541 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 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 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 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 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 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 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 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.

The invention discloses a communication link automatic selection and concurrent transmission method under a helicopter limit condition, and relates to the technical field of aerospace. Comprising the following steps: firstly combing airborne communication link information, retaining four types of links including a satellite, an ultra-short wave, a short wave and a data transfer radio, and collecting parameters by means of a bus controller; filtering the data by using a Kalman filtering algorithm, and removing abnormal data through threshold verification; and finally outputting stabilized data. Determining a task scene and an environment scene based on the stabilization data in combination with airborne control screen task selection and environment sensor data; according to the technical key points, a fault precursor monitoring technology is adopted, a standby link pre-activation technology is combined, a seamless switching mechanism is matched, the defect that an existing system depends on manual switching is overcome, the automatic switching time of the link is shortened to be within 0.5 second, the communication interruption rate under the limiting condition is reduced to be lower than 5% from 30% or above, the real-time communication requirements of rescue, patrol and the like are met, and the real-time communication efficiency is improved. Good use prospects are realized.

The invention relates to the technical field of aerospace, discloses an electric propulsion ignition test anti-sputtering system and method, and solves the problems that an existing vacuum cabin anti-sputtering system is not convenient to disassemble and transport, active cooling is needed in the thruster beam bombardment process, and the thruster ignition test cost is increased. The anti-sputtering system comprises two end socket anti-sputtering targets 1 and a cabin anti-sputtering target 2 which are arranged in a vacuum cabin, the two end socket anti-sputtering targets 1 are arranged on the two sides of the cabin anti-sputtering target 2 in a gapless mode respectively, the bottom of the cabin anti-sputtering target 2 is a plane, the cabin anti-sputtering target 2 is arranged on a guide rail at the top of a butt joint trolley 3, and the butt joint trolley 3 is arranged on the guide rail. The electric thruster anti-sputtering system designed by the invention adopts a modular design, and each module is in butt joint with the vacuum cabin through the pulleys, so that the device provided by the invention is convenient to disassemble, transport and maintain, and the equipment maintenance and transportation cost is reduced.

The invention discloses a multi-spacecraft attitude and orbit coupling all-drive convergence time controller generation method, which is applied to the technical field of aerospace and comprises the following steps: establishing a multi-spacecraft attitude and orbit coupling error system all-drive model based on all-drive characteristics of a tracking error model; wherein the multi-spacecraft attitude and orbit coupling error system all-drive model is a pseudo-linear system model aiming at the stability problem of a multi-spacecraft attitude and orbit coupling error system; aiming at the multi-spacecraft attitude and orbit coupling error system all-drive model, constructing an all-drive convergence time controller based on an all-drive system method, a sliding mode control method and a preset time control method, and adjusting the convergence time of the system based on the all-drive convergence time controller; the all-drive convergence time controller is a controller designed based on a preset time parameter and an adjustable time parameter. According to the method for adjusting the preset time through attitude and orbit coupling, the multiple spacecrafts can maneuver to the expected attitude and orbit within the preset time, formation is completed, and the formation efficiency is improved.

The invention belongs to the technical field of aerospace, discloses a multi-gliding aircraft cooperative formation control method and system, and solves the problems that a multi-gliding aircraft is under-actuated in a high-speed reentry environment, a three-dimensional configuration and collision and obstacle avoidance are not considered in an existing method, and longitudinal errors are not converged. Firstly, a task scene is defined, a three-degree-of-freedom centroid dynamic model is established, and a three-dimensional configuration, collision and obstacle avoidance and control constraints are defined; then, preset time consistency control based on preset performance is designed, and lateral and height formation is achieved; respectively setting task condition correction for the same / different height scenes according to the longitudinal error, and leading a collaborative regulation strategy; setting a collision and obstacle avoidance strategy by using an improved artificial potential field method; and finally, constructing a control framework resolving instruction. According to the method, preset time error convergence, longitudinal configuration maintenance and collision avoidance safety are realized, and the aircraft clustering capability is improved.

The invention provides a satellite irregular telemetry data self-adaptive reconstruction method for continuous time modeling, belongs to the technical field of aerospace, and aims to solve the problems that the precision of nonlinear data reconstruction is reduced, the output dimension is fixed, and the reconstruction precision is low due to the fact that detailed information of data change is easy to lose in traditional satellite telemetry data reconstruction. The method comprises the following steps: calculating an abnormal threshold value of the telemetry data based on the telemetry data; constructing a satellite telemetry data adaptive reconstruction window based on the abnormal threshold, and recording a missing position and missing time; dividing the telemetry data into a training data set and a test data set according to a missing detection result, and establishing a satellite telemetry data reconstruction model based on NODE; inputting the test data set, the missing position and the missing time into the trained satellite telemetry data reconstruction model for data adaptive reconstruction to obtain normally sampled satellite telemetry data; and the calculation is repeated until the maximum number of iterations is reached, and a reconstructed data set of the satellite telemetry data is obtained.

The application discloses a spin-nutation collaborative inhibition method and system for a space tumbling target, and belongs to the technical field of aerospace. The method comprises the following steps: acquiring pose data of a space robot, and controlling the space robot to approach the space tumbling target according to the acquired pose data of the space robot; after the space robot approaches the space tumbling target, the space robot measures the space tumbling target, and acquires real-time state data of the space tumbling target by measuring the space tumbling target; a target stress position prediction model of spin-nutation collaborative inhibition is used to simulate and predict based on the acquired real-time state data of the space tumbling target, and a despinning condition required by a despinning tool in the space robot is determined; the space robot adjusts the pose according to the despinning condition, and controls the despinning tool to contact the space tumbling target, and implements contact despinning. The application can efficiently inhibit the nutation of the target in the initial stage of the despinning task, and effectively reduces the difficulty of the subsequent despinning task.

The invention discloses a coil excitation Hall thruster magnetic circuit structure suitable for a krypton working medium, relates to the technical field of aerospace technology and plasma propulsion, and solves the problems of low ionization efficiency and insufficient characteristic length of an existing Hall thruster after krypton is used. A magnetic conductive inner core is arranged at one end of a bottom plate, and a magnetic conductive cover is arranged at the other end; an inner magnetic pole is arranged at the upper end of the magnetic conductive inner core; an inner coil is arranged between the inner magnetic pole and the bottom plate and corresponds to the outer side of the magnetic conductive inner core; an outer coil is arranged between the outer magnetic pole and the bottom plate and corresponds to the inner side of the magnetic conductive cover; a ceramic channel is arranged between the inner coil and the outer coil; the bottom plate comprises an inner bottom plate and an outer bottom plate, the inner bottom plate and the outer bottom plate are coaxially arranged, and a bottom plate gap is formed between the inner bottom plate and the outer bottom plate. The characteristic length of a magnetic field in a discharge channel can be remarkably prolonged, the krypton ionization efficiency is improved, and meanwhile the structure of the thruster is simplified.

The present application belongs to the field of aerospace technology, and discloses a kind of multi-glider aircraft cooperative formation control method and system, solve the problem of underactuated of multi-glider aircraft in high-speed reentry environment, the problem that three-dimensional configuration and collision avoidance and obstacle avoidance are not considered in existing method and longitudinal error does not converge.Firstly, the task scenario is defined, a three-degree-of-freedom mass center dynamics model is established, and three-dimensional configuration, collision avoidance and obstacle avoidance and control constraints are defined;Then, a preset time consistency control based on preset performance is designed to realize lateral and height formation;For longitudinal error, task condition correction and leader-follower cooperative control strategy are set for same / different height scenarios respectively;Improved artificial potential field method is used to set collision avoidance and obstacle avoidance strategy;Finally, a control framework is constructed to calculate the instruction.The present application realizes preset time error convergence, longitudinal configuration maintenance and collision avoidance safety, and improves the cluster capability of aircraft.

The invention discloses a construction method and system for a low-thrust transfer orbit family between earth-moon translation point orbits, and belongs to the technical field of aerospace. The method comprises the following steps: firstly, solving a time optimal transfer orbit based on a continuous low-thrust spacecraft motion equation; then taking the orbit as a seed, performing forward and reverse phase continuation along a departure orbit and an arrival orbit by adopting a self-adaptive step length numerical continuation method, and generating a time optimal transfer orbit family covering different phases; secondly, mapping the time optimal orbit family into an energy optimal orbit family based on control law consistency, and expanding the transfer time range of the energy optimal orbit family through time continuation; and finally, gradually reducing homotopy parameters by using a homotopy method, detecting and processing thrust switching discontinuity in combination with a switching function, and converting the energy optimal orbit family into a fuel optimal transfer orbit family. According to the method, a continuous and complete orbit family is efficiently generated from a single solution, and systematicness and engineering applicability of orbit design are improved.

The application relates to the field of aerospace technology and discloses a cabin section structure suitable for a recoverable carrier rocket, which comprises a cabin section structure main body, the cabin section structure main body is mainly composed of a thin wall, a high rib and a grid, and the thin wall and the high rib are integrated through the spinning, milling and turning modes to process an aluminum alloy cylinder ingot, so that the design time of the rivet structure is reduced, and the design efficiency is improved. A large number of part assembly designs are needed for a traditional semi-hard shell structure, including rivet design. The rivet design is related to hole making requirements, positioning size and rivet specifications and the like. The design work usually accounts for 30-40% of the workload of a designer. However, the unit wall plate structure of the application cancels the rivet structure, realizes rivet avoidance design, and greatly improves the design efficiency of the structure product.

The present invention discloses a pre-separation shell auxiliary locking mechanism, which belongs to the field of aerospace technology, and includes a lock tongue, an articulated arm, a fixed arm, a spring, a pin, and a base; wherein the articulated arm, the fixed arm, the spring, and the pin are symmetrically arranged in two groups on the base, one end of the articulated arm is hingedly connected to the base through the pin, and a spring is arranged between the other end and the fixed arm, and the fixed arm is fixedly connected to the base. The lock tongue and the articulated arm are provided with corresponding small-angle wedge surfaces and large-angle wedge surfaces. After the shell is installed, the lock tongue contacts the articulated arm through the large-angle wedge surface, and the spring can only be unlocked by using a sufficiently large driving force to compress the spring. The device of the present invention solves the problem that the pre-separation shell is deformed due to stress release after production due to poor rigidity and the shell is deformed due to stress during installation, resulting in the shell shape being out of tolerance and the sealing not meeting the requirements. At the same time, the structure has the advantages of reliable locking, simple structure, small installation space, low cost and good maintainability.

This invention provides a spacecraft escort control method based on zero-sum differential games, belonging to the field of aerospace technology. The control method includes the following steps: obtaining spacecraft escort problem parameters for both the pursuing and defending satellites; establishing a spacecraft escort model based on zero-sum differential games and inputting initial parameters into the model; transforming the spacecraft escort model into a two-point boundary value problem model; using a genetic algorithm to solve the two-point boundary value problem model to obtain the state trajectory within the game time, thus completing the spacecraft escort control. This spacecraft escort control method can effectively solve the spacecraft escort problem under continuous thrust, and is an effective extension of existing continuous thrust escort problem models and methods. The model employs a differential game method described by a zero-sum cost function, thus simplifying the solution of the difficult-to-solve weighted performance index escort problem and improving the solution efficiency of the spacecraft escort game problem.

The utility model discloses a suspension simulator, which belongs to the technical field of aerospace and comprises a shell, a front joint and a rear joint are arranged at the top of the shell, the shell is connected with an aircraft through the front joint and the rear joint, and a first accommodating bin and a second accommodating bin for accommodating internal equipment are sequentially arranged on the shell. A first cover plate matched with the first containing bin is arranged at the top of the first containing bin, and a second cover plate matched with the second containing bin is arranged at the top of the second containing bin. On the basis that the light weight is met, the structural strength is guaranteed, the whole body is formed through assembly, the size accuracy is considered, and a tester can be supported to complete training tasks such as disassembly, maintenance and testing of a hanging object.

The present application relates to the field of aerospace technology, in particular to a kind of spacecraft autonomous control system, including autonomous navigation and return module, autonomous docking module and close-range autonomous rendezvous docking module, by setting autonomous navigation and return module, autonomous docking module and close-range autonomous rendezvous docking module, to realize the full-automatic intelligent navigation of spacecraft in the process of space flight and the autonomous docking of spacecraft and space station, further not need manual intervention, by the autonomous control of spacecraft in the process of space flight, to greatly shorten the working time of astronauts in the small space of spacecraft, improve the work efficiency of astronauts, further solve the technical problems that the control system of existing spacecraft cannot realize automatic control, need manual intervention, to increase the working time of astronauts in the small space of spacecraft, reduce the work efficiency of astronauts.

The application discloses a kind of space target ultra-short arc initial orbit intelligent determination method, belong to aerospace technology field.The application is based on artificial intelligence technology to realize space target ultra-short arc initial orbit determination, there is no strict requirement to complex nonlinear problem, and can develop more complex perturbation condition and initial orbit determination under the operation law of multiple body, coverage is wider;The application generates multiple groups of target orbits randomly and calculates the observation and state quantity of each visible arc segment, constructs the sample data set with ultra-large capacity, which is helpful for neural network to learn the nonlinear relationship between observation and state quantity using rich samples;The application builds and trains long short-term memory network model, determines the optimal network model weight parameter by designing and adjusting hyperparameter, can obtain the fast prediction solution of unknown space target ultra-short arc initial orbit according to observation;The application can meet more space target ultra-short arc initial orbit determination tasks by considering different inputs and outputs.