A method for close-range observation of a service spacecraft's multi-orbit maneuver strategy based on J2 perturbation compensation

Abstract

The present invention discloses a method for close observation of a service spacecraft's multi-orbit maneuvering strategy based on J2 perturbation compensation, which solves the problem of how to quickly improve the accuracy of orbit maneuvering and belongs to the field of close observation of a service spacecraft's multi-orbit maneuvering strategy. The method includes: constructing a motion equation under J2 perturbation compensation in the target orbit coordinate system; constructing a multi-orbit maneuvering strategy model of the service spacecraft through the relative motion propagation relation and the motion equation; inputting the position at the termination moment that meets the distance constraint at the termination moment into the model to output the single-velocity increment, and retaining the single-velocity increments that meet both the illumination angle constraint and the single-velocity increment constraint at the termination moment to obtain a velocity increment solution set; constructing a circular region centered on the space target with a radius that meets the distance constraint at the termination moment; marking the task as successful when the service spacecraft reaches the circular region at the specified moment, and obtaining the multi-orbit maneuvering strategy of the service spacecraft for close observation. The present invention realizes precise and rapid orbit maneuvering.

Description

Technical Field

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

[0002] The actual requirements of space missions have given rise to the research on the problem of proximity observation of orbit maneuvering strategies for on-orbit services, especially in dealing with space targets, such as space debris and repairing failed spacecraft, which is particularly crucial. With the continuous progress of space technology and the frequent increase of space activities, the complexity of the space environment has also increased, which makes protecting space assets and dealing with space targets a common challenge in the space field. When performing these tasks, the service spacecraft needs to be able to accurately execute orbit maneuvers to complete various proximity observation tasks. Traditional research on on-orbit service problems usually based on the continuous thrust model and uses methods such as differential games to find the optimal maneuvering strategy. Although these methods have been quite mature in theory, in practical applications, since the service spacecraft mainly relies on impulsive thrust to execute orbit 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 influence of the J2 perturbation 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 research field of spacecraft orbit maneuvering strategies, although theoretical methods such as differential games provide a theoretical framework for on-orbit service problems, these methods are mostly based on the continuous thrust model, while the actual orbit maneuvers of service spacecraft rely more on impulsive thrust. This difference between the model and the actual operation may make the existing methods face certain limitations when dealing with the on-orbit service problems of impulsive thrust spacecraft. In addition, the existing research mostly focuses on finding the optimal fuel solution under close distance constraints, and less considers the influence of the orbit transfer 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 of the existing research is based on the CW equation, only exploring the maneuvering strategy of the service spacecraft under the two-body dynamics model. For the influence of the J2 perturbation, as the most important perturbation force in the non-spherical perturbation force of the earth, it has a decisive influence on the on-orbit operation of the service spacecraft. It not only causes the long-term change of the orbit elements, but also causes the nutation and precession of the orbit 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 influence of the J2 perturbation in the orbit design and control of the service spacecraft to ensure that the service spacecraft can complete the mission safely and effectively. Summary of the Invention

[0004] To solve the technical problem of how to improve the accuracy of the orbital design and maneuver strategy of a servicing spacecraft, the present invention provides a method for approaching observation of a multi-orbit maneuver strategy of a servicing spacecraft based on J2 perturbation compensation. Through mathematical modeling and optimization techniques, an orbital maneuver strategy is provided for the servicing spacecraft to reach the desired approaching relative state from the current state. This strategy will determine the value range of the required velocity increments in the radial, lateral, and normal directions of the servicing spacecraft to achieve precise orbital maneuvers;

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

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

[0007] The present invention discloses a method for approaching observation of a multi-orbit maneuver strategy of a servicing spacecraft based on J2 perturbation compensation, and the method includes:

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

[0009] Step 2: Based on the relative motion propagation relation formula, derive the relation equation between the position at the termination moment and the single velocity increment of the servicing spacecraft in the target orbit coordinate system through the motion equation, and use the relation equation as the multi-orbit maneuver strategy model of the servicing spacecraft;

[0010] Step 3: Set the constraint conditions that meet the mission requirements, and the constraint conditions are composed of the distance constraint at the termination moment, the illumination angle constraint at the termination moment, and the single velocity increment constraint;

[0011] Step 4: Uniformly sample among all the positions at the termination moment that meet the distance constraint at the termination moment, input all the sampling points into the multi-orbit maneuver strategy model of the servicing spacecraft, output the single velocity increment required for the servicing spacecraft to reach the position at the termination moment, and retain the single velocity increments that simultaneously meet the illumination angle constraint at the termination moment and the single velocity increment constraint to obtain the velocity increment solution set;

[0012] Step 5: Construct a circular area centered on the space target with a radius satisfying the distance constraint at the termination time; display the relative distance between the servicing spacecraft and the space target corresponding to each single velocity increment in the velocity increment solution set through the circular area; mark the mission as successful when the servicing spacecraft reaches the circular area at the specified time; obtain the orbital maneuver strategy of the servicing spacecraft from each single velocity increment marked as mission successful to conduct a close observation of the space target.

[0013] In Step 1, the motion equation is:

[0014]

[0015] where [x, y, z] is the position vector of the servicing spacecraft at time t after the initial time in the target orbit coordinate system, and t is the time from the initial time. represents the state vector of the servicing spacecraft at the initial mission time t0 in the target orbit coordinate system. Among them, 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 servicing spacecraft.

[0016] In Step 2, the relative motion propagation relation is:

[0017]

[0018] where is the velocity vector of the servicing spacecraft at the initial mission time t0 in the target orbit coordinate system, [ΔV x , ΔV y , ΔV z is the single velocity increment of the servicing spacecraft when an impulse increment is applied instantaneously at the initial mission time t0, that is, the velocity increments obtained by the servicing spacecraft in the x, y, and z directions of the target orbit coordinate system, [V x0 , V y0 , V z0 is the velocity of the servicing spacecraft before the impulse increment is applied.

[0019] In Step 2, the relation equation is:

[0020]

[0021]

[0022] where [M, N, U] is the position at the termination time.

[0023] In Step 3, the distance constraint at the termination time is: r1 < r pe < r2; where r peLet \(r\) be the relative distance between the servicing spacecraft and the space target, \(r_1\) be the closest distance of the servicing spacecraft from the space target, and \(r_2\) be the farthest distance of the servicing spacecraft from the space target;

[0024] The illumination angle constraint at the termination time is: \(\theta_1\lt\theta\lt\theta_2\); where \(\theta\) is the illumination angle, i.e., the angle between the vector from the space target to the sun and the vector from the space target to the servicing spacecraft, \(\theta_1\) is the upper limit of the illumination angle, and \(\theta_2\) is the lower limit of the illumination angle;

[0025] The single - impulse velocity increment constraint is: In the formula, \([\Delta V x ,\Delta V y ,\Delta V z \) is the velocity increment obtained by the servicing spacecraft in the \(x\), \(y\), and \(z\) directions of the target orbit coordinate system, ΔV and \(\Delta V_{total}\) represents the total available velocity increment for a single - orbit maneuver of the servicing spacecraft.

[0026] In step four, the method of uniform sampling includes:

[0027] Set the range of the orbit parameters of the servicing spacecraft. With the space target at the termination time as the origin, the azimuth angle, pitch angle, and distance of the servicing spacecraft relative to the space target as the step sizes to ensure the uniformity of sampling.

[0028] In step four, the method of retaining the single - impulse velocity increments that simultaneously meet the illumination angle constraint at the termination time and the single - impulse velocity increment constraint to obtain the velocity increment solution set includes:

[0029] Exclude the single - impulse velocity increments that do not meet the illumination angle constraint at the termination time;

[0030] Calculate the sum of the velocity increments \(\Delta V x ,\Delta V y ,\Delta V z \) obtained in the \(x\), \(y\), and \(z\) directions of the target orbit coordinate system for each remaining single - impulse velocity increment to get the total velocity increment, and exclude the total velocity increments that exceed the threshold to obtain the velocity increment solution set.

[0031] In step five, it also includes the step of selecting a servicing spacecraft orbit maneuver strategy that meets the predefined criteria as the optimal orbit maneuver strategy to perform close - proximity observation on the space target.

[0032] In step five, the predefined criteria are defined based on the mission requirements, including but not limited to selecting the servicing spacecraft orbit maneuver strategy with the minimum velocity increment or the shortest transfer time among the servicing spacecraft orbit maneuver strategies that mark the mission as successful as the optimal orbit maneuver strategy to perform close - proximity observation on the space target.

[0033] In step five, the circular area whose radius satisfies the distance constraint at the termination time is:

[0034] A first circular region with a radius of r1, which is the closest distance between the servicing spacecraft and the space target, and a second circular region with a radius of r2, which is the farthest distance between the servicing spacecraft and the space target. The first circular region and the second circular region are concentric circles.

[0035] The beneficial effects of the present invention are as follows:

[0036] 1. Efficient modeling and optimization of orbit transfer conditions:

[0037] By comprehensively analyzing the current state (position and velocity) of the servicing spacecraft and the target orbit parameters, the present invention constructs a motion equation and a relationship equation including the influence of J2 perturbation to describe the orbit transfer of the space target. The relationship equation not only considers actual constraint conditions such as the distance constraint at the termination moment, the illumination angle constraint at the termination moment, and the single velocity increment constraint, but also specifically incorporates the J2 perturbation to ensure the accuracy and stability of the orbit transfer. By introducing the J2 perturbation, the model of the present invention can more accurately predict the dynamic behavior of the servicing spacecraft in the space environment. Especially in on-orbit servicing problems, this is crucial for the accuracy of orbit transfer, improves the design accuracy of orbit maneuver strategies, and enables the servicing spacecraft to achieve more precise orbit transfer in a complex space environment.

[0038] 2. Precise orbit maneuver strategy based on impulsive thrust and J2 perturbation:

[0039] The present invention deeply analyzes the existing models, and combines the actual conditions of impulsive thrust and the influence of J2 perturbation to design a more accurate and practical orbit maneuver strategy. Compared with traditional methods, the present invention not only solves the optimal control law or a single maneuver strategy, but can calculate all feasible solution sets of velocity increments that meet the conditions in a given scenario, while considering the influence of J2 perturbation. By this method, the design space of orbit maneuver strategies is expanded, the flexibility and practicality of orbit maneuvers are improved, and the accuracy of orbit maneuvers under the influence of J2 perturbation is ensured. The strategy of the present invention can provide multiple orbit maneuver options that meet the conditions for the servicing spacecraft considering the J2 perturbation, so that in actual operation, the most suitable strategy can be flexibly selected according to specific situations, improving the success rate and efficiency of the mission.

[0040] 3. Efficient generation and application of a feasible solution set considering J2 perturbation:

[0041] The method of the present invention can generate all feasible solution sets of velocity increments of orbit maneuver strategies that meet the conditions in a short time considering the J2 perturbation. The velocity increment solution set can more accurately guide the servicing spacecraft to perform orbit maneuvers in a complex space environment, improving the success rate and efficiency of the mission.

[0042] 4. Construction of the Orbital Maneuver Strategy Model for a Service Spacecraft Driven by Impulsive Thrust:

[0043] The present invention proposes an orbital maneuver strategy model for a service spacecraft, 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 impulsive thrust, the present invention expands the limitations of traditional continuous thrust models and provides a more accurate mathematical framework for the complexity analysis and strategy formulation of on-orbit service missions.

[0044] 5. Enhancement of the Flexibility and Practicality of the Orbital Maneuver Strategy:

[0045] The present invention obtains the orbital maneuver strategy of the service spacecraft from the single velocity increment that marks the success of the mission, and significantly improves the mission success rate and the operation flexibility of the service spacecraft for diverse space mission requirements. Through optimized design, this strategy enables the service spacecraft to maneuver efficiently between different orbits, meeting the requirements of multi-orbit maneuver missions. In addition, this strategy comprehensively considers time constraints and fuel consumption, provides an optimal solution for completing orbital maneuvers within a specified time, and ensures the efficient utilization of fuel, thereby reducing the operating cost while ensuring mission efficiency.

[0046] 6. Enhancement of the Intelligence and Proactiveness of Mission Planning:

[0047] The present invention significantly improves the intelligence level of orbital maneuver decision-making and the proactiveness of mission planning by generating a set of velocity increment solutions that meet specific conditions. This intelligence and proactiveness enable mission planning to better adapt to future mission requirement changes, improving the adaptability and flexibility of mission planning. By comprehensively analyzing all feasible solution sets, this research provides a more comprehensive perspective for mission planning, thus maximizing the mission execution effect in a complex and changing space environment. BRIEF DESCRIPTION OF THE DRAWINGS

[0048] The present invention will be further described in detail below with reference to the drawings and embodiments.

[0049] Figure 1 It is a schematic diagram of the target orbital coordinate system in an embodiment of the present invention.

[0050] Figure 2 It is a two-dimensional graphical schematic diagram of the set of velocity increment solutions in an embodiment of the present invention.

[0051] Figure 3 It is a three-dimensional graphical schematic diagram of the set of velocity increment solutions in an embodiment of the present invention. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0052] An approach for close observation of a service spacecraft's multi-orbit maneuver strategy based on J2 perturbation compensation is provided in an embodiment of the present invention. The method includes:

[0053] Step 1: In the target orbit coordinate system, construct the motion equation of the service spacecraft relative to the space target under J2 perturbation compensation.

[0054] Step 2: Based on the relative motion propagation relation, derive the relation equation between the position at the termination time and the single velocity increment of the service spacecraft in the target orbit coordinate system through the motion equation, and use the relation equation as the multi-orbit maneuver strategy model of the service spacecraft.

[0055] Step 3: Set the constraint conditions that meet the mission requirements. The constraint conditions consist of the termination time distance constraint, the termination time illumination angle constraint, and the single velocity increment constraint.

[0056] Step 4: Uniformly sample among all the positions at the termination time that meet the termination time distance constraint, input all the sampling points into the multi-orbit maneuver strategy model of the service spacecraft, output the single velocity increment required for the service spacecraft to reach the position at the termination time, and retain the single velocity increments that meet both the termination time illumination angle constraint and the single velocity increment constraint to obtain the velocity increment solution set.

[0057] Step 5: Construct a circular area centered on the space target with a radius that meets the termination time distance constraint; display the relative distance between the service spacecraft corresponding to each single velocity increment in the velocity increment solution set and the space target through the circular area; when the service spacecraft reaches the circular area at the specified time, mark the mission as successful; obtain the orbit maneuver strategy of the service spacecraft for close observation of the space target from each single velocity increment that marks the mission as successful.

[0058] In Step 1, the motion equation is:

[0059]

[0060] where [x, y, z] is the position vector of the service spacecraft at time t after the initial time in the target orbit coordinate system, and t is the time from the initial time. represents the state vector of the service spacecraft at the initial mission time t0 in the target orbit coordinate system. Among them, 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 orbit inclination of the space target, r ref is the average orbit radius of the space target, R e is the radius of the Earth, J2 is the gravity field harmonic coefficient, and n is the average orbit angular velocity of the service spacecraft.

[0061] In step 2, the relative motion propagation relation is:

[0062]

[0063] In the formula, is the velocity vector of the servicing spacecraft at the initial mission time t0 in the target orbit coordinate system,

[0064] [ΔV x , ΔV y , ΔV z is the single velocity increment of the servicing spacecraft when applying the pulse increment instantaneously at the initial mission time t0, that is, the velocity increments obtained by the servicing spacecraft in the x, y, and z directions of the target orbit coordinate system. [V x0 , V y0 , V z0 is the velocity of the servicing spacecraft before applying the pulse increment.

[0065] In step 2, the relational equation is:

[0066]

[0067]

[0068] In the formula, [M, N, U] is the position at the termination time. This relational equation does not consider the maneuvering strategy under constraints. Given the initial and termination state position vectors, the pulse increment required for orbit transfer can be obtained.

[0069] In step 3, in order to make the strategy more practical and operable, constraint conditions are added during model solving. The constraint conditions involve the proximity to the target point, the solar illumination angle of the target point, and the maximum allowable speed change for a single maneuver. By introducing the constraint conditions, it is ensured that the maneuvering strategy not only conforms to the control characteristics of the servicing spacecraft but also meets the specific requirements of the mission objectives. Among them, the distance constraint at the termination time is: r1 < r pe < r2; In the formula, r pe is the relative distance between the servicing spacecraft and the space target, r1 is the closest distance between the servicing spacecraft and the space target, and r2 is the farthest distance between the servicing spacecraft and the space target;

[0070] The illumination angle constraint at the termination time is: θ1 < θ < θ2; In the formula, θ is the illumination angle, that is, the angle between the vector pointing from the space target to the sun and the vector pointing from the space target to the servicing spacecraft. θ1 is the upper limit of the illumination angle, and θ2 is the lower limit of the illumination angle;

[0071] The single velocity increment constraint is: In the formula, [ΔV x , ΔV y , ΔVz are the velocity increments obtained by the servicing spacecraft in the three directions of x, y, and z in the target orbit coordinate system. ΔV represents the total velocity increment available for a single orbit maneuver of the servicing spacecraft.

[0072] In step four, the method of uniform sampling includes:

[0073] Set the range of the orbit parameters of the servicing spacecraft. Taking the space target at the termination moment as the origin, and using the azimuth angle, pitch angle, and distance of the servicing spacecraft relative to the space target as the step sizes to ensure the uniformity of sampling.

[0074] In step four, the method of obtaining the velocity increment solution set by retaining the single - velocity increments that simultaneously meet the illumination angle constraint at the termination moment and the single - velocity increment constraint includes:

[0075] Exclude the single - velocity increments that do not meet the illumination angle constraint at the termination moment;

[0076] Calculate the sum of the velocity increments ΔV x , ΔV y , ΔV z obtained by each single - velocity increment in the three directions of x, y, and z in the target orbit coordinate system for the remaining single - velocity increments, and exclude the total velocity increments that exceed the threshold to obtain the velocity increment solution set.

[0077] In step five, it also includes the step of selecting the orbit maneuver strategy of the servicing spacecraft that meets the predefined criteria as the optimal orbit maneuver strategy to conduct close - range observation of the space target.

[0078] In step five, the predefined criteria are defined based on the mission requirements, including but not limited to selecting the orbit maneuver strategy of the servicing spacecraft corresponding to the minimum velocity increment or the shortest transfer time among the orbit maneuver strategies that mark the mission as successful as the optimal orbit maneuver strategy to conduct close - range observation of the space target.

[0079] In step five, the circular region whose radius satisfies the distance constraint at the termination moment is:

[0080] The first circular region with a radius of r1, the closest distance between the servicing spacecraft and the space target, and the second circular region with a radius of r2, the farthest distance between the servicing spacecraft and the space target. The first circular region and the second circular region are concentric circles.

[0081] To illustrate the technical solution of the present invention in detail, a specific example is provided as follows:

[0082] The target orbit coordinate system is as Figure 1 shown:

[0083] 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 centroid of the servicing spacecraft and moves along the orbit with it. The x-axis coincides with the geocentric radius vector r of the servicing spacecraft, the y-axis is perpendicular to the x-axis in the orbital plane, and the positive direction points to the direction of motion. The z-axis is determined by the right-hand rule, i.e., along the direction of the orbital angular momentum.

[0084] In a certain mission, at an initial time t0 = 0, the servicing spacecraft and the space debris are located on an orbit at an altitude of 35,779 km. The initial position state of the servicing spacecraft relative to the space debris is shown in Table 1. The servicing spacecraft plans to perform an orbital maneuver through impulsive thrust. The mission objective is to make the servicing spacecraft approach the space debris after t f = 10 h, and the final distance should be between 10 km and 50 km, while the illumination angle is required to be between 0° and 90°. In addition, the velocity increment of each impulsive thrust is controlled, and its maximum value shall not exceed 10 m / s.

[0085] Table 1

[0086]

[0087] To meet the mission requirements, the following constraint conditions are established:

[0088] 0 < θ < 90;

[0089] 10 km < r pe < 50 km;

[0090]

[0091] Through calculation using the relational equation and screening by the constraint conditions, only the solutions that are feasible in actual operation are retained, thus ensuring the practicality of the solutions. Finally, a velocity increment solution set consisting of 128,775 single velocity increments that meet the constraint conditions is determined. These single velocity increments represent the possible paths to mission success in a specific scenario.

[0092] As Figure 2 and Figure 3 shown, by constructing two circular regions centered on the space debris with radii of 10 km and 50 km respectively, the arrival range of the servicing spacecraft under the conditions of meeting the single velocity increment constraint and the illumination angle constraint at the termination time can be visually displayed. When the servicing spacecraft reaches these regions at the specified time, the mission can be considered successful. The termination points of different colors represent different single velocity increments. This visualization method provides an intuitive way to identify and select the optimal orbital maneuver strategy. After adding each single velocity increment that marks the mission success to the original velocity and propagating it through the existing propagation model, the orbital maneuver strategy of the servicing spacecraft can be obtained. The present invention does not improve the existing propagation method and will not be repeated here.

[0093] By using the method described in this example, the velocity increment solution set required for the service spacecraft to approach the space debris after 10 hours under specific initial conditions can be accurately calculated. This method not only optimizes the calculation process of orbit maneuver but also ensures the accurate achievement of mission objectives. Experimental data confirms that the method of the present invention can efficiently handle complex orbit maneuver problems, ensure that the service spacecraft approaches the space debris within the predetermined time, and complete the established tasks under the conditions of satisfying the illumination angle constraint at the termination moment and the distance constraint at the termination moment. In addition, the present invention provides a more direct calculation method for actual operation through a specific relational expression, significantly improving the success rate and efficiency of mission execution.

[0094] The beneficial effects of the embodiments of the present invention are as follows:

[0095] 1. Efficient modeling and optimization of orbit transfer conditions:

[0096] By comprehensively analyzing the current state (position and velocity) of the service spacecraft and the target orbit parameters, the present invention constructs a motion equation and a relational equation including the influence of J2 perturbation to describe the orbit transfer of the space target. This relational equation not only considers practical constraint conditions such as the distance constraint at the termination moment, the illumination angle constraint at the termination moment, and the single velocity increment constraint, but also particularly incorporates the J2 perturbation to ensure the accuracy and stability of orbit 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 on-orbit service problems, this is crucial for the accuracy of orbit transfer, improving the design accuracy of orbit maneuver strategies, and enabling the service spacecraft to achieve more precise orbit transfer in a complex space environment.

[0097] 2. Precise orbit maneuver strategy based on impulsive thrust and J2 perturbation:

[0098] The present invention deeply analyzes the existing models, and combines the actual conditions of impulsive thrust and the influence of J2 perturbation to design a more precise and practical orbit maneuver strategy. Compared with traditional methods, the present invention not only solves the optimal control law or a single maneuver strategy, but can calculate all feasible solution sets of velocity increments that meet the conditions in a given scenario, while considering the influence of J2 perturbation. By this method, the design space of orbit maneuver strategies is expanded, the flexibility and practicality of orbit maneuver are improved, and at the same time, the accuracy of orbit maneuver under the influence of J2 perturbation is ensured. The strategy of the present invention can provide multiple orbit maneuver options that meet the conditions for the service spacecraft considering the J2 perturbation, enabling flexible selection of the most suitable strategy according to specific circumstances in actual operation, and improving the success rate and efficiency of the mission.

[0099] 3. Efficient Generation and Application of the Feasible Solution Set Considering J2 Perturbation:

[0100] The method of the present invention can generate a feasible solution set of the velocity increment of all orbit maneuver strategies that meet the conditions in a short time considering the J2 perturbation. The velocity increment solution set can more accurately guide the servicing spacecraft to perform orbit maneuvers in a complex space environment, improving the success rate and efficiency of the mission.

[0101] 4. Construction of the Orbit Maneuver Strategy Model of the Servicing Spacecraft Driven by Impulsive Thrust:

[0102] The present invention proposes an orbit maneuver strategy model for the servicing spacecraft, which accurately captures the motion characteristics of the servicing spacecraft in the space environment and realizes the simulation of its dynamic behavior. The establishment of this model provides a theoretical support for the design and optimization of on-orbit servicing missions, and as a prediction tool, it enhances the scientificity and accuracy of mission planning. By introducing impulsive 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 servicing missions.

[0103] 5. Enhancement of the Flexibility and Practicality of the Orbit Maneuver Strategy:

[0104] The present invention obtains the orbit maneuver strategy of the servicing spacecraft from a single velocity increment that marks the success of the mission, and significantly improves the success rate of the mission and the operation flexibility of the servicing spacecraft for the changing space mission requirements. Through optimized design, this strategy enables the servicing spacecraft to perform efficient maneuvers between different orbits, meeting the requirements of multi-orbit maneuver missions. In addition, this strategy comprehensively considers time constraints and fuel consumption, provides an optimal solution for completing orbit maneuvers within a specified time, and ensures the efficient utilization of fuel, thereby reducing the operating cost while ensuring the mission efficiency.

[0105] 6. Enhancement of the Intelligence and Proactivity of Mission Planning:

[0106] The present invention significantly improves the intelligent level of orbit maneuver decision-making and the proactivity of mission planning by generating a feasible solution set that meets specific conditions. This intelligence and proactivity enable mission planning to better adapt to the changes in future mission requirements, improving the adaptability and flexibility of mission planning. By comprehensively analyzing all feasible solution sets, this study provides a more comprehensive perspective for mission planning, thus maximizing the mission execution effect in a complex and changing space environment.

[0107] The above are only specific embodiments of the present invention, but the protection scope of the present invention is not limited thereto. Any person skilled in the art within the technical scope disclosed by the present invention can easily think of changes or substitutions, which should all be covered within the protection scope of the present invention. Therefore, the protection scope of the present invention shall be subject to the protection scope of the claims described above.

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 orbital coordinate system, and ΔV represents the total velocity increment available for a single orbit change of the service 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.

Images