A turbine impeller includes: a hub portion coupled to an end of a rotational shaft; a plurality of main blades disposed at intervals on a peripheral surface of the hub portion; and a short blade disposed between two adjacent main blades among the plurality of main blades. An inter-blade flow channel is formed between the two adjacent main blades so that a fluid flows through the inter-blade flow channel from an outer side toward an inner side of the turbine impeller in a radial direction. In a meridional plane, a hub-side end of a leading edge of the short blade is disposed on an inner side, in the radial direction, of a hub-side end of a leading edge of the main blade.

A turbine impeller includes: a hub portion coupled to an end of a rotational shaft; a plurality of main blades disposed at intervals on a peripheral surface of the hub portion; and a short blade disposed between two adjacent main blades among the plurality of main blades. An inter-blade flow channel is formed between the two adjacent main blades so that a fluid flows through the inter-blade flow channel from an outer side toward an inner side of the turbine impeller in a radial direction. In a meridional plane, a hub-side end of a leading edge of the short blade is disposed on an inner side, in the radial direction, of a hub-side end of a leading edge of the main blade.

A moving blade 21d is disposed in a last stage closest to a diffuser 10 among a plurality of stages of a turbine 9 including a turbine rotor 12 and a stationary body 14. The diffuser 10 is connected to an outlet side of the stationary body 14. A distal end of the moving blade 21d is opposed to a seal fin 38 provided in the stationary body 14. The moving blade 21d includes a blade section 26, a cover 27 and a guide 32 provided on a moving blade distal end face 31, which is a surface of the cover 27. The moving blade distal end face 31 extends in a rotation axis direction of the turbine rotor 12, and the guide 32 includes a guide surface 41 located on a side close to the diffuser 10 with respect to the seal fin 38 and formed to incline upward in a direction from the seal fin 38 toward the diffuser 10.

A coupon for replacing a cutout in a hot gas path component of a turbomachine is provided. In one embodiment, the coupon includes a body having an outer surface; and a plurality of grinding depth indicators in the outer surface of the body. In another embodiment, the coupon includes a body having an edge periphery configured to mate with an edge periphery of the cutout, and at least a portion of the edge periphery of the body has a wall thickness greater than a wall thickness of an edge periphery of the cutout. The embodiments may be used together or separately.

A turbine engine that includes a stationary assembly, and a rotor assembly configured to rotate relative to the stationary assembly. The rotor assembly includes a plurality of unitary turbine and fan blades. Each unitary turbine and fan blade includes a single turbine airfoil, a single fan airfoil positioned radially outward from the single turbine airfoil, and a midspan shroud segment defined between the single turbine airfoil and the single fan airfoil.

The invention provides a fitted platform (1) for positioning between two adjacent blades of an aviation turbine engine fan, said platform comprising a flow passage wall (10) made of composite material having a central portion (16) and first and second margins (18) each extending in a longitudinal direction of said wall, each margin extending over a determined distance (D) from the central portion (16) in a transverse direction of said wall, said flow passage wall comprising fiber reinforcement densified by a matrix, the platform being characterized in that the fiber reinforcement present in the central portion (16) presents three-dimensional weaving, and in that the fiber reinforcement present in the first and second margins (18) presents two-dimensional weaving, at least in part. The invention also provides a fan module, a turbine engine, and a method of fabricating such a platform.

A method for manufacturing a blade in composite material having an added metal leading edge for gas turbine aeroengine, includes producing a blade body in composite material including in longitudinal direction, a blade root part, a shank part and an airfoil body part; manufacturing, via additive manufacturing, a leading edge extending in longitudinal direction between a lower end present at the shank part in composite material and an upper end present at the tip; bonding the manufactured leading edge onto the foremost edge portion of the airfoil body of the blade body in composite material.

A gas turbine engine includes a turbine section that has a plurality of turbine vanes. Each of the turbine vanes includes inner and outer platforms and an airfoil section that extends there between. The airfoil section is hollow and rib-less and has a first end at the outer platform and a second end at the inner platform. The airfoil section is tangentially bowed from the first end to the second end with a radius of curvature that is from 17 centimeters to 130 centimeters.

A blade for a turbo machine is provided. The blade for a turbo machine includes an airfoil body extending in a radial direction and including a suction side surface and a pressure side surface opposite to the suction side surface with respect to a circumferential direction extending across the radial direction, and a snubber structure including a first snubber element protruding in the circumferential direction from the suction side surface of the airfoil body and a second snubber element protruding in the circumferential direction from the pressure side surface of the airfoil body, wherein the first snubber element is connected to the suction side surface of the airfoil body by a concave curved first transition portion having a first radius, and the second snubber element is connected to the pressure side surface of the airfoil body by a concave curved second transition portion having a second radius, the first radius being smaller than the second radius.

A blade for turbo machine is provided. The blade for a turbo machine includes an airfoil body extending in a radial direction between a root end and a tip end and including an inner void extending from the root end in the radial direction, a root body integrally formed with the airfoil body, extending from the root end of the air foil body to a bottom end in the radial direction, and including a receiving slot extending from the bottom end in the radial direction and opening into the inner void of the air foil body, and an insert positioned in the receiving slot of the root body and including a plurality of through holes extending in the radial direction to form a fluid connection to the inner void of the air foil body.