Construction and screening method of orbital transfer speed increment required by pulse thrust spacecraft on-orbit service based on CW equation

Abstract

The invention discloses a method for constructing and screening orbital transfer speed increment required by on-orbit service of a pulse thrust spacecraft based on a CW equation, belongs to the field of on-orbit service of spacecrafts, and solves the problem of low precision of a maneuvering orbital transfer strategy of the on-orbit service orbit of the pulse thrust spacecraft. Comprising the following steps: constructing a relative motion equation of a pulsed thrust spacecraft relative to space debris based on a CW equation; inputting the orbital transfer state transfer matrix and the speed transfer matrix into the relative motion equation at the same time to obtain an on-orbit service orbital transfer model; setting a termination state distance constraint, a termination state illumination angle constraint and an orbital transfer speed increment constraint according to task requirements; and inputting all the termination moment position vectors conforming to the termination state distance constraint into the on-orbit service orbital transfer model, and reserving orbital transfer speed increments conforming to the termination state illumination angle constraint and the orbital transfer speed increment constraint to obtain a screened orbital transfer speed increment solution set. According to the invention, the precision of the maneuvering orbital transfer strategy of the on-orbit service orbit of the pulse thrust spacecraft is improved.

Description

Technical Field

[0001] The invention belongs to the technical field of on-orbit service of spacecraft, and relates to a construction and screening method of orbit change speed increment required for on-orbit service of a pulse thrust spacecraft based on a CW equation. Background Art

[0002] The on-orbit servicing problem originates from the actual needs of space missions, especially in the removal of space debris and the repair of failed spacecraft. With the development of space technology and the increase in space activities, the space environment is becoming increasingly complex. The safety of space assets and the handling of space debris have become common challenges faced by the aerospace field. In these tasks, the service spacecraft must be able to perform orbital maneuvers accurately to complete various operational tasks. Traditional on-orbit servicing problems are usually described by continuous thrust models, and the optimal maneuvering strategy is solved by methods such as differential games. These methods are relatively mature in theory, but in practical applications, since the service spacecraft mainly uses pulse thrust for orbital maneuvers, there is a difference between the continuous thrust model and the actual application, which limits its effectiveness in actual space missions.

[0003] Although differential game methods provide theoretical solutions to the on-orbit servicing problem, most of these methods are based on continuous thrust models, while actual service spacecraft orbital maneuvers mostly use pulse thrust. The difference between this model and actual application leads to the limitations of existing methods in dealing with the on-orbit servicing orbit change problem of pulse thrust spacecraft. In addition, existing research mainly focuses on finding the optimal fuel solution under the constraint of approach distance, while ignoring the impact of the choice of orbit change strategy on the final relative position of the service spacecraft, thereby limiting the in-depth understanding of the relative motion law of the service spacecraft and the comprehensive optimization of the maneuvering strategy. Therefore, the existing technology is insufficient in providing feasible orbital maneuvering strategies and ensuring that the service spacecraft completes its mission safely and effectively, especially when considering the impact of actual thrust conditions and strategy selection on the results. Summary of the invention

[0004] In order to solve the technical problem of low accuracy of orbit maneuvering strategy for on-orbit service of pulse thrust spacecraft in the prior art, the present invention provides a construction and screening method of orbit change velocity increment required for on-orbit service of pulse thrust spacecraft based on CW equation, explores the influence of different orbit change schemes on the relative position relationship between two space targets through mathematical modeling and optimization technology, obtains the orbit change velocity increment required for on-orbit service of pulse thrust spacecraft according to the actual conditions of pulse thrust and the constructed on-orbit service orbit change model, obtains the orbit change velocity increment solution set after screening of actual constraint conditions, and designs a more flexible and practical orbit maneuvering strategy from the orbit change velocity increment solution set, and provides an orbit maneuvering strategy for the pulse thrust spacecraft to reach the expected approach relative state from the current state, which can not only optimize the orbit maneuver of the pulse thrust spacecraft, but also provide an approximate range of maneuvering values ​​that meet the mission constraints, determine the orbit change velocity increment value range required for the pulse thrust spacecraft in radial, lateral and normal directions, and improve the efficiency and accuracy of the algorithm to achieve accurate orbit maneuvering. Through the present invention, the relative motion of a pulse thrust spacecraft in an actual space environment can be better understood and predicted, providing a scientific basis for the orbit design and mission planning of the spacecraft.

[0005] The purpose of the present invention is specifically achieved through the following technical solutions:

[0006] The present invention discloses a construction and screening method of a velocity increment for orbit change required for on-orbit service of a pulse thrust spacecraft based on a CW equation, the method comprising:

[0007] Step 1: Based on the CW equation, construct the relative motion equation of the pulse thrust spacecraft relative to the space debris in the target orbit coordinate system;

[0008] Step 2: Input the orbit change state transfer matrix and velocity transfer matrix into the relative motion equation at the same time, and obtain the on-orbit service orbit change model of the relationship between the position vector of the pulse thrust spacecraft at the terminal time in the target orbit coordinate system and the orbit change velocity increment;

[0009] Step 3: Set the terminal state distance constraint, terminal state illumination angle constraint and track change speed increment constraint according to the task requirements;

[0010] Step 4: Input all the terminal time position vectors that meet the terminal state distance constraint into the on-orbit service trajectory change model, and output the trajectory change speed increment; screen and retain the trajectory change speed increments that meet both the terminal state illumination angle constraint and the trajectory change speed increment constraint, and obtain the trajectory change speed increment solution set.

[0011] In step 1, the relative motion equation is:

[0012]

[0013] in, is the state vector of the pulse thrust spacecraft relative to the space debris at the current moment, the state vector includes the position vector and the velocity vector, where [x p (t),y p (t),z p (t)] is the position vector at the current moment, is the velocity vector at the current moment; P represents the pulse thrust spacecraft, t represents the current moment; φ(t,t 0 ) is the relative motion change state transfer matrix, t 0 represents the initial moment; is the state vector of the pulse thrust spacecraft relative to the space debris at the initial moment, where [x p (t 0 ),y p (t 0 ),z p (t 0 )] is the position vector at the initial moment, is the velocity vector at the initial moment; i is the number of pulse thrusts, M is the total number of pulse thrusts, φ v (t,t i ) is the relative motion speed transfer matrix, t i represents the control time when the pulse thrust spacecraft applies the i-th pulse thrust, Δv P (t i ) is the pulse thrust spacecraft at t i The orbit change speed increment that applies pulse thrust at all times.

[0014] In step 2, the track change state transfer matrix is:

[0015]

[0016] Among them, φ(Δt) is the track change state transfer matrix to be input, is the average orbital angular velocity of the pulse thrust spacecraft, μ is the Earth's gravitational constant, r is the orbital radius of the space debris, Δt = tt 0 is the current time t and the initial time t 0 The difference.

[0017] In step 2, the velocity transfer matrix is:

[0018]

[0019] Among them, φ v (Δt i ) is the velocity transfer matrix to be input, is the average orbital angular velocity of the pulse thrust spacecraft, μ is the Earth's gravitational constant, r is the orbital radius of the space debris, Δt i=tt i is the current time t and the control time t when the pulse thrust spacecraft applies the i-th pulse thrust i The difference is , i is the number of pulse thrusts.

[0020] In step 2, the on-orbit service orbit change model is:

[0021]

[0022] In the formula, [x p ,y p ,z p ] is the position vector at the end time, [ΔV x ,ΔV y ,ΔV z ] is the initial time of the task t 0 The orbit change velocity increment of the pulse thrust spacecraft when the pulse increment is applied instantaneously, that is, the orbit change velocity increment obtained by the pulse thrust spacecraft in the three directions of x, y, and z in the target orbit coordinate system; is the average orbital angular velocity of the pulse thrust spacecraft, μ is the Earth's gravitational constant, r is the orbital radius of the space debris, Δt = tt 0 is the current time t and the initial time t 0 The difference, is the state vector of the pulse thrust spacecraft relative to the space debris at the initial moment, where [x p (t 0 ),y p (t 0 ),z p (t 0 )] is the position vector at the initial moment, is the velocity vector at the initial moment; Δt i =tt i is the current time t and the control time t when the pulse thrust spacecraft applies the i-th pulse thrust i The difference is , i is the number of pulse thrusts.

[0023] In step 3, the final state distance constraint is:

[0024] r 1 <r pe <r 2 ;

[0025] Among them, r pe is the relative distance between the pulse thrust spacecraft and the space debris, r 1 is the shortest distance between the pulse thrust spacecraft and the space debris, r 2 The maximum distance between a pulse thrust spacecraft and space debris.

[0026] In step 3, the final state illumination angle constraint is:

[0027] θ 1 <θ<θ 2 ;

[0028] Among them, θ 1 is the lower limit of the illumination angle, θ 2 is the upper limit of the illumination angle, and θ is the illumination angle, that is, the angle between the vector of the space debris pointing to the sun and the vector of the space debris pointing to the pulse thrust spacecraft.

[0029] In step 3, the track change speed increment constraint is:

[0030]

[0031] Where, ΔV x , ΔV y , ΔV z They represent the orbit change velocity increments obtained by the pulse thrust spacecraft in the x, y, and z directions in the target orbit coordinate system, ΔV Represents the total orbit change velocity increment of the pulse thrust spacecraft.

[0032] In step 4, all the terminal time position vectors that meet the terminal state distance constraint are input into the on-orbit service orbit change model, and the method of outputting the orbit change velocity increment includes:

[0033] The orbital parameter range of the pulse thrust spacecraft is set, with the space debris at the termination time as the origin, the azimuth angle, pitch angle and distance of the pulse thrust spacecraft relative to the space debris as the step size, and uniform sampling is performed on all the termination time position vectors that meet the termination state distance constraint. All sampling points are input into the on-orbit service orbit change model, and the orbit change velocity increment is output.

[0034] In step 4, the method of screening and retaining the track change speed increments that meet both the terminal state illumination angle constraint and the track change speed increment constraint to obtain the track change speed increment solution set includes:

[0035] Eliminate the track change speed increment that does not meet the illumination angle constraint of the terminal state;

[0036] Calculate the orbit change speed increment ΔV obtained by each orbit change speed increment in the remaining orbit change speed increment in the three directions of x, y, and z of the target orbit coordinate system x , ΔV y , ΔV z The sum of is the total track change speed increment, and the total track change speed increment exceeding the threshold is excluded to obtain the track change speed increment solution set.

[0037] The fifth step is to construct a circular area with the space debris as the center and a radius that satisfies the distance constraint of the terminal state; the circular area is used to display the relative distance r between the pulse thrust spacecraft and the space debris corresponding to each orbit change velocity increment in the orbit change velocity increment solution set. pe ; When the pulse thrust spacecraft reaches the circular area at the specified time, the mission is marked as successful.

[0038] Among them, the circular area whose radius satisfies the end state distance constraint is:

[0039] The radius is the shortest distance r between the pulse thrust spacecraft and the space debris. 1 The first circular area, and the radius is the maximum distance r of the pulse thrust spacecraft from the space debris 2 The first circular area and the second circular area are concentric circles.

[0040] The beneficial effects of the present invention are:

[0041] The present invention conducts a comprehensive analysis of the current state vector (including position vector and velocity vector) and target orbit parameters of the pulse thrust spacecraft, derives the relationship between the position vector at the termination time and the orbit change velocity increment, and constructs an on-orbit service orbit change model that describes the orbit transfer of the pulse thrust spacecraft. The present invention can provide an optimization scheme for pulse maneuvers under any conditions, enhance the adaptability and flexibility of the on-orbit service of the pulse thrust spacecraft, and can quickly formulate effective maneuvering strategies according to specific mission requirements. The on-orbit service orbit change model takes into account the actual constraints of the terminal state distance constraint, the terminal state illumination angle constraint, and the orbit change velocity increment constraint, ensures the accuracy and stability of the orbit transfer, realizes the accurate description and dynamic simulation of the motion characteristics of the pulse thrust spacecraft, and provides a solid theoretical basis and prediction tool for the design of reasonable on-orbit service tasks.

[0042] The present invention combines the actual conditions of pulse thrust to construct and screen out a set of track change speed increment solutions for designing more accurate and practical orbital maneuvering strategies. Unlike previous methods, the present invention not only solves optimal control or a single maneuvering strategy, but is able to calculate a set of solutions for all track change speed increments that meet the conditions under a given scenario. The present invention introduces scenario constraints, solves the track change amount for each terminal state, and screens feasible solutions based on the terminal state distance constraint, the terminal state illumination angle constraint, and the track change speed increment constraint, providing a set of feasible solutions under given constraints. This not only ensures the feasibility of the proposed maneuvering strategy in actual operation, but also optimizes resource utilization efficiency and provides decision makers with clear choices, thereby improving the success rate of the mission; this method not only expands the design space of orbital maneuvering strategies, but also improves the flexibility and practicality of orbital maneuvers.

[0043] Through the method of the present invention, the orbit change velocity increment solution set of all orbital maneuvering strategies that meet the conditions can be generated in a short time, which not only improves the instant decision-making ability of orbital maneuvers, but also provides more options and foresight for mission planning. This intelligence and foresight enable mission planning to better adapt to future mission requirements and changes, and further improve the overall execution effect of the mission. The solution set can provide a variety of options for pulse thrust spacecraft, so that in actual operation, the most appropriate strategy can be flexibly selected according to specific circumstances, thereby improving the success rate and efficiency of the mission. BRIEF DESCRIPTION OF THE DRAWINGS

[0044] The present invention is further described in detail below based on the accompanying drawings and embodiments.

[0045] Figure 1 It is a schematic diagram of the target orbit coordinate system.

[0046] Figure 2 It is a schematic diagram of the solution set of track change speed increment. DETAILED DESCRIPTION

[0047] Embodiment 1 of the present invention provides a method for constructing and screening an orbit change velocity increment required for on-orbit service of a pulse thrust spacecraft based on a CW equation, the method comprising:

[0048] Step 1: Based on the CW equation, construct the relative motion equation of the pulse thrust spacecraft relative to the space debris in the target orbit coordinate system;

[0049] Step 2: Input the orbit change state transfer matrix and velocity transfer matrix into the relative motion equation at the same time, and obtain the on-orbit service orbit change model of the relationship between the position vector of the pulse thrust spacecraft at the terminal time in the target orbit coordinate system and the orbit change velocity increment;

[0050] Step 3: Set the terminal state distance constraint, terminal state illumination angle constraint and track change speed increment constraint according to the task requirements;

[0051] Step 4: Input all the terminal time position vectors that meet the terminal state distance constraint into the on-orbit service trajectory change model, and output the trajectory change speed increment; screen and retain the trajectory change speed increments that meet both the terminal state illumination angle constraint and the trajectory change speed increment constraint, and obtain the trajectory change speed increment solution set.

[0052] In step 1, the relative motion equation is:

[0053]

[0054] in, is the state vector of the pulse thrust spacecraft relative to the space debris at the current moment, the state vector includes the position vector and the velocity vector, where [x p (t),yp (t),z p (t)] is the position vector at the current moment, is the velocity vector at the current moment; P represents the pulse thrust spacecraft, t represents the current moment; φ(t,t 0 ) is the relative motion change state transfer matrix, t 0 represents the initial moment; is the state vector of the pulse thrust spacecraft relative to the space debris at the initial moment, where [x p (t 0 ),y p (t 0 ),z p (t 0 )] is the position vector at the initial moment, is the velocity vector at the initial moment; i is the number of pulse thrusts, M is the total number of pulse thrusts, φ v (t,t i ) is the relative motion speed transfer matrix, t i represents the control time when the pulse thrust spacecraft applies the i-th pulse thrust, Δv P (t i ) is the pulse thrust spacecraft at t i The orbit change speed increment that applies pulse thrust at all times.

[0055] In step 2, the track change state transfer matrix is:

[0056]

[0057] Among them, φ(Δt) is the track change state transfer matrix to be input, is the average orbital angular velocity of the pulse thrust spacecraft, μ is the Earth's gravitational constant, r is the orbital radius of the space debris, Δt = tt 0 is the current time t and the initial time t 0 The difference.

[0058] In step 2, the velocity transfer matrix is:

[0059]

[0060] Among them, φ v (Δt i ) is the velocity transfer matrix to be input, is the average orbital angular velocity of the pulse thrust spacecraft, μ is the Earth's gravitational constant, r is the orbital radius of the space debris, Δt i =tt i is the current time t and the control time t when the pulse thrust spacecraft applies the i-th pulse thrust i The difference is , i is the number of pulse thrusts.

[0061] In step 2, the on-orbit service orbit change model is:

[0062]

[0063] In the formula, [x p ,y p ,z p ] is the position vector at the end time, [ΔV x ,ΔV y ,ΔV z ] is the initial time of the task t 0 The orbit change velocity increment of the pulse thrust spacecraft when the pulse increment is applied instantaneously, that is, the orbit change velocity increment obtained by the pulse thrust spacecraft in the three directions of x, y, and z in the target orbit coordinate system; is the average orbital angular velocity of the pulse thrust spacecraft, μ is the Earth's gravitational constant, r is the orbital radius of the space debris, Δt = tt 0 is the current time t and the initial time t 0 The difference, is the state vector of the pulse thrust spacecraft relative to the space debris at the initial moment, where [x p (t 0 ),y p (t 0 ),z p (t 0 )] is the position vector at the initial moment, is the velocity vector at the initial moment; Δt i =tt i is the current time t and the control time t when the pulse thrust spacecraft applies the i-th pulse thrust i The difference is , i is the number of pulse thrusts.

[0064] In step 3, the final state distance constraint is:

[0065] r 1 <r pe <r 2 ;

[0066] Among them, r pe is the relative distance between the pulse thrust spacecraft and the space debris, r 1 is the shortest distance between the pulse thrust spacecraft and the space debris, r 2 The maximum distance between a pulse thrust spacecraft and space debris.

[0067] In step 3, the final state illumination angle constraint is:

[0068] θ 1 <θ<θ 2 ;

[0069] Among them, θ 1 is the lower limit of the illumination angle, θ 2 is the upper limit of the illumination angle, and θ is the illumination angle, that is, the angle between the vector of the space debris pointing to the sun and the vector of the space debris pointing to the pulse thrust spacecraft.

[0070] In step 3, the track change speed increment constraint is:

[0071]

[0072] Where, ΔV x , ΔV y , ΔV z They represent the orbit change velocity increments obtained by the pulse thrust spacecraft in the x, y, and z directions in the target orbit coordinate system, ΔV Represents the total orbit change velocity increment of the pulse thrust spacecraft.

[0073] In step 4, all the terminal time position vectors that meet the terminal state distance constraint are input into the on-orbit service orbit change model, and the method of outputting the orbit change velocity increment includes:

[0074] The orbital parameter range of the pulse thrust spacecraft is set, with the space debris at the termination time as the origin and the azimuth, pitch angle and distance of the pulse thrust spacecraft relative to the space debris as the step size to ensure the uniformity of sampling. Uniform sampling is performed in all the termination time position vectors that meet the termination state distance constraint, and all sampling points are input into the on-orbit service orbit change model to output the orbit change velocity increment.

[0075] In step 4, the method of screening and retaining the track change speed increments that meet both the terminal state illumination angle constraint and the track change speed increment constraint to obtain the track change speed increment solution set includes:

[0076] Eliminate the track change speed increment that does not meet the illumination angle constraint of the terminal state;

[0077] Calculate the orbit change speed increment ΔV obtained by each orbit change speed increment in the remaining orbit change speed increment in the three directions of x, y, and z of the target orbit coordinate system x , ΔV y , ΔV z The total track change speed increment is obtained by summing up, and the total track change speed increment exceeding the threshold is excluded to obtain the track change speed increment solution set.

[0078] The fifth step is to construct a circular area with the space debris as the center and a radius that satisfies the distance constraint of the terminal state; the circular area is used to display the relative distance r between the pulse thrust spacecraft and the space debris corresponding to each orbit change velocity increment in the orbit change velocity increment solution set. pe; When the pulse thrust spacecraft reaches the circular area at the specified time, the mission is marked as successful.

[0079] The orbital maneuvering strategy of the pulse thrust spacecraft is obtained by the orbit change speed increment of the successful marking mission; for example: after the single speed increment of each successful marking mission is added to the original speed, the orbital maneuvering strategy of the pulse thrust spacecraft is obtained by propagating through the existing propagation model. The present invention does not improve the existing propagation method and will not be repeated here.

[0080] The pulse thrust spacecraft orbital maneuvering strategy that meets the predefined criteria is selected as the best orbital maneuvering strategy. The predefined criteria are defined based on the mission requirements, including but not limited to selecting the pulse thrust spacecraft orbital maneuvering strategy corresponding to the minimum orbit change speed increment or the shortest transfer time among the pulse thrust spacecraft orbital maneuvering strategies that mark the success of the mission.

[0081] Among them, the circular area whose radius satisfies the end state distance constraint is:

[0082] The radius is the shortest distance r between the pulse thrust spacecraft and the space debris. 1 The first circular area, and the radius is the maximum distance r of the pulse thrust spacecraft from the space debris 2 The first circular area and the second circular area are concentric circles.

[0083] In order to explain the technical solution of the present invention in detail, a specific example is now provided:

[0084] The 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 space debris and moves along the orbit with it. The x-axis coincides with the radius r of the center of the earth of the space debris. The y-axis is perpendicular to the x-axis in the orbital plane and points to the positive direction of motion. The z-axis is determined by the right-hand rule, that is, along the direction of the orbital momentum. The target orbital coordinate system is as follows: Figure 1 As shown:

[0085] In a certain task, at an initial time t 0 = 0, in an orbit with an altitude of 35779 km, there are pulse thrust spacecraft and space debris running, with the space debris as the coordinate origin, and the initial state is shown in Table 1. The pulse thrust spacecraft maneuvers in the form of pulse thrust. The existing mission requires the pulse thrust spacecraft to maneuver at t f = Approaching space debris after 10 hours, the end state distance constraint is required to be in the range of 10km~50km, and the end state illumination angle constraint is 0°~90°. The control component of the orbit change speed increment constraint has an upper limit constraint, which does not exceed 10m / s.

[0086] Table 1

[0087]

[0088] In order to meet the task requirements, the constraints are established as follows:

[0089] 0<θ<90;

[0090] 10km<r pe <50km;

[0091]

[0092] According to the on-orbit servicing orbit change model, for each terminal state that meets the constraints, it is substituted into the model to solve the initial orbit change velocity increment. All maneuvering schemes for this scenario can be obtained. This step is critical because it is directly related to whether a solution that meets the mission requirements can be found. In this way, the orbit change velocity increment required to reach a specific terminal position can be accurately calculated.

[0093] During the solution process, only the initial orbit change speed increments that meet the constraints are stored. This means that only those solutions that are feasible in actual operation will be retained, thus ensuring the practical application value of the solution. Through this method, 128,775 initial orbit change speed increments that meet the constraints were finally screened out, and the solution set of orbit change speed increments was obtained, which represents the possible path to mission success in a given scenario.

[0094] By establishing two circles centered on the space debris (with radii of 10km and 50km respectively), the reach of the pulse thrust spacecraft when the constraints are met can be intuitively demonstrated. When the pulse thrust spacecraft reaches these ranges at a given moment, the mission is considered a success. Figure 2 As shown, different colored end points represent different track change speed increments, which can provide an intuitive way to identify and select the best maneuvering strategy.

[0095] Through this example, all orbit change velocity increments of a pulse thrust spacecraft approaching space debris after 10 hours under given initial conditions can be successfully calculated. This method not only simplifies the calculation process of orbital maneuvers, but also ensures the precise realization of mission objectives. Experimental results show that the method of the present invention can effectively solve complex orbital maneuvering problems, ensure that the pulse thrust spacecraft can approach space debris within the specified time, and complete the mission under constraints. In addition, through the on-orbit service orbit change model, the present invention provides a more direct calculation method in actual operation, which significantly improves the success rate and efficiency of the mission.

[0096] The beneficial effects of the embodiments of the present invention are:

[0097] The present invention conducts a comprehensive analysis of the current state vector (including position vector and velocity vector) and target orbit parameters of the pulse thrust spacecraft, derives the relationship between the position vector at the termination time and the orbit change velocity increment, and constructs an on-orbit service orbit change model that describes the orbit transfer of the pulse thrust spacecraft. The present invention can provide an optimization scheme for pulse maneuvers under any conditions, enhance the adaptability and flexibility of the on-orbit service of the pulse thrust spacecraft, and can quickly formulate effective maneuvering strategies according to specific mission requirements. The on-orbit service orbit change model takes into account the actual constraints of the terminal state distance constraint, the terminal state illumination angle constraint, and the orbit change velocity increment constraint, ensures the accuracy and stability of the orbit transfer, realizes the accurate description and dynamic simulation of the motion characteristics of the pulse thrust spacecraft, and provides a solid theoretical basis and prediction tool for the design of reasonable on-orbit service tasks.

[0098] The present invention combines the actual conditions of pulse thrust to construct and screen out a set of track change speed increment solutions for designing more accurate and practical orbital maneuvering strategies. Unlike previous methods, the present invention not only solves optimal control or a single maneuvering strategy, but is able to calculate a set of solutions for all track change speed increments that meet the conditions under a given scenario. The present invention introduces scenario constraints, solves the track change amount for each terminal state, and screens feasible solutions based on the terminal state distance constraint, the terminal state illumination angle constraint, and the track change speed increment constraint, providing a set of feasible solutions under given constraints. This not only ensures the feasibility of the proposed maneuvering strategy in actual operation, but also optimizes resource utilization efficiency and provides decision makers with clear choices, thereby improving the success rate of the mission; this method not only expands the design space of orbital maneuvering strategies, but also improves the flexibility and practicality of orbital maneuvers.

[0099] Through the method of the present invention, the orbit change velocity increment solution set of all orbital maneuvering strategies that meet the conditions can be generated in a short time, which not only improves the instant decision-making ability of orbital maneuvers, but also provides more options and foresight for mission planning. This intelligence and foresight enable mission planning to better adapt to future mission requirements and changes, and further improve the overall execution effect of the mission. The solution set can provide a variety of options for pulse thrust spacecraft, so that in actual operation, the most appropriate strategy can be flexibly selected according to specific circumstances, thereby improving the success rate and efficiency of the mission.

[0100] 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 method for constructing and screening the orbit change velocity increment required for on-orbit service of a pulse thrust spacecraft based on the CW equation, characterized in that: The method includes: Step 1: Based on the CW equation, construct the relative motion equation of the pulse thrust spacecraft relative to the space debris in the target orbit coordinate system; Step 2: Input the orbit change state transfer matrix and velocity transfer matrix into the relative motion equation at the same time, and obtain the on-orbit service orbit change model of the relationship between the position vector of the pulse thrust spacecraft at the terminal time in the target orbit coordinate system and the orbit change velocity increment; Step 3: Set the terminal state distance constraint, terminal state illumination angle constraint and track change speed increment constraint according to the task requirements; Step 4: Input all the terminal time position vectors that meet the terminal state distance constraint into the on-orbit service trajectory change model, and output the trajectory change speed increment; screen and retain the trajectory change speed increments that meet both the terminal state illumination angle constraint and the trajectory change speed increment constraint, and obtain the trajectory change speed increment solution set.

2. The method according to claim 1, characterized in that In step 1, the relative motion equation is: in, is the state vector of the pulse thrust spacecraft relative to the space debris at the current moment, the state vector includes the position vector and the velocity vector, where [x p (t),y p (t),z p (t)] is the position vector at the current moment, is the velocity vector at the current moment; P represents the pulse thrust spacecraft, t represents the current moment; φ(t,t0) is the relative motion trajectory change state transfer matrix, t0 represents the initial moment; is the state vector of the pulse thrust spacecraft relative to the space debris at the initial moment, where [x p (t0),y p (t0),z p (t0)] is the position vector at the initial moment, is the velocity vector at the initial moment; i is the number of pulse thrusts, M is the total number of pulse thrusts, φ v (t,t i ) is the relative motion speed transfer matrix, t i represents the control time when the pulse thrust spacecraft applies the i-th pulse thrust, Δv P (t i ) is the pulse thrust spacecraft at t i The orbit change speed increment that applies pulse thrust at all times.

3. The method according to claim 1, characterized in that In step 2, the track change state transfer matrix is: Among them, φ(Δt) is the track change state transfer matrix to be input, is the average orbital angular velocity of the pulse thrust spacecraft, μ is the gravitational constant of the Earth, r is the orbital radius of the space debris, and Δt=t-t0 is the difference between the current time t and the initial time t0.

4. The method according to claim 1, characterized in that In step 2, the velocity transfer matrix is: Among them, φ v (Δt i ) is the velocity transfer matrix to be input, is the average orbital angular velocity of the pulse thrust spacecraft, μ is the Earth's gravitational constant, r is the orbital radius of the space debris, Δt i =tt i is the current time t and the control time t when the pulse thrust spacecraft applies the i-th pulse thrust i The difference is , i is the number of pulse thrusts.

5. The method according to claim 1, characterized in that In step 2, the on-orbit service orbit change model is: In the formula, [x p ,y p ,z p ] is the position vector at the end time, [ΔV x ,ΔV y ,ΔV z ] is the orbit change velocity increment of the pulse thrust spacecraft when the pulse increment is applied at the initial moment t0 of the mission, that is, the orbit change velocity increment obtained by the pulse thrust spacecraft in the x, y, and z directions in the target orbit coordinate system; is the average orbital angular velocity of the pulse thrust spacecraft, μ is the Earth's gravitational constant, r is the orbital radius of the space debris, Δt = t-t0 is the difference between the current time t and the initial time t0, is the state vector of the pulse thrust spacecraft relative to the space debris at the initial moment, where [x p (t0),y p (t0),z p (t0)] is the position vector at the initial moment, is the velocity vector at the initial moment; Δt i =tt i is the current time t and the control time t when the pulse thrust spacecraft applies the i-th pulse thrust i The difference is , i is the number of pulse thrusts.

6. The method according to claim 1, characterized in that In step 3, the final state distance constraint is: r1<r pe <r2; Among them, r pe is the relative distance between the pulse thrust spacecraft and the space debris, r1 is the shortest distance between the pulse thrust spacecraft and the space debris, and r2 is the farthest distance between the pulse thrust spacecraft and the space debris; The illumination angle constraint of the terminal state is: θ1<θ<θ2; where θ1 is the lower limit of the illumination angle, θ2 is the upper limit of the illumination angle, and θ is the illumination angle, that is, the angle between the vector of the space debris pointing to the sun and the vector of the space debris pointing to the pulse thrust spacecraft; The track change speed increment constraint is: Where, ΔV x , ΔV y , ΔV z They represent the orbit change velocity increments obtained by the pulse thrust spacecraft in the x, y, and z directions in the target orbit coordinate system, ΔV Represents the total orbit change velocity increment of the pulse thrust spacecraft.

7. The method according to claim 1, characterized in that In step 4, all the terminal time position vectors that meet the terminal state distance constraint are input into the on-orbit service orbit change model, and the method of outputting the orbit change velocity increment includes: The orbital parameter range of the pulse thrust spacecraft is set, with the space debris at the termination time as the origin, the azimuth angle, pitch angle and distance of the pulse thrust spacecraft relative to the space debris as the step size, and uniform sampling is performed on all the termination time position vectors that meet the termination state distance constraint. All sampling points are input into the on-orbit service orbit change model, and the orbit change velocity increment is output.

8. The method according to claim 1, characterized in that In step 4, the method of screening and retaining the track change speed increments that meet both the terminal state illumination angle constraint and the track change speed increment constraint to obtain the track change speed increment solution set includes: Eliminate the track change speed increment that does not meet the illumination angle constraint of the terminal state; Calculate the orbit change speed increment ΔV obtained by each orbit change speed increment in the remaining orbit change speed increment in the three directions of x, y, and z of the target orbit coordinate system x , ΔV y , ΔV z The total track change speed increment is obtained by summing up, and the total track change speed increment exceeding the threshold is excluded to obtain the track change speed increment solution set.

9. The method according to claim 1, characterized in that The fifth step is to construct a circular area with the space debris as the center and a radius that satisfies the distance constraint of the terminal state; the circular area is used to display the relative distance r between the pulse thrust spacecraft and the space debris corresponding to each orbit change velocity increment in the orbit change velocity increment solution set. pe ; When the pulse thrust spacecraft reaches the circular area at the specified time, the mission is marked as successful.

10. The method according to claim 9, characterized in that The circular area whose radius satisfies the end state distance constraint is: A first circular area with a radius of r1 being the closest distance between the pulse thrust spacecraft and the space debris, and a second circular area with a radius of r2 being the farthest distance between the pulse thrust spacecraft and the space debris, the first circular area and the second circular area are concentric circles.

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