A satellite formation autonomous control method based on relative motion fitting

Abstract

A method for autonomous control of satellite formations based on relative motion fitting belongs to the field of spacecraft control technology. The present invention solves the problem that traditional formation control methods cannot meet formation control requirements in an environment where navigation signals are missing. The method of the present invention includes the following steps: Step 1: Whenever the main satellite passes the ascending node, the inter-satellite spacing and direction of multiple discretely distributed sub-satellites relative to the main satellite observed in the previous orbital period are used to perform relative motion fitting, thereby obtaining the coefficients of the sub-satellite relative motion function; Step 2: Based on the coefficients of the sub-satellite relative motion function, the orbital parameter deviation of the sub-satellite relative to the main satellite is obtained; Step 3: Based on the orbital parameter deviation, the control vector and control duration required for sub-satellite formation adjustment are obtained; Step 4: Each sub-satellite is autonomously controlled according to the control vector and control duration obtained in Step 3. The present invention is used for autonomous control of satellite formations.

Description

Technical Field

[0001] The invention relates to a satellite formation autonomous control method based on relative motion fitting, and belongs to the technical field of spacecraft control. Background Art

[0002] Satellite formation flying is a space technology that uses multiple satellites to collaborate to complete specific missions. Compared to traditional single large satellites, satellite formations, through the distributed layout of multiple small satellites, can improve mission resolution, coverage, and redundancy, demonstrating greater flexibility, lower costs, and enhanced fault tolerance when completing complex missions. Formation flying technology is widely used in fields such as Earth observation, space science, deep space exploration, and communication enhancement. For example, in Earth observation and deep space exploration missions, it can achieve scientific goals such as multi-angle observation and interferometry, while also optimizing data relay and communication efficiency.

[0003] However, in an environment without navigation signals, satellite formation flying faces significant challenges, especially in specific scenarios such as deep space exploration, polar observation, or severe electronic interference. Traditional positioning and time synchronization methods based on the Global Navigation Satellite System (GNSS) are difficult to apply. Most traditional satellite formation control methods use the CW equation to control the relative position and velocity obtained by navigation, or use the deviation orbit parameters obtained by navigation to calculate and control the control quantity. However, the use of the CW equation for formation control will cause linearization errors when the formation is at a long distance, and for multi-satellite formation missions under deep space conditions or navigation-constrained conditions, traditional formation control methods cannot meet the formation control requirements. Summary of the Invention

[0004] In order to solve the problem that traditional formation control methods cannot meet formation control requirements in an environment where navigation signals are missing, the present invention provides a satellite formation autonomous control method based on relative motion fitting.

[0005] The present invention provides a satellite formation autonomous control method based on relative motion fitting, the method comprising the following steps:

[0006] Step 1: Whenever the primary star passes the ascending node, the relative motion of multiple discrete satellites relative to the primary star is fitted using the inter-satellite distances and directions observed in the previous orbital period, thereby obtaining the coefficients of the relative motion function of the satellites;

[0007] Step 2: Obtain the orbital parameter deviation of the satellite relative to the main satellite based on the coefficient of the relative motion function of the satellite;

[0008] Step 3: Obtain the control vector and control duration required for satellite formation adjustment based on the orbital parameter deviation;

[0009] Step 4: Each satellite performs autonomous control according to the control vector and control duration obtained in step 3.

[0010] Preferably, in step 1, the relative motion function of the neutron star is:

[0011]

[0012] In the LVHV coordinate system, the relative motion function of the satellite is fitted according to the X, Y, and Z axes respectively;

[0013] Where x, y, z are the intersatellite distances between the satellite and the main satellite in the X, Y, and Z axes in the LVLH coordinate system. is the main star latitude angle, , is the argument of pericenter, is the true anomaly angle, For time;

[0014] is the coefficient to be solved, is the amplitude of the relative motion oscillation on the X axis, is the phase of the relative motion on the X axis relative to the latitude argument oscillation, is the rate of change of the drift distance of relative motion on the Y axis, is the drift distance of relative motion on the Y axis, is the amplitude of the relative motion oscillation on the Y axis, is the phase of the relative motion on the Y axis relative to the true anomaly oscillation, is the amplitude of the relative motion oscillation on the Z axis, is the phase of the relative motion on the Z axis relative to the latitude argument oscillation.

[0015] Preferably, the motion fitting in step 1 adopts a nonlinear least squares method.

[0016] Preferably, in step 2, the orbital parameter deviation of the neutron star relative to the primary star is 、 、 、 、 、 Get it as follows:

[0017]

[0018]

[0019]

[0020]

[0021] Where, is the orbital semi-major axis deviation, is the orbital inclination deviation, is the right ascension deviation of the ascending node, 、 is the non-singular orbital parameter deviation, is the latitude argument deviation;

[0022] is the orbital inclination of the main star, is the distance between the main star and the central celestial body, T is the orbital period of the main star, and n is the orbital angular velocity of the main star;

[0023] is the main star semi-diameter, , is the semi-major axis of the primary star's orbit, is the eccentricity of the main star's orbit;

[0024] are the non-singular orbital parameters of the primary star, , , is the argument of perihelion of the main star;

[0025] Calculated when the primary star passes the ascending node for the k+1th time , Calculated when the primary star passes the ascending node for the first time .

[0026] Preferably, the control vector in step 3 includes a control amount and a latitude argument for applying the control amount;

[0027] Control volume 、 、 Get it as follows:

[0028]

[0029] In the formula is the intermediate parameter, , is the gravitational constant of the central celestial body;

[0030] is the tangential pulse control quantity along the track, , is the orbital angular momentum of the primary star, ; is the total deviation of the semi-major axis that needs to be adjusted, , Add the control amount for the semi-major axis, , Adjust for a given the total duration of the request;

[0031] This control amount is applied to both the pericentric and apocentric celestial points. , the latitude argument of the control quantity applied at the pericentric celestial point , the latitude argument of the control quantity applied at the far-center celestial point ;

[0032] is the pulse control quantity along the track normal direction, , the latitude argument of the control value ;

[0033] is the radial pulse control quantity along the track; , the latitude argument of the control value .

[0034] Preferably, in step 4, the process of each sub-satellite performing autonomous control according to the control vector obtained in step 3 is as follows: normal control, radial control and tangential control are performed in sequence within three consecutive orbital periods, one type of control is performed in each orbital period, and after each type of control is performed, the process returns to steps 1 to 3, and the deviation, control vector and control duration are recalculated before the next type of control is performed, so as to achieve formation adjustment of the sub-satellites with the latitude argument of the main satellite as a reference value when navigation is limited and the relevant orbital information of the sub-satellite cannot be obtained.

[0035] Preferably, the start time and end time of normal control are 、 ; The start and end times of radial control are 、 ; The starting and ending times of the tangential control of the pericentric celestial body are 、 The start and end times of the tangential control of the far-center celestial body are 、 ;

[0036] is the predicted latitude argument of the main satellite when the control variable is applied 、 、 、 Corresponding time, 、 、 are the control time required for normal control, radial control, and tangential control respectively.

[0037] Preferably, the control time required for normal control is:

[0038]

[0039] Where, is the thrust of the satellite thruster, is the mass of the sub-star.

[0040] Preferably, the control time required for radial control is:

[0041]

[0042] Where, is the thrust of the satellite thruster, is the mass of the sub-star.

[0043] Preferably, the control time required for tangential control is:

[0044]

[0045] Where, is the thrust of the satellite thruster, is the mass of the sub-star.

[0046] Beneficial effects of the present invention:

[0047] This paper proposes a method for autonomous satellite formation control based on relative motion fitting, suitable for multi-satellite autonomous formation control missions in deep space or under navigation-constrained conditions. This method uses the discrete inter-satellite ranging and direction information (including inter-satellite spacing and direction) from the previous orbital period to perform relative motion fitting each time a formation satellite passes through the ascending node. The resulting function correlation coefficients are used to directly calculate the corresponding formation satellite orbital parameter deviations. The required control variables and the latitude argument of the applied control variables can then be calculated. Autonomous formation control is achieved through phase drift control and in-plane and out-of-plane flyby control.

[0048] Compared with using GNSS navigation results for formation control, it has the advantages of good autonomy, no need for ground navigation support, and high control accuracy at longer distances. BRIEF DESCRIPTION OF THE DRAWINGS

[0049] Figure 1 is a flow chart of a satellite formation autonomous control method based on relative motion fitting according to the present invention;

[0050] Figure 2 It is a schematic diagram of the LVLH coordinate system;

[0051] Figure 3 is the X-axis fitting result, where Figure 3 (a) is the position curve of LVLH relative to X axis, Figure 3 (b) LVLH system relative X-axis position fitting error curve;

[0052] Figure 4 is the Y-axis fitting result, where Figure 4 (a) is the LVLH relative Y-axis position curve, Figure 4 (b) LVLH system relative Y-axis position fitting error curve;

[0053] Figure 5 is the Z-axis fitting result, where Figure 5 (a) is the relative Z-axis position curve of the LVLH system, Figure 5 (b) LVLH system relative Z-axis position fitting error curve;

[0054] Figure 6 This is a schematic diagram of autonomous control of the control quantity obtained by the method of the present invention;

[0055] Figure 7 is the relative motion relationship diagram of the main star and the satellite, where Figure 7 (a) is the global graph, Figure 7 (b) is the YZ plane diagram, with normal control, where the red dot is the position where normal control is applied. Figure 7 (c) is an XY plane diagram for radial control, where the red dot is the position where radial control is applied. Figure 7 (d) is the XZ plane diagram;

[0056] Figure 8 This is a schematic diagram of the thrust control exerted by the satellite;

[0057] Figure 9 is the orbital semi-major axis deviation curve;

[0058] Figure 10 is the orbital eccentricity deviation curve;

[0059] Figure 11 is the orbital inclination deviation curve;

[0060] Figure 12 is the right ascension deviation curve of the ascending node;

[0061] Figure 13 It is the argument deviation curve of the pericentric celestial body point. DETAILED DESCRIPTION

[0062] The following will clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. 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 making any creative efforts shall fall within the scope of protection of the present invention.

[0063] It should be noted that, in the absence of conflict, the embodiments of the present invention and the features in the embodiments may be combined with each other.

[0064] The present invention will be further described below with reference to the accompanying drawings and specific embodiments, but they are not intended to limit the present invention.

[0065] Specific implementation method 1: Figures 1 to 13 This embodiment describes a method for autonomous control of a satellite formation based on relative motion fitting, which includes the following steps:

[0066] Step 1: Whenever the primary star passes the ascending node, the relative motion of multiple discrete satellites relative to the primary star is fitted using the inter-satellite distances and directions observed in the previous orbital period, thereby obtaining the coefficients of the relative motion function of the satellites;

[0067] Step 2: Obtain the orbital parameter deviation of the satellite relative to the main satellite based on the coefficient of the relative motion function of the satellite;

[0068] Step 3: Obtain the control vector and control duration required for satellite formation adjustment based on the orbital parameter deviation;

[0069] Step 4: Each satellite performs autonomous control according to the control vector and control duration obtained in step 3.

[0070] The present invention takes formation control into consideration, and realizes phase drift control by controlling the semi-major axis of the orbit, and realizes oscillation control of in-plane and out-of-plane quantities by controlling the eccentricity vector, orbit inclination, and right ascension of ascending node. Figure 3 shown.

[0071] LVLH coordinate system see Figure 2 As shown, the main satellite and its sub-satellites fly in formation around a central celestial body. In specific scenarios such as deep space exploration, polar observation, or severe electronic interference, traditional positioning and time synchronization methods based on the Global Navigation Satellite System (GNSS) are difficult to apply. The present invention adopts a nonlinear method for analysis.

[0072] In the LVLH coordinate system, the relative motion law can be expressed by the non-singular relative orbit element deviation as follows:

[0073] (1)

[0074] Where x, y, z are the inter-satellite distances between the satellite and the main satellite in the X, Y, and Z axes in the LVLH coordinate system. is the distance between the main star and the central celestial body, is the semi-major axis of the primary star's orbit, is the eccentricity of the main star's orbit, is the orbital inclination of the main star, is the right ascension of the ascending node of the main star, is the argument of pericenter of the main star, is the true periapsis angle of the main star, is the main star semi-diameter, , is the main star latitude angle, In order to avoid the singularity problem caused by the near-circular orbit, non-singular orbit parameters are used. are the non-singular orbital parameters of the primary star, , The above symbols are all related parameters of the main star. represents the deviation of the relevant parameters between the satellite and the main star, where is the orbital semi-major axis deviation, is the orbital inclination deviation, is the right ascension deviation of the ascending node, is the latitude argument deviation, 、 is the non-singular orbital parameter deviation, 、 Used to characterize the eccentricity vector deviation.

[0075] For the near-circular small deviation assumption, equation (1) can be simplified to the following form:

[0076] (2)

[0077] As the relative motion function of the satellite, the relative motion function of the satellite is fitted along the X, Y, and Z axes in the LVHV coordinate system. That is, three relative motion functions with a total of 8 coefficients are used. The inter-satellite distance (x, y, z) of the satellite relative to the primary star observed in the previous orbital period is obtained by sampling at least 5 discrete sampling points of the satellite within the orbital period (one circle).

[0078] Where, For time;

[0079] is the coefficient to be solved, is the amplitude of the relative motion oscillation on the X axis, is the phase of the relative motion on the X axis relative to the latitude argument oscillation, is the rate of change of the drift distance of relative motion on the Y axis, is the drift distance of relative motion on the Y axis, is the amplitude of the relative motion oscillation on the Y axis, is the phase of the relative motion on the Y axis relative to the true anomaly oscillation, is the amplitude of the relative motion oscillation on the Z axis, is the phase of the relative motion on the Z axis relative to the latitude argument oscillation.

[0080] In step 1, motion fitting adopts nonlinear least squares method.

[0081] The parameters related to the deviation estimation in formula (2) are as follows:

[0082] (3)

[0083] In the formula is the relative deviation orbital angular velocity between the main star and the satellite, n is the orbital angular velocity of the main star, , is the gravitational constant of the central celestial body. Therefore, the relative distance of the Y axis (phase) of the LVLH system can be controlled by controlling the semi-major axis.

[0084] From the above relative motion analysis, it can be seen that in the LVLH coordinate system, the relative motion of the formation can be fitted into the above form, where the above are coefficients to be solved, so we can consider using relative motion fitting to obtain the above unknown parameters and then obtain the orbital deviations of the main satellite and the sub-satellite.

[0085] Step 2: Deviation of the orbital parameters of the neutron star relative to the host star 、 、 、 、 、 Get it as follows:

[0086]

[0087]

[0088] (4)

[0089]

[0090] Formula (4) is derived based on formula (3).

[0091] Where,

[0092] T is the orbital period of the primary star;

[0093] Calculated when the primary star passes the ascending node for the k+1th time , Calculated when the primary star passes the ascending node for the first time .

[0094] The control vector in step 3 includes the control amount and the latitude argument at which the control amount is applied;

[0095] Control volume 、 、 Get it as follows:

[0096]

[0097] In the formula is the intermediate parameter, ;

[0098] is the tangential pulse control quantity along the track, , is the orbital angular momentum of the primary star, ; is the total deviation of the semi-major axis that needs to be adjusted, , Add the control amount for the semi-major axis, , Adjust for a given the total duration of the request;

[0099] This control amount is applied to both the pericentric and apocentric celestial points. , the latitude argument of the control quantity applied at the pericentric celestial point , the latitude argument of the control quantity applied at the far-center celestial point ;

[0100] is the pulse control quantity along the track normal direction, , the latitude argument of the control value ;

[0101] is the radial pulse control quantity along the track; , the latitude argument of the control value .

[0102] The control of dimensional argument deviation adjustment can be achieved indirectly by controlling the semi-major axis. The specific formula is as follows:

[0103]

[0104] Calculate the latitude angle deviation of the satellite relative to the main star After that, the given adjustment Total time required The required additional control amount of the semi-major axis can be calculated by the following formula

[0105]

[0106] Combined with the results obtained in step 2 , get the total deviation of the semi-major axis that actually needs to be controlled .

[0107] In step 4, the process of autonomous control of each sub-satellite according to the control vector obtained in step 3 is as follows: normal control, radial control and tangential control are carried out in sequence within three consecutive orbital periods, and one type of control is performed in each orbital period. After each type of control is completed, the process returns to steps 1 to 3, and the deviation, control vector and control duration are recalculated before the next type of control is performed. This is to achieve formation adjustment of the sub-satellites with the main satellite latitude argument as a reference value when navigation is limited and the sub-satellite-related orbital information cannot be obtained.

[0108] The start and end times of normal control are 、 ; The start and end times of radial control are 、 ; The starting and ending times of the tangential control of the pericentric celestial body are 、 The start and end times of the tangential control of the far-center celestial body are 、 ;

[0109] is the predicted latitude argument of the main satellite when the control variable is applied 、 、 、 Corresponding time, 、 、 are the control time required for normal control, radial control, and tangential control respectively.

[0110] Normal control is to eliminate the deviation of the ascending node's right ascension and orbital inclination deviation , through the latitude argument The applied normal pulse control amount at Take control.

[0111] The control time required for normal control is:

[0112]

[0113] Where, is the thrust of the satellite thruster, is the mass of the sub-star.

[0114] Radial control is used to eliminate the eccentricity vector deviation by latitude argument The applied normal pulse control amount at Take control.

[0115] The control time required for radial control is:

[0116]

[0117] Tangential control is used to eliminate orbital semi-major axis deviation During the orbital period, the satellite and the main satellite will continue to deviate, and the semi-major axis will be controlled by the additional amount. , the total offset of the semi-major axis that needs to be eliminated , by eliminating the semi-major axis deviation and thus eliminating the latitude angle deviation , passing through the pericentric points with a phase interval of 180° ( ), apocentric celestial point ( ) respectively apply a tangential pulse control amount Take control.

[0118] The control time required for tangential control is:

[0119]

[0120] Normal control, radial control, and tangential control are performed in different orbital periods. Normal control is applied first. After normal control is completed, the system returns to steps 1 through 3 to recalculate the deviation, control amount, and control duration before applying radial control. After radial control is completed, the system returns to steps 1 through 3 to recalculate the deviation, control amount, and control duration before applying tangential control.

[0121] A specific example is given below.

[0122] The feasibility of the method is verified by taking a satellite formation control problem of orbiting the moon as a simulation example. The orbital parameters of the main satellite are selected as follows:

[0123] The satellite orbit parameters are selected as follows:

[0124] The control target is the Y-axis of the LVLH coordinate system. After 14 days, the satellite is positioned 10 km ahead of the primary satellite. The satellite mass is set at 100 kg, and it is equipped with an 80 mN thruster for low-thrust control, with a thrust error of 5%. The maximum single-shot thruster operation time is 600 seconds. The primary satellite orbit determination error is 1 km at 1 m / s, the ranging and direction error between the primary and satellite is 3 m at one arc minute, and the thrust direction error is 2°.

[0125] Step 1: Using the observed discrete sampling points, the relative motion fitting results are obtained using nonlinear least squares curve fitting. Figures 3 to 5 shown.

[0126] Step 2: Obtain the orbital parameter deviation of the satellite relative to the main satellite based on the coefficient of the relative motion function of the satellite;

[0127] Based on the fitting results of step 1, the deviation of the primary star is estimated, such as Figures 9 to 13shown.

[0128] Step 3: Obtain the control vector and control duration required for satellite formation adjustment based on the orbital parameter deviation;

[0129] Step 4: Each satellite performs autonomous control according to the control vector and control duration obtained in step 3. Figure 7 and Figure 8 .

[0130] Autonomous control of the satellite formation is achieved through normal, radial, and tangential control. Tangential control implements phase drift control, while normal and radial control enable in-plane and out-of-plane flyby control. The oscillation of the satellites relative to the parent satellite is effectively reduced, and the relative distance in the Y direction can be controlled to 10 km. The autonomous formation control algorithm is effective.

[0131] Although the present invention is described herein with reference to specific embodiments, it should be understood that these embodiments are merely illustrative of the principles and applications of the invention. It should be understood that many modifications may be made to the illustrative embodiments, and that other arrangements may be devised, without departing from the spirit and scope of the invention as defined by the appended claims. It should be understood that the various dependent claims and features described herein may be combined in ways other than those described in the original claims. It should also be understood that features described in conjunction with individual embodiments may be used in conjunction with other described embodiments.

Claims

1. A satellite formation autonomous control method based on relative motion fitting, characterized in that: The method comprises the following steps: Step 1: Whenever the primary star passes the ascending node, the relative motion of multiple discrete satellites relative to the primary star is fitted using the inter-satellite distances and directions observed in the previous orbital period, thereby obtaining the coefficients of the relative motion function of the satellites; Step 2: Obtain the orbital parameter deviation of the satellite relative to the main satellite based on the coefficient of the relative motion function of the satellite; Step 3: Obtain the control vector and control duration required for satellite formation adjustment based on the orbital parameter deviation; Step 4: Each satellite performs autonomous control according to the control vector and control duration obtained in step 3; In step 1, the relative motion function of the neutron star is: In the LVHV coordinate system, the relative motion function of the satellite is fitted according to the X, Y, and Z axes respectively; Where x, y, z are the intersatellite distances between the satellite and the main satellite in the X, Y, and Z axes in the LVLH coordinate system. is the main star latitude angle, , is the argument of pericenter, is the true anomaly angle, For time; is the coefficient to be solved, is the amplitude of the relative motion oscillation on the X axis, is the phase of the relative motion on the X axis relative to the latitude argument oscillation, is the rate of change of the drift distance of relative motion on the Y axis, is the drift distance of relative motion on the Y axis, is the amplitude of the relative motion oscillation on the Y axis, is the phase of the relative motion on the Y axis relative to the true anomaly oscillation, is the amplitude of the relative motion oscillation on the Z axis, is the phase of the relative motion on the Z axis relative to the latitude argument oscillation; Step 2: Deviation of the orbital parameters of the neutron star relative to the host star 、 、 、 、 、 Get it as follows: Where, is the orbital semi-major axis deviation, is the orbital inclination deviation, is the right ascension deviation of the ascending node, 、 is the non-singular orbital parameter deviation, is the latitude argument deviation; is the orbital inclination of the main star, is the distance between the main star and the central celestial body, T is the orbital period of the main star, and n is the orbital angular velocity of the main star; is the main star semi-diameter, , is the semi-major axis of the primary star's orbit, is the eccentricity of the main star's orbit; are the non-singular orbital parameters of the primary star, , , is the argument of perihelion of the main star; Calculated when the primary star passes the ascending node for the k+1th time , Calculated when the primary star passes the ascending node for the first time ; The control vector in step 3 includes the control amount and the latitude argument at which the control amount is applied; Control volume 、 、 Get it as follows: Where, is the intermediate parameter, , is the gravitational constant of the central celestial body; is the tangential pulse control quantity along the track, , is the orbital angular momentum of the primary star, ; is the total deviation of the semi-major axis that needs to be adjusted, , Add the control amount for the semi-major axis, , Adjust for a given the total duration of the request; This control amount is applied to both the pericentric and apocentric celestial points. , the latitude argument of the control quantity applied at the pericentric celestial point , the latitude argument of the control quantity applied at the far-center celestial point ; is the pulse control quantity along the track normal direction, , the latitude argument of the control value ; is the radial pulse control quantity along the track; , the latitude argument of the control value .

2. The method for autonomous satellite formation control based on relative motion fitting according to claim 1, characterized in that: In step 1, motion fitting adopts nonlinear least squares method.

3. The method for autonomous satellite formation control based on relative motion fitting according to claim 2, characterized in that: In step 4, the process of autonomous control of each sub-satellite according to the control vector obtained in step 3 is as follows: normal control, radial control and tangential control are carried out in sequence within three consecutive orbital periods, and one type of control is performed in each orbital period. After each type of control is completed, the process returns to steps 1 to 3, and the deviation, control vector and control duration are recalculated before the next type of control is performed. This is to achieve formation adjustment of the sub-satellites with the main satellite latitude argument as a reference value when navigation is limited and the sub-satellite-related orbital information cannot be obtained.

4. The method for autonomous satellite formation control based on relative motion fitting according to claim 3, characterized in that: The start and end times of normal control are 、 ; The start and end times of radial control are 、 ; The starting and ending times of the tangential control of the pericentric celestial body are 、 The start and end times of the tangential control of the far-center celestial body are 、 ; is the predicted latitude argument of the main satellite when the control variable is applied 、 、 、 Corresponding time, 、 、 are the control time required for normal control, radial control, and tangential control respectively.

5. The method for autonomous satellite formation control based on relative motion fitting according to claim 4, characterized in that: The control time required for normal control is: Where, is the thrust of the satellite thruster, is the mass of the sub-star.

6. The method for autonomous satellite formation control based on relative motion fitting according to claim 4, characterized in that: The control time required for radial control is: Where, is the thrust of the satellite thruster, is the mass of the sub-star.

7. The method for autonomous satellite formation control based on relative motion fitting according to claim 4, characterized in that: The control time required for tangential control is: Where, is the thrust of the satellite thruster, is the mass of the sub-star.

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