82 results about "Spacecraft design" patented technology

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.

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.

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.

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 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 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.