Alstom Technology Limited

Ansaldo Energia Switzerland Ag

General Electric Technology Gmbh

Alstom (switzerland) Ltd

General Electric Company

Asea Brown Boveri Ag

Abb Management Ag

Ansaldo Energia IP UK Limited

Urs Benz

Andrea Ciani

Andre Theuer

Adnan Eroglu

Ewald Freitag

Michael Düsing

Andreas Huber

Franz Joos

Douglas Pennell

Nicolas Noiray

Hans-Christian Mathews

Jost Imfeld

Frank Graf

Johannes Buss

John Philip Wood

Diane Lauffer

Uwe Rüdel

Thorsten Christoph Motzkus

Jaan Hellat

Bruno Schuermans

Mirko Ruben Bothien

Madhavan Narasimhan Poyyapakkam

Michael Thomas Maurer

Marcel Stalder

Paul Marlow

Anton Winkler

Roger Suter

Felix Guethe

Martin Schnieder

Burkhard Schulte-Werning

Remigi Tschuor

Jonas Hurter

Felix Michael Reinert

Jeffrey Alan De Jonge

Fabien Fleuriot

Xianfeng Gao

Christoph Appel

Michael Hutapea

Bogdan Trbojevic

Yang Yang

Elisabetta Carrea

Pirmin Schiessel

Daniel Burri

Hugo Wetter

Felix Baumgartner

Jörg Krückels

Andreas Pfeiffer

Axel Kranzmann

Armin Heger

Felix Güthe

Lothar Schneider

Piotr Siewert

Oliver Konradt

Karolina Krystyna Sobol

Fridolin Heynen

David Walhood

Ivan Lenuzzi

Slawomir Swiatek

Rohit Kulkarni

Christoph Gaupp

Thomas Michael Maurer

Ivo Belina

Stevica Furdek

Anna Maria Olsson

Christoph Welti

Guenter Sybon

Marcel K&ouml;nig

Marijan Beg

Branko Cvitak

Douglas Anthony Pennell

Birute Bunkute

Maurice Mallm

Alexander Sergeevich Myatlev

Andreas De Fazio

The invention concerns a gas turbine having a picture frame, a first vane, and a sealing arrangement to seal a gap between the picture frame and the first vane, the sealing arrangement including two seals arranged in series between the picture frame and the first vane. In exemplary embodiments, one of the seals is a honeycomb seal, a dogbone seal, a hula seal or a piston seal and the other seal is a honeycomb seal, a dogbone seal, a hula seal or a piston seal. A method of supplying cooling fluid to the gap between the picture frame and the first vane is also disclosed.

Sealing arrangements in gas turbines

Combustor device having a twin-shell tubular casing including: an inner tubular member which extends roughly coaxial a longitudinal axis of the combustor device, delimits the combustion chamber and surrounds the burner; and an outer tubular housing which extends roughly coaxial outside of the inner tubular member. An intermediate supporting structure includes an outer annular supporting member, an inner annular supporting member, and a series of three or more oblong connecting beams angularly staggered about the longitudinal axis of combustor device in cantilever manner to stably connect the inner and outer annular supporting members.

Combustor device for a gas turbine engine and gas turbine engine incorporating said combustor device

A fuel lance is disclosed for injecting a gaseous and/or liquid fuel mixed with air into an axial hot gas flow flowing through a sequential combustor of a gas turbine, the fuel lance having at least one finger extending in a longitudinal direction into the axial hot gas flow of the gas turbine essentially perpendicular to the hot gas flow. The finger is configured as a streamlined body which has a streamlined cross-sectional profile. The body has two lateral surfaces essentially parallel to the axial hot gas flow. The body includes an enclosing outer wall having a longitudinally extending air plenum for the distributed introduction of air into the at least one finger. The air plenum is provided with a plurality of distributed impingement cooling holes, such that air exiting through the impingement cooling holes impinges on the inner side of the leading edge region of the body.

Fuel lance cooling for a gas turbine with sequential combustion

A turboengine as disclosed includes an outer wall structure and an inner wall structure, wherein the inner wall structure is provided at a radially inner position with respect to the outer wall structure, and each of the wall structures has a surface, the surfaces being arranged facing each other in the radial direction. At least one guide vane member includes at least one airfoil, a radially inner end and a radially outer end. The inner wall structure and the outer wall structure are jointly provided as a vane carrier unit, wherein the inner wall structure and the outer wall structure are fixedly connected to each other by at least one bridging member extending between the inner wall structure and the outer wall structure.

Turboengine, and vane carrier unit for turboengine

A vortex generating device is provided as a generally airfoil-shaped lobed body. The flow deflection varies along a spanwise extent, such that the body exhibits a corrugated geometry in a trailing edge region, and the trailing edge exhibits an undulating shape. The undulating shape includes at least one corner. The trailing edge may include or may consist of straight trailing edge sections. The trailing edge may exhibit a polygonial waveform shape, in particular a trapezoidal or rectangular waveform shape. The vortex generating device may be provided as a fuel discharge device which is suited to discharge a fuel into a vortex flow generated by the vortex generating device. To this extent, a fuel supply plenum may be provided inside the body and at least one fuel discharge opening which is in fluid communication with the fuel supply plenum may be provided on the trailing edge.

Vortex generating device

A gas turbine unit having a combustor having a liner, a turbine, arranged downstream of the liner along a main flow gas direction and including a plurality of first stage vanes, a rotor cover support located inwardly of the vanes, and a sealing arrangement at a combustor to turbine interface, wherein the sealing arrangement includes a first dogbone seal extending between the rotor cover support and an inner downstream end of the liner or between the rotor cover support and a bulkhead located at the inner downstream end of the liner.

Sealing arrangement on combustor to turbine interface in a gas turbine

The invention concerns a conical hula seal wherein an inner part of the conical hula seal is closer to a conical hula seal longitudinal axis than an outer part of the conical hula seal, and wherein the conical hula seal extends in a circumferential direction relative to the conical hula seal longitudinal axis. A gas turbine which includes the conical hula seal is also disclosed, along with a method of operating a gas turbine which includes a conical hula seal that seals a gap between a first vane and a picture frame, the method including purging the gap with a cooling fluid.

Hula seal

A burner for a combustion chamber of a gas turbine with a mixing and injection device, which includes a limiting wall that defines a gas-flow channel and at least two streamlined bodies. Each streamlined body extends in a first transverse direction into the gas-flow channel, and has two lateral surfaces that are arranged essentially parallel to the main-flow direction. The lateral surfaces are joined to one another at their upstream and downstream sides to form leading and trailing edges of the body, respectively. At least one of the streamlined bodies includes a mixing structure and at least one fuel nozzle at its trailing edge for introducing at least one fuel essentially parallel to the main-flow direction into the flow channel. At least two of the streamlined bodies have different lengths along the first transverse direction such that they may be used for a can combustor.

Burner for a can combustor

The present invention relates to dampers for gas turbines and, for example, to a compensation assembly for a damper of a gas turbine for reducing the pulsations occurring in the combustion chamber. The damper can include a resonator cavity with a neck tube in flow communication with the interior of the combustion chamber, wherein the compensation assembly includes a spherical joint associated to the neck tube and configured to allow relative rotation between the combustion chamber and the resonator cavity, and having a bulb portion disposed around the neck tube and a spherical socket configured to internally host the bulb portion, wherein the spherical socket can have a top collar portion and a bottom collar portion connected to each other.

Compensation assembly for a damper of a gas turbine

A sequential burner for an axial gas turbine comprises: a burner body, which is designed as an axially extending hot gas channel and further comprises a fuel injection device, which extends into said burner body perpendicular to the axial direction. The manufacturing of the burner body is simplified and the fuel injection is stabilized by designing said fuel injection device as a mechanically stiff component, and fixing said fuel injection device to said burner body in order to keep it aligned with said burner body and to stiffen said burner body against creep.

Sequential burner for an axial gas turbine

The present invention relation to a burner for a combustion chamber of a gas turbine with a mixing and injection device. The mixing and injection device includes a limiting wall that defines a gas-flow channel and at least two streamlined bodies, each extending in a first transverse direction into the gas-flow channel. Each streamlined body has two lateral surfaces that are arranged essentially parallel to the main-flow direction, the lateral surfaces being joined to one another at their upstream side to form a leading edge of the body and joined at their downstream side to form a trailing edge of the body. Each streamlined body has a cross-section perpendicular to the first transverse direction that is shaped as a streamlined profile. At least one of the streamlined bodies is provided with a mixing structure and with at least one fuel nozzle located at its trailing edge for introducing at least one fuel essentially parallel to the main-flow direction into the flow channel, wherein at least two of the streamlined bodies have different lengths along the first transverse direction such that they may be used for a can combustor. The invention also relates to a method of using said burner in a gas turbine.

Burner for a combustion chamber of a gas turbine having a mixing and injection device

A combustor arrangement for a gas turbine includes a first burner, a first combustion chamber, a mixer for admixing a dilution gas to the gases leaving the first combustion chamber during operation, a second burner, and a second combustion chamber arranged sequentially in a fluid flow connection. These elements of the combustor arrangement are arranged in a row to form a flow path extending between the first combustion chamber and the second burner. The combustor arrangement includes acentral lance body arranged inside the flow path and extending from the first burner through the first combustion chamber into the mixer and into the second burner, wherein the lance body includes a fuel duct for providing fuel for the first burner and/or for the second burner.

Combustor arrangement for a gas turbine

A centering arrangement includes a ring-like inner part and a ring-like outer part, whereby the outer part surrounds the inner part in a concentric arrangement and with an interspace established between the outer part and the inner part, and whereby the outer part and the inner part are subject to a differential radial expansion. A centering contact element, procures a centering mechanical contact between the outer part and the inner part at a plurality of circumferentially distributed contact points, and is able to deform in order to compensate for differential radial expansion.

Centering arrangement of two parts relative to each other

The invention relates to an annular combustion chamber of a gas turbine having a machine axis. The combustion chamber includes at least two zones. A first zone receives the fuel/air mixture of a plurality of burners. A second zone guides the hot gases being produced by the burners from the first zone to an entrance of a turbine section of said gas turbine. An annular transition liner includes a plurality of liner segments located at the transition between said first zone and second zone. Each of the liner segments includes with respect to the axial hot gas flow a leading edge, a trailing edge, and two sidewalls, and is provided with axial mounting means at the leading and trailing edges, such that the liner segment can be installed in axial direction and is axially fixed by respective segments of the neighboring first zone. Local spacer ribs are provided at the leading edge of the liner segments in order to establish a gap of minimum width between the liner segments and the fixing segments of the neighboring first zone.

Annular combustion chamber of a gas turbine and gas turbine with such a combustion chamber

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Gipf-Oberfrick, Switzerland

Urs Benz