A Distributed Mega-Constellation Maintenance Control Method and System Based on Combinatorial Strategy

Abstract

A distributed mega-constellation maintenance control method and system based on a combined strategy, belonging to the field of spacecraft orbit control technology, solves the problems of high fuel consumption and easy control conflicts in existing collision avoidance methods. The method of this invention includes: integrating multiple objectives such as single-satellite orbit maintenance, in-plane phase maintenance, inter-plane configuration control, and autonomous collision avoidance into a unified framework and resolving conflicts among them; giving collision avoidance the highest decision priority and deeply embedding probabilistic collision risk assessment into the control loop, thereby providing active safety assurance for the constellation in dense orbital environments; optimizing fuel consumption by combining thrust application strategies based on orbital dynamics and phase adjustment strategies for raising the orbit. The design based on the average orbital elements makes it insensitive to high-frequency perturbations, and the control law is stable and reliable. This invention is applicable to distributed autonomous orbit maintenance control of mega-constellations in low Earth orbit, considering autonomous collision avoidance.

Description

Technical Field

[0001] This application relates to the field of spacecraft orbit control technology, and in particular to giant constellation maintenance control technology. Background Technology

[0002] With the rapid development of commercial spaceflight, low-Earth orbit mega-constellations (such as Starlink and OneWeb) have become a research hotspot due to their advantages such as high coverage and low latency. These constellations consist of a large number of satellites (up to thousands or even tens of thousands). Relying entirely on ground stations for orbit maintenance and control would consume enormous human and material resources and would result in poor real-time performance. Therefore, achieving autonomous operation of the constellation is crucial.

[0003] Existing constellation maintenance control methods are mostly designed for traditional small-scale constellations or focus on a single control objective (such as phase maintenance or semi-major axis maintenance). For giant constellations, the control requirements are complex and diverse, including: maintaining the orbital elements (semi-major axis, eccentricity, inclination) of individual satellites, maintaining the phase distribution of satellites within the same orbital plane, maintaining the right ascension (RAAN) spacing between different orbital planes, and most critically, the issue of collision avoidance. These requirements may conflict with each other; for example, the thrust used for phase adjustment may affect the semi-major axis. Currently, there is a lack of a unified control method that can comprehensively consider these diverse requirements and be executed by distributed autonomous decisions from each satellite. Summary of the Invention

[0004] The purpose of this invention is to address the technical problems of existing collision avoidance methods, which typically only perform maneuvers after a high risk is detected on the ground, and whose maneuver strategies are relatively simple and not organically integrated with the normal constellation configuration maintenance control, which may lead to increased fuel consumption or control conflicts. This invention provides a distributed mega-constellation maintenance control method and system based on a combined strategy.

[0005] The technical solution adopted by the present invention to solve the above problems is: a distributed mega-constellation maintenance control method based on a combined strategy, the method comprising: step 1: establishing an orbital dynamics model;

[0006] Step 2: Establish a priority strategy, which is used to determine the satellites that need to be controlled and their control variables for each member satellite in the constellation; specifically including:

[0007] Step 2.1: Design control flag vectors for each member satellite in the giant constellation. This indicates whether the satellite requires control and the amount of control needed:

[0008] (1)

[0009] in, These control flags represent the semi-major axis, eccentricity, orbital inclination, right ascension of the ascending node, phase, and obstacle avoidance control, respectively. The value of each control flag component determines whether the corresponding component requires control.

[0010] Step 2.2: Design control priority flags based on control flags Its components are:

[0011] (2)

[0012] in, For the expected flat semi-major axis of the member satellites, To control the deviation value starting from the horizontal half-major axis, This is the estimated value of the semi-major axis. The expected mean eccentricity of the member satellites, To control the deviation value of the eccentricity, This is the estimated value of the eccentricity. The expected horizontal orbit inclination of the member satellites, To control the deviation value of the track inclination angle. This is an estimated value for the inclination angle of the horizontal track. This indicates the expected rise in the right ascension of the nodes. The control deviation value begins at the right ascension of the ascending intersection. This is the estimated right ascension of the ascending nodes. For the magnitude of the horizontal latitude argument that needs to be adjusted, The deviation value is controlled starting from the latitude angle interval;

[0013] Step 2.3: Control the switching interval as follows In each During the specified period, only member satellites will be affected. Control the portion corresponding to the largest component in the middle;

[0014] Step 3: Based on the priority strategy in Step 2, design autonomous maintenance control strategies for the semi-major axis, eccentricity, orbital inclination, right ascension interval of the ascending and descending nodes, and satellite phase, as well as obstacle avoidance control for satellites on opposite sides.

[0015] The beneficial effects of this invention are:

[0016] This invention provides a distributed mega-constellation maintenance control method based on a combined strategy. Through an innovative "control flag vector" and "adaptive priority decision-making" mechanism, it achieves distributed autonomous operation. This method integrates multiple objectives—single-star orbit maintenance, in-plane phase maintenance, inter-plane configuration control, and autonomous collision avoidance—into a unified framework and intelligently resolves conflicts between them. While ensuring high system autonomy and scalability, it significantly improves overall control efficiency. It assigns collision avoidance the highest decision priority and deeply embeds probabilistic collision risk assessment into the control loop, thus providing proactive safety assurance for the constellation in dense orbital environments. Furthermore, the method effectively optimizes fuel consumption by combining orbital dynamics-based thrust application strategies (such as performing control at specific orbital positions) and "elevation-only" phase adjustment strategies. The design based on the average orbital elements makes it insensitive to high-frequency perturbations, and the control law is stable and reliable. Simulation verification shows that this method can maintain constellation orbit and configuration deviations within a preset range over a long period and successfully avoid collision risks, demonstrating outstanding theoretical and engineering value.

[0017] This invention relates to a distributed autonomous orbit maintenance control method for low Earth orbit (LEO) mega-constellations that takes into account self-collision avoidance. Attached Figure Description

[0018] To more clearly illustrate the technical solution of this application, the drawings used in the embodiments will be briefly introduced below. Obviously, for those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0019] Figure 1 The geometric relationship between the LVLH coordinate system and the NTW coordinate system; Figure 2 This is a schematic diagram showing the location corresponding to the minimum orbital intersection distance; Figure 3 The phase difference between satellites in each plane; Figure 4 This shows the change of the controlled variable deviation of 1388 satellites over time under uncontrolled conditions. Figure 5 The controlled variable deviation of 1388 satellites changes over time under distributed autonomous control (strategy 1); Figure 6 The controlled variable deviation of 1388 satellites changes over time under distributed autonomous control (strategy 2); Figure 7 The mean deviation of the controlled variables of 1388 satellites changes over time; Figure 8 The standard deviation of the controlled variables of 1388 satellites changes over time; Figure 9 This shows the change in the inter-satellite distances of the third group of satellites over time under uncontrolled conditions. Figure 10 This shows the changes in the orbital elements of the sixth group of satellites over time. Figure 11 The priority flag for the 6th group of satellites is the largest component number; Figure 12 The thrust experienced by the sixth group of satellites; Figure 13 This shows the change in inter-satellite distances over time in the sixth group of satellites under uncontrolled conditions. Figure 14 The distance between satellites in group 6 changes over time under both uncontrolled and distributed autonomous control (strategy 1) scenarios. Figure 15 This shows the changes in the orbital elements of the fourth group of satellites over time. Figure 16 The priority flag for the fourth group of satellites is the largest component number; Figure 17 The thrust experienced by the fourth group of satellites; Figure 18 This shows the change in the distance between satellites in group 4 over time under uncontrolled conditions. Figure 19 This shows the change in the distance between satellites in group 4 over time under uncontrolled conditions. Detailed Implementation

[0020] To achieve autonomous constellation configuration maintenance control, the core issue is to construct a distributed cooperative control method that uses local measurement information within the mega-constellation as input. Based on the inputs and requirements of mega-constellation maintenance control, a distributed mega-constellation maintenance control method based on a combined strategy is proposed. The distributed nature is reflected in each member satellite designing its own control strategy based on its own average orbital parameters and the orbital information sensed from other member satellites in the mega-constellation. The combined strategy is reflected in the fact that each member satellite, during autonomous control, must comprehensively consider multiple control requirements and, for potentially conflicting control objectives, design a combined control strategy based on a set adaptive priority.

[0021] The specific implementation method of this embodiment, a distributed mega-constellation maintenance control method based on a combination strategy, includes:

[0022] 3.1 Orbital Dynamics Model

[0023] When the perturbation force is a general case involving both non-conservative and conservative forces, it can be expressed in the LVLH or NTW coordinate system and solved using Gaussian variational equations (GVEs). The LVLH-type Gaussian perturbation equations can be derived from the principle of constant variation; however, to facilitate the incorporation of small thrusts, the transformation relations of the constant roots are further derived, and GVEs are given in the NTW coordinate system. The geometric relationship between the LVLH and NTW coordinate systems is as follows: Figure 1 As shown.

[0024] when and At that time, GVEs are:

[0025]

[0026] in, It is the average orbital angular velocity. and These are the three components of the total acceleration produced by the forces acting on the satellite in the NTW coordinate system. It's the eccentricity. It is the track inclination angle. It is time. It is a semi-major axis. It is the Earth's gravitational constant. It's a true near-point angle. It's closer to the point angle. It is the distance from the satellite to the Earth's center. It is the argument of latitude. It is the right ascension of the ascending node. It is the perigee argument. It is a near-point angle.

[0027] 3.2 Adaptive Priority Strategy Design

[0028] Design control flag vectors for each member satellite in the giant constellation. This indicates whether the satellite requires control and the amount of control needed:

[0029] (1)

[0030] in These represent control flags for semi-major axis, eccentricity, orbital inclination, RAAN (right ascension of the ascending node), phase, and obstacle avoidance control. A value of 0 for each control flag component indicates that the corresponding component does not require control, while a value of 1 indicates that the corresponding component requires control. The method for determining the values ​​of the control flags will be described in sections 3.3 to 3.8.

[0031] Since the components requiring control indicated by the control flags may conflict with each other, meaning the required thrust directions may be opposite, a control priority flag is designed based on the control flags to ensure orderly and rational satellite control. Its components are:

[0032] (2)

[0033] in, For the expected flat semi-major axis of the member satellites, To control the deviation value starting from the horizontal half-major axis, This is the estimated value of the semi-major axis. The expected mean eccentricity of the member satellites, To control the deviation value of the eccentricity, This is the estimated value of the eccentricity. The expected horizontal orbit inclination of the member satellites, To control the deviation value of the track inclination angle. This is an estimated value for the inclination angle of the horizontal track. This indicates the expected rise in the right ascension of the nodes. The control deviation value begins at the right ascension of the ascending intersection. This is the estimated right ascension of the ascending nodes. For the magnitude of the horizontal latitude argument that needs to be adjusted, The starting point for controlling the deviation value is the latitude angle interval.

[0034] Assume the control switching interval is In each During the specified period, only member satellites will be affected. The part corresponding to the largest component is controlled.

[0035] It can be seen from equation (2) that when Sometimes, When a collision risk exists, obstacle avoidance will be prioritized. When there is no collision risk, the choice of which track element to maintain first depends on its degree of deviation from the target. The following sections will introduce the methods for determining the control flags and the control methods for the track elements.

[0036] 3.3 Design of Autonomous Holding Control Strategy for Semi-Long Axis

[0037] The estimated value for the semi-major axis is... The member satellites, assuming their expected semi-major axis is... Define the initial control deviation value of the semi-major axis as follows: That is, there is a need to maintain the semi-major axis when the following conditions are met:

[0038] (3)

[0039] In this case, set control flags This indicates that horizontal semi-major axis maintenance control is required.

[0040] To prevent frequent adjustments and save fuel, the semi-major shaft stop control deviation value is defined as follows: , .exist In the following cases, semi-major axis maintenance control shall be stopped:

[0041] (4)

[0042] At this time, This indicates that the horizontal semi-major axis maintenance is no longer being performed.

[0043] When the highest priority requirement in the priority flag is the horizontal half-major axis control requirement, that is when At this time, the semi-major axis is adjusted by a force along the tangential direction of the track. In the NTW coordinate system, the control force can be expressed as:

[0044] (5)

[0045] in, The magnitude of the thrust applied tangentially along the track. The function is used to determine the direction of the thrust, i.e., according to The sign of the velocity determines whether the thrust is in the positive or negative direction of the velocity.

[0046] In this way, the satellite's semi-major axis can be effectively maintained within the expected range, thereby ensuring orbital stability and the normal operation of the mission.

[0047] 3.4 Autonomous Eccentricity Maintenance Control Strategy

[0048] According to the orbital dynamics equations, forces along the orbital tangential and in the orbital plane normal can both change the satellite's eccentricity. Among these, the coefficient of the tangential thrust acceleration... The coefficient of normal thrust acceleration in the orbital plane They are respectively:

[0049]

[0050] Within one orbital period, the coefficient and They all change periodically. Because... , It can be seen that the maximum efficiency of changing eccentricity with a force along the orbital tangential direction is greater than that with a force along the normal direction in the orbital plane. For low-Earth orbit mega-constellations with an eccentricity on the order of 10 to the negative cube, considering fuel conservation while maintaining eccentricity, eccentricity control with a force along the orbital tangential direction should be used near the perigee and apogee to achieve greater benefits. For the mean eccentricity estimate... The member satellites, assuming their expected eccentricity is... The initial control deviation value of the eccentricity rate is... The maximum permissible angle difference between the true perihelion and the perihelion or apohelion during maintenance is: When satisfied Furthermore, the difference between the satellite's true perihelion and its perigee or apogee angle is less than [missing value]. At times, there is a need to maintain eccentricity, causing .

[0051] Define the eccentricity stop control deviation value as , .exist In the case of, if Or the difference between the satellite's true perihelion and its perigee or apogee is greater than or equal to 1. At that time, eccentricity maintenance is not performed, so that .

[0052] When the highest priority requirement in the priority flag is the eccentricity control requirement, that is when When designing a control strategy, the effects of tangential thrust on the semi-major axis and eccentricity should be comprehensively considered.

[0053] (1) Considering that atmospheric drag will cause the satellite's semi-major axis to continuously decrease, when the satellite's semi-major axis is lower than the expected semi-major axis, or when the satellite's semi-major axis is higher than the expected semi-major axis but the deviation from the expected semi-major axis is less than the initial control deviation value, the semi-major axis should be increased simultaneously when maintaining the eccentricity, that is:

[0054] a. If and When near the perigee, This is to simultaneously improve both the eccentricity and the semi-major axis.

[0055] b. If and When near the apogee, make This reduces eccentricity and increases the semi-major axis.

[0056] (2) If the satellite's horizontal semi-major axis is higher than the expected horizontal semi-major axis and the deviation from the expected semi-major axis is greater than the initial control deviation value, the semi-major axis should be lowered simultaneously when maintaining the eccentricity, that is:

[0057] a. If and When near the perigee, This is to reduce both the eccentricity and the semi-major axis.

[0058] b. If and When near the apogee, make This is to increase the eccentricity and reduce the semi-major axis.

[0059] 3.5 Autonomous Track Inclination Holding Control Strategy

[0060] According to the orbital dynamics equations, a force along the direction of orbital angular momentum can change the satellite's orbital inclination, and the coefficient of thrust acceleration along the direction of orbital angular momentum... for: .

[0061] It can be seen that the coefficient Except for the latitude argument, all other quantities are slow variables. Therefore, considering fuel conservation while maintaining orbital inclination, orbital inclination control can be achieved by using forces along the orbital angular momentum direction near latitude arguments of 180° and 0° to obtain greater benefits. The estimated inclination value for a horizontal orbit is... The member satellites, assuming their expected horizontal orbit inclination is... The initial control deviation value of the track inclination angle is... The maximum permissible angle difference between the satellite's latitude argument and 180° or 0° during maintenance is [missing value]. When satisfied Furthermore, the difference between the satellite's latitude argument and the angle of 180° or 0° is less than... At times, there is a need to maintain the track tilt angle, making .

[0062] Define the track tilt angle stop control deviation value as , .exist In the case of, if Or the difference between the satellite's latitude argument and 180° or 0° is greater than or equal to 180° or 0°. At that time, without maintaining the track tilt angle, let .

[0063] Assume the magnitude of the normal thrust on the orbital plane is Then, when the highest priority requirement in the priority flag is the requirement for horizontal track inclination control, that is... At that time, the track inclination angle is controlled, and the control strategy is as follows:

[0064] (1) When At that time, if the satellite's latitude angle is in Within the range, let If the satellite's latitude argument is within... Within the range, let .

[0065] (2) When At that time, if the satellite's latitude angle is in Within the range, let If the satellite's latitude argument is within... Within the range, let .

[0066] 3.6 Autonomous Maintenance Control Strategy for the Right Ascension Intersection of the Ascending and Floating Nodes

[0067] According to the orbital dynamics equations, a force along the direction of orbital angular momentum can change the right ascension of the satellite's ascending node. The coefficient of the thrust acceleration along the direction of orbital angular momentum... for:

[0068]

[0069] Considering the need to conserve fuel while maintaining the right ascension of the ascending node, greater benefits can be obtained by using forces along the direction of orbital angular momentum near latitude angles of 90° and 270° to control the right ascension of the ascending node.

[0070] The estimated right ascension of the ascending nodes is... For the member satellites, the expected right ascension of the ascending node is calculated based on the average right ascension of the ascending nodes of their left and right adjacent planes:

[0071]

[0072] in, This indicates the expected rise in the right ascension of the nodes. This represents the mean right ascension of the ascending nodes of the member satellites in the left-hand plane. This represents the mean right ascension of the ascending nodes of the member satellites in the right-hand plane.

[0073] Define the right ascension of the ascending intersection as the starting control deviation value. The maximum permissible angle difference between the satellite's latitude argument and 90° or 270° during maintenance is [missing value]. When satisfied Furthermore, the difference between the satellite's current latitude angle and the angle of 90° or 270° is less than... At times, there is a need to maintain the right ascension of the ascending node, causing .

[0074] Define the stopping control deviation value at the ascending node right ascension as: , .exist In the case of, if Or the difference between the satellite's latitude argument and 90° or 270° is greater than or equal to At that time, the right ascension of the ascending node is not maintained, so that .

[0075] When the highest priority requirement in the priority flag is the right ascension interval control requirement at the ascending intersection, that is... At that time, the control strategy is:

[0076] (1) When At that time, if the satellite's latitude angle is in Within the range, let If the satellite's latitude argument is within... Within the range, let .

[0077] (2) When At that time, if the satellite's latitude angle is in Within the range, let If the satellite's latitude argument is within... Within the range, let .

[0078] 3.7 Satellite Phase Autonomous Maintaining Control Strategy

[0079] For controlling satellites within the same plane, the control of the average latitude argument spacing is considered. For mega-constellations, due to satellite failures and the presence of backup satellites, the distribution of satellites within the plane is not uniform, resulting in multiple average latitude argument spacings. Therefore, phase-keeping control cannot be simply considered as a uniform distribution; it requires design and analysis based on different average latitude argument spacings.

[0080] Assuming the expected mean latitude angle interval between the backup satellite and the regular member satellites in the same plane is... The expected mean latitude angle interval between ordinary member satellites is Based on this, the expected mean latitude argument interval vector between satellites in the same plane can be designed as follows:

[0081]

[0082] in, Each component represents a possible expected average latitude argument interval. You need to select manually based on the actual situation.

[0083] Assume the estimated latitude argument of a satellite C in the plane is... The estimated latitude and angle of the satellite D behind it is [value missing]. The estimated value of the average latitude argument interval between the two is:

[0084]

[0085] Assuming the initial control deviation value of the latitude argument interval is Then when At that time, it is necessary to adjust the equal latitude argument interval between the two stars, which will make The expected latitude argument interval that yields the minimum value is denoted as Therefore, for the mean latitude argument interval between satellite C and satellite D, the mean latitude argument that satellite C needs to adjust is:

[0086]

[0087] The required adjustment for the mean latitude angle of satellite D is:

[0088]

[0089] Since each member satellite in the near-circular orbit plane has neighboring satellites in front and behind it, there may be situations where the mean latitude argument intervals in front of and behind a member satellite exceed the initial control deviation value. Therefore, taking satellite C as an example, let's assume that the mean latitude argument interval in front of it is... The rear latitude argument interval is Therefore, the required adjustment to the mean latitude argument is:

[0090]

[0091] in, In order to make The expected average latitude angle interval to achieve the minimum value.

[0092] For satellite C, if we determine and Therefore, it is assumed that the interval between the preceding and following horizontal latitude angles does not need to be adjusted. Adjusting the satellite's orbital altitude can affect its average orbital angular velocity, thereby adjusting the satellite's mean latitude argument. Therefore, adjusting orbital altitude is used for satellite phase adjustment. Two feasible control strategies are provided here:

[0093] (1) Strategy 1: Use thrust in both the positive and negative directions along the track tangentially for control.

[0094] The relationship between the satellite's average orbital angular velocity and its semi-major axis is as follows:

[0095]

[0096] in, Let C be the average orbital angular velocity of satellite C. This is the estimated semi-major axis of satellite C.

[0097] It can be seen that when the orbital altitude of satellite C increases, its average orbital angular velocity decreases; when the orbital altitude of satellite C decreases, its average orbital angular velocity increases. Therefore, The relationship with tangential thrust is as follows:

[0098]

[0099] in, This refers to the tangential thrust along the orbit used by satellite C for latitude-angle interval control.

[0100] To ensure the long-term orbital stability of the satellite, a gradual phasing method is employed, allowing the satellite to gradually reach the required phase over a period of time. This method avoids setting a stop control deviation value directly related to the latitude argument interval; instead, it defines a deviation between the horizontal semi-major axis and its nominal value that allows for horizontal latitude argument interval control. This deviation must satisfy: .

[0101] When the satellite's horizontal semi-major axis is within the allowable range for horizontal latitude argument interval control and the latitude argument interval deviation does not meet the requirements, i.e. and season Otherwise, let .

[0102] When the highest priority requirement in the priority flag is the latitude angle interval control requirement, that is... At this time, a force along the tangential direction of the orbit is used to control the semi-major axis, thereby adjusting the latitude argument interval, which can be represented in the NTW coordinate system as:

[0103] (2) Strategy 2: Use only the thrust in the positive tangential direction along the track for control

[0104] Because atmospheric drag causes the satellite's semi-major axis to continuously decrease, the difference between Strategy 2 and Strategy 1 is that Strategy 2 only uses positive thrust to adjust the mean latitude argument, that is: when and season Otherwise, let .

[0105] When the highest priority requirement in the priority flag is the latitude angle interval control requirement, that is... At this time, a force along the positive tangential direction of the orbit is used to control the semi-major axis, thereby adjusting the latitude argument interval, which can be represented in the NTW coordinate system as: .

[0106] 3.8 Collision Avoidance Control for Non-aligned Satellites

[0107] Before implementing obstacle avoidance control, preliminary screening and proximity analysis are required. For a satellite A in a giant constellation, assuming it can detect a radius of [missing information] through its sensors... Other satellites within the detection range. For detected non-aligned satellites, their average orbital features are estimated, and collision analysis is performed.

[0108] (1) Minimum orbital intersection distance

[0109] MOID is the minimum distance between two closely orbiting satellites and is a widely used parameter in collision risk assessment. Consider satellite A's orbit as follows: The orbit of satellite B, which is located on the opposite side of the orbit, is sensed by satellite A. MOID corresponds to the two closest points on the orbit, such as Figure 2 As shown.

[0110] The true anomaly angle corresponding to MOID can be calculated by analogy with the Gronchi algorithm [1,2]. First, the analytical expression for the square of the distance between two general points on the orbit is derived. In the ECI coordinate system, the orbital components of satellite A are:

[0111]

[0112] in, These are the coordinates of the orbital coordinate system of satellite A, represented in the inertial coordinate system. axis, The unit vector of the axis is the orbital element. The function.

[0113] Similarly, the orbital components of satellite B can be obtained as follows:

[0114] in, These are the coordinates of the orbital coordinate system of satellite A, represented in the inertial coordinate system. axis, The unit vector of the axis is the orbital element. The function.

[0115] The square of the distance between two points on the orbits of satellites A and B can be expressed as:

[0116] in, It is the Euclidean scalar product.

[0117] It can be seen that the above formula depends on and Kepler parameters, the variable being the true anterior angle. and Apply the expression to each and Taking the derivative and setting it to zero, we get two equations, which together form a system of equations:

[0118]

[0119] The solutions to the above equations are stationary points, which may be minima, maxima, or saddle points. Using the Fast Fourier Transform (FFT) method based on Discrete Fourier Transform and Inverse Discrete Fourier Transform, we can solve the two systems of equations to determine the true anterior angle corresponding to each stationary point, where the minimum among the stationary points corresponds to... and This is the true anomaly angle corresponding to the intersection of the orbits of satellites A and B.

[0120] For a circular orbit, two MOID positions are always shared within one orbital period, and their projection onto the celestial sphere is the intersection of the projections of orbits A and B onto the celestial sphere. Furthermore, these two MOID positions are always symmetrical with respect to the Earth's center and have a phase difference of 180°. and The true anterior angles are replaced with the true anterior angles at the intersection points obtained using the Gronchi algorithm, and then converted into position and velocity. The MOID can then be calculated, denoted as . .

[0121] For satellites A and B, when their corresponding Less than the preset threshold In such cases, further filtering of the collision time interval is required. This is because calculating the collision probability over the entire period would incur unnecessary computational overhead.

[0122] (2) Collision time interval analysis

[0123] Typically, the collision can be considered to occur near the MOID location. According to... The time it takes for satellite A to reach the MOID position can be calculated using satellite A's orbital information. ,according to The time it takes for satellite B to reach the MOID position can be calculated using satellite B's orbital information. The time interval between satellite encounters can be defined as the period during which two satellites approach and pass through the MOID position:

[0124]

[0125] in, For time safety margin.

[0126] For each member satellite, to ensure sufficient obstacle avoidance reaction time, calculations are performed for distances greater than 0.25 orbital periods and less than 1 orbital period from the current time. and Considering that if the time difference between the two satellites passing through the intersection line of their orbital planes is too long, a collision will not occur, the calculation... ,when Less than the preset threshold At that time, calculate closest distance between two satellites during the time period and the corresponding time ,when Less than the preset threshold At that time, according to Calculate the collision probability based on the positions and velocities of the two satellites at any given time.

[0127] (3) Collision probability calculation

[0128] Collision probability calculation is the final step in collision risk analysis. Given the characteristics of mega-constellations, self-collisions typically occur near the intersection of their different orbits. Furthermore, from the triggering of a collision warning to the predicted collision time, spacecraft may encounter each other multiple times, requiring multiple collision probability calculations. Therefore, an analytical method is used to quickly calculate the collision probability. Generally speaking, analytical methods are much faster than numerical methods in terms of computational speed.

[0129] First, define a meeting reference frame centered on target B, whose collision probability needs to be calculated: Its three coordinate axes unit vectors are:

[0130]

[0131] in, and These are the velocities of target B and member satellite A, respectively. Assume the encounter plane is a plane. The velocity perpendicular to the relative velocity of member satellite A with respect to target B. Assume the relative position of target B with respect to member satellite A is:

[0132]

[0133] Then its projection in the meeting plane is:

[0134]

[0135] in,

[0136]

[0137] in and It is a unit vector The weight, in relation to Represented in the same frame of reference.

[0138] According to Chan[3]'s method, the original problem is simplified by first assigning the combined covariance to target B (without an envelope) and assigning the combined envelope to member satellite A (without a covariance), thus simplifying the situation where each object has its own spherical envelope and covariance matrix. The combined envelope is centered on member satellite A, and its radius is... It equals the sum of the radii of each envelope. Assuming the error covariances of satellite A and target B are independent, the combined covariance in the meeting reference frame is:

[0139]

[0140] in, and They are respectively and Standard deviation of direction The correlation coefficient is used as the basis for determining the probability of collisions between spacecraft. Subsequently, the probability of collisions between spacecraft can be approximated by a convergent series:

[0141]

[0142] in,

[0143]

[0144] in It is the position where satellite A and target B are closest in the encounter plane, i.e. The coordinates of target B projected onto the meeting coordinate system at time t.

[0145] Member satellite A autonomously calculates the collision probability of all targets within its sensing range using the aforementioned process. If any target has a collision probability greater than [a certain value], [the probability is determined]. The target, therefore, is that member satellite A has an active obstacle avoidance requirement, making... Otherwise, let .

[0146] when At that time, based on the characteristics of a giant constellation, obstacle avoidance was achieved by changing the semi-major axis of member satellite A. Let the semi-major axis of member satellite A be denoted as... Let the semi-major axis of the target to be avoided be denoted as Then, the force required by member satellite A along the orbital tangential direction can be expressed in the NTW coordinate system as:

[0147]

[0148] This method can effectively increase the difference in the semi-major axis between satellites, thereby achieving obstacle avoidance.

[0149] IV. Effects of the Invention:

[0150] 4.1 Numerical Simulation

[0151] (1) Initial value design and parameter selection

[0152] Based on the actual TLE data of 1388 Starlink Shell 1 satellites on November 26, 2023, the initial average orbital elements corresponding to the above satellites at the initial moment of the simulation are given as initial values. The constellation evolution under the proposed autonomous constellation maintenance control is analyzed and compared with the constellation evolution under the uncontrolled condition to verify the effectiveness of the proposed method.

[0153] At the initial moment of the simulation, the distribution of inter-satellite phase differences among the member satellites in each plane is as follows: Figure 3 As shown.

[0154] Therefore, the expected latitude argument interval vector is set as follows:

[0155]

[0156] Furthermore, based on the analysis in section 5.2, we assume that the expected semi-major axis, eccentricity, and orbital inclination of the Shell 1 satellite are as follows: and .

[0157] For a satellite in shell 1, its aerodynamic parameters can be assumed to be: and kg, atmospheric drag model parameters adopted using the NRLMSISE-00 model. Control thrust , .

[0158] Assume the initial control deviations of the semi-major axis, eccentricity, inclination, and right ascension of the ascending intersection of the horizontal track are: The initial control deviation value for the latitude angle interval is: The allowable deviation between the semi-major axis and its nominal value for latitude-angle interval control is: .

[0159] Assume the stopping control deviations for the semi-major axis, eccentricity, inclination, and right ascension of the ascending intersection of the horizontal track are: .

[0160] Assume the obstacle avoidance parameters are as follows: , and .

[0161] (2) Analysis of the overall constellation retention effect

[0162] Under uncontrolled conditions, the changes in the deviations of each controlled variable over time are as follows: Figure 4 As shown in the analysis, under uncontrolled conditions, the orbital semi-major axis deviation and the mean latitude argument interval deviation increase significantly over time, with the maximum semi-major axis deviation reaching approximately 3 km and the maximum mean latitude argument interval deviation reaching approximately 4.5°. The deviations in orbital inclination and right ascension of the ascending node hardly change over time. The satellite's mean semi-major axis and mean eccentricity show a gradual decreasing trend.

[0163] The changes in the deviations of each controlled variable over time under distributed autonomous control (Strategy 1) and distributed autonomous control (Strategy 2) are as follows: Figure 5 and Figure 6 As shown.

[0164] Analysis shows that the effects of distributed autonomous control (Strategy 1) and distributed autonomous control (Strategy 2) on maintaining the constellation configuration are similar. The deviations of the various controlled variables of the member satellites initially decrease relative to the initial situation, and then stabilize within a certain range. Among them, the semi-major axis deviation of the horizontal orbit stabilizes at... Within the range, the deviation of the mean eccentricity rate is stable at Within the range, the inclination deviation of the horizontal track is stable at Within the range, the right ascension deviation of the ascending node satellite is in Within the range, the deviation of the mean latitude angle interval is stable at Within the specified range. This indicates that member satellites whose orbital elements did not meet the requirements maintained the constellation configuration by performing autonomous orbital control.

[0165] Figure 7 The study demonstrates the time-varying mean deviations of controlled variables for 1388 satellites under three conditions: no control, distributed autonomous control (Strategy 1), and distributed autonomous control (Strategy 2). It shows that distributed autonomous control ensures that the mean deviation of all components is consistently less than the stop control deviation. Furthermore, the mean deviations of all components except the eccentricity component are reduced compared to the natural evolution scenario.

[0166] Under distributed autonomous control (Strategy 1), the average semi-major axis deviation of 1388 satellites over 5 days was 188.99m, the average eccentricity deviation was 0.0000301, and the average orbital inclination deviation was... The average of the deviations of the right ascension of the ascending and descending nodes is The average of the mean deviations of the angular intervals of latitude is .

[0167] Under distributed autonomous control (Strategy 2), the average semi-major axis deviation of 1388 satellites over 5 days was 186.61m, the average eccentricity deviation was 0.0000292, and the average orbital inclination deviation was... The average of the deviations of the right ascension of the ascending and descending nodes is The average of the mean deviations of the angular intervals of latitude is .

[0168] It can be seen that the control effects of Strategy 1 and Strategy 2 are similar. However, since Strategy 2 only uses thrust along the positive tangential direction of the orbit to adjust the latitudinal angle interval, it can offset part of the influence of atmospheric drag, thereby saving energy consumption. Therefore, Strategy 2 is a better control method.

[0169] Figure 8 The standard deviations of the controlled variables of 1388 satellites under uncontrolled, distributed autonomous control (strategy 1), and distributed autonomous control (strategy 2) conditions at various time points are shown. The standard deviations of eccentricity and semi-major axis deviation are smaller under uncontrolled conditions than under distributed autonomous control, indicating that the proposed distributed mega-constellation autonomous maintenance control method based on combined strategies can make the semi-major axis more dispersed within the allowable range, reducing the risk of collisions.

[0170] (3) Analysis of the obstacle avoidance effect of constellations

[0171] 1) Collision risk analysis under uncontrolled conditions

[0172] As shown in Table 1, under uncontrolled conditions, three groups of satellites were detected within 5 days that required obstacle avoidance. Among them, the third group of satellites that required obstacle avoidance had a collision probability of 2 times within 5 days that met the criteria for active obstacle avoidance.

[0173] Table 1. Satellite groups with high collision risk under uncontrolled conditions

[0174]

[0175] According to Table 1, among the three groups of satellites that were detected as requiring collision avoidance operations, the third group has the highest maximum collision probability, approximately [missing value]. Therefore, a self-collision risk analysis among member satellites is conducted using Group 3 as an example. The satellites in Group 3 requiring obstacle avoidance include Satellite 219 and Satellite 355. Under uncontrolled conditions, the distance between Satellite 219 and Satellite 355 changes over time as follows: Figure 9 As shown, the highest collision risk for the third group of satellites was detected on day 2.417, with the corresponding collision occurring around day 2.482.

[0176] It can be seen that after the maximum collision risk was detected, the two satellites in the third group first reached the minimum distance position on day 2.449, at which time the minimum distance between the two satellites was approximately 2030.39 km, indicating a relatively low collision risk. On day 2.482, when a collision was predicted to have a higher probability, the two satellites in the third group reached the minimum distance position for the second time, with a distance of approximately 191.012 m between them, making a collision highly likely. Therefore, after detecting the high collision risk on day 2.482 on day 2.417, the satellites should have performed active obstacle avoidance maneuvers within 93.6 minutes.

[0177] Collision risk analysis under uncontrolled conditions reveals a significant risk of self-collision between satellites in adjacent planes within a large low-Earth orbit constellation. Therefore, autonomous configuration maintenance must be coupled with collision avoidance.

[0178] 2) Obstacle avoidance under distributed autonomous control (strategy 1)

[0179] Under the distributed autonomous control (strategy 1), a total of 6 groups of satellites that needed to avoid obstacles were detected within 5 days, as shown in Table 2.

[0180] Table 2. Satellite groups with high collision risk under distributed autonomous control (Strategy 1)

[0181]

[0182] Of the six groups of satellites detected requiring obstacle avoidance maneuvers, the sixth group has the highest probability of collision. Therefore, the analysis will focus on the sixth group as an example. The changes in the horizontal orbital elements of the sixth group of satellites over time are as follows: Figure 10 As shown, the priority flag's largest component number is as follows: Figure 11 As shown, within 5 days, satellite 279 maintained its horizontal semi-major axis and horizontal orbital inclination; satellite 455 maintained its horizontal semi-major axis, horizontal eccentricity, horizontal orbital inclination, and horizontal latitude argument interval; on day 1.944, both satellites performed collision avoidance maneuvers.

[0183] The thrust experienced by the sixth group of satellites is as follows: Figure 12 As shown in the diagram. Analysis indicates that the tangential thrust experienced by satellite 279 is entirely in the positive direction of velocity. However, during the adjustment of the latitude angle interval, satellite 455 sometimes uses thrust in the opposite direction of velocity. In obstacle avoidance maneuvers, satellite 279 uses thrust in the positive direction of velocity to raise its horizontal semi-major axis, while satellite 455 uses thrust in the opposite direction of velocity to lower its horizontal semi-major axis, thereby increasing the distance between the two satellites and achieving collision avoidance.

[0184] The high collision risk detected on day 1.944 corresponds to approximately day 1.973. Without control measures, the distance between the two satellites changes over time as follows: Figure 13 As shown, the closest distance between the two stars is approximately 279.066m.

[0185] Around day 1.973, the distance between the two satellites changes over time under both uncontrolled and distributed autonomous control (Strategy 1) conditions, as shown below. Figure 14 As shown.

[0186] 3) Obstacle avoidance under distributed autonomous control (strategy 2)

[0187] Under distributed autonomous control (Strategy 2), a total of 5 groups of satellites requiring obstacle avoidance were detected within 5 days, as shown in Table 3. Among them, the 4th group of satellites requiring obstacle avoidance performed collision avoidance twice within 5 days. The first collision was detected on day 2.427; without control intervention, the probability of collision would have been... The second detection occurred on day 4.816. Without intervention, the probability of a collision is 0.1011.

[0188] Table 3. Satellite groups with high collision risk under distributed autonomous control (Strategy 2)

[0189]

[0190] Of the five groups of satellites detected as requiring obstacle avoidance maneuvers, the fourth group has the highest probability of collision. Therefore, we will analyze the fourth group as an example. The orbital elements of the fourth group of satellites change over time as follows: Figure 15 As shown, the priority flag's largest component number is as follows: Figure 16 As shown in the figure. Analysis shows that within 5 days, satellite 1127 maintained its horizontal semi-major axis, horizontal eccentricity, and horizontal orbital inclination; satellite 1304 maintained its horizontal semi-major axis and horizontal latitude angle interval; on days 2.427 and 4.816, both satellites performed collision avoidance maneuvers.

[0191] The thrust experienced by the fourth group of satellites is as follows: Figure 17 As shown.

[0192] Analysis shows that when the two satellites in Group 4 were performing autonomous configuration maneuvers, the tangential thrust they used was in the positive velocity direction. During the autonomous obstacle avoidance operation on day 4.881, satellite 1127 used thrust in the positive velocity direction to raise its semi-major axis, while satellite 1304 used thrust in the opposite velocity direction to lower its semi-major axis, thereby increasing the distance between the two satellites and achieving collision avoidance.

[0193] The high collision risk detected on day 4.816 corresponds to approximately day 4.881. Without control measures, the distance between the two satellites changes over time as follows: Figure 18 As shown, the closest distance between the two stars is approximately 88.904m, making a collision highly likely.

[0194] Around day 4.881, the distance between the two satellites changes over time under both uncontrolled and distributed autonomous control (Strategy 2) conditions, as shown below. Figure 19 As shown in the diagram, the closest distance between the two stars increased to approximately 2825.42m, successfully achieving obstacle avoidance.

[0195] Analysis of obstacle avoidance under distributed autonomous control (Strategy 1) and distributed autonomous control (Strategy 2) reveals that both distributed autonomous control methods successfully achieve collision avoidance for the mega-constellation in low Earth orbit. Combining the effects of the two distributed autonomous control strategies on maintaining the autonomous configuration of the mega-constellation, it is clear that, compared to distributed autonomous control (Strategy 1), distributed autonomous control (Strategy 2) can maintain the configuration of the mega-constellation with lower fuel consumption while avoiding collisions, making it a more ideal autonomous control strategy.

[0196] The references cited in this invention are detailed below:

[0197] [1] GRONCHI G F. An algebraic method to compute the critical pointsof the distance function between two keplerian orbits[J]. Celestial Mechanicsand Dynamical Astronomy, 2005, 93: 295-329.

[0198] [2] Analytical model for collision probability assessments with largesatellite constellations[J]. Advances in Space Research, 2023, 72(7): 2515-2534.

[0199] [3] CHAN K F. Spacecraft collision probability[M]. El Segundo, USA:The Aerospace Press, 2008。

Claims

1. A distributed mega-constellation maintenance control method based on a combination strategy, characterized in that, The method includes: Step 1: Establish the orbital dynamics model; Step 2: Establish a priority strategy, which is used to determine the satellites that need to be controlled and their control variables for each member satellite in the constellation; specifically including: Step 2.1: Design control flag vectors for each member satellite in the giant constellation. This indicates whether the satellite requires control and the amount of control needed: (1) in, These control flags represent the semi-major axis, eccentricity, orbital inclination, right ascension of the ascending node, phase, and obstacle avoidance control, respectively. The value of each control flag component determines whether the corresponding component requires control. Step 2.2: Design control priority flags based on control flags Its components are: (2) in, For the expected flat semi-major axis of the member satellites, To control the deviation value starting from the horizontal half-major axis, This is the estimated value of the semi-major axis. The expected mean eccentricity of the member satellites, To control the deviation value of the eccentricity, This is the estimated value of the eccentricity. The expected horizontal orbit inclination of the member satellites, To control the deviation value of the track inclination angle. This is an estimated value for the inclination angle of the horizontal track. This indicates the expected rise in the right ascension of the nodes. The control deviation value begins at the right ascension of the ascending intersection. This is the estimated right ascension of the ascending nodes. For the magnitude of the mean latitude argument that needs to be adjusted, The deviation value is controlled starting from the latitude angle interval; Step 2.3: Control the switching interval as follows In each During the specified period, only member satellites will be affected. Control the portion corresponding to the largest component in the middle; Step 3: Based on the priority strategy in Step 2, design autonomous maintenance control strategies for the semi-major axis, eccentricity, orbital inclination, right ascension interval of the ascending and descending nodes, and satellite phase, as well as obstacle avoidance control for satellites on opposite sides.

2. The distributed mega-constellation maintenance control method based on a combined strategy according to claim 1, characterized in that: In step 3, the autonomous holding control strategy for the semi-major axis specifically includes: The estimated value for the semi-major axis is... The member satellites, assuming their expected semi-major axis is... Define the initial control deviation value of the semi-major axis as follows: There is a need to maintain the semi-major axis when the following conditions are met: (3) In this case, set control flags This indicates that horizontal semi-major axis maintenance control is required; Define the semi-major shaft stop control deviation value as , ;exist In the following cases, semi-major axis maintenance control shall be stopped: (4) make =0 indicates that the maintenance of the semi-major axis is no longer performed; When the highest priority requirement in the priority flag is the horizontal half-major axis control requirement, that is when At this time, a force along the tangential direction of the track is used to adjust the semi-major axis. In the NTW coordinate system, the control force is expressed as: (5) in, The magnitude of the thrust applied tangentially along the track. The function is used to determine the direction of the thrust, i.e., according to The sign of the velocity determines whether the thrust is in the positive or negative direction of the velocity.

3. The distributed mega-constellation maintenance control method based on a combined strategy according to claim 1, characterized in that: In step 3, the autonomous maintenance control strategy for the mean eccentricity specifically includes: The estimated value of the mean eccentricity is The member satellites, assuming their expected eccentricity is... The initial control deviation value of the eccentricity rate is... The maximum permissible angle difference between the true perihelion and the perihelion or apohelion during maintenance is: When satisfied Furthermore, the difference between the satellite's true perihelion and its perigee or apogee angle is less than [missing value]. At times, there is a need to maintain eccentricity, causing ; Define the eccentricity stop control deviation value as , ;exist In the case of, if Or the difference between the satellite's true perihelion and its perigee or apogee is greater than or equal to 1. At that time, eccentricity maintenance is not performed, so that ; When the highest priority requirement in the priority flag is the eccentricity control requirement, that is when When designing a control strategy, the effects of tangential thrust on the semi-major axis and eccentricity should be comprehensively considered. Considering that atmospheric drag causes the satellite's semi-major axis to continuously decrease, when the satellite's semi-major axis is lower than the expected semi-major axis, or when the satellite's semi-major axis is higher than the expected semi-major axis but the deviation from the expected semi-major axis is less than the initial control deviation value, the semi-major axis should be increased simultaneously when maintaining eccentricity, that is: like and When near the perigee, To simultaneously improve eccentricity and semi-major axis; like and When near the apogee, make To reduce eccentricity and increase semi-major axis; If the satellite's horizontal semi-major axis is higher than the expected horizontal semi-major axis and the deviation from the expected semi-major axis is greater than the initial control deviation value, the semi-major axis should be lowered simultaneously when maintaining eccentricity, that is: like and When near the perigee, To simultaneously reduce eccentricity and semi-major axis; like and When near the apogee, make This is to increase the eccentricity and reduce the semi-major axis.

4. The distributed mega-constellation maintenance control method based on a combination strategy according to claim 1, characterized in that: In step 3, the autonomous control strategy for maintaining the inclination angle of the horizontal track specifically includes: The estimated value for the inclination angle of the horizontal track is... The member satellites, assuming their expected horizontal orbit inclination is... The initial control deviation value of the track inclination angle is... The maximum permissible angle difference between the satellite's latitude argument and 180° or 0° during maintenance is [missing value]. When satisfied Furthermore, the difference between the satellite's latitude argument and the angle of 180° or 0° is less than... At times, there is a need to maintain the track tilt angle, making ; Define the track tilt angle stop control deviation value as , ;exist In the case of, if Or the difference between the satellite's latitude argument and 180° or 0° is greater than or equal to 180° or 0°. At that time, without maintaining the track tilt angle, let ; Assume the magnitude of the normal thrust on the orbital plane is Then, when the highest priority requirement in the priority flag is the requirement for horizontal track inclination control, that is... At that time, the track inclination angle is controlled, and the control strategy is as follows: when At that time, if the satellite's latitude angle is in Within the range, let If the satellite's latitude angle is within Within the range, let ; when At that time, if the satellite's latitude angle is in Within the range, let If the satellite's latitude angle is within Within the range, let .

5. The distributed mega-constellation maintenance control method based on a combined strategy according to claim 1, characterized in that: In step 3, the autonomous maintenance control strategy for the right ascension interval of the ascending nodes specifically includes: The estimated right ascension of the ascending nodes is... The member satellites calculate the expected right ascension of the ascending node based on the average right ascension of the ascending nodes of their left and right adjacent planes. ; Define the right ascension of the ascending intersection as the starting control deviation value. The maximum permissible angle difference between the satellite's latitude argument and 90° or 270° during maintenance is [missing value]. When satisfied Furthermore, the difference between the satellite's current latitude angle and the angle of 90° or 270° is less than... At times, there is a need to maintain the right ascension of the ascending node, causing ; Define the stopping control deviation value at the ascending node right ascension as: , ;exist In the case of, if Or the difference between the satellite's latitude argument and 90° or 270° is greater than or equal to At that time, the right ascension of the ascending node is not maintained, so that ; When the highest priority requirement in the priority flag is the right ascension interval control requirement at the ascending intersection, that is... At that time, the control strategy is: when At that time, if the satellite's latitude angle is in Within the range, let If the satellite's latitude angle is within Within the range, let ; when At that time, if the satellite's latitude angle is in Within the range, let If the satellite's latitude angle is within Within the range, let .

6. The distributed mega-constellation maintenance control method based on a combined strategy according to claim 1, characterized in that: In step 3, the autonomous satellite phase-maintaining control strategy specifically includes: Define the deviation between the semi-major axis and its nominal value that allows for latitude-angle interval control. This deviation must satisfy: ; When the satellite's horizontal semi-major axis is within the allowable range for horizontal latitude argument interval control and the latitude argument interval deviation does not meet the requirements, i.e. and At that time, among them Let be the estimated semi-major axis of satellite C. Otherwise, let ; When the highest priority requirement in the priority flag is the latitude angle interval control requirement, that is... At this time, a force along the tangential direction of the orbit is used to control the semi-major axis, thereby adjusting the latitude argument interval, which can be represented in the NTW coordinate system as: , ,in, The orbital tangential thrust is used by Satellite C for latitude angle interval control.

7. The distributed mega-constellation maintenance control method based on a combination strategy according to claim 1, characterized in that: Step 3, the obstacle avoidance control of the non-plane satellite, specifically includes: Calculate the minimum distance between the closely orbiting satellites A and B; Based on the time it takes for satellite A and satellite B to reach the MOID position, obtain the time interval for the satellites to meet; The collision probability of all targets within the perception range is obtained. If there is a target with a collision probability greater than a preset value, it is assumed that member satellite A has an active obstacle avoidance requirement. Otherwise, let ; when At that time, based on the characteristics of a giant constellation, obstacle avoidance was achieved by changing the semi-major axis of member satellite A; the semi-major axis of member satellite A is denoted as... Let the semi-major axis of the target to be avoided be denoted as Then, the force required by member satellite A along the orbital tangential direction can be expressed in the NTW coordinate system as: 。 8. A distributed mega-constellation maintenance control system based on a combined strategy, characterized in that: The system has a program module corresponding to the steps of any one of claims 1-7, and executes the steps in the distributed mega-constellation maintenance control method based on a combination strategy when it is run.

9. A computer device, characterized in that: It includes a memory and a processor, wherein the memory stores a computer program, and when the processor runs the computer program stored in the memory, it performs the steps of the distributed mega-constellation maintenance control method based on a combination strategy as described in any one of claims 1-7.

10. A computer-readable storage medium, characterized in that: The storage medium is used to store a computer program that executes a distributed mega-constellation maintenance control method based on a combination strategy, as described in any one of claims 1-7.

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