A turbine or compressor stage of a gas turbine, the rotor blade having a radially outer shroud (1) which has a sealing fin array having a first sealing fin (3.1) and a second sealing fin (3.2) which is adjacent to the first sealing fin and connected thereto by a first groove base (10) having a circumferential region (13) of maximum radial height, which is located at a first circumferential position is provided. The sealing fin array has a third sealing fin (3.3) adjacent to the second sealing fin and opposite to the first sealing fin, the third sealing fin being connected to the second sealing fin by a second groove base (20) having a circumferential region (23) of maximum radial height, which is located at a second circumferential position different from the first circumferential position.

A system for relieving stress on a turbine rotor blade dovetail in a gas turbine is provided. At least one turbine rotor blade includes a dovetail that is axially insertable into a correspondingly-shaped slot defined in a turbine disk. At least one axially-extending tang is defined on the dovetail. At least one stress relief surface is defined in the at least one tang. The at least one stress relief surface extends along a central portion of a length of the tang. Accordingly, contact between the at least one tang and an inner surface of the slot is precluded, along the central portion, such that stresses generated by radially-directed forces along the central portion are reduced.

An impeller blade formed of a central base portion having a center axis. At least a pair of impeller blades is extended outwardly from the central base portion, with each impeller blade further including an extension. A leading edge is formed on each of the impeller blades, with each leading edge being defined by an outer periphery of the central base portion, and each leading edge having at least a portion that forms a continuous curve with a radius decreasing outwardly toward a radial tip of the corresponding impeller blade, wherein the continuous decreasing radius curve of each leading edge is formed by a portion of an ellipse centered on the center axis, a short axis of the ellipse being equal to a long axis thereof divided by approximately three and one-half. Each blade extension is projected from the corresponding impeller blade opposite of the leading edge.

After a free-form surface is machined on an elongated material 1 with a projection 3 and a blade root 4 held, the holding of the projection 3 is released to release strain generated during machining. Upon release of the holding, the entire elongated material 1 deforms, and the projection 3 moves from a holding position A to a strain-released position B. A re-holding position C obtained by correcting the position B by the deformation amount of the elongated material 1 due to the weight of the elongated material 1 is determined, and the projection 3 is held again at the re-holding position C for further machining the free-form surface on the elongated material 1.

A wind turbine blade 2 comprises a profiled contour including a leading edge 34 and a trailing edge 33 as well as a pressure side and a suction side. The profiled contour is formed by a first shell part 10 and a second shell part 15 being bonded together in a bonding region between the first and the second shell part by a curable bonding means 40. The first and the second shell part 10; 15 are formed in a fiber-reinforced polymer. The wind turbine blade further comprises resistive heating means 50 being arranged in thermal connection with the bonding means 40 such that the resistive heating means 50 supplies heat for curing of the curable bonding means 40 during assembling of the wind turbine blade.

An automated technique for finishing gas turbine engine blades or vanes by generating a bespoke tooling path for each blade or vane. The bespoke tooling path is generated by scanning the aerofoil surface to generate a 3-D electronic representation of the surface. The 3-D electronic surface is then analyzed to identify imperfections or defects, and then a machining path a generated through which the imperfections can be removed. The machining path is determined so as to smoothly blend the surface back to the underlying surface where the imperfections had been present. In this way, the resulting aerofoil, once machined, has optimized aerodynamic performance.

A method of manufacturing a blade for use with a rotary machine includes coupling a suction-side section of a part-span shroud to a suction-side surface of an airfoil of the blade. A first point of the airfoil suction-side surface, defined where a trailing edge of the suction-side section intersects the airfoil suction-side surface, is upstream from a second point, defined where a throat intersects the airfoil suction-side surface. A third point, defined where a leading edge of the suction-side section intersects the airfoil suction-side surface, is downstream from a fourth point, defined at a leading edge of the blade. The method also includes coupling a pressure-side section of the part-span shroud to a pressure-side surface of the airfoil. A fifth point of the airfoil pressure-side surface, defined where a trailing edge of the pressure-side section intersects the airfoil pressure-side surface, is downstream from the first point.

An impeller for a turbo-machine is provided. The impeller includes a metallic base having a back surface and a front surface opposite to the back surface; at least one blade extending on the front surface of the metallic base; a composite shroud attached to the at least one blade such that plural closed paths are formed by the front surface of the metallic base, the at least one blade and the composite shroud; and a connector configured to attach the composite shroud to the metallic base or the at least one blade.

The invention provides a sectional blade for a wind turbine. The blade comprises at least a first blade portion and a second blade portion extending in opposite directions from a joint. Further each blade portion comprises a spar section forming a structural member of the blade and running lengthways. The first blade portion and the second blade portion are structurally connected by at least one spar bridge extending into both blade portions to facilitate joining of said blade portions and the spar bridge joins the spar sections.

Disclosed is a process for producing a blade or vane ring segment for a gas turbine, in particular for an aero engine, and also to a correspondingly produced blade or vane ring segment, the process comprising: forging at least two blanks made of a TiAl material, joining the blanks to form a blade or vane ring by means of an integral connection process, and remachining of the blank composite by material-removing processes.