Method and device for maintaining constellation configuration based on sub-satellite point trajectory drift

Abstract

The invention relates to the technical field of satellite orbit control, and discloses a sub-satellite point trajectory drift-based constellation configuration maintaining method and device, and the method comprises the steps: obtaining the orbit data and the orbit drift distance upper limit of a constellation; according to the orbit data, an orbit drift distance upper limit, a preset orbit drift distance formula and a preset ascending node longitude formula, calculating to obtain a semi-major axis upper limit of a running orbit, thereby respectively calculating to obtain a maneuvering moment and a maneuvering drift distance of each satellite; according to the maneuvering drift distance of each satellite, performing clustering analysis on each satellite to obtain an average maneuvering drift distance corresponding to the satellite group; respectively calculating the speed increment of each satellite at the maneuvering moment based on the orbit data of the constellation and the average maneuvering drift distance corresponding to each satellite group; and when the operation time of each satellite reaches the maneuvering time, adjusting the operation state of each satellite based on the speed increment corresponding to each satellite. The constellation configuration and the constellation performance can be maintained, and the constellation maintenance control accuracy is improved.

Description

Technical Field

[0001] The present invention relates to the field of satellite orbit control technology, and in particular to a constellation configuration maintenance method and device based on sub-satellite point trajectory drift. Background Art

[0002] For the constellation configuration maintenance problem, existing configuration maintenance strategies include configuration bias method and limit cycle method. The configuration bias method mainly considers setting a basic bias value for the initial configuration so that the configuration can obtain a certain degree of compensation in the long-term drift, thereby compensating for the impact of other perturbations. The limit cycle method mainly controls the semi-major axis. After considering the solar pressure, three-body gravity, and atmospheric resistance, the semi-major axis will continue to decay. When it decays to a certain value, it is considered that the original configuration is destroyed, and the semi-major axis is controlled back to the original height. The goal of these methods is to maintain the geometric configuration of the constellation stable. The majority of the optimized variables are the orbital semi-major axis, and a small amount of orbital inclination is considered.

[0003] Most of the current constellation configuration maintenance aims to obtain a stable constellation space geometric configuration rather than a stable constellation performance. The configuration bias method is only applicable to homogeneous constellations (i.e., constellations composed of satellites with the same orbital altitude and inclination). Due to the characteristics of homogeneous constellations, the perturbations exerted on the satellites in the constellation are consistent as a whole, and the bias required by the configuration bias method will not be too large. If the target constellation is a heterogeneous constellation, the bias required by the configuration bias method will differ greatly, which will directly destroy the initial configuration of the constellation. For the limit cycle method, most of the considerations are to control each satellite in the constellation without considering the relative relationship between satellites, that is, it belongs to the absolute configuration control method. This method consumes a lot of energy and is not conducive to saving fuel and reducing costs. Summary of the invention

[0004] The present invention provides a constellation configuration maintenance method and device based on sub-satellite point trajectory drift, which can maintain the constellation configuration and constellation performance and improve the accuracy of constellation maintenance control.

[0005] In order to solve the above technical problems, the present invention provides a constellation configuration maintenance method based on sub-satellite point trajectory drift, comprising:

[0006] Acquire orbital data of a first constellation; wherein the orbital data includes operating orbital data and reference orbital data corresponding to a plurality of satellites in the first constellation;

[0007] Get the upper limit of the orbital drift of the first constellation;

[0008] The upper limit of the semi-major axis of the running orbit is calculated based on the orbit data, the upper limit of the orbit drift, the preset orbit drift formula and the preset ascending node longitude formula;

[0009] Based on the preset ascending node longitude formula, the orbital data and the upper limit of the semi-major axis of the operating orbit, respectively calculating the maneuvering time corresponding to the upper limit of the semi-major axis of the operating orbit and the maneuvering drift amount of each of the satellites;

[0010] According to the maneuvering drift of each satellite, cluster analysis is performed on each satellite to obtain a plurality of satellite clusters, and an average maneuvering drift corresponding to each satellite cluster is calculated;

[0011] Based on the orbital data of the first constellation and the average maneuvering drift corresponding to each of the satellite groups, respectively calculating the first velocity increment of each of the satellites at the maneuvering moment;

[0012] When the operating time of each satellite reaches the maneuvering time, the operating state of each satellite is adjusted based on the first speed increment corresponding to each satellite, so that each satellite enters the phase adjustment orbit from the current operating orbit to maintain the constellation configuration of the first constellation.

[0013] As a preferred solution, the upper limit of the semi-major axis of the running orbit is calculated based on the orbit data, the upper limit of the orbit drift, the preset orbit drift formula and the preset ascending node longitude formula, including:

[0014] Based on the preset orbit drift formula and the preset ascending node longitude formula, a first correlation relationship between the orbit drift and the semi-major axis of the running orbit is obtained;

[0015] Calculate the upper limit of the semi-major axis of the running orbit according to the first correlation, the upper limit of the orbit drift and the orbit data;

[0016] The preset orbit drift formula is:

[0017]

[0018] Where D σ is the orbital drift; R e is the radius of the earth; Δi is the inclination difference between the reference orbit and the running orbit; Δλ N is the difference in longitude between the ascending node of the reference orbit and the running orbit; i1 is the inclination of the reference orbit;

[0019] The preset ascending node longitude formula is:

[0020]

[0021] In the formula, Δλ N is the difference in longitude between the ascending node of the reference orbit and the running orbit; a 20 is the semi-major axis of the orbit; a 10is the semi-major axis of the reference orbit; C D is the atmospheric drag coefficient; A is the satellite's windward area; ω e is the angular velocity of the Earth's rotation; ρ is the atmospheric density; m is the mass of the satellite; μ is the Earth's gravitational constant; u t is the latitude argument at time t; u0 is the initial latitude argument; Δλ N0 is the initial ascending node longitude error.

[0022] As a preferred solution, the maneuvering time corresponding to the upper limit of the semi-major axis of the operating orbit and the maneuvering drift amount of each satellite are calculated based on the preset ascending node longitude formula, the orbital data and the upper limit of the semi-major axis of the operating orbit, respectively, including:

[0023] Based on the preset ascending node longitude formula, the upper limit of the semi-major axis of the running orbit and the orbit data, the maneuvering latitude argument is calculated;

[0024] Based on a preset latitude argument formula, the maneuvering latitude argument and the orbit data, a maneuvering time corresponding to the upper limit of the semi-major axis of the running orbit is calculated;

[0025] The maneuvering drift amount of each satellite corresponding to the upper limit of the semi-major axis of the operating orbit is calculated based on the maneuvering latitude argument, the preset ascending node longitude formula and the orbital data.

[0026] As a preferred solution, the calculation of the maneuvering latitude argument based on the preset ascending node longitude formula, the upper limit of the semi-major axis of the running orbit and the orbit data includes:

[0027] Based on the preset ascending node longitude formula, a second correlation relationship between the latitude argument and the semi-major axis of the orbit is obtained;

[0028] The maneuvering latitude argument is calculated based on the second correlation, the upper limit of the semi-major axis of the operating orbit and the orbital data.

[0029] As a preferred solution, the calculation of the maneuvering time corresponding to the upper limit of the semi-major axis of the running orbit based on the preset latitude argument formula, the maneuvering latitude argument and the orbit data includes:

[0030] Based on a preset latitude argument formula, a third correlation relationship between the latitude argument and the maneuvering time is obtained;

[0031] Calculating a maneuvering time according to the third correlation, the maneuvering latitude argument and the orbital data;

[0032] Wherein, the preset latitude argument formula is:

[0033]

[0034] In the formula, u t is the latitude argument at time t; u0 is the initial latitude argument; t0 is the initial time; μ is the earth's gravitational constant; a 20 is the semi-major axis of the orbit; is the time derivative of the semi-major axis of the orbit.

[0035] As a preferred solution, the maneuvering drift amount of each satellite corresponding to the upper limit of the semi-major axis of the operating orbit is calculated according to the maneuvering latitude argument, the preset ascending node longitude formula and the orbital data, including:

[0036] Obtaining satellite frontal areas and satellite masses corresponding to a plurality of satellites from the satellite basic data in the orbital data;

[0037] The maneuvering latitude argument, the satellite frontal area and satellite mass corresponding to each of the satellites, the operating orbit data and the reference orbit data are respectively substituted into the preset ascending node longitude formula to calculate the maneuvering drift amount of each of the satellites.

[0038] As a preferred solution, the first velocity increment of each satellite at the maneuvering moment is calculated based on the orbital data of the first constellation and the average maneuvering drift corresponding to each satellite group, including:

[0039] Calculating the drift difference of each satellite according to the maneuvering drift of each satellite and the average drift of each satellite group;

[0040] Calculating the semi-major axis of the phase modulation orbit of each satellite according to a preset semi-major axis formula, the drift difference of each satellite and the orbit data;

[0041] Calculating the phase modulation orbit velocity and the reference orbit velocity of each satellite at the maneuvering moment according to the phase modulation orbit semi-major axis of each satellite and the reference orbit semi-major axis;

[0042] Based on the phase-modulated orbital velocity and the reference orbital velocity of each satellite at the maneuvering moment, a first velocity increment of each satellite at the maneuvering moment is calculated.

[0043] As a preferred solution, the calculating of the semi-major axis of the phase modulation orbit of each satellite according to a preset semi-major axis formula, the drift difference of each satellite and the orbit data comprises:

[0044] Calculating the phase modulation period of each satellite based on the drift difference of each satellite, the reference satellite orbit period and the earth's rotation angular velocity;

[0045] Calculating the phase modulation orbit semi-major axis of each satellite according to a preset orbit semi-major axis formula, the phase modulation period of each satellite and the orbit data;

[0046] The semi-major axis formula of the preset orbit is:

[0047]

[0048] Where a is the semi-major axis of the orbit; T is the orbital period; μ is the Earth's gravitational constant.

[0049] As a preferred solution, after each of the satellites enters the phase modulation orbit from the current orbit, the method further includes:

[0050] Determine the upper limit of the semi-major axis of the operational orbit as the semi-major axis of the control orbit of each of the satellites;

[0051] Calculate the second velocity increment and the third velocity increment of each satellite according to the semi-major axis of the phase modulation orbit and the semi-major axis of the control orbit of each satellite;

[0052] The Hohmann orbit change method is adopted to adjust the running speed of each satellite based on the second speed increment and the third speed increment of each satellite, so that each satellite enters the control orbit from the phase modulation orbit.

[0053] Accordingly, the present invention provides a constellation configuration maintenance device based on sub-satellite point trajectory drift, comprising: a first acquisition module, a second acquisition module, an upper limit calculation module, a drift amount calculation module, a cluster analysis module, a velocity increment calculation module and a control module;

[0054] The first acquisition module is used to acquire orbital data of a first constellation; wherein the orbital data includes operating orbital data and reference orbital data corresponding to a plurality of satellites in the first constellation;

[0055] The second acquisition module is used to acquire the upper limit of the orbit drift of the first constellation;

[0056] The upper limit calculation module is used to calculate the upper limit of the semi-major axis of the running orbit according to the orbit data, the upper limit of the orbit drift, the preset orbit drift formula and the preset ascending node longitude formula;

[0057] The drift calculation module is used to calculate the maneuvering time corresponding to the upper limit of the semi-major axis of the running orbit and the maneuvering drift amount of each satellite based on the preset ascending node longitude formula, the orbit data and the upper limit of the semi-major axis of the running orbit;

[0058] The cluster analysis module is used to perform cluster analysis on each of the satellites according to the maneuver drift of each of the satellites, obtain a plurality of satellite clusters, and calculate the average maneuver drift corresponding to each satellite cluster;

[0059] The velocity increment calculation module is used to calculate the first velocity increment of each satellite at the maneuvering moment based on the orbit data of the first constellation and the average maneuvering drift corresponding to each satellite group;

[0060] The control module is used to adjust the operating state of each satellite based on the first speed increment corresponding to each satellite when the operating time of each satellite reaches the maneuvering time, so that each satellite enters the phase adjustment orbit from the current operating orbit to maintain the constellation configuration of the first constellation.

[0061] Compared with the prior art, the embodiments of the present invention have the following beneficial effects:

[0062] The present invention provides a constellation configuration maintenance method based on sub-satellite point trajectory drift. When the operating time of each satellite reaches the maneuvering time, it is considered that the orbit drift of the first constellation where each satellite is located has reached the upper limit of the orbit drift of the first constellation. At this time, the first speed increment of each satellite at the maneuvering time is calculated, and the operating state of each satellite is adjusted based on the first speed increment, so that each satellite enters the phase adjustment orbit from the current operating orbit, which can reduce the orbit drift of the first constellation and maintain the initial constellation configuration of the first constellation. The first speed increment of each satellite at the maneuvering time is calculated based on the cluster analysis of the maneuvering drift of each satellite at the maneuvering time. When calculating the first speed increment, the relative relationship between satellites is considered, so that the constellation performance can be maintained while maintaining the constellation configuration, and the accuracy of constellation maintenance control is improved. BRIEF DESCRIPTION OF THE DRAWINGS

[0063] Figure 1 A schematic flow chart of an embodiment of a constellation configuration maintenance method based on sub-satellite point trajectory drift provided by the present invention;

[0064] Figure 2 A schematic diagram of a track drift amount provided by the present invention;

[0065] Figure 3 A schematic diagram of a change in track drift provided by the present invention;

[0066] Figure 4 A schematic diagram of a flow chart of a satellite clustering method provided by the present invention;

[0067] Figure 5 A schematic diagram of satellite clustering provided by the present invention;

[0068] Figure 6 A schematic flow chart of another embodiment of a constellation configuration maintenance method based on sub-satellite point trajectory drift provided by the present invention;

[0069] Figure 7A schematic structural diagram of an embodiment of a constellation configuration maintaining device based on sub-satellite point trajectory drift provided by the present invention. DETAILED DESCRIPTION

[0070] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0071] The flowcharts shown in the accompanying drawings are only examples and do not necessarily include all the contents and operations / steps, nor must they be executed in the order described. For example, some operations / steps may also be decomposed, combined or partially merged, so the actual execution order may change according to actual conditions.

[0072] Some embodiments of the present invention are described in detail below in conjunction with the accompanying drawings. In the absence of conflict, the following embodiments and features in the embodiments can be combined with each other.

[0073] Example 1

[0074] like Figure 1 FIG. 1 is a flow chart of an embodiment of a constellation configuration maintenance method based on sub-satellite point trajectory drift provided by the present invention. The method includes steps 101 to 107, and each step is specifically as follows:

[0075] Step 101: Acquire orbital data of a first constellation; wherein the orbital data includes operating orbital data and reference orbital data corresponding to a number of satellites in the first constellation.

[0076] In an embodiment of the present invention, the first constellation is a homogeneous constellation or a heterogeneous constellation composed of multiple satellites. The orbital data of the first constellation includes the operating orbit data and reference orbit data corresponding to each satellite in the first constellation. The operating orbit data is the data corresponding to the orbit on which each satellite actually operates, including satellite basic data and orbit basic data. The satellite basic data includes satellite latitude argument, satellite mass and satellite frontal area, etc. The orbit basic data includes the semi-major axis of the operating orbit, the inclination of the operating orbit and the longitude of the ascending node of the operating orbit, etc. The reference orbit data is the operating orbit pre-planned by each satellite, including satellite basic data and orbit basic data. The satellite basic data includes satellite latitude argument, satellite mass and satellite frontal area, etc. The orbit basic data includes the pre-set reference orbit semi-major axis, reference orbit inclination and the longitude of the ascending node of the reference orbit, etc.

[0077] Step 102: Obtain the upper limit of the orbital drift of the first constellation.

[0078] In an embodiment of the present invention, in order to maintain the constellation configuration, each satellite in the first constellation is controlled to operate according to the reference orbit. However, as time goes by, the sub-satellite point trajectory of the actual operating orbit of each satellite will gradually deviate from the sub-satellite point trajectory of the reference orbit. In the actual operation of the satellite, the orbital inclination determines the drift between the actual operating orbit and the reference orbit. Therefore, in order to control the orbital offset of each satellite, it is necessary to control the orbital inclination. The effect of atmospheric drag on the orbital inclination can be ignored, so the orbital inclination can be controlled by controlling the semi-major axis of the orbit. Since the satellite will generate a drift between the reference orbit at every moment of operation, it is necessary to determine the time to control the semi-major axis of each satellite orbit. In this embodiment, when the orbital drift of the satellite reaches the upper limit of the orbital drift, it is determined as the time to control the semi-major axis of the orbit.

[0079] In an embodiment of the present invention, the upper limit of the orbital drift of the first constellation can be determined according to the mission requirements, and the upper limit of the orbital drift is subsequently used as a basic control benchmark. As an example of an embodiment of the present invention, assuming that the mission requirement is: the distance difference between the sensor-to-ground projection and the target point does not exceed 5 km, then 5 km can be used as the upper limit of the orbital drift of the first constellation.

[0080] Step 103: Calculate the upper limit of the semi-major axis of the running orbit based on the orbit data, the upper limit of the orbit drift, the preset orbit drift formula and the preset ascending node longitude formula.

[0081] As a preferred solution of this embodiment, the upper limit of the semi-major axis of the running orbit is calculated based on the orbit data, the upper limit of the orbit drift, the preset orbit drift formula and the preset ascending node longitude formula, including:

[0082] Based on the preset orbit drift formula and the preset ascending node longitude formula, a first correlation relationship between the orbit drift and the semi-major axis of the running orbit is obtained;

[0083] Calculate the upper limit of the semi-major axis of the running orbit according to the first correlation, the upper limit of the orbit drift and the orbit data;

[0084] The preset orbit drift formula is:

[0085]

[0086] Where D σ is the orbital drift; R e is the radius of the earth; Δi is the inclination difference between the reference orbit and the running orbit; Δλ N is the difference in longitude between the ascending node of the reference orbit and the running orbit; i1 is the inclination of the reference orbit;

[0087] The preset ascending node longitude formula is:

[0088]

[0089] In the formula, Δλ N is the difference in longitude between the ascending node of the reference orbit and the running orbit; a 20 is the semi-major axis of the orbit; a 10 is the semi-major axis of the reference orbit; C D is the atmospheric drag coefficient; A is the satellite's windward area; ω e is the angular velocity of the Earth's rotation; ρ is the atmospheric density; m is the mass of the satellite; μ is the Earth's gravitational constant; u t is the latitude argument at time t; u0 is the initial latitude argument; Δλ N0 is the initial ascending node longitude error.

[0090] In the embodiments of the present invention, see Figure 2 , is a schematic diagram of a track drift provided by the present invention, and the track offset is defined as D σ =R e φ, where R e is the radius of the earth, and φ is the angle between the reference orbital plane and the running orbital plane. The expression can be obtained by approximating the spherical triangle and Taylor expansion: Among them, Δi=i2-i1 is the inclination difference between the reference orbit and the running orbit, i1 is the inclination of the reference orbit, Δλ N =λ N1 -λ N2 It represents the difference in longitude between the ascending node of the reference orbit and the running orbit. Therefore, the formula for the preset orbit drift can be obtained. According to the preset orbital drift formula, the orbital drift is determined by the difference in inclination, the difference in the longitude of the ascending node, and the orbital inclination of the reference plane. In practice, the effect of atmospheric drag on the inclination can be ignored, that is, Δi and i1 are both constants, then the orbital drift is given by Δλ N Decide.

[0091] In the embodiment of the present invention, under the influence of atmospheric drag, the semi-major axis of the satellite's orbit will continue to decay, and the change of its decay over time is as follows:

[0092]

[0093] Among them, C D represents the atmospheric drag coefficient; A is the satellite's windward area; is the angular velocity of the satellite; a is the semi-major axis of the orbit; ρ is the atmospheric density; and m is the mass of the satellite.

[0094] Since the semi-major axis of the satellite's orbit decays over time, the change in the satellite's orbital period is as follows:

[0095]

[0096] Among them, u=ω+M represents the latitude argument, ω represents the perigee depression angle, M represents the mean anomaly angle, and μ is the earth's gravitational constant.

[0097] In the embodiment of the present invention, the change of the longitude of the ascending node is defined as:

[0098]

[0099] Substituting the changes of the semi-major axis of the orbit and the changes of the orbit period into the above formula and simplifying it, the formula for the preset ascending node longitude is obtained as follows:

[0100]

[0101] According to the preset ascending node longitude formula, Δλ N is a u t The quadratic function of , therefore, combined with the preset orbital drift formula, the orbital drift D can be obtained σ and the semi-major axis of the orbit a 20 Substituting the upper limit of the orbital drift into the first correlation, the upper limit of the semi-major axis of the orbit can be calculated.

[0102] Step 104: Based on the preset ascending node longitude formula, the orbital data and the upper limit of the semi-major axis of the operating orbit, the maneuvering time corresponding to the upper limit of the semi-major axis of the operating orbit and the maneuvering drift amount of each of the satellites are calculated respectively.

[0103] As a preferred solution of this embodiment, based on the preset ascending node longitude formula, the orbital data and the upper limit of the semi-major axis of the operating orbit, the maneuvering time corresponding to the upper limit of the semi-major axis of the operating orbit and the maneuvering drift amount of each satellite are calculated respectively, including:

[0104] Based on the preset ascending node longitude formula, the upper limit of the semi-major axis of the running orbit and the orbit data, the maneuvering latitude argument is calculated;

[0105] Based on a preset latitude argument formula, the maneuvering latitude argument and the orbit data, a maneuvering time corresponding to the upper limit of the semi-major axis of the running orbit is calculated;

[0106] The maneuvering drift amount of each satellite corresponding to the upper limit of the semi-major axis of the operating orbit is calculated based on the maneuvering latitude argument, the preset ascending node longitude formula and the orbital data.

[0107] As a preferred solution of this embodiment, based on the preset ascending node longitude formula, the upper limit of the semi-major axis of the running orbit and the orbit data, the maneuvering latitude argument is calculated, including:

[0108] Based on the preset ascending node longitude formula, a second correlation relationship between the latitude argument and the semi-major axis of the orbit is obtained;

[0109] The maneuvering latitude argument is calculated based on the second correlation, the upper limit of the semi-major axis of the operating orbit and the orbital data.

[0110] In the embodiment of the present invention, according to the preset ascending node longitude formula, Δλ N is a u t The quadratic function of t and the semi-major axis of the orbit a 20 Substituting the upper limit of the orbital semi-major axis calculated above into the second correlation, the maneuvering latitude argument can be calculated.

[0111] As a preferred solution of this embodiment, based on the preset latitude argument formula, the maneuvering latitude argument and the orbit data, the maneuvering time corresponding to the upper limit of the semi-major axis of the running orbit is calculated, including:

[0112] Based on a preset latitude argument formula, a third correlation relationship between the latitude argument and the maneuvering time is obtained;

[0113] Calculating a maneuvering time according to the third correlation, the maneuvering latitude argument and the orbital data;

[0114] Wherein, the preset latitude argument formula is:

[0115]

[0116] In the formula, u t is the latitude argument at time t; u0 is the initial latitude argument; t0 is the initial time; μ is the earth's gravitational constant; a 20 is the semi-major axis of the orbit; is the time derivative of the semi-major axis of the orbit.

[0117] In the embodiment of the present invention, when the orbital drift of the satellite reaches the upper limit of the orbital drift, the moment of controlling the semi-major axis of the orbit is determined to be the maneuvering moment, and the latitude argument corresponding to this moment is the maneuvering latitude argument. The latitude argument u can be obtained in the preset latitude argument formula t The third correlation between t and the maneuvering time t, so the maneuvering latitude angle calculated above is substituted into the third correlation to calculate the maneuvering time. The maneuvering time is the last time the satellite's orbital drift reaches the upper limit of the orbital drift Dlim time.

[0118] In the embodiments of the present invention, see Figure 3 , is a schematic diagram of the change of orbital drift provided by the present invention, the horizontal axis is the satellite's operating time t, in days, and the vertical axis is the orbital drift D σ , in kilometers. The waveform shows how the orbit drift changes over time. There are two moments when the orbit drift reaches the upper limit of the orbit drift. lim , when the first orbital drift limit D is reached lim After the moment, the orbital drift will not continue to exceed the orbital drift limit D lim , but decreases with the passage of time; in the second lim After the moment, the orbital drift will continue to exceed the orbital drift limit D as time goes by. lim , controlling at this time can reduce the frequency of control and reduce costs. Therefore, the second one that reaches the upper limit of orbit drift D lim The moment is determined as the maneuvering time.

[0119] As a preferred solution of this embodiment, the maneuvering drift amount of each satellite corresponding to the upper limit of the semi-major axis of the operating orbit is calculated according to the maneuvering latitude argument, the preset ascending node longitude formula and the orbit data, including:

[0120] Obtaining satellite frontal areas and satellite masses corresponding to a plurality of satellites from the satellite basic data in the orbital data;

[0121] The maneuvering latitude argument, the satellite frontal area and satellite mass corresponding to each of the satellites, the operating orbit data and the reference orbit data are respectively substituted into the preset ascending node longitude formula to calculate the maneuvering drift amount of each of the satellites.

[0122] In an embodiment of the present invention, basic data of each satellite in the first constellation, including the satellite frontal area and satellite mass, etc., are obtained respectively, and the data of each satellite and the upper limit of the semi-major axis of the orbit and the maneuvering latitude argument obtained by the above calculation are substituted into the preset ascending node longitude formula, so that the maneuvering drift of each satellite at the maneuvering time can be calculated respectively.

[0123] Step 105: performing cluster analysis on the satellites according to the maneuvering drift of each satellite to obtain a plurality of satellite clusters, and calculating the average maneuvering drift corresponding to each satellite cluster.

[0124] In an embodiment of the present invention, each satellite of the first constellation is controlled at the maneuvering moment, and the first speed increment corresponding to each satellite needs to be calculated. If all satellites are individually adjusted according to their respective drift amounts, the required speed increment will be greatly increased. In order to maintain the constellation performance while maintaining the constellation configuration, the relative position relationship of the sub-satellite point trajectory of each satellite needs not to change too much. Therefore, the multiple satellites of the first constellation can be divided into several groups, and the groups can be controlled separately. If all the grouping is performed according to a unified standard, satellites that do not need to be controlled may also participate in the control to increase fuel consumption. Therefore, cluster analysis can be used to reasonably classify satellites.

[0125] In the embodiments of the present invention, see Figure 4 , is a schematic diagram of a satellite clustering method provided by the present invention. In this embodiment, the K-Means algorithm can be used for clustering analysis. First, the number of satellite groups K is set according to the needs of the user, and K center points are initialized according to the number of groups K, and then the distance d between all satellites and each center point is calculated. i,j , d i,j is the distance from the i-th satellite to the j-th center point, i∈[1,n],j∈[1,k], which can be calculated using the Euclidean distance formula. Then, the satellites are grouped according to the distance between each satellite and each center point, and the K center points are recalculated after grouping to determine whether the center points before and after the recalculation have changed. If there is no change, the grouping information is output to obtain K groups of satellite groups. When the multiple satellites of the first constellation are divided into multiple groups of satellite groups, for each satellite group, the average value of the maneuvering drift corresponding to each satellite in the satellite group is calculated to obtain the average maneuvering drift of each satellite group. See Figure 5 , is a satellite clustering schematic diagram provided by the present invention. Assuming that the number of satellite groups is 3, the first group of center points, the second group of center points and the third group of center points are set respectively, and the data points near the center points of each group are classified as satellites of the same satellite group.

[0126] Step 106: Based on the orbital data of the first constellation and the average maneuvering drift corresponding to each of the satellite groups, respectively calculate the first velocity increment of each of the satellites at the maneuvering moment.

[0127] As a preferred solution of this embodiment, based on the orbital data of the first constellation and the average maneuvering drift corresponding to each of the satellite groups, respectively calculating the first velocity increment of each of the satellites at the maneuvering moment includes:

[0128] Calculating the drift difference of each satellite according to the maneuvering drift of each satellite and the average drift of each satellite group;

[0129] Calculating the semi-major axis of the phase modulation orbit of each satellite according to a preset semi-major axis formula, the drift difference of each satellite and the orbit data;

[0130] Calculating the phase modulation orbit velocity and the reference orbit velocity of each satellite at the maneuvering moment according to the phase modulation orbit semi-major axis of each satellite and the reference orbit semi-major axis;

[0131] Based on the phase-modulated orbital velocity and the reference orbital velocity of each satellite at the maneuvering moment, a first velocity increment of each satellite at the maneuvering moment is calculated.

[0132] As a preferred solution of this embodiment, the semi-major axis of the phase modulation orbit of each satellite is calculated according to a preset semi-major axis formula, the drift difference of each satellite and the orbit data, including:

[0133] Calculating the phase modulation period of each satellite based on the drift difference of each satellite, the reference satellite orbit period and the earth's rotation angular velocity;

[0134] Calculating the phase modulation orbit semi-major axis of each satellite according to a preset orbit semi-major axis formula, the phase modulation period of each satellite and the orbit data;

[0135] The semi-major axis formula of the preset orbit is:

[0136]

[0137] Where a is the semi-major axis of the orbit; T is the orbital period; μ is the Earth's gravitational constant.

[0138] In the embodiment of the present invention, when the running time of the satellite reaches the maneuvering time, the phase adjustment maneuvering strategy can be used to control each satellite to enter the phase adjustment orbit. At this time, the orbital semi-major axis of the phase adjustment orbit needs to be calculated, and the following preset semi-major axis formula can be used for calculation:

[0139]

[0140] Where a is the orbital semi-major axis of the phase modulation orbit; T is the orbital period of the phase modulation orbit, which is determined according to the phase that the satellite needs to adjust, and the phase can be determined according to the mean difference between each satellite and the satellite group in which it is located:

[0141]

[0142] Among them, Δλ gi Indicates the drift of the i-th satellite in the g-th group; represents the average maneuver drift of group g.

[0143] In the calculation of δλ jAfter that, the required phase modulation size of the phase modulation orbit can be calculated:

[0144] Δθ=nδt

[0145] δt=δλ i / ω e

[0146] Where n represents the angular velocity of the reference orbit, according to δλ j The numerical value of can determine the control drift direction of the sub-satellite point trajectory. If δλ i > 0, the subsatellite point trajectory needs to be further moved eastward to compensate for the missing drift. If δλ i <0, the sub-satellite point trajectory needs to move westward to reduce the amount of drift. According to the control time δt, the corresponding phase modulation orbit period can be calculated as follows:

[0147] T=T0-δt,δλ i >0

[0148] T=T0+δt,δλ i <0

[0149] where T0 refers to the period of the orbit.

[0150] In the embodiment of the present invention, after the orbital semi-major axis of the phase modulation orbit is obtained according to the above analysis and calculation, the first velocity increment is determined according to the orbital semi-major axis:

[0151]

[0152] ΔV=2|v2-v1|

[0153] Where h1 and h2 are the unit relative angular momentum of the reference orbit and the phase modulation orbit respectively; r a is the semi-major axis of the reference orbit.

[0154] Step 107: When the operating time of each satellite reaches the maneuvering time, based on the first speed increment corresponding to each satellite, the operating state of each satellite is adjusted so that each satellite enters a phase adjustment orbit from a current operating orbit to maintain the constellation configuration of the first constellation.

[0155] In an embodiment of the present invention, after the first velocity increment corresponding to each satellite is calculated, when the operating time of each satellite reaches the maneuvering time, the first velocity increment is increased for each satellite, so that each satellite enters the phase adjustment orbit from the current operating orbit, thereby reducing the orbit drift of the first constellation and maintaining the initial constellation configuration of the first constellation.

[0156] As a preferred solution of this embodiment, after each satellite enters the phase modulation orbit from the current orbit, the method further includes:

[0157] Determine the upper limit of the semi-major axis of the operational orbit as the semi-major axis of the control orbit of each of the satellites;

[0158] Calculate the second velocity increment and the third velocity increment of each satellite according to the semi-major axis of the phase modulation orbit and the semi-major axis of the control orbit of each satellite;

[0159] The Hohmann orbit change method is adopted to adjust the running speed of each satellite based on the second speed increment and the third speed increment of each satellite, so that each satellite enters the control orbit from the phase modulation orbit.

[0160] In an embodiment of the present invention, after each satellite of the first constellation enters the phase adjustment orbit from the current operating orbit through control, in order to ensure that the frequency of constellation configuration maintenance control is reduced, the Hohmann orbit change method is used to move the satellite from the phase adjustment orbit to the control orbit. The Hohmann orbit change includes two stages, one of which is to restore the semi-major axis of the orbit to the initial length, control the relative drift of its sub-satellite point trajectory, and increase the semi-major axis to reduce the impact of its drift. The Hohmann orbit change can be controlled by the following formula:

[0161]

[0162] Where Δv θ1 is the first speed increment corresponding to the first speed pulse; Δv θ2 is the second velocity increment corresponding to the second velocity pulse; r1 and v1 are the semi-major axis and velocity of the orbit just before the first velocity pulse, r2 and v2 are the semi-major axis and velocity just after the second velocity pulse; Δr=r2-r1 is the difference in the semi-major axis before and after the orbit change.

[0163] In this embodiment, each satellite is moved from the phase modulation orbit to the control orbit. Therefore, r1 and v1 correspond to the orbital semi-major axis and velocity of the phase modulation orbit; r2 and v2 correspond to the orbital semi-major axis and velocity of the control orbit. The semi-major axis of the control orbit is determined by the upper limit of the semi-major axis of the operating orbit calculated above. By adjusting the semi-major axis of the actual operating orbit of the satellite, this embodiment can reduce the control frequency so that the orbital drift of each satellite returns to Figure 3 The initial state on the far left, when the orbital drift of each satellite reaches Figure 3 When it reaches the rightmost intersection state, perform phase adjustment automatic control again.

[0164] See also Figure 6 , is a flow chart of another embodiment of the constellation configuration maintenance method based on sub-satellite point trajectory drift provided by the present invention. The control flow of the embodiment of the present invention is: according to the mission requirements, the upper limit D of the sub-satellite point trajectory drift is determined. lim(This drift can comprehensively consider the overall drift of the sub-satellite point trajectory, including the equator and high latitude areas). lim Determine the upper limit of the orbital semi-major axis a 20 Then, the time u required for maneuvering is determined by the semi-major axis attenuation formula and the ascending node longitude drift formula. t , and confirm the maneuver drift Δλ at the equator at that moment. The maneuver drift Δλ of different satellites is clustered using the K-Means algorithm to re-determine the drift of each satellite considering the relative relationship between satellites, and then calculate the speed increment required for control.

[0165] Implementing the above embodiments has the following effects:

[0166] The present invention provides a constellation configuration maintenance method based on sub-satellite point trajectory drift. When the operating time of each satellite reaches the maneuvering time, it is considered that the orbit drift of the first constellation where each satellite is located has reached the upper limit of the orbit drift of the first constellation. At this time, the first speed increment of each satellite at the maneuvering time is calculated, and the operating state of each satellite is adjusted based on the first speed increment, so that each satellite enters the phase adjustment orbit from the current operating orbit, which can reduce the orbit drift of the first constellation and maintain the initial constellation configuration of the first constellation. The first speed increment of each satellite at the maneuvering time is calculated based on the cluster analysis of the maneuvering drift of each satellite at the maneuvering time. When calculating the first speed increment, the relative relationship between satellites is considered, so that the constellation performance can be maintained while maintaining the constellation configuration, and the accuracy of constellation maintenance control is improved.

[0167] Example 2

[0168] See also Figure 7 , is a schematic structural diagram of an embodiment of a constellation configuration maintaining device based on sub-satellite point trajectory drift provided by the present invention, the device comprising a first acquisition module 201, a second acquisition module 202, an upper limit calculation module 203, a drift amount calculation module 204, a cluster analysis module 205, a velocity increment calculation module 206 and a control module 207;

[0169] The first acquisition module 201 is used to acquire orbit data of a first constellation; wherein the orbit data includes operating orbit data and reference orbit data corresponding to a plurality of satellites in the first constellation;

[0170] The second acquisition module 202 is used to obtain the upper limit of the orbit drift of the first constellation;

[0171] The upper limit calculation module 203 is used to calculate the upper limit of the semi-major axis of the running orbit according to the orbit data, the upper limit of the orbit drift, the preset orbit drift formula and the preset ascending node longitude formula;

[0172] The drift calculation module 204 is used to calculate the maneuvering time corresponding to the upper limit of the semi-major axis of the running orbit and the maneuvering drift amount of each satellite based on the preset ascending node longitude formula, the orbit data and the upper limit of the semi-major axis of the running orbit;

[0173] The cluster analysis module 205 is used to perform cluster analysis on each of the satellites according to the maneuver drift of each of the satellites, obtain a number of satellite clusters, and calculate the average maneuver drift corresponding to each satellite cluster;

[0174] The velocity increment calculation module 206 is used to calculate the first velocity increment of each satellite at the maneuvering moment based on the orbit data of the first constellation and the average maneuvering drift corresponding to each satellite group;

[0175] The control module 207 is used to adjust the operating state of each satellite based on the first speed increment corresponding to each satellite when the operating time of each satellite reaches the maneuvering time, so that each satellite enters the phase adjustment orbit from the current operating orbit to maintain the constellation configuration of the first constellation.

[0176] As a preferred solution of this embodiment, the upper limit of the semi-major axis of the running orbit is calculated based on the orbit data, the upper limit of the orbit drift, the preset orbit drift formula and the preset ascending node longitude formula, including:

[0177] Based on the preset orbit drift formula and the preset ascending node longitude formula, a first correlation relationship between the orbit drift and the semi-major axis of the running orbit is obtained;

[0178] Calculate the upper limit of the semi-major axis of the running orbit according to the first correlation, the upper limit of the orbit drift and the orbit data;

[0179] The preset orbit drift formula is:

[0180]

[0181] Where D σ is the orbital drift; R e is the radius of the earth; Δi is the inclination difference between the reference orbit and the running orbit; Δλ N is the difference in longitude between the ascending node of the reference orbit and the running orbit; i1 is the inclination of the reference orbit;

[0182] The preset ascending node longitude formula is:

[0183]

[0184] In the formula, Δλ N is the difference in longitude between the ascending node of the reference orbit and the running orbit; a 20is the semi-major axis of the orbit; a 10 is the semi-major axis of the reference orbit; C D is the atmospheric drag coefficient; A is the satellite's windward area; ω e is the angular velocity of the Earth's rotation; ρ is the atmospheric density; m is the mass of the satellite; μ is the Earth's gravitational constant; u t is the latitude argument at time t; u0 is the initial latitude argument; Δλ N0 is the initial ascending node longitude error.

[0185] As a preferred solution of this embodiment, based on the preset ascending node longitude formula, the orbital data and the upper limit of the semi-major axis of the operating orbit, the maneuvering time corresponding to the upper limit of the semi-major axis of the operating orbit and the maneuvering drift amount of each satellite are calculated respectively, including:

[0186] Based on the preset ascending node longitude formula, the upper limit of the semi-major axis of the running orbit and the orbit data, the maneuvering latitude argument is calculated;

[0187] Based on a preset latitude argument formula, the maneuvering latitude argument and the orbit data, a maneuvering time corresponding to the upper limit of the semi-major axis of the running orbit is calculated;

[0188] The maneuvering drift amount of each satellite corresponding to the upper limit of the semi-major axis of the operating orbit is calculated based on the maneuvering latitude argument, the preset ascending node longitude formula and the orbital data.

[0189] As a preferred solution of this embodiment, based on the preset ascending node longitude formula, the upper limit of the semi-major axis of the running orbit and the orbit data, the maneuvering latitude argument is calculated, including:

[0190] Based on the preset ascending node longitude formula, a second correlation relationship between the latitude argument and the semi-major axis of the orbit is obtained;

[0191] The maneuvering latitude argument is calculated based on the second correlation, the upper limit of the semi-major axis of the operating orbit and the orbital data.

[0192] As a preferred solution of this embodiment, based on the preset latitude argument formula, the maneuvering latitude argument and the orbit data, the maneuvering time corresponding to the upper limit of the semi-major axis of the running orbit is calculated, including:

[0193] Based on a preset latitude argument formula, a third correlation relationship between the latitude argument and the maneuvering time is obtained;

[0194] Calculating a maneuvering time according to the third correlation, the maneuvering latitude argument and the orbital data;

[0195] Wherein, the preset latitude argument formula is:

[0196]

[0197] In the formula, u t is the latitude argument at time t; u0 is the initial latitude argument; t0 is the initial time; μ is the earth's gravitational constant; a 20 is the semi-major axis of the orbit; is the time derivative of the semi-major axis of the orbit.

[0198] As a preferred solution of this embodiment, the maneuvering drift amount of each satellite corresponding to the upper limit of the semi-major axis of the operating orbit is calculated according to the maneuvering latitude argument, the preset ascending node longitude formula and the orbit data, including:

[0199] Obtaining satellite frontal areas and satellite masses corresponding to a plurality of satellites from the satellite basic data in the orbital data;

[0200] The maneuvering latitude argument, the satellite frontal area and satellite mass corresponding to each of the satellites, the operating orbit data and the reference orbit data are respectively substituted into the preset ascending node longitude formula to calculate the maneuvering drift amount of each of the satellites.

[0201] As a preferred solution of this embodiment, based on the orbital data of the first constellation and the average maneuvering drift corresponding to each of the satellite groups, respectively calculating the first velocity increment of each of the satellites at the maneuvering moment includes:

[0202] Calculating the drift difference of each satellite according to the maneuvering drift of each satellite and the average drift of each satellite group;

[0203] Calculating the semi-major axis of the phase modulation orbit of each satellite according to a preset semi-major axis formula, the drift difference of each satellite and the orbit data;

[0204] Calculating the phase modulation orbit velocity and the reference orbit velocity of each satellite at the maneuvering moment according to the phase modulation orbit semi-major axis of each satellite and the reference orbit semi-major axis;

[0205] Based on the phase-modulated orbital velocity and the reference orbital velocity of each satellite at the maneuvering moment, a first velocity increment of each satellite at the maneuvering moment is calculated.

[0206] As a preferred solution of this embodiment, the semi-major axis of the phase modulation orbit of each satellite is calculated according to a preset semi-major axis formula, the drift difference of each satellite and the orbit data, including:

[0207] Calculating the phase modulation period of each satellite based on the drift difference of each satellite, the reference satellite orbit period and the earth's rotation angular velocity;

[0208] Calculating the phase modulation orbit semi-major axis of each satellite according to a preset orbit semi-major axis formula, the phase modulation period of each satellite and the orbit data;

[0209] The semi-major axis formula of the preset orbit is:

[0210]

[0211] Where a is the semi-major axis of the orbit; T is the orbital period; μ is the Earth's gravitational constant.

[0212] As a preferred solution of this embodiment, after each satellite enters the phase modulation orbit from the current orbit, the method further includes:

[0213] Determine the upper limit of the semi-major axis of the operational orbit as the semi-major axis of the control orbit of each of the satellites;

[0214] Calculate the second velocity increment and the third velocity increment of each satellite according to the semi-major axis of the phase modulation orbit and the semi-major axis of the control orbit of each satellite;

[0215] The Hohmann orbit change method is adopted to adjust the running speed of each satellite based on the second speed increment and the third speed increment of each satellite, so that each satellite enters the control orbit from the phase modulation orbit.

[0216] Implementing the above embodiments has the following effects:

[0217] The present invention provides a constellation configuration maintenance method based on sub-satellite point trajectory drift. When the operating time of each satellite reaches the maneuvering time, it is considered that the orbit drift of the first constellation where each satellite is located has reached the upper limit of the orbit drift of the first constellation. At this time, the first speed increment of each satellite at the maneuvering time is calculated, and the operating state of each satellite is adjusted based on the first speed increment, so that each satellite enters the phase adjustment orbit from the current operating orbit, which can reduce the orbit drift of the first constellation and maintain the initial constellation configuration of the first constellation. The first speed increment of each satellite at the maneuvering time is calculated based on the cluster analysis of the maneuvering drift of each satellite at the maneuvering time. When calculating the first speed increment, the relative relationship between satellites is considered, so that the constellation performance can be maintained while maintaining the constellation configuration, and the accuracy of constellation maintenance control is improved.

[0218] The specific embodiments described above further illustrate the purpose, technical solutions and beneficial effects of the present invention. It should be understood that the above description is only a specific embodiment of the present invention and is not intended to limit the scope of protection of the present invention. It is particularly pointed out that for those skilled in the art, any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present invention should be included in the scope of protection of the present invention.

Claims

1. A constellation configuration maintenance method based on sub-satellite point trajectory drift, characterized in that: include: Acquire orbital data of a first constellation; wherein the orbital data includes operating orbital data and reference orbital data corresponding to a plurality of satellites in the first constellation; Get the upper limit of the orbital drift of the first constellation; The upper limit of the semi-major axis of the running orbit is calculated based on the orbit data, the upper limit of the orbit drift, the preset orbit drift formula and the preset ascending node longitude formula; Based on the preset ascending node longitude formula, the orbital data and the upper limit of the semi-major axis of the operating orbit, respectively calculating the maneuvering time corresponding to the upper limit of the semi-major axis of the operating orbit and the maneuvering drift amount of each of the satellites; According to the maneuvering drift of each satellite, cluster analysis is performed on each satellite to obtain a plurality of satellite clusters, and an average maneuvering drift corresponding to each satellite cluster is calculated; Based on the orbital data of the first constellation and the average maneuvering drift corresponding to each of the satellite groups, respectively calculating the first velocity increment of each of the satellites at the maneuvering moment; When the operating time of each satellite reaches the maneuvering time, the operating state of each satellite is adjusted based on the first speed increment corresponding to each satellite, so that each satellite enters the phase adjustment orbit from the current operating orbit to maintain the constellation configuration of the first constellation.

2. The constellation configuration maintenance method based on sub-satellite point trajectory drift according to claim 1, characterized in that: The calculating the upper limit of the semi-major axis of the running orbit according to the orbit data, the upper limit of the orbit drift, the preset orbit drift formula and the preset ascending node longitude formula includes: Based on the preset orbit drift formula and the preset ascending node longitude formula, a first correlation relationship between the orbit drift and the semi-major axis of the running orbit is obtained; Calculate the upper limit of the semi-major axis of the running orbit according to the first correlation, the upper limit of the orbit drift and the orbit data; The preset orbit drift formula is: Where D σ is the orbital drift; R e is the radius of the earth; Δi is the inclination difference between the reference orbit and the running orbit; Δλ N is the difference in longitude between the ascending node of the reference orbit and the running orbit; i1 is the inclination of the reference orbit; The preset ascending node longitude formula is: In the formula, Δλ N is the difference in longitude between the ascending node of the reference orbit and the running orbit; a 20 is the semi-major axis of the orbit; a 10 is the semi-major axis of the reference orbit; C D is the atmospheric drag coefficient; A is the satellite's windward area; ω e is the angular velocity of the Earth's rotation; ρ is the atmospheric density; m is the mass of the satellite; μ is the Earth's gravitational constant; u t is the latitude argument at time t; u0 is the initial latitude argument; Δλ N0 is the initial ascending node longitude error.

3. The constellation configuration maintenance method based on sub-satellite point trajectory drift according to claim 2, characterized in that: The calculating, based on the preset ascending node longitude formula, the orbital data and the upper limit of the semi-major axis of the operating orbit, respectively obtaining the maneuvering time corresponding to the upper limit of the semi-major axis of the operating orbit and the maneuvering drift amount of each satellite comprises: Based on the preset ascending node longitude formula, the upper limit of the semi-major axis of the running orbit and the orbit data, the maneuvering latitude argument is calculated; Based on a preset latitude argument formula, the maneuvering latitude argument and the orbit data, a maneuvering time corresponding to the upper limit of the semi-major axis of the running orbit is calculated; The maneuvering drift amount of each satellite corresponding to the upper limit of the semi-major axis of the operating orbit is calculated based on the maneuvering latitude argument, the preset ascending node longitude formula and the orbital data.

4. The constellation configuration maintenance method based on sub-satellite point trajectory drift according to claim 3, characterized in that: The calculating the maneuvering latitude argument based on the preset ascending node longitude formula, the upper limit of the semi-major axis of the running orbit and the orbit data includes: Based on the preset ascending node longitude formula, a second correlation relationship between the latitude argument and the semi-major axis of the orbit is obtained; The maneuvering latitude argument is calculated based on the second correlation, the upper limit of the semi-major axis of the operating orbit and the orbital data.

5. The constellation configuration maintenance method based on sub-satellite point trajectory drift according to claim 4, characterized in that: The calculating of the maneuvering time corresponding to the upper limit of the semi-major axis of the running orbit based on the preset latitude argument formula, the maneuvering latitude argument and the orbit data comprises: Based on a preset latitude argument formula, a third correlation relationship between the latitude argument and the maneuvering time is obtained; Calculating a maneuvering time according to the third correlation, the maneuvering latitude argument and the orbital data; Wherein, the preset latitude argument formula is: In the formula, u t is the latitude argument at time t; u0 is the initial latitude argument; t0 is the initial time; μ is the earth's gravitational constant; a 20 is the semi-major axis of the orbit; is the time derivative of the semi-major axis of the orbit.

6. The constellation configuration maintenance method based on sub-satellite point trajectory drift according to claim 5, characterized in that: The calculating, according to the maneuvering latitude argument, the preset ascending node longitude formula and the orbital data, the maneuvering drift amount of each satellite corresponding to the upper limit of the semi-major axis of the operating orbit comprises: Obtaining satellite frontal areas and satellite masses corresponding to a plurality of satellites from the satellite basic data in the orbital data; The maneuvering latitude argument, the satellite frontal area and satellite mass corresponding to each of the satellites, the operating orbit data and the reference orbit data are respectively substituted into the preset ascending node longitude formula to calculate the maneuvering drift amount of each of the satellites.

7. The constellation configuration maintenance method based on sub-satellite point trajectory drift according to claim 6, characterized in that: The calculating, based on the orbital data of the first constellation and the average maneuvering drift corresponding to each of the satellite groups, the first velocity increment of each of the satellites at the maneuvering moment comprises: Calculating the drift difference of each satellite according to the maneuvering drift of each satellite and the average drift of each satellite group; Calculating the semi-major axis of the phase modulation orbit of each satellite according to a preset semi-major axis formula, the drift difference of each satellite and the orbit data; Calculating the phase modulation orbit velocity and the reference orbit velocity of each satellite at the maneuvering moment according to the phase modulation orbit semi-major axis of each satellite and the reference orbit semi-major axis; Based on the phase-modulated orbital velocity and the reference orbital velocity of each satellite at the maneuvering moment, a first velocity increment of each satellite at the maneuvering moment is calculated.

8. The constellation configuration maintenance method based on sub-satellite point trajectory drift according to claim 7, characterized in that: The calculating the semi-major axis of the phase modulation orbit of each satellite according to a preset semi-major axis formula, the drift difference of each satellite and the orbit data comprises: Calculating the phase modulation period of each satellite based on the drift difference of each satellite, the reference satellite orbit period and the earth's rotation angular velocity; Calculating the phase modulation orbit semi-major axis of each satellite according to a preset orbit semi-major axis formula, the phase modulation period of each satellite and the orbit data; The semi-major axis formula of the preset orbit is: Where a is the semi-major axis of the orbit; T is the orbital period; μ is the Earth's gravitational constant.

9. The constellation configuration maintenance method based on sub-satellite point trajectory drift according to claim 8, characterized in that: After each of the satellites enters the phase modulation orbit from the current orbit, the method further includes: Determine the upper limit of the semi-major axis of the operational orbit as the semi-major axis of the control orbit of each of the satellites; Calculate the second velocity increment and the third velocity increment of each satellite according to the semi-major axis of the phase modulation orbit and the semi-major axis of the control orbit of each satellite; The Hohmann orbit change method is adopted to adjust the running speed of each satellite based on the second speed increment and the third speed increment of each satellite, so that each satellite enters the control orbit from the phase modulation orbit.

10. A constellation configuration maintenance device based on sub-satellite point trajectory drift, characterized in that: include: A first acquisition module, a second acquisition module, an upper limit calculation module, a drift calculation module, a cluster analysis module, a speed increment calculation module and a control module; The first acquisition module is used to acquire orbital data of a first constellation; wherein the orbital data includes operating orbital data and reference orbital data corresponding to a plurality of satellites in the first constellation; The second acquisition module is used to acquire the upper limit of the orbit drift of the first constellation; The upper limit calculation module is used to calculate the upper limit of the semi-major axis of the running orbit according to the orbit data, the upper limit of the orbit drift, the preset orbit drift formula and the preset ascending node longitude formula; The drift calculation module is used to calculate the maneuvering time corresponding to the upper limit of the semi-major axis of the running orbit and the maneuvering drift amount of each satellite based on the preset ascending node longitude formula, the orbit data and the upper limit of the semi-major axis of the running orbit; The cluster analysis module is used to perform cluster analysis on each of the satellites according to the maneuver drift of each of the satellites, obtain a plurality of satellite clusters, and calculate the average maneuver drift corresponding to each satellite cluster; The velocity increment calculation module is used to calculate the first velocity increment of each satellite at the maneuvering moment based on the orbit data of the first constellation and the average maneuvering drift corresponding to each satellite group; The control module is used to adjust the operating state of each satellite based on the first speed increment corresponding to each satellite when the operating time of each satellite reaches the maneuvering time, so that each satellite enters the phase adjustment orbit from the current operating orbit to maintain the constellation configuration of the first constellation.

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