Alstom Technology Limited

Ansaldo Energia IP UK Limited

Ansaldo Energia Switzerland Ag

General Electric Technology Gmbh

Doosan Enerbility Co., Ltd.

Alstom (switzerland) Ltd

Alstom

Herbert Brandl

Hans-Peter Bossmann

Joerg Krueckels

Ulrich Rathmann

Willy Heinz Hofmann

Carlos Simon-Delgado

Marc Widmer

Philipp Indlekofer

Beat Von Arx

Rudolf Kellerer

Alexander Ralph Beeck

Matthias Hoebel

Alexander Beeck

Eribert Benz

Brian Kenneth Wardle

Alexander Hoffs

Christoph Nagler

Shailendra Naik

Peter Marx

Robert Marmilic

Hartmut Haehnle

Thomas Zierer

Mikhail Pavlov

Christoph Didion

Sascha Justl

Claus Paul Gerdes

Felix Michael Reinert

Igor Tsypkaykin

Erich Kreiselmaier

Ulrich Wellenkamp

Armin Busekros

Marcel Koenig

Andrey Sedlov

Roland Muecke

Jörg Krückels

Guillaume Wagner

Emanuele Facchinetti

Kurt Rubischon

Joerg Krückels

Magali Cochet

Jiwen Tao

Artem Ivanov

Seok Beom Kim

Brammajyosula Ravindra

Sergej Riazantsev

Rene Koehnke

Arrigo Beretta-Müller

Sven Schofer

Joergen Ferber

Ioannis Paralikas

Igor Tsypkakyin

Laura Bogdanic

Andrey Anatolievich Sedlov

Arthur Mateusz Faflik-Brooks

Andrew Wilson

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 concerns a turbine for a gas turbine comprising a blade, a vane and an abradable lip attached to the blade or to the vane, wherein the blade and the vane are separated by a gap and the abradable lip extends part of the distance across the gap. Embodiments include the addition of an abrasive layer attached on the other side of the gap from the abradable lip. A method of manufacturing is also described.

Abradable lip for a gas turbine

The present disclosure generally relates to a guide vane for a gas turbine, and provides for example an innovative guide vane with improved flexibility leading to a reduction of stresses at the interface between the vane platform and the vane carrier. Exemplary embodiments provide only circumferential line contact or point contact between the guide vane and the guide vane carrier.

Gas turbine vane

Disclosed is a turbo-engine component comprising a wall, the wall comprising a hot gas side surface and a coolant side surface. At least one coolant discharge duct is provided in said wall and opening out onto the hot gas side surface at a coolant discharge opening. A coolant flow direction is defined from the interior of the coolant discharge duct towards the discharge opening, the coolant discharge duct further being delimited by a delimiting surface thereof provided inside the wall. The coolant discharge duct has a first cross sectional direction and a second cross sectional direction. The coolant discharge duct is a blind cavity and is closed towards the coolant side surface, and further a dimension of the coolant discharge duct measured in the first cross sectional direction decreases in the coolant flow direction.

Turbo-engine component having outer wall discharge openings

A blading member for fluid flow machines, for example, gas turbine engines. The blading member includes a platform member, at least one airfoil member, and at least one interlock member.

Blading member for a fluid flow machine

The present invention generally relates to a vane for a gas turbine, and more in particular it provides an innovative vane with improved flexibility leading to a reduction of stresses at the transition from the vane trailing edge to the vane platform, without interfering into the cooling scheme of such component. The present invention can increase flexibility of the vane platform by introducing on the vane platform a material cutback confined in the proximity of the trailing edge portion of the vane airfoil.

A wall for a hot gas channel in a gas turbine is described, the wall having a back side and a front side and an impingement sheet having impingement cooling holes, the wall being for exposure to a hot fluid at the front side, and the wall having an array of pins attached to the back side and extending between the back side and the impingement sheet, the wall additionally having a plurality of ribs attached to the back side, each rib extending between two pins to delineate an array of cells on the back side, and/or at least one compartment wall attached to the back side to delineate compartments on the back side. Embodiments include an impingement sheet with impingement cooling holes and cooling exit holes. A gas turbine including the wall is also described.

Wall for a hot gas channel in a gas turbine

The invention concerns a turbine blade comprising a surface, a recess within the surface, and a damping inlay within the recess. The damping inlay comprises a chamber with a damping material, for example particles. The damping inlay should substantially maintain the aerodynamic profile of the blade to enable normal operation. A further embodiment of the invention describes the method of manufacture of a turbine blade with a damping inlay. The method comprises the steps of manufacturing a turbine blade having a surface and a recess in the surface, and providing one or more damping inlays within the recess such that the damping inlay substantially maintains the aerodynamic profile of the blade, the damping inlay comprising a chamber and a damping material disposed within the chamber.

Damping inlay for turbine blades

The invention relates to a turbine blade, including a blade having a front edge and a rear edge, which blade transitions by means of a shaft into a blade root designed for fastening the turbine blade, and including a platform, which is arranged at the lower end of the blade in order to bound a flow channel. The platform is designed as a separate component and can be connected to the blade in a form-fit manner. Flexible application is achieved in that the platform is composed of several individual platform elements, which enclose the blade in the assembled state.

Turbine blade or vane

A rotor blade arrangement (), especially for a gas turbine, which can be fastened on a blade carrier () and includes in each case a blade aerofoil element () and a platform element (), wherein the platform elements () of a blade row form a continuous inner shroud. With such a blade arrangement, a mechanical decoupling, which extends the service life, is achieved by the blade aerofoil element () and the platform element () being formed as separate elements and by being able to be fastened in each case separately on the blade carrier ().

Rotor blade arrangement and gas turbine

An internally cooled airfoil for a rotary machine, for example, a gas turbine engine includes a suction and pressure side wall each extending in an axial direction, i.e. from a leading to a trailing edge of the airfoil. A suction wall sided cooling channel and a pressure wall sided cooling channel extend in the axial direction. A feed chamber is defined between a first and second inner wall for feeding the suction wall and pressure wall sided cooling channel each by at least one through hole inside of the first and second inner wall. The suction wall sided cooling channel and the pressure wall sided cooling channel extend into the trailing edge region separately. The suction wall sided cooling channel and the pressure wall sided cooling channel join before discharging at the trailing edge.

Internally cooled airfoil for a rotary machine

A cooled wall segment in the hot gas path of a gas turbine. The wall segment includes a first surface, exposed to a medium of relatively high temperature, a second surface, exposed to a medium of relatively low temperature, and side surfaces connecting the first and second surface and defining a height of the wall segment. At least one cooling channel for a flow-through of a fluid cooling medium extends through the wall segment. Each cooling channel is provided with an inlet and an outlet for the cooling medium. The at least one cooling channel includes at least two heat transfer sections, a first heat transfer section extending essentially parallel to the first surface at a first distance from the first surface and a second heat transfer section extending essentially parallel to the first surface at a second distance, whereby the second distance is less than the first distance.

Arrangement for cooling a component in the hot gas path of a gas turbine

Growing community of inventors

Waldshut-Tiengen, Germany

Herbert Brandl