Abstract: A gas turbine engine for an aircraft, includes: an engine core having a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan assembly located upstream of the engine core; and a gearbox receiving an input from the core shaft and outputs drive to the fan assembly so as to drive the fan assembly at a lower rotational speed than the core shaft, wherein the fan assembly has fan blades mounted around a hub, the fan blades having blade tips defining an outer diameter of the fan assembly of from around 220 cm to around 400 cm, the hub having slots located around a rim of the hub, each slot receiving a root of a corresponding fan blade, wherein a ratio of a mass of the hub to a total mass of the fan blades is within the range of around 0.45 to around 0.7.

Abstract: The present disclosure relates to a geared turbofan engine.

Abstract: A geared turbofan engine include a gas turbine engine for an aircraft, including: an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan assembly located upstream of the engine core; and a gearbox that receives an input from the core shaft and outputs drive to the fan assembly to drive the fan assembly at a lower rotational speed than the core shaft, the hub including a plurality of slots located around a rim of the hub, each slot receiving a root of a corresponding fan blade, the rim having a minimum radial thickness between a base of each slot and an internal cavity within the hub, wherein the minimum radial thickness is within a range of around 0.5% to around 1.1% of the outer fan diameter.

Abstract: A gas turbine engine for an aircraft, includes: an engine core having a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan assembly located upstream of the engine core; and a gearbox receiving an input from the core shaft and outputs drive to the fan assembly so as to drive the fan assembly at a lower rotational speed than the core shaft, wherein the fan assembly has fan blades mounted around a hub, the fan blades having blade tips defining an outer diameter of the fan assembly of from around 220 cm to around 400 cm, the hub having slots located around a rim of the hub, each slot receiving a root of a corresponding fan blade, wherein a ratio of a mass of the hub to a total mass of the fan blades is within the range of around 0.45 to around 0.7.

Abstract: A fan rotor module for a gas turbine engine, the fan rotor module comprising a drive arm, a fan disc and a windage shield attached at one or more points to a rear portion of the fan disc, wherein the fan rotor module comprises at least one balancing weight disposed at one or more of the points where the windage shield is attached to the rear portion of the fan disc.

Abstract: A windage shield system comprises first and second hollow cylindrical elements, the longitudinal axes of the outer and inner surfaces of each element being parallel and mutually displaced. The elements are mounted to the downstream end of a fan disc such that the longitudinal axes of the inner surface of the first element and the outer surface of the second element coincide with the rotational axis of the disc. The centres of mass of the elements are mountable with azimuthal offsets ?1, ?2 with respect to the fan disc, each being selectable from a large number of values in the range 0° to 360°. The first element is integral with a windage shield. The system allows a fan disc to be provided with a windage shield and the resulting assembly to be balanced at its rear plane in cases where access to the rear of the fan disc is difficult.

Abstract: A windage shield (200?, 200?) for mounting on a fan disc (100?, 100?) of a fan (23) of a gas turbine engine (10), the windage shield (200?, 200?) comprising: a fan disc contacting portion (215?, 215?) adapted to contact and structurally support a rear portion of the fan disc (100?, 100?); wherein the fan disc contacting portion (215?, 215?) includes one or more stiffening elements (205?, 205?) to locally increase the hoop stiffness of the windage shield (200?, 200?).

Abstract: The present disclosure concerns a fan disc assembly, and in particular a fan disc assembly for a gas turbine engine. Embodiments disclosed include a fan disc assembly for a gas turbine engine, comprising: a fan disc with a central bore comprising a bore forward section, a bore aft section and a bore spline between the bore forward and aft sections; a shaft mounted within the central bore of the fan disc, the shaft comprising a shaft forward section connected to the bore forward section, a shaft aft section connected to the bore aft section and a shaft spline between the forward and aft sections and mating with the bore spline; and an aft collar surrounding the shaft aft section and connecting the shaft aft section to the bore aft section to secure the fan disc from radial translation relative to the shaft.

Abstract: The present disclosure concerns a fan disc assembly, and in particular a fan disc assembly for a gas turbine engine. Example embodiments disclosed include a fan disc assembly for a gas turbine engine, comprising: a fan disc with a central bore comprising a bore forward section, a bore aft section and a bore spline between the bore forward and aft sections; a shaft mounted within the central bore of the fan disc, the shaft comprising a shaft forward section connected to the bore forward section, a shaft aft section connected to the bore aft section and a shaft spline between the forward and aft sections and mating with the bore spline; wherein the bore aft section comprises a tapered surface mating against a corresponding tapered surface on the shaft aft section.

Abstract: The present disclosure concerns a fan disc assembly, and in particular a fan disc assembly for a gas turbine engine. Example embodiments disclosed include a fan disc assembly for a gas turbine engine, comprising: a fan disc with a central bore comprising a bore forward section, a bore aft section and a bore spline between the bore forward and aft sections; a shaft mounted within the central bore of the fan disc, the shaft comprising a shaft forward section connected to the bore forward section, a shaft aft section connected to the bore aft section and a shaft spline between the forward and aft sections and mating with the bore spline, wherein an outer radial surface of the bore forward section is secured to an inner radial surface of the shaft forward section with a second interference fit.