An inner ring for supporting variable-pitch blades of an axial compressor stator of a turbomachine includes two or more ring sectors, assembled to form the inner ring, each ring sector including plural housings for receiving a pivot axis of one of the blades. At each junction plane between the ends of two successive ring sectors, an assembly pin and a cavity for receiving the assembly pin are engaged. The assembly pin is fixed on one end one of the two successive ring sectors and the cavity is formed at the end of the other of the two successive ring sectors. The cavity receives the assembly pin with a mounting clearance allowing a mounting by sliding fit, and the assembly pin is made of a material having a coefficient of thermal expansion greater than a coefficient of thermal expansion of a material constituting the successive ring sectors.

An impeller for a centrifugal compressor, the impeller rotatable about a central axis, has: an outer hub body including a first material and extending around the central axis, the outer hub body defining a gaspath face extending from an inlet to an outlet, the gaspath face extending radially away from the central axis from the inlet to the outlet; blades protruding from the gaspath face and circumferentially distributed around the central axis; and an inner hub body extending around the central axis, the inner hub body secured to the outer hub body, the outer hub body axially overlapping and extending around the inner hub body, the inner hub body made of a second material being more cold dwell resistant than the first material.

An airfoil including an airfoil body, a recessed portion of a first depth in a first side of the airfoil body, the recessed portion including a plurality of pockets of a second depth located within the recessed portion and ribs of the first depth located between the pockets, a cover configured to fit into the recessed portion such that an interior surface of the cover engages the ribs and an exterior surface of the cover is about flush with an exterior surface of the first side of the airfoil body, and a high energy beam weld configuration extending through the cover and into the ribs and positioned to attach the cover to the ribs.

Aspects of the disclosure are directed to a sync ring assembly associated with an engine, comprising: a sync ring, and a clevis coupled to the sync ring, where the clevis includes a rib that resists a deflection of the sync ring assembly when the engine is operated, and where the clevis is made of titanium.

A method of manufacturing a hollow aerofoil component 100 for a gas turbine engine 10 comprises joining a first panel 200 to a second panel 300 using bonding, and hot forming the panels into shape. The bonding step and the hot forming step are performed in the same rig, thereby optimizing process time and component quality.

A compressor blade for a gas turbine is provided. The compressor blade has a blade substrate that consists of a metal alloy and has an aluminum diffusion zone on a surface of the blade substrate. In addition, the compressor blade has a hard material coating provided on the surface of the blade substrate. A compressor that has a compressor blade and a method of producing such a compressor blade is also provided.

A fan for a gas turbine engine includes: an annular casing; a disk disposed inside the casing and mounted for rotation about an axial centerline, the disk including a row of fan blades extending radially outwardly therefrom; each of the fan blades including an airfoil having circumferentially opposite pressure and suction sides extending radially in span from a root to a tip, and extending axially in chord between spaced-apart leading and trailing edges, with the airfoils defining corresponding flow passages therebetween for pressurizing air; the row including no more than 21 and no less than 13 of the fan blades; and wherein each of the fan blades has a solidity defined by a ratio of the airfoil chord over a circumferential pitch of the fan blades, measured at 60% of a radial distance from the root to the tip, of less than about 1.6.

Platforms for airfoils in gas turbine engines are provided. The platforms include an aero-structure defining a gaspath surface and a non-gaspath surface, a first securing element arranged to receive a locking pin and secure the aero-structure to a rotor disk, and a first securing element structure extending between the non-gaspath surface and the first securing element, wherein the first securing element structure comprises at least two legs defining a hollow space between the two legs and the non-gaspath surface.

The present disclosure relates to turbine blades adapted for use in gas turbine engines. In particular, this disclosure is directed to turbine blades that include components made from ceramic matrix composite materials and that incorporate abradable materials.

A compressor rotor for a gas turbine engine has blades circumferentially distributed around and extending a span length from a central hub. The blades include alternating first and second blades having airfoils with corresponding geometric profiles. The airfoil of the first blade has a coating varying in thickness relative to the second blade to provide natural vibration frequencies different between the first and the second blades.