62 results about "Spacecraft system" patented technology

The invention discloses a satellite-ground integrated cooperative efficient alignment method based on a large-view-field optical load, relates to the technical field of spacecraft systems, and solves the problems that in the prior art, precision is low, response is slow, the requirements of remote sensing satellite high-speed data transmission for pointing precision and link availability are met, and satellite data transmission efficiency is improved. Comprising the following steps that S1, when a satellite enters a data transmission mode, a laser ground station opens beacon light, and a satellite center computer executes an alignment program; s2, satellite alignment is completed, and a data transmission link between the satellite and the laser ground station is established; and S3, the satellite carries out a data transmission test, and when the data transmission test is completed, the satellite and laser ground station data transmission link is quitted, and the operation is ended.

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.

The invention provides a multi-spacecraft leader following formation control method based on neural adaptation. The method comprises the following steps: firstly, establishing a multi-space aircraft system model with a leader-follower structure; then, designing a robust formation control scheme based on neural self-adaption; firstly, designing a self-adaptive distributed state observer based on local discontinuous interaction for a follower spacecraft; secondly, a distributed robust formation controller based on neural network fitting is designed for the follower spacecraft; and finally, designing an inter-aircraft communication mechanism based on a static event triggering mechanism for the follower spacecraft. According to the method, the problems of local knowing of leader information, limited communication, adverse influence of external disturbance on control performance and the like in the multi-spacecraft formation with a leader-following structure are considered, an effective formation control scheme is designed, and the robustness and reliability of a multi-spacecraft system are remarkably improved.

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 spacecraft system disturbance estimation method with random suppression intensity of measured noise, which belongs to the technical field of disturbance estimation of spacecraft systems, and comprises the following steps of: 1, establishing a nonlinear model of a spacecraft system considering the measured noise, and accurately representing the dynamic behavior of a spacecraft in a complex environment by using the model; 2, a variable-order filter with independent parameters is established, the designed filter has the same cascade structure as the observer, and therefore the parameters of the filter and the observer are mutually independent and not coupled; and step 3, aiming at an observer parameter adjustment problem, adopting an undetermined coefficient method, solving a proposed ESO characteristic equation of filtering and a characteristic equation of an expected configured pole, carrying out coefficient comparison on the two equations, and determining an observer parameter. According to the method, the parameters of the filter are increased along with the increase of the order of the filter, and the sum of the parameters of the observer is equal to the order of the observer, so that the problem of excessive parameterization is avoided, and arbitrary pole assignment is realized.

The invention provides a universal test platform for a GNC (Global Navigation Control) system of a multi-cabin spacecraft, which relates to spacecraft GNC data processing and comprises an application software module suitable for the multi-cabin spacecraft, a dynamics software module, a component interface simulation module, an injection management module and an operation monitoring module, wherein the application software module is used for carrying out universal modification on satellite application software so as to enable the satellite application software to carry out data communication with other modules; the dynamics software module is used for updating state information according to the control instruction; the component interface simulation module is used for performing data communication between the application software module and the dynamics software module; the injection management module is used for controlling the general test platform to complete a simulation test according to the test instruction; the operation monitoring module is used for monitoring telemetry data output by the application software module and the dynamics software module. According to the scheme, the problems of long time consumption and low efficiency of simulation verification of the multi-cabin spacecraft by a traditional hardware platform can be solved.

This invention discloses a multifunctional satellite constellation configuration design optimization method based on a large language model, belonging to the field of spacecraft system design technology. The method includes: setting design variables and fixed variables, where the design variables include the orbital inclination and orbital altitude of satellites in the satellite constellation; constructing an evaluation function based on coverage percentage, ground pixel resolution, average number of visible satellites, and geometric dilution of precision; iteratively generating solutions based on the design variables and fixed variables using the large language model, evaluating the solutions using the evaluation function, and obtaining an optimal satellite constellation configuration design scheme based on the evaluation results. Through the above technical solution, the present invention can provide a satellite constellation configuration design optimization method that takes into account multiple optimization aspects.

A high-efficiency alignment method based on a large field-of-view optical payload, involving spacecraft system technology, solves the problems of low accuracy and slow response in existing technologies, while meeting the requirements of high-speed data transmission from remote sensing satellites for pointing accuracy and link availability, thus improving satellite data transmission efficiency. The method includes the following steps: Step S1, when the satellite enters data transmission mode, the laser ground station turns on the beacon light, and the satellite center computer executes the alignment procedure; Step S2, the satellite alignment is completed, and a data transmission link is established between the satellite and the laser ground station; Step S3, the satellite conducts a data transmission test, and when the test is completed, the data transmission link between the satellite and the laser ground station is terminated, ending the operation.

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 application discloses a predefined time fault-tolerant control method for a rigid spacecraft system and relates to the technical field of aerospace flight control, and comprises the following steps: S1, a kinematics and dynamics model of a fault-tolerant control system of a rigid spacecraft attitude tracking error is established, and system states and control parameters are initialized; S2, a non-singular predefined time sliding mode surface is designed based on an attitude error tracking control system of the rigid spacecraft; and S3, a non-singular adaptive fault-tolerant predefined time controller is designed. The application designs a predefined time preset performance function and a non-singular adaptive fault-tolerant predefined time control method, effectively solves a potential chattering problem in a controller design process, guarantees the steady-state performance and state constraints of the attitude tracking error and guarantees the predefined time convergence of the system in the case that the controller fault, inertia uncertainty and external disturbance exist simultaneously.

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.

A method for intelligently suppressing micro-vibrations of flywheel assemblies for on-orbit satellites. The method relates to the technical field of spacecraft systems, and in particular to the technical field of intelligently suppressing micro-vibrations of flywheel assemblies for on-orbit satellites. The method effectively solves the problem of the lack of quantitative indicators of vibration isolation effects in the temperature control process of a small speed condition coverage of the flywheel assembly, thereby improving computational efficiency and accuracy. The method comprises the following steps: extracting frequency domain disturbance data of the flywheel at a specified speed; obtaining the displacement response of the sensitive component in the frequency domain under unit disturbance; obtaining the displacement response of the sensitive component in the time domain under the disturbance of the flywheel assembly; calculating the actual image shift value; calculating the quantitative characterization of the performance of the vibration isolator under the specified speed condition through the actual image shift value; constructing and training a "speed condition-isolator performance" proxy model; and intelligently controlling the intelligent suppression of micro-vibrations of the flywheel assembly of the on-orbit satellite.

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.

An apparatus for generating a 3D image map comprises a camera, mounted to a spacecraft, configured to capture plurality of images of an object from a plurality of positions about the object. An unstructured light source for illuminating the object to capture the plurality of images of the object. A gimbal mechanism for mounting the camera and the unstructured light source and moving the camera to a plurality of positions. Image processing circuitry configured to process the plurality of images of the object and generate a 3D image map of the of the object responsive to the plurality of image of the object. The image processing circuitry further having a processor configured to implement a shape-from-motion algorithm for generating the 3D image map responsive to the plurality of images of the object.

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.

An apparatus for generating a 3D image map comprises a camera, mounted to a spacecraft, configured to capture plurality of images of an object from a plurality of positions about the object. An unstructured light source for illuminating the object to capture the plurality of images of the object. A gimbal mechanism for mounting the camera and the unstructured light source and moving the camera to a plurality of positions. Image processing circuitry configured to process the plurality of images of the object and generate a 3D image map of the of the object responsive to the plurality of image of the object. The image processing circuitry further having a processor configured to implement a shape-from-motion algorithm for generating the 3D image map responsive to the plurality of images of the object.