Siemens Aktiengesellschaft

Siemens Energy Global Gmbh & Co. Kg

Siemens Aktiengesell Schaft

Siemens Aktiengesellscaft

Fathi Ahmad

Radan Radulovic

Michael Dankert

Christian Lerner

Daniela Koch

Peter Schröder

Eckart Schumann

Giuseppe Gaio

Ralf Müsgen

Christian Kowalzik

Michael Händler

Björn Buchholz

Stefan Schmitt

Christian Kowalski

Günther Walz

Karsten Kolk

Harald Hoell

Hans-Thomas Bolms

Holger Grote

Mirko Milazar

Tobias Buchal

Nihal Kurt

Kevin Kampka

Andreas Heselhaus

Oliver Schneider

Volker Vosberg

Winfried Esser

Knut Halberstadt

Uwe Paul

Thorsten Mattheis

Bernd Vonnemann

Stefan Krause

Thomas-Dieter Tenrahm

Harald Nimptsch

Christian Amann

Silvio-Ulrich Martin

Björn Beckmann

Werner Setz

Oliver Lüsebrink

Markus Gill

Rostislav Teteruk

Stephen A Camillieri

Christian Menke

Stefan Werner Kiliani

Scarlett Fajardo-Reina

Marco Schüler

Markus Lempke

Thomas Hille

Axel Podschwadek

Anne Weidenmüller

Mario Nitsche

Knut Haberstadt

Mathias Stutt

Marco Schuler

Matthias Kuhlee

Andreas Varnholt

Olivier Gand

Albrecht Rieger

Horst-Michael Dreher

Stephan Schwarz

David Johnson

A guide vane ring which is divided into an upper and into a lower guide vane ring half, with an inner ring which is split at least in two, has a substantially U-shaped cross section, and forms a radially outwardly open flow duct which extends in the circumferential direction and is delimited by way of an inner ring bottom wall and two inner ring side walls, and with a multiplicity of vane platforms which receive guide vanes, are arranged along the outer circumference of the inner ring, and have holding webs which in each case project radially inwards, are spaced apart from one another in the axial direction, and engage around the inner ring side walls from the outside, wherein wear elements are inserted into gaps which are present between holding webs and directly adjacently arranged inner ring side walls.

Guide vane ring with wear elements

An aerofoil for a turbine blade, having a leading edge onto which a hot gas can flow and from which a suction-side wall and a pressure-side wall extend to a trailing edge of the aerofoil, the aerofoil extending in a direction transverse thereto from a root-side end having an aerofoil height of 0% to a tip-side end having an aerofoil height of 100%, and having at least two rows of cooling holes arranged along the leading edge, which cooling holes are at a mutual spacing which can be measured perpendicularly to the leading edge. In order to provide a turbine blade which can be used, with reduced cooling complexity, for additional reliable cooling of the leading edge in various operating conditions, the at least two rows of cooling holes are arranged at least in part in a wavy line along the leading edge.

Aerofoil for a turbine blade

A method for producing a cavity in a blade platform of a blade, in particular of a turbine blade, as part of a blade-platform cooling system, wherein the method has the steps of: producing a first bore from a first platform lateral face in the direction of an opposite second platform lateral face, with a first opening in the first platform lateral face being created, and expanding the first bore in a fan-like manner by an electrical discharge machining method, in particular using a wire- or bar-form electrode, such that the first opening, created in the first step, of the first bore represents the starting point of the fan-like expansion. A blade is produced in particular with such a method.

Method for producing a cavity in a blade platform; corresponding blade

Provided is a method for maintaining an axial-flow turbomachine, which has a support structure and a plurality of axially adjacent guide vane rings each of which has a plurality of guide vanes, each guide vane having at least one platform, the platforms being fastened to the support structure by means of retaining elements, wherein facing platform end faces of axially adjacent guide vanes define sealing sections made of a metal material, the sealing sections being associated with each other and jointly forming a sealing assembly.

Method for maintaining a turbomachine

A guide vane for a turbomachine having an outer platform, a blade airfoil protruding from the outer platform, extending in a longitudinal direction and defining a cavity in its interior, an inner platform arranged opposite the outer platform and connected to the blade airfoil, and a connecting tube which passes through the cavity of the blade airfoil in the longitudinal direction. The first free end of the tube is inserted into a through-opening formed in the inner platform and materially bonded to the inner platform, and the second free end of the tube is arranged in a cooling fluid inlet opening formed on the outer platform, away from the edge thereof, and projecting outward from the outer platform. At least one guide is attached to the outer platform and designed to guide the connecting tube in the longitudinal direction in the event of a thermal change in length.

Guide vane having a connecting tube

A turbine rotor blade for a thermal continuous flow machine, wherein a transition region and an aerodynamically shaped turbine blade connected to same, follows a blade foot for securing the turbine rotor blade to a rotor along a notional blade longitudinal axis of the turbine rotor blade from bottom to top. The blade foot has two flat end surfaces facing one another, two contoured side surfaces facing one another and joining the two end surfaces to one another, in which side surface at least one respective carrying edge is formed by creating dovetail- or fir tree-like end surface contour. The carrying edges become free edges or extend into the transition region via concave rounded portions. A channel bordering one of the two end sides is arranged in at least one concave rounded portion, the extension of which channel along the side surface is less than that of the carrying edge.

Turbine rotor blade

A method for producing turbine rotor blades or the base bodies thereof includes a) providing the base body, which has, following one another along a longitudinal axis, a blade root, a blade platform and a blade airfoil, b) sensing a value of at least one parameter of the base body, at least one of the parameters representing a vibrational property of the base body, c) comparing the sensed value with a predetermined target interval, d) if the sensed value lies outside the target interval, reducing the mass of the base body, wherein the reduction of the mass takes place at the blade root and/or on the blade platform by introducing at least one recess and/or by reducing a dimension below the corresponding target value.

Method for producing a base body of a turbine blade

A turbine blade with a ceramic thermal barrier coating system has a substrate designed as a blade platform and as a blade airfoil. On the substrate is a first ceramic layer as a thermal barrier coating, which protects the substrate in the exposed high temperature region and there is locally an increase of the thermal barrier coating for locally reinforcing the thermal barrier. The increase includes a material that is different from the material of the first ceramic layer. The local reinforcement is arranged over the first ceramic layer, without the first ceramic layer having a reduced layer thickness. The local reinforcement is provided at most on 30% of the area of the blade airfoil and is arranged close to a platform extending over the entire pressure side in the direction of flow and with an extent thereto in the radial direction of the blade airfoil is at most 30%.

Local two-layer thermal barrier coating

A turbine blade with aerodynamic aerofoil, which extends from a bottom to a top end and has two aerofoil walls extending between, which transversely thereto extend from a common leading to a common trailing edge, and with at least one platform, which extends transversely in relation to the aerofoil and is arranged at one of the two ends of the aerofoil. A number of openings are provided in the trailing edge between the bottom end and the top end. Either the opening in the trailing edge arranged closest to the platform surface extends into the platform surface or the platform surface facing the aerofoil is lowered locally in the region of that opening that is closest to the platform, and the opening opens out into a portion of the trailing edge that has become free as a result of the lowering.

Turbine blade with a central blowout at the trailing edge

A turbine blade having an internally cooled turbine blade airfoil in which a hollow space is divided by rib elements into at least one cooling duct carrying a coolant, wherein in at least one of the rib elements a separating tear initiating device for producing a separating tear is disposed, which extends at least partially in the longitudinal direction of the at least one rib element.

Turbine blade and turbine

A turbine blade, including: an airfoil and a platform, which has an upper face on which the airfoil is arranged; and at least one lateral face, the lateral face including a slot for insertion of a sealing strip. The transition between the upper face of the platform and the at least one lateral face includes a stepped portion and a beveled portion is provided.

Turbine blade or vane having a stepped and beveled platform edge

A gas turbine includes an insert element for fastening to a ring segment body of a turbine of the gas turbine. The ring segment body has a recess on a hot gas side. The insert element is designed to cover the recess. The insert element includes a cover plate, which has a concavely shaped front side and a back side, and at least one foot arranged on the back side for positioning on the ring segment body. A mounting method includes mounting the insert element where the insert element is fixed to a recess of a ring segment body of a turbine of a gas turbine.

Gas turbine and mounting method

An insert element fastens to an annular segment body of a turbine of a gas turbine. The annular segment body has a recess on a hot-gas side. The insert element is designed to cover the recess and has a concave front side and a rear side having at least one shaped portion for positioning on the annular segment body. An annular segment has the insert element, a gas turbine has the insert element, and a mounting method mounts the insert element.

A coolant bridging line for a gas turbine having inner and outer sides which are separated by a wall, wherein the coolant bridging line extends from a first component of the gas turbine to a second component of the gas turbine is provided herein. The coolant bridging line for a gas turbine provides a further increase in the service life of the components of the gas turbine possible with a degree of efficiency which is nevertheless as high as possible. To this end, the coolant bridging line has means which change the heat transfer between and/or the flow conditions on the inner and outer sides.

Coolant bridging line for a gas turbine, which coolant bridging line can be inserted into a hollow, cooled turbine blade

A process for producing a layer system is provided wherein the layer system has at least a substrate, a ceramic layer, which is applied to a surface structured in a targeted manner, in which process the intermediate layer, in particular the metallic layer, is applied in such a way that the recesses form during the coating. By introducing recesses into a surface, the stresses in the ceramic layer on the metallic substrate are reduced in such a manner that a longer lifespan for the ceramic layer is achieved.

Process for producing a layer system

Growing community of inventors

Kaarst, Germany

Fathi Ahmad