Abstract: An outer diameter shroud for retaining a stator of a gas turbine engine having an axis includes an annular body positioned around the axis, extending in an axial direction, and defining a plurality of slots each configured to receive a portion of one of a plurality of stators. The outer diameter shroud further includes a shroud flange coupled to the annular body, extending in a direction perpendicular to the annular body, and configured to be fastened to a case of the gas turbine engine.

Abstract: An outer diameter shroud for retaining a stator of a gas turbine engine having an axis includes an annular body positioned around the axis, extending in an axial direction, and defining a plurality of slots each configured to receive a portion- of one of a plurality of stators. The outer diameter shroud further includes a shroud flange coupled to the annular body, extending in a direction perpendicular to the annular body, and configured to be fastened to a case of the gas turbine engine.

Abstract: A method of determining the thickness of an internal wall in a gas turbine engine component includes the steps of utilizing flash thermography to measure a complete thickness of a component between an outer wall and at least one enlarged cooling channel at a location where an outer cooling channel is positioned between the outer wall and the at least one enlarged cooling channel and where at least one member spans the cooling channel, such that the thickness is through the member which spans the outer cooling channel. An outer thickness of the component is measured from the outer wall to an outer wall of the outer cooling channel. A thickness is determined from an inner wall of the outer cooling channel to the at least one enlarged cooling channel by subtracting the measured outer thickness from the complete thickness, and also subtracting a known thickness of the outer cooling channel.

Abstract: A method of manufacturing a gas turbine engine component includes providing a core having a brittle feature, supporting the feature with a first meltable material, arranging the core with the first meltable material in a first mold, and surrounding the core and the first meltable material with a second meltable material to provide a component shape. The method also includes coating the second meltable material with a refractory material to produce a second mold, removing the first and second meltable material, and casting a component in the second mold.

Abstract: A method of determining the thickness of an internal wall in a gas turbine engine component includes the steps of utilizing flash thermography to measure a complete thickness of a component between an outer wall and at least one enlarged cooling channel at a location where an outer cooling channel is positioned between the outer wall and the at least one enlarged cooling channel and where at least one member spans the cooling channel, such that the thickness is through the member which spans the outer cooling channel. An outer thickness of the component is measured from the outer wall to an outer wall of the outer cooling channel. A thickness is determined from an inner wall of the outer cooling channel to the at least one enlarged cooling channel by subtracting the measured outer thickness from the complete thickness, and also subtracting a known thickness of the outer cooling channel.

Abstract: A gas turbine engine configured to rotate in a circumferential direction about an axis extending through a center of the gas turbine engine comprises a turbine stage. The turbine stage comprises a disk, a plurality of blades and a mini-disk. The disk comprises an outer diameter edge having slots, an inner diameter bore surrounding the axis, a forward face, and an aft face. The plurality of blades is coupled to the slots. The mini-disk is coupled to the aft face of the rotor to define a cooling plenum therebetween in order to direct cooling air to the slots. In one embodiment of the invention, the cooling plenum is connected to a radially inner compressor bleed air inlet through all rotating components so that cooling air passes against the inner diameter bore.

Abstract: A gas turbine engine configured to rotate in a circumferential direction about an axis extending through a center of the gas turbine engine comprises a turbine stage. The turbine stage comprises a disk, a plurality of blades and a mini-disk. The disk comprises an outer diameter edge having slots, an inner diameter bore surrounding the axis, a forward face, and an aft face. The plurality of blades is coupled to the slots. The mini-disk is coupled to the aft face of the rotor to define a cooling plenum therebetween in order to direct cooling air to the slots. In one embodiment of the invention, the cooling plenum is connected to a radially inner compressor bleed air inlet through all rotating components so that cooling air passes against the inner diameter bore.

Abstract: A turbine engine section has a stationary vane stage and a rotating blade stage. The blade stage is spaced from the vane stage to form an annular chamber therebetween. A manifold delivers a pressurized fluid to the chamber. An outer diameter (OD) sealing system restricts leakage of the pressurized fluid from the chamber. An inner diameter (ID) sealing system restricts leakage of the pressurized fluid from the chamber. A flow guide extends radially between the inner diameter sealing system and the manifold. The flow guide bisects the chamber to form a stationary chamber portion and a rotating chamber portion, the rotating chamber portion at least partially along a disk of the first blade stage.

Abstract: A gas turbine engine includes an active clearance control (ACC) system within the engine casing. The case includes a wall structure with a cast-in cooling configuration through Refractory Metal Core technology which is tailored to address traditional hot spots and improve discrete control of case growth.

Abstract: A method of manufacturing a casting article for use in a casting process includes rapidly forming the casting article out of a metallic material. A ceramic coating is applied to an exterior surface of the casting article.

Abstract: A stator vane segment for a gas turbine engine includes at least one airfoil joined to an outer shroud and an inner platform. A sealing element having a first platform radially inward of the inner platform and an abradable material covering at least a portion of the first platform is integrally joined to the inner platform.

Abstract: A gas turbine engine includes an active clearance control (ACC) system within the engine casing. The case includes a wall structure with a cast-in cooling configuration through Refractory Metal Core technology which is tailored to address traditional hot spots and improve discrete control of case growth.

Abstract: A gas turbine engine having a turbine section cooling system is disclosed which is able to cool the turbine section of the gas turbine engine to an optimal operating temperature range, while at the same time not substantially degrading engine performance. The disclosure provides a number of different structures for doing so including the provision of turning foils, turning holes, and turning grooves within a secondary air flow cavity which turns or directs cooling and parasitic leakage air into the turbine section in a direction in substantial alignment with a gas flow path through the turbine section.