CPC F02C 7/36 (2013.01) [F02K 3/06 (2013.01); F05D 2200/14 (2013.01); F05D 2220/323 (2013.01); F05D 2220/36 (2013.01); F05D 2260/40311 (2013.01)] | 18 Claims |
1. A gas turbine engine for an aircraft comprising:
an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor;
a fan located upstream of the engine core, the fan comprising a plurality of fan blades, the fan having a fan axial centreline;
a gearbox that is configured to:
receive an input from the core shaft, and
output drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft; and
a fan shaft mounting structure arranged to mount the fan shaft within the engine, the fan shaft mounting structure comprising at least two supporting bearings connected to the fan shaft, wherein:
the fan comprises 22, 24 or 26 fan blades;
a system radial bending stiffness is defined as:
a system tilt stiffness is defined as:
and
either: (i) the system radial bending stiffness is greater than or equal to 3.90×106 N/m; or (ii) the system tilt stiffness is greater than or equal to 1.10×105 Nm/rad.
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