Alstom Technology Limited

General Electric Technology Gmbh

Ansaldo Energia Switzerland Ag

Ansaldo Energia IP UK Limited

Alstom Power N.v.

Martin Schnieder

Shailendra Naik

Urs Benz

Beat Von Arx

Joerg Krueckels

Thomas Zierer

Roland Dueckershoff

Jörg Krückels

Tanguy Arzel

Thomas Heinz-Schwarzmaier

Sergey Shchukin

Marcel Koenig

Sacha Parneix

Uwe Rüdel

Chiara Zambetti

Francesco Garbuglia

Volker Eppler

Christoph Appel

Andrew Clifton

Jens Von Wolfersdorf

Günter Filkorn

Helen Marie Saxer-Felici

Roland DÜCKERSHOFF

Hélène Marie Saxer-Felici

Gas turbine unit having an axis parallel to the main gas flow, the gas turbine unit comprising:

Blade to stator heat shield interface in a gas turbine

A turbine vane for a rotary turbomachine has a turbine blade delimited by a concave pressure-side wall and a convex suction-side wall which are connected in the region of a vane front edge which can be assigned to the turbine blade and enclose a cavity which extends in the longitudinal extent of the vane front edge and is delimited on the inner wall by the pressure-side wall and the suction-side wall in the region of the vane front edge and by an intermediate wall which extends in the longitudinal direction to the vane front edge and connects the suction-side wall and the pressure-side wall on the inner wall. The intermediate wall has a perforation at least in sections in the connecting region to the suction-side wall and/or pressure-side wall, in order to increase the elasticity of the intermediate wall.

Turbine blade

A cooled blade is disclosed for a gas turbine that includes a radially extending aerofoil with a leading edge, a trailing edge, a suction side and a pressure side, and wherein a lip overhang is provided on the suction side of the trailing edge The blade also includes a plurality of radial internal flow channels connected via flow bends to form a multi-pass serpentine for a coolant flow, whereby a trailing edge ejection region is provided for cooling the trailing edge, the trailing edge ejection region comprising a trailing edge passage of the multi-pass serpentine running essentially parallel to the trailing edge and being connected over its entire length with a pressure side bleed. An optimized cooling is achieved by also determining the cooling flow from the trailing edge passage to the pressure side bleed by means of a staggered field of pins.

Cooled blade for a gas turbine

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 stator blade for a gas turbine with sequential combustion, has a blade airfoil which extends in the radial direction between a blade tip and a shroud, with cooling passages extending inside the blade airfoil, through which a cooling medium can flow for cooling the blade and can then discharge from the stator blade into the hot gas flow flowing through the turbine. The blade airfoil has a sharply curved shape in space in the radial direction, and three cooling passages, which extend in the radial direction, arranged inside the blade airfoil in series in the hot gas flow direction and are interconnected by deflection regions, which are arranged at ends of the blade airfoil, so that the cooling medium flows through the cooling passages one after the other, with change of direction. The cooling passages follow the curvature of the blade airfoil in space in the radial direction.

Stator blade for a gas turbine and gas turbine having same

A turbine blade or vane includes at least one internal radial channel for the circulation of cooling medium bordered on a pressure side by a pressure side wall and on a suction side by a suction side wall joined at a upstream side at a leading edge and at and downstream side at the trailing edge. At least one exit hole extends through at least one of the pressure side wall or the suction side wall for blowing out of cooling medium from the internal radial channel to a medium surrounding the blade or vane. At least one trailing edge exit hole along the trailing edge has a surfacial exit opening disposed at the pressure side of the trailing edge.

Gas turbine airfoil with shaped trailing edge coolant ejection holes

In a non-destructive method for determining the internal structure of a heat conducting body, such as a cooling structure of a turbine blade, a flow medium is passed through the internal structure and the resultant thermal image on an external surface of the body is registered using a pixelised thermal image detector. Heat transfer coefficients and wall thicknesses of the internal structure are determined by means of a 1-,2-, or 3-dimensional inverse method that includes the numerical modelling of the surface temperatures using initial values for heat transfer coefficients and wall thicknesses and an optimization of the values using an iteration method. In a special variant of the method, the spatial geometry of the internal structure of the body is determined by means of the same inverse method and a geometry model that is optimised by iteration. No prior knowledge of the internal geometry is required.

Method to determine the internal structure of a heat conducting body

A gas turbine is provided that includes a rotor which is rotatable around an axis and equipped with rotor blades, and which is concentrically enclosed at a distance by a casing, which is equipped with stator blades, forming an annular hot gas passage. Rings with stator blades and rotor blades are arranged in a manner alternating in the axial direction. Between adjacent stator blades, heat shield segments are arranged, which delimit the hot gas passage on the outside in a region of the rotor blades and are cooled by impingement cooling where a cooling medium from an outer annular cavity flows into the heat shield segment.

Shroud seal segments arrangement in a gas turbine

A blade for a gas turbine includes a leading edge running in a longitudinal direction substantially radially to an axis of the turbine; a trailing edge running in the longitudinal direction; a blade body disposed between the leading edge and the trailing edge so as to define a pressure side and a suction side; an exit slot disposed in the blade body in an area of the trailing edge and running in the longitudinal direction and configured to discharge a cooling medium from an interior of the blade body; and a row of first and second control elements disposed in the exit slot in a distributed manner in the longitudinal direction and configured to control a mass flow of the cooling medium exiting through the exit slot, the first control elements having a first configuration and the second control elements having a second configuration different from the first configuration.

Blade for a gas turbine

A continuous-flow machine (), in particular a turbine or a compressor, includes at least one stator blade section () which has a plurality of stator blades () which project into a gas path () of a working gas, and at least one adjacent section (), which has a plurality of adjacent elements (), which bound the working gas path () at the side. A gap (), through which a cooling gas can be introduced into the working gas path (), is formed between the stator blades () of the stator blade section () and the adjacent elements () of the adjacent section (). In order to improve the efficiency of the continuous-flow machine (), lateral seals () are formed in the gap () and prevent or at least impede pressure equalization and/or a gas flow in the gap () in the lateral direction ().

Continuous-flow machine, turbine, or compressor

The present invention relates to a component of a flow machine, particularly a turbine blade, which has a cooling channel ( ) through which cooling medium can flow, with at least one deflection ( ) formed by the wall ( ) of the cooling channel, by means of which deflection the flow of the cooling medium from a first channel section ( ) is deflected into a second channel section ( ) situated downstream. At least one flow guiding element ( ) is arranged in the cooling channel in the region of the deflection, and divides the cooling channel into an inner ( ) and an outer ( ) flow channel. In the present component, the inner flow channel ( ) has a constriction in the flow cross section. By the proposed configuration of the cooling channel deflection, the pressure loss brought about by the deflection is minimized and simultaneously a distinctly more homogeneous heat transfer is attained to the cooling medium, without local temperature peaks.

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Ennetbaden, Switzerland

Martin Schnieder