720 results about "Aerospace technology" patented technology

The invention discloses a morning orbit computing power satellite, and belongs to the field of aerospace technology and space infrastructure. The satellite comprises a satellite platform and a computing force load arranged on the satellite platform. The satellite platform is provided with a condensation type energy system and a space pump drive fluid loop system. The condensation type energy system adopts an expandable condensation type solar cell array to generate power, switches a power supply mode according to an illumination period and an earth shadow period of a morning and night orbit, and supplies power to computing force loads and platform equipment. The space pump drive fluid loop system adopts an expandable flexible radiation cooling plate for heat dissipation, heat dissipation is conducted on computing force loads through a fluid loop, and a heating assembly is arranged to prevent a working medium from being frozen under the low-temperature working condition. Through the expandable structure and the orbit environment adaptability design, continuous energy supply and effective heat management are provided for the computing power load in the morning and night orbit environment, and high-power-consumption computing power load in-orbit stable operation is supported.

The invention relates to the technical field of aerospace, in particular to low-altitude traffic flow airspace-oriented real-time planning, and provides a centimeter-level precision detection network covering a low-altitude airspace by integrating multi-dimensional data sources such as radar electromagnetic feature recognition, ADS-B (Automatic Dependent Surveillance-Broadcast) automatic monitoring, Beidou or GPS (Global Positioning System) space-time reference positioning and the like. The three-dimensional trajectory and motion situation of the aircraft are solved in real time, intelligent reconstruction of an airspace sector and adaptive optimization of a flight corridor are realized by adopting a dynamic programming algorithm driven by reinforcement learning based on the real-time pose data of the aircraft, and a flight conflict prediction model constructed by combining a space-time convolutional neural network is used for predicting the flight conflict. According to the method, potential risks can be pre-judged, an optimal avoidance path can be generated, full-process digital management and control from identity verification to airspace authorization can be realized by constructing an aircraft digital identity authentication system and a dynamic access control mechanism, and modules such as a three-dimensional navigation information service module, a low-altitude digital communication private network module and an intelligent early warning and warning system module are integrated. And the guarantee of full-life-cycle service is provided for the low-altitude aircraft.

The invention belongs to the technical field of aerospace, and relates to an unscented batch processing orbit determination method and system for a continuous low-thrust maneuvering satellite. The method comprises the following steps: acquiring a radar observation vector according to radar measured data or satellite ephemeris data in combination with radar site coordinates; obtaining an initial orbit state at a to-be-estimated moment by adopting a Lambert algorithm according to the radar observation vector; generating an initial expansion state quantity according to the initial orbit state at the to-be-estimated moment; a simplified orbit kinetic equation containing earth non-spherical perturbation, atmospheric resistance perturbation and constant small thrust acceleration under an RTN coordinate system is constructed under the ECI coordinate system; based on the simplified orbit kinetic equation, obtaining an expansion state quantity precision value of the to-be-estimated moment through the initial expansion state quantity, the radar observation vector and an unscented batch processing filtering method; and substituting the precision value of the expansion state quantity of the to-be-estimated moment into the simplified orbit kinetic equation to carry out orbit forecasting at other moments. According to the method, the orbit determination precision is remarkably improved, and the influence of single-point observation errors is reduced.

The invention relates to the technical field of aerospace, and discloses an ALOT-based low-altitude intelligent obstacle avoidance route planning system, which comprises a data collection module, a data calculation module, a feedback module, a feedback optimization module and an optimization management module, and is characterized in that the data collection module is responsible for collecting various environmental information required in the flight process of an unmanned aerial vehicle; the data calculation module performs calculation and analysis by using an algorithm based on the data collected by the data collection module to generate a preliminary dynamic route planning model, the feedback module performs real-time monitoring and feedback on a calculation result of the data calculation module and evaluates the flight state of the unmanned aerial vehicle, and the feedback optimization module adjusts a flight strategy according to a feedback result. According to the method, obstacle avoidance and route planning are optimized, optimization measures for adjusting the flight speed and changing the route are put forward in time, an optimization management module manages and implements the optimization measures, it is ensured that the unmanned aerial vehicle flies safely and efficiently according to the optimized route, meanwhile, the implementation effect of the optimization measures is monitored, and effective adjustment is conducted.

The invention discloses a cooperative flight path planning method for offshore aircrafts, and relates to the technical field of aerospace, and the system comprises the following components: S1, a task priority setting step and S2, a cooperative flight path planning step. According to the method, a multi-dimensional task evaluation index system is established, the task priority of a marine aircraft is accurately set, it is ensured that key tasks can be processed preferentially, meanwhile, the weight is dynamically adjusted according to the real-time situation by adopting a dynamic weight self-adaptive search flight path planning algorithm, and the task priority is accurately evaluated. According to the method, the flight path which meets task requirements and is optimized as much as possible is planned for the aircrafts, the efficiency of cooperative operation of the offshore aircrafts is improved, flight path conflicts among the aircrafts can be effectively avoided, and flight safety is ensured.

The invention provides a novel ceramic matrix composite material and metal material combined adjustable turbine blade, and belongs to the technical field of aerospace, the novel ceramic matrix composite material and metal material combined adjustable turbine blade specifically comprises a metal upper edge plate and a metal lower edge plate; the ceramic matrix composite material front edge assembly is fixedly installed between the metal upper edge plate and the metal lower edge plate, and the ceramic matrix composite material front edge assembly serves as a front edge area of the blade; the adjustable metal blade body is rotationally mounted between the metal upper edge plate and the metal lower edge plate; the rear end of the blade body part of the ceramic matrix composite material front edge assembly is in butt joint with the front end of the adjustable metal blade body to form a complete turbine blade body structure. Through the treatment scheme, the problem that in the prior art, cooling of the front edge of the metal blade consumes a large amount of cold air is solved, and the service life of the adjustable turbine blade is prolonged.

The invention discloses a predetermined time expansion state observer generation method and an observation error convergence method, which are applied to the technical field of aerospace, and comprise the following steps: constructing a controller for a second-order nonlinear system according to an all-drive system theory; converting the second-order nonlinear system according to the controller to obtain an all-drive system based on the main controller; wherein the main controller is a controller for controlling the all-wheel-drive system; designing a predetermined time expansion state observer based on the total disturbance and observation error of the all-wheel-drive system; wherein the predetermined time expansion state observer is an observer which limits the upper bound of convergence time of the observer based on a predetermined time parameter. Compared with most existing extended state observers which can only ensure asymptotic convergence and finite time or fixed time convergence, the observer generated based on the predetermined time extended state observer generation method can ensure that the observation error converges within the predetermined time, so that the converged predetermined time is more stable; and the user experience is improved.

The invention belongs to the technical field of aerospace, and discloses an aviation electrical system insulation detection device which comprises a shell and an overall frame serving as the insulation detection device. The display screen is used for displaying measurement data in real time; the function key group integrates the functions of measurement mode switching, data retention and insulation test starting; according to the conversion assembly, through the linkage design of the rotating disc and the rotating switch, the problem that functions cannot be switched when the two hands of a traditional insulation detection device are occupied is thoroughly solved, a worker only needs to rotate the rotating disc with the hand of supporting equipment, the rotating switch can be driven through the driving mechanism, switching of insulation test voltage gears is completed with one hand, and the working efficiency is improved. Compared with a traditional device which needs to frequently put down a connecting line and manually rotate a switch, the design simplifies the operation steps from three steps to one step, shortens the single function switching time, and remarkably improves the efficiency and safety of high-altitude operation or narrow space detection.

The invention relates to the technical field of aerospace, in particular to a multi-objective optimization method for a pod-shaped composite space thin-wall deployable structure. The method comprises the following steps: selecting an optimal sample point for uniformizing the space deployable structure in an interval range of structure parameters; constructing a finite element analysis model of the pod-shaped composite material space thin-wall deployable structure; the method comprises the following steps: by taking a damage failure criterion as a failure mode judgment basis of fibers and a matrix of a pod-shaped composite material, describing an evolution path of material rigidity in a damage development process by utilizing a bilinear rigidity degradation model after failure occurs, and constructing a random forest model based on adaptive particle swarm optimization as an agent model; and a Pareto leading edge optimal solution is obtained by calling the proxy model and is used as an optimal pod rod structure. In this way, the axial bearing capacity and folding performance of the unfolded structure are comprehensively improved, and the optimized structure can show more excellent mechanical performance and adaptability to different scenes in a complex space environment.

The invention relates to an electric servo motor for IBC blade control, a motor system and an aircraft, belongs to the technical field of aerospace, and solves the problems that an electric servo motor system for the IBC blade control technology in the prior art is complex in structure, difficult in layout and poor in matching with blades and the aircraft. The motor comprises a rotor; the rotor comprises a rotating shaft; the rotational inertia of the rotor, the density, diameter and length of the rotor and the density and diameter of the rotating shaft have a relational expression. According to the invention, the matching degree of the motor system, the blades and the aircraft is high, the structure is simple, and the vibration of the blades can be reduced.

The invention relates to a steering engine and piezoelectric patch hybrid driven variable camber wing, belongs to the technical field of aerospace, and solves the problems of poor wing surface deformation continuity and poor transmission effect in the prior art. The plurality of flexible trailing edge deflection modules are arranged in the straight wing main body; the skin structure is arranged on the outer surfaces of the straight wing main body and the flexible trailing edge deflection module; the module connecting structure is connected with the plurality of flexible trailing edge deflection modules; wherein the straight wing main body comprises a front edge D-shaped structure wing box, a plurality of wing ribs arranged on the rear side of the front edge D-shaped structure wing box, and a bearing rear wall arranged on the rear sides of the wing ribs; each flexible trailing edge deflection module comprises a driving rib assembly fixed to the bearing rear wall, driven rib assemblies arranged on the two sides of the driving rib assembly, carbon rods sequentially connected with the driving rib assembly and the driven rib assemblies, an inhaul cable driving mechanism and a piezoelectric driving mechanism, and the driving rib assembly and the driven rib assemblies are sequentially arranged in the longitudinal direction of the variable-camber wing.

The invention relates to the technical field of space flight and aviation, in particular to a low-delay communication control system and method for an analog machine in a fixed flight training course. The control system comprises a multi-source sensor fusion module, a data simplification processing module, a dynamic time delay optimization communication framework, a distributed intelligent decision-making terminal, a virtual-real mapping synchronous controller, a milli delay regulation and control module and an intelligent sensing alarm interaction module. According to the invention, 4G, 5G and satellite double links are integrated, an optimal link is dynamically selected by adopting a Q-Learning algorithm, millisecond-level time delay control is realized, transmission reliability is improved through a packet loss rate compensation algorithm, multi-sensor data is fused based on an FPGA hardware accelerator and an EKF algorithm, and a data compression ratio is improved in combination with Huffman coding and Kalman filtering. The bandwidth requirement is greatly reduced, through software and hardware collaborative optimization, the end-to-end time delay is reduced to be within hundreds of milliseconds, the virtual-real synchronization precision is improved to the centimeter level, and the decision response speed is improved by two times.

The invention belongs to the technical field of aerospace, and relates to a TLE parameter prediction method and system based on TFT multi-source data fusion. The method comprises the following steps: generating a feature data set according to a historical TLE data set of an LEO uncontrolled space target, target static feature data and space weather data; adopting a # imgabs0 # model to obtain an angle parameter at the prediction moment according to the TLE data at the moment before the prediction moment; adopting a # imgabs1 # model to obtain core orbit parameters at a prediction moment according to the feature data set; acquiring a temporary TLE orbit parameter at the prediction moment according to the angle parameter at the prediction moment and the core orbit parameter at the prediction moment; updating a mean anomaly prediction value according to the current prediction step number and the temporary TLE orbit parameter at the previous moment; and replacing the mean anomaly preliminary prediction value in the temporary TLE orbit parameters at the prediction moment with the mean anomaly prediction value to obtain the TLE orbit parameters at the prediction moment. And the precision, real-time performance and efficiency of long-term orbit parameter prediction are improved.

The invention relates to a turbine guide vane structure, a turbine and an aero-engine, belongs to the technical field of aerospace, and solves the problems that in the prior art, a turbine guide vane structure with an impact pipe is prone to inclination of a guide pipe during installation, and the cooling effect is not ideal. The device comprises an impact pipe, wherein a plurality of positioning bosses are arranged on the outer wall surface of the impact pipe; an impact hole is formed in the positioning boss; a cooling through hole is formed in the wall face of the front edge of the impact pipe and used for forming impact cooling on the front edge of the blade. According to the invention, the positioning precision and the cooling effect can be improved.

The invention relates to the technical field of aerospace, in particular to a frame type space storage transfer device with diversified storage capacity and a transfer system. The transfer device comprises a storage device, and the storage device comprises a top plate, a containing device, a side plate moving device, a partition plate moving device and a vertical pushing device; the top plate is arranged on the containing device, the containing device comprises an outer fixing side plate, an inner fixing side plate, a movable side plate and a lead screw fixing side plate which are sequentially arranged at intervals, and movable partition plates are arranged on the outer fixing side plate, the inner fixing side plate, the movable side plate and the lead screw fixing side plate; the side plate moving device is located below the top plate and connected with the inner fixed side plate, the movable side plate and the lead screw fixing side plate, the partition plate moving device is connected with the movable partition plate, and the vertical pushing device is arranged on the inner fixed side plate, the movable side plate and the lead screw fixing side plate. And the transfer platform is located below the storage device and comprises a stator structure and a rotor structure. And the flexibility is improved, the transfer efficiency is high, and the storage space is large.

The invention relates to a main bearing locking nut anti-loosening structure and an aero-engine, belongs to the technical field of aerospace, and solves the problem of unreliable mechanical anti-loosening locking of a locking nut in the prior art. The anti-rotation device comprises a locking nut, a locking plate and an anti-rotation sleeve which are arranged outside a shaft neck of an engine rotating shaft in a sleeving mode. The locking nut fixing section is fixed to the shaft neck, and a first groove is formed in the outer portion of the clamping section in the circumferential direction. The locking plate comprises a locking plate main body, a first tooth and a second tooth; the first tooth is embedded into the first groove to limit the circumferential movement of the locking nut; the anti-rotation sleeve fixing part is connected with the shaft neck; the inner wall face of the anti-rotation part is attached to the outer wall face of the locking nut and the outer wall face of the locking piece. And the end part of the anti-rotation part is provided with a notch which is mutually embedded with the second tooth. Triple locking of the main bearing locking nut can be achieved, locking is reliable, and the structure is simple.

The invention relates to the technical field of aerospace, in particular to a ceramic 3D printing A sandwich structure with a broadband wave-transparent characteristic. According to the ceramic 3D printing A sandwich structure with the broadband wave transmission characteristic, good compatibility of broadband wave transmission, strength-structure accuracy and reliability is achieved, an obvious effect on broadband wave transmission of 1 GHz to 18 GHz is achieved, the dielectric constant and the tangent loss are in a reduction trend under the same frequency band, excellent stability is shown, and the dielectric constant and the tangent loss are reduced under 10 GHz. The dielectric constant of the ceramic 3D printing A sandwich structure with the broadband wave transmission characteristic is reduced by 40% or above compared with that of a full solid structure, the wave transmission rate is increased by 50% or above, the ceramic 3D printing A sandwich structure has excellent electromagnetic performance and mechanical performance, and support is provided for implementation of reliable service application targets such as diversion, heat protection, wave transmission, bearing and rain erosion resistance of hypersonic aircrafts.

The present disclosure relates to the technical field of aerospace, and provides a machining method for high-aspect-ratio model part of ultra-high strength steel for wind tunnel test. The machining method includes the following steps: a) selecting a material; b) performing preliminary treatment, such as forging and solid solution heat treatment, on the material; c) performing rough milling to obtain a wing main body profile, process reference blocks, and grooves and holes with large sizes on a molded surface; d) performing finish milling on all machining features of a wing main body; e) removing all process reference blocks except the first process reference block; f) performing aging treatment when the wing main body is lifted; h) removing a process reference block at a wing main body root; and h) performing shaping treatment on the wing main body.

The present invention belongs to the field of aerospace technology and relates to a digital twin system for low-orbit giant star constellations and a construction method thereof. The system includes a data management module that comprehensively manages the real-time data, vertical data, horizontal data, simulation data, and fusion data of the giant star constellation system; a model management module that dynamically constructs and comprehensively characterizes the geometric, physical, behavioral, rules, and constraint characteristics of complex physical systems from different dimensions, different spatial scales, and different time scales; an information management module utilizes these multi-scale and multi-level characteristic information to perform local to global network reconstruction of the low-orbit giant star constellation system and restore the complete information network; a collaborative coupling mechanism of data interconnection, information exchange, and model interoperability is adopted to achieve efficient and real-time monitoring and management, providing a solution for digital, visual, and intelligent supervision.

The invention provides a low-energy transfer window rapid design method and system for earth-moon DRO, and relates to the technical field of aerospace. Based on a double-circle restrictive four-body model, an earth-moon transfer orbit database is constructed, the earth-moon transfer orbit database comprises a plurality of transfer orbits from a target near-earth orbit to a target long-distance retrograde orbit; determining a target track point from the track distribution diagram corresponding to each track leveraging category according to the track information track distribution diagram of each transfer track and the input near-earth track information and epoch time of the target spacecraft, and correcting the track information of the transfer track corresponding to the target track point based on an ephemeris model, and obtaining a target transfer orbit of the target spacecraft. According to the method provided by the invention, rapid and accurate planning of the earth-moon DRO low-energy transfer window can be realized, the processing time is reduced, and the planning efficiency is improved.

The invention relates to the technical field of aerospace, in particular to digital intelligent airborne edge equipment, which comprises an intelligent sensing and data processing unit, an edge computing and resource optimization unit, a communication and network optimization unit, a decision support and autonomous operation unit and a system integration and test verification unit, the intelligent sensing and data processing unit is composed of a multi-mode sensor fusion module, an automatic modulation identification module, a signal optimization processing module and a real-time data preprocessing module. By integrating advanced technologies such as automatic modulation identification, signal optimization processing, neural network optimization and the like, the intelligent level of the airborne edge equipment is remarkably improved, the processing efficiency of the airborne edge equipment is improved by means of optimizing a signal processing flow, reducing data processing delay and the like, and then through an edge computing resource optimization technology, the processing efficiency of the airborne edge equipment is improved. Dynamic allocation and optimization of computing resources are realized, and the resource utilization efficiency is improved.

The invention discloses a multifunctional integrated interface for on-orbit assembly of modular space facilities, and belongs to the technical field of aerospace. The upper cover plate and the fixed base are connected through a supporting structure to form a circular cavity structure, the driving ring is in driving connection with the pin ring and can drive the pin ring to move up and down, the pin ring is used for achieving alignment positioning of a connector, and the driving ring and the fixed ring are matched through the coupling pin to drive the coupling ring to move up and down. The coupling ring is used for achieving the locking function of the interface, the lower end of the rotating base is in driving connection with the driving ring, the upper end of the rotating base penetrates through the fixing ring and is connected with the electric heating assemblies, the driving ring drives the electric heating assemblies to move up and down through the rotating base, and the electric heating assemblies are used for being matched with the interface on the opposite side to achieve data transmission and power transmission functions. The interface provided by the invention has the functions of mechanical locking, data transmission, power transmission and heat exchange at the same time, and can meet the on-orbit working requirement of a modular space structure.

The invention discloses a high-performance reaction flywheel anti-interference control method, system and device and a medium, relates to the technical field of aerospace, and solves the problems that parameters of a motor model are inaccurate, and time-varying friction torque and load torque bring difficulty to implementation of high-dynamic and high-precision control of a reaction flywheel. The control method comprises the following steps: establishing a kinematics and dynamics model of the reaction flywheel; a mathematical model for controlling the rotating speed of the reaction flywheel is established, the mathematical model considers the uncertainty of model parameters and Td, and Td represents the influence of time-varying friction torque and load torque; designing a robust disturbance observer to estimate the uncertainty of the model parameters and the Td to obtain a disturbance estimation result; and designing a rotating speed loop controller and a current loop controller according to an interference estimation result. According to the method, high-dynamic and high-precision reaction flywheel rotating speed control can be realized under the influence of inaccurate model parameters, time-varying friction torque and load torque.

The invention belongs to the technical field of aerospace, and particularly relates to a hydrogen peroxide solid-liquid rocket engine igniter and an ignition structure. The igniter comprises an igniter shell and an ignition powder column. An igniter cavity extending in the axial direction of the igniter shell is formed in the igniter shell, the igniter cavity comprises an ignition chamber and an ignition chamber which are arranged at intervals in the axial direction and communicate with each other, the igniter shell is provided with a plurality of ignition holes, and the ignition holes are located in the peripheral wall of the igniter shell and communicate with the ignition chamber. The ignition powder column is arranged in the ignition chamber and can generate a violent chemical reaction after making contact with hydrogen peroxide at the normal temperature to form fuel gas, and the fuel gas is sprayed out of the ignition chamber through the ignition holes. The igniter provided by the invention is a miniaturized igniter capable of reacting with hydrogen peroxide and provided with a high-energy solid grain capable of spontaneous combustion, and has the advantages of fast ignition response, high ignition energy, low price, light weight and the like.

The invention relates to the technical field of aerospace, and discloses a multi-engine mounting cabin section and a carrier rocket. The multi-engine installation cabin section comprises a cabin section and a force transmission mechanism, the force transmission mechanism is located in a cabin section body and connected with a front end frame and a rear end frame of the cabin section, and thrust generated by a center engine is transmitted to a force transmission ring through a main cross beam and finally transmitted to the front end frame of the cabin section through a radial beam and a first force transmission beam. Thrust generated by the peripheral engine is transmitted to the radial beam, the rear end frame and the second force transmission beam through the outer cross beam and finally transmitted to the front end frame of the cabin section, the force transmission process is uniform in force transmission, high in efficiency and good in stability, and the carrying efficiency of the carrier rocket can be improved. The invention further provides a carrier rocket which comprises a rocket body and a plurality of engine installation cabin sections installed on the rocket body, and the rocket carrying requirement is met.

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 an earth surface short-wave radiation product production method and device and a long-short-term memory network training method and device, and relates to the technical field of aerospace, and a specific implementation scheme is as follows: extracting a first short-wave radiation data set and a second short-wave radiation data set which are aligned in time from first earth surface short-wave radiation data collected by a satellite; acquiring a photovoltaic short-wave radiation data set which is acquired by a ground photovoltaic radiometer and is aligned with the first short-wave radiation data set in time; determining an albedo data set based on the satellite albedo product and the surface land cover product; obtaining an input short-wave radiation data sequence based on the first short-wave radiation data set, the second short-wave radiation data set, the albedo data set, the photovoltaic short-wave radiation data set and the ground photovoltaic radiometer coordinate vector point; inputting the input short-wave radiation data sequence into a pre-trained long short-term memory network to obtain an initial earth surface short-wave radiation product; and obtaining a target earth surface short-wave radiation product based on the initial earth surface short-wave radiation product.

The invention discloses a spatial distributed observation method based on irregular region geometric feature sampling, and belongs to the technical field of aerospace. The implementation method comprises the following steps: establishing a coordinate system for describing the motion of the spacecraft; establishing a spacecraft dynamic model under the coordinate system; six initial orbit elements of a satellite can be obtained based on existing low-orbit constellation orbit data, the spatial position is calculated through the orbit elements, and then position components are converted into latitude and longitude coordinates, namely sub-satellite point coordinates; defining vertex coordinates of the polygon, and constructing a planar polygon by connecting the vertex coordinates end to end; discretizing the internal region of the closed polygon; on the basis of the kinetic model, on the basis of the coordinates of the sub-satellite point and the center coordinates of the discrete sampling circle, calculating the characteristic distance from the sampling point in the observation area to the sub-satellite point, and on the basis of the size relationship between the characteristic distance and the radius of the observation view field, judging whether the irregular area can be successfully observed in a preset time constraint interval; therefore, spatial distributed observation is realized.

The utility model discloses space junk capturing equipment, and belongs to the technical field of spaceflight. Comprising a metal shell and a controller installed on the metal shell, the metal shell serves as a containing cavity of a mechanical arm, a data collecting unit connected with the controller is arranged on the metal shell, and a mechanical arm fixing frame is installed in the metal shell; the mechanical arm is connected with the controller through the driving module, the fixed end of the mechanical arm is connected with the mechanical arm fixing frame, and the free end of the mechanical arm is stored in the metal shell or extends out of the metal shell in a telescopic and foldable mode. And a manipulator is mounted at the head of the free end of the mechanical arm and is used for capturing space junk fragments. According to the space junk capturing device, a mechanical arm and existing intelligent control are combined, space junk can be accurately captured, efficiency is high, and the defects that traditional capturing equipment cannot be stored, is prone to being lost, is difficult to fold and unfold, and is low in accuracy are overcome.

The invention relates to the technical field of aerospace, and discloses a novel aircraft auxiliary landing stabilizing device and a control system thereof.The novel aircraft auxiliary landing stabilizing device comprises an aircraft body, an undercarriage is rotatably connected to the interior of the aircraft body, a fixing plate is arranged in the undercarriage, and a control center is fixedly connected to the outer wall of one side of the fixing plate; a fixing assembly is fixedly connected to the outer wall of the other side of the fixing plate, an adjusting assembly is fixed in the fixing assembly, an air inlet assembly is fixedly connected to the lower surface of the adjusting assembly, an airflow rotating assembly is fixedly connected to the interior of the air inlet assembly, and a transmission assembly is arranged in the fixing shell. And the outer wall of the transmission assembly is fixedly connected with a machine wheel. In the invention, the electric push rod is controlled by the control center, the rotation angle of the guide plate is finally controlled, and the airplane wheels are finally driven to rotate, so that the friction force at the landing moment can be obviously reduced, and the tire wear and overload risk can be reduced.