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.

Abstract: A gas turbine engine (10) includes: a compressor system comprising a low pressure compressor (15) and a high pressure compressor (16) coupled to low pressure and high pressure shafts, respectively (23, 24); an inner core casing (34) provided radially inwardly of compressor blades (42), 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 (13) coupled to the low pressure shaft via a gearbox (14); wherein the outer core casing comprises a first outer core casing (48) and a second outer core casing (50) 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.

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.

Abstract: A gas turbine engine (10) includes: a compressor system comprising a low pressure compressor (15) and a high pressure compressor (16) coupled to low pressure and high pressure shafts, respectively (23, 24); an inner core casing (34) provided radially inwardly of compressor blades (42), 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 (13) coupled to the low pressure shaft via a gearbox (14); wherein the outer core casing comprises a first outer core casing (48) and a second outer core casing (50) 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.

Abstract: An aircraft gas turbine engine comprises a high pressure compressor coupled to a high pressure turbine by a high pressure shaft, and a low pressure compressor coupled to a low pressure turbine by a low pressure shaft. The engine includes a first shaft bearing configured to support a forward end of the low pressure shaft, the first shaft bearing being mounted to a static structure forward of the high pressure turbine. The engine also includes a second shaft bearing configured to support a rearward end of the low pressure shaft, the second shaft bearing being mounted to a static structure rearward of the high pressure turbine, and forward of the low pressure turbine. The engine further includes a third shaft bearing configured to support the low pressure shaft, the third shaft bearing being mounted to a static structure rearward of the low pressure turbine.

Abstract: A gas turbine engine (10) comprising: a low pressure turbine (19); a fan (13) drivable by the low pressure turbine. A high pressure turbine (18) and a high pressure compressor (16) coupled by a high pressure shaft (24). An epicyclic gearbox (14) in planetary configuration coupled between a low pressure shaft (23) and the fan. The engine having a bypass ratio greater than or equal to 13. Or the fan having a diameter greater than or equal to 85 inches and less than or equal to 170 inches. Or the engine arranged to generate thrust in the range 35,000 lbf to 130,000 lbf.