A turbine assembly includes a first turbine blade circumferentially adjacent to a second turbine blade. The first turbine blade includes a first platform with a first underplatform surface. The second turbine blade includes a second platform with a second underplatform surface. The first underplatform surface and the second underplatform surface at least partially defines a cavity. A damper with a platform rub surface is included adjacent the first underplatform surface and the second underplatform surface. A seal is included with a forward portion, an aft portion, and a damper interface portion therebetween. The damper interface portion extends along a damper underside and the forward portion and the aft portion adjacent the first underplatform surface and the second underplatform surface.

An intermediate element is for a blade/rotor disk connection in a rotor of a fluid flow machine. The intermediate element is adapted to a shape of a blade root of a blade and to a blade root slot in a rotor disk for receiving the blade root such that, when arranged between the blade root and rotor disk, the intermediate element prevents contact between the blade root and rotor disk. The intermediate element has, on an outer surface that faces the rotor disk, at least one protrusion to reduce an air flow parallel to an axis of rotation of the rotor between the rotor disk and the intermediate element; and on an inner surface that faces the blade root, a recess that corresponds to the at least one protrusion.

The present disclosure provides a fir tree coupling for gas turbine engine parts comprising a load beam having a longitudinal axis, a rounded base, a first side, and a second side, wherein the rounded base has a radius of curvature from the first side to the second side, a tooth running parallel to the longitudinal axis and disposed on the first side of the load beam. The fir tree coupling may comprise a channel through the rounded base across a portion of the radius of curvature from the first side to the second side. The channel may comprise a sidewall having a sidewall step cut into a portion of the channel sidewall.

A load absorbing system that may include a rotor blade retention system is provided. The load absorbing system may include a block, a first retainer plate, and a deformable core. The block may be selectively positioned alongside a dovetail groove. The block may have a first face directed away from the blade root and an axially-spaced second face directed toward the blade root. The first retainer plate may be attached to the second face of the block and axially positioned between the block and the axially-directed surface of the blade root. The deformable core may be positioned between the block and the first retainer plate.

A rotor of a multi-staged axial compressor which extends along an axis of rotation. The rotor has a shaft which has rotor blade slots. Rotor blades of the rotor are arranged next to one another in the circumferential direction and are each secured to the rotor blade slots by a blade root to form a respective rotor blade stage. At least two rotor blade stages are provided in axial succession and an interspace slot, extending in the circumferential direction, is provided in the shaft axially between the two rotor blade stages. The rotor blade slots open into the interspace slots and blade roots of the rotor blades are insertable radially into the interspace slots and can be pushed into the rotor blade slots. The rotor has an interspace cover which covers the interspace slots, wherein the interspace cover is designed segmented into interspace cover segments in the circumferential direction.

A platform for a fan of a gas turbine engine includes a body portion, a forward flange at a front of the body portion that includes an engagement feature that interacts with a forward flange of a lug of a hub of a fan of a gas turbine engine, and a pin integral with a rear portion of the body portion. The pin is receivable in an aperture in a rear flange of the lug.

The invention proposes a fan, in particular for a turbomachine of small size such as a jet engine, having a hub ratio which corresponds to the ratio of the diameter of the inner limit of the incoming air stream at the radially inner ends of the leading edges of the fan blades, divided by the diameter of the circle around which the outer ends of the fan blades pass, having a value of between 0.20 and 0.265.

A blade retainer adapted for use in a gas turbine engine is configured to block axial movement of a blade. The blade retainer includes a first brace, a second brace, and a web that extends between the first brace and the second brace. The blade retainer is configured to block axial movement of a root of the blade out of a blade receiver slot.

A rotor assembly includes a rotor having a dovetail slot. The dovetail slot includes a plurality of recesses and a first radially innermost surface. A rotor blade includes an airfoil that extends radially outward from a platform and a dovetail that extends radially inward from the platform. The dovetail includes a plurality of projections extending in opposite directions that are received by the plurality of recesses of the dovetail slot. The dovetail further includes a leading edge surface, a trailing edge surface, and a second radially innermost surface. The second radially innermost surface defines a groove from the leading edge surface to the trailing edge surface. The shim is positioned within the groove and between the first radially innermost surface of the dovetail slot and the second radially innermost surface of the dovetail. The shim extends at least partially radially along both of the leading edge surface and the trailing edge surface to secure the rotor blade within the dovetail slot during various operating conditions of the turbomachine.

The invention relates to a vane (7) comprising: a structure made of composite material, a vane (7) root (16) attachment (9) comprising a base having a radial outer face (20) and in which a slot (24) configured to receive the vane (7) root (16) is formed and two platforms (26), extending on either side of the blade (18) opposite the radial outer face (20), and two stiffening parts (30), added and fixed on the radial outer face (20) on either side of the stilt (17) so as to tightly abut against the stilt (17).