Method and device for converting pulse speed increment into finite thrust orbit control parameters

Abstract

The invention provides a method and device for converting a pulse speed increment into a finite thrust orbit control parameter, and relates to the field of orbit maneuver. Comprising the steps that relevant parameters of satellite orbital maneuver in a pulse mode are obtained, an inertial coordinate system is defined, a spherical coordinate system is defined, and the relevant parameters comprise the satellite dry weight # imgabs0 #, the fuel weight # imgabs1 # before maneuver, the fuel weight # imgabs2 # after maneuver, the orbital pulse maneuver moment # imgabs3 #, the speed increment V of the inertial coordinate system and the positions of a satellite and a celestial body; calculating the thrust direction under the spherical coordinate system according to the related parameters; estimating the thrust when the thruster starts to ignite and shut down according to the related parameters, and calculating the starting duration of the limited thrust of the thruster as an initial value of iterative targeting; and taking the starting time length of the limited thrust as an initial value, and finally generating a limited thrust speed direction and starting time length which are equal to the pulse speed increment through continuous iterative targeting. According to the invention, the accuracy degree of track control is increased.

Description

Technical Field

[0001] The present invention relates to the field of orbital maneuvering, and in particular to a method and a device for converting a pulse velocity increment into a finite thrust orbit control parameter. Background Art

[0002] When designing orbit transfer and orbit maneuvers, especially for spacecraft orbits for lunar exploration or deep space exploration orbits, multiple orbit maneuvers are often required. When designing orbits, the pulse velocity increment method is generally used to complete the orbit transfer, that is, the pulse time is the optimal point for orbit change, and the orbit maneuver is completed in an instant. However, in actual engineering tasks, the velocity increment of the orbit maneuver depends on the thruster being turned on for a certain period of time. This requires that the design value of the pulse velocity increment be accurately converted into orbit control parameters such as the engine start-up time, start-up duration, and thrust direction in actual engineering applications. However, due to the change in engine thrust during the maneuver, roughly using the direction of the pulse velocity increment as the limited thrust velocity direction will lead to deviations, and the required thrust duration may also be inconsistent with the actual engineering application. Summary of the invention

[0003] In order to solve the above technical problems, the present invention provides a method and device for converting pulse velocity increments into limited thrust orbit control parameters, which use iteration to accurately convert pulse velocity increments into orbit control parameters such as the start-up time, start-up duration, and thrust direction of the limited thrust mode.

[0004] In order to achieve the above object, the present invention adopts the following technical solution:

[0005] A method for converting pulse velocity increments into finite thrust orbit control parameters comprises the following steps:

[0006] Step 1: Obtain the relevant parameters of the satellite orbit maneuver in pulse mode, define the inertial coordinate system, define the spherical coordinate system, and the relevant parameters include the satellite dry weight , fuel weight before maneuvering , fuel weight after maneuvering , orbital pulse maneuver time , velocity increment V of the inertial coordinate system and the position of satellites and celestial bodies;

[0007] Step 2: Calculate the thrust direction in the spherical coordinate system based on the relevant parameters obtained in step 1;

[0008] Step 3: Estimate the thrust of the thruster when it starts ignition and shuts down based on the relevant parameters obtained in step 1, and calculate the startup time of the thruster's limited thrust as the initial value for iterative target shooting;

[0009] Step 4: Based on the limited thrust start-up time calculated in step 3 as the initial value, after continuous iterative targeting, the limited thrust speed direction and start-up time equivalent to the pulse speed increment are finally generated.

[0010] The present invention also provides a device for converting pulse velocity increment into limited thrust orbit control parameters, comprising the following modules:

[0011] Parameter acquisition module, obtains the relevant parameters of satellite orbit maneuvers in pulse mode, defines the inertial coordinate system, defines the spherical coordinate system, and the relevant parameters include the satellite dry weight , fuel weight before maneuvering , fuel weight after maneuvering , orbital pulse maneuver time , velocity increment V of the inertial coordinate system and the position of satellites and celestial bodies;

[0012] A thrust direction calculation module calculates the thrust direction in the spherical coordinate system based on the acquired relevant parameters;

[0013] The startup time calculation module estimates the thrust size when the thruster starts ignition and shuts down based on the relevant parameters obtained, and calculates the startup time of the thruster with limited thrust as the initial value of iterative target shooting;

[0014] The iteration module takes the startup time of the limited thrust as the initial value, and after continuous iterative targeting, it finally generates the limited thrust speed direction and startup time equivalent to the pulse speed increment.

[0015] The present invention also provides an electronic device, comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, wherein when the processor executes the program, the steps of the above-mentioned method of converting a pulse velocity increment into a finite thrust orbit control parameter are implemented.

[0016] The present invention also provides a non-transitory computer-readable storage medium having a computer program stored thereon, which, when executed by a processor, implements the steps of the above-mentioned method of converting pulse velocity increments into finite thrust orbit control parameters.

[0017] Compared with the prior art, the present invention has the following beneficial effects:

[0018] The present invention accurately converts the pulse velocity increment of maneuvers during orbit transfer into orbit control parameters such as the start-up time, start-up duration, and thrust direction of the limited thrust mode, and at the same time, through continuous iterative calculation, the effect of the final limited thrust maneuver is close to the effect of the pulse maneuver. Since the pulse maneuver effect is theoretically the optimal effect, the conversion result of the present invention makes the limited thrust maneuver better approach the theoretical result in actual engineering applications, increases the accuracy of orbit control, and is of great significance in particular to deep space exploration projects that require precise orbit control. BRIEF DESCRIPTION OF THE DRAWINGS

[0019] Figure 1 A flow chart of a method for converting a pulse velocity increment into a finite thrust orbit control parameter according to an embodiment of the present invention;

[0020] Figure 2 It is a schematic diagram of the inertial coordinate system;

[0021] Figure 3 Schematic diagram of the judgment and iteration process in step 4. DETAILED DESCRIPTION

[0022] In order to make the purpose, technical solutions and advantages of the present invention more clearly understood, the present invention is further described in detail below in conjunction with the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the present invention and are not intended to limit the present invention. In addition, the technical features involved in the various embodiments of the present invention described below can be combined with each other as long as they do not conflict with each other.

[0023] like Figure 1 As shown, a method for converting a pulse velocity increment into a limited thrust orbit control parameter disclosed in this embodiment is used to convert a pulse velocity increment into orbit control parameters such as a startup time, startup duration, and thrust direction of a limited thrust mode, and includes the following steps:

[0024] Step 1: Obtain the relevant parameters of the satellite orbit maneuver in pulse mode and the satellite and celestial body position parameters, including:

[0025] Step 1.1: Define an inertial coordinate system (an inertial coordinate system with the central celestial body as the origin). Generally, the inertial coordinate system with the central celestial body of the satellite orbit as the origin; define a spherical coordinate system with the satellite's center of mass as the origin, and the azimuth It is the angle between the projection of the inertial velocity vector on the inertial coordinate plane and the x-axis of the inertial coordinate system. It is positive when measured in the clockwise direction. is the angle between the inertial velocity vector and the inertial coordinate plane, such as Figure 2 shown.

[0026] Taking a far-moon maneuver of the probe in the lunar orbit as an example, the lunar-centered inertial coordinate system is defined as the coordinate system for calculating the orbital parameters, and the spherical coordinate system with the satellite mass point as the origin is defined as the coordinate system for calculating the finite thrust direction.

[0027] Step 1.2: Obtain satellite dry weight, fuel weight before maneuver, fuel weight after maneuver, orbit pulse maneuver time, inertial coordinate system velocity increment V, and satellite and celestial body positions.

[0028] In this embodiment, the satellite dry weight is 54.2 kg, the fuel weight before maneuver is 10.27 kg, the fuel weight after maneuver is 6.15 kg, the orbit pulse maneuver time is 25 Mar 2024 14:11:00.131 UTCG (Universal Coordinated Time), and the inertial coordinate system velocity increment V=[-108.73 m / sec 86.2962 m / sec -35.7555 m / sec].

[0029] Step 2: Calculate the thrust direction in the spherical coordinate system based on the parameters obtained in step 1, including:

[0030] Step 2.1: The components of the velocity increment V in the inertial coordinate system on the X, Y, and Z axes are V1, V2, and V3 respectively, and the azimuth angle of the spherical coordinate system is recorded as , determine and calculate the azimuth of the spherical coordinate system .

[0031] like and ,but ;

[0032] like and ,but ;

[0033] like ,but ;

[0034] If the calculated ,but ;

[0035] For example, according to V=[-108.73 m / sec 86.2962 m / sec -35.7555 m / sec], V1=-108.73 m / sec, V2=86.2962 m / sec, V3=-35.7555 m / sec, , , then we can calculate .

[0036] Step 2.2: The components of the velocity increment V in the inertial coordinate system on the X, Y, and Z axes are V1, V2, and V3 respectively, and the pitch angle of the spherical coordinate system is recorded as , determine and calculate the pitch angle of the spherical coordinate system :

[0037] ;

[0038] For example, according to V=[-108.73 m / sec 86.2962 m / sec -35.7555 m / sec], V3=-35.7555m / sec, , calculated .

[0039] Step 2.3: Get the azimuth of the thrust direction in the spherical coordinate system , pitch angle , speed increment size .in, Represents the magnitude of a vector.

[0040] Specifically, according to the above calculation, it is obtained that the thrust direction azimuth angle in the spherical coordinate system is 141.56°, the pitch angle is -14.44°, and the velocity increment is 143.34 m / sec.

[0041] Step 3: Estimate the thrust of the thruster when it starts ignition and shuts down based on the parameters obtained in step 1, and calculate the startup time of the thruster's limited thrust as the initial value for iterative target shooting, including:

[0042] Step 3.1: Calculate the thrust coefficient of the thruster according to the fuel mass before the maneuver, and then calculate the thrust magnitude at the ignition time according to the thrust coefficient of the thruster;

[0043] The calculation of the thrust coefficient of the thruster and the thrust magnitude at the ignition time is related to the properties of the thruster. In this embodiment, it is assumed that the properties of the thruster are calculated by the following formula:

[0044] Thruster thrust coefficient ;

[0045] in, is the fuel weight before maneuvering;

[0046] Thrust size ;

[0047] Taking 10.2689 kg, the thrust coefficient of the thruster is calculated to be 1.0992, and the thrust magnitude at the ignition time is 18.47 N.

[0048] Step 3.2: Calculate the thrust coefficient of the thruster based on the fuel mass after the maneuver, and then calculate the thrust magnitude at shutdown based on the thrust coefficient;

[0049] The calculation of thrust coefficient and thrust magnitude at ignition time is related to the properties of the thruster. In this example, it is assumed that the thruster properties are calculated by the following formula:

[0050] Thruster thrust coefficient ;

[0051] in, is the fuel weight after maneuvering;

[0052] Thrust size ;

[0053] When the value is 6.1475 kg, the thrust coefficient of the thruster is calculated to be 0.8892, and the thrust magnitude at shutdown is 15.43 N.

[0054] Step 3.3: Calculate the duration of the thruster's limited thrust startup as the initial value for iterative targeting.

[0055] Assume the thrust at startup is , the thrust at shutdown time is , the fuel mass before maneuvering is , the fuel mass after maneuvering is , the satellite dry weight is , the limited thrust startup time is for:

[0056] ;

[0057] The calculated limited thrust startup time is 513.25 seconds.

[0058] Step 4: Based on the limited thrust start-up time calculated in step 3 as the initial value, after continuous iterative target practice, the limited thrust orbit control parameters equivalent to the pulse velocity increment are finally generated, such as Figure 3 As shown, including:

[0059] Step 4.1: Taking the original orbital maneuver as the target, reset the variables of the three directions of the original velocity increment to the azimuth angle of the limited thrust direction. , pitch angle , limited thrust start-up time , boot time ,in is the orbital pulse maneuver time.

[0060] The original target of the orbital maneuver was a perigee of 450km, a perigee amplitude of 97°, a perigee time of 27 Mar 2024 18:00:00.000 UTCG, and an orbital inclination of 122.5°. The target of the orbital maneuver remained unchanged, and the variables in the three directions of the velocity increment were reset to the azimuth of the limited thrust direction of 141.56°, the pitch angle of -14.44°, the limited thrust startup time of 513.25sec, and the startup time of 25 Mar 2024 14:06:43.506 UTCG. The original fourth variable true anomaly remained unchanged at the time of the maneuver.

[0061] Step 4.2: Target shooting to obtain a new limited thrust start-up time and the azimuth of the new limited thrust velocity direction and pitch angle .

[0062] Through target practice, the new limited thrust startup time was obtained as 598.36 seconds, the new limited thrust speed direction azimuth angle was 141.56°, and the pitch angle was -14.44°.

[0063] Step 4.3: Convert the direction of the finite thrust in the new spherical coordinate system to the direction of the velocity in the inertial coordinate system , The components on the three axes are , , .

[0064] ,like ,but is positive, otherwise it is negative;

[0065] ,like ,but is positive, otherwise it is negative;

[0066] ,like ,but is positive, otherwise it is negative;

[0067] , ,therefore is negative, is positive, is negative.

[0068] According to the above three formulas, we can get the absolute values ​​of the three quantities, such as , and the signs of each quantity are obtained. The three components together give the direction of velocity ,therefore =[-0.7832 0.6217 -0.2494].

[0069] Step 4.4: Calculate the midpoint of the limited thrust start time , calculate the angle between the finite thrust velocity direction and the pulse velocity direction .

[0070] in, Represents the magnitude of a vector.

[0071] The midpoint of the limited thrust start-up time is calculated to be 25 Mar 2024 14:11:42.686 UTCG, and the angle between the limited thrust speed direction and the pulse speed direction is calculated to be approximately 0°.

[0072] Step 4.5: Determine whether the midpoint of the limited thrust start time is consistent with the orbital pulse maneuvering time, and determine whether the limited thrust speed direction is consistent with the pulse speed direction, including:

[0073] like (Unit: sec) and (unit: deg), the limited thrust velocity direction and duration are obtained by converting the pulse velocity increment; if the conditions are not met, the original orbital maneuver target is taken as the target, and the three velocity increment variables are reset to the azimuth of the limited thrust direction. , Pitch angle , Limited thrust startup time , boot time , re-shoot until a result that satisfies the conditions is obtained.

[0074] when The time is 25 Mar 2024 14:11:42.686 UTCG, and the power-on time is 25 Mar 2024 14:06:43.506 UTCG. The time difference is 42.55 seconds. =42.55 sec, not satisfied and The original orbital maneuver target remains unchanged, and the three velocity increment variables are reset to the new limited thrust velocity direction azimuth angle of 141.56° and pitch angle of -14.44°. The new limited thrust startup time is 598.36 sec, and the startup time is set to 25 Mar 2024 14:06:00.951 UTCG.

[0075] Repeat steps 4.2 to 4.5 to obtain a new limited thrust speed direction azimuth angle of 141.56° and a pitch angle of -14.44°. The new limited thrust startup time is 598.36 seconds. The midpoint of the limited thrust startup time is 25 Mar 2024 14:11:00.131 UTCG. The angle between the limited thrust speed direction and the pulse speed direction is about 0°. =0 sec, , the conditions are met.

[0076] It can be seen that with a limited thrust speed direction azimuth angle of 141.56° and a pitch angle of -14.44°, the ignition started at 25Mar 2024 14:06:00.951 UTCG, and the total power-on time was 598.36 seconds. The results of this limited thrust maneuver are equivalent to those of the pulse speed increment maneuver.

[0077] The embodiment of the present invention further provides a device for converting pulse velocity increment into limited thrust orbit control parameters, comprising the following modules:

[0078] Parameter acquisition module, obtains the relevant parameters of satellite orbit maneuvers in pulse mode, defines the inertial coordinate system, defines the spherical coordinate system, and the relevant parameters include the satellite dry weight , fuel weight before maneuvering , fuel weight after maneuvering , orbital pulse maneuver time , velocity increment V of the inertial coordinate system and the position of satellites and celestial bodies;

[0079] A thrust direction calculation module calculates the thrust direction in the spherical coordinate system based on the acquired relevant parameters;

[0080] The startup time calculation module estimates the thrust size when the thruster starts ignition and shuts down based on the relevant parameters obtained, and calculates the startup time of the thruster with limited thrust as the initial value of iterative target shooting;

[0081] The iteration module takes the startup time of the limited thrust as the initial value, and after continuous iterative targeting, it finally generates the limited thrust speed direction and startup time equivalent to the pulse speed increment.

[0082] An embodiment of the present invention also provides an electronic device, including a memory, a processor, and a computer program stored in the memory and executable on the processor, wherein when the processor executes the program, the steps of the above-mentioned method of converting a pulse velocity increment into a finite thrust orbit control parameter are implemented.

[0083] An embodiment of the present invention further provides a non-transitory computer-readable storage medium having a computer program stored thereon. When the computer program is executed by a processor, the steps of the above-mentioned method of converting a pulse velocity increment into a finite thrust orbit control parameter are implemented.

[0084] It should be understood by those skilled in the art that the embodiments of the present invention may be provided as methods, systems, or computer program products. Therefore, the present invention may take the form of a complete hardware embodiment, a complete software embodiment, or an embodiment combining software and hardware. Moreover, the present invention may take the form of a computer program product implemented on one or more computer-usable storage media (including but not limited to disk storage, CD-ROM, optical storage, etc.) containing computer-usable program codes. The solutions in the embodiments of the present invention may be implemented in various computer languages, for example, object-oriented programming language Java and interpreted scripting language JavaScript, etc.

[0085] The present invention is described with reference to flowcharts and / or block diagrams of methods, devices (systems), and computer program products according to embodiments of the present invention. It should be understood that each process and / or block in the flowchart and / or block diagram, as well as the combination of processes and / or blocks in the flowchart and / or block diagram, can be implemented by computer program instructions. These computer program instructions can be provided to a processor of a general-purpose computer, a special-purpose computer, an embedded processor, or other programmable data processing device to produce a machine, so that the instructions executed by the processor of the computer or other programmable data processing device generate instructions for implementing the processes in the flowchart and / or block diagram. Figure 1 A process or multiple processes and / or boxes Figure 1 A device that provides the functions specified in a block or multiple blocks.

[0086] These computer program instructions may also be stored in a computer-readable memory capable of directing a computer or other programmable data processing device to operate in a specific manner, so that the instructions stored in the computer-readable memory produce an article of manufacture comprising an instruction device, which implements the process Figure 1 A process or multiple processes and / or boxes Figure 1 A function specified in one or more boxes.

[0087] These computer program instructions can also be loaded onto a computer or other programmable data processing device so that a series of operating steps are executed on the computer or other programmable device to produce a computer-implemented process, thereby providing instructions for implementing the process. Figure 1 A process or multiple processes and / or boxes Figure 1 The steps for the functions specified in one or more boxes.

[0088] Although the preferred embodiments of the present invention have been described, those skilled in the art may make other changes and modifications to these embodiments once they have learned the basic creative concept. Therefore, the appended claims are intended to be interpreted as including the preferred embodiments and all changes and modifications that fall within the scope of the present invention.

[0089] Obviously, those skilled in the art can make various changes and modifications to the present invention without departing from the spirit and scope of the present invention. Thus, if these modifications and variations of the present invention fall within the scope of the claims of the present invention and their equivalents, the present invention is also intended to include these modifications and variations.

Claims

1. A method for converting pulse velocity increment into finite thrust orbit control parameters, characterized in that: The following steps are involved: Step 1: Obtain the relevant parameters of the satellite orbit maneuver in pulse mode, define the inertial coordinate system, define the spherical coordinate system, and the relevant parameters include the satellite dry weight , fuel weight before maneuvering , fuel weight after maneuvering , orbital pulse maneuver time , velocity increment V of the inertial coordinate system and the position of satellites and celestial bodies; Step 2: Calculate the thrust direction in the spherical coordinate system based on the relevant parameters obtained in step 1; Step 3: Estimate the thrust of the thruster when it starts ignition and shuts down based on the relevant parameters obtained in step 1, and calculate the startup time of the thruster's limited thrust as the initial value for iterative target shooting; Step 4: Based on the limited thrust start-up time calculated in step 3 as the initial value, after continuous iterative targeting, the limited thrust speed direction and start-up time equivalent to the pulse speed increment are finally generated.

2. A method for converting pulse velocity increment into limited thrust orbit control parameters according to claim 1, characterized in that: The step 1 comprises the following steps: Step 1.1: Define the inertial coordinate system, taking the central celestial body of the satellite orbit as the origin of the inertial coordinate system; define the spherical coordinate system, taking the satellite's center of mass as the origin, the azimuth angle is the angle between the projection of the inertial velocity vector on the inertial coordinate plane and the x-axis of the inertial coordinate system, and the clockwise direction is measured as a positive value, and the pitch angle is the angle between the inertial velocity vector and the inertial coordinate plane; Step 1.2: Get the satellite dry weight , fuel weight before maneuvering , fuel weight after maneuvering , orbital pulse maneuver time , the velocity increment V of the inertial coordinate system and the position of the satellite and celestial body.

3. A method for converting pulse velocity increment into limited thrust orbit control parameters according to claim 1, characterized in that: The step 2 comprises the following steps: Step 2.1: The components of the velocity increment V in the inertial coordinate system on the X, Y, and Z axes are V1, V2, and V3 respectively, and the azimuth angle of the spherical coordinate system is recorded as , determine and calculate the azimuth of the spherical coordinate system ; Step 2.2: The components of the velocity increment V in the inertial coordinate system on the X, Y, and Z axes are V1, V2, and V3 respectively, and the pitch angle of the spherical coordinate system is recorded as , determine and calculate the pitch angle of the spherical coordinate system ; Step 2.3: The thrust direction azimuth in the spherical coordinate system is , the azimuth is , the thrust is .

4. A method for converting pulse velocity increment into limited thrust orbit control parameters according to claim 3, characterized in that: The step 2.1 comprises: like and ,but ; like and ,but ; like ,but ; If the calculated ,but ; Among them, the components of the velocity increment V of the inertial coordinate system on the X, Y, and Z axes are V1, V2, and V3 respectively.

5. A method for converting pulse velocity increment into limited thrust orbit control parameters according to claim 4, characterized in that: The step 2.2 comprises: ； in, is the pitch angle of the spherical coordinate system, is the speed increment size, Represents the magnitude of a vector.

6. A method for converting pulse velocity increment into limited thrust orbit control parameters according to claim 5, characterized in that: The step three comprises the following steps: Step 3.1: Calculate the thrust coefficient of the thruster based on the fuel mass before the maneuver, and then calculate the thrust magnitude at the ignition time based on the thrust coefficient; Step 3.2: Calculate the thrust coefficient of the thruster based on the fuel mass after the maneuver, and then calculate the thrust magnitude at shutdown based on the thrust coefficient; Step 3.3: Calculate the duration of the thruster's limited thrust startup as the initial value for iterative targeting.

7. A method for converting pulse velocity increment into limited thrust orbit control parameters according to claim 6, characterized in that: The step 3.3 comprises: The thrust at startup is , the thrust at shutdown time is , the fuel mass before maneuvering is , the fuel mass after maneuvering is , the satellite dry weight is , the limited thrust startup time is for: 。 8. The method for converting pulse velocity increment into limited thrust orbit control parameters according to claim 1, characterized in that: The step 4 comprises the following steps: Step 4.1: Taking the original orbital maneuver as the target, reset the variables of the three directions of the original inertial velocity increment to the azimuth angle of the limited thrust direction. , pitch angle , thrust duration , boot time ,in is the orbital pulse maneuver time; Step 4.2: Target shooting to obtain a new limited thrust start-up time and the new limited thrust velocity direction azimuth and pitch angle ; Step 4.3: Convert the direction of the finite thrust in the new spherical coordinate system to the direction of the velocity in the inertial coordinate system , The components on the three axes are , , ; Step 4.4: Calculate the midpoint of the limited thrust start time , calculate the angle between the finite thrust velocity direction and the pulse velocity direction ; Step 4.5: Determine whether the midpoint of the limited thrust start-up time is consistent with the orbital pulse maneuvering time, and determine whether the limited thrust speed direction is consistent with the pulse speed direction.

9. A method for converting pulse velocity increment into limited thrust orbit control parameters according to claim 8, characterized in that: The step 4.3 includes: ,like ,but is positive, otherwise it is negative; ,like ,but is positive, otherwise it is negative; ,like ,but is positive, otherwise it is negative; , ,therefore is negative, is positive, is negative.

10. The method for converting pulse velocity increment into limited thrust orbit control parameters according to claim 8, characterized in that: The step 4.5 comprises: like sec and deg, the limited thrust speed direction, start-up time and start-up duration are obtained by pulse speed increment conversion; if the conditions are not met, the original orbital maneuver target is taken as the target, and the three variables are reset to the azimuth of the limited thrust direction , pitch angle , thrust duration , boot time , target shooting to get a new limited thrust start time and the new limited thrust velocity direction azimuth and pitch angle , repeat steps 4.1 to 4.5 until a result that satisfies the conditions is obtained.

11. A device for converting pulse velocity increments into limited thrust orbit control parameters, characterized in that: Includes the following modules: Parameter acquisition module, obtains the relevant parameters of satellite orbit maneuvers in pulse mode, defines the inertial coordinate system, defines the spherical coordinate system, and the relevant parameters include the satellite dry weight , fuel weight before maneuvering , fuel weight after maneuvering , orbital pulse maneuver time , velocity increment V of the inertial coordinate system and the position of satellites and celestial bodies; A thrust direction calculation module calculates the thrust direction in the spherical coordinate system based on the acquired relevant parameters; The startup time calculation module estimates the thrust size when the thruster starts ignition and shuts down based on the relevant parameters obtained, and calculates the startup time of the thruster with limited thrust as the initial value of iterative target shooting; The iteration module takes the startup time of the limited thrust as the initial value, and after continuous iterative targeting, it finally generates the limited thrust speed direction and startup time equivalent to the pulse speed increment.

12. An electronic device comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, wherein: When the processor executes the program, the steps of the method for converting pulse velocity increments into limited thrust orbit control parameters as described in any one of claims 1 to 10 are implemented.

13. A non-transitory computer-readable storage medium having a computer program stored thereon, wherein when the computer program is executed by a processor, the steps of the method for converting a pulse velocity increment into a finite thrust orbit control parameter as claimed in any one of claims 1 to 10 are implemented.

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