Siemens Aktiengesellschaft

Siemens Energy Global Gmbh & Co. Kg

Siemens Power Generation, Inc.

Hans-Thomas Bolms

Peter Tiemann

Alexander Ralph Beeck

Fathi Ahmad

Andreas Heselhaus

Christian Lerner

Michael Händler

Stefan Baldauf

Jan Münzer

Ralf Müsgen

Michael Scheurlen

Michael Strassberger

Dirk Anding

Bernhard Hoffschmidt

Stefan Irmisch

Robert W Sunshine

Torsten Koch

Lutz Wolkers

Roland Häbel

Thomas Podgorski

Burkhard Bischoff-Beiermann

Thomas Schulenberg

Dieter Brillert

Wayne Giddens

Harald Hoell

Juergen Hermeler

Boris Dobrzynski

Stefan A Irmisch

A guide vane for a turbomachine, having a blade airfoil and at least one platform, to which the blade airfoil is connected. A cooling channel system is provided for cooling the platform and the blade airfoil. The platform has, on the side thereof facing the airfoil, at least one sealing lip for sealing to a rotating system of the turbomachine. At least one cooling channel extends through the sealing lip, which cooling channel forms part of the cooling channel system.

Guide vane for a turbomachine

A gas turbine having a turbine stage is presented. The gas turbine has a guide vane that can be operated cooled by a cooling medium, and a cooling medium feed device for feeding the cooling medium into the interior of the guide van, which, in the region of its rear edge, has on the pressure side at least one cooling medium outlet opening, through which the cooling medium can flow out of the interior of the guide vane into the main flow. The cooling medium feed device has a mass flow control device for controlling the mass flow through the at least one cooling medium outlet opening, with which the mass flow through the at least one cooling medium outlet opening can be increased during operation of the gas turbine under part-load compared to operation of the gas turbine under full load.

Process for cooling the turbine stage and gas turbine having a cooled turbine stage

A method for producing a hole with side-delimiting flanks in a component using a laser beam is provided. A laser beam is directed onto the component. A side flank of the hole is traced with the laser beam. A partial volume of the hole is formed by vaporizing the component material. This is repeated until the whole volume is formed. The laser beam may be oriented so that it includes an angle of more than 8&#xb0; with the traced side flank.

Method for producing a hole with side-delimiting flanks in a component

There is described a method with which, on the basis of a three-dimensional recording of a component to be reworked, its surface configuration can be determined and stored temporarily so that, after it has been coated, it can be produced in its original surface form, or in its surface form then required, in certain regions, i.e. locally in the area of film-cooling openings. An especially precise and quick three-dimensional recording can be achieved by the use of the triangulation method. In this case, a reference pattern depicted on the component by a projector is recorded by two camera arranged at an angle. From the images from the cameras, the coordinates describing the surface three-dimensionally can then be determined by a control system using the triangulation method.

Method for producing a gas turbine component which is to be coated, with exposed holes, device for carrying out the method, and coatable turbine blade with film cooling holes

A turbine blade is presented including a vane that runs along a blade axis and a platform region, which is located at the root of the vane and has a platform that extends transversally to the blade axis. An improved cooling of a platform region and the transition region is provided from a turbine blade to a platform of a turbine blade. The platform includes a first platform wall that does not support the vane and a second platform wall that supports the vane. At the root of the vane and over the course of the transition region from the turbine blade to the platform, the first platform wall is aerodynamically curved and the course of the second platform wall has a receding shoulder in relation to the first platform wall, as a continuation of the vane.

Turbine blade and gas turbine equipped with a turbine blade

The invention relates to a turbine blade comprising a vane that runs along a blade axis and a platform region, located at the root of the vane having a platform that extends transversally to the blade axis. The aim of the invention is to configure a delimitation of a flow channel of a gas turbine in the simplest possible manner. Therefore, the platform is configured by an elastic sheet metal part that rests on the vane. Said part leads to a gas turbine comprising a flow conduit that runs along an axis of the gas turbine, said conduit having an annular cross-section for a working medium and a second vane stage that is situated downstream of a first vane stage, which runs along the axis.

There is described a vane wheel of a turbine comprising at least one vane, the footing thereof being held on a wheel disk. At least one cooling channel is arranged between the wheel disk and the vane footing. The vane wheel having a plurality of turbulators is embodied on at least one of the walls of the cooling channel, the turbulators being configured in such a way that the turbulence and thus the heat transfer of a cooling fluid flowing through the cooling channel are increased.

Vane wheel of turbine comprising a vane and at least one cooling channel

The invention relates to a turbine blade comprising a vane that runs along a blade axis and a platform region, which is located at the root of the vane having a platform that extends transversally to the blade axis. The aim of the invention is to configure a delimitation of a flow channel of a gas turbine in the simplest possible manner. Therefore, the platform is configured by an elastic sheet metal part that rests on the vane. Said part leads to a gas turbine comprising a flow conduit that runs along an axis of the gas turbine, said conduit having an annular cross-section for a working medium and a second vane stage that is situated downstream of a first vane stage, which runs along the axis.

A turbine blade (), with a root section (), with a tip section () and with a blade leaf () which is provided with a number of coolant ducts () through which a coolant (K) is capable of flowing, outlet ducts () which issue in outlet ports () branching off, in the leading edge region () of the blade leaf (), from a coolant duct () running essentially in the longitudinal direction (L) of the turbine blade () and spaced apart from the leading edge (), the outlet ports () being arranged along at least two rows oriented essentially parallel to the leading edge (), and the outlet ducts () being oriented, in the region of their respective outlet port (), obliquely with respect to the longitudinal direction (L) of the turbine blade (), in such a way that the coolant (K) flowing out in a root-side subsection () of each row possesses, in the region of the outlet ports (), a velocity component pointing toward the tip section () of the turbine blade (), and the coolant (K) flowing out in a tip-side subsection (), continguous thereto, of each row has a velocity component pointing toward the root section (), is designed for a particularly reliable and uniform cooling of the leading edge region (), at the same time with the requirement for coolant (K) being kept particularly low. For this purpose, according to the invention, the transitional points () at which the orientation of the outlet ducts () changes are arranged so as to be offset relative to one another in the longitudinal direction (L) in each case for two adjacent rows.

Turbine blade

The invention relates to a temperature resistant layered structure comprising a substrate and a porous layer arranged on the substrate having a pore defined by a wall, and a ceramic coating on an interior surface of the wall. The invention also relates to a layered turbine component arrangement comprising a substrate having a cooling passage adapted to allow a cooling gas medium to pass through the substrate and a porous layer arranged on the substrate, the porous layer having cooling passages formed by gas-permeable inter-connections between pores in the porous layer.

Layer structure and method for producing such a layer structure

A coolable segment for a turbomachine such as a combustion turbine, which turbomachine is operated with a hot fluid. The segment comprises a cooling wall extending in an axial direction and in a circumferential direction orthogonal to the axial direction; a hot fluid surface to be exposed to the hot fluid. Between the wall and the surface a cooling structure is arranged which is permeable to cooling fluid and provides cooling surfaces for cooling by heat transmission through radiation. The cooling structure comprises either a netting, in particular a wire netting or a porous material.

Coolable segment for a turbomachine and combustion turbine

Embodiments of the invention relate to a system and method for controlling the size of gaps in a turbine engine. In many instances, it is desirable to minimize the size of the gaps between neighboring rotating and stationary components in a turbine engine, such as between a disc cover plate and a proximate pre-swirler. According to embodiments of the invention, each component can be provided with a sealing surface. The sealing surfaces can be angled relative to the axis of rotation. The sealing surfaces are spaced from each other so as to form a gap therebetween. The sealing surfaces may or may not be substantially parallel. As a result of such configuration, the size of the gap can be controlled by axial and radial movement of the components. For example, the gap between the cover plate and the pre-swirler can be adjusted by axially movement of the rotor.

Gap control system for turbine engines

A turbine vane, especially a turbine vane of the last stages, respectively includes a lower area which is radially and externally arranged, an upper area which is radially and internally arranged, and a radial cooling air channel extending between the upper area and the lower area. Cooling air can be introduced into the channel via an inlet in the lower area, and can be at least partially discharged via an outlet in the upper area. The cooling channel includes a radial inner channel through which the cooling air flows from the lower area to the upper area, and an outer channel which is adjacent to the inner channel on the circumferential side thereof. The outer channel communicates with the inner channel and includes an outlet which is arranged in the lower area. Part of the cooling air flows back in the direction of the lower area via the outer channel and emerges via the outlet.

Turbine vane system

The invention pertains to a seal element ( ) for sealing a gas-path leakage-gap ( ) between spaced apart first and second components ( ) of a turbo machinery ( ), which first and second components ( ) having opposing inner and outer surfaces ( ) situated outside the leakage-gap ( ). The seal element ( ) comprises a generally gas impervious sealing member ( ) and a layer ( ) comprising ceramic fibers. The layer ( ) covers at least partially said sealing element ( ) and defines a sealing surface ( ) for contacting outside the gap ( ) the outer surfaces ( ) of the first and second components ( ) of the turbo machinery ( ). The inventions further pertains to a combustion turbine ( ), which comprises plurality of components ( ) disposed in axial and circumferential direction, said components including guide-blade plates ( ) of guide blades ( ) and wall components ( ), which are spaced apart in the circumferential direction or the axial direction by a leakage-gap ( ) sealed off by a seal element ( ).

Seal element for sealing a gap and combustion turbine having a seal element

A component can be subjected to hot gas. At least one duct is provided which can be subjected to a cooling fluid. The duct is bounded by two first walls opposite to one another. The walls include turbulators with the same direction of inclination. In order to avoid constrictions, the turbulators of the first wall have a different angle of inclination relative to a flow direction of the cooling fluid to the turbulators of the second wall.

Growing community of inventors

Mülheim an der Ruhr, Germany

Hans-Thomas Bolms