Orbit maneuver planning method and system based on spacecraft orbital transfer primitive

Abstract

The invention relates to the field of spacecraft orbital maneuver planning, and provides an orbital maneuver planning method based on spacecraft orbital transfer primitives by performing feature analysis and recognition on the spacecraft orbital transfer primitives, orbital transfer tasks are integrated into independent orbital transfer primitives, and the primitives can be randomly combined to realize specific space tasks. Therefore, the implementation complexity of orbit maneuvering planning is reduced, and the efficiency is improved. The planning method comprises the following steps: S1, in a complex orbit maneuver planning task, dividing a flight process into a plurality of basic sections, and combining the plurality of basic sections to form a complete orbit maneuver planning task; s2, each basic section is divided into an initial condition primitive, a stop condition primitive and an orbit primitive, and orbit maneuvering task planning is completed by freely combining all the primitives. The method is suitable for maneuvering planning tasks in complex orbits.

Description

Technical Field

[0001] The present invention relates to the field of spacecraft orbit maneuver planning, and particularly to an orbit maneuver planning method based on spacecraft orbit change elements. Background Art

[0002] Orbit maneuver planning is one of the core technologies in space missions. Its importance is not only reflected in ensuring the mission success rate, but also deeply affects the lifespan of spacecraft, mission costs, and the sustainable utilization of future space resources.

[0003] However, orbit maneuver planning is a very complex task, involving complex dynamic environments, optimization requirements of multiple objectives and constraints, real-time computing requirements, the impact of uncertainties, and the challenges of collaborative missions. Each aspect has its own unique difficulties, and it is necessary to comprehensively apply technologies such as mathematical modeling, optimization algorithms, control theory, and high-performance computing to solve them.

[0004] In summary, orbit maneuver planning has a decisive impact on the success or failure of space missions, but realizing orbit maneuver planning requires knowledge in multiple fields and is relatively complex to achieve. Summary of the Invention

[0005] The purpose of the present invention is to propose an orbit maneuver planning method based on spacecraft orbit change elements by analyzing and identifying the characteristics of spacecraft orbit change elements. By integrating orbit change tasks into independent orbit change elements, these elements can be arbitrarily combined to achieve specific space missions, thereby reducing the implementation complexity of orbit maneuver planning and improving efficiency.

[0006] To achieve the above purpose, the present invention provides the following technical solutions:

[0007] The present invention proposes an orbit maneuver planning method based on spacecraft orbit change elements by analyzing and identifying the characteristics of spacecraft orbit change elements. The planning method includes the following steps:

[0008] Step S1: In a complex orbit maneuver planning task, divide the flight process into multiple basic segments, and the multiple basic segments are combined to form a complete orbit maneuver planning task;

[0009] Step S2: Divide each basic segment into an initial condition element, a stop condition element, and an orbit element, and complete the orbit maneuver task planning by freely combining these elements.

[0010] Further, the above initial condition element is used to set a reference spacecraft, set the initial phase position and velocity, and set the initial epoch time when the scene starts.

[0011] Further, the above stop condition element is used to stop orbit extrapolation according to the stop condition input by the user.

[0012] Furthermore, the above-mentioned stopping conditions include stopping according to time, stopping according to duration, or stopping at a specific epoch moment, stopping according to the shape of the chasing star's orbit, stopping at the apogee or perigee, stopping according to the position of the chasing star, stopping at a specific distance from the central target, and stopping when crossing the orbital plane.

[0013] Furthermore, the above-mentioned orbital primitives include elliptical fly-around primitive, circular fly-around primitive, spiral fly-around primitive, fixed-point hovering primitive, water-drop hovering primitive, VBar following primitive, straight-line jump primitive, and curved-line jump primitive.

[0014] Furthermore, the primitive feature of the above-mentioned elliptical fly-around primitive is that the projection of the relative motion of the accompanied spacecraft with respect to the reference spacecraft in the orbital plane is an elliptical trajectory with a ratio of the major axis to the minor axis of 2:1;

[0015] The primitive feature of the circular fly-around primitive is that multiple maneuvers are performed, and the distances of the points on the relative motion trajectory from the central target are equal;

[0016] The primitive feature of the spiral fly-around primitive is that from the motions in the x, y, and z directions, an elliptical trajectory drifting along the x direction can be obtained;

[0017] The primitive feature of the fixed-point hovering primitive is that the observing star is located at any specific point and maintains a fixed relative distance from the target star;

[0018] The primitive feature of the water-drop hovering primitive is the configuration naturally formed by the relative motion trajectory in the satellite drifting state, which has symmetry and revisit characteristics;

[0019] The primitive feature of the VBar following primitive is that the observing star and the target star maintain a fixed front-back distance in the x direction;

[0020] The primitive feature of the straight-line jump primitive is that the transfer trajectory from the starting point to the ending point is a straight line and multiple maneuvers are required;

[0021] The primitive feature of the curved-line jump primitive is that the transfer trajectory from the starting point to the ending point is a curve and only two maneuvers at the start and end are required.

[0022] Furthermore, the above-mentioned step S2 is specifically as follows:

[0023] Step S21: Set the initial condition primitive: Set the reference spacecraft, the VVLH coordinate system centered on the reference spacecraft, set the initial relative position and velocity, and the initial epoch moment;

[0024] Step S22: Set the stop condition primitive: Select the stop condition as duration stop and set the duration so that, based on the initial conditions, when this primitive ends after maintaining the duration, the relative position and velocity at this moment are output to the elliptical hovering primitive;

[0025] Step S23: Set the elliptical hovering primitive: Set the semi-major axis of the hover and the transfer time, and determine whether the conditions for elliptical hovering are met based on the output of the stop condition primitive. If they are met, perform the hover; if not, perform two maneuvers until the conditions are met;

[0026] Step S24: Set the stop condition primitive again: Select the stop condition as duration stop and set the duration so that our spacecraft exits the hover from any position on the ellipse. When this primitive ends, the relative position and velocity at this moment are output to the circular hovering primitive;

[0027] Step S25: Set the circular hovering primitive: Set the circular hovering radius, the number of navigation points, and the transfer time respectively so that our spacecraft completes the circular hover according to the output of the stop condition primitive;

[0028] Step S26: Set the stop condition primitive once again: Select the stop condition as duration stop and set the duration so that our spacecraft completes a full circle of hover according to the set conditions.

[0029] The orbit maneuver planning method based on spacecraft orbit transfer primitives described in the present invention can be fully implemented by computer software. Therefore, correspondingly, the present invention also provides an orbit maneuver planning system based on spacecraft orbit transfer primitives. The system includes a storage device, and the storage device is used to execute the orbit maneuver planning method and steps proposed above.

[0030] The present invention also provides a computer-readable storage medium. A computer program is stored on the computer-readable storage medium, and when the computer program is run by a processor, the processor executes the orbit maneuver planning method based on spacecraft orbit transfer primitives described in any one of the above.

[0031] The present invention also provides a computer device. The device includes a memory and a processor. A computer program is stored in the memory, and when the processor runs the computer program stored in the memory, the processor executes the orbit maneuver planning method based on spacecraft orbit transfer primitives described in any one of the above.

[0032] The beneficial effects of the present invention are:

[0033] 1. By analyzing and identifying the characteristics of the spacecraft orbit transfer primitive, the present invention proposes an orbit maneuver planning method based on the spacecraft orbit transfer primitive. By integrating the orbit transfer tasks into individual independent orbit transfer primitives, these primitives can be arbitrarily combined to achieve specific space missions, thereby reducing the implementation complexity of orbit maneuver planning and improving efficiency.

[0034] Furthermore, the free combination of the orbit transfer primitives proposed by the present invention has the advantages of high-efficiency integration and flexible expansion. Its positive effects are not only reflected in shortening the development cycle and improving mission reliability, but also in reducing the complexity for space mission designers when facing orbit maneuver planning, enabling them not to deal with code details such as orbit calculation and optimization algorithms, and improving the efficiency of space mission design.

[0035] The present invention is applicable to maneuver planning tasks in complex orbits. BRIEF DESCRIPTION OF THE DRAWINGS

[0036] In order to more clearly illustrate the specific embodiments of the present invention or the technical solutions in the prior art, the following will briefly introduce the drawings required for the description of the specific embodiments or the prior art. Obviously, the drawings in the following description are some embodiments of the present invention. For those of ordinary skill in the art, without creative efforts, other drawings can also be obtained based on these drawings.

[0037] Figure 1 is the attribute setting of the initial state primitive proposed by the present invention;

[0038] Figure 2 are the selectable stop conditions of the stop condition primitive proposed by the present invention;

[0039] Figure 3 is the subdivision of the orbit primitive proposed by the present invention;

[0040] Figure 4 is the detailed flowchart of a specific embodiment described by the present invention;

[0041] Figure 5 is the schematic diagram of elliptical fly-around described by the present invention;

[0042] Figure 6 is the schematic diagram of circular fly-around described by the present invention. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0043] The following further elaborates on the specific implementation manners of the present invention in conjunction with the accompanying drawings. The following implementation manners will help those skilled in the art further understand the present invention, but do not limit the present invention in any form. It should be noted that for those of ordinary skill in the art, without departing from the concept of the present invention, several changes and improvements can still be made, and these all fall within the protection scope of the present invention.

[0044] It should be understood that when used in this specification and the appended claims, the term "comprising" indicates the presence of the described features, wholes, steps, operations, elements, and / or components, but does not exclude the presence or addition of one or more other features, wholes, steps, operations, elements, components, and / or their combinations.

[0045] Embodiment 1. In this embodiment, by analyzing and identifying the characteristics of spacecraft orbit transfer elements, an orbit maneuver planning method based on spacecraft orbit transfer elements is proposed. The planning method includes the following steps:

[0046] Step S1: In a complex orbit maneuver planning task, the flight process is divided into multiple basic segments, and multiple basic segments are combined to form a complete orbit maneuver planning task.

[0047] Step S2: Each basic segment is divided into an initial condition element, a stop condition element, and an orbit element. By freely combining each element, the orbit maneuver task planning is completed.

[0048] The orbit maneuver planning method based on spacecraft orbit transfer elements proposed in this embodiment integrates the orbit transfer tasks into individual independent orbit transfer elements. These elements can be arbitrarily combined to achieve specific space missions, thereby reducing the implementation complexity of orbit maneuver planning and improving efficiency.

[0049] Embodiment 2. Refer to Figure 1 and Figure 2 to illustrate this embodiment. This embodiment specifically describes the initial condition element and the stop condition element proposed in the above Embodiment 1.

[0050] Initial condition element: As Figure 1 shown, it is used to set a reference spacecraft, set the initial phase position and velocity, and set the initial epoch time when the scenario starts.

[0051] Stop condition element: It is used to stop the orbit extrapolation according to the stop condition input by the user. The stop conditions set in this embodiment are: As Figure 2 shown, stop according to time, stop according to the duration, or stop at a specific epoch time; stop according to the orbit shape of the chasing star, stop at the apogee or perigee; stop according to the position of the chasing star, stop at a specific distance from the central target, or stop when crossing the orbital plane.

[0052] Embodiment 3. Refer to Figure 3 To describe this embodiment, this embodiment specifically describes the orbital element proposed in Embodiment 1 above;

[0053] As Figure 3 shown, the orbital elements are subdivided according to different space missions into: elliptical fly-around element, circular fly-around element, spiral fly-around element, fixed-point hover element, water-drop hover element, VBar following element, linear jump element, and curvilinear jump element. Each element can be freely combined to form the flight process of the orbital maneuver planning task in a "building block" manner.

[0054] Among them, the elliptical fly-around element is specifically:

[0055] The relative motion of a circular orbit at close range without external forces can be described by the following CW equation:

[0056]

[0057] Solving the above equation, the solution of the CW equation is:

[0058]

[0059] Taking the derivative of the position component expression with respect to time, the relative velocity expression can be obtained:

[0060]

[0061] When z 00 = 0, and the other two initial conditions are satisfied then the elliptical motion becomes an elliptical motion around the origin of the coordinate system, that is, around the target spacecraft as the center.

[0062] Element feature: The projection of the motion of the accompanying spacecraft relative to the reference spacecraft in the orbital plane is an elliptical trajectory with a ratio of the major axis to the minor axis of 2:1.

[0063] The circular fly-around element is specifically:

[0064] Element feature: Multiple maneuvers are performed, and the distances from the points on the relative motion trajectory to the central target are equal.

[0065] The spiral fly-around element is specifically:

[0066] Spiral fly-around is a form of fly-around in a different plane. Compared with natural elliptical fly-around, spiral fly-around is non-periodic. According to the C-W equation, if the initial state is as follows:

[0067]

[0068] Then the position equation can be obtained as:

[0069]

[0070] From the motion equations in the y and z directions, the motion equation in the oyz plane can be obtained:

[0071]

[0072] Elementary characteristics: From the motion equation in the x direction, it can be seen that there are non-periodic terms in the motion in the x direction, which is the reason for the non-periodicity of the natural spiral hovering trajectory. The motions in the y and z directions form an ellipse equation, forming an elliptical trajectory. From the motions in the three directions of x, y, and z, an elliptical trajectory drifting along the x direction can be obtained, that is, the natural spiral hovering trajectory.

[0073] The specific form of VBar following the element is:

[0074] When x = x0 ≠ 0 and the rest of the initial values are all zero, that is It becomes a fixed point (x0, 0).

[0075]

[0076] Elementary characteristics: The observation star and the target star maintain a fixed front-back distance in the x direction.

[0077] The specific form of the fixed-point hovering element is:

[0078] Elementary characteristics: The observation star is located at any specific point and maintains a fixed relative distance from the target star.

[0079] The specific form of the water-drop hovering element is:

[0080] Elementary characteristics: "Water-drop" hovering is a type of configuration naturally formed by the relative motion trajectory in the satellite drifting state. Its shape is like a water drop and has obvious symmetry and revisit characteristics

[0081] The specific form of the straight-line jump element is:

[0082] Elementary characteristics: The transfer trajectory from the starting point to the ending point is a straight line and multiple maneuvers are required.

[0083] The specific form of the curve jump element is:

[0084] Elementary characteristics: The transfer trajectory from the starting point to the ending point is a curve and only two maneuvers at the start and the end are required.

[0085] Based on the above specific descriptions of the initial condition element, the stop condition element, and the orbit element in the embodiments, step S2 can be illustrated by examples;

[0086] Step S21: Set the initial condition primitive: Set the reference spacecraft, and set the VVLH coordinate system centered on the reference spacecraft, and set the initial relative position and velocity and the initial epoch time;

[0087] Step S22: Set the stop condition primitive: Select the stop condition as duration stop and set the duration, so that the duration is maintained based on the initial conditions. When this primitive ends, the relative position and velocity at this moment are output to the elliptical fly-around primitive;

[0088] Step S23: Set the elliptical fly-around primitive: Set the semi-major axis of the fly-around and the transfer time, and judge whether the conditions for elliptical fly-around are met according to the output of the stop condition primitive. If they are met, perform fly-around. If not, perform two maneuvers until they are met;

[0089] Step S24: Set the stop condition primitive again: Select the stop condition as duration stop and set the duration, so that our spacecraft exits the fly-around from any position on the ellipse. When this primitive ends, the relative position and velocity at this moment are output to the circular fly-around primitive;

[0090] Step S25: Set the circular fly-around primitive: Set the circular fly-around radius, the number of navigation points, and the transfer time respectively, so that our spacecraft completes the circular fly-around according to the output of the stop condition primitive;

[0091] Step S26: Set the stop condition primitive once again: Select the stop condition as duration stop and set the duration, and our spacecraft completes a full circle of fly-around according to the set conditions.

[0092] Embodiment 4. Refer to Figures 4 to 6 Describe this embodiment. This embodiment proposes a specific implementation method for elaborately explaining an orbit maneuver planning method based on spacecraft orbit-changing primitives proposed in Embodiment 1;

[0093] Objective: Realize the transition from elliptical fly-around to circular fly-around.

[0094] Scenario description: Our spacecraft is named "Fly-around Star", and the enemy spacecraft is named "Target Star". "Fly-around Star" first performs elliptical fly-around on "Target Star" and then connects to circular fly-around.

[0095] The specific steps are as Figure 4 shown:

[0096] Step1: Set the initial state primitive:

[0097] Set the reference spacecraft as: "Target Star", then all subsequent relative coordinate systems are the VVLH coordinate systems centered on "Target Star".

[0098] Set the initial relative position velocity as: X0 = [-15, 0, 0, 0, 0, 0]

[0099] Set the initial epoch time as: 2023 / 10 / 01_00:00:00 UTCG

[0100] Step2: Set the stop condition primitive:

[0101] Select the stop condition as "duration stop" and set the duration as: 600 s.

[0102] Detailed description: The meaning of this primitive is to keep the "fly-by satellite" for 600 s under the determined initial state, and then perform subsequent actions. When this primitive ends, the relative position velocity at this moment will be output to the next primitive.

[0103] Step3: Set the elliptical fly-by primitive:

[0104] Set the semi-major axis of the fly-by as: 2 km.

[0105] Set the transfer time as: 8608 s.

[0106] Detailed description: From the output of the previous "stop condition primitive", first judge whether the condition for elliptical fly-by is met In this embodiment, it does not meet the condition. Therefore, it is necessary to perform two maneuvers until it is met, and the interval time between the two maneuvers is the input parameter "transfer time". The schematic diagram of the elliptical fly-by is as Figure 5 shown.

[0107] Step4: Set the stop condition primitive:

[0108] Select the stop condition as "duration stop" and set the duration as: 7173 s.

[0109] Detailed description: The meaning of this step is that when the "fly-by satellite" performs elliptical fly-by relative to the "target satellite" and reaches the set duration (7173 s), the fly-by will stop, so that the "fly-by satellite" can exit the fly-by from any position on the ellipse, and output the relative position velocity at this moment. In this embodiment, after flying for 7173 s, the "fly-by satellite" runs to the long semi-axis point in front of the "target satellite" ( Figure 5 the D point in

[0110] Step5: Set the circular fly-by primitive:

[0111] Set the radius of the circular fly-by as: 4 km.

[0112] Set the number of navigation points as: 6.

[0113] Set the transfer time as: 0 s.

[0114] Detailed description: Approximating a circle with a regular N-sided polygon, which is a regular hexagon in this embodiment. Taking the apogee point of the elliptical fly-around withdrawal as the first navigation point, applying a pulse to enter the second navigation point, and continuing in sequence until 6 pulses are completed to return to the starting point, realizing a circular fly-around with a radius of 4 km, as Figure 6 shown.

[0115] Step6: Set the stop condition primitive:

[0116] Select the stop condition as "duration stop", and set the duration to: 5739 s.

[0117] Detailed description: The significance of this step is that when the "fly-around satellite" performs a circular fly-around relative to the "target satellite" and reaches the set duration (5739 s), the fly-around will stop, so that the "fly-around satellite" can exit the fly-around from any position on the circle and output the relative position and velocity at that moment. In this embodiment, after flying around for 5739 s, the "fly-around satellite" completes a full circle of fly-around.

[0118] Step7: Set an arbitrary orbit primitive:

[0119] If other fly-around tasks need to be executed, a certain orbit primitive can be directly added and the input can be set.

[0120] Embodiment 5. A trajectory maneuver planning method based on spacecraft orbit change primitives described in any one of the above Embodiments 1 to 3 can be fully implemented by computer software. Therefore, correspondingly, this embodiment proposes a trajectory maneuver planning system based on spacecraft orbit change primitives, and the system includes:

[0121] A storage device for dividing the flight process into multiple basic segments in complex trajectory maneuver planning tasks, and combining multiple basic segments to form a complete trajectory maneuver planning task;

[0122] A storage device for dividing each basic segment into an initial condition primitive, a stop condition primitive, and an orbit primitive, and completing the trajectory maneuver task planning by freely combining each primitive.

[0123] Embodiment 6. This embodiment provides a computer-readable storage medium, on which a computer program is stored. When the computer program is run by a processor, it executes the trajectory maneuver planning method based on spacecraft orbit change primitives described in any one of the above embodiments.

[0124] Embodiment 7. This embodiment provides a computer device, which includes a memory and a processor. A computer program is stored in the memory. When the processor runs the computer program stored in the memory, the processor executes the trajectory maneuver planning method based on spacecraft orbit change primitives described in any one of the above embodiments.

[0125] A computer device provided by this embodiment. The hardware devices in this part are of general models and are not shown in the form of drawings. The system includes a processor and a memory. The processor and the memory can be connected through a bus or other means. The memory, as a non-transitory computer-readable storage medium, can be used to store non-transitory software programs, non-transitory computer-executable programs, and modules, as well as corresponding program instructions / modules. The processor executes various functional applications and data processing of the processor by running the non-transitory software programs, instructions, and modules stored in the memory, so as to implement the orbital maneuver planning method and steps based on the spacecraft orbit change primitive in the above method embodiments.

[0126] The above is only the embodiment of the present invention and does not limit the present invention. For those skilled in the art, the present invention can have various changes and modifications. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present invention shall be included within the scope of the claims of the present invention.

Claims

1. An orbital maneuver planning method based on spacecraft orbit-changing primitive elements, characterized in that The method is as follows: S1: Divide the flight process into multiple basic segments, and multiple basic segments are combined to form a complete orbital maneuver planning task; S2: Divide each basic segment into an initial condition primitive, a stop condition primitive, and an orbit primitive. By freely combining each primitive, complete the orbital maneuver task planning.

2. The orbital maneuver planning method based on the spacecraft orbit transfer primitive according to claim 1, wherein The initial condition primitive is used to set the reference spacecraft, set the initial phase position and velocity, and set the initial epoch time when the scenario starts.

3. The orbital maneuver planning method based on spacecraft orbit-changing elements according to claim 1, wherein The stop condition primitive is used to stop the orbit extrapolation according to the stop condition input by the user.

4. The orbital maneuver planning method based on the spacecraft orbit transfer primitive according to claim 1, wherein The stop conditions include stopping according to time, stopping according to duration, or stopping at a specific epoch time, stopping according to the chasing star's orbit shape, stopping at the apogee or perigee, stopping according to the position of the chasing star, stopping at a specific distance from the central target, and stopping when crossing the orbital plane.

5. The orbital maneuver planning method based on the spacecraft orbit transfer primitive according to claim 1, characterized in that, The orbit primitives include an elliptical fly-around primitive, a circular fly-around primitive, a spiral fly-around primitive, a fixed-point hover primitive, a water-drop hover primitive, a VBar following primitive, a straight-line jump primitive, and a curve jump primitive.

6. The orbital maneuver planning method based on the spacecraft orbit transfer primitive according to claim 5, wherein The primitive feature of the elliptical fly-around primitive is that the projection of the relative motion of the accompanied spacecraft with respect to the reference spacecraft in the orbital plane is an elliptical trajectory with a ratio of the major axis to the minor axis of 2:1; The primitive feature of the circular fly-around primitive is to perform multiple maneuvers, and the distances of the points on the relative motion trajectory from the central target are equal; The primitive feature of the spiral fly-around primitive is that from the motions in the x, y, and z directions, an elliptical trajectory drifting along the x direction can be obtained; The primitive feature of the fixed-point hover primitive is that the observing star is located at any specific point and keeps a fixed relative distance from the target star; The primitive feature of the water-drop hover primitive is the configuration naturally formed by the relative motion trajectory in the satellite drifting state, with symmetry and revisit characteristics; The primitive feature of the VBar following primitive is that the observing star and the target star keep a fixed front-back distance in the x direction; The primitive feature of the straight-line jump primitive is that the transfer trajectory from the starting point to the ending point is a straight line, and multiple maneuvers are required; The primitive feature of the curve jump primitive is that the transfer trajectory from the starting point to the ending point is a curve, and only two maneuvers are required at the start and the end; 7. The orbital maneuver planning method based on the spacecraft orbit transfer primitive according to claim 6, wherein Specifically, S2 is as follows: S21: Set the initial condition primitive: Set the reference spacecraft, and set the VVLH coordinate system centered on the reference spacecraft, set the initial relative position and velocity, and the initial epoch time; S22: Set the stop condition primitive: Select the stop condition as stopping according to duration and set the duration. Based on the initial conditions, maintain the duration. When this primitive ends, output the relative position and velocity at this moment to the elliptical fly-around primitive; S23: Set the elliptical fly-around primitive: Set the fly-around semi-major axis and the transfer time. Judge whether the conditions for performing elliptical fly-around are met through the output of the stop condition primitive. If met, perform fly-around. If not met, perform two maneuvers until met; S24: Set the stop condition primitive again: Select the stop condition as stopping according to duration and set the duration. The spacecraft exits the fly-around from any position on the ellipse. When this primitive ends, output the relative position and velocity at this moment to the circular fly-around primitive; S25: Setting the circle fly primitive: setting the circle fly radius, the number of navigation points and the transfer time respectively. The spacecraft completes the circle fly according to the output of the stop condition primitive; S26: Set the stop condition primitive again: select the stop condition as duration stop and set the duration. The spacecraft completes a complete orbit according to the set conditions.

8. An orbit maneuver planning system based on spacecraft orbit change elements, characterized in that, The system includes a storage device, which is used to execute the method and steps described in claim 1.

9. A computer-readable storage medium, characterized in that, The computer-readable storage medium stores a computer program, which, when executed by a processor, executes the orbital maneuvering planning method based on spacecraft trajectory change primitives as described in any one of claims 1 to 7.

10. A computer device, characterized in that, The device includes a memory and a processor, wherein a computer program is stored in the memory. When the processor runs the computer program stored in the memory, the processor executes the orbital maneuvering planning method based on the spacecraft orbit change primitive as described in any one of claims 1 to 7.

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