42 results about "Spacecraft system" patented technology

The invention discloses a six-degree-of-freedom multi-spacecraft control experiment platform and method based on semi-physics-virtuality, belongs to the technical field of spacecraft system and control system simulation, and constructs a multi-spacecraft system composed of semi-physics spacecrafts and pure virtual spacecrafts. The attitude of the semi-physical spacecraft is simulated by a simulator mounted on a three-dimensional attitude air floating table, is acquired by a sensor and is converted into a rotation matrix form, and a translation part of the semi-physical spacecraft is virtually calculated through a simulation integrator; the pose of the virtual spacecraft is virtually simulated by an SE (3) integrator with high-guarantee geometric characteristics; the system takes the ROS as a communication framework, and realizes state and instruction release by adopting a custom message; and ensuring that the time system of each spacecraft is consistent with the real time through a time synchronization mechanism. The platform supports cooperative operation of a semi-physical spacecraft and a pure virtual spacecraft, and has the characteristics of flexible structure, high simulation precision and good expandability.

The invention discloses a distributed spacecraft low-communication attitude fine control method based on event triggering, and aims to solve the problems of high communication frequency, serious resource waste and difficulty in effective compensation of multi-source interference in the prior art. According to the method, a spacecraft attitude deep coupling model is constructed, and the dynamic influence of compound interference such as centroid change and flexible vibration is comprehensively considered; designing an event triggering mechanism based on the control signal variation, and updating a control instruction as required to reduce the communication frequency; and a self-adaptive composite controller is constructed, interference and communication errors are estimated and compensated on line, and the attitude control precision is ensured. According to the method, only 75 times of communication is triggered in 50-second simulation, meanwhile, the attitude tracking error is controlled within 5 * 10 <-3 > rad, and the method has the advantages of being low in communication load, high in control precision, high in anti-interference capacity and the like and is suitable for a distributed spacecraft system with strict requirements for attitude stability and communication resource consumption.

Various embodiments comprise a multi-agent spacecraft system for rendezvous and proximity operations. The spacecraft system comprises a chief spacecraft and a deputy spacecraft. The chief spacecraft maneuvers to the orbit of a Resident Space Object (RSO) and deploys a deputy spacecraft to rendezvous with the RSO. The deputy spacecraft maneuvers to the RSO and senses the RSO during the rendezvous and proximity operations maneuver. The deputy spacecraft generates a Three-Dimensional (3D) surface reconstruction of the RSO based on the sensing. The deputy spacecraft determines a roll rate and a relative trajectory of the RSO based on the sensing. The deputy spacecraft returns to the chief spacecraft.

The invention discloses a transient performance enhancement-oriented multi-spacecraft attitude cooperative tracking control method, which provides a fixed time controller through an inversion control strategy, realizes quick response of a multi-spacecraft attitude system, and ensures that a consistency error is converged to a sufficiently small domain within fixed time. A fuzzy weight updating law is designed, dependence on system model information is reduced, and adaptability and high robustness are achieved. According to the asymmetric fractional obstacle Lyapunov function, adherence of multiple spacecrafts to time-varying constraints is guaranteed, meanwhile, the problem that the control amplitude related to a large constraint boundary in a traditional logarithmic obstacle Lyapunov function is too large is solved, and therefore the proposed method can process constrained and unconstrained scenes. A saturation threshold event triggering strategy is designed, a multi-spacecraft system is allowed to select three different triggering strategies by adjusting parameters, the transmission frequency of control signals is flexibly changed, and control bus resources are reasonably distributed.

The application provides a variable configuration spacecraft system and method for non-cooperative target capture, and belongs to the field of deep space exploration. The method comprises the following steps: 1) a rigid combination flies to the vicinity of the target; 2) the combination is spun to be expanded into a ring belt; 3) the ring belt envelopes the target; and 4) the ring belt encloses the target. Large space debris and small celestial bodies both belong to non-cooperative targets, and cannot provide special interfaces for assisting capture. It is extremely difficult to develop general control mechanisms and methods to realize reliable capture and stable connection of various targets. In order to solve the problem, the application utilizes the inter-satellite connection to expand a plurality of controllable satellites into a ring-shaped multi-body tether satellite formation system, and carries out ring belt enveloping and then enclosing the tether to realize capture of a non-cooperative target with large scale, complex geometric shape, great uncertainty of motion state and inertia information according to a predetermined trajectory. The application has the advantages of high applicability, flexibility, high intelligence, reusability and small size.

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.

This invention discloses a distributed, event-triggered, low-communication fine-grained attitude control method for spacecraft, aiming to overcome the problems of high communication frequency, severe resource waste, and difficulty in effectively compensating for multi-source interference in existing technologies. This method constructs a deeply coupled attitude model for the spacecraft, comprehensively considering the dynamic effects of combined interferences such as center-of-mass changes and flexural vibrations; designs an event-triggered mechanism based on changes in control signals to update control commands as needed, thereby reducing the communication frequency; and constructs an adaptive composite controller to estimate and compensate for interference and communication errors online, ensuring attitude control accuracy. In a 50-second simulation, this method triggers only 75 communications while controlling the attitude tracking error to within 5 × 10⁻⁶. ‑3 Within rad, it has the advantages of low communication load, high control precision, and strong anti-interference capability, and is suitable for distributed spacecraft systems with strict requirements for attitude stability and communication resource consumption.

The invention discloses a shared memory protection method based on a multi-core controller, and belongs to the technical field of spacecraft system design. The method is applied to a shared memory protection system based on a multi-core controller, and the system comprises a first data cache block and a second data cache block which are arranged in a first core controller, and a third data cache block and a fourth data cache block which are arranged in a second core controller; the method comprises the following steps: transmitting cached inter-satellite task input data to a fourth data cache block by utilizing a second data cache block, so that the fourth data cache block performs inter-satellite task calculation; and receiving an inter-satellite calculation result transmitted by the third data cache block by using the first data cache block, and telemetering and downloading cached data to the ground. According to the method, the conflict problem of shared memory data access of the multi-core computer can be solved.

The invention discloses a spacecraft system parameter on-orbit identification method in an extension rod fault state. The method comprises the following steps: constructing an attitude dynamic model of a spacecraft system rigid-flex coupling configuration in a coiled extension rod expansion fault state; establishing an expression of the torque of the extension rod vibration to the spacecraft, and substituting the torque expression into the attitude dynamics model to obtain an expression of the system control torque represented by model parameters and attitude parameters; a system control moment is given and applied to a spacecraft rigid body through an actuator of the spacecraft, so that the spacecraft rotates, and corresponding attitude parameter measurement values are obtained; sending the obtained attitude parameter measured value to the ground, expressing the estimated value of the system control torque by using the attitude parameter measured value and the estimated value of the model parameter, and defining the system identification error as a function of the system control torque and the estimated value thereof; and determining a to-be-identified parameter based on the model parameter, and searching the to-be-identified parameter capable of enabling the system identification error function to be minimum as an optimized parameter identification result.

Vehicle or spacecraft system for an autonomous or semi-autonomous vehicle, comprising at least one flow sensor arranged on the outside of the vehicle, an electrically connected electronic evaluation unit, and a control, steering, braking or stabilization device coupled to the evaluation unit for automatic counter-control on the vehicle side in the event of external air or flow effects.

This invention discloses a model- and data-driven fault diagnosis method for deep space exploration spacecraft, belonging to the field of deep space exploration technology. Addressing the shortcomings of traditional methods in model adaptability, data utilization, and fault location, the method constructs a four-layer fusion architecture comprising the entity layer, model layer, data layer, and application layer of the deep space exploration spacecraft system, achieving dynamic integration and interaction between the mechanistic model and multi-source telemetry data. Through a mechanistic-data hybrid model, real-time system state prediction, fault injection simulation, multi-source data comparative analysis, and health status assessment are achieved. A corresponding software prototype has been developed, possessing functions such as fault mode library management, telemetry-simulation data-driven comparison, 3D visualization, and simulation verification of fault handling strategies. This method can effectively assist in the rapid and accurate location and handling decisions of on-orbit faults, improving the safety and reliability of deep space exploration missions.

The application discloses a thruster fault identification method and device based on a sliding mode observer. The method comprises the following steps: designing a sliding mode observer based on a spacecraft dynamics model; calculating real three-axis angular velocities of the spacecraft in a current control period based on gyro outputs of the spacecraft system; estimating estimated three-axis angular velocities of the spacecraft in the current control period based on the sliding mode observer; determining a residual function in the current control period based on the real three-axis angular velocities and the estimated three-axis angular velocities; determining whether a thruster fault exists in the current control period based on a relationship between the residual function and a preset threshold; if the thruster fault exists, calculating an estimated thruster fault matrix in the current control period based on the residual function, and calculating an included angle between the estimated thruster fault matrix in the current control period and each column of a predetermined thruster fault matrix, and determining a thruster corresponding to a column with the minimum included angle as a fault thruster. The application can locate the fault to each thruster.

The data flow design method for flexible information processing of remote sensing constellations belongs to the field of spacecraft system overall design technology. This method first calculates the data rate of different types of data content, including: (1) calculating the original image data rate based on detector size, quantity, and imaging frame rate; (2) calculating the small area windowing data rate based on window size and windowing data frame rate; (3) calculating the target point's adjacent area slice data rate based on threshold crossing rate; and (4) calculating the target trajectory information data rate based on single target trajectory information and target quantity. Then, the data transmission content is flexibly configured, and information overlap strategies at different levels are formulated. This enables dual-satellite information fusion for target positioning and tracking tasks under different link transmission capacity constraints. For randomly triggered multi-target tracking tasks, the data transmission content is adaptively adjusted according to link margin, enabling dual-satellite positioning and tracking under different levels of information fusion processing conditions.

A spacecraft system may include a first spacecraft having a chassis and a first antenna moveably mounted to the chassis, and a second spacecraft having a chassis, wherein the second spacecraft is adjacent the first spacecraft during launch to define a stacked configuration, and a passive restraint system defined between the first and second spacecraft that is configured to substantially prevent movement of the first antenna in a launch position while in a launch state when the first and second spacecraft are in a stacked configuration.

The application discloses a spacecraft software component requirement analysis and modeling method based on a dynamic data flow graph, relates to the technical field of spacecraft integrated electronics, and solves the problem that a traditional data flow graph is difficult to express the data transmission relationship between internal modules of a complex spacecraft system component and components. The method comprises the following steps: a dynamic data flow graph is constructed for data in spacecraft software; component requirement analysis is performed on the dynamic data flow graph to obtain components in the data flow graph; component outline design is performed on the components in the data flow graph; and a component model is established for the components according to the component outline design, which is used for component verification in a subsequent visual assembly process with tool support.

Provided are a spacecraft control system, a spacecraft control method, and a server device capable of quickly transmitting control data to a spacecraft. The spacecraft system includes: a plurality of spacecrafts; a plurality of small communicators that transmit control data related to imaging operations of the spacecrafts to the spacecrafts; and at least one ground station that receives, from the spacecrafts, captured image data related to images captured by the spacecrafts and space position data for the spacecrafts, and transmits identification data for the small communicators to the spacecrafts, in which each of the spacecrafts includes at least a first reception device that receives the control data transmitted from the small communicators, a control device that controls an imaging operation of the spacecraft on the basis of the control data received by the first reception device, and an imaging device that executes the imaging operation.

The invention discloses a low-cost constrained multidisciplinary robust optimization design method, and belongs to the field of spacecraft manufacturing and application. The implementation method comprises the following steps: constructing an aircraft multidisciplinary model; a numerical simulation model and a finite element analysis model are respectively constructed for pneumatic and structural subjects, and a proxy model is constructed for the pneumatic and structural subjects in combination with Bayesian learning. And constructing an aircraft manufacturing cost prediction model based on knowledge distillation. Based on a global sensitivity equation GSE of a deterministic domain coupling multidisciplinary system, input random input and model cognition uncertainty are introduced, expressions of mean values and covariances of subject coupling variables and system responses are derived respectively, and a rapid semi-analytical multidisciplinary random and model cognition hybrid uncertainty propagation method is established. Multidisciplinary uncertainty analysis is achieved, and aircraft multidisciplinary system performance response is obtained. And constructing a multi-objective multidisciplinary optimization model with optimal performance and lowest cost, and solving by adopting an NSGA-II algorithm to obtain a final aircraft design scheme.

The invention relates to the technical field of spacecraft system-level testing, in particular to a spacecraft multi-user navigation function closed-loop testing system and method, which can realize high-precision comprehensive verification of the navigation function of a spacecraft system and consider the characteristics of multiple tested users, high precision requirement and the like. Aiming at multi-user test requirements of tested spacecrafts, accurate test of a high-dynamic multi-user GNSS simulation source based on a cloud architecture is adopted, multi-spacecraft multi-user simultaneous association test is supported, functional tests such as absolute positioning, relative positioning and trajectory prediction are realized, the system architecture is flexible, the number of user channels is large, and multi-user requirements of complex spacecrafts are supported. According to the method, high-precision comprehensive verification of the navigation function of a spacecraft system can be realized, the characteristics of multiple tested users, high precision requirement and the like are considered, and a high-dynamic multi-user GNSS simulation source spacecraft closed-loop test based on a cloud architecture becomes an effective solution.

The invention discloses a semi-analytical coupling multidisciplinary global sensitivity analysis method considering cost and performance, and belongs to the field of spacecraft manufacturing and application. Defining an aircraft multidisciplinary system, and determining an optimal mean value and variance of input variables; the mean value and the variance are substituted into the agent model, the mean value and the covariance of all coupling variables and subject output responses are calculated, and the mean value and the variance of system output responses are obtained; the mean value and the variance are responded according to the system; carrying out linear normalization on the mean value to obtain a normalized performance index; based on the normalized performance index and the weight factor, obtaining a final dimensionless system response y eta; based on the mean value and the variance, constructing a variable cost-mean value and variance mapping relation table; establishing an input-oriented main sensitivity coefficient MSI, looking up the mapping relation table to obtain costs ci of different input variables and a mean value and a variance under the costs, and obtaining y eta through the mean value and the variance; and the ci and the y eta are substituted into the MSI to respectively obtain system response sensitive to the input variable cost and the input variable cost sensitive to the system response.

The invention discloses a SysML language-based general spacecraft system modeling method. The method comprises the following steps of: 1) establishing an organization structure of a target spacecraft system design model layer by layer in a packet form; 2) engineering task and demand analysis; 3) analyzing functions of the spacecraft system; 4) designing an overall preliminary scheme; 5) designing a subsystem scheme; and 6) integrating the overall final scheme. According to the method, a system engineering method based on SysML is adopted, an organization structure, a modeling process and detailed modeling requirements of a spacecraft system design model are provided, demand capture and analysis, system reuse and system optimization and evaluation are supported, digital modeling of a large-scale complex general spacecraft system is supported, and development efficiency and development quality of a spacecraft are improved.

The invention discloses a retractable platform of a space on-orbit satellite, and relates to the technical field of spacecraft systems and overall technologies, the retractable platform comprises a frame and a folding platform, a retractable part capable of being telescopically adjusted in the axis direction of the frame is arranged in the frame, and the extending end of the retractable part can extend out of the frame; the folding platform comprises a plurality of folding plates capable of being turned over, and every two folding plates are connected through a connecting component; the positions of the even number of connecting components on the folding platform are kept fixed on the telescopic piece, and the end part, which is not fixed on the telescopic piece, of the folding platform and the even number of connecting components, which is not fixed on the telescopic piece, of the folding platform can completely or partially slide on the telescopic piece in the axis direction of the telescopic piece; the even number of connecting components on the folding platform are jointly provided with a telescopic mechanism, and the frame is further provided with a driving mechanism for driving the telescopic mechanism. The operation platform is simple in structure and high in reliability, and the problem that an operation platform is needed for in-orbit unmanned cargo taking and placing of a freight satellite or in-orbit maintenance of an aircraft is effectively solved.

The invention discloses a spacecraft relative motion obstacle avoidance control method based on an anti-interference control obstacle function, and the method comprises the steps: 1, initializing a spacecraft system state and sampling time through building a spacecraft relative motion model with external disturbance; 2, designing an unknown dynamic estimator, and adopting a low-pass filter to obtain estimation of external disturbance; 3, constructing a relative motion tracking error of the spacecraft, designing a Lyapunov function, and designing a nominal controller by adopting a Lyapunov stability theory; and 4, constructing an obstacle function of the spacecraft and the obstacle, designing an obstacle avoidance controller according to the safety constraint of the control obstacle function, and finally integrating the obstacle avoidance controller to actual control input. According to the invention, formalized guarantee of safety is realized, and a robust item based on an unknown dynamic estimator is embedded into a control obstacle function constraint condition, so that a system state can be strictly prevented from entering a dangerous area even under continuous disturbance, and a provable obstacle avoidance method is provided.

Spacecraft systems, methods and devices comprising satellite upgrade modules or hosted payload modules that facilitate the deployment of small, deployable satellites from a base station to perform various potential missions including, for example, space domain awareness (SDA) missions to inspect and characterize resident space objects in orbit.

The invention provides a linear rotation two-degree-of-freedom space momentum wheel with a mass center position adjusting function, which is characterized by comprising an outer stator, a rotor, an inner stator, a rotation driving unit, a linear driving unit, a sliding bearing, an outer iron core type LVDT (Linear Variable Differential Transformer) and a linear rotation two-degree-of-freedom eddy current damper, the outer stator, the rotor and the inner stator are arranged layer by layer from outside to inside, the outer stator and the rotor are provided with the rotary driving unit and the sliding bearing, and the linear driving unit and the outer iron core type LVDT are arranged between the inner stator and the rotor; linear rotation two-degree-of-freedom eddy current dampers are installed on the side faces of the outer stator, the rotor and the inner stator, and circumferential damping force or axial damping force opposite to the movement direction of the rotor is allowed to be generated through the linear rotation two-degree-of-freedom eddy current dampers. The momentum wheel has two degrees of freedom of straight line and rotation, and the rotating wheel body (rotor) can move in the axial direction, so that the mass center of a spacecraft system is changed, and disturbance torque is minimized.

To provide a spacecraft system for achieving highly efficient power transmission.SOLUTION: The system includes a photovoltaic power generation satellite having a solar panel that receives sunlight to obtain electric power, an electromagnetic wave transmission device that converts the electric power into an electromagnetic wave and transmits the electromagnetic wave, and a control device that controls a trajectory and an attitude control device and a transmission angle of the electromagnetic wave transmission device and is installed on a space side, and a ground facility 3 having a reception facility and an information processing device that receives the electromagnetic wave from the electromagnetic wave transmission device and converts the electromagnetic wave into electric power and is installed on a ground side. The information processing device calculates a sun-satellite-receiving facility angle formed by the sun and the receiving facility around the solar power generation satellite based on the position of the receiving facility, the orbit information of the solar power generation satellite, and the position of the sun, calculates the amount of power received by the receiving facility based on the position of the receiving facility, the attitude of the solar power generation satellite, the orbit information of the solar power generation satellite, and the position and the angle of the sun to calculate and output the orbit or the attitude of the solar power generation satellite, and controls the orbit or the attitude using the orbit and attitude control device.SELECTED DRAWING: Figure 1

The invention belongs to the technical field of spacecraft attitude control, and particularly relates to a control-distribution integrated full-distributed spacecraft attitude cooperative control method. The method comprises the following steps: step 1, establishing a discrete linear spacecraft attitude kinematics and dynamics model; 2, designing a combined spacecraft system model prediction control method; and 3, solving the overall optimization problem of the combined spacecraft system based on model predictive control in the step 2 by adopting a completely distributed attitude control distribution method. According to the method, a central controller does not need to design a total control law, distribution errors in a double-layer controller structure mode can be avoided, and the calculation and solving process of a control strategy is simplified; on the basis of a communication network diagram, under the condition that only information of adjacent modules is received, all the modules solve self control strategies, complete distribution is achieved, and all the satellite modules cooperate with one another to achieve overall optimization of performance indexes of a combined spacecraft system.

Systems and methods related to optical navigation systems for spacecrafts are disclosed herein. According to the optical navigation system, several types of sensors are combined into a compact optical sensor system, and the function of each type of sensor is kept. An optical navigation system uses optical sensors, artificial intelligence, automatic dimming, and other functions to perform navigation functions, such as sun sensing, horizon sensing, and star tracking. Artificial intelligence may include deep neural networks that can provide complex machine vision applications after being trained. The optical navigation system may be organized into modular slices to increase customizability. The types of slices may include optical slices, network switched slices, visual navigation slices, and inertial navigation system switched slices. Software packages for the navigation system are also customizable, reducing user cost and complexity to the extent required only by a given application.

This invention discloses a disturbance estimation method for spacecraft systems with arbitrary suppression strength of measurement noise, belonging to the field of disturbance estimation technology for spacecraft systems. The method includes: Step 1: Establishing a nonlinear model of the spacecraft system considering measurement noise, using this model to accurately characterize the dynamic behavior of the spacecraft in complex environments; Step 2: Establishing a variable-order filter with independent parameters. The designed filter has the same cascaded structure as the observer, which makes the filter and observer parameters independent and uncoupled; Step 3: For the observer parameter adjustment problem, the method of undetermined coefficients is adopted. By solving the characteristic equation of the proposed filter's ESO and the characteristic equation of the desired pole placement, the coefficients of the two equations are compared to determine the observer parameters. In this method, the filter parameters increase with the filter order, and the sum of the observer parameters equals the observer order, avoiding the problem of over-parameterization and achieving arbitrary pole placement.

The invention discloses a multi-spacecraft distributed energy storage consistency control method, and belongs to the field of spacecraft power supply control. The invention aims to solve the problem of energy and power coordination control when the battery capacity and the battery available power of each spacecraft in a spacecraft system are different. Comprising the following steps: when an unknown load is applied to a distributed system of multiple spacecrafts, a leader node of the distributed system calculates the total power of the current load according to the newly added output power of a node generating additional output power, and distributes output power to each node according to a power distribution coefficient; based on communication topology among spacecrafts, each node calculates and obtains the average energy storage SOC of the system in a distributed mode through information interaction of energy storage SOCs among neighbor nodes; and each node adjusts the output power according to the difference value between the average energy storage SOC of the system and the energy storage SOC of the node, so that the energy storage SOC of each node tends to be consistent, and the energy storage consistency control is realized. According to the invention, energy storage consistency control between spacecrafts is realized.

The application relates to a spacecraft system simulation collision damage judgment system and method based on a real-time judgment process, which comprises the following: a spacecraft system simulation business calculation module, which is used for calculation and output according to business demand and task targets, and realizes external data and clock transceiving through a spacecraft system simulation business calculation service; a plurality of damage judgment task modules; a simulation display module, which is used for interface design of input data and interface display of output data, and realizes external data transceiving and clock synchronization through a simulation display service; and a simulation scheduling support framework service module, which is interactively connected with the spacecraft system simulation business calculation module, the damage judgment task modules and the simulation display module, and is used for data receiving and distribution scheduling and clock synchronization control of the system. The application can form quantitative evaluation of target damage and objectively improve the evaluation result.