291results about "Cosmonautic safety/emergency devices" patented technology

Structures based upon periodic cellular materials that provide a potential for defeating combinations of both air blast loading and ballistic attack either sequentially or simultaneously, or combination of both. The cellular structures may also be configured to meet the stiffness and strength support requirements of particular vehicle or other applications, systems or structures. The armor is therefore potentially able to support normal service loads and defeat blast and ballistic threats when necessary. The structure provides for using efficient load support capabilities of the material (without a high armor protection level) in low threat conditions, as well as the ability to modify the system to increase its level protection to a desired or required level. This would reduce the weight of the protection system in normal (low threat) conditions which reduces vehicle wear and tear, as well as cost savings in fabrication of applicable structures or systems.

A multilayer insulation with an array of rigid posts is provided. In particular, the posts are comprised of multiple post elements that are interconnected to radiation shields or sheets comprising the layers of the integrated multi-layer insulation structure. The post elements maintain spacing between adjacent sheets, thus maintaining a volume between the sheets. The post elements within a post can be attached to one another to form an integrated post structure. Moreover, neighboring posts can be interconnected to one another by beams. The post elements can also be configured so that, when the IMLI structure is not subjected to atmospheric pressure, the elements within a post are separated from one another to form gaps, thereby reducing heat transfer between layers.

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.

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.

A multi-layer tile material produced from layers of alumina enhanced thermal barrier material having different densities. The insulation layers are bound together by a high strength, high temperature alumina or silica binder having a coefficient of thermal expansion similar to that of the insulation layers. Use of the multi-layered tile allows the problems of tile slumping and of insufficient heat management associated with low density alumina enhanced thermal barrier tile to be overcome.

The invention relates to the field of the protective structures of spacecrafts, and discloses a lightweight filled composite protective structure for space debris, comprising a first layer of protective screen which is a corrugated board, a second layer of protective screen which is a corrugated board / density gradient metal foam; and spacing is left among the first layer of protective screen, the second layer of protective screen and a stiffener structure. The first layer of protective screen is in a corrugated shape, and the corrugated protective screen can effectively decompose debris cloud produced when the space debris with different incident angles collide with the protective structure at high speed, so that the damage energy is scattered. The corrugated surface plate of the second layer of protective screen decomposes the debris cloud again produced when the space debris collide with the first layer of protective structure, and a density gradient metal foam filling layer can further crush and melt the hit space debris, reduce the speed of the cloud debris, and minimize the damage to the external structure of a spacecraft and a sub-system.

The invention is a payload fairing separation system that is resettable and operates in two stages. The system includes releasable, resettable seam connectors that hold the fairing together in combination with releasable point connectors. At the time of separation, the seam connectors release first, and then the point connectors release. This allows the fairing to separate cleanly at a precise time, without using pyrotechnic charges, and with low shock to the payload.

A protective cover for a space suit, suitable for use in lunar or Mars visits which provides protection to the space suit against dust and organic contaminants while withstanding the rigors of contact with lunar or Martian rocks and equipment and not degrade or tear. The cover can be easily donned and doffed without contaminating the space suit. High strength coated fabrics which are functional in the thermal extremes of the lunar and Mars environments are suitable materials for the protective covers.

A fuel delivery system for a gas turbine engine includes a catalytic device for treating fuel to increase the usable cooling capability of an endothermic fuel. The catalytic device operates to treat and decompose components within in the fuel to render the fuel non-coking beyond 250° F. The catalytic device includes material that initiates reactions, and decomposition of coke forming components within the fuel to non-coke forming components within the fuel.

The invention discloses a spacecraft potential active control device based on a helicon wave plasma, comprising a discharge chamber, a magnet and a potential control electrode. The magnet is externally arranged on one end of the discharge chamber; a radio-frequency antenna is winded outside the discharge cavity and is adjacent to the magnet; the position of the discharge chamber working medium gas inlet of a discharge chamber working medium gas pipe is not beyond the front end of the adjacent radio-frequency antenna; the potential control electrode which is electrically connected with the conductive structural element of a spacecraft is arranged in the discharge cavity; a radio-frequency source is used for providing power supply for the radio-frequency antenna; and the working medium gas is ionized into the plasma, under the action of the electromagnetic field generated by the radio-frequency antenna and the magnetic field generated by the magnet. The invention further discloses the application of the spacecraft potential active control device based on the helicon wave plasma. The spacecraft potential active control device based on the helicon wave plasma has the advantages of low power consumption, light weight of used working medium, high potential control efficiency, high starting frequency and is superior to the prior art in the spacecraft potential control and the orbit energization effect risk control.

The invention discloses a pyramidal micro-truss sandwich plate bearing and thermal protection integrated structure with flow channels, which includes at least one structural unit, the structural unit includes an upper panel, a lower panel and a core, and the core includes multiple A micro-truss unit cell, the micro-truss unit cell is formed by connecting the tops of two pyramid-shaped lattice unit cells, and the pyramid-shaped lattice unit cell is a pyramid formed by four rods with circular cross-sections type structure, the left and right ends of the upper panel are respectively provided with flow channel wall plates extending towards the direction of the core. The present invention can improve the bearing capacity, defect resistance and impact protection capacity of the present invention at the same time. The flow channel is arranged near the heat source to further improve the heat dissipation effect. By adjusting the flow rate of the cooling liquid, the temperature of the lower panel can be efficiently controlled in one In an appropriate range, it is more conducive to controlling the sealing, reducing the amount of fluid used to reduce the overall weight coefficient of the structure, and the best performance can be obtained by adjusting the size of each part.

A cryogenic fuel tank assembly 10 is provided comprising a cryogenic fuel tank wall 22. A foam assembly 24 is affixed to the cryogenic fuel tank wall 22, the foam assembly 24 having an inner surface 30 and an outer surface 32. A first solid film 40 bonded to the outer surface 32 to provide a uniform outer bonding surface 42. A thermal protection system assembly 38 is bonded to the uniform outer bonding surface 42.

A cryogenic fuel tank assembly 10 is provided comprising a cryogenic fuel tank wall 22. A foam assembly 24 is affixed to the cryogenic fuel tank wall 22, the foam assembly 24 having an inner surface 30 and an outer surface 32. A first solid film 40 bonded to the outer surface 32 to provide a uniform outer bonding surface 42. A thermal protection system assembly 38 is bonded to the uniform outer bonding surface 42.

Advanced spaceport information management methods and systems are disclosed. In one embodiment, a method includes coupling a test system to the payload and transmitting one or more test signals that emulate an anticipated condition from the test system to the payload. One or more responsive signals are received from the payload into the test system and are analyzed to determine whether one or more of the responsive signals comprises an anomalous signal. At least one of the steps of transmitting, receiving, analyzing and determining includes transmitting at least one of the test signals and the responsive signals via a communications link from a payload processing facility to a remotely located facility. In one particular embodiment, the communications link is an Internet link from a payload processing facility to a remotely located facility (e.g. a launch facility, university, etc.).

The invention relates to a material, in particular for a thermal insulation layer, with increased thermal stability, a low heat conductivity and a large thermal coefficient of expansion. According to the invention, said material comprises lanthanides, in particular the elements La, Ce, Nd, Yb, Lu, Er or Tm, which preferably occur as a mixture in a Perovskite structure. Said thermal insulation layer is particularly suitable for replacing thermal insulation layers comprising yttrium stabilized zirconium oxides (YSZ) as the thermal stability thereof is given as well over 1200° C.

A process for evaluating space systems for safety assurance and residual risk to the flight crew. The process includes a success oriented System Safety phase which attempts to reduce the probability of failure with respect to the loss of a crew member as low as practical; a failure oriented SPACESAFE phase which assumes that at least one Safety Critical Subsystem has failed and attempts to engineer a risk mitigation design minimize adverse effects on the crew; and an Integration phase which complimentary integrates the System Safety phase with the SPACESAFE phase. Such process allows for increased flight crew safety by minimizing risk of a failures that contribute to loss of a crew member.

A launch vehicle upper-stage escape system is described that allows a crew capsule or a payload capsule to be safely and rapidly separated from a launch vehicle in the event of an emergency using the upper stage main engine for propulsion. During the initial portion of the flight the majority of the propellant mass for the upper stage is stored in the lower stage. This minimizes the mass of the upper stage allowing the upper stage main engine to provide sufficient acceleration to lift the capsule off of the launch vehicle and to move the capsule away from the launch vehicle to a safe distance with sufficient speed in the event of an emergency. It can also be used to lift the crew or payload capsule to a sufficient height for recovery systems to be employed successfully in the event of an on-pad or low-altitude launch emergency.

A satellite thermal management system for dissipating excess heat generated within a satellite. The thermal management system includes a central platform for mounting heat dissipating components. A first thermal radiator panel and a second thermal radiator panel are mounted to the central platform. The first thermal radiator panel is arranged to face East when the satellite is in orbit. The second thermal radiator panel is arranged to face West when the satellite is in orbit. The first and second thermal radiator panels are thermally coupled to the central platform and are configured to radiate heat dissipated by the heat dissipating component mounted on the central platform.

Thermal module (1) for use on a spacecraft to control thermal loads coming from a heat source (3) comprising a two-phase loop system (5) and a heat rejection system (13), the two-phase loop system (5) comprising a thermal collector (7), a heat flow regulator (10), a by-pass line (6) and a condenser (11), the condenser (11) and the heat rejection system (13) being thermally coupled, such that the heat flow regulator (10) of the two-phase loop system (5) redirects part of the thermal loads from the heat source (3) to the condenser (11), from which the heat rejection system (13) directs said thermal loads to a heat sink (14), the temperature of the heat source (3) being regulated by bypassing another part of the thermal loads back to the thermal collector (7) through the by-pass line (6) in a proportional manner, to avoid the heat source (3) overcooling, being the heat rejection system (13) designed based on the hottest possible conditions for the spacecraft mission.

An apparatus having a composite land-based structure with a first carbon nanotube infused material and a second carbon nanotube infused material. The first and second carbon nanotube infused materials each having a range of carbon nanotube loading selected to provide different functionalities.

A geosynchronous Solar Power Satellite System is created by an artificial gravity, closed ecology, multiple use structure in low earth orbit that manufactures modular solar power panels and transmitter arrays. This facility takes empty fuel tanks and expended rocket boosters from launch vehicles that are sent into low earth orbit, and re-manufactures them into structural components. These components are mated to solar cells that are launched from earth. The modular solar panels are transported to geosynchronous orbit by vehicles with ion engines, where the panels are mated to other solar panels to collect power. Structural components are also mated to transmitter elements launched from earth. These are likewise transported to geosynchronous orbit. They are mated to the solar power collecting panels and they beam the collected power back to earth.

A heat transfer assembly for a spacecraft is disclosed. The assembly includes an equipment panel having an east end and a west end. East and west radiator panels are coupled to the east and west ends, respectively, of the equipment panel. The assembly also includes a plurality of flexible heat pipes each having a first rigid tube thermally coupled to the east radiator panel, a second rigid tube coupled to the equipment panel, a third rigid tube thermally coupled to the west radiator panel, a first flexible tube sealingly coupled between the first and second rigid tubes, and a second flexible tube sealingly coupled between the second and third rigid tubes. The equipment panel is configured to retain one or more equipment modules in thermal contact with the second rigid tube of at least one of the plurality of flexible heat pipes.

A method for developing and utilizing a Space Cruiser vehicle efficiently combines resources from relevant end users in industry and the military. The preferred vehicle has an elongated conical shell with an elliptical transverse cross section. The cabin of the vehicle preferably has a circular transverse cross section, leaving space between the cabin and the shell to store vehicle fuel. The cabin is insulated and capable of supporting occupants in a shirt-sleeve, air-breathing environment whether in the Earth's atmosphere or in outerspace. The vehicle is capable of being launched from conventional aircraft, and also being stored in and launched from a space shuttle or orbiting space station.

The invention discloses a space debris prevention structure of a spacecraft, which comprises a flat and vertical inner board, a corrugated outer board and a metal foam sandwich layer, wherein the metal foam sandwich layer is arranged between the flat and vertical inner board and the corrugated outer board. The space debris prevention structure of a spacecraft disclosed by the invention can effectively resolve the debris clouds generated by impacting the prevention structure at a high speed by the space debris of different incidence angles and scatter the damage energy of the debris clouds; the filled metal foam can further crash and melt the impacted space debris and reduce the speed of the debris clouds; when the debris clouds pass through an inner flat and vertical metal panel, the inner panel further scatters the damage energy of the debris clouds, reduces the speed of the debris clouds again and stops the residual space debris after passing through the metal foam layer to a certain extend, so as to effectively reduce the damages of the space debris to the external structure of a spacecraft and a sub system.