Abstract: The invention relates to a method for producing integrally bladed gas turbine rotors, especially of integrally bladed turbine rotors, wherein rotating blades (10) are joined to a rotor base (11) with the intermediate disposal of an adapter (14) in such a way that first an adapter is joined to a connecting section (15) of each rotating blade to be connected to the rotor base, and that subsequently each rotating blade (10) is joined to a connecting section (16) of rotor base (11) via adapter (14) connected to said blade. According to the invention, each adapter (14) is connected to connecting section (15) of the respective rotating blade (10) by welding, each rotating blade (10) being connected to the respective connecting section of the rotor base by inductive high-frequency pressure welding via adapter (14) connected to the blade.

Abstract: A method for producing a component of a turbomachine is disclosed. The method includes a) layer-by-layer deposition of a powder component material onto a component platform in a region of a buildup and joining zone, where the deposition takes place in accordance with layer information of the component to be produced; b) local layer-by-layer fusion or sintering of the powder component material by energy supplied in the region of the buildup and joining zone, where the buildup and joining zone is heated to a temperature just below a melting point of the powder component material; c) layer-by-layer lowering of the component platform by a predefined layer thickness; and d) repetition of steps a) to c) until the component is finished. A device for producing a component of a turbomachine is also disclosed.

Abstract: A blade-disk unit for a turbine engine includes a disk and a plurality of blades which are integrally connected to the disk, a welding seam being disposed in the joining region between the blade and the disk, which welding seam is situated entirely either in the blade or in the disk.

Abstract: A method for replacing a blade (12) of a rotor (2) having integral blades, wherein a new blade (22) is joined to a base (34) arranged on a separation surface (28) on the main rotor body, and to a rotor repaired or mended according to said method.

Abstract: A method for hardfacing a metal component surface (14, 16), especially a shroud surface of a turbine blade made of a TiAl alloy, with at least one metal material (18, 20), in particular a Co—Cr alloy. The hardfacing coating is produced separately from the component surface and is then joined to the component surface in a high-temperature soldering process. A turbine blade including such a hardfacing coating, primarily in a shroud region (2).

Abstract: The present invention relates to a method for replacing an inner disk element of a blade integrated disk wherein the blade integrated disk has a plurality of recesses in the form of axial through-bores in the joining area. The method begins with filling the axial through-bores, after which the inner disk element of the blade integrated disk is separated from the remaining blade ring. Next a new inner disk element is inserted into the remaining blade ring and the new inner disk element and the remaining blade ring are joined. Finally the axial through-bores are restored.

Abstract: The invention relates to a method for producing integrally bladed gas turbine rotors, especially of integrally bladed turbine rotors, wherein rotating blades (10) are joined to a rotor base (11) with the intermediate disposal of an adapter (14) in such a way that first an adapter is joined to a connecting section (15) of each rotating blade to be connected to the rotor base, and that subsequently each rotating blade (10) is joined to a connecting section (16) of rotor base (11) via adapter (14) connected to said blade. According to the invention, each adapter (14) is connected to connecting section (15) of the respective rotating blade (10) by welding, each rotating blade (10) being connected to the respective connecting section of the rotor base by inductive high-frequency pressure welding via adapter (14) connected to the blade.

Abstract: A method for producing a blisk (bladed disk) or a bling (bladed ring) of a gas turbine is disclosed. The method includes the following steps: a) producing a turbine blade by joining a blade to an adapter element consisting of a metal material that is suitable for fusion welding, the adapter element being used to form a blade root of the turbine blade, and b) joining the turbine blade to a rotor disk consisting of a metal material that is suitable for fusion welding or to a rotor ring consisting of a metal material that is suitable for fusion welding in such a manner that the turbine blade is arranged on the outer periphery of the rotor disk or of the rotor ring. A component of a gas turbine or of a high-pressure or low-pressure compressor, especially a blisk or bling, and a turbine blade are also disclosed.

Abstract: The invention relates to a device for the production or repair of a blade for a gas turbine, said device comprising: a holding device for holding a blade or an aerofoil of a blade; a lowering mechanism for lowering the blade or the aerofoil; a container for receiving power and/or for receiving the blade or the aerofoil; a heating device; and a radiation source, wherein the heating device is formed as an induction coil.

Abstract: A method for repairing turbine blades by replacing at least a part of the blade profile, the method having: a) manufacture of a replacement blade part; b) separation of the damaged area with a standardized cutting plane, leaving behind a remaining blade; c) matching of the replacement blade part to the actual geometry of the remaining blade; d) connecting or joining of the replacement blade part to the remaining blade by soldering of at least one web and welding of an outer contour.

Abstract: The invention relates to a procedure for repair of high pressure turbine blades of an aircraft engine, with the steps: —separating a damaged section of the high pressure turbine blade; and—generating a section to replace the separated section on or upon the high pressure turbine blade by means of laser beam generation from the powder bed, as well as a high pressure turbine blade.

Abstract: A method is for repairing turbo machine blades, e.g., gas turbine blades on integrally bladed gas turbine rotors. The method includes at least the following steps: a) providing a turbo machine blade to be repaired or an integrally bladed rotor to be repaired having a damaged blade section; b) removing the damaged blade section from the turbo machine blade or from the integrally bladed rotor resulting in a blade stump; c) clamping the blade stump at least in the region of a flow leading edge and a flow trailing edge of the same in a cooling device; d) applying respectively a punctiform welding bead in the region of the flow leading edge and a flow trailing edge of the blade stump by laser powder build-up welding, the welding beads protruding at least over the flow leading edge and the flow trailing edge; e) applying material by laser powder build-up welding between the welding beads in the region of the or each removed blade section for restoring the respective blade section.

Abstract: A gas turbine rotor and a method for manufacturing a gas turbine rotor having integral blading, is disclosed. A rotor blade has a blade pan and a blade footing connected to the blade pan. The blade footing is designed with a V-shaped cross section in at least some portions and the blade footing is welded to a rotor mount by capacitor discharge welding.

Abstract: A method and apparatus for the fabrication and/or repair of an integrally bladed rotor, specifically of an integrally bladed gas turbine rotor, is disclosed. At least one rotor blade is joined to a main rotor body for fabrication and/or repair. The joining is carried out by diffusion welding.

Abstract: The invention relates to a rotor blade for gas turbine rotors for manufacturing gas turbine rotors having integral blading. The rotor blade has a blade pan and a blade footing connected to the blade pan. According to this invention, for manufacturing a gas turbine rotor having integral blading, the blade footing is adapted by capacitor discharge welding such that the blade footing is designed with a V-shaped cross section in at least some portions.

Abstract: The invention relates to a method of producing an aperture in a metallic component, in which the aperture comprises, at least in certain portions, a non-cylindrically formed funnel, extends from a first surface to a second surface of a component wall and is formed by a laser beam. The invention improves the dimensional accuracy and reduces the roughness of the surface of the aperture and/or funnel in that, by appropriate choice of the laser parameters, the metal is predominantly removed by sublimation during the formation of the funnel and the funnel is formed by laser removal of layers of a substantially constant thickness.

Abstract: The invention relates to a method of producing an aperture (2) in a metallic component (1), in which the aperture (2) comprises, at least in certain portions, a non-cylindrically formed funnel (9), extends from a first surface (5) to a second surface (6) of a component wall (3) and is formed by a laser beam (10), it being attempted to improve the dimensional accuracy and reduce the roughness of the surface of the aperture and/or funnel in that, by appropriate choice of the laser parameters, the metal is predominantly removed by sublimation during the formation of the funnel (9) and the funnel (9) is formed by laser removal of layers (11) of a substantially constant thickness (FIG. 3).

Abstract: A worked nozzle ring for a gas turbine, in particular an aircraft engine, having at least one shroud and at least one blade sheet arranged thereon, whereby, in order to create a worked nozzle ring that is easy to manufacture, the blade sheet comprises, at the frontal side, at least one two-dimensionally profiled connecting space region that is placed in a profiled opening formed in the shroud by beam cutting and is connected therewith.

Abstract: A worked nozzle ring for a gas turbine, in particular an aircraft engine, having at least one shroud and at least one blade sheet arranged thereon, whereby in order to create a worked nozzle ring that is easy to manufacture, the blade sheet comprises, at the frontal side, at least one two-dimensionally profiled connecting space region that is placed in a profiled opening formed in the shroud by means of beam cutting and is connected therewith.

Abstract: In a method of repairing worn blade tips of compressor or turbine blades, the worn blade tip (4) is first removed from the remaining blade portion (2) at a standardized height (h). The actual geometry of the end of the remaining blade portion (2) at the height (h) is measured, for example by optical digital image processing means. A repair part (4') is cut from a plate or sheet of repair material having a thickness (d) essentially corresponding to the difference between the nominal finished height (H) of the repaired turbine blade (1) and the standardized height (h) of the remaining blade portion (2) undergoing repair. The repair part (4') is cut to have a contour corresponding to the measured actual geometry of the end of the remaining blade portion (2) at the standardized cut-off height (h), for example using a laser cutting process CNC-controlled by the acquired data representing the actual blade geometry.