563results about "Cosmonautic environmental control arrangement" patented technology

A lightweight photovoltaic concentrator is disclosed. The concentrator comprises at least two sections each comprising a cell panel, a radiator panel and a reflector panel. In one embodiment, the concentrator system of the present invention comprises at least two hinged sections each comprising a solar cell panel, a flat radiator panel, and a curved reflective concentrator panel. The solar cell panel comprises at least one photovoltaic cell for generating electrical power in response to radiation. The solar cell panel is aligned with the radiator panel at an angle less than 180 degrees but not less than 90 degrees facing the reflective concentrator panel. In another embodiment, the cell panels on adjoining sections are angled in opposite directions with respect to the radiator panels and wherein the reflective concentrator panels are located on opposite sides of the radiator panels. A heat pipe or loop heat pipe is located in the radiator panel to dissipate heat from the solar cells.

A novel, high-efficiency, extremely light-weight, inflatable refractive solar concentrator for space power is described. It consists of a flexible Fresnel lens, flexible sides, and a back surface, together enclosing a volume of space which can be filled with low pressure gas to deploy the concentrator on orbit. The back surface supports the energy receiver/converter located in the focal region of the Fresnel lens. The back surface can also serve as the waste heat radiator. Prior to deployment, the deflated flexible lens and sides are folded against the back surface to form a flat, low-volume package for efficient launch into space. The inflatable concentrator can be configured to provide either a line focus or a point focus of sunlight. The new inflatable concentrator approach will provide significant advantages over the prior art in two different space power areas: photovoltaic concentrator arrays and high-temperature solar thermal conversion systems. Photovoltaic concentrator arrays using the new inflatable lens will be much lighter than prior space concentrator arrays. In addition, for photovoltaic concentrator arrays, the new inflatable lens will eliminate the need for a fragile glass superstrate to support the lens, substantially improving robustness of the lens. Solar thermal concentrator arrays using the new inflatable lens will be much lighter than prior art space concentrators which used parabolic mirrors. In addition, for solar thermal applications in space, the new inflatable lens will eliminate the need for high surface accuracy, which has been a significant problem for prior art concentrators.

A system for oxygen, hydrogen and carbon mass regeneration and recycling for breathing, and fuel/energy generation purposes, especially for fuel cells and rocket motors, by combination and integration of a photoelectrolytically powered electrochemical and gas handling system with one or more fuel cells.

An In Orbit Transportation & Recovery System (IOSTAR(TM)) (10) is disclosed. One preferred embodiment of the present invention comprises a space tug powered by a nuclear reactor (19). The IOSTAR(TM) includes a collapsible boom (11) connected at one end to a propellant tank (13) which stores fuel for an electric propulsion system (12). This end of the boom (11) is equipped with docking hardware (14) that is able to grasp and hold a satellite (15) and as a means to refill the tank (13). Radiator panels (16) mounted on the boom (11) dissipate heat from the reactor (19). A radiation shield (20) is situated next to the reactor (19) to protect the satellite payload (15) at the far end of the boom (11). The IOSTAR(TM) (10) will be capable of accomplishing rendezvous and docking maneuvers which will enable it to move spacecraft between a low Earth parking orbit and positions in higher orbits or to other locations in our Solar System.

A high temperature sealing system is operative to seal an interface between adjacent hot structures and to minimize parasitic flow between such structures that move relative to one another in-plane or out-of-plane. The sealing system may be used to seal thrust-directing ramp structures of a reusable launch vehicle and includes a channel and a plurality of movable segmented sealing elements. Adjacent ramp structures include edge walls which extend within the channel. The sealing elements are positioned along the sides of the channel and are biased to engage with the inner surfaces of the ramp structures. The segmented sealing elements are movable to correspond to the contour of the thrust-directing ramp structures. The sealing system is operative to prevent high temperature thrust gases that flow along the ramp structures from infiltrating into the interior of the vehicle.

The invention relates to a combined thermal protection and surface temperature control apparatus. In one embodiment, a combined thermal protection and surface temperature control apparatus comprises a porous member having an entrance side, and a separate exit side. One or more coolant entrance channels extend through the entrance side, extend part-way through the porous member, and end within the porous member before reaching the exit side. Conversely, one or more coolant exit channels begin within the porous member, extend through a portion of the porous member, and extend through the exit side. The coolant entrance and exit channels may be parallel and may alternate. The channels may only extend across a portion of the thickness of the porous member.

An In Orbit Transportation & Recovery System (IOSTAR™) (10) is disclosed. One preferred embodiment of the present invention comprises a space tug powered by a nuclear reactor (19). The IOSTAR™ includes a collapsible boom (11) connected at one end to a propellant tank (13) which stores fuel for an electric propulsion system (12). This end of the boom (11) is equipped with docking hardware (14) that is able to grasp and hold a satellite (15) and as a means to refill the tank (13). Radiator panels (16) mounted on the boom (11) dissipate heat from the reactor (19). A radiation shield (20) is situated next to the reactor (19) to protect the satellite payload (15) at the far end of the boom (11). The IOSTAR™ (10) will be capable of accomplishing rendezvous and docking maneuvers which will enable it to move spacecraft between a low Earth parking orbit and positions in higher orbits or to other locations in our Solar System.

A thermal protection system suitable as a heat shield for a space vehicle includes a flexible insulation mat that is adhesively bonded onto the outer skin of the space vehicle, a cover layer made of at least one ply of a ceramic fiber composite material secured onto the insulation mat, and a coating of an inorganic material applied onto an outer surface of the ceramic cover layer. The cover layer is secured to the insulation mat by adhesive bonding, stitching with ceramic threads or the like, or by a snap fastener button connection. The ceramic fiber composite materials of the cover layer include inorganic fibers which are preferably oxides, carbides or mixtures thereof embedded in a matrix of the same compositional components, whereby the components can be selected appropriately depending on the requirements of the particular application. The overall thermal protection system remains flexible, lightweight, and highly effective for thermal shielding, so that it may be used not only on the relatively cooler portions of the space vehicle, but also on the most highly thermally loaded areas such as the nose and the leading edges of wings and the like.

The invention relates to an actuating mechanism of a spacecraft for integrating the heat control and a liquid momentum wheel, comprising a liquid storage device, a circulating pump assembly, a heat-dissipating public loop, a heating assembly, a heating public loop and an irradiator/momentum generator integration device, wherein the circulating pump assembly is used for driving liquid working substance to flow; the heat-dissipating public loop includes a first flow rate distribution valve, a first heat exchanging device and a first bypassing branch, wherein the first heat exchanging device is connected in parallel with the first bypassing branch; the heating assembly heats fluid working substance for heating the heat-dissipating public loop; the heating public loop includes a second flow rate distribution valve, a second heat exchanging device and a second bypassing branch, wherein the second heat exchanging device is connected in parallel with the second bypassing branch; the irradiator/momentum generator integration device comprises at least two different pipelines which are connected in parallel and a device which controls the distribution of the liquid momentum in the different pipelines, and are used for dispersing heat out of the spacecraft and providing large enough and changeable angular momentum; the heat-dissipating public loop, the heating public loop and the irradiator/momentum generator integration device are sequentially connected and then in series connected with the liquid storage device and the circulating pump assembly to form a working substance circulating loop, and the liquid working substance flows in the working medium circulating loop.

The invention provides an ultra-high speed aircraft thermal protection and drag reduction method and system. The method includes that a cold source is arranged inside the cavity of an ultra-high speed aircraft, a plurality of micropores are arranged in the wall surface of the cavity, and the cold source is ejected out of the micropores in a high-pressure air state under the action of driving force to form an air film on the outer surface of the cavity. The air film can perform thermal protection on the ultra-high speed aircraft, viscous drag of the aircraft with external air can be effectively reduced, thermal barrier is favorably alleviated or eliminated, security of the ultra-high speed aircraft is improved, service life is prolonged, and a good application prospect is achieved. A cold source driving unit comprising an air pump, a cold source storer and a buffer is adopted in the system, the air pump supplies compressed air to the cold source storer during the operation state, the cold source enters the buffer to be vaporized under the action of air pressure, and the air is ejected from the micropores to form the air film under the action of pressure.

The invention provides a deployable radiator for a space vehicle such as a satellite. The radiator of the invention has two radiating faces, one of which is insulating over a portion of its area so that during launch and while the satellite is being moved onto station, with the radiator in the stored position, it naturally serves to keep the highest possible temperature on board the vehicle.

The invention discloses a composite radiation refrigeration film. The composite radiation refrigeration film comprises a top layer and a reflection layer arranged under the top layer. The reflection layer has high reflectivity for at least a part of solar radiation. The top layer comprises one or more polymers, and the polymer has an emissivity of not less than 80% at a wave band of 7-14 microns.The top layer comprises a first emission layer close to the reflection layer and a second emission layer far away from the reflection layer. The first emission layer comprises a first polymer and a plurality of first foam holes formed in the first polymer, and the second emission layer comprises a second polymer and a plurality of second foam holes formed in the second polymer, wherein the diameter of the first foam holes is 1-20 microns, and the diameter of the second foam holes is 1-200 nm.

A sensor system for monitoring for the presence of contamination with one or more contaminating biological, chemical and/or radioactive agents on a terrain surface, and creating a contamination map thereof. The sensor system includes a plurality of sensor pods, a ground station, an airborne dispensing portion, and an airborne monitoring portion. The sensor pods include a pod housing, a descent slowing airfoil, a detector unit for detection of the contaminating agent and outputting contaminating agent data, a processor for processing the contaminating agent data, a GPS unit for determining the position of the sensor, and a transmitter for transmitting contamination agent data and position data to the airborne monitoring portion. The airborne monitoring portion receives the transmitted data from the sensor pods, and relays the data to the ground station, where the contamination map is made available.

The invention provides a star sensor sub-low temperature independent thermal control device of an orbit transfer vehicle. The device comprises an imported star sensor A, a domestic star sensor B, thermistors (2), heaters (3), light shields (4), F46 films (5), an S781 white paint thermal control coating (6), mounting supports (7), a self support (8), heat insulation pads (8), a plurality of layers of heat insulation components (10) and thermal protection layers (11). Thermal resistance between stars and the star sensors is increased through the heat insulation pads, heat leakage compensation is greatly reduced, and a foundation is laid for independent temperature control over the star sensors. According to the sub-low temperature design requirement, the S781 white paint thermal control coating is adopted to increase the temperature adaptability of a star sensor motive track. The heat insulation components and the protection layers are adopted, and the radiation influence of a 25N engine on the imported star sensor A is effectively prevented. The device is convenient to implement, high in reliability and adaptability and small in total occupied resource, and the heat radiation of the engine can be effectively prevented.

A spacecraft instrument thermal control method and system providing scalable thermal control of on-board instrument temperature. Adapted for a spacecraft with bus supporting instrumentation and thermal radiator panels, cooling is carried out by one or more active coolers, such as cryocoolers, each mounted to the spacecraft at a radiator spatially separated from the instrument, with cold side of the cooler being thermally coupled to an instrument focal plane or other instrument location requiring cryogenic cooling and the warm side coupled to the radiator. A closed loop temperature control system measures the temperature of the controlled portion of the instrument, and adjusts active cooler drive signals to maintain a specified set point temperature.