Satellite orbit transfer strategy planning method, device and equipment and storage medium

Abstract

The application discloses a satellite orbit transfer strategy planning method and device, equipment and a storage medium. The method comprises the following steps: initializing an initial time of a satellite and performing orbit period extrapolation to update a t time to an extrapolated time; initializing an orbit transfer flag as not completed, calculating a first orbit parameter of the satellite at the t time and a second orbit parameter of a target virtual satellite at the t time, and determining a relative orbit parameter of the satellite to the target virtual satellite; determining each orbit parameter correction flag according to the relative orbit parameter and a preset orbit threshold value, and determining a corresponding orbit control sub-mode based on the each orbit parameter correction flag; when the orbit transfer flag is completed, determining a firing parameter of the satellite based on the orbit control sub-mode, and performing orbit transfer maneuver and orbit transfer phase adjustment processing. The scheme can quickly select the corresponding orbit control sub-mode according to the current orbit transfer requirement, quickly generate the calculation and simulation verification of the electric propulsion orbit transfer parameter, and make the satellite accurately enter the predetermined orbit position.

Description

Technical Field

[0001] This invention generally relates to the field of satellite orbit control technology, and specifically to a satellite orbit change strategy planning method, apparatus, equipment, and storage medium. Background Technology

[0002] With the rapid development of satellite communication technology, electric propulsion technology has also developed rapidly. It is widely used in satellite orbit control. Due to the characteristics of high specific impulse, low propellant consumption and stronger orbit change capability, electric propulsion systems are increasingly used on satellites to achieve orbital maneuver control.

[0003] Currently, related technologies employ direct or indirect optimization methods to plan electric propulsion orbit change strategies, such as genetic algorithms. However, these strategies are time-consuming and struggle to model limitations such as satellite attitude pointing, attitude maneuverability, single ignition duration, and ground shadow. This can easily lead to situations where the solution fails to converge to the optimal one, resulting in low accuracy in satellite orbit changes. Summary of the Invention

[0004] In view of the above-mentioned defects or deficiencies in the prior art, it is desirable to provide a satellite orbit change strategy planning method, apparatus, equipment and storage medium.

[0005] In a first aspect, the present invention provides a satellite orbit change strategy planning method, the method comprising:

[0006] Initialize the satellite's initial time and extrapolate its orbital period to update time t to the extrapolated time; t>0;

[0007] The initial orbit change flag is set to incomplete. The first orbital parameters of the satellite at time t and the second orbital parameters of the target virtual satellite at time t are calculated. Based on the first orbital parameters and the second orbital parameters, the relative orbital parameters of the satellite to the target virtual satellite are determined.

[0008] Based on the relative track parameters and the preset track threshold, determine the correction flags for each track parameter, and based on the correction flags for each track parameter, determine the completion results of the corresponding track control sub-mode and track change flag.

[0009] Based on the completion result of the track change indicator and the track control sub-mode, track change maneuvering and track change phase adjustment processing are performed.

[0010] In one embodiment, the relative orbital parameters include: relative semi-major axis root, relative eccentricity root, relative inclination root, relative ascending node right ascension root, relative perigee argument root, relative latitude argument root, and relative eccentricity vector root.

[0011] Based on the relative orbit parameters and preset orbit threshold values, determine the correction flags for each orbit parameter, including:

[0012] When the absolute value of the relative semi-major axis flat root is greater than the preset relative semi-major axis flat root, the semi-major axis flat root correction flag is determined to be 1; otherwise, the semi-major axis flat root revision flag is determined to be 0.

[0013] When the absolute value of the relative eccentricity root is greater than the preset relative eccentricity root, the eccentricity root correction flag is determined to be 1; otherwise, the eccentricity root correction flag is determined to be 0.

[0014] When the absolute value of the relative perigee argument root is greater than the preset relative perigee argument root, the perigee argument root correction flag is set to 1; otherwise, the perigee argument root correction flag is set to 0.

[0015] When the absolute value of the relative tilt angle root is greater than the preset relative tilt angle root, the tilt angle root correction flag is determined to be 1; otherwise, the tilt angle root correction flag is determined to be 0.

[0016] When the absolute value of the relative ascending node right ascension root is greater than the preset relative ascending node right ascension root, the ascending node right ascension root correction flag is set to 1; otherwise, the ascending node right ascension root correction flag is set to 0.

[0017] In one embodiment, the orbital parameter correction flags further include a first correction flag and a second correction flag, and the method further includes:

[0018] When the eccentricity square root correction flag is 1 or the perigee argument square root correction flag is 1, the first correction flag is determined to be 1; otherwise, the first correction flag is determined to be 0.

[0019] When the first correction flag is 1 and the semi-major axis flat root correction flag is 1, the second correction flag is determined to be 1; otherwise, the second correction flag is determined to be 0.

[0020] In one embodiment, the corresponding orbit control sub-mode is determined based on the orbit parameter correction flags, including:

[0021] Determine whether the second correction flag is 1;

[0022] If the second correction flag is 1, the corresponding track control sub-mode is determined according to the inclination angle root correction flag and the ascending node right ascension root correction flag.

[0023] If the second correction flag is not 1, the corresponding track control sub-mode is determined according to the correction flag of the semi-major axis root, the correction flag of the inclination root, and the correction flag of the ascending node right ascension root.

[0024] In one embodiment, the corresponding track control sub-mode is determined based on the revision flag of the semi-major axis root, the inclination root correction flag, and the ascending node right ascension root correction flag, including:

[0025] Determine whether the correction flag for the flat root of the semi-major axis is 1;

[0026] If the correction flag for the semi-major axis flat root is 1, then the corresponding track control sub-mode is determined based on the correction flag for the inclination angle flat root and the correction flag for the right ascension of the ascending node.

[0027] If the correction flag for the semi-major axis flat root is not 1, then the corresponding track control sub-mode is determined based on the first correction flag, the inclination angle flat root correction flag, and the ascending node right ascension flat root correction flag.

[0028] In one embodiment, based on the completion result of the track change indicator and the track control sub-mode, track change maneuvering and track change phase adjustment processing are performed, including:

[0029] When the completion result of the orbit change indicator is "completed", calculate the relative latitudinal argument between the satellite and the target orbit;

[0030] The absolute value of the relative latitude argument is compared with the preset latitude argument;

[0031] When the absolute value of the relative latitude argument is not greater than the preset latitude argument, the phase adjustment time is adjusted using an optimization algorithm, and the orbit change flag is re-initialized to incomplete.

[0032] Based on the first and second orbit parameters, the corresponding orbit control sub-mode is determined again, and orbit change maneuvers and orbit change phase adjustment processes are performed until the absolute value of the relative latitude angle is not greater than the preset latitude angle.

[0033] In one embodiment, the track control sub-mode includes at least one of the following:

[0034] The combined corrections for the semi-major axis root, eccentricity root, and perigee argument root, as well as the individual corrections for the tilt root, are performed alternately.

[0035] The combined corrections to the semi-major axis, eccentricity, and perigee argument, as well as the individual correction to the right ascension of the ascending node, are performed alternately.

[0036] Perform joint corrections to the semi-major axis root, eccentricity root, and perigee argument root;

[0037] Alternately perform individual correction of the semi-major axis root and individual correction of the camber root;

[0038] The corrections for the semi-major axis root and the ascending node right ascension root are performed alternately.

[0039] Perform separate correction of the semi-major axis root;

[0040] The combined correction for the root of eccentricity and the root of the perigee angle, as well as the individual correction for the root of tilt, are performed alternately.

[0041] The joint correction of the eccentricity square root and the perigee argument square root, as well as the individual correction of the ascending node right ascension square root, are performed alternately.

[0042] Perform joint corrections to the square root of eccentricity and the square root of the perigee angle;

[0043] Perform individual correction of the tilt angle to flatten the root;

[0044] Perform a separate correction of the ascending node, right ascension, and root.

[0045] Secondly, embodiments of this application provide a satellite orbit change strategy planning device, which includes:

[0046] The initialization module is used to initialize the satellite's initial time and extrapolate its orbital period to update time t to the extrapolated time; t>0;

[0047] The calculation module is used to initialize the orbit change flag to incomplete, calculate the first orbit parameters of the satellite at time t and the second orbit parameters of the target virtual satellite at time t, and determine the relative orbit parameters of the satellite to the target virtual satellite based on the first orbit parameters and the second orbit parameters.

[0048] The determination module is used to determine the correction flags for each track parameter based on the relative track parameters and the preset track threshold values, and to determine the corresponding track control sub-mode based on the correction flags for each track parameter.

[0049] The orbit change module is used to determine the ignition parameters of the satellite based on the orbit control sub-mode when the orbit change flag is completed, and to perform orbit change maneuvers and orbit change phase adjustment processing.

[0050] Thirdly, embodiments of this application provide a computer device, including a memory, a processor, and a computer program stored in the memory and executable on the processor. When the processor executes the program, it implements the satellite orbit change strategy planning method as described in the first aspect above.

[0051] Fourthly, embodiments of this application provide a computer-readable storage medium having a computer program stored thereon, the computer program being used to implement the satellite orbit change strategy planning method as described in the first aspect above.

[0052] The satellite orbit change strategy planning method, apparatus, device, and storage medium provided in this application initialize the initial time of the satellite and extrapolate the orbit period to update time t to the extrapolated time. The orbit change flag is initialized as incomplete. The first orbit parameters of the satellite at time t and the second orbit parameters of the target virtual satellite at time t are calculated. Based on the first and second orbit parameters, the relative orbit parameters of the satellite to the target virtual satellite are determined. Then, based on the relative orbit parameters and preset orbit thresholds, correction flags for each orbit parameter are determined. Based on each correction flag, the corresponding orbit control sub-mode is determined. When the orbit change flag is complete, the ignition parameters of the satellite are determined based on the orbit control sub-mode, and orbit change maneuvers and orbit change phase adjustment processing are performed. Compared with existing technologies, this technical solution accurately determines the relative orbital parameters through the first and second orbital parameters, enabling the use of a low-thrust orbital maneuver during satellite orbit change. It allows for the rapid selection of the corresponding orbit control sub-mode based on current orbit change requirements, and the quick generation and simulation verification of electric propulsion orbit change parameters. This ensures the satellite accurately enters the predetermined orbit position (assuming a target virtual satellite always operates at this position), avoiding the convergence problems of traditional direct optimization and indirect algorithm optimization which are easily affected by initial values. It also meets the requirement of rapidly completing the electric propulsion orbit change strategy optimization calculation within a limited time, improving the accuracy of satellite orbit change. Furthermore, it is applicable to orbit changes from any initial orbit to a higher orbit. Attached Figure Description

[0053] Other features, objects, and advantages of this application will become more apparent from the following detailed description of non-limiting embodiments with reference to the accompanying drawings:

[0054] Figure 1 A flowchart illustrating the satellite orbit change strategy planning method provided in this application embodiment;

[0055] Figure 2 A schematic diagram of the semi-major axis root variation curve provided in the embodiments of this application;

[0056] Figure 3 A schematic diagram of the eccentricity root variation curve provided in the embodiments of this application;

[0057] Figure 4 A schematic diagram of the perigee argument square root variation curve provided in the embodiments of this application;

[0058] Figure 5 A schematic diagram of the relative parking reference track inclination angle variation curve provided in the embodiments of this application;

[0059] Figure 6 A schematic diagram of the relative ascending node right ascension and root variation curve provided in the embodiments of this application;

[0060] Figure 7 This is a flowchart illustrating the method for determining the corresponding track control sub-mode based on correction flags for each track parameter provided in an embodiment of this application.

[0061] Figure 8 This is a flowchart illustrating the satellite orbit change strategy planning method according to an embodiment of this application;

[0062] Figure 9 This is a schematic diagram of the structure of the satellite orbit change strategy planning device provided in the embodiments of this application;

[0063] Figure 10 This is a schematic diagram of the structure of a computer device provided in an embodiment of this application. Detailed Implementation

[0064] The present application will now be described in further detail with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are merely illustrative of the invention and not intended to limit it. Furthermore, it should be noted that, for ease of description, only the parts relevant to the invention are shown in the accompanying drawings.

[0065] It should be noted that, unless otherwise specified, the embodiments and features described in this application can be combined with each other. This application will now be described in detail with reference to the accompanying drawings and embodiments.

[0066] It's understandable that satellite constellations, due to the large number of satellites and the pursuit of low cost, typically equip each satellite with a single small-thrust electric thruster, with thrust generally in the tens of millinewtons. The orbital maneuvering time from launch to operational orbit can range from tens to hundreds of days, and the acceleration generated by the electric thruster is roughly on the order of magnitude of the satellite's environmental perturbation acceleration. Currently, related technologies employ direct or indirect optimization methods to plan electric thruster orbital maneuvering strategies, such as genetic algorithms. However, these strategies are time-consuming and struggle to model constraints such as satellite attitude pointing, attitude maneuvering capabilities, single ignition duration, and Earth shadow, making it prone to failing to converge to the optimal solution and resulting in low accuracy in satellite orbital maneuvers.

[0067] To address the aforementioned shortcomings, this application provides a satellite orbit change strategy planning method, apparatus, device, and storage medium. Compared with existing technologies, this technical solution accurately determines relative orbit parameters through first and second orbit parameters, enabling the use of a low-thrust orbit change method during satellite orbit change. It allows for rapid selection of the corresponding orbit control sub-mode based on current orbit change requirements, and rapid generation and simulation verification of electric propulsion orbit change parameters. This ensures the satellite accurately enters the predetermined orbit position (assuming a target virtual satellite always operates at this position), avoiding the convergence problems of traditional direct optimization and indirect algorithm optimization, which are easily affected by initial values. Simultaneously, it meets the requirement of rapidly completing electric propulsion orbit change strategy optimization calculations within a limited time, improving the accuracy of satellite orbit changes. Furthermore, it is applicable to orbit changes from any initial orbit to a higher orbit.

[0068] For ease of understanding and explanation, the following will use... Figures 1 to 10 This application provides a detailed description of the satellite orbit change strategy planning method, apparatus, equipment, and storage medium provided in its embodiments.

[0069] Figure 1 The diagram shown is a schematic flowchart of a satellite orbit change strategy planning method according to an embodiment of this application. This method can be executed by a satellite orbit change strategy planning device. Figure 1 As shown, the method includes:

[0070] S101. Initialize the satellite's initial time and extrapolate its orbital period to update time t to the extrapolated time; t>0.

[0071] It should be noted that the satellite orbits N times per day, completing ground orbit determination in the first and second orbits, performing electric propulsion orbit change strategy optimization calculations, collision risk analysis, and orbit change parameter annotation in the third orbit, and completing orbit maneuvers from the fourth to the Nth orbit.

[0072] Specifically, time t is defined as the initial time t0. Then, the phase adjustment time DTphase is set to 0 seconds, and the satellite orbit is extrapolated from time t by multiple orbital periods. Time t is then updated to the extrapolated time; t > 0 and t can be a positive integer. The multiple orbital periods can be set based on the number of orbits the satellite completes ground orbit determination, electric propulsion orbit change strategy, collision risk analysis, and orbit change parameter uploading optimization calculations. For example, if the satellite completes ground orbit determination in the first and second orbits, and completes electric propulsion orbit change strategy optimization calculations and orbit change parameter uploading in the third orbit, then the multiple orbital periods can be three orbital periods.

[0073] S102. Initialize the orbit change flag as incomplete, calculate the first orbit parameters of the satellite at time t and the second orbit parameters of the target virtual satellite at time t, and determine the relative orbit parameters of the satellite to the target virtual satellite based on the first orbit parameters and the second orbit parameters.

[0074] It should be noted that the aforementioned target virtual satellite refers to a satellite virtually operating in a preset orbital position. The aforementioned first orbital parameter refers to the parameter corresponding to the satellite's current orbit at time t, the second orbital parameter refers to the parameter corresponding to the target virtual satellite's predetermined orbital position at time t, and the relative orbital parameter refers to the relative parameter between the satellite's orbit and the target virtual satellite's orbit.

[0075] Specifically, the orbit change flag Iot = 0 can be defined to indicate that the orbit change is not completed, and the orbit change flag Iot = 1 can be defined to indicate that the orbit change is completed. The first orbital parameters of the satellite at time t and the second orbital parameters of the target virtual satellite at time t are calculated respectively. The first orbital parameters include the root mean square of the first semi-major axis, the root mean square of the first eccentricity, the root mean square of the first inclination, the root mean square of the first ascending node right ascension, the root mean square of the first perigee argument, the root mean square of the first latitude argument, and the root mean square of the first eccentricity vector. The second orbital parameters include the root mean square of the second semi-major axis, the root mean square of the second eccentricity, the root mean square of the second inclination, the root mean square of the second ascending node right ascension, the root mean square of the second perigee argument, the root mean square of the second latitude argument, and the root mean square of the second eccentricity vector.

[0076] Wherein, the root of the first semi-major axis of the satellite at time t is The first eccentricity square root is The first inclination angle is the root. The first ascending node, right ascension, is at the root. The first perigee argument square root is The square root of the first latitude argument is The square root of the first eccentricity vector

[0077] The root of the second semi-major axis of the target virtual satellite at time t is The second eccentricity square root is The second perigee argument is square root The root of the second eccentricity vector The root of the second inclination angle of the target virtual satellite at time t is The second ascending node, right ascension, is at the root. The square root of the second latitude argument is

[0078] After determining the first and second orbital parameters, the relative orbital parameters of the satellite to the target virtual satellite can be determined based on these parameters. These relative orbital parameters may include the root mean square of the relative semi-major axis, the root mean square of the relative eccentricity, the root mean square of the relative inclination, the root mean square of the relative right ascension of the ascending node, the root mean square of the relative perigee argument, the root mean square of the relative latitude argument, and the root mean square of the relative eccentricity vector. Please refer to [link to relevant documentation]. Figures 2-6As shown, schematic diagrams are presented sequentially of the horizontal root variation curves for the semi-major axis, eccentricity, perigee argument, inclination relative to the parking reference orbit, and right ascension relative to the ascending node. The relative orbital parameters can be calculated using the following formula:

[0079] Among them, the root of the relative semi-major axis of the satellite relative to the target virtual satellite is:

[0080]

[0081] The square root of the relative eccentricity of the above satellite relative to the target virtual satellite is:

[0082]

[0083]

[0084]

[0085] The square root of the perigee argument of the above satellite relative to the target virtual satellite is:

[0086]

[0087] The root of the relative inclination angle between the above satellite and the target virtual satellite is:

[0088]

[0089] The relative ascending node right ascension root of the above satellite relative to the target virtual satellite is:

[0090]

[0091] The square root of the relative latitude argument of the above satellite relative to the target virtual satellite is:

[0092]

[0093] In this embodiment, the relative orbital parameters of the satellite to the target virtual satellite can be accurately determined based on the first orbital parameters and the second orbital parameters, providing data guidance information for the subsequent determination of the orbital control sub-mode, so that the satellite can accurately enter the predetermined orbital position by changing its orbit.

[0094] S103. Based on the relative track parameters and preset track thresholds, determine the correction flags for each track parameter, and based on each track parameter correction flag, determine the completion results of the corresponding track control sub-mode and track change flag.

[0095] The aforementioned preset orbital thresholds are custom-set according to actual needs and may include preset relative semi-major axis root, preset relative eccentricity root, preset relative perigee argument root, preset relative inclination root, preset relative ascending node right ascension root, and preset relative latitude argument root. The preset relative semi-major axis root can be a relative semi-major axis root control threshold, denoted as... The preset relative eccentricity root can be a relative eccentricity root control threshold, expressed as: The preset relative tilt angle root can be a relative tilt angle root control threshold, expressed as: The preset relative ascending node right ascension root can be used as a control threshold for the relative ascending node right ascension root, denoted as: The preset relative perigee angle root can be used as the relative perigee angle root control threshold, expressed as: The preset relative latitude argument root can be used as the relative latitude argument root control threshold, expressed as: in and The threshold value is greater than 0.

[0096] Specifically, in the process of determining the correction flag for each track parameter based on the relative track parameters and the preset track threshold value, it can be that when the absolute value of the relative semi-major axis root is greater than the preset relative semi-major axis root, that is... Sure The correction flag for the semi-major axis root is set to Ia = 1; otherwise, the correction flag for the semi-major axis root is set to Ia = 0. When the absolute value of the relative eccentricity root is greater than the preset relative eccentricity root, i.e. The determined eccentricity root correction flag is Ie = 1; otherwise, the determined eccentricity root correction flag is Ie = 0. When the absolute value of the relative perigee argument root is greater than the preset relative perigee argument root, i.e. Set the perigee argument root correction flag to Iw = 1; otherwise, set the perigee argument root correction flag to Iw = 0. When the absolute value of the relative tilt root is greater than the preset relative tilt root, i.e. Sure The inclination angle square root correction flag is Ii = 1; otherwise, it is determined. The correction flag for the right ascension root of the relative ascending node is Ii = 0. When the absolute value of the right ascension root of the relative ascending node is greater than the preset right ascension root of the relative ascending node, i.e. Sure Set the right ascension and square root of the ascending node as the correction flag IW = 1; otherwise, determine... The right ascension root correction flag for the ascending node is IW = 0.

[0097] Furthermore, when the eccentricity square root correction flag is 1 and the perigee argument square root correction flag is 1, the first correction flag is determined to be 1; otherwise, the first correction flag is determined to be 0. When the first correction flag is 1 and the semi-major axis square root correction flag is 1, the second correction flag is determined to be 1; otherwise, the second correction flag is determined to be 0.

[0098] Specifically, if or This means that when the eccentricity square root correction flag is 1 or the perigee argument square root correction flag is 1, then... The first correction flag Iew = 1; otherwise, Iew = 0. If the semi-major axis square root correction flag Ia equals 1 and the first correction flag Iew equals 1, then The second correction flag Iaew = 1, otherwise Iaew = 0;

[0099] It should be noted that the above-mentioned track control sub-modes include at least one of the following:

[0100] The first orbit control sub-mode can be represented by alternating the combined correction of semi-major axis root, eccentricity root, and perigee argument root, as well as the individual correction of inclination root. and By alternately applying the combined corrections to the semi-major axis square root, eccentricity square root, and perigee argument square root, as well as the individual corrections to the ascending node right ascension square root, this second orbital control sub-mode can be represented as follows: and By performing joint corrections to the semi-major axis root, eccentricity root, and perigee argument root, the third orbital control sub-mode can be represented as follows: Alternating between individually correcting the semi-major axis root and individually correcting the inclination root, this fourth track control sub-mode can be represented as alternating completion. and Alternating between correcting the semi-major axis root and correcting the ascending node right ascension root separately, this fifth orbital control sub-mode can be represented as alternating completion. and Performing a separate correction to the semi-major axis root can represent the sixth track control mode as complete. By alternately performing the combined correction for eccentricity square root, perigee argument square root, and individual correction for tilt square root, this seventh orbital control mode can be represented as an alternating process. and By alternately performing the combined corrections for the square root of eccentricity and the square root of the perigee argument, as well as the separate corrections for the square root of the right ascension of the ascending node, this eighth orbital control sub-mode can be represented as an alternating process. and By performing joint corrections to the square root of eccentricity and the square root of the perigee angle, the ninth orbital control sub-mode can be represented as a complete... Performing a separate correction of the tilt angle root allows the tenth orbital control sub-mode to be represented as complete. By performing a separate correction to the right ascension of the ascending node, the eleventh orbital submode can be represented as complete.

[0101] In one embodiment, this application also provides a specific implementation method for determining the corresponding orbit control sub-mode based on correction flags for each orbit parameter; please refer to [link to relevant documentation]. Figure 7 As shown, the method includes:

[0102] S201. Determine whether the second correction flag is 1.

[0103] S202. If the second correction flag is 1, determine the corresponding track control sub-mode based on the inclination angle root correction flag and the ascending node right ascension root correction flag.

[0104] S203. If the second correction flag is not 1, determine the corresponding track control sub-mode based on the correction flag of the semi-major axis root, the correction flag of the inclination root, and the correction flag of the right ascension root of the ascending node.

[0105] Specifically, please see Figure 8 As shown, in the process of determining the corresponding track control sub-mode, it can first be determined whether the second correction flag Iaew is 1; if the second correction flag Iaew is 1, then it can be determined whether the inclination root correction flag Ii is 1; if the inclination root correction flag Ii is equal to 1, then the first track control sub-mode is selected.

[0106] If the inclination root correction flag Ii is not equal to 1, i.e., it is 0, then it is determined whether the ascending node right ascension root correction flag IW is 1. If the ascending node right ascension root correction flag IW is 1, then the second orbit control sub-mode is selected; if the ascending node right ascension root correction flag IW is not 1, i.e., it is 0, then the third orbit control sub-mode is selected.

[0107] Specifically, in determining the corresponding track control sub-mode based on the revision flag of the semi-major axis horizontal root, the inclination horizontal root correction flag, and the ascending node right ascension horizontal root correction flag, it can be first determined whether the correction flag of the semi-major axis horizontal root is 1; if the correction flag of the semi-major axis horizontal root is 1, then the corresponding track control sub-mode is determined based on the inclination horizontal root correction flag and the ascending node right ascension horizontal root correction flag; if the correction flag of the semi-major axis horizontal root is 0, then the corresponding track control sub-mode is determined based on the first correction flag, the inclination horizontal root correction flag, and the ascending node right ascension horizontal root correction flag.

[0108] If the second correction flag Iaew is not equal to 1 (i.e., it is 0), then the semi-major axis root correction flag Ia is checked to see if it is equal to 1. If the semi-major axis root correction flag Ia is equal to 1, then the inclination root correction flag Ii is checked to see if it is equal to 1. If the inclination root correction flag Ii is equal to 1, then the fourth rail control sub-mode is selected. If the inclination root correction flag Ii is not equal to 1 (i.e., it is 0), then the ascending node right ascension root correction flag IW is checked to see if it is equal to 1. If the ascending node right ascension root correction flag IW is equal to 1, then the fifth rail control sub-mode is selected. If the ascending node right ascension root correction flag IW is not equal to 1 (i.e., it is 0), then the sixth rail control sub-mode is selected.

[0109] If the correction flag for the semi-major axis is determined to be 0 (i.e., not 1), then the first correction flag (Iew) is checked to see if it equals 1. If Iew equals 1, then the inclination correction flag (Ii) is checked to see if it equals 1. If Ii equals 1, then the seventh rail control mode is selected. If Ii is not 1 (i.e., not 1), then the ascending node right ascension correction flag (IW) is checked to see if it equals 1. If IW equals 1, then the eighth rail control mode is selected. If IW is not 1 (i.e., not 0), then the ninth rail control mode is selected.

[0110] If the first correction flag Iew is not equal to 1 (i.e., it is 0), then it is determined whether the inclination root correction flag Ii is 1. If the inclination root correction flag Ii is equal to 1, then the tenth track control mode is selected. If the inclination root correction flag Ii is not equal to 1 (i.e., it is 0), then it is determined whether the ascending node right ascension root correction flag IW is 1. If the ascending node right ascension root correction flag IW is 1, then the eleventh track control mode is selected. If the ascending node right ascension root correction flag IW is not 1 (i.e., it is 0), then the track change flag Iot is updated to 1 to complete the process.

[0111] S104. Based on the completion results of the track change markers and the track control sub-mode, perform track change maneuvers and track change phase adjustment processing.

[0112] It should be noted that the above ignition parameters are used to characterize the parameters required for the satellite during the ignition process.

[0113] In this step, after selecting the corresponding orbit control sub-mode based on the orbit parameter correction flag, it can be determined whether the orbit change flag Iot is incomplete. When the orbit change flag Iot = 0, indicating incompleteness, the satellite's ignition parameters are calculated according to the selected orbit control sub-mode, a collision risk analysis is performed, the orbit change strategy is applied, and the orbit change maneuver from the 4th to the Nth orbit is completed. Time t is updated as the orbit change maneuver completion time, and the orbit change flag Iot = 0 is defined as incompleteness. Then, the first orbit parameters of the satellite at time t and the second orbit parameters of the target virtual satellite at time t are calculated. Based on the first and second orbit parameters, the relative orbit parameters of the satellite to the target virtual satellite are determined. Based on the relative orbit parameters and the preset orbit threshold, the correction flags for each orbit parameter are determined. Based on each orbit parameter correction flag, the corresponding orbit control sub-mode is determined, and the orbit change flags are updated until the orbit change flag Iot = 1.

[0114] When the orbit change flag Iot = 1, indicating that the orbit change is complete, the relative latitude argument between the satellite and the target orbit can be calculated, and the absolute value of the relative latitude argument can be compared with the preset latitude argument. When the absolute value of the relative latitude argument is not greater than the preset latitude argument, the phase adjustment time is adjusted using an optimization algorithm, the orbit change flag is reset to incomplete, and the corresponding orbit control sub-mode is determined again based on the first orbit parameters and the second orbit parameters, and orbit change maneuvers and orbit change phase adjustment processing are performed until the absolute value of the relative latitude argument is not greater than the preset latitude argument.

[0115] Specifically, in the process of comparing the absolute value of the relative latitude argument with the preset latitude argument, it can be to determine whether... If not, meaning the absolute value of the relative latitude argument is greater than the preset latitude argument, then the orbit change and phase adjustment are considered complete. If yes, meaning the absolute value of the relative latitude argument is not greater than the preset latitude argument, then the phase adjustment time DTphase is adjusted using an optimized algorithm, and the orbit change flag is re-initialized to incomplete. Then, based on the first and second orbit parameters, the corresponding orbit control sub-mode is determined again, and orbit change maneuvers and orbit change phase adjustment are performed until the absolute value of the relative latitude argument is not greater than the preset latitude argument. At this point, the satellite's orbit change and phase adjustment are complete.

[0116] In one embodiment, the satellite's orbital angular velocity and orbital period can be obtained, and can be expressed by the following formula:

[0117] Orbital angular velocity:

[0118] Orbital period:

[0119] Wherein, the root of the first semi-major axis of the satellite at time t is

[0120] After determining the corresponding orbit control sub-mode based on the correction flags for each orbital parameter, an adjustment strategy can be obtained by adjusting the semi-major axis, eccentricity, perigee argument, right ascension of the ascending node, and inclination. The adjustment strategy is as follows:

[0121] Assuming the satellite is equipped with only one thruster installed on its -X panel, and the satellite maintains a three-axis Earth-oriented attitude, the thrust generated by the thruster is entirely along the flight direction, and the magnitude of the thrust is F. t .

[0122] Among them, the combined corrections for the semi-major axis root, eccentricity root, and perigee argument root are... The trajectory change calculation includes determining the theoretical ignition duration, the actual ignition duration, and the latitude argument of the midpoint of the ignition arc segment. The theoretical ignition duration, the actual ignition duration, and the latitude argument of the midpoint of the ignition arc segment are expressed by the following formulas:

[0123] Theoretical ignition time:

[0124] Actual ignition time:

[0125] Latitude and angle of the midpoint of the ignition arc:

[0126] Then we can determine that the thrust direction is along the direction of the satellite's flight speed.

[0127] Optionally, for correcting the root of the semi-major axis The trajectory change calculation method includes determining the theoretical ignition duration and the actual ignition duration, expressed by the following formula:

[0128] Theoretical ignition time:

[0129] Actual ignition time: K can be taken as 0.5 to 0.8;

[0130] The first ignition state and the first ignition duration are determined, which can be expressed by the following formula: Determining the first ignition state may include determining the latitude argument of the midpoint of the ignition arc: u1=0°;

[0131] First ignition duration:

[0132] Then it can be determined that the thrust direction is along the direction of the satellite's flight speed.

[0133] And determine the second ignition state and the second ignition duration. Determining the second ignition state may include determining the latitude argument of the midpoint of the ignition arc: u2 = 180°.

[0134] Second ignition duration:

[0135] The thrust direction was determined to be along the direction of the satellite's flight speed.

[0136] Optionally, for the joint correction of the semi-major axis root and the eccentricity root... The trajectory change calculation method includes determining the theoretical ignition duration, the actual ignition duration, and the latitude argument of the midpoint of the ignition arc segment, which is expressed by the following formula:

[0137] Theoretical ignition time:

[0138] Actual ignition time:

[0139] Latitude and angle of the midpoint of the ignition arc:

[0140] Then the thrust direction was determined to be along the direction of the satellite's flight speed.

[0141] Optionally, for correcting the root of the dip angle The trajectory change calculation method includes determining the theoretical ignition duration, the actual ignition duration, and the latitude argument of the midpoint of the ignition arc segment, which can be expressed by the following formula:

[0142] Theoretical ignition time:

[0143] Actual ignition time:

[0144] Latitude and angle of the midpoint of the ignition arc: u mid =0°

[0145] The thrust direction can be determined based on the root of the relative inclination angle between the satellite and the virtual satellite relative to the target, as expressed by the following formula:

[0146] if The direction of the thrust is along the negative normal to the orbital plane;

[0147] if The direction of the thrust is along the normal to the orbital plane.

[0148] Optionally, for separately correcting the ascending node, right ascension, and root... The trajectory change calculation method can include determining the theoretical ignition duration, the actual ignition duration, and the latitude argument of the midpoint of the ignition arc segment, which can be expressed by the following formula:

[0149] Theoretical ignition time:

[0150] Actual ignition time:

[0151] Latitude and angle of the midpoint of the ignition arc: u mid =900

[0152] The thrust direction can be calculated using the right ascension root of the ascending node relative to the virtual satellite of the target, as shown by the following formula:

[0153] if The direction of the thrust is along the negative normal to the orbital plane;

[0154] if The direction of the thrust is along the normal to the orbital plane.

[0155] Furthermore, after calculating the specific ignition parameters based on the selected orbit control sub-mode, the orbit change strategy can be applied, and the satellite completes the orbit change maneuver from the 4th to the Nth orbit. Then, time t is updated to the end time of the Nth orbit change maneuver, and the process proceeds to steps 3 to 7 for iterative execution until the orbit change is marked as complete.

[0156] For example, assume that the root of the semi-major axis of the initial orbit is 7476 km, the root of the eccentricity is 0.004, the root of the inclination is 86.64°, the root of the right ascension of the ascending node is -0.15°, the root of the argument of perigee is 270°, and the root of the true perigee is 270°; the root of the semi-major axis of the target orbit is 7503 km, the root of the eccentricity is 0.0012, the root of the inclination is 86.5°, the root of the right ascension of the ascending node is 0°, the root of the argument of perigee is 90°, and the root of the true perigee is 270°.

[0157] Using the method described in this paper, the maneuver from the initial orbit to the target orbit can be completed within 70 days, with the inclination deviation correction completed on day 21, the orbit ascent completed on day 25, the frozen orbit established on day 49, and the right ascension deviation correction of the ascending node completed on day 70.

[0158] The orbit change strategy in this step involves minimal computation, allowing for iterative calculation of the orbit change strategy and collision risk analysis within a short timeframe. Furthermore, the orbit change strategy avoids algorithm divergence issues and is suitable for orbit changes from any initial orbit to a higher orbit. By optimizing the electric propulsion orbit change strategy into 11 sub-strategies (i.e., orbit control sub-modes), the corresponding sub-strategy can be quickly selected based on the current orbit change requirements. This facilitates rapid calculation and simulation verification of the electric propulsion orbit change parameters, avoiding the convergence issues of traditional direct and indirect algorithm optimizations, which are easily affected by initial values. Simultaneously, it meets the requirement of rapidly completing the electric propulsion orbit change strategy optimization calculation within a limited timeframe.

[0159] The satellite orbit change strategy planning method provided in this application initializes the initial time of the satellite and extrapolates the orbit period to update time t as the extrapolated time. The orbit change flag is initialized as incomplete. The first orbit parameters of the satellite at time t and the second orbit parameters of the target virtual satellite at time t are calculated. Based on the first and second orbit parameters, the relative orbit parameters of the satellite to the target virtual satellite are determined. Then, based on the relative orbit parameters and preset orbit thresholds, correction flags for each orbit parameter are determined. Based on each correction flag, the corresponding orbit control sub-mode is determined. When the orbit change flag is complete, the ignition parameters of the satellite are determined based on the orbit control sub-mode, and orbit change maneuvers and orbit change phase adjustment processing are performed. Compared with existing technologies, this technical solution accurately determines the relative orbital parameters through the first and second orbital parameters, enabling the use of a low-thrust orbital maneuver during satellite orbit change. It allows for the rapid selection of the corresponding orbit control sub-mode based on current orbit change requirements, and the quick generation and simulation verification of electric propulsion orbit change parameters. This ensures the satellite accurately enters the predetermined orbit position (assuming a target virtual satellite always operates at this position), avoiding the convergence problems of traditional direct optimization and indirect algorithm optimization which are easily affected by initial values. It also meets the requirement of rapidly completing the electric propulsion orbit change strategy optimization calculation within a limited time, improving the accuracy of satellite orbit change. Furthermore, it is applicable to orbit changes from any initial orbit to a higher orbit.

[0160] It should be noted that although the operations of the method of the present invention are described in a specific order in the accompanying drawings, this does not require or imply that these operations must be performed in that specific order, or that all of the operations shown must be performed to achieve the desired result. On the contrary, the steps depicted in the flowchart may be performed in a different order. Additionally or alternatively, certain steps may be omitted, multiple steps may be combined into one step, and / or one step may be broken down into multiple steps.

[0161] on the other hand, Figure 9 This is a schematic diagram of a satellite orbit change strategy planning device provided in an embodiment of this application. This device can be a component within a computer device, such as… Figure 9 As shown, the device 400 includes:

[0162] The acquisition module 410 is used to initialize the initial time of the satellite and extrapolate the orbital period to update time t to the extrapolated time; t>0;

[0163] The calculation module 420 is used to initialize the orbit change flag to incomplete, calculate the first orbit parameters of the satellite at time t and the second orbit parameters of the target virtual satellite at time t, and determine the relative orbit parameters of the satellite to the target virtual satellite based on the first orbit parameters and the second orbit parameters.

[0164] The determination module 430 is used to determine the correction flags for each track parameter based on the relative track parameters and the preset track threshold values, and to determine the corresponding track control sub-mode based on each track parameter correction flag.

[0165] The orbit change module 440 is used to determine the satellite's ignition parameters based on the orbit control sub-mode when the orbit change flag is completed, and to perform orbit change maneuvers and orbit change phase adjustment processing.

[0166] In one embodiment, the determining module 430 is specifically used for:

[0167] When the absolute value of the relative semi-major axis flat root is greater than the preset relative semi-major axis flat root, the semi-major axis flat root correction flag is set to 1; otherwise, the semi-major axis flat root revision flag is set to 0.

[0168] When the absolute value of the root of relative eccentricity is greater than the preset root of relative eccentricity, the root of eccentricity correction flag is set to 1; otherwise, the root of eccentricity correction flag is set to 0.

[0169] When the absolute value of the relative perigee argument root is greater than the preset relative perigee argument root, the perigee argument root correction flag is set to 1; otherwise, the perigee argument root correction flag is set to 0.

[0170] When the absolute value of the relative tilt angle root is greater than the preset relative tilt angle root, the tilt angle root correction flag is set to 1; otherwise, the tilt angle root correction flag is set to 0.

[0171] When the absolute value of the relative ascending node right ascension root is greater than the preset relative ascending node right ascension root, the ascending node right ascension root correction flag is set to 1; otherwise, the ascending node right ascension root correction flag is set to 0.

[0172] In one embodiment, the determining module 430 is further configured to:

[0173] When the eccentricity square root correction flag is 1 or the perigee argument square root correction flag is 1, the first correction flag is determined to be 1; otherwise, the first correction flag is determined to be 0.

[0174] When the first correction flag is 1 and the semi-major axis root correction flag is 1, the second correction flag is determined to be 1; otherwise, the second correction flag is determined to be 0.

[0175] In one embodiment, the determining module 430 is further configured to:

[0176] Determine if the second correction flag is 1;

[0177] If the second correction flag is 1, the corresponding track control sub-mode is determined based on the inclination angle root correction flag and the ascending node right ascension root correction flag.

[0178] If the second correction flag is not 1, determine the corresponding track control sub-mode based on the correction flags for the semi-major axis root, the inclination root, and the ascending node right ascension root.

[0179] In one embodiment, the determining module 430 is further configured to:

[0180] Determine if the correction flag for the flat root of the semi-major axis is 1.

[0181] If the correction flag for the semi-major axis root is 1, then the corresponding track control sub-mode is determined based on the correction flags for the inclination root and the right ascension root of the ascending node.

[0182] If the correction flag for the semi-major axis root is not 1, then the corresponding track control sub-mode is determined based on the first correction flag, the inclination root correction flag, and the ascending node right ascension root correction flag.

[0183] In one embodiment, the track-changing module 440 is specifically used for:

[0184] When the completion result of the orbit change indicator is "completed", calculate the relative latitude argument between the satellite and the target orbit;

[0185] Compare the absolute value of the relative latitude argument with the preset latitude argument;

[0186] When the absolute value of the relative latitude argument is not greater than the preset latitude argument, the phase adjustment time is adjusted using an optimized algorithm, and the orbit change flag is re-initialized to incomplete.

[0187] Based on the first and second orbital parameters, the corresponding orbital control sub-mode is determined again, and orbital maneuvering and orbital phase adjustment are performed until the absolute value of the relative latitude angle is not greater than the preset latitude angle.

[0188] In one embodiment,

[0189] The track control sub-mode includes at least one of the following:

[0190] The combined corrections for the semi-major axis root, eccentricity root, and perigee argument root, as well as the individual corrections for the tilt root, are performed alternately.

[0191] The combined corrections to the semi-major axis, eccentricity, and perigee argument, as well as the individual correction to the right ascension of the ascending node, are performed alternately.

[0192] Perform joint corrections to the semi-major axis root, eccentricity root, and perigee argument root;

[0193] Alternately perform individual correction of the semi-major axis root and individual correction of the camber root;

[0194] The corrections for the semi-major axis root and the ascending node right ascension root are performed alternately.

[0195] Perform separate correction of the semi-major axis root;

[0196] The combined correction for the root of eccentricity and the root of the perigee angle, as well as the individual correction for the root of tilt, are performed alternately.

[0197] The joint correction of the eccentricity square root and the perigee argument square root, as well as the individual correction of the ascending node right ascension square root, are performed alternately.

[0198] Perform joint corrections to the square root of eccentricity and the square root of the perigee angle;

[0199] Perform individual correction of the tilt angle to flatten the root;

[0200] Perform a separate correction of the ascending node, right ascension, and root.

[0201] Compared with the prior art, the satellite orbit change strategy planning device provided in this application accurately determines the relative orbit parameters through the first orbit parameters and the second orbit parameters. This allows for the use of a low-thrust orbit change method during satellite orbit change. Based on the current orbit change requirements, the device can quickly select the corresponding orbit control sub-mode, rapidly generate calculations and simulation verifications of electric propulsion orbit change parameters, and ensure that the satellite accurately enters the predetermined orbit position (assuming that a target virtual satellite always operates in this orbit position). This avoids the problem that the convergence of traditional direct optimization and indirect algorithm optimization is easily affected by the initial value and cannot converge. At the same time, it meets the requirement of quickly completing the electric propulsion orbit change strategy optimization calculation within a limited time, improves the accuracy of satellite orbit change, and is applicable to orbit change from any initial orbit to a higher orbit.

[0202] On the other hand, the computer device provided in the embodiments of this application includes a memory, a processor, and a computer program stored in the memory and executable on the processor. When the processor executes the program, it implements the satellite orbit change strategy planning method as described above.

[0203] The following is for reference. Figure 10 , Figure 10 This is a schematic diagram of the structure of a computer system according to an embodiment of this application.

[0204] like Figure 10 As shown, the computer system 600 includes a central processing unit (CPU) 601, which can perform various appropriate actions and processes based on programs stored in read-only memory (ROM) 602 or programs loaded from storage portion 603 into random access memory (RAM) 603. The RAM 603 also stores various programs and data required for the operation of the system 600. The CPU 601, ROM 602, and RAM 603 are interconnected via a bus 604. An input / output (I / O) interface 605 is also connected to the bus 604.

[0205] The following components are connected to I / O interface 605: an input section 606 including a keyboard, mouse, etc.; an output section 607 including a cathode ray tube (CRT), liquid crystal display (LCD), etc., and speakers, etc.; a storage section 608 including a hard disk, etc.; and a communication section 609 including a network interface card such as a LAN card, modem, etc. The communication section 609 performs communication processing via a network such as the Internet. A drive 610 is also connected to I / O interface 605 as needed. A removable medium 611, such as a disk, optical disk, magneto-optical disk, semiconductor memory, etc., is installed on drive 610 as needed so that computer programs read from it can be installed into storage section 608 as needed.

[0206] Specifically, according to embodiments of this application, the processes described above with reference to the flowcharts can be implemented as computer software programs. For example, embodiments of this application include a computer program product comprising a computer program carried on a machine-readable medium, the computer program containing program code for performing the methods shown in the flowcharts. In such embodiments, the computer program can be downloaded and installed from a network via communication section 603, and / or installed from removable medium 611. When the computer program is executed by central processing unit (CPU) 601, it performs the functions defined above in the system of this application.

[0207] It should be noted that the computer-readable medium shown in this application can be a computer-readable signal medium or a computer-readable storage medium, or any combination of the two. A computer-readable storage medium can be, for example,—but not limited to—an electrical, magnetic, optical, electromagnetic, infrared, or semiconductor system, apparatus, or device, or any combination thereof. More specific examples of a computer-readable storage medium may include, but are not limited to: an electrical connection having one or more wires, a portable computer disk, a hard disk, random access memory (RAM), read-only memory (ROM), erasable programmable read-only memory (EPROM or flash memory), optical fiber, portable compact disk read-only memory (CD-ROM), optical storage device, magnetic storage device, or any suitable combination thereof. In this application, a computer-readable storage medium can be any tangible medium containing or storing a program that can be used by or in conjunction with an instruction execution system, apparatus, or device. In this application, a computer-readable signal medium can include a data signal propagated in baseband or as part of a carrier wave, carrying computer-readable program code. Such propagated data signals can take various forms, including but not limited to electromagnetic signals, optical signals, or any suitable combination thereof. Computer-readable signal media can also be any computer-readable medium other than computer-readable storage media, which can send, propagate, or transmit a program for use by or in connection with an instruction execution system, apparatus, or device. The program code contained on the computer-readable medium can be transmitted using any suitable medium, including but not limited to: wireless, wire, optical fiber, RF, etc., or any suitable combination thereof.

[0208] The flowcharts and block diagrams in the accompanying drawings illustrate the architecture, functionality, and operation of possible implementations of systems, methods, and computer program products according to various embodiments of this application. In this regard, each block in a flowchart or block diagram may represent a module, segment, or portion of code containing one or more executable instructions for implementing a specified logical function. It should also be noted that in some alternative implementations, the functions indicated in the blocks may occur in a different order than those indicated in the drawings. For example, two consecutively indicated blocks may actually be executed substantially in parallel, and they may sometimes be executed in reverse order, depending on the functions involved. It should also be noted that each block in the block diagrams and / or flowcharts, and combinations of blocks in the block diagrams and / or flowcharts, can be implemented using a dedicated hardware-based system that performs the specified function or operation, or using a combination of dedicated hardware and computer instructions.

[0209] The units or modules described in the embodiments of this application can be implemented in software or hardware. The described units or modules can also be housed in a processor; for example, they can be described as: a processor including: an initialization module, a calculation module, a determination module, and an orbit-changing module. The names of these units or modules do not necessarily limit the specific unit or module itself. For example, the initialization module can also be described as "used to initialize the initial time of the satellite and extrapolate the orbital period to update time t to the extrapolated time; t>0;".

[0210] In another aspect, this application also provides a computer-readable storage medium, which may be included in the electronic device described in the above embodiments; or it may exist independently and not assembled into the electronic device. The aforementioned computer-readable storage medium stores one or more programs, which, when used by one or more processors, execute the satellite orbit change strategy planning method described in this application:

[0211] Initialize the satellite's initial time and extrapolate its orbital period to update time t to the extrapolated time; t>0;

[0212] The initial orbit change flag is set to incomplete. The first orbital parameters of the satellite at time t and the second orbital parameters of the target virtual satellite at time t are calculated. Based on the first orbital parameters and the second orbital parameters, the relative orbital parameters of the satellite to the target virtual satellite are determined.

[0213] Based on the relative track parameters and the preset track threshold, a correction flag for each track parameter is determined, and based on the correction flag for each track parameter, the corresponding track control sub-mode is determined.

[0214] When the orbit change flag is completed, the ignition parameters of the satellite are determined based on the orbit control sub-mode, and orbit change maneuvers and orbit change phase adjustment processes are performed.

[0215] In summary, the satellite orbit change strategy planning method, apparatus, device, and storage medium provided in this application initialize the initial time of the satellite and extrapolate the orbit period to update time t as the extrapolated time. The orbit change flag is initialized as incomplete. The first orbit parameters of the satellite at time t and the second orbit parameters of the target virtual satellite at time t are calculated. Based on the first and second orbit parameters, the relative orbit parameters of the satellite to the target virtual satellite are determined. Then, based on the relative orbit parameters and preset orbit thresholds, correction flags for each orbit parameter are determined. Based on each correction flag, the corresponding orbit control sub-mode is determined. When the orbit change flag is complete, the ignition parameters of the satellite are determined based on the orbit control sub-mode, and orbit change maneuvers and orbit change phase adjustment processing are performed. Compared with existing technologies, this technical solution accurately determines the relative orbital parameters through the first and second orbital parameters, enabling the use of a low-thrust orbital maneuver during satellite orbit change. It allows for the rapid selection of the corresponding orbit control sub-mode based on current orbit change requirements, and the quick generation and simulation verification of electric propulsion orbit change parameters. This ensures the satellite accurately enters the predetermined orbit position (assuming a target virtual satellite always operates at this position), avoiding the convergence problems of traditional direct optimization and indirect algorithm optimization which are easily affected by initial values. It also meets the requirement of rapidly completing the electric propulsion orbit change strategy optimization calculation within a limited time, improving the accuracy of satellite orbit change. Furthermore, it is applicable to orbit changes from any initial orbit to a higher orbit.

[0216] The above description is merely a preferred embodiment of this application and an explanation of the technical principles employed. Those skilled in the art should understand that the scope of the invention involved in this application is not limited to technical solutions formed by specific combinations of the above-described technical features, but should also cover other technical solutions formed by arbitrary combinations of the above-described technical features or their equivalents without departing from the inventive concept. For example, technical solutions formed by substituting the above features with (but not limited to) technical features with similar functions disclosed in this application.

Claims

1. A satellite orbit change strategy planning method, characterized in that, The method includes: Initialize the satellite's initial time and extrapolate its orbital period to update time t to the extrapolated time; t>0; The initial orbit change flag is set to incomplete. The first orbital parameters of the satellite at time t and the second orbital parameters of the target virtual satellite at time t are calculated. Based on the first orbital parameters and the second orbital parameters, the relative orbital parameters of the satellite to the target virtual satellite are determined. Based on the relative track parameters and the preset track threshold, determine the correction flags for each track parameter, and based on the correction flags for each track parameter, determine the completion results of the corresponding track control sub-mode and track change flag. Based on the completion result of the track change indicator and the track control sub-mode, track change maneuver and track change phase adjustment processing are performed; The relative orbital parameters include: relative semi-major axis root, relative eccentricity root, relative inclination root, relative ascending node right ascension root, relative perigee argument root, relative latitude argument root, and relative eccentricity vector root. Based on the relative orbit parameters and preset orbit threshold values, determine the correction flags for each orbit parameter, including: When the absolute value of the relative semi-major axis flat root is greater than the preset relative semi-major axis flat root, the semi-major axis flat root correction flag is determined to be 1; otherwise, the semi-major axis flat root revision flag is determined to be 0. When the absolute value of the relative eccentricity root is greater than the preset relative eccentricity root, the eccentricity root correction flag is determined to be 1; otherwise, the eccentricity root correction flag is determined to be 0. When the absolute value of the relative perigee argument root is greater than the preset relative perigee argument root, the perigee argument root correction flag is set to 1; otherwise, the perigee argument root correction flag is set to 0. When the absolute value of the relative tilt angle root is greater than the preset relative tilt angle root, the tilt angle root correction flag is determined to be 1; otherwise, the tilt angle root correction flag is determined to be 0. When the absolute value of the relative ascending node right ascension root is greater than the preset relative ascending node right ascension root, the ascending node right ascension root correction flag is set to 1; otherwise, the ascending node right ascension root correction flag is set to 0.

2. The method according to claim 1, characterized in that, The orbital parameter correction flags further include a first correction flag and a second correction flag, and the method further includes: When the eccentricity square root correction flag is 1 or the perigee argument square root correction flag is 1, the first correction flag is determined to be 1; otherwise, the first correction flag is determined to be 0. When the first correction flag is 1 and the semi-major axis flat root correction flag is 1, the second correction flag is determined to be 1; otherwise, the second correction flag is determined to be 0.

3. The method according to claim 2, characterized in that, Based on the aforementioned track parameter correction flags, the corresponding track control sub-mode is determined, including: Determine whether the second correction flag is 1; If the second correction flag is 1, the corresponding track control sub-mode is determined according to the inclination angle root correction flag and the ascending node right ascension root correction flag. If the second correction flag is not 1, the corresponding track control sub-mode is determined according to the correction flag of the semi-major axis root, the correction flag of the inclination root, and the correction flag of the ascending node right ascension root.

4. The method according to claim 3, characterized in that, Based on the revision flag of the semi-major axis root, the correction flag of the inclination root, and the correction flag of the right ascension root of the ascending node, the corresponding track control sub-mode is determined, including: Determine whether the correction flag for the flat root of the semi-major axis is 1; If the correction flag for the semi-major axis flat root is 1, then the corresponding track control sub-mode is determined based on the correction flag for the inclination angle flat root and the correction flag for the right ascension of the ascending node. If the correction flag for the semi-major axis flat root is not 1, then the corresponding track control sub-mode is determined based on the first correction flag, the inclination angle flat root correction flag, and the ascending node right ascension flat root correction flag.

5. The method according to claim 1, characterized in that, Based on the completion result of the track change indicator and the track control sub-mode, track change maneuvering and track change phase adjustment processing are performed, including: When the completion result of the orbit change indicator is "completed", calculate the relative latitudinal argument between the satellite and the target orbit; The absolute value of the relative latitude argument is compared with the preset latitude argument; When the absolute value of the relative latitude argument is not greater than the preset latitude argument, the phase adjustment time is adjusted using an optimization algorithm, and the orbit change flag is re-initialized to incomplete. Based on the first and second orbit parameters, the corresponding orbit control sub-mode is determined again, and orbit change maneuvers and orbit change phase adjustment processes are performed until the absolute value of the relative latitude angle is not greater than the preset latitude angle.

6. The method according to claim 5, characterized in that, The track control sub-mode includes at least one of the following: The combined corrections for the semi-major axis root, eccentricity root, and perigee argument root, as well as the individual corrections for the tilt root, are performed alternately. The combined corrections to the semi-major axis, eccentricity, and perigee argument, as well as the individual correction to the right ascension of the ascending node, are performed alternately. Perform joint corrections to the semi-major axis root, eccentricity root, and perigee argument root; Alternately perform individual correction of the semi-major axis root and individual correction of the camber root; The corrections for the semi-major axis root and the ascending node right ascension root are performed alternately. Perform separate correction of the semi-major axis root; The combined correction for the root of eccentricity and the root of the perigee angle, as well as the individual correction for the root of tilt, are performed alternately. The joint correction of the eccentricity square root and the perigee argument square root, as well as the individual correction of the ascending node right ascension square root, are performed alternately. Perform joint corrections to the square root of eccentricity and the square root of the perigee angle; Perform individual correction of the tilt angle to flatten the root; Perform a separate correction of the ascending node, right ascension, and root.

7. A satellite orbit change strategy planning device, characterized in that, The device includes: The initialization module is used to initialize the satellite's initial time and extrapolate its orbital period to update time t to the extrapolated time; t>0; The calculation module is used to initialize the orbit change flag to incomplete, calculate the first orbit parameters of the satellite at time t and the second orbit parameters of the target virtual satellite at time t, and determine the relative orbit parameters of the satellite to the target virtual satellite based on the first orbit parameters and the second orbit parameters. The determination module is used to determine the correction flags for each track parameter based on the relative track parameters and the preset track threshold values, and to determine the corresponding track control sub-mode based on the correction flags for each track parameter. The orbit change module is used to determine the ignition parameters of the satellite based on the orbit control sub-mode when the orbit change flag is completed, and to perform orbit change maneuvers and orbit change phase adjustment processing. The relative orbital parameters include: relative semi-major axis root, relative eccentricity root, relative inclination root, relative ascending node right ascension root, relative perigee argument root, relative latitude argument root, and relative eccentricity vector root. The determining module is used for: When the absolute value of the relative semi-major axis flat root is greater than the preset relative semi-major axis flat root, the semi-major axis flat root correction flag is determined to be 1; otherwise, the semi-major axis flat root revision flag is determined to be 0. When the absolute value of the relative eccentricity root is greater than the preset relative eccentricity root, the eccentricity root correction flag is determined to be 1; otherwise, the eccentricity root correction flag is determined to be 0. When the absolute value of the relative perigee argument root is greater than the preset relative perigee argument root, the perigee argument root correction flag is set to 1; otherwise, the perigee argument root correction flag is set to 0. When the absolute value of the relative tilt angle root is greater than the preset relative tilt angle root, the tilt angle root correction flag is determined to be 1; otherwise, the tilt angle root correction flag is determined to be 0. When the absolute value of the relative ascending node right ascension root is greater than the preset relative ascending node right ascension root, the ascending node right ascension root correction flag is set to 1; otherwise, the ascending node right ascension root correction flag is set to 0.

8. A computer device comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, characterized in that, When the processor executes the computer program, it implements the satellite orbit change strategy planning method as described in any one of claims 1-6.

9. A computer-readable storage medium having a computer program stored thereon, the computer program being used to implement the satellite orbit change strategy planning method as described in any one of claims 1-6.

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