31 results about "Spacecraft formation" patented technology

The invention relates to a spacecraft collaborative formation safety control method and system based on an adjustable self-adaptive control barrier function. The method comprises the following steps: establishing a six-degree-of-freedom model for describing relative motion of a spacecraft formation; multiple safety key constraints which need to be met by the spacecraft formation are established, and the constraints are formalized into a zero-hyperlevel set function of a state variable; a disturbance observer is designed, and unknown lumped disturbance existing in the six-degree-of-freedom model is estimated and compensated online; a safety controller is constructed, the safety controller comprises an adjustable self-adaptive control barrier function used for generating a safety constraint boundary changing along with time, and parameters of the adjustable self-adaptive control barrier function are adjusted online through the auxiliary dynamic system; and constructing a quadratic programming problem, and solving to obtain an optimal safety control instruction which minimizes the deviation between the actual control input and the nominal control input. Compared with the prior art, the method has the advantages of improving the task success rate and safety of the spacecraft formation in a complex, dynamic and uncertain environment and the like.

The invention discloses an earth-moon translation point orbit spacecraft formation relative motion configuration design and control method, and belongs to the technical field of spacecraft orbit dynamics and control. The method comprises the following steps: firstly, constructing a double-layer dynamic model consisting of a circular restrictive three-body problem model and a high-precision ephemeris model, and defining an orbit amplitude difference and a phase difference as configuration parameters; then, according to different track amplitude difference and phase difference combinations, three types of natural relative motion configurations are set, a periodic track is rapidly generated under a circular restrictive three-body problem model, and multi-circle correction is conducted through a high-precision ephemeris model and two-stage multiple targeting; a hierarchical constraint model is established for a controlled task surrounding a fixed position, a plane and a space, and a control pulse is solved by adopting a two-stage optimization strategy. According to the method, efficient design and stable control of formation configuration in a complex dynamic environment are realized, and the method is suitable for various translation point orbit tasks.

The application relates to the technical field of multi-spacecraft cooperative control, and discloses a multi-spacecraft grouping consistency restraint control method and device under optimized FLS, which comprises the following steps: acquiring the network topology of a heterogeneous multi-spacecraft system and constructing a multi-subgroup directed communication topology network, constructing a spacecraft attitude control system model based on modified Rodrigues parameters; approximating unknown nonlinear terms in the model through a fuzzy logic system, constructing a fuzzy logic system based on a gradient descent algorithm and Armijo line search criterion; constructing a restraint control term and using a backstepping method to construct a virtual control law, combining the output of the fuzzy logic system based on the gradient descent algorithm and the Armijo line search criterion to construct an actual control input law, so that grouping consistency restraint control of the heterogeneous multi-spacecraft system is realized. The application can improve real-time approximation precision and adaptive step length adjustment capability while guaranteeing grouping consistency, and can enhance topology applicability and task cooperation flexibility.

The present application relates to the field of aerospace technology, specifically to a kind of multi-pulse parallel construction method of gravitational wave detection formation considering timing optimization, each spacecraft of formation simultaneously departs from virtual center, each spacecraft is moved to the preset specified position by applying multiple pulses, so as to form a kind of spacecraft formation that can be used for space gravitational wave detection, by giving the steps of calculating the initial position and velocity of each spacecraft and applying initial control to reach the initial position and velocity, the state quantity of each spacecraft at each time in the formation construction process is calculated, so as to provide reference for the design of satellite formation task in actual engineering application, and the accuracy of the linearized relative motion dynamics model established by the multi-pulse parallel space gravitational wave detection formation is improved by differential correction, the error in the relative motion dynamics model is reduced, the initial state set in advance can be effectively reached, and the working efficiency of formation component is improved.

The application provides a spacecraft formation trajectory tracking distributed control method and related equipment, comprising: constructing a relative motion PH equation of each spacecraft in a target spacecraft formation; converting the relative motion PH equation of each spacecraft respectively to obtain a relative motion trajectory tracking error PH equation of the target spacecraft as a tracking error system of the target spacecraft formation; based on the tracking error system, constructing an expected Hamilton function that couples the relative position error of each spacecraft in the target spacecraft formation to the tracking error system; designing a control law of the target spacecraft formation according to the expected Hamilton function and calculating a value of the distributed cooperative control law of the target spacecraft formation; performing trajectory tracking distributed control on the target spacecraft formation according to the value of the distributed cooperative control law; the spacecraft formation can accurately and quickly reach the expected motion trajectory and maintain the configuration, and the distributed cooperative control of the spacecraft cluster network is realized.

A fully distributed control method for heterogeneous spacecraft formations considering time-varying topology and time-varying time delays is proposed. The method includes establishing an orbital dynamics model of the accompanying satellite formation with distinct input time delays and obtaining its state-space equations; establishing a signal model of the master spacecraft to be tracked and obtaining its state-space equations; constructing explicit fully distributed observers for each accompanying satellite to obtain the master spacecraft's state information constrained by time-varying communication time delays; designing explicit linear state feedback gains and feedforward gains respectively; and using the state feedback gains, feedforward gains, and fully distributed observers, establishing a corresponding fully distributed control protocol for the accompanying satellite formation system affected by time-varying topology and time-varying time delays to ensure that each accompanying satellite completes its tracking mission of the master spacecraft. This invention exhibits good control performance, ensuring stable operation of the heterogeneous spacecraft formation system and achieving the task of accompanying satellites tracking the master spacecraft.

The application provides a simplified attitude distribution cooperative control method for a directed communication leader-follower spacecraft formation, which comprises the following steps: firstly, describing the simplified attitude motion of the spacecraft in a unit celestial coordinate system, and converting the simplified attitude motion into a translation of the terminal point of a pointing vector on a unit celestial sphere; secondly, designing a distributed cooperative virtual control quantity under a directed communication topology according to the consistency theory, so as to ensure that the pointing vector of the follower spacecraft converges to a fixed pointing direction in the inertial space determined by the pointing vector of the leader spacecraft, and then calculating an ideal angular velocity vector of the follower spacecraft according to the virtual control quantity; thirdly, taking the difference between the angular velocity vector of the follower spacecraft and the ideal angular velocity vector as a sliding mode variable, and designing a sliding mode surface; and finally, designing a distributed cooperative control law for the follower spacecraft based on a super-twisting second-order sliding mode control method. The application can ensure that the pointing vector of the follower spacecraft in the directed communication formation converges to the pointing vector of the leader spacecraft fixed in the inertial space under a disturbance environment, and reduces the communication amount of the formation.

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

This invention discloses a method for modeling and simulating the attitude-orbit coupling of spacecraft formations, comprising: establishing coupling models of the desired attitude and orbital parameters of the lead satellite, the actual attitude and orbital parameters of the lead satellite, the desired attitude and orbital parameters of the follower satellite, and the actual attitude and orbital parameters of the follower satellite; calculating the desired attitude and orbital parameters of the lead / follower satellites and the actual attitude and orbital parameters of the lead / follower satellites respectively based on the models; and performing flight formation control based on the calculated parameters to achieve formation flight of the follower satellites relative to the lead satellite. This invention can be used in satellite controller design. On the one hand, by modeling the attitude and orbital parameter information using dual quaternions, this method can more accurately establish a satellite dynamics model; on the other hand, by using the parameters of the lead satellite to determine the parameters of the follower satellites, the follower satellites can more accurately maintain their relative positional relationship with the lead satellite.

The application particularly relates to a spacecraft formation flight task planning method based on cooperation of a large language model and a professional tool set, and comprises the following steps: constructing a three-layer task planning architecture of the professional tool set, the LLM and a high-level dynamic execution engine; constructing the professional tool set for the spacecraft formation flight; decomposing a user instruction into a meta-task sequence through a complex task decomposition model; and calling the tool set to execute the meta-task in sequence through a task sequence execution model. The application provides a more efficient, flexible and intelligent solution with scene generalization capability for the spacecraft formation flight task planning problem.

For maneuvering a spacecraft, a method calculates a virtual point that represents a plurality of spacecraft orbiting in a spacecraft formation. The method iteratively calculates a desired trajectory relative to the virtual point for a given spacecraft of the plurality of spacecraft using a relative orbital elements (ROE) state x of the virtual point and the given spacecraft. Model predictive control maintains the given spacecraft within an ROE constraint for a given time interval subject to a fuel consumption. In response to determining a drift trajectory of the given spacecraft is outside of a trajectory tolerance of the desired trajectory, maneuvering the given spacecraft to within the trajectory tolerance using a maneuver trajectory determined using the model predictive control.

The application discloses a kind of under DoS attack based on dynamic event triggering's spacecraft formation system's self-adapting consistency control method.The method is first based on multi-agent consistency model theory, establishes the multi-agent system model of multi-agent spacecraft formation system, then consider based on time series DoS attack, by the frequency and duration of DoS attack are analyzed and researched, to solve the consistency problem of multi-agent spacecraft formation under DoS attack, then design a dynamic event triggering mechanism to reduce communication transmission frequency, and it is proved that there is no Zeno behavior existence, finally based on Lyapunov stability theory analysis, design a kind of event triggered distributed adaptive controller, solve the consistency problem of multi-spacecraft formation.The method is applied to multi-spacecraft formation system, guarantee the normal operation of system under DoS attack.

The invention relates to a detection task planning method and device, equipment, a medium and a product, and the method comprises the steps: screening a first observation target of which the absolute value of the yellow latitude is greater than a solar avoidance angle from to-be-observed targets, and obtaining a target observation set; determining a first observation parameter of each first observation target in the target observation set, and determining first observation time consumption of the spacecraft formation for each first observation target based on the first observation parameter; determining a first observation sequence on the sun-earth translation point task orbit based on the first observation time consumption; wherein the first observation sequence is a sequence in which the beam combiner observes the first observation target according to a yellow channel nearest principle; processing the first observation sequence through a convex optimization algorithm to obtain a target sequence serial plan; wherein the target sequence serial planning is used for indicating the light collector to be matched with the beam combiner to observe the first observation target under the optimal fuel track. According to the invention, distributed detection of a plurality of targets by the spacecraft formation can be realized.

The application discloses a spacecraft attitude preset time cooperative control method based on reinforcement learning, carries out mathematical description of a multi-spacecraft attitude cooperative control problem, constructs a preset time distributed observer, guarantees that a follower obtains observation information of a state of a leader within preset time, determines a preset time performance function, quantitatively characterizes convergence time, transient performance and steady-state performance constraints of a cooperative tracking error, carries out error transformation based on an obstacle function, converts the cooperative tracking error system subjected to preset performance constraints into an unconstrained system, determines a performance index function and a corresponding Hamilton-Jacobi-Bellman equation, solves partial derivatives about optimal control to obtain an expression form of optimal control input about an optimal function, and designs an approximate optimal controller under an evaluation network framework by using a reinforcement learning method. The application can guarantee that a spacecraft formation system meets preset convergence time, transient and steady-state performance, and considers energy consumption problems.

The application discloses a completely distributed state observation method with a time convergence, and belongs to the technical field of multi-agent system cooperative control and state estimation. The application solves the problems that the existing method needs to depend on global network topology information and cannot realize controllable real convergence time and poor robustness. The application can make each follower independently and efficiently reconstruct the state and derivative of a leader which cannot be directly obtained without any global network information, and ensure that the estimation error accurately converges to zero within the real time set by a user. Moreover, the state observation method of the application is suitable for dynamic systems with external disturbance and model uncertainty, has good robustness and engineering application prospect, and is suitable for application scenes, such as spacecraft formation, unmanned aerial vehicle cluster and robot cooperation, which need high-precision synchronization and robustness. The method of the application can be applied to multi-agent system cooperative control and state estimation.

The invention provides a data-driven spacecraft formation configuration reconstruction method and device, and belongs to the technical field of spaceflight. According to the method, the current relative motion state of each slave spacecraft in a plurality of slave spacecrafts in the spacecraft formation, the target relative motion state after configuration reconstruction and the state auto-encoder are combined to construct the control law of each slave spacecraft for realizing spacecraft formation configuration reconstruction, and the process improves the calculation efficiency of solving the control law and improves the calculation efficiency of the spacecraft formation configuration reconstruction. The method is low in computing power requirement, is easy to achieve on-satellite deployment, achieves autonomous planning, can improve the reconstruction efficiency of the spacecraft formation configuration, and facilitates the spacecraft formation to complete tasks.

The application discloses an inter-satellite baseline high-precision measurement and estimation method, and is characterized in that the method comprises the following steps: S1. Estimating and calibrating GNSS measurement information through a star sensor and laser ranging information; S2. Determining a measurement noise variance matrix R, and estimating in combination with satellite motion characteristics and noise characteristics; and S3. Adopting an interactive multiple model unscented Kalman filtering method to perform information processing. In the application, the problems that a measurement scale of an existing method does not meet long-interference baseline measurement requirements, measurement information is single, measurement precision is unknown, and an information processing method is not suitable are overcome, and a method for realizing fusion of GNSS, laser link and star sensor measurement information by using an interactive multiple model unscented Kalman filtering is provided, so that inter-satellite high-precision real-time baseline measurement information is obtained, the demand of future multi-spacecraft formation cooperation for high-precision inter-satellite measurement information is met, and mm-level inter-satellite baseline measurement precision is achieved.

The application discloses a spacecraft formation configuration decoupling control method and belongs to the technical field of satellite formation control. The application aims at solving the problems of serious approximation error accumulation and high fuel consumption of a spacecraft formation control method based on a CW equation. The method comprises the following steps: calculating relative orbit elements of a sub-satellite and a main satellite, obtaining a radial relative motion equation, a trace relative motion equation and a normal relative motion equation of the sub-satellite and the main satellite, and obtaining a modified equation; determining a velocity pulse according to a dynamic equation between the relative orbit elements and the velocity pulse; determining a required change amount of the relative orbit elements, and then determining a required change amount of a relative orbit inclination vector, a required change amount of a relative semi-major axis, a required change amount of a relative eccentricity vector and a required change amount of a relative angular distance; controlling the relative orbit inclination vector by adopting a single-pulse normal control strategy; controlling the relative semi-major axis and the relative eccentricity vector by adopting a double-pulse trace pulse strategy; and controlling the relative angular distance by adopting a multi-period trace control strategy. The application is used for formation control.

The invention relates to the technical field of space on-orbit robot control, in particular to a reconfigurable rigid-flexible coupling space on-orbit assembly robot hierarchical control method and system, and the control system comprises a spacecraft formation control system and a rigid-flexible coupling robot control system. According to the rigid-flexible coupling robot control system, the working stroke of a robot can be changed through reconfiguration of spacecraft formation re-planning, and on the basis of an upper-layer task planner, a controller fuses multi-perception information of vision, force sense and position to achieve hierarchical control over a flexible cable driving robot and a mechanical arm. The control method is oriented to space large-size component assembly tasks, the robot hierarchical controller fused with multi-source information is designed by decoupling robot kinematics and dynamics and combining operation scene requirements, the overall operation efficiency and precision of the robot are improved, and the adaptability of the robot to different on-orbit tasks is enhanced.

This invention proposes a spacecraft formation flight control method utilizing the null-space projection behavior method. First, a spacecraft formation dynamics model and its communication topology network model are established, and control target levels are defined. Second, a desired trajectory is generated based on the null-space behavior control method. Considering mass uncertainty, actuator saturation, and external disturbances, a controller is designed to track the generated desired trajectory. This invention monitors the relative positions and velocities between spacecraft in real time, adjusts the control inputs of each spacecraft, ensures the connectivity of the communication network is not disrupted, avoids collisions with obstacles and other spacecraft, and can achieve the desired configuration of the spacecraft formation according to mission requirements, improving mission efficiency and accuracy. Simultaneously, considering complex environmental factors such as mass uncertainty, actuator saturation, and external disturbances, it achieves efficient, safe, and autonomous control of spacecraft formation flight.

The invention provides a dynamic full-link simulation method and system for space gravitational wave detection spacecraft formation. A multi-physical field model, a full-link noise model and drag-free control are fused to carry out system-level simulation. The method comprises the steps that firstly, multi-physics field models such as a temperature field, a magnetic field and a self-gravitational field are established, optical measurement noise and residual acceleration noise models are combined and input to a spacecraft formation orbit dynamics and drag-free control system, then the drag-free control model feeds satellite state information after control back to the multi-physics field models, system-level closed-loop simulation is carried out, and the system-level closed-loop simulation is carried out. State changes of the spacecraft are dynamically analyzed, scientific measurement data are output to evaluate the sensitivity of the spacecraft formation space gravitational wave detection system, a multi-physical field noise transmission rule among an instrument, the spacecraft and the environment is disclosed, and space gravitational wave detection spacecraft formation dynamic full-link simulation is achieved.

The application discloses a kind of impact dangerous area calculation method with single-sided constraint spacecraft formation, for the safety problem of single-sided constraint spacecraft formation system, focus on the dangerous area of tether impact, load impact etc.For impact risk, load motion model is established based on Hill equation, and the whole state space is analyzed in combination with simple cell mapping algorithm.Tension calculation formula is derived at the moment of tether impact, and it is calculated that tension peak value is mainly concentrated in the region where the initial time load velocity direction is consistent with the direction of tether length.Simulation results show that impact dangerous area is related to the initial kinetic energy of system.The application can provide technical support for impact warning of spacecraft formation system.

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

The invention discloses a spacecraft formation configuration decoupling control method, and belongs to the technical field of satellite formation control. The invention aims to solve the problems of serious approximate error accumulation and high fuel consumption of a spacecraft formation control method based on a CW equation. Comprising the following steps: calculating a relative orbital element of a sub-satellite and a main satellite, obtaining a radial, trace and normal relative motion equation of the sub-satellite and the main satellite, and obtaining a corrected equation; determining a speed pulse according to a kinetic equation between the relative orbital element and the speed pulse; determining the variable quantity required by the number of relative orbits, and further determining the variable quantity required by a relative orbit inclination angle vector, the variable quantity required by a relative semi-major axis, the variable quantity required by a relative eccentricity rate vector and the variable quantity required by a relative angular distance; adopting a monopulse normal control strategy to control the relative orbit inclination angle vector; adopting a double-pulse trace pulse strategy to control the relative semi-major axis and the relative eccentricity rate vector; and controlling the relative angular distance by adopting a multi-period trace direction control strategy. The method is used for formation control.

The application provides a leader-follower spacecraft formation simplified attitude distributed control method with pointing constraints, which comprises the following steps: first, a simplified attitude error model between the spacecraft is established; second, a pointing constraint model of each follower spacecraft is constructed; third, an artificial potential field of each follower spacecraft is designed according to the pointing constraint, the simplified attitude error and a communication topology; fourth, a Lyapunov function is constructed based on the artificial potential field function, and a distributed cooperative virtual control quantity of each follower spacecraft is designed; and fifth, the difference between the angular velocity vector of each follower spacecraft and the virtual control quantity is taken as a sliding mode variable, a sliding mode surface is constructed, and a distributed cooperative sliding mode control law of each follower spacecraft is designed. The application can realize that the pointing vector of the follower spacecraft converges to the pointing vector of the leader spacecraft fixed in the inertial space under the condition of the undirectional communication topology, and can meet the pointing constraint of the follower spacecraft in the attitude maneuvering process.

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

For spacecraft control, a method iteratively measures a drifting spacecraft trajectory of a spacecraft relative to a bounding box, wherein the spacecraft has a fixed maximum thrust, a minimum thrust duration, a thrust duration that is an integer multiple of the minimum thrust duration, and a fixed update period. In response to the determining the spacecraft does not satisfy a trajectory condition, the method iteratively calculates maneuver controls wherein there are a number of guidance trajectories within the thruster duration. The method maneuvers the spacecraft along the guidance trajectories with the maneuver controls.

The application discloses a spacecraft formation event-triggered attitude coordination control method considering communication time delay, comprising the following steps: according to the communication topology structure of a spacecraft formation system, an attitude dynamics model of the spacecraft formation is established, which can consider external disturbance, model uncertainty and communication time delay; based on the attitude dynamics model, a nonlinear observer is constructed, which can observe the collective nonlinear term of the spacecraft formation system; based on the nonlinear observer, an event-triggered attitude coordination control law and an event-triggering mechanism are designed; according to the event-triggered attitude coordination control law, the attitude and angular velocity of the formation spacecraft are controlled, and according to the event-triggering mechanism, it is determined whether the corresponding event-triggered attitude coordination control law of the formation spacecraft needs to be updated and whether the state information of the formation spacecraft needs to be sent to the adjacent formation spacecraft. The application can consider the influence of disturbance, uncertainty and communication time delay, save the energy, communication and computing resources of the system while ensuring good control performance.

This invention discloses a lossless convexification method for stochastic optimal control of gravitational wave spacecraft formation attitude alignment, comprising: setting relevant parameters under the mission scenario, including the spacecraft's moment of inertia, maximum thrust torque, initial and target states, pointing constraints, and perturbation coefficient matrix; establishing a stochastic spacecraft attitude dynamics model; describing multiple constraints with uncertainties using chance constraints, including torque constraints, taboo constraints, and mandatory constraints; using the Schur complement theorem to transform the non-convex parts of the constraints into linear matrix inequalities (LMIs), and adding virtual control variables and penalty terms to obtain a convex problem that can be directly solved using CVX; this invention aims to achieve stochastic optimal control of gravitational wave spacecraft formation attitude alignment through convex optimization methods.

The invention relates to a micro-vibration evaluation method of a spatially distributed optical interferometer and a virtual device, and is applied to the technical field of aerospace. According to the invention, the disturbed type of the spatial distributed light interferometer running along with the spacecraft formation is obtained. And based on the disturbed type, obtaining a pose influence parameter of the spatial distributed light interferometer. And determining a pose parameter of a movable mirror surface of the spatial distributed light interferometer in the observation time period based on the pose influence parameter. And simulation imaging is carried out based on the pose parameters, and the imaging ambiguity of the spatial distributed light interferometer in the observation time period is determined. And according to the imaging ambiguity and the pose parameter, determining a micro-vibration evaluation result of the spatial distributed light interferometer. By determining the relationship between the imaging result and the pose state of the movable mirror surface under different disturbance factors, the influence on the imaging quality when the spatial distributed optical interferometer generates micro-vibration can be analyzed. Therefore, the imaging quality of the spatially distributed optical interferometer can be improved.