76 results about "Nano satellite" patented technology

The invention discloses a micro-nano calibration satellite constellation cooperative control method based on GA and DRL coupling optimization, and relates to the field of satellite networking and constellation optimization design. The method comprises the following steps: constructing a micro-nano satellite constellation task model based on constellation information data; performing global optimization on the constellation parameter vector according to the fitness function in a full constellation parameter space by adopting a genetic algorithm based on the micro-nano satellite constellation task model to obtain an initial constellation configuration set; determining a strategy parameter set according to the initial constellation configuration set by adopting a deep learning method; and carrying out in-orbit deployment according to the strategy parameter set, carrying out state estimation on each satellite, carrying out dynamic updating on the satellites based on the satellite execution control quantity, and completing task allocation and maneuvering negotiation through inter-satellite links by adopting a satellite distributed cooperative strategy so as to realize constellation-level execution control. According to the invention, constellation-level optimal configuration and autonomous cooperative control can be realized, and the in-orbit calibration capability of the micro-nano satellite constellation is improved.

The invention discloses an assembly line and an assembly method for batch production of stacked micro-nano satellites, belongs to the technical field of batch production of satellites, and particularly relates to the assembly line and the assembly method suitable for batch production of the stacked micro-nano satellites. The problems that an existing satellite is low in production (assembly) efficiency, high in cost and long in period are solved. The method comprises the following steps: performing single-machine thermal control implementation and component assembly to form a single-machine component; cabin section assembling is conducted, and a complete cabin section module is formed; carrying out platform cabin closing to form a complete whole-satellite platform; butting the upper cabin and the lower cabin of the load to form a satellite in a middle state; and solar wing assembly is carried out to form a final whole satellite in a to-be-launched state. The assembly line for batch production of the stack type micro-nano satellites and the assembly method are suitable for batch production of the stack type micro-nano satellites.

The invention discloses an omnidirectional space group target detection method based on micro-nano satellite group cooperation. The method comprises the following steps: carrying out observation configuration initialization on a micro-nano satellite group; establishing an intelligent detection model, and initializing observation star and decision star detection parameters and the intelligent detection model; the observation star performs target detection through the intelligent detection model to form a single-frame detection result; the decision-making star receives the single-frame detection result of each observation star to obtain summarized single-frame detection information; the decision star performs cross-view collaborative target association according to the summarized single-frame detection information, and updates a potential target list; and the decision star performs multi-view multi-frame collaborative target detection according to a single-frame detection result and the potential target list, updates a suspicious target list, and finally obtains suspicious target information. The method has the capability of autonomous information collaboration of seven micro-nano satellites, realizes omnibearing coverage target detection of a 4pi space above 200km, and has the capability of cross-view target association and multi-angle group target distinguishing during collaborative detection.

The invention provides a calibration performance evaluation system and a calibration performance evaluation method suitable for a single performance index, and relates to the technical field of spaceflight. The method comprises the steps of receiving a remote control instruction, identifying a task intention and triggering an evaluation mode; acquiring a real-time physical state of a satellite, and calculating a dynamic calibration reference value of a single efficiency index based on a physical model; obtaining an actual measurement result of the index probe type measurement, and calculating the deviation between the actual measurement result and a reference value; and combining the task intention and the deviation adjustment evaluation parameter, and outputting a self-calibration calibration evaluation result. According to the method, the problems that dynamic response is lagged and errors are introduced by manual intervention in traditional multi-dimensional evaluation can be solved, rapid and reliable calibration of a single efficiency index in space-based micro-nano satellite networking is achieved, and safe switching of high-value services is guaranteed.

The invention discloses a micro-nano satellite attitude semi-physical simulation experiment demonstration system, which comprises a control system, a semi-physical simulation machine and a three-axis turntable used for installing a tested micro-nano satellite, and is characterized in that through the cooperation of the control system, the semi-physical simulation machine and the three-axis turntable, a solution-control-feedback-iteration closed-loop process is formed; the specific operation steps of the demonstration system are as follows: S1, initialization and parameter calculation; s2, issuing and driving a target attitude; s3, physical state feedback and closed loop iteration are carried out; through cooperation of the control system, the semi-physical simulation machine and the three-axis rotary table and superposition disturbance of the semi-physical simulation machine, unbiased mapping of physical motion-digital model-space environment is realized, attitude dynamic characteristics of a micro-nano satellite in a real space can be reproduced, simulation fidelity is greatly improved, and the simulation precision of the micro-nano satellite is greatly improved. And the high-precision micro-nano satellite attitude simulation requirement is met.

The invention provides a low-cost storage and calculation integrated intelligent management system applied to a low-orbit commercial micro-nano satellite. The low-cost storage and calculation integrated intelligent management system comprises a ZYNQ core processing unit, a GPU computing power unit, a storage unit 1, a storage unit 2 and an interface unit, the ZYNQ core processing unit is used for real-time data acquisition, low-delay control and hardware-level anti-radiation design, and transmission of data streams on a satellite and scheduling of tasks on the satellite are realized; the power management unit and the GPU computing power unit jointly realize an intelligent power management function and a satellite-borne image processing function; the GPU computing power unit is responsible for executing a complex algorithm task and implementing dynamic optimization design, and is used for responding to a ZYNQ instruction to pre-process satellite data and schedule a computing task; the storage unit 1 and the storage unit 2 are responsible for storing historical data, algorithm model parameters and redundant backups; the storage and calculation integrated intelligent management equipment is low in cost, low in power consumption and high in reliability, and the industrial application barrier of on-orbit real-time intelligent processing of commercial micro-nano satellites is broken through.

The application discloses a capillary type ion liquid electric spray thruster with electrically controlled thrust, and belongs to the technical field of micro-nano satellite micro propulsion. The application comprises a propellant tank, an upper shell, a constant force spring, a limiting plate, a capillary and an extraction electrode. The inner and outer walls of the capillary are covered with a dielectric layer with a micron level thickness, and the inner wall is covered with a hydrophobic layer with a nanometer level thickness. The extraction electrode is fixed through an insulating screw; and the upper shell is fixed with the propellant tank. The application is based on the principle of electric wetting, uses an electric field to control the active supply of ion liquid and simultaneously adjusts the thrust size of the thruster. The application can not only save complicated fluid control components such as high-precision pumps, valves and pipelines, but also can couple the supply process and the emission process of the ion liquid, simplify the operation process of the thruster system and facilitate the miniaturization of the thruster system. The application is suitable for resistance compensation, position keeping and attitude control of low-power micro-nano satellites.

The invention relates to an ultra-large field-of-view dual-waveband aperture multiplexing type optical remote sensing camera, and belongs to the technical field of optical remote sensing. The optical remote sensing camera comprises a visible light optical system and an infrared optical system which share a primary mirror and a secondary mirror. The visible light optical system comprises a primary mirror, a secondary mirror, a visible light third mirror, a reflex mirror and a visible light focal plane in the light path direction; the infrared optical system comprises a primary mirror, a secondary mirror, an infrared third mirror and an infrared focal plane in the light path direction. According to the ultra-large field-of-view dual-waveband aperture multiplexing type optical remote sensing camera, all-weather ground investigation and surveying and mapping can be realized; primary mirrors and secondary mirrors of the visible light optical system and the infrared light optical system can be reused, the focal lengths of the two systems are different, the view fields of the visible light optical system and the infrared light optical system can be 20 degrees or above, and the ultrahigh energy utilization rate of the system is guaranteed through the total reflection mode. The camera is especially suitable for being used as a high-resolution camera of a wide-area detection, small-size and multi-spectrum integrated micro-nano satellite.

This invention discloses a composite attitude control method for the maneuvering phase of a microsatellite. The method includes: constructing a novel composite actuator based on a zero-momentum gear train and a two-dimensional translational mechanism; establishing and optimizing attitude rotational dynamics and slider translational dynamics models considering inertial principal axis offset and additional disturbance torque; designing an exponentially convergent disturbance observer to monitor thrust eccentricity torque in real time; and designing an attitude control loop MPC (Model Predictive Control) controller and a thrust eccentricity torque compensation loop discrete PID (Proportional-Integral-Derivative) controller to form a dual-closed-loop composite attitude control system. This control method can effectively observe slowly time-varying thrust eccentricity torque, and features high control accuracy and no saturation in the momentum gear train, making it suitable for long-term thrust eccentricity disturbances in microsatellite maneuvering missions.

This invention relates to a method for inverting remote sensing data from overlapping strong and weak absorption lines of greenhouse gases, belonging to the field of satellite remote sensing technology. Specifically, it includes the following steps: Step 1: Constructing an optical system, including a satellite platform, camera 1, and camera 2; camera 1 and camera 2 are mounted at the bottom of the satellite platform; camera 1 and camera 2 are mirrored; the imaging spectral band of camera 1 moves from the weak absorption spectral band to the strong absorption spectral band along the positive push-broom direction to the negative push-broom direction; the imaging spectral band of camera 2 is in the opposite direction; the optical system moves horizontally from left to right to achieve horizontal push-brooming of greenhouse gases on the ground, obtaining a dual-camera fused image; the dual-camera differential absorption spectroscopy algorithm is used to invert the dual-camera fused image to obtain the greenhouse gas column concentration; this invention, based on the spectral characteristics of strong and weak absorption lines of greenhouse gases, adopts a detector focal plane segmentation method, and the dual cameras achieve differential absorption and inversion in the same region, which can guide the design of greenhouse gas detection micro / nano satellite payloads.

The invention discloses a spacecraft force bearing and multi-layer flexible circuit integrated structure based on 3D printing. The spacecraft force bearing and multi-layer flexible circuit integrated structure is integrally formed by a metal force bearing layer, an electric insulation coating layer and a multi-layer flexible circuit functional layer through an in-situ continuous printing process, wherein the metal force bearing layer is made of metal additive materials. The metal bearing layer is a lightweight metal gradient dot matrix shell with complex internal functional channels; the electric insulation coating layer is sprayed or printed on the surface of the metal bearing layer in a surface modification binding force enhancing mode. And the multi-layer flexible circuit function layer is directly printed on the insulating layer across the curved surface and the edge. Physical and functional deep integration of a spacecraft main force bearing structure, a thermal management path, an electromagnetic compatibility substrate, electrical insulation and a high-low frequency circuit is realized, and the systematic problems of large size, heavy weight, multiple interfaces and low reliability of a traditional discrete design are solved. And the method is particularly suitable for platforms with strict requirements on space, weight and reliability, such as micro-nano satellites and the like.

A kind of space micro debris monitoring micro-nano satellite based on double-layer film sensor.The satellite carries out bidirectional verification through non-vision in-situ measurement of double-layer film sensor and visual measurement of monitoring camera, carries out monitoring on space micro debris with size less than 1 cm, obtains data information such as velocity vector of debris, and provides real-time and accurate micro debris cloud data.The normal direction of satellite film is the same as the direction of satellite speed, which increases the probability of collision monitoring between satellite and space micro debris.The folding and unfolding mechanism adopted by the present application has small folding volume and is convenient to unfold, and can be used as a debris monitoring device and satellite de-orbit device commonly installed in micro-nano satellite module, and has double functions.The locking-releasing mechanism adopted by the present application can realize locking of the whole mechanism only by using small pre-tightening force, and the design of burning rope release can reduce the separation impact to the minimum, and improve the stability and safety of the system.

The invention relates to the technical field of space laser communication, and discloses a laser communication device based on autonomous learning regulation and control of channel characteristics, which comprises an autonomous regulation and control communication transmitting end, an autonomous regulation and control communication receiving end and an alignment capture tracking module, the device can realize self-adaptive switching between rates V1-V4 and modulation formats M1 / M2 in a complex channel environment, and supports a quick rollback strategy under the conditions of communication out-of-step and link fading; compared with the prior art, miniaturization and full-duplex self-adaptive regulation and control capability of the laser communication device are realized, the link stability and the channel utilization rate are remarkably improved, and the requirements of micro-nano satellites on low-power-consumption and high-reliability laser communication are met.

A kind of micro-nano satellite module rapid integration method belongs to the field of relay satellite platform, there is standardization backplane and standardization module, standardization backplane is provided with bus area and slot area, bus area includes uniform specification power bus and data bus, slot area is arranged with multiple standardization slots, standardization module is inserted into standardization slot.The structure of standardization slot is uniform, and a uniform standard data electrical interface is provided inside, each standardization slot corresponds to fixed physical port and logical address / node number, or unified address configuration is realized through slot coding, according to functional task requirement, select the standardization module that meets functional task requirement, and build the satellite system that meets functional task requirement.The present application significantly shortens the development cycle, greatly reduces the development cost, greatly improves the system flexibility and scalability, improves the system reliability and maintainability, and promotes standardization and industrialization.

A rapid integration method for micro-nano satellite modules belongs to the field of relay satellite platforms, a standardized backboard and standardized modules are provided, the standardized backboard is provided with a bus area and a slot area, the bus area comprises a power bus and a data bus which are unified in specification, a plurality of standardized slots are arranged in the slot area, and the standardized modules are inserted into the standardized slots. The standardized slots are uniform in structure, data electric appliance interfaces with uniform standards are arranged in the standardized slots, each standardized slot corresponds to a fixed physical port and a logic address / node number, or uniform address configuration is achieved through slot position coding, and standardized modules meeting function task requirements are selected according to the function task requirements. And a satellite system meeting functional task requirements is constructed. According to the method, the research and development period is remarkably shortened, the research and development cost is greatly reduced, the flexibility and expandability of the system are greatly improved, the reliability and maintainability of the system are improved, and standardization and industrialization are promoted.

The application discloses a kind of inductively coupled plasma micro-propeller of double antenna structure, and its discharge system includes gas source, air inlet valve, antenna, discharge tube and radio frequency power supply.Gas source is directly connected with discharge tube by valve, wherein the valve can control the air intake during discharge;Then two identical helical antennas are connected in series at a certain distance and are wound on the same discharge tube, the output end of radio frequency power supply is connected to the end of antenna away from air inlet through impedance matcher, and the other end of radio frequency power supply and antenna is connected to the same ground.The application has simple structure, small space size and mass, and only needs to work stably at low power and low air intake, realizing micro-newton to millinewton level thrust.It is suitable for equipping in most of the current constellation micro-satellites without propulsion system of nanosatellite and cubic satellite.In addition, the length and spacing of the two antennas in the system are adjustable, and the size of the propeller can be adjusted according to the specific requirements of micro-satellite.

This invention relates to a method, system, and device for dynamically adjusting the protection threshold of a micro / nano satellite battery, belonging to the field of power management technology. The method includes: real-time acquisition of voltage, current, and temperature parameters; real-time estimation of the health status value of each individual battery cell based on the acquired voltage, current, and temperature parameters; using the estimated health status value as a query index to query a preset mapping table between health status and protection thresholds to obtain a dynamic protection threshold for at least one protection type corresponding to the current health status value; comparing the voltage, current, and temperature parameters with the dynamic protection thresholds respectively to determine if any parameter exceeds the corresponding dynamic protection threshold; and executing a preset battery protection action when any parameter exceeds the corresponding dynamic protection threshold. This invention can sense the battery's own health status and adaptively adjust the protection threshold according to its aging degree.

The utility model discloses a sliding unfolding device for a micro-nano satellite, and relates to the technical field of satellites. According to the technical key points, the satellite comprises a satellite main body, solar wing plates are hinged to the opposite sides of the satellite main body respectively, and elastic driving pieces are arranged at the joints of the solar wing plates and the satellite main body; a telescopic supporting unit is installed in a shell of the satellite body, the top of the supporting unit is connected with a horizontal linkage rod, and the two ends of the horizontal linkage rod are slidably connected with the shell of the satellite body in the vertical direction. The two ends of the horizontal linkage rod are each provided with a clamping head, and the two solar wing plates are each provided with a clamping groove. An adjusting unit capable of controlling the position of the horizontal linkage rod at the top of the supporting unit in the vertical direction is arranged on the upper side of the middle position of the horizontal linkage rod. According to the unfolding device, the solar wing plate can be rapidly unfolded according to needs, and components in a satellite can be effectively prevented from being damaged by high temperature when the unfolding device is used.

The application discloses an on-orbit use method of a green single-component thruster, and the single-component thruster structure comprises five parts of an injector, an anti-inversion functional material, a catalytic bed, a nozzle and a thermal control assembly, wherein the thermal control assembly comprises an armored heater, an armored thermistor and a multilayer thermal insulation assembly. The application provides the on-orbit use method of the green single-component thruster containing the anti-inversion functional material, and solves the technical problem of the warm start life of the green single-component thruster through reasonable working condition design, and the green single-component thruster is successfully applied to micro-nano satellites.

The invention relates to a micro-nano satellite group distributed cooperative multi-target tracking method, and belongs to the technical field of space target tracking. Through inter-satellite cooperative networking construction, single-satellite weak and small target detection, single-satellite multi-target tracking, cross-view-field track association, multi-satellite distributed information fusion and the like, the target tracking number of satellite group cooperation can be greater than or equal to 12, and support is provided for the micro-nano satellite group in the fields of space situation awareness, on-orbit target monitoring and the like. According to the method, the problems of poor robustness of a centralized architecture, insufficient cross-view-field tracking precision and the like in micro-nano satellite group multi-target tracking are solved, the real-time performance and reliability of space target tracking are improved, and the requirement for accurate and stable tracking of multiple targets in a complex space environment is met.

This invention relates to the field of spacecraft mission planning technology and discloses a method for autonomous mission planning of micro- and nano-satellite reconnaissance based on predictive spatiotemporal heat maps. The method includes: generating a predictive spatiotemporal heat map characterizing the predicted severity and observational uncertainty of a disaster area by fusing multi-source observation data with a disaster evolution physics model through Gaussian process regression of physical information; constructing a rolling temporal control planning model coupled with satellite attitude, energy, and orbital constraints based on the predictive spatiotemporal heat map for each satellite; solving the planning model online using a sequential convex optimization algorithm to obtain the optimal observation mission sequence and attitude maneuver strategy for each satellite; dynamically updating the observation value of the planned area through a shared heat map based on a coverage suppression mechanism, guiding neighboring satellites to automatically turn to unobserved high-value areas to achieve distributed collaborative mission allocation; and updating the predictive spatiotemporal heat map in a closed loop using newly acquired observation data.

The application discloses a method for adjusting and controlling the phase of a low-orbit micro-nano satellite constellation by means of atmospheric differential drag, and belongs to the field of on-orbit phase control of a constellation. The method aims at solving the problem of low phase control precision of the existing atmospheric differential drag control satellite constellation method. The method comprises the following steps: establishing a single-satellite atmospheric drag dynamics model to obtain the corresponding relationship between the single-satellite control input and the control command; determining the state constraint condition of other single satellites based on a reference satellite, combining the energy constraint and the task constraint, and adopting a simulated annealing algorithm to perform constellation slot allocation, and taking the shortest control time as the target to obtain the optimal result of the constellation slot allocation; calculating an initial control command matrix; taking the lowest adaptive function value and meeting the target control precision as the control target, optimizing the initial control command matrix to obtain an optimal control command matrix, so as to obtain the optimal satellite control input, and performing the angular acceleration control of the single satellite, thereby realizing the phase adjustment of the same-orbit constellation. The application is used for the phase adjustment control of the constellation.

The application discloses a kind of micro-nano satellite angular momentum autonomous management method, belongs to satellite attitude control technical field.The application is in view of the problem that existing satellite angular momentum management adopts open loop method, and autonomous correction capability is poor.The application comprises: satellite is identified through on-orbit flywheel rotating speed data environment torque parameter autonomously, and parameter that changes speed is faster is updated autonomously along with task time extension;Only through the attitude bias of relative sun orientation and sail rotation, angular momentum accumulation inhibition and angular momentum unloading two task objectives can be realized by manipulating environment torque, without active control torque intervention such as magnetic torque device, jet unloading;The whole process does not need artificial intervention, and angular momentum autonomous management is realized.The application can be autonomously carried out when satellite has no task, and reliability is good.

The invention provides an engineering plastic wire suitable for FDM additive manufacturing of a circuit board substrate, the engineering plastic wire comprises engineering plastic resin, chopped carbon fiber and a modification auxiliary agent, the engineering plastic resin accounts for 40-80% of the mass of the 3D printing wire, the chopped carbon fiber accounts for 10-20% of the mass of the 3D printing wire, and the balance is the modification auxiliary agent. The 3D printing wire material is prepared from the high-performance engineering plastic, the short carbon fibers and the modification additive, the problem that an existing FDM wire material for additive manufacturing cannot be directly used for additive manufacturing forming of parts such as spacecraft electronic device PCBs is directly solved, and the research, development and application processes of integral additive manufacturing of micro-nano satellites and other small spacecrafts are greatly promoted; meanwhile, the linear expansion coefficient of the material is matched with that of a common circuit material, the performance is extremely excellent, and the material can be used for aerospace and other fields with printed circuit board additive manufacturing requirements in severe environments.

A nanosatellite configured for launching from a cube satellite launcher is disclosed. The nanosatellite has a lower portion sized to the fit within the launcher housing, and has top, bottom, left, right, front and rear sides forming a payload housing. The top side has a top surface configured to engage with a front end of the spring of the launcher. The nanosatellite has an upper portion extending from the top surface of the lower portion, and is sized to fit within the spring of the launcher. A control subsystem contained within the upper portion. The nanosatellite may have one or more a solar panel arrays connected to the lower portion. In some embodiments, the solar arrays are movable between a stowed position against the lower portion and a deployed position extending outwardly from the lower portion. An additional 1U-3U of payload may be connected to the lower portion.

The invention discloses a micro-nano satellite solid-liquid mixed thruster arc ignition circuit and an ignition control strategy. The arc ignition circuit comprises a control module, an ignition module, a signal acquisition module, an ignition switch circuit, a secondary power supply and a bus. The ignition control strategy comprises a three-section type ignition protection strategy and two ignition control strategies. The state of the system is monitored and managed in real time, so that stray current between two ignition charge electrodes and in a front-end circuit can be effectively eliminated, and controllable ignition is realized. A good solution is provided for the design of a high-reliability and high-safety solid-liquid mixed arc ignition system in a complex space environment.

The integrated nanosatellite data processing platform comprises a nanosatellite and includes a sensor module bidirectionally connected to an on-board computer. The on-board computer is bidirectionally connected to a communication module bidirectionally connected to a communication antenna. The communication antenna is connected to a ground antenna bidirectionally connected to a communication equipment unit of a ground station. The ground station is bidirectionally connected to a control server bidirectionally connected to a database. The database is bidirectionally connected to a market portal server, a developer portal server, and an application server. The control server, market portal server, developer portal server, and application server are bidirectionally connected to a client station.

The invention discloses an L-band high-orbit SAR full-link radiometric calibration method based on satellite-ground cooperation, and the method comprises the following steps: S1, obtaining a hardware gain function changing along with a space angle based on echo data and RCS information between a micro-nano satellite carrying an active transponder and a high-orbit SAR satellite; s2, a pre-trained deep learning model is utilized to obtain software gains changing along with working parameters; and S3, performing high-precision radiometric calibration on the input to-be-calibrated high-orbit SAR image by using a full-link gain model obtained based on a hardware gain function and a software gain, and inverting the absolute radar cross-sectional area sigma of the target. According to the method, the problems of high-sea far-field calibration difficulty, long calibration period, uncalibrated ground processing software gain, obvious atmospheric effect influence and the like of a high-orbit SAR satellite can be solved, the separation calibration of the gain of a whole link from a space antenna to the ground processing can be realized, and the difficulty of software and hardware gain coupling in a traditional method is solved.

The invention provides a cabin plate butt-joint locking structure for mass production of micro-nano satellites and a use method of the cabin plate butt-joint locking structure, and belongs to the field of satellite structures. The problems that a traditional screw connection mode is low in final assembly efficiency, poor in operation reliability, inconvenient to test and maintain quickly, difficult to meet the batch production requirement and the like are solved. The cabin comprises a cabin plate, locking execution units and lock seats, the cabin plate is divided into a bottom plate, a top plate and a plurality of side plates, the side plates are connected with the bottom plate through hinges, the locking execution units and the lock seats are arranged on the two butt joint edges of the adjacent side plates respectively, and the locking execution units and the lock seats are arranged on the two butt joint edges of the side plates and the top plate respectively. The adjacent side plates and the side plates and the top plate are locked in a butt joint mode through matching of the locking execution units and the lock seats. The method is mainly used for mass production of micro-nano satellites.

The application discloses a docking mechanism and method capable of repeatedly unlocking micro-nano satellites, belongs to the technical field of on-orbit satellite capturing, and comprises an active docking mechanism and a passive docking mechanism; the active docking mechanism and the passive docking mechanism are respectively installed on two spacecrafts, an electromagnetic locking mechanism in a taper-rod type structure is used, electromagnetic lock cores installed on the active docking mechanism and the passive docking mechanism are matched with electromagnetic lock valves, two spacecrafts are guided to accurately dock and repeated locking and unlocking are realized. The taper-rod type docking is adopted, attitude angle control is relatively flexible, the capturing range is relatively large, good buffering performance is obtained by using a spring damper, and the passive mechanical locking mode, the laser beam intersection sensor technology and the electromagnetic device are combined, the reliability of the docking process is improved, and the docking mechanism has the functions of electric transmission and signal transmission.