Alstom Technology Limited

Ansaldo Energia Switzerland Ag

General Electric Technology Gmbh

Ansaldo Energia IP UK Limited

Sven Olmes

Thomas Zierer

Carlos Simon-Delgado

Stefan Rofka

Jonas Hurter

Sascha Justl

Sergei Riazantsev

Ulrich Robert Steiger

Giovanni Cataldi

Wilhelm Reiter

Thomas Heinz-Schwarzmaier

Thomas Kramer

Armin Busekros

Sergey Shchukin

Holger Kiewel

Axel Heidecke

Christophe Simonet

Ronald Wifling

Rene Waelchli

Oliver Joseph Taheny

Meisam Sistaninia

Luca Alberti

Isaac-Eliseo Gomez-Lor-Lopez

Daniel Jemora

Eugenio Rossi

Pancrazio Abate

Nikolas Wiedemann

Thomas F Kramer

Uwe Wolfgang Ruedel

Phillipe Lott

Hyazinth Immanuel Hennel

A method is disclosed for cooling a gas turbine having a turbine, wherein a rotor, which rotates about a machine axis, carries a plurality of rotating blades, which are mounted by blade roots and extend with their airfoils into a hot gas path of the gas turbine. The rotor is concentrically surrounded by a turbine vane carrier carrying a plurality of stationary vanes, whereby the rotating blades and the stationary vanes are arranged in alternating rows in axial direction. An extended lifetime with external cooling is achieved by providing first and second cooling systems for the turbine.

Method for cooling a gas turbine and gas turbine for conducting said method

A compressor assembly, and more in general relates to a compressor for a gas turbine providing a solution that teaches to locate within a cavity formed by the outer casing of the compressor and the inner vane carrier a separator element, or membrane, such to divide the cavity into two sub-cavities. This advantageously results in a more flexible design with respect to the positioning of the flange blow-off extractor and to the cavity sizing, as the flange position is not necessarily the boundary for the flow anymore as it would be without the separator element.

Compressor assembly for gas turbine

A method and a cooling system for cooling blades of at least one blade row in a rotary flow machine includes an axial flow channel which is radially limited on the inside by a rotor unit and at the outside by at least one stationary component, the blades are arranged at the rotary unit and provide a shrouded blade tip facing radially to said stationary component. Pressurized cooling air is fed through from radially outside towards the tip of each of said blades in the at least one blade row, and the pressurized cooling air enters the blades through at least one opening at the shrouded blades' tip.

Method and cooling system for cooling blades of at least one blade row in a rotary flow machine

A gas turbine includes a rotor concentrically surrounded by a casing, with an annular hot gas channel axially extending between the rotor and the casing. The rotor includes a plurality of blades arranged annularly on the rotor. Each of the blades is mounted with a root in a respective axial slot on a rim of the rotor radially extending with an airfoil into the hot gas channel and adjoining with an axially oriented root surface to an annular rim cavity. A cooling device is provided at the root of each blade to receive cooling air injected into the rim cavity through stationary injectors. An optimized cooling is achieved by providing an essentially plane root surface and the cooling device includes a scoop for capturing and redirecting at least part of the injected cooling air, which scoop is a recess with respect to the root surface.

Gas turbine and turbine blade for such a gas turbine

A gas turbine engine includes a compressor with rotor blades having roots connected into seats of a compressor drum. The rotor blade roots and/or the compressor drum have longitudinal passages for a cooling fluid, connecting higher pressure areas to lower pressure areas of the gas turbine engine.

Gas turbine engine and method for cooling the compressor of a gas turbine engine

A turbomachine (), in particular a gas turbine, has a stator () and a rotor () and with at least one axial sealing device () arranged between them. On the stator side, the sealing device () has a radially stepped sealing contour () with regions () projecting and retracting in the direction of the rotor (), on the rotor side a plurality of sealing fins () projecting in the direction of the stator () being arranged, which engage in each case into adjacent retracting regions () of the stator-side sealing contour (). On the rotor side, at least one additional fin () projecting in the direction of the stator () is provided, which is positioned between two adjacent sealing fins () arranged on the rotor side and which lies opposite a projecting region of the stator-side sealing contour ().

Turbo machine

A method for axial thrust control of a gas turbine, and a gas turbine with a device for controlling axial thrust are provided. A gas turbine, with regard to aerodynamic forces and pressure forces, which exert an axial force upon the rotor, is configured such that at no-load and low partial load it has a negative thrust, and at high load it has a positive thrust. In order to ensure a resulting positive thrust upon the thrust bearing within the entire load range of the gas turbine, an additional thrust is applied in a controlled manner. The additional thrust for example can be controlled in dependence upon the gas turbine load. The resulting thrust force at full load is consequently less than in the case of a conventionally designed gas turbine without thrust balance.

Gas turbine with axial thrust balance

A gap seal () radially seals a gap () which extends axially and radially between two blades () of a turbomachine which are adjacent in the circumferential direction (). The two blades () have, in each case, an axially extending longitudinal slot (), which is open towards the gap (), on its respective blade root (). A band-form or strip-form sealing element () engages with its longitudinal sides () in the two longitudinal slots () and bridges the gap (). The one blade () has a projection () on its blade root (), which projects from the blade root () in the circumferential direction () and extends in the circumferential direction () and radially, at least in the region of the respective longitudinal slot (), and bridges an axial longitudinal end of the gap () in the process. The other blade () has a step-shaped recess () on its blade root (), complementary to the projection () of the one blade () and in which the projection engages. The projection () has a transverse slot () which extends in the circumferential direction () and is open towards the gap (). The sealing element () engages with an end face () in the transverse slot ().

Gap seal in blades of a turbomachine

A gas turbine includes a compressor for compressing the combustion air, at least one combustion chamber, in which a fuel is burned while compressed combustion air is supplied, and at least one first turbine arranged downstream of the combustion chamber and in which the hot combustion gases from the combustion chamber are expanded to perform work. Part of the compressed combustion air is branched off, cooled in a cooling air cooler, brought to a lower cooling air pressure by a second turbine and supplied to the gas turbine for cooling purposes.

Gas turbine with heat exchanger for cooling compressed air and preheating a fuel

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Windisch, Switzerland

Sven Olmes