Efficient gas turbine engine installation and operation

A gas turbine engine for an aircraft comprising:

• an engine core comprising a turbine comprising a plurality of turbine blades, a compressor comprising a plurality of compressor blades, and a combustor;
• a fan comprising a plurality of fan blades;
• and
• a gearbox that receives an input from a core shaft that is connected to at least a part of the turbine, the gearbox configured to drive the fan at a lower rotational speed than a rotational speed of the core shaft, wherein:
• a bypass duct is defined radially outside the engine core, with a leading edge of a splitter defining a point at which flow splits into a core flow and a bypass flow during operation of the gas turbine engine;
• a core compressor aspect ratio is defined as a ratio of an axial distance between the leading edge of the splitter and a leading edge of a tip of a downstream-most compressor blade to a radius of the leading edge of the splitter, the radius of the leading edge of the splitter being measured from a centerline axis of the gas turbine engine, the core compressor aspect ratio being in a range of 1.7 to 3.4;
• a hub-to-tip ratio defined as a ratio of a radius of a fan blade of the plurality of fan blades at the hub to the radius of the fan blade at the tip is in a range of 0.27 to 0.35;
• and
• a bypass ratio defined as a ratio of a mass flow rate of the bypass flow through the bypass duct to the mass flow rate of the core flow through the engine core at cruise conditions is in a range of 13 to 15.