28 results about "Flexible spacecraft" patented technology

The invention belongs to the technical field of spacecraft control, and discloses a flexible spacecraft attitude control method and related equipment based on an all-drive system theory, and the method comprises the steps: building an attitude tracking error model containing flexible vibration, external interference and rotational inertia uncertainty, the model is converted into a second-order all-wheel-drive system model with a lumped disturbance term; converting the second-order all-drive system model into a linear steady-state closed-loop system containing a lumped disturbance term, and constructing a preset time sliding mode control law capable of enabling the state of the linear steady-state closed-loop system to converge within a preset time; constructing an RBF disturbance observer for performing approximation estimation on the lumped disturbance item; according to the second-order all-drive system model, the preset time sliding mode control law and the RBF disturbance observer, an RBF preset time pseudo-linear sliding mode controller is constructed; the flexible spacecraft is controlled by the pseudo-linear sliding mode controller according to the RBF preset time; therefore, the method has the advantages of realizing preset time convergence and improving attitude control precision and stability.

A method for composite attitude control of a spacecraft with nonlinear sloshing and large flexible appendages disclosed by the present invention belongs to the technical field of spacecraft control. The implementation method of the present invention is as follows: various non-conservative disturbance torques acting on the spacecraft during on-orbit operation are characterized by Te; the attitude motion of the spacecraft body at any angle is described by Euler quaternions; the elastic vibration of 1 to m flexible appendages is described by the assumed mode discretization method; the motion description of the equivalent spherical pendulum model of liquid sloshing is improved by using the split variable technique; a large-range motion dynamics model of a liquid-filled spacecraft with large flexible appendages is established; the spacecraft attitude dynamics model is linearized; an output feedback attitude control law for a liquid-filled flexible spacecraft designed according to the Lyaponuv stability theorem; an input shaping-output feedback composite attitude controller for a liquid-filled flexible spacecraft designed based on the output feedback attitude control law of the liquid-filled flexible spacecraft and the Euler quaternion input command, thereby realizing active vibration suppression and improving the accuracy of spacecraft composite attitude control.

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 relates to the field of spacecraft attitude control, in particular to a load optical axis pointing control system and method for a truss type flexible spacecraft, and the system comprises a truss type flexible spacecraft structure platform, an attitude control module and a load optical axis pointing control module. The attitude control module is used for central star attitude angle fine adjustment and preliminary adjustment of optical load primary mirror optical axis direction. The load optical axis direction control module is used for optical axis direction fine adjustment and micro-vibration suppression. Two independent closed loops of central star attitude angle closed-loop control and load optical axis pointing closed-loop control are constructed, input signal sources, control objects and execution mechanisms of the two control loops are completely decoupled, and the problem of control effect divergence caused by mutual interference in traditional coupling control is solved in principle.

The application discloses an anti-unwinding flexible spacecraft attitude tracking composite control method, first, based on the quaternion method, a flexible spacecraft attitude tracking error kinematics and dynamics model is established; then, the flexible accessory vibration mode, external environmental disturbance and model uncertainty are regarded as a lumped disturbance, a generalized proportional integral observer is designed, and the lumped disturbance is observed and estimated; finally, a non-singular terminal sliding mode surface is constructed, the lumped disturbance observation and estimation compensation of the generalized proportional integral observer is taken as a feedforward signal, and a feedforward-feedback anti-unwinding attitude tracking composite controller is designed. Through determining appropriate control gain, the closed-loop attitude tracking error system is globally finite time stable, and the anti-unwinding effect can be realized. The application can realize finite time attitude tracking of the flexible spacecraft, has high attitude tracking precision and strong robustness.

The application belongs to the technical field of satellite attitude control, and relates to a large flexible spacecraft distributed fault-tolerant attitude and vibration control method, a large flexible spacecraft rigid-flexible coupling dynamics model with distributed actuators is established by using the Lagrange method, a hierarchical Stackelberg game control framework is established by taking an attitude control subsystem as a leader layer and a vibration suppression subsystem and a fault-tolerant control subsystem as follower layers, performance functions of the three subsystems are defined, and through game of spacecraft distributed fault-tolerant attitude and vibration control, strategy cooperation and dynamic balance between at least two control targets are realized; the application realizes three-axis attitude stabilization and low-frequency vibration suppression of a target spacecraft, and shows stronger fault-tolerant capability under node failure.

The invention discloses a high-precision attitude maneuver control method for a flexible spacecraft, and the method comprises the steps: considering the fault and saturation conditions of an execution mechanism for a large-inertia spacecraft containing a flexible accessory, firstly building a rigid-flexible coupling dynamic model of the spacecraft, and obtaining the first three-order modal frequency and damping ratio of the spacecraft; secondly, designing a fast robust input shaper according to modal and damping ratio information, and taking a result obtained by convolution of the fast robust input shaper and reference input as control input; and then considering uncertainty such as inertia and external interference of the spacecraft, designing a self-adaptive sliding mode controller, and approaching the uncertainty of the system by adopting a Chebyshev inequality, thereby realizing high-precision attitude control of the spacecraft in finite time. According to the method, the stability of the spacecraft in the attitude maneuver process can be improved, meanwhile, vibration of the flexible component is restrained, and the control performance of the system is improved.

The present invention discloses a full - drive attitude saturation control method for a flexible spacecraft with input constraints, which includes the following steps: Step 1, based on the full - drive system theory method, according to the attitude dynamics and kinematics models of the flexible spacecraft, convert the spacecraft attitude model into a full - drive system control model; Step 2, design an attitude controller architecture for the full - drive system control model of the flexible spacecraft in Step 1; Step 3, for the linear feedback part of the control law in Step 2, the parameter matrices A0 and A1 are determined by the "direct parameter method" under the full - drive system theory framework; Step 4, for the non - linear term compensation part of the control law in Step 2, an extended state observer is used for comprehensive observation and estimation; Step 5, for the control law in Step 2, considering the physical limitations of the control actuators, design a saturation function to ensure that the control input torque is within a reasonable range. The present invention can achieve rigid - body attitude and flexible vibration suppression within the capabilities of the actuators.

The application discloses a flexible spacecraft attitude control method and system, and the method comprises the following steps: establishing a flexible spacecraft attitude control model and a flexible accessory model; according to the homogeneous system theory, the flexible spacecraft attitude control model and the flexible accessory model are simplified for multiple times to obtain a simplified flexible spacecraft attitude control model and a simplified flexible accessory model; according to the coupling relationship between the simplified flexible spacecraft attitude control model and the simplified flexible accessory model, a full drive model is obtained; the output equation in the flexible accessory model is combined to design an observer, the modal variable in the flexible accessory model is estimated through the observer, the full drive model is combined, and a controller is designed based on a high-order full drive system method; the controller is substituted into the full drive model to obtain a linear closed-loop system; according to a parameterized design method, a parameterized controller is obtained, and a stable linear closed-loop system is obtained. The application can stabilize the flexible spacecraft attitude control system.

The present invention discloses a collaborative attitude control and distributed vibration suppression method for ultra-large-scale flexible spacecraft. The method includes the following steps: Step 1, construct a rigid-flexible coupling dynamics model of an ultra-large-scale spacecraft with multiple flexible structures; Step 2, decompose the coupling dynamics model based on singular perturbation theory to obtain slow-varying attitude and fast-varying vibration subsystems respectively; Step 3, use an adaptive switching and input-saturated terminal sliding mode controller to complete high-precision and high-stability attitude control; Step 4, use distributed piezoelectric actuators based on consensus theory and feedback collaborative controllers to complete high-precision and fast vibration suppression. The beneficial effects of the present invention are as follows: precise dimensionality reduction of the ultra-large-scale flexible spacecraft model is achieved; attitude and vibration active controllers are designed relatively independently; the accuracy and stability of attitude control and vibration suppression of the ultra-large-scale flexible spacecraft are improved simultaneously in a shorter time.

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.

The invention belongs to the technical field of satellite attitude control, and relates to a distributed fault-tolerant attitude and vibration control method for a large flexible spacecraft, which comprises the following steps of: establishing a rigid-flexible coupling dynamic model of the large flexible spacecraft with a distributed actuator by adopting a Lagrange method, and taking an attitude control subsystem as a leader layer; the vibration suppression subsystem and the fault-tolerant control subsystem establish a layered Stackelberg game control framework for the follower layer, meanwhile, the performance functions of the three subsystems are defined, and strategy collaboration and dynamic equilibrium between at least two control targets are achieved through the game of spacecraft distributed fault-tolerant attitude and vibration control; according to the method, three-axis attitude stability and low-frequency vibration suppression of the target spacecraft are realized, and higher fault-tolerant capability is shown under the condition of node failure.

The invention provides a zero-gravity unloading method for a ground test of a load platform integrated flexible spacecraft, and the method comprises the steps: arranging gravity unloading mechanisms, such as an air floatation mechanism and a suspension mechanism, at an optical load primary mirror, an optical load secondary mirror, a central star, a primary mirror section truss, a central star truss and the like of the load platform integrated flexible spacecraft; therefore, the gravity of the spacecraft in a ground test environment is balanced. Compared with the prior art, the method has the advantages that a feasible technical scheme for unloading of a load platform integrated flexible spacecraft ground test is provided, structural deformation of the flexible spacecraft can be effectively controlled, internal force of a spacecraft truss structure in a ground gravity environment is reduced, normal actuation of various mechanisms installed on the truss is ensured, and the reliability of the flexible spacecraft is improved. And normal development of various on-orbit working condition verification of the ground test is ensured.

The present invention discloses a hybrid power dynamics modeling method and system based on a long short-term memory neural network, which relates to the field of modeling technology, constructs a hybrid power dynamics model of a flexible spacecraft; adjusts the physical parameters of the dynamics model within a set threshold range to obtain a plurality of adjusted dynamics models; performs programming calculations on the hybrid power dynamics model and the plurality of adjusted dynamics models to obtain a reference parameter matrix and a plurality of difference parameter matrices; subtracts each difference parameter matrix from the reference parameter matrix to obtain a difference value set; trains a long short-term memory neural network based on each difference parameter matrix and each difference set to obtain a trained long short-term memory neural network; predicts the actual modeling error based on the trained long short-term memory neural network, and then offsets the uncertainty error, improves the modeling accuracy, and provides a good foundation for subsequent vibration control.

The multi-node distributed flexible spacecraft configuration and intelligent attachment trajectory planning method disclosed by the application belong to the field of spacecraft guidance and control. The spacecraft configuration has the advantages of light weight, high energy absorption efficiency, avoidance of contact collision between rigid structure and spacecraft surface, facilitation of multi-point distributed detection, and modularization and rapid assembly, and can avoid rebound escape during spacecraft attachment. The multi-node distributed flexible spacecraft intelligent attachment trajectory planning method adopts a linear spring model to describe the flexible connection constraint force generated by the deformation of the flexible air bag; a multi-node distributed flexible spacecraft cooperative attachment trajectory planning model under multiple constraint conditions is constructed, and is converted into a standard quadratic programming form to reduce the complexity of the planning problem; an original dual neural network algorithm is adopted to construct a trajectory planning solution framework with state coupling at each control node, and the attachment trajectory of each control node is solved, and the safety and stability of the overall spacecraft attachment are improved through distributed cooperation.

The present invention relates to an on-orbit estimation method for spacecraft mass characteristic parameters based on thrusters, including: spacecraft actuator configuration, establishment of a dynamic model, establishment of a kinematic model, establishment of an attitude measurement component model, on-orbit estimation of the three-axis moment of inertia based on thrusters, on-orbit estimation of the center of mass in the steady state mode and the orbital maneuver mode, and on-orbit estimation of the spacecraft mass. The present invention solves the problems of on-orbit estimation of the spacecraft moment of inertia, center of mass, and mass, and lays a solid technical foundation for the on-orbit redesign or correction of attitude and orbit control algorithms that rely on mass characteristic parameters. Through simulation analysis, this method is reasonable and effective, with a simple algorithm, small computational load, low fuel consumption during implementation, and is easy to implement in engineering. At the same time, a flexible model is added to the simulation of this method, so it is also applicable to flexible spacecraft.

This invention provides a method and apparatus for analyzing the coupling characteristics of asteroid impactors based on SPH (Smooth Particle Hydrodynamics), belonging to the field of spacecraft dynamics modeling. The method includes: establishing a liquid sloshing model described by the SPH method; SPH represents smooth particle hydrodynamics; calculating the liquid sloshing forces and moments in a non-inertial frame based on the liquid sloshing model; establishing a distributed parameter model of the flexible spacecraft; integrating the liquid sloshing forces and moments in the non-inertial frame into the distributed parameter model of the flexible spacecraft to establish a dynamic model of the liquid-filled flexible spacecraft; and, based on the dynamic model of the liquid-filled flexible spacecraft, analyzing the coupling characteristics of the central rigid body attitude and position, flexible appendage vibration, and liquid fuel sloshing as system states. This invention provides a new approach and reference for modeling liquid-filled flexible spacecraft.

The invention discloses an asymmetric flexible spacecraft equivalent dynamic modeling method considering thermal deformation. The method comprises the following steps: establishing an equivalent beam model for a space truss structure, and establishing a coordinate system; describing the displacement of any point on the periodic units of the truss structure; analyzing stress, strain and structural volume of the periodic unit along the neutral axis direction to obtain thermal deformation strain energy and axial strain energy of the periodic unit after thermal deformation; based on the total strain energy decomposition, obtaining an elastic matrix and a stress vector; superposing the kinetic energy of each rod piece in the periodic unit to obtain the total kinetic energy of the periodic unit, and sorting and decomposing an inertia matrix of the periodic unit; based on a finite element method, obtaining a mass element matrix, a linear stiffness element matrix and a geometric stiffness element matrix of each periodic element; and assembling to obtain a corresponding total mass matrix, a total linear stiffness matrix and an initial stress stiffness matrix so as to obtain an equivalent beam damping-free vibration equation. According to the method, the existing equivalent continuum modeling theory is perfected, and the precision of the equivalent model is improved.

The invention discloses a flexible spacecraft control system space-ground consistency test device and test method based on uncertainty quantization, and belongs to the technical field of intelligent manufacturing equipment industry and space-ground consistency test. In order to solve the problem of space-ground consistency inspection and evaluation of the flexible spacecraft, the space-ground consistency inspection and evaluation device comprises an air flotation robot and a flexible part, the air flotation robot is used for simulating a central rigid body of the flexible spacecraft, and the flexible part is used for simulating an unfolded solar panel; one end of the flexible part is connected to an upper-layer supporting column of the air flotation robot, the other end of the flexible part is free and is additionally provided with a weight, fiber grating sensors are distributed on the flexible part to measure flexible vibration, and a piezoelectric ceramic piece is installed at the end, connected to the air flotation robot, of the flexible part and used for driving control over vibration of the flexible part. According to the invention, the space-ground consistency evaluation is carried out on the designed flexible spacecraft control test system, and the influence level of an uncertainty source on the space-ground consistency is given based on uncertainty quantitative analysis.

This application belongs to the field of spacecraft control technology and discloses a flexible spacecraft attitude control method and related equipment based on the theory of all-drive systems. The method includes: establishing an attitude tracking error model that includes the uncertainty of flexible vibration, external disturbances, and rotational inertia, and transforming it into a second-order all-drive system model with lumped disturbance terms; converting the second-order all-drive system model into a linear time-invariant closed-loop system with lumped disturbance terms, and constructing a preset-time sliding mode control law that enables the state of the linear time-invariant closed-loop system to converge within a preset time; constructing an RBF disturbance observer to approximate the estimation of the lumped disturbance terms; constructing an RBF preset-time pseudo-linear sliding mode controller based on the second-order all-drive system model, the preset-time sliding mode control law, and the RBF disturbance observer; and controlling the flexible spacecraft according to the RBF preset-time pseudo-linear sliding mode controller. This method has the advantages of achieving preset-time convergence and improving attitude control accuracy and stability.

A flexible spacecraft attitude active fault-tolerant control method without angular velocity measurement is proposed to solve the problems of actuator faults, flexible appendage vibrations and external disturbances in the process of spacecraft on-orbit attitude maneuver. The method considers the actuator faults, space environment external disturbances and flexible appendage vibrations, establishes the spacecraft attitude kinematics and dynamics model, designs the fixed-time sliding mode observer based on cosine sliding mode function and the flexible vibration observer, uses the estimation results of the two types of observers to rewrite the dynamics model equivalently, and designs the adaptive learning observer to estimate the comprehensive uncertainty caused by actuator faults, angular velocity estimation error, flexible vibration estimation error and external disturbances. Finally, the specified time fault-tolerant controller is designed according to the estimated value of the comprehensive uncertainty. The method has strong fault-tolerant ability, good flexible vibration suppression performance and robustness to external disturbances.

The invention discloses an attitude and orbit integrated fault-tolerant control method, system, medium and equipment for a flexible spacecraft, and the method comprises the steps: building a geocentric inertial coordinate system and a spacecraft body coordinate system, and carrying out the modeling of a rigid body and a flexible solar panel of the flexible spacecraft through a finite element theorem, obtaining a dynamic formula and a kinematics formula of the single-satellite spacecraft; establishing a relative kinematics equation between the tracking spacecraft and the target spacecraft; performing mathematical modeling on the fault type of the actuator, wherein the mathematical modeling comprises a gain fault, a deviation fault and a clamping stagnation fault; constructing a nonsingular terminal sliding mode surface, wherein the sliding mode surface comprises a relative dual quaternion error item and an auxiliary sliding mode surface item; and deriving a spacecraft system error kinematical equation based on the sliding mode surface, deriving a control moment calculation formula in combination with an actuator output moment expression, generating a control instruction according to the control moment formula, and driving the tracking spacecraft to realize integrated fault-tolerant control of attitude and orbit.

The application provides a spacecraft on-orbit attitude control and structure dynamics coupling simulation method and system, establishes a simplified three-dimensional model of the spacecraft; defines analysis steps, torques and concentrated forces, studies attitude control and flexible structure vibration suppression equations and compiles the equations into UAMP subprograms, simulates attitude control torques and flexible structure vibration suppression dynamics; and finally evaluates whether the attitude angle accuracy and the flexible structure vibration suppression effect of the spacecraft reach the expectation. The application considers the influence of the motion and deformation of each flexible component of the spacecraft on the attitude of the central rigid body, the mass center displacement, reveals the structure dynamics characteristics of the large flexible spacecraft under the on-orbit attitude control condition, and provides a new technical approach for the evaluation of the structure mechanics performance, reliability and pointing accuracy of the spacecraft.

The invention discloses a quantitative stability attitude tracking control method, device and equipment of a flexible spacecraft, belongs to the technical field of flexible spacecrafts, and particularly relates to quantitative stability attitude tracking control of the flexible spacecraft. The problem that an existing flexible spacecraft attitude control method cannot give consideration to the convergence speed and the quantitative stability of precision is solved. The method comprises the following steps: representing the attitude of the spacecraft by using a modified Rodrigues parameter, and obtaining an attitude dynamic model of the flexible spacecraft. The quantitative stability attitude tracking control method, device and equipment of the flexible spacecraft are suitable for quantitative stability attitude tracking control of the flexible spacecraft.

This invention relates to the field of on-orbit satellite structural condition monitoring technology, specifically disclosing an intelligent assessment method for satellite structural condition based on fiber optic sensors. The method includes: Step S1, conducting ablation simulations of the rubber material on the surface of a flexible spacecraft to obtain different damage states and corresponding strain data; Step S2, training a time recurrent neural network (LSTM) using the different damage states and corresponding strain data to obtain a classification prediction model; Step S3, inputting the data to be predicted into the classification prediction model to obtain the predicted damage state output by the model. This invention achieves intelligent assessment of laser ablation damage modes of flexible satellite surface materials through fiber Bragg grating sensing networks and long short-term memory (LSTM) network technology.

The invention relates to a hierarchical reinforcement learning cooperative control method for a large flexible spacecraft, and belongs to the field of spacecraft attitude control. The method comprises the following steps: 1) converting a large flexible spacecraft attitude control system into an interconnection system composed of a decentralized multi-rigid-body attitude control subsystem and an overall structure vibration control subsystem; the hierarchical reinforcement learning cooperative control method for the large-scale flexible spacecraft has the advantages that high-precision pointing control of the large-scale flexible spacecraft can be realized, and the robustness of the system is enhanced. The hierarchical reinforcement learning cooperative control method for the large-scale flexible spacecraft has the advantages that the hierarchical reinforcement learning cooperative control method for the large-scale flexible spacecraft can be implemented by the aid of the hierarchical reinforcement learning cooperative control method for the large-scale flexible spacecraft, and the hierarchical reinforcement learning cooperative control method for the large-scale flexible spacecraft can be implemented by the hierarchical reinforcement learning cooperative control method;

The application provides a method and system suitable for full-physical simulation of attitude dynamics of a flexible spacecraft, and relates to the field of spacecraft dynamics and control technology, and the method comprises the following steps: a static pressure gas lubrication three-axis air bearing table is used to simulate the dynamic environment of the spacecraft in on-orbit flight; the balance of the table body is coarsely adjusted by adding or reducing the weight blocks of the table body, and the slider on the guide rail on the table is finely adjusted to adjust the center of mass of the table body to the origin; inertia identification is performed on the table body to obtain the three-axis rotational inertia of the air bearing table, and a required scale reduction coefficient of the test is determined; a disturbance simulator is used to simulate the flexible disturbance torque generated in the jet process; and the attitude angle of the table body is measured according to different measurement ranges, the influence of the disturbance torque generated by the flexible accessory on the spacecraft is analyzed, and the flexible disturbance torque generator is used to simulate the disturbance torque generated by the vibration of the flexible accessory on the spacecraft, so that the center of mass can be kept unchanged, and the table overturning phenomenon can be avoided; and the application can also provide a reference basis for the design and demonstration of the attitude control scheme of the spacecraft.

The application relates to a flexible spacecraft attitude inner mode control method under external interference, in particular to a flexible spacecraft attitude tracking method, and particularly relates to a flexible spacecraft attitude maneuver control method in space operation under unknown frequency disturbance, and belongs to the technical field of spacecraft attitude control. In the case that unknown external disturbance exists and uncertain parameters exist in inertial information, a compound attitude controller based on the combination of an inner mode principle and an adaptive control method is provided, on the basis of realizing high-precision attitude control, asymptotic modal vibration suppression and unknown external disturbance suppression can be realized, the size of a control moment can be adjusted through corresponding parameters, and the compound attitude controller has actual engineering application value and theoretical research value.