A hub assembly for construction of a fan includes a hub plate having a front face, a rear face, and an annular surface extending between the front face and the rear face. The hub plate defines a hub assembly axis extending through the front face and the rear face which is configured to coincide with the fan axis. The hub assembly further includes a plurality of elongated root sections that are circumferentially spaced apart and extend radially from the annular surface. The elongated root sections have a substantially circular cross-section. The hub plate and the plurality of elongated spokes are unitarily formed.

Low-pressure turbine rotors for a thermoelectric power station include an interconnection of slender blades. The rotor includes blades bearing caps being roughly in the shape of a chevron.

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 rotor assembly is provided for a gas turbine engine. This rotor assembly includes a first rotor disk, a second rotor disk, a plurality of rotor blades and a plurality of disk mounts. 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. Each of the disk mounts connects the first rotor disk and the second rotor disk together. The disk mounts include a first disk mount that further supports the first rotor blade.

A platform for a gas turbine engine according to an example of the present disclosure includes, among other things, a platform body that extends in a circumferential direction between first and second sidewalls to define a gas path surface, and opposed rows of flexible retention tabs that extend from the platform body and are dimensioned to wedge the platform between adjacent airfoils.

A hub assembly for construction of a fan includes a hub plate having a front face, a rear face, and an annular surface extending between the front face and the rear face. The hub plate defines a hub assembly axis extending through the front face and the rear face which is configured to coincide with the fan axis. The hub assembly further includes a plurality of elongated root sections that are circumferentially spaced apart and extend radially from the annular surface. The elongated root sections have a substantially circular cross-section. The hub plate and the plurality of elongated spokes are unitarily formed.

An assembly for a gas turbine engine includes an airfoil that includes an airfoil section that extends from a root section. The airfoil section extends between a leading edge and a trailing edge in a chordwise direction and extends between a tip portion and the root section in a radial direction. The airfoil section defines a pressure side and a suction side separated in a thickness direction. A platform is dimensioned to receive a retention pin to mount the platform to a rotatable hub. The platform includes a plurality of composite layers that define an internal cavity. A filler includes a stacked composite structure in the internal cavity that extends between the plurality of composite layers.

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.

A rotor assembly for a gas turbine engine includes a rotatable hub that has a main body and an array of annular flanges that extend about an outer periphery of the main body to define an array of annular channels. An array of airfoils are circumferentially distributed about the outer periphery. Each one of the airfoils includes an airfoil section extending from a root section received in the annular channels. Retention pins extend through the root section of a respective one of the airfoils and through the array of annular flanges to mechanically attach each root section to the hub.