80 results about "Spacecraft system" patented technology

Systems, apparatus, and method for momentum management for a spacecraft are disclosed. An example spacecraft includes a body; an appendage moveably coupled to the body; a thruster carried by the body; machine-readable instructions; and processor circuitry to execute the machine-readable instructions to detect a change in angular momentum at the spacecraft during firing of the thruster; and in response to the detection, cause the appendage to move relative to the body to adjust a center of mass of the spacecraft relative to a net thrust vector associated with the thruster.

The invention discloses a communication optical axis autonomous correction method based on tracking point nutation, which comprises the following steps: controlling a fine-following fast reflector to scan, monitoring the output power of an erbium-doped fiber amplifier (EDFA) within scanning residence time, searching the position of the maximum power by adopting a sliding window average algorithm, and recording the position; and after scanning for one period, returning to the tracking point with the maximum recorded power, thereby realizing rapid and accurate calibration of the tracking optical axis of the laser communication terminal. According to the invention, a closed loop is formed based on tracking point nutation and a pre-amplification output signal, so that high-precision correction of a camera tracking point is realized, the stability of precise tracking of the system can be improved, and the signal-to-noise ratio received by the laser communication system can also be improved; the traditional receiving optical axis calibration convergence speed and precision are improved; higher real-time performance and automation degree are achieved, relevant data can be acquired and processed in real time, and correction of a receiving optical axis is achieved; the method has wide applicability and expandability, and can be applied to other spacecraft systems needing high-precision tracking control.

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.

The present invention provides a pure three-axis magnetic control preset performance attitude control method for an on-orbit spacecraft, comprising the following steps: Step 1: providing a spacecraft control torque calculation formula when a single magnetic torquer is used as an actuator; Step 2: providing a mathematical model of an attitude tracking system; Step 3: designing an attitude controller with preset performance to complete the attitude control with preset performance. The beneficial effects of the present invention are as follows: the pure three-axis magnetic control preset performance attitude control method for an on-orbit spacecraft of the present invention only uses a magnetic torquer as an actuator, and considering the requirements for static and dynamic performance required by the spacecraft in an actual working environment, a preset performance controller is designed to achieve the control of the dynamic steady-state performance of the attitude response of the spacecraft, thereby improving the safety and performance reliability of the magnetic control spacecraft system, and having broad application prospects in the fields of spacecraft docking and separation, on-orbit assembly, formation flying, rescue of faulty spacecraft, and de-orbiting of abandoned spacecraft.

The invention discloses a predefined time fault-tolerant control method for a rigid spacecraft system, and relates to the technical field of aerospace flight control, and the method comprises the following steps: S1, building kinematics and dynamics models of a fault-tolerant control system for attitude tracking errors of a rigid spacecraft, and initializing a system state and control parameters; s2, designing a non-singular predefined time sliding mode surface based on an attitude error tracking control system of the rigid spacecraft; and S3, designing a non-singular adaptive fault-tolerant predefined time controller, and effectively solving the potential buffeting problem in the design process of the controller based on the predefined time preset performance function and the non-singular adaptive fault-tolerant predefined time control method. Under the condition that controller faults, inertia uncertainty and external interference exist at the same time, the steady-state performance and state constraint of attitude tracking errors are guaranteed, and system predefined time convergence is guaranteed.

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 relates to a simulation deduction-oriented scene breakpoint reset scheduling method, which comprises the following steps of: S1, slicing simulation deduction data according to time periods, and compressing and storing the slices based on the simulation deduction data; s2, identifying a key breakpoint of the simulation scene; s3, loading a simulation scene data fragment where the key breakpoint is located to a simulation engine for pre-deduction to obtain pre-deduction data; and S4, carrying out simulation scene baseline merging. The method is suitable for debugging the simulation scene of the large-scale spacecraft system, and can solve the problem that the simulation scene of the large-scale spacecraft system cannot be effectively debugged.

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.

A spacecraft attitude preset performance control method based on an improved preset finite time function comprises the following steps: step 1, determining a control object and a control target according to a motion law of a spacecraft: designing a control law so that the control object and the control target are controlled under the influence of saturation factors of an execution mechanism; a spacecraft system attitude quaternion error vector part qev and an angular velocity error We of a spacecraft body coordinate system track an expected signal; 2, constructing a class sliding mode variable S, a self-adaptive boundary preset finite time function and auxiliary variables e and v; 3, designing a control law to enable the variable S to move in an envelope formed by a preset finite time function # imgabs0 #; and 4, checking and improving the performance of the preset finite time controller. According to the method, the problem of singularity easily occurring when an existing preset finite time control method faces internal and external interference and the problem that the angular velocity is too large in the large-angle maneuvering process are solved, meanwhile, the actual physical limitation of an executing mechanism is considered, and it can be guaranteed that higher reliability is obtained.

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.

A method for node sensitivity analysis affecting the degradation of the communication performance of a giant constellation. The present invention relates to the field of spacecraft system simulation. It solves the problem that in the prior art, there is a lack of a method for measuring the sensitivity of the constellation to changes in the number of satellites, which brings changes in performance such as communication delay. The method includes: reading the parameter values under different two-line element data segments according to the format of the two-line element data record and saving them to the SGP4 orbit model structure variable; dividing different sub-constellations within the giant constellation according to the orbital inclination or the semi-major axis of the orbit; setting that the satellites within each orbital plane form a routing domain, setting the gateway on / off condition, setting that the satellite nodes can communicate within the domain, and restricting the connectivity of the nodes in different domains; updating the connectivity between the nodes, outputting the on / off situation and delay between the nodes; and performing routing and outputting the routing result to complete the node sensitivity analysis. It is also applicable to the field of measuring the communication delay of giant constellations.

The invention discloses a multi-field fusion digital prototype development system and a digital prototype construction method thereof, and belongs to the technical field of digital prototypes. According to the method, by introducing the FACE architecture, professional models in other fields are subjected to virtual modeling again, and multi-scale and multi-dimensional models are unified, so that the professional models in different subject fields can be integrated comprehensively, and a complete digital prototype model is obtained. Meanwhile, after virtualization modeling, the model can be decoupled from the operation platform, and multiplexing of the model on different operation platforms is realized, so that the model can seamlessly run through a system design stage and a system implementation stage, repeated development and test of the model are avoided, and research and development efficiency of a spacecraft system can be further improved.

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 non-zero and dynamic Stackelberg pursuit game control method for a spacecraft system, and aims to solve the pursuit game problem of a spacecraft in the fields of space exploration, military reconnaissance and the like. According to the method, a traditional zero-sum game model is broken through, a non-zero-sum dynamic game framework is adopted, and the cooperation and competition relation between purchasers and escapers is reflected more practically. Effective control of a spacecraft pursuit game is realized through five steps of establishing a mathematical model, designing a control strategy, converting a cost function, solving a matrix Riccati equation and solving an optimal control gain.

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.

The present invention provides a simulation method, system and device for a spacecraft system driven by a time-event hybrid, belonging to the field of aerospace technology. First, a simulation framework is designed by combining time-driven and event-driven. Then, a control instruction set is established: the control instruction set is divided into a general instruction set and a dynamic instruction set according to the urgency of the event occurrence. Next, a component model of the spacecraft simulation system is established, and the solution order of each model at the same time point is determined. Subsequently, an instruction distribution model and an instruction execution model are established. Finally, based on the time-driven software operation framework, solutions are carried out at each time point to complete the simulation of the spacecraft system driven by the time-event hybrid.

Intelligent In-orbit Micro-vibration Suppression System for Satellites. It belongs to the technical field of spacecraft systems, specifically involving the technical field of vibration suppression for remote sensing satellites. It solves the problem that the existing variable-parameter satellite vibration isolators cannot be actually controlled and applied. The system includes a variable-parameter satellite vibration isolator, a central control unit, and a power distribution and thermal control unit; the central control unit: includes a rotational speed acquisition module, a parameter calculation module, and a temperature control module; the power distribution and thermal control unit: includes a first power distribution and thermal control unit and a second power distribution and thermal control unit; the second power distribution and thermal control unit receives the temperature control command from the temperature control module and supplies power to the heating wire in the variable-parameter satellite vibration isolator according to the temperature control command. The system of the present invention can be applied in the fields of remote sensing satellite assembly and manufacturing, remote sensing satellite vibration suppression, and remote sensing satellite imaging system design.

The invention discloses a manned spacecraft distributor circuit modeling method based on a new generation power distribution switch, and belongs to the field of spacecraft system simulation. According to requirements of system energy matching simulation, the method realizes simulation of working steady-state characteristics and a dynamic process of the distributor by establishing a contactor switch module, an SSPC intelligent switch module and a signal interface module, and adopts a modular modeling method to realize simulation of the working steady-state characteristics and the dynamic process of the distributor. On the basis of an established contactor switch module, an SSPC intelligent switch module and a signal interface module, a power distributor model is constructed according to system construction, and can be used for accurately simulating complex characteristics of a novel power distributor.

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.