Blade root receptacle for receiving a blade root of a rotor blade of a turbomachine. The blade root receptacle, for radially bearing in a form-fitting manner on the blade root, has a supporting flank which, in terms of a rotation axis, at least in proportions faces radially inward, wherein the supporting flank is provided with a convexity which, when viewed in an axially perpendicular section, at least in portions has a convex shape and, also when viewed in an axially parallel section, at least in portions has a convex shape.

The present invention relates to a pressure wall for a centrifugal pump for fluid having substantially the shape of a disc, the disc-shaped pressure wall having a central axis, the pressure wall comprising: a top surface; and a bottom surface opposing the top surface; wherein the top surface includes an inner surface section and an outer surface section, wherein the inner surface section extends radially from the central axis and is recessed to form a central recess; and wherein the outer surface section includes an inner circumferential edge portion and an outer circumferential edge portion, wherein the inner circumferential edge portion is located closer to the central axis than the outer circumferential edge portion, and wherein the outer circumferential edge portion is located higher than the inner circumferential edge portion with respect to a plane perpendicular to the central axis and passing through the inner circumferential edge portion.

An integrated bladed rotor of a gas turbine engine is provided. The integrated bladed rotor includes a hub having a rotation axis and a radially outer platform relative to the rotation axis, and a plurality of blades extending radially outwardly from the outer platform of the hub. The blades are integrally formed with the hub to define a monolithic component with the hub. Two or more of the blades each include: an airfoil including a groove formed in an outer surface of the airfoil to mitigate crack propagation, and a root fillet providing a transition between the outer platform of the hub and the airfoil.

The film cooling structure includes a wall part and a cooling hole inclined such that an outlet is positioned rearward of an inlet. The cooling hole includes a straight-tube part and a diffuser part. The diffuser part includes a flat surface, a curved surface curved rearward and forming, together with the flat surface, a semicircular or semi-elliptical channel cross section larger than that of the straight-tube part, a first section and a second section extending from the first section toward the outlet. In the first section, an area of the channel cross section increases as it approaches the outlet. In the second section, the area of the channel cross section increases as it approaches the outlet at an increase rate smaller than that of the first section or is constant. The diffuser part has a width equal to or twice greater than the depth of the diffuser part.

Curved beams stacked structures-compliant shrouds for gas turbine engines are disclosed. An example shroud assembly comprising a plurality of concave curved beams, a plurality of convex curved beams, and a plurality of bumpers, wherein a first concave curved beam of the plurality of concave curved beams is inversely coupled to a first convex curved beam of the plurality of convex curved beams, a second concave curved beam of the plurality of concave curved beams, inversely coupled to a second convex curved beam of the plurality of convex curved beams, the first and second concave curved beams configured to stack on top of the first and second convex curved beams, a first bumper of the plurality of bumpers coupled to the first and second concave curved beams, and a second bumper of the plurality of bumpers coupled to the first and second convex curved beams.

A sealing device according to at least one embodiment includes not less than three arc-shaped fins arranged in an axial direction. The arc-shaped fins include: a first fin which is one of two outermost fins located on an outermost side in the axial direction; a second fin disposed adjacent to the first fin in the axial direction; and at least one third fin disposed opposite to the first fin across the second fin in the axial direction. It is preferable that the third fin is disposed to be inclined with respect to a radial direction such that a tip end portion is located on a side of the first fin in the axial direction relative to a base end portion, and the third fin has a larger inclination angle than the first fin or the second fin with respect to the radial direction.

A gas turbine has blades. A blade may have a leading edge; a trailing edge; a pressure side and a suction side, which extend between the leading edge and the trailing edge. The blade has, along the leading edge, a leading edge profile with profile portions, each of which, along its profile portion length, transitioning, proceeding from a depression, into an elevation via a first transition portion and back into a next depression via a second transition portion. An apex of the elevation of a profile portion is arranged in an asymmetric manner in relation to the profile portion length, in such a way that the first transition portion has a first transition length and the second transition portion has a second transition length. The first transition length and the second transition length are different lengths.

A gas turbine has blades. A blade may have a leading edge; a trailing edge; a pressure side and a suction side, which extend between the leading edge and the trailing edge. The blade has, along the leading edge, a leading edge profile with profile portions, each of which, along its profile portion length, transitioning, proceeding from a depression, into an elevation via a first transition portion and back into a next depression via a second transition portion. An apex of the elevation of a profile portion is arranged in an asymmetric manner in relation to the profile portion length, in such a way that the first transition portion has a first transition length and the second transition portion has a second transition length. The first transition length and the second transition length are different lengths.

A damper seal for a gas turbine engine includes a damper body extending in a first direction between a leading edge portion and a trailing edge portion, extending in a second direction between first and second sidewalls, and extending in a third direction between a convex outer damper face and a concave inner damper face. The inner damper face establishes a damper pocket. The leading and trailing edge portions slope inwardly from opposite ends of the damper body to bound the damper pocket in the first direction. The first and second sidewalls extend from the leading edge portion to the trailing edge portion and slope inwardly from opposite sides of the damper body to bound the damper pocket in the second direction. The outer damper face is pre-formed according to a first predetermined geometry that substantially corresponds to a second predetermined geometry of a platform undersurface bounding a neck pocket of an airfoil. A method of damping for a gas turbine engine is also disclosed.

A turbine component includes a body having a pair of spaced walls, with at least one of the walls for facing a fluid flow when mounted in a gas turbine engine. There are a plurality of wall cooling passages having a generally boomerang shape such that a peak apex is spaced from the wall and an indent apex is adjacent to the wall, with the plurality of wall cooling passages having interior sides extending from the peak apex toward the wall to define a corner. Outer sides extend from the corners with a component away from the wall and to the indent apex. A gas turbine engine is also disclosed.