to perform constraint checks on all candidate lunar-Earth transfer trajectory schemes, eliminate candidate lunar-Earth transfer trajectory schemes that do not meet the constraints, and obtain a set of feasible candidate lunar-Earth transfer trajectory schemes. The trajectory initial value generation module uses multiple indicators to comprehensively score and filter the set of feasible candidate lunar-Earth transfer trajectory schemes to obtain the preferred candidate transfer orbit set, and performs local encrypted search on the preferred candidate transfer orbit set to obtain the preferred candidate transfer trajectory initial value set. The trajectory correction module is used to perform precise trajectory propagation and multiple target shooting iteration correction under the full perturbation high-precision dynamic model, using each preferred candidate initial value in the preferred candidate transfer trajectory initial value set as the initial solution, to obtain a high-precision corrected trajectory set. The control command output module performs convergence checks and engineering constraint verifications on the high-precision corrected trajectory set, eliminates non-converged corrected trajectories, and obtains a high-precision feasible trajectory set. The high-precision feasible trajectory set is comprehensively evaluated based on the total orbit change, flight time, terminal corridor margin, control execution margin, and mission preference. The high-precision feasible trajectory with the best comprehensive evaluation is selected as the final lunar-Earth transfer trajectory, and the corresponding control commands are output.

10. An electronic device, comprising: At least one processor; as well as A memory that is communicatively connected to the at least one processor; The memory stores a computer program that can be executed by the at least one processor, characterized in that, when the computer program is executed by the at least one processor, it implements a method for jointly planning the departure time and orbit change pulse of a lunar-Earth transfer mission as described in any one of claims 1-8.

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