Rolls-royce Corporation

James M Pointon

Stephen John Bradbrook

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 (D), and a root leading edge of the first compressor rotor blade defines a second diameter (D). A first ratio (D:D) of the first diameter (D) to the second diameter (D) is greater than or equal to 1.4.

Gas turbine engine with optimized fan, core passage inlet, and compressor forward stage diameter ratios

A gas turbine engine () includes: a compressor system comprising a low pressure compressor () and a high pressure compressor () coupled to low pressure and high pressure shafts, respectively (); an inner core casing () provided radially inwardly of compressor blades (), and an outer core casing provided outwardly of compressor blades, the inner core casing and outer core casing defining a core working gas flow path (B) therebetween; a fan () coupled to the low pressure shaft via a gearbox (); wherein the outer core casing comprises a first outer core casing () and a second outer core casing () spaced radially outwardly from the first outer core casing, and wherein at an axial plane (E) of an inlet to the high pressure compressor, the second outer core casing has an inner radius at least 1.4 times the inner radius of the first outer core casing.

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Bristol, United Kingdom

James M Pointon