43 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.

The invention discloses a flexible constraint multi-agent formation coordination optimal control method based on hierarchical training reinforcement learning, and belongs to the field of flexible constraint multi-agent control. The method comprises the following steps: constructing a dynamic model of the flexible constraint multi-agent formation system; decomposing the control problem of the flexible constraint multi-agent formation system into a tether control layer and a node spacecraft control layer; training a tether control strategy in a tether control layer environment; freezing the trained tether control strategy, taking the tether control strategy as a part of a node spacecraft layer environment, and providing a real-time tether state; training a node spacecraft control strategy in a node spacecraft control layer environment; and deploying the training convergence tether control strategy and the node spacecraft control strategy to an actual system or a simulation environment to realize coordinated optimal control of the flexible constraint multi-agent formation. The method autonomously learns and realizes an approximate optimal cooperative control strategy coping with a strong coupling dynamics system under the multi-target constraint.

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.

The application discloses a kind of management methods of large batch simulation of spacecraft control system, including the data structure of specification simulation case, storage forms simulation case node file;Through establishing simulation case management tool, load and analyze simulation case node file, realize the management and difference comparison of simulation case node;Batch generates simulation case node file, and independently carries out simulation.The present application is based on the engineering demand of spacecraft control system scheme simulation verification, can carry out the design generation of simulation case, and the designed simulation case is easy to artificial interpretation, easy to quickly reuse and modify.Can carry out the centralized management and automatic execution of massive simulation case, with the simulation verification ability of large batch task scene of unattended, ensure the high reliability and high efficiency of spacecraft control system simulation verification.

The application discloses a spacecraft control method and system considering residual vibration, and the method comprises the following steps: based on Euler rotation theorem, an Euler axis expression of a spacecraft attitude maneuver process is constructed; a dynamic model of a flexible spacecraft is established; a residual vibration expression of a flexible accessory in the flexible spacecraft is constructed; according to the characteristics and threshold limit of the residual vibration, a trajectory planning is carried out, and a spacecraft angular acceleration trajectory planning curve is generated; and according to the Euler axis expression of the spacecraft attitude maneuver process, the dynamic model of the flexible spacecraft and the spacecraft angular acceleration trajectory planning curve, a control loop is composed. The system comprises an Euler axis calculation module, a dynamic model construction module, a residual vibration construction module, an angular acceleration trajectory planning module and an overall control module. By using the application, the vibration generated in the maneuver process can be inhibited. The application can be widely applied in the field of spacecraft control.

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 application relates to a method for establishing a high-stability inertial reference of a satellite star sensor, belonging to the technical field of spacecraft control, which comprises the following steps: integrating the star sensor and a load accelerometer; using a separated two-stage light shield to suppress the stray light of the star sensor in the cabin; using a sectional heat control coating outside the two-stage light shield, and performing anti-stray-light treatment on the mounting surface of the light shield; calculating the inertial attitude of the load accelerometer by using the star sensor output, and establishing an inertial attitude model of the load accelerometer; calibrating the installation matrix of the star sensor based on the load accelerometer, and establishing a unified inertial measurement reference between the star sensor and the load accelerometer; and establishing a unified inertial measurement reference between multiple star sensors.

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.

The application discloses a kind of control system simulation method and system of on-orbit spacecraft, the method is applied to the host computer of the simulation system of on-orbit spacecraft control system, system includes sensor simulation unit, GNC controller simulation unit, actuator simulation unit, dynamics model and telemetry data storage unit;Method includes: for the first control cycle, GNC controller simulation unit constructs the initial parameter of simulation system based on the telemetry data of preset time in telemetry data storage unit, and carries out the simulation calculation of this control cycle based on initial parameter;For other control cycles, all execute: judge whether to use telemetry data to participate in the simulation calculation of this control cycle;If yes, then GNC controller simulation unit calls the telemetry data of corresponding time, and carries out the simulation calculation of this control cycle based on the telemetry data of calling;If no, then carry out the simulation calculation of this control cycle based on measurement state.This application can improve the accuracy of simulation system.

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 application discloses a spacecraft control system fault identification method and device based on multi-source information fusion. The method comprises the following steps: multi-source fusion of data in the development and testing process of a spacecraft and historical data in orbit to obtain a plurality of fault samples; each fault sample is provided with a fault label, and each fault label is used for representing a fault mode corresponding to the fault sample; each fault sample is normalized to obtain a normalized fault sample; the normalized fault sample is subjected to clustering processing based on a preset clustering algorithm to obtain a plurality of clusters and a fault mode of each cluster; each cluster comprises at least one core object; for in-orbit data at any moment, the distance between the in-orbit data and the core object in each cluster is calculated to determine the cluster to which the in-orbit data belongs based on the calculated distance; and the fault mode corresponding to the cluster to which the in-orbit data belongs is determined as the identification result of the in-orbit data. The application can improve the accuracy of spacecraft fault identification.

A kind of graphical multi-configuration transformation spacecraft control system simulation program generation method, including combination dynamics model construction method;Fast model encapsulation and model library construction method;Multi-configuration transformation spacecraft control system modeling method;Model automatic wiring and layout method;Simulation program automatic generation method.The present application can be standardized and centrally managed to C / C++ language written basic model (including sensor model, controller model, actuator model, dynamics model, environment model etc.), form reusable model library, using digital means, with graphical interactive operation mode, can quickly build multi-configuration transformation spacecraft control system simulation model, automatically generate the control system simulation program supporting multi-configuration transformation, the method has been applied in space station development process, solved the problem that space station multi-configuration simulation is difficult to maintain, significantly improves development efficiency.

A kind of spacecraft control system user demand automatic generation method based on C code, algorithm encapsulation is carried out using C compiling technology;Then state flow chart and program flow chart are built, and user demand simulation C code is edited to carry out controller design;Then the user edited user demand simulation C code is encapsulated into a C++ class that meets the interface requirements of mathematical simulation verification platform;Finally, the C code is automatically generated into a user demand Word document according to the template format, and the C language variables specified by the user for replacement are replaced into mathematical symbols, and the mathematical functions existing in the rule table are replaced into mathematical calculation expressed by mathematical symbol table. The present application realizes the encapsulation of mature and reliable C code, which is used as a standard component for writing software user demand, solves the problems of manual writing of spacecraft control system user demand, no effective tool for error checking, and inconsistency between user demand and simulation verification code.

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 invention discloses a rapid attitude planning method for a spacecraft under multiple constraints, and belongs to the technical field of spacecraft control. The method comprises the steps of establishing a quaternion attitude model, constructing a photosensitive load and communication antenna pointing constraint, designing a multi-constraint index evaluation function, sampling a feasible angular velocity set by adopting a dynamic window algorithm, and selecting an optimal control instruction to realize attitude maneuver. According to the method, multiple constraint conditions can be processed in real time through the dynamic window method, rapid, smooth and safe attitude planning is achieved in combination with the multi-target evaluation function, and the real-time performance, autonomy and robustness of spacecraft attitude control are remarkably improved.

The application discloses a spacecraft control information determination method, device, equipment, medium and product, and relates to the technical field of spacecrafts. The method obtains first state information, constraint information, control input information and second state information of an associated spacecraft of the spacecraft; determines a constraint violation value according to the first state information and the constraint information; determines a degradation value according to the constraint violation value, the first state information and the control input information; when the degradation value or the control input information exceeds a set threshold value, adjusts a degradation reference trajectory according to the constraint violation value and the degradation value, and determines target control input information in combination with the first state information and the second state information; otherwise, determines updated control input information according to the first state information, the second state information and the constraint violation value, and determines the target control input information according to the first state information, the constraint violation value and the updated control input information. The embodiment of the application realizes the determination of the control input information in the case that the constraints cannot be satisfied simultaneously.

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.