Patent ID: 10677169

Abstract:
The present disclosure relates to a geared turbofan engine. Example embodiments include a gas turbine engine (10) for an aircraft, comprising: an engine core (11) comprising a turbine (19), a compressor (14), and a core shaft (26) connecting the turbine (19) to the compressor (14); a fan assembly (23) located upstream of the engine core (11); and a gearbox (30) that receives an input from the core shaft (26) and outputs drive to the fan assembly (23) so as to drive the fan assembly (23) at a lower rotational speed than the core shaft (26), wherein the fan assembly (23) comprises a plurality of fan blades (41) mounted around a hub (42), the fan blades (41) having blade tips defining an outer diameter of the fan assembly (23) of from around 220 cm to around 400 cm, the hub (42) comprising a plurality of slots (51) located around a rim (52) of the hub (42), each slot (51) receiving a root of a corresponding fan blade (41), the rim (52) having a minimum radial thickness between a base of each slot (51) and an internal cavity (54) within the hub (42), wherein the minimum radial thickness is within a range of around 0.5% to around 1.1% of the outer fan diameter.