A device and method for verifying orbital deflection caused by laser ablation of asteroids in orbital space

Abstract

The present application provides an apparatus and method for verifying orbital deflection caused by laser ablation of asteroids in orbital space, the apparatus comprising: a meteorite ball and an apparatus body; wherein the meteorite ball is used to simulate an asteroid; the apparatus body comprises: a pulse laser for emitting laser light to the meteorite ball; a clamping and releasing spinning mechanism for rotating the meteorite ball at a set speed; a laser rangefinder for measuring the distance and angle between the apparatus body and the meteorite ball; an attitude and orbit control system for adjusting the distance and angle between the apparatus body and the meteorite ball according to the measurement results of the laser rangefinder; a visible light camera for taking photos or recording videos of the meteorite ball, monitoring the ablation of the meteorite ball and the relative angle between the apparatus body and the meteorite ball; an accelerometer for measuring and recording acceleration data of the motion of the apparatus body. The advantages of the present application are: the experimental scene is similar to the scene environment of the actual application of orbital deflection caused by laser ablation of asteroids; and it has strong feasibility for engineering implementation.

Description

Technical Field

[0001] The present application belongs to the field of asteroid defense, and specifically relates to a device and method for in-orbit space verification of laser ablation of an asteroid to cause orbital deflection. Background Art

[0002] There are a large number of asteroids in the solar system, among which near-Earth asteroids refer to those with a perihelion distance q less than 1.3AU (1AU = 1.496×10 8 km), an asteroid whose orbit is close to or intersects with the Earth. Near-Earth asteroid impacts are a major catastrophic threat facing human society. In history, there have been many near-Earth asteroid impacts on the Earth, causing changes in the Earth's climate and environment, and even triggering species extinction. The scientific community generally believes that 65 million years ago, a near-Earth asteroid with a diameter of about 10km hit the Earth, leading to the extinction of 75% of the world's species, including dinosaurs. In the history of human development, many regional asteroid impacts have brought great harm to human production and life. The Younger Dryas event 12,000 years ago may have led to the extinction of the Clovis people in North America; the meteorite shower event in Qingyang, Gansu in the Ming Dynasty may have caused large-scale casualties; the Tunguska explosion in 1908 burned more than 2000m 2 The 2013 Chelyabinsk incident injured nearly 1,500 people and damaged 3,000 houses. From a statistical point of view, the possibility of near-Earth asteroid impacts in the future is very high. Asteroids of 5 to 10 meters will form fireballs when they hit the Earth, which will happen almost every year. Asteroids of 20 meters will explode in the atmosphere, causing disasters at the town level, which will happen about once every 100 years. Asteroids of 50 meters will cause large-scale city-level disasters when they hit the Earth, which will happen about once every 1,000 years. Asteroids of 140 meters in diameter will cause small-scale national disasters when they hit the Earth, which will happen about once every 20,000 years. Carrying out asteroid defense is an inevitable requirement for the continuation of human civilization. It is necessary to conduct systematic, forward-looking and targeted research on asteroid impact events from the perspectives of probability and time of occurrence, size and scale of asteroids, and technical characteristics of defense measures.

[0003] Asteroid defense technology can be divided into three categories according to the length of warning time and the size of the target asteroid: (1) For asteroids with short warning time and large size, use nuclear explosions to destroy or change their orbits; (2) For asteroids with long warning time and large size or short warning time and small size, use spacecraft to directly impact and change their orbits; (3) For asteroids with very long warning time, use long-term forces to change their orbits, that is, use contact or non-contact forces to cause a small speed change, which gradually and slowly evolves into a larger orbit change.

[0004] Laser ablation asteroid driving technology is a long-term defense technology. The basic principle is to use laser to irradiate the surface of the asteroid, and use the reaction force brought by the ablation and vaporization of the surface material or even the plasma jet to change its orbit, thereby achieving the defense purpose. Compared with other defense technologies, its advantages are: (1) It is flexible and can be deployed in low Earth orbit, geosynchronous orbit, the moon or the Earth-Moon Lagrange point; (2) It can act for a long time and at a long distance to avoid collision between spacecraft and asteroids; (3) Since the source of the reaction force is the material of the asteroid itself, the energy acquisition is almost infinite; (4) The unique energy transfer process of the laser will not change the orbit of the spacecraft itself, reducing the requirements for orbit maintenance control, and the accuracy of the on-orbit control and disposal of asteroids is relatively high; (5) It can avoid the disintegration of asteroids and prevent the threat of secondary impact; (6) Especially in terms of defending asteroids with long warning time, compared with other long-term defense technologies, it does not need to consider the attachment of asteroids, avoiding the technical risks of complex mechanisms such as physical connection and landing equipment, and has a more efficient deflection efficiency.

[0005] At present, domestic and foreign scholars have carried out pulsed laser ablation experiments on the ground, mainly focusing on ablation fragments experiments and ablation experiments with asteroid simulants (meteorites, etc.) as targets. The efficiency of pulsed laser ablation drive is determined by the impulse of the backjet hot material flow generated during laser loading, and the impulse depends on the impulse coupling coefficient of the laser irradiated object material. The impulse coupling coefficient reflects the ability to obtain the impulse of the object by consuming unit laser energy, and is an important parameter to characterize the efficiency of laser ablation drive. Assume that the energy of a single laser pulse is E. For a regular plane target, the impulse coupling coefficient C m C m =I A / E, where: I A The impulse generated by a single laser pulse. In a vacuum environment, the impulse generated by pulsed laser irradiation and ablation of target materials is about 10 to 100×10 -6 N·s level, so the measurement of tiny impulses becomes a major difficulty in evaluating the driving efficiency of pulsed laser ablation. At present, there are three main methods for experimentally measuring impulse coupling coefficients, namely impulse pendulum method, air cushion guide method and torsion pendulum interferometry. Impulse pendulum measurement is the most widely used method for testing tiny impulses. This method is based on the principle of laser interference. It generates an optical path difference in the two optical paths of the interference device to form an interference ring with alternating light and dark. The impulse transmitted to the material target by the laser is calculated by measuring the number of interference fringes within a specific period of the impulse pendulum swing. The air cushion guide method calculates the impulse by directly measuring the moving distance of the effect target within a specific time after laser loading, and reduces the friction resistance of the effect target during movement through the air cushion guide. The torsion pendulum interferometry method converts the measurement of the backjet impulse into the measurement of the micro-vibration of the torsion pendulum, and indirectly calculates the size of the impulse by giving the angle change of the torsion pendulum rotation through the laser interference method.

[0006] At present, there has been a relatively extensive experimental study on the impulse coupling coefficient of different types of material objects in a simulated vacuum environment. The impulse coupling coefficient can be obtained by calculating the impulse generated by laser ablation of the object surface through experiments. Limited by the output energy of laboratory pulsed lasers, the magnitude of the impulse generated by ablation is generally on the order of 10 –5 N·s or even smaller, so a special torsion pendulum type micro-impulse device is needed for measurement. In the experiment, asteroid meteorites are selected as test samples to obtain results close to the actual ablation of asteroids. Ground experiment results show that the impulse coupling coefficient C m has a maximum relationship with the laser power density I injected into the surface: when I < I0 (I0 is the optimal laser power density), the surface of the meteorite undergoes gasification until plasma formation, and the gasification mechanism plays a major role in most intervals, and the impulse coupling coefficient increases non-linearly with the increase of the power density; when I > I0, the surface plasma formation mechanism plays a major role. As the power density continues to increase, the shielding effect of the plasma on the laser energy becomes more obvious, and the impulse coupling coefficient decreases slowly with the increase of the power density. Therefore, C m = f(I) is a piecewise complex non-linear function. American scholars such as Phipps et al. have carried out a large number of impulse coupling coefficient measurement experiments on laser ablation of space debris, and through experimental data analysis, they have proposed a curve fitting method, that is:

[0007]

[0008] Based on this general form, combined with the discrete point values of specific experimental data, the key parameters A (control quantity), B (control quantity), I0 (optimal laser power density), C max (the maximum impulse coupling coefficient corresponding to the ablation material) can be obtained, so as to give the specific function model of the impulse coupling coefficient of asteroid meteorites.

[0009] Up to now, domestic and foreign scholars have achieved a large number of research results in the aspects of laser ablation driving law, theoretical model of action effect and engineering implementation plan.

[0010] However, due to the extremely weak effect of single-pulse laser, complex 10 -6 N·s magnitude micro-impulse measurement devices are often used in domestic and foreign ground experiments or scheme designs, or the scale of the laser system is increased to the kilojoule level and the meter-level emission aperture to amplify the action effect, resulting in difficulties in carrying out low-cost technical experiments in orbit, so that all the currently publicly known work is based on ground experiment data, and there is no laser ablation asteroid in-orbit technology experiment globally. As a result, the key parameters A, B, I0, C in the impulse coupling coefficient function model, which is公认 as the most core of the laser ablation effect internationally maxThe determination is still obtained using the fitting of ground experimental data, and the applicability of laser ablation-driven defense technology in the real environment of space has not yet been tested. Summary of the invention

[0011] The purpose of this application is to overcome the defect that the existing technology cannot conduct low-cost technical experiments in orbit.

[0012] In order to achieve the above-mentioned purpose, the present application proposes a device for verifying orbital deflection caused by laser ablation of an asteroid in orbit, the device comprising: a meteorite ball and a device body; wherein,

[0013] The meteorite ball is a hollow sphere used to simulate an asteroid;

[0014] The device body comprises:

[0015] A pulse laser, used for emitting laser light toward the meteorite ball according to a set rule;

[0016] A clamping and releasing spinning mechanism is used to rotate the meteorite ball at a set rotation speed;

[0017] A laser rangefinder, used to measure the distance and angle between the device body and the meteorite ball;

[0018] An attitude and orbit control system, used for adjusting the distance and angle between the device body and the meteorite ball according to the measurement results of the laser rangefinder;

[0019] A visible light camera, used to take photos or videos of the meteorite ball, monitor the ablation of the meteorite ball and the relative angle between the device body and the meteorite ball;

[0020] The accelerometer is used to measure and record the acceleration data of the movement of the device body.

[0021] As an improvement of the above device, the clamping, releasing and rotating mechanism includes a clamping, releasing system and a rotating system;

[0022] The clamping and releasing system is used to clamp the meteorite ball and release the meteorite ball when the meteorite ball reaches a set rotation speed;

[0023] The spinning system is used to drive the meteorite ball to rotate at a set speed.

[0024] As an improvement of the above device, the clamping and releasing system comprises: a first micro-stepping motor, a first clamping movable rod, a second clamping movable rod, a rotating gear, a first bearing clamping rotating block and a second bearing clamping rotating block; wherein,

[0025] The first micro-stepping motor is arranged inside the device body to drive the rotating gear to rotate;

[0026] One end of the first clamping movable rod and the second clamping movable rod are both located inside the device body and clamp the rotating gear, and the rotating gear drives the two clamping movable rods to move closer to or away from each other; the other ends of the first clamping movable rod and the second clamping movable rod are both extended out of the device body;

[0027] The first bearing clamping rotating block is arranged at one end of the first clamping movable rod outside the device body; the second bearing clamping rotating block is arranged at one end of the second clamping movable rod outside the device body; the first bearing clamping rotating block and the second bearing clamping rotating block clamp the meteorite ball.

[0028] As an improvement of the above device, the spinning system includes a second micro-stepping motor;

[0029] The second micro-stepping motor is fixed to one end of the first clamping movable rod outside the device body, and the rotating shaft thereof passes through the first clamping movable rod and is fixedly connected to the first bearing pressing rotating block, thereby driving the first bearing pressing rotating block to rotate;

[0030] The axis of the rotating shaft of the second micro-stepping motor is perpendicular to the movement direction of the device.

[0031] As an improvement of the above device, the outer diameter of the meteorite ball is ≥42mm and the wall thickness is ≥0.214mm.

[0032] The present application also provides a method for verifying orbital deflection caused by laser ablation of an asteroid in orbital space, which is implemented based on the above-mentioned device, and the method includes:

[0033] Step 1: placing the device in a geocentric orbit in space;

[0034] Step 2: The attitude and orbit control system adjusts the attitude of the device so that the meteorite ball faces the wind and the overall roll axis of the device is parallel to the flight direction;

[0035] Start the clamping and releasing spinning mechanism to make the meteorite ball spin at the set speed, and then release it in place;

[0036] Turn on the laser rangefinder and visible light camera, use the attitude and orbit control system to aim at the center of the meteorite ball in real time, and maintain the aiming posture;

[0037] The attitude and orbit control system pushes the device body to move away from the meteorite ball to a set distance position along the flight direction, and keeps flying at the same distance and speed as the meteorite ball in real time;

[0038] Step 3: The pulse laser emits laser towards the meteorite ball according to the set rules;

[0039] The laser rangefinder, visible light camera and attitude and orbit control system cooperate to keep the device body and the meteorite ball at the same distance and angle;

[0040] The accelerometer monitors and records the acceleration data of the movement of the device body in real time.

[0041] As an improvement of the above method, the pulse laser emits laser to the meteorite ball according to a set rule, including:

[0042] In each experiment, the pulse laser fires multiple pulses of laser at the meteorite in the sequence of the set single pulse laser energy values ​​at the set repetition frequency, and each firing process lasts for the set duration;

[0043] Set the time interval and conduct multiple experiments.

[0044] As an improvement of the above method, the set single pulse laser energy value sequence is 0.5, 1, 1.5, 2, 2.5, 3, 3.5, 4, 4.5, 5 mJ;

[0045] The set repetition frequency is 10 Hz;

[0046] The setting duration is 30 minutes.

[0047] 9. The method for verifying orbital deflection caused by laser ablation of asteroids in orbit according to claim 6, characterized in that the method of starting the clamping and releasing spinning mechanism to make the meteorite ball spin at a set speed and then releasing it in situ comprises:

[0048] Step A1: clamp and release the spinning mechanism to clamp the meteorite ball and keep the meteorite ball still;

[0049] Step A2: Clamp and release the spinning mechanism to spin the meteorite ball to 20±10° / s, and keep rotating for the set holding time;

[0050] Step A3: Clamp and release the rotation mechanism to release the meteorite ball in situ, and maintain the output of stable rotational power to the meteorite ball during the release process.

[0051] As an improvement of the above method, the set distance is 3100±100 mm.

[0052] Compared with the prior art, the advantages of this application are:

[0053] 1. The experiment is conducted in orbit, and has a relatively consistent space environment and application scenario with the actual application of laser ablation-driven asteroid orbit deflection defense technology in the future in terms of vacuum degree, vacuum environment, environmental physical composition, gravity level, etc. The implementation of the experiment can more directly verify the feasibility of laser ablation-driven asteroid orbit deflection defense technology, and provide more accurate theoretical design technical parameters for laser ablation-driven asteroid orbit deflection defense technology, so that more accurate asteroid orbit deflection effects can be obtained in specific defense missions.

[0054] 2. The experiment uses a series of ingenious and simple system integration design schemes, such as small-volume, low-energy pulse lasers, simple laser rangefinders, tiny high-precision accelerometers, spherical hollow asteroid meteorites, conventional micro-nano satellite platforms, multi-pulse long-time laser ablation strategies, and convenient measurement methods of acceleration curve integral velocity increments. These enable the in-orbit technology verification experiment of laser ablation driven asteroid orbit deflection to be implemented at low cost, quickly and efficiently. All experimental components and measurement methods can be obtained through conventional technical channels, without the need for any difficult technical research, and have strong engineering implementation feasibility. BRIEF DESCRIPTION OF THE DRAWINGS

[0055] Figure 1 The figure shows the operation status of the device for verifying the orbit deflection caused by laser ablation of asteroids in orbit space;

[0056] Figure 2 The following is an actual image of the ground pulsed laser ablation experiment; where Fe is iron; graphite is asbestos;

[0057] Figure 3 Shown is a diagram of the configuration of a chondrite shell;

[0058] Figure 4 Shown is a stereoscopic view of the clamping, releasing and rotating mechanism;

[0059] Figure 5 Shown is a top view of the clamping release and rotating mechanism;

[0060] Figure 6 The figure shows the working process of the clamping, releasing and rotating mechanism;

[0061] Figure 7 The figure shows the first step of the in-orbit space verification of the method of laser ablation of asteroids to cause orbital deflection;

[0062] Figure 8 The figure shows the second step of the method for verifying the orbital deflection caused by laser ablation of asteroids in orbital space;

[0063] Fig. 9 The third step of the method for verifying the orbital deflection caused by laser ablation of asteroids in orbit is shown;

[0064] Fig.10 The figure shows the effect of the device for verifying the orbital deflection by laser ablation of asteroids in orbital space when it is not in operation;

[0065] Fig.11 The figure shows the internal structure of the device used for in-orbit space verification of laser ablation of asteroids to cause orbital deflection;

[0066] Fig.12 The figure shows a top view of the internal structure of the device used for in-orbit space verification of laser ablation of asteroids to cause orbital deflection;

[0067] Fig.13 The figure shows the design diagram of the micro-nano satellite platform structural subsystem;

[0068] Fig.14 Shown is the spatial layout diagram of the attitude and orbit control subsystem.

[0069] Figure ID:

[0070] 1. First micro stepper motor 2. Second micro stepper motor

[0071] 3. First clamping movable rod 4. Second clamping movable rod

[0072] 5. Rotate the gear 6. Press the first bearing against the rotating block

[0073] 7. Second bearing pressing block DETAILED DESCRIPTION

[0074] The technical solution of the present application is described in detail below with reference to the accompanying drawings.

[0075] like Figure 1 As shown, the present invention provides an apparatus and method for in-orbit space verification of laser ablation of asteroids to cause orbital deflection, the method has the characteristics of low implementation cost, consistency with actual defense mission scenarios and space environments, compact and light pulse laser, flexible and optional orbit entry schemes, and strong engineering adaptability.

[0076] The present invention provides a device and method for verifying in-orbit space the orbit deflection caused by laser ablation of an asteroid.

[0077] 1. Experimental principle design:

[0078] According to the impulse coupling coefficient C of the pulsed laser ablation driving target m formula (E is the laser energy, m is the target mass, ΔV is the target velocity increment), and the general function form of the impulse coupling coefficient and the single-pulse laser power density (I is the laser power density):

[0079]

[0080] The C is calculated using the ΔV of the meteorite produced by pulsed laser ablation measured on orbit. m , adjust the pulse laser single pulse energy output 10 groups of different energy values, and get 10 groups of different C m , I discrete data points, and finally fitted to obtain the key parameters A, B, I0, C max In order to amplify the ΔV value of the meteorite to a measurable level, E is the sum of the energy of multiple pulses of laser in a single experiment, that is, E = eTf (e is the energy of a single pulse laser, T is the time of a single experiment, and f is the laser frequency).

[0081] 2. Overall design of experimental methods:

[0082] The asteroid is simulated by a stony (S-type) asteroid meteorite. The experimental device consists of a micro-nano satellite platform and a meteorite ball and other experimental components. The meteorite ball is fixed on the outside of the satellite platform. The experimental device is launched into orbit (it can be launched alone; it can also be launched as a piggyback; it can also be launched to the space station with the main mission, and then the space station's robotic arm will transport the experimental device from the cabin to the outside of the cabin and release it in situ. The atmospheric resistance will make it away from the space station) and after it flies normally in orbit, its clamping and releasing spin mechanism will stabilize the spin of the meteorite ball and then release it in situ. The micro-nano satellite platform will be away from the meteorite ball at a certain distance, and a pulse laser will be fired at the meteorite ball to ablate the meteorite ball to drive it to produce a velocity increment, causing it to deviate from the orbit.

[0083] Since the single pulse laser effect is extremely weak (complex 10 -6 N·s-level tiny impulse measurement device, or increase the scale of the laser system to the kilojoule level and meter-level emission aperture to amplify the effect, making it difficult to conduct low-cost technical experiments on orbit). In order to amplify the ablation drive effect to a measurable state, according to the formula A sufficiently large ΔV requires a large C m E multiplied by the smaller m, so that C m The maximum value of 1.735×10 -5 N·s / J, assuming that the mass of the meteorite ball in this experiment is 0.1kg, and considering the resource limitation of micro-nano satellite, a single pulse maximum energy of 5mJ laser is selected. Then, the maximum instantaneous ΔV of the meteorite ball produced by the single pulse laser ablation is only 8.675×10 -4 mm / s, cannot be measured by simple means.

[0084] Experiments on the ground using nanosecond pulsed laser ablation have shown that the ablation process takes microseconds from the material surface to the generation of vaporization and plasma jet plumes until the reaction effect of the jet plumes disappears. Figure 2As shown in the figure, the pulse laser frequency used in the device of the present application is 10Hz, that is, a pulse laser is emitted once every 100ms, and the interval between two adjacent pulse laser emissions is 5 orders of magnitude higher than the entire process time of a single pulse laser ablation. Therefore, the continuous emission of pulse laser for ablation does not affect the ablation process of a single pulse laser, and is a linear superposition of the ablation of a single pulse laser. Therefore, the formula Can be equivalent to (e is the single pulse laser energy, T is the single experiment time, and f is the laser frequency). Therefore, this experiment is designed as follows: according to the single pulse energy change range of 0.5mJ to 5mJ, each time it increases by 0.5mJ, a total of 10 experiments are carried out, and the ablation experiment time for each energy value is 30 minutes (i.e. 18,000 pulses). Finally, it can be obtained that after the continuous action of the multi-pulse laser designed in this experiment, the final ΔV of the meteorite ball corresponding to the single pulse energy value of 0.5mJ to 5mJ in each single experiment is estimated to be 1.56 to 15.6mm / s, which is 4 to 5 orders of magnitude higher than that of a single pulse, so that it can be measured by simple and low-cost means: in a single experiment, the micro-nano satellite implements the same-distance and equal-speed tracking and aiming control strategy for the meteorite ball, and the meteorite ΔV can be obtained by integrating the full-course measurement curve using the accelerometer of the star itself.

[0085] The satellite platform needs to track the meteorite at the same speed and at the same distance as the meteorite ball. The relationship between the various design parameters of the pulse laser system is as follows: 2 λz (D is the laser emission aperture, d is the diameter of the target spot, a is the Gaussian beam with π / 4, M 2 The beam quality is 1.5, λ is the wavelength, 1064nm, z is the distance from the laser emission end to the target), and the single pulse laser power density is defined as (t is the single pulse laser pulse width, which is 5ns), we can get According to the method of obtaining ΔV in this experiment, if the satellite platform does not actively maintain a constant relative distance with the meteorite ball, the meteorite ball will gradually move away from the satellite platform under the action of continuous pulse lasers, that is, z in the formula changes continuously, causing d to change continuously, and I to change continuously, so C calculated by inverse calculation of ΔV is m It no longer corresponds to a unique I, so parameter fitting cannot be performed. Only by ensuring that z remains unchanged can the unique C corresponding to the continuous superposition of single-pulse lasers be obtained. mTherefore, it is necessary to design a low-cost satellite platform to track the meteorite ball at the same speed and at the same distance under limited resources. The specific method for realizing the equidistant and same-speed tracking of the satellite platform and the meteorite ball is as follows: when the distance value measured by the laser rangefinder reaches 3100±100mm (the selection of this distance value and error is a comprehensive consideration of the collision risk of the general satellite platform and the distance control accuracy), and is fed back to the satellite platform, the cold air nozzle at the front of the satellite platform stops working, and the solid cold air micro-propulsion module at the rear of the satellite platform is in a standby working state. The front nozzle and the rear propulsion module cooperate with each other in the feedback and comparison process of the distance value of the laser rangefinder, and the thrust centers of both pass through the center of mass of the satellite platform, and the thrust direction is parallel to the roll axis of the satellite platform, thereby realizing the real-time equidistant and same-speed tracking state of the satellite platform and the meteorite ball, so that the distance between the satellite platform and the meteorite ball is always kept within the range of 3100±100mm.

[0086] 3. Key parameters of experimental methods and component design:

[0087] 1) Laser key parameter design:

[0088] The key design parameters of pulsed lasers include: single pulse energy e, wavelength λ, pulse width t, repetition frequency f, beam quality M 2 , the diameter of the light spot at the emitting end D, and the distance z from the target.

[0089] In order to minimize the size, weight, and system complexity of the laser, combined with the experience of laser energy values ​​selected in ground laser ablation experiments, the maximum single-pulse laser energy is selected as 5mJ (if it is increased to 10mJ, a refrigerant module needs to be added, which will increase the weight of the laser); wavelength λ, pulse width t, repetition frequency f, beam quality M 2 The selection of the four values ​​is based on the conventional design experience of the laser, and is determined as follows: wavelength λ is 1064nm, pulse width t is 5ns, repetition frequency f is 10Hz, and beam quality M is 1000nm. 2 Determined as 1.5 according to the Gaussian beam; the target distance z is taken as 3100 mm according to the above results.

[0090] According to the relationship between the various design parameters of the pulse laser system, Dd = aM 2 λz (D is the laser emission aperture, d is the diameter of the target spot, a is the Gaussian beam with π / 4, M 2 The beam quality is 1.5, λ is the wavelength, 1064nm, z is the distance from the laser emission end to the target), and the single pulse laser power density is defined as (t is the single pulse laser pulse width, which is 5ns) For the formula Only I and D need to be determined. For the value of I, we can choose I0=1.489×10 9 W / cm 2 At this point, only D remains to be determined. Obviously, the determination of D depends entirely on the selection of the e value. If the e value is selected as the maximum single-pulse laser energy of 5mJ, the meaning of the D value is: the D value and other selected design parameters enable the pulse laser to obtain the optimal target power density I0 for ablation experiments at a fixed distance of 3100mm from the target when the energy value is 5mJ. The final measured C m corresponds to 5mJ, and at this time C m The value is the highest point of the impulse coupling coefficient-power density curve. Then, when the laser energy value takes a series of values ​​less than 5mJ, the corresponding power density values ​​must all be less than I0. Then the corresponding C m The values ​​must all be on the complete curve C m The left side of the maximum value will not get C m Any discrete point value to the right of the maximum value, that is, the slowly decreasing section of the final curve, will not be fitted. Based on the above reasons, in order to obtain experimental discrete point values ​​on both sides of the highest point of the curve, e cannot be taken as 5mJ when calculating D, but should be taken as the middle value of 2.5mJ. In this way, when the laser energy value is less than or greater than 2.5mJ, the discrete value of the impulse coupling coefficient corresponding to the power density less than or greater than the optimal power density I0 can be obtained through experiments, and finally the impulse coupling coefficient-power density curve that first rises nonlinearly and then slowly decreases can be completely fitted. Therefore, when e is taken as 2.5mJ, the calculated D is rounded to 20mm.

[0091] 2) Key parameter design of meteorite ball:

[0092] The meteorite ball is made of aluminum alloy. In order to maximize the velocity increment of the meteorite ball after being driven by laser ablation, and considering that the density of S-type asteroid meteorites is very close to that of aluminum alloy metal, it has reached 2700kg / m 3 Too small a mass will result in a too small volume, making it difficult to track and aim, so the mass of the meteorite ball was selected to be 0.1kg.

[0093] After the selection and spatial layout of the two groups of laser rangefinders in the overall design, the distance between the two laser emission ports of each group of laser rangefinders is 42mm. In order to enable the four ranging laser beams to irradiate the surface of the meteorite ball and realize real-time ranging, it is determined that the diameter of the meteorite ball must be ≥42mm. Combined with the mass calculation formula of the meteorite ball thin-walled shell design configuration (where R is the outer radius of the spherical shell and r is the inner radius of the spherical shell), as well as the feasible processing wall thickness, it was finally determined that the outer diameter of the meteorite spherical shell configuration was 80mm and the wall thickness was 2mm.

[0094] Considering the continuous pulse ablation method of this experiment, it is necessary to perform a composite calculation on whether the shell of the meteorite ball will be burned through after 180,000 pulses of laser ablation. The mass loss data of laser ablation of S-type asteroid meteorites was obtained based on the results of ground ablation experiments in relevant literature: the single pulse laser energy value used in the experiment was 5J, and the ablation mass produced by 300 times was 0.05g. When converted to the single pulse laser energy value of 5mJ in this experiment, it is equivalent to 300,000 times of 5mJ single pulse laser to produce an ablation mass of 0.05g. In this experiment, the energy of the 10 single pulse lasers changed from 0.5mJ to 5mJ, a total of 180,000 pulses, all converted to a unified 5mJ energy value, which is equivalent to a total of 99,000 5mJ single pulse lasers. Calculated in proportion to the ground test in the literature, this experiment will ablate a total of 0.0165g of mass.

[0095] If the meteorite ball is always in a three-axis stable static state, then the equivalent 99,000 5mJ single pulse lasers will ablate the same point, and the formula Dd = aM 2 λz is calculated to obtain a pulse laser to target spot diameter d of 0.29 mm. The spherical shell corresponding to this diameter spot is an approximate cylinder, and the mass of this cylinder is It is far less than 0.0165g, indicating that under this attitude strategy, the spherical shell of the meteorite will be burned through during the experiment, resulting in the overall failure of the experiment.

[0096] In order to prevent the same position of the meteorite ball from being ablated by the laser all the time, its release posture is designed to be a single-axis spin release posture, with its spin axis parallel to the pitch or yaw axis of the satellite platform. The calculation formula for the ablation mass of the meteorite ball in the spin state is established to obtain the appropriate spin angular velocity. The meteorite ball configuration modeling is as follows: Figure 3 shown.

[0097] In order to minimize the number of ablations of the same part of the meteorite ball, it should be rotated by an angle corresponding to the diameter of the laser spot during the 100ms interval between adjacent pulse lasers. At this time, the spin angular velocity is the optimal angular velocity. Figure 3 The outer radius of the meteorite ball is R and the inner radius is r.

[0098] Then the angle corresponding to a laser spot is Spend,

[0099] The corresponding spin angular velocity is Degrees / second.

[0100] Then the number of laser pulses emitted during the time it takes for the meteorite to rotate once is indivual.

[0101] According to the calculation basis obtained in the above calculation of ablation mass, the number of equivalent pulses corresponding to this experiment converted to 5mJ is 99,000, so the number of meteorite spheres required to emit 99,000 pulses of laser is: cycles, that is, the number of times the same part of the meteorite ball is repeatedly ablated is also Second-rate.

[0102] In order to prevent the spherical shell of the meteorite from being burned through, a redundant design is adopted, and the mass of the cylinder corresponding to each spot is twice the total mass of the ablation, and the equation is:

[0103]

[0104] When the diameter of the meteorite ball is 80mm, the solution of r is 39.786mm when substituted into the above ablation mass calculation equation. The initial design r according to the shell mass formula is 38mm, which is less than 39.786mm and meets the ablation mass requirements. The results obtained based on the ablation mass equation show that the required shell wall thickness is only 0.214mm under a 1x redundancy factor. Substituting the outer radius of the meteorite ball 40mm into the formula The optimal spin angular velocity was obtained to be 4.156° / s, which means that at this spin angular velocity, the corresponding shell wall thickness must be 0.214mm to avoid burning through. In other words, the number of pulsed lasers corresponding to one rotation of the meteorite ball is 866.2, and the equivalent number of repetitions of ablation of the same part is 111.7 times.

[0105] Considering that the actual design wall thickness of the spherical shell is 2mm, which is 9.34 times the required wall thickness of 0.214mm, the rotation speed of the meteorite ball can be increased, that is, the number of repeated ablations of the same part is increased, and the maximum rotation speed of the meteorite ball can be 9.34 times the optimal rotation speed of 4.156° / s, that is, 37.4° / s. At this time, according to the 1x redundancy coefficient, the 2mm thick spherical shell of the meteorite ball will be completely consumed by ablation. Therefore, in order to reduce the design complexity and speed accuracy index of the clamping release spinning mechanism, the design rotation speed of the meteorite ball is taken as Rounded to 20° / s, after comparison with the optimal and maximum speeds, the speed error can be taken as ±10° / s. This design value will minimize the device selection requirements for the clamping and releasing rotating mechanism, and achieve low-cost and high-reliability implementation of the experiment.

[0106] 3) Clamping, releasing and rotating mechanism design:

[0107] The clamping release and rotating mechanism comprises a first clamping movable rod 3, a second clamping movable rod 4, a first bearing pressing rotating block 6, a second bearing pressing rotating block 7, a first micro-stepping motor 1, a second micro-stepping motor 2 and a rotating gear 5. Figure 4 and Figure 5 shown.

[0108] The first clamping movable rod 3 and the second clamping movable rod 4 are both installed on the structural components of the satellite platform. They can slide and unfold along the two sides of the satellite platform under the action of the first micro-stepping motor 1 driving the rotating gear 5, thereby realizing the in-situ release of the meteorite ball.

[0109] The first bearing clamping rotating block 6 and the second bearing clamping rotating block 7 are respectively installed at the end parts outside the cabin of the first clamping movable rod 3 and the second clamping movable rod 4. They are provided with rotating bearings, which can flexibly rotate relative to the clamping movable rods. The material used is a non-metallic polytetrafluoroethylene material with a certain toughness, and a heat-insulating multi-layer material is wrapped on the outside of the rotating block to prevent its thermal expansion and contraction deformation. The first bearing clamping rotating block 6 installed on one side with the second micro-stepping motor 2 is fixedly connected to the output shaft of the second micro-stepping motor 2, and can be driven to rotate by the stepping motor; the second bearing clamping rotating block 7 installed on the other side rotates relative to the second clamping movable rod 4 through its own bearing, and rotates passively when the meteorite ball starts to spin.

[0110] The first micro-stepping motor 1 and the second micro-stepping motor 2 are small and mature 28HS28 model micro-stepping motors, with an angular accuracy of 0.09°, a speed of 0-600 / min, which can be set to the required speed in advance, and a torque of 0.08Nm, which is enough to drive the meteorite ball to spin. The mass is 90g, the fuselage height (excluding the output shaft length) is 28mm, and the power consumption is 3W. The second micro-stepping motor 2 is installed at the end of the first clamping active rod 3 outside the cabin to drive the meteorite ball to spin; the first micro-stepping motor 1 is installed at the tail ends of the two clamping active rods in the satellite platform cabin to drive the rotating gear 5 to drive the two clamping active rods to separate to both sides.

[0111] The rotating gear 5 is installed on the output shaft of the first micro-stepping motor 1 in the satellite platform cabin and is fixedly connected thereto. Its gear teeth match the rack gear teeth at the tail ends of the two clamping movable rod cabins. Driven by the first micro-stepping motor 1, the two clamping movable rods can be driven to separate to both sides.

[0112] The working process of the clamping and releasing rotating mechanism is as follows: Figure 6As shown, first, in the stage of clamping the meteorite ball, the first micro-stepping motor 1 is not started, so that the rotating gear 5 and the tail end racks of the first clamping movable rod 3 and the second clamping movable rod 4 are closely matched, ensuring that the first clamping movable rod 3 and the second clamping movable rod 4 firmly clamp the meteorite ball. Secondly, in the stage of spinning the meteorite ball, the second micro-stepping motor 2 is turned on and started, driving the bearing pressing the first bearing pressing rotating block 6 to rotate, thereby driving the meteorite ball to rotate by friction, and at the same time, the meteorite ball drives the second bearing pressing rotating block 7 on the other side to rotate relative to the clamping movable rod, and the symmetrical bearing pressing rotating blocks on both sides ensure the stability of the meteorite ball's spinning state. Finally, when the second micro-stepping motor 2 reaches the specified speed and drives the meteorite ball to rotate for a certain period of time, the first micro-stepping motor 1 drives the rotating gear 5 to drive the first clamping movable rod 3 and the second clamping movable rod 4 to separate to both sides to realize that the meteorite ball in the spinning state is released in situ, and at the same time, the second micro-stepping motor 2 still maintains the specified speed to ensure the spinning stability of the meteorite ball at the moment of release.

[0113] Example 1

[0114] 1. Experimental task process:

[0115] Step 1: (Take the release into orbit from the space station as an example) The space station's robotic arm transports the experimental device from the cabin to the outside, making its windward side parallel to the orbit direction and releasing it in situ. The experimental device is in a dormant state and decelerates to descend to orbit under the action of atmospheric resistance. When it is about 10 kilometers away from the space station, it is ready to start the experiment. Figure 7 As shown:

[0116] Step 2: After the experimental device reaches the designated position, it is turned on. First, the overall attitude is adjusted through the attitude and orbit control system so that the meteorite ball is facing the wind and the overall roll axis of the device is parallel to the flight direction; then, the clamping and releasing spinning mechanism makes the meteorite spin to 20±10° / s, and then releases it in situ; at the same time, four sets of laser rangefinders and visible light cameras are turned on, and the attitude and orbit control system is used to aim the center of the meteorite ball in real time and maintain the aiming attitude. The visible light camera is used to observe the four laser rangefinder light spots on the meteorite ball in real time to determine that the meteorite ball has not been lost and the satellite platform is always in an active tracking state; finally, the solid cold gas micro-propulsion module is turned on, and the cold gas nozzle at the front of the device works. The laser rangefinder measures data to make the satellite platform slowly move away from the meteorite ball along the flight direction to a distance of 3100±100mm from it, and the attitude and orbit control system is used to keep flying at the same distance and speed as the meteorite ball in real time. Figure 8 shown.

[0117] Step 3: Turn on the pulse laser and fire 10 pulse lasers at the meteorite in the order of single pulse laser energy values ​​of 0.5, 1, 1.5, 2, 2.5, 3, 3.5, 4, 4.5, and 5 mJ. Each firing process lasts 30 minutes. The pulse laser frequency is designed to be 10 Hz, that is, each firing process fires a total of 18,000 pulse lasers at the meteorite ball. In 10 experiments, a total of 180,000 pulse lasers are fired. A certain time interval is used for data differentiation and status adjustment of the satellite platform and pulse laser. Fig. 9 shown.

[0118] Step 4: During each experiment, the satellite platform always maintains a distance of 3100±100mm from the meteorite ball through the tracking and aiming system, and the laser is aimed at the center of the meteorite ball. The visible light camera observes the four laser rangefinder spots on the meteorite ball throughout the process to determine that the meteorite ball has not been lost and the satellite platform is always in an active tracking and aiming state. In addition, the camera takes pictures or records videos of the state of the laser ablation meteorite ball to obtain image data of the laser ablation driving effect in the orbital environment. The satellite platform records the equidistant and uniform tracking acceleration curve of the satellite platform during the experiment through an accelerometer, and integrates the acceleration process curve at the end of each experiment to obtain the final velocity increment ΔV of the meteorite ball after each experiment. A square windward plate structure is designed at the rear of the satellite platform to keep the satellite platform and the meteorite ball have the same surface quality, so that the atmospheric resistance factor is completely offset.

[0119] Step 5: After all 10 laser ablation meteorite ball experiments are completed, the satellite platform will transmit 10 sets of ΔV as well as the original data of the accelerometer and laser rangefinder, and the image and video data of the visible light camera to the ground. The ground data processing system will perform parameter fitting and related data processing based on the experimental principles.

[0120] 2. Design of key components of the experimental device:

[0121] 1) Overall design:

[0122] The complete design of the experimental device is as follows Fig.10 , Fig.11 and Fig.12 As shown in the figure, the device consists of two parts: experimental components and micro-nano satellite platform. The experimental components include: pulse laser, meteorite ball, clamping and releasing spinning mechanism, laser rangefinder, accelerometer, visible light camera; the micro-nano satellite platform includes conventional structure, attitude and orbit control, thermal control, energy and other subsystems.

[0123] The total weight of the experimental device is close to 10 kg, and the outer envelope volume is (excluding the rear windward square plate): 488 mm long, 148 mm wide, and 190 mm high. The side length of the windward square plate is 400 mm.

[0124] 2) Laser rangefinder design:

[0125] The laser rangefinder uses the LDM40 OEM 40-meter high-precision laser ranging module. The laser rangefinder has a repeatability of ±0.3mm, stable ranging performance, high reliability, small size, and weighs only 42g. Its resolution can reach 0.1mm. The spot diameter is 6mm at 10 meters. According to the design of this experiment, the distance between the satellite platform and the meteorite ball target is 3100mm. According to the linear change law of the spot, the diameter of the laser spot irradiated to the surface of the meteorite ball is 2mm.

[0126] 3) Accelerometer design:

[0127] The accelerometer is used to measure and monitor the acceleration state of the device. The accelerometer uses the FH2000G model quartz flexible military accelerometer. The measurement resolution of the accelerometer is 5ug and the range is ±35g. The rated thrust of the orbital control solid cold gas micro-propulsion module used in this experiment is 50mN. Considering that the total design mass of the satellite platform is 10kg, the thrust acceleration generated is 5mg, which can be accurately measured by this accelerometer; the measurement cutoff frequency of the accelerometer is ≤2.5kHz. The minimum pulse impulse of the orbital control solid cold gas micro-propulsion module used in this experiment is 5×10 -4 N·s, according to the thrust value of 50mN, it can be concluded that the minimum pulse time length is 0.01s, that is, the maximum frequency of acceleration generated by thrust is 100Hz, which is less than the cutoff frequency of 2.5kHz of the accelerometer and can be fully measured by the accelerometer. The accelerometer is small in size, weighs only 80g, and consumes only 0.36W of power. It is installed on the structural components in the satellite platform cabin and is fixed to it, and the measuring axis is parallel to the roll axis of the satellite platform.

[0128] 4) Visible light camera design:

[0129] The visible light camera uses a simplified version of a multifunctional status monitor in a near-space exploration project, retaining only the module with visible light camera function, which is installed on the front structural component of the satellite platform. The visible light camera will collect images and videos of the experimental phenomenon of pulsed laser on-orbit ablation of asteroid meteorites for subsequent related research; and the visible light camera will further judge the tracking effect of the satellite platform on the meteorite ball and whether it has lost the tracking by imaging the four light spots irradiated on the meteorite ball by the laser rangefinder, providing the most direct evidence for whether the experimental data is true and reliable.

[0130] The field of view of the visible light camera is designed to be 20°. After calculation, the market range at the distance where the meteorite ball is located is a circular area with a diameter of 1090mm, which can fully ensure that the meteorite ball appears in the field of view of the visible light camera; the resolution of the visible light camera is 1920×1080 pixels, so the area of ​​one pixel in the area where the meteorite ball is located is 0.45mm2. According to the aforementioned laser rangefinder irradiating a circular light spot with a diameter of 3mm on the surface of the meteorite ball, it can be known that the number of pixels occupied by a light spot is 15 pixels, which can be clearly identified by the visible light camera.

[0131] 5) Micro-nano satellite platform structure subsystem design:

[0132] The micro-nano satellite platform structure is designed around the main functions of this experiment and the spatial layout of the main experimental components. It adopts a rectangular structure envelope configuration similar to that of a cubic satellite, plus a square windward plate structure located at the rear of the platform. The main function of the square windward plate structure is to keep the satellite platform and the meteorite ball with the same surface-to-mass ratio parameters, so that the atmospheric drag factors of the two during their in-orbit flight can be deducted during the experiment, reducing the error effect of the space atmospheric environment on the laser ablation drive experimental measurement. According to the surface-to-mass ratio calculation formula, the surface-to-mass ratio of the meteorite ball can be obtained as follows: According to the overall design plan of the satellite platform, the total mass of the satellite platform is 10kg, so the required side length of the square windward surface is 400mm.

[0133] The structural load-bearing design adopts a box-plate configuration, with metal structural parts embedded in the internal equipment layout for mechanical installation of the equipment, and the rest are thin non-metallic honeycomb panel structures or carbon fiber hollow panel structures to minimize the overall structural mass. Fig.13 shown.

[0134] 6) Design of attitude and orbit control subsystem of micro-nano satellite platform:

[0135] The attitude and orbit control subsystem is responsible for adjusting and determining the overall attitude of the experimental device after it is released by the space station's robotic arm and reaches a specified distance away from the space station, as well as maintaining and determining the overall attitude of the satellite platform throughout the experiment, and implementing the tracking and aiming drive between the satellite platform and the meteorite ball during the laser ablation drive experiment.

[0136] The attitude and orbit control subsystem consists of a solid cold gas micro-propulsion module (including two cold gas nozzles installed at the front of the satellite platform, and the cold gas pipeline is deployed inside the platform and extends to the front of the platform and connected to the cold gas nozzles), an offset momentum wheel, two star sensors, two sun sensors, and three magnetic torquers.

[0137] The spatial layout of the attitude and orbit control subsystem is as follows: Fig.14The main part of the solid cold gas micro-propulsion module is installed at the rear of the satellite platform, with its thrust nozzle facing the opposite direction of the platform's flight direction, and the thrust direction passes through the platform's center of mass and is parallel to the platform's roll axis. Two cold gas nozzles are installed outside the front cabin of the satellite platform, symmetrically installed, with the nozzles facing the same direction as the platform's flight direction, and the thrust resultant direction passes through the platform's center of mass and is parallel to the platform's roll axis; an offset momentum wheel is installed in the platform cabin, and its rotation axis is perpendicular to the thrust resultant direction of the two cold gas nozzles at the front of the platform. The offset momentum wheel and the two cold gas nozzles (thrust The satellite platform can adjust its pitch / yaw axis attitude when the forces are not equal. Two star sensors are installed in the front area of ​​the platform cabin, and their probes are extended out of the structural cabin and exposed to the outside of the platform. The two star sensors are installed vertically at 90 degrees, and the plane they form is perpendicular to the platform roll axis. Two sun sensors are installed on the outside of the front and rear cabins of the satellite platform, pointing to the forward and backward directions of flight respectively. Three magnetic torquers are installed inside the satellite platform, and the central axes of the three magnetic torquers are parallel to the pitch / yaw / roll axes of the satellite platform.

[0138] The solid cold air micro-propulsion module adopts mature shelf products, and its performance parameters are shown in the following table:

[0139] Table 1 Performance parameters of solid cold gas micro propulsion module

[0140]

[0141] The device and method for verifying orbital deflection caused by laser ablation of asteroids in orbital space proposed in this application have the following advantages:

[0142] (1) A laser ablation-driven asteroid orbit deflection technology verification test method is proposed, which uses an asteroid meteorite as the laser ablation object and a small-caliber and ultra-low single-pulse energy value pulse laser as the ablation laser. It has low cost, strong engineering feasibility, and can be quickly implemented on-orbit.

[0143] (2) A multi-pulse, long-duration laser ablation driving experimental strategy was adopted, which effectively amplified the extremely weak ablation driving effect of a low-energy single-pulse laser, greatly increasing the magnitude of the velocity increment corresponding to the impulse generated, and thus enabling it to be measured in a simple manner.

[0144] (3) A simple combined equidistant and same-speed tracking design using a laser rangefinder and a solid-cooled gas micro-propulsion module was proposed. In the experimental process, the equidistant and same-speed tracking acceleration curve was integrated using a high-precision accelerometer to equivalently obtain the velocity increment generated by the laser ablation driving impulse acting on the target.

[0145] Finally, it should be noted that the above embodiments are only used to illustrate the technical solution of the present application and are not intended to limit it. Although the present application is described in detail with reference to the embodiments, a person skilled in the art should understand that any modification or equivalent replacement of the technical solution of the present application does not depart from the spirit and scope of the technical solution of the present application and should be included in the scope of the claims of the present application.

Claims

1. An on-orbit space verification device for laser ablation of asteroids to cause orbital deflection, characterized in that: The device comprises: a meteorite ball and a device body; wherein, The meteorite ball is a hollow sphere used to simulate an asteroid; The device body comprises: A pulse laser, used for emitting laser light toward the meteorite ball according to a set rule; A clamping and releasing spinning mechanism is used to rotate the meteorite ball at a set speed; A laser rangefinder, used to measure the distance and angle between the device body and the meteorite ball; An attitude and orbit control system, used for adjusting the distance and angle between the device body and the meteorite ball according to the measurement results of the laser rangefinder; A visible light camera, used to take photos or videos of the meteorite ball, monitor the ablation of the meteorite ball and the relative angle between the device body and the meteorite ball; and An accelerometer, used to measure and record acceleration data of the movement of the device body; The clamping, releasing and rotating mechanism comprises a clamping, releasing system and a rotating system; The clamping and releasing system is used to clamp the meteorite ball and release the meteorite ball when the meteorite ball reaches a set rotation speed; The spinning system is used to drive the meteorite ball to rotate at a set speed.

2. The device for verifying the orbital deflection caused by laser ablation of asteroids in orbit space according to claim 1, characterized in that: The clamping and releasing system comprises: a first micro-stepping motor, a first clamping movable rod, a second clamping movable rod, a rotating gear, a first bearing clamping rotating block and a second bearing clamping rotating block; wherein, The first micro-stepping motor is arranged inside the device body to drive the rotating gear to rotate; One end of the first clamping movable rod and the second clamping movable rod are both located inside the device body and clamp the rotating gear, and the rotating gear drives the two clamping movable rods to move closer to or away from each other; the other ends of the first clamping movable rod and the second clamping movable rod are both extended out of the device body; The first bearing clamping rotating block is arranged at one end of the first clamping movable rod outside the device body; the second bearing clamping rotating block is arranged at one end of the second clamping movable rod outside the device body; the first bearing clamping rotating block and the second bearing clamping rotating block clamp the meteorite ball.

3. The device for verifying the orbital deflection caused by laser ablation of asteroids in orbit space according to claim 2, characterized in that: The spinning system includes a second micro-stepping motor; The second micro-stepping motor is fixed to one end of the first clamping movable rod outside the device body, and the rotating shaft thereof passes through the first clamping movable rod and is fixedly connected to the first bearing pressing rotating block, thereby driving the first bearing pressing rotating block to rotate; The axis of the rotating shaft of the second micro-stepping motor is perpendicular to the movement direction of the device.

4. The device for verifying the orbital deflection caused by laser ablation of asteroids in orbit space according to claim 1, characterized in that: The outer diameter of the meteorite ball is ≥42mm and the wall thickness is ≥0.214mm.

5. A method for verifying orbital deflection caused by laser ablation of an asteroid in orbit, implemented based on the device of any one of claims 1 to 4, comprising: Step 1: placing the device in a geocentric orbit in space; Step 2: The attitude and orbit control system adjusts the attitude of the device so that the meteorite ball faces the wind and the overall roll axis of the device is parallel to the flight direction; Start the clamping and releasing spinning mechanism to make the meteorite ball spin at the set speed, and then release it in place; Turn on the laser rangefinder and visible light camera, use the attitude and orbit control system to aim at the center of the meteorite ball in real time, and maintain the aiming posture; The attitude and orbit control system pushes the device body to move away from the meteorite ball to a set distance position along the flight direction, and keeps flying at the same distance and speed as the meteorite ball in real time; Step 3: The pulse laser emits laser towards the meteorite ball according to the set rules; The laser rangefinder, visible light camera and attitude and orbit control system cooperate to keep the device body and the meteorite ball at the same distance and angle; The accelerometer monitors and records the acceleration data of the movement of the device body in real time.

6. The method for verifying orbital deflection caused by laser ablation of an asteroid in orbit in space according to claim 5, characterized in that: The pulse laser emits laser to the meteorite ball according to the set rules, including: In each experiment, the pulse laser fires multiple pulses of laser at the meteorite in the sequence of the set single pulse laser energy values ​​at the set repetition frequency, and each firing process lasts for the set duration; Set the time interval and conduct multiple experiments.

7. The method for verifying orbital deflection by laser ablation of an asteroid in orbit space according to claim 6, characterized in that: The set single pulse laser energy value sequence is 0.5, 1, 1.5, 2, 2.5, 3, 3.5, 4, 4.5, 5 mJ; The set repetition frequency is 10 Hz; The setting duration is 30 minutes.

8. The method for verifying orbital deflection caused by laser ablation of an asteroid in orbit space according to claim 5, characterized in that: The method of starting the clamping and releasing spinning mechanism to make the meteorite ball spin at a set speed and then releasing it in situ includes: Step A1: clamp and release the spinning mechanism to clamp the meteorite ball and keep the meteorite ball still; Step A2: Clamp and release the spinning mechanism to spin the meteorite ball to , and keep rotating to set the holding time; Step A3: Clamp and release the rotation mechanism to release the meteorite ball in situ, and maintain the output of stable rotational power to the meteorite ball during the release process.

9. The method for verifying orbital deflection caused by laser ablation of an asteroid in orbit in space according to claim 5, characterized in that: The set distance is 3100±100 mm.

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