Elliptical reference orbit plane circular satellite formation configuration design method

Abstract

The invention discloses an elliptical reference orbit plane circle satellite formation configuration design method. The method comprises the steps of S1, selecting the orbit number of plane circle formation reference satellites; s2, selecting an elliptical orbit reference satellite RIC coordinate system; s3, establishing a relative kinetic equation of the elliptical reference orbit satellite formation flight; s4, establishing a periodic solution of the linear relative kinetic equation of the elliptic reference orbit; s5, giving out constraint conditions of the elliptical orbit plane circle formation configuration; s6, designing plane circle formation parameters; s7, calculating a relative position speed parameter of the accompanying satellite at the initial moment in the RIC coordinate system of the reference satellite; s8, calculating position and speed parameters of the accompanying satellite in the geocentric inertial coordinate system at the initial moment; and S9, calculating the orbital elements accompanying the initial moment of the satellite, and completing the design of the plane circle formation satellite configuration.

Description

Technical Field

[0001] The invention belongs to the technical field of satellite technology integrated application system, and in particular relates to a method for designing a configuration of a plane circular satellite formation in an elliptical reference orbit. Background Art

[0002] Satellite formation refers to a number of spacecraft that use natural orbital characteristics or slight control to fly in a certain configuration in space and work together to form a virtual instrument to perform specific functions. It is a new type of satellite comprehensive application technology, and the corresponding satellite groups are called distributed satellites or formation satellites.

[0003] The proposal of satellite formations has opened up a brand new space for the application of satellites, especially small satellites. It breaks the concept of traditional spacecraft combination by hardware, proposes an innovative idea of ​​combining spacecraft according to mission requirements, expands the application scope of small satellites, and can complete many tasks that were impossible for spacecraft in the past, bringing revolutionary changes to the future aerospace field.

[0004] The classic satellite formation configuration is designed using the CW equation. Common configurations include serial formation, space circular formation, and horizontal circular formation. However, the CW equation is only applicable when the reference satellite orbit is a circular orbit.

[0005] Based on the CW equation, a 2:1 elliptical formation configuration can be designed in the reference satellite orbit plane, but a planar circular formation configuration cannot be designed. Summary of the invention

[0006] The present invention aims to solve the technical problems existing in the background technology and to provide a method for designing a plane circular satellite formation configuration in an elliptical reference orbit. The method establishes the relative dynamic equations of the elliptical reference orbit satellite formation flight and the periodic solution of the linearized relative dynamic equations of the elliptical reference orbit, combines the plane circular formation configuration constraints, takes the plane circular formation radius as the input parameter, and completes the design of the plane circular formation configuration parameters; converts the plane circular configuration parameters into the relative position and velocity of the companion satellite relative to the reference satellite at the initial moment in the RIC coordinate system of the elliptical orbit reference satellite, and calculates the position, velocity and orbital elements of the companion satellite in the geocentric inertial coordinate system to complete the construction of the plane circular formation configuration.

[0007] In order to solve the technical problem, the technical solution of the present invention is:

[0008] A method for designing a circular satellite formation configuration in an elliptical reference orbit plane, the method comprising:

[0009] S1: Select the orbital elements of the reference satellite of the plane circle formation;

[0010] S2: Select the elliptical orbit reference satellite RIC coordinate system;

[0011] S3: Based on steps S1 and S2, relative dynamic equations of satellite formation flying in an elliptical reference orbit are established;

[0012] S4: Based on steps S1 and S2, a periodic solution of the linearized relative dynamic equation of the elliptical reference orbit is established;

[0013] S5: Constraints for constructing the elliptical orbital planar circle formation configuration;

[0014] S6: Design plane circle formation parameters;

[0015] S7: Based on the steps 3 and 4 and the plane circle formation parameters, the relative position velocity parameters of the companion satellite in the reference satellite RIC coordinate system at the initial moment are calculated;

[0016] S8: Based on steps S1 and S7, calculate the position and velocity parameters of the companion satellite in the geocentric inertial coordinate system at the initial moment;

[0017] S9: Based on the position and velocity parameters obtained in step S8, the orbital elements of the companion satellite at the initial moment are calculated to complete the satellite configuration design of the plane circular formation.

[0018] Further, the step S1 refers to the satellite orbit element number selection method, including:

[0019] S101: reference satellite orbit semi-major axis a, value range [6678km, 43000km], including low Earth orbit, medium Earth orbit and high Earth orbit;

[0020] S102: reference satellite orbit eccentricity e, value range [0.01, 0.99];

[0021] S103: Reference satellite orbit inclination angle i, value range [0°, 180°];

[0022] S104: right ascension Ω of the reference satellite ascending node, with a value range of [0°, 360°];

[0023] S105: reference satellite perigee angle ω, value range [0°, 360°];

[0024] S106: reference satellite true anomaly f, value range [0°, 360°];

[0025] S107: reference satellite mean anomaly M, value range [0°, 360°];

[0026] S108: reference satellite anomaly angle E, value range [0°, 360°];

[0027] S109: Any combination of S101-S108.

[0028] Further, the step S2 selects the elliptical orbit reference satellite RIC coordinate system, including:

[0029] S201: origin of coordinate system, reference satellite mass center;

[0030] S202: In the x-axis direction of the coordinate system, the center of the earth points to the center of mass of the reference satellite;

[0031] S203: coordinate system z-axis direction, reference satellite orbit normal direction;

[0032] S204: The y-axis direction of the coordinate system is determined by the right-hand rule.

[0033] Further, the step S3 comprises:

[0034] The mathematical model of the relative motion of satellite formations in an elliptical reference orbit is:

[0035]

[0036] where μ is the Earth's gravitational constant, a is the semi-major axis of the reference satellite orbit selected by S1, e is the eccentricity of the reference satellite orbit selected by S1, f is the true anomaly angle of the reference satellite orbit selected by S1, [xyz] T is the projection of the position vector of the companion satellite in the formation relative to the reference satellite on the x, y, and z axes of the RIC coordinate system, are the second-order derivatives of x, y, and z with respect to time.

[0037] Further, the step S4 comprises:

[0038] The mathematical model of the periodic solution of the linearized relative dynamics equation of the elliptical reference orbit is:

[0039]

[0040]

[0041]

[0042]

[0043] Where, e is the eccentricity of the reference satellite orbit selected by S1, f is the true anomaly angle of the reference satellite orbit selected by S1, [xy z] T is the projection of the position vector of the companion satellite in the formation relative to the reference satellite on the x, y, and z axes of the RIC coordinate system, [x′y′z′] T is the first-order derivative of x, y, and z relative to the true perigee angle, and the subscript 0 is the initial moment.

[0044] Furthermore, the step S5 constructs the constraint conditions for the elliptical orbit plane circle formation configuration, including:

[0045] The constraints of the elliptical orbit plane circle formation configuration are:

[0046] x 2 +y 2 =R 2

[0047] Where R is the radius of the plane circle formation, and x and y are the projections of the position vector of the companion satellite in the formation relative to the reference satellite on the x and y axes of the RIC coordinate system.

[0048] Further, the step S6 designs the plane circle formation parameters, including:

[0049] S601: Select the radius R of the plane circle formation, where R is in the range of [0.1km, 200km];

[0050] S602: The calculation formula of the position vector of the center of the plane circle formation relative to the reference satellite in the reference satellite RIC coordinate system is:

[0051]

[0052] Where R is the radius of the plane circle formation selected by S601, e is the eccentricity of the reference satellite orbit selected by S1, + represents the positive direction of the y-axis of the reference satellite RIC coordinate system, and - represents the negative direction of the y-axis of the reference satellite RIC coordinate system.

[0053] Further, the step S7 comprises:

[0054] S701: In the reference satellite RIC coordinate system, the calculation formula for the initial position of the companion satellite in the x direction is:

[0055] x 0 =Rsinf 0

[0056] Where R is the radius of the plane circle formation selected by S601, f 0 The initial true anomaly of the reference satellite orbit selected for S1;

[0057] S702: In the reference satellite RIC coordinate system, the calculation formula for the initial position of the companion satellite in the y direction is:

[0058]

[0059] Where R is the radius of the plane circle formation selected by S601, f 0is the initial true anomaly of the reference satellite orbit selected by S1, e is the eccentricity of the reference satellite orbit selected by S1, + represents the positive direction of the y-axis of the reference satellite RIC coordinate system, and - represents the negative direction of the y-axis of the reference satellite RIC coordinate system;

[0060] S703: In the reference satellite RIC coordinate system, the calculation formula for the initial position of the companion satellite in the y direction is:

[0061] z 0 =0

[0062] S704: In the reference satellite RIC coordinate system, the calculation formula for the initial velocity of the companion satellite in the x direction is:

[0063]

[0064] Where R is the radius of the plane circle formation selected by S601, μ is the Earth's gravitational constant, a is the semi-major axis of the reference satellite orbit selected by S1, e is the eccentricity of the reference satellite orbit selected by S1, and f 0 The initial true anomaly of the reference satellite orbit selected for S1;

[0065] S705: In the reference satellite RIC coordinate system, the calculation formula for the initial velocity of the companion satellite in the y direction is:

[0066]

[0067] S706: In the reference satellite RIC coordinate system, the calculation formula for the initial position of the companion satellite in the z direction is:

[0068]

[0069] Further, the step S8 comprises:

[0070] S801: The calculation formula of the direction cosine matrix from the reference satellite geocentric inertial coordinate system to the RIC coordinate system is:

[0071]

[0072] Where r and v are the position vector and velocity vector of the reference satellite's center of mass in the geocentric inertial coordinate system at the initial moment, respectively;

[0073] S802: The calculation formula of the position vector of the companion satellite in the Earth-centered inertial coordinate system is:

[0074] r B =r+R T [x 0 y 0 z 0 ] T

[0075] Where r is the position vector of the reference satellite mass center in the geocentric inertial coordinate system at the initial moment, R T is the transpose of the direction cosine matrix from the reference satellite geocentric inertial coordinate system to the RIC coordinate system in S801, [x 0 y 0 z 0 ] T is the relative position vector of the companion satellite in S7 at the initial time in the reference satellite RIC coordinate system;

[0076] S803: The velocity vector calculation formula in the Earth-centered inertial coordinate system of the companion satellite is:

[0077]

[0078] Where v is the position vector of the reference satellite mass center in the geocentric inertial coordinate system at the initial moment, R T is the transpose of the direction cosine matrix from the reference satellite geocentric inertial coordinate system to the RIC coordinate system in S801, [x 0 y 0 z 0 ] T is the relative position vector of the companion satellite in S7 at the initial time in the reference satellite RIC coordinate system, is the relative velocity vector of the companion satellite in S7 at the initial moment in the reference satellite RIC coordinate system, μ is the earth's gravitational constant, a is the semi-major axis of the reference satellite orbit selected by S1, e is the eccentricity of the reference satellite orbit selected by S1, and f ... 0 The initial true anomaly of the reference satellite orbit selected for S1.

[0079] Further, the step S9 includes:

[0080] S901: The calculation formula of the companion satellite orbital angular momentum is:

[0081] h=r B ×v B

[0082] where r B 、v B are the position vector and velocity vector of the satellite's mass center in the geocentric inertial coordinate system at the initial moment in S8, respectively;

[0083] S902: The calculation formula for the companion satellite orbit eccentricity is:

[0084]

[0085] e=|e|

[0086] where μ is the Earth's gravitational constant, h is the orbital angular momentum of the companion satellite in S901, and r B 、vB are the position vector and velocity vector of the satellite mass center in the geocentric inertial coordinate system at the initial moment in S8, respectively, and |·| represents the vector modulus;

[0087] S903: The calculation formula for the companion satellite orbit inclination is:

[0088]

[0089] where h is the orbital angular momentum of the companion satellite in S901, h z is the component of h along the z axis in the geocentric inertial coordinate system;

[0090] S904: The calculation formula for the right ascension of the companion satellite upgrade point is:

[0091] Ω=arctan(h x ,-h y )

[0092] where h x 、h y are the x-axis and y-axis components of the companion satellite orbital angular momentum h in S901 in the geocentric inertial coordinate system;

[0093] S905: The calculation formula for the semi-major axis of the companion satellite orbit is:

[0094]

[0095] Where μ is the Earth's gravitational constant, h is the modulus of the companion satellite's orbital angular momentum h in S901, and e is the companion satellite's orbital eccentricity in S902;

[0096] S906: The calculation formula for the companion satellite perigee angle is:

[0097] i n =[0 0 1]×h

[0098]

[0099] Where h is the orbital angular momentum of the companion satellite in S901, and e is the eccentricity vector of the companion satellite in S902;

[0100] S906, the calculation formula of the true anomaly angle of the companion satellite is:

[0101]

[0102] where r B is the position vector of the companion satellite mass center in the geocentric inertial coordinate system at the initial moment in S8, e is the companion satellite eccentricity vector in S902, and |·| represents the vector modulus.

[0103] Compared with the prior art, the advantages of the present invention are:

[0104] The present invention proposes a method for designing a plane circle satellite formation configuration in an elliptical reference orbit. The proposed method can be used to design a satellite orbit for plane circle formation flight in an elliptical reference orbit, breaking through the traditional method of designing a formation satellite configuration based on the CW equation, and designing a new formation configuration. The designed plane circle formation configuration can be used in, but not limited to, the fields of space target observation, remote sensing of ground resources, and close inspection of target satellites.

[0105] Accuracy and systematicity: Detailed steps and mathematical models are used to ensure the accuracy of the design and calculations, from selecting orbital parameters to calculating initial conditions and determining the number of orbital elements.

[0106] Wide applicability: This method covers a variety of orbit types, including low Earth orbit, medium orbit and high orbit, and is suitable for satellite formation design with various orbital eccentricities and orbital inclinations.

[0107] Utilization of reference system: By introducing the RIC coordinate system, the description of the relative motion of the companion satellite with respect to it is more intuitive, and the derivation and solution of the dynamic model are simplified.

[0108] Establishment of periodic solutions: Using the periodic solutions of the linearized relative dynamics equations helps to understand and design the periodic motion characteristics of the formation, making it more feasible to maintain the formation under dynamic conditions.

[0109] Design flexibility: By introducing multiple parameter selections and calculation methods in the steps, the method allows customized design of specific formation configurations according to different mission requirements.

[0110] Improved efficiency: Through systematic steps and clear sequences, designers can efficiently complete formation design, reducing trial and error costs and the need for complex calculations.

[0111] In general, the present invention provides an efficient and accurate solution for satellite formation design in elliptical orbits through clear steps and detailed mathematical models, which is particularly suitable for space missions that require precise control and long-term stability. These advantages are of great significance in improving the success rate and service life of spacecraft formation missions. BRIEF DESCRIPTION OF THE DRAWINGS

[0112] Figure 1 , a flow chart of a satellite configuration design for an elliptical orbit plane circular formation according to an embodiment of the present invention;

[0113] Figure 2 , an example diagram of a left plane circle formation configuration according to an embodiment of the present invention;

[0114] Figure 3, an example diagram of a right-side plane circle formation configuration according to an embodiment of the present invention;

[0115] Figure 4 , an example diagram of a double-sided plane circle formation configuration according to an embodiment of the present invention. DETAILED DESCRIPTION

[0116] The specific implementation mode of the present invention is described below in conjunction with embodiments:

[0117] It should be noted that the structures, proportions, sizes, etc. shown in this specification are only used to match the contents disclosed in the specification so that people familiar with this technology can understand and read them, and are not used to limit the conditions under which the present invention can be implemented. Any structural modification, change in proportional relationship or adjustment of size should still fall within the scope of the technical content disclosed in the present invention without affecting the effects and purposes that can be achieved by the present invention.

[0118] At the same time, the terms such as "upper", "lower", "left", "right", "middle" and "one" cited in this specification are only for the convenience of description and are not used to limit the scope of implementation of the present invention. Changes or adjustments to their relative relationships should be regarded as the scope of implementation of the present invention without substantially changing the technical content.

[0119] Embodiment 1:

[0120] The present invention provides a method for designing a formation configuration of a plane circle satellite in an elliptical reference orbit. By establishing the relative dynamic equation of the formation flight of the elliptical reference orbit satellite and the periodic solution of the linearized relative dynamic equation of the elliptical reference orbit, the method breaks through the traditional method of designing the formation satellite configuration based on the CW equation. Combined with the plane circle formation configuration constraint, the plane circle formation radius is used as the input parameter to complete the design of the plane circle formation configuration parameters, convert the plane circle configuration parameters into the relative position and speed of the companion satellite relative to the reference satellite at the initial moment in the RIC coordinate system of the elliptical orbit reference satellite, and calculate the orbital elements of the companion satellite to design a new formation configuration.

[0121] According to this method, a left-side plane circle formation configuration, a right-side plane circle formation configuration, and a double-side plane circle formation configuration can be quickly designed. This configuration is used for, but not limited to, the observation of space targets, remote sensing of ground resources, close inspection of target satellites, and other fields.

[0122] This embodiment takes an elliptical reference orbit formation satellite as an example to design the left plane circle, right plane circle, and double-side plane circle formation configurations. The elliptical reference orbit elements at the initial moment are: orbit semi-major axis 20000km, orbit eccentricity 0.6, orbit inclination 30°, upgrade point right ascension 0°, perigee 0°, true anomaly 0°, and plane circle radius 10km.

[0123] like Figure 1 As shown, one embodiment of the present invention provides a flow chart for designing a satellite configuration of an elliptical orbit plane circular formation, and the specific steps include:

[0124] S1. Select the orbital elements of the reference satellite for the plane circle formation.

[0125] In one embodiment of the present invention, the elements of the elliptical reference orbit at the initial moment are: orbit semi-major axis 20000km, orbit eccentricity 0.6, orbit inclination 30°, upgrade point right ascension 0°, perigee angle 0°, true anomaly angle 0°

[0126] S2. Select the elliptical orbit reference satellite RIC coordinate system.

[0127] In one embodiment of the present invention, the origin of the RIC coordinate system is the reference satellite's center of mass, the coordinate system's x-axis direction is from the center of the earth to the reference satellite's center of mass, the coordinate system's z-axis direction is the reference satellite's orbit normal direction, and the coordinate system's y-axis direction is determined by the right-hand rule.

[0128] S3. Establish the relative dynamic equations of satellite formation flying in elliptical reference orbit.

[0129] In one embodiment of the present invention, the relative kinetic equation is:

[0130]

[0131] S4. Establish the periodic solution of the linearized relative dynamic equations of the elliptical reference orbit.

[0132] In one embodiment of the present invention, the periodic solution of the linearized relative dynamic equation of the elliptical reference orbit is:

[0133]

[0134]

[0135]

[0136]

[0137] S5. Give the constraints on the plane circle formation configuration of the elliptical orbit.

[0138] In one embodiment of the present invention, the constraints of the planar circle formation configuration are:

[0139] x 2 +y 2 =R 2

[0140] S6. Design plane circle formation parameters.

[0141] In one embodiment of the present invention, the radius R of the plane circle formation is 10 km;

[0142] The calculation formula of the position vector of the center of the plane circle formation relative to the reference satellite in the reference satellite RIC coordinate system is:

[0143]

[0144] In one embodiment of the present invention, the center position vector of the plane circle formation is + represents the right-side plane circle formation configuration, - represents the left-side plane circle formation configuration, and ± represents the bilateral plane circle formation configuration.

[0145] S7. Calculate the relative position velocity parameters of the companion satellite in the reference satellite RIC coordinate system at the initial moment.

[0146] S8. Calculate the position and velocity parameters of the companion satellite in the geocentric inertial coordinate system at the initial moment.

[0147] In one embodiment of the present invention, the position speed parameter calculation method;

[0148] The initial position of the companion satellite in the x direction is calculated as follows:

[0149] x 0 =Rsinf 0

[0150] The initial position of the companion satellite in the y direction is calculated as follows:

[0151]

[0152] The initial position of the companion satellite in the y direction is calculated as follows:

[0153] z 0 =0

[0154] The initial velocity of the companion satellite in the x direction is calculated as:

[0155]

[0156] The initial velocity of the companion satellite in the y direction is calculated as:

[0157]

[0158] The initial position of the companion satellite in the z direction is calculated as follows:

[0159]

[0160] S9. Calculate the orbital elements of the companion satellite at the initial moment and complete the satellite configuration design of the plane circular formation.

[0161] In one embodiment of the present invention, the designed left plane circle formation configuration is as follows Figure 2 shown.

[0162] In one embodiment of the present invention, the designed right plane circle formation configuration is as follows Figure 3 shown.

[0163] In one embodiment of the present invention, the designed double-sided plane circle formation configuration is as follows Figure 4 shown.

[0164] Those skilled in the art will appreciate that embodiments of the present invention may be provided as methods, systems, or computer program products. Therefore, the present invention may take the form of a complete hardware embodiment, a complete software embodiment, or an embodiment combining software and hardware. Moreover, the present invention may take the form of a computer program product implemented on one or more computer-usable storage media (including but not limited to disk storage, CD-ROM, optical storage, etc.) containing computer-usable program code.

[0165] The present invention is described with reference to flowcharts and / or block diagrams of methods, devices (systems), and computer program products according to embodiments of the present invention. It should be understood that each process and / or block in the flowchart and / or block diagram, as well as the combination of processes and / or blocks in the flowchart and / or block diagram, can be implemented by computer program instructions. These computer program instructions can be provided to a processor of a general-purpose computer, a special-purpose computer, an embedded processor, or other programmable data processing device to produce a machine, so that the instructions executed by the processor of the computer or other programmable data processing device generate instructions for implementing the processes in the flowchart and / or block diagram. Figure 1 A process or multiple processes and / or boxes Figure 1 A device that provides the functions specified in a block or multiple blocks.

[0166] These computer program instructions may also be stored in a computer-readable memory capable of directing a computer or other programmable data processing device to operate in a specific manner, so that the instructions stored in the computer-readable memory produce an article of manufacture comprising an instruction device, which implements the process Figure 1 A process or multiple processes and / or boxes Figure 1 A function specified in one or more boxes.

[0167] These computer program instructions can also be loaded onto a computer or other programmable data processing device so that a series of operating steps are executed on the computer or other programmable device to produce a computer-implemented process, thereby providing instructions for implementing the process. Figure 1 A process or multiple processes and / or boxes Figure 1 A step that specifies a function in one or more boxes.

[0168] The preferred embodiments of the present invention are described in detail above, but the present invention is not limited to the above embodiments, and various changes can be made within the knowledge scope of ordinary technicians in this field without departing from the purpose of the present invention.

[0169] Many other changes and modifications may be made without departing from the concept and scope of the present invention.It should be understood that the present invention is not limited to the specific embodiments, and the scope of the present invention is defined by the appended claims.

Claims

1. A method for designing a circular satellite formation configuration in an elliptical reference orbit plane, characterized in that: The method comprises: S1: Select the orbital elements of the reference satellite of the plane circle formation; S2: Select the elliptical orbit reference satellite RIC coordinate system; S3: Based on steps S1 and S2, relative dynamic equations of satellite formation flying in an elliptical reference orbit are established; S4: Based on steps S1 and S2, a periodic solution of the linearized relative dynamic equation of the elliptical reference orbit is established; S5: Constraints for constructing the elliptical orbital planar circle formation configuration; S6: Design plane circle formation parameters; S7: Based on the steps 3 and 4 and the plane circle formation parameters, the relative position velocity parameters of the companion satellite in the reference satellite RIC coordinate system at the initial moment are calculated; S8: Based on steps S1 and S7, calculate the position and velocity parameters of the companion satellite in the geocentric inertial coordinate system at the initial moment; S9: Based on the position and velocity parameters obtained in step S8, the orbital elements of the companion satellite at the initial moment are calculated to complete the satellite configuration design of the plane circular formation.

2. The method for designing a circular satellite formation configuration in an elliptical reference orbit plane according to claim 1, characterized in that: The step S1 refers to the satellite orbit element number selection method, including: S101: reference satellite orbit semi-major axis a, value range [6678km, 43000km], including low Earth orbit, medium Earth orbit and high Earth orbit; S102: reference satellite orbit eccentricity e, value range [0.01, 0.99]; S103: Reference satellite orbit inclination angle i, value range [0°, 180°]; S104: right ascension Ω of the reference satellite ascending node, with a value range of [0°, 360°]; S105: reference satellite perigee angle ω, value range [0°, 360°]; S106: reference satellite true anomaly f, value range [0°, 360°]; S107: reference satellite mean anomaly M, value range [0°, 360°]; S108: reference satellite anomaly angle E, value range [0°, 360°]; S109: Any combination of S101-S108.

3. The method for designing a circular satellite formation configuration in an elliptical reference orbit plane according to claim 1, characterized in that: The step S2 selects the elliptical orbit reference satellite RIC coordinate system, including: S201: origin of coordinate system, reference satellite mass center; S202: In the x-axis direction of the coordinate system, the center of the earth points to the center of mass of the reference satellite; S203: coordinate system z-axis direction, reference satellite orbit normal direction; S204: The y-axis direction of the coordinate system is determined by the right-hand rule.

4. The method for designing a circular satellite formation configuration in an elliptical reference orbit plane according to claim 1, characterized in that: The step S3 comprises: The mathematical model of the relative motion of satellite formations in an elliptical reference orbit is: where μ is the Earth's gravitational constant, a is the semi-major axis of the reference satellite orbit selected by S1, e is the eccentricity of the reference satellite orbit selected by S1, f is the true anomaly angle of the reference satellite orbit selected by S1, [x yz] T is the projection of the position vector of the companion satellite in the formation relative to the reference satellite on the x, y, and z axes of the RIC coordinate system, is the second-order derivative of x, y, and z with respect to time.

5. The method for designing a circular satellite formation configuration in an elliptical reference orbit plane according to claim 1, characterized in that: The step S4 comprises: The mathematical model of the periodic solution of the linearized relative dynamics equation of the elliptical reference orbit is: Where, e is the eccentricity of the reference satellite orbit selected by S1, f is the true anomaly angle of the reference satellite orbit selected by S1, [x yz] T is the projection of the position vector of the companion satellite in the formation relative to the reference satellite on the x, y, and z axes of the RIC coordinate system, [x′y′z′] T is the first-order derivative of x, y, and z relative to the true perigee angle, and the subscript 0 is the initial moment.

6. The method for designing a circular satellite formation configuration in an elliptical reference orbit plane according to claim 1, characterized in that: The step S5 constructs the constraint conditions of the elliptical orbit plane circle formation configuration, including: The constraints of the elliptical orbit plane circle formation configuration are: x 2 +y 2 =R 2 Where R is the radius of the plane circle formation, and x and y are the projections of the position vector of the companion satellite in the formation relative to the reference satellite on the x and y axes of the RIC coordinate system.

7. The method for designing a circular satellite formation configuration in an elliptical reference orbit plane according to claim 1, characterized in that: The step S6 designs the plane circle formation parameters, including: S601: Select the radius R of the plane circle formation, where R is in the range of [0.1km, 200km]; S602: The calculation formula of the position vector of the center of the plane circle formation relative to the reference satellite in the reference satellite RIC coordinate system is: Where R is the radius of the plane circle formation selected by S601, e is the eccentricity of the reference satellite orbit selected by S1, + represents the positive direction of the y-axis of the reference satellite RIC coordinate system, and - represents the negative direction of the y-axis of the reference satellite RIC coordinate system.

8. The method for designing a circular satellite formation configuration in an elliptical reference orbit plane according to claim 1, characterized in that: The step S7 comprises: S701: In the reference satellite RIC coordinate system, the calculation formula for the initial position of the companion satellite in the x direction is: x0=Rsinf0 Where R is the radius of the plane circle formation selected by S601, and f0 is the initial true anomaly angle of the reference satellite orbit selected by S1; S702: In the reference satellite RIC coordinate system, the calculation formula for the initial position of the companion satellite in the y direction is: Where R is the radius of the plane circle formation selected by S601, f0 is the initial true anomaly of the reference satellite orbit selected by S1, e is the eccentricity of the reference satellite orbit selected by S1, + represents the positive direction of the y-axis of the reference satellite RIC coordinate system, and - represents the negative direction of the y-axis of the reference satellite RIC coordinate system; S703: In the reference satellite RIC coordinate system, the calculation formula for the initial position of the companion satellite in the y direction is: z0＝0 S704: In the reference satellite RIC coordinate system, the calculation formula for the initial velocity of the companion satellite in the x direction is: Where R is the radius of the plane circle formation selected by S601, where μ is the Earth’s gravitational constant, a is the semi-major axis of the reference satellite orbit selected by S1, e is the eccentricity of the reference satellite orbit selected by S1, and f0 is the initial true anomaly angle of the reference satellite orbit selected by S1; S705: In the reference satellite RIC coordinate system, the calculation formula for the initial velocity of the companion satellite in the y direction is: S706: In the reference satellite RIC coordinate system, the calculation formula for the initial position of the companion satellite in the z direction is:

9. The method for designing a circular satellite formation configuration in an elliptical reference orbit plane according to claim 1, characterized in that: The step S8 comprises: S801: The calculation formula of the direction cosine matrix from the reference satellite geocentric inertial coordinate system to the RIC coordinate system is: Where r and v are the position vector and velocity vector of the reference satellite's center of mass in the geocentric inertial coordinate system at the initial moment, respectively; S802: The calculation formula of the position vector of the companion satellite in the Earth-centered inertial coordinate system is: r B =r+R T [x0 y0 z0] T Where r is the position vector of the reference satellite mass center in the geocentric inertial coordinate system at the initial moment, R T is the transpose of the direction cosine matrix from the reference satellite geocentric inertial coordinate system to the RIC coordinate system in S801, [x0 y0 z0] T is the relative position vector of the companion satellite in S7 at the initial time in the reference satellite RIC coordinate system; S803: The velocity vector calculation formula in the Earth-centered inertial coordinate system of the companion satellite is: Where v is the position vector of the reference satellite mass center in the geocentric inertial coordinate system at the initial moment, R T is the transpose of the direction cosine matrix from the reference satellite geocentric inertial coordinate system to the RIC coordinate system in S801, [x0 y0 z0] T is the relative position vector of the companion satellite in S7 at the initial time in the reference satellite RIC coordinate system, is the relative velocity vector of the companion satellite in S7 in the reference satellite RIC coordinate system at the initial moment, μ is the earth's gravitational constant, a is the semi-major axis of the reference satellite orbit selected by S1, e is the eccentricity of the reference satellite orbit selected by S1, and f0 is the initial true anomaly angle of the reference satellite orbit selected by S1.

10. The method for designing a circular satellite formation configuration in an elliptical reference orbit plane according to claim 1, characterized in that: The step S9 comprises: S901: The calculation formula of the companion satellite orbital angular momentum is: h=r B ×v B where r B 、v B are the position vector and velocity vector of the satellite's mass center in the geocentric inertial coordinate system at the initial moment in S8, respectively; S902: The calculation formula for the companion satellite orbit eccentricity is: e=|e| where μ is the Earth's gravitational constant, h is the orbital angular momentum of the companion satellite in S901, and r B 、v B are the position vector and velocity vector of the satellite mass center in the geocentric inertial coordinate system at the initial moment in S8, respectively, and |·| represents the vector modulus; S903: The calculation formula for the companion satellite orbit inclination is: where h is the orbital angular momentum of the companion satellite in S901, h z is the component of h along the z axis in the geocentric inertial coordinate system; S904: The calculation formula for the right ascension of the companion satellite upgrade point is: Ω=arctan(h x ,-h y ) where h x 、h y are the x-axis and y-axis components of the companion satellite orbital angular momentum h in S901 in the geocentric inertial coordinate system; S905: The calculation formula for the semi-major axis of the companion satellite orbit is: Where μ is the Earth's gravitational constant, h is the modulus of the companion satellite's orbital angular momentum h in S901, and e is the companion satellite's orbital eccentricity in S902; S906: The calculation formula for the companion satellite perigee angle is: i n =[0 0 1]×h Where h is the orbital angular momentum of the companion satellite in S901, and e is the eccentricity vector of the companion satellite in S902; S906, the calculation formula of the true anomaly angle of the companion satellite is: where r B is the position vector of the companion satellite mass center in the geocentric inertial coordinate system at the initial moment in S8, e is the companion satellite eccentricity vector in S902, and |·| represents the vector modulus.

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