173 results about "Spaceflight" patented technology

This invention discloses a method for measuring the integrity of a spaceborne embedded system based on link symbol topology. In the static injection phase, a linker script is configured to inject symbols at critical memory segment boundaries, generating hierarchical hybrid fingerprints and redundantly persisting them to non-volatile storage. In the dynamic verification phase, a lightweight program is started before the operating system takes over, performing hierarchical dual-dimensional and SEU-resistant temporal redundancy verification, supporting multi-container dedicated isolation verification. In the linkage blocking phase, anomalies are forcibly blocked and primary / backup linkage is triggered, collecting eight types of fault contexts and supporting on-orbit updates. This invention solves the problems of traditional hash methods failing to detect layout anomalies and high time consumption, reducing complexity to O(K), achieving sub-millisecond-level measurement decoupled from image volume, adapting to spaceborne radiation-resistant, cloud-native, and redundancy scenarios, and meeting the high reliability requirements of aerospace.

The application provides a kind of aerospace vehicle assembly workshop intelligent dynamic scheduling method and system, method includes steps S1: by real-time connection assembly workshop central control system to obtain real-time multi-source data;Step S2: training multi-evaluation network coordination safety near end strategy optimization model, optimal control sequence in limited time domain is solved at each sampling time, only the first control action is adopted, the next sampling time is optimized based on last control action and state, after completing all control action optimization, reward model is based on neural network back update;Repeat the above process until convergence;Step S3: the multi-evaluation network coordination safety near end strategy optimization model trained is deployed to carry out workshop state update and rescheduling decision, generate and issue new scheduling plan.The application realizes the real-time response to the abnormal disturbance such as response emergency order, equipment failure, realizes the safe, efficient, flexible dynamic self-adapting scheduling of aerospace vehicle assembly workshop.

This invention belongs to the field of aerospace composite material manufacturing technology, specifically relating to a method and equipment for forming thermoplastic composite strain energy rods. First, prepreg is laid up to form a preform, which is then pressed by a flatbed pre-pressing device to form a preform sheet intermediate. Subsequently, the preform intermediate is fed into a hot-pressing module with a central heating and two-end cooling gradient temperature field. It undergoes melting and molding in a high-temperature zone and controlled cooling in a low-temperature zone to achieve final shaping. Finally, the formed rod is cooled and shaped, then pulled out at a uniform speed by a traction device and coiled and collected by a winding device, achieving continuous production. The hot-pressing module of the supporting hot-pressing equipment is equipped with a zoned heating and integrated cooling system, which can precisely construct a gradient temperature field and control the cooling rate. This invention, by using gradient temperature control to precisely regulate the material's crystallization behavior and release internal stress, effectively ensures the straightness, dimensional stability, and consistent deployment performance of the strain energy rod, perfectly meeting the manufacturing requirements of strain energy rods for large deployable structures in spacecraft.

The application discloses a remote control torch ignition system based on wireless instruction and fault self-diagnosis function and relates to the technical field of space propulsion. The application comprises a remote control center and a field control unit. The remote control center is used for generating an encrypted ignition instruction package containing an ignition instruction sequence and adjustable parameters and sending the package through a wireless data link, simultaneously receiving and displaying control feedback and diagnosis data from the field. The field control unit is arranged on an engine test stand and is electrically connected with a torch igniter and its matched fuel valve, oxidant valve and ignition power supply, used for receiving, decrypting and executing the encrypted ignition instruction package to control the corresponding valves and ignition power supply to act in sequence. The system further comprises a multi-dimensional data acquisition module used for synchronously acquiring multi-channel physical quantity time sequence signals reflecting the working state of the igniter during the ignition process. Through embedded multi-dimensional time sequence fusion diagnosis, the application can quantitatively evaluate the performance gradual state of the torch igniter in real time.

1. The name of the design product: Launch Vehicle (Chirei 6). 2. The use of the design product: used for space transportation, sending the payload such as artificial earth satellites, manned spacecraft, space stations or space probes into the predetermined space orbit. 3. The design points of the design product: the combination of shape, pattern and color. 4. The picture or photo that best shows the design points: perspective view. 5. The design claimed contains color.

The application relates to a modular magnetic reconfiguration parallel robot joint system for on-orbit autonomous assembly of a spacecraft, and belongs to the technical field of space equipment. The system is aimed at solving the problems of poor flexibility, low reconfiguration efficiency, difficulty in balancing high precision and high stiffness, and low reliability of existing space parallel robots. The technical scheme comprises a Halbach array magnetic ball and an electromagnetic locking ring of a magnetic joint module, a carbon fiber telescopic connecting rod micro servo motor and a harmonic reducer of a distributed drive branch chain, a multi-sensor data fusion unit configuration optimization algorithm module and a real-time control unit of an intelligent control system. The system can realize rapid joint mode switching, improve task response and execution efficiency, autonomously switch between precision mode and high stiffness mode to meet different operation scene requirements, adopt non-contact magnetic transmission to avoid friction and wear and pollution, improve system service life and reliability, and realize modular design for convenient maintenance and upgrading.

A space-based repeatable tolerance-tolerant docking ground simulation test device is disclosed. This invention addresses the limitations of existing ground simulation technologies in accurately simulating docking errors, faithfully reproducing microgravity environments, and simulating tunnel conditions. The invention includes a ball-joint servo-oscillating frame. A six-degree-of-freedom attitude adjustment component is located at the bottom of the ball-joint servo-oscillating frame. A standardized interface prototype is mounted on the six-degree-of-freedom attitude adjustment component, and a servo component is mounted on the standardized interface prototype. One end of a steel wire rope is connected to the servo component, and the other end of the steel wire rope passes through a pulley system fixed to the ball-joint servo-oscillating frame and is connected to a counterweight. This invention belongs to the field of aerospace engineering ground simulation test technology.

The utility model discloses a detachable aircraft fuselage keel joint belongs to air space aerospace spare part design technical field, include: upper connecting plate, lower connecting plate, positioning pin subassembly and hydraulic locking mechanism, the upper connecting plate is equipped with the taper positioning hole and the wedge guide groove, the lower connecting plate is equipped with the force sleeve of coaxial with taper positioning hole and the dovetail boss of with guide groove matching, positioning pin subassembly is by taper's head pin shaft and self -lubricating bushing constitutes, the head pin shaft is through positioning hole with force sleeve, hydraulic locking mechanism contains high pressure oil cavity, piston push and locking bolt, the piston push is driven after oil pressure and presses the side of dovetail boss, locking bolt is locked after the piston push. The utility model's keel joint has the advantages such as high positioning accuracy, locking reliable, convenient to dismantle.

The embodiment of the application provides a satellite structure, and relates to the technical field of spaceflight equipment. The satellite structure comprises: a main frame; a cabin plate; and a connecting assembly arranged between the main frame and the cabin plate. The cabin plate is movably arranged on the main frame through the connecting assembly and is fixed to any position after movement through the connecting assembly. The satellite structure can improve the test efficiency of electronic equipment during ground heat dissipation test.

The application discloses a non-cooperative target satellite rendezvous orbit design method based on active disturbance rejection control and belongs to the technical field of space orbit control. In order to solve the problem of modeling and simulation of relative motion between a mobile satellite and a target satellite under the condition that the parameters of the target satellite are unknown, the application calculates initial position and velocity components in the first geodetic coordinate system of the earth center according to prior orbit parameters of the mobile satellite and the target satellite; a dynamic and kinematic model of the mobile satellite and the target satellite is established; a state equation of relative motion between the mobile satellite and the target satellite is constructed by taking relative motion position, velocity and disturbance acceleration as independent variables; an extended state observer is adopted to estimate state variables; a mobile satellite orbit motion control model is established based on a proportional and differential control framework; and disturbance acceleration is compensated by adopting active disturbance rejection design to realize the rendezvous orbit design of the non-cooperative target satellite. The application does not need to accurately obtain orbit motion characteristics of the target satellite, and the method has strong adaptability.

The application discloses a kind of non-cooperative target centroid intelligent positioning method and system fusing multi-dimensional physical constraint.The method is directed to sparse, noisy point cloud data, and constructs a comprehensive evaluation function containing four-dimensional information of geometry, orbital dynamics, time sequence continuity and surface physical properties;Using a two-stage solution framework of surrogate model and hybrid optimization, first, a lightweight neural network surrogate model is used with a differential evolution algorithm for global coarse search, and then switching to a high-fidelity orbit model combined with an adaptive particle swarm optimization algorithm for local fine search;Finally, the optimal state estimation and its covariance matrix, evaluation decomposition and other decision support information are output.The application improves the accuracy, robustness and computational efficiency of centroid positioning, and enhances the interpretability of the results, suitable for on-orbit servicing, space debris removal and other high-risk space missions.

The application discloses a data synchronization system, a data synchronization method and a software synchronization deployment method, and belongs to the technical field of aerospace data synchronization, wherein the data synchronization system comprises a ground test, launch and control system, an on-rocket system and a satellite communication system; the ground test, launch and control system is arranged in a ground control center and is used for rocket testing, launch process control and remote measurement data processing, storage and display; the on-rocket system is used for realizing the control of the rocket by completing the transmission and reception of test, launch and control instructions, navigation guidance and attitude control; the satellite communication system is used for providing communication between the ground test, launch and control system and the on-rocket system; the satellite communication system comprises a satellite and a ground station with a ground gateway; and the ground station is used for packing and unpacking received and transmitted data. The application can realize the rapid deployment of on-rocket software in a satellite wireless communication mode, and the data transmission is not affected by signal interference caused by bad sites and weather.

The application provides a station task scheduling file generation method and device, and relates to the field of spaceflight measurement and control. The method comprises the following steps: generating corresponding station task setting events and task demand parameters according to a spacecraft instruction; sequentially performing inter-tracking arc segment merging and intra-tracking arc segment merging on the station task setting events to obtain merged events; and generating a station task scheduling file according to the spacecraft instruction, the merged events and corresponding task demand parameters. The application can automatically derive station task setting events based on a spacecraft instruction, and match and output the station task setting events to a station task scheduling file, so as to realize automatic output of measurement and control station task demand parameters.

This invention relates to the field of aerospace propulsion technology, and in particular to an oxygen-resistant Hall thruster electric propulsion system, comprising a Hall thruster, a radio frequency neutralizer, a power processing unit, a working propellant storage and supply system, and an integrated electronic control system. This invention establishes an electrical connection between the Hall thruster and the radio frequency neutralizer, directly connecting the collecting electrode to the negative terminal of a high-voltage power supply, eliminating the holding electrode and its independent power supply branch. Simultaneously, it combines a dual-path working propellant supply structure with oxidation-resistant conductive materials for the anode and collecting electrode. This allows the radio frequency neutralizer to rapidly establish a plasma electron source without the need for traditional hot cathode preheating. After the Hall thruster ejects the ion beam, electrons are automatically extracted based on the potential gradient. This not only reduces the complexity of the system's power supply structure and improves the discharge stability and electron compensation capability in low-purity working propellants and oxygen-containing environments, but also enhances the system's oxygen-resistant lifetime and rapid start-up performance in low Earth orbit oxygen-rich environments.

The application discloses a kind of W-VO2-based middle far infrared spaceflight thermal control multilayer film and its preparation and application method.The film includes aluminium substrate, inorganic medium layer, W-VO2 functional layer and protective layer in turn from bottom to top.The medium layer is used to constitute interference cavity;W-VO2 functional layer is deposited using magnetron sputtering method and is annealed, realizes phase change crystallization and the dynamic control of infrared emissivity;Protective layer is used to protect functional layer from environmental oxidation and mechanical damage, while the influence on infrared performance is limited.By adjusting the proportion of W element doping and the thickness of each layer, the phase transition temperature can be adjusted in the range of 15~68℃, and the emissivity modulation rate of 2~25μm wave band is 0.5-0.7, the visible reflectivity is 0.55-0.65.No obvious oxidation occurs after being placed in air for two weeks.The film has simple structure, controllable preparation process and stable performance, and is suitable for spacecraft cabin, battery panel, optical load and other aspects.

The application discloses a satellite remote control instruction unique identification generation method and system, and belongs to the field of space flight vehicle measurement and control communication.The method comprises the following steps: configuring the same seed value in advance at the satellite end and the ground end; when generating any remote control instruction, the ground end generates a unique identification according to the seed value and the serial number of the remote control instruction by using a specified deterministic UUID generation algorithm; after receiving the remote control instruction and analyzing the remote control sequence counter, the satellite end generates a unique identification according to the same seed value and the serial number of the remote control instruction by using the same deterministic UUID generation algorithm as the ground end; and the unique identification between the ground end and the satellite end is used to realize cross-system full-process association.The application does not increase any link bandwidth cost, can complete SHA / MD5 / HMAC calculation by occupying a small amount of computing resources, and meets the requirements of high reliability and high consistency in spaceflight.

The utility model discloses a temperature impact and hot vacuum test integrated equipment, including the vacuum chamber and sealing cover plate of being equipped with vacuum channel, wherein, the vacuum chamber is equipped with respectively and is sealed interval arrangement heating sub -chamber and refrigeration sub -chamber, the positioning transplant module of measured workpiece is still equipped with in the vacuum chamber, is used for installing positioning measured workpiece, and realizes the displacement switching of measured workpiece between heating sub -chamber and refrigeration sub -chamber simultaneously, the utility model discloses can provide hot vacuum test condition and temperature impact test condition simultaneously, has greatly reduced the period of measured workpiece to carry out temperature impact test and hot vacuum test, and has reduced the installation positioning error of measured workpiece in the test installation process, has effectively protected measured workpiece, is very suitable for being used in the temperature impact test and hot vacuum test in spaceflight material and spare part (for example space photovoltaic module) reliability test.

This paper presents a sparse grid polynomial-based orbital game situation analysis method, belonging to the field of aerospace orbital adversarial and intelligent decision analysis technology. Addressing the problems of lack of systematic representation of multi-source uncertainties, low computational efficiency of traditional quantification methods, and difficulty in achieving multi-objective equilibrium optimization in existing high-orbit multi-to-multi dynamic adversarial scenarios, this paper constructs a POMDP multi-to-multi dynamic game model incorporating uncertainties in high-orbit dynamics and control execution, strategy behavior, and communication delay. A surrogate model is established using sparse grid sampling and polynomial chaotic expansion, and global sensitivity analysis is performed. Based on this, a multi-objective optimization model is constructed, and the Pareto optimal policy set is solved. This achieves game strategy generation that balances performance, robustness, and economy, and is applicable to spacecraft strategy optimization and space situation analysis in high-orbit multi-to-multi dynamic adversarial missions.

This application relates to the field of orbital dynamics design and simulation technology, and provides a method and apparatus for free transformation between spatial coordinate systems for solar system spacecraft design and simulation. This method transforms the position and velocity of different coordinate systems to the celestial coordinate system for origin transformation, and then rotates to the target coordinate system. This allows position and velocity to be arbitrarily transformed between coordinate systems included in the coordinate transformation model library. Furthermore, when expanding to a new coordinate system, only the transformation function between the new coordinate system and the co-centered celestial coordinate system is needed, without having to calculate the transformation functions between the new coordinate system and all other coordinate systems separately. This achieves effective organization of scattered coordinate system transformation relationships, facilitating the design and simulation of spacecraft missions and reducing the workload generated by transformations between multiple spatial coordinate systems.

This invention provides a streaming compression upgrade method for onboard equipment firmware for commercial aerospace, relating to the field of spacecraft on-orbit maintenance technology. The ground-based system uses the DEFLATE algorithm to compress the firmware, extracting an identifier header including the magic number and original size, and segmenting the compressed stream into independent data packets according to link capacity. The onboard system receives data using a pre-set input circular buffer and a circular history buffer, driving an interruptible and recoverable decompression state machine for streaming decoding. Decoded bytes are synchronously written to non-volatile memory and the circular history buffer. When input is exhausted, the entire running state is saved and paused; it is precisely restored upon the arrival of new data. The advantages of this invention are: total RAM usage is less than 40KB, enabling simultaneous reception, decompression, and writing, significantly shortening narrowband link transmission time, and effectively solving the upgrade challenges of limited memory, short transit windows, and easy link interruptions for commercial low-Earth orbit small satellites.

This invention relates to a method and system for optimizing coating trajectory to ensure uniformity of spacecraft coating thickness. The method includes: acquiring a model of the surface to be coated, generating a regular grid initial coating curve and setting constraint parameters; mapping the curve to a complex surface using a pseudo-projection algorithm, and generating a fully covered, non-intersecting initial trajectory after trajectory correction; based on an experimentally calibrated coating deposition model, constructing a multi-objective optimization function that minimizes coating thickness deviation while considering trajectory smoothness and operation time, using trajectory waypoint coordinates, coating gun attitude quaternions, and coating segment duration as optimization variables, and applying constraints such as coating distance, velocity, and attitude; using an interior-point solver for iterative optimization, limiting variable iteration deviations to ensure model accuracy; and outputting an directly executable optimized trajectory. This invention can automatically adapt to the coating requirements of complex surfaces, significantly improve coating thickness uniformity, control deviations within the allowable range of aerospace processes, and effectively improve operational accuracy and efficiency.

The application relates to the field of rubber, in particular to a super-wide-temperature-range rubber for spaceflight and a preparation method thereof. The raw material composition of the rubber is as follows in terms of weight parts: 80-140 parts of natural rubber, 9-17 parts of a plasticizer, 2-4 parts of bis-[gamma-(triethoxysilane) propyl] tetrasulfide, 1-3 parts of an antioxidant, 30-60 parts of carbon black, 25-30 parts of modified white carbon black, 3-10 parts of sulfur and 1-4 parts of zinc oxide. The rubber prepared by the application has high tensile strength and tear strength, low change rate of the tensile strength and tear strength after high-temperature treatment, can effectively pass the low-temperature brittleness test, and has low high-temperature and low-temperature compression permanent deformation, and the effect is excellent.

The present application relates to the technical field of aerospace artificial intelligence, and provides a space debris collision risk probability prediction method based on deep learning, which takes ESA Kelvins real rendezvous data as a benchmark, and uses Kessler to generate a synthetic CDM sequence to expand the training sample to improve the coverage of high risk and complex situations; at the same time, a plurality of update records and their associated related events are expressed as a dynamic heterogeneous graph, a dynamic graph neural network is used to learn the law of risk evolution over time and the similarity between events can be optionally used for knowledge transfer, so that the continuous regression prediction result of the final collision risk probability of the rendezvous event can still be stably output under the condition of missing and noise, and the prediction is quickly refreshed when new CDM arrives through the local graph incremental update mechanism, which significantly improves the accuracy, robustness and online availability of risk prediction.

The present application relates to the field of space folding and unfolding mechanism, and in particular to a solar wing synchronous unfolding structure. The solar wing synchronous unfolding structure comprises a first synchronous wheel, a second synchronous wheel, a first hinge, a second hinge and a connecting rope. The first hinge is used to connect an inner panel and a spacecraft body, and the second hinge is used to connect the inner panel and an outer panel. The first synchronous wheel is coaxially arranged with the first hinge and can drive the inner panel to rotate around the first hinge. The second synchronous wheel is coaxially arranged with the second hinge and can drive the outer panel to rotate around the second hinge. The first synchronous wheel and the second synchronous wheel are connected by the connecting rope. The overall structure of the present application is simple in form, light in weight and reliable in unfolding. Compared with the traditional synchronous unfolding mechanism, the synchronous unfolding mechanism of the present application can be placed on the inner side of the panel or on the outer side of the panel, and is more flexible to use.

This invention discloses a reinforcement learning-based object-scenario data-real fusion evaluation method for aerospace equipment, belonging to the field of aerospace equipment testing and evaluation technology. The method includes: constructing a joint evolution model containing a dynamic model of the aerospace equipment and a parameterized environmental model; defining coupling operators and a joint state space; using physical measurement data, calculating deviations and updating coupling operator parameters through a residual network; instantiating dual agents for alternating iterative training, where the scenario generation agent aims to minimize the object's safety margin; integrating the control law under test, dynamically adjusting the scenario parameter search range based on the task success rate; calculating the dynamic safety margin, and outputting a quantitative report including extreme scenario parameter distribution and safety margin statistics. This invention solves the problems of traditional testing scenarios relying on manual intervention, poor data-real consistency, and subjective evaluation, achieving full coverage of extreme operating conditions in testing scenarios and quantitative evaluation of safety boundaries.

The application belongs to the technical field of aerospace electronic integrated circuits, and particularly relates to a satellite load FPGA configuration data error correction method and system resisting single event upsets, a double-layer error correction method of "bit-symbol" based on RS decoding operator error correction of a finite field algebra on the basis of a "two out of three" majority voting logic, which can correct single particle independent upsets and multiple copy common mode failures at the same time, and solve the traditional TMR error locking problem. Meanwhile, when configuration data is written into a non-volatile memory, a cross-sector spatial staggered dispersion storage mechanism is introduced, so that the symbols in the same RS code word are separated as much as possible in the physical position, so as to convert the aggregation physical damage into dispersed symbol level errors.