Patent ID: 11047339

Abstract:
An aircraft gas turbine engine comprises a fan coupled to a fan drive turbine, the fan being configured to provide a bypass flow (B) and a core flow (A) in use. The engine includes a reduction gearbox which couples the fan to the fan drive turbine and a core compressor arrangement. The core compressor arrangement has a core inlet at an upstream end of a core gas flow passage (A) defined by radially inner and outer walls, and at least a first compressor rotor blade provided at an upstream end of the compressor arrangement. The radially inner wall of the core inlet defines a first diameter (DINLET), and a root leading edge of the first compressor rotor blade defines a second diameter (DCOMP). A first ratio (DINLET:DCOMP) of the first diameter (DCOMP) to the second diameter (DCOMP) is greater than or equal to 1.4.