26 results about "Space vehicle control" patented technology

The invention discloses a six-degree-of-freedom multi-spacecraft control experiment platform and method based on semi-physics-virtuality, belongs to the technical field of spacecraft system and control system simulation, and constructs a multi-spacecraft system composed of semi-physics spacecrafts and pure virtual spacecrafts. The attitude of the semi-physical spacecraft is simulated by a simulator mounted on a three-dimensional attitude air floating table, is acquired by a sensor and is converted into a rotation matrix form, and a translation part of the semi-physical spacecraft is virtually calculated through a simulation integrator; the pose of the virtual spacecraft is virtually simulated by an SE (3) integrator with high-guarantee geometric characteristics; the system takes the ROS as a communication framework, and realizes state and instruction release by adopting a custom message; and ensuring that the time system of each spacecraft is consistent with the real time through a time synchronization mechanism. The platform supports cooperative operation of a semi-physical spacecraft and a pure virtual spacecraft, and has the characteristics of flexible structure, high simulation precision and good expandability.

The invention discloses a method for predicting the remaining service life of a control moment gyroscope based on parameter fusion, and relates to the technical field of health management of spacecraft attitude control systems. According to the method, optimal fusion of measurement parameters and implicit parameters is realized through adversarial learning feature extraction and a channel attention mechanism, and the method is particularly suitable for high-precision residual life prediction of the spacecraft control moment gyroscope under variable working conditions. The method comprises four main steps of data preparation and multi-source input, adversarial learning feature extraction, SE-Attention feature fusion and three-stage training optimization. In the data preparation step, original time sequences of voltage and current of a rotor motor are collected through a built-in electrical sensor of a CMG, and key physical parameters of a full-film lubrication friction coefficient, lubricating oil viscosity, a bearing clearance and contact stiffness are inverted based on an electromechanical coupling model; in the adversarial learning step, a game framework of a feature extractor and a working condition discriminator is constructed, and working condition-independent robust feature extraction is realized through a gradient inversion layer; in the SE-Attention feature fusion step, importance weights are adaptively distributed to different feature channels through an extrusion-excitation mechanism; and the three-stage training optimization step adopts a preheating-confrontation-fine tuning strategy to ensure the network convergence. The method solves the problem of domain adaptation, has the technical advantages of strong robustness, high precision, strong generalization, interpretability and self-adaptation, and can realize high-precision RUL prediction under variable working conditions.

This invention relates to a test device and verification method for the space-ground consistency of a flexible spacecraft control system based on uncertainty quantification, belonging to the fields of intelligent manufacturing equipment industry and space-ground consistency verification technology. To address the problem of space-ground consistency verification and evaluation for flexible spacecraft, this invention includes an air-floating robot and a flexible component. The air-floating robot simulates the central rigid body of the flexible spacecraft, and the flexible component simulates the deployed solar panels. One end of the flexible component is connected to the upper support column of the air-floating robot, while the other end is free and equipped with a weight. Fiber optic grating sensors are distributed on the flexible component to measure flexible vibrations. A piezoelectric ceramic plate is installed at the end of the flexible component connected to the air-floating robot for driving and controlling the vibration of the flexible component. This invention evaluates the space-ground consistency of the designed flexible spacecraft control test system and, based on uncertainty quantification analysis, provides the impact level of uncertainty sources on space-ground consistency.

The invention discloses a control distribution method for a spacecraft with an eight-thruster structure, and belongs to the technical field of spacecraft control. The problem that after the number of thrusters of an existing spacecraft is reduced, decoupling control requirements of force / torque in all directions cannot be met is solved. Comprising the steps that eight thrusters are arranged on a spacecraft, so that the eight thrusters are distributed in an axial symmetry mode in a structure installation coordinate system, nozzles face an evasion limiting area, and the axes of the nozzles evade the centroid deviation range of a whole-satellite structure center area; for the combined control of the expected single-axis force and the expected three-axis torque, establishing a relational expression by taking the longest total ignition duration of the four thrusters as a target, and calculating the ignition duration of each thruster; and for combined control of the expected triaxial force and the expected triaxial torque, six thrusters are selected from the eight thrusters to establish a relational expression between the ignition duration and the expected triaxial force and the expected triaxial torque, and the ignition duration of the six thrusters is calculated. The method is used for ignition control of the spacecraft.

The invention relates to an intelligent synthesis-oriented spacecraft component equipment description method, and belongs to the technical field of architecture design. Abstract modeling and unified description are carried out on management operation of common component equipment in a spacecraft control subsystem, an application example is given, and the development efficiency and quality of spaceflight embedded software intelligent composite codes can be improved. According to the method, firstly, component equipment composition of an existing spacecraft control subsystem is combed, then general and difference demand analysis is carried out on component equipment management functions of the spacecraft control subsystem, then a software unified equipment model oriented to control system components is established, and finally a specific example is given. The method can quickly adapt to software codes formed by different parts of the control system, the reuse capability of the control software is improved, and then intelligent synthesis of the software is supported.

This application provides an electromagnetic spacecraft control method, device, product, and electronic device, relating to the field of spacecraft control technology. This method can obtain the directed connected topology network of an electromagnetic spacecraft formation system and the dynamic equations of each electromagnetic spacecraft. Based on the dynamic equations and the directed connected topology network, a distributed control protocol is constructed. When the state update error of the distributed control protocol meets preset conditions, position control is performed on each electromagnetic spacecraft based on the distributed control protocol. This method can save communication resources while precisely controlling the electromagnetic spacecraft and can solve the actuator saturation problem involved in maintaining the formation configuration of a multi-electromagnetic spacecraft formation system. Compared with related technologies, this method considers the inherent characteristics of electromagnetic spacecraft, eliminating the need to trigger communication at every moment; communication only needs to be triggered when the state update error of the distributed control protocol meets preset conditions.

The invention provides a high-precision inter-satellite pointing control method fusing laser interference measurement, and relates to the field of spacecraft control. The method comprises the following steps: firstly, carrying out validity judgment and processing on an attitude error value provided by the measurement of the laser interferometer in the period, carrying out attitude angular velocity differential calculation, and when the attitude error value provided by the measurement of the laser interferometer in the period is valid, carrying out reference error calculation in combination with an attitude error value provided by orbit extrapolation and on-satellite attitude determination; then, according to attitude error value validity provided by measurement of the laser interferometer in the current period and the last period, on the basis of a reference introduction / direct introduction selection mark set on the ground, determining to implement reference introduction calculation or direct introduction calculation, calculating an attitude control quantity and an angular velocity control quantity, and finally, according to the attitude control quantity and the angular velocity control quantity, calculating the attitude control quantity and the angular velocity control quantity. An attitude control moment is calculated by using a control law and is implemented by an execution mechanism. According to the invention, high-precision pointing error measurement of the laser interferometer is introduced, so that high-precision inter-satellite pointing control of satellites is realized.

This invention discloses a method and apparatus for real-time onboard determination of interference in the field of view of a star sensor, belonging to the field of spacecraft control technology. The method includes: pre-determining the obstruction rotation range of the dual-axis solar array's field of view based on its geometry, installation position, and rotation method, as well as the star sensor's field of view angle, installation position, and orientation; determining that the star sensor's field of view is obstructed by the dual-axis solar array when the rotation angles of the A-axis and B-axis of the dual-axis solar array are within the obstruction rotation range; acquiring the rotation angle measurements of the A-axis and B-axis of the dual-axis solar array in real-time on-orbit, and determining whether the star sensor's field of view is obstructed by the dual-axis solar array based on whether the rotation angle measurements are within the obstruction rotation range; and determining whether the star sensor is affected by solar interference based on the change of the solar vector in the satellite orbital system; thus obtaining the result of interference in the star sensor's field of view. This invention enables real-time onboard determination of the star sensor's field of view interference results.

A kind of stop method of space vehicle target access optimization task based on local search and exploration enhancement belongs to space vehicle control technical field.To solve the problem of optimization termination control with robustness and self-adaptation, the present application designs optimization termination criterion as when the maximum prediction standard deviation of Gaussian process regression surrogate model in current local area is lower than preset threshold, terminate the iteration optimization of bayesian optimization;The design optimization termination criterion is embedded into BO;The prediction standard deviation of GPR surrogate model is used to measure the cognitive uncertainty of GPR surrogate model to objective function at a point;Set the judgment condition to judge whether the exploration enhancement ability of acquisition function is enhanced and whether local evaluation is triggered.The present application realizes the optimization termination control with robustness and self-adaptation.

The invention discloses a satellite autonomous rendezvous control method based on adaptive time sequence hidden space learning, and relates to the technical field of deep reinforcement learning and spacecraft control. According to the method, a satellite long-distance rendezvous task is modeled into a partially observable Markov decision process, and the core innovation is that the decision time step length is converted from a fixed hyper-parameter to a key decision variable which can be autonomously optimized by an intelligent agent. A continuous step size scaling factor is introduced by expanding an action space, so that end-to-end collaborative optimization of a control strategy and a decision time sequence is realized; designing a self-adaptive time sequence decision-making mechanism based on hidden state perception, and enabling an intelligent agent to dynamically adjust decision-making frequency according to a task stage; and constructing a multi-target composite reward function containing step length efficiency reward and punishment, and guiding the intelligent agent to intelligently balance the control precision and task time consumption under multiple constraints. Experiments prove that the method is obviously superior to a fixed step length baseline algorithm in task success rate, fuel economy and time efficiency, shows strong generalization ability, and provides an efficient and robust solution for intelligent autonomous control of spacecrafts.

The invention discloses a spacecraft adaptive sliding mode saturation formation control method and system based on SE (3), and belongs to the technical field of spacecraft control. The method comprises the following main processes: modeling a system model by utilizing SE (3), and describing the attitude and orbit of a spacecraft; performing error design based on attitude dynamics to obtain a configuration error function; designing a self-adaptive sliding mode saturation controller based on kinematics and dynamics models of a rigid body dynamics system of SE (3) in combination with a configuration error function; a periodic event triggering strategy is provided, and the control rate of the spacecraft is periodically updated. According to the method, event trigger control and periodic sampling data control are combined through the self-adaptive sliding mode saturation controller, the Zeno phenomenon is avoided so as to reduce resource consumption, the unified lower bound of interaction time is deduced, high-precision control of the spacecraft during on-orbit task execution is guaranteed, and the method has the advantages of being simple in structure and convenient to operate. And meanwhile, the formation collision phenomenon is also avoided.

Provided are a spacecraft control system, a spacecraft control method, and a server device capable of quickly transmitting control data to a spacecraft. The spacecraft system includes: a plurality of spacecrafts; a plurality of small communicators that transmit control data related to imaging operations of the spacecrafts to the spacecrafts; and at least one ground station that receives, from the spacecrafts, captured image data related to images captured by the spacecrafts and space position data for the spacecrafts, and transmits identification data for the small communicators to the spacecrafts, in which each of the spacecrafts includes at least a first reception device that receives the control data transmitted from the small communicators, a control device that controls an imaging operation of the spacecraft on the basis of the control data received by the first reception device, and an imaging device that executes the imaging operation.

The invention provides a micro-nano satellite multi-constraint fine control method based on actuator dynamics, and belongs to the field of spacecraft control, and the method comprises the steps: considering the electromechanical characteristics of a reaction wheel, and building a micro-nano satellite attitude deep coupling model containing the dynamic characteristics of the reaction wheel; aiming at reaction wheel axle friction and counter electromotive force interference, designing a double-interference observer to accurately estimate the reaction wheel axle friction and counter electromotive force interference; designing a nonlinear state-dependent obstacle function to convert a constraint state into a tracking control variable according to the star rotating speed and reaction wheel current physical constraint; and a composite controller is constructed by combining a double-interference observer and a nonlinear state-dependent barrier function, and the design of the high-precision controller of the micro-nano satellite is completed. According to the method, the problem of various physical constraints in micro-nano satellite control can be solved, the interference of the reaction wheel is compensated, and the method has the characteristics of high reliability and high accuracy.

This invention provides a method and system for coordinated attitude control of ultra-low Earth orbit (UEO) satellites based on aerodynamics, relating to the field of spacecraft control technology. The method includes: obtaining the relative position error of the secondary satellite and the attitude error of the primary satellite based on real-time state data; determining the control mode of the secondary satellite based on the relative position error and the attitude error of the primary satellite; when the control mode is position control, constructing a sliding mode surface and solving it using a composite reaching law to determine the target deflection angle; when the control mode is attitude control, determining the desired control torque of the secondary satellite, constructing a first-order decaying differential equation based on the desired control torque, and determining the target deflection angle using a local Jacobian matrix and the first-order decaying differential equation; and driving the aerodynamic actuators of the secondary satellite to perform coordinated attitude control actions relative to the primary satellite based on the target deflection angle. This invention improves the control accuracy for aerodynamic UEO satellite attitude control.

The invention relates to a robust adaptive dynamic planning method for spacecraft attitude control, and belongs to the technical field of spacecraft control. According to the method, the neural network identifier is designed, online identification and reconstruction are carried out on unknown dynamic items caused by inertia parameter uncertainty in spacecraft attitude dynamics, the attitude tracking optimal control law based on adaptive dynamics is constructed on the basis, and under the condition that inertia parameter uncertainty exists, the attitude tracking optimal control law is optimized. The balance between tracking precision and control energy consumption is realized, and the applicability of the attitude control method in complex space tasks is improved.

The invention provides a random MPC control method driven by reinforcement learning and inspired by terminal strategy optimization, and relates to the field of on-orbit avoidance control, and the method comprises the steps: constructing a discrete relative dynamics model; generating expert demonstration data through stochastic model predictive control (SMPC); defining the state and action of reinforcement learning and designing an instant reward function; training a terminal strategy by adopting an offline reinforcement learning algorithm, determining a reinforcement learning state in a prediction time domain and generating corresponding reference control input to obtain an initial reference control sequence; and a random model prediction control SMPC framework is adopted, an initial reference control sequence is introduced into a target function, solving optimization is performed to obtain a current control quantity, and prediction and optimization at the next moment are propelled in a rolling manner. According to the method, through fusion of reinforcement learning and random model prediction control, the spacecraft control has better adaptability and robustness when dealing with orbit perturbation and environment uncertainty, and the overall control performance of spacecraft space debris avoidance is remarkably improved.

The invention discloses a fuzzy adaptive sliding mode control method and system suitable for a satellite flexible solar wing system, and belongs to the technical field of spacecraft control. Aiming at the nonlinear vibration problem caused by structural flexibility, joint gaps and abrasion of a satellite flexible solar wing and the defects of poor trajectory tracking precision and the like caused by insufficient robustness, obvious buffeting and insufficient adaptive capacity of an existing control method, the method comprises the following steps of: constructing a nonlinear dynamic model for accurately representing complex dynamic characteristics of a system; and designing a fuzzy adaptive sliding mode control method to realize high-precision control of the attitude of the flexible solar wing. The method has the strong robustness of sliding mode control and the dynamic adaptation capability of fuzzy self-adaption, the inherent defects of a traditional control method are effectively overcome, the system operation stability and control precision are improved, and technical support is provided for reliable control of spacecraft flexible components.

The application discloses a kind of dual-axis solar wing control method and device giving consideration to flexibility inhibition, belong to spacecraft control technical field.Method includes: according to the unit vector of sun in satellite orbit system, the nominal angle of A axis and B axis is calculated respectively;Wherein, the nominal angle of B axis is after discretization processing and according to the flexible vibration inhibition interval determined;The flexible vibration inhibition interval is the angle avoidance interval set to avoid resonance frequency;According to the nominal angle and the rotation angle measurement value of A axis and B axis, the control error of A axis and B axis is calculated;According to satellite imaging state and the state of dual-axis solar wing, the rotation of A axis and B axis is controlled using control error, to eliminate the sun error of dual-axis solar wing.The application can avoid the resonance interval of driving frequency and axial vibration frequency, reduce resonance risk.

The application discloses a kinetic energy interceptor attitude control method and system with specified time convergence, and the method comprises the following steps: step one, kinetic energy interceptor attitude control system modeling; step two, preset performance function design; step three, obstacle Lyapunov function design; step four, virtual controller design; step five, RBF adaptive estimation law design; step six, controller design; step seven, control moment adopts PWPF pulse width modulation discretization. The system comprises the following modules: state equation construction module, performance function preset module, obstacle Lyapunov function construction module, virtual controller construction module, adaptive estimation law construction module, final controller construction module and discrete modulation module. The kinetic energy interceptor attitude control with specified time can be realized by designing the preset performance function and the controller. The application can be widely applied in the technical field of spacecraft control.

This invention relates to the field of spacecraft control technology and discloses a spacecraft formation maintenance control system and its control method. The invention maintains the relative position between two spacecraft by setting a flexible connector between them and using a drive assembly to control the movement of a mass block slidably fitted onto the flexible connector. The return impulse of the mass block at both ends of the flexible connector generates opposing forces on the two control platforms to maintain the relative position between the two spacecraft. This eliminates the need for continuous consumption of propellant or working fluid, thus reducing resource consumption during formation maintenance. The invention enables control over the movement of the mass block, facilitating fine adjustment of the maintenance control force and improving the accuracy of maintaining the relative position of the formation. It is suitable for long-term, low-disturbance, and high-precision spacecraft formation maintenance control scenarios.

The application provides a spacecraft injection data analysis and verification method and device, which can be used in the technical field of spacecraft control. The method comprises the following steps: obtaining data source code of injection data; the injection data is generated according to preset input parameters; using the configuration information data dictionary of the injection data constructed in advance to analyze the data source code to obtain an analysis result; and verifying the injection data according to the preset input parameters and the analysis result. The spacecraft injection data analysis and verification method and device provided by the application effectively shield the differences of different spacecraft injection formats and realize the standardized analysis and verification of remote control injection data.

The application discloses a space debris cleaning mechanism based on electromagnetic ejection and a debris cleaning method, and relates to the technical field of spacecraft control research.The application is characterized in that flexible grippers are symmetrically arranged on the two sides of a service spacecraft, the flexible grippers are connected to corresponding spinning wheels arranged on brushless motors through ropes, and electromagnetic coils are arranged on the side walls of the flexible grippers and the inner side walls of the service spacecraft corresponding to the flexible grippers.In use, the two electromagnetic coils are oppositely arranged, and the current of the two electromagnetic coils is in the same direction, so that the flexible grippers are ejected to the space debris; meanwhile, the brushless motors generate different resistances according to the orientation of the debris, so that the spinning wheels rotate the ropes to stretch the flexible grippers to the space debris, and then the rudders and electromagnets arranged on the flexible grippers are used to adsorb and capture the space debris, thereby avoiding the problem that the spring is not easy to control and unstable in use, and the efficiency of capturing the debris in space is high, and the high-quality cleaning of the debris in the space can be ensured.

The application discloses a kind of based on diagnosability spacecraft fault diagnosis measurement data autonomous optimization method, comprising: obtaining the state equation of spacecraft control system;According to the state equation of spacecraft control system, system dynamic expression is constructed;According to system dynamic expression, system dynamic expression determined by diagnostic measurement data is constructed;According to the system dynamic expression determined by diagnostic measurement data, the random characteristics corresponding to diagnostic measurement data combination is obtained;According to the random characteristics corresponding to diagnostic measurement data combination, the diagnosability analysis result corresponding to diagnostic measurement data combination is obtained;According to the diagnosability analysis result corresponding to diagnostic measurement data combination, diagnostic measurement data autonomous screening logic is obtained;According to diagnostic measurement data autonomous screening logic, the optimal diagnostic measurement data combination is obtained.The application effectively reduces the amount of fault diagnosis calculation by reducing the amount of data required for diagnosis, realizes the autonomous optimization of spacecraft diagnostic measurement data.

The application provides a micro-nano satellite multi-constraint fine control method based on actuator dynamics, and belongs to the field of spacecraft control, and comprises the following steps: considering the electromechanical characteristics of the reaction wheel, an attitude deep coupling model of the micro-nano satellite containing the dynamic characteristics of the reaction wheel is established; aiming at the wheel shaft friction and back electromotive force interference of the reaction wheel, a double interference observer is designed to accurately estimate the wheel shaft friction and back electromotive force interference; aiming at the physical constraints of the satellite rotation speed and the reaction wheel current, a nonlinear state-dependent barrier function is designed to convert the constraint state into a tracking control variable; a composite controller is constructed by combining the double interference observer and the nonlinear state-dependent barrier function, and the high-precision controller design of the micro-nano satellite is completed. The application can overcome the multi-class physical constraint problem in the control of the micro-nano satellite, compensate the interference of the reaction wheel, and has the characteristics of high reliability and high accuracy.

This invention discloses a dragless spacecraft control method based on a combination of BP neural network and PID control, belonging to the field of dynamics and control. The method includes the following steps: establishing a dynamic model of the dragless spacecraft and the test mass; performing PID negative feedback control based on the error between the real-time attitude information of the spacecraft and the test mass and the target value; constructing a BP neural network to adjust the three parameters of the PID controller based on the current control results, outputting a suitable control force; and iteratively optimizing the structure of the BP neural network to achieve high-precision control during spacecraft operation. This invention introduces a BP neural network into traditional PID control, enabling precise control of multi-body, multi-degree-of-freedom systems in space. In the scientific model of gravitational wave detection, it solves the coupling problem between different loops of the spacecraft and the test mass, exhibits good adaptability to random disturbances in space, and meets high-precision requirements.

This invention belongs to the field of spacecraft control technology, specifically relating to a spacecraft formation attitude-orbit coupling predetermined time control method based on dual quaternions. First, a spacecraft relative motion attitude-orbit coupling control system is established. Then, a predetermined time attitude-orbit coupling controller for the spacecraft formation is designed. Finally, based on the coupled dynamics model, the attitude-orbit dual quaternion and dual velocity of each spacecraft are output. These are fed back to the input of the coupling control system, and the differences between these errors and the given desired attitude-orbit dual quaternion and desired velocity are calculated to obtain the dual quaternion error and dual velocity error. Based on these two error states, the predetermined time controller for the spacecraft formation attitude-orbit coupling is calculated, and the predetermined time result of the spacecraft formation attitude-orbit coupling is output. This invention considers model uncertainties and external disturbances by constructing an extended state observer and a predetermined time attitude-orbit coupling controller to achieve predetermined time stability of the spacecraft formation system.