Abstract: A method for repairing a shroud segment of a gas turbine, in particular of an aircraft engine, having the steps: a) provision of a shroud segment to be repaired; b) removal of a damaged radially inner section of the shroud segment, so that after the removal of the damaged section the shroud segment has an inner radius; c) manufacture of a replacement section for the shroud segment, the replacement section having an outer radius matched to the inner radius of the shroud segment; d) alignment of the replacement section and the shroud segment; and e) joining of the replacement section to the shroud segment, the replacement section and the shroud segment being first welded to one another for this purpose at their edges in gas-tight fashion in a vacuum, and the surfaces of the replacement section and the shroud segment being subsequently diffusion-bonded to one another by hot isostatic pressing.

Abstract: A method is provided for joining blades to blade roots or rotor disks when manufacturing and/or repairing gas turbine blades or integrally bladed gas turbine rotors. A blade and a blade root or rotor disk that is to be joined to the blade are supplied, and provided with a thickening in sections that are to be joined together. The sections of the blade and the blade root or rotor disk that are to be joined together are machined so as to form recesses and; c) are then aligned relative to each other, opposite recesses defining at least one groove-shaped seam preparation. The blade and the blade root or rotor disk are joined in the area of the or each seam preparation by means of laser power build-up welding and the joined blade and blade root or rotor disk are machined so as to provide a gas turbine blade or an integrally bladed gas turbine rotor having a predefined geometrical profile.

Abstract: A process for manufacturing a honeycomb seal is disclosed. The process includes the following steps: a) the honeycomb structure is brought to the desired three-dimensional form and is fixed in this form; b) then the honeycomb structure is filled with a filler material; c) then the honeycomb structure filled with the filler material is machined on a side on which a carrier element is arranged, such that this side has the desired contour and that the end edges of the honeycomb structure are sealed flushly with the filler compound; d) then the side contoured in this way is coated with a coating; e) next the carrier element is applied to the coating, preferably by laser powder build-up welding.

Abstract: The present technology generally relates to a method for producing a composite component, preferably, a composite gas turbine component, wherein the composite component exhibits a first component section made of a magnesium-based material or an aluminum-based material and a second component section made of a high-strength material, preferably, made of an iron-based or nickel-based or titanium-based or cobalt-based material.

Abstract: A sealing device is e.g., for a gas turbine, such as an aircraft engine. The sealing device, is used for sealing a gap between a rotor and a stator, e.g., for sealing a gap between radially inside ends of fixed vanes and a rotor and/or a gap between radially outer ends of rotating rotor blades and a fixed housing, a honeycomb seal including a plurality of honeycomb seal cells being assigned to the stator, and the honeycomb seal cells being separated from one another by walls. At least the walls of the honeycomb seal cells that extend transversely to the direction of rotation of the rotor are placed radially at a slant in the direction of rotation of the rotor.

Abstract: The invention relates to a method for the connecting of components, namely for the connecting of at least one pipe-shaped or rod-shaped component with a wall-type component, with the following steps: a) providing a wall-type component (11), whereby the wall-type component (11) comprises a sandwich-type structure of two plate-shaped or pipe-shaped elements (12, 13) with a layer (14) of an expandable foamable metal material positioned therebetween; b) providing at least one pipe-shaped or rod-shaped component (10); c) introducing at least one opening in the wall-type component (11) for the receiving of the or each pipe-shaped or rod-shaped component (10); d) inserting of the or each pipe-shaped or rod-shaped component (10) in a corresponding opening of the wall-type component (11); e) heating of the arrangement of the wall-type component (11) and the or each pipe-shaped or rod-shaped component (10) with foaming-up expansion of the expandable foamable metal material for the connecting of the or each pipe-shaped

Abstract: The invention relates to a method for repairing a structural component, especially a stator-side structural component of a gas turbine such as a housing or a guide vane ring, whereby a damaged section is separated-out from the structural component, and whereby a new section that replaces the damaged as well as separated-out section is fixedly or rigidly connected with the structural component by welding. According to the invention, the damaged section is separated-out from the structural component to be repaired in such a manner so that the length of a separating seam and therewith a later weld seam is minimized, whereby depending on the material thickness distribution along the separating seam, for providing a most uniform possible material thickness along the later weld seam, material is removed from the structural component, and whereby after the connecting of the structural component with the new section, at least the removed material is renewed by laser powder deposit welding.

Abstract: A milling tool which is used to mill recesses, in particular, circular-shaped grooves, in a workpiece is provided. The milling tool (10) is disposed on a disk-shaped or plate-shaped base body (14) and on at least one cutting body (15) arranged on the outer periphery of the base body (14). The cutting body or each cutting body (15) is inclined in relation to the disk-shaped base body or the plate-shaped base body (14).

Abstract: The invention relates to a gas turbine component, namely static gas turbine component, especially for an aircraft engine. According to the invention, the static gas turbine component is formed at least region-wise of metal foam.

Abstract: A method for producing gas turbine components, in particular blades, blade sections or rotors with integral blades for a jet engine is provided. The method comprises at least the following steps: preparation of at least one metallic powder and at least one expanding agent; mixing of the metallic powder or each metallic powder with the expanding agent or each expanding agent; compression of the resultant mixture to form at least one semi-finished product; expansion of the semi-finished product or each semi-finished product by heating to achieve a defined degree of expansion; termination of the expansion process by cooling, once the defined degree of expansion has been reached.

Abstract: A method for connecting dynamically loaded components, especially gas turbine components is provided. At least two components (10, 11; 14, 15) to be interconnected are connected by means of laser powder build-up welding.

Abstract: A method is for the cladding of blade tips of compressor blades of a gas turbine power plant, using a blank, which is made up of a solder foil provided with a metal layer having embedded hard particles and which is adapted to the geometry of the blade tip that is to be clad, which blank is melted onto the blade tip while applying pressure, and in the process obtains, impressed into it, a ribbed or pimpled structure.

Abstract: A method is provided for restoring geometrically complex components having damage or defects in regions of limited access, in which damaged or defective pieces are cut out of the components and replacement pieces of the same kind are welded in, the cutting and the welding being performed with the aid of laser energy. Using a CNC-controlled laser beam, a damaged or defective piece is cut out of a component. Using a laser beam and the same path data, modified by a correction if necessary, a replacement piece is cut out of a repair part. Using a laser beam and the same path data as in A, the replacement piece is welded into the component. A space-saving special adapter is used in the three foregoing machining steps. During the first three machining steps, the component remains in the same clamped support.

Abstract: A honeycomb seal which is manufactured by powder metallurgical injection molding is disclosed. The honeycomb seal is composed preferably of several segments. Each segment is embodied as a single piece and has a base element as well as honeycomb elements that are embodied as a single piece with the base element.

Abstract: A method for holding a metallic component of a gas turbine, which is to be connected, includes providing a metallic component having a surface and at least one machining or joining surface, providing a case having a cutout, which has an inner surface, for receiving the component, positioning the component in the case in such a manner that its surface is surrounded, at a distance, by the inner surface of the case, so as to form a closed volume and so that its machining or joining surface does not face the volume, filling the volume with a foamable material, foaming the material and cooling so as to form a dimensionally stable foam, structure. An apparatus for holding the metallic component and a method for connecting the metallic component to a further component are also provided.

Abstract: A method is available for repairing metallic components such as gas turbines, which have mechanically weight-carrying grooves, where the longitudinal edges of the grooves which exist on the component surface are largely in a plane and their actual contour deviates, at least locally, due to removal and/or displacement the component material from the desired contour. The method begins with abrasive blasting of the groove surfaces, followed by closing of the open groove ends by build-up welding, and subsequent filling of the grooves with soldering powder. The soldering powder is heated to liquefaction, then cooled and solidified in a vacuum, and the desired contour of the groove is then machined.

Abstract: The invention relates to a blade and a rotor for a gas turbine and to a method of joining blade parts of a gas turbine, a blade-body section (1) and at least one further blade part (4, 16, 21) being provided, and the blade-body section (1) being provided with at least one end joining area (2 , 3) and the further blade part (4, 16, 21) being provided with an end joining area (6, 18, 24), an inductor (10) being arranged so as to be adjacent to the joining areas (2, 3, 6, 18, 24) of the blade-body section (1) and of the second blade part (4, 16, 21), the joining areas (2, 3, 6, 18, 24) of the blade-body section (1) and of the further blade part (4, 16, 21) being positioned essentially in alignment with and at a distance from one another, and the blade-body section (1) and the second blade part (4, 16, 21) being welded to one another by energizing the inductor (10) with a high-frequency current and by bringing them together with their joining areas (2, 3, 6, 18, 24) touching one another (FIG. 1).

Abstract: A method is disclosed for reconditioning blades for turbo-machinery that have a material deficiency resulting from wear and tear, damage, or dimensions that are smaller than specified, in which material is removed, at least from parts of the blade, and is replaced with new material. The actual geometry of the blade is measured at certain specific points and is stored. The desired geometry, or the desired profile in its desired position, is determined via data processing technology. The desired geometry is compared with the actual geometry, and in cases of significant deviations in position, is adjusted to coincide with the actual geometry. The adjusted desired geometry is mathematically reduced by a defined amount of coating to be removed. The calculated net geometry is actually produced via the removal of material, at least over part of the blade, and new material is applied to the finished net geometry, until the desired profile is achieved.

Abstract: A method for repairing/manufacturing an integrally bladed rotor for a turbine or turbo engine, in which, in the case of a repair, at least one vane section to be replaced is removed leaving a stub section to which a replacement vane section is to be welded, and, in the case of manufacturing, a carrier to be provided with vanes has stubs projecting over its peripheral surface, to each of which a vane is to be welded, includes the steps of: arranging an inductor around the front face of the stub section or stub; positioning a replacement vane section or vane at a small radial distance from and aligned with the stub section or stub; and welding the replacement vane section or vane to the stub section or stub, respectively, in a protective gas atmosphere by exciting the inductor with high-frequency current and moving the opposing heated part surfaces together.

Abstract: A method is disclosed for reconditioning blades for turbo-machinery that have a material deficiency resulting from wear and tear, damage, or dimensions that are smaller than specified, in which material is removed, at least from parts of the blade, and is replaced with new material. The actual geometry of the blade is measured at certain specific points and is stored. The desired geometry, or the desired profile in its desired position, is determined via data processing technology. The desired geometry is compared with the actual geometry, and in cases of significant deviations in position, is adjusted to coincide with the actual geometry. The adjusted desired geometry is mathematically reduced by a defined amount of coating to be removed. The calculated net geometry is actually produced via the removal of material, at least over part of the blade, and new material is applied to the finished net geometry, until the desired profile is achieved.