23 results about "Space techniques" patented technology

The invention belongs to the technical field of lunar exploration space, and particularly relates to a manned lunar rover driving operation platform. The platform comprises a central processing module, a main display screen, an operation panel, a foldable force feedback control module, an auxiliary control panel, a sensing module and a voice control module. A comfortable and intelligent driving operation space is created for a lunar rover driver, intelligent perception and panoramic display of the surrounding lunar surface environment are provided, the driving experience is optimized, and the reliability and safety of driving operation are guaranteed.

The invention discloses a monitoring method and system based on GNSS distributed solution, and belongs to the technical field of space technology and satellite application, and the method comprises the steps that S1, a base station and monitoring stations receive GNSS data corresponding to a monitoring area sent by a satellite positioning system, and the base station sends the received GNSS data and coordinate information of the base station to a plurality of monitoring stations; s2, the plurality of monitoring stations calculate three-dimensional displacement data according to the GNSS data from the satellite positioning system and the GNSS data and the coordinate information from the base station, and send the three-dimensional displacement data to the base station according to a preset uploading frequency; s3, the base station uploads the three-dimensional displacement data sent by the monitoring station to an early warning platform in real time; and S4, the early warning platform detects the deformation trend and the abnormal change in the area in real time, generates an early warning signal when the abnormal deformation is found, and sends the early warning signal to the user terminal in real time. A distributed GNSS deformation monitoring method is adopted, and dense deployment of distributed nodes is utilized, so that a larger area can be covered, and blind areas are reduced.

This invention discloses a spacecraft configuration-maintaining attitude-orbit coupling cooperative iterative learning control method, belonging to the field of space technology. The implementation method of this invention is as follows: a six-degree-of-freedom dynamic model is established based on the Lie group SE(3) framework, fully considering the attitude-orbit coupling problem caused by thruster installation errors in actual engineering, improving configuration-maintaining accuracy, and avoiding the unwinding problem caused by traditional six-degree-of-freedom dual quaternion representation methods. Based on this, a cooperative iterative learning control method is designed to achieve, in the case of an unknown precise perturbation model, offset the influence of perturbation on configuration maintenance through iterative learning, taking into account both adjustment time and steady-state error, achieving a short adjustment time while maintaining a small steady-state error. That is, high-precision six-degree-of-freedom configuration maintenance of the spacecraft is achieved through attitude-orbit coupling cooperative iterative learning control. This invention can shorten the adjustment time, reduce steady-state error, and improve the speed and accuracy of spacecraft configuration maintenance without relying on a precise perturbation model.

The application discloses a kind of only angle measurement lunar space Lagrange point orbit type intelligent identification method, belong to space technology field.The application establishes only angle measurement observation model, sets only angle measurement observation error covariance matrix;Establish the motion model of space Lagrange point orbit spacecraft in lunar space;According to the number of space Lagrange point orbit in lunar space, randomly generate space Lagrange point orbit in lunar space and recursion length, generate only angle measurement observation vector;The sample point required for training orbit type identification deep neural network is generated;Based on deep neural network, the fast and efficient identification of space Lagrange point orbit in lunar space is realized, and then the initial orbit determination and refined orbit estimation of space Lagrange point orbit in lunar space are realized.The application has the advantages of not depending on physical model, high identification accuracy, small amount of calculation, high efficiency and good robustness.The application is beneficial to improve the space situation awareness capability in lunar space, improve the determination accuracy and calculation speed of space Lagrange point orbit in lunar space.

The invention discloses a spacecraft configuration maintenance attitude-orbit coupling collaborative iterative learning control method, and belongs to the technical field of space. The implementation method comprises the steps that a six-degree-of-freedom kinetic model is established based on a Lie group SE (3) frame, the attitude and orbit coupling problem caused by thruster installation errors in actual engineering is fully considered, the configuration keeping precision is improved, and the unwinding problem caused in a traditional six-degree-of-freedom dual quaternion representation method is avoided. On the basis, a collaborative iterative learning control method is designed, under the condition that an accurate disturbance model is unknown, the influence of disturbance on configuration keeping is counteracted through iterative learning, adjustment time and steady-state errors are considered, the adjustment time is short, meanwhile, the small steady-state errors can be achieved, and the stability of the system is improved. That is to say, the high-precision six-degree-of-freedom configuration of the spacecraft is maintained through attitude and orbit coupling collaborative iterative learning control. The method does not depend on an accurate disturbance model, the adjustment time is shortened, steady-state errors are reduced, and the speed and precision of spacecraft configuration keeping are improved.

The application discloses an interferometric constellation orbit optimization method based on a second-order state transition tensor and belongs to the field of space technology. The method of the application is as follows: an orbit dynamics model and a configuration stability index model of an interferometric constellation spacecraft are established, an orbit deviation propagation expression is given based on a second-order state transition tensor, and a second-order analytical mapping expression of configuration stability of the interferometric constellation and initial orbit deviation is derived; an iterative deviation of the initial state of the spacecraft is given; an initial value of the constellation configuration with strong stability is obtained through iteration, the constellation configuration satisfying long-term stability is obtained through recursive initial value, the interferometric constellation is realized to be long-term stable according to the constellation configuration, and the detection precision of the interferometric constellation is improved. The application has the advantages of high-precision space measurement constellation configuration stability, high efficiency, is favorable for fast iterative optimization of the high-precision space measurement constellation configuration, and is favorable for improving the stability and detection performance of the high-precision space measurement constellation configuration.

This invention discloses a method for analyzing the configuration stability of a high-precision space measurement constellation based on decoupled state transition tensors, belonging to the field of space technology. The method involves: decoupling the influence of different perturbations on configuration stability; establishing a mapping relationship between the spacecraft's orbital state and configuration stability indices; substituting the obtained orbital state transition tensor expression for perturbation decoupling into the method to obtain the state transition tensor expression for the configuration stability indices; analyzing and obtaining the influence of different perturbations on configuration stability, thereby supporting the optimization of the high-precision space measurement constellation configuration, improving the performance of the high-precision space measurement constellation system, and increasing detection accuracy. This invention has the advantages of high accuracy and high efficiency in analyzing the configuration stability of high-precision space measurement constellations, which is beneficial for the rapid iterative optimization of high-precision space measurement constellation configurations and for improving the stability and detection performance of high-precision space measurement constellation configurations.

The present invention relates to the field of space technology and vibration control, and in particular to a vibration suppression control method for a space microgravity vibration isolation system driven by both knowledge and data. First, a dynamic model of a space microgravity vibration isolation system is constructed and used as an initial knowledge-driven module. Then, a knowledge-driven nominal preset performance control law is designed; based on the constructed vibration isolation dynamics model, a preliminary vibration suppression control strategy is designed. Secondly, a data-driven vibration suppression strategy and update law are designed. Through the data-driven module and real-time monitored vibration data, a neural network algorithm is used to approximate the uncertainty and disturbance conditions in the current environment. Furthermore, sample generation and network training based on the nominal controller are carried out to simulate the system dynamics behavior under different disturbances. Finally, online training and application based on dual drive of knowledge and data are performed. The present invention can perform adaptive adjustments for different vibration sources and provide efficient and accurate vibration isolation control.

The invention relates to the technical field of space, in particular to an interference measurement constellation error sensitivity analysis method based on a proxy model. The method comprises the following steps: setting a space interference measurement constellation orbit coordinate system and related parameters; constructing a spacecraft orbit recursion model in the interference measurement constellation; constructing a spatial interference measurement constellation configuration stability index model, a 0-order cutting point of the agent model, a first-order item of the agent model, a second-order item of the agent model and an interference measurement constellation error sensitivity agent model; and giving an initial deviation based on an interference measurement constellation error sensitivity agent model, predicting a configuration stability index evolution condition, and analyzing the interference measurement constellation error sensitivity based on the predicted configuration stability index evolution. By adopting the interference measurement constellation error sensitivity analysis method based on the proxy model, the calculation precision and the calculation efficiency of error sensitivity analysis are effectively improved, the calculation mode is simple, the robustness is good, and the order is convenient to expand.

The application discloses a high-precision interferometric measurement constellation error evolution method based on high-order characteristics and belongs to the field of space technology. The application realizes the method as follows: setting the orbit states of three spacecrafts in an inertial system, setting the orbit dynamics model and initial state of each spacecraft, selecting the characteristic direction of the initial orbit error of the spacecraft, establishing a high-order model of the characteristic direction, establishing a high-order inverse model of the characteristic direction, setting the initial value of the orbit error evolution semi-analytical expression based on the high-order characteristics, calculating the orbit error evolution semi-analytical expression based on the high-order characteristics, constructing the orbit error evolution semi-analytical expression of the three spacecrafts, constructing a configuration stability index model, establishing a high-precision interferometric measurement constellation error evolution model based on the characteristic direction, giving an initial deviation, calculating the configuration stability index deviation at a given moment, and realizing high-precision interferometric measurement constellation error evolution analysis based on the calculated configuration stability index deviation at the given moment.

A six-degree-of-freedom finite-time space rendezvous control method under multiple constraints belongs to the field of space technology. The implementation method of the present invention is as follows: a six-degree-of-freedom dynamic model is established based on the Lie group SE (3) framework, so that the spacecraft rendezvous process can satisfy the position end constraint while also satisfying the attitude end constraint, fully considering the attitude-orbit coupling problem caused by the thruster installation error in actual engineering, improving the spacecraft rendezvous accuracy, and avoiding the unwinding problem caused by the traditional six-degree-of-freedom dual quaternion representation method. A finite-time sliding mode control strategy is designed so that the rendezvous process can satisfy the mission time window constraint and improve the error convergence speed. In order to solve the potential collision risk in the rendezvous process and meet the safety constraints of the rendezvous mission, an artificial potential function is introduced to achieve collision avoidance in the rendezvous process and improve the safety of the spacecraft rendezvous process. The present invention can shorten the rendezvous time, improve the spacecraft space rendezvous accuracy, and improve the safety of the spacecraft rendezvous process.

The invention discloses a six-degree-of-freedom finite time space intersection control method under multiple constraints, and belongs to the technical field of space. According to the implementation method, a six-degree-of-freedom kinetic model is established based on a Lie group SE (3) framework, so that the spacecraft rendezvous process can meet position end constraints and attitude end constraints, the attitude-orbit coupling problem caused by thruster installation errors in actual engineering is fully considered, the spacecraft rendezvous precision is improved, and the spacecraft rendezvous performance is improved. And the unwinding problem caused by a traditional six-degree-of-freedom dual quaternion representation method is avoided. A finite time sliding mode control strategy is designed, so that the rendezvous process can meet task time window constraints, and the error convergence speed is increased. In order to solve potential collision risks in the rendezvous process and meet safety constraints of rendezvous tasks, an artificial potential function is introduced, collision avoidance in the rendezvous process is achieved, and safety of the spacecraft rendezvous process is improved. The rendezvous time can be shortened, the space rendezvous precision of the spacecraft is improved, and the safety of the rendezvous process of the spacecraft is improved.

The invention relates to a system and method for measuring and error analyzing air-floating friction coefficient based on active feedback, and belongs to the field of space technology. The invention proposes an active feedback air-floating friction coefficient measurement method, which can accurately measure the startup and steady-state friction coefficients of a spacecraft in an air-suspended state, thereby obtaining the air-floating friction interference term in the dynamic control model, making the dynamic model more accurate and improving the effectiveness of ground verification tests of spacecraft.

The invention discloses a transfer orbit design method in on-orbit service of a synchronous orbit satellite. The problem of how to effectively image a high-orbit target satellite with low fuel consumption is solved. Belongs to the technical field of space. Comprising the following steps: determining total transfer time according to different initial states of a service satellite and a target satellite; constructing a transfer orbit mathematical model based on the total transfer time to obtain the running time of the service satellite on the initial orbit, the transfer orbit and the approaching orbit in the transfer process and the speed increment for completing the Horman transfer; taking a relative position vector included angle and an illumination angle of two stars at the end of intersection as an objective function, taking total transfer time as a constraint condition, taking the operation time of a service star in an initial orbit and an approaching orbit as an optimization variable, optimizing the objective function by using a genetic algorithm based on an HPOP model, and correcting parameters of a transfer orbit mathematical model to obtain a transfer orbit mathematical model; obtaining a transfer orbit mathematical optimization model; according to the method, effective imaging of the high-orbit target with low burnup is realized through the mathematical optimization model of the transfer orbit.

This invention relates to a method for verifying a structured startup module for spaceborne software, belonging to the field of space technology. The invention verifies the structured startup module of spaceborne software, checking whether the software's operating state is consistent with the design under conditions such as initial power-on, reset, system disconnection, and power-off, whether it can transition to the normal operating process, and whether it can guarantee the normal operation of the spacecraft. This invention verifies, on the one hand, the rationality of the software design and whether it can achieve the expected results; on the other hand, it verifies the correctness of the software logic, verifying the operating logic of the computer upon initial power-on, as well as the correctness of fault logic and process jumps, ensuring the stable operation of the spacecraft in orbit.

The invention discloses a space station material uplink and downlink system with high task adaptability and low cost, and belongs to the technical field of space. Based on the track tugboat which has the configuration two-dimensional expansion capacity and can be repeatedly used in orbit, the system can ascend to pressurize cargoes, large out-cabin loads and cabin sections, descend waste and return useful materials, and has the capacity of rescuing manned airships and independently flying space station cabin sections. The method has the advantages of being high in task adaptability and low in engineering operation and emergency disposal cost.

This invention relates to unified platform for nanosatellite systems which will find application in the field of space technology and satellite communications. The created platform consists of communication buses, to which a power supply module, an on-board computer and a communication module including a UHF transceiver are bidirectionally connected. The platform also includes deployable solar panels, which are connected to the power supply module and to the attitude determination and control module, which is connected in both directions to the communication buses and. The on-board computer is also connected to additional interfaces. The communication module is bidirectionally connected to an antenna unit, including a UHF antenna, which is connected to the additional interface.

This invention discloses a hierarchical task planning method and system for precise space operations, belonging to the field of space technology. To address the large variance of gradient estimates and poor algorithm convergence when policy networks learn multiple tasks, a high-level policy planning method based on a dynamics estimator and Monte Carlo tree search is constructed. The method specifically includes three steps: offline data acquisition, asynchronous high- and low-level policy optimization, and model evaluation. This method improves algorithm convergence, helps conserve onboard computing resources for on-orbit servicing space operations, and meets practical engineering needs.

The invention relates to the technical field of space technology application, and provides an inter-satellite natural semantic communication method based on a big language model, which comprises the following steps: S1, task decomposition and natural semantic coding: decomposing a multi-source instruction including a ground task request and an inter-satellite cooperation signal based on the big language model LLM, and matching functions of other nodes on an inter-satellite link, forming a natural language instruction required by execution; s2, carrying out natural semantic inter-satellite channel coding, and packaging semantics into a data format conforming to a spatial transmission specification; s3, carrying out natural semantic inter-satellite channel decoding, and recovering natural semantics from spatial transmission data; and S4, natural semantic decoding and task execution: understanding natural semantics through a large language model LLM and a local knowledge base, and calling a local tool output instruction to execute a task through a model context protocol MCP. According to the method, anthropomorphic domain-specific language interaction is taken as a core, and flexible expression and intelligent collaboration of inter-satellite complex intentions are realized.

The invention discloses a high-precision interference measurement constellation error evolution method based on high-order characteristics, and belongs to the technical field of space. The implementation method comprises the following steps: setting orbit states of three spacecrafts under an inertial system, and setting an orbit dynamic model and an initial state of each spacecraft; selecting a characteristic direction of an initial orbit injection error of the spacecraft, establishing a high-order model of the characteristic direction, and establishing a high-order inverse model of the characteristic direction; setting an orbit error evolution semi-analytical expression initial value based on high-order characteristics; calculating an orbit error evolution semi-analytical expression based on high-order features; constructing orbit error evolution semi-analytical expressions of the three spacecrafts; constructing a configuration stability index model; establishing a high-precision interference measurement constellation error evolution model based on the characteristic direction; and giving an initial deviation, calculating a configuration stability index deviation at a given moment, and realizing high-precision interference measurement constellation error evolution analysis based on the calculated configuration stability index deviation at the given moment.

The present application relates to the field of space technology test evaluation, and more particularly to a method for analyzing necessity of on-orbit verification of load based on space environment, comprising obtaining test requirement data of space technology to be verified, the test requirement data including environment parameter requirement data, on-orbit operation requirement data, manned spacecraft on-orbit platform resource requirement data, and cost and risk data related to the test; based on the obtained test requirement data, respectively calculating environment irreplaceability comprehensive index N E , human-in-the-loop dependence comprehensive index D H , platform support adaptability comprehensive index A P , and comprehensive benefit gain index G B ; based on the four indexes calculated, according to a preset three-level judgment rule, outputting a manned spacecraft on-orbit test necessity judgment conclusion. The present application quantifies the multi-factor coupled environment irreplaceability, human-machine cooperation irreplaceability and platform matching degree, provides objective decision support for test planning, and optimizes the on-orbit resource allocation of manned spacecraft.

The invention discloses a satellite-borne software structured startup module verification method, and belongs to the technical field of space. According to the method, verification of the structural starting module of the satellite-borne software is completed, and whether the software operation state is consistent with the design, whether the software can be transferred to a normal working process and whether the normal operation of a spacecraft can be ensured are verified under the working conditions of initial power-on, reset, generator tripping, power failure and the like of the software. On one hand, software design rationality is verified, and whether design can achieve an expected effect or not is verified; on the other hand, the software logic correctness is verified, the operation logic of first-time power-on of the computer, the fault logic and the process jump correctness are verified, and stable on-orbit operation of the spacecraft is guaranteed.

This invention relates to the field of space technology experiment evaluation, and more particularly to a method for analyzing the necessity of on-orbit verification of payloads based on the space environment. The method includes acquiring experimental requirement data for the space technology to be verified, including environmental parameter requirements, on-orbit operational requirements, manned spacecraft on-orbit platform resource requirements, and cost and risk data related to the experiment. Based on the acquired experimental requirement data, a comprehensive index of environmental irreplaceability is calculated. N E Human-loop dependency comprehensive index D H Platform compatibility comprehensive indicators A P and comprehensive benefit gain indicators G B Based on the four calculated indicators, and according to the preset three-level judgment rules, the necessity judgment conclusion for the on-orbit test of the manned spacecraft is output. This invention quantifies the unsimulability of the multi-factor coupled environment, the irreplaceability of human-machine collaboration, and the platform matching degree, providing objective decision support for test planning and optimizing the on-orbit resource allocation of manned spacecraft.