General Electric Company

Siemens Energy, Inc.

Siemens Aktiengesellschaft

Mikro Systems, Inc.

Siemens Energy Global Gmbh & Co. Kg

Other

United States of America As Represented by the Secretary of the Air Force

Siemens Aktiengesellshaft

Mikro Sysytems, Inc.

Ching-Pang Lee

Kok-Mun Tham

Robert Alan Johnson

Chander Prakash

Aspi Rustom Wadia

Nesim Abuaf

Gm Salam Azad

Ramgopal Darolia

David Glenn Cherry

Ronald Scott Bunker

Robert Joseph Orlando

Steven Robert Brassfield

Wayne Charles Hasz

Paul Hadley Vitt

John Joseph Marra

Hsin-Pang Wang

Robert Edward Schafrik

Kattalaicheri Srinivasan Venkataramani

Eric Alan Estill

Jay Alan Morrison

Vincent Paul Laurello

Kevin Samuel Klasing

Thomas Ory Moniz

Brian Keith

Benjamin E Heneveld

Steven Koester

Wenfeng Lu

Eric Joseph Schroeder

Gary Brian Merrill

Melvin Robert Jackson

Bin Wei

William Randolph Stowell

John Howard Starkweather

Glenn Herbert Nichols

Nicholas F Martin, Jr

Jan Christopher Schilling

Stephen Joseph Ferrigno

Harvey M Maclin

Bangalore Aswatha Nagaraj

Jeffrey Donald Clements

Joseph David Rigney

Marc Thomas Edgar

Ramesh Subramanian

John Frederick Ackerman

Christian X Campbell

Lawrence Butler

Andrew Jeremiah Burns

Jeremy Clyde Bailey

Paul Stuart Wilson

Darek Tomasz Zatorski

Jon Conrad Schaeffer

Lawrence Bernard Kool

Canan Uslu Hardwicke

Ji-Cheng Zhao

Andrew Jay Skoog

Tingfan Pang

Jane Ann Murphy

Kevin Richard Kirtley

Jiping Zhang

Monty Lee Shelton

Matthew D Montgomery

Reinhard Schilp

Thomas Walter Rentz

Harold Ray Hansel

Brian Alan Norton

Darryl S Eng

Marshall Gordon Jones

Venkat Subramaniam Venkataramani

Andrew Breeze-Stringfellow

Kurt David Murrow

George A Coffinberry

Mark Daniel Gorman

Todd Jay Rockstroh

Ann Melinda Ritter

David Edwin Budinger

Richard William Albrecht, Jr

Robert J Corsmeier

Yevgeniy P Shteyman

Peter Schröder

Richard Colin Charron

David Allen Little

Scott Hunter

Ambrose A Hauser

Melvin Bobo

Patrick Jarvis Scoggins

Liang Jiang

Gary L Leonard

Erich Alois Krammer

John Peter Heyward

Gulcharan Singh Brainch

Gary Eugene Wheat

Rajasingh Schwartz Israel

John Thomas Leman

John Christopher Brauer

Christopher W Ross

Leslie Eugene Leeke, Jr

Erwan Baleine

Wilbur Douglas Scheidt

Michael Joseph Danowski

Stefan Tschirren

A turbine vane in a gas turbine engine includes an inner platform, an outer platform, and a vane airfoil positioned therebetween. The vane airfoil includes a first cooling passage extending between the outer platform and the inner platform, and a second cooling passage extending between the outer platform and the inner platform. The second cooling passage is arranged downstream of the first cooling passage with respect to a flow direction. The turbine vane includes a jumper tube disposed between the second cooling passage and the inner platform. The jumper tube includes an inlet, an outlet, and a tube wall enclosing a hollow interior. The inlet is positioned a distance within the second cooling passage. The outlet is positioned at least partially through an aperture of the inner platform.

Turbine vane in gas turbine engine

A turbine airfoil includes a trailing edge coolant cavity between a pressure sidewall and a suction sidewall. The trailing edge coolant cavity is positioned adjacent to and extending out to a trailing edge of the turbine airfoil. The interior includes an internal arrangement comprising an array of discrete fins formed aft of the trailing edge coolant cavity along the trailing edge.

Turbine airfoil with trailing edge features and casting core

A flow inducer assembly and a method for cooling turbine blades of a gas turbine engine are presented. The gas turbine engine includes a rotor disk having circumferentially distributed disk grooves and turbine blades. Each turbine blade includes a blade root inserted into blade mounting section of the disk groove. Seal plates are attached to an aft side circumference of the rotor disk. The flow inducer assembly is integrated to each seal plate at a side facing away from the rotor disk. The flow inducer assembly is configured to function as a paddle due to rotation of the rotor disk and the seal plate therewith during operation of the gas turbine engine to drive ambient air as a cooling fluid into the disk cavity and enter inside of the turbine blade from the blade root for cooling the turbine blade.

Cover plate with flow inducer and method for cooling turbine blades

A blade () for a turbine engine that includes an internal cooling system () formed from at least one cavity () positioned within a generally elongated airfoil (). A squealer tip () and at least one densified oxide dispersion strengthened layer () extend radially from a radially outer tip cap () of the blade (), the tip cap () having a tip cap upper surface ().

Turbine blade with squealer tip and densified oxide dispersion strengthened layer

A trailing edge cooling feature for a turbine airfoil () includes a plurality of pins (-) positioned in an airfoil interior () toward the trailing edge), each extending from the pressure side () to the suction side () and further being elongated in a radial direction (R). The pins (-) are arranged in multiple radial rows (A-L) spaced along the chordal axis (), with the pins (-) in each row (A-L) being interspaced to define coolant passages (-) therebetween. A row of radially spaced apart partition walls () are positioned aft of the pins (-). Each partition wall () extends from the pressure side () to the suction side () and is elongated in a generally axial direction, extending along the chordal axis () to terminate at the trailing edge (). Axially extending coolant exit slots () are defined in the interspaces between adjacent partition walls (-) that direct coolant exiting a last row (L) of pins () to be discharged from the airfoil () into a hot gas path.

Turbine airfoil with trailing edge cooling featuring axial partition walls

A turbine airfoil () includes a trailing edge coolant cavity () located in an airfoil interior () between a pressure sidewall () and a suction sidewall (). The trailing edge coolant cavity () is positioned adjacent to a trailing edge () of the turbine airfoil () and is in fluid communication with a plurality of coolant exit slots () positioned along the trailing edge (). At least one framing passage () is formed at a span-wise end of the trailing edge coolant cavity (). The airfoil () further includes framing features (A-B,A-B) located in the framing passage (). The framing features are configured as ribs (A-B,A-B) protruding from the pressure sidewall () and/or the suction sidewall (). The ribs (A-B,A-B) extend partially between the pressure sidewall () and the suction sidewall ().

Turbine airfoil with trailing edge framing features

A blade airfoil for a turbine engine that includes an internal multiple pass serpentine flow cooling circuits with a leading edge circuit and a trailing edge circuit. The entrance of a cavity in the leading edge circuit has a narrowing of a cavity width that expands further downstream to a consistent cavity width similar to the cavity width of the rest of the leading edge circuit.

Turbine rotor airfoil and corresponding method for reducing pressure loss in a cavity within a blade

An integrated airfoil and platform cooling system () for a turbine rotor blade () includes an inlet () located at the root () for receiving a supply of a coolant (K), and at least one cooling leg () fluidly connected to the inlet () and configured for conducting the coolant (K) in a radially outboard direction. The cooling leg () is defined at least partially by a span-wise extending internal cavity () within a blade airfoil (). An entrance of the cooling leg () comprises a flow passage () that extends radially outboard and laterally into a blade platform (), so as to direct a radially outboard flowing coolant (K) to impinge on an inner side () of a radially outer surface () of the blade platform (), before leading the coolant (K) into the cooling leg ().

Turbine rotor blade with airfoil cooling integrated with impingement platform cooling

A shroud assembly for a turbine engine includes a seal for sealing a gap between a first mate face of a first shroud segment and a second mate face of a circumferentially adjacent second shroud segment. The seal is received in first and second slots formed respectively on the first and second mate faces. The first and second slots extend axially between a leading edge and a trailing edge of the respective shroud segment. The first slot is open at the leading and the trailing edges while the second slot is open at the leading edge and closed at the trailing edge. The seal has axially extending first and second sides which are receivable respectively within the first and second slots. The seal has an axial length substantially equal to tan axial length of the shroud segments and has a cutout on the second side at a trailing edge end of the seal.

Sealing arrangement between turbine shroud segments

A gas turbine engine airfoil includes an outer wall including a suction side, a pressure side, a leading edge, and a trailing edge, the outer wall defining an interior chamber of the airfoil. The airfoil further includes cooling structure provided in the interior chamber. The cooling structure defines an interior cooling cavity and includes a plurality of cooling fluid outlet holes, at least one of which is in communication with a pressure side cooling circuit and at least one of which is in communication with a suction side cooling circuit. At least one of the pressure and suction side cooling circuits includes: a plurality of rows of airfoils, wherein radially adjacent airfoils within a row define segments of cooling channels. Outlets of the segments in one row align aerodynamically with inlets of segments in an adjacent downstream row such that the cooling channels have a serpentine shape.

Cooling configuration for a gas turbine engine airfoil

A core structure () includes a first core element () including a leading edge section (), a tip section (), and a turn section () joining the leading edge and tip sections (). The first core element () is adapted to be used to form a leading edge cooling circuit () in a gas turbine engine airfoil (). The leading edge cooling circuit () includes a cooling fluid passage () having a leading edge portion () formed by the first core element leading edge section (), a tip portion () formed by the first core element tip section (), and a turn portion () formed by the first core element turn section (). Each of the leading edge portion (), the tip portion (), and the turn portion () of the cooling fluid passage () are formed concurrently in the airfoil () by the first core element ().

Gas turbine airfoil including integrated leading edge and tip cooling fluid passage and core structure used for forming such an airfoil

A turbine component including a shrouded airfoil with a flow conditioner configured to direct leakage flow and coolant to be aligned with main hot gas flow is provided. The flow conditioner is positioned on a shroud base radially adjacent to the tip of the airfoil and includes a ramped radially outer surface positioned further radially inward than a radially outer surface of the shroud base. The ramped radially outer surface extends from a first edge to a second edge in a direction generally from the suction side to the pressure side of the airfoil, such that the first edge is positioned further radially inward than the second edge. Multiple coolant ejection holes are positioned on the ramped radially outer surface. The coolant ejection holes are connected fluidically to an interior of the airfoil.

Shrouded turbine blade

A compressor () configured for use in a gas turbine engine () and having a rotor assembly () with a pumping system () positioned on a rotor drum () to counteract reverse leakage flow at a gap () formed between one or more stator vane tips () and a radially outer surface () of the rotor drum (). The pumping system () may be from pumping components () positioned radially inward of one or more stator vane tips () to reduce, if not completely eliminate, reverse leakage flow at the stator vane tips (). In at least one embodiment, the pumping component () may be formed from one or more cutouts () in the outer surface () of the rotor drum (). In another embodiment, the pumping component () may be formed from at least one pumping fin () extending from the radially outer surface () of the rotor drum (). In at least one embodiment, rows () of pumping components () may be aligned with rows () of stator vanes () within the compressor ().

Compressor usable within a gas turbine engine

Turbine and compressor casing/housing abradable component embodiments for turbine engines, have abradable surfaces with ridges projecting from the abradable surface, separated by grooves. The ridges have one or both sidewalls inclined against the opposing turbine blade tip rotational direction for redirecting and/or dissipating blade tip gap leakage airflow energy. In some embodiments the ridge tip and/or groove base have inclined profiles for redirecting airflow leakage away from the blade tip gap. In some embodiments, the inclined ridge tip profile provides a progressive wear zone that increases abradable surface area as the inclined ridge is abraded by the rotating blade tip.

Turbine abradable layer with inclined angle surface ridge or groove pattern

A turbine blade including a pressure sidewall () and a suction sidewall (), and at least one partition rib () extends between the pressure and suction sidewalls () to define a serpentine cooling path () having adjacent cooling channels () extending in the spanwise direction (S) within the airfoil (). A flow turning guide structure () extends around an end of the at least one partition rib () and includes a first element () extending from the pressure sidewall () to a lateral location in the cooling path between the pressure and suction sidewalls (), a second element () extending from the suction sidewall () to the lateral location in the cooling path between the pressure and suction sidewalls (). The first and second elements () include respective distal edges () that laterally overlap each other at the lateral location.

Turbine blade with a non-constraint flow turning guide structure

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Cincinnati, OH, United States of America

Ching-Pang Lee