Alstom Technology Limited

General Electric Technology Gmbh

Ansaldo Energia Switzerland Ag

Alstom

Alstom (switzerland) Ltd

Ansaldo Energia IP UK Limited

Shailendra Naik

Ulrich Rathmann

Martin Schnieder

Brian Kenneth Wardle

Gordon David Anderson

Reinhard Fried

Herbert Brandl

Maxim Konter

Alexander Beeck

Beat Von Arx

Willy Heinz Hofmann

Gregory Edwin Vogel

Peter Marx

Thomas Zierer

Jonas Reiter

Peter Ernst

Christoph Didion

Stefan Schlechtriem

Marc Henze

Andre Saxer

Roland Dueckershoff

Sacha Parneix

Gaurav Pathak

Frank Gersbach

Ivan Luketic

Ruben Valiente

Helene Saxer-Felici

Christian Sommer

Helen Marie Saxer-Felici

Rainer Bauer

Jan Ehrhard

Allewis A Greninger

Gregory Vogel

Arrigo Beretta-Mueller

André Saxer

Roland DÜCKERSHOFF

Hélène Marie Saxer-Felici

An airfoil for a working fluid path of a turboengine extends along a spanwidth direction from a base to a tip. An aerodynamic body thereof includes a suction side surface, a pressure side surface, a leading edge, a trailing edge and a tip, the tip of the aerodynamic body having a tip cross section and a cross-sectional contour circumscribing the tip cross section. A rim extends to the tip of the airfoil and follows the cross-sectional contour on the pressure side, the suction side and extends over the leading edge of the airfoil, the rim delimiting a tip cavity which is open at the tip of the airfoil. The rim is further open at the trailing edge of the airfoil such that the tip cavity is open at the trailing edge of the airfoil.

Airfoil having a tip capacity

The present invention relates to an airfoil for a gas turbine, including an improved turbulator arrangement formed on an inner cooling channel of the airfoil. According to preferred embodiments of the invention, in order to ensure a constant angle of the cooling flow inside the channel relative to each turbulator, the angle formed between the turbulator and the vertical axis is advantageously adapted, in the curved area, for every single turbulator. Furthermore, the same principle may be applied to all the cooling channels present within the airfoil.

Airfoil with improved cooling

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 includes a rotor having a rotor groove and a rotor bore extending through the rotor, the rotor bore having a diffuser-shaped rotor bore exit. A blade is attached to the rotor and includes a blade tip having at least one dust hole. An airfoil has a leading edge and a trailing edge extending along a longitudinal axis of the blade between a lower end of the airfoil and the blade tip. A blade root is disposed at the lower end of the airfoil and is configured to be removably disposed in the rotor groove. The blade root includes a blade inlet having a cross sectional area that exceeds a cross sectional area of the rotor bore in at least one direction. A hollow blade core is disposed in the airfoil and extends along the longitudinal axis of the blade between the blade root and the blade tip.

Gas turbine

A turbine blade of an axial turbine includes internal cooling fluid passages with radially outwardly extending passages connected to holes in the blade root. The holes are generally core printouts providing stability to the core during the casting process, but are not needed and need to be closed to guarantee the functioning of the cooling system. This is achieved by at least one covering plate. The plate is held by at least two slots located at the root of the turbine blade. Thus, the supply holes for cooling fluid located at the root section are closed by a simple mechanical device, e.g., a plate that does not require any subsequent brazing/welding operations. In addition, the plate is removable to facilitate inspection/cleaning, or further processing of the blade at service intervals.

Turbine 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

A blade for a gas turbine includes an airfoil extending in radial direction of the turbine or longitudinal direction of the blade, respectively, between a platform and a blade tip. The airfoil is bordered across the airfoil by a leading edge and a trailing edge and has a suction side and a pressure side. At the trailing edge a first cooling passage runs parallel to the trailing edge from the platform to the blade tip in the interior of the airfoil. The cooling passage is supplied with a cooling air flow from the platform side, and from which cooling air is discharged through a plurality of cooling holes arranged all over the blade. For such a blade the cooling is optimized by providing a first cooling passage, the passage area of which is tapered in radial direction by between 35% and 59%.

Cooled blade for a gas turbine, method for producing such a blade, and gas turbine having such a blade

A turbine component () includes a root (), a tip (), and an airfoil portion () having a leading and a trailing edge (), an external suction side and pressure side wall () between the leading and trailing edge. The walls enclose a central cavity (-) for the passage of cooling air, the cavity being partitioned into a leading edge- and a trailing edge region () by at least one longitudinally extending first web () connecting the suction side wall with the pressure side wall and a second longitudinally extending web (), connecting the first web with the suction side wall, thereby defining a first and second entry chamber (). The first web () is provided with at least one cross-over hole (H) between the first entry chamber and the leading edge chamber (), whereas the second web has no openings.

Gas turbine airfoil

A gas turbine airfoil () includes a pressure sidewall () and a suction sidewall (), extending from a root to a tip and from a leading edge region to a trailing edge and having at least one cooling passage between the pressure sidewall () and the suction sidewall () for cooling air to pass through and cool the airfoil from within. One or several of the cooling passages () extend along the leading edge of the airfoil () and several film cooling holes () extend from the internal cooling passages () along the leading edge region to the outer surface of the leading edge region. The film cooling holes () each have a shape that is diffused in a radial outward direction of the leading edge of the airfoil () at least over a part of the length of the film cooling hole (). Improved cooling in the leading edge region can be achieved because the cooling holes () have a principal axis (), and the shape is asymmetrically diffused in that it is diffused in the radial outward direction from the principal axis () along a forward inclination axis (), and it is additionally diffused in a second lateral direction from the principal axis () along a lateral inclination axis ().

Gas turbine airfoil with leading edge cooling

The invention describes a device for separating dust and dirt out of flowing media. A curved flow path () is imposed on the flow (), and heavy foreign bodies are separated out of the main flow by centrifugal force. The partial stream required to carry away the foreign bodies is aftertreated in a filter () and returned in purified form. Unlike with conventional filters, which have to process the entire flow of media, pressure losses during the separation of dust are minimized, and on the other hand no medium is lost for removing the dust load.

Device for separating dust and dirt out of flowing media

A hot gas path assembly, suitable for use in the hot gas path of a gas turbine, has as a hot gas duct wall an impact-cooled gas-impermeable element and a transpiration-cooled gas permeable element. The gas-permeable element is a run-on covering for the sealing tip, and the gas-impermeable element is a blade foot of a turbine blade. Coolant is led in series through an impact-cooling element to cool the gas-impermeable element, and through the gas-permeable element for transpiration cooling and, if appropriate, also cools the sealing tip. Coolant thus is utilized particularly efficiently. Subdividing walls are arranged for the lateral subdivision of the coolant path, particularly in the circumferential direction, into segments. Because of the subdivision, in the event of damage to the gas-permeable element in one segment, the other segments remain essentially uninfluenced. Redundant cooling orifices may ensure coolant flow even when flow resistance in a transpiration-cooled element rises.

Hot gas path assembly

In an emergency cooling system () for a component () which is subject to thermal load in operation, in, e.g., a turbine, the component () has a wall () which, in operation, is acted on by heat on a first wall side () and is acted on by a flow of cooling fluid () on a second wall side (). The wall () has at least one emergency cooling opening () which is closed off by a plug () and through which cooling fluid flows from the second wall side () to the first wall side () when the plug () is absent. The plug () is designed so as to melt at a predetermined temperature. The plug () is a body which is produced separately from the component (), the plug () being inserted into the emergency cooling opening (), in which it is connected to the component ().

Emergency cooling system for a thermally loaded component

A gas turbine arrangement includes a rotor, a hot-gas passage, through which a hot-gas stream () flows while the gas turbine is operating, at least one first row of turbine blades or vanes (), which have an airfoil() with a suction side () and a pressure side (), and a platform (), and at least one second row of turbine blades or vanes (), which are arranged upstream of the first row of turbine blades or vanes () in the direction of the hot gases (), and in the axial direction of the rotor and likewise have a platform (). There is at least one seal (), through which a cooling air leakage stream () emerges while the gas turbine is operating, between the first row of turbine blades or vanes () and the second row of turbine blades or vanes (), in the region of the respective platforms (). An element guides the leakage stream () of cooling air along the surface of the platform () to the pressure side () of the first row of turbine blades or vanes ().

Gas turbine arrangement

A heat shield of a turbo machine is provided with a microstructure. In a preferred embodiment, the microstructure element comprises a plateau, which is arranged on a rib in such a way that the structure has a &ldquo;T&rdquo;-shaped cross section. The ribs are preferably embodied as plates set on edge on the heat shield and are aligned with their surface perpendicular to the circumferential direction (U) of the turbo machine. This results in low bending dimensional rigidity in the circumferential direction. In this way, it is possible to accommodate scraping of a component involved in relative motion in the circumferential direction without plastic deformation. Moreover this arrangement gives the maximum possible resistance to a leakage flow. When used at high temperatures, it is furthermore advantageous to provide ways to enable a coolant to be supplied to the microstructure.

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Gebenstorf, Switzerland

Shailendra Naik