Alstom Technology Limited

Alstom (switzerland) Ltd

Andreas Boegli

John Fernihough

Alexander Stankowski

James Ritchie

Slawomir Slowik

Matthias Hoebel

Alexander Beeck

Alexander Schnell

Bernd Fehrmann

Klaus Germerdonk

Bruno Benedetti

Christopher John Hulme

Alexander Mahler

Giampiero Antonelli

Klaus Erich Schneider

Wolfgang Mueller

Markus Schmid

Patrick Wolfgang Schnedler

Heinrich Stuessi

Christopher Hulme

Process of masking cooling holes of a gas turbine component with an external surface, comprising a cavity and a plurality of cooling holes before coating the gas turbine component, comprising the steps of first applying a mask material to the cooling holes so that the cooling holes are filled at least closest to the external surface, whereby the mask material contains a substance which fluoresces under ultraviolet light and a filler material. Then the mask material within the cooling holes is thickening. An inspection using ultraviolet light to locate any unwanted residual mask material on the external surface is carried out and unwanted residual mask material is removed before the coating is applied to the external surface of the component and the masked cooling holes. In the end the mask material is removed from the cooling holes.

Masking material for holes of a component

A blade for a turbo machine has a blade section and a shroud element terminating the blade section in the blade section longitudinal direction. The blade section has a suction face and a pressure face. The shroud element extends essentially at right angles to the blade section longitudinal direction and has a first platform section projecting beyond the blade section and a second platform section projecting beyond the blade section. The platform sections are asymmetric with respect to one another. At least the first platform section of the shroud element is arranged at an additional inclination angle with respect to a normal alignment of the first platform section, with the additional inclination angle being in the opposite direction to the bending torque which acts on the first platform section during operation.

Blade with shroud

A guide vane ring of a turbomachine, in particular of an axial-throughflow turbine, includes a vane group () having a plurality of vanes (). The vane group () has a vane root () which has two flanges (). The vane group () is fastened to a vane carrier by means of the flanges (). In order to reduce stresses in the vanes () which occur during deformations of the vane carrier, one flange () has on a front end portion () and on a rear end portion (), radially on the inside and radially on the outside, in each case a contact zone () which bears against the vane carrier. The other flange () has on one end portion (), radially on the inside, a contact zone () which bears against the vane carrier, and is spaced apart from the vane carrier radially on the outside. Moreover, this flange () has on the other end portion (), radially on the outside, a contact zone () which bears against the vane carrier, and is spaced apart from the vane carrier radially on the inside. Furthermore, the flanges (), between their end portions (), are spaced apart from the vane carrier radially on the inside and radially on the outside.

Guide vane ring of a turbomachine and associated modification method

An automated machining process for machining of an obstructed passage () of an article () includes the steps of deriving the position and orientation of the passage () from automated processing of images of the passage () and additional information from a distance measurement device and saving positions and orientations of the passage () as local coordinates with respect to a reference coordinate system attached to the material that surrounds the passage (). The saved positions and orientations of the passage () are used for subsequent removal of unwanted material from the obstructed passage ().

Automated adaptive machining of obstructed passages

A turbomachine blade is disclosed having a shroud element, wherein plastic deformations and lifting of the shroud element on one side result during operation under centrifugal load. This load may result in high-temperature creep of the blade. A sealing strip which is arranged on the shroud element can be configured with a thickness varying in a circumferential direction. The mass of the shroud element and thus the asymmetrical centrifugal load and the lifting of the shroud element on one side resulting therefrom can be reduced by material removal at regions lying on the outside in the circumferential direction.

Turbomachine blade

A method of reconditioning a rotating blade () with a shroud () following the operation in a gas turbine, includes the application of a barrier () placed on the edge of the shroud () facing the edge of a shroud () of an adjacent blade having creep deformation. The barrier () prevents hot gas ingestion into a shroud cavity (). It is pre-fabricated or built-up layer-by-layer by welding or laser cladding or other methods. Additionally, the mass of the shroud () is reduced in order to re-establish the initial mass of the shroud. A thermal barrier coating () is additionally applied to the surfaces of the shroud () that are exposed to the hot gas of the turbine. The method significantly increases the lifetime of the blade ().

Method of reconditioning a turbine blade

Process of masking cooling holes of a gas turbine component

A process of masking cooling holes () of a gas turbine component () with an external surface (), including a cavity () and a plurality of cooling holes () before coating the gas turbine component (), includes first applying a mask material () to the cooling holes () so that the cooling holes are filled at least closest to the external surface (), whereby the mask material () contains a filler material. Then the mask material () within the cooling holes () is thickening and the coating () is applied to the external surface () of the component () and the masked cooling holes (). In the end remaining thickened mask material () is removed from the cooling holes ().

It is disclosed a method of protecting a local area of components () from the effects of thermochemical or mechanical processes carried out on the surface () of the component. A masking material () containing at least one filler material is applied to the local area so that the local area is protected by the masking material (). This is at least partially polymerized on the local area. Subsequently the thermochemical or physical processes on the surface () of the component () are carried out after which the polymerized masking material () is removed from the local area of the component ().

Method of protecting a local area of a component

It is disclosed a method of detecting and quantifying subsurface defects () in an article () made of high strength non magnetisable materials after the use in a high temperature environment. A crack () or gap on a surface () of the article () is brazed and after the brazing operation the crack () or any remaining braze defect or subsurface crack () is detected and quantified by means of a multifrequency eddy current system.

Non-destructive method of detecting defects in braze-repaired cracks

A nondestructive method for determining the thickness of a metallic protective layer determined by detection of a different type of intervening layer which is situated between the metallic protective layer and the metallic base material, by ultrasound, eddy-current or (pulsed) thermography. The intervening layer may comprise a nonmetallic layer of oxides or carbides, a metallic layer of pure nickel or a layer with depleted Al and Ti content in that region of the base material which is close to the surface. This method can be used for turbine blades in order to determine the layer thickness, for example during the test phase of coating, and to avoid the known destruction of the turbine blade, which is labor- and cost-intensive.

NONDESTRUCTIVE METHOD FOR DETERMINING THE THICKNESS OF A METALLIC PROTECTIVE LAYER ON A METALLIC BASE MATERIAL VIA A DIFFERENT TYPE OF LAYER BETWEEN THE METALLIC PROTECTIVE LAYER AND THE METALLIC BASE MATERIAL

A method of correcting deformed turbine components ( ) in an incompressible medium ( ) is achieved via means of a pressure wave ( ). The turbine component ( ) is rigidly fixed to a matrix ( ) containing a specified mold ( ) and is put into the incompressible medium ( ). The method may advantageously work at ambient temperature. It is quick, reproducible and cost-effective.

Growing community of inventors

Vogelsang-Turgi, Switzerland

Andreas Boegli