Y-shaped Formation Configuration and Reconfiguration Method of Low Earth Orbit Satellites

Abstract

The present invention provides a Y-shaped formation configuration and reconstruction method for low-earth orbit satellites, comprising the following steps: Step 1: Take one satellite as the main satellite and design its orbital parameters according to mission requirements; Step 2: Take the remaining satellites except the main satellite as auxiliary satellites. The auxiliary satellites take the main satellite as the intermediate reference position of the formation, and design the orbital parameters of each auxiliary satellite according to the orbital parameters of the main satellite, so that each auxiliary satellite forms a Y-shaped formation with the main satellite in the XOY plane of the orbital coordinate system of the main satellite; Step 3: When each auxiliary satellite is operating normally in orbit, perform formation maintenance control on each auxiliary satellite; Step 4: When one of the auxiliary satellites in the formation fails, reconstruct the Y-shaped formation configuration by moving the remaining auxiliary satellites on the same side. Through the Y-star formation configuration of low-earth orbit satellites, the present invention can effectively solve the engineering requirements of passive microwave payloads for the size of large-aperture equivalent antennas, thereby achieving high-resolution ground detection.

Description

Technical Field

[0001] The present invention relates to the technical field of satellite formation flying and formation reconfiguration, and in particular, to a Y-shaped formation configuration and reconfiguration method for low-earth orbit satellites. Background Art

[0002] To achieve high-resolution passive microwave detection, the detection aperture of the payload of a satellite needs to reach the order of 100 meters, which is obviously not achievable in engineering. Therefore, the method of satellite formation flying is adopted to meet the requirement of an equivalent large-scale detection aperture for microwave payloads, so as to achieve the purpose of high-resolution ground detection.

[0003] Since satellite formation is completed by the cooperation of multiple satellites, the failure of a satellite often leads to the inability to maintain the formation shape, and the on-board payload cannot continue to complete the detection task. Therefore, the formation reconfiguration method is one of the effective ways to solve the problem of partial satellite failure in the formation. It can not only ensure that the satellite formation and payload detection are not affected, but also improve the reliability of satellite formation flying and task completion, and extend the mission execution period.

[0004] The patent document with the publication number CN113468671A discloses a method, system, device and storage medium for designing a regular triangular formation of 3N space satellites, including the following steps: obtaining the number N of single-side satellites in the regular triangular formation and the formation scale d; calculating and determining the flying-around radii and phase angles of all satellites on the triangular formation configuration according to the geometric relation formula; calculating the initial relative positions and velocities of all satellites by using the first-order condition of the flying-around circle and the period matching condition; calculating the orbital parameters of the 3N satellites according to the central orbital parameters of the formation to complete the configuration initialization deployment task. The patent document with the publication number CN109885087B discloses a method for close formation of two micro-nano satellites, which is used to realize the close formation of two micro-nano satellites, including: establishing a coordinate system for the two-satellite formation, establishing a relative motion equation, initializing the formation configuration of the two satellites, carrying out inter-satellite communication and maintaining the formation configuration of the two satellites. The patent document with the publication number CN110244767B discloses an optimization of formation configuration reconfiguration using the finite element method. The formation reconfiguration time is divided into a series of equally spaced intervals by using the finite element method, and the problem of solving the reconfiguration path is transformed into an optimization problem. During the optimization process, according to the numerical iteration method, the optimal orbital control quantity and the corresponding reconfiguration orbital state quantity in each time sub-interval are determined, and considering the minimum total control fuel consumption criterion, control timing and safety factors such as avoiding collisions between spacecraft, the control acceleration applied to the spacecraft by the small-thrust engine in the optional sub-interval and the reconfiguration orbital state quantity are iteratively corrected again. However, the above patent documents are all about the simple formation configurations of a small number of satellites or formation flying in the same orbital plane, which are different from the technical solutions of the present invention.

[0005] The patent document with the publication number CN103676955B discloses a satellite autonomous control system for realizing distributed formation flight. The autonomous control system consists of six channels and is embedded in the satellite controller. Under the existing satellite control system, it can generate orbit control commands in advance to provide preparation time for engine catalytic bed heating, attitude maneuvering, ground verification, etc. However, this patent document still focuses on the simple formation configurations of a small number of satellites or formation flight within the same orbital plane, which is different from the technical solution of the present invention. Summary of the Invention

[0006] Aiming at the deficiencies in the prior art, the object of the present invention is to provide a Y-shaped formation configuration and reconstruction method for low-earth orbit satellites.

[0007] A Y-shaped formation configuration and reconstruction method for low-earth orbit satellites according to the present invention includes the following steps:

[0008] Step 1: Design the orbital parameters of a satellite as the main satellite according to mission requirements.

[0009] Step 2: Take the satellites other than the main satellite as auxiliary satellites. The auxiliary satellites use the main satellite as the intermediate reference position of the formation. Design the orbital parameters of each auxiliary satellite according to the orbital parameters of the main satellite, so that each auxiliary satellite and the main satellite form a Y-shaped formation within the XOY plane of the orbital coordinate system of the main satellite.

[0010] Step 3: When each auxiliary satellite is operating normally in orbit, perform formation maintenance control on each auxiliary satellite.

[0011] Step 4: When one of the auxiliary satellites in the formation fails, reconstruct the Y-shaped formation configuration by the movement of the remaining auxiliary satellites on the same side.

[0012] Preferably, in the above Step 1, the orbital parameters of the main satellite include the following parameters: semi-major axis a0, eccentricity e0, orbital inclination i0, right ascension of the ascending node Ω0, argument of perigee ω0, and true anomaly f0.

[0013] Among them, the unit of the semi-major axis a0 is kilometers, the eccentricity e0 has no unit, and the units of the orbital inclination i0, right ascension of the ascending node Ω0, argument of perigee ω0, and true anomaly f0 are all degrees.

[0014] Preferably, in the above Step 2, the satellite orbital coordinate system is defined as follows:

[0015] The Z-axis points from the satellite's center of mass O to the Earth's center of mass O e , the X-axis is along the satellite's flight direction and perpendicular to the Z-axis, the Y-axis forms a right-hand rule with the X-axis and Z-axis, and the main satellite and each auxiliary satellite are all in the XOY plane.

[0016] Preferably, in the step 2, the included angle between adjacent sides of the Y-shaped formation is 120°, one of the sides is along the satellite flight direction, and the included angles between the other two sides and the flight direction are both 60°;

[0017] Each side includes three auxiliary satellites arranged in a straight line. The auxiliary satellites on each side are equidistant from each other, and the distance is L.

[0018] Preferably, in the step 2, for the three satellites on one side along the flight direction, their orbital parameters are designed as follows:

[0019] The semi-major axes of the three satellites are a1, a2, and a3 respectively, and the semi-major axes a1 = a2 = a3 = a0; the eccentricities of the three satellites are e1, e2, and e3 respectively, and the eccentricities e1 = e2 = e3 = e0; the orbital inclinations of the three satellites are i1, i2, and i3 respectively, and the orbital inclinations i1 = i2 = i3 = i0; the right ascensions of the ascending nodes of the three satellites are Ω1, Ω2, and Ω3 respectively, and the right ascensions of the ascending nodes Ω1 = Ω2 = Ω3 = Ω0; the arguments of perigee of the three satellites are ω1, ω2, and ω3 respectively, and the arguments of perigee ω1 = ω2 = ω3 = ω0;

[0020] The true anomalies of the three satellites are f1, f2, and f3 respectively, and the true anomalies f1 = f0 - (180*L) / (πR), f2 = f0 - (180*2L) / (πR), f3 = f0 - (180*3L) / (πR), where R is the average radius of the Earth.

[0021] Preferably, in the step 2, for the six satellites on the two sides with an included angle of 60° with the flight direction, their orbital parameters are designed as follows:

[0022] The semi-major axes of the three satellites on one side are a4, a5, and a6 respectively, and the semi-major axes of the three satellites on the other side are a7, a8, and a9 respectively. The semi-major axes a4 = a5 = a6 = a7 = a8 = a9 = a0;

[0023] The eccentricities of the three satellites on one side are e4, e5, and e6 respectively, and the eccentricities of the three satellites on the other side are e7, e8, and e9 respectively. The eccentricities e4 = e5 = e6 = e7 = e8 = e9 = e0;

[0024] The orbital inclinations of the three satellites on one side are i4, i5, and i6 respectively; the orbital inclinations of the three satellites on the other side are i7, i8, and i9 respectively. The orbital inclinations i4 = i7 = i0 ± (180*cos30°*L) / (πR), i5 = i8 = i0 ± (180*cos30°*2L) / (πR), i6 = i9 = i0 ± (180*cos30°*3L) / (πR);

[0025] The right ascensions of the ascending nodes of the three satellites on one side are Ω4, Ω5, and Ω6 respectively, and the right ascensions of the ascending nodes of the three satellites on the other side are Ω7, Ω8, and Ω9 respectively. The right ascensions of the ascending nodes are Ω4 = Ω5 = Ω6 = Ω7 = Ω8 = Ω9 = Ω0;

[0026] The arguments of perigee of the three satellites on one side are ω4, ω5, and ω6 respectively, and the arguments of perigee of the three satellites on the other side are ω7, ω8, and ω9 respectively. The arguments of perigee are ω4 = ω5 = ω6 = ω7 = ω8 = ω9 = ω0;

[0027] The true anomalies of the three satellites on one side are f4, f5, and f6 respectively, and the true anomalies of the three satellites on the other side are f7, f8, and f9 respectively. The true anomalies are f4 = f7 = f0 + (0.5 * 180 * L) / (πR), f5 = f8 = f0 + (0.5 * 180 * 2L) / (πR), f6 = f9 = f0 + (0.5 * 180 * 3L) / (πR), where R is the average radius of the Earth.

[0028] Preferably, in step 3, the on - satellite installed orbit control thrusters are used to timely adjust the respective orbit parameters of each secondary satellite, so that each orbit parameter is maintained within the allowable error threshold range centered on the design value.

[0029] Preferably, in step 4, if one satellite on one side along the flight direction fails, the changed orbit parameters of the remaining two secondary satellites are as follows: the true anomalies are f0 - (180 * 1.5L) / (πR), f0 - (180 * 3L) / (πR), where R is the average radius of the Earth.

[0030] Preferably, in step 4, if one satellite on one side with an included angle of 60° with the flight direction fails, the changed orbit parameters of the remaining two secondary satellites are as follows:

[0031] The orbital inclinations are i0 - (180 * cos30° * 1.5L) / (πR), i0 - (180 * cos30° * 3L) / (πR) or i = i0 + (180 * cos30° * 1.5L) / (πR), i0 + (180 * cos30° * 3L) / (πR);

[0032] The true anomalies are f0 + (0.5 * 180 * 1.5L) / (πR), f0 + (0.5 * 180 * 3L) / (πR), where R is the average radius of the Earth.

[0033] Compared with the prior art, the present invention has the following beneficial effects:

[0034] 1. Through the Y-shaped formation configuration of low-earth orbit satellites, the present invention can effectively meet the engineering requirements of passive microwave payloads for the equivalent antenna size of large apertures, thereby achieving high-resolution ground detection.

[0035] 2. Through the reconstruction method of the Y-shaped formation configuration, the present invention can ensure that the satellite formation and payload detection are not affected in the event of the failure of individual satellites, thereby improving the reliability of satellite formation flight and mission completion and extending the mission execution period.

[0036] 3. The novel formation configuration and reconstruction method proposed in the present invention can not only meet the requirements of microwave payload missions, but also adapt to the normal execution of missions in the event of the failure of some satellites, improve the reliability of satellite formation flight and mission completion, and extend the mission execution period. BRIEF DESCRIPTION OF THE DRAWINGS

[0037] Other features, objects, and advantages of the present invention will become more apparent by reading the following detailed description of non-limiting embodiments with reference to the accompanying drawings:

[0038] Figure 1 is a flowchart of the steps of the Y-shaped formation configuration and reconstruction method of low-earth orbit satellites of the present invention;

[0039] Figure 2 is a schematic diagram of the Y-shaped formation configuration of low-earth orbit satellites of the present invention;

[0040] Figure 3 is a schematic diagram of the Y-shaped formation reconstruction method of low-earth orbit satellites of the present invention. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0041] The present invention will be described in detail below with reference to specific embodiments. The following embodiments will help those skilled in the art to further understand the present invention, but do not limit the present invention in any form. It should be noted that those of ordinary skill in the art can make several changes and improvements without departing from the concept of the present invention. These all belong to the protection scope of the present invention.

[0042] Example 1:

[0043] As Figures 1 to 3 shown, this embodiment provides a Y-shaped formation configuration and reconstruction method for low-earth orbit satellites, including the following steps:

[0044] Step 1: Design the orbital parameters of a satellite as the main satellite according to mission requirements; the orbital parameters of the main satellite include the following parameters: semi-major axis a0, eccentricity e0, orbital inclination i0, right ascension of the ascending node Ω0, argument of perigee ω0, and true anomaly f0.

[0045] Among them, the unit of the semi-major axis \(a_0\) is kilometers, the eccentricity \(e_0\) has no unit, and the units of the orbital inclination \(i_0\), the right ascension of the ascending node \(\Omega_0\), the argument of perigee \(\omega_0\), and the true anomaly \(f_0\) are all degrees.

[0046] Step 2: Take the satellites other than the main satellite as auxiliary satellites. The auxiliary satellites use the main satellite as the middle reference position of the formation. Design the orbital parameters of each auxiliary satellite according to the orbital parameters of the main satellite, so that each auxiliary satellite forms a Y-shaped formation with the main satellite in the XOY plane of the orbital coordinate system of the main satellite. The satellite orbital coordinate system is defined as follows:

[0047] The Z-axis points from the satellite's center of mass \(O\) to the Earth's center of mass \(O\) e , the X-axis is along the satellite's flight direction and perpendicular to the Z-axis, the Y-axis forms a right-hand rule with the X-axis and the Z-axis, and the main satellite and each auxiliary satellite are all in the XOY plane;

[0048] The included angle between adjacent sides of the Y-shaped formation is \(120^{\circ}\), one of the sides is along the satellite's flight direction, and the included angles between the other two sides and the flight direction are both \(60^{\circ}\);

[0049] Each side includes three auxiliary satellites arranged in a straight line, and the distances between the auxiliary satellites on each side are equal, with a distance of \(L\);

[0050] For the three satellites on one side along the flight direction, their orbital parameters are designed as follows:

[0051] The semi-major axes of the three satellites are \(a_1\), \(a_2\), and \(a_3\) respectively, and the semi-major axes \(a_1 = a_2 = a_3 = a_0\); the eccentricities of the three satellites are \(e_1\), \(e_2\), and \(e_3\) respectively, and the eccentricities \(e_1 = e_2 = e_3 = e_0\); the orbital inclinations of the three satellites are \(i_1\), \(i_2\), and \(i_3\) respectively, and the orbital inclinations \(i_1 = i_2 = i_3 = i_0\); the right ascensions of the ascending nodes of the three satellites are \(\Omega_1\), \(\Omega_2\), and \(\Omega_3\) respectively, and the right ascensions of the ascending nodes \(\Omega_1 = \Omega_2 = \Omega_3 = \Omega_0\); the arguments of perigee of the three satellites are \(\omega_1\), \(\omega_2\), and \(\omega_3\) respectively, and the arguments of perigee \(\omega_1 = \omega_2 = \omega_3 = \omega_0\);

[0052] The true anomalies of the three satellites are \(f_1\), \(f_2\), and \(f_3\) respectively, and the true anomalies \(f_1 = f_0-(180*L) / (πR)\), \(f_2 = f_0-(180*2L) / (πR)\), \(f_3 = f_0-(180*3L) / (πR)\), where \(R\) is the average radius of the Earth;

[0053] For the six satellites on the two sides with an included angle of \(60^{\circ}\) with the flight direction, their orbital parameters are designed as follows:

[0054] The semi-major axes of the three satellites on one side are \(a_4\), \(a_5\), and \(a_6\) respectively, and the semi-major axes of the three satellites on the other side are \(a_7\), \(a_8\), and \(a_9\) respectively. The semi-major axes \(a_4 = a_5 = a_6 = a_7 = a_8 = a_9 = a_0\);

[0055] The eccentricities of three satellites on one side are e4, e5, and e6 respectively, and the eccentricities of three satellites on the other side are e7, e8, and e9 respectively. The eccentricity e4 = e5 = e6 = e7 = e8 = e9 = e0;

[0056] The orbital inclinations of three satellites on one side are i4, i5, and i6 respectively; the orbital inclinations of three satellites on the other side are i7, i8, and i9 respectively. The orbital inclination i4 = i7 = i0 ± (180 * cos30° * L) / (πR), i5 = i8 = i0 ± (180 * cos30° * 2L) / (πR), i6 = i9 = i0 ± (180 * cos30° * 3L) / (πR);

[0057] The right ascensions of the ascending nodes of three satellites on one side are Ω4, Ω5, and Ω6 respectively, and the right ascensions of the ascending nodes of three satellites on the other side are Ω7, Ω8, and Ω9 respectively. The right ascension of the ascending node Ω4 = Ω5 = Ω6 = Ω7 = Ω8 = Ω9 = Ω0;

[0058] The arguments of perigee of three satellites on one side are ω4, ω5, and ω6 respectively, and the arguments of perigee of three satellites on the other side are ω7, ω8, and ω9 respectively. The argument of perigee ω4 = ω5 = ω6 = ω7 = ω8 = ω9 = ω0;

[0059] The true anomalies of three satellites on one side are f4, f5, and f6 respectively, and the true anomalies of three satellites on the other side are f7, f8, and f9 respectively. The true anomalies are f4 = f7 = f0 + (0.5 * 180 * L) / (πR), f5 = f8 = f0 + (0.5 * 180 * 2L) / (πR), f6 = f9 = f0 + (0.5 * 180 * 3L) / (πR), where R is the average radius of the Earth.

[0060] Step 3: When each auxiliary satellite is operating normally in orbit, perform formation maintenance control on each auxiliary satellite; timely adjust each orbital parameter of each auxiliary satellite through the on - satellite installed orbit control thrusters to keep each orbital parameter within the allowable error threshold range centered on the design value.

[0061] Step 4: When one auxiliary satellite in the formation fails, reconstruct the Y - shaped formation configuration by the movement of the remaining auxiliary satellites on the same side; if one satellite on one side along the flight direction fails, the orbital parameters changed by the remaining two auxiliary satellites are as follows: the true anomalies are f0 - (180 * 1.5L) / (πR), f0 - (180 * 3L) / (πR), where R is the average radius of the Earth;

[0062] If one satellite on one side with an included angle of 60° with the flight direction fails, the orbital parameters changed by the remaining two auxiliary satellites are as follows:

[0063] The orbital inclinations are respectively \(i_0 - \frac{(180\times\cos30^{\circ}\times1.5L)}{\pi R}\), \(i_0 - \frac{(180\times\cos30^{\circ}\times3L)}{\pi R}\) or \(i = i_0 + \frac{(180\times\cos30^{\circ}\times1.5L)}{\pi R}\), \(i_0 + \frac{(180\times\cos30^{\circ}\times3L)}{\pi R}\);

[0064] The true anomalies are respectively \(f_0 + \frac{(0.5\times180\times1.5L)}{\pi R}\), \(f_0 + \frac{(0.5\times180\times3L)}{\pi R}\), where \(R\) is the average radius of the Earth.

[0065] Figure 1 It is the flowchart of the steps of the Y-shaped formation configuration and reconstruction method for the low-Earth orbit satellite of the present invention;

[0066] Figure 2 It is a schematic diagram of the Y-shaped formation configuration of a low-Earth orbit satellite in this embodiment. As Figure 2 shown, the total number of satellites is 10. There is 1 main satellite in the middle, and the remaining 9 are auxiliary satellites. The auxiliary satellites take the main satellite as the reference position of the formation and are arranged in a Y shape in the XOY plane of the orbital coordinate system. Each side of the Y-shaped formation includes 3 auxiliary satellites arranged in a straight line, the inter-satellite distance is equal, the included angle between adjacent sides is 120°, and one of the sides is along the flight direction.

[0067] Figure 3 It is a schematic diagram of the Y-shaped formation reconstruction method of a low-Earth orbit satellite in this embodiment. As Figure 3 shown, assume that the auxiliary satellite at the farthest position in the lower right corner fails, so as to carry out the reconstruction of the satellite formation configuration, and make the two auxiliary satellites on the same side perform corresponding position supplementary movements. The auxiliary satellite 2 moves to the position of the original farthest failed satellite, and the auxiliary satellite 1 moves to the middle position between the farthest failed satellite and the middle main satellite.

[0068] Example 2:

[0069] Those skilled in the art can understand this embodiment as a more specific description of Embodiment 1.

[0070] As Figures 1 to 3 shown, this embodiment provides a Y-shaped formation configuration and reconstruction method for a low-Earth orbit satellite, including the following steps:

[0071] Step 1: Design the orbital parameters of the main satellite as follows: semi-major axis \(a_0 = 7171.23\) km, eccentricity \(e_0 = 0\), orbital inclination \(i_0 = 60^{\circ}\), right ascension of the ascending node \(\Omega_0 = 50^{\circ}\), argument of perigee \(\omega_0 = 0^{\circ}\), and the true anomaly \(f_0\) changes with time.

[0072] Step 2: Take the satellites other than the main satellite as secondary satellites. The secondary satellites use the main satellite as the intermediate reference position of the formation. Design the orbital parameters of each secondary satellite according to the orbital parameters of the main satellite, so that each secondary satellite and the main satellite form a Y-shaped formation in the XOY plane of the orbital coordinate system of the main satellite.

[0073] The included angle between adjacent sides of the Y-shaped formation is 120°. One of the sides is along the satellite flight direction, and the included angles between the other two sides and the flight direction are both 60°. Each side includes three secondary satellites arranged in a straight line. The distance between the secondary satellites on each side is equal. That is, if the length of each side is designed to be 90m, then the inter-satellite distance on each side is 30m.

[0074] For the 3 satellites on one side along the flight direction, their orbital parameters are designed as follows: semi-major axis a = 7171.23 km, eccentricity e = 0, orbital inclination i = 60°, right ascension of the ascending node Ω = 50°, argument of perigee ω = 0°, true anomaly are f = f0 - 0.00027°, f0 - 0.00054°, f0 - 0.00081° respectively, where R is the average radius of the Earth, 6371.23 km.

[0075] For the 6 satellites on the 2 sides with an included angle of 60° with the flight direction, their orbital parameters are designed as follows: semi-major axis a = 7171.23 km, eccentricity e = 0, orbital inclinations are i = 59.99977° and 60.00023°, 59.99953° and 60.00047°, 59.99930° and 60.00070° respectively, right ascension of the ascending node Ω = 50°, argument of perigee ω = 0°, true anomaly are f = f0 + 0.000135°, f0 + 0.000270°, f0 + 0.000405° respectively, where R is the average radius of the Earth, 6371.23 km.

[0076] Step 3: When each secondary satellite is operating normally in orbit, formation maintenance control needs to be carried out on each secondary satellite; each secondary satellite adjusts each orbital parameter in a timely manner through the on-board orbit control thruster, so that each orbital parameter is maintained within the allowable error threshold of 5m centered on the designed value.

[0077] Step 4: When one of the secondary satellites in the formation fails, the Y-shaped formation configuration is reconstructed by the movement of the remaining secondary satellites on the same side. If one satellite on one side along the flight direction fails, the orbital parameters changed by the remaining 2 secondary satellites are as follows: true anomaly are f = f0 - 0.00040°, f0 - 0.00081° respectively, where R is the average radius of the Earth, 6371.23 km.

[0078] If one satellite on one side with an included angle of 60° with the flight direction fails, the orbital parameters changed by the remaining two auxiliary satellites are as follows: the orbital inclinations are i = 59.99965°, 59.99930° or i = 60.00035°, 60.00070° respectively, and the true anomaly are f = f0 + 0.000202°, f0 + 0.000405° respectively, where R is the average radius of the earth, 6371.23 km.

[0079] The present invention can effectively meet the engineering requirements of passive microwave payloads for the equivalent antenna size of large apertures, improve the reliability of satellite formation flight and mission completion, and extend the mission execution period.

[0080] The specific embodiments of the present invention have been described above. It should be understood that the present invention is not limited to the above specific embodiments, and those skilled in the art can make various changes or modifications within the scope of the claims, which do not affect the essence of the present invention. Without conflict, the embodiments of the present application and the features in the embodiments can be arbitrarily combined with each other.

Claims

1. A Y-shaped formation configuration and reconstruction method for low-earth orbit satellites, characterized in that, It includes the following steps: Step 1: Take a satellite as the main satellite and design its orbital parameters according to mission requirements; Step 2: Take the remaining satellites except the main satellite as auxiliary satellites. The auxiliary satellites use the main satellite as the intermediate reference position of the formation. Design the orbital parameters of each auxiliary satellite according to the orbital parameters of the main satellite, so that each auxiliary satellite forms a Y-shaped formation with the main satellite in the XOY plane of the orbital coordinate system of the main satellite; Step 3: When each auxiliary satellite operates normally in orbit, perform formation maintenance control on each auxiliary satellite; Step 4: When one of the auxiliary satellites in the formation fails, reconstruct the Y-shaped formation configuration by the movement of the remaining auxiliary satellites on the same side; In the above Step 4, if one satellite on one side along the flight direction fails, the orbital parameters changed by the remaining two auxiliary satellites are as follows: the true anomaly is respectively f0-(180*1.5L) / (πR), f0-(180*3L) / (πR), where R is the average radius of the earth; In the above Step 4, if one satellite on one side with an angle of 60° to the flight direction fails, the orbital parameters changed by the remaining two auxiliary satellites are as follows: The orbital inclinations are respectively i0-(180*cos30°*1.5L) / (πR), i0-(180*cos30°*3L) / (πR) or i=i0+(180*cos30°*1.5L) / (πR), i0+(180*cos30°*3L) / (πR); The true anomalies are respectively f0+(0.5*180*1.5L) / (πR), f0+(0.5*180*3L) / (πR), where R is the average radius of the earth.

2. The Y-shaped formation configuration and reconstruction method for low-earth orbit satellites according to claim 1, characterized in that, In the above Step 1, the orbital parameters of the main satellite include the following parameters: semi-major axis a0, eccentricity e0, orbital inclination i0, right ascension of the ascending node Ω0, argument of perigee ω0, true anomaly f0; Among them, the unit of the semi-major axis a0 is kilometer, the eccentricity e0 has no unit, and the units of the orbital inclination i0, right ascension of the ascending node Ω0, argument of perigee ω0, and true anomaly f0 are all degrees.

3. The Y-shaped formation configuration and reconstruction method for low-earth orbit satellites according to claim 1, characterized in that, In the above Step 2, the satellite orbital coordinate system is defined as follows: The Z-axis points from the satellite's centroid O to the Earth's centroid O e , the X-axis is along the satellite's flight direction and perpendicular to the Z-axis, the Y-axis forms a right-hand rule with the X-axis and Z-axis, and the main satellite and each auxiliary satellite are all in the XOY plane.

4. The Y-shaped formation configuration and reconstruction method for low-earth orbit satellites according to claim 1, characterized in that, In the above Step 2, the included angle between adjacent sides of the Y-shaped formation is 120°, one of the sides is along the satellite flight direction, and the included angles between the other two sides and the flight direction are both 60°; Each side includes three auxiliary satellites arranged in a straight line, and the auxiliary satellites on each side are equidistant from each other, and the distance is L.

5. The Y-shaped formation configuration and reconstruction method for low-earth orbit satellites according to claim 1, characterized in that, In the above Step 2, for the three satellites on one side along the flight direction, their orbital parameters are designed as follows: The semi-major axes of the three satellites are respectively a1, a2, a3, and the semi-major axis a1 = a2 = a3 = a0; the eccentricities of the three satellites are respectively e1, e2, e3, and the eccentricity e1 = e2 = e3 = e0; the orbital inclinations of the three satellites are respectively i1, i2, i3, and the orbital inclination i1 = i2 = i3 = i0; the right ascensions of the ascending nodes of the three satellites are respectively Ω1, Ω2, Ω3, and the right ascension of the ascending node Ω1 = Ω2 = Ω3 = Ω0; the arguments of perigee of the three satellites are respectively ω1, ω2, ω3, and the argument of perigee ω1 = ω2 = ω3 = ω0; The true anomalies of the three satellites are f1, f2, and f3 respectively, where the true anomaly f1 = f0 - (180*L) / (πR), f2 = f0 - (180*2L) / (πR), f3 = f0 - (180*3L) / (πR), and R is the average radius of the Earth.

6. The Y-shaped formation configuration and reconstruction method for low-earth orbit satellites according to claim 1, characterized in that, In step 2, for the six satellites on the two sides with an included angle of 60° with the flight direction, their orbital parameters are designed as follows: The semi-major axes of the three satellites on one side are a4, a5, and a6 respectively, and the semi-major axes of the three satellites on the other side are a7, a8, and a9 respectively. The semi-major axes a4 = a5 = a6 = a7 = a8 = a9 = a0; The eccentricities of the three satellites on one side are e4, e5, and e6 respectively, and the eccentricities of the three satellites on the other side are e7, e8, and e9 respectively. The eccentricities e4 = e5 = e6 = e7 = e8 = e9 = e0; The orbital inclinations of the three satellites on one side are i4, i5, and i6 respectively, and the orbital inclinations of the three satellites on the other side are i7, i8, and i9 respectively. The orbital inclinations i4 = i7 = i0 ± (180*cos30°*L) / (πR), i5 = i8 = i0 ± (180*cos30°*2L) / (πR), i6 = i9 = i0 ± (180*cos30°*3L) / (πR); The right ascensions of the ascending nodes of the three satellites on one side are Ω4, Ω5, and Ω6 respectively, and the right ascensions of the ascending nodes of the three satellites on the other side are Ω7, Ω8, and Ω9 respectively. The right ascensions of the ascending nodes Ω4 = Ω5 = Ω6 = Ω7 = Ω8 = Ω9 = Ω0; The arguments of perigee of the three satellites on one side are ω4, ω5, and ω6 respectively, and the arguments of perigee of the three satellites on the other side are ω7, ω8, and ω9 respectively. The arguments of perigee ω4 = ω5 = ω6 = ω7 = ω8 = ω9 = ω0; The true anomalies of the three satellites on one side are f4, f5, and f6 respectively, and the true anomalies of the three satellites on the other side are f7, f8, and f9 respectively. The true anomalies are f4 = f7 = f0 + (0.5*180*L) / (πR), f5 = f8 = f0 + (0.5*180*2L) / (πR), f6 = f9 = f0 + (0.5*180*3L) / (πR), where R is the average radius of the Earth.

7. The Y-shaped formation configuration and reconstruction method for low-earth orbit satellites according to claim 1, characterized in that, In step 3, the on-board orbit control thrusters are used to adjust the various orbital parameters of each auxiliary satellite in a timely manner, so that each orbital parameter is maintained within the allowable error threshold range centered on the designed value.

Images