GEO Satellite Hybrid Propulsion Transfer Phase Orbit Change Method and Device

Abstract

This invention relates to a method and apparatus for orbit change during the transfer phase of a GEO satellite using hybrid chemical-electric propulsion. The method includes: dividing the transfer phase of the GEO satellite into a first transfer phase and a second transfer phase; the initial state of the first transfer phase is the state at the moment of satellite-launcher separation, and the final state is determined based on the total propellant of the chemical thrusters and the propellant consumption during the synchronization phase; the initial state of the second transfer phase is the final state of the first transfer phase, and the final state is the target orbit; the first transfer phase uses chemical propulsion for orbit change: the orbit change process of the first transfer phase is divided into multiple transition orbit changes, and for each transition orbit change, the orbit control target value is determined based on preset telemetry and control constraints and the double-pulse assumption to complete the orbit change of the first transfer phase; the second transfer phase uses electric propulsion for orbit change: the thrust direction of the electric thrusters is adjusted by adjusting the in-plane control angle and the out-of-plane control angle to complete the orbit change of the second transfer phase. This scheme is computationally simple and can quickly complete satellite orbit change.

Description

Technical Field

[0001] This invention relates to the field of satellite orbit control technology, and in particular to a method and apparatus for orbit change during the transfer phase of a GEO satellite with hybrid chemical and electric propulsion. Background Technology

[0002] When GEO satellites perform orbit changes, they typically use the 490N engine for the transfer phase, a process that is fast and accurate. However, if the 490N engine fails, other orbit-changing methods must be employed, such as using the chemical or electric propulsion systems typically found on communication satellites.

[0003] Currently, while related technologies can complete orbital maneuvers using chemical or electric propulsion, the transfer time is long and the computational load is large, which is not conducive to autonomous implementation on the satellite.

[0004] Therefore, there is an urgent need for a method and device for orbit change during the transfer phase of GEO satellites using hybrid propulsion of chemical and electric propulsion to solve the above-mentioned technical problems. Summary of the Invention

[0005] This invention provides a method and apparatus for orbit change during the transfer phase of a GEO satellite using hybrid propulsion (chemical and electric). The calculations are simple and the satellite orbit change can be completed quickly.

[0006] In a first aspect, embodiments of the present invention provide a method for orbit change during the transfer phase of a GEO satellite using hybrid chemical and electric propulsion, comprising:

[0007] The transfer phase of the GEO satellite is divided into a first transfer phase and a second transfer phase. The initial state of the first transfer phase is the state at the moment of satellite-launch separation, and the final state of the first transfer phase is determined based on the total propellant of the chemical thrusters and the propellant consumption during the synchronization phase. The initial state of the second transfer phase is the final state of the first transfer phase, and the final state of the second transfer phase is the target orbit.

[0008] For the first transfer segment, chemical propulsion is used for orbit change: the orbit change process of the first transfer segment is divided into multiple transition orbit changes. For each transition orbit change, the orbit control target value is determined based on the preset measurement and control constraints and the double pulse assumption to complete the orbit change of the first transfer segment. The orbit control target value includes the semi-major axis of the track, the track inclination, the velocity increment, the number of ignition cycles, and the geographical longitude of the ignition point.

[0009] For the second transfer segment, electric propulsion is used for trajectory change: the thrust direction of the electric propulsion is adjusted by adjusting the in-plane control angle and the out-of-plane control angle to complete the trajectory change of the second transfer segment.

[0010] Secondly, embodiments of the present invention also provide a transfer segment orbit-changing device for GEO satellites with hybrid chemical and electric propulsion, comprising:

[0011] The division unit is used to divide the transfer segment of the GEO satellite into a first transfer segment and a second transfer segment; the initial state of the first transfer segment is the state at the moment of satellite-rocket separation, and the final state of the first transfer segment is determined based on the total propellant of the chemical thrusters and the propellant consumption during the synchronization segment; the initial state of the second transfer segment is the final state of the first transfer segment, and the final state of the second transfer segment is the target orbit;

[0012] The chemical propulsion orbit change unit is used to perform orbit change using chemical propulsion for the first transfer segment: the orbit change process of the first transfer segment is divided into multiple transition orbit changes. For each transition orbit change, the orbit control target value is determined based on preset measurement and control constraints and the double pulse assumption to complete the orbit change of the first transfer segment. The orbit control target value includes the semi-major axis of the track, the track inclination angle, the velocity increment, the number of ignition cycles, and the geographical longitude of the ignition point.

[0013] The electric propulsion trajectory-changing unit is used to perform trajectory-changing for the second transfer section using electric propulsion: by adjusting the in-plane control angle and out-of-plane control angle, the thrust direction of the electric propulsion is adjusted to complete the trajectory-changing for the second transfer section.

[0014] Thirdly, embodiments of the present invention also provide an electronic device, including a memory and a processor, wherein the memory stores a computer program, and when the processor executes the computer program, it implements the method described in any embodiment of this specification.

[0015] Fourthly, embodiments of the present invention also provide a computer-readable storage medium having a computer program stored thereon, which, when executed in a computer, causes the computer to perform the methods described in any embodiment of this specification.

[0016] This invention provides a method and apparatus for orbit change during the transfer phase of a GEO satellite using hybrid chemical-electric propulsion. By dividing the transfer phase into a first transfer phase and a second transfer phase, chemical propulsion and electric propulsion are processed separately, simplifying calculations and facilitating engineering implementation. Furthermore, for each transition orbit change using chemical propulsion, solutions are performed based on preset constraints and the bipulse assumption. This transforms the solution process of complex dynamic models into a geometric problem, decomposing a multi-step programming problem into multiple single-step optimization problems. This not only optimizes the results but also reduces computational load and increases orbit change speed. Therefore, this solution is computationally simple and can quickly complete satellite orbit changes. Attached Figure Description

[0017] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0018] Figure 1 This is a schematic diagram of the transfer segment orbit change method for GEO satellites with hybrid chemical and electric propulsion provided in an embodiment of the present invention;

[0019] Figure 2 This is a hardware architecture diagram of an electronic device provided in an embodiment of the present invention;

[0020] Figure 3 This is a structural diagram of a transfer segment orbit-changing device for a GEO satellite with hybrid electric propulsion, provided in an embodiment of the present invention. Detailed Implementation

[0021] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are some embodiments of the present invention, but not all embodiments. All other embodiments obtained by those skilled in the art based on the embodiments of the present invention without creative effort are within the scope of protection of the present invention.

[0022] Please refer to Figure 1 This invention provides a method for orbit change during the transfer phase of a GEO satellite using hybrid chemical and electric propulsion, the method comprising:

[0023] Step 100: Divide the transfer phase of the GEO satellite into a first transfer phase and a second transfer phase; the initial state of the first transfer phase is the state at the moment of satellite-rocket separation, and the final state of the first transfer phase is determined based on the total propellant of the chemical thrusters and the propellant consumption during the synchronization phase; the initial state of the second transfer phase is the final state of the first transfer phase, and the final state of the second transfer phase is the target orbit.

[0024] Step 102: For the first transfer segment, chemical propulsion is used for orbit change: The orbit change process of the first transfer segment is divided into multiple transition orbit changes. For each transition orbit change, the orbit control target value is determined based on the preset measurement and control constraints and the double pulse assumption to complete the orbit change of the first transfer segment. The orbit control target value includes the semi-major axis of the track, the track inclination angle, the velocity increment, the number of ignition cycles, and the geographical longitude of the ignition point.

[0025] Step 104: For the second transfer segment, electric propulsion is used to change the trajectory: the thrust direction of the electric propulsion is adjusted by adjusting the control angle inside and outside the track plane to complete the trajectory change of the second transfer segment.

[0026] In this embodiment, by dividing the transfer segment into a first transfer segment and a second transfer segment, chemical propulsion and electric propulsion are treated separately, simplifying the calculations and making it more conducive to engineering implementation. Furthermore, for each transition orbit change using chemical propulsion, solutions are performed based on preset constraints and the bipulse assumption. This transforms the solution process of the complex dynamic model into a geometric problem, breaking down a multi-step programming problem into multiple single-step optimization problems. This not only optimizes the results but also reduces the computational load and increases the orbit change speed. Therefore, this scheme is computationally simple and can quickly complete satellite orbit changes.

[0027] The following description Figure 1 The execution method for each step is shown.

[0028] First, for step 100, the transfer segment of the GEO satellite is divided into the first transfer segment and the second transfer segment.

[0029] In this step, the initial state of the first transfer segment is the state at the moment of satellite-rocket separation, i.e., the satellite's initial orbital semi-major axis, initial orbital inclination, initial velocity, and initial geographic longitude at the moment of separation. The final state of the first transfer segment is determined by using all remaining propellant after reserving chemical propellant for the synchronization segment, thus obtaining the orbital transfer range for chemical propulsion and defining the final state of the first transfer segment. Therefore, the orbital semi-major axis, orbital inclination, and satellite velocity at the end of the first transfer segment can be calculated based on the amount of propellant used. The final state of the second segment is the target orbit, i.e., the state where the satellite has completed its orbital maneuver and is orbiting in a circular orbit.

[0030] By dividing the transfer stage into a first transfer stage and a second transfer stage, chemical propulsion and electric propulsion are treated separately, simplifying calculations and making them more conducive to engineering implementation.

[0031] Then, for step 102, for the first transfer stage, chemical propulsion is used to change course.

[0032] In some implementations, the target orbit control value for each transition orbit change is determined based on preset measurement and control constraints and the bipulse assumption, including:

[0033] Predetermine the semi-major axis step size and the orbital inclination step size;

[0034] For each transition track change, execute:

[0035] The final state of the first transfer segment is taken as the final state of the current transition trajectory change. Based on the initial and final states of the current transition trajectory change, the semi-major axis step size and the track inclination step size are used as the step size of the exhaustive method to assume that the current transition trajectory change is multiple double-pulse schemes. For each double-pulse scheme, it is determined whether it meets the preset constraint conditions. If it does, it is determined as a scheme to be confirmed; if not, it is removed. The total speed increment of each scheme to be confirmed and the speed increments of the first and second pulses in the scheme to be confirmed are calculated. The scheme to be confirmed with the smallest total speed increment is determined as the target scheme. The track semi-major axis, track inclination, speed increment, number of ignition cycles, and geographical longitude of the ignition point corresponding to the first pulse in the target scheme are taken as the track control target values ​​for the current transition trajectory change.

[0036] The semi-major axis and inclination of the current transition orbit change are used as the initial state for the next transition orbit change, and the next transition orbit change is performed until the last transition orbit change is reached; where, for the first transition orbit change, its initial state is the state at the moment of separation of the star and the rocket.

[0037] For the final transition orbit change, the orbital control target values ​​for the second pulse of the previous transition orbit change are the orbital semi-major axis, orbital inclination, velocity increment, number of ignition cycles, and geographical longitude of the ignition point.

[0038] In this step, during the transition orbit change, the semi-major axis and inclination angle of the orbit are used as independent variables. The step size for the semi-major axis and the step size for the inclination angle can be 10 km and 0.1°, respectively. Then, using these step sizes as the step size for the exhaustive method, each transition orbit change is assumed to be one of multiple two-pulse schemes. For example, if the initial state of a certain transition orbit change is at point A and the final state is at point E, with step sizes of 10 km and 0.1°, points B, C, and D can be determined. The semi-major axes of the orbits at points A, B, C, D, and E are 800, 810, 820, 830, and 840 km, respectively, and the orbit inclination angles are 80°, 70.9°, 70.8°, 70.7°, and 70.6°, respectively. The resulting two-pulse schemes include three options: ABE, ACE, and ADE. Then, based on the aforementioned preset constraints and calculation process, the target scheme is determined from these three schemes.

[0039] In some embodiments, the preset measurement and control constraints include a first constraint and a second constraint; wherein,

[0040] The first constraint is: in the double-pulse scheme, the change in the semi-major axis and the change in the tilt angle of the first pulse are less than the preset maximum change in the semi-major axis and the maximum change in the tilt angle, respectively, and the velocity increment of the first pulse is not less than the preset minimum velocity increment and not greater than the preset maximum velocity increment.

[0041] The second constraint is that the geographical longitude of the ignition point corresponding to the first pulse and the geographical longitude of the ignition point corresponding to the second pulse in the dual-pulse scheme are both within the preset measurement and control range.

[0042] For each of the two-pulse schemes, it is determined whether it meets the preset constraints, including:

[0043] Determine whether the dual-pulse scheme satisfies the first constraint condition. If not, remove the dual-pulse scheme. If yes, calculate the geographical longitude of the ignition point of the first pulse and the geographical longitude of the ignition point of the second pulse in the dual-pulse scheme, and determine whether the calculated geographical longitude of the ignition point satisfies the second constraint condition. If yes, determine the dual-pulse scheme as a scheme to be confirmed. If not, adjust the corresponding number of ignition cycles and the measurement and control longitude range until the dual-pulse scheme satisfies the second constraint condition.

[0044] In this step, for the first constraint, the preset maximum semi-major axis change, maximum tilt angle change, and maximum speed increment are determined by setting the ignition duration. The ignition duration is determined according to actual needs and is not specifically limited here, while limiting the minimum speed increment can ensure that the dual-pulse process has a solution.

[0045] The second constraint is designed to accommodate different fixed-point longitudes, i.e., the position the satellite should reach after the orbit change. This scheme can ensure that the geographical longitude of the ignition point of each pulse meets the second constraint by adjusting the telemetry and control constraints for each ignition or adjusting the number of ignition cycles, thus adapting to different fixed-point longitudes.

[0046] Specifically, the geographical longitude of the ignition point for each pulse is determined in the following way:

[0047] The orbital drift rate corresponding to the current sub-pulse is calculated using the following formula:

[0048]

[0049] The geographical longitude of the ignition point of the current subpulse is calculated based on the orbital drift rate, using the following formula:

[0050] λ i =λ i-1 +D i-1 NCircle i-1

[0051] Where, when i = 1, λ1 = λ0 + D0NCircle0

[0052] In the formula, D i Let a be the orbital offset rate of the i-th pulse. i Let a be the semi-major axis of the orbit of the i-th pulse. s= 42164.2 km, μ is the gravitational constant, λ i Let λ be the geographical longitude of the ignition point of the i-th pulse. i-1 Let NCircle be the geographical longitude of the ignition point of the (i-1)th pulse. i-1 λ0 is the number of ignition cycles for the (i-1)th pulse; when i = 1, λ0 is the geographical longitude of the satellite's first arrival at the ascending node, D0 is the drift rate corresponding to the semi-major axis of the initial orbit, and NCircle0 is the number of ignition cycles set from the geographical longitude λ0 of the first arrival at the ascending node to the first ignition; i = 1, 2, ..., Nfire, where Nfire is the number of transition orbit changes.

[0053] It should be noted that i in the above formula refers to the number of pulses in the entire first transfer stage track change process. For example, for the first track change in a certain transition track change, it may be the 5th time in the total number of pulses, then i = 5; the number of ignition cycles for each pulse can be adjusted according to actual needs.

[0054] Furthermore, in some implementations, the total velocity increment for each scheme to be verified, as well as the velocity increments of the first and second pulses within that scheme, are determined as follows:

[0055] The satellite velocities before, after, before, and after the first pulse are determined based on the vitality equation. The calculation formula for the vitality equation is as follows:

[0056]

[0057] In the formula, V is the satellite velocity, μ is the gravitational constant, a is the orbital radius, and r is the satellite radius vector.

[0058] Based on the satellite velocities before and after the first pulse, and the corresponding orbital inclinations before and after the first pulse, the first velocity increment of the first pulse is determined.

[0059] The formula for calculating the first velocity increment dV1 is:

[0060]

[0061] Based on the satellite velocities before and after the second pulse, and the orbital inclinations before and after the first pulse, the second velocity increment of the second pulse is determined; whereby...

[0062] The formula for calculating the second velocity increment dV2 is:

[0063]

[0064] The total velocity increment of the dual-pulse scheme is determined based on the first velocity increment and the second velocity increment. The formula for calculating the total velocity increment dV of the dual-pulse scheme is as follows:

[0065] dV = dV1 + dV2

[0066] In the above formula, V 10 V 1f V 20 and V 2f These are the satellite velocities before the first pulse, after the first pulse, before the second pulse, and after the second pulse, respectively; Incl 10 Incl 1f Incl 20 and Incl 2f These are the orbital inclination angles before the first pulse, after the first pulse, before the second pulse, and after the second pulse, respectively.

[0067] The velocity increment for each transition maneuver is calculated using the method described above. The sum of these velocity increments gives the total velocity increment required for the entire first transfer phase. A smaller total velocity increment results in less propellant consumption. This scheme conserves propellant during each transition maneuver, thereby increasing the number of chemical propulsion maneuvers and accelerating the maneuver speed.

[0068] After the chemical propulsion orbit change process is completed, the final state of chemical propulsion (i.e., the orbital elements corresponding to the satellite) is taken as the initial state for the start of the electric propulsion phase, and the target orbit is taken as the final state of the electric propulsion phase.

[0069] This scheme decomposes the thrust of the electric thruster into two components: in-plane and orbital normal. The thrust direction is adjusted by changing the in-plane control angle and the out-of-plane control angle. The specific process is as follows:

[0070] Within the preset apogee range, adjust the in-plane control angle using the following formula:

[0071]

[0072] Within the preset track lifting and lowering point range, adjust the track out-of-plane control angle using the following formula:

[0073]

[0074] u′=W+θ

[0075] In the formula, α is the in-plane control angle, θ is the true anomaly angle, the value of θ is determined based on mission time and thruster requirements, e is the eccentricity, β is the out-of-plane control angle, u′ is the intermediate value, and w is the perigee argument.

[0076] In this step, the perigee altitude can be increased by adjusting the in-plane control angle, and the track inclination angle can be decreased by adjusting the out-of-plane control angle. The preferred θ value is [0, 2π], which allows for operation across the entire arc, thereby reducing the track change transfer time.

[0077] like Figure 2 , Figure 3 As shown, this embodiment of the invention provides a transfer phase orbit-changing device for GEO satellites using hybrid electric-chemical propulsion. The device embodiment can be implemented through software, hardware, or a combination of both. From a hardware perspective, as... Figure 2 The diagram shown is a hardware architecture diagram of an electronic device for a GEO satellite's hybrid propulsion transfer phase orbit change device, provided in an embodiment of the present invention. Besides... Figure 2 In addition to the processor, memory, network interface, and non-volatile memory shown, the electronic device in the embodiment may also include other hardware, such as a forwarding chip responsible for processing packets. Taking software implementation as an example, such as... Figure 3 As shown, a device in a logical sense is formed by the CPU of the electronic device in which it is located reading the corresponding computer program from the non-volatile memory into the memory for execution.

[0078] This embodiment provides a transfer segment orbit-changing device for a GEO satellite with hybrid chemical and electric propulsion, comprising:

[0079] The dividing unit 300 is used to divide the transfer segment of the GEO satellite into a first transfer segment and a second transfer segment; the initial state of the first transfer segment is the state at the moment of satellite-rocket separation, and the final state of the first transfer segment is determined based on the total propellant of the chemical thruster and the propellant consumption during the synchronization segment; the initial state of the second transfer segment is the final state of the first transfer segment, and the final state of the second transfer segment is the target orbit;

[0080] The chemical propulsion orbit change unit 302 is used to perform orbit change using chemical propulsion for the first transfer segment: the orbit change process of the first transfer segment is divided into multiple transition orbit changes. For each transition orbit change, the orbit control target value is determined based on preset measurement and control constraints and the double pulse assumption to complete the orbit change of the first transfer segment. The orbit control target value includes the semi-major axis of the track, the track inclination angle, the velocity increment, the number of ignition cycles, and the geographical longitude of the ignition point.

[0081] The electric propulsion trajectory-changing unit 304 is used to perform trajectory changing for the second transfer section using electric propulsion: by adjusting the control angle in the track plane and the control angle outside the track plane, the thrust direction of the electric propulsion is adjusted to complete the trajectory changing for the second transfer section.

[0082] It is understood that the structures illustrated in the embodiments of the present invention do not constitute a specific limitation on a transfer phase orbit-changing device for a GEO satellite with hybrid electric-chemical propulsion. In other embodiments of the present invention, a transfer phase orbit-changing device for a GEO satellite with hybrid electric-chemical propulsion may include more or fewer components than illustrated, or combine some components, or split some components, or have different component arrangements. The illustrated components may be implemented in hardware, software, or a combination of software and hardware.

[0083] The information interaction and execution process between the modules in the above-mentioned device are based on the same concept as the method embodiment of the present invention, and the specific details can be found in the description of the method embodiment of the present invention, and will not be repeated here.

[0084] This invention also provides an electronic device, including a memory and a processor. The memory stores a computer program, and when the processor executes the computer program, it implements a transfer segment orbit change method for GEO satellite chemical-electric hybrid propulsion according to any embodiment of this invention.

[0085] This invention also provides a computer-readable storage medium storing a computer program. When executed by a processor, the computer program causes the processor to perform a transfer segment orbit change method for GEO satellite chemical-electric hybrid propulsion according to any embodiment of this invention.

[0086] Specifically, a system or apparatus equipped with a storage medium may be provided, on which software program code implementing the functions of any of the embodiments described above is stored, and the computer (or CPU or MPU) of the system or apparatus may read and execute the program code stored in the storage medium.

[0087] In this case, the program code read from the storage medium can itself implement the function of any of the above embodiments, and therefore the program code and the storage medium storing the program code constitute part of the present invention.

[0088] Examples of storage media used to provide program code include floppy disks, hard disks, magneto-optical disks, optical disks (such as CD-ROM, CD-R, CD-RW, DVD-ROM, DVD-RAM, DVD-RW, DVD+RW), magnetic tapes, non-volatile memory cards, and ROMs. Alternatively, program code can be downloaded from a server computer via a communication network.

[0089] Furthermore, it should be clear that not only can the program code read by the computer be executed, but also the operating system or other components operating on the computer can be instructed based on the program code to perform some or all of the actual operations, thereby realizing the function of any of the embodiments described above.

[0090] Furthermore, it is understood that the program code read from the storage medium is written to the memory set in the expansion board inserted into the computer or to the memory set in the expansion module connected to the computer. Then, based on the instructions of the program code, the CPU or other components installed on the expansion board or expansion module execute some and all of the actual operations, thereby realizing the function of any of the above embodiments.

[0091] It should be noted that, in this document, relational terms such as "first" and "second" are used only to distinguish one entity or operation from another, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Furthermore, the terms "comprising," "including," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such a process, method, article, or apparatus. Without further limitations, an element defined by the phrase "comprising one..." does not exclude the presence of other identical elements in the process, method, article, or apparatus that includes said element.

[0092] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, and not to limit them; although the present invention has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that modifications can still be made to the technical solutions described in the foregoing embodiments, or equivalent substitutions can be made to some of the technical features; and these modifications or substitutions do not cause the essence of the corresponding technical solutions to deviate from the spirit and scope of the technical solutions of the embodiments of the present invention.

Claims

1. A method for orbit change during the transfer phase of a GEO satellite using hybrid electric propulsion, characterized in that, include: The transfer phase of the GEO satellite is divided into a first transfer phase and a second transfer phase. The initial state of the first transfer phase is the state at the moment of satellite-launch separation, and the final state of the first transfer phase is determined based on the total propellant of the chemical thrusters and the propellant consumption during the synchronization phase. The initial state of the second transfer phase is the final state of the first transfer phase, and the final state of the second transfer phase is the target orbit. For the first transfer segment, chemical propulsion is used for orbit change: the orbit change process of the first transfer segment is divided into multiple transition orbit changes. For each transition orbit change, the orbit control target value is determined based on the preset measurement and control constraints and the double pulse assumption to complete the orbit change of the first transfer segment. The orbit control target value includes the semi-major axis of the track, the track inclination, the velocity increment, the number of ignition cycles, and the geographical longitude of the ignition point. For the second transfer segment, electric propulsion is used for trajectory change: the thrust direction of the electric propulsion is adjusted by adjusting the in-plane control angle and the out-of-plane control angle to complete the trajectory change of the second transfer segment; The determination of the track control target value for each transition track change based on preset measurement and control constraints and the bipulse assumption includes: Predetermine the semi-major axis step size and the orbital inclination step size; For each transition track change, execute: The final state of the first transfer segment is taken as the final state of the current transition trajectory change. Based on the initial and final states of the current transition trajectory change, the semi-major axis step size and the track inclination step size are used as the step size of the exhaustive method to assume that the current transition trajectory change is multiple double-pulse schemes. For each double-pulse scheme, it is determined whether it meets the preset constraint conditions. If it does, it is determined as a scheme to be confirmed; if not, it is removed. The total speed increment of each scheme to be confirmed and the speed increments of the first and second pulses in the scheme to be confirmed are calculated. The scheme to be confirmed with the smallest total speed increment is determined as the target scheme. The track semi-major axis, track inclination, speed increment, number of ignition cycles, and geographical longitude of the ignition point corresponding to the first pulse in the target scheme are taken as the track control target values ​​for the current transition trajectory change. The semi-major axis and inclination of the current transition orbit change are used as the initial state for the next transition orbit change, and the next transition orbit change is performed until the last transition orbit change is reached; where, for the first transition orbit change, its initial state is the state at the moment of separation of the star and the rocket. For the final transition orbit change, the orbital control target values ​​for the second pulse of the previous transition orbit change are the orbital semi-major axis, orbital inclination, velocity increment, number of ignition cycles, and geographical longitude of the ignition point.

2. The method according to claim 1, characterized in that, The preset measurement and control constraints include a first constraint and a second constraint; wherein, The first constraint is: in the double-pulse scheme, the change in the semi-major axis and the change in the tilt angle of the first pulse are less than the preset maximum change in the semi-major axis and the maximum change in the tilt angle, respectively, and the velocity increment of the first pulse is not less than the preset minimum velocity increment and not greater than the preset maximum velocity increment. The second constraint is that the geographical longitude of the ignition point corresponding to the first pulse and the geographical longitude of the ignition point corresponding to the second pulse in the dual-pulse scheme are both within the preset measurement and control range. For each of the two-pulse schemes, it is determined whether it meets the preset constraints, including: Determine whether the dual-pulse scheme satisfies the first constraint condition. If not, remove the dual-pulse scheme. If yes, calculate the geographical longitude of the ignition point of the first pulse and the geographical longitude of the ignition point of the second pulse in the dual-pulse scheme, and determine whether the calculated geographical longitude of the ignition point satisfies the second constraint condition. If yes, determine the dual-pulse scheme as a scheme to be confirmed. If not, adjust the corresponding number of ignition cycles and the measurement and control longitude range until the dual-pulse scheme satisfies the second constraint condition.

3. The method according to claim 2, characterized in that, The geographical longitude of the ignition point for each pulse is determined as follows: The orbital drift rate corresponding to the current sub-pulse is calculated using the following formula: The geographical longitude of the ignition point of the current subpulse is calculated based on the orbital drift rate, using the following formula: Where, when i=1, In the formula, Let be the orbital offset rate of the i-th pulse. Let be the semi-major axis of the orbit of the i-th pulse. km, The gravitational constant, Let be the geographical longitude of the ignition point of the i-th pulse. Let be the geographical longitude of the ignition point of the (i-1)th pulse. This represents the number of ignition cycles for the (i-1)th pulse; when i=1, The geographical longitude of the satellite's first arrival at the ascending node. The drift rate corresponding to the semi-major axis of the initial orbit. The geographical longitude of the first arrival at the ascending node The number of ignition cycles set from the first ignition onwards; i = 1, 2, ..., Nfire, where Nfire is the number of transition trajectories.

4. The method according to claim 1, characterized in that, The total velocity increment for each scheme to be confirmed, as well as the velocity increments of the first and second pulses within that scheme, are determined as follows: The satellite velocities before the first pulse, after the first pulse, before the second pulse, and after the second pulse were determined based on the vitality equation. Based on the satellite velocities before and after the first pulse, and the orbital inclinations before and after the first pulse, the first velocity increment of the first pulse is determined. Based on the satellite velocities corresponding to the second pulse and the pulse after the second pulse, and the orbital inclinations corresponding to the first pulse and the pulse after the first pulse, the second velocity increment of the second pulse is determined. The total velocity increment of the dual-pulse scheme is determined based on the first velocity increment and the second velocity increment.

5. The method according to claim 4, characterized in that, First speed increment The calculation formula is: Second speed increment The calculation formula is: Total velocity increment of the dual-pulse scheme The calculation formula is: In the formula, , , and These are the satellite velocities before the first pulse, after the first pulse, before the second pulse, and after the second pulse, respectively. , , and These are the orbital inclination angles before the first pulse, after the first pulse, before the second pulse, and after the second pulse, respectively.

6. The method according to claim 1, characterized in that, The method of adjusting the thrust direction of the electric thruster by adjusting the in-plane control angle and the out-of-plane control angle includes: Within the preset apogee range, adjust the in-plane control angle using the following formula: Within the preset track lifting and lowering point range, adjust the track out-of-plane control angle using the following formula: In the formula, The control angle in the track plane, For true near point angle, The value is determined based on mission time and thruster requirements. Eccentricity; For out-of-plane control angle, The median value. This is the argument of perigee.

7. A transfer segment orbit-changing device for GEO satellites with hybrid electric-chemical propulsion, characterized in that, include: The division unit is used to divide the transfer segment of a GEO satellite into a first transfer segment and a second transfer segment; The initial state of the first transfer segment is the state at the moment of separation of the spacecraft and the launch vehicle, and the final state of the first transfer segment is determined based on the total propellant of the chemical thrusters and the propellant consumption during the synchronization segment; the initial state of the second transfer segment is the final state of the first transfer segment, and the final state of the second transfer segment is the target orbit. The chemical propulsion orbit change unit is used to perform orbit change using chemical propulsion for the first transfer segment: the orbit change process of the first transfer segment is divided into multiple transition orbit changes. For each transition orbit change, the orbit control target value is determined based on preset measurement and control constraints and the double pulse assumption to complete the orbit change of the first transfer segment. The orbit control target value includes the semi-major axis of the track, the track inclination angle, the velocity increment, the number of ignition cycles, and the geographical longitude of the ignition point. The electric propulsion trajectory-changing unit is used to perform trajectory-changing for the second transfer section using electric propulsion: the thrust direction of the electric propulsion is adjusted by adjusting the in-plane control angle and the out-of-plane control angle to complete the trajectory-changing for the second transfer section; The determination of the track control target value for each transition track change based on preset measurement and control constraints and the bipulse assumption includes: Predetermine the semi-major axis step size and the orbital inclination step size; For each transition track change, execute: The final state of the first transfer segment is taken as the final state of the current transition trajectory change. Based on the initial and final states of the current transition trajectory change, the semi-major axis step size and the track inclination step size are used as the step size of the exhaustive method to assume that the current transition trajectory change is multiple double-pulse schemes. For each double-pulse scheme, it is determined whether it meets the preset constraint conditions. If it does, it is determined as a scheme to be confirmed; if not, it is removed. The total speed increment of each scheme to be confirmed and the speed increments of the first and second pulses in the scheme to be confirmed are calculated. The scheme to be confirmed with the smallest total speed increment is determined as the target scheme. The track semi-major axis, track inclination, speed increment, number of ignition cycles, and geographical longitude of the ignition point corresponding to the first pulse in the target scheme are taken as the track control target values ​​for the current transition trajectory change. The semi-major axis and inclination of the current transition orbit change are used as the initial state for the next transition orbit change, and the next transition orbit change is performed until the last transition orbit change is reached; where, for the first transition orbit change, its initial state is the state at the moment of separation of the star and the rocket. For the final transition orbit change, the orbital control target values ​​for the second pulse of the previous transition orbit change are the orbital semi-major axis, orbital inclination, velocity increment, number of ignition cycles, and geographical longitude of the ignition point.

8. A computing device comprising a memory and a processor, wherein the memory stores a computer program, and the processor, when executing the computer program, implements the method as described in any one of claims 1-6.

9. A computer-readable storage medium having a computer program stored thereon, which, when executed in a computer, causes the computer to perform the method of any one of claims 1-6.

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