Combustor-turbine seal interface for gas turbine engine

A combustor-turbine seal interface is provided for deployment within a gas turbine engine. In one embodiment, the combustor-turbine assembly a combustor, a turbine nozzle downstream of the combustor, and a first compliant dual seal assembly. The first compliant dual seal assembly includes a compliant seal wall sealingly coupled between the combustor and the turbine nozzle, a first compression seal sealingly disposed between the compliant seal wall and the turbine nozzle, and a first bearing seal generally defined by the compliant seal wall and the turbine nozzle. The first bearing seal is sealingly disposed in series with the first compression seal.

TURBINE BLADE PLATFORM

A turbine blade assembly is provided. The turbine blade assembly comprises a turbine blade comprising a cavity, and a blade platform supporting the turbine blade, the cavity extending into the blade platform. The blade platform comprises an upper surface adjacent the turbine blade and a lower surface comprising a first rib, the cavity extending into the first rib, the first rib coupled to the lower surface, tapering as it extends away from the turbine blade, and comprising a first port extending from the cavity to the upper surface.

Turbine component with shaped cooling pins

A turbine component with shaped cooling pins is provided. The turbine component includes at least one cooling circuit defined within the turbine component, the at least one cooling circuit in fluid communication with a source of a cooling fluid. The turbine component includes at least one shaped cooling pin disposed in the at least one cooling circuit. The at least one shaped cooling pin has a first end and a second end extending along an axis. The first end has a first curved surface defined by a minor diameter and the second end has a second curved surface defined by a major diameter. The first curved surface is upstream in the cooling fluid and the minor diameter is less than the major diameter.

Axially-split radial turbines and methods for the manufacture thereof

Embodiments of an axially-split radial turbine, as are embodiments of a method for manufacturing an axially-split radial turbine. In one embodiment, the method includes the steps of joining a forward bladed ring to a forward disk to produce a forward turbine rotor, fabricating an aft turbine rotor, and disposing the forward turbine rotor and the aft turbine rotor in an axially-abutting, rotationally-fixed relationship to produce the axially-split radial turbine.

Axially-split radial turbines

Embodiments of an axially-split radial turbine, as are embodiments of a method for manufacturing an axially-split radial turbine. In one embodiment, the method includes the steps of joining a forward bladed ring to a forward disk to produce a forward turbine rotor, fabricating an aft turbine rotor, and disposing the forward turbine rotor and the aft turbine rotor in an axially-abutting, rotationally-fixed relationship to produce the axially-split radial turbine.

Dual alloy turbine rotors and methods for manufacturing the same

Dual alloy turbine rotors and methods for manufacturing the same are provided. The dual alloy turbine rotor comprises an assembled blade ring and a hub bonded to the assembled blade ring. The assembled blade ring comprises a first alloy selected from the group consisting of a single crystal alloy, a directionally solidified alloy, or an equi-axed alloy. The hub comprises a second alloy. The method comprises positioning a hub within a blade ring to define an interface between the hub and the blade ring. The interface is a non-contacting interface or a contacting interface. The interface is enclosed by a pair of diaphragms. The interface is vacuum sealed. The blade ring is bonded to the hub after the vacuum sealing step.

DUAL ALLOY TURBINE ROTORS AND METHODS FOR MANUFACTURING THE SAME

Dual alloy turbine rotors and methods for manufacturing the same are provided. The dual alloy turbine rotor comprises an assembled blade ring and a hub bonded to the assembled blade ring. The assembled blade ring comprises a first alloy selected from the group consisting of a single crystal alloy, a directionally solidified alloy, or an equi-axed alloy. The hub comprises a second alloy. The method comprises positioning a hub within a blade ring to define an interface between the hub and the blade ring. The interface is a non-contacting interface or a contacting interface. The interface is enclosed by a pair of diaphragms. The interface is vacuum sealed. The blade ring is bonded to the hub after the vacuum sealing step.

Methods and tooling assemblies for the manufacture of metallurgically-consolidated turbine engine components

Embodiments of a method for manufacturing a turbine engine component are provided, as are embodiments of a thermal growth constraint tool for the manufacture of turbine engine components. In one embodiment, the method includes the steps of obtaining a plurality of arched pieces, arranging the plurality of arched pieces in a ring formation, and bonding the plurality of arched pieces together to produce a monolithic ring by heating the ring formation to a predetermined bonding temperature while constraining the outward radial growth thereof.

DUAL ALLOY BLADED ROTORS SUITABLE FOR USAGE IN GAS TURBINE ENGINES AND METHODS FOR THE MANUFACTURE THEREOF

Dual alloy bladed rotors are provided, as are methods for manufacturing dual alloy bladed rotors. In one embodiment, the method includes arranging bladed pieces in a ring formation such that contiguous bladed pieces contact along shank-to-shank bonding interfaces. The ring formation is positioned around a hub disk, which is contacted by the bladed pieces along a shank-to-hub bonding interface. A metallic sealing material is deposited between contiguous bladed pieces utilizing, for example, a laser welding process to produce an annular seal around the ring formation. A hermetic cavity is then formed, which is circumferentially bounded by the annular seal and which encloses the shank-to-shank and shank-to-hub bonding interface. Afterwards, a Hot Isostatic Pressing process is performed during which the ring formation and the hub disk are exposed to elevated pressures external to the hermetic cavity sufficient to diffusion bond the shank-to-shank and shank-to-hub bonding interface.

Impingement baffle for gas turbine engine

An impingement baffle for directing a cooling fluid onto a target surface includes a baffle body having a first end opposite a second end, and a first side opposite a second side. The second side is spaced a distance apart from the target surface, with the distance varying from the first end to the second end. The baffle body defines impingement holes that extend through the baffle body from the first side to the second side. The impingement holes are spaced apart along the baffle body to receive the cooling fluid. The impingement baffle includes tubular extensions coupled to the second side. Each tubular extension is in fluid communication with a respective one of the impingement holes to direct the cooling fluid onto the target surface. Each tubular extension extends for a length from the second side, and the length of each tubular extension is based on the distance.

Directed cooling for rotating machinery

A rotating machine includes a hub portion, wherein the hub portion comprises a forward face and an aft face. The rotating machine further includes a cooling channel formed on either the forward face or the aft face and configured to direct cooling air to a location on the rotating machine, wherein the cooling channel extends from a radially inner location along the face to a radially outer location along the face, and wherein the cooling channel is configured as a recess formed into an outer surface of the face.

TURBINE TIP SHROUD ASSEMBLY WITH PLURAL SHROUD SEGMENTS HAVING INTER-SEGMENT SEAL ARRANGEMENT

A shroud assembly for a gas turbine engine includes a plurality of shroud segments that are attached to a shroud support with an inter-segment joint defined between shroud segments. The shroud assembly also includes a cooling flow path cooperatively defined by the shroud support and the first shroud segment. The cooling flow path includes an internal cooling passage within the shroud segments. The cooling flow path includes an outlet chamber configured to receive flow from the internal cooling passage. The shroud assembly additionally includes a seal arrangement that extends across the inter-segment joint. The seal arrangement, the first shroud segment, and the second shroud segment cooperatively define a seal chamber that is enclosed.

Turbine shroud assemblies for gas turbine engines

A turbine shroud assembly includes a shroud support and a shroud segment. The shroud support structure includes a forward support rail and an aft support rail. The forward support rail includes forward first engagement structures and the aft support rail includes aft first engagement structures. The shroud segment includes a forward segment rail and an aft segment rail. The forward segment rail includes forward second engagement structures positioned on a forward segment rail periphery and the aft segment rail includes second engagement structures positioned on an aft segment rail periphery. The forward first engagement structures radially and circumferentially engage with the forward second engagement structures and the aft first engagement structures radially and circumferentially engage with the aft second engagement structures to radially and circumferentially interlock the shroud segment to the shroud support structure.

Turbine scroll assembly for gas turbine engine

A gas turbine engine includes a compressor section and a combustion section with a scroll, a scroll baffle, a combustor, and a combustor case. The scroll defines an interior scroll flow path. The scroll baffle surrounds the scroll to define a scroll cooling passage. The combustor case surrounds the combustor and the scroll baffle to define a collector space. Moreover, the engine includes a turbine section with a turbine rotor and a turbine rotor blade shroud that includes a shroud cooling passage. The compressor flow path is fluidly connected to the scroll for cooling the scroll. Also, the scroll cooling passage is fluidly connected to the shroud cooling passage for cooling the turbine rotor blade shroud. Furthermore, the shroud cooling passage is fluidly connected to the collector space. Flow from the collector space flows into the combustor, along the interior scroll flow path, toward the turbine rotor.

DIRECT TRANSFER AXIAL TANGENTIAL ONBOARD INJECTOR SYSTEM (TOBI) WITH SELF-SUPPORTING SEAL PLATE

An apparatus for cooling turbine blades in a turbine engine including a direct transfer axial tangential onboard injector (TOBI) for a turbine rotor and a self-supporting seal plate disposed on a rotating disk for the turbine engine. The TOBI includes a plurality of openings emanating a flow of cooling air. The self-supporting seal plate comprises a plurality of shaped cooling holes in fluid communication with the flow of cooling air emanating from the TOBI. The rotating disk includes a plurality of turbine blade slots formed therein. The plurality of cooling holes are in fluid communication with the plurality of turbine blade slots for directing the flow of cooling air to provide cooling to the plurality of turbine blades. The plurality of openings, the plurality of cooling holes and the plurality of turbine blade slots are in axial alignment and optimized to minimize radial and hoop stresses.

TURBINE TIP SHROUD ASSEMBLY WITH PLURAL SHROUD SEGMENTS HAVING INTER-SEGMENT SEAL ARRANGEMENT

A shroud assembly for a gas turbine engine includes a plurality of shroud segments that are attached to a shroud support with an inter-segment joint defined between shroud segments. The shroud assembly also includes a cooling flow path cooperatively defined by the shroud support and the first shroud segment. The cooling flow path includes an internal cooling passage within the shroud segments. The cooling flow path includes an outlet chamber configured to receive flow from the internal cooling passage. The shroud assembly additionally includes a seal arrangement that extends across the inter-segment joint. The seal arrangement, the first shroud segment, and the second shroud segment cooperatively define a seal chamber that is enclosed.

AIRFOIL WITH LEADING EDGE CONVECTIVE COOLING SYSTEM

An airfoil includes a leading edge and an opposing trailing edge. The airfoil includes a pressure sidewall and an opposing suction sidewall. A leading edge cavity is defined between the pressure sidewall and the suction sidewall. The leading edge cavity has a first end opposite the leading edge and a second end defined at a rib. The airfoil includes at least one pin structure defined in the leading edge cavity between the first end and the second end. The at least one pin structure includes a main body and a first branch. The main body is coupled to the second end and extends toward the first end. The first branch extends from the main body toward the first end.

Method of producing an abrasive tip for a turbine blade

A method of producing an abrasive tip for a turbine blade includes producing or obtaining a metal powder that is mixed with an abrasive ceramic powder and producing or obtaining a metallic mold that is in the shape of an airfoil. The metallic mold includes a hollow interior portion. The method further includes sealing the metal and ceramic powder mixture within the hollow interior portion of the metallic mold under vacuum and subjecting the sealed mold to a hot isostatic pressing process. The hot isostatic pressing process compacts and binds the metal and ceramic powder mixture together into a solid article in the shape of the airfoil. Still further, the method includes slicing the solid article into a plurality of airfoil-shaped slices and bonding one slice of the plurality of airfoil-shaped slices to a tip portion of a turbine blade.

Turbine shroud support coupling assembly

A coupling assembly for a turbine shroud is provided. The coupling assembly comprises a rotatable positioning block having a first surface, and a biasing spring having a second surface, the second surface generally facing the first surface, and the biasing spring adapted to exert a force toward the positioning block when compressed.

Turbine nozzle with reduced leakage feather seals

A turbine nozzle for a gas turbine engine includes a plurality of nozzle segments that are configured to be assembled into a full ring such that each one of the plurality of nozzle segments is adjacent to another one of the plurality of nozzle segments. Each one of the plurality of nozzle segments includes an endwall segment and a nozzle vane. The turbine nozzle includes a feather seal interface defined by endwall segments of adjacent ones of the plurality of nozzle segments. The feather seal interface is defined along an area of reduced pressure drop through a pressure field defined between adjacent nozzle vanes of the plurality of nozzle segments to reduce leakage through the plurality of nozzle segments. The turbine nozzle includes a feather seal received within the feather seal interface that cooperates with the feather seal interface to reduce leakage through the plurality of nozzle segments.

Turbine nozzle assembly including radially-compliant spring member for gas turbine engine

Embodiments of a turbine nozzle assembly are provided for deployment within a gas turbine engine (GTE) including a first GTE-nozzle mounting interface. In one embodiment, the turbine nozzle assembly includes a turbine nozzle flowbody, a first mounting flange configured to be mounted to the first GTE-nozzle mounting interface, and a first radially-compliant spring member coupled between the turbine nozzle flowbody and the first mounting flange. The turbine nozzle flowbody has an inner nozzle endwall and an outer nozzle endwall, which is fixedly coupled to the inner nozzle endwall and which cooperates therewith to define a flow passage through the turbine nozzle flowbody. The first radially-compliant spring member accommodates relative thermal movement between the turbine nozzle flowbody and the first mounting flange to alleviate thermomechanical stress during operation of the GTE.

Turbine nozzles and methods of manufacturing the same

Turbine nozzles and methods of manufacturing the turbine nozzles are provided. In an embodiment, by way of example only, a turbine nozzle includes a first ring, a vane, and a first joint. The first ring comprises a single unitary component and having a first opening and including a first metal alloy. The vane includes a first end disposed in the first opening and includes a second metal alloy. The first joint is formed in the first opening between the first ring and the vane and includes a first braze layer and an oxide layer. The first braze layer is disposed adjacent to the oxide layer, and the first braze layer and the oxide layer are disposed between the first ring and the vane.

TURBINE NOZZLES AND METHODS OF MANUFACTURING THE SAME

A turbine nozzle is provided and includes a first ring having a first microstructure, a vane extending from the first ring, a first porous zone between the first ring and the vane that is more porous than the first microstructure to attenuate thermo-mechanical fatigue cracking between the vane and the first ring. Methods of manufacturing the turbine nozzle are also provided.

Diverging-converging cooling passage for a turbine blade

A turbine blade and a radial turbine having at least one blade is provided. The turbine blade includes a trailing edge and a leading edge opposite the trailing edge. The turbine blade also includes a cooling passage defined internally within the turbine blade. The cooling passage is in fluid communication with a source of cooling fluid via a single inlet to receive a cooling fluid. The cooling passage diverges at a first point downstream from the single inlet into at least two branches that extend along the at least one blade from the first point to a second point near a tip of the leading edge and the cooling passage converges at the second point.

Turbine nozzles and methods of manufacturing the same

A turbine nozzle is provided and includes a first ring having a first microstructure, a vane extending from the first ring, a first porous zone between the first ring and the vane that is more porous than the first microstructure to attenuate thermo-mechanical fatigue cracking between the vane and the first ring. Methods of manufacturing the turbine nozzle are also provided.

Methods for processing bonded dual alloy rotors including differential heat treatment processes

Methods for processing bonded dual alloy rotors are provided. In one embodiment, the method includes obtaining a bonded dual alloy rotor including rotor blades bonded to a hub disk. The rotor blades and hub disk are composed of different alloys. A minimum processing temperature (TDISK_PROCESS_MIN) for the hub disk and a maximum critical temperature for the rotor blades (TBLADE_MAX) is established such that TBLADE_MAX is less than TDIsK_PROCESS_MIN. A differential heat treatment process is then performed during which the hub disk is heated to processing temperatures equal to or greater than TDISK_PROCESS_MIN, while at least a volumetric majority of each of the rotor blades is maintained at temperatures below TBLADE_MAX. Such a targeted differential heat treatment process enables desired metallurgical properties (e.g., precipitate hardening) to be created within the hub disk, while preserving the high temperature properties of the rotor blades and any blade coating present thereon.

COMBUSTOR-TURBINE SEAL INTERFACE FOR GAS TURBINE ENGINE

A combustor-turbine seal interface is provided for deployment within a gas turbine engine. In one embodiment, the combustor-turbine assembly a combustor, a turbine nozzle downstream of the combustor, and a first compliant dual seal assembly. The first compliant dual seal assembly includes a compliant seal wall sealingly coupled between the combustor and the turbine nozzle, a first compression seal sealingly disposed between the compliant seal wall and the turbine nozzle, and a first bearing seal generally defined by the compliant seal wall and the turbine nozzle. The first bearing seal is sealingly disposed in series with the first compression seal.

METHODS FOR MANUFACTURING A TURBINE NOZZLE WITH SINGLE CRYSTAL ALLOY NOZZLE SEGMENTS

Methods for manufacturing a turbine nozzle are provided. A plurality of nozzle segments is formed. Each nozzle segment comprises an endwall ring portion with at least one vane. The plurality of nozzle segments are connected to an annular endwall forming a segmented annular endwall concentric to the annular endwall with the at least one vane of each nozzle segment extending between the segmented annular endwall and the annular endwall. 1. A method for manufacturing a turbine nozzle comprising:forming a plurality of nozzle segments, each nozzle segment comprising an endwall ring portion with at least one vane, the at least one vane having a free end portion;connecting the free end portion of the at least one vane of each nozzle segment of the plurality of nozzle segments to an annular endwall, with the endwall ring portion of each nozzle segment of the plurality of nozzle segments forming a segmented annular endwall concentric to the annular endwall with the at least one vane of each nozzle segment extending between the segmented annular endwall and the annular endwall; andforming the annular endwall prior to the connecting by separately casting the annular endwall as one-piece,wherein the step of connecting the plurality of nozzle segments comprises brazing the free end portion of the at least one vane of each nozzle segment to the annular endwall.2. The method of claim 1 , wherein the step of forming a plurality of nozzle segments comprises forming the plurality of nozzle segments with a single crystal material.3. The method of claim 1 , wherein the step of forming a plurality of nozzle segments comprises forming by casting.4. The method of claim 1 , wherein the step of forming a plurality of nozzle segments comprises forming a plurality of singlet nozzle segments claim 1 , doublet nozzle segments claim 1 , triplet nozzle segments claim 1 , quadruplet nozzle segments claim 1 , or combinations thereof.5. The method of claim 1 , further comprising the step of ...

Airfoil with leading edge convective cooling system

An airfoil includes a leading edge and an opposing trailing edge. The airfoil includes a pressure sidewall and an opposing suction sidewall. A leading edge cavity is defined between the pressure sidewall and the suction sidewall. The leading edge cavity has a first end opposite the leading edge and a second end defined at a rib. The airfoil includes at least one pin structure defined in the leading edge cavity between the first end and the second end. The at least one pin structure includes a main body and a first branch. The main body is coupled to the second end and extends toward the first end. The first branch extends from the main body toward the first end.

AXIALLY-SPLIT RADIAL TURBINES AND METHODS FOR THE MANUFACTURE THEREOF

Embodiments of an axially-split radial turbine, as are embodiments of a method for manufacturing an axially-split radial turbine. In one embodiment, the method includes the steps of joining a forward bladed ring to a forward disk to produce a forward turbine rotor, fabricating an aft turbine rotor, and disposing the forward turbine rotor and the aft turbine rotor in an axially-abutting, rotationally-fixed relationship to produce the axially-split radial turbine.

Methods for manufacturing a turbine nozzle with single crystal alloy nozzle segments

Methods for manufacturing a turbine nozzle are provided. A plurality of nozzle segments is formed. Each nozzle segment comprises an endwall ring portion with at least one vane. The plurality of nozzle segments are connected to an annular endwall forming a segmented annular endwall concentric to the annular endwall with the at least one vane of each nozzle segment extending between the segmented annular endwall and the annular endwall.

Cooling circuit with shaped cooling pins

A cooling circuit to receive a cooling fluid includes at least one shaped cooling pin disposed in the cooling circuit. The at least one shaped cooling pin has a first end and a second end extending along an axis. The first end has a first curved surface defined by a minor diameter and the second end has a second curved surface defined by a major diameter. The first curved surface is to be upstream in the cooling fluid and the minor diameter is less than the major diameter.

TURBINE SCROLL ASSEMBLY FOR GAS TURBINE ENGINE

A gas turbine engine includes a compressor section and a combustion section with a scroll, a scroll baffle, a combustor, and a combustor case. The scroll defines an interior scroll flow path. The scroll baffle surrounds the scroll to define a scroll cooling passage. The combustor case surrounds the combustor and the scroll baffle to define a collector space. Moreover, the engine includes a turbine section with a turbine rotor and a turbine rotor blade shroud that includes a shroud cooling passage. The compressor flow path is fluidly connected to the scroll for cooling the scroll. Also, the scroll cooling passage is fluidly connected to the shroud cooling passage for cooling the turbine rotor blade shroud. Furthermore, the shroud cooling passage is fluidly connected to the collector space. Flow from the collector space flows into the combustor, along the interior scroll flow path, toward the turbine rotor. 1. A gas turbine engine comprising:a compressor section that defines a compressor flow path;a combustion section that includes a scroll, a scroll baffle, a combustor, and a combustor case, the scroll defining an interior scroll flow path, the scroll baffle surrounding at least part of the scroll to define a scroll cooling passage between the scroll baffle and the scroll, the combustor case surrounding the combustor and at least part of the scroll baffle to define a collector space between the combustor case and the scroll baffle; anda turbine section with a turbine rotor and a turbine rotor blade shroud, the turbine rotor blade shroud including a shroud cooling passage;the compressor section being coupled to the combustion section with the compressor flow path fluidly connected to the scroll cooling passage to direct flow from the compressor flow path along the scroll cooling passage to cool the scroll;the combustion section being coupled to the turbine section with the scroll cooling passage fluidly connected to the shroud cooling passage to direct flow from the ...

GAS TURBINE ENGINE COMPONENTS HAVING SEALED STRESS RELIEF SLOTS AND METHODS FOR THE FABRICATION THEREOF

Embodiments of a gas turbine engine component having sealed stress relief slots are provided, as are embodiments of a gas turbine engine containing such a component and embodiments of a method for fabricating such a component. In one embodiment, the gas turbine engine includes a core gas flow path, a secondary cooling flow path, and a turbine nozzle or other gas turbine engine component. The component includes, in turn, a component body through which the core gas flow path extends, a radially-extending wall projecting from the component body and into the secondary cooling flow path, and one or more stress relief slots formed in the radially-extending wall. The stress relief slots are filled with a high temperature sealing material, which impedes leakage between the second cooling and core gas flow paths and which fractures to alleviate thermomechanical stress within the radially-extending wall during operation of the gas turbine engine. 1. A gas turbine engine , comprising:a core gas flow path;a secondary cooling flow path; and a component body through which the core gas flow path extends;', 'a radially-extending wall projecting from the component body into the secondary cooling flow path;', 'one or more stress relief slots formed in the radially-extending wall; and', 'a high temperature sealing material filling the one or more stress relief slots and impeding leakage between the secondary cooling flow path and the core gas flow path, the high temperature sealing material fracturing to alleviate thermomechanical stress within the radially-extending wall during operation of the gas turbine engine., 'a gas turbine engine component, comprising2. The gas turbine engine of wherein the gas turbine engine component comprises a turbine nozzle claim 1 , and wherein the component body comprises:an inner endwall;an outer endwall circumscribing the inner endwall; anda plurality of circumferentially-spaced vanes extending between the inner and outer endwalls.3. The gas turbine ...

AXIALLY-SPLIT RADIAL TURBINES

Embodiments of an axially-split radial turbine, as are embodiments of a method for manufacturing an axially-split radial turbine. In one embodiment, the method includes the steps of joining a forward bladed ring to a forward disk to produce a forward turbine rotor, fabricating an aft turbine rotor, and disposing the forward turbine rotor and the aft turbine rotor in an axially-abutting, rotationally-fixed relationship to produce the axially-split radial turbine. 1. An axially-split radial turbine , comprising: a forward disk; and', 'a forward bladed ring circumscribing and metallurgically bonded to the forward disk; and, 'a forward turbine rotor, comprisingan aft turbine rotor disposed axially adjacent to and rotationally fixed relative to the forward turbine rotor.2. The axially-split radial turbine of wherein the forward bladed ring comprises a plurality of forward bladed pieces metallurgically-consolidated into the forward bladed ring.3. The axially-split radial turbine of further comprising:a plurality of blade cooling passages extending in the plurality of forward bladed pieces; andan inner disk cavity formed within the axially-split radial turbine and in fluid communication with the plurality of blade cooling passages.4. The axially-split radial turbine of wherein the plurality of forward bladed pieces are individually cast from a single crystal alloy claim 2 , and wherein the forward disk is composed of a non-single crystal alloy.5. The axially-split radial turbine of wherein the aft turbine rotor comprises:an aft bladed ring produced from a plurality of aft bladed pieces each including at least one aft blade segment; andan aft disk joined to the aft bladed ring.6. The axially-split radial turbine of further comprising:a plurality of blade cooling passages extending in the plurality of aft bladed pieces; andan inner disk cavity formed within the axially-split radial turbine and in fluid communication with the plurality of blade cooling passages.7. The axially ...

TURBINE ASSEMBLIES WITH IMPINGEMENT COOLING

A gas turbine engine assembly includes a housing including an annular duct wall that at least partially defines a mainstream hot gas flow path; a stator assembly with a stator vane extending into the mainstream gas flow; and a turbine rotor assembly upstream of the stator assembly and defining a turbine cavity with the stator assembly. The turbine rotor assembly includes a rotor disk having a forward side and an aft side, a rotor platform positioned on a periphery of the rotor disk, the rotor platform defining an aft flow discourager, a rotor blade mounted on the rotor platform extending into the mainstream gas flow, and an aft seal plate mounted on the aft side of the rotor disk. The aft seal plate has a radius such that the aft seal plate protects the rotor disk from hot gas ingestion.

HIGH TEMPERATURE THERMAL PROCESS SYSTEMS

A thermal process system includes a retort assembly, a heating assembly, and a vessel housing. The retort assembly includes a retort chamber and is configured to substantially form a containment boundary to contain one or more gases in the retort chamber during a thermal process. The heating assembly includes one or more heating elements and is configured to heat the retort chamber. The vessel housing is positioned around the retort chamber and the one or more heating elements and configured to form a pressure boundary to maintain a pressure within the retort chamber and reduce a pressure across the retort chamber.

GAS TURBINE ENGINES WITH ABRADABLE TURBINE SEAL ASSEMBLIES

A turbine section of a gas turbine engine includes a housing, a rotor assembly, and a seal assembly. The rotor assembly includes a rotor disk, a rotor platform coupled to the rotor disk, and a rotor blade extending from the rotor platform into the mainstream hot gas flow path. The stator assembly includes a stator platform with a stator vane that extends from the stator platform into the mainstream hot gas flow path. The seal assembly includes a first flow discourager extending in a first direction from the rotor platform, a second flow discourager extending in a second direction from the stator platform, the first flow discourager axially overlapping the second flow discourager such that the second flow discourager is interior to the first flow discourager in a radial direction, a hard coating applied to the first flow discourager, and an abradable coating applied to the second flow discourager. 1. A turbine section of a gas turbine engine , comprising:a housing including an annular duct wall that at least partially defines a mainstream hot gas flow path configured to receive mainstream hot gas flow;a rotor assembly including a rotor disk, a rotor platform coupled to the rotor disk, and a rotor blade extending from the rotor platform into the mainstream hot gas flow path;a stator assembly positioned adjacent to the rotor assembly and forming a turbine disk cavity with the rotor disk of the rotor assembly, the stator assembly including a stator platform with a stator vane that extends from the stator platform into the mainstream hot gas flow path; and a first flow discourager extending in a first direction from the rotor platform,', 'a second flow discourager extending in a second direction from the stator platform, the first flow discourager axially overlapping the second flow discourager such that the second flow discourager is interior to the first flow discourager in a radial direction,', 'a hard coating applied to the first flow discourager, and', 'an abradable ...

METHOD OF PRODUCING AN ABRASIVE TIP FOR A TURBINE BLADE

A method of producing an abrasive tip for a turbine blade includes producing or obtaining a metal powder that is mixed with an abrasive ceramic powder and producing or obtaining a metallic mold that is in the shape of an airfoil. The metallic mold includes a hollow interior portion. The method further includes sealing the metal and ceramic powder mixture within the hollow interior portion of the metallic mold under vacuum and subjecting the sealed mold to a hot isostatic pressing process. The hot isostatic pressing process compacts and binds the metal and ceramic powder mixture together into a solid article in the shape of the airfoil. Still further, the method includes slicing the solid article into a plurality of airfoil-shaped slices and bonding one slice of the plurality of airfoil-shaped slices to a tip portion of a turbine blade.

COMPOSITE TURBINE DISC ROTOR FOR TURBOMACHINE

A rotor for a turbomachine and a method of manufacturing the same. The method includes providing a lug with a lug body and an interface material disposed on the lug body. The method also includes friction welding the lug to a hub member via the interface material to define a projected structure for an outer radial area of a disc assembly of the rotor. The projected structure is configured to support a first side of a rotor blade of the rotor in cooperation with a second projected structure of the disc assembly supporting a second side of the rotor blade. The lug body and the hub member are made from different materials.

GAS TURBINE ENGINE ASSEMBLIES WITH RECIRCULATED HOT GAS INGESTION

A gas turbine engine assembly includes a housing including an annular duct wall that at least partially defines a mainstream hot gas flow path configured to receive mainstream hot gas flow. The assembly further includes a stator assembly including a stator vane that extends into the mainstream hot gas flow path and a turbine rotor assembly downstream of the stator assembly that includes a turbine disk and a turbine blade extending from the turbine disk into the mainstream hot gas flow path. The stator assembly and turbine assembly define a turbine disk cavity, and the turbine disk cavity includes a recirculation cavity configured to recirculate gas ingested from the mainstream hot gas flow path back into the mainstream hot gas flow path.

Radial turbine rotor for gas turbine engine

A radial turbine rotor associated with an engine includes a disk, and a plurality of blades spaced apart about a perimeter of the disk. Each blade includes a forward end, an aft end and a root. The radial turbine rotor includes a plurality of sectors, with each sector coupled to the root of a respective blade of the plurality of blades. Each sector of the plurality of sectors defines a first surface configured to contact a working fluid and a second surface configured to be coupled to the disk, and each sector of the plurality of sectors defines at least one pocket between the first surface and the second surface proximate the forward end that extends toward the aft end. The radial turbine rotor includes a feather seal slot defined between adjacent sectors of the plurality of sectors proximate the first surface.

Gas turbine engine assemblies with recirculated hot gas ingestion

A gas turbine engine assembly includes a housing including an annular duct wall that at least partially defines a mainstream hot gas flow path configured to receive mainstream hot gas flow. The assembly further includes a stator assembly including a stator vane that extends into the mainstream hot gas flow path and a turbine rotor assembly downstream of the stator assembly that includes a turbine disk and a turbine blade extending from the turbine disk into the mainstream hot gas flow path. The stator assembly and turbine assembly define a turbine disk cavity, and the turbine disk cavity includes a recirculation cavity configured to recirculate gas ingested from the mainstream hot gas flow path back into the mainstream hot gas flow path.

IMPINGEMENT BAFFLE FOR GAS TURBINE ENGINE

An impingement baffle for directing a cooling fluid onto a target surface includes a baffle body having a first end opposite a second end, and a first side opposite a second side. The second side is spaced a distance apart from the target surface, with the distance varying from the first end to the second end. The baffle body defines impingement holes that extend through the baffle body from the first side to the second side. The impingement holes are spaced apart along the baffle body to receive the cooling fluid. The impingement baffle includes tubular extensions coupled to the second side. Each tubular extension is in fluid communication with a respective one of the impingement holes to direct the cooling fluid onto the target surface. Each tubular extension extends for a length from the second side, and the length of each tubular extension is based on the distance.

METHODS FOR MANUFACTURING A TURBINE NOZZLE WITH SINGLE CRYSTAL ALLOY NOZZLE SEGMENTS

Methods for manufacturing a turbine nozzle are provided. A plurality of nozzle segments is formed. Each nozzle segment comprises an endwall ring portion with at least one vane. The plurality of nozzle segments are connected to an annular endwall forming a segmented annular endwall concentric to the annular endwall with the at least one vane of each nozzle segment extending between the segmented annular endwall and the annular endwall. 1. A method for manufacturing a turbine nozzle comprising:forming a plurality of nozzle segments, each nozzle segment comprising an endwall ring portion with at least one vane; andconnecting the plurality of nozzle segments to an annular endwall forming a segmented annular endwall concentric to the annular endwall with the at least one vane of each nozzle segment extending between the segmented annular endwall and the annular endwall.2. The method of claim 1 , wherein the step of forming a plurality of nozzle segments comprises forming the plurality of nozzle segments with a single crystal material.3. The method of claim 1 , wherein the step of forming a plurality of nozzle segments comprises forming by casting.4. The method of claim 1 , wherein the step of forming a plurality of nozzle segments comprises forming a plurality of singlet nozzle segments claim 1 , doublet nozzle segments claim 1 , triplet nozzle segments claim 1 , quadruplet nozzle segments claim 1 , or combinations thereof.5. The method of claim 1 , further comprising the step of processing at least one nozzle segment of the plurality of nozzle segments prior to the connecting step claim 1 , wherein the step of processing comprises applying a protective coating to at least one nozzle segment of the plurality of nozzle segments claim 1 , the annular endwall claim 1 , or both.6. The method of claim 1 , further comprising the step of processing at least one nozzle segment of the plurality of nozzle segments prior to the connecting step claim 1 , wherein the step of ...

TURBINE TIP SHROUD ASSEMBLY WITH PLURAL SHROUD SEGMENTS HAVING INTERNAL COOLING PASSAGES

A shroud assembly for a gas turbine engine includes a shroud support and a plurality of shroud segments that are attached to the shroud support. The shroud segment includes an internal cooling passage. 1. A shroud assembly for a gas turbine engine , the shroud assembly configured to receive a cooling fluid flow , the shroud assembly comprising:a shroud support that extends arcuately about an axis; and an internal cooling passage that extends through the at least one of the plurality of shroud segments;', 'at least one inlet for receiving and directing the cooling fluid flow into the internal cooling passage; and', 'at least one outlet for outputting the cooling fluid flow from the internal cooling passage to a backflow cavity of the shroud assembly;, 'a plurality of shroud segments that are attached to the shroud support and that are arranged annularly about the axis at different circumferential positions with respect to the axis, at least one of the plurality of shroud segments includingthe internal cooling passage being substantially hermetically sealed from the at least one inlet to the at least one outlet.2. The shroud assembly of claim 1 , wherein the internal cooling passage includes a first chamber and a second chamber claim 1 , the internal cooling passage including a first impingement aperture directed into the first chamber claim 1 , the internal cooling passage including a second impingement aperture directed into the second chamber claim 1 , the internal cooling passage configured to direct the cooling fluid flow from the first impingement aperture into the first chamber and downstream into the second chamber via the second impingement aperture.3. The shroud assembly of claim 2 , wherein the shroud segment includes an inner diameter surface configured to oppose a turbine blade as the turbine blade rotates about the axis;wherein at least one of the first chamber and the second chamber is partly defined by a backside surface that is disposed opposite the ...

TURBINE ROTOR SEAL PLATE WITH INTEGRAL FLOW DISCOURAGER

A seal plate for a turbine engine is provided. The turbine engine has a central axis. The seal plate comprises a forward face extending radially from the central axis. The forward face comprises a flat portion extending circumferentially around the central axis, and a front flow discourager coupled to the flat portion and extending outward from the flat portion.

SHROUDED BONDED TURBINE ROTORS AND METHODS FOR MANUFACTURING THE SAME

Methods are provided for manufacturing a shrouded bonded turbine rotor. A shrouded blade ring is formed. The shrouded blade ring is formed by bonding a unitary shroud ring to an assembled blade ring or assembling a plurality of shrouded turbine blade segments. The shrouded blade ring is bonded to a hub. The shrouded bonded turbine rotors are also provided. The shrouded bonded turbine rotor comprises a shrouded blade ring and a shroud. The shrouded blade ring comprises a plurality of turbine blade segments and a shroud. Each turbine blade segment comprises an airfoil portion including an airfoil having a root and a tip. The shroud covers the tip of each airfoil in the shrouded blade ring. A hub is bonded with the shrouded blade ring. 1. A method for manufacturing a shrouded bonded turbine rotor , the method comprising the steps of:forming a shrouded blade ring comprising bonding a unitary shroud ring to an assembled blade ring or assembling a plurality of shrouded turbine blade segments; andbonding the shrouded blade ring to a hub.2. The method of claim 1 , wherein the step of forming a shrouded blade ring comprises bonding the unitary shroud ring to the assembled blade ring and the method further comprises the step of assembling a plurality of turbine blade segments into the assembled blade ring prior to forming the shrouded blade ring.3. The method of claim 1 , wherein the step of forming a shrouded blade ring comprising bonding the unitary shroud ring to the assembled blade ring comprises diffusion bonding the unitary shroud ring to the assembled blade ring.4. The method of claim 2 , wherein the step of assembling a plurality of turbine blade segments and bonding the unitary shroud ring to the assembled blade ring are performed substantially simultaneously.5. The method of claim 4 , wherein the step of assembling a plurality of turbine blade segments and bonding the unitary shroud ring to the assembled blade ring are performed by compression in a single heating ...

TURBINE NOZZLE WITH REDUCED LEAKAGE FEATHER SEALS

A turbine nozzle for a gas turbine engine includes a plurality of nozzle segments that are configured to be assembled into a full ring such that each one of the plurality of nozzle segments is adjacent to another one of the plurality of nozzle segments. Each one of the plurality of nozzle segments includes an endwall segment and a nozzle vane. The turbine nozzle includes a feather seal interface defined by endwall segments of adjacent ones of the plurality of nozzle segments. The feather seal interface is defined along an area of reduced pressure drop through a pressure field defined between adjacent nozzle vanes of the plurality of nozzle segments to reduce leakage through the plurality of nozzle segments. The turbine nozzle includes a feather seal received within the feather seal interface that cooperates with the feather seal interface to reduce leakage through the plurality of nozzle segments. 1. A turbine nozzle for a gas turbine engine , comprising:a plurality of nozzle segments that are configured to be assembled into a full ring such that each one of the plurality of nozzle segments is adjacent to another one of the plurality of nozzle segments, each one of the plurality of nozzle segments including an endwall segment and a nozzle vane;a feather seal interface defined by endwall segments of adjacent ones of the plurality of nozzle segments, the feather seal interface defined along an area of reduced pressure drop through a pressure field defined between adjacent nozzle vanes of the plurality of nozzle segments; anda feather seal received within the feather seal interface that cooperates with the feather seal interface to reduce leakage through the plurality of nozzle segments.2. The turbine nozzle of claim 1 , wherein the area of reduced pressure drop is proximate a pressure side of one of the adjacent nozzle vanes such that the feather seal interface is defined proximate the pressure side.3. The turbine nozzle of claim 2 , wherein the feather seal interface ...

TURBINE SHROUD SUPPORT COUPLING ASSEMBLY

A coupling assembly for a turbine shroud is provided. The coupling assembly comprises a rotatable positioning block having a first surface, and a biasing spring having a second surface, the second surface generally facing the first surface, and the biasing spring adapted to exert a force toward the positioning block when compressed.

TURBINE COMPONENT WITH SHAPED COOLING PINS

A cooling circuit to receive a cooling fluid includes at least one shaped cooling pin disposed in the cooling circuit. The at least one shaped cooling pin has a first end and a second end extending along an axis. The first end has a first curved surface defined by a minor diameter and the second end has a second curved surface defined by a major diameter. The first curved surface is to be upstream in the cooling fluid and the minor diameter is less than the major diameter. 1. A cooling circuit adapted to receive a cooling fluid , comprising:at least one shaped cooling pin disposed in the cooling circuit, the at least one shaped cooling pin having a first end and a second end extending along an axis, the first end having a first curved surface defined by a minor diameter and the second end having a second curved surface defined by a major diameter, the first curved surface is configured to be upstream in the cooling fluid and the minor diameter is less than the major diameter.2. The cooling circuit of claim 1 , wherein the first curved surface is spaced apart from the second curved surface by a length.3. The cooling circuit of claim 1 , wherein the first curved surface and the second curved surface are interconnected by a pair of surfaces defined by a pair of planes substantially tangent to a respective one of the first curved surface and the second curved surface.4. The cooling circuit of claim 2 , wherein claim 2 , in cross-section claim 2 , the second curved surface is defined by a first circle having a first center point and the first curved surface is defined by a second circle having a second center point claim 2 , and the length is defined between the first center point and the second center point.5. The cooling circuit of claim 4 , wherein the at least one shaped cooling pin comprises a plurality of shaped cooling pins that are arranged in a pattern that includes at least one row of a first sub-plurality of the plurality of shaped cooling pins and at least ...

Composite turbine disc rotor for turbomachine

A rotor for a turbomachine and a method of manufacturing the same. The method includes providing a lug with a lug body and an interface material disposed on the lug body. The method also includes friction welding the lug to a hub member via the interface material to define a projected structure for an outer radial area of a disc assembly of the rotor. The projected structure is configured to support a first side of a rotor blade of the rotor in cooperation with a second projected structure of the disc assembly supporting a second side of the rotor blade. The lug body and the hub member are made from different materials.

Turbine blade platform

A turbine blade assembly is provided. The turbine blade assembly comprises a turbine blade comprising a cavity, and a blade platform supporting the turbine blade, the cavity extending into the blade platform. The blade platform comprises an upper surface adjacent the turbine blade and a lower surface comprising a first rib, the cavity extending into the first rib, the first rib coupled to the lower surface, tapering as it extends away from the turbine blade, and comprising a first port extending from the cavity to the upper surface.

DIRECTED COOLING FOR ROTATING MACHINERY

A rotating machine includes a hub portion, wherein the hub portion comprises a forward face and an aft face. The rotating machine further includes a cooling channel formed on either the forward face or the aft face and configured to direct cooling air to a location on the rotating machine, wherein the cooling channel extends from a radially inner location along said face to a radially outer location along said face, and wherein the cooling channel is configured as a recess formed into an outer surface of said face. 1. A rotating machine comprising:a hub portion, wherein the hub portion comprises an outer circumference, an inner circumference, a forward face, and an aft face;a ring portion, wherein the ring portion comprises an inner circumference that is metallurgically bonded to the outer circumference of the hub portion along a circumferential bond line; anda cooling channel formed on either the forward face or the aft face and configured to direct cooling air to the bond line, wherein the cooling channel extends from a radially inner location along said face to a radially outer location along said face, and wherein the cooling channel is configured as a recess formed into an outer surface of said face.2. The rotating machine of claim 1 , wherein the cooling channel comprises a radially inner portion defined from the radially inner location along said face to a mid-point location along said face claim 1 , and a radially outer portion defined from the mid-point location along said face to the radially outer location along said face claim 1 , wherein the radially inner portion is angled against a direction of rotation of the rotating machine as the cooling channel extends radially outward.3. The rotating machine of claim 1 , wherein the ring portion comprises a forward end and an aft end claim 1 , wherein either the forward end or the aft end comprises a circumferential flange extending outward from said end and positioned radially above the bond line claim 1 , ...

BLADED GAS TURBINE ENGINE ROTORS HAVING DEPOSITED TRANSITION RINGS AND METHODS FOR THE MANUFACTURE THEREOF

Bladed Gas Turbine Engine (GTE) rotors including deposited transition rings are provided, as are embodiments of methods for manufacturing bladed GTE rotors. In one embodiment, the method includes providing an outer blade ring having an inner circumferential surface defining a central opening, and depositing a deposited transition ring on the inner circumferential surface of the outer blade ring. The outer blade ring can be a full bladed ring or an annular grouping of individually-fabricated bladed pieces. After deposition of the transition ring, a hub disk is inserted into the central opening such that the transition ring extends around an outer circumferential surface of the hub disk. The transition ring is then bonded to the outer circumferential surface of the hub disk utilizing, for example, a hot isostatic pressing technique to join the transition ring and the outer blade ring thereto.

GAS TURBINE ENGINES WITH TURBINE ROTOR BLADES HAVING IMPROVED PLATFORM EDGES

A turbine rotor blade is provided. The turbine rotor blade includes a root, a platform coupled to the root, and an airfoil extending from the platform. The platform has a leading edge, a trailing edge, a suction side edge, and a pressure side edge. The pressure side edge includes a first concave portion. 1. A turbine rotor blade , comprising:a root;a platform coupled to the root, the platform having a leading edge, a trailing edge, a suction side edge, and a pressure side edge, wherein the pressure side edge includes a first concave portion; andan airfoil extending from the platform.2. The turbine rotor blade of claim 1 , wherein the first concave portion is concave relative to an axial direction.3. The turbine rotor blade of claim 1 , wherein the suction side edge is parallel to the pressure side edge.4. The turbine rotor blade of claim 1 , wherein the airfoil has a pressure side wall and a suction side wall claim 1 , and wherein the pressure side edge of the platform is non-parallel to the pressure side wall of the airfoil.5. The turbine rotor blade of claim 1 , wherein the pressure side edge further includes a first convex portion.6. The turbine rotor blade of claim 1 , wherein the pressure side edge is continuously curved from the leading edge to the trailing edge.7. A method for producing a turbine rotor blade claim 1 , comprising the following steps:performing a plastic analysis of a blade platform of a baseline turbine rotor blade with a first side edge and a second side edge;identifying a first area with plasticity greater than a predetermined limit on the first side edge;modifying the baseline turbine rotor blade to result in an intermediate turbine rotor blade by removing the first area from the first side edge and adding a second area, corresponding in size and shape to the first area, to the second side edge; andfinishing the intermediate turbine rotor blade as a final turbine rotor blade.8. The method of claim 7 , further comprising claim 7 , after the ...

TURBINE NOZZLES AND METHODS OF MANUFACTURING THE SAME

Turbine nozzles and methods of manufacturing the turbine nozzles are provided. In an embodiment, by way of example only, a turbine nozzle includes a first ring, a vane, and a first joint. The first ring comprises a single unitary component and having a first opening and including a first metal alloy. The vane includes a first end disposed in the first opening and includes a second metal alloy. The first joint is formed in the first opening between the first ring and the vane and includes a first braze layer and an oxide layer. The first braze layer is disposed adjacent to the oxide layer, and the first braze layer and the oxide layer are disposed between the first ring and the vane.

TURBINE TIP SHROUD ASSEMBLY WITH PLURAL SHROUD SEGMENTS HAVING INTERNAL COOLING PASSAGES

A shroud assembly for a gas turbine engine includes a shroud support and a plurality of shroud segments that are attached to the shroud support. The shroud segment includes an internal cooling passage.

Gas turbine engines with abradable turbine seal assemblies

A turbine section of a gas turbine engine includes a housing, a rotor assembly, and a seal assembly. The rotor assembly includes a rotor disk, a rotor platform coupled to the rotor disk, and a rotor blade extending from the rotor platform into the mainstream hot gas flow path. The stator assembly includes a stator platform with a stator vane that extends from the stator platform into the mainstream hot gas flow path. The seal assembly includes a first flow discourager extending in a first direction from the rotor platform, a second flow discourager extending in a second direction from the stator platform, the first flow discourager axially overlapping the second flow discourager such that the second flow discourager is interior to the first flow discourager in a radial direction, a hard coating applied to the first flow discourager, and an abradable coating applied to the second flow discourager.

Turbine tip shroud assembly with plural shroud segments having inter-segment seal arrangement

A shroud assembly for a gas turbine engine includes a plurality of shroud segments that are attached to a shroud support with an inter-segment joint defined between shroud segments. The shroud assembly also includes a cooling flow path cooperatively defined by the shroud support and the first shroud segment. The cooling flow path includes an internal cooling passage within the shroud segments. The cooling flow path includes an outlet chamber configured to receive flow from the internal cooling passage. The shroud assembly additionally includes a seal arrangement that extends across the inter-segment joint. The seal arrangement, the first shroud segment, and the second shroud segment cooperatively define a seal chamber that is enclosed.

Gas turbine engine components having sealed stress relief slots and methods for the fabrication thereof

Embodiments of a gas turbine engine component having sealed stress relief slots are provided, as are embodiments of a gas turbine engine containing such a component and embodiments of a method for fabricating such a component. In one embodiment, the gas turbine engine includes a core gas flow path, a secondary cooling flow path, and a turbine nozzle or other gas turbine engine component. The component includes, in turn, a component body through which the core gas flow path extends, a radially-extending wall projecting from the component body and into the secondary cooling flow path, and one or more stress relief slots formed in the radially-extending wall. The stress relief slots are filled with a high temperature sealing material, which impedes leakage between the second cooling and core gas flow paths and which fractures to alleviate thermomechanical stress within the radially-extending wall during operation of the gas turbine engine.

Methods for processing bonded dual alloy rotors including differential heat treatment processes

Methods for processing bonded dual alloy rotors are provided. In one embodiment, the method includes obtaining a bonded dual alloy rotor including rotor blades bonded to a hub disk. The rotor blades and hub disk are composed of different alloys. A minimum processing temperature (T DISK _ PROCESS _ MIN ) for the hub disk and a maximum critical temperature for the rotor blades (T BLADE _ MAX ) is established such that T BLADE _ MAX is less than T DISK _ PROCESS _ MIN . A differential heat treatment process is then performed during which the hub disk is heated to processing temperatures equal to or greater than T DISK _ PROCESS _ MIN , while at least a volumetric majority of each of the rotor blades is maintained at temperatures below T BLADE _ MAX . Such a targeted differential heat treatment process enables desired metallurgical properties (e.g., precipitate hardening) to be created within the hub disk, while preserving the high temperature properties of the rotor blades and any blade coating present thereon.

DUAL ALLOY GAS TURBINE ENGINE ROTORS AND METHODS FOR THE MANUFACTURE THEREOF

Dual alloy Gas Turbine Engine (GTE) rotors and method for producing GTE rotors are provided. In one embodiment, the method include includes arranging bladed pieces in an annular grouping or ring formation such that shank-to-shank junctions are formed between circumferentially-adjacent bladed pieces. A first or bonding alloy is deposited along the shank-to-shank junctions utilizing a localized fusion deposition process to produce a plurality of alloy-filled joints, which join the bladed pieces in a bonded blade ring. The bonding alloy is preferably selected to have a ductility higher than and a melt point lower than the alloy from which the bladed pieces are produced. After deposition of the first alloy and formation of the alloy-filled joints, a hub disk is inserted into the central opening of the bonded blade ring. The hub disk and blade ring are then bonded utilizing, for example, a Hot Isostatic Pressing process. 1. A method for manufacturing a bladed Gas Turbine Engine (GTE) rotor , the method comprising:arranging bladed pieces in a ring formation such that shank-to-shank junctions are formed between circumferentially-adjacent bladed pieces;depositing a first alloy along the shank-to-shank junctions utilizing a localized fusion deposition process to produce a plurality of alloy-filled joints bonding the bladed pieces in a bonded blade ring having a central opening;inserting a hub disk into the central opening of the bonded blade ring; andbonding the bonded blade ring to the hub disk.2. The method of wherein the plurality of alloy-filled joints are spaced around a circumference of the bonded blade ring.3. The method of wherein the dual alloy GTE rotor comprises an axial turbine rotor claim 2 , and wherein the alloy-filled joints are spaced around an inner circumference of the bonded blade ring.4. The method of further comprising maintaining the bladed pieces in the ring formation utilizing tooling while depositing the first alloy along the shank-to-shank ...

METHODS FOR PROCESSING BONDED DUAL ALLOY ROTORS INCLUDING DIFFERENTIAL HEAT TREATMENT PROCESSES

Methods for processing bonded dual alloy rotors are provided. In one embodiment, the method includes obtaining a bonded dual alloy rotor including rotor blades bonded to a hub disk. The rotor blades and hub disk are composed of different alloys. A minimum processing temperature (T) for the hub disk and a maximum critical temperature for the rotor blades (T) is established such that Tis less than T. A differential heat treatment process is then performed during which the hub disk is heated to processing temperatures equal to or greater than T, while at least a volumetric majority of each of the rotor blades is maintained at temperatures below T. Such a targeted differential heat treatment process enables desired metallurgical properties (e.g., precipitate hardening) to be created within the hub disk, while preserving the high temperature properties of the rotor blades and any blade coating present thereon.

Bladed gas turbine engine rotors having deposited transition rings and methods for the manufacture thereof

Bladed Gas Turbine Engine (GTE) rotors including deposited transition rings are provided, as are embodiments of methods for manufacturing bladed GTE rotors. In one embodiment, the method includes providing an outer blade ring having an inner circumferential surface defining a central opening, and depositing a deposited transition ring on the inner circumferential surface of the outer blade ring. The outer blade ring can be a full bladed ring or an annular grouping of individually-fabricated bladed pieces. After deposition of the transition ring, a hub disk is inserted into the central opening such that the transition ring extends around an outer circumferential surface of the hub disk. The transition ring is then bonded to the outer circumferential surface of the hub disk utilizing, for example, a hot isostatic pressing technique to join the transition ring and the outer blade ring thereto.

TURBINE SHROUD ASSEMBLIES FOR GAS TURBINE ENGINES

A turbine shroud assembly includes a shroud support and a shroud segment. The shroud support structure includes a forward support rail and an aft support rail. The forward support rail includes forward first engagement structures and the aft support rail includes aft first engagement structures. The shroud segment includes a forward segment rail and an aft segment rail. The forward segment rail includes forward second engagement structures positioned on a forward segment rail periphery and the aft segment rail includes second engagement structures positioned on an aft segment rail periphery. The forward first engagement structures radially and circumferentially engage with the forward second engagement structures and the aft first engagement structures radially and circumferentially engage with the aft second engagement structures to radially and circumferentially interlock the shroud segment to the shroud support structure.

TURBINE NOZZLE ASSEMBLY INCLUDING RADIALLY-COMPLIANT SPRING MEMBER FOR GAS TURBINE ENGINE

Embodiments of a turbine nozzle assembly are provided for deployment within a gas turbine engine (GTE) including a first GTE-nozzle mounting interface. In one embodiment, the turbine nozzle assembly includes a turbine nozzle flowbody, a first mounting flange configured to be mounted to the first GTE-nozzle mounting interface, and a first radially-compliant spring member coupled between the turbine nozzle flowbody and the first mounting flange. The turbine nozzle flowbody has an inner nozzle endwall and an outer nozzle endwall, which is fixedly coupled to the inner nozzle endwall and which cooperates therewith to define a flow passage through the turbine nozzle flowbody. The first radially-compliant spring member accommodates relative thermal movement between the turbine nozzle flowbody and the first mounting flange to alleviate thermomechanical stress during operation of the GTE.

Turbine tip shroud assembly with plural shroud segments having internal cooling passages

A shroud assembly for a gas turbine engine includes a shroud support and a plurality of shroud segments that are attached to the shroud support. The shroud segment includes an internal cooling passage.

DUAL ALLOY TURBINE ROTORS AND METHODS FOR MANUFACTURING THE SAME

Dual alloy turbine rotors and methods for manufacturing the same are provided. The dual alloy turbine rotor comprises an assembled blade ring and a hub bonded to the assembled blade ring. The assembled blade ring comprises a first alloy selected from the group consisting of a single crystal alloy, a directionally solidified alloy, or an equi-axed alloy. The hub comprises a second alloy. The method comprises positioning a hub within a blade ring to define an interface between the hub and the blade ring. The interface is a non-contacting interface or a contacting interface. The interface is enclosed by a pair of diaphragms. The interface is vacuum sealed. The blade ring is bonded to the hub after the vacuum sealing step. 1. A method for manufacturing a dual alloy turbine rotor , the method comprising the steps of:positioning a hub within a blade ring to thereby define an interface between the hub and the blade ring, the interface comprising a non-contacting interface or a contacting interface;enclosing the interface with a pair of diaphragms;vacuum sealing the interface; andbonding the blade ring to the hub after the vacuum sealing step.2. The method of claim 1 , wherein:the blade ring comprises an assembled blade ring having an inner annular surface and a first coefficient of thermal expansion;the hub has an outer peripheral surface and a second coefficient of thermal expansion higher than the first coefficient of thermal expansion;wherein the step of positioning a hub within the blade ring comprises spacing an outer peripheral surface of the hub apart from the inner annular surface of the assembled blade ring to form the non-contacting interface defined by a gap; and wherein the step of vacuum sealing the interface reduces the gap as the assembled blade ring expands less than the hub.3. The method of claim 1 , wherein:the blade ring comprises an assembled blade ring having an inner annular surface and a first coefficient of thermal expansion;the hub has an outer ...

TURBINE NOZZLE WITH REDUCED LEAKAGE FEATHER SEALS

A turbine nozzle for a gas turbine engine includes a plurality of nozzle segments that are configured to be assembled into a full ring such that each one of the plurality of nozzle segments is adjacent to another one of the plurality of nozzle segments. Each one of the plurality of nozzle segments includes an endwall segment and a nozzle vane. The turbine nozzle includes a feather seal interface defined by endwall segments of adjacent ones of the plurality of nozzle segments. The feather seal interface is defined along an area of reduced pressure drop through a pressure field defined between adjacent nozzle vanes of the plurality of nozzle segments to reduce leakage through the plurality of nozzle segments. The turbine nozzle includes a feather seal received within the feather seal interface that cooperates with the feather seal interface to reduce leakage through the plurality of nozzle segments.

Direct transfer axial tangential onboard injector system (TOBI) with self-supporting seal plate

An apparatus for cooling turbine blades in a turbine engine including a direct transfer axial tangential onboard injector (TOBI) for a turbine rotor and a self-supporting seal plate disposed on a rotating disk for the turbine engine. The TOBI includes a plurality of openings emanating a flow of cooling air. The self-supporting seal plate comprises a plurality of shaped cooling holes in fluid communication with the flow of cooling air emanating from the TOBI. The rotating disk includes a plurality of turbine blade slots formed therein. The plurality of cooling holes are in fluid communication with the plurality of turbine blade slots for directing the flow of cooling air to provide cooling to the plurality of turbine blades. The plurality of openings, the plurality of cooling holes and the plurality of turbine blade slots are in axial alignment and optimized to minimize radial and hoop stresses.

DIVERGING-CONVERGING COOLING PASSAGE FOR A TURBINE BLADE

A turbine blade and a radial turbine having at least one blade is provided. The turbine blade includes a trailing edge and a leading edge opposite the trailing edge. The turbine blade also includes a cooling passage defined internally within the turbine blade. The cooling passage is in fluid communication with a source of cooling fluid via a single inlet to receive a cooling fluid. The cooling passage diverges at a first point downstream from the single inlet into at least two branches that extend along the at least one blade from the first point to a second point near a tip of the leading edge and the cooling passage converges at the second point. 1. A turbine blade , comprising:a trailing edge;a leading edge opposite the trailing edge; anda cooling passage defined internally within the turbine blade, the cooling passage in fluid communication with a source of cooling fluid via a single inlet to receive a cooling fluid, and the cooling passage diverges at a first point downstream from the single inlet into at least two branches that extend along the at least one blade from the first point to a second point near a tip of the leading edge and the cooling passage converges at the second point.2. The turbine blade of claim 1 , wherein the turbine blade has a pressure side opposite a suction side claim 1 , the pressure side and the suction side extending from a first surface to a second surface claim 1 , and a respective one of the at least two branches extends along the pressure side adjacent to the first surface and the other of the at least two branches extends along the suction side adjacent to the first surface.3. The turbine blade of claim 1 , further comprising a cross flow path that fluidly interconnects the at least two branches at the leading edge.4. The turbine blade of claim 1 , wherein the cooling passage includes a tip flow passage defined at the tip of the leading edge that is fluidly coupled to the cooling passage near the second point.5. The turbine ...

Gas turbine engines with turbine rotor blades having improved platform edges

A turbine rotor blade is provided. The turbine rotor blade includes a root, a platform coupled to the root, and an airfoil extending from the platform. The platform has a leading edge, a trailing edge, a suction side edge, and a pressure side edge. The pressure side edge includes a first concave portion.

Turbomachine compressor scroll with load-carrying inlet vanes

A compressor scroll housing for use in conjunction with turbo-machinery, particularly applicable in aircraft. The scroll housing can include a plurality of scroll vanes arrayed around the scroll housing. Scroll vanes, integrally formed with the scroll housing, carry stress load on the scroll housing, including the load from fluid pressure within the scroll and carcass loading from the engine. The plurality of scroll vanes adapted for guiding flow of fluid from an inlet to an outlet while supporting the scroll housing. Chord length and cross sectional area of each scroll vane can be sized to maintain an equal stress in all scroll vanes. A method of making the scroll housing for use with an impeller connected to an engine is disclosed, as well as a method of operating turbo-machinery including supporting a load on the scroll housing with scroll vanes while maintaining an equal stress on each scroll vane.

Dual alloy turbine rotors and methods for manufacturing the same

Dual alloy turbine rotors and methods for manufacturing the same are provided. The dual alloy turbine rotor comprises an assembled blade ring and a hub bonded to the assembled blade ring. The assembled blade ring comprises a first alloy selected from the group consisting of a single crystal alloy, a directionally solidified alloy, or an equi-axed alloy. The hub comprises a second alloy. The method comprises positioning a hub within a blade ring to define an interface between the hub and the blade ring. The interface is a non-contacting interface or a contacting interface. The interface is enclosed by a pair of diaphragms. The interface is vacuum sealed. The blade ring is bonded to the hub after the vacuum sealing step.

TURBINE COMPONENT WITH SHAPED COOLING PINS

A turbine component with shaped cooling pins is provided. The turbine component includes at least one cooling circuit defined within the turbine component, the at least one cooling circuit in fluid communication with a source of a cooling fluid. The turbine component includes at least one shaped cooling pin disposed in the at least one cooling circuit. The at least one shaped cooling pin has a first end and a second end extending along an axis. The first end has a first curved surface defined by a minor diameter and the second end has a second curved surface defined by a major diameter. The first curved surface is upstream in the cooling fluid and the minor diameter is less than the major diameter. 1. A turbine component , comprising:at least one cooling circuit defined within the turbine component, the at least one cooling circuit in fluid communication with a source of a cooling fluid; andat least one shaped cooling pin disposed in the at least one cooling circuit, the at least one shaped cooling pin having a first end and a second end extending along an axis, the first end having a first curved surface defined by a minor diameter and the second end having a second curved surface defined by a major diameter, the first curved surface is upstream in the cooling fluid and the minor diameter is less than the major diameter.2. The turbine component of claim 1 , wherein the first curved surface is spaced apart from the second curved surface by a length.3. The turbine component of claim 1 , wherein the first curved surface and the second curved surface are interconnected by a pair of surfaces defined by a pair of planes substantially tangent to a respective one of the first curved surface and the second curved surface.4. The turbine component of claim 2 , wherein claim 2 , in cross-section claim 2 , the second curved surface is defined by a first circle having a first center point and the first curved surface is defined by a second circle having a second center point ...

Turbine tip shroud assembly with plural shroud segments having inter-segment seal arrangement

A shroud assembly for a gas turbine engine includes a plurality of shroud segments that are attached to a shroud support with an inter-segment joint defined between shroud segments. The shroud assembly also includes a cooling flow path cooperatively defined by the shroud support and the first shroud segment. The cooling flow path includes an internal cooling passage within the shroud segments. The cooling flow path includes an outlet chamber configured to receive flow from the internal cooling passage. The shroud assembly additionally includes a seal arrangement that extends across the inter-segment joint. The seal arrangement, the first shroud segment, and the second shroud segment cooperatively define a seal chamber that is enclosed.

Method to modify an airfoil internal cooling circuit

Methods are provided for modifying an airfoil internal cooling circuit that include a flow path configured to direct air through the airfoil in a direction and the airfoil having a leading edge, a trailing edge, and a first and a second wall therebetween, each wall having an inner and an outer surface, the inner surfaces defining a cavity and having features forming at least a portion of the internal cooling circuit. The methods may include the steps of forming a pilot hole through the airfoil first and second walls at a predetermined location, forming an insert hole based on the predetermined location, the insert hole enveloping the pilot hole and configured to receive at least a portion of an insert configured to modify the internal cooling circuit flow path, placing the insert into the insert hole, and bonding the insert to the airfoil first and second walls.

Turbine wheels, turbine engines including the same, and methods of fabricating turbine wheels with improved bond line geometry

Turbine wheels, turbine engines, and methods of fabricating the turbine wheels are provided. An exemplary method includes fabricating a turbine wheel that includes a rotor disk and a plurality of turbine blades operatively connected to the rotor disk through a blade mount. The method includes locating a cooling passage within a blade mount preliminary configuration and a cooling inlet on a surface of the blade mount preliminary configuration. A rotor disk bonding surface geometry and a blade mount bonding surface geometry are designed based upon a stress analysis of the turbine wheel and locations of the cooling passage and cooling inlet. A rotor disk production configuration and a blade mount production configuration are generated based upon the preliminary configurations. A blade mount and a rotor disk are formed based upon the production configurations. A blade ring including a plurality of blade mounts is formed and bonded to the rotor disk.

METHODS FOR PROCESSING BONDED DUAL ALLOY ROTORS INCLUDING DIFFERENTIAL HEAT TREATMENT PROCESSES

Methods for processing bonded dual alloy rotors are provided. In one embodiment, the method includes obtaining a bonded dual alloy rotor including rotor blades bonded to a hub disk. The rotor blades and hub disk are composed of different alloys. A minimum processing temperature (T) for the hub disk and a maximum critical temperature for the rotor blades (T) is established such that Tis less than T. A differential heat treatment process is then performed during which the hub disk is heated to processing temperatures equal to or greater than T, while at least a volumetric majority of each of the rotor blades is maintained at temperatures below T. Such a targeted differential heat treatment process enables desired metallurgical properties (e.g., precipitate hardening) to be created within the hub disk, while preserving the high temperature properties of the rotor blades and any blade coating present thereon. 1. A method for processing a bonded dual alloy rotor that includes rotor blades composed of a first alloy that are bonded to a hub disk composed of a second alloy different than the first alloy , the method comprising the steps of:{'sub': DISK_PROCESS_MIN', 'BLADE_MAX, 'performing a differential heat treatment process on the bonded dual alloy rotor during which the hub disk is heated to processing temperatures equal to or greater than a minimum processing temperature (T), while at least a volumetric majority of each of the rotor blades is maintained at temperatures below a maximum critical temperature (T),'}{'sub': BLADE_MAX', 'DISK_PROCESS_MIN, 'wherein Tis less than T.'}2. The method of wherein performing the differential heat treatment process comprises:{'sub': 'DISK_PROCESS_MIN', 'heating the hub disk to processing temperatures equal to or greater than Tover a peak heating period; and'}{'sub': DISK_PROCESS_MIN', 'BLADE_MAX, 'controlling heat transfer to and from the bonded dual alloy rotor during the differential heat treatment process such the magnitude of any ...

TURBINE WHEELS, TURBINE ENGINES INCLUDING THE SAME, AND METHODS OF FABRICATING TURBINE WHEELS WITH IMPROVED BOND LINE GEOMETRY

Turbine wheels, turbine engines, and methods of fabricating the turbine wheels are provided. An exemplary method includes fabricating a turbine wheel that includes a rotor disk and a plurality of turbine blades operatively connected to the rotor disk through a blade mount. The method includes locating a cooling passage within a blade mount preliminary configuration and a cooling inlet on a surface of the blade mount preliminary configuration. A rotor disk bonding surface geometry and a blade mount bonding surface geometry are designed based upon a stress analysis of the turbine wheel and locations of the cooling passage and cooling inlet. A rotor disk production configuration and a blade mount production configuration are generated based upon the preliminary configurations. A blade mount and a rotor disk are formed based upon the production configurations. A blade ring including a plurality of blade mounts is formed and bonded to the rotor disk.

TURBINE WHEELS, TURBINE ENGINES INCLUDING THE SAME, AND METHODS OF FABRICATING TURBINE WHEELS WITH IMPROVED BOND LINE GEOMETRY

Turbine wheels, turbine engines, and methods of fabricating the turbine wheels are provided. An exemplary method includes fabricating a turbine wheel that includes a rotor disk and a plurality of turbine blades operatively connected to the rotor disk through a blade mount. The method includes locating a cooling passage within a blade mount preliminary configuration and a cooling inlet on a surface of the blade mount preliminary configuration. A rotor disk bonding surface geometry and a blade mount bonding surface geometry are designed based upon a stress analysis of the turbine wheel and locations of the cooling passage and cooling inlet. A rotor disk production configuration and a blade mount production configuration are generated based upon the preliminary configurations. A blade mount and a rotor disk are formed based upon the production configurations. A blade ring including a plurality of blade mounts is formed and bonded to the rotor disk.

Dual alloy gas turbine engine rotors and methods for the manufacture thereof

Dual alloy Gas Turbine Engine (GTE) rotors and method for producing GTE rotors are provided. In one embodiment, the method include includes arranging bladed pieces in an annular grouping or ring formation such that shank-to-shank junctions are formed between circumferentially-adjacent bladed pieces. A first or bonding alloy is deposited along the shank-to-shank junctions utilizing a localized fusion deposition process to produce a plurality of alloy-filled joints, which join the bladed pieces in a bonded blade ring. The bonding alloy is preferably selected to have a ductility higher than and a melt point lower than the alloy from which the bladed pieces are produced. After deposition of the first alloy and formation of the alloy-filled joints, a hub disk is inserted into the central opening of the bonded blade ring. The hub disk and blade ring are then bonded utilizing, for example, a Hot Isostatic Pressing process.

Integral compressor housing of gas turbine engines

A turbine engine compressor design utilizing multiple component integration, thereby reducing the number of required engine components. In conventional compressor designs, a multiple component system makes it difficult to predict the structural behaviors due to thermal and mechanical loading during transient conditions. The compressor design of the present invention has three main parts: a forward bearing housing, a bell-mouth (heat shield) and a coupled impeller shroud/diffuser. Such a design achieves the design objectives of the present invention, including reducing weight, reducing cost, minimizing tolerance build up and improving aerodynamic performance by utilizing multiple component integration for multiple modes of engine operation.

Turbine scroll assembly for gas turbine engine

A gas turbine engine includes a compressor section and a combustion section with a scroll, a scroll baffle, a combustor, and a combustor case. The scroll defines an interior scroll flow path. The scroll baffle surrounds the scroll to define a scroll cooling passage. The combustor case surrounds the combustor and the scroll baffle to define a collector space. Moreover, the engine includes a turbine section with a turbine rotor and a turbine rotor blade shroud that includes a shroud cooling passage. The compressor flow path is fluidly connected to the scroll for cooling the scroll. Also, the scroll cooling passage is fluidly connected to the shroud cooling passage for cooling the turbine rotor blade shroud. Furthermore, the shroud cooling passage is fluidly connected to the collector space. Flow from the collector space flows into the combustor, along the interior scroll flow path, toward the turbine rotor.

Method to augment heat transfer using chamfered cylindrical depressions in cast internal cooling passages

An air-cooled turbine blade is provided having an airfoil shape defined by a convex suction side wall, a concave pressure side wall, a leading edge, a trailing edge, a root and a tip, each wall including an interior surface that defines an interior with the edges, root and tip. The blade includes a plurality of independent cooling circuit flow paths within the blade interior and a roughened surface. The roughened surface is formed on an interior surface of at least one of the convex suction side wall and the concave pressure side wall, the roughened surface comprising a plurality of cylindrical depressions. Each depression includes a cylindrical wall coupled to a bottom wall by a chamfered edge formed therebetween, and the roughened surface defines at least a portion of one of the flow paths of the plurality of independent cooling circuit flow paths. Methods for forming the blade are also provided.

TURBINE SCROLL ASSEMBLY FOR GAS TURBINE ENGINE

A gas turbine engine includes a compressor section and a combustion section with a scroll, a scroll baffle, a combustor, and a combustor case. The scroll defines an interior scroll flow path. The scroll baffle surrounds the scroll to define a scroll cooling passage. The combustor case surrounds the combustor and the scroll baffle to define a collector space. Moreover, the engine includes a turbine section with a turbine rotor and a turbine rotor blade shroud that includes a shroud cooling passage. The compressor flow path is fluidly connected to the scroll for cooling the scroll. Also, the scroll cooling passage is fluidly connected to the shroud cooling passage for cooling the turbine rotor blade shroud. Furthermore, the shroud cooling passage is fluidly connected to the collector space. Flow from the collector space flows into the combustor, along the interior scroll flow path, toward the turbine rotor. 1. A gas turbine engine with a longitudinal axis comprising:a compressor section that defines a compressor flow path;a combustion section that includes a scroll, a scroll baffle, and a combustor, the scroll defining an interior scroll flow path, the scroll baffle surrounding at least part of the scroll to define a scroll cooling passage between the scroll baffle and the scroll;a turbine section with a turbine rotor and a turbine rotor blade shroud, the turbine rotor blade shroud including a shroud cooling passage;the compressor flow path fluidly connected to the scroll cooling passage via a mid-frame plenum, the mid-frame plenum extending at least partly about the longitudinal axis in a circumferential direction, the mid-frame plenum defining a mid-frame flow path from the compressor flow path, through the mid-frame plenum, and to the scroll cooling passage, the mid-frame flow path directed in a radially inboard direction, through the mid-frame plenum, and in a radially outboard direction into the scroll cooling passage; andthe combustion section being coupled to ...

METHODS FOR FABRICATING HIGH TEMPERATURE CASTABLE ARTICLES AND GAS TURBINE ENGINE COMPONENTS

Embodiments of a method are provided for fabricating a high temperature castable article, such as a casting mold or a core. In one embodiment, the method includes the steps of providing a suspension containing a plurality of ceramic particles and a photo-curable monomer; photo-curing selected portions of the suspension to produce a green body coated, at least partially, with uncured suspension; and removing the uncured suspension from the green body under process conditions at which the viscosity of the uncured suspension is reduced. The process conditions include exposing the green body to one of the group consisting of: (i) microwave energy sufficient to excite a dipole moment of the uncured suspension and (ii) sonic energy sufficient to induce shear-thinning of the uncured suspension. The green body is then heat treated to produce the high temperature castable article. 1. A method for fabricating a high temperature castable article , comprising:providing a suspension comprising a plurality of ceramic particles and a photo-curable monomer;photo-curing selected portions of the suspension to produce a green body coated, at least partially, with uncured suspension;removing the uncured suspension from the green body under process conditions at which the viscosity of the uncured suspension is reduced, the process conditions comprising exposing the green body to one of the group consisting of (i) microwave energy sufficient to excite a dipole moment of the uncured suspension and (ii) sonic energy sufficient to induce shear-thinning of the uncured suspension; andheat treating the green body to produce the high temperature castable article.2. A method according to further comprising the step of casting a gas turbine engine component utilizing claim 1 , at least in part claim 1 , the high temperature castable article.3. A method according to wherein step of removing comprises treating the green body with a solvent.4. A method according to wherein the step of removing ...

Turbine wheels, turbine engines including the same, and methods of fabricating turbine wheels with improved bond line geometry

Turbine component with shaped cooling pins

Turbine rotor seal plate with integral flow discourager

A seal plate for a turbine engine (310) is provided. The turbine engine (310) has a central axis (450). The seal plate comprises a forward face extending radially from the central axis (450). The forward face comprises a flat portion extending circumferentially around the central axis (450), and a front flow discourager (512) coupled to the flat portion and extending outward from the flat portion.

Turbine nozzle assembly including radially-compliant spring member for gas turbine engine

Embodiments of a turbine nozzle assembly (58) are provided for deployment within a gas turbine engine (GTE 20) including a first GTE-nozzle mounting interface (101). In one embodiment, the turbine nozzle assembly includes a turbine nozzle flowbody, a first mounting flange (98) configured to be mounted to the first GTE-nozzle mounting interface, and a first radially-compliant spring member (131) coupled between the turbine nozzle flowbody and the first mounting flange (98). The turbine nozzle flowbody has an inner nozzle endwall (92) and an outer nozzle endwall (90), which is fixedly coupled to the inner nozzle endwall (92) and which cooperates therewith to define a flow passage (96) through the turbine nozzle flowbody. The first radially-compliant spring member (131) accommodates relative thermal movement between the turbine nozzle flowbody and the first mounting flange (98) to alleviate thermomechanical stress during operation of the GTE (20).

Turbine tip shroud assembly with plural shroud segments having internal cooling passages

A shroud assembly for a gas turbine engine includes a shroud support and a plurality of shroud segments that are attached to the shroud support. The shroud segment includes an internal cooling passage.

Turbine tip shroud assembly with plural shroud segments having internal cooling passages

A shroud assembly for a gas turbine engine includes a shroud support and a plurality of shroud segments that are attached to the shroud support. The shroud segment includes an internal cooling passage.

Cooled turbine blade platform

A turbine blade assembly (100) is provided. The turbine blade assembly (100) comprises a turbine blade comprising a cavity (138), and a blade platform (120) supporting the turbine blade, the cavity (138) extending into the blade platform (120). The blade platform (120) comprises an upper surface (128) adjacent the turbine blade and a lower surface (130) comprising a first rib (200), the cavity (138) extending into the first rib (200), the first rib (200) coupled to the lower surface (130), tapering as it extends away from the turbine blade, and comprising a first port (132) extending from the cavity (138) to the upper surface (128).

Turbomachine compressor scroll with load-carrying inlet vanes

A compressor scroll housing for use in conjunction with turbo-machinery, particularly applicable in aircraft. The scroll housing can include a plurality of scroll vanes arrayed around the scroll housing. Scroll vanes, integrally formed with the scroll housing, carry stress load on the scroll housing, including the load from fluid pressure within the scroll and carcass loading from the engine. The plurality of scroll vanes adapted for guiding flow of fluid from an inlet to an outlet while supporting the scroll housing. Chord length and cross sectional area of each scroll vane can be sized to maintain an equal stress in all scroll vanes. A method of making the scroll housing for use with an impeller connected to an engine is disclosed, as well as a method of operating turbo-machinery including supporting a load on the scroll housing with scroll vanes while maintaining an equal stress on each scroll vane.

Methods for manufacturing a turbine nozzle with single crystal alloy nozzle segments

Turbine tip shroud assembly with plural shroud segments having inter-segment seal arrangement

A shroud assembly for a gas turbine engine includes a plurality of shroud segments that are attached to a shroud support with an inter-segment joint defined between shroud segments. The shroud assembly also includes a cooling flow path cooperatively defined by the shroud support and the first shroud segment. The cooling flow path includes an internal cooling passage within the shroud segments. The cooling flow path includes an outlet chamber configured to receive flow from the internal cooling passage. The shroud assembly additionally includes a seal arrangement that extends across the inter-segment joint. The seal arrangement, the first shroud segment, and the second shroud segment cooperatively define a seal chamber that is enclosed.

Turbine shroud assemblies for gas turbine engines

A turbine shroud assembly includes a shroud support and a shroud segment. The shroud support structure includes a forward support rail and an aft support rail. The forward support rail includes forward first engagement structures and the aft support rail includes aft first engagement structures. The shroud segment includes a forward segment rail and an aft segment rail. The forward segment rail includes forward second engagement structures positioned on a forward segment rail periphery and the aft segment rail includes second engagement structures positioned on an aft segment rail periphery. The forward first engagement structures radially and circumferentially engage with the forward second engagement structures and the aft first engagement structures radially and circumferentially engage with the aft second engagement structures to radially and circumferentially interlock the shroud segment to the shroud support structure.

Turbine tip shroud assembly with plural shroud segments having internal cooling passages

Dual alloy turbine rotors and methods for manufacturing the same

Dual alloy turbine rotors and methods for manufacturing the same are provided. The dual alloy turbine rotor comprises an assembled blade ring and a hub bonded to the assembled blade ring. The assembled blade ring comprises a first alloy selected from the group consisting of a single crystal alloy, a directionally solidified alloy, or an equi-axed alloy. The hub comprises a second alloy. The method comprises positioning a hub within a blade ring to define an interface between the hub and the blade ring. The interface is a noncontacting interface or a contacting interface. The interface is enclosed by a pair of diaphragms. The interface is vacuum sealed. The blade ring is bonded to the hub after the vacuum sealing step.

Cooled turbine blade platform

Spring biased shroud retention system for gas turbine engine

A system for coupling a shroud to a case associated with a gas turbine engine includes the case having a mounting pad, and the shroud having a surface that faces the case. The system includes a load spreader having a spreader surface in contact with the surface of the shroud and a locator pin coupled to the mounting pad and the load spreader to couple the shroud to the case. The system includes a load spreader retainer coupled to the load spreader. The load spreader retainer is to distribute a force to the load spreader. The system includes a biasing system coupled about the mounting pad that includes a spring arm configured to apply the force to the load spreader retainer to maintain the spreader surface of the load spreader in contact with the surface of the shroud.