A method for constructing and selecting the orbit change velocity increment required for on-orbit service of pulse thrust spacecraft based on CW equation

Abstract

The present invention discloses a method for constructing and screening the orbital transfer velocity increment required for on-orbit servicing of pulsed-thrust spacecraft based on the CW equation, belonging to the field of on-orbit servicing of spacecraft, and solving the problem of low accuracy of the orbital maneuvering and transfer strategy of pulsed-thrust spacecraft during on-orbit servicing; it includes: based on the CW equation, constructing the relative motion equation of the pulsed-thrust spacecraft relative to the space debris; simultaneously inputting the transfer state matrix and the velocity transfer matrix of the orbital transfer into the relative motion equation to obtain the on-orbit servicing orbital transfer model; setting the termination state distance constraint, the termination state illumination angle constraint, and the orbital transfer velocity increment constraint according to the mission requirements; inputting all the position vectors at the termination time that meet the termination state distance constraint into the on-orbit servicing orbital transfer model, and retaining the orbital transfer velocity increments that meet the termination state illumination angle constraint and the orbital transfer velocity increment constraint to obtain the screened set of orbital transfer velocity increments. The present invention improves the accuracy of the orbital maneuvering and transfer strategy of pulsed-thrust spacecraft during on-orbit servicing.

Description

Technical Field

[0001] The present invention belongs to the technical field of on-orbit servicing of spacecraft, and relates to a method for constructing and screening the orbital transfer velocity increment required for on-orbit servicing of impulse-thrust spacecraft based on the CW equation. Background Art

[0002] The problem of on-orbit servicing originated from the actual needs in space missions, and is particularly important in tasks such as space debris removal and repair of failed spacecraft. With the development of space technology and the increase of space activities, the space environment has become increasingly complex, and the safety of space assets and the treatment of space debris have become common challenges in the aerospace field. In these tasks, the servicing spacecraft must be able to precisely perform orbital maneuvers to complete various mission operations. Traditional on-orbit servicing problems are usually described by continuous thrust models and the optimal maneuvering strategies are solved by methods such as differential games. These methods are relatively mature in theory, but in practical applications, since the servicing spacecraft mainly uses impulse 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 methods such as differential games provide solutions to on-orbit servicing problems in theory, most of these methods are based on continuous thrust models, while the orbital maneuvers of actual servicing spacecraft mostly use impulse thrust. This difference between the model and the actual application leads to limitations of existing methods in dealing with the orbital transfer problems of impulse-thrust spacecraft in on-orbit servicing. In addition, existing research mainly focuses on finding the fuel-optimal solution under close distance constraints, while ignoring the impact of the orbital transfer strategy selection on the final relative position of the servicing spacecraft, thus limiting the in-depth understanding of the relative motion law of the servicing spacecraft and the comprehensive optimization of the maneuvering strategy. Therefore, the existing technology has deficiencies in providing feasible orbital maneuvering strategies and ensuring that the servicing spacecraft can complete tasks safely and effectively, especially when considering the actual thrust conditions and the impact of strategy selection on the results. Summary of the Invention

[0004] To solve the technical problem of low accuracy of the orbital maneuver and orbit transfer strategy for pulsed-thrust spacecraft in on-orbit servicing in the prior art, the present invention provides a method for constructing and screening the required orbit transfer velocity increment of a pulsed-thrust spacecraft in on-orbit servicing based on the CW equation. Through mathematical modeling and optimization techniques, the influence of different orbit transfer schemes on the relative position relationship between two space targets is explored. According to the actual conditions of pulsed thrust and combined with the constructed on-orbit servicing orbit transfer model, the required orbit transfer velocity increment of the pulsed-thrust spacecraft in on-orbit servicing is obtained. After screening by actual constraint conditions, the solution set of the orbit transfer velocity increment is obtained. A more flexible and practical orbital maneuver strategy can be designed from the solution set of the orbit transfer velocity increment, providing an orbital maneuver strategy for the pulsed-thrust spacecraft to reach the desired approaching relative state from the current state. This strategy can not only optimize the orbital maneuver of the pulsed-thrust spacecraft but also provide a rough range of maneuver values that meet the mission constraints, determine the value ranges of the required orbit transfer velocity increments of the pulsed-thrust spacecraft in the radial, transverse, and normal directions, improve the efficiency and accuracy of the algorithm, and achieve precise orbital maneuver. Through the present invention, the relative motion of the pulsed-thrust spacecraft in the actual space environment can be better understood and predicted, providing a scientific basis for the orbit design and mission planning of spacecraft.

[0005] The object of the present invention is specifically realized through the following technical solutions:

[0006] The present invention discloses a method for constructing and screening the required orbit transfer velocity increment of a pulsed-thrust spacecraft in on-orbit servicing based on the CW equation. The method includes:

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

[0008] Step 2: Input the orbit transfer state transition matrix and the velocity transfer matrix into the relative motion equation at the same time to obtain the on-orbit servicing orbit transfer model of the relationship between the position vector at the termination moment of the pulsed-thrust spacecraft in the target orbit coordinate system and the orbit transfer velocity increment;

[0009] Step 3: Set the termination state distance constraint, the termination state illumination angle constraint, and the orbit transfer velocity increment constraint according to the mission requirements;

[0010] Step 4: Input all the position vectors at the termination moment that meet the termination state distance constraint into the on-orbit servicing orbit transfer model and output the orbit transfer velocity increment; screen and retain the orbit transfer velocity increments that meet both the termination state illumination angle constraint and the orbit transfer velocity increment constraint to obtain the solution set of the orbit transfer velocity increment.

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

[0012] ;

[0013] Wherein, is the state vector of the pulsed-thrust spacecraft relative to the space debris at the current moment. The state vector includes the position vector and the velocity vector. Among them, is the position vector at the current moment, is the velocity vector at the current moment; P represents the pulsed-thrust spacecraft, t represents the current moment; is the relative motion orbit-transfer state-transition matrix, t 0 represents the initial moment; is the state vector of the pulsed-thrust spacecraft relative to the space debris at the initial moment. Among them, is the position vector at the initial moment, is the velocity vector at the initial moment; i is the number of pulsed thrusts, M is the total number of pulsed thrusts, is the relative motion velocity-transfer matrix, t i represents the control moment when the pulsed-thrust spacecraft applies the i -th pulsed thrust, is the orbit-transfer velocity increment when the pulsed-thrust spacecraft applies the pulsed thrust at t i moment.

[0014] In step 2, the orbit-transfer state-transition matrix is:

[0015] ;

[0016] Among them, is the to-be-input orbit-transfer state-transition matrix, is the average orbital angular velocity of the pulsed-thrust spacecraft, is the Earth's gravitational constant, r is the orbital radius of the space debris, is the current moment t minus the initial moment t 0.

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

[0018] ;

[0019] Among them, is the to-be-input velocity-transfer matrix, is the average orbital angular velocity of the pulsed-thrust spacecraft, is the Earth's gravitational constant, is the orbital radius of the space debris, is the current moment t minus the moment when the pulsed-thrust spacecraft applies the iControl moment of the secondary pulse thrust t i The difference of i is the number of pulse thrusts.

[0020] In step 2, the orbital service orbit transfer model is:

[0021] ;

[0022] In the formula, is the position vector at the termination moment, is the initial moment of the mission t The orbit transfer velocity increment of the pulse thrust spacecraft when the pulse increment is instantaneously applied at 0, that is, the orbit transfer velocity increments obtained by the pulse thrust spacecraft in the three directions of x , y , z in the target orbit coordinate system; is the average orbital angular velocity of the pulse thrust spacecraft, is the gravitational constant of the earth, r is the orbital radius of the space debris, is the current moment t and the initial moment t The difference of 0, is the state vector of the pulse thrust spacecraft relative to the space debris at the initial moment, where, is the position vector at the initial moment, is the velocity vector at the initial moment; is the current moment t and the control moment when the pulse thrust spacecraft applies the i th pulse thrust t i The difference of i is the number of pulse thrusts.

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

[0024] ;

[0025] Among them, is the relative distance between the pulse thrust spacecraft and the space debris, is the closest distance between the pulse thrust spacecraft and the space debris, is the farthest distance between the pulse thrust spacecraft and the space debris.

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

[0027] ;

[0028] Among them, is the lower limit value of the illumination angle, is the upper limit value of the illumination angle, is the illumination angle, that is, the angle between the vector pointing from the space debris to the sun and the vector pointing from the space debris to the pulsed thrust spacecraft.

[0029] In step three, the constraint of the orbit transfer velocity increment is:

[0030] ;

[0031] wherein, respectively represent the orbit transfer velocity increments obtained by the pulsed thrust spacecraft in the three directions of the target orbit coordinate system x , y , z , and represents the total orbit transfer velocity increment of the pulsed thrust spacecraft.

[0032] In step four, the method of inputting all the position vectors at the termination time that meet the termination state distance constraint into the on-orbit service orbit transfer model and outputting the orbit transfer velocity increment includes:

[0033] Set the orbit parameter range of the pulsed thrust spacecraft. Taking the space debris at the termination time as the origin, and using the azimuth angle, pitch angle, and distance of the pulsed thrust spacecraft relative to the space debris as the step sizes, uniformly sample among all the position vectors at the termination time that meet the termination state distance constraint, and input all the sampling points into the on-orbit service orbit transfer model to output the orbit transfer velocity increment.

[0034] In step four, the method of screening and retaining the orbit transfer velocity increments that simultaneously meet the termination state illumination angle constraint and the orbit transfer velocity increment constraint to obtain the solution set of the orbit transfer velocity increment includes:

[0035] Exclude the orbit transfer velocity increments that do not meet the termination state illumination angle constraint;

[0036] Calculate the sum of the orbit transfer velocity increments obtained by each orbit transfer velocity increment in the three directions of the target orbit coordinate system x , y , z among the remaining orbit transfer velocity increments to obtain the total orbit transfer velocity increment, and exclude the total orbit transfer velocity increments that exceed the threshold to obtain the solution set of the orbit transfer velocity increment.

[0037] It further includes step five: constructing a circular region centered on the space debris with a radius that meets the termination state distance constraint; displaying the relative distance between the pulsed thrust spacecraft corresponding to each orbit transfer velocity increment in the solution set of the orbit transfer velocity increment and the space debris through the circular region ; when the pulsed thrust spacecraft reaches the circular region at the specified time, mark the mission as successful.

[0038] wherein, the circular region with a radius that meets the termination state distance constraint is:​

[0039] The radius is the closest distance between the pulsed-thrust spacecraft and the space debris. of the first circular region, and the radius is the farthest distance between the pulsed-thrust spacecraft and the space debris of the second circular region. The first circular region and the second circular region are concentric circles.

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

[0041] By comprehensively analyzing the current state vectors (including position vector and velocity vector) of the pulsed-thrust spacecraft and the target orbit parameters, the present invention deduces the relationship between the position vector at the termination moment and the orbit transfer velocity increment, and constructs an on-orbit service orbit transfer model for describing the orbit transfer of the pulsed-thrust spacecraft. The present invention can provide an optimization scheme for pulsed maneuvers under any conditions, enhance the adaptability and flexibility of the pulsed-thrust spacecraft for on-orbit services, and can quickly formulate an effective maneuver strategy according to specific mission requirements. The on-orbit service orbit transfer model takes into account the actual constraint conditions such as the termination state distance constraint, the termination state illumination angle constraint, and the orbit transfer velocity increment constraint, ensures the accuracy and smoothness of the orbit transfer, realizes the accurate description and dynamic simulation of the motion characteristics of the pulsed-thrust spacecraft, and provides a solid theoretical basis and prediction tool for designing a reasonable on-orbit service mission.

[0042] Combined with the actual conditions of pulsed thrust, the present invention constructs and screens a solution set of orbit transfer velocity increments for designing more accurate and practical orbit maneuver strategies. Different from previous methods, the present invention not only solves the optimal control or a single maneuver strategy, but can calculate the solution set of all orbit transfer velocity increments that meet the conditions under a given scenario. The present invention introduces scenario constraints, solves the orbit transfer amount for each termination state, and screens the feasible solutions according to the termination state distance constraint, the termination state illumination angle constraint, and the orbit transfer velocity increment constraint, providing a set of feasible solution sets under the given constraint conditions. This not only ensures the feasibility of the proposed maneuver strategy in actual operation, but also optimizes the resource utilization efficiency, and provides a clear choice for decision-makers, thereby improving the success rate of the mission; this method not only expands the design space of the orbit maneuver strategy, but also improves the flexibility and practicality of the orbit maneuver.

[0043] Through the method of the present invention, the solution set of orbit transfer velocity increments of all orbit maneuver strategies that meet the conditions can be generated in a short time, which not only improves the immediate decision-making ability of orbit maneuvers, but also provides more choices 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 multiple choices for the pulsed-thrust spacecraft, so that the most suitable strategy can be flexibly selected according to the specific situation in actual operation, thereby improving the success rate and efficiency of the mission. Brief Description of the Drawings

[0044] The present invention will be further described in detail below with reference to the 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 the orbit transfer velocity increment. Specific Embodiments

[0047] Embodiment 1 of the present invention provides a method for constructing and screening the orbit transfer velocity increment required for on-orbit servicing of a pulsed-thrust spacecraft based on the CW equation. The method includes:

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

[0049] Step 2: Input the orbit transfer state transition matrix and the velocity transfer matrix into the relative motion equation simultaneously to obtain an on-orbit servicing orbit transfer model that shows the relationship between the position vector at the termination time and the orbit transfer velocity increment of the pulsed-thrust spacecraft in the target orbit coordinate system;

[0050] Step 3: Set the termination state distance constraint, the termination state illumination angle constraint, and the orbit transfer velocity increment constraint according to the mission requirements;

[0051] Step 4: Input all the position vectors at the termination time that meet the termination state distance constraint into the on-orbit servicing orbit transfer model, and output the orbit transfer velocity increment; screen and retain the orbit transfer velocity increments that meet both the termination state illumination angle constraint and the orbit transfer velocity increment constraint to obtain the solution set of the orbit transfer velocity increment.

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

[0053] ;

[0054] Wherein, is the state vector of the pulsed-thrust spacecraft relative to the space debris at the current time. The state vector includes a position vector and a velocity vector. Among them, is the position vector at the current time, is the velocity vector at the current time; P represents the pulsed-thrust spacecraft, t represents the current time; is the relative motion orbit transfer state transition matrix, t 0 represents the initial time; is the state vector of the pulsed-thrust spacecraft relative to the space debris at the initial time. Among them, is the position vector at the initial time, is the velocity vector at the initial moment; i is the number of impulse thrusts, M is the total number of impulse thrusts, is the relative motion velocity transfer matrix, t i represents the control moment when the impulse thrust spacecraft applies the i th impulse thrust, is the impulse thrust spacecraft at t i The orbital velocity increment at the moment of applying the impulse thrust.

[0055] In step two, the orbit transfer state matrix is:

[0056] ;

[0057] Among them, is the orbit transfer state matrix to be input, is the average orbital angular velocity of the impulse thrust spacecraft, is the Earth's gravitational constant, r is the orbital radius of the space debris, is the current moment t and the initial moment t The difference from 0.

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

[0059] ;

[0060] Among them, is the velocity transfer matrix to be input, is the average orbital angular velocity of the impulse thrust spacecraft, is the Earth's gravitational constant, is the orbital radius of the space debris, is the current moment t and the impulse thrust spacecraft applies the i th impulse thrust control moment t i The difference, i is the number of impulse thrusts.

[0061] In step two, the on-orbit service orbit transfer model is:

[0062] ;

[0063] In the formula, is the position vector at the termination moment, is the mission initial moment t 0 The orbital velocity increment of the impulse thrust spacecraft when applying the impulse increment instantaneously, that is, the impulse thrust spacecraft in the target orbit coordinate systemx , y , z The velocity increments for orbit transfer obtained in three directions; is the average orbital angular velocity of the pulsed-thrust spacecraft, is the gravitational constant of the Earth, r is the orbital radius of the space debris, is the current time t and the initial time t 0 difference, is the state vector of the pulsed-thrust spacecraft relative to the space debris at the initial time, where, is the position vector at the initial time, is the velocity vector at the initial time; is the current time t and the pulsed-thrust spacecraft applies the i th t i control time of the pulsed thrust i is the number of pulsed thrusts.

[0064] In step three, the termination state distance constraint is:

[0065] ;

[0066] where, is the relative distance between the pulsed-thrust spacecraft and the space debris, is the closest distance of the pulsed-thrust spacecraft from the space debris, is the farthest distance of the pulsed-thrust spacecraft from the space debris.

[0067] In step three, the termination state illumination angle constraint is:

[0068] ;

[0069] where, is the lower limit value of the illumination angle, is the upper limit value of the illumination angle, is the illumination angle, that is, the angle between the vector pointing from the space debris to the sun and the vector pointing from the space debris to the pulsed-thrust spacecraft.

[0070] In step three, the orbit transfer velocity increment constraint is:

[0071] ;

[0072] where, respectively represent the orbit transfer velocity increments obtained by the pulsed-thrust spacecraft in the x , y , z three directions in the target orbit coordinate system, Represents the total delta-v of the pulsed-thrust spacecraft for orbit transfer.

[0073] In step four, for all the position vectors at the termination time that meet the termination state distance constraint, the method of inputting them into the on-orbit servicing orbit transfer model and outputting the delta-v includes:

[0074] Set the range of the orbital parameters of the pulsed-thrust spacecraft. Taking the space debris at the termination time as the origin, use the azimuth angle, pitch angle, and distance of the pulsed-thrust spacecraft relative to the space debris as the step sizes to ensure uniform sampling. Perform uniform sampling among all the position vectors at the termination time that meet the termination state distance constraint, and input all the sampling points into the on-orbit servicing orbit transfer model to output the delta-v.

[0075] In step four, the method of screening and retaining the delta-v that meets both the termination state illumination angle constraint and the delta-v constraint to obtain the set of delta-v solutions includes:

[0076] Exclude the delta-v that does not meet the termination state illumination angle constraint;

[0077] Calculate the sum of the delta-v obtained in each of the three directions of the remaining delta-v in the target orbital coordinate system x 、 y 、 z to obtain the total delta-v, and exclude the total delta-v that exceeds the threshold to obtain the set of delta-v solutions. It also includes step five: Construct a circular region centered on the space debris with a radius that meets the termination state distance constraint; display the relative distance between the pulsed-thrust spacecraft corresponding to each delta-v in the set of delta-v solutions and the space debris through the circular region

[0078] ; When the pulsed-thrust spacecraft reaches the circular region at the specified time, mark the mission as successful. Obtain the orbital maneuver strategy of the pulsed-thrust spacecraft from the delta-v that marks the mission as successful; for example: After adding each single delta-v that marks the mission as successful to the original velocity and propagating it through the existing propagation model, the orbital maneuver strategy of the pulsed-thrust spacecraft can be obtained. The present invention does not improve the existing propagation method and will not repeat it here.

[0079] Select the orbital maneuver strategy of the pulsed-thrust spacecraft that meets the predefined criteria as the optimal orbital maneuver strategy. The predefined criteria are defined based on the mission requirements, including but not limited to selecting the orbital maneuver strategy of the pulsed-thrust spacecraft corresponding to the minimum delta-v or the shortest transfer time among the orbital maneuver strategies of the pulsed-thrust spacecraft that mark the mission as successful.

[0080]

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

[0082] The radius is the closest distance between the pulse thrust spacecraft and the space debris. The first circular area, and the radius is the maximum distance of the pulse thrust spacecraft from the space debris 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. O It is fixed to the center of mass of the space debris and follows its orbit. x Axis and space debris r coincide, y The axis is perpendicular to the orbital plane x Axis, pointing in the direction of motion is positive, z The axis is determined by the right-hand rule, that is, along the direction of the orbital momentum, and the target orbital coordinate system is as follows Figure 1 As shown:

[0085] In a task, at a certain initial moment t 0=0, in an orbit with an altitude of 35779km, there are pulse thrust spacecraft and space debris, 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. t f =10h later, it approaches the space debris, requiring the end state distance constraint to be in the range of 10km~50km, and the end state illumination angle constraint to be 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] ;

[0090] ;

[0091] ;

[0092] According to the orbit service orbit transfer model, for each terminal state that meets the constraints, substitute it into the model to solve the initial orbit transfer velocity increment. All maneuvering schemes in this scenario can be obtained. This step is crucial because it directly relates to whether a solution that meets the mission requirements can be found. In this way, the orbit transfer velocity increment required to reach a specific terminal position can be accurately calculated.

[0093] During the solution process, only the initial orbit transfer velocity 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 transfer velocity increments that meet the constraints are finally selected, obtaining the solution set of the orbit transfer velocity increment, which represents the possible paths for mission success in the given scenario.

[0094] By establishing two circles centered on the space debris (with radii of 10 km and 50 km respectively), the arrival range of the pulsed-thrust spacecraft when meeting the constraints can be visually displayed. When the pulsed-thrust spacecraft reaches within these ranges at a given moment, it is considered that the mission is successful. As Figure 2 shown, the termination points of different colors represent different orbit transfer velocity increments, and this way can provide an intuitive way to identify and select the optimal maneuvering strategy.

[0095] Through this example, all the orbit transfer velocity increments of the pulsed-thrust spacecraft approaching the space debris 10 hours later under the given initial conditions can be successfully calculated. This method not only simplifies the calculation process of orbit maneuvering but also ensures the accurate achievement of the mission objectives. The experimental results show that the method of the present invention can effectively solve complex orbit maneuvering problems, ensure that the pulsed-thrust spacecraft can approach the space debris within the specified time, and complete the mission under the constraints. In addition, through the orbit service orbit transfer model, the present invention provides a more direct calculation means in actual operation, significantly improving the success rate and efficiency of the mission.

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

[0097] The present invention comprehensively analyzes the current state vector (including the position vector and velocity vector) of a pulsed-thrust spacecraft and the target orbit parameters, derives the relationship between the position vector at the termination moment and the orbit-transfer velocity increment, and constructs an on-orbit service orbit-transfer model for describing the orbit transfer of a pulsed-thrust spacecraft. The present invention can provide an optimization scheme for pulsed maneuvers under any conditions, enhance the adaptability and flexibility of the on-orbit service of pulsed-thrust spacecraft, and can quickly formulate an effective maneuver strategy according to specific mission requirements. The on-orbit service orbit-transfer model takes into account the actual constraint conditions such as the termination-state distance constraint, the termination-state illumination angle constraint, and the orbit-transfer velocity increment constraint, ensures the accuracy and smoothness of orbit transfer, realizes the accurate description and dynamic simulation of the motion characteristics of pulsed-thrust spacecraft, and provides a solid theoretical basis and prediction tool for designing reasonable on-orbit service missions.

[0098] The present invention combines the actual conditions of pulsed thrust to construct and screen a solution set of orbit-transfer velocity increments for designing more accurate and practical orbit maneuver strategies. Different from previous methods, the present invention not only solves the optimal control or a single maneuver strategy, but can calculate the solution set of all orbit-transfer velocity increments that meet the conditions in a given scenario. The present invention introduces scenario constraints, solves the orbit-transfer amount for each termination state, and screens the feasible solutions according to the termination-state distance constraint, the termination-state illumination angle constraint, and the orbit-transfer velocity increment constraint, providing a set of feasible solution sets under given constraint conditions. This not only ensures the feasibility of the proposed maneuver strategy in actual operation, but also optimizes the resource utilization efficiency, and provides a clear choice for decision-makers, thus improving the success rate of the mission; this method not only expands the design space of orbit maneuver strategies, but also improves the flexibility and practicality of orbit maneuvers.

[0099] By the method of the present invention, the solution set of orbit-transfer velocity increments of all orbit maneuver strategies that meet the conditions can be generated in a short time, which not only improves the instant decision-making ability of orbit maneuvers, but also provides more choices 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 multiple choices for pulsed-thrust spacecraft, so that the most suitable strategy can be flexibly selected according to specific situations in actual operation, thus improving the success rate and efficiency of the mission.

[0100] The above is only the specific implementation manner 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, and all should be covered by the protection scope of the present invention. Therefore, the protection scope of the present invention should be subject to the protection scope of the claimed rights.

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, and the state vector includes a position vector and a velocity vector, wherein, is the current position vector, is the velocity vector at the current moment; P represents a pulse thrust spacecraft, t Indicates the current moment; is the relative motion change state transfer matrix, t 0 represents the initial time; is the state vector of the pulse thrust spacecraft relative to the space debris at the initial moment, where, is the initial position vector, is the velocity vector at the initial moment; i is the number of pulse thrusts, M is the total number of pulse thrusts, is the relative motion velocity transfer matrix, t i Indicates the impulse thrust spacecraft applies i The control moment of the secondary pulse thrust, For pulse thrust spacecraft 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: ； in, is the orbit change state transfer matrix to be input, is the average orbital angular velocity of the pulse thrust spacecraft, is the Earth's gravitational constant, r is the orbital radius of the space debris, For the current moment t With the initial moment The difference.

4. The method according to claim 1, characterized in that In step 2, the velocity transfer matrix is: ； in, is the velocity transfer matrix to be input, is the average orbital angular velocity of the pulse thrust spacecraft, is the Earth's gravitational constant, r is the orbital radius of the space debris, For the current moment t With pulse thrust spacecraft exerting i Control time of sub-pulse thrust t i The difference, 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, is the position vector at the end time, The initial time of the task t The orbit change velocity increment of the pulse thrust spacecraft when the pulse increment is applied at the instant 0, that is, the velocity increment of the pulse thrust spacecraft in the target orbit coordinate system x , y , z The track change speed increments obtained in three directions; is the average orbital angular velocity of the pulse thrust spacecraft, is the Earth's gravitational constant, r is the orbital radius of the space debris, For the current moment t With the initial moment t The difference of 0, is the state vector of the pulse thrust spacecraft relative to the space debris at the initial moment, where, is the initial position vector, is the velocity vector at the initial moment; For the current moment t With pulse thrust spacecraft exerting i Control time of sub-pulse thrust t i The difference, 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: ;in, is the relative distance between the pulse thrust spacecraft and the space debris, is the shortest distance between a pulse thrust spacecraft and space debris, is the maximum distance between a pulse thrust spacecraft and space debris; The final state illumination angle constraint is: ;in, is the lower limit of the illumination angle, is the upper limit of the illumination angle, is the illumination angle, i.e., 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: ;in, They represent the pulse thrust spacecraft in the target orbit coordinate system. x , y , z The track change speed increments obtained in three directions are: 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 coordinates of each orbit change speed increment in the target orbit coordinate system x , y , z The track change speed increments obtained in three directions 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 between the pulse thrust spacecraft and the space debris corresponding to each orbit change velocity increment in the orbit change velocity increment solution set. r 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: The radius is the closest distance between the pulse thrust spacecraft and the space debris. The first circular area, and the radius is the maximum distance of the pulse thrust spacecraft from the space debris The first circular area and the second circular area are concentric circles.

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