Abstract: The present disclosure is directed to a system and method for detecting distortion in a component of a gas turbine engine. In one embodiment, the system includes at least one detectable datum marker configured at least partially within the component. Further, the system includes a non-destructive inspection tool configured to monitor a location of the detectable datum marker over time. Thus, distortion of the component can be detected as a function of movement of the location of the detectable datum.

Abstract: The present disclosure is directed to a system and method for repairing an abradable material coated on a casing of a gas turbine engine. The system includes an articulating guide configured to fit into an access port of the gas turbine engine. Further, the articulating guide has a proximal end and a distal end. The system also includes a repair tool configured at a distal end of the articulating guide. The repair tool includes a body having a proximal end and a shaped distal end, with the shaped distal end extending away from the body. Thus, the shaped distal body is configured to trench out an area of the abradable material comprising a defect. The system also includes a filler material for filling the trenched out area.

Abstract: The present disclosure is directed to an improved system and method for repairing a bend in a turbine blade of a turbine of a gas turbine engine. The system includes an articulating guide configured to fit into an access port of the turbine. The articulating guide includes a proximal end and a distal end. The system also includes a repair tool configured at the distal end of the articulating guide. Further, the repair tool is configured to fit over the turbine blade. Thus, the repair tool is configured to bend the turbine blade to an unbent position while the turbine blade is secured within the turbine.

Abstract: A method for remotely stopping a crack in a component of a gas turbine engine is provided. The method can include inserting an integrated repair interface attached to a cable delivery system within a gas turbine engine; positioning the tip adjacent to a defect within a surface of the component; temporarily attaching the tip adjacent to the defect within the surface on the component; supplying a new material to the area to fill the defect; and heating the new material to fuse the new material to the component within the defect.

Abstract: A method for remotely stopping a crack in a component of a gas turbine engine is provided. The method can include inserting an integrated repair interface attached to a cable delivery system within a gas turbine engine; positioning the tip adjacent to a defect defined on a surface on the component; and locally heating the base of the defect. A method is also provided for clamping a crack defined between a first surface and a second surface of a component of a gas turbine engine. The method can include attaching a strap over the crack such that a first end of the strap is attached to the first surface of the component and the second end of the strap is attached to the second surface of the component.

Abstract: Methods for remotely joining material a surface area of a component of a gas turbine engine are provided. The method can include inserting an integrated repair interface attached to a cable delivery system within a gas turbine engine; positioning a tip of the integrated repair interface adjacent to a defect defined within a surface of the component; temporarily attaching the tip adjacent to the defect within the surface of the component; and supplying a new material to the area to fill the defect.

Abstract: Methods provided for remotely stopping a crack in a component of a gas turbine engine are provided, along with methods of remotely cleaning a surface area of a component of a gas turbine engine. The method can include inserting an integrated repair interface attached to a cable delivery system within a gas turbine engine; positioning the tip adjacent to a defect within a surface of the component; temporarily attaching the tip adjacent to the defect within the surface on the component; and drilling a hole into the base of the defect. An integrated repair interface is also provided.

Abstract: Methods for material build-up on a tip of a blade of a gas turbine engine are provided. The method can include inserting a material supply adjacent to the tip and directing a laser onto the interface of the material supply and the tip such that the material supply melts and attaches to the tip. Methods are also provided for remotely stopping a crack in a component of a gas turbine engine. The method can include inserting an integrated repair interface attached to a cable delivery system within a gas turbine engine; positioning the tip adjacent to a defect within a surface of the component; supplying a new material to the fillable area to fill the defect; and directing a laser to the new material within the fillable area to fuse the new material to the component within the defect.

Abstract: A casing for a turbo-machine at least partially defines a flow path for a working fluid through or around one or more of a compressor section, a combustor assembly, or a turbine section. The casing defines an inner surface and the inner surface defines a plurality of debris routing channels. The plurality of debris routing channels are configured to route debris in a working fluid within the casing towards a debris collection mechanism.

Abstract: In one aspect, a method for performing in situ repairs of internal components of a gas turbine engine may generally include inserting a repair tool through an access port of the gas turbine engine such that the repair tool includes a tip end positioned within the gas turbine engine and a material supply end positioned outside the gas turbine engine. The method may also include positioning the tip end of the repair tool adjacent to a defect of an internal component of the gas turbine engine, wherein the defect defines a fillable volume along a portion of the internal component. In addition, the method may include supplying a filler material to the tip end of the repair tool and expelling the filler material from the tip end of the repair tool such that the fillable volume is at least partially filled with the filler material.

Abstract: A rotor blade formed via additive manufacturing is provided. The rotor blade includes an airfoil and a coupled component. The airfoil includes a plurality of fused layers of a first material formed via additive manufacturing and defines a leading edge and a tip at a distal end. The coupled component includes a plurality of fused layers of a second material formed via additive manufacturing. An interlocking transition zone includes a plurality of projections alternately extending from the airfoil and the coupled component, respectively, to undetachably couple the airfoil and the coupled component.

Abstract: Turbine nozzles are provided for gas turbine engines. The turbine nozzle includes an arcuate inner band; an arcuate outer band; and a nozzle vane disposed between the arcuate inner band and the arcuate outer band. The radially inner end of the nozzle vane is attached to the arcuate inner band through an interlocking transition zone comprising a plurality of projections alternately extending from the radially inner end of the nozzle vane and the arcuate inner band, respectively, to undetachably couple the nozzle vane and the arcuate inner band. Optionally, the radially outer end of each nozzle vane is also attached to the arcuate outer band through an interlocking transition zone.

Abstract: A method of joining two components using additive manufacturing is provided. The method includes forming a first component made of a first material; forming an interlocking transition zone from the first material and a second material; and forming a second component made of the second material. The interlocking transition zone includes a plurality of projections alternately extending from the first component and the second component, respectively, to undetachably couple the first component and the second component.

Abstract: A bladed disk for a gas turbine engine is provided. The bladed disk includes a rotor disk, a plurality of rotor blades, and an interlocking transition zone. The rotor disk includes a plurality of fused layers of a first material formed via additive manufacturing and defines an outer rim. The plurality of rotor blades includes a plurality of fused layers of a second material formed via additive manufacturing. The interlocking transition zone includes a plurality of projections alternately extending from the outer rim of the rotor disk and the plurality of rotor blades, respectively, to undetachably couple the rotor disk and the plurality of rotor blades.

Abstract: Methods for repairing surface of a metal substrate are provided, which can include preparing the surface of the metal substrate for repair; melt attaching a base layer onto the surface of the metal substrate; fusing a plurality of first layers of a first material via additive manufacturing to the base coating; forming an interlocking transition zone via additive manufacturing from the first material and a second material; and fusing a plurality of second layers of the second material via additive manufacturing on the interlocking transition zone. The interlocking transition zone can have a plurality of projections alternately extending from the plurality of first layers and the plurality of second layers, respectively, to undetachably couple the plurality of first layers to the plurality of second layers. A repaired metal substrate is also provided.

Abstract: A turbomachine component includes a body having an exterior surface and an interior surface, an internal cavity defined by the interior surface, and a reactivity neutralizing member arranged within the internal cavity. The reactivity neutralizing member is configured and disposed to neutralize turbomachine combustion products on the interior surface of the body.

Abstract: A bond layer may be applied to the substrate of an article and a first layer may be applied to the bond layer by thermal spray. A second layer may be applied above the first layer by slurry coating.

Abstract: Methods suitable for reducing corner cracking of environmental barrier coating systems. The methods include forming at least one layer of a first portion of the environmental barrier coating system on at least a first region of the component, performing a first heat treatment, forming at least one layer of a second portion of the environmental barrier coating system on a second region of the component, and performing a second heat treatment to the component. Edges of the first portion and edges of the second portion form an angle of at most 60° with respect to a surface of the first region of the component. During the second heat treatment, the second portion of the EBC system expands and the expansion of the second portion of the EBC system is at least partially constrained by the first portion of the EBC system to reduce tensile interfacial stresses in the EBC system.

Abstract: A gear set may include a gear having improved operating performance by including location specific strength and wear properties, more specifically the gear may include a hub that may be made from a first material and a plurality of gear teeth that may be made from a second material and mounted to the hub.

Abstract: A casing for a turbo-machine at least partially defines a flow path for a working fluid through or around one or more of a compressor section, a combustor assembly, or a turbine section. The casing defines an inner surface and the inner surface defines a plurality of debris routing channels. The plurality of debris routing channels are configured to route debris in a working fluid within the casing towards a debris collection mechanism.