Alstom Technology Limited

Doosan Enerbility Co., Ltd.

Ansaldo Energia Switzerland Ag

Alstom (switzerland) Ltd

General Electric Technology Gmbh

Alstom

Ulrich Rathmann

Joerg Krueckels

Willy Heinz Hofmann

Herbert Brandl

Erhard Kreis

Shailendra Naik

Naresh Aluri

Alexander Ralph Beeck

Alexander Beeck

Beat Von Arx

Christoph Nagler

Ulrich Robert Steiger

Thomas Heinz-Schwarzmaier

Stefan Schlechtriem

Sacha Parneix

Helene Saxer-Felici

Shailandra Naik

Hans Wettstein

Reinhard Fried

Maxim Konter

Michael Thomas Maurer

Carlos Simon-Delgado

Franklin Marie Genin

Peter Marx

Klaus Knapp

Markus Oehl

Peter Ernst

Armin Busekros

Ulrich Wellenkamp

Igor Tsypkaykin

Daniel Burri

Bruno Benedetti

Andre Saxer

Roland Dueckershoff

Christof Pfeiffer

Thorsten Christoph Motzkus

Ronald Wifling

Kaspar Loeffel

Nicolas Tran

Michael Huber

Andreas Kieninger

Stefan Florjancic

Willy H Hoffmann

Holger Lübcke

Jan Ehrhard

Igor Tsypkaykine

Seok Beom Kim

Willy Hoffmann

Roland DÜCKERSHOFF

Thorsten Motzkus

Hans E Wettstein

A flow body for a gas turbine includes an airfoil extending along a radial direction between a platform end and a tip which has a tip surface. The airfoil is formed of a first metal material and comprises an inner cavity for receiving a gaseous cooling fluid. The flow body further includes a squealer tip protruding from the tip surface of the tip and extending along a circumference of the tip so that the squealer tip at least partially surrounds the tip surface. The squealer tip is formed from a second metal material and includes a plurality of internal cooling cavities that are separated from each other within the squealer tip, wherein each of the internal cooling cavities is in fluid communication with the inner cavity via one or more fluid passages.

Flow body for a gas turbine, gas turbine, method for manufacturing a flow body for a gas turbine, and method for repairing a flow body of a gas turbine

A method for manufacturing a blade for a gas turbine includes forming a blade body, forming a groove in an outer surface of the blade body, positioning a cover on the blade body such that it covers the groove and such that an outer surface of the cover forms a continuous surface with the outer surface of the blade body, and joining the cover to the blade body so that the cover and the groove define a cooling channel.

Method for manufacturing a blade for a gas turbine, turbine blade and gas turbine

The present technique presents a gas turbine blade for re-using cooling air, a turbomachine assembly having the blade, and a gas turbine having the turbomachine assembly. The blade includes a platform and an airfoil extending from the platform. The airfoil includes a pressure surface, a suction surface, a leading edge and a trailing edge. The platform includes a pressure side, a suction side, a leading-edge side and a trailing-edge side, disposed towards the pressure surface, the suction surface, the leading edge and the trailing edge of the airfoil, respectively. The suction side of the platform includes a part of the upper surface and a suction-side lateral surface of the platform. At least a part of an edge between the suction-side lateral surface and the upper surface of the platform comprises a chamfer part.

Gas turbine blade for re-using cooling air and turbomachine assembly and gas turbine comprising the same

A turbine blade is provided. The turbine blade may include an airfoil having an airfoil tip, a leading edge, a trailing edge, and a pressure side and a suction side extending from the leading edge to the trailing edge and defining an airfoil cavity, a squealer tip arranged at the airfoil tip part and comprising a trailing edge tip portion disposed at the trailing edge of the airfoil and a pressure side rail and a suction side rail meeting at the trailing edge tip portion and defining a squealer tip pocket at the airfoil tip, and at least one tip cooling hole disposed at the squealer tip pocket to provide cooling air from the airfoil cavity to the squealer tip pocket, wherein the trailing edge tip portion of the squealer tip includes a chamfer disposed towards the pressure side of the airfoil and a groove extending from the squealer tip pocket to the chamfer to provide cooling air from the squealer tip pocket to the chamfer.

Trailing edge tip cooling of blade of a gas turbine blade

The present invention provides an airfoilwith the squealer tip cooling systemfor a turbine bladeat the blade tip, wherein the squealer tip cooling systemcomprises a cooling passagearranged within a squealer tip, wherein the cooling passageat least partly extends toward a terminal endof the squealer tip, and a pocketat a lateral surfaceof the squealer tip, open externally and extending inwardly at least partly across the cooling passage. The pocketintersects the cooling passageand the pocketcomprises an impingement surfacefacing the cooling passage, on which a cooling medium expelled through the cooling passageimpinges before being discharged externally through the pocket

Airfoil with a squealer tip cooling system for a turbine blade, a turbine blade, a turbine blade assembly, a gas turbine and a manufacturing method

The present technique presents a bladefor a gas turbine. The bladeincludes an airfoilhaving an airfoil tip partand a pressure sideand a suction sidemeeting at a leading edgeand a trailing edgeand defining an internal spaceof the airfoil. A squealer tipis arranged at the airfoil tip part. The squealer tipcomprises a suction side rail. The suction side railcomprises a chamfer partand at least one squealer tip cooling hole. The chamfer partcomprises a chamfer surface. An outletof the at least one squealer tip cooling holeis disposed at the chamfer surface

Technique for cooling squealer tip of a gas turbine blade

The invention relates to a can-combustor for a can-annular combustor arrangement in a gas turbine. The can combustor includes an essentially cylindrical casing with an axially upstream front panel and an axially downstream outlet end. The can combustor further includes a number of premixed burners, extending in an upstream direction from said front panel and having a burner exit, supported by this front panel, for supplying a fuel/air mixture into a combustion zone inside the casing. Up to four premixed burners are attached to the front panel in a substantially annular array. Each burner has a conical swirl generator and a mixing tube to induce a swirl flow of said fuel/air mixture.

Can combustor for a can-annular combustor arrangement in a gas turbine

A combustor arrangement with a front panel, a combustor liner, and a carrier structure element is provided for carrying the front panel and the combustor liner, wherein the combustor arrangement further includes a fastening system for connecting the front panel, the combustor liner, and the carrier structure element to one another. The fastening system includes at least one elastic connection element, the latter being fixedly connected to the carrier structure element and extending therefrom to the combustor liner and to the front panel. The elastic connection element is further fixedly connected to the combustor liner and/or the front panel such as to clamp the front panel, the combustor liner, and the carrier structure element to one another in a substantially fluid tight manner.

Combustor arrangement with fastening system for combustor parts

A front panel for a combustor has a hot side and a cold side and at least one reception adapted for receiving a combustor part. The front panel has a double-wall design with a hot-side wall and a cold-side wall. The hot-side wall defines a hot-side downstream surface of the front panel. The cold-side wall defines a cold-side upstream surface of the front panel. The hot-side wall and the cold-side wall are axially spaced from one another, extend parallel to one another, and are connected to one another by an outer side wall.

Combustor front panel

The invention relates to a transition piece assembly for a gas turbine. A transition piece having one end adapted for connection to a gas combustor and an opposite end adapted for connection to a first turbine stage. The transition piece having at least one external liner and at least one internal liner. The internal liner forms the hot gas flow channel. A first section of the transition piece assembly upstream of a first turbine stage has a plurality of cooling apertures. Cooling medium through the cooling apertures enters the plenum, which is created between external and internal liners and a cooling medium flows along at least the first section of the transition piece assembly. At least one second section upstream of the first section with respect to the hot gas flow has at least one additional air inlet system.

Combustor cooling structure

A device on a shroud which is provided on a turbine rotor blade tip has a cutting structure designed like a line of ribs, locally projecting over the shroud radially to the rotational axis around which the turbine rotor blade is rotatably arranged. The sealing structure has a longitudinal extent (S) oriented in the direction of rotation (U) of the turbine rotor blade, tapers with increasing radial distance from the shroud, and terminates with a flat end face (St) which radially faces away from the turbine rotor blade, and can be divided into five interrelated surfaces.

Sealing structure on a shroud of a turbine blade

A wall structure for restricting a hot-gas path in a gas turbine or in a combustion chamber including a first wall segment having a first face, and a second wall segment having a second face. Each wall segment disposed adjacent to each other so as to define an inside of the wall segments and an outside of the wall segments. The first and second faces being disposed so as to form a gap leading from the inside to the outside. Each of the faces having a receiving groove. The wall structure further including a sealing element having a thickness and disposed in the receiving grooves so as to bridge the gap. The first face having a recess and the second wall segment having a projection that extends from the second face and protrudes into the recess in the first face. The projection has a tapered cross section.

Wall structure for limiting a hot gas path

A lance of a gas turbine burner includes an inner duct having inner nozzles; an annular outer duct encircling the inner duct and having outer nozzles; and a spacer disposed between a terminal portion of the inner duct and a terminal portion of the outer duct, the spacer being fixed to both the inner duct and the outer duct.

Lance of a gas turbine burner

A strip seal and method of configuration thereof for sealing two adjacent non-rotating gas turbine hot gas components exposed to pressure pulsations. The strip seal has at least two clamping projections distributed along discrete points of the strip seal length that extend out from a pressure face of the strip seal. The location of the projections are defined by two ratios. The first ratio, of less than 25, is the strip seal length extending free of clamping projections from any one of the ends of the strip seal to a clamping projection to the ratio of strip seal thickness. The second ratio, of less than 200, is the ratio of the strip seal length extending free of clamping projections between any two projections to the thickness of the strip seal. This seal arrangement ameliorates detrimental effects of induced resonance.

Strip seal and method for designing a strip seal

An arrangement between blade elements in a blade row in a turbine is described. Each blade element has at least one shroud element and a blade airfoil which abuts on, and is connected to, the shroud element, and essentially extends radially with regard to a principal axis of the blade row. When installed, the shroud element sides, which extend circumferentially, abut on the respectively adjacent shroud element of the respectively adjacent blade element, each forming an essentially radial gap. At least one blade element has a projection which projects into the shroud element of the abutting blade element and extends in the circumferential direction, and at least one blade element has a recess which accommodates such a projection. In the region of the projection or recess there is a stepped region of the radial gap, and the guiding of the radial gap in this stepped region is a labyrinth seal.

Shiplap arrangement

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Baden, Switzerland

Ulrich Rathmann