53results about "Cosmonautic components" patented technology

Exemplary embodiments described herein may include lightweight, low stow volume solar concentrator.

The invention relates to a system for full-supercooling large-flow liquid methane filling and a using method of the system. The full-supercooling large-flow liquid methane filling system is stable, reliable, high in automation degree, capable of dynamically adjusting the supercooling temperature and flow of liquid methane and high in accuracy, and comprises a storage box, a pressurization module, a recovery module, a methane filling module, a control module and a nitrogen filling module, and the pressurization module comprises a methane pressurization tank; the methane filling module comprises a methane filling tank, the nitrogen filling module comprises a nitrogen filling tank and a subcooler, the recovery module comprises a methane recovery tank, and the control module pressurizes the methane recovery tank and the methane filling tank by controlling the pressurization module and adjusts the filling flow and temperature of methane to prevent condensation by controlling the methane filling module and the nitrogen filling module. Flow stabilization and backflow are carried out by controlling the methane filling module and the recovery module.

The invention discloses a liquid oxygen circulation supercooling system, and relates to the technical field of liquid oxygen storage management, liquid oxygen is stored in a liquid oxygen storage tank, the liquid oxygen enters a self-pressurization vaporizer from a vaporization liquid inlet pipe to be vaporized, oxygen generated by vaporization enters the liquid oxygen storage tank from a vaporization steam exhaust pipe, so that the pressure in the liquid oxygen storage tank is increased, and the liquid oxygen is discharged into a pipe needing to be cooled after being increased; the supercooling process is realized when the liquid oxygen in the pipe needing to be cooled flows through the cooler, and the liquid oxygen flows back to the liquid oxygen storage tank again after being supercooled; the device is applied before a launching task starts, liquid oxygen in the liquid oxygen storage tank is supercooled through continuous liquid oxygen circulating flow, compared with supercooling when liquid oxygen is injected into a rocket, supercooling is conducted in advance in the storage stage, injection peak flow does not need to be matched, the size of the device is smaller, and the manufacturing cost is lower.

A header rail 10 for transporting gas cylinders (30, fig 3) on lorries has an elongate non-metal (e.g. composite such as glass / carbon fibre) central core 12, with metal collars 14 spaced apart along t

The invention relates to a storage system (1) for a pressurized gaseous fuel, comprising at least two pressurized gas containers (2), which are coupled to one another and to filling and withdrawal connections via a line system (6), and at least one temperature-control device (15), wherein the temperature-control device (15) is arranged in the line system (6). The storage system according to the invention is characterized in that the temperature-control device (15) is arranged in a connecting line (13) between at least two of the pressurized gas containers (2).

A number of techniques, structures, and materials for thermally insulating a cryogenic propellant tank barrel section are presented. Thermal insulation for a tank, such as tanks for use in space flight missions, may be a dual wall vacuum jacket system where one wall is the propellant tank wall and the other wall is a metallic foil that is welded to portions of the propellant tank wall. The metallic foil may act as a vapor barrier and a vacuum seal. A structural and breathable insulation that is durable and functional in a range from cold to relatively high temperatures may be used to maintain a standoff distance between the two walls. The volume between the propellant tank wall and the metallic foil may be purged with a gas that has a liquefaction temperature greater than that of the cryogenic propellant that will be in the propellant tank during operations.

A double-wall tank comprising an inner vessel surrounded by an outer jacket. The double-wall tanks comprising arrangements for coupling an inner vessel relative to an outer jacket which include roller bearings between the inner vessel and the outer jacket. Also an aircraft having at least one double-wall tank with the roller bearings.

The application provides a gravity unloading attitude adjusting device and method suitable for heavy load assembly, four gravity unloading units are arranged on a carrying vehicle, a single gravity unloading unit comprises a plane air floating support mechanism, a support column, a vertical hydraulic cylinder and a spherical air floating support mechanism; a load platform is supported by the four gravity unloading units, the load platform is driven by an attitude adjusting driving unit to carry out six-degree-of-freedom position adjustment, so that the assembly body reaches a target assembly position. In the position adjustment process, the plane air floating support mechanism provides translational freedom and rotation freedom around the vertical direction, the four vertical hydraulic cylinders and the spherical air floating support mechanisms thereon jointly provide tilting freedom in any direction and vertical movement freedom, and support force is provided to offset the gravity of the load platform and its load, gravity unloading is realized, the driving load of the attitude adjusting driving unit is greatly reduced, the accuracy and feedback speed of the attitude adjusting process are provided, and oscillation is effectively avoided or reduced.

The application discloses a micro electric propulsion storage and supply system, characterized in that the system comprises a storage module, a pressure regulating module, a flow regulating module, a power module and a control module, wherein the storage module, the pressure regulating module and the flow regulating module are sequentially connected through stainless steel pipelines, and the modules and the pipelines are connected through standard sleeves; the power module supplies power to the pressure regulating module and the flow regulating module, and the control module realizes pressure and flow control. The system adopts a combination of a miniature mechanical pressure reducing valve and a pressure sensor as the pressure regulating module, realizes smaller volume and mass relative to a Bang-Bang type, and has lower cost relative to a proportional type, so that balance is achieved between the reduction of the volume and mass of the system and economy; the flow regulating module adopts a pasting mode of a capillary tube and a silicon rubber heating belt, so that the silicon rubber heating belt can be directly pasted to the metal capillary tube, and an outer shell of the capillary tube is not needed, thereby reducing the volume and mass of the module and improving the heating efficiency.

The invention discloses a liquid oxygen methane carrier rocket low-temperature filling system and method, and belongs to the technical field of aerospace power. Comprising a filling target container, a measurement control system, a methane filling system, a liquid oxygen filling system and an oxygen supercooling system. The methane filling system is communicated with the methane storage tank and is used for filling methane into the methane storage tank; the liquid oxygen filling system is communicated with the liquid oxygen storage tank and is used for providing liquid oxygen for the oxygen storage tank; the oxygen supercooling system is connected with the liquid oxygen filling system so as to provide supercooled liquid oxygen; the measurement control system is connected with the methane filling system, the liquid oxygen filling system and the oxygen supercooling system through a control circuit so as to monitor and control the operations thereof; the system adopts a tank car extrusion propellant filling mode. The device has the characteristics of low cost, short construction period, adjustable filling flow and controllable temperature. The rear end of methane filling pipeline equipment is remotely controlled, and safety is guaranteed. The system has engineering application value for multiple flight tests of a reusable rocket.

An object of the system according to the embodiment is to analyze a signal or a radio wave from outer space and estimate information regarding an unknown region or a living body.SOLUTION: A system includes a signal collection unit, a generation AI analysis unit, and an estimation unit. The signal collection unit collects signals and radio waves from outer space. The generation AI analysis unit analyzes the signals and radio waves collected by the signal collection unit in the generation AI. The estimation unit estimates information on an unknown region of the universe or a living body on the basis of a result analyzed by the generation AI analysis unit.SELECTED DRAWING: Figure 1

The objective is to reduce the effects of Coriolis force in an artificial gravity generator. [Solution] An artificial gravity generating device comprising: a driving source 1; a rotating shaft member 2 that is rotated by the driving source 1; an arm 3 attached to the rotating shaft member 2; and a first box-body 6 that can accommodate a person and is connected to the arm 3 via a joint member 4 that can rotate in any direction and can be tilted in any direction by the joint member 4, thereby reducing the effects of the Coriolis force acting during rotation; and wherein the rotation of the rotating shaft member 2 causes the first box-body 6 to rotate around the rotating shaft member 2, thereby generating centrifugal force.

Provided is a system, method and device for payload positioning. The system includes a rover configured to travel to a site and hold the payload, a sensor module attached to the rover for obtaining sensor data of the site, a robotic arm for equipping and stowing away the sensor module, attaching to a connection point of the payload, and controlling a position of the payload when attached to the connection point.

This invention relates to a mass-optimized geogrid system and method for stabilizing regolith and other granular soils in extraterrestrial and extreme Earth environments. Mass reduction is achieved both locally—within individual geogrid layers through optimization of lattice geometry, variable mesh density, and reinforcement at stress concentrations—and globally, by varying the spacing and distribution of layers according to structural demands. The geogrid may be fabricated from materials that are locally sourced or compositionally traceable to in-situ resource utilization (ISRU) processes in environments such as the Moon, Mars, or terrestrial polar and desert regions where conventional construction materials are limited or costly to transport. The method includes restructuring the regolith by size-sorting and compaction to produce a particle-size distribution proportionally matched to the mesh openings, enhancing mechanical interlock and shear resistance. Structural efficiency can be tuned and verified through physical testing and finite-element analysis for scalable, repeatable construction.

The invention provides a satellite on-orbit assembly structure relative pose resolving method and system, and the method comprises the steps: constructing a satellite on-orbit assembly structure relative pose measurement system through two sets of five-degree-of-freedom measurement equipment, and building a mathematical model of a three-dimensional displacement measurement device and a two-dimensional deflection angle measurement device; a measurement model of five-degree-of-freedom measurement equipment is deduced and established based on a geometrical relationship, a pose conversion parameter linear resolving model is obtained through processing according to a pose conversion small spinor theory, pose conversion parameters are iteratively calculated through a least square method, and the method is suitable for pose measurement and resolving of most satellite in-orbit assembly or adjustment structures. Good application prospects are realized.

The application provides a lunar surface protection device and relates to the technical field of space protection.The lunar surface protection device provided by the application adopts a modular design combining a composite material plate and a prefabricated framework, the prefabricated framework is manufactured on the earth, a shape memory hinge is used as a folding node, the prefabricated framework is folded to a compressed form, the volume size is greatly reduced, the lunar surface protection device is convenient for transportation between the earth and the moon, and the carrying pressure of a carrier rocket is reduced.The lunar surface protection device fully utilizes lunar in-situ resources, completes the manufacturing of the composite material plate, realizes the combination of the manufacturing on the earth and the moon, can reduce the operation steps of manufacturing the protection structure in the extreme environment on the lunar surface, and improves the manufacturing efficiency of the protection structure.

Valve (1) for a fluid reservoir, comprising an inner portion configured to be arranged inside the reservoir (3), the inner portion comprising at least one portion of a sensor (34, 35) for measuring at least one parameter characterising the fluid, the valve (1) being capable of receiving information from a computer (5), external to the valve (1), characterised in that the valve (1) also comprises an outer portion configured to be arranged outside the reservoir (3), the outer portion comprising an internal computer (7) comprising: - a first acquisition means (9) configured to acquire data from at least one sensor (34, 35), - a communication means (11) using a wired digital communication interface capable of communicating bidirectionally with the external computer (5), - a control means (13) configured to take into account information received from the external computer (5) and data from at least one sensor (34, 35) for controlling at least one actuator (100) of the valve (1), the at least one actuator (100) being connected to the control means (13) by at least one electrical connection (15), and - a means (17) for measuring at least one current parameter in the electrical connection (15).

The vessel includes first and second hollow shells including first and second inner surfaces, respectively, for receiving a portion of a pressure vessel (PV). The first hollow shell includes a fibrous layer at least partially impregnated with a resin and an energy dissipative material substantially concentric with the first inner surface and disposed between the first inner surface and the fibrous layer. The second hollow shell includes a fibrous layer at least partially impregnated with a resin and an energy dissipative material substantially concentric with the second inner surface and disposed between the second inner surface and the second fibrous layer. The first and second hollow shells are attachable to one another to define a volume for at least partially enclosing the PV.

Systems and methods for manufacturing modular bricks from regolith on the surface of the moon for radiation protection and lunar habitation are provided. For example, regolith can be used to manufacture modular bricks that form maintainable and reusable structures that can provide protection from radiation on the lunar surface. In some embodiments, hollow bricks can be filled with loose regolith to leverage its radiation protection capabilities by increasing an amount of material within the structure while reducing the amount of processed material, thereby increasing the radiation protection afforded by the structure. The bricks can be shaped like hexagonal prisms to promote stacking with one another in the absence of binding materials while also allowing for tolerance during positioning. Manufacturing of the bricks can involve melting lunar regolith in a heating device, casting it into hollow hexagonal prism bricks, filling these bricks with regolith, and assembling them into structures.

The invention relates to a tank container system (100), in particular for pressure tank containers and very particularly for high-pressure tank containers for storing a gaseous fuel for driving a vehicle. The invention also relates to a system for accommodating such a tank container system. According to a first aspect, the stated problem is solved by a tank container system (100) for mounting on a vehicle, which tank container system comprises a first and a second carrier (110, 120), wherein the first and the second carrier (110, 120) comprise brackets (131, 132) for a tank container (150), wherein the brackets (131, 132) are each provided in pairs, wherein a first bracket (131) of each bracket pair is located on the first carrier (110) and a second bracket (132) of each bracket pair is located on the second carrier (120), wherein each bracket pair comprises a floating bearing (135) and a fixed bearing (136). The brackets (131, 132) are designed to be interchangeable, wherein the first carrier (110) and the second carrier (120) each comprise a device (111, 121) in which the first bracket (131) and the second bracket (132) are mounted so as to be displaceable and fixable in the longitudinal direction of the associated carrier (110, 120).

The application discloses a multi-spacecraft rendezvous orbit interval determination method, and belongs to the technical field of aerospace. The method is as follows: according to the task scale and the carrying capacity of a spacecraft mother ship, the maximum number of sub-spacecraft that can be carried by the spacecraft mother ship is set. The deployment orbit height of the spacecraft mother ship is set, a Lambert problem of the sub-spacecraft released by the spacecraft mother ship and a target spacecraft is established, the sub-spacecraft rendezvous range boundary is solved according to the speed increment capacity of the sub-spacecraft, and a sub-spacecraft rendezvous range database is constructed with the orbit plane angle as input and the sub-spacecraft rendezvous range boundary as output. The minimum rendezvous transfer phase range of all the sub-spacecraft carried by a single spacecraft mother ship is calculated by interpolation using the sub-spacecraft rendezvous range database, and the initial deployment orbit interval covered by the spacecraft mother ship is determined. The process of determining the spacecraft mother ship deployment range is repeated until the corresponding spacecraft mother ship orbit interval of all the target spacecraft is determined.

An isolable double-walled insulation tank for a launch vehicle includes a hollow open cylinder part having upper and lower end portions that are open in an upright state; an upper cap part for sealing the upper end portion of the open cylinder part; and a lower cap part for sealing the lower end portion of the open cylinder part. The open cylinder part includes a flat plate part having a flat outer peripheral surface; an uneven part slidably fitted and coupled to the flat plate part; and a plurality of load-dispersing spaces which are ring-shaped according to a portion of the uneven part being spaced from the flat plate part, and which are arranged to be spaced apart in the height direction of the open cylinder part.

The invention relates to a hydrogen supply device and an aircraft. The hydrogen supply device comprises a hydrogen tank (52) and at least one hydrogen circuit (54), the hydrogen circuit (54) comprising:-at least one first section (56) positioned in a gas-tight first container (76), the first section (56) comprising at least one item of equipment (66) of a first category; -at least one second section (58) positioned in a gas-tight, removable second container (78), the second section (58) comprising at least one item of equipment (68) of a second category; -at least one connection section (60) connecting the first section (56) and the second section (58), the connection section (60) being positioned below the first container (76) and below the second container (78), or positioned below the first container (76) and between the second container (78) and the hydrogen tank (52). The solution helps to improve the compactness of the device and facilitate the disassembly and handling of the detachable second container (78).

The application relates to the technical field of emergency disposal of aerospace low-temperature chemical fuel leakage, in particular to a mobile emergency disposal system for aerospace low-temperature chemical fuel leakage and a disposal method thereof. The mobile emergency disposal system for aerospace low-temperature chemical fuel leakage comprises a powder conveying assembly, a water conveying assembly and a first water-powder film-forming spraying element. The powder conveying assembly comprises a pneumatic driving element and a first sending tank. The first sending tank is used for storing water-absorbing resin composite material. The pneumatic driving element and the first sending tank and the first sending tank and the first water-powder film-forming spraying element are communicated through gas conveying pipes. The water conveying assembly comprises a hydraulic driving element and a first water conveying pipe. The first water-powder film-forming spraying element is communicated with the hydraulic driving element through the first water conveying pipe. When the water-absorbing resin composite material is combined with water, a gel state water film can be formed. The application has the effects of improving the emergency processing capacity of aerospace low-temperature chemical fuel leakage, controlling the accident in the budding and initial state in time, and reducing the possibility of secondary disasters such as fire and explosion.

Storage system for storing a cryogenic medium, in particular for storing hydrogen, comprising a storage container (1) for receiving the medium, wherein a first extraction line forms a fluid-conducting connection from the interior of the storage container (1) to a consumer connection (10) for connecting a consumer, wherein at least a first controllable line shut-off valve (6) and a first heat exchanger (3) are arranged in the first extraction line, wherein a second extraction line, different from the first extraction line, forms a fluid-conducting connection from the interior of the storage container (1) to a consumer connection (10) for connecting the same consumer, wherein at least a second controllable line shut-off valve (7) and a further first heat exchanger (3) are arranged in the second extraction line, characterized in thatthat the first withdrawal line additionally has a third controllable line shut-off valve (9) immediately before the consumer connection (10) and the second withdrawal line additionally has another third controllable line shut-off valve (9) immediately before the consumer connection (10).

The application discloses a kind of light weight liquid hydrogen conveying device and method, it is related to liquid hydrogen conveying technical field.The liquid hydrogen storage tank inner cavity of device is equipped with liquid hydrogen pump, top and gas cavity are communicated;Liquid hydrogen pump outlet pipeline is located in the top of liquid hydrogen pump and is communicated with the liquid hydrogen conveying pipeline equipped with liquid hydrogen vaporizer after vertically passing through gas cavity;Liquid hydrogen conveying pipeline located in the upstream of liquid hydrogen vaporizer is connected with the inlet of pressurized vaporizer by pressurized liquid hydrogen pipeline, and the outlet of pressurized vaporizer is communicated with one side of gas cavity by pressurized hydrogen pipeline equipped with second hydrogen regulating valve.The application uses liquid hydrogen pump to provide power for liquid hydrogen conveying, when stopping, using the gravity of liquid hydrogen itself carries out liquid seal in pipeline, and with the smaller volume normal temperature hydrogen gas stop valve, power equipment inlet pipeline is cut off;Low temperature gas cavity is used as heat buffer, and the vaporization of liquid hydrogen in liquid hydrogen pump outlet pipeline is slowed down, the sealing property of liquid seal is guaranteed, and liquid hydrogen pump is operated in liquid hydrogen environment.

The invention provides a rocket high-pressure gas cylinder mounting device and a rocket device, belongs to the technical field of rocket devices, and aims to solve the problem that an existing gas cylinder mounting device cannot effectively adapt to axial extension after gas cylinders are inflated. Comprising an upper gas cylinder connecting assembly and a lower gas cylinder connecting assembly, an upper support in the gas cylinder upper connecting assembly is connected with a stringer and a skin, a first square through hole is formed between the upper support and an upper gas outlet fixing piece, a square groove section of an upper gas outlet of a high-pressure gas cylinder penetrates through the first square through hole, and the length of the square groove section of the upper gas outlet is larger than that of the first square through hole. A lower support in the gas cylinder lower connecting assembly is connected with a stringer, a skin and a lower ring frame, a second square through hole is formed between the lower support and a lower gas outlet fixing piece, a square groove section of a lower gas outlet of the high-pressure gas cylinder penetrates through the second square through hole, and the length of the square groove section of the lower gas outlet is equal to that of the second square through hole. The device can adapt to inflation deformation of the gas cylinder, and the rigidity self-adaption effect is achieved.

Device and method for restricting a moving part. Device (100) and method for restricting a moving part (13) during acceleration, with a mass (110) and a component (120). The mass (110) moves inertia between a first position and a second position in response to the acceleration, while the component moves between a release position of the moving part (13) when the mass (110) is in said first position and a restricted position of the moving part (13) when the mass (110) is in said second position. Figure for the abstract: Fig. 1B.

Exemplary embodiments described herein may include lightweight, low stow volume solar concentrator.

An assembly is configured for use in a module that is configured to hold a plurality of vessels, at least one of the plurality of vessels including a neck. The assembly includes a bracket and a mounting ring. The bracket has an opening and includes a plate and a collar. The plate has a fixed position along the longitudinal axis, and the collar extends from the plate. The mounting ring includes a second aperture configured to surround the neck, and the mounting ring comprises an outer surface configured to move within the opening along the longitudinal axis. In another aspect, a module is configured for attachment to a plurality of pressure vessels and includes a frame and a bracket. A method is described for supporting a pressure vessel.