844 results about "Aerospace technology" patented technology

The invention discloses a comprehensive performance prediction and structural parameter optimization method for a hemispherical resonator, and belongs to the technical field of aerospace inertia. According to the method, a high-precision finite element simulation model of the ideal hemispherical harmonic oscillator is established based on a multi-region grid uniform division method; aiming at the condition that each structure parameter corresponds to different structure sizes, calculating through a finite element model of a control variable to obtain a data set of different performance parameters of the hemispherical resonator; based on a BP neural network algorithm optimized by a particle swarm optimization algorithm, after the data set is divided into a training set and a test set, a comprehensive performance parameter prediction model of the hemispherical harmonic oscillator is established; and taking the PSO-BP neural network model as the input of a second-generation non-dominated sorting genetic algorithm, and performing multi-objective optimization on the structural parameters of the hemispherical harmonic oscillator by taking higher thermoelastic damping and larger minimum frequency difference between the interference modal and the PSO-BP neural network model as optimization objectives. The use performance of the hemispherical resonator is improved, and the performance requirement of the aerospace technology field on the high precision of the hemispherical resonator gyroscope is met.

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 relates to the technical field of aerospace, in particular to a heat insulation coating failure detection system. Comprising a spray gun, a heat insulation plate, a heat dissipation device, piezoelectric ceramics and a signal receiving and processing device, the heat insulation plate is vertically arranged on the surface, plated with a heat insulation coating, of the to-be-detected plate body and divides the to-be-detected plate body into a first part with a large area and a second part with a small area, the spray gun is used for spraying flames to the first part, the heat dissipation device comprises a heat dissipation bottom plate, and the heat dissipation bottom plate is fixedly arranged on the edge, away from the heat insulation plate, of the second part. The heat dissipation bottom plate is provided with a groove, the piezoelectric ceramic is fixedly arranged at the bottom of the groove, the piezoelectric ceramic is connected with the heat insulation coating through the heat dissipation bottom plate so as to receive sound signals generated by internal damage of the heat insulation coating, and the signal receiving and processing device is electrically connected with the piezoelectric ceramic and used for receiving and processing electric signals generated by the piezoelectric ceramic. The embodiment of the invention provides a thermal insulation coating failure detection system which can detect the failure process of a thermal insulation coating.

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 relates to the technical field of aerospace, in particular to a fixed-point vibration absorption method based on distributed magnetic suspension flywheel axial momentum exchange. The method comprises the following steps: establishing a magnetic suspension rotor dynamical model comprising axial translation and axial rotation of a rotor; determining a magnetic suspension flywheel group control quantity model according to the magnetic suspension rotor dynamic model; determining a star active control force and a star active control moment according to the magnetic suspension flywheel group control quantity model and the star dynamic model; under the constraint of the star active control force and the star active control torque, determining a force / torque distribution matrix of the magnetic suspension flywheel group; according to the force / torque distribution array of the magnetic suspension flywheel group, determining the axial vibration force and the torque action quantity of each flywheel; and on the basis of the magnetic suspension rotor dynamic model, according to the axial vibration force and the torque action quantity of each flywheel, determining a driving current to suppress vibration. According to the scheme, load precision stability and ultra-static performance can be achieved.

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 is applied to the technical field of aerospace, and particularly discloses a satellite attitude and orbit integrated control device for step-length control thrusters, the satellite attitude and orbit integrated control device comprises a body, the mass center coordinate system of the body is 0-XYZ, a third thruster and a fourth thruster are rotatably mounted on the XOZ face of the body, a rotating meshing mechanism is arranged on the outer surface of the body, and the third thruster and the fourth thruster are meshed with each other through the rotating meshing mechanism. The posture of the body can be accurately adjusted by rotating the meshing mechanism. According to the satellite attitude and orbit integrated control device for controlling the thruster through the step length, when the body needs to be subjected to orbit adjustment, a nozzle in a certain axial direction is adjusted, so that the thrust direction of the nozzle is perpendicular to the axial direction of the advancing direction, the speed is provided, and orbit transfer is conducted according to the Kepler law; when the satellite needs to be subjected to attitude adjustment around a certain axis, the nozzles are adjusted to be parallel to the bottom surface, the two engines on the same plane are started to generate torque, the satellite is made to roll to adjust the attitude, and when the body needs attitude and orbit control, the engine step length can be adjusted to control thrust working time adjustment.

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 embodiment of the invention relates to the technical field of aerospace, in particular to an intelligent track prediction method based on Transform-LSTM (Long Short Term Memory), and the method comprises the steps: obtaining the information of a multi-sensor track intelligent processing system, and carrying out the preprocessing, association and fusion, and obtaining the original track data; wherein the original track data comprises original latitude and longitude data and original height data; converting the original track data into a unified Cartesian coordinate system, and performing normalization processing on the track data after space coordinate conversion and the original height data by using maximum and minimum normalization functions; inputting the normalized track data and the normalized height data into a Transform network and an LSTM (Long Short Term Memory) network of a pre-constructed track intelligent prediction model respectively, and extracting and learning spatial position features and time sequence features; and finally, an output network of the intelligent track prediction model outputs a track prediction result based on the spatial position features and the time sequence features, so that more accurate track prediction is realized.

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 present disclosure provides a method, device, equipment and medium for calculating right ascension and declination based on Euler angles, which relates to the field of aerospace technology. The method includes: obtaining the initial observation pointing parameters of a satellite camera, the first rotation Euler angle of the coordinate system unit basis vector corresponding to the initial observation pointing to the body coordinate system unit basis vector corresponding to the satellite platform, and the second rotation Euler angle of the coordinate system unit basis vector corresponding to the satellite platform to the coordinate system unit basis vector corresponding to the final observation pointing; calculating the final observation pointing parameters of the satellite camera according to the initial observation pointing parameters, the first rotation Euler angle and the second rotation Euler angle; calculating the right ascension and declination of the satellite camera according to the final observation pointing parameters. The method can quickly and accurately solve the celestial right ascension and declination from the initial pointing of the camera, and can fully consider various rotation sequences of aerospace engineering, with a wide range of applications and strong applicability.

The invention discloses a conical surface dimensional space deformable honeycomb mechanism, and belongs to the technical field of aerospace. The conical surface dimensional space deformable honeycomb mechanism comprises an annular space honeycomb mechanism and a driving system, and the annular space honeycomb mechanism is deformed along a conical surface with a fixed taper by controlling the driving system. The annular space honeycomb mechanism comprises a plurality of honeycomb units which are of concave hexagonal structures, each honeycomb unit comprises two beams with different lengths and four ribs with the same length and structure, the beams are connected through hinges, and the ribs are connected through hinges; the driving system comprises a main body frame, a motor, a synchronous belt wheel, a trapezoidal lead screw and a clamping jaw. The annular space honeycomb mechanism is designed, and fitting of a target conical surface can be achieved by reasonably designing the sizes of the rod pieces in the honeycomb units; meanwhile, a driving system of the honeycomb mechanism is designed, and the honeycomb mechanism can deform along the conical surface with the fixed taper by controlling the driving system.

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 discloses a maneuvering spacecraft orbit determination method for adaptive state noise compensation, and belongs to the technical field of aerospace. The implementation method comprises the following steps: establishing a general dynamical model and an angle-only observation model for unknown mutation of the random nonlinear dynamical system; predicting a system state and a covariance matrix based on the dynamic model; calculating a system measurement residual error and a covariance matrix at the tk + 1 moment, and representing the matching degree of a prediction result of the system based on the kinetic model and the maneuver-free model through an index function lambda k + 1; comparing the index function at the tk + 1 moment with a mean value of the index functions from t1 to tk, when the difference between an evaluation index and the mean value is greater than a threshold value, judging that the state of the system is abruptly changed, and performing state noise compensation on a system prediction result; and calculating orbit determination algorithm gain based on a system state noise compensation result, weighting and outputting a tk + 1 system state and a covariance estimation value based on a system prediction result # imgabs0 # noise compensation result # imgabs1 # and a measurement result zk + 1, and iteratively solving a state estimation value at each moment so as to realize high-precision orbit determination of the maneuvering spacecraft.

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 invention relates to the technical field of aerospace, and discloses a micro-nano satellite attitude control method based on deterministic strategy gradient and PID, and the method comprises the following steps: obtaining the pitching theta, the rolling psi and the yaw angle # imgabs0 # of a satellite through a sensitive sensor; the required control moment is converted into the attitude angle speed # imgabs 1 #, the angles theta t and psi t of the satellite in the current state are obtained according to an attitude kinematics model, and the satellite attitude quaternion error qe is determined according to the difference of the satellite attitude angles determined by a sensitive sensor through # imgabs 2 #. According to the micro-nano satellite attitude control method based on the deterministic strategy gradient and the PID, a better control strategy can be learned by adopting a deterministic strategy gradient algorithm, and by combining the deterministic strategy gradient algorithm with the traditional PID, more accurate adjustment can be carried out for the complex attitude change of the micro-nano satellite, and attitude deviation is reduced.

The invention provides a method and device for controlling a spacecraft to avoid multistage collision, and belongs to the technical field of aerospace. According to the method, each collision object in a plurality of collision objects of the spacecraft is used as a repulsive force source, and according to a dynamic mapping matrix of the spacecraft and each collision object at a predicted collision moment, a repulsive force acceleration generated by each collision object to the spacecraft is calculated; and taking the sum of the repulsive force accelerations of the plurality of collision objects as a control acceleration for controlling the spacecraft to avoid the target collision object. The method not only can ensure that the spacecraft avoids the target collision object, but also can ensure that the spacecraft does not collide with other collision objects.

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 relates to the technical field of aerospace, and provides a multi-stage pneumatic telescopic mechanism and a rocket recovery landing leg device.The multi-stage pneumatic telescopic mechanism comprises a plurality of barrels which are arranged in a sliding and sleeving mode from inside to outside; the outermost cylinder is connected with an end cover, and the end cover is provided with an unfolding inflation hole; in every two adjacent barrels, a vent hole is formed in the end, away from the end cover, of the outer barrel, a locking groove is formed in the inner wall of the outer barrel, and a locking component is arranged at the end, close to the end cover, of the inner barrel. The locking component is switched from the locking state to the unlocking state through the vent hole of the outermost cylinder, the inner cylinder can slide to contract, and when the two adjacent cylinders are in the contraction state, the vent holes of the two adjacent cylinders communicate with each other, so that the locking state of the locking component is changed, and the service life of the locking component is prolonged. The corresponding locking grooves are unlocked by all the locking parts by means of the vent holes of the outermost cylinder body; the extension of the multi-stage barrel body is realized by unfolding the inflation hole, so that the extension and retraction of the multi-stage barrel body can be realized through the vent hole in the outermost barrel body and the unfolding inflation hole, and the design complexity of an air path is reduced.