Honeywell International Inc.

Bradley Reed Tucker

Jason Smoke

Ardeshir Riahi

Shezan Kanjiyani

Rajiv Rana

Ed Zurmehly

Mark C Morris

Gregory O Woodcock

Lorenzo Crosatti

Stony Kujala

Alexander MirzaMoghadam

Natalie Wali

John Gintert

Yogendra Y Sheoran

Derek Raybould

Harry Lester Kington

Cristopher Frost

Paul Chipko

David Chou

John Taylor Pearson

Zedic Daniel Judd

Terrel Kuhn

Daniel Ryan

Bob Mitlin

Bill Macelroy

Mike O'Brien

John Hogan

Clay R Johnstun

Louis P Begin

Daniel Hicks

Dan Crites

Douglas Hollingshead

Alexander Mirzamoghadam

An asymmetric inlet particle separator for a gas turbine engine includes an inlet having a first cross-sectional shape, and a duct downstream of the inlet. The duct includes a bend upstream from a splitter, a scavenge branch and an engine airflow branch. The splitter is outside of a line of sight from the inlet and the splitter separates the scavenge branch from the engine airflow branch. The asymmetric inlet particle separator includes an annulus downstream of the engine airflow branch configured to be coupled to the gas turbine engine. The annulus has a second cross-sectional shape, and the engine airflow branch transitions from the first cross-sectional shape to the second cross-sectional shape.

Asymmetric inlet particle separator for gas turbine engine

Embodiments of a turbine nozzle having slip joints impregnated by an oxidation-resistant sealing material are provided, as are embodiments of methods for the manufacture of turbine nozzles. In one embodiment, the method includes providing a support ring, a slip joint ring substantially concentric with the support ring and radially spaced apart therefrom, and a plurality of vanes fixedly coupled to the support ring. The plurality of vanes extends radially from the support ring into a plurality of circumferentially-spaced slots provided in the slip joint ring to form a plurality of slip joints therewith. The plurality of slip joints are impregnated with a silicon-modified aluminide sealing material. The silicon-modified aluminide sealing material impedes gas flow into the radial slip joints during operation of the turbine nozzle, while also fracturing to permit relative radial movement between the plurality of vanes and the slip joint ring along the plurality of slip joints.

Turbine nozzles with slip joints impregnated by oxidation-resistant sealing material and methods for the production thereof

A tie shaft for a rotating group of an engine core includes a cylindrical body having an internal surface and an external surface and extending between a forward end and an aft end. The tie shaft further includes a first group of internal grooves on the internal surface of the cylindrical body proximate to the forward end and a second group of internal grooves on the internal surface of the cylindrical body proximate to the aft end.

Gas turbine engines with internally stretched tie shafts

Turbine nozzles and cooling systems for cooling slip joints therein are provided. The turbine nozzle has an endwall, a vane coupled to the endwall, a slip joint, and a plurality of airfoil quenching holes that cooperate with a plurality of endwall cooling holes. The vane comprises a leading edge and a trailing edge interconnected by a pressure sidewall and a suction sidewall and an end portion. The slip joint is between the end portion and the endwall. The airfoil quenching holes are defined through the pressure sidewall in the end portion. The endwall cooling holes are defined through the endwall along the pressure sidewall and in proximity to the leading edge. The airfoil quenching holes and endwall cooling holes are disposed adjacent the slip joint.

Turbine nozzles and cooling systems for cooling slip joints therein

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.

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

A turbine nozzle assembly includes an inner circumferential support platform, an outer circumferential support platform, and a plurality of airfoil vanes disposed between the inner circumferential support platform and the outer circumferential support platform. The turbine nozzle assembly further includes a plurality of impingement plates disposed along a radially outer surface of the outer circumferential support platform or a radially inner surface of the inner circumferential support platform, and a plurality of gap-maintaining features disposed between the plurality of outer or inner circumferential support platforms and the plurality of impingement plates. Each gap-maintaining feature of the plurality of gap-maintaining features is provided at a height such that a cooling air flow space is maintained between the plurality of outer or inner circumferential support platforms and the plurality of impingement plates.

Turbine nozzles and methods of manufacturing the same

A turbine section includes a stator assembly having an inner diameter end wall, an outer diameter end wall, and a stator vane; a turbine rotor assembly including a rotor blade extending into the mainstream gas flow path; a housing including an annular shroud that circumscribes the rotor blade and at least partially defines the mainstream hot gas flow path; a first baffle arranged to define a first cavity with the outer diameter end wall of the stator assembly; a second baffle; and a third baffle arranged to define a second cavity with the second baffle and a third cavity with the shroud. The first cavity is fluidly coupled to the second cavity and the second cavity is fluidly coupled to the third cavity such that cooling air flows from the first cavity to the second cavity and from the second cavity to the third cavity.

Turbine sections of gas turbine engines with dual use of cooling air

Stationary airfoils configured to form an improved slip joint in bi-cast turbine engine components and the turbine engine components including the same are provided. The stationary airfoil for a bi-cast turbine engine component comprises a leading edge and a trailing edge interconnected by a pressure sidewall and a suction sidewall. An end portion is shaped with a pair of opposing flanges to form a slip joint with a shroud ring in the bi-cast turbine engine component and to define an interlocking feature. The slip joint permits radial movement of the stationary airfoil relative to the shroud ring due to thermal differential expansion and contraction.

Stationary airfoils configured to form improved slip joints in bi-cast turbine engine components and the turbine engine components including the same

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.

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

A coated component is provided comprising a silicon-based substrate and a braze layer overlying the silicon-based substrate. The braze layer comprises silicon, tantalum, and a metal element having substantially the same melt temperature with silicon as tantalum has with silicon. The braze layer further comprises ceramic particles. Protective coatings are also provided.

Protective coatings and coated components comprising the protective coatings

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.

Gas turbine engines with abradable turbine seal assemblies

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.

Turbine assemblies with impingement cooling

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.

Combustor-turbine seal interface 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 nozzle assembly including radially-compliant spring member for gas turbine engine

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.

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Chandler, AZ, United States of America

Bradley Reed Tucker