Abstract: A toroidal pressure vessel storing pressurized gas defines an outwardly exposed axially extending structural part of an automotive vehicle. Heat transfer from outside the vehicle directly to the pressure vessel advantageously offsets gas-expansion refrigeration effect when gas is used from the vessel.

Abstract: The invention comprises a combination boundary layer control system that utilizes both suction and cryogenic wall cooling to reduce aircraft drag. The invention is particulary useful in circumstances where liquid hydrogen or other cryogenic fuels are used to operate the aircraft.In the preferred embodiment of the invention, a network of D-tube ducts are provided that are fluidly connected to a cryogenic fluid source. These cryogenic fluid ducts are selectively fed cryogenic fluid to reduce aircraft surrface temperature and promote adhesion of boundary layer air to the aircraft. An air duct system is also provided to apply suction to the aircraft surface in order to remove the boundary layer at low altitude. The suction is applied by means of a multitude of minute holes in the aircraft skin which are fluidly connected to the air duct system.

Abstract: The present invention relates to a reusable thermal protection member for an object subject to great heating, e.g. an aircraft or spacecraft.The member comprises in per se known manner a refractory shield and an insulating layer between the refractory shield and the structure of the craft. According to the invention the refractory shield comprises at least one thin film of refractory material having an adequate flexibility to adapt to the shape of the structure of the craft in the area to be protected.Application to the thermal protection of aircraft or spacecraft.

Abstract: A method of manufacturing a dielectric material for use in spacecraft, comprising the operation of coating a dielectric substrate with a layer comprising a metal oxide dispersed in a solution of a polymer material of a thickness such that when the polymer solvent has evaporated the resultant layer has a thickness comparable with the expected maximum range of incident electrons.

Abstract: In order to dissipate aerodynamic heating in hypersonic aircraft, while at the same time providing structural support for the leading edge of an airfoil or the nose cone in a manner that avoids transmittal of thermal bending loads into the aircraft, a structural cooling unit is formed of a load bearing structure. The load bearing structure is of hollow construction defining a fluid flow path therethrough and has a fluid inlet at one end of the fluid flow path and a fluid outlet at the other end of the fluid flow path which comprises a tortuous path through the load bearing structure. Additionally, for purposes of forming a structural load path, the load bearing structure is formed into an elongated tubular configuration having at least one substantially continuous heat exchanging surface for aerodynamic heat dissipation.

Abstract: An ultra hypersonic aircraft having aerodynamic heat transferring pipes forming an airframe structure is shown. The heat transformer pipes include a high pressure steam generator located at the forward section thereof and a low pressure steam condensator located at the rearward section thereof. Low pressure water is circulated between the rearward section and the forward section for heat balancing of the air frame structure and for operating the room air conditioning system of an aircraft. A steam pressure reactant turbine, which drives an electric generator, is used as a boundary between the high and low pressure steam sections. The electric generator provided electrical and power supply for the infrared or arc heaters, and electric heaters recessed in the ceiling and wall panels of an acrodynamic lift-thrust generating channel for back-heating of the turbo-ram oval thrust stream in the thrust generating channel of a hypersonic aircraft.

Abstract: Aircraft thermal management is optimized by integration of normally separate airframe and engine fuel transfer/heat sink systems into a composite heat sink and secondary power source. Fuel temperature is controlled to limits set by coking and varnishing within the engine fuel nozzles. Fuel in excess of engine demand is returned to the aircraft's main fuel tank after being cooled by either aerodynamic surfaces or by a "fuel to ram-air" heat exchanger. Airframe and engine systems are cooled through use of fuel as coolant passing through heat exchangers mounted in the feed line from the "cool" main tank to the engines. The Engine Electronic Control Unit (ECU) is cooled by the airframe Environmental Control System (ECS) increasing its operational reliability. Controlling engine burn fuel to nozzle temperature limits results in lower main tank fuel temperature, less fuel boiloff, smaller sized ECS, greater heat sink capacity and less dependence on supplemental ram air.

Abstract: A cooled liner assembly for a nozzle has a plurality of longitudinally aligned support hinges secured to a support plate. A thin plate liner of arcuate shape is hinged to the support hinges with cooling air supplied between the liner and support plate. The liner is lightweight resisting the differential pressure of 25 psi in membrane hoop stresses, with components experiencing bending moments remote from the hot gas inside the nozzle.

Abstract: A cooling arrangement for the trailing edge of a nozzle flap 18 has a trailing edge liner 30,54 and a baffle plate 32,55. Air from plenum 22 passes through openings 38,54 to cool the trailing edge liner. Expansion of the nozzle liner 18 causes abutment seal 52 to open, increasing cooling air in proportion to the temperature of the nozzle liner.

Abstract: An integral cooling system for high temperature missile structures which is essentially a passive cooling system integrated into the missile structure. An evaporative material carried in the missile structure is ejected by vapor pressure through openings in the missile's outer skin and cools the missile as it flows along the outer surface thereof.

Abstract: Avionics systems are cooled by circulating the heated air through an onboard closed loop system which includes a heat transfer envelope comprised of a liner and the fuselage of the flight vehicle. The liner keeps a constant, thermodynamically optimal passage height between the liner and the fuselage, and is contoured in order to maintain that optimal height between the liner and internal structural support members which abut the fuselage and pass through the heat transfer envelope. The liner is comprised of flexible interlocking liner sections, each of which is produced from the same mold using heat and flame resistant material and subsequently trimmed to precisely fit into its particular area. Each liner section has a convex ridge which fits over an interal structural support member to enclose the member within the heat exchange envelope. Each convex ridge has transverse flexible ribs which accept and removably grasp the support member, thereby removably attaching the liner section to the support member.

Abstract: A temperature control device for a jet aircraft engine installation that includes a jet engine and a nacelle for supporting and enclosing the jet engine. The jet engine has a primary exhaust nozzle from which the engine produces a rearwardly extending primary exhaust plume of hot gases. The primary exhaust plume has a centerline and a generally cylindrical shape about its centerline. Each engine installation includes a nacelle-associated structure, such as a support strut, that extends in a generally fore and aft direction above the centerline of the primary exhaust plume. The temperature control device comprises a deflector formed at the lowermost surface of the nacelle-associated structure. The deflector includes a sheet-like lower surface having a lateral extent greater than the lateral extent of the nacelle-associated structure immediately above the deflector, and extends longitudinally aft of the primary exhaust nozzle.

Abstract: The present invention discloses a ceramic reusable externally applied thermal protection system (TPS). The system functions by using the composite device created by combining an upper shell, thermal insulation and lower tile components shown in FIG. 2. The upper shell effectively separates its primary functions as a toughened outer-surface and load carrier. In a preferred embodiment, a structurally strong ceramic/ceramic upper shell is manufactured from ceramic cloth upon which additional ceramic material has been deposited and maintains the configuration of the upper shell. The lower tile is composed of lightweight ceramic normally acceptable in this application. One or more lightweight rigid or flexible insulation materials are used in the cavity created by combining the upper shell and lower tile. The assembly of the overall tile is facilitated by attachment means effective to withstand the temperature and stress conditions.

Abstract: A direct air cooling system 12 for electronics carried by aircraft. The cooling system provides a submerged air scoop 16 which directs outside air to several electronic modules 14. The air flows through passages 76 in the modules 14 which are adjacent to circuit boards 74, 78 bearing discrete electronic components 88. A foil layer 86 and aluminum fin stock help transfer heat from the electronic components to the directed air. Heated air is then exhausted through exhaust port 24 or 28.

Abstract: An edge situated in high speed flow of medium is cooled by a tubular heat pipe having a longitudinal axis situated parallel to the edge. The exterior surface of the evaporator section of the heat pipe forms the edge. The heat pipe contains a working fluid existing in both gaseous and liquefied states. Heat input along the edge evaporates working fluid in the liquid state located in the evaporator section, and the heat is transferred transversely across the longitudinal axis of the heat pipe by the flow of vapor to the condenser section. The vapor condenses in the condenser section and releases its latent heat of vaporization. The condensate flows to the evaporator section by means of wicks composed of capillary material. The heat released into the walls of the condenser section is transferred through the walls by conduction and into another heat transfer device, for example, an active cooling apparatus or the evaporator sections of abutting heat pipes of conventional tubular configuration.

Abstract: A replaceable tip for an aircraft leading edge and which has a metallic abrasion shield glove removably mounted to the leading edge of the aircraft, fasteners for securing the abrasion shield glove to the leading edge of the aircraft, and a non-metallic resilient cushion insert fixedly attached to the abrasion shield glove so that the resilient cushion insert forms a dampening arrangement which reduces the amount of damage to the aircraft leading edge when the replaceable leading edge tip is hit.

Abstract: A structural component which is coolable for use under high thermal load conditions, such as a turbine blade, has a metallic support core with cooling ducts separated by lands in its surface. The core and its cooling ducts and lands are enclosed by an inner layer of metal felt and an outer layer of heat insulating ceramic material which partially penetrates into the metal felt to form a bonding zone between the felt and the ceramic material. Thus, any heat passing through the ceramic layer is introduced into the large surface area of the metal felt enabling the latter to efficiently introduce the heat into a cooling medium flowing in the ducts, thereby preventing thermal loads from adversely affecting the metal core to any appreciable extent.

Abstract: The invention concerns a composite assembly forming a thermal protection or dissipation screen comprising a hot wall and a cold wall, wherein it includes between said hot wall and said cold wall at least two adjacent elements called vaporizer superheaters of a refractory material, each comprising within itself a socalled reactive substance capable of releasing at a given temperature specific to each element a non flammable gas or vapor, each of said elements being separated from the other by a wall sealed from the temperature of release of vapor or gas of the element located the closest to said hot wall but capable of becoming porous as soon as such temperature is exceeded, each element due to the presence of said reactive substance evolving from the vaporizer function to the superheater function depending on development of the thermal aggression to which it is submitted.

Abstract: In a hollow airfoil, thin metal baffle sheets are bonded to opposing suction and pressure sides of a longitudinally extending cooling air cavity. The longitudinally extending downstream edges of these opposing baffle sheets are closely spaced apart defining a cooling air outlet slot for the cavity which may, for example, feed cooling air to the trailing edge portion of the airfoil. If the rearward edges of either baffle sheet become unbonded from the inside cavity wall the edge of the baffle sheet might bend inwardly and close the outlet slot. To prevent this possibility a plurality of pedestals extend outwardly from each side of the cavity wall opposite the rearward edge of each baffle sheet and extending substantially to the inwardly facing surface of that baffle sheet trapping it against the wall.

Abstract: Liquid hydrogen stored aboard a space vehicle, and constituting a fuel for use in the vehicle's main rocket motor, is introduced, in a gaseous state, into an inflatable aerodynamic braking member, for moving it from a collapsed stored position outwardly into an inflated operational position. Such member provides a large area braking surface which extends generally laterally outwardly from the space vehicle. During reentry the space vehicle is oriented to reenter nozzle end first. Fuel is consumed within the rocket motor to produce gases which are discharged out through the rocket nozzle to provide a cooling layer of gases forwardly of the large area braking surface. A space vehicle equipped with such an aerodynamic braking system is operated to reenter into the atmosphere. The drag produced by the braking member slows the space vehicle down and then the space vehicle is set into a low earth orbit.

Abstract: A novel heat management system for aircraft is provided which is based on the aircraft fuel as the heat exchange medium and comprises a dedicated thermal reserve fuel tank for containing refrigerated aircraft fuel (or other expendable liquid) which is cooled by heat exchange with the primary aircraft fuel flow to the engines, a fuel line loop for conducting fuel through a plurality of heat exchangers associated with cooling loops for heat generating aircraft systems, and a central microprocessor based controller for controlling fuel flow and temperature throughout the fuel loop system and into the aircraft engine.

Abstract: Cooling for an electronic package mounted on an aircraft engine housed in a nacelle is achieved by connecting the package to the inlet of the engine and overboard of the nacelle so that substantially continuous air flow over the electronic package will ensue from either direction depending on the varying pressure gradients adjacent the points where connections are made and the conditions exists during the flight envelope. This negates the necessity of providing positive pumping devices for effectuating the flow of the cooling air.

Abstract: A space vehicle 10 is steered from outer space toward the earth's atmosphere main rocket nozzle end first. An inflatable braking member 28 is deployed around the vehicle 10 by introducing gas into it. The main rocket motor 16 is ignited to produce a substantially throttled stream of gases. This stream provides a cooling layer forwardly of the braking member 28 to protect the member 28 from heating by friction with the atmosphere. The reduction of the velocity of the vehicle 10 is controlled by directly varying the drag of the vehicle 10 to compensate for variations in the density of the atmosphere. The drag is varied by varying the thrust of the motor 16 and/or varying the shape of the member 28. A backup method is to deflate the member 28 in the atmosphere. The member 28 may be jettisoned after the vehicle 10 leaves the atmosphere to enter a low earth orbit.

Abstract: Spacecraft are subject to electron bombardment resulting in electrical charge build-up on dielectric materials, such as thermal blankets, used to control heat gain and loss from the craft. The invention provides a multiple-layer material including inner and outer thermally insulating layers and an intermediate conductive layer, the outer insulating layer having a thickness such that many of the bombarding electrons will come to rest at or near the conductive layer and the resulting charge conducted away. Material may be in the form of a blanket with e.g. Kapton inner and outer layers or in the form of a glass tile. Material is also usable in other situations where electrical charge build-up may be a problem, e.g. for insulating explosive liquid containers.

Abstract: A thermal dissipation screen includes a hot wall and a cool wall and between the hot wall and the cool wall, a supple material of the layer, mat, fiber network type, used as a support or reinforcement for a pulverulent material capable of absorbing and retaining a liquid to produce a paste or a gel of a high thixotropic efficiency, such material constituting a hydrator-overheater element in the presence of the liquid evolving from the hydrating function to the overheating function, depending on the development of the thermal aggression.

Abstract: A reusable, externally applied thermal protection system is disclosed and functions by utilizing shell tile structure 10 which effectively separates its primary functions as an insulator and load absorber. Tile 10 consists of structurally strong upper and lower metallic shells 12,16 manufactured from materials meeting the thermal and structural requirements incident to tile 10 placement on the spacecraft. A lightweight, high temperature package of insulation 26 is utilized in upper shell 12, while a lightweight, low temperature insulation 28 is utilized in lower shell 16. Assembly of tile 10, which is facilitated by self-locking mechanism 20, may occur subsequent to installation of lower shell 16 on the spacecraft structural skin 30.

Abstract: A laser hardened casing structure for a missile or the like is described herein which comprises, in a preferred embodiment, a pair of thermally protective layers with a laser hardened barrier layer sandwiched therebetween and comprising a heavy metal bearing resin impregnated carbon fabric, all of said layers being adhesively joined and secured to the outer surface of the casing structure to be protected.

Abstract: A heat resisting canopy which utilizes a mosaic overlay of quartz or heat resisting glass elements adhesively bonded to a transparent organic shell. A transparent adhesive bonds each mosaic element to the organic canopy shell. A plurality of metallic ribbons, egg-crated together to form a lattice, stabilizes the mosaic structure, and individual elements are retained by a metallic wire mesh which covers the mosaic layer and which is attached to the canopy shell.

Abstract: A spacecraft modular system with two components of a cooling system, one in each replacement hardware module and one in the satellite station, with a coupling means in the replacement module in the form of heat sink which collects heat from the heat generating hardware in the replacement module. The heat sink incorporates heat sink fins which interdigitize with coolant tubes in the satellite coolant system and make a mechanical interface with the fins in response to the satellite coolant system pressure for a thermally efficient heat exchange from the replacement module to the satellite coolant system. In another embodiment, replacement module coolant tubes interdigitize directly with the satellite coolant tubes in a fluid pressure responsive thermally efficient mechanical contact to provide a fluid-to-fluid mechanical coupling.

Abstract: A ramjet powered vehicle having a bleed slot in the subsonic diffuser, for the ramjet combustor, which supplies air to a turbine for supplying power to vehicle accessories. A vortex tube is provided between the bleed slot and the turbine to provide a cool air supply. Radar absorbing material is positioned adjacent ramjet inlet with a cooling duct provided to supply cooling air from the vortex tube to the ramjet inlet cowl lip to provide cooling for the radar absorbing material.

Abstract: The invention is for a system which provides total cooling for an aircraft airframe which is designed to fly in the speed range of Mach 2 to Mach 8. The system eliminates the necessity of shielding an aircraft airframe constructed of material such as aluminum. Cooling is accomplished by passing a coolant through the aircraft airframe, the coolant acting as a carrier to remove heat from the airframe. The coolant is circulated through a heat pump and a heat exchanger which together extract essentially all of the added heat from the coolant. The heat is transferred to the aircraft fuel system via the heat exchanger and the heat pump. The heat extracted from the coolant is utilized to power the heat pump. The heat pump has associated therewith power turbine mechanism which is also driven by the extracted heat. The power turbines are utilized to drive various aircraft subsystems, the compressor of the heat pump, and provide engine cooling.

Abstract: A base vent assembly for entry space vehicles having an access cover utilized in conjunction with a plurality of spaces and deflector member so as to establish a labyrinth vent adjacent the access opening of the entry vehicle. The labyrinth vent is created by carefully forming a passageway between the edge of the access opening and the access cover using spacers and carefully shaping of the innerface therebetween. Consequently a decrease in weight over a non-vented structure is created.

Abstract: A finlet injector for external burning fuel injection is provided to obtain a higher performance through a combination of more optimum air-to-fuel ratio and better air/fuel mixing. This is done by capturing (mixing with) a greater quantity of air for a given fuel injection pressure and flow rate. The device is thermally protected and for most flight regimes has minimum drag.

Abstract: The ablative response at predetermined locations in a carbon-carbon reinforced composite material are selectively modified by one or more implants of a compatible, non-carbon material disposed at selected sites in the material. The implants of non-carbon material, which should have a melting point below the sublimation temperature of the surrounding carbon matrix and thermal expansibility having a force below the strength of the surrounding carbon matrix, preferably are arranged essentially parallel to one another at selected axial sites in the direction of expected thermal flux.

Abstract: A disintegratable aerodynamic brake is disclosed for use on a re-entry vehicle for retarding vehicle speed upon re-entry. The brake is connected to the travelling vehicle and is fabricated from a refractory material and is configured to progressively fuse and disintegrate as it becomes heated by friction as the vehicle encounters increasing atmospheric resistance upon reentry thereinto, so as to retard vehicle speed.

Abstract: Slip cast, fused silica shields such as radomes and antenna windows with ir pores filled with graded-Z radiation absorbing particles, give protection against radiation damage, improve the ablation characteristics and do not impair the microwave properties of the shielding.

Abstract: Sheets of metal foil superalloy heat-resisting metal foil, aluminized mylar or other thermal insulators are superposed one upon the other and deployed over the surface to be protected. A "dead space" thermal insulation such as a vacuum or simply a high resistance physical separation, is provided between surfaces. Physical separation between surfaces may be effected by flaps, integral to the insulator material, which are actuated by an aerodynamic or forced airstream or by the static energy stored in uncoiled flaps. Deployment is effected, only when needed, either by automatic or manually controlled reels which dispense the insulating material.

Abstract: The ablative response at predetermined locations in a carbon-carbon reinforced composite material are selectively modified by one or more implants of a compatible, non-carbon material disposed at selected sites in the material. The implants of non-carbon material, which should have a melting point below the sublimation temperature of the surrounding carbon matrix and thermal expansibility having a force below the strength of the surrounding carbon matrix, preferably are arranged essentially parallel to one another at selected axial sites in the direction of expected thermal flux.

Abstract: A cooling arrangement for a gas turbine exhaust nozzle is provided which ludes a plurality of heat shield panels slidably positioned on tracks mounted on the nozzle casing. The panels are spaced apart to permit thermal expansion and include impingement baffles which are attached thereto intermediate the panels and the casing. Holes in the impingement baffles direct cooling air against the interior surface of the panels. The cooling air is directed by suitable means such as slots or film holes to form a film of cooling air along the surface of the panels exposed to hot gases.

Abstract: A streamlined electronic-component-containing pod is mounted on the exterior of a military aircraft in subjection to the airflow past the aircraft during flight. An evaporator in the pod is close coupled thermally to the electronic components for cooling. The outer skin of the pod acts as a condenser for return of liquid refrigerant to the evaporator during normal flight operations. During short-term high-speed dash operation the outer skin of the pod becomes too hot for condensing refrigerant vapor, which then becomes condensed by a heat sink. At this time the refrigerant at the outer skin remaining in the vapor phase acts as thermal insulation.

Abstract: A cooled nosetip arrangement for a vehicle designed to travel at hypersonic velocities is taught herein. We provide a nosetip having one or more orifices at its forwardmost location, but instead of utilizing an arrangement for expelling a liquid coolant, which would necessitate a considerable amount of plumbing and tankage, we utilize a novel gas generator arrangement enabling a coolant gas to be selectively generated and then expelled from such nosetip orifice(s). In this manner we utilize a coolant that advantageously remains in a chemical form until it is to be actually utilized.

Abstract: A laminate thermal control coating for spacecraft comprising a layer of solar-radiation stable film, a layer of particulate-radiation stable film applied to the upper surface of the solar-radiation stable film, and a layer of reflecting material applied to the lower surface of the solar-radiation stable film. The coating experiences no increase in solar radiation absorptance (the proportion of radiant energy absorbed) upon exposure to particular or solar radiation as the particulate radiation is substantially absorbed in the particulate-radiation stable layer and the solar radiation partially absorbed by the particulate-radiation stable layer is transmitted by the solar-radiation stable film to the reflecting material which reflects it back through the laminate and into space.

Abstract: A missile nose cone removable protective cover having a rigid outer shell of cross-linked polyethylene with a semirigid backing of polyurethane foam. A shaped linear charge is positioned in a serpentine pattern between the rigid outer shell and the foam backing. The shaped linear charge is detonated by means of a detonator in an epoxy junction box in the forward end of the nose cover.

Abstract: An auger device is used to attach rigidized surface insulation to a spacecraft. The auger is preferably screwed into an insulation tile which has been predrilled. The augertile combination is then fastened to the spacecraft using an attachment screw which penetrates the spacecraft skin and which is secured by a blind end fastener. In an alternate method, the auger is incorporated in the insulation tile when the latter is fabricated.