Abstract: A method of repairing a tip of an airfoil can comprise: removing a coating from the tip of the airfoil; preparing a repair area for bonding, the repair area including the tip and a portion of a concave side of the airfoil; applying an epoxy film and an abrasive component to the repair area; and bonding the epoxy film and the abrasive component to the repair area.

Abstract: A method is provided involving a fairing for a turbine engine. During this method, a detail is provided. The detail includes a carrier and an exterior layer bonded to the carrier. The carrier is configured from or otherwise includes fiber-reinforced composite material. The exterior layer is configured from or otherwise includes polymer material. The detail is arranged with the fairing. The fairing includes an exterior side and an edge. The detail covers and extends along at least a portion of the exterior side. The detail wraps at least partially around the edge. The carrier and at least a first overhang portion of the exterior layer are bonded to the fairing.

Abstract: Aspects of the disclosure are directed to an inlet cap, comprising: a first lug slot located proximate a perimeter of the inlet cap, and a second lug slot located proximate the perimeter of the inlet cap and located peripherally adjacent to the first lug slot, where the first lug slot includes a first ply, the second lug slot includes a second ply that is different from the first ply, and at least a third ply is common to the first lug slot and the second lug slot.

Abstract: A method of immobilizing a contaminant within a component includes identifying the contaminant. The component is thermally processed via a specific thermal cycle related to the contaminant. The specific thermal cycle is performed within an inert environment relative to a substrate. The specific thermal cycle is configured to effect a controlled evaporation of volatiles of the contaminant and a controlled coking of remaining contaminant.

Abstract: An annular bushing includes an annular cylindrical portion with a flange that defines a fastener aperture along an axis. A multiple of spacers extend from an outer surface of the annular cylindrical portion and the flange. The multiple of spacers are operable to control a bond line thickness.

Abstract: A system and method for nose cone edge delamination repair is provided. A nose cone may comprise an outer skirt having an aft edge, a flange, and a pin landing. The system and method may comprise trimming the aft edge of the nose cone to form a repaired edge. A first layer of fiberglass may be placed over repaired edge. A second layer of fiber glass may be placed over the first layer and the repaired edge. The first layer and second layer may be trimmed to create a spot face to allow for lower tolerances when installing the nose cone into a gas turbine engine.

Abstract: Aspects of the disclosure are directed to milling a nose of a first shield of a blade to leave at least one strip of the first shield coupled to a blade body, subsequent to the milling, applying a cryogenic technique to the blade to weaken a bond between the first shield and the blade body, and subsequent to the applying of the cryogenic technique, removing the at least one strip of the first shield from the blade body.

Abstract: A method of remanufacturing a liner panel for a gas turbine engine includes removing a bushing from a damaged component; and molding the bushing with a material charge. A liner panel for a gas turbine engine includes a forward fan exit case liner panel with a donor bushing from a damaged forward fan exit case liner panel. A liner panel for a gas turbine engine includes a forward fan exit case liner panel molded from a material charge that includes a multiple of layers. At least one of the multiple of layer includes a discontinuity.

Abstract: Aspects of the disclosure are directed to an inlet cap, comprising: a first lug slot located proximate a perimeter of the inlet cap, and a second lug slot located proximate the perimeter of the inlet cap and located peripherally adjacent to the first lug slot, where the first lug slot includes a first ply, the second lug slot includes a second ply that is different from the first ply, and at least a third ply is common to the first lug slot and the second lug slot.

Abstract: Aspects of the disclosure are directed to milling a nose of a first shield of a blade to leave at least one strip of the first shield coupled to a blade body, subsequent to the milling, applying a cryogenic technique to the blade to weaken a bond between the first shield and the blade body, and subsequent to the applying of the cryogenic technique, removing the at least one strip of the first shield from the blade body.

Abstract: A system and method for nose cone edge delamination repair is provided. A nose cone may comprise an outer skirt having an aft edge, a flange, and a pin landing. The system and method may comprise trimming the aft edge of the nose cone to form a repaired edge. A first layer of fiberglass may be placed over repaired edge. A second layer of fiber glass may be placed over the first layer and the repaired edge. The first layer and second layer may be trimmed to create a spot face to allow for lower tolerances when installing the nose cone into a gas turbine engine.

Abstract: An annular bushing includes an annular cylindrical portion with a flange that defines a fastener aperture along an axis. A multiple of spacers extend from an outer surface of the annular cylindrical portion and the flange. The multiple of spacers are operable to control a bond line thickness.

Abstract: A method of immobilizing a contaminant within a component includes identifying the contaminate; and thermally processing the component via a specific thermal cycle related to the contaminate. A method of repairing a component includes immobilizing a contaminant within the component; and weld repairing the component subsequent to immobilizing the contaminant. An aerospace component includes a honeycomb structure with a thermally decomposed contaminate and a weld repair.

Abstract: A method of remanufacturing a liner panel for a gas turbine engine includes removing a bushing from a damaged component; and molding the bushing with a material charge. A liner panel for a gas turbine engine includes a forward fan exit case liner panel with a donor bushing from a damaged forward fan exit case liner panel. A liner panel for a gas turbine engine includes a forward fan exit case liner panel molded from a material charge that includes a multiple of layers. At least one of the multiple of layer includes a discontinuity.

Abstract: A method of repairing a part, which has a first side, a second side opposite of the first side, and an aperture that makes the part susceptible to cracking, includes the steps of forming a slot in the first side of the part near the aperture, inserting a stiffener into the slot, and attaching at least one repair ply onto the first side of the part to cover the stiffener.

Abstract: A method of repairing a turbine engine component includes providing a turbine engine component having a first surface and a second surface. A porous structure extends from the first surface to the second surface. The first surface is exposed to a first pressure and the second surface is exposed to a second pressure. The first pressure is higher than the second pressure. A difference between the first pressure and the second pressure is used to pass a cleaning liquid through the porous structure from the first surface to the second surface.

Abstract: A method and system of repairing an engine part having a cracked metal skin bonded to a honeycomb core includes removing a portion of the honeycomb that corresponds to a cracked portion of the metal skin. A crack in the metal skin originates from an edge of the engine part. A portion of the honeycomb is removed to form a slot in the honeycomb beginning at the edge of the part. A doubler is inserted into the slot in the honeycomb and is attached to an underside of the metal skin to form a lap joint between the doubler and the metal skin. In one embodiment, the portion of the metal skin that includes the crack remains. In an alternative embodiment, the cracked portion of the metal skin is removed and a metal patch is attached to the doubler such that the patch replaces the removed metal skin.

Abstract: A repair device and method for repairing a forward frame v-blade on an aircraft engine are described in which the repair device replaces a section of the forward frame v-blade. In one embodiment, the repair device includes a supplemental part, which is secured to the forward frame v-blade. The supplemental part includes a central web and a pair of legs with a leg thickness that is greater than the thickness of corresponding flanges on the forward frame v-blade. Fasteners such as rivets are used to couple the supplemental part to the forward frame v-blade.

Abstract: A doubler assembly repairs at least one damaged aperture of a composite flange. The doubler assembly includes an insert, an adhesive, and a support. The insert is positioned in the damaged aperture to provide compressive load transfer through the composite flange. The adhesive is positioned directly on at least a portion of the composite flange proximate the damaged aperture. The support is positioned over the adhesive.

Abstract: An assembly includes a composite flange, an insert, an adhesive and a support. The composite flange has at least one damaged aperture. The insert is positioned in the damaged aperture and extends therefrom to provide compressive load transfer through the composite flange. The adhesive is positioned directly on at least a portion of the composite flange proximate the damaged aperture. The support is positioned over the adhesive and contacts the insert.