Abstract: A method is provided for machining an airfoil section (12) of a turbine blade or vane produced by a casting process. The airfoil section (12) has an outer wall (18) delimiting an airfoil interior having one or more internal cooling passages (28). The method involves: receiving design data pertaining to the airfoil section (12), including a nominal outer airfoil form (40N) and nominal wall thickness (TN) data; generating a machining path by determining a target outer airfoil form (40T), the target outer airfoil form (40T) being generated by adapting the nominal outer airfoil form (40N) such that a nominal wall thickness (TN) is maintained at all points on the outer wall around the one or more internal cooling passages (28) in a subsequently machined airfoil section; and machining an outer surface (18a) of the airfoil section (12) produced by the casting process according to the generated machining path, to remove excess material to conform to the generated target outer airfoil form (40T).

Abstract: A gas turbine blade having a casted metal airfoil, the airfoil has a main wall defining at least one interior cavity, having a first side wall and a second side wall, which are coupled to each other at a leading edge and a trailing edge, extending in a radial direction from a blade root to a blade tip and defining a radial span from 0% at the blade root to 100% at the blade tip. The main airfoil has a radial span dependent chord length defined by a straight line connecting the leading edge and the trailing edge as well as a radial span dependent solidity ratio of metal area to total cross-sectional area. Solidity ratios in a machined zone of the airfoil from 80% to 85% of span are below 35%, in particular all solidity ratios in the zone.

Abstract: A bladed rotor system includes first and second sets of blades with respective airfoils each having at least one internal cavity. The airfoils of both the first and second sets of blades have identical outer shapes defined by an outer surface of an outer wall of the respective airfoils. The airfoils of the first set of blades are distinguished from the airfoils of the second set of blades by a geometry and/or position of at the least one internal cavity, which is unique to blades of a given set. The natural frequency of a blade of the first set differs from the natural frequency of a blade of the second set by a predetermined amount. The blades of the first set and the second set are alternately arranged in a periodic fashion in said circumferential row, to provide a frequency mistuning to stabilize flutter of the blades.

Abstract: A bladed rotor system includes first and second sets of blades with respective airfoils each having at least one internal cavity. The airfoils of both the first and second sets of blades have identical outer shapes defined by an outer surface of an outer wall of the respective airfoils. The airfoils of the first set of blades are distinguished from the airfoils of the second set of blades by a geometry and/or position of at the least one internal cavity, which is unique to blades of a given set. The natural frequency of a blade of the first set differs from the natural frequency of a blade of the second set by a predetermined amount. The blades of the first set and the second set are alternately arranged in a periodic fashion in said circumferential row, to provide a frequency mistuning to stabilize flutter of the blades.

Abstract: A gas turbine blade having a casted metal airfoil, the airfoil has a main wall defining at least one interior cavity, having a first side wall and a second side wall, which are coupled to each other at a leading edge and a trailing edge, extending in a radial direction from a blade root to a blade tip and defining a radial span from 0% at the blade root to 100% at the blade tip. The main airfoil has a radial span dependent chord length defined by a straight line connecting the leading edge and the trailing edge as well as a radial span dependent solidity ratio of metal area to total cross-sectional area. Solidity ratios in a machined zone of the airfoil from 80% to 85% of span are below 35%, in particular all solidity ratios in the zone.

Abstract: Provided is a method for determining the external and internal geometry of a component with at least one cavity, wherein a component to be measured is provided with at least one cavity, the external geometry of the component is determined by carrying out a 3D scan, the wall thickness of at least one section of the component is determined using ultrasound, the internal and external component geometry of at least one section of the component in particular is determined using x-ray computer tomography, and the data obtained by means of the 3D scan, the ultrasound wall thickness measurement, and in particular the x-ray computer tomography is combined, wherein the internal geometry of the component in the region of the at least one section measured using ultrasound is reconstructed from the external geometry data of the 3D scan and the data of the ultrasound wall thickness measurement.

Abstract: A method is provided for machining an airfoil section (12) of a turbine blade or vane produced by a casting process. The airfoil section (12) has an outer wall (18) delimiting an airfoil interior having one or more internal cooling passages (28). The method involves: receiving design data pertaining to the airfoil section (12), including a nominal outer airfoil form (40N) and nominal wall thickness (TN) data; generating a machining path by determining a target outer airfoil form (40T), the target outer airfoil form (40T) being generated by adapting the nominal outer airfoil form (40N) such that a nominal wall thickness (TN) is maintained at all points on the outer wall around the one or more internal cooling passages (28) in a subsequently machined airfoil section; and machining an outer surface (18a) of the airfoil section (12) produced by the casting process according to the generated machining path, to remove excess material to conform to the generated target outer airfoil form (40T).

Abstract: A turbine blade for a turbomachine, having a blade airfoil with a blade tip region and a sealing region which has an auxetic material and is arranged and designed such that, when transitioning from a rest state to an operating state of the turbine blade, the blade tip region of the turbine blade expands in a first direction perpendicular to a longitudinal axis of the blade airfoil. An auxetic material is used for a turbine blade for sealing a gas path during operation of a turbomachine.

Abstract: Provided is a machining center, in particular a 5-axis milling machine, for machining an elongate workpiece, in particular a turbine blade, including an elongate machine bed extending in a longitudinal direction, a machining device movably retained on the machine bed, two clamping devices, which are retained on the machine bed in such a way that the clamping devices lie opposite each other in the longitudinal direction and which are designed to clamp the free ends of the workpiece and to rotate the clamped workpiece about a workpiece axis of rotation, wherein at least one clamping device can be moved in the longitudinal direction of the machine bed and at least one clamping device has a drive device designed to rotate the clamped workpiece, in particular a stepper motor, and a steady rest, which is retained on the machine bed and positioned between the clamping devices.

Abstract: Ceramics or ceramic layer systems can be machined due to the use of high speeds of an EDM electrode. Disclosed is a method and an EDM device for carrying out the method for drilling ceramics or ceramic-coated components, which uses an EDM electrode and wherein the EDM electrode is rotated at a speed of at least 1000 rpm, in particular of at least 10,000 rpm, for drilling through the ceramic or through the ceramic-coated component.

Abstract: A method in which a hollow component is produced by the method steps consisting in casting and joining a sheet metal, wherein the hollow component can include thin walls and has a high recession in geometry is provided.

Abstract: A method of forming a component for a gas turbine engine, including: casting a component around a ceramic core, wherein the ceramic core forms a pilot channel (40) in the component, the pilot channel oriented from a base (176) to a tip (20) of the component; sinking an ECM electrode into the pilot channel; and enlarging the pilot channel to form an inner surface of an external wall (120) of the component via electro-chemical machining, wherein a contour (94) of the inner surface is different than a contour of the pilot channel.