Abstract: A combustor assembly for a gas turbine engine includes an outer liner that at least partially defines a combustion chamber. The outer liner extends between an aft end and a forward end generally along an axial direction within an outer casing of the gas turbine engine. The outer liner includes an inner surface facing the combustion chamber and an outer surface facing away from the combustion chamber. The combustor assembly includes a radial damper assembly disposed against the outer liner at a plurality of circumferential points.

Abstract: The present disclosure generally relates to partial integrated core-shell investment casting molds that can be assembled into complete molds. Each section of the partial mold may contain both a portion of a core and portion of a shell. Each section can then be assembled into a mold for casting of a metal part. The partial integrated core-shell investment casting molds and the complete molds may be provided with filament structures corresponding to cooling hole patterns on the surface of the turbine blade or the stator vane, which provides a leaching pathway for the core portion after metal casting. The invention also relates to core filaments that can be used to supplement the leaching pathway, for example in a core tip portion of the mold.

Abstract: Partial integrated core-shell investment casting molds that can be assembled into complete molds are provided herein. Each section of the partial mold may contain both a portion of a core and portion of a shell. Each section can then be assembled into a mold for casting of a metal part. The partial integrated core-shell investment casting molds and the complete molds may be provided with filament structures corresponding to cooling hole patterns on the surface of the turbine blade or the stator vane, which provides a leaching pathway for the core portion after metal casting. Core filaments that can be used to supplement the leaching pathway, for example in a core tip portion of the mold are also provided herein.

Abstract: The present disclosure generally relates to partial integrated core-shell investment casting molds that can be assembled into complete molds. Each section of the partial mold may contain both a portion of a core and portion of a shell. Each section can then be assembled into a mold for casting of a metal part. The partial integrated core-shell investment casting molds and the complete molds may be provided with filament structures corresponding to cooling hole patterns on the surface of the turbine blade or the stator vane, which provides a leaching pathway for the core portion after metal casting. The invention also relates to core filaments that can be used to supplement the leaching pathway, for example in a core tip portion of the mold.

Abstract: Partial integrated core-shell investment casting molds that can be assembled into complete molds are provided herein. Each section of the partial mold may contain both a portion of a core and portion of a shell. Each section can then be assembled into a mold for casting of a metal part. The partial integrated core-shell investment casting molds and the complete molds may be provided with filament structures corresponding to cooling hole patterns on the surface of the turbine blade or the stator vane, which provides a leaching pathway for the core portion after metal casting. Core filaments that can be used to supplement the leaching pathway, for example in a core tip portion of the mold are also provided herein.

Abstract: A combustor assembly for a gas turbine engine includes an outer liner that at least partially defines a combustion chamber. The outer liner extends between an aft end and a forward end generally along an axial direction within an outer casing of the gas turbine engine. The outer liner includes an inner surface facing the combustion chamber and an outer surface facing away from the combustion chamber. The combustor assembly includes a radial damper assembly disposed against the outer liner at a plurality of circumferential points.

Abstract: A gas turbine engine and combustor assembly is generally provided. The gas turbine engine may include an outer casing while the combustor assembly may include a liner and a damper assembly. The liner may at least partially define a combustion chamber extending between an aft end and a forward end generally along an axial direction within the outer casing. The liner may include an inner surface facing the combustion chamber and an outer surface facing away from the combustion chamber. The damper assembly may extend between the outer casing and the outer surface of the liner. The damper assembly may include a selectively separable support and damper spring. The damper spring may be disposed between the support and the liner.

Abstract: The present disclosure generally relates to partial integrated core-shell investment casting molds that can be assembled into complete molds. Each section of the partial mold may contain both a portion of a core and portion of a shell. Each section can then be assembled into a mold for casting of a metal part. The partial integrated core-shell investment casting molds and the complete molds may be provided with filament structures corresponding to cooling hole patterns on the surface of the turbine blade or the stator vane, which provides a leaching pathway for the core portion after metal casting. The invention also relates to core filaments that can be used to supplement the leaching pathway, for example in a core tip portion of the mold.

Abstract: A turbine blade sealing structure for a ceramic matrix composite component is provided. The turbine sealing structure includes at least one top ply abutting a top portion of a shank. The at least one top ply extends out over the top portion of the shank. The turbine sealing structure includes at least one side ply abutting a side portion of the shank. The at least one side ply extends along a side portion of the shank. The at least one top ply and at least one side ply form an angel wing and a skirt with on the shank of the ceramic matrix composite component.

Abstract: A rotor blade assembly for a rotor of a gas turbine engine having an axis of rotation includes a shank portion formed from a ceramic matrix composite (CMC) material. The rotor blade assembly also includes a platform portion formed from a substantially similar CMC material as that of the shank portion. The platform portion is coupled to the shank portion. The platform portion and the shank portion cooperate to at least partially define two opposing side portions of the rotor blade assembly. The opposing side portions are angularly separated with respect to the axis of rotation. The rotor blade assembly further includes a damper retention apparatus. The damper retention apparatus is coupled to the shank portion. The damper retention apparatus includes at least one angled bracket apparatus extending toward a circumferentially adjacent rotor blade assembly.

Abstract: A clearance control apparatus for a gas turbine engine including an annular turbine case having opposed inner and outer surfaces; an annular manifold surrounding a portion of the turbine case, the manifold including: an inlet port in fluid communication with the manifold and the outer surface of the turbine case, and an exit port; and a bypass pipe having an upstream end coupled to the exit port, a downstream end coupled to a low-pressure sink, and a valve disposed between upstream and downstream ends, the valve selectively moveable between a first position which blocks flow between the upstream and downstream ends, and a second position which permits flow between the upstream and downstream ends.

Abstract: Methods for fabricating a component of a gas turbine engine are provided. In one embodiment, the method includes molding a CMC material to form a first portion of the gas turbine engine component, processing the first portion to form a first assembly, preparing the first assembly and a second portion of the gas turbine engine component for processing, and processing the first assembly and second portion to form a second assembly. In another embodiment, the method includes processing a first plurality of CMC plies to form a first assembly; positioning the first assembly and a second plurality of CMC plies on a tool for processing, the first assembly defining a first plane, the second plurality of plies defining a second plane, wherein the second plane is perpendicular to the first plane; and processing the first assembly and the second plurality of plies to form a second assembly.

Abstract: The present disclosure generally relates to integrated core-shell investment casting molds including a main core portion, a core tip portion, and a shell portion with at least one cavity between the core portion and the shell portion. The cavity defines the shape of a cast component upon casting and removal of the ceramic mold. These molds also provide filament structures corresponding to cooling hole patterns in the surface of the turbine blade or the stator vane, which provide a leaching pathway for the core portion after metal casting. At least two ceramic tip filaments connect the core tip portion and the shell portion and eliminate the need for tip pins or a shell lock to hold the tip plenum core in place during casting. The invention also relates to core filaments that can be used to supplement the leaching pathway, for example in a core tip portion of the mold.

Abstract: The present disclosure generally relates to integrated core-shell investment casting molds that provide a filament structure corresponding to a cooling hole pattern in the surface of the turbine blade or stator vane, which provide a leaching pathway for the core portion after metal casting. The invention also relates to core filaments that can be used to supplement the leaching pathway, for example in a core tip portion of the mold.

Abstract: A gas turbine engine and combustor assembly is generally provided. The gas turbine engine may include an outer casing while the combustor assembly may include a liner and a damper assembly. The liner may at least partially define a combustion chamber extending between an aft end and a forward end generally along an axial direction within the outer casing. The liner may include an inner surface facing the combustion chamber and an outer surface facing away from the combustion chamber. The damper assembly may extend between the outer casing and the outer surface of the liner. The damper assembly may include a selectively separable support and damper spring. The damper spring may be disposed between the support and the liner.

Abstract: Ceramic matrix composite (CMC) components and methods for forming CMC components of gas turbine engines are provided. In one embodiment, a CMC component for a gas turbine engine includes an inner wall defining a first inner surface; an outer wall defining a second inner surface; and a nozzle extending from the inner wall to the outer wall. The inner wall, outer wall, and nozzle are integrally formed from a CMC material such that the inner wall, outer wall, and nozzle are a single unitary component. An exemplary method for forming a CMC component includes laying up a plurality of plies of a CMC material; processing the plurality of plies to form a green state component; firing the green state component; and densifying the fired component to produce a final unitary component. The unitary component comprises a combustor liner portion and a combustor discharge nozzle stage portion.

Abstract: A shroud for a gas turbine is provided. The shroud can include an outer box having a first wall and a second wall circumferentially spaced apart from the first wall. An inner box is positioned within the outer box. The inner box defines one or more passageways extending therethrough and a first chamber. The inner box and the outer box collectively define a second chamber. The one or more passageways defined by the inner box fluidly couple the first chamber and the second chamber. The first wall of the outer box defines a first boss and first notch, and the second wall of the outer box defines a second boss and notch. A gas turbine is also provided that can include a plurality of such turbine shrouds.

Abstract: Methods for fabricating a component of a gas turbine engine are provided. In one embodiment, the method includes molding a CMC material to form a first portion of the gas turbine engine component, processing the first portion to form a first assembly, preparing the first assembly and a second portion of the gas turbine engine component for processing, and processing the first assembly and second portion to form a second assembly. In another embodiment, the method includes processing a first plurality of CMC plies to form a first assembly; positioning the first assembly and a second plurality of CMC plies on a tool for processing, the first assembly defining a first plane, the second plurality of plies defining a second plane, wherein the second plane is perpendicular to the first plane; and processing the first assembly and the second plurality of plies to form a second assembly.

Abstract: A process for producing airfoil components containing ceramic-based materials and having a tip cap. The process entails forming an airfoil portion of the component from an airfoil portion material that contains a precursor of a ceramic-based material. The airfoil portion material defines concave and convex walls of the airfoil portion, and the concave and convex walls define a tip region of the airfoil portion and at least one cavity within the airfoil portion. At least a first ply is formed that contains a precursor of a ceramic-based material, and the first ply at least partially closes the cavity at the tip region of the airfoil portion. The airfoil portion material and the first ply are then cured so that the first ply forms a tip cap that closes the cavity and the precursors of the airfoil portion material and first ply are converted to the ceramic-based materials thereof.

Abstract: Hybrid turbine airfoil components containing a ceramic material, in which detailed features of the components are formed of materials other than ceramic materials. The components include a first component formed of a ceramic-based material and a second component formed of a metallic material. The first component comprises an airfoil portion and a nub, and the second component is separately formed and attached to the first component by casting the metallic material around the nub of the first component. The second component includes a platform portion between the airfoil portion and the nub of the first component and a dovetail portion on the nub of the first component. Each of the platform and dovetail portions has at least one off-axis geometric feature that results in the second component having a more complex geometry than the first component.