Turbine engine

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;
• an inner core casing provided radially inwardly of the compressor blades of the compressor system;
• and
• an outer core casing surrounding the compressor system, the inner core casing and the outer core casing defining a core working gas flow path therebetween, the outer core casing 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 gas path radius is defined as the outer radius of the core gas flow path at the axial position of the first flange connection, and a gas path ratio of:
• first
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• flange
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• radius
• gas
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• path
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• radius
• is equal to or greater than 1.10.