A shaft for a gas turbine engine includes an inner contour with a stiffening rib that defines a stiffened wall thickness related to a nominal wall thickness according to a ratio between about 1.125-2.1.

A rotor assembly is provided for a gas turbine engine. This rotor assembly includes a first rotor disk, a second rotor disk and a plurality of rotor blades. The first rotor disk is configured to rotate about a rotational axis. The second rotor disk is configured to rotate about the rotational axis. The rotor blades are arranged circumferentially around the rotational axis. Each of the rotor blades is mounted to the first rotor disk and to the second rotor disk. The rotor blades include a first rotor blade. The first rotor blade includes an attachment projecting axially along the rotational axis into a first pocket in the first rotor disk and a second pocket in the second rotor disk. The attachment has a dovetail cross-sectional geometry when viewed in a plane perpendicular to the rotational axis. A portion of the first rotor disk extends circumferentially across and covers the attachment.

An airfoil of a gas turbine engine includes a leading edge and an opposed trailing edge defining a chord between the leading edge and the trailing edge, wherein the chord has a chord length. A concave surface is between the leading edge and the trailing edge, which includes a first portion proximal the leading edge of the airfoil and a second portion proximal the trailing edge of the airfoil, wherein the first portion of the concave surface includes about 10% to about 50% of the chord length. An erosion-resistant ceramic, cermet or intermetallic coating is on the second portion of the concave surface, which includes a coating leading edge pattern. The first portion of the concave surface is free of the erosion-resistant coating.

A coupling adapted for use with a gas turbine engine includes a first flange, a second flange, and a seal assembly. The first flange is coupled with an exhaust end of the gas turbine engine. The second flange is coupled with ducting downstream of the gas turbine engine. The seal assembly extends between the first flange and the second flange and is configured to block the gases from passing radially between the first flange and the second flange.

An airfoil for a turbine engine can include an outer wall bounding an interior and defining a pressure side and a suction side extending between a leading edge and a trailing edge to define a chord-wise direction and extending between a root and a tip to define a span-wise direction. A plurality of outlets and a plurality of scalloped portions can extend along the outer wall.

A turbine engine can comprise a fan section, compressor section, a combustion section, and a turbine section in axial flow arrangement. At least one of the fan section and compressor section can include an airfoil with a leading edge, and a plurality of riblets can be arranged on the leading edge.

A turbine ring assembly including ring sectors forming a turbine ring and a ring support structure, each ring sector having, along a section plane defined by an axial direction and a radial direction of the ring, a portion forming an annular base with, in the radial direction an inner face defining the inner face of the ring and an outer face from which a first and a second attachment tabs protrude, the structure including a central shroud from which a first and a second radial clamps protrude between which the first and second attachment tabs of each ring sector are maintained. It includes a first and a second annular flanges removably fastened to the first radial clamp of the central shroud and separated from each other by a contact abutment.

An aerofoil component (10) defining an in use leading edge (16) and a trailing edge (18), the leading edge (16) comprising at least one serration (22) defining a sawtooth profile projecting in a generally upstream direction, each serration (22) comprising a plurality of rectangular-wave projections (26) extending in an upstream chordwise direction from the sawtooth profile.

A component for a gas turbine engine may comprise a cooling channel. A first trip strip may extend from a wall of the cooling channel. A leg of the first trip strip and a first plane orthogonal to a direction of airflow through the cooling channel may form a first angle. A second trip strip may be downstream from the first trip strip and extending from the wall of the cooling channel. A leg of the second trip strip and a second plane orthogonal to the direction of airflow through the cooling channel may form a second angle different from the first angle.

A rotor assembly is provided for a gas turbine engine. This rotor assembly includes a first rotor disk, a second rotor disk and a plurality of rotor blades. The first rotor disk is configured to rotate about a rotational axis. The second rotor disk is configured to rotate about the rotational axis. The rotor blades are arranged circumferentially around the rotational axis. Each of the rotor blades is mounted to the first rotor disk and to the second rotor disk. The rotor blades include a first rotor blade. The first rotor blade includes an attachment projecting axially along the rotational axis into a first pocket in the first rotor disk and a second pocket in the second rotor disk. The attachment has a dovetail cross-sectional geometry when viewed in a plane perpendicular to the rotational axis. A portion of the first rotor disk extends circumferentially across and covers the attachment.