132results about "Cosmonautic components" patented technology

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

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.

A diffuser for separating a gas component and a liquid component of a liquid-gas mixture to reduce slosh in vehicle systems is described herein. The diffuser includes a wall with a gas end, a liquid end, and multiple vanes partially extending from the liquid side to the gas side. The vanes are spaced apart from each other by a given distance to permit the liquid to flow toward the liquid end via capillary action.

Systems and methods for deposition-based manufacturing and quantum spin separation in space environments are disclosed herein. A method of manufacturing a film in a space environment includes melting a metal in a melting pot using concentrated solar energy and pushing the melted metal through small openings in the melting pot to produce a liquid stream. The method further includes dispersing the liquid stream into liquid drops of the melted metal and depositing the liquid drops onto a mold surface.

The invention relates to a double-wall storage tank. The double-wall storage tank comprises an inner container surrounded by an outer sheath. More specifically, the present invention relates to a double-walled storage tank comprising an arrangement for coupling an inner container relative to an outer jacket.

The utility model relates to the technical field of aerospace, in particular to a servo actuator zero position adjusting tool which comprises a guide rail, a plurality of sliding blocks, a plurality of sets of measuring pins and a vernier caliper. The upper surface of the guide rail is provided with a downwards-concave sliding groove, and each sliding block is located in the sliding groove and slides in the sliding groove in the extending direction of the sliding groove. Each group of measuring pins is provided with two same pin shafts, the lower end of one pin shaft is fixed to the upper surface of one sliding block, the upper end of each pin shaft vertically extends upwards, and the two pin shafts of each group of measuring pins are used for clamping the two ends of a servo actuator of a rocket; one end of the vernier caliper is fixed to the outer side surface of one pin shaft of one group of measuring pins, and the other end of the vernier caliper is clamped to the outer side surface of the other pin shaft of the group of measuring pins. According to the invention, the accuracy of zero distance measurement of the servo actuator of the rocket can be ensured, so that the success of rocket launching and aerospace tasks is ensured.

A collapsible space enclosure system is disclosed, designed for efficient transport and deployment in the field of mechanical engineering. The system described addresses the technical problem of cumbersome and inefficient collapsible structures by providing a system that integrates a telescopic lifting central section with an upper and lower portion for balanced vertical retraction and extension. The system includes folding panels connected to the central structure, configured to fold in two for compact shipment. Actuators enable continuous 360° rotation and balanced vertical movement. Parallelogram linkage assemblies facilitate simultaneous lateral retraction and vertical lifting. Retractable strut channel assemblies allow for the installation of add-on retractable spaces, increasing floor area. This system is particularly useful for military, commercial, and recreational applications, offering a streamlined and versatile solution for rapid deployment and compact storage.

The invention discloses a composite material pressure container suitable for a liquid oxygen environment and a preparation method of the composite material pressure container, and belongs to the technical field of composite materials. The composite material pressure container suitable for the liquid oxygen environment sequentially comprises a thin-wall metal lining layer, a transition layer and a fiber winding composite material layer from inside to outside, the transition layer comprises fluorine-containing polyarylether; the fiber winding composite material layer is formed by carbon fiber woven cloth dipped with composite epoxy resin and carbon fiber yarns on the lining layer in a continuous annular and spiral cross winding manner; matrix resin in the composite epoxy resin comprises bromine-containing epoxy resin and / or chlorine-containing epoxy resin. By arranging the thin-wall metal lining layer, the transition layer and the fiber winding composite material layer, each layer of structure has good liquid oxygen compatibility; and by arranging the transition layer, the combining capacity of the thin-wall metal lining layer and the fiber winding composite material layer is further improved, and then the overall mechanical property and anti-permeation performance of the pressure container are improved.

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 relates to the technical field of rocket power tests, and discloses a liquid oxygen supply system and a liquid oxygen supply method.The liquid oxygen supply system is characterized in that a liquid oxygen storage tank is provided with an extrusion gas inlet and a liquid oxygen inlet and outlet; the gas extrusion subsystem extrudes liquid oxygen in the liquid oxygen storage tank; the liquid oxygen filling subsystem is used for filling liquid oxygen into a liquid oxygen storage tank; the main regulating valve is connected with the liquid oxygen inlet and outlet; the input ends of the plurality of parallel branches are connected with the main regulating valve, and the output ends of the plurality of branches provide liquid oxygen for test pieces at different stations respectively; a branch regulating valve is arranged on each branch; the flow and pressure of each branch are adjusted by adjusting the opening degree of the main adjusting valve and adjusting the opening degree of the branch adjusting valves. The flow of each branch is adjusted by adjusting the opening degree of the adjusting valve of each branch, so that the resistance of each branch is as close as possible, and the high-precision control of the flow and pressure of test pieces at different stations is achieved; and multi-path equal-pressure equal-flow liquid oxygen distribution is realized. And the main regulating valve is combined to regulate the total flow.

The invention provides a liquid nitrogen heating, vaporizing and pressurizing system, a pressurizing control method and a liquid rocket. The liquid nitrogen heating, vaporizing and pressurizing system comprises a first-stage oxygen tank, a first-stage fuel tank and a liquid nitrogen tank, and the liquid nitrogen tank is arranged in the first-stage oxygen tank; the liquid nitrogen tank is sequentially connected with a liquid nitrogen conveying pipeline, a first liquid nitrogen heating and vaporizing unit, an oxygen tank pressurizing main pipeline and an oxygen tank energy dissipater arranged in the primary oxygen tank; the liquid nitrogen tank is sequentially connected with a liquid nitrogen conveying pipeline, a second liquid nitrogen heating and vaporizing unit, a fuel tank pressurizing main pipeline and a fuel tank energy dissipater arranged in the first-stage fuel tank; the liquid nitrogen box is sequentially connected with a liquid nitrogen conveying pipeline, a first liquid nitrogen heating and vaporizing unit, an oxygen box pressurizing main pipeline and a liquid nitrogen box pressurizing pipeline. The three storage tanks, namely the oxygen tank, the fuel tank and the liquid nitrogen tank, are pressurized by heating and vaporizing the liquid nitrogen, so that the cost of the pressurization system is reduced.

A storage system for storing a cryogenic medium, in particular, for storing hydrogen. The storage system includes storage container for receiving the cryogenic medium, at least one pipe projecting from outside the storage container into the storage container, and a shut-off valve in fluidic communication with the at least one pipe. The at least one pipe is closed at an end thereof facing away from the storage container and is open at another end thereof located in the storage container. The shut-off valve is moveable between an open operating state in which an inner space of the at least one pipe is in fluidic communication with an inner space of the storage container, and a closed operating state in which the inner space of the at least pipe is not in fluidic communication with the inner space of the storage container.

An isolable double-walled insulation tank for a launch vehicle comprises: 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, wherein 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 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 present invention relates to a reinforced bulging tank of a launch vehicle and a manufacturing method therefor. The tank is made of metal material with good plasticity. Firstly, the tank is manufactured by welding annealed plastic metal materials, wherein internal longitudinal reinforcing ribs and internal annular frames / annular plates of barrel sections are welded with a barrel section shell by a low energy laser welding method, and tank bottoms are progressively welded by a plurality of conical sections. Weldments are strengthened and formed through internal pressure bulging under the constraint of external tooling. The outer walls of the barrel sections are not radially deformed under the constraint of a plurality of small width metal rings, and the metal rings are not connected to each other axially. The strength of the materials can be greatly improved by a deep cooling bulging technology.

The invention relates to the technical field of propellant storage and management, in particular to a liquid oxygen tank for a space gas station, the liquid oxygen tank comprises a pressure-resistant shell, a mixing channel assembly, a superconducting coil assembly and an exhaust port, the mixing channel assembly is of a double-layer thin-wall circular tube structure and comprises an inner-layer circular tube and an outer-layer circular tube; one end of the inner-layer circular tube is in clearance fit with an assembly hole in the inner wall of the pressure-resistant shell, and the other end of the inner-layer circular tube extends out of the pressure-resistant shell to form a liquid inlet; the outer-layer circular pipe integrally surrounds the outer part of the inner-layer circular pipe, and one end, close to the liquid inlet, of the outer-layer circular pipe extends out of the pressure-resistant shell to form a liquid outlet; the superconducting coil assembly is wound on the outer wall of the inner-layer circular tube; and the exhaust port is formed in the pressure-resistant shell. According to the device, liquid oxygen can be injected or discharged in a microgravity environment, a container for storing and managing a liquid oxygen propellant is provided for a space gas station, and therefore on-orbit liquid oxygen discharging management and repeated liquid oxygen filling of a spacecraft are achieved.

A system that will enable humans to live on the moon and utilize the moon for the benefit of Earth. [Solution] Utilizing vertical and horizontal caverns on the moon. A roof-like cosmic ray protection plate 9 is installed above the elevator 5 and staircase 4 installed in the vertical cavern 2 on the moon to block radiation, meteorites, etc. 8. Since the moon receives light not only from the sun but also from the earth, if this is converted into a solar or geophotovoltaic power generation panel, it can also be used to generate electrical energy for use on the moon. Similar to the vertical cavern protection plate 9, an upper horizontal cavern protection plate 10 is installed above the horizontal cavern 6 on the moon to prevent cosmic rays, etc. 8 from entering the cavity 6. 11 is the staircase entrance. The staircase 4 can be a spiral staircase or a spiral staircase. 12 is the exterior entrance / exit of the elevator, 13 is the elevator car, and 14 is the entrance / exit of the elevator inside the cavity. [Effects] Humans can live on the moon safely, building houses, towns, factories, and hospitals without fear of cosmic rays or ultraviolet rays. It can also supply energy to Earth, making up for the Earth's energy shortage, without polluting the Earth's environment.

A storage system for storing a cryogenic medium, in particular for storing hydrogen, comprising a storage container (1) for receiving the medium, wherein at least one pipe (21) projects from outside the storage container (1) into the storage container (1), wherein the pipe (21) is closed at its end facing away from the storage container (1) and is open at its end located in the storage container (1), wherein a shut-off valve (22) is arranged on the pipe (21) such that the interior of the pipe (21) is fluid-conductingly connected to the interior of the storage container (1) when the shut-off valve (22) is open and the interior of the pipe (21) is not fluid-conductingly connected to the interior of the storage container (1) when the shut-off valve (22) is closed.

The invention relates to a method for operating a fluid-flowable component of a fluid storage system. The invention further relates to a fluid-flowable component of a gas storage system. It also relates to a controller for controlling the system component and to a vehicle with a gas storage system that includes the fluid-flowable component operated by the method according to the invention. The method according to the invention for operating a fluid-flowable component of a fluid storage system of a vehicle, wherein the fluid-flowable component is heatable, is characterized in that the heating of the fluid-flowable component is controlled in a situation-dependent manner, taking into account an expectation of a future change in the situation.

1. A valve (1) for a fluid reservoir, comprising: an inner part arranged inside the reservoir (3), the inner part including at least a part of a sensor (34, 35) for measuring at least one parameter characterizing the fluid; the valve (1) being capable of receiving information originating from an external computer (5); the valve (1) comprising an outer part arranged outside the reservoir (3), the outer part including first acquisition means (9) for acquiring data originating from the sensors (34, 35); communication means (11) using a digital cable communication interface for two-way communication with the computer (5); operation means (13) configured to take into account the information received from the computer (5) and the data originating from the sensors (34, 35) to operate an actuator (100) of the valve (1), the actuator (100) being coupled to the actuator (100) by at least one electrical connection (15); and means (17) for measuring at least one current parameter within the electrical connection (15).

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).

Techniques and materials provide for thermal insulation for tanks, such as tanks for use in space flight missions. Microporous glass blocks may be used as an insulative bridge between spray-on foam insulation and a stainless steel wall of a tank, for example. Blocks of microporous glass may be mechanically attached to the interior of the tank, after which spray-on foam insulation is applied directly to the blocks. The bridging insulation provided by the microporous glass blocks allows the temperature of the stainless steel tank wall to exceed the usable temperature of the spray-on foam insulation by preventing a substantial portion of heat of the stainless steel wall from reaching the spray-on foam insulation. An insulated tank may instead comprise a tank wall that is a sandwich-structured composite layer including an exterior skin, an interior skin, and a core that presents a relatively long thermal path, for improved thermal insulation.

The invention relates to a terminating device of a fibre composite material structure, wherein the terminating device comprises a fibre composite material connecting portion for connecting a tubular fibre composite material structure to a connecting device, wherein the connecting portion has at least one fibre deflecting member in its interior, wherein the course of the long fibres from the fibre composite material structure follows the shape of a fibre deflection portion of a fibre deflection member so that their fibre direction is deflected at the fibre deflection portion, and wherein the long fibres do not completely wrap around the fibre deflection members with which they are associated, the fibre deflection members being made of fibre composite material. According to the invention, the connecting device has a first element and a second element, wherein the second element has a contact area with the fibre composite material structure on its outer circumference and a contact area with the first element on its inner circumference, wherein each of the elements has a conical shape in the contact area with the other element. Advantageously, the conical shape of the elements of the connecting device can create a wedge, which causes the elements of the connecting device to generate a radial preload in the fibre composite material after the fibre composite material has been applied.

In one embodiment, a controlled closed-ecosystem development system (CCEDS) includes one or more a closed ecological systems (CESs) each having one or more controlled ecosystem modules (CESMs). Each CESM can have a biome containing at least one organism, and equipment comprising one or more of sensors, actuators, or components that are associated with the biome. A controller operates the equipment to effect transfer of material among CESMs to optimize one or more of organism variety, size, population, capacity, or sustainability of a biome of at least one CESM.

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.

Various embodiments relate generally to additive manufacturing and construction techniques to form structures with embodiments directed to computer software and systems, and control systems, and, more specifically, to a computing and a mechanical platform configured to implement local material to form a structure by selecting or filtering a subset of particulate that is deposited in a form at which adaptive levels of energy are applied to construct structures in-situ in a vacuum additively (e.g., three-dimensionally, or in “3D”), whereby adaptive levels of energy may be generated by one or more lasers and may be configurable to control temperatures associated with, for example, crystallization of indigenous particulate.

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 present invention discloses a method for designing an orbit of a kinetic impact mission on a near-Earth asteroid with the optimal deflection effect, comprising the following steps: S1, selecting the initial parameters of the spacecraft, including the departure and arrival times of the spacecraft; S2, designing the spacecraft orbit as a splicing of three segments, namely a large elliptical Earth parking orbit, an Earth escape orbit starting from the large elliptical orbit, and an interplanetary transfer orbit, and calculating the orbit parameters; S3, evaluating the deflection effect of the target near-Earth asteroid; S4, traversing all possible departure and arrival times of the spacecraft, repeating steps S1 to S3, drawing a contour map of the deflection distance versus the departure and arrival times of the spacecraft, and determining the optimal spacecraft launch window from it after eliminating unreasonable times. By adopting a large elliptical Earth parking orbit and screening the spacecraft launch window, the present invention enables the mass and relative velocity of the spacecraft to be relatively large when impacting the asteroid, thereby achieving the optimal asteroid deflection effect.