45 results about "Space environment" patented technology

Embodiments of the present application provide a space target orbit prediction method, a prediction device, an electronic device and a medium. The method comprises: obtaining two-line orbital elements of a space target at a current time, space environment data and perturbation coefficients, the perturbation coefficients being coefficients in a space target orbit prediction model; correcting the two-line orbital elements and / or the perturbation coefficients based on a pre-trained parameter correction model, the two-line orbital elements, the perturbation coefficients and the space environment data, to obtain corrected data, the corrected data comprising corrected two-line orbital elements and / or corrected perturbation coefficients; and predicting orbit information of the space target at a preset future time based on the space target orbit prediction model and the corrected data. The method makes the predicted orbit information more accurate.

The application provides a satellite communication power adaptive regulation method, system and device, which divides the space where satellite communication belongs into multiple space grid areas according to a preset space division scale; acquires satellite sensing data, space weather data and near-earth weather data in each space grid area to determine communication link loss evaluation data corresponding to each space grid area; generates space interference summary information according to the loss evaluation data corresponding to each space grid area and location data thereof; and adjusts the transmission power of satellite communication in each space grid area according to the space interference summary information. The method and system of the application collect weather data in the space area and the near-earth area thereof to determine loss evaluation data, and realize dynamic adaptive adjustment of satellite transmission power in each space grid area based on the loss evaluation data, thereby improving the reliability of satellite communication in a complex space environment.

The application provides a satellite in-orbit anomaly and space environment correlation method based on density distribution drift, comprising the following steps: S1: determining a typical repetitive anomaly of an in-orbit spacecraft, on the basis of which, it can be considered that the event has consistent environmental correlation characteristics, so that it can be statistically analyzed together and further analyzed; S2: using the occurrence of the abnormal event as a condition, the space environment data is divided into two parts, and the probability density distribution of the two parts is counted; S3: according to the difference between the probability density distribution of the two parts in S2, the influence of the space environment is formed to form an explainable correlation characteristic, and the correlation analysis is completed. The method is suitable for analyzing the in-orbit repetitive anomaly phenomenon, and the possible correlation characteristics between the change characteristics of the space environment and a kind of repetitive in-orbit anomaly phenomenon are analyzed by using the distribution characteristic construction method in statistics.

This invention discloses a ground-based experimental system and method for onboard autonomous maneuvering mission planning. It utilizes a ground-based telemetry, tracking, and command (TT&C) simulation module to manage TT&C resource constraints, satellite-ground status, and mission requirements based on the test scenario. An onboard computer equivalent module performs autonomous mission planning based on mission instructions generated by the ground-based TT&C simulation module, generating control commands or mission execution sequences. A spacecraft dynamics simulation module determines the satellite's orbit, attitude, and visibility based on the control commands. A space environment simulation module generates sensor simulation data based on the orbit and attitude. A navigation sensor simulation module initiates or terminates the mission based on control commands and generates measurement data based on the sensor simulation data, sending it to the onboard computer equivalent module. A payload execution simulation module simulates the payload's operation process based on the mission execution sequence. This allows for the simulation and comprehensive verification of the entire mission planning algorithm, from mission injection to execution feedback.

The application discloses a ground test system and method for on-orbit autonomous mobile task planning, which comprises a ground measurement and operation control simulation module, a satellite computer equivalent module, a spacecraft dynamics simulation module, a space environment simulation module, a navigation sensor simulation module and a load execution simulation module. The ground measurement and operation control simulation module generates a task instruction based on a test scene, a measurement and operation control resource constraint, a satellite-ground state and a task demand. The satellite computer equivalent module generates a control instruction or a task execution sequence based on the task instruction generated by the ground measurement and operation control simulation module. The spacecraft dynamics simulation module determines a satellite orbit, an attitude and visibility based on the control instruction. The space environment simulation module generates sensor simulation data based on the orbit and the attitude. The navigation sensor simulation module starts or terminates a task based on the control instruction and generates measurement data based on the sensor simulation data and sends the measurement data to the satellite computer equivalent module. The load execution simulation module simulates a load working process based on the task execution sequence. The application can simulate and verify the whole process of a task planning algorithm from task injection to execution feedback.

This invention provides a zero-gap holding linear actuator and its installation method, belonging to the field of spacecraft structures. The zero-gap holding linear actuator includes: an inner cylinder, an outer cylinder, a base, an elastic plate, and an electric telescopic rod assembly. The inner cylinder, outer cylinder, base, and elastic plate are all made of carbon fiber or Invar. The elastic plate is disposed within the base and fixed at one end to the bottom of the inner cylinder. The inner cylinder is nested within the outer cylinder, and an interference fit is achieved through the elastic plate. The electric telescopic rod assembly engages with a limiting hole at the bottom of the inner cylinder via a limiting screw, achieving anti-rotation and push-pull actuation. This actuator achieves ultra-low thermal deformation through low-expansion materials, and the electric telescopic rod does not require power for holding, solving the problems of large on-orbit thermal deformation and difficulty in achieving zero gap in existing spacecraft actuators. It has advantages such as high stability, compact structure, and strong adaptability to the space environment, and is suitable for on-orbit attitude adjustment of high-precision antennas or payloads in spacecraft.

A method and system for optimizing on-board orbit mission sequences based on real-time windward area is disclosed. The system includes a data configuration module, an area mapping module, a high-precision prediction module, and an autonomous feedback optimization module. The data configuration module receives mission sequences and space environment characteristic parameters, and synchronizes the satellite's real-time orbit vector. The area mapping module analyzes the target attitude based on the attitude parameters in the mission sequence and calculates the equivalent windward area corresponding to each mission time period in real time by combining the velocity vector. The high-precision prediction module executes the attitude-coupled variable area HPOP algorithm, incorporating the dynamically calculated equivalent windward area as a key perturbation term into the atmospheric drag calculation, and outputs orbit prediction results reflecting the attitude adjustment impact. The autonomous feedback optimization module performs consistency judgment on the difference between two adjacent recursive results. If the difference does not meet the preset requirements, it autonomously adjusts the mission execution time and mission execution parameters in the mission sequence and feeds them back to the area mapping module for iterative calculation.

The application discloses a kind of multi-sky-oriented space environment data three-dimensional coordinate adaptive conversion and rendering method and system, it is related to space environment data visualization technical field.The application establishes the scene radius and real radius of each sky by establishing sky parameter registry, proposes double radius parameter adaptive coordinate conversion formula, so that the same set of conversion logic adapts to earth, moon, Mars and other arbitrary sky.Using three-dimensional scene graph parent-child level mechanism, particle system is mounted as sky grid subnode, and data points automatically follow sky movement.The application proposes a dynamic near clipping plane adaptive algorithm based on sight distance, and solves the problem of small sky invisibility through double camera synchronous setting strategy.Independent instantiation of multi-sky particle system and on-demand switching in single frame are realized.The application solves the technical problems of coordinate conversion and sky coupling, height mapping distortion and small sky invisibility in the prior art, and is suitable for three-dimensional visualization of multi-sky space environment model results.

The application discloses an anti-radiation flexible perovskite battery and a preparation method thereof, and belongs to the technical field of space photovoltaics. The battery is based on a polycrystalline diamond / PEDOT:PSS flexible composite transparent conductive substrate, and a battery functional layer and a top electrode are further deposited. The substrate has excellent anti-radiation property, ultrahigh thermal conductivity, high light transmittance and good flexibility, and the efficiency retention rate is greater than or equal to 94% under 10 MeV proton irradiation, so that the space environment stability is significantly improved, and the battery is suitable for lightweight and high-reliability photovoltaic application scenarios such as low earth orbit, high orbit and deep space exploration.

The application relates to a small-size high-reliability automatic light switch suitable for a space environment and belongs to the field of light switches for space vehicles. The light switch is a key component of a micro-nano optical sensor for a spacecraft. The micro-nano optical sensor for the spacecraft has strict weight and size limits, and has extremely high requirements on the weight and the size of the light switch. In addition, the light switch is in a vacuum environment under the space environment, the working temperature range changes greatly, and the light switch cannot be maintained and replaced, so the reliability is also extremely high. The application provides the small-size high-reliability automatic light switch suitable for the space environment, the structure is simple, the size is small, the reliability is high, and the light switch is convenient to operate, so that the micro-nano optical sensor for the spacecraft is ensured to be used on an orbit.

The present application relates to the technical field of space physics and space environment detection, and relates to a multi-layer nested electrostatic analyzer, which comprises a shielding grid, an electrostatic deflection device, a trajectory guiding device, a hemispherical multi-layer concentric grid electrode, a microchannel plate, a position sensitive anode, an electronic processor and a device shell; the device shell is arranged on the electronic processor, the shell opening is covered by the shielding grid, and the electrostatic deflection device, the trajectory guiding device, the hemispherical multi-layer concentric grid electrode, the microchannel plate and the position sensitive anode are arranged in the shell; the present application reduces two-dimensional voltage scanning required by traditional electrostatic analysis type low-energy charged particle measurement to one-dimensional voltage scanning, can greatly save voltage scanning time, improve the time resolution of low-energy charged particle measurement, and has wide application demand in the detection of small-scale structures in space.

The application discloses a kind of network equipment temperature backup methods suitable for space environment, it is related to manned spacecraft overall design field.This method is configured in switch main backup equipment identical with software and hardware, main backup equipment shares network port connector, but with independent power conversion module, 1553B bus communication module, network data exchange module and network interface module.Working, main backup equipment simultaneously power-on start 1553B bus communication module, network data exchange module, but only start the network interface module of main device and forward data;Spacecraft information system periodically synchronizes the network configuration information of main device to backup device periodically.When main device fails, disconnect the power supply of main device and start the network interface module of backup device, and backup device is based on the network configuration of synchronization and takes over data forwarding.The application solves the problems of long recovery time of cold backup and complex hot backup system.

This application discloses a method for maintaining high vacuum under high-speed neutral particle flow field extraction, belonging to the field of low-Earth orbit space environment simulation technology. The method involves coarsely pumping the vacuum container to reduce the pressure below a first pressure value; activating the container's cryogenic pump and the transition section's cryogenic pump to reduce the pressure below a second pressure value, establishing a high-vacuum background environment; activating the gas-helium refrigeration system to cool the cryogenic cold plate, lowering its temperature below a first temperature threshold and its pressure below a third pressure value, forming high-pumping-rate vacuum maintenance conditions; allowing a neutral particle source to extract a high-speed neutral particle flow field through the transition section; and continuously evacuating the introduced gas molecules during the continuous extraction of the particle flow using the container's cryogenic pump, the transition section's cryogenic pump, and the cryogenic cold plate to maintain the pressure below a fourth pressure value. This application solves the technical problem of maintaining high vacuum in a container under high-speed neutral particle flow field extraction.

This disclosure relates to a method and apparatus for analyzing the flux of space environment indicators on the surface of a spacecraft. The method includes: constructing a three-dimensional simulation model of the spacecraft based on its basic parameter information; obtaining a first included angle based on the spacecraft's operational parameter information and the three-dimensional simulation model, wherein the first included angle is the angle between the normal vector of the spacecraft's outer surface and the direction associated with the space environment indicators; determining the space environment indicator influencing factors based on the first included angle; and determining the flux of space environment indicators on the spacecraft's outer surface based on the space environment indicator influencing factors. This disclosure can accurately and rapidly calculate the flux of space environment indicators on the outer surface of a spacecraft.

This invention belongs to the field of satellite navigation and space environment monitoring technology, and relates to a dual-frequency TEC inversion optimization method based on real-time altimetry data. Based on GNSS ground-based dual-frequency observations, it integrates ionospheric altimetry data and absolute signal deviation products to achieve high-precision conversion of STEC (Slant Total Electron Content) to VTEC (Vertical Total Electron Content). The method includes data preprocessing, TEC inversion, cycle slip detection, deviation correction, and mapping correction, and generates ROTI (Rate of TEC Index, i.e., ionospheric scintillation characterization parameter) products. By introducing a single-layer model for correcting the effective ionospheric height, the ability to characterize ionospheric disturbances is improved.

This invention relates to a ground flight control and operation support system for a very low Earth orbit (UEO) satellite constellation. It incorporates predicted space environment information, satellite attitude change information, and orbit control information into the high-precision satellite orbit prediction process, effectively ensuring higher orbit prediction accuracy for the UEO constellation and maximizing the assistance to ground control stations during satellite transits in capturing and tracking the satellite. The satellite attitude and orbit control system simulation module performs real-time simulations of the impact of attitude control and orbit control on orbital maneuvers, thereby achieving higher orbit control accuracy and providing better assurance for the precise execution of missions by the UEO constellation. Automatic scheduling of Earth remote sensing and data transmission tasks, along with ground system fault identification and handling, and collision warning, provide more intelligent support for the operation of the UEO constellation.

This invention discloses a ground-based experimental system and method for onboard autonomous maneuvering mission planning. The system utilizes a first computer to manage telemetry and control resource constraints, satellite-ground status, and mission requirements based on the test scenario. The onboard computer performs autonomous mission planning based on mission instructions generated by the ground-based telemetry, tracking, and command (TT&C) simulation module, generating control commands or mission execution sequences. A dynamics simulator determines the satellite's orbit, attitude, and visibility based on the control commands. A space environment simulation module generates sensor simulation data based on the orbit and attitude. A navigation sensor simulation module initiates or terminates the mission based on control commands and generates measurement data based on the sensor simulation data, which is then sent to the onboard computer. Finally, a payload execution simulation module simulates the payload's operational process based on the mission execution sequence. This system enables comprehensive simulation and verification of the entire mission planning algorithm, from mission injection to execution feedback.

This invention provides a hierarchical estimation-based joint autonomous orbit determination method and system for LEO constellations using GNSS / inter-satellite links. First, GNSS data and inter-satellite link data collected by each satellite are acquired. Then, these data are input into local filters to obtain the state estimates and covariance matrices of each satellite. Next, the state estimates and covariance matrices are input into a main filter to obtain the overall reference error estimate of the constellation. Finally, the overall reference error estimate is fed back to each local filter to correct the state estimates of each satellite, resulting in the final high-precision orbit parameters and clock error parameters for each satellite. This invention decomposes the global orbit determination problem into local estimation and low-dimensional reference error estimation through a hierarchical estimation architecture. Combined with an observation model based on an inter-satellite link closed loop and an adaptive weight adjustment mechanism, it effectively suppresses overall constellation reference drift, significantly reduces computational complexity, and enhances the system's robustness in complex space environments.

The application provides a shape memory material-based metamorphic satellite structure and relates to the technical field of space satellites. The metamorphic satellite structure comprises a variable-configuration regular pentagon module and a variable-configuration regular hexagon module made of shape memory material, the variable-configuration regular pentagon module is connected with the variable-configuration regular hexagon module to form a spherical polyhedral structure, and the variable-configuration regular pentagon module and the variable-configuration regular hexagon module are both used for synchronous contraction or synchronous expansion by temperature change. The structure is optimized through material-structure collaborative design, the characteristics of the multi-stable structure that can still maintain a stable geometric configuration after load removal are utilized, the configuration stability of the system is fundamentally improved, and the dependence on single material characteristics is reduced; even when the shape memory polymer and the composite material thereof cause incomplete phase change due to temperature fluctuation, the system can still maintain a target configuration, and the reliability in an extreme space environment is significantly improved.

This invention discloses a special cable for use in space environments, comprising, from the inside out: an inner conductor assembly, including multiple signal conductor cores, a first shielding and protective layer covering the periphery of the multiple signal conductor cores, and a first filler material filling the spaces between the multiple signal conductor cores; the multiple signal conductor cores are stranded into a cable, each signal conductor core including a double-strand signal conductor and a second shielding and protective layer covering the periphery of the double-strand signal conductor; an outer conductor assembly, surrounding the inner conductor assembly, including multiple outer conductor cores and a second filler material filling the spaces between the multiple outer conductor cores and the inner conductor assembly; a double-layer heterogeneous electromagnetic shielding layer covering the periphery of the outer conductor assembly; and a gradient functional thermal control sheath layer covering the periphery of the double-layer heterogeneous electromagnetic shielding layer. This invention is applicable to ultra-long-duration on-orbit missions of 10–15 years, such as deep space exploration, high-orbit satellites, and manned space stations, and has significant engineering application value and technological advancement.

The application discloses a space spacecraft AI chip-oriented target detection network anti-radiation reinforcement method and belongs to the technical field of neural network anti-radiation. The method comprises the following steps: obtaining a target detection network model; respectively adopting Dropout enhancement and exponential bit quantization optimization models; and adopting the optimized model to perform target detection in a space environment. The application can not only significantly improve the robustness of a neural network in a space radiation environment, but also maintain the reasoning accuracy and speed of the model, solves the reliable operation problem of a deep learning model in a space environment, and optimizes the stability and task success rate of a spacecraft intelligent system. The application proposes a training-reasoning collaborative optimization strategy for the single event upset problem caused by high-energy particle radiation in a space environment, and provides reliable protection for AI application on a spacecraft.

This invention relates to the field of nuclear reactor engineering technology and discloses a Stirling engine device for a space environment gas-cooled reactor, comprising: a casing defining an internal cavity; and an insulating baffle disposed within the cavity and dividing it into an inner chamber, an outer chamber, a connecting chamber, and a crankshaft chamber. The inner chamber contains a reactor heat source, a gas distribution piston, and a power piston arranged sequentially at intervals. The outer chamber contains a regenerator and a cooler arranged adjacent to each other. By integrating the reactor heat source within the Stirling engine casing, the reactor heat transfer system is eliminated, eliminating the need for heat pipes or other heat transfer circuits. Simultaneously, the rotating crankshaft drives a self-sustaining cooling device, passively operating to propel the cooling medium from the cooler into a droplet generator, where it is transformed into fine droplets. During droplet flight, the droplets radiate heat to the space environment and then return to the cooler. This simplifies the structure, reduces the overall size and weight of the engine, and makes it suitable for various applications requiring continuous and reliable power output, such as deep space exploration.

The application specifically relates to a spacecraft attitude weakening-takeover control method based on deep reinforcement learning, which comprises the following steps: in the weakening phase, an interference strategy is learned by a double-delay deep deterministic policy gradient algorithm agent to quickly consume the fuel of a target spacecraft; in the takeover phase, a takeover torque output by a service spacecraft is determined by a linear quadratic regulator; when the takeover torque output by the service spacecraft makes the attitude and angular velocity of a combination converge to a target attitude angle and a target angular velocity respectively, it is determined that the service spacecraft completes the attitude takeover of the target spacecraft; and the method has high robustness and can effectively cope with control challenges in a complex space environment by combining the adaptability of the double-delay deep deterministic policy gradient algorithm with the stability of the linear quadratic regulator.

The application discloses an electric propulsion system, an electric propulsion gas supply device and a gas supply method. The electric propulsion gas supply device comprises a control module, a low-temperature gas suction module, a refrigeration module and a gas storage module. The low-temperature gas suction module is used for sublimating gas into solid working medium after contacting with the gas in the space environment, and also converts the solid working medium into fluid working medium into the gas storage module. The refrigeration module is used for refrigerating the low-temperature gas suction module to make the low-temperature gas suction module sublimate the contacted gas into solid working medium. The control module is used for controlling the refrigeration module to stop refrigerating the low-temperature gas suction module when the solid working medium in the low-temperature gas suction module is greater than a preset value, and controlling the low-temperature gas suction module to convert the solid working medium in the low-temperature gas suction module into fluid working medium. The gas storage module is used for storing the fluid working medium. The application can absorb the gas in the low orbit, can not carry working medium or carry a small amount of working medium, can prolong the on-orbit operation time of the satellite, and can also ensure the accurate control of the operation attitude and operation track of the satellite.

This invention relates to a programmable thermal protection ceramic-metal lattice superstructure for spacecraft in extreme space environments, belonging to the field of spacecraft thermal protection technology. It comprises multiple body-centered cubic lattice cells arranged periodically in three-dimensional space to form an isotropic three-dimensional network structure. Each body-centered cubic lattice cell includes eight lattice connecting rods, which are rod-shaped structures. The eight connecting rods intersect at a central point, and any two adjacent connecting rods have the same angle relative to the central point, forming a cubic unit configuration body-centered cubic lattice cell. This invention solves the problem of traditional thermal protection structures failing to simultaneously address heat dissipation capacity and thermal shock protection, providing an innovative solution for thermal protection of spacecraft in extreme space environments.

This invention discloses a probiotic nutritional preparation for enhancing astronauts' immunity and its preparation method, belonging to the field of biomedical technology. The preparation consists of lyophilized Bifidobacterium lactis BL-Immuno97 strain powder, dextran, sodium butyrate, galactooligosaccharides, trehalose, skim milk powder, and microcrystalline cellulose. The mass ratio of β-glucan to sodium butyrate is 3:1 to 4:1, and the viable count is not less than 10 billion CFU / g. The preparation method employs pH-stat controlled fermentation, liquid nitrogen instantaneous freezing, and programmed lyophilization processes to ensure a high viable bacteria survival rate. This invention aims to solve the problems of weak immune targeting, poor activity stability, and process incompatibility of existing probiotic preparations in the space environment. This preparation can maintain a high viable bacteria count for a long time in orbit, significantly improving sIgA levels, NK cell activity, and intestinal flora diversity, effectively maintaining the systemic immune function of astronauts.

The application relates to the technical field of satellite autonomous control, and discloses a satellite round game pursuit-escape game decision method based on Alpha-Zero, a terminal and a medium. The method sets physical constants and game rules, establishes a relative orbit motion model of satellites of both sides of pursuit and escape based on a C-W equation, and thus constructs a satellite round game pursuit-escape game environment; a strategy value neural network based on an Alpha-Zero framework is constructed, and when impulse orbit change decision needs to be made in each round, Monte Carlo tree search is performed based on the satellite round game pursuit-escape game environment and the strategy value neural network; the parameters of the strategy value neural network are iteratively optimized through self-game iteration, and an optimal strategy value neural network is obtained; after the real-time collected satellite orbit state is input into the optimal network after feature coding, an impulse orbit change strategy is generated and the satellite is controlled to execute. The application realizes efficient autonomous game of the satellite in a complex space environment.

This application relates to the field of space environment simulation technology, specifically to a device and method for obtaining ultra-high vacuum in a lunar regolith water ice detection time-of-flight mass spectrometer. The device includes a vacuum tank, a pumping system, a heating furnace, a pressure stabilizing chamber, and a lunar regolith water ice detection time-of-flight mass spectrometer. The vacuum tank is a large-volume vacuum tank equipped with a sealed door. The heating furnace is connected to the pressure stabilizing chamber via a first valve. The pressure stabilizing chamber is connected to the lunar regolith water ice detection time-of-flight mass spectrometer via a small orifice. The pumping system includes a first molecular pump, a dry pump, and a second molecular pump. This application places the second molecular pump used in the lunar regolith water ice detection time-of-flight mass spectrometer within a large-volume vacuum tank, enabling the acquisition of an ultra-high vacuum sufficient for the lunar regolith water ice detection time-of-flight mass spectrometer to operate normally, allowing the internal vacuum level of the mass spectrometer to be reduced to 10⁻⁶. ‑5 Pa level.