United Technologies Corporation

Raytheon Technologies Corporation

Other

Metro Industries, Incorporated

Rtx Corporation

Dayco Products, Inc.

Nelson, Jack

Lewis, Charles M.

Oliver, Larry R.

Hollaway, Gerald C., Jr.

Hill, James D., Jr.

James D Hill

Gabriel L Suciu

Brian D Merry

William K Ackermann

James W Norris

Nathan A Snape

David R Pack

Frederick M Schwarz

Ioannis Alvanos

Robert J Welch

Julian Partyka

Jesse M Chandler

Jeffrey C Olson

Joel H Wagner

Robert R Steele

Gopal Das

Bernard J Raver

Michael G McCaffrey

Mark F Zelesky

Joseph Brent Staubach

Wesley K Lord

Matthew A Devore

Ram Ranjan

Glenn Levasseur

Cheryl L Kniffin

Katherine McGillivray

James W Shell

Wendell V Twelves

Stephen Karl Kramer

Anthony R Bifulco

Eric A Hudson

John H Welsch

Craig A Nordeen

Shankar S Magge

Willard J Sickles

Lane Mikal Thornton

Allan R Penda

Simon Pickford

Jonathan Ortiz

Michael J Cuozzo

Pitchaiah Vijay Chakka

Gregory M Dolansky

Charles M Lewis

Chris J Niggemeier

Gary M Stetson

Anthony M Zimmerman

Claude R Stogner

Steven J Laporte

Roger F Blinn

Justin J Phillips

Jeffery F Perlak

Jeffrey F Perlak

Jeffery Perlak

Larry R Oliver

Jack Nelson

Gerald C Hollaway

Gabriel L Sucui

Nathan D Snape

William K Ackerman

Shankar S Maggee

A gas turbine engine includes a propulsor with a power turbine, a power turbine shaft extending forward therefrom defining a centerline axis, and a fan driven by the power turbine shaft. The fan is aligned with the centerline axis forward of the power turbine and is operatively connected to be driven by the power turbine through the power turbine shaft. A gas generator operatively connected to the propulsor is included downstream from the fan and forward of the power turbine, wherein the gas generator defines a generator axis offset from the centerline axis. The gas generator is operatively connected to the power turbine to supply combustion products for driving the power turbine.

Offset cores for gas turbine engines

A gas turbine engine comprises a compressor section and a turbine section, the compressor section having a last compressor stage. High pressure cooling air is tapped from a location downstream of the last compressor stage and passed through a heat exchanger. Lower pressure air passes across the heat exchanger to cool the high pressure cooling air. A housing surrounds the compressor section and the turbine section and there being a space radially outwardly of the housing, and a mixing chamber received in the space radially outwardly of the housing, the mixing chamber receiving the high pressure cooling air downstream of the heat exchanger, and further receiving air at a temperature higher than a temperature of the high pressure cooling air downstream of the heat exchanger. Mixed air from the mixing chamber is returned into the housing and utilized to cool at least the turbine section.

External mixing chamber for a gas turbine engine with cooled turbine cooling air

An aircraft thermal management system includes a first fluid system containing a first fluid, a fluid loop containing a thermally neutral heat transfer fluid, a second fluid system containing a second fluid, a first heat exchanger configured to transfer heat from the first fluid to the thermally neutral heat transfer fluid, and a second heat exchanger configured to transfer heat from the thermally neutral heat transfer fluid to the second fluid. The fluid loop is configured to provide the thermally neutral heat transfer fluid to the first heat exchanger at a pressure that matches the pressure of the first fluid.

Aircraft thermal management system

A blade for a gas turbine engine includes a body that includes an airfoil that extends in a radial direction from a 0% span position near an airfoil base to a 100% span position at an airfoil tip. The airfoil has a leading edge and a trailing edge that define the true chord length. The airfoil includes a first portion near the airfoil base with a first density and a second portion near the airfoil tip with a second density. The second density is less than the first density. The second portion includes an increasing true chord length in the radial direction. The second portion is in the range of 90% span to 100% span.

Gas turbine engine blade with variable density and wide chord tip

A rotor assembly of a gas turbine engine may be spoked and includes a rotor and a shell. The rotor has a rotor disk and a plurality of blades each having a platform attached to the rotor disk and with a first channel defined radially between the platforms and the rotor disk. The shell projects aft of the rotor and includes inner and outer walls with a passage defined therebetween. The passage is in fluid communication with the first channel and, together, form part of a secondary flowpath for cooling of adjacent components. The rotor assembly may further include a structure located radially inward of the rotor disk and shell. The structure defines a supply conduit for flowing air from the passage and into a rotor bore defined at least in part by adjacent rotor disks. The entering air, being pre-heated when flowing through the channel and passage, warms the bore and reduces thermal gradients, thus thermal fatigue, across the rotor disk.

Secondary flowpath system for a gas turbine engine

A turbine injection system for a gas turbine engine includes a first end operable to receive air from a heat exchanger, a second end operable to distribute mixed cooling air to a turbine stage, an opening downstream of said first end and a mixing plenum downstream of said first end and said opening. The opening provides a direct fluid pathway into said turbine injection system.

Turbine stage cooling

A gas turbine engine according to the present disclosure includes a first compressor and a first turbine for driving the first compressor. A core section includes a second compressor and a second turbine for driving the second compressor. A third turbine is arranged fluidly downstream of the first turbine and the second turbine and configured to drive a power take-off. A first duct system is arranged fluidly between the low-pressure compressor and the core section. The first duct system is arranged to reverse fluid flow before entry into the core section.

Advanced distributed engine architecture-design alternative

A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a first compressor having a first overall pressure ratio, and a second compressor having a second overall pressure ratio. A ratio of the first overall pressure ratio to the second overall pressure ratio is greater than or equal to about 2.0. Further, a section of the gas turbine engine includes a thermally isolated area.

Gas turbine engine having section with thermally isolated area

A thermal energy exchange system for cooling air of a gas turbine engine includes a heat exchanger located at a diffuser of the gas turbine engine. The diffuser is positioned axially between a compressor and a combustor of the gas turbine engine. A fuel source is operably connected to the heat exchanger to direct a flow of fuel through the heat exchanger via a fuel pipe and toward a fuel nozzle of the combustor. An airflow inlet directs a cooling airflow through the heat exchanger to reduce an airflow temperature via thermal energy exchange between the cooling airflow and the flow of fuel. An airflow outlet directs the cooling airflow from the heat exchanger toward one or more of components of the turbine to cool the one or more components.

Diffuser mounted fuel-air heat exchanger

An assembly includes a valve housing and a valve element such as a poppet valve. The valve housing includes a tubular duct and an annular valve seat disposed within the tubular duct. A first flowpath includes an inner bore of the annular valve seat. A second flowpath includes an aperture formed between the annular valve seat and the tubular duct. The poppet valve is configured to engage the annular valve seat and substantially close the first flowpath when the poppet valve is in a first position. The poppet valve is further configured to disengage the annular valve seat and at least partially open the first flowpath when the poppet valve is in a second position.

Multi-flowpath fluid control valve

An airfoil may comprise an airfoil body having a leading edge and a trailing edge. A heat pipe may be disposed within the airfoil. The heat pipe may include a vaporization section and a condensation section. The vaporization section may be disposed within the airfoil body and may be configured to remove heat from the trailing edge. The second cooling apparatus may be disposed within the airfoil body and may be configured to remove heat from the leading edge.

Hybrid airfoil cooling

Air mixing systems for gas turbine engines include a heat exchanger, a first extraction conduit fluidly coupled to an inlet of the heat exchanger, a second extraction conduit fluidly coupled to an outlet of the heat exchanger, an injection conduit fluidly coupled to the second extraction conduit, an onboard injector supply chamber fluidly coupled to the injection conduit, and an onboard injector fluidly coupled to the onboard injector supply chamber.

Air mixing systems having mixing chambers for gas turbine engines

An inlet manifold for a multi-tube pulse detonation engine includes a vaneless diffuser disposed in a first zone to collect an air discharged from a compressor; a vaned diffuser including a plurality of guide vanes disposed in a second zone to slow the air from the compressor; a plenum disposed in a third zone located next to second zone to provide the air from the compressor to chambers; and a splitter disposed in a fourth zone to split the air from the compressor into an airflow required by each pulse detonation tube for detonation.

Inlet manifold for multi-tube pulse detonation engine

A turbine casing may comprise a casing body a heat pipe disposed in the casing body. The heat pipe may include a vaporization section and a condensation section. The vaporization section may be located forward the condensation section. The vaporization section may be located in a high pressure turbine region of the casing body. The condensation section may be located in a low pressure turbine region of the casing body.

Turbine casing cooling structure

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Tolland, CT, United States of America

James D Hill