Abstract: Aspects of the embodiments set forth a sacrificial plug system including a component having a surface and at least one cooling hole in the surface; a sacrificial plug integrally formed with the component and integrally formed in the at least one cooling hole, where the sacrificial plug includes a top portion; a cover portion; and a bottom portion, the bottom portion integrally formed, engaged to, and connected to at least one cooling hole. The sacrificial plug system also includes at least one connective member integrally formed with the bottom portion of the sacrificial plug and integral with an inner wall of each respective at least one cooling hole; each at least one connective member being severable from the respective inner wall when a force is applied to the top portion, thus permitting the sacrificial plug to be removed from the at least one respective cooling hole.

Abstract: Various embodiments include a metal-repaired ceramic matrix composite (CMC) article, and a method of repairing a CMC article having a defect. Particular embodiments include a method including: removing a defect-containing portion of the CMC article; forming at least one opening in a remaining portion of the CMC article; preparing a metal repair preform for replacing at least the removed portion of the CMC article, wherein a portion of the metal repair preform complements the at least one opening; and attaching the metal repair preform to the remaining portion of the CMC article.

Abstract: A ceramic matrix composite turbine nozzle includes a primary outer nozzle platform; a primary inner nozzle platform; and an airfoil-shaped body extending between the primary inner and primary outer nozzle platforms. The body includes core plies defining a cavity; composite wrap plies circumscribing the core plies and defining an airfoil shape; a secondary outer nozzle platform in contact with the primary outer nozzle platform; and a secondary inner nozzle platform in contact with the primary inner nozzle platform. Each composite wrap ply has two layers of unidirectional fibers oriented transverse to each other and has first and second longitudinal edges. The first and second longitudinal edges are cut into fingers, which are folded in a transverse direction away from a turbine nozzle longitudinal axis and are interleaved between platform plies to define the secondary inner nozzle platform and the secondary outer nozzle platform.

Abstract: A coating includes: at least 34.9 percent by mass silicon dioxide; at least 9.1 percent by mass aluminum oxide; and at least 16.1 percent by mass yttrium oxide.

Abstract: A turbine apparatus is disclosed including a first article and a second article disposed between the first article and a hot gas path of a turbine. The first article includes at least one first article cooling channel in fluid communication with and downstream from a cooling fluid source, and the second article includes at least one second article cooling channel in fluid communication with and downstream from the at least one first article cooling channel. A method for redundant cooling of the turbine apparatus is disclosed including flowing a cooling fluid from the cooling fluid source through at least one first article cooling channel, exhausting the cooling fluid from the at least one first article cooling channel into at least one second article cooling channel, and flowing the cooling fluid through the at least one second article cooling channel.

Abstract: A method may include: in a used metal turbomachine blade including a root including a shank, a platform coupled to the shank and an airfoil coupled to the platform, removing the airfoil, leaving a remaining base including the platform, the shank and the root. The method may also form a radially extending opening through the platform into the shank, and insert a ceramic shank nub extending from a ceramic airfoil into the radially extending opening of the remaining base. The ceramic airfoil is fixedly attached to the remaining base. The method allows reuse of the metal shank while providing the lower cooling requirements of a ceramic airfoil.

Abstract: An airfoil having a wall structure including a plurality of spaced walls for improved cooling and lifetime is disclosed. The airfoil and walls are made by additive manufacturing. The airfoil includes an exterior wall, an intermediate wall, and an interior wall each separated from adjacent walls by a plurality of standoff members; a plurality of outer cooling chambers defined between the exterior and intermediate walls, the chambers partitioned by an outer partition; a plurality of intermediate cooling chambers defined between the intermediate and interior walls, the chambers partitioned by an intermediate partition; a thermal barrier coating on each of the exterior wall and the intermediate wall; a first plurality of impingement openings through the intermediate wall; a second plurality of impingement openings through the interior wall; and a plurality of cooling passages through the exterior wall.

Abstract: An embodiment of an independent cooling circuit for selectively delivering cooling fluid to a component of a gas turbine system includes: at least one coolant feed channel fluidly coupled to a supply of cooling fluid; and an interconnected circuit of cooling channels, including: an interconnected circuit of cooling channels embedded within an exterior wall of the component; an impingement plate; and a plurality of feed tubes connecting the impingement plate to the exterior wall of the component and fluidly coupling a supply of cooling fluid to the interconnected circuit of cooling channels; wherein the cooling fluid flows through the plurality of feed tubes into the interconnected circuit of cooling channels only in response to a formation of a breach in the exterior wall of the component that exposes at least one of the cooling channels.

Abstract: A method may include: in a used metal turbomachine blade including a root including a shank, a platform coupled to the shank and an airfoil coupled to the platform, removing the airfoil, leaving a remaining base including the platform, the shank and the root. The method may also form a radially extending opening through the platform into the shank, and insert a ceramic shank nub extending from a ceramic airfoil into the radially extending opening of the remaining base. The ceramic airfoil is fixedly attached to the remaining base. The method allows reuse of the metal shank while providing the lower cooling requirements of a ceramic airfoil.

Abstract: An insert system for an airfoil plenum includes a first insert and a second insert that include a plurality of impingement openings defined therein. The first insert includes a forward-facing inlet opening. The second insert includes a neck portion having a radial-facing inlet opening, an aft opening, and a cavity in flow communication between the radial-facing inlet opening and the aft opening. The second insert is sized for insertion into the plenum radially through a plenum inlet such that the neck portion is positioned in the plenum inlet. The first insert is sized for insertion into the second insert radially through the radial-facing inlet opening. When the neck portion is positioned in the plenum inlet, the first insert is moveable aftward through the aft opening into an installed position such that the forward-facing inlet opening opens into the cavity.

Abstract: A flow regulating system for increasing a flow of cooling fluid supplied to a cooling system of a component of a gas turbine system is provided. The flow regulating system includes: a pneumatic circuit embedded within a section of the component, the pneumatic circuit including a set of interconnected pneumatic passages; and a pressure-actuated switch fluidly coupled to the pneumatic circuit. The pressure-actuated switch is activated in response to a formation of a breach in the section of the component and an exposure of at least one of the pneumatic passages of the pneumatic circuit embedded in the section of the component. The activation of the pressure-actuated switch increases the flow of cooling fluid supplied to the cooling system of the component.

Abstract: A turbine shroud assembly including an outer shroud arranged within a turbine and further comprising opposed extending portions. The turbine shroud assembly further including an inner shroud shielding the outer shroud from a gas flowing along a gas path within the turbine during operation of the turbine and comprising opposed first and second arcuate portions extending around and in direct contact with a corresponding extending portion of the outer shroud for supporting the inner shroud from the outer shroud. The turbine shroud assembly further including a spline seal extending between the first and second arcuate portions and positioned between the inner shroud and the outer shroud. The turbine shroud assembly further including at least one of the inner shroud, the outer shroud and the spline seal including at least one protrusion for maintaining positive retention of the spline seal during non-operation of the turbine.

Abstract: A nozzle assembly is disclosed, including a CMC nozzle shell, a nozzle spar, and an endwall. The CMC nozzle shell includes a CMC composition and an interior cavity. The nozzle spar is partially disposed within the interior cavity and includes a metallic composition, a cross-sectional conformation, a plurality of spacers protruding from the cross-sectional conformation, the plurality of spacers contacting the CMC nozzle shell, and a spar cap. The endwall includes at least one surface in lateral contact with the spar cap and maintains a lateral orientation of the CMC nozzle shell and the nozzle spar relative to the endwall. The lateral orientation maintains a predetermined throat area of the nozzle assembly. A method for forming the nozzle assembly includes inserting the nozzle spar into the interior cavity, rotating the CMC nozzle shell and the nozzle spar laterally relative to the endwall, and maintaining the lateral orientation.

Abstract: A gas turbine that includes a rotor blade that includes an airfoil. The airfoil may include non-integral base and top portions. The airfoil may include a pin connector connecting the top portion to the base portion. The pin connector may include: a tab extending from the top portion; a complimentary slot for receiving the tab formed in the base portion; an elongated pin cavity formed through an interior region of the airfoil, where the pin cavity intersects the slot to divide the pin cavity into first and second pin cavity segments; a tab aperture formed through the tab, where the tab aperture is positioned so align with the pin cavity upon the tab being received within the slot; and a locking pin that extends continuously through the first segment of the pin cavity, the tab aperture, and the second segment of the pin cavity.

Abstract: A gas turbine that includes a rotor blade that includes an airfoil. The airfoil is defined between a pressure face and a laterally opposed suction face, the pressure face and the suction face extending axially between opposite leading and trailing edges and radially between an outboard tip and a connection with a root of the rotor blade. The airfoil may include non-integral portions in which: a base portion of the airfoil is made from a first material; and a top portion of the airfoil is made from a second material. The airfoil may include a dovetail joint connecting the top portion to the base portion. The dovetail joint may include a dovetail extending from the top portion being received within a complementary dovetail groove formed in the base portion.

Abstract: A turbine component assembly is disclosed, including a first component, a second component, and a circumferentially oriented flat spring. The first component is arranged to be disposed adjacent to a hot gas path, and includes a ceramic matrix composite composition. The second component is adjacent to the first component and arranged to be disposed distal from the hot gas path across the first component. The circumferentially oriented flat spring is disposed on and directly contacting the second component and directly contacting and supporting the first component as a compliant contact interface between the first component and the second component. The circumferentially oriented flat spring provides a radial spring compliance between the first component and the second component.

Abstract: A method for forming a coated turbine component and a coated turbine component is provided. The method includes a step of providing a component having a substrate including a trailing edge face. The method further includes a step of applying a thermal barrier coating or environmental barrier coating selectively to the substrate to form a discontinuous transition from a hot gas path surface at the trailing edge face to discourage hot gas flow along the trailing edge face.

Abstract: A method may include: in a used metal turbomachine blade including a root including a shank, a platform coupled to the shank and an airfoil coupled to the platform, removing the airfoil, leaving a remaining base including the platform, the shank and the root. The method may also form a radially extending opening through the platform into the shank, and insert a ceramic shank nub extending from a ceramic airfoil into the radially extending opening of the remaining base. The ceramic airfoil is fixedly attached to the remaining base. The method allows reuse of the metal shank while providing the lower cooling requirements of a ceramic airfoil.

Abstract: A gas turbine that includes a rotor blade that includes an airfoil. The airfoil may include non-integral portions in which: a base portion is made from a first material; and a top portion is made from a second material. The airfoil may include a connector securing the base portion to the top portion of the airfoil, wherein the connector comprises a wire-lock connector that includes: a tab extending from one of the base portion and the top portion; a complimentary slot for receiving the tab formed in the other one of the base portion and the top portion; a first groove formed in a side of the tab; a second groove formed in a side of the slot; a retaining aperture formed cooperatively via an alignment of the first and second grooves upon the tab being received into the slot; and a retaining wire housed within the retaining aperture.

Abstract: Various embodiments include gas turbine seals and methods of forming such seals for advanced gas path (AGP) heat transfer designed gas turbine engines. In some cases, a turbine includes: a first arcuate component adjacent to a second arcuate component, each arcuate component formed of a ceramic matrix composite and including a seal slot located in an end face, each seal slot having a plurality of axial surfaces and radially facing surfaces extending from opposite ends of the axial surfaces and an overhanging seal assembly disposed in the seal slot. The overhanging seal assembly including an intersegment seal including a plurality of seal segments defining two vertical seal segments and a plurality of horizontal seal segments and a first overhanging positioner shim seal and a second overhanging positioner shim seal, each including an extended portion. The overhanging seal assembly disposed in the seal slot so as to provide position constrainment of the two vertical seal segments.