Siemens Energy, Inc.

Siemens Aktiengesellschaft

Siemens Westinghouse Power Corporation

Other

Siemens Energy Global Gmbh & Co. Kg

Siemens Canada Limited

David James Wiebe

Timothy A Fox

Richard Colin Charron

Nicholas F Martin, Jr

Jerome H Katy

Alexander Ralph Beeck

Matthew D Montgomery

Stephen A Ramier

Robert H Bartley

Daniel W Garan

Adam Wallace

Mark L Adamson

Yadollah Naghian

George P Liang

Jay Alan Morrison

Yevgeniy P Shteyman

Robert James Bland

David Allen Little

Walter Ray Laster

William Richard Ryan

John Carl Glessner

Jose L Rodriguez

Todd A Ebert

Reinhard Schilp

James S Marshall

Robert T Brooks

Malberto Gonzalez

David M Ritland

Cheryl A Schopf

Walter Marussich

Bruce Loren Smith

Brian J Wessell

Andrew Carlson

Richard Seleski

Michael H Koenig

John W Finneran

Jonathan M Leagon

Brian Harry Terpos

Adam C Pela

Robert Watson Spitzer

Jeffery W Samuelson

Erick J Deane

David D North

Benjamin G Hettinger

Scott W Smith

Harry Patat

Brian Micklos

Christopher Scogin

Kyle C Stoker

Lawrence P Budnick

Jr Nicholas F Martin

David Dealba North

A flexible damper () for turbine blades () includes a plurality of segments () positioned together in a substantially linear pattern, each segment () having a first side (), a second side (), a top side (), a bottom side (), a length (), a width (), and a thickness ().

Flexible damper for turbine blades

A removably attachable snubber assembly for turbine blades includes a turbine blade airfoil including a trailing edge and a leading edge joined by a pressure side and a suction side to provide an outer surface extending in a radial direction to a tip. At least one snubber attachment platform is integrally formed onto the outer surface of the turbine blade airfoil. The at least one snubber attachment platform includes an interlocking mechanism. A snubber is removably attachable to the at least one snubber attachment platform, the snubber including a first end, a second end, a trailing edge, a leading edge, a snubber length, and a snubber width. The snubber also includes a removable attachment mechanism on at least one of the first end and the second end that connects with the interlocking mechanism on the at least one snubber attachment platform.

Removably attachable snubber assembly

A method of making an aero-derivative gas turbine engine () is provided. A combustor outer casing () is removed from an existing aero gas turbine engine (). An annular combustor () is removed from the existing aero gas turbine engine. A first row of turbine vanes () is removed from the existing aero gas turbine engine. A can annular combustor assembly () is installed within the existing aero gas turbine engine. The can annular combustor assembly is configured to accelerate and orient combustion gasses directly onto a first row of turbine blades of the existing aero gas turbine engine. A can annular combustor assembly outer casing () is installed to produce the aero-derivative gas turbine engine (). The can annular combustor assembly is installed within an axial span () of the existing aero gas turbine engine vacated by the annular combustor and the first row of turbine vanes.

Method of making an aero-derivative gas turbine engine

A turbine assembly having an outer casing (), an inner structural ring (), and an annular gas path () defined between outer and inner flow path walls () for conducting a gas flow through the turbine assembly. A plurality of structural struts () are spaced apart in a circumferential direction, each strut () including a strut body () extending in a radial direction for supporting the inner structural ring () to the outer casing (). A first strut end () at a radially outer end of the strut body () is detachably attached to the outer casing () with a first fastener structure () engaging the outer casing (), and a second strut end () at a radially inner end of the strut body () is detachably attached to the inner structural ring () with a second fastener structure () engaging the inner structural ring ().

Turbine assembly with detachable struts

A transition duct system () for delivering hot-temperature gases from a plurality of combustors in a combustion turbine engine is provided. The system includes an exit piece () for each combustor. The exit piece may include an arcuate connecting segment (). An arcuate ceramic liner () may be inwardly disposed onto a metal outer shell () along the arcuate connecting segment of the exit piece. Structural arrangements are provided to securely attach the ceramic liner in the presence of substantial flow path pressurization. Cost-effective serviceability of the transition duct systems is realizable since the liner can be readily removed and replaced as needed.

Transition duct system with arcuate ceramic liner for delivering hot-temperature gases in a combustion turbine engine

A transition duct system for routing a gas flow in a combustion turbine engine is provided. The transition duct system includes one or more converging flow joint inserts forming a trailing edge at an intersection between adjacent transition ducts. The converging flow joint insert may be contained within a converging flow joint insert receiver and may be disconnected from the transition duct bodies by which the converging flow joint insert is positioned. Being disconnected eliminates stress formation within the converging flow joint insert, thereby enhancing the life of the insert. The converging flow joint insert may be removable such that the insert can be replaced once worn beyond design limits.

Converging flow joint insert system at an intersection between adjacent transitions extending between a combustor and a turbine assembly in a gas turbine engine

A turbine airfoil is provided with at least one insert positioned in a cavity in an airfoil interior. The insert extends along a span-wise extent of the turbine airfoil and includes first and second opposite faces. A first near-wall cooling channel is defined between the first face and a pressure sidewall of an airfoil outer wall. A second near-wall cooling channel is defined between the second face and a suction sidewall of the airfoil outer wall. The insert is configured to occupy an inactive volume in the airfoil interior so as to displace a coolant flow in the cavity toward the first and second near-wall cooling channels. A locating feature engages the insert with the outer wall for supporting the insert in position. The locating feature is configured to control flow of the coolant through the first or second near-wall cooling channel.

Turbine airfoil having near-wall cooling insert

A transition duct system () for delivering hot-temperature gases from a plurality of combustors in a combustion turbine engine is provided. The system includes an exit piece () for each combustor. The exit piece may include a straight path segment () for receiving a gas flow from a respective combustor. A straight ceramic liner () may be inwardly disposed onto a metal outer shell () along the straight path segment of the exit piece. Structural arrangements are provided to securely attach the ceramic liner in the presence of substantial flow path pressurization. Cost-effective serviceability of the transition duct systems is realizable since the liner can be readily removed and replaced as needed.

Transition duct system with straight ceramic liner for delivering hot-temperature gases in a combustion turbine engine

A transition duct system () for delivering hot-temperature gases from a plurality of combustors in a combustion turbine engine is provided. The system includes an exit piece () for each combustor. The exit piece may include a straight path segment () and an arcuate connecting segment (). A respective straight metal liner () and an arcuate metal liner () may be each inwardly disposed onto a metal outer shell () along the straight path segment and the arcuate connecting segment () of the exit piece. Structural arrangements are provided to securely attach the respective liners in the presence of substantial flow path pressurization. Cost-effective serviceability of the transition duct systems is realizable since the liners can be readily removed and replaced as needed.

Transition duct system with metal liners for delivering hot-temperature gases in a combustion turbine engine

A gas turbine engine ducting arrangement (), including: an annular chamber () configured to receive a plurality of discrete flows of combustion gases originating in respective can combustors and to deliver the discrete flows to a turbine inlet annulus, wherein the annular chamber includes an inner diameter () and an outer diameter (); an outer diameter mounting arrangement () configured to permit relative radial movement and to prevent relative axial and circumferential movement between the outer diameter and a turbine vane carrier (); and an inner diameter mounting arrangement () including a bracket () secured to the turbine vane carrier, wherein the bracket is configured to permit the inner diameter to move radially with the outer diameter and prevent axial deflection of the inner diameter with respect to the outer diameter.

Gas turbine structural mounting arrangement between combustion gas duct annular chamber and turbine vane carrier

A circumferentially extending sealing band is located within sealing band receiving slots formed in adjacent turbine engine disks. The sealing band includes a plurality of seal strips forming overlap joints defined by overlapping end portions, each formed by a tongue portion extending from a seal face of one seal strip past a seal face of the adjacent seal strip, along a radially inward facing side of the adjacent seal strip. A joint gap is defined within at least one overlap joint between the seal face of the one seal strip and the seal face of the adjacent strip, and a spacer is affixed to the one seal strip and is located at a position within the joint gap between the seal faces to limit circumferential movement of the seal faces toward each other.

Belly band seal with circumferential spacer

A casing for a can annular gas turbine engine, including: a compressed air section () spanning between a last row of compressor blades () and a first row of turbine blades (), the compressed air section () having a plurality of openings () there through, wherein a single combustor/advanced duct assembly () extends through each opening (); and one top hat () associated with each opening () configured to enclose the associated combustor/advanced duct assembly () and seal the opening (). A volume enclosed by the compressed air section () is not greater than a volume of a frustum () defined at an upstream end () by an inner diameter of the casing at the last row of compressor blades () and at a downstream end () by an inner diameter of the casing at the first row of turbine blades ().

Casing for a gas turbine engine

A combustor apparatus defining a combustion zone where air and fuel are burned to create high temperature combustion products. The combustor apparatus comprises an outer wall including a fuel inlet opening for receiving a fuel feed pipe. A coupling assembly is engaged with the fuel feed pipe at the fuel inlet opening to attach the fuel feed pipe to the outer wall. A fuel injection system is located in the interior volume of the outer wall and comprises fuel supply structure including a fuel feed block having a fuel intake passage aligned with the outlet portion of the fuel feed pipe. A coupling fastener is engaged against an exterior outer face of the fuel feed block to create a sealed coupling for containing fuel passing from the fuel feed pipe into the fuel feed block, and to secure the fuel feed block relative to the coupling assembly.

Combustor apparatus in a gas turbine engine

A turbine exhaust casing having an outer casing, an inner casing, an annular exhaust gas path defined between outer and inner flow path walls, and a turbine exhaust casing cavity located radially outward and radially inward from the gas path. A plurality of structural struts support the inner casing to the outer casing, and a fairing surrounds each of the struts in an area extending between the outer and inner flow path walls. A first purge air path extends through at least one of the struts for conducting purge cooling air radially inward to the inner casing, and a second purge air path extends through the strut for further conducting the purge cooling air radially outward to provide a flow of purge air to a location of the exhaust casing cavity radially outward from the outer flow path wall.

Purge and cooling air for an exhaust section of a gas turbine assembly

A combustor apparatus defines a combustion zone where air and fuel are burned to create high temperature combustion products. The combustor apparatus includes an outer wall, coupling structure on the outer wall adjacent to a fuel inlet opening thereof, a fuel injection system, a fuel feed assembly, and a fitting member. The fuel injection system provides fuel to be burned in the combustion zone. The fuel supply structure includes a threaded inner surface formed from a first material. The fuel feed assembly includes a fuel feed pipe that extends through the fuel inlet opening in the outer wall and has an outlet portion formed from the first material and that is threadedly engaged with the fuel supply structure, and an inlet portion affixed to the outlet portion and formed from a second material. The fitting member secures the fuel feed assembly relative to the outer wall.

Growing community of inventors

Orlando, FL, United States of America

David James Wiebe