Patent ID: 10598022

Abstract:
A gas turbine engine (10) for an aircraft comprises an engine core (11) comprising: a compressor system comprising a first, lower pressure, compressor (14), and a second, higher pressure, compressor (15); and an outer core casing (70) surrounding the compressor system. The gas turbine engine further comprises a fan (23) located upstream of the engine core (11), the fan comprising a plurality of fan blades. The outer core casing comprises: a first flange connection (60) arranged to allow separation of the outer core casing (70) at an axial position of the first flange connection (60), the first flange connection (60) having a first flange radius (104), wherein the first flange connection (60) is the first flange connection that is downstream of an axial position defined by the axial midpoint between the mid-span axial location on the trailing edge of the most downstream aerofoil of the first compressor (14) and the mid-span axial location on the leading edge of the most upstream aerofoil of the second compressor (15). A fan blade mass ratio of: is equal to or less than 19.0 mm/lb.