Service spacecraft multi-orbit maneuvering strategy approaching observation method based on J2 perturbation compensation

Abstract

The invention discloses a service spacecraft multi-orbit maneuvering strategy approaching observation method based on J2 perturbation compensation, solves the problem of how to quickly improve orbit maneuvering accuracy, and belongs to the field of service spacecraft multi-orbit maneuvering strategy approaching observation. Comprising the following steps: constructing a motion equation under J2 item perturbation compensation under a target orbit coordinate system; constructing a service spacecraft multi-orbit maneuvering strategy model through the relative motion propagation relational expression and the motion equation; inputting the termination moment position conforming to the termination moment distance constraint into the model to output a single speed increment, and reserving the single speed increment conforming to the termination moment illumination angle constraint and the single speed increment constraint at the same time to obtain a speed increment solution set; constructing a circular region which takes the space target as a center and has a radius meeting a termination moment distance constraint; when the service spacecraft arrives at the circular area at a specified moment, marking the task successfully, and obtaining a service spacecraft multi-orbit maneuvering strategy for approaching observation; according to the invention, accurate and rapid track maneuvering is realized.

Description

Technical Field

[0001] The present invention belongs to the technical field of approach observation of a service spacecraft multi-orbit maneuvering strategy, and relates to an approach observation method of a service spacecraft multi-orbit maneuvering strategy based on J2 perturbation compensation. Background Art

[0002] The actual needs of space missions have led to the study of orbital maneuvering strategies for on-orbit servicing and close observation problems, especially in dealing with space targets, such as space debris and repairing failed spacecraft. With the continuous advancement of space technology and the increase in the frequency of space activities, the complexity of the space environment has also increased, making the protection of space assets and the handling of space targets a common challenge faced by the aerospace field. When performing these tasks, the service spacecraft needs to be able to perform orbital maneuvers accurately to complete various close observation tasks. Traditional on-orbit servicing problem research is usually based on the continuous thrust model and uses methods such as differential games to find the optimal maneuvering strategy. Although these methods are quite mature in theory, in practical applications, since the service spacecraft mainly relies on pulse thrust to perform orbital maneuvers, there is a difference between the continuous thrust model and the actual operation, which limits its application effect in actual space missions. In addition, considering the impact of J2 perturbations on the orbit, this perturbation must also be taken into account in orbit design and maneuvering strategies to improve the accuracy and reliability of the mission.

[0003] In the field of spacecraft orbital maneuvering strategies, although theoretical methods such as differential games provide a theoretical framework for on-orbit servicing problems, these methods are mostly based on continuous thrust models, while actual service spacecraft orbital maneuvers rely more on pulse thrust. The difference between this model and actual operation may cause existing methods to face certain limitations when dealing with pulse thrust spacecraft on-orbit servicing problems. In addition, existing studies mostly focus on finding the optimal fuel solution under the constraint of approach distance, and rarely consider the impact of orbit change strategy selection on the final relative position of the service spacecraft, which may limit the in-depth understanding of the relative motion law of the service spacecraft and the comprehensive optimization of the maneuvering strategy. At the same time, most existing studies are based on the CW equation, and only explore the maneuvering strategy of the service spacecraft under the two-body dynamics model. The impact of J2 perturbation, as the most important perturbation force of the non-spherical earth, has a decisive influence on the on-orbit operation of the service spacecraft. It not only causes long-term changes in orbital elements, but also causes nutation and precession of the orbital plane. These effects are crucial for precise orbit maintenance, formation flight of service spacecraft, and mission execution. Therefore, it is necessary to fully consider the impact of J2 perturbation in the orbit design and control of the service spacecraft to ensure that the service spacecraft completes its mission safely and effectively. Summary of the invention

[0004] In order to solve the technical problem of how to improve the accuracy of the service spacecraft orbit design and maneuvering strategy, the present invention provides a service spacecraft multi-orbit maneuvering strategy approach observation method based on J2 perturbation compensation, and provides a service spacecraft with an orbit maneuvering strategy from the current state to the expected approach relative state through mathematical modeling and optimization technology. This strategy will determine the range of velocity increments required by the service spacecraft in the radial, lateral and normal directions to achieve accurate orbital maneuvers;

[0005] The goal of this technology is to develop a more flexible, fast and practical orbital maneuvering strategy based on the actual conditions of pulse thrust. This method aims to optimize the orbital maneuver of the service spacecraft and provide an approximate range for the maneuvering parameters that meet the mission constraints to improve the efficiency and accuracy of the algorithm. In addition, considering the impact of the J2 perturbation on the orbit, this technology establishes the equation of motion and relationship equations under the J2 perturbation to improve the accuracy of the orbital design and maneuvering strategy. Through this technology, the relative motion of the service spacecraft in the actual space environment can be more accurately understood and predicted, providing a solid scientific basis for the orbital design and mission planning of the service spacecraft.

[0006] The purpose of the present invention is specifically achieved through the following technical solutions:

[0007] The present invention discloses a service spacecraft multi-orbit maneuver strategy approach observation method based on J2 perturbation compensation, the method comprising:

[0008] Step 1: In the target orbit coordinate system, construct the motion equation of the service spacecraft relative to the space target under the J2 perturbation compensation;

[0009] Step 2: Based on the relative motion propagation relation, the relation equation between the terminal position of the service spacecraft in the target orbit coordinate system and the single velocity increment is derived through the motion equation, and the relation equation is used as the multi-orbit maneuver strategy model of the service spacecraft;

[0010] Step 3: Setting constraints that meet the task requirements, wherein the constraints consist of a termination time distance constraint, a termination time illumination angle constraint, and a single speed increment constraint;

[0011] Step 4: uniformly sample all the termination time positions that meet the termination time distance constraint, input all the sampling points into the service spacecraft multi-orbit maneuvering strategy model, output the single velocity increment required for the service spacecraft to reach the termination time position, retain the single velocity increment that meets both the termination time illumination angle constraint and the single velocity increment constraint, and obtain the velocity increment solution set;

[0012] Step 5: Construct a circular area with the space target as the center and a radius that satisfies the distance constraint at the termination time; use the circular area to display the relative distance between the service spacecraft and the space target corresponding to each single velocity increment in the velocity increment solution set; when the service spacecraft arrives at the circular area at the specified time, the marking mission is successful; the orbital maneuvering strategy of the service spacecraft is obtained from each successful single velocity increment of the marking mission to conduct close observation of the space target.

[0013] In step 1, the equation of motion is:

[0014]

[0015]

[0016]

[0017] Where [x, y, z] is the position vector of the service spacecraft after time t in the target orbital coordinate system, and t is the time from the initial time; represents the state vector of the service spacecraft at the initial time t0 of the mission in the target orbital coordinate system, where the state vector includes the position vector and the velocity vector, [x0, y0, z0] is the position vector, is the velocity vector; c is the perturbation coefficient, and n is the average orbital angular velocity of the service spacecraft.

[0018] In step 2, the relative motion propagation relationship is:

[0019]

[0020] In the formula, is the velocity vector of the service spacecraft at the initial time t0 of the mission in the target orbit coordinate system,

[0021] [ΔV x ,ΔV y ,ΔV z ] is the single velocity increment of the service spacecraft when the pulse increment is applied at the initial moment t0 of the mission, that is, the velocity increment obtained by the service spacecraft in the three directions x, y, and z of the target orbit coordinate system, [V x0 ,V y0 ,V z0 ] is the velocity of the service spacecraft before the pulse increment is applied.

[0022] In step 2, the relationship equation is:

[0023]

[0024]

[0025]

[0026] Where [M, N, U] is the end time position.

[0027] In step 3, the end time distance constraint is: r1<r pe <r2; where r pe is the relative distance between the service spacecraft and the space target, r1 is the shortest distance between the service spacecraft and the space target, and r2 is the farthest distance between the service spacecraft and the space target;

[0028] The illumination angle constraint at the termination time is: θ1<θ<θ2; where θ is the illumination angle, that is, the angle between the vector of the space target pointing to the sun and the vector of the space target pointing to the service spacecraft, θ1 is the upper limit of the illumination angle, and θ2 is the lower limit of the illumination angle;

[0029] The single velocity increment constraint is: In the formula, [ΔV x ,ΔV y ,ΔV z ] is the velocity increment obtained by the service spacecraft in the x, y, and z directions of the target orbit coordinate system, ΔV Represents the total velocity increment available for a single orbit change of the serviced spacecraft.

[0030] In step 4, the uniform sampling method includes:

[0031] The orbital parameter range of the service spacecraft is set, with the space target at the termination time as the origin and the azimuth, pitch angle and distance of the service spacecraft relative to the space target as the step size to ensure the uniformity of sampling.

[0032] In step 4, the single speed increment that satisfies both the illumination angle constraint at the termination time and the single speed increment constraint is retained, and the method for obtaining the speed increment solution set includes:

[0033] Eliminate the single speed increment that does not meet the illumination angle constraint at the end time;

[0034] Calculate the velocity increment ΔV obtained by each single velocity increment in the remaining single velocity increments in the three directions of the target orbit coordinate system x, y, and z x , ΔV y , ΔV z The sum of is the total speed increment, and the total speed increment exceeding the threshold is excluded to obtain the speed increment solution set.

[0035] Step five also includes the step of selecting a service spacecraft orbital maneuvering strategy that meets predefined standards as the best orbital maneuvering strategy for close-in observation of the space target.

[0036] In step five, the predefined criteria are defined based on mission requirements, including but not limited to selecting the service spacecraft orbital maneuvering strategy corresponding to the smallest velocity increment or the shortest transfer time among the service spacecraft orbital maneuvering strategies that mark the success of the mission as the optimal orbital maneuvering strategy for close observation of the space target.

[0037] In step 5, the circular area whose radius satisfies the distance constraint at the end time is:

[0038] A first circular area with a radius of r1 being the closest distance between the service spacecraft and the space target, and a second circular area with a radius of r2 being the farthest distance between the service spacecraft and the space target, the first circular area and the second circular area are concentric circles.

[0039] The beneficial effects of the present invention are:

[0040] 1. Efficient modeling and optimization of orbital transfer conditions:

[0041] The present invention constructs the motion equation and relational equation including the influence of J2 perturbation by comprehensively analyzing the current state (position and velocity) of the service spacecraft and the target orbital parameters to describe the orbital transfer of the space target. The relational equation not only takes into account the actual constraints such as the distance constraint at the termination time, the illumination angle constraint at the termination time and the single velocity increment constraint, but also specifically incorporates the J2 perturbation to ensure the accuracy and stability of the orbital transfer. By introducing the J2 perturbation, the model of the present invention can more accurately predict the dynamic behavior of the service spacecraft in the space environment, especially in the orbital service problem, which is crucial to the accuracy of the orbital transfer, improves the design accuracy of the orbital maneuvering strategy, and enables the service spacecraft to achieve more accurate orbital transfer in a complex space environment.

[0042] 2. Precise orbital maneuvering strategy based on impulse thrust and J2 perturbation:

[0043] The present invention deeply analyzes the existing models, and combines the actual conditions of pulse thrust and the influence of J2 perturbations to design a more accurate and practical orbital maneuvering strategy. Compared with traditional methods, the present invention not only solves the optimal control law or a single maneuvering strategy, but can calculate all feasible solutions of velocity increments that meet the conditions in a given scenario, while considering the influence of J2 perturbations. Through this method, the design space of orbital maneuvering strategies is expanded, and the flexibility and practicality of orbital maneuvers are improved, while ensuring the accuracy of orbital maneuvers under the influence of J2 perturbations. The strategy of the present invention can provide a variety of orbital maneuvering options that meet the conditions for the service spacecraft while considering the J2 perturbation, so that in actual operation, the most appropriate strategy can be flexibly selected according to the specific situation, thereby improving the success rate and efficiency of the mission.

[0044] 3. Efficient generation and application of feasible solution sets considering J2 perturbation:

[0045] The method of the present invention can generate a feasible solution set of velocity increments for all orbital maneuvering strategies that meet the conditions in a short time while considering the J2 perturbation. The velocity increment solution set can more accurately guide the service spacecraft to perform orbital maneuvers in a complex space environment, thereby improving the success rate and efficiency of the mission.

[0046] 4. Construction of orbital maneuvering strategy model for service spacecraft driven by pulse thrust:

[0047] The present invention proposes a service spacecraft orbital maneuvering strategy model, which accurately captures the motion characteristics of the service spacecraft in the space environment and realizes the simulation of its dynamic behavior. The establishment of this model provides theoretical support for the design and optimization of on-orbit service missions, and as a prediction tool, it enhances the scientificity and accuracy of mission planning. By introducing pulse thrust, the present invention expands the limitations of the traditional continuous thrust model and provides a more accurate mathematical framework for the complexity analysis and strategy formulation of on-orbit service missions.

[0048] 5. Enhanced flexibility and practicality of orbital maneuvering strategies:

[0049] The present invention obtains the orbital maneuvering strategy of the service spacecraft from the single velocity increment of the successful marking mission, and realizes a significant improvement in the mission success rate and the operational flexibility of the service spacecraft in response to the changing space mission requirements. Through optimized design, the strategy enables the service spacecraft to perform efficient maneuvers between different orbits, meeting the requirements of multi-orbit maneuvering missions. In addition, the strategy comprehensively considers time constraints and fuel consumption, provides an optimal solution for completing orbital maneuvers within the specified time, and ensures efficient use of fuel, thereby reducing operating costs while ensuring mission efficiency.

[0050] 6. Improved intelligence and foresight in mission planning:

[0051] This invention significantly improves the intelligence level of orbital maneuvering decision-making and the foresight of mission planning by generating velocity increment solutions that meet specific conditions. This intelligence and foresight enable mission planning to better adapt to changes in future mission requirements and improve the adaptability and flexibility of mission planning. By comprehensively analyzing all feasible solutions, this study provides a more comprehensive perspective for mission planning, thereby maximizing the effectiveness of mission execution in a complex and changing space environment. BRIEF DESCRIPTION OF THE DRAWINGS

[0052] The present invention is further described in detail below based on the accompanying drawings and embodiments.

[0053] Figure 1It is a schematic diagram of the target orbit coordinate system in the example of the present invention.

[0054] Figure 2 It is a two-dimensional graphical schematic diagram of the velocity increment solution set in the example of the present invention.

[0055] Figure 3 It is a three-dimensional graphical schematic diagram of the velocity increment solution set in the example of the present invention. DETAILED DESCRIPTION

[0056] An embodiment of the present invention provides a service spacecraft multi-orbit maneuver strategy approach observation method based on J2 perturbation compensation, the method comprising:

[0057] Step 1: In the target orbit coordinate system, construct the motion equation of the service spacecraft relative to the space target under the J2 perturbation compensation;

[0058] Step 2: Based on the relative motion propagation relation, the relation equation between the terminal position of the service spacecraft in the target orbit coordinate system and the single velocity increment is derived through the motion equation, and the relation equation is used as the multi-orbit maneuver strategy model of the service spacecraft;

[0059] Step 3: Setting constraints that meet the task requirements, wherein the constraints consist of a termination time distance constraint, a termination time illumination angle constraint, and a single speed increment constraint;

[0060] Step 4: uniformly sample all the termination time positions that meet the termination time distance constraint, input all the sampling points into the service spacecraft multi-orbit maneuvering strategy model, output the single velocity increment required for the service spacecraft to reach the termination time position, retain the single velocity increment that meets both the termination time illumination angle constraint and the single velocity increment constraint, and obtain the velocity increment solution set;

[0061] Step 5: Construct a circular area with the space target as the center and a radius that satisfies the distance constraint at the termination time; use the circular area to display the relative distance between the service spacecraft and the space target corresponding to each single velocity increment in the velocity increment solution set; when the service spacecraft arrives at the circular area at the specified time, the marking mission is successful; the orbital maneuvering strategy of the service spacecraft is obtained from each successful single velocity increment of the marking mission to conduct close observation of the space target.

[0062] In step 1, the equation of motion is:

[0063]

[0064]

[0065]

[0066] Where [x, y, z] is the position vector of the service spacecraft after time t in the target orbital coordinate system, and t is the time from the initial time; represents the state vector of the service spacecraft at the initial time t0 of the mission in the target orbital coordinate system, where the state vector includes the position vector and the velocity vector, [x0, y0, z0] is the position vector, is the velocity vector; c is the perturbation coefficient, i ref is the average orbital inclination of the space target, r ref is the average orbital radius of the space object, R e is the radius of the earth, J2 is the harmonic coefficient of the gravity field, and n is the average orbital angular velocity of the service spacecraft.

[0067] In step 2, the relative motion propagation relationship is:

[0068]

[0069] In the formula, is the velocity vector of the service spacecraft at the initial time t0 of the mission in the target orbit coordinate system,

[0070] [ΔV x ,ΔV y ,ΔV z ] is the single velocity increment of the service spacecraft when the pulse increment is applied at the initial moment t0 of the mission, that is, the velocity increment obtained by the service spacecraft in the three directions x, y, and z of the target orbit coordinate system, [V x0 ,V y0 ,V z0 ] is the velocity of the service spacecraft before the pulse increment is applied.

[0071] In step 2, the relationship equation is:

[0072]

[0073]

[0074]

[0075] Where [M, N, U] is the position at the end time. This relationship equation does not consider the maneuvering strategy under constraints. Given the initial and final state position vectors, the pulse increment required for track change can be obtained.

[0076] In step 3, in order to make the strategy more realistic and operational, constraints were added when solving the model. The constraints involved the proximity of the target point, the sun angle of the target point, and the maximum speed change allowed for a single maneuver. By introducing constraints, it is ensured that the maneuver strategy not only meets the control characteristics of the service spacecraft, but also meets the specific requirements of the mission objectives; among them, the termination time distance constraint is: r1<r pe <r2; where r pe is the relative distance between the service spacecraft and the space target, r1 is the shortest distance between the service spacecraft and the space target, and r2 is the farthest distance between the service spacecraft and the space target;

[0077] The illumination angle constraint at the termination time is: θ1<θ<θ2; where θ is the illumination angle, that is, the angle between the vector of the space target pointing to the sun and the vector of the space target pointing to the service spacecraft, θ1 is the upper limit of the illumination angle, and θ2 is the lower limit of the illumination angle;

[0078] The single velocity increment constraint is: In the formula, [ΔV x ,ΔV y ,ΔV z ] is the velocity increment obtained by the service spacecraft in the x, y, and z directions of the target orbit coordinate system, ΔV Represents the total velocity increment available for a single orbit change of the serviced spacecraft.

[0079] In step 4, the uniform sampling method includes:

[0080] The orbital parameter range of the service spacecraft is set, with the space target at the termination time as the origin and the azimuth, pitch angle and distance of the service spacecraft relative to the space target as the step size to ensure the uniformity of sampling.

[0081] In step 4, the single speed increment that satisfies both the illumination angle constraint at the termination time and the single speed increment constraint is retained, and the method for obtaining the speed increment solution set includes:

[0082] Eliminate the single speed increment that does not meet the illumination angle constraint at the end time;

[0083] Calculate the velocity increment ΔV obtained by each single velocity increment in the remaining single velocity increments in the three directions of the target orbit coordinate system x, y, and z x , ΔV y , ΔV z The sum of is the total speed increment, and the total speed increment exceeding the threshold is excluded to obtain the speed increment solution set.

[0084] Step five also includes the step of selecting a service spacecraft orbital maneuvering strategy that meets predefined standards as the best orbital maneuvering strategy for close-in observation of the space target.

[0085] In step five, the predefined criteria are defined based on mission requirements, including but not limited to selecting the service spacecraft orbital maneuvering strategy corresponding to the smallest velocity increment or the shortest transfer time among the service spacecraft orbital maneuvering strategies that mark the success of the mission as the optimal orbital maneuvering strategy for close observation of the space target.

[0086] In step 5, the circular area whose radius satisfies the distance constraint at the end time is:

[0087] A first circular area with a radius of r1 being the closest distance between the service spacecraft and the space target, and a second circular area with a radius of r2 being the farthest distance between the service spacecraft and the space target, the first circular area and the second circular area are concentric circles.

[0088] In order to explain the technical solution of the present invention in detail, a specific example is now provided:

[0089] The target orbit coordinate system is Figure 1 As shown:

[0090] An orbital coordinate system is established with the space debris as the origin. The origin O of the coordinate system is fixedly connected to the center of mass of the service spacecraft and moves along the orbit with it. The x-axis coincides with the geocentric radius r of the service spacecraft. The y-axis is perpendicular to the x-axis in the orbital plane, pointing in the positive direction of motion. The z-axis is determined by the right-hand rule, that is, in the direction of the orbital angular momentum.

[0091] In a certain mission, at an initial time t0 = 0, the service spacecraft and the space debris are located in an orbit with an altitude of 35779 km. The initial position of the service spacecraft relative to the space debris is shown in Table 1. The service spacecraft is planned to perform orbital maneuvers through pulse thrust. The mission goal is to make the service spacecraft at t f = Approaching space debris after 10 hours, and the final distance must be between 10 km and 50 km, and the illumination angle is required to be between 0° and 90°. In addition, the speed increment of each pulse thrust is controlled, and its maximum value shall not exceed 10 m / s.

[0092] Table 1

[0093]

[0094] In order to meet the task requirements, the constraints are established as follows:

[0095] 0<θ<90;

[0096] 10km<r pe <50km;

[0097]

[0098] By calculating the relational equation and screening through constraints, we ensure that only solutions that are feasible in actual operations are retained, thereby ensuring the practicality of the solution. Finally, 128,775 single speed increments that meet the constraints are determined to form the speed increment solution set. These single speed increments represent possible paths for mission success in specific scenarios.

[0099] like Figure 2 and Figure 3 As shown, by constructing two circular areas with the space debris as the center and the radius of 10 kilometers and 50 kilometers respectively, the reach of the service spacecraft can be intuitively displayed under the conditions of satisfying the single velocity increment constraint and the illumination angle constraint at the termination time. When the service spacecraft reaches these areas at the specified time, the mission can be considered successful. The end points of different colors represent different single velocity increments. This visualization method provides an intuitive way to identify and select the best orbital maneuvering strategy. After adding each single velocity increment that marks the success of the mission to the original velocity, the orbital maneuvering strategy of the service spacecraft can be obtained by propagating through the existing propagation model. The present invention does not improve the existing propagation method, which will not be repeated here.

[0100] By using the method described in this example, it is possible to accurately calculate the velocity increment solution set required for the service spacecraft to approach the space debris after 10 hours under a specific initial state. This method not only optimizes the calculation process of orbital maneuvers, but also ensures the accurate achievement of mission objectives. Experimental data confirms that the method of the present invention can efficiently handle complex orbital maneuvering problems, ensure that the service spacecraft approaches the space debris within the predetermined time, and complete the established mission under the conditions of satisfying the illumination angle constraint at the termination time and the distance constraint at the termination time. In addition, the present invention provides a more direct calculation method for actual operations through specific relational expressions, which significantly improves the success rate and efficiency of mission execution.

[0101] The beneficial effects of the embodiments of the present invention are:

[0102] 1. Efficient modeling and optimization of orbital transfer conditions:

[0103] The present invention constructs the motion equation and relational equation including the influence of J2 perturbation by comprehensively analyzing the current state (position and velocity) of the service spacecraft and the target orbital parameters to describe the orbital transfer of the space target. The relational equation not only takes into account the actual constraints such as the distance constraint at the termination time, the illumination angle constraint at the termination time and the single velocity increment constraint, but also specifically incorporates the J2 perturbation to ensure the accuracy and stability of the orbital transfer. By introducing the J2 perturbation, the model of the present invention can more accurately predict the dynamic behavior of the service spacecraft in the space environment, especially in the orbital service problem, which is crucial to the accuracy of the orbital transfer, improves the design accuracy of the orbital maneuvering strategy, and enables the service spacecraft to achieve more accurate orbital transfer in a complex space environment.

[0104] 2. Precise orbital maneuvering strategy based on impulse thrust and J2 perturbation:

[0105] The present invention deeply analyzes the existing models, and combines the actual conditions of pulse thrust and the influence of J2 perturbations to design a more accurate and practical orbital maneuvering strategy. Compared with traditional methods, the present invention not only solves the optimal control law or a single maneuvering strategy, but can calculate all feasible solutions of velocity increments that meet the conditions in a given scenario, while considering the influence of J2 perturbations. Through this method, the design space of orbital maneuvering strategies is expanded, and the flexibility and practicality of orbital maneuvers are improved, while ensuring the accuracy of orbital maneuvers under the influence of J2 perturbations. The strategy of the present invention can provide a variety of orbital maneuvering options that meet the conditions for the service spacecraft while considering the J2 perturbation, so that in actual operation, the most appropriate strategy can be flexibly selected according to the specific situation, thereby improving the success rate and efficiency of the mission.

[0106] 3. Efficient generation and application of feasible solution sets considering J2 perturbation:

[0107] The method of the present invention can generate a feasible solution set of velocity increments for all orbital maneuvering strategies that meet the conditions in a short time while considering the J2 perturbation. The velocity increment solution set can more accurately guide the service spacecraft to perform orbital maneuvers in a complex space environment, thereby improving the success rate and efficiency of the mission.

[0108] 4. Construction of orbital maneuvering strategy model for service spacecraft driven by pulse thrust:

[0109] The present invention proposes a service spacecraft orbital maneuvering strategy model, which accurately captures the motion characteristics of the service spacecraft in the space environment and realizes the simulation of its dynamic behavior. The establishment of this model provides theoretical support for the design and optimization of on-orbit service missions, and as a prediction tool, it enhances the scientificity and accuracy of mission planning. By introducing pulse thrust, the present invention expands the limitations of the traditional continuous thrust model and provides a more accurate mathematical framework for the complexity analysis and strategy formulation of on-orbit service missions.

[0110] 5. Enhanced flexibility and practicality of orbital maneuvering strategies:

[0111] The present invention obtains the orbital maneuvering strategy of the service spacecraft from the single velocity increment of the successful marking mission, and realizes a significant improvement in the mission success rate and the operational flexibility of the service spacecraft in response to the changing space mission requirements. Through optimized design, the strategy enables the service spacecraft to perform efficient maneuvers between different orbits, meeting the requirements of multi-orbit maneuvering missions. In addition, the strategy comprehensively considers time constraints and fuel consumption, provides an optimal solution for completing orbital maneuvers within the specified time, and ensures efficient use of fuel, thereby reducing operating costs while ensuring mission efficiency.

[0112] 6. Improved intelligence and foresight in mission planning:

[0113] This invention significantly improves the intelligence level of orbital maneuvering decision-making and the foresight of mission planning by generating velocity increment solutions that meet specific conditions. This intelligence and foresight enable mission planning to better adapt to changes in future mission requirements and improve the adaptability and flexibility of mission planning. By comprehensively analyzing all feasible solutions, this study provides a more comprehensive perspective for mission planning, thereby maximizing the effectiveness of mission execution in a complex and changing space environment.

[0114] The above is only a specific embodiment of the present invention, but the protection scope of the present invention is not limited thereto. Any person skilled in the art can easily think of changes or substitutions within the technical scope disclosed by the present invention, which should be included in the protection scope of the present invention. Therefore, the protection scope of the present invention should be based on the protection scope of the claims.

Claims

1. A multi-orbit maneuver strategy approach observation method for service spacecraft based on J2 perturbation compensation, characterized in that: The method includes: Step 1: In the target orbit coordinate system, construct the motion equation of the service spacecraft relative to the space target under the J2 perturbation compensation; Step 2: Based on the relative motion propagation relation, the relation equation between the terminal position of the service spacecraft in the target orbit coordinate system and the single velocity increment is derived through the motion equation, and the relation equation is used as the multi-orbit maneuver strategy model of the service spacecraft; Step 3: Setting constraints that meet the task requirements, wherein the constraints consist of a termination time distance constraint, a termination time illumination angle constraint, and a single speed increment constraint; Step 4: uniformly sample all the termination time positions that meet the termination time distance constraint, input all the sampling points into the service spacecraft multi-orbit maneuvering strategy model, output the single velocity increment required for the service spacecraft to reach the termination time position, retain the single velocity increment that meets both the termination time illumination angle constraint and the single velocity increment constraint, and obtain the velocity increment solution set; Step 5: Construct a circular area with the space target as the center and a radius that satisfies the distance constraint at the termination time; use the circular area to display the relative distance between the service spacecraft and the space target corresponding to each single velocity increment in the velocity increment solution set; when the service spacecraft arrives at the circular area at the specified time, the marking mission is successful; the orbital maneuvering strategy of the service spacecraft is obtained from each successful single velocity increment of the marking mission to conduct close observation of the space target.

2. The method according to claim 1, characterized in that In step 1, the equation of motion is: Where [x, y, z] is the position vector of the service spacecraft after time t in the target orbital coordinate system, and t is the time from the initial time; represents the state vector of the service spacecraft at the initial time t0 of the mission in the target orbital coordinate system, where the state vector includes the position vector and the velocity vector, [x0, y0, z0] is the position vector, is the velocity vector; c is the perturbation coefficient, and n is the average orbital angular velocity of the service spacecraft.

3. The method according to claim 2, characterized in that In step 2, the relative motion propagation relationship is: In the formula, is the velocity vector of the service spacecraft at the initial time t0 of the mission in the target orbit coordinate system, [ΔV x ,ΔV y ,ΔV z ] is the single velocity increment of the service spacecraft when the pulse increment is applied at the initial moment t0 of the mission, that is, the velocity increment obtained by the service spacecraft in the three directions x, y, and z of the target orbit coordinate system, [V x0 ,V y0 ,V z0 ] is the velocity of the service spacecraft before the pulse increment is applied.

4. The method according to claim 3, characterized in that In step 2, the relationship equation is: Where [M, N, U] is the end time position.

5. The method according to claim 1, characterized in that In step 3, the end time distance constraint is: r1<r pe <r2; where r pe is the relative distance between the service spacecraft and the space target, r1 is the shortest distance between the service spacecraft and the space target, and r2 is the farthest distance between the service spacecraft and the space target; The illumination angle constraint at the termination time is: θ1<θ<θ2; where θ is the illumination angle, that is, the angle between the vector of the space target pointing to the sun and the vector of the space target pointing to the service spacecraft, θ1 is the upper limit of the illumination angle, and θ2 is the lower limit of the illumination angle; The single velocity increment constraint is: In the formula, [ΔV x ,ΔV y ,ΔV z ] is the velocity increment obtained by the service spacecraft in the x, y, and z directions of the target orbit coordinate system, ΔV Represents the total velocity increment available for a single orbit change of the serviced spacecraft.

6. The method according to claim 1, characterized in that In step 4, the uniform sampling method includes: The orbital parameter range of the service spacecraft is set, with the space target at the termination time as the origin and the azimuth, pitch angle and distance of the service spacecraft relative to the space target as the step size to ensure the uniformity of sampling.

7. The method according to claim 1, characterized in that In step 4, the single speed increment that satisfies both the illumination angle constraint at the termination time and the single speed increment constraint is retained, and the method for obtaining the speed increment solution set includes: Eliminate the single speed increment that does not meet the illumination angle constraint at the end time; Calculate the velocity increment ΔV obtained by each single velocity increment in the remaining single velocity increments in the three directions of the target orbit coordinate system x, y, and z x , ΔV y , ΔV z The sum of is the total speed increment, and the total speed increment exceeding the threshold is excluded to obtain the speed increment solution set.

8. The method according to claim 1, characterized in that Step five also includes the step of selecting a service spacecraft orbital maneuvering strategy that meets predefined standards as the best orbital maneuvering strategy for close-in observation of the space target.

9. The method according to claim 8, characterized in that In step five, the predefined criteria are defined based on mission requirements, including but not limited to selecting the service spacecraft orbital maneuvering strategy corresponding to the smallest velocity increment or the shortest transfer time among the service spacecraft orbital maneuvering strategies that mark the success of the mission as the optimal orbital maneuvering strategy for close observation of the space target.

10. The method according to claim 1, characterized in that In step 5, the circular area whose radius satisfies the distance constraint at the end time is: A first circular area with a radius of r1 being the closest distance between the service spacecraft and the space target, and a second circular area with a radius of r2 being the farthest distance between the service spacecraft and the space target, the first circular area and the second circular area are concentric circles.

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