82 results about "Spacecraft design" patented technology

The invention discloses a method and a device for analyzing an abnormal state of a spacecraft in the operating process. The method comprises the following steps of: comparing telemeasuring data in a time period to be detected with historical data in a normal state so as to determine an abnormal condition of the telemeasuring data; under the condition that the telemeasuring data is abnormal, according to a correlation coefficient of sequences corresponding to the two groups of data, determining abnormality degree information of corresponding parameters of the telemeasuring data; according to spacecraft design data, establishing an incidence matrix of the corresponding parameters of the telemeasuring data and stand-alone equipment; according to associated data corresponding to the incidence matrix and the abnormality degree information, determining a fault probability of each piece of stand-alone equipment so as to analyze the possibility that the abnormal condition occurs to each piece of stand-alone equipment. The method and the device utilize the telemeasuring data so as to analyze the abnormal condition of the telemeasuring data, and solve the problem that due to shortage of a method for analyzing the abnormal state of the spacecraft in the operating process, a fault part cannot be determined when the spacecraft is abnormal in the operating process.

The invention discloses a collimation measurement system for attitudes and angles among spacecraft devices based on the combination between a robot and a theodolite. The system herein includes a robot, a laser tracker, a laser tracking target (T-MAC), a robot terminal tool, and the like. The system searches a to-be-tested datum cube mirror which is disposed on a spacecraft device by conducting mode identification, and computes the relations of phase position and direction of the datum cube mirror with respect to the theodolite. The laser tracker is intended for calibrating the relative direction relation in a coordinate system of respective spacecraft devices and integrating the measurement results of the theodolite at different measurement positions to the same coordinate system. Based onthe calibration relation and the relative relation, the laser tracker is guided to real-timely track the robot terminal tool and establish the relative relation between the laser track and the robotterminal tool. And eventually, the attitude relation matrix of the spacecraft is computed. According to the invention, the automatic measurement of the attitude relation among different devices is realized, the measurement efficiency can reach one time per half-minute, the measurement precision can be higher than 30'', on-site measurement flexibility is higher, and construction and measurement indifferent places can be much easier.

The invention discloses a spacecraft test system based on assemblies and scripts. The spacecraft test system based on the assemblies and the scripts comprises a test case editing environment module and a test case executing environment module, wherein the test case editing environment module comprises a test assembly library, a script function library and a test case imaging interface; the test assembly library comprises test assemblies; the script function library comprises script functions; the test case imaging interface arranges the test assemblies, the script functions and / or script placeholders by dragging according to a test case, sets parameters of the arranged test assemblies and edits scripts for the script placeholders so as to form the test case; and the test case executing environment module analyses the test case so as to form an executable sequence of the test case and executes the test case sequence in the testing process to finish the test. The invention also discloses a corresponding spacecraft test method. The test scripts and the test assemblies are combined to realize automatic test on a spacecraft.

The Dual Spacecraft Deployment System (DSDS) is a concept developed for deploying Dual Satellites into independent orbits from a single launch vehicle. The Deployment Concept, Deployment System and Spacecraft Design with Representative Instruments are presented herein. The design employs an innovative half-hex design for the spacecraft buses which allows the spacecraft to be accommodated in a parallel mounted configuration on the launch vehicle. The Spacecraft are arranged on the launch vehicle with a central deployment structure with a deployment mechanism which allows independent lateral deployment of two spacecraft. The Spacecraft Design can have a large volume, e.g., sized to accommodate Spacecraft Sub-systems and Instrument Systems. This concept allows the launch vehicle to accommodate two large Spacecraft deployed to different orbits with a single launch doubling the efficiency of the launch vehicle.

This invention is a launch-to-space, refuel and return system which comprises: a unique circular and straight maglev track and maglev sled spacecraft launcher; a single stage-to-LEO spacecraft (Mother Spacecraft); an orbiting space platform (OSP); a currently utilized Space Shuttle is altered to improve its safety, range, longevity and versatility which serves as a prototype for the Mother Spacecraft design. The Mother Spacecraft is releasably attached to the maglev sled and propulsion for the maglev sled is supplied by magnetic propulsion and on-board rocket propulsion. The collapsible OSP can fit into the Space Shuttle storage bay and is remotely deployed. The OSP is a refueling depot for the Mother Spacecraft, the Space Shuttle, as well as a rendezvous point for various space missions. A single, liquid, non-cryogenic rocket fuel is standardized throughout the entire system which eliminates: all vertical launches; all expendables; all solid rocket propellants.

The invention discloses a spacecraft general assembly three-dimensional design model and a BOM association import system and method. The import system comprises a design data import and conversion module which is used for importing and processing design data and performing component construction so as to perform light weight model conversion on the three-dimensional design model; a design import data management module which is used for associating the version of the light weight model and the component and establishing product structure; a design import data controlled and changing module which is used for carrying out three-level approval on the imported design data and publishing the design data after the data passing through the approval. The spacecraft general assembly three-dimensional design model and the BOM association import system and method form a unified data source facing general assembly technology design, containing product structure, three-dimensional design model and structural attribute to realize association of three-dimensional design model and the BOM structure, increase the efficiency of spacecraft design information transmission and guarantee the product quality and reduce investment in human resource.

The invention belongs to the technical field of design, and dynamics simulation and control of high-precision spacecraft, and in particular relates to a simulated analysis platform for thermal disturbance responses of the spacecraft. The platform comprises a data input modeling module, a flexible attachment on-orbit thermal analysis module, an equivalent thermal load derivation module, a modal analysis module, a coupling dynamics modeling module, a coupling dynamics model solution module and a post-processing module in sequence; the simulated analysis platform builds a coupling dynamics model of large flexible satellite-borne attachments and stars, and provides a special simulated analysis platform for coupling responses of the large flexible satellite-borne attachments and the stars. The simulated analysis platform can be used for solving posture variations of the spacecraft with complex flexible attachments caused by thermally induced dynamic deformations of the flexible attachments, can provide a simple and effective model, method and tool for forecast and assessment of thermally induced vibration responses of the spacecraft with the flexible attachments under actual conditions, and can serve as a reference for a satellite design process.

An inventive autonomous active manoeuvring method and system 1 for spinning spacecraft is provided having a capability to enhance the AOCMS performance of passive spinning satellites and to fulfil the emerging autonomy requirements applicable to new generation satellites. In broad terms, the invention resides in (a) the overall concept of providing autonomous execution of spin axis re-orientation manoeuvring for spinning spacecraft designed and executed autonomously on-board the spacecraft by the AOCMS and (b) in the proposed strategy set in place to execute the re-orientation manoeuvres with respect to the handling of residual nutation. Advantageously, the provision of coupling nutation avoidance manoeuvres with active nutation damping 8 on board the spacecraft reduces / minimises the manoeuvre settling time required to return the spacecraft to the steady state pointing performance, while not imposing constraints upon the particular spacecraft inertia sensor properties.

The invention discloses a halo orbit detector structure of a sun-earth system and an attitude directing thereof, belonging to the design field of spacecrafts. A digital antenna is fixedly arranged on the +xb surface of a detector body; a shaft of the digital antenna is along the xb shaft direction; solar arrays are fixedly arranged on the + / -yb surface of the detector, the axial line is vertical to the + / -yb surface of the detector body, and the included angle between the normal and the -xb shaft is theta which is the installation angle of the solar array and is calculated by average halo orbit information. The xb shaft of the detector body is directional to the ground during the running time of the detector. The normal of the solar array directs to the sun, after the theta is selected, the solar incident angle of the solar array changes within a certain range and is 90 degrees on average in the orbit cycle. The halo orbit detector structure has the advantages that the long-term direction of the digital antenna to the ground is realized by means of one fixed digital antenna and one pair of fixed solar arrays, the solar arrays are directional to the sun for a long term through the digital antenna with low digital code rate, moving parts of the detector are few, the controllability and reliability of the detector are greatly improved, and the control difficulty is reduced.

The invention relates to a system and a method for spacecraft semi-physical simulation experiments based on a rotor craft. The system comprises a ground station system and a simulated aircraft. The ground station system comprises a computer simulation system, a wireless communication device, a rotorcraft hand-control manipulator. The simulated aircraft comprises a rotor craft, a three-freedom-degree holder, a force transducer, and a load. The three-freedom-degree holder is fixedly connected under the rotor craft. The load is hung on an inner shaft of the holder. The force transducer is arranged between the holder and the load. Three-freedom-degree mass center motion of the rotor craft simulates three-freedom-degree mass center motion of a spacecraft. Three-freedom-degree rotation of the holder simulates three-freedom-degree posture motion of a spacecraft. The load simulates six-freedom-degree motion of the spacecraft. The force transducer measures stress between the holder and the load, and is used to compensate influence of load gravitational moment. The system and the method greatly reduce realization difficulty, and are advantaged by low development cost and simple and feasible.

A composite panel provides structural strength and rigidity for modular assembly of spacecraft while serving the dual purposes of structure and heat transfer for thermal management of an environment for equipment, such as a spacecraft. Instruments, gimbals, surveillance, imaging, detectors, and the like may be mounted in a spacecraft designed and constructed from standard panels to provide the structural and heat transfer requirements to support the onboard equipment. Extremely small temperature differentials required by the panels support a substantially isothermal perimeter for the structure, able to sink heat from any location on a panel, transport it to a rejection site, and reject it to the cold environment of space, thus easing the task of design and packaging of instrumentation and infrastructure of satellites and other spacecraft. Two-phase mass transport of fluids inside the panels aids high heat flux rates with minimal temperature differentials, even with reinforced, composite, polymeric materials for the structural elements of the panels.

The invention relates to a spreadsheet-based fast customization and integration method for a satellite design system and belongs to the field of design of spacecrafts. The method includes the first step of sorting and summarizing the general flow of designing remote sensing satellites or other types of spacecrafts on the basis of a traditional process expression mode, the second step of conducting mapping to generate a DSM matrix, conducting statistics on DSM items and conducting structured quantitative analysis item by item to obtain a perfect parameter correlation list, and the third step of utilizing the spreadsheet-based middleware technology to achieve storage of a spreadsheet in a database, establishing four modules through existing functions of middleware of the spreadsheet and then achieving fast customization of the spacecraft design system, wherein the four modules are an engineering application module, a DSM modeling module, a project management module and an import and export module. According to the method, the existing spacecraft design process is mapped to generate a DSM, the DSM is stored into the database for maintenance, the Excel spreadsheet is used as a front-end interface, and consequently a foundation is laid for follow-up optimization.

Transport and in-space assembly of spacecraft from not-qualified-for-launch, low cost parts / modules. Any number of spacecraft or parts thereof are built inexpensively for space usage when they are transported in a manner that does not subject them to stresses of the launch environment. Such spacecraft / parts are transported whole, or in component pieces, in protective containment, i.e., qualified containers which meet all of the applicable safety and performance requirements for containers launched into space. These whole or component pieces of spacecraft may then be tested and deployed and / or assembled and deployed by either robot or human activity. The spacecraft, when at times modular, are assembled so as to form larger structures in space. This piecewise method of construction provides a significant increase in cost-effectiveness concerning both radical spacecraft design and spacecraft-to-launch vehicle qualification processes.

The invention relates to a geometry and behavior integrated three-dimensional modeling method for a spacecraft. After the invention fully researches the geometric shape and behavior characteristics of the spacecraft, the invention provides the method by adopting layering modeling, model abstraction technique and computer graphics three-dimensional modeling to be combined with a molding theory. The method is simple and efficient, the model can be modified easily, the components can be transplanted and reused, and the method can be used for behavior modeling for the spacecraft, provides auxiliary means for designing and manufacturing the spacecraft, can realize reasonable unification between the three-dimensional geometrical model and the behavioral model of the spacecraft, can be applied to full life cycle and full system comprehensive management of complex aerospace products, analyzes complex products on the system layer, and supports development of the complex system from top to bottom.

The invention relates to a satellite design data transmission method based on a DSM and workflow and belongs to the field of design of spacecrafts. The method includes the first step of sorting and summarizing the general flow of designing satellites or other types of spacecrafts on the basis of a traditional process expression mode, the second step of conducting mapping to generate a DSM matrix, conducting statistics on DSM items and conducting structured quantitative analysis item by item to obtain a perfect parameter correlation list, and the third step of utilizing the spreadsheet-based middleware technology to achieve storage of a spreadsheet in a database, and conducting retrieval to generate the satellite coupling design process reflecting engineering practice, conducting dynamic display on the correlation between all satellite design activities through a friendly human-computer interface, and finally conducting automatic data flow to generate a design report. As a result, a foundation is laid for follow-up optimization.

The invention relates to a simulation analysis platform for thermally-induced micro-vibration response of a satellite-borne flexible accessory, and belongs to the technical field of high-precision spacecraft design as well as dynamical simulation and control. The simulation analysis platform for thermally-induced micro-vibration response of the satellite-borne flexible accessory comprises a data inputting and modeling module, an on-orbit thermal analysis module, an equivalent thermal load derivation module, a high-precision modal analysis module, a high-precision vibration response analysis module and a post-processing module, all of which are connected in a sequence; the output of the previous module serves as an input of the next module; and the analysis platform carries out analysis and simulation research on thermal vibration of an engineering-practical large-scale satellite-borne flexible accessory and research on corresponding software for calculation, and realizes the purpose of rapidly and efficiently obtaining the thermally-induced micro-vibration of large-scale flexible components and coupling vibration response between the components and a satellite. Meanwhile, the simulation analysis platform can promote engineering application, testing and verification of a thermal deformation and thermally-induced micro-vibration simulation technology of the large-scale satellite-borne flexible components, can drive progress of related support technical research and can accumulate a technological base for subsequent research of novel satellite types.

A spacecraft architecture and accompanying standard allows for the creation of a spacecraft using an assortment of modules that comply with the standard. The standard preferably includes both mechanical and electrical compatibility criteria. To assure physical / mechanical compatibility, the structure of each module is constrained to be compatible with any other compatible module. To minimize the interference among modules, the extent of each module in select dimensions is also constrained. To assure functional compatibility, a common communication format is used to interface with each module, and each public-function module is configured to respond to requests for function capabilities that it can provide to other functions. Each module is preferably designed to provide structural support to the assemblage of modules, and an anchor module is provided or defined for supporting the entire assemblage and coupling the assemblage to other structures, such as a launch vehicle.

The invention discloses a multi-spacecraft task intelligent planning and control method for a complex dynamic environment, and the method comprises the steps: randomly generating a large amount of sample data based on a three-pulse indirect transfer model through combining with a reserved task orbit dynamic environment of a spacecraft; estimating the transfer trajectory parameter of the potentialtarget by using the neural network with complete learning and combining the actual task scene of the spacecraft, and substituting the estimated value of the neural network into the constructed task comprehensive performance evaluation model to calculate the single-satellite task performance; performing multi-spacecraft task coordination allocation by using a distributed bidding and tendering market mechanism to obtain a scheduling scheme with optimal global task efficiency; and the task spacecraft designing an accurate transfer track in a complex dynamic environment according to the task allocation result, generating a control instruction and controlling the spacecraft to complete a task. According to the method, the defects of poor planning real-time performance and low control precisionwhen an existing method is used for processing a multi-satellite task under the condition of a complex dynamic environment are overcome.

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 relates to a DSM-based satellite coupling design flow structured analysis and retrieval method and belongs to the field of design of spacecrafts. The method includes the first step of sorting and summarizing the general flow of designing remote sensing satellites or other types of spacecrafts on the basis of a traditional process expression mode, the second step of conducting mapping to generate a DSM matrix, conducting statistics on DSM items and conducting structured quantitative analysis item by item to obtain a perfect parameter correlation list, and the third step of utilizing the spreadsheet-based middleware technology to achieve storage of a spreadsheet in a database, and retrieving the coupling design process of the satellites according to the storage information about DSM correlation subitems in the database. According to the DSM-based satellite coupling design flow structured analysis and retrieval method, on the basis of the existing design process of the spacecrafts, the design process is mapped to generate a DSM, the DSM is stored into the database for maintenance, the Excel spreadsheet is used as a front-end interface, and consequently a foundation is laid for follow-up optimization.

The invention discloses a three-dimensional process digital prototyping system oriented to spacecraft final assembly and a three-dimensional process digital prototyping construction method. The three-dimensional process digital prototyping system comprises a design data importing module, a process digital prototype management module and a process digital prototype issuing module. The design data importing module is used for receiving and converting design data in an upstream design system. The process digital prototype management module is used for generating a process digital prototype view, reconstructing product structures of the converted design data and filling in structured product attributes. The process digital prototype issuing module is used for issuing a process digital prototype which is approved. According to the three-dimensional process digital prototyping system oriented to the spacecraft final assembly and the three-dimensional process digital prototyping construction method, process data sources are uniform, transmission efficiency of design information of a spacecraft is improved, and a solid foundation is built for achieving the three-dimensional structured final assembly of the spacecraft.

A spacecraft space firm docking method relates to tiny spacecraft space docking methods. According to the spacecraft space firm docking method, a microreactor coated with reactive adhesive is coated on the surface of a flexible unfolding device; the flexible unfolding device is folded and stored in a spacecraft launching cabin in an enclosed mode during when the spacecraft launching cabin is launched; when the spacecraft launching cabin is in a pre-selected orbit, the enclosed device is opened to enable the flexible unfolding device to extend out and be unfolded slowly; when the flexible unfolding device touches another spacecraft, a permanent firm bond is achieved through the reactive adhesive released by the broken microreactor. The spacecraft space firm docking method utilizes the flexible slow unfolding device to buffer the flight speed and the docking force, and utilizes the reactive adhesive which reacts in the orbit in the space to bond the two spacecrafts firmly as a whole to further achieve permanent docking of the two spacecrafts in the space, so that the flexibility, the reliability and the success rate of spacecraft docking can be improved and the difficulty of the docking mode is reduced.

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 a truss-type all-flexible spacecraft structure platform, belonging to the technical field of spacecraft structure design. The spacecraft structure platform adopts an expandabletruss structure formed by splicing a plurality of basic truss structure units as a main bearing structure, and a deployable solar array is laid on top of the truss structure to form a plane that canobtain sunlight. Various equipment loads carried on the spacecraft are installed by hanging under the truss or being arranged in cavities inside the basic truss units. Compared with the traditional spacecraft structure platform, the rigidity of the structure system is improved, the mass of the structure system is reduced, and the payload ratio of the system is greatly improved. In addition, the truss-type all-flexible spacecraft structure platform also has good expansibility, and can be assembled into various different configurations flexibly and conveniently to meet the needs of spacecraft design tasks of different scales.

A composite combination for launching a payload into space including a hypersonic spacecraft carrying the payload and a carrier aircraft, the spacecraft design to take off from the ground on the aircraft and then, upon separation from the carrier aircraft, to propel the payload on its own toward an assigned trajectory. The combination further includes a separation mechanism for separating the vehicle from the aircraft, and measuring elements for measuring physical data operative in the evaluation of the safety of a phase of separation of the vehicle and the aircraft, which data is evaluated by a computer component to authorize a prohibit activation of the separation mechanism.

The invention provides a reconfigurable single board pico satellite system. The system relates to the field of spacecraft design and development. The system is based on a printed circuit board satellite structure, a function module required for space missions of a pico satellite is integrated on a single board, logic and control are all implemented by using FPGA, meanwhile an FPGA-based single board satellite is developed by using IP cores of processors, and the fault tolerance of the processor can also be improved through a software secondary development mode; according to the system, a bus controller is connected with various modules through scheduling, the control of a computer system to each subsystem is achieved, the system has the advantages of being simple in structure, high n integration, small in size and low in power consumption and cost, the FPGA is used for achieving the condition that the IP cores of the processors of different architectures are provided by different manufacturers to achieve the reconfigurable properties of the processors, and an anti-radiation technology is used for secondary development of the IP cores of the processors to improve the fault toleranceof the processors; according to a pico satellite sensor and a measurement function, the FPGA is used for completing hardware implementation of different types of algorithms.

The invention discloses a film spacecraft, and belongs to the technical field of spacecraft design. The thin-film spacecraft comprises a multifunctional body platform, a thin-film propelling device, athin-film camera, a thin-film antenna, a thin-film battery piece, a thin-film sensor, a thin-film actuator and the like. The thin film propelling device stretches the six thin film sails into a hugeregular hexagon sail surface in a supporting rod stretching mode, flight power is obtained under the action of sunlight pressure, and the thin film camera can conduct high-resolution imaging on an observed target. According to the film spacecraft, multiple functions of thin film propulsion, thin film imaging, thin film communication, thin film power generation, thin film sensing and actuation andthe like are integrated; the film spacecraft can complete long-term voyage and large-range space flight tasks, the advantages of a thin film structure technology and a multifunctional structure technology are fused, through depth integrated design of the structure and loads, weight reduction, energy increase and function ratio improvement of the spacecraft structure are further achieved, and the requirements of spacecrafts for high performance and long service life can be better met.

The invention relates to a sealed cabin wall structure of a manned spacecraft. The sealed cabin wall structure comprises a first partition plate (1), a first dot matrix layer (2), a second dot matrixlayer (3) and a third partition plate (4) which are sequentially arranged; the first dot matrix layer (2) and the second dot matrix layer (3) are both composed of tightly arranged cell structures (6);the sealed cabin wall structure further comprises a second separation plate (5), and the second partition plate (5) is positioned between the first dot matrix layer (2) and the second dot matrix layer (3); the size of a single cell structure (6) forming the first dot matrix layer (2) is smaller than that of a single cell structure (6) forming the second dot matrix layer (3); and the thickness ofthe first dot matrix layer (2) is larger than that of the second dot matrix layer (3). The cabin wall structure integrates multiple functions, and the requirements for weight reduction and efficiencyimprovement of spacecraft design are met.

The invention discloses a large-irradiation-area atomic oxygen beam simulation system. The system comprises a water-cooled electromagnetic field system, a microwave generation and transmission system, a molybdenum neutralization plate, a water cooling system, a vacuum system and a target base mechanism, wherein the water-cooled electromagnetic field system can generate a strong magnetic mirror field which can reach 1,200 gausses; the frequency of the microwave generation and transmission system is 2.45GHz and the power of the microwave generation and transmission system can reach 1,000W; the microwave generation and transmission system can excite working oxygen to ionize in the strong magnetic mirror field to generate and restrain high-density oxygen plasma; major axis length and minor axis length of the molybdenum neutralization plate are 200mm and 140mm respectively; the molybdenum neutralization plate is used for accelerating, collecting and reflecting positive ions in the plasma to form a directional atomic oxygen beam of which the sectional area diameter is more than 150mm; the water cooling system can guarantee that an atomic oxygen source can uninterruptedly and stably workfor 100 hours; the vacuum system is used for guaranteeing a vacuum condition for normal work of atomic oxygen; and the target table mechanism is used for placing a test sample in an atomic oxygen beam region. The system guarantees that a device can continuously and stably run in a high atomic oxygen beam state through measures in aspects of cooling, sealing, material selection and the like. By using the system, the atomic oxygen environmental adaptation testing efficiency of an aerospace material can be greatly improved, the test period is shortened, the test cost is reduced, the technical bottleneck of component-level atomic oxygen test can be broken through, and a simulation means is provided for more effectively verifying reliability of spacecraft design.