Abstract: A tooling assembly and method for forming a flange of a composite component is provided. The tooling assembly includes a mold configured to receive a composite material. The composite material may include a plurality of fabric plies impregnated with resin. The fabric plies may be debulked as they are being laid using a plurality of debulking tools, with the radius of each debulking tool growing as additional fabric plies are laid. A gas-permeable bagging material may be placed along the edge of the composite material to restrict resin flow while allowing for outgassing. The tooling assembly may further include a flange shoe tool that is joined with the mold to form the flange of the composite material. The mold and flange shoe tool may define a chamber and venting passageways that allow gases such as volatile compounds to escape while the composite component is being cured.

Abstract: Method includes forming a preform utilizing a polyimide resin-impregnated fiber-reinforced layers; removing solvent from the system at initial vacuum, pressure, and temperature conditions for an initial time interval sufficient to remove substantially all the solvent; imidizing the polyimide resin system under second vacuum, pressure, and temperature conditions for a second time interval sufficient to substantially completely imidize the polyimide resin; consolidating the preform following imidization under third vacuum, pressure, and temperature conditions and including applying pressure to the preform when the preform is at a predetermined temperature; and solidifying the preform under fourth vacuum, pressure, and temperature conditions to provide a cured laminate structure having a shape of a turbine engine component.

Abstract: A method of repairing a thermal barrier coating (16) on a component (10) designed for use in a hostile thermal environment, such as turbine, combustor and augmentor components of a gas turbine engine. The method more particularly involves repairing a thermal barrier coating (16) on a component (10) that has suffered localized spallation (20) of the thermal barrier coating (16). After cleaning the surface area (22) of the component (10) exposed by the localized spallation (20), a ceramic paste (24) comprising a ceramic powder in a binder is applied to the surface area (22) of the component (10). The binder is then reacted to yield a ceramic-containing repair coating (26) that covers the surface area of the component and comprises the ceramic powder in a matrix of a material formed when the binder was reacted.

Abstract: A method of repairing a thermal barrier coating (16) on a component (10) designed for use in a hostile thermal environment, such as turbine, combustor and augmentor components of a gas turbine engine. The method more particularly involves repairing a thermal barrier coating (16) on a component (10) that has suffered localized spallation (20) of the thermal barrier coating (16). After cleaning the surface area (22) of the component (10) exposed by the localized spallation (20), a ceramic paste (24) comprising a ceramic powder in a binder is applied to the surface area (22) of the component (10). The binder is then reacted to yield a ceramic-containing repair coating (26) that covers the surface area of the component and comprises the ceramic powder in a matrix of a material formed when the binder was reacted.

Abstract: A protective coating and coating method for protecting a thermal barrier coating (TBC) on a component, such as a component of a gas turbine engine. The protective coating comprises alumina particles in a silica-containing matrix, and may be substantially homogeneous or formed of multiple layers having different compositions. The composition and relative amounts of alumina and matrix material in the protective coating enable the coating to react with molten compounds containing calcia, magnesia, alumina and/or silica (CMAS), forming a compound with a melting temperature that is significantly higher than CMAS. As such, infiltration of molten CMAS into the TBC is significantly reduced or entirely avoided.

Abstract: A method of repairing a thermal barrier coating on a component designed for use in a hostile thermal environment, such as turbine, combustor and augmentor components of a gas turbine engine. The method more particularly involves repairing a thermal barrier coating on a component that has suffered localized spallation of the thermal barrier coating. After cleaning the surface area of the component exposed by the localized spallation, a ceramic paste comprising a ceramic powder in a binder is applied to the surface area of the component. The binder is then reacted to yield a ceramic-containing repair coating that covers the surface area of the component and comprises the ceramic powder in a matrix of a material formed when the binder was reacted. The binder is preferably a ceramic precursor material that can be converted immediately to a ceramic or allowed to thermally decompose over time to form a ceramic, such that the repair coating has a ceramic matrix.