195 results about "Spacecraft system" patented technology

The invention discloses an embedded real-time emulation and fault simulation system based on multiple data buses, comprising upper computer simulation management software and an embedded real-time emulation and fault simulation device. The upper computer simulation management software guides the user to perform emulation configuration and fault setting aiming at a fault-tolerant control system to be tested through a human-computer interface, and generates fault emulation data of multiple redundancy equipment. The embedded real-time emulation and fault simulation device outputs the fault emulation data in real-time and synchronously through the multiple data buses, simulates the fault of the multiple redundancy equipment and outputs the fault data stream. The upper computer and a lower computer communicate mutually through an Ethernet to achieve the downloading of the information configuration, the fault emulation data and the emulation control command of the upper computer, and the uploading of the working condition of the lower computer. The invention overcomes the defects of poor universality and lower synchronous precision of the traditional multiple equipment fault simulation device based on the emulation, and is suitable for the test and the verification of the different redundancy fault-tolerant designs of the complex spacecraft system with numerous equipment and frequent data exchange.

The invention discloses a ground testing system for on-track micro vibration of a spacecraft, which comprises a vibration isolating foundation, a supporting truss with a framework structure, a suspended positioning platform, a gravity suspended compensation unit, an integrated control subsystem, a micro vibration measurement subsystem and a test data processing and analyzing subsystem, wherein the vibration isolating foundation is connected with the ground, the supporting truss with the framework structure is fixedly supported on the vibration isolating foundation, the suspended positioning platform is supported on the supporting truss, the gravity suspended compensation unit is hung on the suspended positioning platform, and the integrated control subsystem is used for controlling the positioning and the fixation of the suspended positioning platform via coordination of sensors in various positions and performing the overall integrated control on the state of the system. By adopting the system, the ground simulation of the on-track micro vibration environment, the measurement of micro vibration and the data processing can be realized for the spacecraft at the system level and the component level, the technical level of testing the micro vibration of the spacecraft can be improved, and the demands for testing and verification of the on-track micro vibration environment in research on the spacecraft, particularly the high-precision satellite, can be satisfied.

A network architecture and protocol provides “plug and play” spacecraft capabilities. A distributed-control architecture is used, wherein each component operates semi-autonomously, and interacts with other components on a task / resource level. Each component announces its requirement for system resources as a request to the network, and components that can provide some or all of the requested resources respond to the request. An arbitration device centralizes and coordinates requests for critical and / or singular resources, such as requests for a specific orientation of the spacecraft. To facilitate such distributed control, requests are made in advance of the requirement for the resource, and include a time interval during which the resource is required. A configuration and test system is provided to process the mission requirements and provide a set of components that can be configured to satisfy the requirements.

The invention discloses a distributed spacecraft ground simulation system and a method for realizing same, belonging to the technical field of space flight and aviation. The invention aims at solving the problems that the configuration of the simulation platform of a prior distributed spacecraft is solidified and the expandability is bad. The invention adopts a one-dimensional rotational base air floated platform which is based on a plane two-dimensional parallel motion and is perpendicular to a parallel motion plane. By configuring a cooling air thrustor and a counteractive flywheel as an executing mechanism and configuring a high-precision optical fibre gyro, an x-axle accelerometer and a Y-axle accelerometer as a sensing component, the definition of a high-precision relative position and the standardization of an initiate posture can be realized by a high-precision local GPS positioning system, and other hardware systems can be configured according to different tasks, thereby forming a multi-spacecraft ground simulation system. The invention can realize the ground simulation multi-spacecraft system by configuring different practical hardwares or simulators according to the task of the spacecraft, thereby having strong expansion capability and applicability.

The invention relates to a method for controlling an index time-varying slide mode of flexible spacecraft characteristic shaft attitude maneuver, and belongs to the technical field of spacecraft control. The method comprises the steps that firstly, a system dynamically equivalent model, a dynamic model and a flexible vibration model are established under a spacecraft system, then, the vibration frequency and the damping ratio parameter of a closed loop system with the index time-varying slide mode control law are calculated, and a single-shaft multi-modality filtering input shaping device with a characteristic shaft as a rotary shaft is designed according to the designing method of the single-shaft input shaping device to restrain flexible vibration in three-shaft motion. Meanwhile, a state observer is designed to estimate flexible modal information in real time, and the method for controlling an output feedback index time-varying slide mode is formed. At last, saturability analysis is conducted on control torque so as to satisfy the physical saturation constraint of the control torque. By means of the method, the application range of existing input shaping is expanded, the input shaping technology is expanded from single-shaft maneuver to three-shaft maneuver, the self-robustness of filter input shaping is enhanced, and the purpose that the attitude maneuver path of the spacecraft is the shortest is achieved.

A spacecraft system and method includes a platform with a dock and an umbilical payout device. A robot is connected to an umbilical paid out by the umbilical payout device and is repeatedly deployable from the dock. The robot includes one or more imagers, an inertial measurement unit, and a plurality of thrusters. A command module receives image data from the one or more robot imagers and orientation data from the inertial measurement unit. An object recognition module is configured to recognize one or more objects from the received image data. The command module determines the robot's orientation with respect to an object and issues thruster control commands to control movement of the robot based on the robot's orientation. The combination of the space platform and robot on umbilical line can be used for towing another object to different orbital location, inspection including self-inspection of the robot carrying platform and for robotic servicing.

The invention discloses a self-correcting redundancy switching mechanism for a spacecraft system and a verification method thereof. The invention adds the redundancy switching mechanism with the feature of self-correcting into the spacecraft system and provides the verification method thereof. The subsystems of the each spacecraft are communicated by a 1553B bus, and are respectively communicatedwith a ground monitor and control system by an Ethernet communication module. The ground monitor and control system immediately detects the fault state of the spacecraft system, and injects fault as well as a control instruction, thereby being capable of significantly improving the high efficiency and the safety of the redundancy switching mechanism of a complex system. By accumulating and optimizing expert opinions, the invention can reduce unnecessary backup switch, has significant research meaning for improving the long-time endurance capability, can immediately switch the fault of a key sensitive part, prevents from an uncontrolled state, and guarantees the safe operation of the system. The method has significant meaning for realizing the reliability research of a spacecraft autonomous intelligent control system.

The invention discloses a 1553B bus network emulation system applied to a spacecraft system, which belongs to the field of aerospace and computer emulation. The 1553B bus network emulation system comprises a 1553B bus, the Ethernet, a bus network tester, a bus terminal tester, a digital phosphor oscilloscope, a bus separation switch, an emulation terminal, a bus system performance estimation computer and a bus interface control file configuration computer, and is used for emulating and analyzing the information transmission and system running of an actual spacecraft bus system. The 1553B bus network emulation system replaces conventional cable connection with a 1553B bus network, realizes the digitalized transmission and comprehensive utilization of network system information and improves the reliability and electromagnetic compatibility of the system by equipment with standard digital interfaces in combination with Ethernet communication, meets 1553B specifications, realizes coexistence in one network emulation system, can be used for a plurality of stages of analysis, development, test and the like of the spacecraft system and is of great importance to researches on the reliability of the actual spacecraft system.

The invention discloses an attitude control method for a spacecraft networked system based on event trigger, belonging to the field of system servo control for space robot networked control systems. The method comprises the following steps: firstly arranging a suitable transition process for the desired attitude of the system by designing a tracking differentiator to avoid a serious overshoot on the output of the system caused by an over-large initial error, and at the same time, acquiring a differential signal of the desired value for preparation for subsequent controller design; then, considering the network transmission signals, designing an event-triggered expansion state observer by using an attitude-angle measurement signal output by an event trigger mechanism, estimating the nonlinear uncertainties formed by the state in the spacecraft system, coupling, external disturbances and the like in real time, and compensating the estimated values of the nonlinear uncertainties into an error feedback control rate to form a composite controller. Finally, the method provided by the invention prevents nonlinear factors such as internal and external interferences from adversely affectingthe system, improves the robustness of the system, and provides guarantee for the smooth completion of space operations tasks.

An improved radiation dosimeter device in an improved radiation dosimeter system provides a DC analog output voltage that is proportional to the total ionizing dose accumulated as a function of time at the location of the dosimeter in a host spacecraft, so as to operate in a system bus voltage range common to spacecraft systems with the output being compatible with conventional spacecraft analog inputs, while the total dose is measured precisely by continually monitoring the energy deposited in a silicon test mass accumulating charge including charge contribution prior to radiation threshold detection for improved measurement of the total accumulated charge with the dosimeters being daisy-chained and distributed about the spacecraft for providing a spacecraft dose profile about the spacecraft using the improved radiation dosimeter system.

The invention belongs to the technical field of spacecraft systems and is mainly applied to liquid propulsion systems or solid-liquid mixing propulsion systems adopting high-saturated vapor pressure propulsive agents. In order to solve the problems that the temperature of propulsive agents and the pressure in a storage box are seriously reduced in the working process of a self-pressurization supply system, a self-pressurization scheme based on regenerative cooling is provided. The self-pressurization supply system is composed of a safety valve, a storage box, a filling pipeline, a filling valve, a one-way valve, a cooling channel, a liquid accumulating cavity, a main supply pipeline, a coolant pipeline, a burning agent pipeline, a main valve, a pressurization pipeline and the like. Propulsive agents cool the wall of a combustor and acquire heat at the same time. The propulsive agents subjected to heat exchange and vaporization return to the storage box to pressurize the storage box. According to the self-pressurization supply system based on regenerative cooling, the pressure intensity of the storage box can be basically maintained to be stable before the liquid propulsive agents in the storage box are completely discharged, and accordingly the effect that the thrust is basically stable is ensured.

The invention discloses a model-based automated test method for a spacecraft system and belongs to the field of system testing. The method comprises the steps of firstly building a window tree model of a tested spacecraft system through a formalization method, classifying behavior states of the tested spacecraft system during operation according to business processes, and constructing a system behavior state transition diagram; secondly traversing a window tree to generate a test case set, running the test case set in a tested spacecraft digital model and validation prototype system, setting tracking points, obtaining system behavior states and recording a state transition sequence; and finally checking whether the obtained state transition sequence obeys a state transition rule or not through a model detection algorithm, thereby automatically validating overall design of the tested spacecraft system and interface design among modules. According to the method, the test cost of the spacecraft system and the whole research and development cycle of the spacecraft system are reduced and shortened by adopting a model-based test technology; and the quality and accuracy of spacecraft system testing are improved by adopting the formalization method and the model detection algorithm.

The invention provides an active disturbance rejection control method for cluster spacecraft electromagnetic traction, complexity of a cluster spacecraft system and model uncertainty are fully considered, an electromagnetic force / torque model and a relative motion dynamics equation are established, so that the active disturbance rejection control method formed by a most speed tracking differentiator, a non-linear extension observer and a non-linear feedback control law has good working performance, a coupled problem of a dynamic model is effectively solved, high precise electromagnetic traction control can be realized, and the effectiveness and the Lyapunov asymptotic stability of the method are verified; compared with the control performance of a finite time control method, the control method has better rapidity, vibration resistance ability and disturbance rejection, and the control method is simple in design and has engineering realizability.

The invention discloses an imaging type method for correcting earth oblateness by an earth sensor, which comprises the following steps: extracting a coordinate of an image point of an earth contour point on a photoelectric detector of the sensor and computing a unit vector of the image point in a spacecraft system; assuming the earth is a standard sphere, computing a geocentric vector VEB, and acquiring a rolling angle and a pitch angle Theta of the spacecraft according to the geocentric vector VEB; estimating a yaw angle Psi of the spacecraft at the current moment according to the posture of the spacecraft at previous moment, simultaneously acquiring a current position and a coordinate system transformation matrix of the spacecraft from a spacecraft posture orbit control system, and computing an azimuth angle Sigma of the earth contour point; computing corresponding viewing angle radius Eta according to the azimuth angle Sigma of the earth contour point; and finally, using the viewing angle radius Eta of the earth contour point to remove the influence on the earth oblateness, and acquiring the corrected geocentric vector. The method has simple operation, and can provide more accurate geocentric vector.

Provided is an anti-interference attitude control method based on the friction characteristics of a reaction wheel. To solve the problems about zero-crossing friction and low attitude control precision of a reaction wheel, the method includes the steps: establishing a spacecraft coupling kinetic equation containing the friction characteristics of the reaction wheel, designing a friction interference estimator because the reaction wheel brings friction interference, and estimating the friction of the reaction wheel in real time; designing a norminal controller to inhibit friction interference estimated errors and environment interference in a spacecraft system; and combining the norminal controller with the friction interference estimator, and designing a composite layered anti-interference controller to achieve spacecraft anti-interference attitude control under the influence of multi-source interference. According to the method, actuating mechanism dynamics are additionally applied to the spacecraft kinetic equation so as to better analyze the problem of the decreased control precision due to the characteristics of an actuating mechanism. The method can be used for high-precision attitude control over high-precision earth observation satellites, space telescopes and other spacecrafts in the aerospace field.

The invention discloses an attitude control and fuel sloshing suppression method for a liquid-filled spacecraft system. The method includes the steps that a model of the liquid-filled spacecraft system is built according to the Lagrange-Euler equation, and a reduced-order model of the system is obtained and divided into two subsystems to be respectively designed; the feedback control design is conducted on the first subsystem, and the state of the system serves as external input of the second subsystem; the second subsystem represented in a form of an under-actuated system is made into a standard form, the characteristics of the system are analyzed, and then a controller is designed through a sliding mode control method suitable for the under-actuated system. When liquid in a storage tank of a spacecraft is sloshed, attitude control over the spacecraft and liquid sloshing suppression can be achieved at the same time, eventually the whole spacecraft system can be kept stable to finish preset actuation actions, and therefore flight safety of the spacecraft is guaranteed.

The invention discloses an electric signal feature extraction and recognition system oriented to aircraft flying process, particularly relates to a collection, processing, extraction and recognition system for electric signals of an aircraft. The system takes the electric signals of the aircraft as an object, is based on the process oriented to flying tasks and is suitable for convention testing of the aircraft at each stage and in various modes and combined testing of multiple aircrafts. By the system, dynamic data, which are more comprehensive and are based on the flying process, of the electric signals can be accumulated at a comprehensive testing stage and serve as a criterion to judge whether electric performance of the aircraft is normal or not, and the dynamic data exceed corresponding remote measuring indicators of a tested object in the aspects of accuracy and instantaneity; as a key link for verifying the aircraft, compared with conventional remote measuring of the aircraft, monitoring accuracy and rate of the system are improved remarkably on a number level, an efficient means is provided for comprehensive testing of the aircraft oriented for complex flying process, and level of judging electric performance working state of the aircraft on a system level is improved.

The invention discloses a spacecraft multi-target heuristic ordering task planning method and belongs to the technical field of aerospace. The method describes the composition, the resources, the subsystem functions and various constraint conditions required to be satisfied of a spacecraft according to subsystem internal constraint characteristics and in consideration of spacecraft structure, taskrequirements, equipment status and spacecraft capabilities, in view of the complex constraint of the spacecraft system and the coupling of system state information, describes multiple parallel subsystems of the spacecraft by using a timeline, establishes the internal state transition diagram of the subsystem, performs multi-target heuristic ordering task planning according to the constraint information and the state transition diagram between tasks, determines the orders of the tasks according to the symbol of an evaluation value, guides a planning search direction according to a multi-targetheuristic ordering result, finally outputs a multi-target heuristic ordering task planning solution, that is completes task planning of the spacecraft, reduces search space and improves the efficiency of task planning.

The invention discloses a sputtering target device for reducing the plume sputtering pollution of an electric thruster during the system-level ignition test on a spacecraft powered on by the electric thruster. The device structurally comprises two parts, namely a sputtering main target and a shielding cover. The main target is directly facing the nozzle of the thruster. The material of the main target is composed of the low-sputtering-rate material graphite. The back of the main target is provided with a pipeline system. The pipeline system is in close contact with the main target, so that the pipeline system is used for cooling the main target through the cooling liquid. The pipeline system is also used for heating the main target through the hot water. The shielding cover is arranged on the main target and is used for guiding and capturing sputtering product particles, reducing the escaping probability of the sputtering product particles, and lowering the sputtering pollution. Meanwhile, the requirement of the vacuum degree is considered. The surface material of the shielding cover adopts the same material as the main target, so that the good avidity of sputtered product particles is guaranteed. The capture of sputtering products is facilitated. The device can be used for the system-level ignition test on the spacecraft powered on by the electric thruster, and serves as a plume sputtering protection means.

Systems and methods are disclosed employing electric propulsion stationkeeping in a cyclical manner to better match the cyclical pattern of power generated by the solar array system. For a typical orbit design, e.g. a geostationary orbit, North-South stationkeeping can be intermittently suspended, tolerating some additional drift but yielding in a very significant reduction in the required solar power system. If necessary, stationkeeping can be supplemented with a chemical thrusters during off periods for the electric propulsion. Because of this, the overall electrical power margin for the solar array system design can be reduced without compromising the mission performance.

The present invention provides a spacecraft system-level single event upset effect analysis method based on fault propagation. The method comprises spatial radiation environmental effect estimation, system single event fault propagation modeling and device-level, single-machine-level and system-level oriented single event upset fault rate calculation. The method solves the systematic and quantitative problems of estimating the effect of a single event effect on a sensitive system during existing spacecraft design and can provide a necessary reference for selection of spacecraft components and optimization of an anti-radiation design scheme of a single machine and a system.

The invention provides a dynamic sliding mode attitude tracking control method and system for a flexible spacecraft. The method comprises the steps of S1, establishing error posture quaternion-based kinematic equation and kinetic equation of the flexible spacecraft; and S2, by introducing a dynamic switching function, designing a dynamic sliding mode attitude tracking control law of the attitude tracking problem of the flexible spacecraft, and designing a finite time convergence-based robust differentiator to estimate a partial state of a flexible spacecraft system. The dynamic sliding mode attitude tracking control method and system has the beneficial effects that the problem of buffeting caused by a traditional sliding mode control law symbol function is effectively inhibited through thedesign of the switching function; and the dynamic sliding mode attitude tracking control method can enable the spacecraft system to carry out attitude tracking.

The invention relates to a spaceborne remote control and telemetering terminal, system and method based on an FPGA. The terminal comprises an interface management module, a data processing module, a receive data cache module and a transmit data cache module, wherein the interface management module is suitable for respectively receiving a remote control instruction and the first telemetering data delivered by an external first responder and an ADC collector; the data processing module is suitable for analyzing the remote control instruction and writing the remote control instruction to the receive data cache module according to the type of the remote control instruction; and the data processing module is also suitable for reading the second telemetering data of the transmit data cache module, and is connected to an external second responder through the interface management module. According to the spaceborne remote control and telemetering terminal, system and method disclosed by the invention, the stability and compatibility of a spacecraft system when processing the remote control instruction and transmitting the telemetering data can be improved.

The invention discloses an infrared measurement based system and a method thereof for determining the relative state of an inner satellite, belonging to the technical field of spacecraft system technologies and measuring instruments. The state determining system comprises an inner satellite, an outer satellite cavity, an infrared positioning system and an inner satellite locking and releasing mechanism, wherein the inner satellite and the outer satellite cavity are subjected to surface processing, have different emission properties and can present images with different gray scales in the infrared positioning system under the condition of no illumination so as to achieve the function of acquiring the three-dimensional motion state of the inner satellite and finish a key measuring task of an inner formation under the condition of not generating disturbing force.

The invention discloses a separated module spacecraft system based on state perception. The separated module spacecraft system comprises a plurality of module spacecrafts, wherein all the module spacecrafts perform network state conversion according to interaction of the module spacecrafts and other module spacecrafts to form a cluster network, and the network state includes the master node state, the slave node state and the module spacecraft master node state. Therefore, the cluster network formed by different spacecrafts and having strong self organizing ability and adaptive ability is achieved in the space environment. The invention further discloses a self-networking method of the separated module spacecraft system based on state perception, and networking of the cluster network among the spacecrafts is achieved.

The invention discloses a satellite system multidisciplinary optimization method based on multi-model fusion, and belongs to the field of spacecraft system design. The implementation method comprises the following steps: establishing a satellite system key subject analysis model, and optimizing a design variable by taking the satellite system quality as an objective function; in order to improve the solving efficiency of the system scheme optimization problem, a high-precision model and a low-precision model of the geosynchronous orbit transfer subject and the satellite platform structure subject are established, and a Co-kriging agent model is used for fusing model data with different precisions;, sequence filling sampling is carried out through multi-objective optimization with the expected improvement degree and the feasibility probability being comprehensively considered, then the design space is fully explored, and the Co-kriging agent model is updated and managed, so that a scheme which meets the task requirements of the satellite system and has the minimum total mass of the satellite is efficiently obtained, the calculation cost of satellite system optimization is reduced, and the optimization efficiency is improved. The method is helpful for solving the technical problems in other related engineering fields.

A spacecraft system is provided, including a measurement device configured to measure information associated with position of a spacecraft, a filter configured to process the measured information and to provide estimated pre-maneuver information associated with position of the spacecraft and estimated post-maneuver information associated with position of the spacecraft, a propagator configured to predict post-maneuver information associated with position of the spacecraft based upon the estimated pre-maneuver information associated with position of the spacecraft and a model of a maneuver of the spacecraft, an error calculator configured to calculate an acceleration error based upon the estimated post-maneuver information associated with position of the spacecraft and the predicted post-maneuver information associated with position of the spacecraft, and a maneuver archive configured to store the acceleration error for updating the model of the maneuver. The filter, the propagator, the error calculator and the maneuver archive may be implemented in software and / or onboard the spacecraft.

The invention discloses an MDO method of spacecraft local control. The method comprises the steps of receiving user demand information and recording a task source corresponding to the user demand information; generating an overall design index and engineering constraint; subjecting subsystems of a spacecraft to compatible decoupling to generate optimization objectives of the subsystems; and obtaining design parameters of the subsystems to form an overall scheme. The invention also provides a corresponding MDO system of the spacecraft local control. By fully utilizing an interactive cooperation mechanism in a spacecraft system to optimize the design of a large-scale complex system and based on the idea that local minimum serves overall optimization, the MDO method and system of the spacecraft local control can simplify the system optimization problem; compared with a traditional spacecraft design method, has the advantages of being well-targeted, clearly-structured and good in operability; and is a multidisciplinary optimization method applicable to engineering practice, and has some engineering use values.

The invention discloses an input saturation spacecraft unwinding-free attitude tracking control method. The input saturation spacecraft unwinding-free attitude tracking control method includes the steps: calculating the error attitude according to a given command attitude and a practical attitude, then designing a control law according to a backstepping method, including designing a virtual control law and designing a feedback control law by using nonlinear damping, and finally using an extended state observer to estimate the lumped disturbance item in real time so as to obtain an attitude tracking control quantity u; and controlling the spacecraft by means of the control quantity u. The spacecraft system controlled by the input saturation spacecraft unwinding-free attitude tracking control method can track the command attitude with high precision under the condition that the moment of inertia is unknown, and the external disturbance exists, and the input is saturated. Compared with atraditional input saturation attitude control method, the input saturation spacecraft unwinding-free attitude tracking control method has anti-interference and strong robustness, and provides an effective scheme for engineering implementation of attitude tracking control.