246 results about "Planet Mars" patented technology

The invention provides a phased multi-rate self-adaptive measurement and control system of a mars exploration deep space spacecraft, which is characterized by comprising an X-frequency range deep space responder, a high stable frequency source, an X-frequency range solid state power amplifier, an X-frequency range high-power traveling wave tube, an output multiplexer, an input low noise amplification receiving unit, a wide-beam low-gain receiving antenna, a wide-beam low-gain transmitting antenna, an intermediate-gain transmitting antenna, a two-dimensional driven high-gain measurement and control antenna and a waveguide switch. According to the invention, the problems of communication with the ground and measurement and control of the deep space explosion spacecraft are solved in the whole process that the deep space explosion spacecraft is launched from the earth, flies from the earth to the mars cruising segment and flies around the mars. The phased multi-rate self-adaptive measurement and control system has an excellent effect for the mars exploration deep space spacecraft to complete the measurement communication control from the earth orbit flight to the mars orbit flight. The phased multi-rate self-adaptive measurement and control system can solve the key difficult points of measurement and control of the mars deep space spacecraft and has a certain effect on the aspect of improving the reliability of the deep space spacecraft.

The invention aims at providing a Mars wind tunnel with sand and dust simulation. The wind tunnel is horizontally arranged in a vacuum container and comprises a collector, a stable section, a contraction section, a test section, an ejector section, an outlet diffusion section, an ejector, and a first diffusion section. The test section is connected with the first diffusion section; the first diffusion section is connected with the ejector section; the ejector section is connected with the diffusion section, a transitional equal straight section is arranged between the ejector section and the diffusion section, and the ejector is arranged at the inlet of the ejector section. The diffusion angle of the first diffusion section is not more than 5 degrees, and the cross section area of the inlet of the ejector section is larger than that of the outlet of the test section. Compared with the prior art, simulation of the wind speed of 5-100 m / s in a large-span range can be achieved under the air pressure of 100 Pa only by consuming a small amount of air, the flow field of the test section is uniform and stable, and the sand and dust environment of the Mars can be simulated under the low air pressure.

The invention discloses a parsing homotopy method for generating a Martian atmosphere entry phase longitudinal accessible region, and belongs to the field of deep space exploration. The parsing homotopy method comprises the following steps that a Martian atmosphere entry phase longitudinal dynamic model is built; based on the parsing homotopy method, known auxiliary problems are parsed from an optimum solution, and by constructing homotopy parameters, a maximum flying range is subjected to continuation; based on the homotopy method, the known auxiliary problems are parsed with the optimum solution, and through constructing the homotopy parameters, a minimum flying range is subjected to continuation and obtained; based on the parsing homotopy method, the known auxiliary problems are parsedwith the optimum solution, and by constructing the homotopy parameters, a maximum umbrella opening height is subjected to continuation and obtained; based on the parsing homotopy method, according tothe maximum flying range, the minimum flying range and the maximum umbrella opening height which are obtained through continuation, and by constructing the homotopy parameters, the longitudinal accessible region is obtained through continuation. The parsing homotopy method for generating the Martian atmosphere entry phase longitudinal accessible region can avoid co-state initial value conjecture and internal point accessibility verification, and thus improves the generating efficiency of the accessible region.

The invention discloses a combined laser induced spectrum system for detecting Mars substance components. The system comprises a cassette telescope, an ICCD spectrometer, a laser range finder, a main controller and data analysis system, a multichannel digital delay pulser, a 1064nm beam expander, a 1064nm high-energy pulse laser, a frequency-multiplication beam expanding assembly and an auxiliary light machine. The Mars substance components can be detected based on the system combined with a laser induced plasma spectroscopy technology and a laser induce time domain resolution fluorescence spectroscopy technology. The system has the beneficial effects that a combined laser spectroscopy method is provided and is combined with an LIBS (laser induced breakdown spectroscopy) detection and LIF (laser induced fluorescence) time domain resolution spectroscopy technology, and respective advantages are played, so that the detection range of substance elements on the surface of Mars can be effectively expanded. A same laser source and a same detector are adopted in the two laser spectroscopy technologies under the condition that the detection range of the elements can be effectively expanded, so that the compact and convenient performances of the whole system are kept.

The invention relates to a simulation method of a Mars object for deep space exploration, which belongs to the field of deep space exploration application. In order to solve the problems in the prior art of lacking of a method for simulating the location, size, contour, imaging orientation and imaging gray level of Mars and existing technical gaps in engineering practices, the simulation method of the Mars object for deep space exploration is proposed. The simulation method includes transforming a center coordinate of Mars from a heliocentric ecliptic coordinate system to a display plane coordinate system and a projector coordinate system, calculating an imaging size of Mars according to an imaging field of view of a Mars explorer and the relative distance between Mars and the Mars explorer, simulating illuminated areas of Mars which are illuminated by the sun and unilluminated areas which are not illuminated by the sun, calculating an imaging orientation of the contour of Mars according to the relation of relative positions among the sun, Mars and the Mars explorer, and mapping a magnitude of Mars to the gray level of a computer to display on an interface. The simulation method of the Mars object for deep space exploration is applicable to the simulation software of deep space exploration.

The invention provides a combined autonomous navigation method of a mars exploration approach section. The combined autonomous navigation method comprises the following steps: the step 1 of using a spacecraft-mounted sun sensor to measure the viewing direction of the sun and using a spacecraft-mounted solar-spectrum velocity-measurement navigation sensor to measure the relative movement velocity of a mars explorer to the sun; the step 2 of using a wireless responder to obtain distance and relative velocity information with a mars surrounding device; the step 3 of establishing an orbital dynamics equation and establishing a navigation system model under a J2000.0 heliocentric ecliptic inertial coordinate system and using a corresponding navigation filtering algorithm to obtain autonomous navigation position and velocity information. The combined autonomous navigation method disclosed by the invention avoids depending on ground wireless information, can achieve autonomous navigation of a spacecraft in a flying process by only needing optical information from the sun and the wireless information from the mars surrounding device, can establish a system equation of an autonomous navigation system of the explorer, utilizes different navigation filter algorithms to achieve real-time estimation of navigation information of the explorer and is simple, easy, practical and free of time delay.

The embodiment of the invention provides a Mars exploration method and a Mars probe. The Mars exploration method comprises the steps that the Mars probe is launched from the earth, lifts off and enters an earth Mars transfer orbit, orbit trimming is conducted, and the Mars probe enters a Mars atmospheric layer; pneumatic braking is conducted through the Mars atmospheric layer, and the Mars probe enters and parks on a large elliptical orbit around the Mars; when the Mars probe parks on the large elliptical orbit around the Mars, dust storm on the Mars is monitored; and when the situation that the dust storm on the Mars is formed is determined, the Mars probe enters the Mars atmospheric layer to sample the dust storm on the Mars. According to the scheme, the cost can be reduced during Mars exploration.

The invention discloses a method for predicting the surface flow transition position of a Mars entering device, and the method comprises the steps: carrying out computational fluid mechanics simulation based on a Mars atmosphere model by employing a thermochemical non-equilibrium model, and obtaining the streaming field data of the Mars entering device; generating a transition prediction grid by taking the molded line of the Mars entering device as a reference, and interpolating the streaming field data into the transition prediction grid to obtain transition prediction flow field data; calculating the total enthalpy of the flow field based on the transition prediction flow field data, obtaining the outer edge position of the boundary layer, carrying out the integral calculation on a first Reynolds number based on the momentum loss thickness in the boundary layer in the normal direction of the wall surface, and calculating a second Reynolds number based on the roughness height; and taking the first Reynolds number and the second Reynolds number as transition criteria, respectively calculating a first transition criterion and a second transition criterion, and obtaining the transition position of the surface flow of the Mars entering device based on the first transition criterion and the second transition criterion. Thus, the surface flow transition position of the Mars entering device can be predicted via the method.

The invention provides solar flare TDOA (Time Difference Of Arrival) measurement, anintegrated navigation method and an integrated navigation system, which belong to the field of spacecraft autonomous navigation. The solar flare TDOA measurement comprises the following steps of acquiring flare arrival time directly from the sun and flare arrival time reflected by the Mars; calculating a difference value between the two flare arrival times; building a solar flare TDOA measurement model. The integrated navigation method comprises the following steps of building a direction-findingnavigation model, building the solar flare TDOA measurement model, utilizing a filter to filter, and adopting the direction-finding model when the solar flare TDOA is not acquired. The solar flare TDOA measurement, and theintegrated navigation method and theintegrated navigationsystem provided by the invention solve the problem of large radial deviation of direction- finding navigation, are high in positioning precision, and have lower requirement on an instrument and ephemeris. Therefore, the solar flare TDOA measurement, and the integrated navigation method and the integrated navigation system provided by the invention has the important practical significance on spacecraftautonomous navigation.

The invention provides a Mars probe secondary brake capture control method based on Newton iteration. During capturing and braking, after a Mars detector breaks down, an on-board computer is reset, then, the Mars probe is subjected to secondary braking by adopting a Newton iteration-based capturing and braking optimization method, so that the detector can smoothly enter a target orbit in the Marscapturing and braking process, and a secondary capturing and braking strategy can be autonomously designed for faults in the capturing and braking period. Besides, a safety zone for capturing and braking is also arranged and comprises the shortest ignition duration and the longest ignition duration of braking and capturing; when the on-board time is within the interval of the shortest ignition duration and the longest ignition duration, the speed increment reaches the nominal speed increment or the on-board time reaches the maximum ignition duration, the engine shutdown operation is executed,so that the Mars detector can form an annular fire track during braking and capturing and cannot impact Mars due to overlong ignition time, and the safety of Mars braking and capturing is guaranteed.

The invention discloses a comprehensive environment heat balance test system for a Mars detector single-machine product. The system comprises a Mars environment simulation chamber, a solar irradiationsimulation system, a Mars atmosphere simulation system, a Mars surface wind simulation system, a temperature control system and a test piece support system, wherein the Mars environment simulation chamber is a horizontal and capsule-shaped cabin body, container gates which can be opened are arranged at the two ends of the cabin body, and the Mars environment simulation chamber is mainly used forforming a closed space without obvious leakage and maintaining a special atmospheric environment in the cabin. The system can provide comprehensive stress environment tests of five parameters including gas atmosphere, wind speed, pressure, solar irradiation and environment temperature on the surface of Mars, and is easy to operate and good in generalizability.

The invention discloses a star sensor and gyro combined calibration method suitable for Mars detection, and a Mars detector is allowed to complete the combined calibration of the star sensor and the gyro under any attitude reference in a full life cycle. In the steady-state flight process of the Mars detector, a calibration algorithm is constructed under the inertial reference according to measurement models of the star sensor and the gyroscope. Compared with the prior art, the star sensor and gyroscope combined calibration method suitable for the Mars detection has the advantages that the Mars detector can perform combined calibration of the star sensor and the gyroscope in any attitude in an inertial space, the calibration arc section is improved, and the reliability of the combined calibration of the star sensor and the gyroscope is improved.

The invention relates to a star sensor suitable for Mars detection and a gyro on-orbit combined use method, and relates to the field of Mars detector star sensor and gyro use. The method comprises: performing fault diagnosis on the four star sensors with different priorities in sequence, and judging whether a gyroscope needs to be used for outputting the attitude of the Mars detector or not; whenthe gyroscope is used for outputting the posture of the Mars detector, restarting the four star sensors in sequence according to the priority in real time, and achieving troubleshooting; when the starsensor with faults removed exists, switching the attitude of the Mars detector output by the star sensor in real time; keeping the attitude of the Mars detector output by the gyroscope after the starsensor with no troubleshooting is determined. According to the method, in the whole flight process of the Mars probe, the star sensor can be automatically diagnosed and reconstructed and has the automatic restart capacity after all faults occur, and gyro integration is automatically switched to provide high-precision attitude information under the condition that no star sensor is available.

The invention relates to a hybrid laser radar for Mars atmosphere detection. The hybrid laser radar comprises a single set of optical transceiver module for obtaining back scattering signals from Mars atmosphere molecules and dust; a coherent balance detection module which mixes the local oscillation light and the atmospheric echo signal into an intermediate frequency electric signal and converts the intermediate frequency electric signal into a high-speed analog electric signal, so a Mars atmospheric wind field is obtained through data processing; a single-photon direct detection module which is used for measuring photon electric signals of each channel to obtain photon count of each channel; and a data processing module which obtains the concentration of polarized light and Mars aerosol particles. The laser power and the echo signal stability are calibrated by using the signal-to-noise ratio of coherent detection and the photon number of direct detection, and the method is used for detecting the content of carbon dioxide gas. The hybrid laser radar has the advantages of large dynamic range and small data processing calculation amount by using single photon direct detection, and also has the advantages of high coherent detection sensitivity and good weak signal amplification performance, and the problem of multi-atmosphere parameter detection of the Mars atmosphere detection laser radar is solved under the condition that the load weight and complexity are not increased.

The invention relates to the technical field of deep space exploration, and discloses a deep space communication system suitable for a phobos detection task. On one hand, a Mars polar satellite running on a special polar orbit is arranged on the Mars side, obstacles between the detector and the Mars polar satellite are always prevented from appearing to block decimetric wave communication. The purpose of uninterrupted full period of the decimetric wave communication link from the phobos to the Mars is achieved; on the other hand, at least three revolution satellites around the sun are used asinter-planet relay satellites, the purpose that the laser wave communication link between the planets is not interrupted in the whole period can be achieved; and at least three geosynchronous satellites are used as satellite-ground relay satellites, so that the purpose of uninterrupted full period of the satellite-ground Ku wave communication link can be achieved, the purpose of uninterrupted fullperiod of the complete communication link from the phobos to the ground is finally achieved, and the phobos detection task is ensured not to be affected by communication interruption.

The invention discloses a multi-directional driving task autonomous management control method suitable for Mars detection. The method comprises the following steps: S1, according to driving constraints of a solar wing mounting and driving shaft of a Mars detector, a driving shaft of a directional antenna and a driving shaft of an X relay antenna, establishing a small-amplitude dynamic attitude reference capable of realizing multi-directional coverage on a nominal attitude reference in combination with an on-orbit sun vector, an earth vector and a lander vector, and providing a feasible attitude for autonomous management of multiple rotating parts; and S2, autonomously and dynamically tracking and pointing to respective targets according to a dynamic attitude reference established in the S1, a solar wing rotation range constraint, a directional antenna rotation range constraint and an X relay rotation range constraint, and autonomously feeding back the dynamic adjustment of the attitudereference after the rotation constraint range is exceeded so as to realize simultaneous pointing of multiple rotations. The method is simple and feasible, and the pressure of on-orbit time-sharing work of different rotating parts is reduced.

The invention discloses an autonomous dormancy and awakening control method for a Mars rover, provides a basis for selecting an autonomous dormancy strategy under a relatively sufficient energy working condition or an autonomous dormancy strategy under an energy shortage working condition on a device, and improves the use flexibility of the autonomous dormancy strategy. According to the peak electric quantity of the Mars rover in each Mars day, whether the Mars rover needs to enter a dormant state or carry out a minimum working mode is judged, and the Mars rover can autonomously recover to a normal working mode from the minimum working mode after the energy condition is improved; the method is suitable for the Mars rover to sleep as soon as possible after sand and dust are detected on thefire surface, sufficient electric quantity can be kept after the Mars rover sleeps, and a basis is provided for charging in the minimum working mode after subsequent awakening to restore energy balance. In combination with the peak electric quantity of the Mars rover in each Mars day and the current real-time electric quantity of the Mars rover, whether the Mars rover needs to enter dormancy is judged together; the strategy is suitable for being used after awakening, and aims to use the storage battery as much as possible to maintain power supply without dormancy after awakening, and protect the storage battery from being over-discharged at the same time.

The invention provides a Mars rover moving system traction performance test device, andbelongs to the technical field of Mars detection test equipment. The test device comprises a support base for supporting a tested Mars rover; aresistance moment loading mechanism which is used for being connected with wheels of the tested Mars rover and loading resistance moment to the wheels of the tested Marsrover; a traction loading mechanism which is used for being connected with a body of a tested Mars rover and applying traction force to thetested Mars rover in the walking process of the Mars rover; and a vacuum tank and temperature control mechanism which is used for containing the tested Mars rover and controlling the control vacuum degree and temperature of the test environment. The resistancemomentloading mechanism and the traction force loading mechanism are both arranged in the vacuum tank and temperature control mechanism. Compared with the prior art, the performance test of the Mars rover moving system on the ground can be realized, the wheel load and the wheel resistance moment required by different geomorphic characteristics can be simulated, and the requirements of 0-3000N traction force loading technical indexes and loading precision can be met.

The invention provides a backflow-type wind tunnel device for simulating a Mars low-pressure low-density dust storm environment. The device comprises a backflow-type wind tunnel, an ejector, and an jetting device; the backflow-type wind tunnel comprises a stable segment, a contraction segment, a test segment, a first diffusion segment, a second diffusion segment, a mixing segment, a third diffusion segment, and a large-angle diffusion segment; a first corner is arranged between the first diffusion section and the second diffusion section in a communicated manner; a second corner is arranged between the second diffusion section and the mixing section in a communicated manner;a third corner and a fourth corner are sequentially arranged between the third diffusion section and the large-angle diffusion section in a communicated manner;a corner flow deflector is arranged in each corner; the ejector is arranged in the mixing section; and the jetting device is arranged in the contraction section. The ejector is combined with the backflow-type wind tunnel, the wind speed required by the test section can be achieved through the small inlet pressure of the ejector; the mode of combining the large contraction ratio and the large-angle diffusion section is adopted, so that the airflow quality is improved, losses are reduced, the overall size of the wind tunnel is reduced, and the vacuumizing cost is reduced.

The invention relates to a midway correction strategy making and implementing method suitable for Mars detection. The method comprises the following steps that: S1, an error transfer model of a detector from any moment to the moment when the detector enters a Mars influence ball is established according to a limited three-body gravity orbit kinetic model, the orbit deviation of the detector arriving at the Mars at the current moment is calculated according to the error transfer model, and when the orbit deviation exceeds a deviation preset value, the method shifts to S2; S2, the speed error ofthe detector is corrected at the current moment by adopting a differential correction method with the parameter of entering the B-plane of the orbit of the Mars taken as a target, and iterative computation is performed with the error transfer model in the S1 to obtain pulse speed increment at the current moment; and S3, according to the pulse speed increment in the step S2, an orbit control duration and an orbit control ignition direction are obtained. With the method adopted, the autonomous formulation and implementation of a midway correction strategy under limited computing resources are realized. The method is simple and feasible, and Mars detection reliability under the condition of the absence of ground support is improved.

An autonomous mobile mechanically deployed spaceport is disclosed to provide a self-leveling stable landing pad for lunar and Martian descending and ascending spacecraft and which alleviates rocket plume blast effects upon surface soils and volatiles during spacecraft descent and ascent. The autonomous mobile mechanically deployed spaceport is a mechanically deployed unfolding flying landing pad to allow spacecraft to land on the Moon and Mars and alleviates rocket plume blast effects upon surface soils and volatiles.

The invention provides a high-precision Mars high-rise atmospheric wind field and density inversion method. The specific implementation method comprises the following steps: constructing a relationship between atmospheric resistance acceleration entering a ball and parameters such as atmospheric wind speed and density, determining trajectory attitude parameters entering the ball by using inertialcomponents such as an accelerometer and a gyroscope, realizing atmospheric wind field parameter inversion based on an atmospheric resistance direction measured by the accelerometer, and performing modeling based on the atmospheric resistance modulus value and the atmospheric resistance coefficient to realize atmospheric density inversion. The method has the advantages of high precision and low cost; the effective load only comprises one three-axis accelerometer and one gyroscope, a high-precision result can be obtained based on high-precision data given by a precision instrument; the method can be realized based on a micro-nano satellite platform, a micro-nano satellite detection satellite group is established on the basis, and the full-coverage detection of the Mars high-rise atmosphere can be realized.

The invention relates to an autonomous thrust system and method for the safety of the separation of a Mars exploration surrounding device. The system comprises a large thrust system, a small thrust system, and a non-thrust direction small thrust system, and can work in three working modes; under the first air injection mode, the large thrust system is used for the orbit control of the Mars exploration surrounding device; the small thrust system is adopted to control the Y-axis and Z-axis attitudes of a Mars detection surrounding device body coordinate system; in the first air injection mode, when the thruster in the small thruster system breaks down, the first air injection mode is switched to the second air injection mode; the large thrust system is adopted to control the Y-axis and Z-axis attitudes of the track of the Mars exploration surrounding device and the body coordinate system; in the first air injection mode, when the large thruster breaks down, the first air injection mode is switched to the third air injection mode; the large thrust system is adopted to control the Y-axis and Z-axis attitudes of the track of the Mars detection surrounding device and the body coordinatesystem; and in the three modes, the non-thrust direction small thrust system is adopted to control the X-axis attitude of the Mars detection surround device body coordinate system.

The invention provides a Mars rover moving system vibration loading mechanism and a performance testing device and belongs to the technical field of Mars detection testing equipment. The mechanismis used for carrying out vibration loading on a tested Mars rover and comprises a first vibration platform, a vibration loading assembly and a second vibration platform which are sequentially arranged from top to bottom, the first vibration platform is used for being arranged at the bottom of a wheel of the tested Mars rover, the vibration loading assembly is arranged between the first vibration platform and the second vibration platform, and the vibration loading assembly is connected with the second vibration platform. Compared with the prior art, the amplitude and frequency of the wheels are adjusted through the eccentric distance adjusting structure, single-degree-of-freedom vibration loading of the tested Mars rover within the range of 0-10 Hz is achieved, the mechanismis simple, testingis accurate, and precision is high.

The invention relates to a leg type landing and buffering device capable of being repeatedly used. The leg type landing and buffering device is characterized in that foot pads are connected with support leg main supporting columns through ball hinge connecting heads, metal rubber blocks are arranged between ball hinges and the foot pads, one-way dampers and long springs are arranged between innersleeves and outer sleeves, the one-way dampers do not have damp in the compression process and have damp in the stretching process, the long springs compress so that impact kinetic energy can be stored, after the maximum compression amount is reached, the one-way dampers work, the springs slowly restore to the original lengths, so that the attitudes of landing legs are restored, a load box is connected with an upper frame and a lower frame through a CR-type steel wire rope shock absorber, and is connected with an outer cylinder through a GS-type steel wire rope shock absorber, and the outer cylinder is connected with an inner cylinder through adjustable dampers and short springs. According to the leg type landing and buffering device, multiple shock isolation mechanisms are formed, so thatthe shock isolation efficiency is improved, and the leg type landing and buffering device can be repeatedly used, and is suitable for being used in landing and detection activities of planets such asthe earth, the moon, the mars and the minor planet.

The invention discloses a Mars probe autonomous orbit determination method by considering a sun vector, and the method enables a Mars probe to complete the orbit determination just through a sun sensor and a Mars sensor carried by the Mars probe in a sun-around flight process of a ground fire transfer section without depending on the ground support. The Mars detector performs orbit extrapolation according to the orbit recursion model; while the Mars sensor is used for measuring the position of the Mars probe relative to the Mars, solar vector measurement is introduced to further restrain the orbit extrapolation error. Compared with the prior art, the orbit determination method has the beneficial effects that the Mars detector can obtain stable observed quantity under the condition that a tracking Mars target is lost or a Mars sensor breaks down, and the reliability of orbit autonomous determination is improved.

The invention discloses an autonomous management method suitable for a propulsion system in Mars exploration. Faults of the propulsion system are divided into attitude control thruster leakage, rail control pipeline overpressure, pressure reducing valve static pressure overpressure, propulsion system underpressure and attitude control pipeline overpressure; the priority of the fault processing ofthe attitude control thruster leakage is the highest; the priorities of the fault processing of the other four faults are the same, and the fault processing is carried out in sequence according to a fault occurrence time sequence. According to the autonomous management method for the attitude control thruster leakage, a current thruster is switched out, so that other thrusters can be switched in.According to the autonomous management method for the rail control pipeline overpressure, a valve on a rail control pipeline is opened and closed. According to the autonomous management method for thepressure reducing valve static pressure overpressure, the pressure of a pressure reducing valve is reduced through gas path pressure balance. According to the autonomous management method of the propulsion system underpressure, a storage tank is pressurized; according to the autonomous management method for the attitude control pipeline overpressure, pressure relief is conducted on an attitude control pipeline. According to the method, the safety and reliability of the propulsion system in the whole Mars detection process are improved.

The invention provides a planetary exploration measurement and control communication system, which is used for ultra-remote communication scenes including Mars and Jumur exploration, and comprises two power amplifiers which are configured to be respectively provided with a high transmission power and a low transmission power, the low transmission power is used for near-field communication and protecting ground deep space measurement and control system equipment, the high transmission power is used for far-field communication, and the high transmission power is used for near-field communication and protecting ground deep space measurement and control system equipment. Enough emission EIRP is ensured; four pairs of antennas, wherein each pair of antennas is configured to be capable of transmitting and receiving signals at the same time; the high-isolation duplexer is configured to ensure that each antenna can be used for transmitting and receiving signals at the same time; wherein the four antennas are selected, and the output powers of the two power amplifiers are selected, so that measurement and control communication in various distance ranges is respectively carried out.