315 results about "Spacecraft attitude control" patented technology

The invention relates to a spacecraft attitude integral sliding mode fault tolerance control method taking consideration of a performer fault and provides a robustness attitude active fault tolerance control method based on an integral sliding mode surface for problems of the performer fault, external disturbance and control moment amplitude limits in a spacecraft attitude control process. The method comprises steps that firstly, a spacecraft attitude dynamics model taking consideration of the performer fault and containing external disturbance is established; secondly, on the condition that a performer is not in fault, a designed nominal controller can guarantee system stability, and input saturation amplitude limits can be easily satisfied through adjusting controller parameters; lastly, the fault information is introduced to design an integral sliding mode controller, robustness of external disturbance and the performer fault can be effectively improved, system stability is analyzed on the basis of an Lyapunov method. The method is advantaged in that stability of the attitude control system is guaranteed when a spacecraft operating on orbit generates the performer fault, and relatively strong fault tolerance capability and external disturbance robustness are realized.

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 present invention provides a method for disturbance compensation based on a sliding mode disturbance observer for a spacecraft with a large flexible appendage, comprising steps of: a) building a spacecraft attitude control system; b) constructing an external system, the external system being incorporated with an uncertain portion of a damping matrix of a flexible appendage of the spacecraft; the external system being incorporated with an uncertain portion of a rigidity matrix of the flexible appendage of the spacecraft and describing a sum of flexible vibration and environmental disturbance; c) configuring a sliding mode disturbance observer for estimating the value of the sum of flexible vibration and environmental disturbance; d) compounding a nominal controller with the sliding mode disturbance observer in step c) to obtain a compound controller; the compound controller compensating for the sum of flexible vibration and environmental disturbance.

A modular control moment gyroscope (CMG) system for a spacecraft attitude control system (ACS) is formed by a plurality of CMG modules, wherein each CMG module has a modular enclosure design that is identical to that of the other CMG modules, such that the plurality of CMG modules are mountable in a spacecraft array bus structure in any desired one of multiple array configurations.

Provided are a fault detector and a fault detection method for an attitude control system (ACS) of a spacecraft. The fault detector includes a first interacting multiple model (IMM) fault detection block for generating a normal model filter of the plurality of actuators and a plurality of upper level filters including fault model filters corresponding to the respective actuators, and detecting faults of the plurality of actuators using an IMM estimation technique from the plurality of upper level filters, and a second IMM fault detection block for generating a plurality of lower level filters each including a fault type model filter of the fault-detected actuator in the first IMM fault detection block, and detecting a fault type of the failed actuator using the IMM estimation technique.

Provided is a fault tolerant control method for a spacecraft equipped with a flywheel based on robust adaptation. The method comprises five steps of: 1, establishing a spacecraft kinematical equation; 2, establishing a spacecraft kinetic equation; 3, performing desired angular velocity locus design; 4, performing controller output torque design; and 5, performing numerical simulation. Under the circumstance that a FDD device is not used, the method online estimates a triaxial torque fault factor and designs a robust adaptive fault tolerant controller. The research aims at enriching spacecraft PFTCS methods and providing technical support for spacecraft attitude control in the future.

The invention relates to a spacecraft fault tolerant control method based on an angular speed observer. The method comprises a step of considering an actuator fault and external disturbance borne by a spacecraft, and establishing a spacecraft attitude control system dynamics and kinematics model, a step of designing a finite time observer to estimate the attitude angular velocity information of a spacecraft, and a step of constructing a robust controller based on the obtained angular velocity information, and at the same time considering control amount input saturation and designing an auxiliary system such that the spacecraft has robustness for the actuator fault without angular velocity information. According to the method, the stability of an attitude control system is ensured when an on-orbit working spacecraft has the actuator fault and even the attitude angular velocity information is unknown, and the method has high control precision, high fault tolerance ability and robustness to external disturbance.

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 invention relates to a rapid stable joint control method for attitude maneuver of spacecraft and belongs to the field of spacecraft attitude control and vibration control. The method is invented by absorbing advantages in spacecraft attitude maneuver trajectory planing technology and input shaping technology, and allows spacecraft to maneuver in task-required time and keep attitude quickly stable within index required values after maneuvering. Torque outputting capacity of a spacecraft attitude control actuator and maximum angular velocity maneuverability of spacecraft are further considered comprehensively, so that output torque of the spacecraft attitude control actuator is easy to implement.

The invention discloses a spacecraft attitude control semi-physical simulation system which relates to a spacecraft control semi-physical simulation system. The invention solves the problems of poor universality and systematicness in the prior spacecraft attitude control semi-physical simulation system. The simulation signal input or output end of a real-time simulation machine of the system is connected with the simulation signal output or input end of an onboard spacecraft computer; the wireless signal output or output end of a system wireless communication module is connected with the wireless signal input or output end of the real-time simulation machine; the signal output end of a ground wireless communication module is connected with the signal input end of a data storage module; the data signal output end 1 of the data storage module is connected with the data signal input end of a flight simulation display module; and the data signal output end 2 of the data storage module is connected with the data signal input end of an animation and graphic display module. The invention is suitable for simulation of spacecraft attitude control.

The invention relates to an integrated test simulation system for a spacecraft attitude control algorithm and belongs to the computer simulation technology and application technical field. The integrated test simulation system includes a controller, a measurement model, an actuator model, a spacecraft attitude dynamic model, an antenna model, a sailboard model and a load model, wherein the controller includes a mode management module, a path optimization module, a closed loop control module and an accessory monitoring module. According to the system of the invention, the components of the system are configured based on a universal satellite platform; modeling is performed on system equipment related to typical attitude control and dynamic characteristics; and a simulation on-orbit simulation environment supporting modularized applications can be provided for the control algorithm; near-earth orbit satellite mission modes are preset in the mode management module, and a user can select one kind of near-earth orbit satellite mission mode to fast establish a simulation scene, and test data is of practical significance; the path optimization module provides a function for planning paths of spacecraft attitude maneuver, and therefore, the performance of a closed-loop control mechanism can be improved.

The invention discloses a rigid spacecraft performer multi-fault diagnosis and fault tolerance control method. According to the method, the kinetic and dynamic model of a rigid spacecraft attitude control system is put forward, a fault model for existing of both the performance failure fault and the deviation fault of a rigid spacecraft is established, and then a fault detection observer adopting the adaptive threshold technology and a fault estimation observer based on the adaptive technology are respectively established so that online real-time detection and estimation of the fault time and the concrete situation of the fault can be realized, and finally a backstepping sliding mode fault tolerance controller is designed according to fault information established by the fault estimation observer. Attitude stabilization control of the rigid spacecraft under the condition of the performer efficiency damage and the deviation fault can be realized, and the influence of external disturbance on the system and the observers can also be considered in the design process. Besides, the fault detection observer and the fault estimation observer can be independently designed so that the engineering application can be more easily implemented.

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 provides a method for performance evaluation of a spacecraft attitude control ground simulation system. According to the method, the credibility of a dynamic system, a kinematics system, a measurement system and a control system is studied based on the simulation system credibility theory, the overall credibility of the spacecraft attitude control ground simulation system is obtained through comprehensive calculation, and then the performance of the spacecraft attitude control ground simulation system is evaluated. According to the evaluation method, the credibility of the ground simulation system and an actual satellite attitude control system is considered from all aspects, simulation data are fully utilized, and the credibility of the ground simulation system is described objectively and quantitatively; the theoretical basis is reliable, the calculation process is simple, the application range is wide, and a new idea and reference are provided for performance evaluation study of other simulation systems.

The invention discloses a semi-physical simulation system for attitude control of a star-arrow integrated spacecraft, which relates to the semi-physical simulation system for control of the star-arrow integrated spacecraft. The invention solves the problem that the conventional semi-physical simulation system for the attitude control of the spacecraft cannot simulate the spacecraft from a launching point to an active operation stage attitude control system. A simulation signal input or output end of a real-time simulator is connected with a simulation signal output or input end of an onboard computer; a wireless signal input or output end of a stage wireless communication module is connected with a wireless signal output or input end of the real-time stimulator; and a signal output or input end of a ground wireless communication module is connected with a signal input or output end of a ground control module. The system is suitable for the stimulation of the attitude control of the star-arrow integrated spacecraft.

The invention provides a testing device and a method applicable for spacecraft high-stability pointing control. The testing device comprises a base, an air bearing, an instrument platform, a spacecraft pointing unit, a spacecraft attitude control unit, a platform data instruction transmitting-receiving system, a mark point, a target measuring head unit, a reflection prism, a counteractive momentum wheel system, an image processing unit, a photoelectric auto-collimator, a laser tracker, an off-platform data instruction transmitting-receiving system and a camera. According to the device and the method, the characteristics of a micro-interference moment environment are simulated by fully using the air bearing; as the high-precision position and attitude measuring techniques of equipment such as the image measuring unit, the laser tracker and the photoelectric auto-collimator are combined, the goal of spacecraft high-stability pointing control test can be realized; the precision is high; the stability is good; and the engineering realization is facilitated. The device and the method can be used for a high-precision pointing control test on spacecraft such as various high-precision weather and military reconnaissance satellites and space-based offense and defense platforms.

The present invention relates to a spacecraft attitude fault-tolerant control method based on an iterative-learning disturbance observer. In order to solve the problem that actuator faults and external disturbances might occur during the attitude control process of the on-orbit operation of a spacecraft in the prior art, the invention provides an active fault-tolerant control method based on the iterative-learning observer. The method comprises the steps of firstly, establishing a dynamic model for the attitude control system of a spacecraft in considering the fault of an actuator in the attitude control system of the spacecraft and the external disturbance in the space of the spacecraft; secondly, estimating and designing a generalized disturbance torque of the iterative-learning disturbance observer composed of the actuator fault information and the external disturbance; finally, based on the estimated generalized interference, designing an active fault-tolerant controller. Based on the method, the stability of the attitude control system of the spacecraft, when the actuator of the spacecraft has a fault and the external disturbance exists in the space of the spacecraft, can be ensured. Meanwhile, the precision requirement of an actual control system is met. The method is strong in fault-tolerant ability and good in robustness for the external disturbance.

The invention relates to a magnetically suspended gyroscope flywheel which can be used as a three-free degree attitude control actuator and a two-free degree attitude sensor of a spacecraft like a satellite and the like, and the magnetically suspended gyroscope flywheel mainly comprises a flywheel rotor, an axial magnetic bearing, a radial magnetic bearing, a protection bearing, a displacement sensor, a seal cover, a mounting shaft, a base, a motor, and other parts. The magnetic bearing is adopted for carrying out five-free degree suspension on the flywheel rotor, the flywheel rotor is driven to rotate around a rotating shaft by the motor, and the axial free degree control torque is outputted; the flywheel rotor is controlled to realize the torsion around an X axis or a Y axis of a housing by the axial magnetic bearing, and the radial two-free degree control torque is outputted; and the two-free degree attitude angular velocity of the spacecraft relative to inertia space can be sensed. The magnetically suspended gyroscope flywheel integrates the three-free degree attitude control actuator and the two-free degree attitude sensor of the spacecraft into a whole, thereby reducing the volume of an attitude control system of the spacecraft, reducing the weight of the attitude control system and being particularly applicable to being applied as the attitude control system of the spacecraft like the satellite and the like.

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 provides an inertial system spacecraft attitude control / angular momentum management method including four steps. The invention aims to solve the problems of control moment gyro angular momentum accumulation caused by gravitation gradient moment and other interference moments in an inertial system, the gravitation gradient moment is adopted to balance the gestures, and a space station angular momentum management controller based on pole assignment is designed. A space station linear model is established under the inertial system, the infeasibility of the inertial system angular momentum management in the pitch axis direction is analyzed, the pitch axis is decoupled from a rolling / yaw axis, and the CMG angular momentum in the pitch axis direction is not restrained, constant disturbance, disturbance being one-time of the track frequency, and disturbance being twice of the track frequency are brought into a state equation to suppress the influence to the pitch axis gestures, and a linear quadratic algorithm based on pole assignment is adopted to solve a feedback gain matrix. The algorithm prevents the selection of cost matrix Q, and based on the requirement of the system performance, the closed-loop poles can be configured to a specified area at the left side of a complex plane imaginary axis, and at the end, the feasibility of the algorithm is verified by the simulation results.

The invention discloses a method for accurately and quickly manipulating a variable speed control moment spinning top group. According to the method, the condition that the rotor rotating speed is changeable of the VSCMG is fully utilized. According to the moment needed by an attitude control task, the operation modes of the VSCMG are divided into the CMG / MW mixed operation mode and the independent MW operation mode. Through reasonable switching of the two operation modes, the large-moment requirements for tasks such as attitude maneuver control of the VSCMG can be satisfied through one VSCMG and the high-precision small-moment requirement for tasks such as attitude stable control can be satisfied through one VSCMG under the condition that an additional actuating mechanism does not need to be configured, so that the problem which cannot be solved when a single SGCMG or a single MW group is used for conducting spacecraft attitude control is well solved.

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 relates to a parameter selection method adopting an onboard control moment gyroscope group vibration-isolating platform and belongs to the field of the attitude control and the vibration and shaking control of a spacecraft. (1) an idea of the cube vibration-isolating platform is adopted, the cube vibration-isolating platform is arranged between a configuration consisting of a plurality of control moment gyroscopes and a satellite body and the high frequency vibration caused by a plurality of control moment gyroscopes is isolated to the greatest extent; and (2) by reasonably selecting parameters of the onboard control moment gyroscope group vibration-isolating platform, the problem of effectiveness loss of other onboard systems, which is caused after the vibration-isolating platform is applied on a satellite, is avoided.

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 provides a spacecraft formation attitude limited time cooperation control method under a complicated constraint condition. The spacecraft formation attitude limited time cooperation control method comprises steps that a formation topological structure is determined, and the current attitude value of the spacecraft formation is acquired; a spacecraft attitude control equation is converted into a system error tracking equation; a system state observer is invoked to estimate the angular speeds and the attitude values of the various members of the spacecraft formation, and then angular speed estimation values and attitude estimation values are acquired; the angular speed estimation values and the attitude estimation values of the members of the adjacent formation are coupled to acquire control slip form variables; the slip form variables are transmitted to a controller, and the controller is used to convert the control slip form variables to control signals according to a preset slip form control algorithm. By adopting the method provided by the invention, the accurate control over the spacecraft formation attitude under the complicated constraint condition is realized.

The invention relates to a multi-stage attitude control method for a remote sensing micro-nano satellite, and belongs to the technical field of spacecraft attitude control. A three-axis magnetic moment device and a momentum wheel set are adopted as an attitude control execution component, and the attitude control method for the four main stages of recemization after satellite-rocket separation, counterglow capture, counterglow orientation and counter-earth remote sensing is designed, and an attitude mobile control method for the two stages of counterglow orientation and counter-earth remote sensing is designed. According to the recemization stage, the large angular rate of the satellite obtained after satellite-rocket separation is eliminated, and the basic controlled attitude is built; inthe counterglow capture and counterglow orientation stages, it is ensured that a sun sailboard is in the charging state for a long time; in the counter-earth remote sensing stage, it is ensured thatthe remote sensing photographing task is completed. According to the method, traditional PD control is adopted as the foundation, on the basis of feedforward control, quaternion is used for representing attitude errors, the method has the advantages of being rapid in satellite maneuvering response, high in steady state robustness and high in control precision, and the task requirement for a remotesensing micro-nano satellite is met.

The invention which discloses a precision compensation method for an area APS digital sun sensor belongs to the spacecraft attitude control measure system field. The precision compensation method comprises the small-angle incidence angle precision compensation and the wide-angle incidence angle precision compensation. The precision compensation method comprises the following steps: 1, carrying out primary rotation correction on the solar faculae center coordinate value which is obtained through the calculation of a sun sensor image processing chip and is input by the sun sensor image processing chip; 2, carrying out secondary correction on the coordinate value of the wide-angle incidence angle, and utilizing the rotarily-corrected coordinate value to do the coordinate mapping of a refraction model; 3, calculating through an uniaxial higher-order polynomial fitting process to obtain the solar angle tangent value, utilizing the rotarily-corrected coordinate value to calculate if the incidence angle is a small angle, and utilizing the secondarily-corrected coordinate value to calculate if the incidence angle is a large angle; and 4, finding the arc tangent to obtain the sun angle value. The precision compensation method which is suitable for the precision compensation of the area APS array digital large-field sun sensor has the advantages of major error factor compensation and high precision.

A momentum-control system comprising a plurality of momentum actuators and a platform upon which the plurality of momentum actuators are mounted. The momentum control system further comprises a plurality of active struts mounted on the bottom side of the platform. The active struts are configured to produce a force to steer the plurality of momentum actuators and the platform to produce forces and moments for spacecraft attitude control and disturbance suppression.

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.