Abstract: A cooling system (18) for supplying an even distribution of cooling air within a payload chamber (14) of a launch vehicle is disclosed. The cooling system (18) includes an elongate duct (20), a diverter (22), and a plurality of partitions (24a, 24b, 24c . . . ). The diverter (22) is attached to a distal end (21) of the elongate duct (20) by the plurality of partitions (24a, 24b, 24c . . . ). A turning vane (26) integral with two of the plurality of partitions (24a, 24b, 24c . . . ) extends into the elongate duct (20) dividing part of the elongate duct (20) into two separate halves. Cooling air passes through the elongate duct (20) and strikes the diverter (22) which evenly distributes the cooling air within the payload chamber (14) of the launch vehicle.

Abstract: A system and method for attaching an annular structural component to an aerospace vehicle is provided. The system includes a plurality of fittings circumferentially affixed to an engagement portion of the vehicle, and a plurality of supports. Each support has a first end connected to a corresponding fitting, and a second end having an elongate slot. A flange is configured to extend inwardly from the structural component and into engagement with the second end of each support. The flange has an elongate slot located at each support. Each flange slot is non-parallel to an open communication with each respective support slot. A plurality of pins connects the supports to the flange. Each pin extends through a corresponding flange slot and respective support slot to constrain the structural component, while attaching the structural component to the vehicle during use of the vehicle.

Abstract: A supersonic flight aircraft having a longitudinally forwardly extending fuselage having an axis in the direction of flight, and a wing, and which comprises the wing extending generally laterally relative to the axis, and having a leading edge angled forward or rearwardly relative to a normal to the axis at an angle , and the wing having leading edge sharpness defined by upper and lower wing surfaces, which taper toward the leading edge to define an angle .delta., closely proximate the leading edge at all spanwise locations; the angle and sharpness .delta.

Abstract: An element made of a reinforced low density heat protective material including an elastomer and/or a silicon resin matrix loaded with organic and/or inorganic components, wherein the reinforcement is formed of glass or ceramic thread sections or the like or organic materials fitted in the mass of the matrix along directions approximately orthogonal to at least one of the faces of the element and being flush, at least at one of their extremities, with at least one of the faces. The method for obtaining this element, includes, after molding to the general dimensions and shapes of the element with the aid of a loaded silicon matrix, placing the thread sections by stitching.

Abstract: A cooling device for an aircraft turboshaft engine includes a boundary layer suction device formed by a perforated external wall portion of the aircraft, a collector on the inside of the wall portion for receiving the air flowing in through the perforations, and a discharge device discharging the air from the collector. The cooling device is completed by a heat exchanger which is disposed in the path of the air flow through the collector, a pump for pumping fluid to be cooled through the heat exchanger from a circuit of the engine, and a regulator regulating the fluid flow through the exchanger to control the temperature of the fluid. The fluid which is cooled may be from the lubrication circuit, the fuel circuit, or an air circuit of the engine, and the boundary layer suction device is preferably disposed in the pod of the engine.

Abstract: The invention relates to a system for removing the heat that is incident upon an aircraft, and especially a supersonic aircraft, during flight. The incident heat is generated by the constant heating effect of aerodynamic friction on the outer skin of the fuselage, by compression heating at the ram or stagnation point and stagnation edges of the aircraft fuselage, and by the heating of external air as it is compressed to cabin pressure. In the present system, the used cabin air is exhausted out of the cabin through a space between the cabin interior panels and the fuselage outer skin, where the airflow absorbs the heat incident through the outer skin and is heated to approximately the temperature of the outer skin. Next, the heated airflow is preferably first filtered through a filter and is then expanded and cooled through an expansion turbine. The output shaft of the turbine can drive a generator to recover useful electrical energy.

Abstract: A system for eliminating distortion of optical ports during supersonic flight includes an electrohydraulic pump which is controlled by a potentiometer actuated by expansion and contraction of a gas that acts on a piston which is mounted in a cylinder. The cylinder is mounted in thermal contact with a frame which supports one or more layers of optical glass. The electrohydraulic pump operates to pump cooling fluid into the frame thereby cooling the frame and the optical glass when the gas expands as a result of heating during supersonic flight.

Abstract: A vehicle having a ceramic radome includes a vehicle body having an opening therein, a ceramic radome sized to cover the opening of the vehicle body, and an attachment structure joining the radome to the vehicle body to cover the opening. The attachment structure has a substantially cylindrically symmetric compliant metallic transition element disposed structurally between the radome and the body. The radome is joined to a first end of the transition element by a first brazed butt-joint. A primary seal is disposed between the transition element and the vehicle body. A spanner nut overlies the transition element and has a first engagement to the transition element. A threaded engagement between the spanner nut and the vehicle body allows the radome/transition element assembly to be attached and tightened to the vehicle body and to be later removed.

Abstract: A liquid coolant heat exchange system for use in semimonocoque aircraft includes an arcuate planar heat sink fixed along a radius of curvature R.sub.1 by forming members, a flexible arcuate planar spreader plate having a radius R.sub.2, such that R.sub.1 >R.sub.2, a heat exchange tube for transferring heat from the liquid coolant to the heat sink and means for attaching the spreader plate to the forming members to hold the spreader plate in contact with the heat sink.

Abstract: A liquid coolant heat exchange system for use in semimonocoque aircraft includes an arcuate planar heat sink fixed along a radius of curvature R.sub.1 by forming members, a flexible arcuate planar spreader plate having a radius R.sub.2, such that R.sub.1 >R.sub.2, a heat exchange tube for transferring heat from the liquid coolant to the heat sink and means for attaching the spreader plate to the forming members to hold the spreader plate in contact with the heat sink.

Abstract: The non-explosive core of a reentry projectile is fixedly positioned within hollow casing of the projectile at a location maximizing conversion and transfer of kinetic energy to an earth bound target in response to the projectile's impact at a hypersonic velocity and at a steep impact angle to the surface of the earth. The hollow casing is formed of a material capable of withstanding high temperatures, therefore, not requiring any cooling and allowing for the hollow casing to be free of heat transfer medium that might otherwise impede the desired quick release of the core upon impact of the projectile. The core is in the form of a single dense metallic slug having a mass establishing a center of gravity and moments of inertia for the projectile as a kinetic energy warhead corresponding to that of an explosive or nuclear warhead without weapon system modification.

Abstract: A spacecraft multilayer insulation blanket is made from multiple sheets of polyimide e.g. KAPTON. A preferred six sheet embodiment employs three carbon loaded sheets (26, 32, 34) separated by two dielectric sheets (28, 30). A third dielectric sheet (36) insulates the carbon loaded sheets from the spacecraft body. The outermost carbon loaded sheet (32) has a high resistivity and is primarily transmissive while the other two carbon loaded sheets (26, 34) have a low resistivity and are primarily reflective.

Abstract: The nacelle of the aircraft is purged and the sidewall of the vector nozzles of the gas turbine engine in the nacelle is cooled by utilizing the engine fan air to drive ejector pumps mounted in series flow relationship, a rectangular mixing zone and a diffuser with a rectangular cross section and a step in the exhaust nozzle gaspath. This results in minimizing the amount of fan air for driving the system and enhancing engine thrust and enhanced sidewall cooling because the temperature of the mixed air is lower than heretofore known systems.

Abstract: A method and apparatus are disclosed for providing an aircraft thermal protection system for hypersonic cruise and space launch vehicles. A flexible outer skin formed from a metal super alloy is designed to flex and accommodate thermal growth in the vehicle structure. The flexible super alloy skin is made from a plurality of hexagonal shaped cups which are welded together at the edges in a honeycomb type of array with thermal expansion gaps provided between the outermost surfaces of the hexagonal cups. Gap covers extend across the thermal expansion gaps to reduce aerodynamic drag. The flexible outer skin extends over hexagonal shaped, high temperature ceramic blocks, which provide both an insulation layer and support for the outer skin. The flexible outer skin distributes airloads across various ones of the rigid ceramic blocks.

Abstract: A system for removing and/or preventing ice build-up on an aircraft includes at least one heat exchanger mounted within an aperture in the tailpipe of a gas turbine engine. The heat exchanger includes an inner plate having a surface which faces inwardly into the tailpipe and which is in direct contact with the heated exhaust gas in the tailpipe, and an outer plate with a surface facing outwardly from the tailpipe and in direct contact with a cooler air stream passing around the turbine engine core. A series of heat exchange fins are formed on the inner plate to direct exhaust gas across the inner plate, and a series of heat exchange fins are formed on the outer plate to direct the cooler air across the outer plate. Heat transfer fluid flows between the inner plate and the outer plate and absorbs heat energy from the exhaust gas. The heat transfer fluid flows through a filter, accumulator and pump system to heat transfer structure integral with the exposed portions of the aircraft.

Abstract: A transpiration cooling system for avoiding overheating of an airfoil is provided. The airfoil is provided with a plurality of apertures and a source of pressurized fluid for providing a flow of fluid through the apertures to establish an aerodynamic radius. The aerodynamic radius of curvature of leading edge is sufficiently greater than the mechanical radius of curvature of the leading edge that peak heat flux is independent of the mechanical radius of curvature. The mechanical radius of curvature is preferably less than 50% of the aerodynamic radius of curvature during hypersonic operation. Preferably the mechanical radius of curvature is the smallest allowed by the fabrication method (i.e., a knife edge), such as being less than about 0.02 inches, preferably less than about 0.01 inches. The transpiration blowing rate can be adjusted so that the blowing rate and aerodynamic radius of curvature are relatively low except during periods of maximum heat flux, such as the shock-on-lip point.

Abstract: A liquid coolant heat exchange system for use in semimonocoque aircraft includes an arcuate planar heat sink fixed along a radius of curvature R, by forming members, a flexible arcuate planar spreader plate having a radius R.sub.2, such that R.sub.1 >R.sub.2, a heat exchange tube for transferring heat from the liquid coolant to the heat sink and means for attaching the spreader plate to the forming members to hold the spreader plate in contact with the heat sink.

Abstract: In a seeker head assembly in a guided missile, particularly a high-speed guided missile, the seeker head (22) is covered by a dome (28) which is rotatably mounted in the structure and is rotated. The dome (28) is circumferentially surrounded by the structure (10) of the guided missile except for a window section (14). A coolant is supplied to an interspace (46) formed between dome (28) and structure (10) of the guided missile.

Abstract: A transpiration cooling system for avoiding overheating of an airfoil is provided. The airfoil is provided with a plurality of apertures and a source of pressurized fluid for providing a flow of fluid through the apertures to establish an aerodynamic radius. The aerodynamic radius of curvature of leading edge is sufficiently greater than the mechanical radius of curvature of the leading edge that peak heat flux is independent of the mechanical radius of curvature. The mechanical radius of curvature is preferably less than 50% of the aerodynamic radius of curvature during hypersonic operation. Preferably the mechanical radius of curvature is the smallest allowed by the fabrication method (i.e., a knife edge), such as being less than about 0.02 inches, preferably less than about 0.01 inches. The transpiration blowing rate can be adjusted so that the blowing rate and aerodynamic radius of curvature are relatively low except during periods of maximum heat flux, such as the shock-on-lip point.

Abstract: A projectile, particularly a kinetic energy projectile, whose nose cone is cooled due to the provision of an ablation layer and a thermal insulation layer (8, 9) on the exterior of the metal hood (4) of the projectile nose cone, and due to a highly thermally conductive medium in the spaces (5, 6) between the hood (4) and the payload (3) particularly a penetrator, of the projectile.

Abstract: A device for thermal protection of a structure (22) is provided. In one embodiment the device comprises a retention layer (12) impregnated by a fluid substance that is normally in the liquid state. The fluid substance is resupplied to the retention layer via a porous membrane that is impermeable to the substance when it is in the liquid state but is permeable thereto when it is in the vapor state. The membrane is interposed between the retention layer and a supply space for the substance in its liquid state adjacent the structure to be protected.

Abstract: A nonsacrifical laser hardened structure and method protects against incident high energy irradiation. A very thin outer face sheet with a reflective outer surface is laminated to a foam core which is further laminated to a support structure. The foam core includes liquid and vapor passages. A supply of liquid working fluid impregnates the foam core. A portion of incident irradiation energy is first reflected from the outer face sheet and substantially all the remaining irradiation energy conducted through the face sheet to be absorbed at the back of the face sheet into heat of vaporization of the liquid working fluid. Additional liquid working fluid is supplied to the back of the face sheet by capillary action through the foam core, and additional liquid working fluid is resupplied to the foam core through the liquid passages. Working fluid vapor is removed through the foam vapor passages.

Abstract: Circulation of temperature controlled fluid through cavities in structure stiffening members of hypersonic aerospace vehicles precludes destructive temperature gradients in such members and preserves stress residuals for functional requirements.

Abstract: A leading edge construction for reducing drag on an airfoil, the construction includes a relatively thin solid plate extending forwardly from the air foil, the plate having an exposed top surface and an exposed bottom surface, and a relatively sharply radiused forward edge forming the leading edge of the air foil. The construction further includes slots adjacent to the top and bottom surfaces of the plates for injecting coolant over the top and bottom surfaces of the plate, toward the leading edge, to provide active cooling of the plate for removing stagnant heat lead.

Abstract: A dual mode semi-passive nosetip cooling system for a hypersonic weapon. A cone shaped capsule or pressure vessel is positioned within the nosetip and it is filled with a sublimating material in solid state. The capsule has a nozzle formed adjacent its forward end that directs a steady impingement stream of sublimation material to the back side of the nosetip shell in its stagnation region when the sublimating material is being heated by the surface heat on the nosetip itself. The concept is based on utilizing the advantages of both a subliminating and impingement cooling system designs.

Abstract: Enclosure of reception of electronic components is embarked in a missile. The missile is first fixed to an aircraft carrier and to follow a carried flight, then after being fired, to follow a free flight. The enclosure is located inside of an exterior casing. Sectors sensitive to the acceleration are provided, on this side of a given acceleration, in order to form a thermal bridge between the enclosure and the casing and, beyond the given acceleration in free flight, to break that thermal bridge and to insulate the enclosure from the exterior. In free flight, the electronic components of the enclosure are no longer subject to a source of thermal energy.

Abstract: A hybrid active cooling system for hypersonic vehicle nosecaps and leading edges integrates a forced convection cooling system with a transpiration cooling system. A pair of concentric shells, hemispherical for a nosecap and cylindrical for a leading edge, are separated by pin fins to form a first manifold for passage of a convection coolant. A second manifold for a transpirant coolant is located behind the first manifold. The pin fins comprise a nonporous structural outer shell surrounding a porous core. The porous cores are open to the inside of the second manifold and to the outside of the first manifold. The outside of the first manifold forms part of the outer surface of the nosecap or leading edge. Convection coolant is pumped through the first manifold to provide forced convection cooling. Transpirant coolant pumped into the second manifold travels through the porous cores of the pin fins to be injected into the atmosphere to provide transpiration cooling.

Abstract: The present invention provides for the integration of structural cooling and fuel treatment within hypersonic vehicles. This is achieved by channeling a hydrocarbon fuel to a portion of the aircraft structure and imparting via the heat sink effect sufficient heat to cause a pyrolysis of the hydrocarbon fuel. After pyrolysis, the resulting fractions are then utilized as a fuel by the vehicle.

Abstract: An insulative panel is attachable to a structure to thermally protect the structure. The insulative panel includes a piece of a compliant, porous material. A support base is bonded to a first side of the piece of porous material, the support base including means for attaching the support base to the structure that is to be protected. A skin made of a thermally resistant material is bonded to a second side of the piece of porous material. Preferably, the porous material is a piece of porous ceramic foam having an apparent density of less than about 25 percent, the support base is a piece of nonporous ceramic material, and the skin is a ceramic thermal barrier coating that is plasma sprayed onto the porous material.

Abstract: A thermal protection device for protecting walls of structures exposed to ablative flow at high temperatures in the presence of vibrations comprises, embedded in a thermally insulative matrix, a refractory armature formed by a fringed mat comprising a meshed part exposed to the ablative flow and fringes intended to be directed towards the wall of the structure to be protected. The device further comprises an open-weave refractory woven material parallel to the meshed part of the armature and through which the fringes pass.

Abstract: An inlet ramp for an aircraft engine is constructed as a movable panel which is internally cooled by fluid coolant. At a selected time, the coolant flow is terminated allowing the panel to overheat and melt away so as to expose an underlayer of ablative material. The ablative material then begins to melt away thereby cooling the underlying surface of the aircraft via ablation.

Abstract: A thermal protection coating for a surface having a reinforcement embedded in a thermally insulative material as a main layer. This layer is formed from a succession of refractory fibrous reinforcements parallel to each other and inclined at an angle to the surface being protected. Insulative layers are located between the reinforcements. An insulating sublayer is located between the surface and the main layer and is compatible with the main layer and the surface.

Abstract: An integral structure and thermal protection system, particularly designed for hypersonic aerospace vehicles, is comprised of a hard external shell or outer face sheet formed of a ceramic matrix, such as silicon carbide, and a rigid insulator core in the form of foamed ceramic, such as silicon carbide, the outer face sheet being integrally connected to the insulator core. A prime strucutral material, such as an aircraft structural member, is integrally connected as by bonding or brazing to the insulator core, the core being attached to the prime structural material or aircraft structural member by an inner face sheet forming the outer skin of such structural member.

Abstract: A heat source, which may be on a high speed vehicle, may be cooled by transferring thermal energy from the heat source to an endothermic fuel decomposition catalyst to heat the catalyst to a temperature sufficient to crack at least a portion of an endothermic fuel stream. The endothermic fuel is selected from the group consisting of isoparaffinic hydrocarbons, blends of normal and isoparaffinic hydrocarbons, and conventional aircraft turbine fuels. The heated endothermic fuel decomposition catalyst is contacted with the endothermic fuel stream at a liquid hourly space velocity of at least about 10 hr.sup.-1 to cause the endothermic fuel stream to crack into a reaction product stream.

Abstract: A cooling system for a hypersonic aircraft is disclosed which enables hypersonic flight using non-cryogenic fuels. The cooling system positions a primary heat exchanger at an external location on the aircraft which remains relatively cool during hypersonic flight. A working fluid is passed through the primary heat exchanger to the hot parts of a supersonic combustion ram jet engine.

Abstract: The invention is a propellant utilization system for flight vehicles operating in both the subsonic and supersonic speed ranges. In particular, the invention uses steam and water produced in the system to replace hydrogen as a coolant and for power generation. It also provides for the utilization or oxygen obtained from air captured by the air inlet instead of oxygen stored in the vehicle.

Abstract: A fin stabilized projectile in which measures are taken for increasing the thermal resistance of its guide fins against the uncontrollable danger of burning or melting due to overheating of the material as a result of air friction during the flight of the projectile. Each guide fin is composed, at least in the region of its leading edge of a heat resistant fiber material or a heat resistant composite fiber material. The fins may be composed entirely of the heat resistant material or the leading edges of the fins, the leading and outer edges of the fins, or the entire front fin portion may be configured as a strip of the fiber material in front of and fastened to a metal body member comprising the remainder of the respective guide fin. Moreover, the connecting region between any such strip and the metal body member is preferably configured with a sloped, and thus as a contact area which has a larger surface area. The strip is glued, riveted and/or screwed to the body member in the connection region.

Abstract: Disclosed is an apparatus and a method for cooling a rotatable convergent side disk in a two-dimensional nozzle for a gas turbine engine, in which coolant is supplied and exhausted through concentric conduits, the cooling requirements of a liner assembly are reduced by use of a thermal barrier coating, and exhaust gas from the nozzle is prevented from aspirating into the side disk by use of a liner which is impervious to the exhaust gas.

Abstract: The present invention relates to a strut assembly which may be selectively projected and retracted with respect to associated structure of a solid or liquid propellant rocket of missile, with such structure comprising one of, for example, a nozzle thereof or the body thereof. The invention struct may be projected either internally of the nozzle or externally of the missile body, as desired. Projection and retraction of the strut are accomplished through use of fluid actuation means, the fluid mode of means of which comprises coolant which is also directed in various locations on the inventive strut assembly to cool the strut assembly and to protect it and adjacent structures from particulate impaction, heat and thermal damage, etc.

Abstract: A crew protection system designed to isolate an aircraft crew and provide a means for safely evacuating an aircraft, particularly hypersonic vehicles, in emergency situations. The crew compartment is designed within the forebody of the main vehicle. The forebody is designed so as to separate cleanly from the main vehicle as by generation of a circumferential structural break, in the event of a catastrophic failure or explosion of the main vehicle fuel tank. In the event of such an uncommanded failure of the main tank and propellant combustion, resulting in the above noted structural failure, the resultant pressure impulse thrusts the forebody containing the crew compartment clear from the remainder of the vehicle. The provision of a system of baffles in the forebody between the crew compartment and the fuel tank, and the shape of the forebody and the materials used in its construction mitigate the effects of the blast pressure impulse on the crew.

Abstract: A braided fiber reinforced composite, particularly a braided graphite fiber reinforced copper composite leading edge heat exchanger 18 is disclosed, formed of inlet and outlet conduits 22 and 24 and means forming a plurality of passages 26 interconnecting the conduits, as a single integral unit. The heat exchanger is formed by providing inlet and outlet manifolds 36 and 38, successively interconnecting a plurality of mandrels 40, 52 and 56 transversely between the manifolds and braiding reinforcing fibers 32 preferably having high elevated temperature strength and high thermal conductivity, e.g. graphite fibers, around each of the mandrels and the manifolds, the braiding progressing in a direction along the length of the manifolds. The resulting braided preform 58 is consolidated by introducing a matrix material of high thermal conductivity, such as copper or copper alloy, e.g.

Abstract: A buffering portion of boundary layer flow along a flight vehicle surface diverted into an internal cavity at a downstream edge of a surface mounted window isolating recirculating flow of the diverted fluid through the cavity while it is cooled therewithin. Outflow of the cooled buffering fluid from at an upstream edge of the window is conducted into the buffering portion of the boundary layer flow to cool the window and prevent damage thereof by aerodynamically generated heating of the surface under high velocity flight conditions.

Abstract: The invention provides a system for selective reduction of infrared signature of a vehicle subjected to aerodynamic heating. Liquid coolant under pressure vaporizes in porous sections of the skin of the vehicle to transpiration-cool the skin. Adjacent downstream solid skin sections are film-cooled by the vapor introduced in the boundary layer. Coolant flow control is achieved by pressurizing the liquid coolant, by modulating flow control valves, and, optionally, by configuring the porous section to operate as a cavitating venturi, or combinations of these methods.

Abstract: A primary load bearing structure particularly adapted for hypersonic aerospace vehicles utilizing a cryogenic liquid fuel, and comprised of a plurality of rigid load bearing, fluid passageways or tubes connected together in the form of a panel structure. Such panel structure and the passageways or the tubes thereof form the outside surface of the aerospace vehicle and are also positioned adjacent to or integral with a wall of the cryogenic liquid fuel tank. During flight, cryogenic liquid fuel is passed through the tubes to provide cooling for the external surface of the vehicle. When the aerospace vehicle is on the ground, the load bearing tubes can be evacuated to provide insulation for the cryogenic liquid fuel tank to thus prevent formation of frost and liquid air on the outside surface of the tank, as well as on the external surface of the aerospace vehicle and prevent heat transfer to the tank.

Abstract: Heat pipe (10) includes an acute angle leading wedge shape form (38) with a radiused leading edge (26). Engine fuel supply (65) cools the trailing wedge shape form (40). The intense heat flux at the leading edge is cleanly dissipated and spread over substantial heat pipe cooling surface (64, 66). A coating (36) on the external surface of the leading edge wedge shape form is formed of a layer of iridium and a layer of oxide.

Abstract: A method and apparatus for eliminating heat build-up caused by air friction on wings and other air foils, especially common in high-speed aircraft. As an aircraft is moving, a small portion of air in the airstream moving by the wing or other air foil is passed through an air intake valve on the wing, into a Venturi section and then into a chamber within the wing or frame where the air is compressed naturally. Each Venturi section consists of a converging and diverging nozzle which creates a substantial cooling effect when the high speed air passes through it and is compressed and released. The intake valve spins or flips when hit by the airstream so it alternately intakes and exhausts the air through the Venturi section, thereby causing a double cooling effect. A second embodiment uses valves, one on the leading edge for air intake and the second near the back of the lower surface of the wing for air exhaust.

Abstract: A radome type nose cone shell subjected to aerodynamic heating during protile travel, is ablatively cooled by evaporation of metal coolant from a solidified coating thereof adhering to the external surface of a porous section of the cone shell through which the coolant seeps from internal storage at a rate regulated by formation of the solidified coating. The evaporating coolant from the coating mixes with the air in the flow stream to form a protective a boundary layer extending downstream from the porous section of the nose cone shell. The coolant is melted by said aerodynamic heating during internal storage within the nose cone shell for seepage outflow from the porous section and has the requisite heats of vaporation and fusion to effect alternate freezing and thawing thereof in response to withdrawal of heat by the cooling process for establishment and maintenance of the flow regulating coating.

Abstract: A high temperature heat shield system for re-entry vehicles and high speed ircraft is assembled from separate insulating panels made of metal foil layers diffusion soldered together and with a ceramic fiber material filled into interspaces between the metal foil layers. In this arrangement structural metal foil layers carry mechanical loads while separate reflecting foil layers and the ceramic fiber material are non-load bearing.

Abstract: An improved means for convection cooling the skin surfaces of an atmospheric flight craft such as are subject to air friction-induced overheating when the craft operates at high speed, featuring employment of multiple layered, spaced apart skin members separated and relatively braced by web and post members. Fresh coolant fluid is simultaneously flushed against and contiguously covers a multiplicity of relatively small patches of the inside surface of the outer skin member to be protected, and the spent coolant from each patch is separately removed therefrom and returned to the craft's refrigeration facility. The web and post members also act to direct the pattern of coolant flow so as to provide a much improved heat transfer coefficient than can be attained by other coolant path configurations when an equal pressure drop is applied.

Abstract: A fiber composite containing a system of reinforcing fibers that is impregnated with a resin matrix. The resin matrix includes a phenolic-based resin and a phosphorus-containing polymer. The composite exhibits low heat release rate and total heat release when subjected to combustion conditions including a radiant heat source.