Abstract: A composite structure has a first layer connect to a second layer to form a layered structure. The second layer has a plurality of base portions abutting the first layer and a plurality of beaded portions protruding from the plurality of base portions. Each beaded portion of the plurality of beaded portions defines a channel between the first layer and the second layer. A method for manufacturing a composite structure includes depositing composite material over a tool to form a second layer. The second layer has a plurality of beaded portions and a plurality of base portions. The method includes distributing a plurality of mandrels over the second layer to define channels in the plurality of beaded portions. The method includes depositing composite material over the second layer and the plurality of mandrels to form a first layer. The method includes bonding the first layer to the second layer.

Abstract: A part of a spacecraft, for instance an upper stage, is configured to re-enter into the atmosphere and to be reused for several missions. The part is equipped with a reusable kit comprising non-ablative heat shields, for example constructed with an outer surface formed of Ceramic Matrix Composites, such as Ultra High Temperature Ceramic Matrix Composites, and optionally at least one decelerator and/or at least one parachute and/or parafoil system.

Abstract: One or more temperature control units are distributed throughout an environment to provide localized heating and cooling. The temperature control units include a thermal storage system including one or more substances of high latent heat capacity, a heat distribution surface, a solid-state heat pump positioned between the thermal storage system and the heat distribution surface, an environmental sensing module including a proximity sensor, and unit control modules in communication with the solid-state heat pumps. The solid-state heat pumps are individually controllable so that heating and cooling can be provided simultaneously from separate heat pumps in the same temperature control unit, or separate temperature control units in the same environment.

Abstract: An aircraft propulsion unit includes an electric motor, at least one accessory unit used for operating the electric motor, an inverter module, the inverter module including a plurality of inverters for powering the electric motor and the at least one accessory unit, and a cooling system coupled to the electric motor and the inverter module, the cooling system comprising a coolant path for circulating a coolant through or adjacent to the electric motor and the at least one accessory unit.

Abstract: A thermal control system may transfer energy (directly or after a delay) to a thrusting device that can be used to slow a reentry vehicle entering a gaseous atmosphere from orbit. The thermal control system may mitigate the heating of the vehicle by transferring heat generated by the viscous interaction between the vehicle and high-altitude portions of a planetary atmosphere to a working fluid. This working fluid may then be routed through and/or ejected through one or more nozzles aligned to produce thrust in a direction that opposes the forward motion of this vehicle. This counter thrust may help to slow the reentry vehicle and reduce the amount of kinetic energy that can be converted into thermal energy. The working fluid may also be stored to use for propulsion after the reentry vehicle slows below hypersonic velocities.

Abstract: An aerospace vehicle including: a body, wherein the body is configured to generate heat during operation; a coating disposed over at least a portion of the body, the coating being configured to shift a frequency of at least one wavelength of the heat generated by the body from a first frequency to a second frequency having higher transmissivity relative to a neighboring medium surrounding the body as compared to the first frequency.

Abstract: An aircraft propulsion unit includes an electric motor, at least one accessory unit used for operating the electric motor, an inverter module, the inverter module including a plurality of inverters for powering the electric motor and the at least one accessory unit, and a cooling system coupled to the electric motor and the inverter module, the cooling system comprising a coolant path for circulating a coolant through or adjacent to the electric motor and the at least one accessory unit.

Abstract: A hypersonic aircraft includes one or more leading edge assemblies that are designed to manage thermal loads experienced at the leading edges during high speed or hypersonic operation. Specifically, the leading edge assemblies may include an outer wall tapered to a leading edge or stagnation point. The outer wall may define a vapor chamber and a capillary structure within the vapor chamber for circulating a working fluid in either liquid or vapor form to cool the leading edge. In addition, a dual-modal cooling structure can enhance heat transfer from the outer wall at the leading edge to the outer wall within the condenser section of the vapor chamber.

Abstract: An electromagnetic-wave-transmissive module of a vehicle radar is provided to minimize a dielectric impact reflection effect, which occurs when an electromagnetic wave radiated from an antenna is transmitted through a radome and a transmissive cover The electromagnetic-wave-transmissive module includes one or more of a radome covering the antenna and a transmissive cover, which is disposed to be spaced apart from a front side of the antenna and through which a radio wave radiated from the antenna and then transmitted through the radome is subsequently transmitted. At least one coating layer, which includes PTFE and which has a dielectric permittivity higher than the dielectric permittivity of air and lower than the dielectric permittivity of the radome and the transmissive cover, is formed on the surface of at least one of the radome and the transmissive cover.

Abstract: A hypersonic aircraft includes one or more leading edge assemblies that are designed to manage thermal loads experienced at the leading edges during high speed or hypersonic operation. Specifically, the leading edge assemblies may include an outer wall tapered to a leading edge or stagnation point. The outer wall may define a vapor chamber and a capillary structure within the vapor chamber for circulating a working fluid in either liquid or vapor form to cool the leading edge. In addition, a thermal energy storage reservoir positioned within the vapor chamber contains a phase change material for absorbing thermal energy.

Abstract: A hypersonic aircraft includes one or more leading edge assemblies that are designed to cool the leading edge of certain portions of the hypersonic aircraft that are exposed to high thermal loads, such as extremely high temperatures and/or thermal gradients. Specifically, the leading edge assemblies may include an outer wall tapered to a leading edge or stagnation point. A coolant supply may be in fluid communication with at least one fluid passageway that passes through the outer wall to deliver a flow of cooling fluid to the stagnation point. In addition, a nose cover is positioned at least partially over or within the at least one fluid passageway and is formed from a material that ablates or melts when the leading edge is exposed to a predetermined critical temperature, the nose cover being configured for restricting the flow of coolant until the nose cover is ablated or melted away.

Abstract: A method for building an aerodynamic structure, an aerodynamic structure, and a vehicle that includes the aerodynamic structure are provided. The method includes providing a structure with at least one substantially-flat exterior surface. The method also includes attaching blocks of rigid fibrous insulation to the at least one substantially-flat outer surface of the structure. Outward-facing surfaces of the blocks of rigid fibrous insulation extend past a target outer mold line of a final aerodynamic shape. The method also includes machining the outward-facing surfaces of the attached blocks to the outer mold line.

Abstract: An aircraft propulsion system with a drag reduction portion adapted to reduce skin friction on at least a portion of the external surface of an aircraft. The drag reduction portion may include an inlet to ingest airflow. The aircraft may also have an internally cooled electric motor adapted for use in an aerial vehicle. The motor may have its stator towards the center and have an external rotor. The rotor structure may be air cooled and may be a complex structure with an internal lattice adapted for airflow. The stator structure may be liquid cooled and may be a complex structure with an internal lattice adapted for liquid to flow through. A fluid pump may pump a liquid coolant through non-rotating portions of the motor stator and then through heat exchangers cooled in part by air which has flowed through the rotating portions of the motor rotor. The drag reduction portion and the cooled electric motor portion may share the same inlet.

Abstract: A system for fabricating an ablative, 3D fiber-woven thermal protection material, with porosity 0.5-15 percent, reduced thermal conductivity, very low thermal recession, high glass transition temperature, high frontface-backface temperature difference, relatively high mass density, and significant compression strength and tensile strength.

Abstract: A method includes absorbing heat generated at a flight vehicle using a heat sink, where the heat sink includes a matrix. The matrix includes a porous structure having multiple pores at least partially filled with one or more phase change materials. The method also includes converting the matrix into an insulator as the one or more phase change materials change state and exit the porous structure due to the absorbed heat. The matrix with the one or more phase change materials could include an alcogel, and the insulator could include an aerogel. The matrix could reside within a pressurized container that includes at least one seal. The at least one seal can fail due to increased pressure within the pressurized container as the heat is absorbed by the heat sink in order to allow the one or more phase change materials to exit the porous structure.

Abstract: A system and method for restricting thermal transfer between an ambient space and a substructure include providing an inner mold line side (IML) surface to the substructure. A plurality of clevises extends upward from the IML surface. An outer mold line side (OML) surface is interposed between the ambient space and the IML surface. The OML surface includes at least one horizontally oriented wing and a plurality of horizontally oriented shoulders. The wing selectively engages at least two of the shoulders. A plurality of lugs extends downward from the OML surface. At least one aperture extends longitudinally through the OML surface to selectively permit fluid communication between the ambient space and the IML surface. A plug selectively substantially fills the aperture and is selectively secured to the IML surface. A fastener, configured to selectively secure the OML surface to the IML surface, extends horizontally through the clevises and the lugs.

Abstract: An optical arrangement in an optical system, such as a microlithographic projection exposure apparatus, includes: at least one heat-emitting subsystem which emits heat during the operation of the optical system; a first heat shield which is arranged such that it at least partly absorbs the heat emitted by the heat-emitting subsystem; a first cooling device which is in mechanical contact with the first heat shield and is designed to dissipate heat from the first heat shield; and a second heat shield which at least partly absorbs heat emitted by the first heat shield. The second heat shield is in mechanical contact with a cooling device that dissipates heat from the second heat shield.

Abstract: The disclosure herein relates to a measurement system to measure characteristics of the velocity vector of an aircraft in relation to a surrounding air mass, the measurement system comprising—a frontal surface of the aircraft, two primary sensors, each being fixed to the frontal surface of the aircraft and able to deliver an output value relating to the deformation experienced by the sensor, and a processing unit able to receive the output values and able to calculate the angle of attack and/or the velocity of the aircraft on the basis of these output values.

Abstract: A vehicle ablator system (“VAS”) is described that includes a reinforcing core and ablative material that has been packed and cured into the reinforcing core. In general, the VAS includes the reinforcing core that is attached to a vehicle structure such as, for example, a reentry vehicle or hypersonic vehicle prior to the ablative material being packed into the reinforcing core.

Abstract: In some examples, a method includes identifying a damaged area in a ceramic matrix composite coating of an in-service component; applying a restoration slurry to the damaged area of the ceramic matrix composite coating, wherein the restoration slurry comprises a liquid carrier and a restoration coating material; drying the restoration slurry to form a dried restoration slurry; and heat treating the dried restoration slurry to form a restored portion of the ceramic matrix composite coating. In some examples, an assembly may include a component including a substrate and a coating on the substrate, where the coating defines a damaged portion; masking around the damaged portion on undamaged portions of the coating; and a restoration slurry in the damaged portion, wherein the restoration slurry comprises a liquid carrier and a restoration coating material.

Abstract: A fastener assembly is disclosed for mounting a component to a bracket attached to an inner surface of the skin of an aircraft wing such that the component is thermally and electrically isolated from the bracket is disclosed. The component and bracket each have mounting holes and the fastener assembly includes a fastening element having a shaft to extend through the holes, first and second isolating members mountable on the shaft on either side of the bracket such that the bracket is sandwiched between the first and second isolating members.

Abstract: A launch system comprises a nose section comprising a nose coupling surface, a tail section comprising a tail coupling surface facing the nose coupling surface and a mast coupling the nose and tail sections. The mast is configured to expand and retract to displace the nose and tail sections within a range of distances from one another. In a retracted state, the nose and tail sections are either structurally coupled to one another at the nose and tail coupling surfaces or structurally coupled to at least one integrated module located between the nose and the tail sections.

Abstract: A deployable aerodynamic decelerator structure includes a ring member disposed along a central axis of the aerodynamic decelerator, a plurality of jointed rib members extending radially from the ring member and a flexible layer attached to the plurality of rib members. A deployment device is operable to reconfigure the flexible layer from a stowed configuration to a deployed configuration by movement of the rib members and a control device is operable to redirect a lift vector of the decelerator structure by changing an orientation of the flexible layer.

Abstract: The thermal protection system disclosed herein is suitable for use with a spacecraft such as a reentry module or vehicle, where the spacecraft has a convex surface to be protected. An embodiment of the thermal protection system includes a plurality of heat resistant panels, each having an outer surface configured for exposure to atmosphere, an inner surface opposite the outer surface and configured for attachment to the convex surface of the spacecraft, and a joint edge defined between the outer surface and the inner surface. The joint edges of adjacent ones of the heat resistant panels are configured to mate with each other to form staggered joints that run between the peak of the convex surface and the base section of the convex surface.

Abstract: Launch vehicles with fixed and deployable deceleration surfaces and associated systems and methods are disclosed. A system in accordance with a particular embodiment includes a launch vehicle that has a first end and a second end generally opposite the first end, and is elongated along a vehicle axis extending between the first and second ends. The vehicle carries an exposed outwardly facing surface having a first region positioned or positionable to have a first cross-sectional area generally normal to the vehicle axis toward the first end of the vehicle, and a second region positioned or positionable to have a second cross-sectional area generally normal to the vehicle axis toward the second end of the vehicle. The first cross-sectional area is less than the second cross-sectional area. The system can further include a propulsion system carried by the launch vehicle and having at least one nozzle positioned toward the first end of the vehicle to launch the launch vehicle.

Abstract: A portable thermal-control system adapted to support space-related research and exploration. Embodiments of the present invention assist in preventing overheating of small payloads being transported from an orbiting space vehicle to a planetary surface by small atmospheric-entry vehicles. Other embodiments of the present invention provide thermal control within an extra-vehicular activity (EVA) suit. Each embodiment utilizes at least one phase-change material, cooled significantly below the freezing temperature, to absorb heat.

Abstract: Two algorithms are disclosed that, with the use of a 3-axis accelerometer, will be able to determine the angles of attack, sideslip and roll of a capsule-type spacecraft prior to entry (at very high altitudes, where the atmospheric density is still very low) and during entry. The invention relates to emergency situations in which no reliable attitude and attitude rate are available. Provided that the spacecraft would not attempt a guided entry without reliable attitude information, the objective of the entry system in such case would be to attempt a safe ballistic entry. A ballistic entry requires three controlled phases to be executed in sequence: First, cancel initial rates in case the spacecraft is tumbling; second, maneuver the capsule to a heat-shield-forward attitude, preferably to the trim attitude, to counteract the heat rate and heat load build up; and third, impart a ballistic bank or roll rate to null the average lift vector in order to prevent prolonged lift down situations.

Abstract: An attachment system for mounting a tile of a thermal protection system to a substructure comprises a post assembly configured to interconnect the tile to the substructure. The post assembly is configured to allow relative sliding movement between the tile and the substructure along a plane which is generally parallel to the tile. The attachment system facilitates installation and removal of the tile from the substructure for inspection, maintenance and repair of the tile and/or the substructure.

Abstract: A satellite recovery apparatus having a base configured to be attached to a satellite and having a pressurized gas chamber and a valve assembly. A plurality of walls are hingedly attached to the base and movable between a closed position, wherein the walls define a volume above the base, and an open position, wherein the walls are disposed spaced away from the base. A heat shield is attached to the base. A deployable decelerator is attached to the heat shield and has an outer perimeter with an expandable torus that is operably connected to the valve assembly such that the expandable torus can be pressurized from the chamber. Pressurization of the expandable torus deploys the decelerator assembly from a non-deployed position within the volume to a deployed high-drag position.

Abstract: A combination structural support and thermal protection system is provided. A cured preceramic polymer matrix reinforced with carbon fibers can be used alone or in a multiple component system.

Abstract: A fairing includes three sections: an aft section, a middle section, and a nose section. The aft section includes a first surface having a generally constant conic angle (i.e., a “boat-tail angle”) with respect to the longitudinal axis such that the aft section tapers to a first end configured to attach to a generally cylindrical body. The middle section, which intersects and is axially aligned with the aft section, has a second surface characterized by a constant elliptical cross-section along a plane orthogonal to the longitudinal axis. The nose section intersects and is axially aligned with the middle section. The nose section is further defined by four generally concave trianguloid surfaces, each extending from the middle section to a common apex intersecting the longitudinal axis, wherein each adjacent pair of trianguloid surfaces intersects at an edge (e.g., an incurvate edge).

Abstract: A method of applying sealant to a narrow, deep gap in a surface includes covering the gap to form a covered channel, inserting a flared end of a dispensing nozzle into the covered gap at a first location, and applying a vacuum to the covered gap at a second location while using the nozzle to inject the sealant into the gap at the first location. This method may be used to seal gaps between ablative elements of a heat shield.

Abstract: A thermal protection system capable of withstanding impact loads without a detrimental effect to its main function, i.e. a protection of the vehicle from high temperature during re-entry or flight in the atmosphere. The system consists of the outer ceramic matrix composite shield, a ceramic tile sandwiched between the shield and the skin of the vehicle and fasteners joining the system and the protected structure. The ceramic matrix composite fasteners are capable of working in the same high-temperature environment as the rest of thermal protection system. They are also strong and stiff due to the presence of ceramic fibers. Cracks in the outer shield produced by impact cannot propagate into the ceramic tile since the shield and tile materials are discontinuous along the interface. The tile cannot disintegrate from the vehicle since it is pressed to its skin by the outer shield. The system can easily be disassembled for inspection and/or repair.

Abstract: A thermal protection system for use upon an air or space vehicle includes a plurality of thermal protection panels formed of a suitable ceramic matrix composition or other thermally protective material each panel having a continuous outer surface free of penetrations or apertures. A plurality of cooperating fasteners are secured between the outer surface of a host vehicle and the undersurface of the thermal protection panels. The fasteners utilize a latching mechanism to secure the panels to the surface of the host air or space vehicle without requiring penetrations or irregularities in the panel surfaces.

Abstract: An electrostatic dissipative paint having a pigment. The pigment includes a composition according to the following formula: (Mg1?x,Znx)Ga2+yO4?? wherein the value of x is a value from about 0 to 1, y is a value from about 0 to 0.04 and ? is in the range of about 0 to about 0.03. An electrostatic dissipative coating system, a method for making an electrostatic dissipative coating, and a method for protecting a spacecraft is also disclosed.

Abstract: A method of forming a heat shield that involves thermally stabilizing a plurality of phenolic microspheres; mixing the thermally stabilized phenolic microspheres with a phenolic resin to form a phenolic ablative material; compressing the phenolic ablative material into a honeycomb core; and allowing the phenolic ablative material to cure.

Abstract: A pyrolyzing flexible ablator comprising a flexible substrate and a pyrolyzing material that will pyrolize upon exposure to a heat flux greater than 5 W/cm2, the pyrolyzing material being bonded to the flexible substrate.

Abstract: An apparatus for heat shielding and deceleration of a spacecraft comprises an unfoldable shield of individual panels made of a high temperature resistant fiber reinforced ceramic and pivotally mounted on the outer structure of the spacecraft. An unfolding ring slides axially along the body of the spacecraft. The panels are pivotally connected to the unfolding ring by compression struts acting as toggle levers. Tension springs pull the unfolding ring axially along the spacecraft body to toggle out the compression struts so as to outwardly pivotally deploy the panels. The unfoldable shield can be arranged on the forward end or the rear end of the spacecraft relative to its flight direction upon entry into the atmosphere.

Abstract: A system for reducing plasma induced communication disruption utilizes an electrophilic gas injectant, bent reentry vehicle nose shaping and a number of flaps configured to provide 3-axis steering capability. Injecting electrophilic gas injectant into plasma accomplishes reduction of plasma induced communication disruption by removing free electrons in the plasma which causes the communication disruption. For reentry vehicles, a bent biconic nose also reduces plasma effect during reentry. By combining electrophilic gas injection and bent nose shaping, the system reduces and/or eliminates plasma induced communication disruption, such as, GPS communication blackout, for a reentry vehicle during reentry. Furthermore, the 3-axis steering capability renders the reentry vehicle more maneuverable. Performance of the reentry vehicle during reentry, such as, target accuracy, is thus improved.

Abstract: A space based orbital kinetic energy weapon system and method of using same is provided. The space based orbital kinetic energy weapon system includes a satellite having a control system configured to maintain an orbit in outer space around the earth and to deorbit the satellite on a desired trajectory corresponding to an earth based target upon a command, and a projectile object operably connected to the satellite. The projectile object includes a dense mass and a heat shield operably surrounding the mass such that at least a portion of the mass survives reentry into the earth's atmosphere and strikes the earth based target delivering its kinetic energy.

Abstract: A space vehicle and its launcher are capable of low-cost launch, safe maneuvering, plus powered go-around for landing.

Abstract: A lifting body aircraft suitable for atmospheric flight and/or as a reentry vehicle. The craft has a substantially flat upper surface, a lower surface with a doubly convex first section and a flat sloping second section. Chines may be provided between the upper and lower surfaces. The doubly convex first section allows the craft to have its center of gravity forward of its longitudinal center line. The flat sloping second half of the lower surface and the substantially flat upper surface form the aft end of an airfoil. A pair of vertical stabilizers enhance stability and include rudders which, along with a pair of elevons, provide steerage. The craft may glide or it may have an engine or rockets for thrust. The result is an extremely stable lifting body design that is well suited for launch or conventional take off, insertion, orbital operations, reentry, atmospheric flight, and conventional landings.

Abstract: A detection-grid is disclosed that is part of a vehicle's thermal protection layer, such as that of a space shuttle. A hybrid digital/analog system detects electrical changes in the detection grid caused by mechanical trauma to a vehicle's external surface. The system produces timely and useful display of such events. Furthermore, with redundant verification of such real-time data, the vehicle can detach from other apparatus, such as an external fuel tank or booster rockets, to execute pre-planned glide or descent scenarios maximizing a crew's and vehicle's safe return before proceeding to orbit. The detection-grid ablates off during re-entry of a regular mission.

Abstract: An experiment system with six different re-entry experiment locations for testing high temperature re-entry materials, creating new thermal protection systems, proving innovative new concepts for spacecraft exterior surfaces and the incremental development of next generation aerospace materials. A commercial transportation system to and from orbit provides a 24-hour return cycle for the experiments on a surface actually re-entering the earth's atmosphere. Now using existing doors, hatches and other points on the reusable launch vehicle's exterior, the actual re-entry environment is experienced by test specimens with quick turn around for a wide variety of different re-entry temperatures ranges for broad testing and development purposes. The reusable launch vehicle launches, remains in orbit for 24 hours and returns to provide an actual test environment for the exterior experiment system.

Abstract: An experiment system with six different re-entry experiment locations for testing high temperature re-entry materials, creating new thermal protection systems, proving innovative new concepts for spacecraft exterior surfaces and the incremental development of next generation aerospace materials. A commercial transportation system to and from orbit provides a 24-hour return cycle for the experiments on a surface actually re-entering the earth's atmosphere. Previously expensive arc jet wind tunnels attempted to simulate the re-entry temperatures and ever changing re-entry flow environment for researchers. Now using existing doors, hatches and other points on the reusable launch vehicle's exterior, the actual re-entry environment is experienced by test specimens with quick turn around for a wide variety of different re-entry temperatures ranges for broad testing and development purposes.

Abstract: The present invention blocks and/or attenuates the upstream travel of acoustic disturbances or sound waves from a flight vehicle or components of a flight vehicle traveling at subsonic speed using a local injection of a high molecular weight gas. Additional benefit may also be obtained by lowering the temperature of the gas. Preferably, the invention has a means of distributing the high molecular weight gas from the nose, wing, component, or other structure of the flight vehicle into the upstream or surrounding air flow. Two techniques for distribution are direct gas injection and sublimation of the high molecular weight solid material from the vehicle surface. The high molecular weight and low temperature of the gas significantly decreases the local speed of sound such that a localized region of supersonic flow and possibly shock waves are formed, preventing the upstream travel of sound waves from the flight vehicle.

Abstract: Methods and an apparatus for a lifeboat of a space station in orbit to be docked with and used in combination with a separately launched logistics module having an upper stage propulsion capability as a space craft for human pilots and passengers to be flown to orbits and trajectories in deep space beyond low Earth orbit including the Moon, at the lowest practical cost.

Abstract: A composite thermal protection structure, for applications such as atmospheric re-entry vehicles, that can withstand temperatures as high as 3600° F. The structure includes an exposed surface cap having a specially formulated coating, an insulator base adjacent to the cap with another specially formulated coating, and one or more pins that extend from the cap through the insulator base to tie the cap and base together, through ceramic bonding and mechanical attachment. The cap and insulator base have corresponding depressions and projections that mate and allow for differences in thermal expansion of the cap and base.

Abstract: The invention discloses systems for mounting heat shields onto a spacecraft. It shows a number of ways to attach heat shield tiles in a way that allows for and accommodates the thermal expansion and contraction of the tiles without overstressing them and without loosing them due to delamination of adhesives. Shown also are various fasteners, which are flexible in several directions and which should preferably be oriented to provide the least resistance to the expected deformations of the tiles.

Abstract: A self-transpiring hot skin for a hypersonic or reusable space vehicle that can provide protection to the vehicle during short periods of abnormally high heat flux (either planned in the flight profile or an off-nominal event). The hot skin includes a ceramic composite structure having an internal cavity that is coupled either to the insulating layer or directly to the support structure of the hypersonic vehicle. The internal cavity includes a material system that vaporizes, sublimes or decomposes into a gas when the temperature exceeds the upper temperature capability of the composite material. The gas transpires through the outer layer of the composite material to provide cooling to the outer layer below the upper temperature capability. Cooling may occur both by conduction of heat from the composite material to the transpiring gas and by the interaction of the transpiring gas with the boundary layer of hypersonic flow over the outer surface, leading to a reduction of the heat flux entering the surface.