A fan for a gas turbine engine, the fan comprising a first set of fan blades, each fan blade comprising a root and an axial retention feature provided on the root at a first position, and a second set of fan blades, each fan blade comprising a root and an axial retention feature provided on the root at a second position. Relative to the respective blade on which the respective retention feature is provided, the first position is different to the second position.

A rotor disc of a gas turbine rotor assembly is provided. The rotor disc may support a plurality of blades attached thereto. The rotor disc may have a plurality of fixing members defined therein through its peripheral surface and circumferentially spaced apart from one another. The fixing members may extend axially from a front end portion to a rear end portion of the disc. Profiled slots may be defined between pairs of adjacent ones of the fixing members and may be configured to receive complementary profiled blade root portions. A plurality of disc lips may project axially forward from a surrounding surface of the disc in the front end portion. A disc lip may be disposed at a tip portion of a fixing member, adjacent a leading edge of the disc to minimize air leakage at a disc/blade interface. A rotor assembly with such rotor disc and a plurality of blades attached thereto is also provided.

A gas turbine that includes a rotor blade that includes an airfoil. The airfoil may include non-integral base and top portions. The airfoil may include a pin connector connecting the top portion to the base portion. The pin connector may include: a tab extending from the top portion; a complimentary slot for receiving the tab formed in the base portion; an elongated pin cavity formed through an interior region of the airfoil, where the pin cavity intersects the slot to divide the pin cavity into first and second pin cavity segments; a tab aperture formed through the tab, where the tab aperture is positioned so align with the pin cavity upon the tab being received within the slot; and a locking pin that extends continuously through the first segment of the pin cavity, the tab aperture, and the second segment of the pin cavity.

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).

An assembly for receiving in a radially extending bucket groove (31) of a turbine disc (22), the assembly including a turbine blade having an elongate body (26) of aerofoil cross-section, a root portion (24) at one end of the elongate body (26) and a tip at the other. The root portion (24) configured to be retained in the bucket groove (31). A heat shield (27) configured to be received between an end of the root portion (24) and a radially inner surface of the bucket groove (31) and to interlock with the root portion (24) in a manner which deters separation in a radial direction.

Fixing set for securely assembling and locking a blade shaped element, for instance a turbine blade, on a support, for instance a turbine rotor. The blade shaped element is adapted to be inserted into a groove of the external surface of the support and the fixing set includes first, second and third locking elements adapted to be inserted into the groove, between the lower end surface of the root portion of the blade shaped element and the groove base surface. The first, second and third locking elements are adapted to securely lock the blade shaped element on the support and are further adapted to allow easy and simple installation by an operator with no need of additional working and extra machining.

A turbine blade includes a blade having a first end, and a second end opposite the first end. A tip is located at an outer radial portion of the blade. A blade dovetail complements a corresponding dovetail slot in a turbine rotor. The dovetail has a body, and a plurality of projections extending from the body in opposing directions for complementing a plurality of recesses in the corresponding dovetail slot. A tapered groove extends through the body from the first end to the second end. The tapered groove has a tapered profile such that a first depth of the tapered groove near the first end is greater than a second depth of the tapered groove near the second end. The tapered profile gradually transitions from the first depth to the second depth, so that the tapered groove is open at a bottom surface and sized to engage a shim.

A turbine blade has a first end, and an opposing second end. A base includes a dovetail for complementing a corresponding dovetail slot in a rotor. The dovetail has a body and a plurality of projections extend from the body in opposing directions. A tapered groove extends through the body from the first end to the second end, and has a tapered profile such that a first depth of the tapered groove near the first end is greater than a second depth of the tapered groove near the second end. The tapered profile gradually transitions from the first depth to the second depth. The tapered groove is open at a bottom surface of the body and is sized to engage a shim. The tapered groove includes a flat section near the first end or the second end, and the flat section has a constant depth.

A turbine blade assembly for a gas turbine engine is provided. The turbine blade assembly having: a rotor; a plurality of turbine blades mounted to the rotor, each of the plurality of turbine blades having a forward tab that contacts a front face of the rotor and an aft tab; a plurality of tabs that extend from an aft surface of the rotor; and a retaining ring that engages each aft tab of the plurality of turbine blades and each of the plurality of tabs that extend from the aft surface of the rotor.

A rotor disc of a gas turbine rotor assembly is provided. The rotor disc may support a plurality of blades attached thereto. The rotor disc may have a plurality of fixing members defined therein through its peripheral surface and circumferentially spaced apart from one another. The fixing members may extend axially from a front end portion to a rear end portion of the disc. Profiled slots may be defined between pairs of adjacent ones of the fixing members and may be configured to receive complementary profiled blade root portions. A plurality of disc lips may project axially forward from a surrounding surface of the disc in the front end portion. A disc lip may be disposed at a tip portion of a fixing member, adjacent a leading edge of the disc to minimize air leakage at a disc/blade interface. A rotor assembly with such rotor disc and a plurality of blades attached thereto is also provided.