136 results about "Flexible spacecraft" patented technology

The present invention provides an emulation system having a control system that allows the testing of a satellite control system with all of its hardware in place, i.e. fully integrated. The emulation methodology is applicable to the case of either a rigid spacecraft or a flexible spacecraft, provided that the spacecraft's sensors and actuators are stowed to the rigid part of spacecraft in the case of a flexible spacecraft. Practically, the latter condition is not restrictive, as the actuators and sensors are usually placed rigidly in the satellite bus, while the satellite solar panels constitute the flexible elements. The control system is used to tune the mass properties and dynamic behaviour of a rigid ground-spacecraft in a 1-G environment to those of a flight-spacecraft in 0-G. A six-axis force / moment sensor is placed at an interface of the ground-spacecraft and a manipulator. Signals received from the force / moment sensor, and in some cases signals relating to the position and velocity of manipulator joints, are received into the control system.

A composite anti-jamming attitude control method for a flexible spacecraft, which is characterized in that it includes the following steps: First, by considering the vibration of the flexible attachment, the change of the spacecraft moment of inertia caused by the expansion of the flexible attachment and the space environment disturbance torque on the attitude In order to control the influence of control, a flexible spacecraft dynamics model including neutral uncertain dynamic items and external equivalent disturbance variables is established; secondly, in view of the serious influence of structural vibration on the stability of the spacecraft, and the large existence of flexible accessories such as sailboards, With the characteristics of flexibility and low damping, the PPF active vibration controller is constructed to reduce the impact of vibration modes on the spacecraft body; thirdly, the H∞ anti-jamming controller is designed to suppress vibrations from flexible mechanisms such as sailboards and extension rods. The disturbance caused by the change of the spacecraft rotational inertia caused by the deployment and the bounded disturbance such as the space environment disturbance moment; finally, based on the convex optimization algorithm, the composite anti-jamming output feedback attitude and the vibration composite controller are solved; Design and other advantages, can be used for high stability control of flexible spacecraft.

The invention relates to a flexible satellite locus linearization attitude control method based on a disturbance observer. The invention aims at solving problems that a single locus linearization control method is poor in capability of inhibiting interference, is poorer in robustness, and does not consider external interference and the impact from flexible accessories. The method comprises the steps: employing Euler angles for describing attitudes of a spacecraft, employing an idea of equivalent disturbance, and building a flexible spacecraft dynamics and kinetics equation; solving the pseudo-inverse of a controlled object under the condition of neglecting equivalent disturbance, designing a quasi-differentiator of a specific type, and obtaining the nominal control of an expected locus; and designing a linear time varying adjuster through proportion-integration control. The method gives consideration to the influence of equivalent disturbance, designs the disturbance observer, and guarantees the asymptotic convergence of a tracking error of a flexible spacecraft. The method improves the anti-interference capability of a system, and improves the robustness of the system. The method is used in the attitude control field of flexible satellites.

The invention belongs to the field of spacecraft control and relates to a fuzzy singular perturbation modeling and robust attitude control method for complex flexible spacecraft, namely a robust combined control method for fusing static output feedback control and output integration. The method comprises the following steps of: establishing an uncertain continuous fuzzy singular perturbation model and a standard discrete fuzzy singular perturbation model according to a dynamic model and a kinematic model of the spacecraft in combination with fuzzy logic and singular perturbation technology; and designing a robust controller combined by a static output feedback controller and an output integrator by a spectral norm and linear matrix inequality (LMI) method and resolving a group of LMIs which are unrelated to a perturbation parameter so as to obtain a controller parameter and solve an ill-conditioned problem caused by the perturbation parameter and the problem of difficulty in selection of an initial value in an LMI resolving static output feedback controller gain method. Through the method, flexible vibration and external interference can be overcome effectively, and control effects such as high response speed, high attitude control accuracy, high anti-jamming capability and high robust performance are achieved.

The invention relates to a method for controlling an index time-varying slide mode of flexible spacecraft characteristic shaft attitude maneuver, and belongs to the technical field of spacecraft control. The method comprises the steps that firstly, a system dynamically equivalent model, a dynamic model and a flexible vibration model are established under a spacecraft system, then, the vibration frequency and the damping ratio parameter of a closed loop system with the index time-varying slide mode control law are calculated, and a single-shaft multi-modality filtering input shaping device with a characteristic shaft as a rotary shaft is designed according to the designing method of the single-shaft input shaping device to restrain flexible vibration in three-shaft motion. Meanwhile, a state observer is designed to estimate flexible modal information in real time, and the method for controlling an output feedback index time-varying slide mode is formed. At last, saturability analysis is conducted on control torque so as to satisfy the physical saturation constraint of the control torque. By means of the method, the application range of existing input shaping is expanded, the input shaping technology is expanded from single-shaft maneuver to three-shaft maneuver, the self-robustness of filter input shaping is enhanced, and the purpose that the attitude maneuver path of the spacecraft is the shortest is achieved.

The invention relates to a method of designing a multi-channel linear active disturbance rejection controller, comprising the following steps: S1, rewriting the dynamics and kinematics equations of a flexible spacecraft, and getting a form suitable for designing an active disturbance rejection controller; S2, designing a third-order linear extended state observer to estimate the system state quantity and the total internal and external disturbances; and S3, designing a multi-channel linear active disturbance rejection controller by use of generalized disturbance estimated by the observer. The method has the advantages possessed by the traditional active disturbance rejection control method, does not need to know the accurate mathematical model of the system, and has characteristics such as less overshoot, high precision, strong adaptability, high stability and strong robustness. In addition, the control law is improved for spacecrafts with inertial gyro failure on the basis of an active disturbance rejection controller, the control law is designed by using the observed state of the observer to replace the attitude angle and attitude angular velocity information of a spacecraft, and a fault-tolerant control method is provided for the condition of gyro failure.

The invention provides a T-S fuzzy model-based flexible spacecraft multi-objective integrated control method. According to the method, a T-S fuzzy dynamic model of a flexible spacecraft is established, and the universal approximation of the T-S fuzzy model of the spacecraft is proved; the uncertainty of the inertia of the spacecraft which is caused by relative movement of flexible parts, and various kinds of space disturbance torques are considered; a control-performance LMI description and multi-objective integrated LMI method is adopted; and a robust H-infinity state feedback controller which can make a closed-loop system meet pole constraints and control input constraints is designed based on the T-S fuzzy model of the spacecraft. As indicated by numerical value simulation results, the designed state feedback control system has the advantages of short dynamic adjustment time, fast response, small overshoot and high steady-state accuracy, and can effectively inhibit the vibration of the flexible parts caused by attitude variation, and has high robustness and adaptability for the uncertainty of the model of the spacecraft.

The invention relates to a flexible spacecraft attitude control system and a flexible spacecraft attitude control method in allusion to flywheel low-speed friction. The system comprises six modules which are a flexible spacecraft dynamics real-time simulation module with actuating mechanism characteristics, a flexible spacecraft kinematics real-time simulation module, an attitude measurement module, an attitude determination module, an attitude control module and an executing mechanism module, wherein the flexible spacecraft dynamics real-time simulation module with the executing mechanism characteristics comprises a spacecraft body, a flywheel and flexible appendage dynamics; the flexible spacecraft kinematics real-time simulation module can select different types of attitude description modes according to task requirements; the attitude measurement module can select different real sensors and sensor simulators according to the task requirements; the attitude control module comprises a conventional PID control method, a robust control method and the flexible spacecraft attitude control method in allusion to flywheel low-speed friction, and switching can be carried out according to the task requirements of the system; and the executing mechanism module comprises a real flywheel and a thruster simulator.

The present invention provides a method for attitude control based on finite time friction estimation for a flexible spacecraft. The control method includes the following steps: a. introducing spacecraft flywheel friction disturbance into a spacecraft dynamics system, and establishing a flexible spacecraft dynamics system with flywheel friction disturbance; b. converting the flexible spacecraft dynamics system with flywheel friction disturbance into a state-space form; c. constructing a flywheel friction disturbance estimator; d. constructing a flexible appendage vibration disturbance observer; and e. combining the flywheel friction disturbance estimator in the step c and the flexible appendage vibration disturbance observer in the step d with a nominal controller to obtain a compound controller; the compound controller compensating for flywheel friction according to an estimated value of a flywheel friction moment; and the compound controller compensating for flexible appendage vibration disturbance according to an estimated value of flexible appendage vibration disturbance.

The invention provides a method for controlling a flexible structure and a self-adaptive changing structure by a radial basis function (RBF) neural network, belonging to the field of aviation. The method aims at solving the problem that the existing method can not preferably solve the conflict between the shake of a solar sailboard and the high-precision control target of an attitude control system. The method comprises the following steps: an E1 input forming module is used for converting an inputted expected satellite attitude angle theta d into a response uE1, and outputting the response uE1 to a nominal system and a flexible spacecraft; the nominal system is used for outputting expected satellite attitude information xm (t), and the flexible spacecraft is used for outputting practical satellite attitude information x (t) to obtain an error e (t) by comparing the xm (t) with the x (t); a sliding film face control module is used for obtaining a proper sliding film face s according to the error e (t), and transmitting the s to the RBF neural network and a self-adaptive locoregional control module; the self-adaptive locoregional control module is used for outputting a self-adaptive locoregional control u* to the RBF neural network; and the RBF neural network is used for obtaining and adjusting a locoregional control un and an adding result between the un and the uE1 according to the s and the u* to control the satellite attitude of the flexible spacecraft to achieve an expected value.

The invention discloses a composite anti-interference controller comprising measurement and input time delay for a flexible spacecraft and relates to attitude control of the flexible spacecraft under time-varying delay and multi-source interference. The method comprises the following steps of: establishing a dynamical model of the flexible spacecraft; constructing a composite anti-interference controller comprising measurement and input time delay, estimating and performing feed-forward compensation on the anti-interference controller according to an interference observer comprising measurement time delay with interference design caused by vibration of flexible accessories, and designing a state feedback H8 controller to suppress the anti-interference controller according to the norm-bounded interference; designing gain of the interference observer to guarantee the stability of an interference estimated error dead time delay equation according to the 3 / 2 stability theorem; and finally, designing the gain of the state feedback H8 controller for a composite control system comprising measurement and input time delay based on a convex optimization algorithm, so that the system is stabilized, and a certain H8 performance is met. The composite anti-interference controller has the advantages of high interference resistance, convenience in design and the like and can be used for the attitude control of the flexible spacecraft comprising measurement and input time delay.

The invention discloses a flexible spacecraft underactuated system based on the switching control method and an attitude control method of the flexible spacecraft underactuated system based on the switching control method and belongs to the technical field of spacecraft attitude control. According to the flexible spacecraft underactuated system based on the switching control method and the attitude control method of the flexible spacecraft underactuated system based on the switching control method, a dynamical model of a flexible spacecraft with an execution mechanism completely ineffective is established, coupling, caused by elastic vibration of a flexible accessory, with a rigid body is taken as an uncertainty of the system, a dynamical model of a conversion system is a standard nonlinear model, all states are divided into three sliding model surfaces according to the layering sliding model idea, an equivalent control component is designed for the first layer of sliding model surface according to the Filippov equivalent theorem, a switching control law is designed by means of the second layer of sliding model surface and the third layer of sliding model surface, and a final control law is formed by combining equivalent control input and the switching control law. By the adoption of the flexible spacecraft underactuated system based on the switching control method and the attitude control method of the flexible spacecraft underactuated system based on the switching control method, the defect that in the prior art, a flexible spacecraft can not operate normally under the condition that an actuator becomes ineffective completely during operating of the flexible spacecraft is overcome, and operating reliability of a flexible spacecraft attitude control system is improved.

The invention discloses an aircraft thermally-induced vibration dynamics response evaluation method. A flexible aircraft thermally-induced micro-vibration coupling dynamics analysis method is adopted, so that response higher in accuracy can be acquired when compared with a simplified aircraft thermally-induced micro-vibration non-coupling dynamics modeling method which is currently used and composed of a two-dimensional cantilever and a central rigid body, analysis of influence on aircraft posture, especially high-resolution remote sensing satellite imaging resolution, by flexible accessory thermally-induced micro-vibration dynamics environment is better facilitated.

The present invention provides a method for refined attitude control based on output feedback for a flexible spacecraft. The control method comprises the following steps of: a) building a flexible spacecraft dynamical system Σ1, converting the flexible spacecraft dynamical system Σ1 into a flexible spacecraft dynamical system Σ2, and incorporating spacecraft rigid-flexible coupling dynamic disturbance into the flexible spacecraft dynamical system Σ2; b) constructing an external system Σ3, and describing the rigid-flexible coupling dynamic disturbance through the external system Σ3; c) configuring a disturbance observer for estimating the value of the rigid-flexible coupling dynamic disturbance; d) configuring a dynamic output feedback H∞ controller; e) compounding the disturbance observer in step c) with the dynamic output feedback H∞ controller in step d) to obtain a flexible spacecraft refined attitude control system Σ6; the flexible spacecraft refined attitude control system Σ6 compensating for the rigid-flexible coupling dynamic disturbance through the estimated value.

The invention provides a flexible satellite neural network backstepping sliding mode attitude control method, relates to a flexible spacecraft attitude control method, and aims at solving a problem of perturbation generated by plate flexible vibration and antenna rotation and a problem of enhancement of steady state precision and stability of existing attitude control methods. According to the method, a flexible satellite attitude kinetic model is established firstly according to a spacecraft, and then a model formula is processed; a sliding mode attitude controller (which is expressed in the specification) based on a backstepping method is designed; then an RBF neural network approach (eta+htau)sgn(sigma) is adopted; the controller (which is expressed in the specification) is designed; and finally a complete attitude controller (which is expressed in the specification) is obtained. A three-axis attitude controller is designed respectively according to the aforementioned process. The flexible satellite neural network backstepping sliding mode attitude control method is suitable for the field of flexible spacecraft attitude control.

The invention discloses a method for simultaneously realizing attitude maneuver and suppressing vibration of a spacecraft with a flexible accessory, relating to a control method of a spacecraft with a flexible accessory and solving the problems of jet resource waste and vibration suppression incapability of the traditional attitude maneuver control method. The method for simultaneously realizing attitude maneuver and suppressing the vibration comprises the following steps of: (1) measuring the vibration modal information of a flexible spacecraft structure; (2) generating four types of jet switching sequences; (3) setting a minimum action time of the switching sequences; (4) acquiring an attitude angle and the angular velocity of the spacecraft; (5) setting an expected angle value of the attitude maneuver and generating a continuous control moment quantity of the attitude maneuver of the spacecraft according to the attitude angle and the angular velocity; (6) executing the four types of jet switching sequences according to a jet control logic and generating a non-linear switching command acted on a jet thrustor of the spacecraft; and (7) realizing the attitude maneuver of the spacecraft with the flexible accessory to the expected angle value. The invention is suitable for the field of the control of spacecrafts with flexible accessories.

The invention relates to a strong interference resistance control method of a flexible spacecraft attitude control system, targeting a flexible spacecraft attitude control system having a nonlinear dynamic state, flexible vibration and measurement noise multi-source interference. The strong interference resistance control method of the flexible spacecraft attitude control system comprises steps of establishing a multi-source interference resistance model according to a nonlinear euler angle kinematics and posture dynamics of the flexible spacecraft to finish multi-source interference mathematic representation and modeling, designing an interference observer on the basis of mathematic representation and modeling to perform estimation on flexible vibration, designing an extended state observer to perform estimation on a system state and a non-linear dynamic state according to output of an interference observer, designing a strong interference resistance controller according to the interference observer and the extended state observer and solving an unknown gain on the basis of pole configuration to finish design of a strong interference resistance controller according to estimation values of the interference observer and the extended state observer. The strong interference resistance control method of the flexible spacecraft attitude control system is strong in interference resistance, high in control accuracy and can be applied to high accuracy control of the flexible spacecraft.

The invention discloses a wind tunnel test data correction method for a near-space high-aspect ratio flexible spacecraft. The wind tunnel test data correction method comprises a basic state data correction of wind tunnel test, a horizontal tail installation angle adjustment and a data correction of wind tunnel test pilot. The wind tunnel test data correction method combines the prior test conditions for wind tunnel tests in China and completes the correction of the wind tunnel test data of rigid body shape by means of aeroelastic and fluid mechanics analysis. The wind tunnel test data correction method can provide accurate aerodynamic data for near-space high-aspect ratio flexible spacecraft control rate design and flight simulation and ensure that the aircraft can stably flight under the condition of large deformation near the space.

The invention discloses a flexible spacecraft active fault tolerant control method based on dynamic output feedback control, which belongs to the field of spacecraft attitude control. The method comprises the steps that a dynamic equation of a flexible spacecraft attitude control system is transformed into a general state space equation; a fault model with additive sensor measurement offset is established; a fault detection and identification module composed of an unknown input observer and a filter is established; real-time detection and on-line estimation are carried out on unknown sensor fault; and a fault tolerant controller based on dynamic output feedback is designed based on acquired fault estimation information. According to the invention, when a flexible spacecraft can normally reach an expected posture in the case of an additive sensor measurement offset fault; the influence of disturbance generated by modeling uncertainty and flexible accessory is considered in a design process; and a fault diagnosis and identification module and the fault tolerant controller can be independently designed, and can be easily realized by a project.

The present invention discloses a flexible spacecraft sensor fault adjustment method, belonging to the field of aerospace flight control technology. A fault estimation module and a fault tolerance controller are mainly employed by the flexible spacecraft sensor fault adjustment method, and the method comprises the steps that: a filter and an adaptive observer are employed to form a fault estimation module; fault estimation information and sensor output signals are employed to establish the fault tolerance controller through combination of adaptive integral sliding mode control technology; through adoption of the Lyapunov stability theory, the asymptotic stability of a system is proved in a condition of generation of sensor faults; and finally, a simulation experiment result proves the validity of the method provided by the invention. The fault tolerance control problem when sensor measurement offset faults are generated in an operation process of a flexible spacecraft is solved so thataccurate control of an attitude angle in the sensor fault condition is achieved and the system has a high tolerance capability for the faults; and moreover, the method fully takes the consideration of disturbances generated by spacecraft parameter uncertainty and flexible accessories in the design process and is suitable for engineering application.

The invention relates to a flexible spacecraft attitude control method for flywheel saturation and friction characteristics and aims to solve a problem of influence of flywheel saturation and friction characteristics on spacecraft control precision. According to the method, a spacecraft coupling kinetic equation containing flywheel saturation and friction characteristics is firstly constructed; secondly, a friction interference estimator according to flywheel friction characteristics is established; thirdly, a flexible vibration observer is designed according to interference caused by flexible accessory vibration; fourthly, an anti-saturation controller is designed for inhibition according to friction interference estimation errors and flexible vibration interference observation errors; and lastly, through solving control gain of the anti-saturation controller, the friction interference estimator and the flexible vibration observer, a composite layered anti-interference controller is designed, and spacecraft anti-interference attitude control under the action of multi-source interference influence is accomplished. The method is advantaged in that attitude control precision of spacecrafts employing a flywheel execution mechanism can be remarkably improved, and the method can be applied to high precision attitude control on spacecrafts such as high precision earth observation satellites in the aerospace field and space telescopes.

The present invention provides a rigid-flexible liquid coupling system attitude controller and maneuvering path combination optimization method. The method comprises the following steps of: establishing a dynamical model of a liquid-filled flexible spacecraft; obtaining an angular acceleration curve, an angular speed curve and obtained multi-section angular position curve of the liquid-filled flexible spacecraft, and planning the attitude maneuvering path of the liquid-filled flexible spacecraft; calculating the expression of each section of the curve in the multi-section curve; employing a PDcontrol system to perform attitude control of the liquid-filled flexible spacecraft; and performing combined optimization of the parameters of the liquid-filled flexible spacecraft controller and themaneuvering path. The method reduces the motivation of the attitude maneuvering for the flexible accessory vibration and liquid shaking so as to achieve large-angle rapid maneuvering and rapid and stable control for the liquid-filled flexible spacecraft.

The invention provides a dynamic sliding mode attitude tracking control method and system for a flexible spacecraft. The method comprises the steps of S1, establishing error posture quaternion-based kinematic equation and kinetic equation of the flexible spacecraft; and S2, by introducing a dynamic switching function, designing a dynamic sliding mode attitude tracking control law of the attitude tracking problem of the flexible spacecraft, and designing a finite time convergence-based robust differentiator to estimate a partial state of a flexible spacecraft system. The dynamic sliding mode attitude tracking control method and system has the beneficial effects that the problem of buffeting caused by a traditional sliding mode control law symbol function is effectively inhibited through thedesign of the switching function; and the dynamic sliding mode attitude tracking control method can enable the spacecraft system to carry out attitude tracking.

The invention discloses a model-free self-adaptive control method for an attitude of an assembled spacecraft. The method comprises the following steps: S1) establishing a full-format dynamic equivalent linear data model for an attitude movement of a general assembled spacecraft; S2) designing a parameter self-adaptive estimating equation and estimating a pseudo gradient for the pseudo gradient in the full-format dynamic equivalent linear data model; and S3) establishing an error criterion function according to the dynamic equivalent linear data model, and designing a weighting one-step forward controller through a minimizing error criterion function. The dynamic equivalent linear model is established for a general discrete time nonlinear system, and a modified projection algorithm is adopted for performing online parameter estimation for the unknown variable in the dynamic equivalent linear model, so that the attitude determination accuracy of a flexible spacecraft movement simulator is increased.

The invention relates to a flexible spacecraft ground control test platform and a test method, which relate to the field of aerospace, and solve the problem that a realistic simulated test can be performed on a ground control system difficultly by the existing flexible spacecraft ground control test platform. The test platform comprises an air flotation rotary table, a rigid main body, a flexible accessory, equipment above the rotary table and equipment under the rotary table, wherein the equipment under the rotary table comprises a ground control host, a ground database, a rotary table motion measuring device and a set of wireless communication devices; and the equipment above the rotary table comprises a sensor and an xPC main controller. The test method is implemented on the basis of the test platform. The flexible spacecraft ground control test platform and the test method lay a foundation for developing the ground control system of a flexible spacecraft.

The invention discloses a super-size flexible spacecraft dispersion cooperative control method, and aims to achieve high stability and vibration inhibition control on a super-size flexible spacecraft. The method comprises the following steps: S1, dividing a super-size flexible spacecraft control system into a spacecraft posture control subsystem and a flexible attachment vibration subsystem, and designing corresponding partial robust controllers for the spacecraft posture control subsystem and the flexible attachment vibration subsystem respectively; and S2, designing a coordinate controller for overall properties of the super-size flexible spacecraft control system. The method has the advantages that according to the dynamic characteristics of the super-size flexible spacecraft, a dispersion cooperative control method is adopted, a partial structure is stably controlled through dispersion stability, overall high-precision performance indexes are achieved through the cooperative controller, high appointing precision and stability of postures and deformation control precision of a flexible component are achieved, and the method is widely applied to high-precision high-stability appointing control on a large-size flexible structure.

The invention relates to a path planning based flexible spacecraft attitude adaptive control method. First softening operation on an expected angle of a spacecraft is performed based on an SMPPa attitude maneuvering path planning method; on-line parameter identification is performed by utilizing a gradient descent method based on a flexible spacecraft mathematical model based on the feature modeltheory, and a three-axis model feature parameter is determined; a control moment is determined according to the determined flexible spacecraft mathematical model and the feature parameter thereof, theflexible spacecraft attitude is controlled through the control moment, and finally the steps are repeated until the angle reaches the expected spacecraft angle. According to the path planning based flexible spacecraft attitude adaptive control method, the influence of flexible accessory vibration on control performances is restrained, attitude control steady accuracy and dynamic characteristics are improved, and the path planning based flexible spacecraft attitude adaptive control method is suitable for the flexible spacecraft having three-axis coupling nonlinear characteristics.

An embodiment of the invention provides an attitude control method of a flexible spacecraft. The attitude control method comprises the steps of: regarding a vector portion of an error attitude quaternion at a current moment as variable parameters of a variable gain controller, and combining a control gain restriction condition to obtain control gain at the current moment; regarding the error attitude quaternion, an error attitude angular speed and a control moment at the current moment as input parameters of a dimensionality reduction observer, and acquiring estimated values of modal coordinates of a flexible attachment at the current moment and estimated values of parameter variables; and regarding the vector portion of the error attitude quaternion at the current moment, the control gain, an error attitude angular speed, the estimated values of the modal coordinates and the estimated values of the parameter variables as input of the variable gain controller, and outputting a controlmoment at a next moment. The attitude control method enables a system which is externally disturbed to complete high-precision attitude maneuvering tasks, and achieves the purpose of limiting the control gain through introducing additional linear matrix inequality constraints.