Method and device for converting pulse velocity increment into finite thrust orbit control parameter

Abstract

The present invention provides a method and device for converting a pulsed velocity increment into finite thrust orbit control parameters, and the present invention relates to the field of orbital maneuvers. It includes: obtaining relevant parameters of satellite orbital maneuvers in pulsed mode, defining an inertial coordinate system, defining a spherical coordinate system, and the relevant parameters include satellite dry weight, fuel weight before maneuver, fuel weight after maneuver, orbital pulsed maneuver time, inertial coordinate system velocity increment V, and satellite and celestial body positions; calculating the thrust direction in the spherical coordinate system based on the relevant parameters; estimating the thrust magnitudes at the start and shutdown of the thruster based on the relevant parameters, calculating the on-time of the thruster with finite thrust as the initial value for iterative shooting; using the on-time of the finite thrust as the initial value, through continuous iterative shooting, finally generating a finite thrust velocity direction and on-time equivalent to the pulsed velocity increment. The present invention improves the accuracy of orbit control.

Description

Technical Field

[0001] The present invention relates to the field of orbital maneuver, and specifically relates to a method and device for converting impulse velocity increment into finite thrust orbit control parameters. Background Art

[0002] When designing orbital transfer and orbital maneuver, especially for spacecraft orbits going to lunar exploration or deep space exploration orbits, multiple orbital maneuvers are often required. When conducting orbital design, the method of impulse velocity increment is generally used to complete orbital transfer first, that is, the impulse moment is the optimal point for orbit change, and the orbital maneuver is completed instantaneously. However, in actual engineering tasks, the velocity increment of orbital maneuver is completed by the thruster being turned on for a certain period of time. This requires accurately converting the designed value of impulse velocity increment into orbit control parameters such as the engine start time, start duration, and thrust direction in actual engineering applications. However, due to the change of engine thrust during the maneuver, roughly taking the direction of impulse velocity increment as the direction of finite thrust velocity will cause deviation, and the required thrust duration may also be inconsistent with actual engineering applications. Summary of the Invention

[0003] To solve the above technical problems, the present invention provides a method and device for converting impulse velocity increment into finite thrust orbit control parameters, and accurately converts the impulse velocity increment into orbit control parameters such as the start time, start duration, and thrust direction in the finite thrust mode by iteration.

[0004] To achieve the above object, the present invention adopts the following technical solutions:

[0005] A method for converting impulse velocity increment into finite thrust orbit control parameters, comprising the following steps:

[0006] Step 1: Obtain relevant parameters of satellite orbital maneuver in impulse mode, define an inertial coordinate system, define a spherical coordinate system, and the relevant parameters include the dry weight of the satellite , the fuel weight before maneuver , the fuel weight after maneuver , the orbital impulse maneuver moment , the inertial coordinate system velocity increment V and the positions of the satellite and the celestial body;

[0007] Step 2: Calculate the thrust direction in the spherical coordinate system according to the relevant parameters obtained in Step 1;

[0008] Step 3: Estimate the thrust magnitudes at the start and shutdown of the thruster according to the relevant parameters obtained in Step 1, calculate the start duration of the finite thrust of the thruster, and use it as the initial value of iterative shooting;

[0009] Step 4: Using the startup duration of the finite thrust calculated in Step 3 as the initial value, through continuous iterative shooting, finally generate the finite thrust velocity direction and startup duration equivalent to the pulse velocity increment.

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

[0011] A parameter acquisition module, which acquires relevant parameters for satellite orbit maneuvering in pulse mode, defines an inertial coordinate system, defines a spherical coordinate system, and the relevant parameters include the satellite dry weight , the fuel weight before maneuver , the fuel weight after maneuver , the orbit pulse maneuvering moment , the inertial coordinate system velocity increment V and the positions of the satellite and celestial body;

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

[0013] A startup duration calculation module, which estimates the thrust magnitudes at the start and shutdown of the thruster based on the acquired relevant parameters, and calculates the startup duration of the finite thrust of the thruster as the initial value for iterative shooting;

[0014] An iteration module, which uses the startup duration of the finite thrust as the initial value, and through continuous iterative shooting, finally generates the finite thrust velocity direction and startup duration equivalent to the pulse velocity increment.

[0015] The present invention also provides an electronic device, including a memory, a processor, and a computer program stored on the memory and executable on the processor. When the processor executes the program, it implements the steps of the method for converting a pulse velocity increment into finite thrust orbit control parameters described above.

[0016] The present invention also provides a non-transitory computer-readable storage medium, on which a computer program is stored. When the computer program is executed by a processor, it implements the steps of the method for converting a pulse velocity increment into finite thrust orbit control parameters described above.

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

[0018] The present invention accurately converts the pulse velocity increment for maneuvering during orbit transfer into orbit control parameters such as the startup moment, startup duration, and thrust direction in the finite thrust mode. At the same time, through continuous iterative calculation, the effect of the final finite thrust maneuver approaches the effect of the pulse maneuver. Since the pulse maneuver effect is theoretically the optimal effect, the conversion result of the present invention enables the finite thrust maneuver to better approximate the theoretical result in practical engineering applications, increasing the accuracy of orbit control, which is particularly significant for deep space exploration projects that require precise orbit control. Brief Description of the Drawings

[0019] Figure 1 It is a flowchart of the method for converting the pulse velocity increment into the finite-thrust orbit control parameters according to the embodiment of the present invention;

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

[0021] Figure 3 It is a schematic diagram of the judgment and iteration process in Step 4. Detailed Embodiments

[0022] In order to make the objectives, technical solutions and advantages of the present invention clearer and more understandable, the present invention will be further described in detail below with reference to the drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the present invention and are not used 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] As Figure 1 shown, a method for converting the pulse velocity increment into the finite-thrust orbit control parameters disclosed in this embodiment is used to convert the pulse velocity increment into orbit control parameters such as the start time, start duration, and thrust direction of the finite-thrust mode, and includes the following steps:

[0024] Step 1: Obtain relevant parameters of the satellite orbit maneuver in the pulse mode and parameters such as the positions of the satellite and celestial bodies, including:

[0025] Step 1.1: Define an inertial coordinate system (an inertial coordinate system with the central celestial body as the origin), generally taking 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 angle is the angle between the projection of the inertial velocity vector on the plane of the inertial coordinate system and the x-axis of the inertial coordinate system, and the clockwise direction is measured as positive, and the pitch angle is the angle between the inertial velocity vector and the plane of the inertial coordinate system, as Figure 2 shown.

[0026] Taking a far-side lunar maneuver of a detector in a lunar orbit as an example, define the lunar-centered inertial coordinate system as the coordinate system for calculating orbit parameters, and define the spherical coordinate system with the satellite's mass point as the origin as the coordinate system for calculating the finite-thrust direction.

[0027] Step 1.2: Obtain parameters such as the satellite's dry weight, fuel weight before maneuver, fuel weight after maneuver, orbit pulse maneuver time, inertial coordinate system velocity increment V, and the positions of the satellite and celestial bodies.

[0028] In this embodiment, the dry weight of the satellite is 54.2 kg, the fuel weight before maneuver is 10.27 kg, the fuel weight after maneuver is 6.15 kg, the orbital pulse maneuver time is 25 Mar 2024 14:11:00.131 UTCG (Coordinated Universal Time), and the velocity increment V in the inertial coordinate system is [-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 denoted as , judge and calculate the azimuth angle .

[0031] If and , then ;

[0032] If and , then ;

[0033] If , then ;

[0034] If the calculated , then ;

[0035] For example, based on 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 the calculated .

[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 denoted as , judge and calculate the pitch angle :

[0037] ;

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

[0039] Step 2.3: Obtain the azimuth angle of the thrust direction in the spherical coordinate system , the pitch angle , the magnitude of the velocity increment . Among them, represents the modulus of the vector.

[0040] Specifically, according to the above calculations: the azimuth angle of the thrust direction in the spherical coordinate system is 141.56°, the pitch angle is -14.44°, and the magnitude of the velocity increment is 143.34 m / sec.

[0041] Step Three: Estimate the thrust magnitudes at the start and shutdown of the thruster based on the parameters obtained in Step One, and calculate the on-time duration of the thruster with finite thrust as the initial value for iterative shooting, including:

[0042] Step 3.1: Calculate the thruster thrust coefficient based on the fuel mass before the maneuver, and then calculate the thrust magnitude at the ignition moment based on the thruster thrust coefficient;

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

[0044] Thruster thrust coefficient ;

[0045] Among them, is the fuel weight before the maneuver;

[0046] Thrust magnitude ;

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

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

[0049] The calculation of the thruster thrust coefficient and the thrust magnitude at the ignition moment 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] Among them, is the fuel weight after maneuver;

[0052] Thrust magnitude ;

[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 the shutdown moment is 15.43 N.

[0054] Step 3.3: Calculate the on - time of the thruster with finite thrust as the initial value for iterative shooting.

[0055] Let the thrust magnitude at the start - up moment be , and the thrust magnitude at the shutdown moment be , the fuel mass before maneuver be , the fuel mass after maneuver be , the dry weight of the satellite be , and the on - time of the finite thrust be is:

[0056] ;

[0057] The calculated on - time of the finite thrust is 513.25 sec.

[0058] Step Four: Taking the on - time of the finite thrust calculated in Step Three as the initial value, through continuous iterative shooting, finally generate finite - thrust orbit control parameters equivalent to the pulsed velocity increment, as shown in Figure 3 , including:

[0059] Step 4.1: Taking the original target of orbit maneuver as the goal, reset the variables in the three directions of the original velocity increment to the azimuth angle of the finite - thrust direction, the pitch angle , the on - time of the finite thrust , and the start - up moment , where is the moment of orbital pulsed maneuver.

[0060] The original target of orbit maneuver is 450 km at the periselene point, the argument of periselene is 97°, the periselene time is 27 Mar 2024 18:00:00.000 UTCG, the orbital inclination is 122.5°. The target of orbit maneuver remains unchanged. Reset the variables in the three directions of the velocity increment to the azimuth angle of the finite - thrust direction as 141.56°, the pitch angle as - 14.44°, the on - time of the finite thrust as 513.25 sec, the start - up moment as 25 Mar 2024 14:06:43.506 UTCG, and the true anomaly of the original fourth variable at the maneuver moment remains unchanged.

[0061] Step 4.2: Conduct shooting to obtain the new finite-thrust startup duration and the azimuth angle of the new finite-thrust velocity direction and the pitch angle .

[0062] Through shooting, the new finite-thrust startup duration is obtained as 598.36 sec, the azimuth angle of the new finite-thrust velocity direction is 141.56°, and the pitch angle is -14.44°.

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

[0064] , if , then is positive, otherwise it is negative;

[0065] , if , then is positive, otherwise it is negative;

[0066] , if , then is positive, otherwise it is negative;

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

[0068] According to the above three formulas, the absolute value magnitudes of the three quantities are obtained. For example, , and the signs of each quantity are also obtained. The three components together obtain the velocity direction , so = [-0.7832 0.6217 -0.2494].

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

[0070] Among them, represents the modulus of the vector.

[0071] The midpoint of the calculated finite-thrust startup time is 25 Mar 2024 14:11:42.686 UTCG, and the calculated angle between the finite-thrust velocity direction and the impulse velocity direction is approximately 0°.

[0072] Step 4.5: Determine whether the midpoint of the finite-thrust startup time is consistent with the orbital impulse maneuver moment, and determine whether the finite-thrust velocity direction is consistent with the impulse velocity direction, including:

[0073] If (unit: sec) and (unit: deg), then the finite-thrust velocity direction and duration are obtained by converting the impulse velocity increment; if the conditions are not met, then taking the original orbital maneuver target as the target, reset the three velocity increment variables to the azimuth angle of the finite-thrust direction, the pitch angle , the finite-thrust startup duration , the startup moment , and re-shoot until a result that meets the conditions is obtained.

[0074] When the moment is 25 Mar 2024 14:11:42.686 UTCG, the startup moment is 25 Mar 2024 14:06:43.506 UTCG, and the time difference is 42.55 seconds, then = 42.55 sec, not meeting the and conditions. The original orbital maneuver target remains unchanged. Reset the three velocity increment variables to the new azimuth angle of the finite-thrust velocity direction 141.56° and the pitch angle -14.44°, the new finite-thrust startup duration 598.36 sec, and set the startup moment to 25 Mar 2024 14:06:00.951 UTCG.

[0075] Repeat steps 4.2 to 4.5, and obtain the new azimuth angle of the finite-thrust velocity direction 141.56° and the pitch angle -14.44°, the new finite-thrust startup duration 598.36 sec. The midpoint of the finite-thrust startup time is 25 Mar 2024 14:11:00.131 UTCG, the angle between the finite-thrust velocity direction and the impulse velocity direction is approximately 0°, = 0 sec, , meeting the conditions.

[0076] It can be seen from this that when starting ignition at the azimuth angle of 141.56° and pitch angle of -14.44° of the finite thrust velocity direction and at the start time of 25Mar 2024 14:06:00.951 UTCG, and when the total startup time is 598.36 sec, the results of this finite thrust maneuver are equivalent to those of the impulsive velocity increment maneuver.

[0077] The embodiment of the present invention also provides a device for converting an impulsive velocity increment into finite thrust orbit control parameters, including the following modules:

[0078] A parameter acquisition module, which acquires relevant parameters of satellite orbit maneuver in the impulsive mode, defines an inertial coordinate system, defines a spherical coordinate system, and the relevant parameters include the satellite dry weight , the fuel weight before maneuver , the fuel weight after maneuver , the orbit impulsive maneuver time , the inertial coordinate system velocity increment V and the positions of the satellite and the celestial body;

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

[0080] A startup duration calculation module, which estimates the thrust magnitudes at the start and shutdown of the thruster based on the acquired relevant parameters, calculates the startup duration of the thruster's finite thrust, and uses it as the initial value for iterative shooting;

[0081] An iteration module, which uses the startup duration of the finite thrust as the initial value, and through continuous iterative shooting, finally generates a finite thrust velocity direction and startup duration equivalent to the impulsive velocity increment.

[0082] The embodiment of the present invention also provides an electronic device, including a memory, a processor, and a computer program stored on the memory and executable on the processor. When the processor executes the program, it implements the steps of the method for converting an impulsive velocity increment into finite thrust orbit control parameters described above.

[0083] The embodiment of the present invention also provides a non-transitory computer-readable storage medium, on which a computer program is stored. When the computer program is executed by a processor, it implements the steps of the method for converting an impulsive velocity increment into finite thrust orbit control parameters described above.

[0084] Those skilled in the art should understand that the embodiments of the present invention can be provided as a method, a system, or a computer program product. Therefore, the present invention can take the form of a completely hardware embodiment, a completely software embodiment, or an embodiment combining software and hardware aspects. Moreover, the present invention can take the form of a computer program product implemented on one or more computer-usable storage media (including but not limited to disk memory, CD-ROM, optical memory, etc.) containing computer-usable program code. The solutions in the embodiments of the present invention can be implemented in various computer languages. For example, object-oriented programming languages such as Java and interpreted scripting languages such as JavaScript, etc.

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

[0086] These computer program instructions can also be stored in a computer-readable memory that can direct a computer or other programmable data processing devices to work in a specific manner, such that the instructions stored in the computer-readable memory generate a manufactured article including an instruction means, and the instruction means implements the functions specified in Figure 1 one flow or multiple flows and / or blocks Figure 1 one block or multiple blocks.

[0087] These computer program instructions can also be loaded onto a computer or other programmable data processing devices, such that a series of operation steps are executed on the computer or other programmable devices to generate a computer-implemented process, and thus the instructions executed on the computer or other programmable devices provide steps for implementing the functions specified in Figure 1 one flow or multiple flows and / or blocks Figure 1 one block or multiple blocks.

[0088] Although the preferred embodiments of the present invention have been described, those skilled in the art can make additional changes and modifications once they know the basic creative concepts. Therefore, the appended claims are intended to be construed as including the preferred embodiments and all changes and modifications falling within the scope of the present invention.

[0089] Obviously, those skilled in the art can make various modifications and variations 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 equivalent technologies, 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 It includes the following steps: Step 1: Obtain the relevant parameters of satellite orbital maneuver in pulse mode, define the inertial coordinate system, define the spherical coordinate system, and the relevant parameters include the dry weight of the satellite , the fuel weight before maneuver , the fuel weight after maneuver , the orbital pulse maneuver time , the inertial coordinate system velocity increment V and the positions of the satellite and the celestial body; 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 magnitudes at the start and shutdown of the thruster based on the relevant parameters obtained in Step 1, and calculate the on-time duration of the thruster's finite thrust as the initial value for iterative shooting; Step 4: Using the on-time duration of the finite thrust calculated in Step 3 as the initial value, through continuous iterative shooting, finally generate a finite thrust velocity direction and on-time duration equivalent to the pulsed velocity increment.

2. The method for converting a pulse speed increment into a finite thrust orbit control parameter according to claim 1, characterized in that, The said Step 1 includes the following steps: Step 1.1: Define an inertial coordinate system, an 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, the azimuth angle being the angle between the projection of the inertial velocity vector on the inertial coordinate system plane and the x-axis of the inertial coordinate system, measured as positive in the clockwise direction, and the pitch angle being the angle between the inertial velocity vector and the inertial coordinate system plane; Step 1.2: Obtain the satellite's dry weight , the fuel weight before maneuver , the fuel weight after maneuver , the orbital impulsive maneuver time , the inertial coordinate system velocity increment V and the positions of the satellite and celestial body.

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

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

5. A method for converting pulse speed increment into finite thrust orbit control parameters according to claim 4, characterized in that The said Step 2.2 includes: ； wherein, is the pitch angle of the spherical coordinate system, is the magnitude of the velocity increment, represents the magnitude of the vector.

6. A method for converting a pulse speed increment into a finite thrust orbit control parameter according to claim 5, characterized in that, The said Step 3 includes the following steps: Step 3.1: Calculate the thruster thrust coefficient based on the fuel mass before the maneuver, and then calculate the thrust magnitude at the ignition moment based on the thrust coefficient; Step 3.2: Calculate the thruster thrust coefficient based on the fuel mass after the maneuver, and then calculate the thrust magnitude at the shutdown moment based on the thrust coefficient; Step 3.3: Calculate the on-time duration of the thruster's finite thrust as the initial value for iterative shooting.

7. A method for converting a pulse speed increment into a finite thrust trajectory control parameter according to claim 6, characterized in that The said Step 3.3 includes: The thrust at startup is , and the thrust at shutdown is . The fuel mass before maneuver is , and the fuel mass after maneuver is . The dry mass of the satellite is . The startup duration of the finite thrust is : 。 8. A method for converting pulse speed increment into finite thrust orbit control parameters according to claim 1, characterized in that The said Step 4 includes the following steps: Step 4.1: Taking the original target of orbital maneuver as the goal, reset the variables of the original inertial velocity increment in three directions to the azimuth angle of the finite thrust direction , pitch angle , thrust duration , starting moment , where is the moment of orbital pulse maneuver; Step 4.2: Perform targeting to obtain the new limited-thrust startup duration and the new azimuth angle of the limited-thrust velocity direction and pitch angle ; Step 4.3: Convert the finite thrust direction in the new spherical coordinate system into the velocity direction in the inertial coordinate system , The components on the three axes are respectively , , ; Step 4.4: Calculate the midpoint of the finite-thrust startup time , calculate the angle between the finite-thrust velocity direction and the pulse velocity direction ; Step 4.5: Judge whether the midpoint of the finite thrust on-time is consistent with the orbital pulsed maneuver moment, and judge whether the finite thrust velocity direction is consistent with the pulsed velocity direction.

9. A method for converting pulse speed increment into finite thrust orbit control parameters according to claim 8, characterized in that The said Step 4.3 includes: , if , then is positive, otherwise it is negative; , if , then is positive, otherwise negative; , if , then is positive, otherwise it is negative; , , so is negative, is positive, is negative.

10. A method for converting a pulse speed increment into a finite thrust trajectory control parameter according to claim 8, characterized in that, The said Step 4.5 includes: If sec and deg, the finite thrust velocity direction, startup time, and startup duration are obtained by converting the pulse velocity increment; if the conditions are not met, the target of the original orbital maneuver is used as the target, and the three variables are reset to the azimuth angle of the finite thrust direction, the pitch angle , the thrust duration , the startup time , and the new finite thrust startup duration is obtained by target shooting as well as the new azimuth angle of the finite thrust velocity direction and the pitch angle . Steps 4.1 to 4.5 are repeated until a result that meets the conditions is obtained.

11. A device for converting pulse velocity increment into finite thrust orbit control parameters, characterized in that It includes the following modules: The parameter acquisition module acquires relevant parameters for satellite orbital maneuver in pulse mode, defines an inertial coordinate system, defines a spherical coordinate system, and the relevant parameters include the satellite dry mass , the fuel weight before maneuver , the fuel weight after maneuver , the orbital pulse maneuver time , the inertial coordinate system velocity increment V and the positions of the satellite and celestial body; A thrust direction calculation module, which calculates the thrust direction in the spherical coordinate system based on the obtained relevant parameters; An on-time duration calculation module, which estimates the thrust magnitudes at the start and shutdown of the thruster based on the obtained relevant parameters, and calculates the on-time duration of the thruster's finite thrust as the initial value for iterative shooting; An iterative module, which uses the on-time duration of the finite thrust as the initial value, through continuous iterative shooting, finally generates a finite thrust velocity direction and on-time duration equivalent to the pulsed velocity increment.

12. An electronic device, comprising a memory, a processor, and a computer program stored on the memory and executable on the processor, characterized in that, When the processor executes the said program, it implements the steps of the method for converting pulsed velocity increment into finite thrust orbit control parameters as described in any one of claims 1 - 10.

13. A non-transitory computer-readable storage medium, on which a computer program is stored, characterized in that when the computer program is executed by a processor, it implements the steps of the method for converting pulsed velocity increment into finite thrust orbit control parameters as described in any one of claims 1 - 10.

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