Gas turbine engine with improved bending resistance

A gas turbine engine for an aircraft comprising:

• an engine core comprising:
• a compressor system with compressor blades comprising respective aerofoils, the compressor system comprising a first, lower pressure, compressor, and a second, higher pressure, compressor;
• and
• an outer core casing surrounding the compressor system and comprising:
• a first flange connection arranged to allow separation of the outer core casing at an axial position of the first flange connection, the first flange connection having a first flange radius, wherein the first flange connection 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 and the mid-span axial location on the leading edge of the most upstream aerofoil of the second compressor;
• and
• a fan located upstream of the engine core, the fan comprising a plurality of fan blades and having a fan diameter;
• wherein a fan diameter ratio of:
• first
• 
• 
• flange
• 
• 
• radius
• fan
• 
• 
• diameter
• is equal to or greater than 0.125.