Abstract: A penetrator device has an outer housing of non-conductive, insulating material having a through bore, at least one conductive pin formed in one or two parts extending through the housing and having a first end portion and a second end portion extending out of the respective first and second ends of the housing, a first cladding layer bonded over the first end portion of the pin to form a first bonded assembly, a second cladding layer bonded over the second end portion of the pin to form a second bonded assembly, and the material of the first and second cladding layer comprising a corrosion resistant conductive material different from the pin material.

Abstract: A penetrator device has an outer housing of non-conductive, insulating material having a through bore, at least one conductive pin formed in one or two parts extending through the housing and having a first end portion and a second end portion extending out of the respective first and second ends of the housing, a first cladding layer bonded over the first end portion of the pin to form a first bonded assembly, a second cladding layer bonded over the second end portion of the pin to form a second bonded assembly, and the material of the first and second cladding layer comprising a corrosion resistant conductive material different from the pin material.

Abstract: Aqueous precursor solutions are described that comprise at least one monazite-based material precursor, at least one xenotime-based material precursor or a combination thereof; and a plurality of fine suspended particles of an oxide material. Contemplated oxide composites, as described herein, comprise a plurality of fibers surrounded by at least one monazite or xenotime-based material, wherein the oxide composite has nearly a fully dense matrix. Contemplated embodiments disclosed herein provides a method for producing an oxide composite with nearly fully dense matrix and with all fibers surrounded by a monazite- or xenotime-based material that prevents embrittlement at temperatures at least as high as 1200° C.

Abstract: Aqueous precursor solutions are described that comprise at least one monazite-based material precursor, at least one xenotime-based material precursor or a combination thereof; and a plurality of fine suspended particles of an oxide material. Contemplated oxide composites, as described herein, comprise a plurality of fibers surrounded by at least one monazite or xenotime-based material, wherein the oxide composite has nearly a fully dense matrix. Contemplated embodiments disclosed herein provides a method for producing an oxide composite with nearly fully dense matrix and with all fibers surrounded by a monazite- or xenotime-based material that prevents embrittlement at temperatures at least as high as 1200° C.

Abstract: A protective hybrid composite for a rotor blade is based on the use of tape cast ceramic layers densified by pre-ceramic polymer infiltration methods and laminated together with polymer matrix composite prepregs, with or without an embedded metallic mesh, to form a conforming helicopter blade cladding that is laminated to the blade surface for added erosion protection. The hybrid composite is fabricated to net shape and laminated to the blade using either an adhesive or a polymer composite prepreg inner layer. Installation is accomplished by a standard composite fabrication method of vacuum bagging the blade while the system is laminated to its surface. Repair methods based on removal of ceramic tiles is facilitated by incorporation of a metallic mesh element laminated beneath the ceramic tiles that can be used to heat the tile and decrease its adhesion strength.

Abstract: A protective hybrid composite for a rotor blade is based on the use of tape cast ceramic layers densified by pre-ceramic polymer infiltration methods and laminated together with polymer matrix composite prepregs, with or without an embedded metallic mesh, to form a conforming helicopter blade cladding that is laminated to the blade surface for added erosion protection. The hybrid composite is fabricated to net shape and laminated to the blade using either an adhesive or a polymer composite prepreg inner layer. Installation is accomplished by a standard composite fabrication method of vacuum bagging the blade while the system is laminated to its surface. Repair methods based on removal of ceramic tiles is facilitated by incorporation of a metallic mesh element laminated beneath the ceramic tiles that can be used to heat the tile and decrease its adhesion strength.

Abstract: A temperature tolerant hook and loop attachment, a method of forming a sheet of the hooks and, a method of insulating the skin of a flight vehicle. Temporary loops are formed in a fabric containing temperature tolerant fiber tows, e.g., the tows may be carbon, a metal, a carbide such as carbon silicide, a nitride, or an oxide. The temporary loops are stiffened (e.g., with resin, metal or ceramic), and severed to form temperature tolerant fiber composite hooks. The sheet may be cut and permanently applied, for example, to the skin of a spacecraft or aircraft. A fibrous material, e.g., fibrous insulation or batting, may be pressed in place or formed into the hooks, or the fibrous material may be attached to another structure and pressed in place for a temperature tolerant hook and loop attachment.

Abstract: A protective hybrid composite for a rotor blade is based on the use of tape cast ceramic layers densified by pre-ceramic polymer infiltration methods and laminated together with polymer matrix composite prepregs, with or without an embedded metallic mesh, to form a conforming helicopter blade cladding that is laminated to the blade surface for added erosion protection. The hybrid composite is fabricated to net shape and laminated to the blade using either an adhesive or a polymer composite prepreg inner layer. Installation is accomplished by a standard composite fabrication method of vacuum bagging the blade while the system is laminated to its surface. Repair methods based on removal of ceramic tiles is facilitated by incorporation of a metallic mesh element laminated beneath the ceramic tiles that can be used to heat the tile and decrease its adhesion strength.

Abstract: To produce a ceramic foam-filled structural sandwich panel, a coating of a pre-ceramic slurry is applied on a preform. The preform includes a foam template sandwiched between a plurality of panels. In addition, the coating is cured to the preform, the preform is modified, and the coating is converted to a ceramic.

Abstract: An exemplary morphable ceramic composite structure includes a flexible ceramic composite skin and a truss structure attached to the skin. The truss structure can morph shape of the skin from a first shape to a second shape that is different than the first shape. The flexible ceramic composite skin may include a single-layer of three-dimensional woven fabric fibers and a ceramic matrix composite. The truss structure may include at least one actuatable element or an actuator may move a portion of the truss structure from a first position to a second position. A cooling component may be disposed in thermal communication with the skin. The cooling component may include thermal insulation or a cooling system that circulates cooling fluid in thermal communication with the skin. The morphable ceramic composite structure may be incorporated into any of an air inlet, combustor, exhaust nozzle, or control surfaces of a hypersonic aircraft.

Abstract: To produce a ceramic foam-filled structural sandwich panel, a coating of a pre-ceramic slurry is applied on a preform. The preform includes a foam template sandwiched between a plurality of panels. In addition, the coating is cured to the preform, the preform is modified, and the coating is converted to a ceramic.

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.

Abstract: Monazite or xenotime-based blanket coatings that stiffen ceramic fabrics without causing embrittlement at temperatures of at least as high as 2400° F. are provided. Methods for making the coatings are also provided. The methods comprise the synthesis of high purity, monazite and xenotime powders with the stoichiometric ratio of metal to phosphorous of about 1:1.

Abstract: Monazite or xenotime-based blanket coatings that stiffen ceramic fabrics without causing embrittlement at temperatures of at least as high as 2400° F. are provided. Methods for making the coatings are also provided. The methods comprise the synthesis of high purity, monazite and xenotime powders with the stoichiometric ratio of metal to phosphorous of about 1:1.

Abstract: Monazite or xenotime-based blanket coatings that stiffen ceramic fabrics without causing embrittlement at temperatures of at least as high as 2400° F. are provided. Methods for making the coatings are also provided. The methods comprise the synthesis of high purity, monazite and xenotime powders with the stoichiometric ratio of metal to phosphorous of about 1:1.

Abstract: A woven preform for a ceramic composite has a plurality of layers and structural members. The plurality of layer are of woven yarns of fibrous material. The structural members extend between the layers. The layers and members define interlayer spaces. One or more of the layers may have a plurality of openings extending therethrough. Low density ceramic insulation made be deposited in the interlayer spaces via a slurry that enters the preform, or the preform after it has been made a part of a composite, through the openings. The carrier of the slurry exits the preform, leaving the randomly packed fibers in the interlayer spaces. The structural members may be walls that, along with the layers, define channels. The channels may be used to direct fluid through so as the composite functions as an insulator. The channels may be directed in the warp direction for achieving increased benefits.