Orbit design method for continuous intersection of multiple spacecrafts on same orbital plane

Abstract

The invention relates to an orbit design method for continuous rendezvous of multiple spacecrafts on the same orbit plane, which comprises the following steps: setting a task scene that a service satellite enters a service orbit from a current orbit to continuously rendezvous n target satellites on an orbit plane of a constellation satellite; determining orbit parameters of a service satellite according to the task scene, wherein the orbit parameters of the service satellite comprise a service orbit perigee geocentric distance, a service orbit semi-major axis and an apogee geocentric distance; and determining the speed increment and time for the service satellite to enter the service orbit. The invention provides an orbit design method for continuous rendezvous of multiple spacecrafts on the same orbit plane, which comprises the following steps of: firstly, setting a continuous rendezvous task scene in an environment of uniformly distributed spacecrafts in the same orbit plane by considering constellation configuration constraint; influence factors such as environment interference and orbit perturbation are considered, orbit transfer times and speed increment are reduced through design of service satellite orbit parameters and starting point positions, and continuous high-precision rendezvous of multiple spacecrafts in the same orbit plane by a single satellite is achieved.

Description

Technical Field

[0001] The invention relates to the technical field of orbital rendezvous, and in particular to an orbital design method for continuous rendezvous of multiple spacecraft on the same orbital plane. Background Art

[0002] The mass deployment and large-scale application of satellite constellations have promoted the development of space missions such as communications, navigation, and remote sensing. However, the high-frequency and high-density deployment of satellites and their failure and deorbiting have brought an unignorable impact on the operational safety of space assets, becoming an important factor restricting the development of the aerospace field. In order to ensure the long-term and stable operation of satellite constellations in complex space environments, real-time observation and on-orbit maintenance of target spacecraft are required.

[0003] Traditional methods often use a single satellite service platform to perform space maintenance on a single satellite. The technology is highly mature and effective, but when a large number of satellites need to be observed and maintained, the cost of observation and maintenance is too high. For multi-spacecraft rendezvous scenarios, the relevant technologies mainly focus on the mission planning technology for multi-spacecraft rendezvous and docking. With time, energy and other constraints as optimization goals, the mission planning problem is converted into a dynamic traveling salesman problem for solution, or heuristic algorithms such as genetic algorithms are used for calculation and solution. The calculation method is complex and the calculation efficiency is low. Summary of the invention

[0004] The technical problem to be solved by the present application is that the calculation method for multi-spacecraft rendezvous scenarios in the relevant technology is complex and has low calculation efficiency.

[0005] The present application provides a method for designing a trajectory for continuous rendezvous of multiple spacecraft on the same orbital plane, comprising the following steps:

[0006] Setting a mission scenario, wherein the mission scenario is that the service satellite enters the service orbit from the current orbit and continuously rendezvouses with n target satellites on the orbital plane of the constellation satellite;

[0007] The orbital parameters of the service satellite are determined according to the mission scenario, wherein the orbital parameters of the service satellite include the perigee distance r of the service orbit sp 、Service orbit semi-major axis a s , and the apogee geocentric distance r sa ;

[0008] The velocity increment and time for the servicing satellite to enter the servicing orbit are determined.

[0009] In one embodiment, the current orbit of the service satellite is identical to the orbit of the constellation satellite in terms of orbital inclination and ascending node right ascension.

[0010] In one embodiment, the orbital parameters of the service satellite are determined according to the mission scenario, and the orbital parameters of the service satellite include the perigee distance r of the service orbit. sp 、Service orbit semi-major axis a s , and the geocentric distance r of the apogee of the service orbit sa include:

[0011] Determine the service orbit perigee distance r of the service satellite sp , taking the semi-major axis of the constellation satellite orbit as the perigee geocentric distance r of the service orbit sp ;

[0012] Determine the semi-major axis a of the service orbit of the service satellite s ,

[0013] The calculation formula is:

[0014] Among them, P s is the service orbit period; μ = is the Earth's gravitational constant;

[0015] Determine the geocentric distance r of the service orbit apogee of the service satellite sa ,

[0016] The calculation formula is: sa =2a s -r sp .

[0017] In one embodiment, the service orbit period P s The calculation formula is:

[0018] Among them, P O is the target satellite period, and n is the number of target satellites.

[0019] In one embodiment, determining the velocity increment of the service satellite into the service orbit comprises:

[0020] The transfer orbit is designed by the Hohmann transition method. The transfer orbit is coplanar with the service orbit, and the orbit radius of the service satellite is the perigee distance r of the transfer orbit. tp , taking the apogee of the service orbit as the apogee of the transfer orbit, the distance from the geocenter to the apogee r ta ;

[0021] The velocity increment of the service satellite from the current orbit to the transfer orbit is:

[0022]

[0023] The velocity increment of the service satellite from the transfer orbit into the service orbit is:

[0024]

[0025] In one embodiment, determining the time when the service satellite enters the service orbit includes:

[0026] During the first rendezvous, the time taken by the constellation satellite to move from the current position to the rendezvous position is equal to the time taken by the service satellite to move to the rendezvous position, wherein the time taken by the service satellite to move to the rendezvous position is the time taken by the service satellite to maneuver from the current orbit to the apogee of the service orbit plus half of the service orbit period.

[0027] In one embodiment, the orbit design method for continuous rendezvous of multiple spacecraft on the same orbital plane further includes: correcting the right ascension deviation of the ascending node of the service orbit;

[0028] The correcting the right ascension deviation of the ascending node of the service orbit includes: calculating the speed increment required to correct the deviation according to the orbit elements of the service satellite.

[0029] In one embodiment, when only the right ascension of the ascending node is corrected without affecting the orbital inclination, the velocity increment required to correct the deviation is:

[0030]

[0031] Where ΔΩ is the correction value of the ascending node of the service orbit, u * The new orbit and the original orbit have an angle of u at latitude * The point of intersection is ω, ω is the perigee angular distance, p is the orbital semi-diameter, u * = 90° or 270° for orbit control at the northernmost or southernmost end of the orbit.

[0032] In one embodiment, the orbit design method for continuous rendezvous of multiple spacecraft on the same orbital plane further includes: controlling the rendezvous accuracy;

[0033] The controlling rendezvous accuracy includes: taking the relative position of the service satellite and the constellation satellite during the rendezvous process as a constraint condition, and iteratively solving the velocity increment satisfying the constraint condition through the Newton-Raphson method.

[0034] In one implementation, the calculation formula for the relative position deviation during the intersection process is:

[0035]

[0036] In the formula, x is the control variable, b is the terminal constraint, f(x) represents the mapping relationship between the control variable and the terminal state, g(x) is the error equation, and g′(xk ) is the Jacobian matrix.

[0037] The beneficial effects brought by the technical solution provided in the embodiments of the present application include:

[0038] The present application provides an orbit design method for continuous rendezvous of multiple spacecraft on the same orbital plane. First, the constellation configuration constraints are considered to set a continuous rendezvous mission scenario in an environment where spacecraft are evenly distributed on the same orbital plane. Then, considering the deviation caused by influencing factors such as environmental interference and orbital perturbations, the orbit parameters and power-on points of the service satellite are designed to reduce the number of orbit changes and speed increments, thereby reducing the requirements for the attitude and orbit control system of the service satellite. The rendezvous opportunities occur periodically, which is convenient for the design of the on-board mission planning system, so as to meet the requirements of completing the observation and maintenance of all targets on a certain orbital plane of the constellation with the support of a small number of service satellites. BRIEF DESCRIPTION OF THE DRAWINGS

[0039] In order to more clearly illustrate the technical solutions in the embodiments of the present application, the drawings required for use in the description of the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present application. For ordinary technicians in this field, other drawings can be obtained based on these drawings without creative work.

[0040] Figure 1 The present invention is a flowchart of a method for orbit design for continuous rendezvous of multiple spacecraft on the same orbital plane in one embodiment of the present invention.

[0041] Figure 2 It is a schematic diagram of the along-orbit rendezvous of multiple spacecraft facing the same orbital plane in one embodiment of the present invention.

[0042] Figure 3 Schematic diagram of perigee period and nodal period in one embodiment of the present invention.

[0043] Figure 4 FIG. 4 is an iterative flow chart of the semi-major axis of the service orbit in one embodiment of the present invention.

[0044] Figure 5 The flowchart of controlling intersection accuracy in one embodiment of the present invention is shown. DETAILED DESCRIPTION

[0045] In order to enable those skilled in the art to better understand the solution of the present application, the technical solution in the embodiments of the present application will be clearly and completely described below in conjunction with the drawings in the embodiments of the present application. Obviously, the described embodiments are only part of the embodiments of the present application, not all of the embodiments. Based on the embodiments in the present application, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of this application.

[0046] like Figure 1 and Figure 2 As shown, Figure 1 The present invention is a flowchart of a method for orbit design for continuous rendezvous of multiple spacecraft on the same orbital plane in one embodiment of the present invention. Figure 2 It is a schematic diagram of the along-orbit rendezvous of multiple spacecraft facing the same orbital plane in one embodiment of the present invention.

[0047] This embodiment provides a trajectory design method for continuous rendezvous of multiple spacecraft on the same orbital plane, comprising the following steps:

[0048] Step S1, setting a mission scenario, where the mission scenario is that the service satellite enters the service orbit from the current orbit and continuously rendezvouses with n target satellites on the orbital plane of the constellation satellite;

[0049] Step S2: Determine the orbital parameters of the service satellite according to the mission scenario. The orbital parameters of the service satellite include the perigee distance r of the service orbit. sp 、Service orbit semi-major axis a s , and the apogee geocentric distance r sa ;

[0050] Step S3: Determine the velocity increment and time for the service satellite to enter the service orbit.

[0051] This embodiment provides an orbit design method for continuous rendezvous of multiple spacecraft on the same orbital plane. First, the constellation configuration constraints are considered to set a continuous rendezvous mission scenario in an environment where spacecraft are evenly distributed on the same orbital plane. Then, considering the deviation caused by influencing factors such as environmental interference and orbital perturbations, the orbital parameters and power-on points of the service satellite are designed to reduce the number of orbit changes and speed increments, thereby reducing the requirements for the attitude and orbit control system of the service satellite. The rendezvous opportunities appear periodically, which is convenient for the design of the on-board mission planning system, so as to meet the requirements of completing the observation and maintenance of all targets on a certain orbital plane of the constellation with the support of a small number of service satellites.

[0052] Each step is described and explained in detail below.

[0053] Please also refer to Figure 3 , Figure 3 Schematic diagram of perigee period and nodal period in one embodiment of the present invention.

[0054] The basic principle of the method provided in this application:

[0055] A satellite constellation is a collection of satellites that are launched into orbit and can work normally. It is usually a satellite network composed of some satellite rings configured in a certain way. The applicant considers that the satellite constellation configuration presents certain regularities in space. For example, the orbits of giant constellations are often circular orbits with the same orbital altitude and the same inclination to the reference plane. The ascending nodes of each orbit are evenly distributed at equal intervals, and the satellites on each orbital plane (i.e., constellation satellites) are also evenly distributed at equal intervals.

[0056] Since the service satellite and the constellation satellite are in the same orbital plane, the orbital inclination and right ascension of the ascending node are the same between the service satellite and the constellation satellite. Since the satellites of a giant constellation are often distributed in equal phases in the same orbital plane, in order to achieve the rendezvous of a single service satellite with all the constellation satellites in the orbital plane, when performing along-track rendezvous, the time for the service satellite to run one cycle should be an integer multiple of the time for the constellation satellite to run one phase. The orbital period of the service satellite is denoted as P s , the number of satellites in the same orbital plane is n, and the orbital period is P O , they satisfy the following relationship: Where k is an integer and is relatively prime to n. In order to shorten the time for the service satellite to continuously rendezvous with satellites on the same orbital plane, the value of k should not be too large. The orbital period of the Kepler orbit (i.e., the orbit without perturbation) is: Where a is the semi-major axis of the orbit, μ = 309860036 × 10 14 m 3 / s 2 is the Earth's gravitational constant.

[0057] The orbital parameters of the service satellite are determined by designing the orbital period of the service satellite. Satellites are affected by the environment in space. For low-orbit satellites, the J2 perturbation term is the main influencing factor. When the J2 perturbation term is considered, the orbit is not a closed curve, so there are different definitions of the orbital period. Figure 3 As shown, P 1 and P 2 is the perigee connecting the two circles, B 1 and B 2 It is the ascending node of two consecutive circles. 1 To B 2 The time is called the nodal period, denoted by P nod It is the most practical and clearly measurable period. The latitude argument u=ω+θ (ω is the perigee argument, θ is the true anomaly) changes from 0 to 2π, and the mean latitude argument H=M+ω (ω is the perigee argument, M is the mean anomaly) also changes from 0 to 2π. The corresponding calculation formula is:

[0058]

[0059] In the formula, R Eis the average radius of the Earth, a is the semi-major axis of the orbit, e is the orbital eccentricity, i is the orbital inclination, μ=309860036×10 14 m 3 / s 2 is the Earth's gravitational constant. This improves the calculation accuracy of the orbital elements of the service satellite.

[0060] At the same time, the J2 term will also cause the right ascension Ω of the ascending node to change, and the average change rate is:

[0061]

[0062] The unit is rad / s. Due to the difference in the semi-major axis of the orbits of the service satellite and the constellation satellite, the inconsistent change rate of the ascending node of the service satellite and the constellation satellite may cause deviations in the orbital parameters of the service satellite and the constellation satellite, affecting the rendezvous accuracy. Based on the average change rate and time, the right ascension correction of the ascending node can be obtained. Orbital control is used during the rendezvous process to improve the rendezvous accuracy and meet the mission requirements.

[0063] Step S1, setting a mission scenario, where the mission scenario is that the service satellite enters the service orbit from the current orbit and continuously rendezvouses with n target satellites on the orbital plane of the constellation satellite.

[0064] Specifically, for example, if 18 satellites with equal phase distribution at an altitude of 550km in a constellation in space need regular maintenance, and considering the safe distance in space, the service satellite is at an altitude of 500km in the same orbital plane when on standby, the service satellite orbit design is carried out.

[0065] In one embodiment, the current orbit of the serving satellite is identical to the orbit of the constellation satellite in terms of orbital inclination and right ascension of the ascending node.

[0066] The above scheme facilitates the subsequent design of the orbital parameters of the service satellite.

[0067] In one embodiment, step S2, determining the orbital parameters of the service satellite according to the mission scenario, the orbital parameters of the service satellite include the perigee distance r of the service orbit. sp 、Service orbit semi-major axis a s , and the geocentric distance r of the apogee of the service orbit sa include:

[0068] Step S21: Determine the service orbit perigee distance r of the service satellite sp , taking the semi-major axis of the constellation satellite orbit as the distance from the service orbit perigee to the center of the earth sp .

[0069] Step S22: Determine the semi-major axis a of the service orbit of the service satellite s ,

[0070] The calculation formula is:

[0071] Among them, P s is the service orbit period; μ= is the earth's gravitational constant.

[0072] Specifically, according to the principle of this application, when considering the J2 perturbation, there is a small difference between the intersection period and the Kepler orbit period. At this time, the semi-major axis of the service orbit is solved by an iterative method. In order to improve the iterative convergence speed, the semi-major axis of the Kepler orbit is used as the initial value of the iteration. The iterative process is as follows: Figure 4 shown.

[0073] In one embodiment, the service orbit period P s The calculation formula is:

[0074] Among them, P O is the target satellite period, and n is the number of target satellites.

[0075] Intersection period P nod The calculation formula is:

[0076]

[0077] In the formula, R E is the average radius of the Earth, a is the semi-major axis of the orbit, e is the orbital eccentricity, i is the orbital inclination, μ=309860036×10 14 m 3 / s 2 is the Earth's gravitational constant.

[0078] Step S23: Determine the geocentric distance r of the service orbit apogee of the service satellite sa ,

[0079] The calculation formula is: sa =2a s -r sp .

[0080] Through the above scheme, the parameters of the service track are accurately designed.

[0081] In one embodiment, in step S3, determining the velocity increment of the service satellite entering the service orbit includes:

[0082] The Hohmann transition method is used to design the transfer orbit. The transfer orbit is coplanar with the service orbit, and the radius of the service satellite orbit is the perigee distance r of the transfer orbit. tp , taking the apogee of the service orbit as the apogee of the transfer orbit, the distance from the center of the earth ta ;

[0083] The velocity increment of the service satellite from the current orbit to the transfer orbit is:

[0084]

[0085] The velocity increment of the service satellite from the transfer orbit to the service orbit is:

[0086]

[0087] Through the above scheme, the transfer orbit is designed using the Hohmann transition method with minimal energy consumption.

[0088] In one embodiment, in step S3, determining the time when the service satellite enters the service orbit includes:

[0089] During the first rendezvous, the time it takes for the constellation satellite to move from its current position to the rendezvous position is equal to the time it takes for the service satellite to move to the rendezvous position, where the time it takes for the service satellite to move to the rendezvous position is the time it takes for the service satellite to maneuver from its current orbit to the apogee of the service orbit plus half the service orbit period.

[0090] Through the above scheme, the time for the service satellite to move to the rendezvous position is the time for the service satellite to maneuver from the current orbit to the apogee of the service orbit plus half of the service orbit period, which satisfies the first rendezvous constraint between the target orbit constellation satellite and the service satellite, that is, determines the moment when the service satellite starts orbital maneuvering.

[0091] In one embodiment, the orbit design method for continuous rendezvous of multiple spacecraft on the same orbital plane further includes: step S4, correcting the right ascension deviation of the ascending node of the service orbit;

[0092] Step S4, correcting the right ascension deviation of the ascending node of the service orbit includes: calculating the velocity increment required to correct the deviation according to the orbit elements of the service satellite.

[0093] During along-orbit rendezvous, the semi-major axis deviation causes the rate of change of the right ascension of the orbital ascending node to be inconsistent, resulting in inconsistent orbital planes during the rendezvous of multiple spacecraft, and the rendezvous accuracy may deviate. Through the above scheme, the right ascension deviation of the ascending node is corrected before the next rendezvous, and the speed increment required to correct the deviation is calculated according to the orbital elements of the service satellite, thereby improving the rendezvous accuracy.

[0094] In one embodiment, when only the right ascension of the ascending node is corrected without affecting the orbital inclination, the velocity increment required to correct the deviation is

[0095]

[0096] Where ΔΩ is the correction value of the ascending node of the service orbit, u * The new orbit and the original orbit have an angle of u at latitude * The point of intersection is ω, ω is the perigee angular distance, p is the orbital semi-diameter, u *= 90° or 270° for orbit control at the northernmost or southernmost end of the orbit.

[0097] Through the above scheme, the deviation caused by the drift of the ascending node can be corrected to meet the general fine requirements of the mission constraints.

[0098] In one embodiment, the orbit design method for continuous rendezvous of multiple spacecraft on the same orbital plane further includes: step S5, controlling the rendezvous accuracy;

[0099] Step S5, controlling the rendezvous accuracy, includes: taking the relative positions of the service satellite and the constellation satellite during the rendezvous process as constraints, and iteratively solving the velocity increment that satisfies the constraints through the Newton-Raphson method.

[0100] Specifically, the relative position (such as elevation, azimuth, angular velocity, distance, etc.) of the service satellite and the target satellite is observed and maintained with the required stability conditions as the constraint conditions. The Newton-Raphson method does not need to consider the iteration direction, has fast convergence speed, good convergence, simple calculation, and is easy to implement. The specific process is as follows Figure 5 shown.

[0101] In one embodiment, the calculation formula of the relative position deviation during the intersection process is:

[0102] Δg(x)=f(x)-b=0;

[0103] In the formula, x is the control variable, b is the terminal constraint, f(x) represents the mapping relationship between the control variable and the terminal state, g(x) is the error equation, and g′(x k ) is the Jacobian matrix.

[0104] Using the velocity increment in the three-axis directions of the satellite orbit system, the extrapolated time, etc. as the control quantities, and the relative position of the rendezvous point as the target quantity for iteration, the rendezvous accuracy can be effectively improved to meet higher precision mission requirements.

[0105] Through the above scheme, according to the orbital parameters of the constellation satellites and the constraints of the rendezvous process, taking the orbital plane as the unit, through detailed design of the orbital parameters of the service satellite and the orbital rendezvous accuracy, continuous high-precision rendezvous of multiple spacecraft within the same orbital plane by a single satellite can be achieved.

[0106] Through the above scheme, the accuracy requirements of different tasks and the service satellite energy are taken into consideration, and different rendezvous point iterative control algorithms are used to meet the various constraints of the rendezvous process, thereby improving the adaptability of the method.

[0107] In the description of the present application, it should be noted that the terms "upper", "lower", etc. indicate orientations or positional relationships based on the orientations or positional relationships shown in the accompanying drawings, and are only for the convenience of describing the present application and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore cannot be understood as a limitation on the present application. Unless otherwise clearly specified and limited, the terms "installed", "connected", and "connected" should be understood in a broad sense, for example, it can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be a direct connection, or it can be an indirect connection through an intermediate medium, or it can be a connection between the two elements. For those of ordinary skill in the art, the specific meanings of the above terms in the present application can be understood according to the specific circumstances.

[0108] It should be noted that, in this application, relational terms such as "first" and "second" are only used to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Moreover, the terms "include", "comprise" or any other variants thereof are intended to cover non-exclusive inclusion, so that a process, method, article or device including a series of elements includes not only those elements, but also other elements not explicitly listed, or also includes elements inherent to such process, method, article or device. In the absence of further restrictions, the elements defined by the sentence "comprise a ..." do not exclude the presence of other identical elements in the process, method, article or device including the elements.

[0109] In the description of the embodiments of the present application, unless otherwise specified, “ / ” means or, for example, A / B can mean A or B; the “and / or” in the text is merely a description of the association relationship of associated objects, indicating that three relationships may exist, for example, A and / or B can mean: A exists alone, A and B exist at the same time, and B exists alone. In addition, in the description of the embodiments of the present application, “multiple” refers to two or more than two.

[0110] The above is only a specific implementation of the present application, so that those skilled in the art can understand or implement the present application. Various modifications to these embodiments will be apparent to those skilled in the art, and the general principles defined herein can be implemented in other embodiments without departing from the spirit or scope of the present application. Therefore, the present application will not be limited to the embodiments shown herein, but will conform to the widest range consistent with the principles and novel features applied for herein.

Claims

1. A trajectory design method for continuous rendezvous of multiple spacecraft on the same orbital plane, characterized in that: The following steps are involved: Setting a mission scenario, wherein the mission scenario is that the service satellite enters the service orbit from the current orbit and continuously rendezvouses with n target satellites on the orbital plane of the constellation satellite; The orbital parameters of the service satellite are determined according to the mission scenario, wherein the orbital parameters of the service satellite include the perigee distance r of the service orbit sp 、Service orbit semi-major axis a s , and the apogee geocentric distance r sa ; The velocity increment and time for the servicing satellite to enter the servicing orbit are determined.

2. The orbit design method for continuous rendezvous of multiple spacecraft on the same orbital plane as claimed in claim 1, characterized in that: The current orbit of the service satellite is identical to the orbit of the constellation satellite in terms of orbital inclination and ascending node right ascension.

3. The orbit design method for continuous rendezvous of multiple spacecraft on the same orbital plane as claimed in claim 2, characterized in that: The orbital parameters of the service satellite are determined according to the mission scenario, and the orbital parameters of the service satellite include the perigee distance r of the service orbit sp 、Service orbit semi-major axis a s , and the geocentric distance r of the apogee of the service orbit sa include: Determine the service orbit perigee distance r of the service satellite sp , taking the semi-major axis of the constellation satellite orbit as the perigee geocentric distance r of the service orbit sp ; Determine the semi-major axis a of the service orbit of the service satellite s , The calculation formula is: Among them, P s is the service orbit period; μ = is the Earth's gravitational constant; Determine the geocentric distance r of the service orbit apogee of the service satellite sa , The calculation formula is: sa =2a s -r sp .

4. The orbit design method for continuous rendezvous of multiple spacecraft on the same orbital plane as claimed in claim 3, characterized in that: The service orbit period P s The calculation formula is: Among them, P O is the target satellite period, and n is the number of target satellites.

5. The orbit design method for continuous rendezvous of multiple spacecraft on the same orbital plane as claimed in claim 1, characterized in that: Determining the velocity increment of the service satellite entering the service orbit comprises: The transfer orbit is designed by the Hohmann transition method. The transfer orbit is coplanar with the service orbit, and the orbit radius of the service satellite is the perigee distance r of the transfer orbit. tp , taking the apogee of the service orbit as the apogee of the transfer orbit, the distance from the geocenter to the apogee r ta ; The velocity increment of the service satellite from the current orbit to the transfer orbit is: The velocity increment of the service satellite from the transfer orbit into the service orbit is:

6. The orbit design method for continuous rendezvous of multiple spacecraft on the same orbital plane as claimed in claim 1, characterized in that: Determining the time when the service satellite enters the service orbit comprises: During the first rendezvous, the time taken by the constellation satellite to move from the current position to the rendezvous position is equal to the time taken by the service satellite to move to the rendezvous position, wherein the time taken by the service satellite to move to the rendezvous position is the time taken by the service satellite to maneuver from the current orbit to the apogee of the service orbit plus half of the service orbit period.

7. The orbit design method for continuous rendezvous of multiple spacecraft on the same orbital plane as claimed in claim 1, characterized in that: The orbit design method for continuous rendezvous of multiple spacecraft on the same orbital plane further includes: correcting the right ascension deviation of the ascending node of the service orbit; The correcting the right ascension deviation of the ascending node of the service orbit includes: calculating the speed increment required to correct the deviation according to the orbit elements of the service satellite.

8. The orbit design method for continuous rendezvous of multiple spacecraft on the same orbital plane as claimed in claim 7, characterized in that: When only the right ascension of the ascending node is corrected without affecting the orbital inclination, the velocity increment required to correct the deviation is: u * =90°, or * =270°, Where ΔΩ is the correction value of the ascending node of the service orbit, u * The new orbit and the original orbit have an angle of u in latitude. * The point of intersection is ω, ω is the perigee angular distance, p is the orbital semi-diameter, u * = 90° or 270° for orbit control at the northernmost or southernmost end of the orbit.

9. The orbit design method for continuous rendezvous of multiple spacecraft on the same orbital plane as claimed in claim 1 or 7, characterized in that: The orbit design method for continuous rendezvous of multiple spacecraft on the same orbital plane further includes: controlling the rendezvous accuracy; The controlling rendezvous accuracy includes: taking the relative position of the service satellite and the constellation satellite during the rendezvous process as a constraint condition, and iteratively solving the velocity increment satisfying the constraint condition through the Newton-Raphson method.

10. The orbit design method for continuous rendezvous of multiple spacecraft on the same orbital plane as claimed in claim 9, characterized in that: The calculation formula of relative position deviation during the intersection process is: Δg(x) = f(x)-b = 0; In the formula, x is the control variable, b is the terminal constraint, f(x) represents the mapping relationship between the control variable and the terminal state, g(x) is the error equation, and g′(x k ) is the Jacobian matrix.

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