Alstom Technology Limited

General Electric Technology Gmbh

Ansaldo Energia IP UK Limited

Jörg Krückels

Urs Benz

Hans-Peter Bossmann

Martin Schnieder

Brian Kenneth Wardle

Thomas Heinz-Schwarzmaier

Jose Anguisola McFeat

Uwe Rüdel

Marc Henze

Bruno Stephan

Christoph Appel

Thomas Sommer

Tanguy Arzel

Rene Koehnke

Laura Bogdanic

Milan Pathak

Martin Schneider

Jose Anguisola Mcfeat

Roland DÜCKERSHOFF

A gas turbine including a combustion chamber and a first row of guide vanes, arranged essentially directly downstream thereof, of a turbine. The outer and/or inner limitation of the combustion chamber defined by at least one outer and/or inner heat shield, mounted on at least one combustion chamber structure arranged radially outside and/or inside. The hot gases flow path in the region of the guide vane row being restricted radially on the outside and/or inside by an outer and/or inner vane platform, mounted at least indirectly on at least one turbine carrier. A minimal gap size directly upstream of the first row of guide vanes is achieved by mounting at least indirectly on the turbine carrier at least one mini heat shield, arranged upstream of the first row of guide vanes and essentially adjacent the vane platform and in the flow direction between the heat shield and the vane platform.

Transitional region for a combustion chamber of a gas turbine

A blade () for a gas turbine has a blade airfoil (), the blade wall () of which encloses an interior space (). For cooling the blade wall (), the blade wall () includes a cooling arrangement () which has a radial passage () extending in the longitudinal direction of the blade and from which a multiplicity of cooling passages (), extending in the blade wall (), branch in the transverse direction, and from which a multiplicity of film-cooling holes () are led to the outside in the transverse direction. Particularly efficient cooling is made possible by the distribution of the film-cooling holes () along the radial passage () being selected independently of the distribution of the cooling passages () along the radial passage ().

Gas turbine blade and method for producing a blade

A cooled blade for a gas turbine includes a blade airfoil extending between leading and trailing edges in a flow direction and on suction and pressure sides is delimited by a wall, which include an interior space in which cooling air flows towards the trailing edge in the flow direction and discharges to the outside in the region of the trailing edge. The pressure-side wall terminates at a distance in front of the trailing edge in the flow direction, forming a pressure-side lip, such that the cooling air discharges from the interior space on the pressure side. Multiple ribs subdivides the interior space, parallel to the flow direction, into a plurality of parallel cooling passages which create a high pressure drop and in which turbulators are arranged for increasing cooling. Before the outlet, multiple flow barriers are provided in the cooling air flow path, distributed transversely to the flow direction.

Cooled blade for a gas turbine

A gas turbine is provided including a compressor, a combustion chamber, a guide vane row and a rotor airfoil row. The guide vane row includes a plurality of guide vane airfoils having a blade and an inner platform. The ratio between the pitch and the leading edge diameter of the guide vane airfoils is between 6.3-7.6 and the ratio between the platform length and the leading edge diameter of the guide vane airfoils is between 4.0-5.5.

Gas turbine

Growing community of inventors

Birmenstorf, Switzerland

Jörg Krückels