Other

Siemens Energy, Inc.

Doosan Enerbility Co., Ltd.

Siemens Aktiengesellschaft

S & J Deisgn, LLC

Darryl S Eng

David M Day

Christian X Campbell

Jinuk Kim

Kwangil Kim

Jeff Greenberg

Ching-Pang Lee

John Joseph Marra

Barry J Brown

Gene Chong

Jason Kopko

JaeBin Lee

Todd A Ebert

Jae Y Um

Daniel O Davies

Andres Jaramillo

Christopher K Rawlings

Friedrich T Rogers

Kibaek Kim

Bryan C Bernier

Thomas F Pechette

Harry Patat

IkSang Lee

Hongseung Seo

YeJee Kim

Hyunwoo Joo

Younggi Mun

A gas turbine nozzle assembly of a gas turbine is provided. The turbine nozzle assembly may include a turbine nozzle extending from an inner diameter to an outer diameter and having an airfoil-shaped cross section having a leading edge and a trailing edge, and a pressure side and a suction side each of which extends from the leading edge to the trailing edge, wherein the turbine nozzle may include a hollow airfoil including a plurality of cavities positioned in the airfoil, an insert positioned in one or more of the plurality of cavities of the hollow airfoil, a plurality of cover plates, at least one of which is positioned at one of the inner diameter and at the outer diameter, and a plurality of impingement pans, at least one of which is positioned at one of the inner diameter and at the outer diameter.

Turbine nozzle assembly and gas turbine including the same

An internal core profile for a turbine nozzle airfoil of a gas turbine is provided. The turbine nozzle may include an airfoil core having an uncoated nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in Table 1, wherein the X, Y, and Z coordinates are distances in inches measured in a Cartesian coordinate system, the corresponding X and Y coordinates, when connected by a smooth continuous arc, define one of a plurality of airfoil core profile sections at each Z distance, and the plurality of airfoil core profile sections, when joined together by smooth continuous arcs, define an airfoil core shape.

Internal core profile for a turbine nozzle airfoil

An internal core profile for a second stage turbine nozzle airfoil of a gas turbine is provided. The turbine nozzle may include an airfoil core having an uncoated nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in Table 1, wherein the X, Y, and Z coordinates are distances in inches measured in a Cartesian coordinate system, the corresponding X and Y coordinates, when connected by a smooth continuous arc, define one of a plurality of airfoil core profile sections at each Z distance, and the plurality of airfoil core profile sections, when joined together by smooth continuous arcs, define an airfoil core shape.

A gas turbine nozzle assembly of a gas turbine is provided. The turbine nozzle assembly may include a turbine nozzle extending from an inner platform to an outer platform and having an airfoil-shaped cross section having a leading edge and a trailing edge, and a pressure side and a suction side each of which extends from the leading edge to the trailing edge, wherein the turbine nozzle may include a plurality of vanes attached to the inner and outer platforms and the inner platform having an attached first and second endfaces and a flow surface surrounding opposing ends of a vane of the plurality of vanes, the flow surface terminating circumferentially at the first and second endfaces and terminating axially at forward and aft edges, and the inner platform may include a platform corner portion comprising the flow surface attached to the first endface at the forward edge and attached to the second endface at the aft edge.

Turbine nozzle and gas turbine including the same

An airfoil profile for a second stage turbine nozzle of a gas turbine is provided. The turbine nozzle may include an airfoil portion having an uncoated nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in Table 1, wherein the X, Y, and Z coordinates are distances in inches measured in a Cartesian coordinate system, the corresponding X and Y coordinates, when connected by a smooth continuous arc, define one of a plurality of airfoil profile sections at each Z distance, and the plurality of airfoil profile sections, when joined together by smooth continuous arcs, define an airfoil shape.

An airfoil profile for a turbine nozzle of a gas turbine is provided. The turbine nozzle may include an airfoil portion having an uncoated nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in Table 1, wherein the X, Y, and Z coordinates are distances in inches measured in a Cartesian coordinate system, the corresponding X and Y coordinates, when connected by a smooth continuous arc, define one of a plurality of airfoil profile sections at each Z distance, and the plurality of airfoil profile sections, when joined together by smooth continuous arcs, define an airfoil shape.

Airfoil profile for a turbine nozzle

A shroud cooling system configured to cool a shroud adjacent to an airfoil within a gas turbine engine is disclosed. The turbine engine shroud may be formed from shroud segments that include a plurality of cooling air supply channels extending through a forward shroud support for impingement of cooling air onto an outer radial surface of the shroud segment with respect to the inner turbine section of the turbine engine. The channels may extend at various angles to increase cooling efficiency. The backside surface may also include various cooling enhancement components configured to assist in directing, dispersing, concentrating, or distributing cooling air impinged thereon from the channels to provide enhanced cooling at the backside surface. The shroud cooling system may be used to slow down the thermal response by isolating a turbine vane carrier from the cooling fluids while still providing efficient cooling to the shroud.

Shroud cooling system for shrouds adjacent to airfoils within gas turbine engines

A floating brush seal in a rim cavity of a turbine in a gas turbine engine, where the floating brush seal includes a seal holder in which the floating brush seal floats, and a expandable seal that fits within two radial extending seal slots that maintains a seal with radial displacement of the floating brush seal and the seal holder.

Turbine with radial acting seal

Platforms () span between turbine blades () on a disk (). Each platform may be individually mounted to the disk by a pin attachment (). Each platform () may have a rotationally rearward edge portion () that underlies a forward portion () of the adjacent platform (). This limits centrifugal bending of the rearward portion of the platform, and provides coolant sealing. The rotationally forward edge (A,B) of the platform overlies a seal element () on the pressure side () of the forwardly adjacent blade, and does not underlie a shelf on that blade. The pin attachment allows radial mounting of each platform onto the disk via tilting () of the platform during mounting to provide mounting clearance for the rotationally rearward edge portion (). This facilitates quick platform replacement without blade removal.

Turbine blade and non-integral platform with pin attachment

A triple hook ring segment including forward, midsection and aft mounting hooks for engagement with respective hangers formed on a ring segment carrier for supporting a ring segment panel, and defining a forward high pressure chamber and an aft low pressure chamber on opposing sides of the midsection mounting hook. An isolation plate is provided on the aft side of the midsection mounting hook to form an isolation chamber between the aft low pressure chamber and the ring segment panel. High pressure air is supplied to the forward chamber and flows to the isolation chamber through crossover passages in the midsection hook. The isolation chamber provides convection cooling air to an aft portion of the ring segment panel and enables a reduction of air pressure in the aft low pressure chamber to reduce leakage flow of cooling air from the ring segment.

Cooling system for three hook ring segment

A turbine airfoil (E-H) extends from a shank (E-H). A platform (E-H) brackets or surrounds a first portion of the shank (E-H). Opposed teeth () extend laterally from the platform (E-H) to engage respective slots () in a disk. Opposed teeth () extend laterally from a second portion of the shank () that extends below the platform (E-H) to engage other slots () in the disk. Thus the platform (E-H) and the shank (E-H) independently support their own centrifugal loads via their respective teeth. The platform may be formed in two portions (E-H,E-H), that are bonded to each other at matching end-walls () and/or via pins (G) passing through the shank (E-H). Coolant channels () may pass through the shank beside the pins (G).

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Palm Beach Gardens, FL, United States of America

Darryl S Eng