An orbit design method for continuous rendezvous of multiple spacecrafts on the same orbit surface

Abstract

The application relates to an orbit design method for continuous rendezvous of multiple spacecrafts on the same orbit plane, comprising the following steps: setting a task scene, the task scene being that a service satellite enters a service orbit to continuously rendezvous with n target satellites on the orbit plane of constellation satellites; determining orbit parameters of the service satellite according to the task scene, the orbit parameters of the service satellite comprising a perigee geocentric distance of the service orbit, a semi-major axis of the service orbit and an apogee geocentric distance; and determining a velocity increment and a time of the service satellite entering the service orbit. The application provides an orbit design method for continuous rendezvous of multiple spacecrafts on the same orbit plane, which firstly considers constellation configuration constraints, sets a continuous rendezvous task scene in the environment of uniformly distributed spacecrafts on the same orbit plane, and then considers environmental interference, orbit perturbation and other influencing factors, so that the number of orbit transfer and the velocity increment are reduced by designing the orbit parameters of the service satellite and the starting point, and single-satellite continuous high-precision rendezvous with multiple spacecrafts on the same orbit plane is realized.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of orbit rendezvous, and particularly relates to an orbit design method for continuous rendezvous of multiple space vehicles on the same orbit plane. BACKGROUND

[0002] Batch deployment and large-scale application of satellite constellations promote the development of space missions such as communication, navigation, remote sensing, etc., but high-frequency and high-density deployment and failure of satellites off-orbit have brought an impact on the operation safety of space assets that cannot be ignored, and have become an important factor restricting the development of the aerospace field. In order to ensure that the satellite constellation works stably in the complex space environment for a long time, it is necessary to observe and maintain the target spacecraft in orbit in real time.

[0003] Traditional methods often use a single satellite service platform to perform space maintenance on a single satellite, and the technology is mature and effective, but when the number of satellites to be observed and maintained is huge, the cost of observation and maintenance is too high. For the multi-spacecraft rendezvous scene, in related technologies, the task planning technology for multi-spacecraft rendezvous docking is mainly studied, and time, energy and other constraints are taken as optimization objectives, and the task planning problem is converted into a dynamic traveling salesman problem for solving, or a heuristic algorithm such as genetic algorithm is used for calculation and solving. The calculation method is complex and the calculation efficiency is low. SUMMARY

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

[0005] The orbit design method for continuous rendezvous of multiple space vehicles on the same orbit plane provided by the embodiments of the present application comprises the following steps:

[0006] A task scene is set, and the task scene is that a service satellite enters a service orbit to perform continuous rendezvous on n target satellites on the orbit plane of the satellite constellation;

[0007] Orbit parameters of the service satellite are determined according to the task scene, and the orbit parameters of the service satellite include a service orbit perigee distance r sp , a service orbit semi-major axis a s , and an apogee distance r sa ;

[0008] The velocity increment and time of the service satellite entering the service orbit are determined.

[0009] In one embodiment, the current orbit of the service satellite is the same as the orbit of the satellite constellation in orbit inclination and right ascension of ascending node.

[0010] In an embodiment, the orbit parameters of the service satellite include a service orbit perigee r sp , a service orbit semi-major axis a s , and a service orbit apogee r sa .

[0011] The service orbit perigee r sp of the service satellite is determined as the constellation satellite orbit semi-major axis a sp .

[0012] The service orbit semi-major axis a s of the service satellite is determined as:

[0013] The calculation formula is:

[0014] wherein P s is a service orbit period; μ = 3.986004418 × 10

[0015] The service orbit apogee r sa of the service satellite is determined as:

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

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

[0018] wherein P O is a target satellite period, and n is a number of target satellites.

[0019] In an embodiment, the determination of the velocity increment of the service satellite entering the service orbit includes:

[0020] A transfer orbit is designed by using the Hohmann transfer method, the transfer orbit is coplanar with the service orbit, the service satellite orbit radius is used as the perigee r tp of the transfer orbit, and the service orbit apogee is used as the apogee r ta of the transfer orbit.

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

[0022]

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

[0024]

[0025] In an embodiment, the determining the time of the service satellite entering the service orbit comprises:

[0026] At the first time of the rendezvous, the time of the constellation satellite running from the current position to the rendezvous position is equal to the time of the service satellite running to the rendezvous position, wherein the time of the service satellite running to the rendezvous position is the time of the service satellite maneuvering from the current orbit to the apogee of the service orbit plus half of the period of the service orbit.

[0027] In an embodiment, the method for designing the orbit of the multi-spacecraft continuous rendezvous in the same orbit plane further comprises: correcting the right ascension of the ascending node of the service orbit;

[0028] The correcting the right ascension of the ascending node of the service orbit comprises: calculating the velocity increment required for the correction according to the orbit elements of the service satellite.

[0029] In an embodiment, when only the right ascension of the ascending node is corrected without affecting the inclination of the orbit, the velocity increment required for the correction is:

[0030]

[0031] In the formula, ΔΩ is the correction amount of the right ascension of the ascending node of the service orbit, u * is the intersection of the new orbit and the original orbit at the point with the latitude amplitude u * , ω is the argument of perigee, p is the semi-major axis of the orbit, and u * = 90° or 270° is the orbit control at the most northern or southern end of the orbit.

[0032] In an embodiment, the method for designing the orbit of the multi-spacecraft continuous rendezvous in the same orbit plane further comprises: controlling the rendezvous accuracy;

[0033] The controlling the rendezvous accuracy comprises: iteratively solving the velocity increment satisfying the constraint condition by the Newton-Raphson method with the relative position of the service satellite and the constellation satellite in the rendezvous process as the constraint condition.

[0034] In an embodiment, the calculation formula of the relative position deviation in the rendezvous 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 amount and the terminal state, g(x) is the error equation, and g′(xk ) is a Jacobian matrix.

[0037] The technical scheme provided by the embodiment of the application has the beneficial effects that include:

[0038] The application provides a method for orbit design for continuous rendezvous of multiple spacecrafts on the same orbit plane, which first considers constellation configuration constraints, and sets a continuous rendezvous task scenario under the environment of uniformly distributed spacecrafts on the same orbit plane; then, in the case of deviation caused by environmental interference, orbit perturbation and other factors, the orbit parameters of the service satellite and the starting point are designed to reduce the number of orbit transfers and the speed increment, reduce the requirements on the attitude and orbit control system of the service satellite, periodically generate rendezvous opportunities, facilitate the design of the on-board mission planning system, and meet the observation, maintenance and other tasks of all targets on a certain orbit plane of the constellation supported by a small number of service satellites. BRIEF DESCRIPTION OF DRAWINGS

[0039] In order to more clearly illustrate the technical solutions in the embodiments of the application, the following will briefly introduce the drawings needed to be used in the embodiment description. Obviously, the drawings in the following description are only some embodiments of the application, and other drawings can be obtained by those skilled in the art without creative effort.

[0040] Figure 1 The figure is a flowchart of the orbit design method for continuous rendezvous of multiple spacecrafts on the same orbit plane in an embodiment of the application.

[0041] Figure 2 The figure is a schematic diagram of on-orbit rendezvous of multiple spacecrafts on the same orbit plane in an embodiment of the application.

[0042] Figure 3 The figure is a schematic diagram of the perigee period and the node period in an embodiment of the application.

[0043] Figure 4 The figure is an iterative flowchart of the semi-major axis of the service orbit in an embodiment of the application.

[0044] Figure 5 The figure is a flowchart of controlling the rendezvous accuracy in an embodiment of the application. DETAILED DESCRIPTION

[0045] In order to make the personnel in the technical field better understand the application scheme, the technical solutions in the embodiments of the application will be described clearly and completely in the following with reference to the drawings in the embodiments of the application. Obviously, the described embodiments are only some of the embodiments of the application, not all. Based on the embodiments in the application, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the application.

[0046] As Figure 1 and Figure 2 shown, wherein, Figure 1 is a flow chart of the orbit design method for the continuous rendezvous of multiple spacecrafts on the same orbit plane in an embodiment of the present application. Figure 2 is a schematic diagram of the on-orbit rendezvous of multiple spacecrafts on the same orbit plane in an embodiment of the present application.

[0047] The embodiment provides an orbit design method for the continuous rendezvous of multiple spacecrafts on the same orbit plane, comprising the following steps:

[0048] Step S1, setting a task scenario, the task scenario being that a service satellite enters a service orbit to perform continuous rendezvous on n target satellites on the orbit plane of a satellite constellation;

[0049] Step S2, determining orbit parameters of the service satellite according to the task scenario, the orbit parameters of the service satellite including a perigee distance from the earth center r sp , a semi-major axis a s of the service orbit, and an apogee distance from the earth center r sa ;

[0050] Step S3, determining a velocity increment and a time of the service satellite entering the service orbit.

[0051] The orbit design method for the continuous rendezvous of multiple spacecrafts on the same orbit plane provided by the embodiment firstly considers constellation configuration constraints, sets a continuous rendezvous task scenario in the environment of uniformly distributed spacecrafts on the same orbit plane, and then, in the case of deviations caused by environmental interference, orbit perturbation and other factors, designs the orbit parameters of the service satellite and the starting point, so as to reduce the number of orbit transfers and the velocity increment, reduce the requirements on the attitude and orbit control system of the service satellite, periodically appear the rendezvous opportunities, facilitate the design of the on-board mission planning system, and meet the observation, maintenance and other tasks of all targets on a certain orbit plane of a satellite constellation supported by a small number of service satellites.

[0052] The steps are described and explained in detail as follows.

[0053] Please refer to Figure 3 , Figure 3 is a schematic diagram of the perigee period and the node period in an embodiment of the present application.

[0054] The present application provides the basic principle of the method:

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

[0056] Since the service satellite and the constellation satellites are in the same orbital plane, their orbital inclinations and right ascension of the ascending node are the same. Because giant constellation satellites are often distributed with equal phase within the same orbital plane, to achieve rendezvous between a single service satellite and all constellation satellites within that plane, the time it takes for the service satellite to complete one orbit must be an integer multiple of the time it takes for the constellation satellites to complete one phase during in-orbit rendezvous. Let the orbital period of the service satellite be 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: In the formula, k is an integer and coprime to n. To shorten the time for continuous rendezvous between the service satellite and satellites on the same orbital plane, the value of k should not be too large. The orbital period for a Kepler orbit (i.e., an unperturbed orbit) 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 a service satellite are determined by designing its orbital period. Satellites are affected by the space environment; for low-Earth orbit satellites, the J2 perturbation is the primary influencing factor. When considering the J2 perturbation, the orbit is not a closed curve, thus different definitions of the orbital period are possible. For example... Figure 3 As shown, P1 and P2 are the perigees connecting two orbits, and B1 and B2 are the ascending nodes connecting two consecutive orbits. The time from B1 to B2 is called the nodal period, denoted by P. nod This represents the most practically significant and clearly measurable period. The latitudinal argument u = ω + θ (ω is the argument of perigee, θ is the true anomaly) changes from 0 to 2π, and the mean latitudinal argument H = M + ω (ω is the argument of perigee, M is the mean anomaly) also changes from 0 to 2π. The corresponding calculation formulas are:

[0058]

[0059] In the formula, R E Let be the Earth's average radius, 'a' be the semi-major axis of the orbit, 'e' be the orbital eccentricity, 'i' be the orbital inclination, and μ = 309860036 × 10⁻⁶. 14 m 3 / s 2 This is the Earth's gravitational constant. This improves the accuracy of calculating the orbital elements of service satellites.

[0060] At the same time, term J2 will also cause a change in the right ascension Ω of the ascending node, with an average rate of change of:

[0061]

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

[0063] Step S1: Set the mission scenario. The mission scenario is that the service satellite enters the service orbit from the current orbit and continuously rendezvous with n target satellites on the orbital plane of the constellation satellites.

[0064] Specifically, for example: suppose there are 18 satellites of equal phase distribution at an altitude of 550km in a constellation in space that need to be maintained regularly. Considering the space safety distance, the service satellites are in the same orbital plane at an altitude of 500km when they are on standby. Conduct orbit design for the service satellites.

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

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

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

[0068] Step S21: Determine the perigee distance r of the service satellite's service orbit. sp Using the semi-major axis of the constellation satellite orbit as the geocentric distance r of the perigee of the service orbit. 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 The service orbital period; μ = is the Earth's gravitational constant.

[0072] Specifically, according to the principle of the present application, it is known that when the J2 perturbation is considered, 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 convergence speed of the iteration, the semi-major axis of the Kepler orbit is taken as the initial value of the iteration, and the iteration process is as shown in the following formula (1). Figure 4

[0073] In an embodiment, the calculation formula of the service orbit period P s is as follows:

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

[0075] The calculation formula of the intersection period P nod is as follows:

[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 eccentricity of the orbit, i is the inclination of the orbit, and μ = 309860036 × 10 14 m 3 / s 2 is the earth's gravitational constant.

[0078] Step S23, determining the service orbit apogee distance from the center of the earth r sa ,

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

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

[0081] In an embodiment, in step S3, the speed increment of the service satellite entering the service orbit includes:

[0082] The Hohmann transfer method is used to design the transfer orbit, the transfer orbit is coplanar with the service orbit, the semi-major axis of the service satellite orbit is taken as the perigee distance from the center of the earth r tp , and the apogee of the service orbit is taken as the apogee distance from the center of the earth r ta .

[0083] The speed increment of the service satellite entering the transfer orbit from the current orbit is:

[0084]

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

[0086]

[0087] Through the above scheme, the energy consumption is minimum by using the Hohmann transfer method to design the transfer orbit.

[0088] In an embodiment, the time when the service satellite enters the service orbit is determined in step S3.

[0089] At the first time of the rendezvous, the time for the constellation satellite to run from the current position to the rendezvous position is equal to the time for the service satellite to run to the rendezvous position, wherein the time for the service satellite to run to the rendezvous position is the time for the service satellite to maneuver from the current orbit to the far point of the service orbit plus half of the period of the service orbit.

[0090] Through the above scheme, the time for the service satellite to run to the rendezvous position is the time for the service satellite to maneuver from the current orbit to the far point of the service orbit plus half of the period of the service orbit, which meets the first time rendezvous constraint of the target orbit constellation satellite and the service satellite, that is, the time when the service satellite starts to maneuver is determined.

[0091] In an embodiment, the orbit design method for the multiple spacecraft continuous rendezvous facing the same orbit plane further comprises: 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 comprises: calculating the velocity increment required for correction deviation according to the orbit elements of the service satellite.

[0093] Due to the semi-major axis deviation, the rate of change of the right ascension of the ascending node is inconsistent when the spacecrafts rendezvous in the same orbit, which leads to the inconsistency of the orbit planes during the multiple spacecraft rendezvous, and the rendezvous accuracy may be deviated. Through the above scheme, the right ascension deviation of the ascending node is corrected before the next rendezvous, and the velocity increment required for correction deviation is calculated according to the orbit elements of the service satellite, thereby improving the rendezvous accuracy.

[0094] In an embodiment, when only the right ascension of the ascending node is corrected without affecting the orbit inclination, the velocity increment required for correction deviation is

[0095]

[0096] In the formula, ΔΩ is the correction amount of the right ascension of the ascending node of the service orbit, u * is the intersection of the new orbit and the original orbit at the latitude amplitude u * , ω is the perigee angle distance, p is the orbit half-pitch, and u * = 90° or 270° is to perform orbit control at the northernmost or southernmost end of the orbit.

[0097] By the above scheme, the deviation caused by the drift of ascending node is corrected, and general fine requirements of task constraints can be met.

[0098] In an embodiment, the orbit design method for the continuous rendezvous of multiple spacecrafts on the same orbit plane further comprises: step S5, controlling rendezvous accuracy;

[0099] Step S5, controlling rendezvous accuracy comprises: taking the relative position of the service satellite and the constellation satellite in the rendezvous process as a constraint condition, and iteratively solving the velocity increment meeting the constraint condition by the Newton-Raphson method.

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

[0101] In an embodiment, the calculation formula of the relative position deviation in the rendezvous process is:

[0102] Δg(x)=f(x)-b=0;

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

[0104] Taking the velocity increment of the three-axis direction of the satellite orbit system, the extrapolation time and the like as the control variable, and the relative position of the rendezvous point as the target variable for iteration, the rendezvous accuracy can be effectively improved, and the task requirements of higher accuracy can be met.

[0105] By the above scheme, according to the orbit parameters of the constellation satellite and the rendezvous process constraint condition, the orbit plane is taken as a unit, the orbit parameters of the service satellite and the orbit rendezvous accuracy are designed in detail, and the single satellite is realized for the continuous high-precision rendezvous of multiple spacecrafts on the same orbit plane.

[0106] By the above scheme, considering the accuracy requirements of different tasks and the energy of the service satellite, different rendezvous point iteration control algorithms are used to meet various constraints in the rendezvous process, and the adaptability of the method is improved.

[0107] In the description of the present application, it needs to be explained that the terms "upper", "lower" and the like indicate the orientation or positional relationship shown in the drawings, and are only for the convenience of describing the present application and simplifying the description, and do not indicate or imply that the devices or elements referred to must have a particular orientation, be constructed and operated in a particular orientation, and therefore cannot be understood as a limitation on the present application. Unless otherwise explicitly specified and limited, the terms "mounting", "connection", "connection" should be broadly understood, for example, it can be fixedly connected, or it can be detachably connected, or integrally connected; it can be mechanically connected, or it can be electrically connected; it can be directly connected, or it can be indirectly connected through an intermediate medium, or it can be the internal communication of two elements. For those skilled in the art, the specific meaning of the above terms in the present application can be understood according to the specific circumstances.

[0108] It should be noted that in the present application, relational terms such as "first" and "second" and the like are used only to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply that there is any such actual relationship or order between these entities or operations. Moreover, the terms "include", "contain" or any other variant thereof are intended to cover non-exclusive inclusion, so that the process, method, article or device including a series of elements not only includes those elements, but also includes other elements not explicitly listed or inherent to such process, method, article or device. Without more limitations, the element defined by the statement "including a" does not exclude the presence of another identical element in the process, method, article or device including the element.

[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; "and / or" in the text only describes the association relationship of the associated objects, which means that there can be three relationships, for example, A and / or B, which means that there are three cases of A alone, A and B, and B alone. In addition, in the description of the embodiments of the present application, "multiple" means two or more than two.

[0110] The above is only a specific embodiment of the present application, which enables those skilled in the art to understand or implement the present application. Various modifications of 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 these embodiments shown herein, but will conform to the widest scope consistent with the principles and novel features applied herein.

Claims

1. A method for orbit design for continuous rendezvous of multiple spacecraft on the same orbital plane, characterized in that, Includes the following steps: The mission scenario is defined as follows: a service satellite enters a service orbit from its current orbit onto the orbital plane of the constellation satellites. The target satellites will rendezvous continuously; The orbital parameters of the serving satellite are determined based on the mission scenario. These orbital parameters include the perigee distance from the Earth's center to the service orbit. , service track semi-major axis and the distance from the geocentric point of apogee ; Determine the velocity increment and time of the service satellite entering its service orbit; The ascending nodes of each orbit are evenly distributed at equal intervals, and the satellites on each orbital plane are also evenly distributed at equal intervals. The current orbit of the service satellite is the same as the orbit of the constellation satellites in terms of orbital inclination and right ascension of the ascending node, and the service orbit period is... The calculation formula is: ,in, For the target satellite period; Using the semi-major axis of the constellation satellite orbit as the perigee distance of the service orbit ; The semi-major axis of the service track is solved using an iterative method, with the Kepler track semi-major axis as the initial value for the iteration. The steps of the iterative process are as follows: After the iteration process begins, input the initial value. , , ;in, The orbital frequency angle; Calculate eccentricity ; Calculated based on the intersection period formula ;in, The period of the intersection point; Calculate the deviation ; Determine whether If so, output If not, the iteration process ends; otherwise, it will... Input, recalculate eccentricity Where eps is the tolerance error and k is the correction factor; Determine the geocentric distance of the service satellite's apogee. ; The transfer orbit is designed using the Hohmann transition method, with the transfer orbit coplanar with the service orbit. The perigee distance of the transfer orbit is the radius of the service satellite's orbit. The geocentric distance of the apogee of the service orbit is used as the apogee of the transfer orbit. ; The time it takes for the service satellite to reach the rendezvous position is the time it takes for the service satellite to maneuver from its current orbit to the apogee of its service orbit plus half of the service orbit period.

2. The orbit design method for continuous rendezvous of multiple spacecraft on the same orbital plane as described in claim 1, characterized in that, The orbital parameters of the serving satellite are determined according to the mission scenario, including the perigee distance of the serving orbit. , service track semi-major axis and the geocentric distance from the apogee of the service orbit. include: Determine the perigee distance of the service satellite's service orbit. The semi-major axis of the constellation satellite orbit is used as the geocentric distance of the perigee of the service orbit. ; Determine the semi-major axis of the service orbit of the service satellite. , The calculation formula is: ; in, For service orbit cycle; μ The gravitational constant of Earth; Determine the geocentric distance from the apogee of the service orbit of the service satellite. , The calculation formula is: .

3. The orbit design method for continuous rendezvous of multiple spacecraft on the same orbital plane as described in claim 1, characterized in that, The determination of the velocity increment for the service satellite to enter its service orbit includes: The transfer orbit is designed using the Hohmann transition method, and the transfer orbit is coplanar with the service orbit, with the perigee distance of the transfer orbit being the radius of the service satellite's orbit. The apogee of the service orbit is used as the geocentric distance of the apogee of the transfer orbit. ; The velocity increment of the service satellite as it enters the transfer orbit from the current orbit is: ； The velocity increment of the service satellite as it enters the service orbit from the transfer orbit is: 。 4. The orbit design method for continuous rendezvous of multiple spacecraft on the same orbital plane as described in claim 1, characterized in that, The determination of the time when the service satellite enters its service orbit includes: During the first rendezvous, the time it takes for the constellation satellites to travel from their current positions to the rendezvous positions is equal to the time it takes for the service satellites to travel to the rendezvous positions. The time it takes for the service satellites to travel to the rendezvous positions is the time it takes for the service satellites to maneuver from their current orbits to the apogee of their service orbits plus half of the service orbit period.

5. The orbit design method for continuous rendezvous of multiple spacecraft on the same orbital plane as described in claim 1, characterized in that, The orbit design method for continuous rendezvous of multiple spacecraft on the same orbital plane also includes: correcting the right ascension deviation of the ascending node of the service orbit; The correction of the right ascension deviation of the ascending node of the service orbit includes: calculating the velocity increment required to correct the deviation based on the orbital elements of the service satellite.

6. The orbit design method for continuous rendezvous of multiple spacecraft on the same orbital plane as described in claim 5, 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: or ; In the formula, To serve the right ascension correction of the ascending node of the orbit, The new orbit and the original orbit have a latitude angle of 0. The points intersect. Angular distance from perigee For the track semi-major, To enable orbit control at the northernmost or southernmost point of the track.

7. The orbit design method for continuous rendezvous of multiple spacecraft on the same orbital plane as described in claim 1 or 5, characterized in that, The orbit design method for continuous rendezvous of multiple spacecraft on the same orbital plane also includes: controlling rendezvous accuracy; The controlled rendezvous accuracy includes: using the relative positions of the service satellite and the constellation satellites during the rendezvous process as constraints, and iteratively solving for the velocity increments that satisfy the constraints using the Newton-Raphson method.

8. The orbit design method for continuous rendezvous of multiple spacecraft on the same orbital plane as described in claim 7, characterized in that, The formula for calculating the relative position deviation during the rendezvous process is: ; ; In the formula, It is a control variable. For terminal constraints, This represents the mapping relationship between control variables and terminal states. The error equation is... Let X be a Jacobian matrix. k Let X be the current set of control variables. k+1 This is a set of control variables updated using the Newton-Raphson method.

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