A method for orbit deflection of near-Earth asteroid kinetic impact based on resonance orbit

Abstract

The present invention provides a method for deflecting the kinetic impact orbit of a near-Earth asteroid based on a resonant orbit, including: determining the launch window using a pock-chop diagram, setting the range of optimization variables, calculating the orbits of an observer and an impactor according to a dynamic model, and obtaining the position velocity parameters of the mission node; optimizing using a model combining a genetic algorithm and a local search to obtain a mission orbit model with the largest asteroid deflection distance. The specific impact process is: launching an impactor and an observer from the Earth with one rocket and two satellites into the asteroid resonant orbit, first the observer flies over the asteroid to perform close-range approach detection; then the impactor hits the asteroid; finally the observer revisits the flyby to evaluate the effect. After that, the observer revisits the flyby of the asteroid every other resonance cycle to obtain more accurate and reliable data. This method uses a rocket to complete the "detection-impact-evaluation" mission in 3 to 5 years, and can achieve low-cost and high-efficiency completion of the kinetic impact on-orbit verification mission.

Description

Technical Field

[0001] The present invention belongs to the field of aerospace asteroid defense, and in particular relates to a near-Earth asteroid kinetic impact orbit deflection method based on a resonance orbit. Background Art

[0002] The on-orbit verification mission of kinetic impact defense against asteroids usually requires multiple visits to the asteroid at intervals to conduct detection, impact, and evaluation. It often requires the use of multiple launch vehicles to launch multiple probes, which places high demands on the transportation cost and mission orbit.

[0003] At present, there is only one specially designed mission in the world to verify the orbit deflection of an asteroid by kinetic impact. The AIDA (Asteroid Impact & Deflection Assessment) mission jointly implemented by the United States and Europe uses two launch vehicles to launch two spacecraft to complete the impact (DART) and effect evaluation (HERA). It uses ground-based optical telescopes and radars to measure the changes in luminosity and the orbit of the asteroid, releases a CubeSat on orbit to monitor the impact process, and the HERA mission launched separately to conduct a detailed evaluation of the momentum transfer and orbit deflection effects.

[0004] However, this mission mode has the following problems: First, it is only applicable to large-sized asteroids with rich characteristic information. For a large number of small-sized asteroids, it is impossible to obtain their detailed characteristic information using ground-based equipment. If effective characteristic information cannot be obtained before the impact, it is difficult to ensure accurate hitting of the asteroid, and it is not conducive to the evaluation of momentum transfer and orbit deflection effects. Second, due to the limited carrying capacity of the rocket and the satellite's orbit change capability at the time, and the fact that it takes a certain amount of time for the asteroid to have a more obvious deflection effect after the impact, it is difficult for one rocket to complete the requirements of impact deflection and effect evaluation. Two rockets are required to carry out the impact and effect evaluation tasks respectively at an interval of time, which is costly. Summary of the invention

[0005] The purpose of the present invention is to overcome the shortcomings of traditional kinetic impact on-orbit verification and evaluation missions, such as the need for multiple rocket launches, high cost, inability to provide sufficient characteristic information support for precise impacts on small asteroids without close-in detection before impact, and difficulty in accurately evaluating momentum transfer factors and orbital deflection effects through direct kinetic impacts.

[0006] In order to achieve the above-mentioned purpose, the present invention proposes a near-Earth asteroid kinetic impact orbit deflection method based on a resonant orbit, which is characterized by using a pock-chop diagram to determine the launch window, setting the range of optimization variables, calculating the orbits of the observer and the impactor according to the dynamic model, and obtaining the position and velocity parameters of the mission node; optimizing the model combining a genetic algorithm with a local search to obtain the mission orbit model with the maximum asteroid deflection distance, and determining the launch time and the impact time.

[0007] As an improvement of the above method, the dynamic model considers a two-body gravitational model centered on the sun and a third-body gravitational perturbation model including eight planets, Pluto and the moon. The calculation formula is as follows:

[0008]

[0009] Where r is the distance from the heliocenter of the spacecraft or asteroid; u is the gravitational constant; d i is the heliocentric distance of the ith planet; r′ i represents the position vector from the i-th planet to the spacecraft or asteroid; i ranges from 1 to 10, representing Mercury, Venus, Earth, Mars, Jupiter, Saturn, Uranus, Neptune, Pluto and the Moon respectively; a RE It represents the perturbation acceleration caused by the relativistic effect, and is the correction of Newtonian mechanics by general relativity in high-speed problems.

[0010] As an improvement of the above method, the kinetic model is adopted, and the optimization variable selected is the launch time t d and the impact time t i ;

[0011]

[0012] The optimization target is the deflection distance r of the asteroid CA , the constraints consider the launch C3, the spacecraft mass m at impact sc (t i ), the launch site declination De and the asteroid apparent magnitude V at the time of impact:

[0013]

[0014] subject to C3<50,De>28.5,m sc (t i )>750,V<24

[0015] After determining the launch window using the pock-chop diagram, set the optimization variable t with reference to the launch window d and t iThe following penalty function is set in the model to constrain the optimization: Considering the rocket carrying capacity constraint, the launch C3 cannot exceed 50km 2 / s 2 Considering the launch site constraints, the launch declination cannot be less than 28.5°; considering the deflection effect, the impactor mass cannot be less than 750kg; considering ground observability, the asteroid magnitude at the time of impact cannot be greater than 24; Under the above constraints, optimize t d and t i , find the solution with the largest deflection distance;

[0016] An optimization algorithm combining genetic algorithm with local optimal search is adopted. In the genetic algorithm, a local optimal search is performed in the interval where the best individual is located every several generations, and the best individual in this generation is replaced by the result of the local optimal search, and is optimized as the parent in the next generation of genetic algorithm.

[0017] As an improvement of the above method, the specific process of the optimization algorithm combining the genetic algorithm with the local optimal search is as follows:

[0018] In the first step, the initial population is generated within the given range of the optimization variable;

[0019] The second step is to determine whether the constraints are met. If they are met, continue; if they are not met, directly use the penalty function to calculate the optimization target;

[0020] The third step is to use a high-precision orbit recursion model to calculate the near-Earth distance of the asteroid after deflection;

[0021] The fourth step is to determine whether the number of generations is a multiple of 5. If so, a local optimal search is performed, and the optimal individuals of this generation are replaced and output as the parent population of the next generation; otherwise, the next generation calculation is directly entered;

[0022] The fifth step is to determine whether the termination conditions are met. If yes, stop and output the results; if not, perform a new round of optimization.

[0023] As an improvement on the above method, after the launch window is determined, the orbit transfer model is used to determine the mission node times for detection, impact, and evaluation.

[0024] As an improvement to the above method, the orbit transfer model is designed by using the conic section splicing method.

[0025] As an improvement of the above method, the method includes:

[0026] 1. Use the pock-chop diagram to determine the launch window;

[0027] 2. Launch two satellites into a resonant orbit with one rocket: Launch an impactor and an observer into the asteroid's resonant orbit with one rocket;

[0028] 3. The observer flies over the asteroid for detection;

[0029] 4. The impactor hits the asteroid, and the observer monitors the impact process;

[0030] 5. The observer revisits the flyby and evaluates the impact effect. Every other resonance period, the observer revisits the flyby of the asteroid for observation and evaluation.

[0031] The present invention also provides a computer device, comprising a memory, a processor, and a computer program stored in the memory and executable on the processor, wherein the processor implements any of the above methods when executing the computer program.

[0032] The present invention also provides a computer-readable storage medium, wherein the computer-readable storage medium stores a computer program, and when the computer program is executed by a processor, the processor executes any of the methods described above.

[0033] Compared with the prior art, the advantages of the present invention are:

[0034] 1. Asteroid resonant orbit design method based on pock-chop diagram, which can quickly lock the launch window that can realize the resonant orbit and reduce the amount of calculation;

[0035] 2. A method for close-in detection, kinetic impact and deflection assessment of near-Earth asteroids based on a resonant orbit is proposed. The "detection-impact-assessment" mission can be completed in 3 to 5 years using a single rocket, which can achieve low-cost and high-efficiency on-orbit verification of kinetic impact;

[0036] 3. The mission will conduct close-in detection of the asteroid before impact, providing sufficient characteristic information support for accurate impact of small-sized asteroids, and will be able to know the asteroid's size, density and other physical properties in advance, which is conducive to accurately calculating the kinetic energy transfer factor and evaluating the deflection effect.

[0037] 4. During the impact, space-based observers monitor the impact process, while ground-based telescopes can also observe the impact situation. After the impact, joint observation between the ground and the sky can be achieved, with space-based observers revisiting and flying over the object regularly, and ground-based telescopes observing when they are visible, which is conducive to accurately evaluating the deflection effect. BRIEF DESCRIPTION OF THE DRAWINGS

[0038] Figure 1 Shown is a schematic diagram of the operation of the method of impacting a near-Earth asteroid;

[0039] Figure 2 Shown is a flow chart of the method for impacting a near-Earth asteroid;

[0040] Figure 3 Shown is a pock-chop image of asteroid 2019 VL5;

[0041] Figure 4 Shown is a pock-chop image of asteroid 2005 ES70;

[0042] Figure 5 Shown is a schematic diagram of the transfer model;

[0043] Figure 6 Shown is the optimization algorithm flow chart. DETAILED DESCRIPTION

[0044] The present invention addresses the problems that traditional kinetic energy on-orbit verification missions require multiple rocket launches, which are costly, the lack of close-in detection before impact, which cannot provide sufficient characteristic information support for accurate impacts with small-sized asteroids, and the difficulty in accurately evaluating momentum transfer factors and orbit deflection effects through direct kinetic impacts. By utilizing the characteristics of a resonant orbit, the present invention realizes a mission mode of "probe (close-in detection to obtain physical and chemical properties) - impact (kinetic impact to deflect the asteroid orbit) - evaluation (joint space-ground effect evaluation)" that can be completed by a single rocket in a relatively short period of time.

[0045] The technical solution of the present invention is described in detail below with reference to the accompanying drawings.

[0046] like Figure 1 and Figure 2 As shown, the method of the present invention is specifically:

[0047] An impactor and an observer will be launched from Earth with one rocket into the asteroid's resonant orbit. The mission will first fly over the asteroid and conduct close-range detection. Visible light / infrared cameras and spectrometers will be used to obtain the asteroid's precise orbit, topography, spectral type, reflectivity, size, shape, rotation and other information. A cubesat can then be released to measure the asteroid's mass.

[0048] Then the impactor and observer separate before impact. The impactor impacts, and the observer monitors the impact process, obtains the precise orbit of the asteroid, conducts imaging and spectral detection, and obtains the crater morphology and the distribution of ejecta and material composition.

[0049] Finally, the observer will revisit the flyby to obtain the precise orbit, crater morphology, and sputtering distribution, and evaluate the momentum transfer factor and orbit deflection effect. After that, the observer will revisit the flyby every other resonance period to obtain more accurate and reliable data, which will help to more accurately evaluate the momentum transfer factor and orbit deflection effect.

[0050] In the present invention, the design of the resonant orbit first requires drawing the pock-chop diagram of the asteroid, finding a preliminary launch window based on the pock-chop diagram and judging whether the asteroid is suitable as a mission target. After determining the launch window of the mission target, the orbit transfer model is used to determine the mission nodes of detection, impact and evaluation, and a suitable dynamic model is selected to calculate each orbit to obtain parameters such as the position and velocity at the mission node. With reference to the preliminary launch window determined by the pock-chop diagram, the range of the optimization variable is set, and the model combining the genetic algorithm and the local search is used for optimization to obtain the mission orbit model with the maximum deflection distance of the asteroid under the constraints of carrying capacity and the like.

[0051] The calculation model used in the present invention is as follows:

[0052] 1) Draw a pock-chop diagram of the asteroid

[0053] For asteroids with an orbital period that resonates with the Earth's revolution period, low orbital uncertainty, and more observational data, and with a size between 20 and 50 meters, a pock-chop diagram is drawn based on the spacecraft launch time, arrival time at the asteroid, launch of C3, transfer orbit period, and the speed of the spacecraft relative to the asteroid when flying over the asteroid, to find a preliminary launch window. Based on the launch window and the known physical properties and orbital parameters of the asteroid, it is determined whether the asteroid is suitable as a target.

[0054] Taking asteroids 2019VL5 and 2005ES70 as examples, the launch time is January 2025-December 2026, the transfer time is 50-365 days, and the pock-chop diagrams with resonance ratios with the Earth are 1:1 and 3:2 respectively. Figure 3 and Figure 4 .

[0055] 2) Orbital transfer model

[0056] The transfer model of the present invention is specifically as follows Figure 5 As shown: at the launch moment, the spacecraft escapes from the Earth and enters the asteroid resonance orbit, flies over the asteroid for the first time at the flyby detection moment, and hits the asteroid at the impact moment after a resonance cycle. At the same time, the observer monitors the impact process, and then after another resonance cycle, the observer revisits the flyby of the asteroid to evaluate its deflection effect.

[0057] The present invention adopts the conic section splicing method to design the orbit transfer model. The basic idea of ​​the conic section splicing method is to simplify the orbit of the spacecraft into several two-body problems according to the boundary of the influence sphere for preliminary design. Within the influence sphere of each planet, it is considered that it is only affected by the gravity of the planet. The spacecraft moves around the planet. At the boundary of the influence sphere, the connection of the two trajectories is unified by taking the terminal state of the previous segment as the initial state of the next segment. The conic section splicing method is usually used for the preliminary design of deep space orbits, which can achieve the purpose of improving design efficiency. Figure 5 As shown, the present invention solves the Lambert problem in the phase from Earth launch to flyby detection, and the initial and final velocities can be directly solved from the initial and final positions and the transfer time. There are many common simplified numerical solution methods. The designs of other mission phases of the present invention are recursively deduced using a high-precision orbital dynamics model.

[0058] 3) Orbital dynamics model

[0059] The present invention adopts a high-precision orbital dynamics numerical model, which mainly considers the two-body gravitational model centered on the sun; the third-body gravitational perturbation model including the eight planets and Pluto and the moon, and the relativistic effect. The high-precision dynamics model expression is as follows:

[0060]

[0061] where r is the distance from the heliocenter of the spacecraft or asteroid; u is the gravitational constant; d i is the heliocentric distance of the ith planet; r′ i represents the position vector from the i-th planet to the spacecraft or asteroid; i ranges from 1 to 10, representing Mercury, Venus, Earth, Mars, Jupiter, Saturn, Uranus, Neptune, Pluto and the Moon respectively; a RE It represents the perturbation acceleration caused by the relativistic effect, which is the correction of Newtonian mechanics by general relativity in high-speed problems. The ephemeris of the planet and the moon are obtained from the JPL DE430 ephemeris. The initial position of the asteroid refers to the data published on the JPL website (https: / / www.jpl.nasa.gov / ), and the position at the time of impact is obtained by orbital recursion from the initial position.

[0062] 4) Optimize the model

[0063] Using the kinetic model described above, the optimization variable selected by the present invention is the launch time t d and the impact time t i ;

[0064]

[0065] The optimization target is the deflection distance r of the asteroid CA , the constraints consider the launch C3, the spacecraft mass m at impactsc (t i ), the launch site declination De and the asteroid apparent magnitude V at the time of impact:

[0066]

[0067] subject to C3<50,De>28.5,m sc (t i )>750,V<24

[0068] After the launch window is preliminarily determined using the pock-chop diagram, the optimization variable t is set with reference to the launch window. d and t i The following penalty function is set in the model to constrain the optimization: Considering the rocket carrying capacity constraint, the launch C3 cannot exceed 50km 2 / s 2 ; Considering the launch site constraints, the launch declination cannot be less than 28.5°; considering the deflection effect, the impactor mass cannot be less than 750kg; considering ground observability, the asteroid magnitude at the time of impact cannot be greater than 24. Under the above constraints, optimize t d and t i , find the solution with the largest deflection distance.

[0069] The present invention adopts the optimization algorithm of genetic algorithm + local optimal search. In the genetic algorithm, the optimal individual interval is searched locally every 5 generations, and the optimal individual in the current generation is replaced by the result of the local optimal search, and is optimized as the parent in the next generation of genetic algorithm. Figure 6 As shown:

[0070] The specific process is:

[0071] In the first step, the initial population is generated within the given range of the optimization variable;

[0072] The second step is to determine whether the constraints are met. If they are met, continue; if they are not met, directly use the penalty function to calculate the optimization target;

[0073] The third step is to use a high-precision orbit recursion model to calculate the near-Earth distance of the asteroid after deflection;

[0074] The fourth step is to determine whether the number of generations is a multiple of 5. If so, a local optimal search is performed, and the optimal individuals of this generation are replaced and output as the parent population of the next generation; otherwise, the next generation calculation is directly entered;

[0075] The fifth step is to determine whether the termination conditions are met. If yes, stop and output the results; if not, perform a new round of optimization.

[0076] After completing the impact detection and assessment of the asteroid, the kinetic energy transfer factor can be calculated using the following formula:

[0077]

[0078] where Δv Ast is the change in the asteroid's velocity before and after the impact, β is the kinetic energy transfer factor, m KI is the impactor mass, m Ast is the mass of the asteroid, v KI and v Ast are the speeds of the impactor and the asteroid at the time of impact, respectively. In an ideal completely elastic collision, β is usually taken as 1; however, in reality, different types of asteroids have different materials, surface shapes, and distribution of ejecta after impact. It is necessary to obtain the precise mass (density), rotation information, surface shape and other parameters of the asteroid through flyby detection before the impact, and revisit the flyby asteroid after the impact to obtain information such as its ejecta distribution and orbital deflection to calculate the precise kinetic energy transfer factor.

[0079] The present invention aims at the problems that the traditional kinetic energy on-orbit verification mission requires multiple rockets with high cost, the close detection before impact cannot provide sufficient characteristic information support for the precise impact of small-sized asteroids, and the direct kinetic energy impact is difficult to accurately evaluate the momentum transfer factor and orbit deflection effect. A near-Earth asteroid close detection, kinetic energy impact and deflection evaluation method based on a resonant orbit is proposed. First, the asteroid pock-shop diagram is drawn to quickly lock the launch window. Then, the observer and the impactor are launched into the asteroid resonant orbit with one rocket and two satellites. The observer performs flyby detection on the asteroid before the impact, and knows the physical properties of the asteroid such as size and density in advance, which is conducive to the precise impact of small-sized asteroids, and is convenient for accurately analyzing and calculating the key parameter "kinetic energy transfer factor" of the mission and evaluating the deflection effect. After separation from the observer, the impactor impacts the asteroid, and the observer monitors the impact process and records data. After the impact, the observer revisits the flyby asteroid regularly to evaluate the impact effect. This mission mode only requires one rocket, and the "detection-impact-evaluation" can be completed within 3 to 5 years, which can complete the kinetic energy impact on-orbit verification mission at low cost and high efficiency. The present invention establishes a high-precision orbital dynamics model and an optimization model of genetic algorithm and local search, and performs simulation optimization and performance evaluation on the mission. The simulation results show that, taking the constraints of carrier and launch site into consideration, the resonance ratios of asteroids 2019VL5 and 2005ES70 with the Earth are 1:1 and 3:2 respectively. The former uses the Long March 3C launch vehicle (launch C3 is 1.37 km 2 / s 2, corresponding to a carrying capacity of 2400kg), it can complete a "exploration-impact-assessment" mission within 3 years, with an impact speed of up to 9km / s. The asteroid can be deflected up to 7745km during the first flyby revisit after the impact; the latter uses the Long March 3B carrier rocket (9.4km for launching C3 2 / s 2 , corresponding to a carrying capacity of 2700kg), it can complete a "exploration-impact-assessment" mission within 5 years, with an impact speed of up to 10km / s, and the asteroid deflection distance can reach 804km during the first flyby revisit after the impact. In summary, the near-Earth asteroid close detection, kinetic impact and deflection assessment based on the resonant orbit is a mission mode that can improve mission efficiency and make up for the lack of ground-based observations.

[0080] The present invention may also provide a computer device, comprising: at least one processor, a memory, at least one network interface and a user interface. The various components in the device are coupled together through a bus system. It is understood that the bus system is used to achieve connection and communication between these components. In addition to the data bus, the bus system also includes a power bus, a control bus and a status signal bus.

[0081] The user interface may include a display, a keyboard, or a pointing device (eg, a mouse, a trackball, a touch pad, or a touch screen).

[0082] It is understood that the memory in the disclosed embodiments of the present application may be a volatile memory or a non-volatile memory, or may include both volatile and non-volatile memories. Among them, the non-volatile memory may be a read-only memory (ROM), a programmable read-only memory (PROM), an erasable programmable read-only memory (EPROM), an electrically erasable programmable read-only memory (EEPROM), or a flash memory. The volatile memory may be a random access memory (RAM), which is used as an external cache. By way of example and not limitation, many forms of RAM are available, such as static random access memory (SRAM), dynamic random access memory (DRAM), synchronous dynamic random access memory (SDRAM), double data rate synchronous dynamic random access memory (DDRSDRAM), enhanced synchronous dynamic random access memory (ESDRAM), synchronous link dynamic random access memory (SLDRAM), and direct memory bus random access memory (DRRAM). The memories described herein are intended to include, but are not limited to, these and any other suitable types of memories.

[0083] In some embodiments, the memory stores the following elements, executable modules or data structures, or a subset thereof, or an extended set thereof: an operating system and applications.

[0084] The operating system includes various system programs, such as a framework layer, a core library layer, a driver layer, etc., which are used to implement various basic services and process hardware-based tasks. The application includes various application programs, such as a media player (Media Player), a browser (Browser), etc., which are used to implement various application services. The program for implementing the method of the embodiment of the present disclosure can be included in the application.

[0085] In the above embodiment, the processor may also call a program or instruction stored in the memory, specifically, a program or instruction stored in an application program, and is used to:

[0086] Execute the steps of the above method.

[0087] The above method can be applied to a processor or implemented by a processor. The processor may be an integrated circuit chip with signal processing capabilities. In the implementation process, each step of the above method can be completed by an integrated logic circuit of hardware in the processor or an instruction in the form of software. The above processor may be a general processor, a digital signal processor (Digital Signal Processor, DSP), an application specific integrated circuit (Application Specific Integrated Circuit, ASIC), a field programmable gate array (Field Programmable Gate Array, FPGA) or other programmable logic devices, discrete gates or transistor logic devices, discrete hardware components. The above-disclosed methods, steps and logic block diagrams can be implemented or executed. The general processor may be a microprocessor or the processor may also be any conventional processor, etc. The steps of the above-disclosed method can be directly embodied as a hardware decoding processor to execute, or the hardware and software modules in the decoding processor are combined to execute. The software module can be located in a mature storage medium in the field such as a random access memory, a flash memory, a read-only memory, a programmable read-only memory or an electrically erasable programmable memory, a register, etc. The storage medium is located in the memory, and the processor reads the information in the memory and completes the steps of the above method in combination with its hardware.

[0088] It is understood that the embodiments described in the present invention can be implemented by hardware, software, firmware, middleware, microcode or a combination thereof. For hardware implementation, the processing unit can be implemented in one or more application specific integrated circuits (ASIC), digital signal processors (DSP), digital signal processing devices (DSPD), programmable logic devices (PLD), field programmable gate arrays (FPGA), general processors, controllers, microcontrollers, microprocessors, other electronic units for performing the functions described in the present application or a combination thereof.

[0089] For software implementation, the technology of the present invention can be implemented by executing the functional modules (such as procedures, functions, etc.) of the present invention. The software code can be stored in a memory and executed by a processor. The memory can be implemented in the processor or outside the processor.

[0090] The present invention may also provide a non-volatile storage medium for storing a computer program. When the computer program is executed by a processor, each step in the above method embodiment can be implemented.

[0091] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention and are not intended to limit the present invention. Although the present invention is described in detail with reference to the embodiments, it should be understood by those skilled in the art that any modification or equivalent replacement of the technical solutions of the present invention does not depart from the spirit and scope of the technical solutions of the present invention and should be included in the scope of the claims of the present invention.

Claims

1. A method for deflecting the orbit of a near-Earth asteroid kinetic impact based on a resonance orbit, characterized in that: The launch window is determined using the pock-chop diagram, the range of the optimization variables is set, the orbits of the observer and the impactor are calculated based on the dynamic model, and the position and velocity parameters of the mission node are obtained; the model combining the genetic algorithm and the local search is used for optimization to obtain the mission orbit model with the maximum asteroid deflection distance, and the launch time and impact time are determined; The dynamic model considers a two-body gravitational model centered on the sun and a third-body gravitational perturbation model including eight planets, Pluto, and the moon. The calculation formula is as follows: Where r is the distance from the heliocenter of the spacecraft or asteroid; u is the gravitational constant; d i is the heliocentric distance of the ith planet; r' i represents the position vector from the i-th planet to the spacecraft or asteroid; i ranges from 1 to 10, representing Mercury, Venus, Earth, Mars, Jupiter, Saturn, Uranus, Neptune, Pluto, and the Moon respectively; a RE It represents the perturbation acceleration caused by the relativistic effect, which is the correction of Newtonian mechanics by general relativity in high-speed problems. Using the dynamic model, the optimization variable selected is the launch time t d and the impact time t i ; The optimization target is the asteroid's near-Earth distance r CA , the constraints consider the launch C3, the spacecraft mass m at impact sc (t i ), the launch site declination De, the asteroid apparent magnitude V at the time of impact: subject to C3<50,De>28.5,m sc (t i )>750,V<24 After determining the launch window using the pock-chop diagram, set the optimization variable t with reference to the launch window d and t i The following penalty function is set in the model to constrain the optimization: Considering the rocket carrying capacity constraint, the launch C3 cannot exceed 50km 2 / s 2 Considering the launch site constraints, the launch declination cannot be less than 28.5°; considering the deflection effect, the impactor mass cannot be less than 750kg; considering ground observability, the asteroid magnitude at the time of impact cannot be greater than 24; Under the above constraints, optimize t d and t i , find the solution with the largest deflection distance; The optimization algorithm combines genetic algorithm with local optimal search. In the genetic algorithm, the local optimal search is performed in the interval where the best individual is located every several generations. The best individual in this generation is replaced by the result of the local optimal search and optimized as the parent in the next generation of genetic algorithm. The specific process of the optimization algorithm combining the genetic algorithm with the local optimal search is as follows: In the first step, the initial population is generated within the given range of the optimization variable; The second step is to determine whether the constraints are met. If they are met, continue; if they are not met, directly use the penalty function to calculate the optimization target; The third step is to use a high-precision orbit recursion model to calculate the near-Earth distance of the asteroid after deflection; The fourth step is to determine whether the number of generations is a multiple of 5. If so, a local optimal search is performed, and the optimal individuals of this generation are replaced and output as the parent population of the next generation; otherwise, the next generation calculation is directly entered; The fifth step is to determine whether the termination conditions are met. If yes, stop and output the results; if not, perform a new round of optimization; The method comprises: (1) Determine the launch window using the pock-chop diagram; (2) Launching two satellites into a resonant orbit with one rocket: Launching an impactor and an observer into a resonant orbit with one rocket; (3) An observer flies over an asteroid to conduct exploration; (4) The impactor strikes the asteroid, and the observer monitors the impact process; (5) The observer revisits the flyby and evaluates the impact effect. Every other resonance period, the observer revisits the flyby of the asteroid for observation and evaluation.

2. The near-Earth asteroid kinetic impact orbit deflection method based on a resonance orbit according to claim 1, characterized in that: After determining the launch window, the orbit transfer model is used to determine the mission node times for detection, impact, and evaluation.

3. The near-Earth asteroid kinetic impact orbit deflection method based on a resonance orbit according to claim 2, characterized in that: The track transfer model is designed using a conic section splicing method.

4. A computer device comprising a memory, a processor and a computer program stored in the memory and executable on the processor, characterized in that: When the processor executes the computer program, the method according to any one of claims 1 to 3 is implemented.

5. A computer-readable storage medium, characterized in that: The computer-readable storage medium stores a computer program, which, when executed by a processor, causes the processor to perform the method according to any one of claims 1 to 3.

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