Honeywell International Inc.

Alliedsignal Inc.

Vibrant Corporation

Mark C Morris

Donald G Godfrey

Daniel Cregg Crites

Jason Smoke

Ardeshir Riahi

David R Waldman

Harry Lester Kington

Steve Halfmann

Steven Whitaker

Andy Szuromi

Daniel Ryan

Jeff Howe

Benjamin Dosland Kamrath

Jorge Alvarez

Malak Fouad Malak

Don Mittendorf

Joseph Jensen

William J Howe

Francis Carbonell

Steve Starr

Craig A Wilson

Thomas E Strangman

Luis A Tapia

Kin Poon

Brian Hann

Alan Tiltman

Bob Mitlin

Karl Fleer

George W Wolfmeyer

Daira Legzdina

Bradley Reed Tucker

Yiping Hu

Milton Ortiz

Eric Bridges

Jude Miller

Surendra Singh

Brian G Baughman

John Meier

Christopher David Gatto

James Neumann

Nnawuihe A Okpara

John McClintic

Douglas P Freiberg

Stuart Andrew Harman

Joseph Yanof

William C Baker

Partha Das

Jon Kettinger

Khosro Molla Hosseini

Greg Heitland

Mamballykalathil Menon

Dave Narasimhan

Yogendra Y Sheoran

Rodolphe Dudebout

David K Jan

Charles S Lo

Richard Bryce Fox

William F Hehmann

Yong Weon Kim

Donald L Mittendorf

Justin C Mickelsen

Donald L Palmer

Vincent Chung

Lorenzo Crosatti

Alexander S Kozlov

Peter Hall

Raymond Gage

Vighneswara Rao Kollati

Richard Bye

Alexander MirzaMoghadam

Paul Kurlak

Blake Petersen

George Reimer

Mark McNair

Brian Cottrell

Robert Hogan

Adrian R Allan

George E Zurmehly

Bradley Volkmann

Brent Ludwig

James J Cobb

Paul Couey

Andy Kinney

Brandan Wakefield

Zach Rogers

Tom Strangman

James Kroeger

Ryon James Stanley

Craig W McKeever

Douglas Hall

Vivek Agarwal

Michael G Volas

Michael K Bischoff

Jeffrey D Harrison

Manuel Sandoval

Michael Vinup

Michael Ryan Wedig

Klaus Helmut Schwarz

Donlad G Godfrey

Laurie Strom

Systems provide enhanced cooling of tip regions of turbine blades. A cooling system includes a turbine blade with a root, a base, a tip, a leading edge, a trailing edge, an outer structure that includes a pressure side wall, a suction side wall and a tip cap at its tip. A core in the blade defines a leading channel that ejects cooling air through the leading edge. A primary tip flag channel ejects cooling air through the tip and the trailing edge. A secondary tip flag channel ejects cooling air through the trailing edge. The primary tip flag channel and the secondary tip flag channel promote cooling of the tip due to the locations in the tip area and their delivery of low temperature and high pressure cooling air thereto.

Turbine blade double tip flag cooling system

Gas turbine engines and methods for reducing spool thrust in a combustor section thereof are provided. The engines include an impeller, a shaft coupled to the impeller, an inner cavity downstream of the impeller, a thrust bearing assembly mechanically coupled to the shaft, a seal assembly having a seal land, and a controller. The controller is configured to detect a loss of oil event to the thrust bearing assembly and, in response, actuate the seal assembly to move the seal land toward a downstream face of the impeller to form a continuous axial seal about the inner cavity, vent gases from the inner cavity to decrease pressure therein to an extent less than an operating gas pressure thereof, and maintain the axial seal and the pressure within the inner cavity and thereby reduce the axial loads on the shaft.

Gas turbine engines and methods for reducing spool thrust in a combustor section thereof

Gas turbine engines and methods for reducing spool thrust in a turbine section thereof are provided. The engines include a turbine disk, a shaft coupled to the turbine disk, an inner cavity downstream of the turbine disk, a thrust bearing assembly mechanically coupled to the shaft, a seal assembly having a seal land, and a controller. The controller is configured to detect a loss of oil event to the thrust bearing assembly and, in response: actuate the seal assembly to move the seal land toward a downstream face of the turbine disk to form a continuous axial seal about the inner cavity, supply a pressurized gas to the inner cavity to increase pressure in therein to an extent greater than an operating pressure thereof, and maintain the axial seal and the pressure within the inner cavity and thereby reduce the axial loads on the shaft.

Gas turbine engines and methods for reducing spool thrust in a turbine section thereof

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.

Composite turbine disc rotor for turbomachine

A turbine vane airfoil and a turbine nozzle that includes the turbine vane airfoil. The turbine vane airfoil includes an airfoil defining a chamber proximate a surface, and an impingement tube. The impingement tube includes a first wall and a second wall. The first wall defines a plurality of first apertures, and each has a first upstream surface portion opposite a first downstream surface portion. The second wall defines a plurality of second apertures, and each has a second upstream surface portion opposite a second downstream surface portion. Each of the first apertures cooperate with a respective one of the second apertures to direct a cooling fluid onto the surface, and the first downstream surface portion of at least one of the first apertures includes a first point that is collinear with a second point of the second downstream surface portion of the respective one of the second apertures.

Turbine nozzle with dust tolerant impingement cooling

A turbine vane includes an airfoil that extends from an inner diameter to an outer diameter, and from a leading edge to a trailing edge. The turbine vane includes an inner platform coupled to the airfoil at the inner diameter. The turbine vane includes a cooling system defined in the airfoil including a first conduit in proximity to the leading edge to cool the leading edge and a second conduit to cool the trailing edge. The first conduit has an inlet at the outer diameter to receive a cooling fluid and an outlet portion that is defined at least partially through the inner platform. The first conduit includes a plurality of cooling features that extend from a first surface of the first conduit, and the first surface of the first conduit is opposite the leading edge.

Turbine vane with dust tolerant cooling system

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.

Impingement baffle for gas turbine engine

A turbine vane includes an airfoil that extends from an inner diameter to an outer diameter, and from a leading edge to a trailing edge. The turbine vane includes an inner platform coupled to the airfoil at the inner diameter. The turbine vane includes a cooling system defined in the airfoil including a first conduit in proximity to the leading edge to cool the leading edge and a second conduit to cool the trailing edge. The first conduit has an inlet at the outer diameter to receive a cooling fluid and an outlet portion that is defined at least partially through the inner platform. The first conduit includes a plurality of cooling features that extend between a first surface and a second surface of the first conduit, and the first surface of the first conduit is opposite the leading edge.

A turbine nozzle for a gas turbine engine includes an outer endwall, and an inner endwall spaced apart from the outer endwall. The turbine nozzle includes at least one airfoil coupled between the inner endwall and the outer endwall, and a compliant joint defined in at least one of the outer endwall and the inner endwall. The compliant joint includes at least one partially-fused seal that is configured to restrict a flow of fluid through the at least one of the outer endwall and the inner endwall and to form a slip-joint above a predetermined threshold stress. The slip-joint is configured to maintain a radial position of the at least one of the outer endwall and the inner endwall.

Turbine nozzle with compliant joint

A heat exchanger includes a plurality of interconnected separator members that respectively include a first surface and an opposite second surface. The separator members respectively include an array of wave features. Also, the separator members are stacked and disposed in an alternating arrangement with the first surfaces of adjacent separator members facing each other and attached at the respective wave features, and with the second surfaces of adjacent separator members facing each other and attached at the respective wave features. The heat exchanger also includes a plurality of first flow passages for first fluid flow and second flow passages for second fluid flow. The second fluid and the first fluid are configured to exchange heat through the separator members.

Prime surface heat exchanger with contoured separator members

A turbine blade includes an airfoil that has a tip region that extends from the leading edge toward the trailing edge, and the tip region is bounded by a wall that extends at an angle. The leading edge has a leading edge cooling circuit that is defined from the platform to a tip flag channel, and the leading edge cooling circuit is in fluid communication with the tip flag channel. The pressure side includes at least one tip dust hole defined through the wall proximate the pressure side, and the at least one tip dust hole has an inlet and an outlet. The airfoil has at least one rib defined on the wall that extends at a second angle to direct the particles and a portion of the cooling fluid into the inlet.

Turbine blade with dust tolerant cooling system

A turbomachine includes a shroud and a rotor, which includes first and second blades. A first blade tip and a second blade tip respectively include a base and a first layer. The second blade tip also includes an abrasive second layer layered over the respective first layer. The first layer has a lower material hardness than the shroud. The second layer has a lower thermal stability than the shroud and the first layer. The rotor performs a grind operation and, subsequently, a post-grind operation. The second layer, in the grind operation, contacts and removes material from the shroud, and wears away, thereby revealing the first layer of the second blade tip for the post-grind operation. The first layer of the first blade tip is spaced apart with at least some radial clearance from the shroud in the grind and post-grind operations.

Rotor assembly for in-machine grinding of shroud member and methods of using the same

An engine component includes a body having an internal surface and an external surface, the internal surface at least partially defining an internal cooling circuit. The component further includes a plurality of cooling holes formed in the body and extending between the internal cooling circuit and the external surface of the body. The plurality of cooling holes includes a first cooling hole with a metering portion and a diffuser portion extending from the metering portion to the external surface of the body.

Engine components with cooling holes having tailored metering and diffuser portions

A heat exchanger includes a separator member that divides a first flow passage from a second flow passage. The heat exchanger also includes a plurality of first hollow members that extend across the first flow passage at respective non-orthogonal angles. The plurality of first hollow members are fluidly connected to the second flow passage. Moreover, the heat exchanger includes a plurality of second hollow members that extend across the second flow passage at respective non-orthogonal angles. The plurality of second hollow members are fluidly connected to the first flow passage.

Heat exchanger with interspersed arrangement of cross-flow structures

A turbine component includes a hot wall, a cold wall spaced apart from the hot wall and a conduit defined between the hot wall and the cold wall. A cooling system is defined in the conduit. The cooling system includes a plurality of cooling pins, each including a first end having a first cross-sectional area and a second end having a second cross-sectional area. Each cooling pin includes a body extending between the first end and the second end, with a pin leading edge defined along the body from the first end to the second end. The pin leading edge is defined by a first diameter and a pin trailing edge is defined by a second diameter. At least one first cooling pin has the first end coupled to the hot wall and the second end coupled to the cold wall with a fillet.

Turbine component with dust tolerant cooling system

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Phoenix, AZ, United States of America

Mark C Morris