United Technologies Corporation

Raytheon Technologies Corporation

Rtx Corporation

Michael G McCaffrey

John R Farris

Ross Wilson

Eric A Hudson

Edwin Otero

Jeffrey Michael Jacques

Tracy A Propheter-Hinckley

Bryan P Dube

Scott Daniel Virkler

Benjamin F Hagan

Mark F Zelesky

Raymond C Surace

Gabriel L Suciu

Brian D Merry

Grant O Cook, Iii

Matthew A Devore

John E Holowczak

Matthew Andrew Hough

Daniel Bernard Kupratis

Brandon W Spangler

Zaffir A Chaudhry

Dilip M Shah

Alan David Cetel

Frederick M Schwarz

Wayde R Schmidt

Joel H Wagner

Timothy M Davis

Jonathan Jeffery Eastwood

Shankar S Magge

Paul M Lutjen

Thomas N Slavens

San Quach

Marc J Muldoon

Paul Sheedy

Fanping Sun

Thomas E Clark

William K Ackermann

Conway Chuong

James D Hill

Andrew Joseph Lazur

Mark John Rogers

Ioannis Alvanos

Anthony R Bifulco

Howard J Liles

Sameh Dardona

William M Barker

David Wasserman

Tania Bhatia Kashyap

John E Paul

Theodore W Hall

Alan W Stoner

Michael G Abbott

Christopher Corcoran

Stanley J Funk

Douglas Michael Berczik

Shelton O Duelm

Venkatarama K Seetharaman

Slade R Culp

Elizabeth F Vinson

John J Korzendorfer

David Richard Griffin

Brandon T Rouse

Mark Borja

Thomas Almy

Pedro Rivero

Brian R Pelletier

Stephen G Pixton

Kevin L Rugg

Nathan L Messersmith

Peter Balawajder

George T Suljak, Jr

Joseph F Englehart

Thomas J Praisner

Brooks E Snyder

Andrew S Aggarwala

Alan C Barron

Hoyt Y Chang

William M Proscia

Christopher King

Joseph T Caprario

Andrew D Burdick

Daniel J Whitney

Joseph C Burge

Bryan H Farrar

Shahram Amini

Graham Ryan Philbrick

Daniel R Sabatino

Jesse M Carr

Tuan David Vo

Andrew J Murphy

Christopher W Robak

Henry K Webster

Richard Seleski

Lee A Hoffman

Caroline Karanian

Karl A Mentz

Greg C Ojard

Kristine Marie Carnavos

Ronald R Gagnon, Jr

Christopher William Moore

An assembly is provided for a turbine engine. This assembly includes an engine core configured to drive rotation of a propulsor rotor. The engine core includes a first compressor section, a second compressor section, a flowpath, a bleed port, an inner passage and an outer passage. The first compressor section and the second compressor section are arranged axially along an axis. The flowpath extends longitudinally through the first compressor section and the second compressor section. The bleed port fluidly couples the flowpath to the inner passage and the outer passage in parallel. The bleed port is located longitudinally along the flowpath between the first compressor section and the second compressor section. The inner passage is arranged radially between the flowpath and the outer passage.

Bleeding core air from a turbine engine core flowpath

A dirt blocker including a support structure disposed within a gas turbine engine, the support structure defining an upstream control volume proximate a forward portion of the gas turbine engine and a downstream control volume proximate an aft portion of the gas turbine engine, the downstream control volume being opposite the upstream control volume relative to the support structure, a flow passage formed through the support structure, the flow passage configured to fluidly couple the upstream control volume with the downstream control volume; a radial contact wall formed from the support structure in fluid communication with the upstream control volume, the radial contact wall configured to intercept debris entrained within cooling air within the gas turbine engine; and a stagnation zone fluidly coupled with the flow passage, the stagnation zone configured to reduce momentum of the debris.

High pressure turbine dirt blocker

An assembly is provided for a turbine engine. This assembly includes an engine core extending axially along an axis. The engine core includes a compressor section, a combustor, a diffuser structure, a diffuser plenum and a plurality of separators. The combustor is arranged within the diffuser plenum. The combustor includes a combustion chamber and a combustor wall between the combustion chamber and the diffuser plenum. The diffuser structure includes a plurality of diffuser passages. Each of the diffuser passages fluidly couples the compressor section to a respective one of the separators. Each of the separators includes a first outlet into the diffuser plenum and a second outlet into the combustion chamber.

Separating airflows within a turbine engine

An assembly for a gas turbine engine includes at least one rotational assembly, an engine static structure, a compressor, one or more compressed air loads, and a particulate separator assembly. The at least one rotational assembly includes a shaft, a bladed compressor rotor, and a bladed turbine rotor. The engine static structure includes an engine case assembly. The engine case assembly surrounds the at least one rotational assembly. The compressor includes the bladed compressor rotor. The compressor is configured to form a compressed air flow. The one or more compressed air loads are disposed within the engine case assembly. The particulate separator assembly includes a plurality of particulate separators. The plurality of particulate separators are disposed outside of the engine case assembly. The plurality of particulate separators are configured to separate particulate from the compressed air flow and direct the compressed air flow to the one or more compressed air loads.

Particulate separator assembly for a gas turbine engine

A fan has a rotating bulge defining a radially outermost point of the rotating cone relative to a rotational axis. The rotating bulge is defined by a radially outwardly extending forward portion leading to the radially outermost point on the bulge and then a radially inwardly extending portion aft of the radially outermost point of the bulge leading into the core engine path within the splitter housing. The core splitter housing has a leading edge, and the bulge radially outermost point being upstream of the leading edge of the splitter housing. The bulge radially outermost point is a first radial distance from a rotational axis, and the leading edge of the splitter housing being at a second radial distance of the rotational axis, and a ratio of the first radial distance to the second radial distance being at least 0.95.

Gas turbine engine with core debris protection

An inlet protector for a vane cooling air passage inlet including a vane including a vane cooling air inlet; an inlet protector fluidly coupled to the vane cooling air inlet, the inlet protector comprising a protector body having a discharge portion and a suction portion, an interior flow passage fluidly coupling the suction portion with the discharge portion; the discharge portion fluidly coupled with the vane cooling air inlet; and a suction inlet formed in the suction portion, wherein the suction inlet is located distally from the vane cooling air inlet.

Inlet protector for vane coupling hole

A turbine for a gas turbine engine includes, among other things, a shaft rotatable about a longitudinal axis, a turbine rotor including one or more rows of turbine blades and a disk assembly coupled to the shaft. The disk assembly includes one or more disks each having an attachment region extending radially between an inner boundary and an outer boundary, the outer boundary is established by an outer periphery of the respective disk, the attachment region defines an array of slots distributed about the outer periphery, each of the slots extends radially inwardly from the outer boundary to the inner boundary, and each of the slots is dimensioned to receive a root section of a respective one of the turbine blades to mount the turbine blades to the disk assembly.

Airfoil attachment for turbine rotor

An airfoil includes a CMC airfoil wall that defines a platform, a root that extends radially from the platform, and an airfoil section that extends radially from the platform opposite the root. The CMC airfoil wall includes a wishbone-shaped fiber layup structure. The wishbone-shaped fiber layup structure includes first and second arms that merge into a single leg. The first and second arms are comprised of a network of fiber tows, and the single leg includes fiber tows from each of the first and second arms that are interwoven in the single leg. The single leg forms a portion of the root, and the first and second arms extend in opposite directions from the single leg and form a portion of the platform.

CMC airfoil with wishbone shaped fiber layup structure forming portion of platform and root

An assembly is provided for a turbine engine. This assembly includes an engine core extending along an axis. The engine core includes a first compressor section, a second compressor section, a flowpath, a bleed port, an inner passage, an outer passage and a flow diverter. The flowpath extends longitudinally through the first compressor section and the second compressor section. The bleed port fluidly couples the flowpath to the inner passage and the outer passage in parallel. The bleed port is located longitudinally along the flowpath between the first compressor section and the second compressor section. The flow diverter is located at an inlet into the bleed port from the flowpath. The flow diverter is configured to move between a first position and a second position.

Regulating core air bled from a turbine engine core flowpath

An assembly for an aircraft engine includes a variable area nozzle. The variable area nozzle includes a nozzle wall, a nozzle sleeve, an actuation system and a flowpath extending axially along an axis through the variable area nozzle and radially between the nozzle wall and the nozzle sleeve. The nozzle sleeve includes a support and a liner. The support includes a first mount and a second mount axially spaced from the first mount. The liner extends axially along and circumferentially about the support. The liner is axially and circumferentially anchored to the support through the first mount. The liner is circumferentially anchored to the support through the second mount. The actuation system is coupled to the support. The actuation system is configured to move the nozzle sleeve axially along the axis relative to the nozzle wall.

Variable area nozzle for an aircraft engine

A variable area nozzle assembly includes a fixed center plug, a fixed outer nozzle, and an inner nozzle. The fixed center plug includes a seal. The inner nozzle is disposed between the fixed center plug and the fixed outer nozzle. The inner nozzle includes an inner nozzle body. The inner nozzle body forms a primary duct between the inner nozzle and the fixed outer nozzle. The inner nozzle body forms a secondary duct and a gap. The gap is a nozzle outlet of the secondary duct. The secondary duct is disposed between the inner nozzle body and the fixed center plug upstream of the seal. The inner nozzle body forms a plurality of apertures. The inner nozzle body is translatable between a first position and a second position. In the first position, the primary duct has a first cross-sectional area and the plurality of apertures are isolated from the gap by the seal. In the second position, the primary duct has a second cross-sectional area and the plurality of apertures are connected in fluid communication with the gap.

Variable area nozzle assembly for an aircraft propulsion system

A variable area nozzle assembly includes a fixed center plug, a fixed outer nozzle, and an inner nozzle. The fixed center plug includes a plug body and a protrusion. The fixed outer nozzle extends about the fixed center plug. The inner nozzle extends about the fixed center plug. The inner nozzle is disposed between the fixed center plug and the fixed outer nozzle. The inner nozzle forms a primary duct between the inner nozzle and the fixed outer nozzle. The primary duct includes an exit plane. The inner nozzle body forms a secondary duct and a gap. The gap is a nozzle outlet of the secondary duct. The gap includes a gap inlet and a gap outlet. The gap inlet is disposed upstream of the protrusion. The gap outlet is disposed downstream of the protrusion. The inner nozzle body is translatable between a first position and a second position. In the first position, the primary duct has a first area at the exit plane and the inner nozzle body is disposed at the protrusion. In the second position, the primary duct has a second area at the exit plane and the inner nozzle body is spaced from the protrusion.

Tangential onboard injector assemblies include a TOBI and a TOBI blocker assembly arranged upstream of an inlet of the TOBI. The TOBI blocker assembly includes a first blocker plate upstream from the inlet and defining a blocker cavity between the first blocker plate and the TOBI. The first blocker plate includes a blocking portion and a filter portion to permit airflow therethrough. A second blocker plate is arranged within the blocker cavity to obstruct a flow of air containing particulate matter and a gap is present between the second blocker plate and the first blocker plate to permit flow of air around the second blocker plate. A third blocker plate is arranged downstream from the second blocker plate within the blocker cavity and includes a feed hole. Air within a region between the second blocker plate and the third blocker plate flows through the feed hole into the TOBI.

Blocker assembly for tangential onboard injectors

A method includes providing a ceramic insert on a mandrel, the mandrel and the ceramic insert together define a peripheral working surface, forming a fiber preform by wrapping a fiber layer around the mandrel and the ceramic insert so as to conform to the peripheral working surface, removing the mandrel from the fiber preform to leave a cavity in the fiber preform, the ceramic insert remaining in the fiber preform and bordering the cavity, and densifying the fiber preform with a ceramic matrix to form a gas turbine engine component.

Engine article with ceramic insert and method therefor

A blade outer airseal (BOAS) assembly of a gas turbine engine includes a BOAS segment, and a BOAS carrier located radially outboard of the BOAS segment relative to an engine central longitudinal axis. The BOAS segment secured to the BOAS carrier. The BOAS assembly is selectably radially movable during operation of the gas turbine engine. One or more secondary retention features are configured to limit the radial movement of the BOAS assembly.

Failsafe blade outer airseal retention

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Windsor, CT, United States of America

Michael G McCaffrey