Ansaldo Energia Switzerland Ag

Alstom Technology Limited

Asea Brown Boveri Ag

Abb Management Ag

Ansaldo Thomassen B.v.

Paul Marlow

Urs Benz

Marc Widmer

Jakob Keller

Michael Thomas Maurer

Joerg Krueckels

Jost Imfeld

Selma Zahirovic

Robert Marmilic

Eduard Brühwiler

Thomas Heinz-Schwarzmaier

Robin McMillan

Fabien Fleuriot

Martin Valk

Oliver Konradt

Joerg Krückels

Dalibor Vrbetic

Ark Jan Tijssen

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

An annular Helmholtz damper for a gas turbine can combustor, the annular Helmholtz damper having an axis; an inner wall and an outer wall concentrically arranged with respect to the axis to define an annular damping volume arranged around a can combustor; a front and rear circumferential plates for closing the annular damping volume upstream and downstream; at least one intermediate circumferential plate arranged between the front and the rear plates for dividing the annular damping volume in a main and a secondary volume; and a plurality of intermediate drain holes passing the intermediate circumferential plate and configured for draining collected liquid from the main volume to the secondary volume.

Annular Helmholtz damper for a gas turbine can combustor

A gas turbine transition duct includes: an inner tubular body, defining a transition channel and having a first upstream end and a first downstream end for coupling to a can combustor and to a turbine section of a gas turbine assembly, respectively; an outer tubular body, arranged around the inner tubular body and having a second upstream end at the first upstream end of the inner tubular body and a second downstream end at the first downstream end of the inner tubular body; wherein a convective cooling channel is defined between the inner tubular body and the outer tubular body, the convective cooling channel having an inlet between the first downstream end and the second downstream end; and wherein the outer tubular body is continuous between the second upstream end and the second downstream end.

Gas turbine transition duct

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

A gas turbine is provided having a secondary combustion chamber and a first guide vane row of a low-pressure turbine, the row being arranged directly downstream of the chamber. The radially outer boundary of the secondary combustion chamber is formed by at least one outer wall segment, which is secured on at least one support element arranged radially outwardly. The flow path of the hot gases is bounded radially outwardly, in the region of the guide vane row, by an outer platform which is secured at least indirectly on at least one guide vane support. A substantially radially extending gap-shaped cavity having a width in the range of 1-25 mm in the axial direction in the inlet region is arranged between the wall segment and the outer platform. At least one step element, which reduces the width by at least 10% in at least one step, extending substantially perpendicularly to the direction of flow of the hot gas in the cavity, is arranged in the inlet region.

Transitional region for a secondary combustion chamber of a gas turbine

The invention refers to an arrangement for sealing an open cavity against hot gas entrainment. The open cavity is arranged between two axially adjacent stationary components limiting radially a hot gas path of a rotary flow machine, of which at least the downstream component carries at least one airfoil extending radially into said hot gas path. A multitude of air jets is directed into the open cavity such that hot gas flowing over the open cavity is prevented from penetrating into the open cavity completely. The invention has an arrangement of supply conduits for air jets which are directed into the open cavity with a jet momentum such that hot gas is induced not to enter the open cavity and forming at least one hot gas vortex close to the hot gas path within the open cavity and preventing hot gas from penetrating into the open cavity beyond the extension of the hot gas vortex.

Arrangement for sealing an open cavity against hot gas entrainment

The burner arrangement for a gas turbine comprises at least one burner (2), which is arranged in the plenum of the gas turbine and leads with an inner injection space into a combustion chamber and to which compressed air is admitted on the outside from a compressor stage of the gas turbine. Furthermore, a fuel lance (9) for the alternative feeding of liquid and/or gaseous fuels is allocated to the burner (2), which fuel lance (9) has a central liquid-fuel tube (19) and a pilot gas tube (11) concentrically surrounding the liquid-fuel tube (19), the tubes (11, 19) ending in associated outlet openings (22, 23, 23a-c) in a lance head (12) at the tip of the fuel lance (9). In order to provide protection from an ingress of hot gases and/or hot liquid fuel into the outlet openings (22, 23, 23a-c) of the tubes (11, 19) when these tubes are not being used, an additional opening (17) is provided in the lance head (12) in the immediate vicinity of the outlet openings (22, 23, 23a-c), through which additional opening (17) compressed air flows out into the injection space (8). Furthermore, means are provided which direct compressed air from the plenum, surrounding the burner (2), to the additional opening (17).

Burner arrangement for a gas turbine for preventing the ingress of

A compressor for a gas turbine includes a gas seal chamber to seal the space between the guide vanes and the rotor vanes from ambient air. A seal casing is positioned at a radially inner side of the inlet casing adjacent to an inlet duct and the rotor blades with an annular gap between the seal casing and the rotor shaft. The annular gap communicates for air flow with the inlet duct. A plurality of ribs are mounted on the one or both of the seal casing and rotor shaft to extend into the annular gap to provide flow resistance. The seal casing defines therewithin a sealing air chamber is connected to the inlet space and the annular gap to permit air from the inlet space to flow to the annular gap and to the inlet duct.

Growing community of inventors

Baden, Switzerland

Paul Marlow