Abstract: A gas turbine engine for an aircraft includes an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor, a fan located upstream of the engine core, and a gearbox. The turbine is the lowest pressure turbine of the engine, and has a turbine length that is the distance between the root of the most upstream blade of the turbine at its leading edge and the root of the most downstream blade of the lowest pressure turbine at its trailing edge. The engine core further includes three bearings arranged to support the core shaft. The three bearings include a forward bearing and two rearward bearings, with a minor span being defined as the axial distance between the two rearward bearings. A ratio of the minor span to a turbine length is equal to or less than 1.05.

Abstract: A gas turbine engine for an aircraft includes an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor, and a fan including a plurality of fan blades located upstream of the engine core. The fan has a fan diameter in the range from 240 cm to 280 cm. The turbine is the lowest pressure turbine of the engine and the compressor is the lowest pressure compressor of the engine. The turbine includes a total of three sets of turbine blades. The engine core further includes three bearings arranged to support the core shaft. The three bearings include a forward bearing and two rearward bearings located downstream of a leading edge of a lowest pressure turbine blade of the turbine at a root of the blade.

Abstract: An engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor, the turbine being the lowest pressure turbine of the engine and the compressor being the lowest pressure compressor of the engine; a fan located upstream of the engine core; and a gearbox that receives an input from the core shaft and outputs drive to the fan. The engine core further includes three bearings arranged to support the core shaft, the three bearings including a forward bearing and two rearward bearings, and wherein the forwardmost rearward bearing has a bearing stiffness in the range of 30 kN/mm to 100 kN/mm, the bearing stiffness being defined by the radial displacement caused by the application of a radial force at the axial centrepoint of the bearing.

Abstract: A gas turbine engine for an aircraft includes an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor, and a fan including a plurality of fan blades located upstream of the engine core. The fan has a fan diameter in the range from 240 cm to 280 cm. The turbine is the lowest pressure turbine of the engine and the compressor is the lowest pressure compressor of the engine. The turbine includes a total of three sets of turbine blades. The engine core further includes three bearings arranged to support the core shaft. The three bearings include a forward bearing and two rearward bearings located downstream of a leading edge of a lowest pressure turbine blade of the turbine at a root of the blade.

Abstract: An aircraft gas turbine engine has an engine core with a turbine, compressor, and core shaft connecting the turbine to the compressor, a fan upstream of the engine core; and a gearbox. The engine core has three bearings, one forward, two rearward, to support the core shaft, a minor span being the axial distance between the two rearward bearings. A first blade to bearing ratio of the minor span divided by the product of the mass, radius at mid-height, and the square of the angular velocity at cruise for a blade of the lowest pressure set may have a value in the range from 2.0×10?6 to 7.5×10?6 kg?1·rad?2·s2. A second blade to bearing ratio of the minor span divided by the product of mass and radius at mid-height for a blade of the lowest pressure set may have a value in the range from 0.8 to 6.0 kg?1.

Abstract: A gas turbine engine includes an engine core, a fan located upstream of the engine core, a nacelle surrounding the engine core and defining a bypass duct, and a fan outlet guide vane (OGV) extending radially across the bypass duct between an outer surface of the engine core and an inner surface of the nacelle. The engine core includes a compressor system, and an outer core casing surrounding the compressor system and including a first flange connection arranged to allow separation of the outer core casing at an axial position of the first flange connection. An axial midpoint of a radially inner edge of the fan OGV is defined as the fan OGV root centrepoint. A fan OGV root position to fan diameter ratio of: an ? ? axial ? ? distance ? ? between ? ? the ? ? first ? ? flange ? ? connection ? ? and the ? ? fan ? ? OGV ? ? root ? ? centrepoint the ? ? fan ? ? diameter is equal to or less than 0.33.

Abstract: A gas turbine engine for an aircraft includes an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; and a fan located upstream of the engine core. The engine core further includes three bearings arranged to support the core shaft including a forward bearing and two rearward bearings, with a minor span defined as the distance between the two rearward bearings. A minor span to turbine length ratio is equal to or less than 1.05.

Abstract: A gas turbine engine includes an engine core, a fan located upstream of the engine core, a nacelle surrounding the engine core and defining a bypass duct, and a fan outlet guide vane (OGV) extending radially across the bypass duct between an outer surface of the engine core and an inner surface of the nacelle. The engine core includes a compressor system, and an outer core casing surrounding the compressor system and including a first flange connection arranged to allow separation of the outer core casing at an axial position of the first flange connection. An axial midpoint of a radially inner edge of the fan OGV is defined as the fan OGV root centrepoint. A fan OGV root position to fan diameter ratio of: an ? ? axial ? ? distance ? ? between ? ? the ? ? first ? ? flange ? ? connection ? ? and the ? ? fan ? ? OGV ? ? root ? ? centrepoint the ? ? fan ? ? diameter is equal to or less than 0.33.

Abstract: An engine core including a turbine, compressor, and core shaft connecting the turbine and compressor, the turbine being the lowest pressure turbine, the core shaft having a running speed range from 1500-6200 rpm, and the compressor being the lowest pressure compressor; a fan located upstream of the engine core; and a gearbox receiving an input from the core shaft and outputs drive to the fan. The engine core further includes three bearings arranged to support the core shaft, the three bearings including a forward bearing and two rearward bearings, the core shaft having a length between the forward and the rearmost rearward bearing ranging from 1800-2900 mm, and a minor span between two rearward bearings ranging from 250-350 mm, wherein there is no primary resonance of the core shaft between the forward and forwardmost rearward bearing within the running speed range of the core shaft.

Abstract: A gas turbine engine for an aircraft has an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor, the turbine being the lowest pressure turbine of the engine and the compressor being the lowest pressure compressor of the engine; a fan located upstream of the engine core; and a gearbox that receives an input from the core shaft and outputs drive to the fan. The engine core further has three bearings arranged to support the core shaft, and two rearward bearings, and wherein the forward most rearward bearing has a bearing stiffness defined by the radial displacement caused by the application of a radial force at the axial centerpoint of the bearing, and wherein a stiffness ratio of the bearing stiffness at the forward most rearward bearing to the minor span is in the range from 0.08 to 0.5 kN/mm2.

Abstract: A gas turbine engine for an aircraft has an engine core comprising turbine, compressor, and core shaft connecting the turbine to the compressor, the turbine being the lowest pressure turbine of the engine, and having turbine blades, and the compressor being the lowest pressure compressor of the engine; fan located upstream of the engine core; and gearbox that receives an input from the core shaft and outputs drive to the fan. The engine core further has three bearings arranged to support the core shaft, the three bearings having two rearward bearings located downstream of the leading edge of the lowest pressure turbine blade of the turbine at the root of the blade, and/or, when the turbine comprises four sets of turbine blades, downstream of the trailing edge of a turbine blade of the third set of turbine blades from the front of the turbine, at the root of the blade.

Abstract: An aircraft gas turbine engine has an engine core having a turbine, compressor, and core shaft connecting the turbine to the compressor, the turbine being the lowest pressure turbine and having a turbine length being the distance between the roots of the most upstream turbine blade at its leading edge and of the most downstream turbine blade trailing edge, and the compressor being the lowest pressure compressor of the engine; a fan located upstream of the engine; and a gearbox receiving an input from the core shaft and outputting drive to the fan. The engine core has three bearings to support the core shaft, the three bearings having a forward bearing and two rearward bearings, with a minor span defined as the distance between the two rearward bearings, and wherein further a minor span to turbine length ratio of: minor ? ? span tu ? ? rbine ? ? length is equal to or less than 1.05.

Abstract: A gas turbine engine for an aircraft has an engine core comprising turbine, compressor, and core shaft connecting the turbine to the compressor, the turbine being the lowest pressure turbine of the engine, and having turbine blades, and the compressor being the lowest pressure compressor of the engine; fan located upstream of the engine core; and gearbox that receives an input from the core shaft and outputs drive to the fan. The engine core further has three bearings arranged to support the core shaft, the three bearings having two rearward bearings located downstream of the leading edge of the lowest pressure turbine blade of the turbine at the root of the blade, and/or, when the turbine comprises four sets of turbine blades, downstream of the trailing edge of a turbine blade of the third set of turbine blades from the front of the turbine, at the root of the blade.

Abstract: A gas turbine engine includes an engine core including: a compressor system including first, lower pressure, compressor, and a second, higher pressure, compressor; and an outer core casing surrounding the compressor system and including a first flange connection arranged to allow separation of the outer core casing at an axial position of the first flange connection, 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; a nacelle surrounding the engine core and defining a bypass duct between the engine core and the nacelle; wherein an axial midpoint of the radially outer edge is defined as the fan OGV tip centrepoint.

Abstract: A gas turbine engine includes an engine core and a fan located upstream of the engine core. The engine core includes: a compressor system including first, lower pressure compressor, and second, higher pressure compressor; and an outer core casing surrounding the compressor system and including 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 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. A fan diameter ratio of: first ? ? flange ? ? radius fan ? ? diameter is equal to or greater than 0.125.

Abstract: An engine core including a turbine, compressor, and core shaft connecting the turbine and compressor, the turbine being the lowest pressure turbine, the core shaft having a running speed range from 1500-6200 rpm, and the compressor being the lowest pressure compressor; a fan located upstream of the engine core; and a gearbox receiving an input from the core shaft and outputs drive to the fan. The engine core further includes three bearings arranged to support the core shaft, the three bearings including a forward bearing and two rearward bearings, the core shaft having a length between the forward and the rearmost rearward bearing ranging from 1800-2900 mm, and a minor span between two rearward bearings ranging from 250-350 mm, wherein there is no primary resonance of the core shaft between the forward and forwardmost rearward bearing within the running speed range of the core shaft.

Abstract: A gas turbine engine for an aircraft has an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor, the turbine being the lowest pressure turbine of the engine and the compressor being the lowest pressure compressor of the engine; a fan located upstream of the engine core; and a gearbox that receives an input from the core shaft and outputs drive to the fan. The engine core further has three bearings arranged to support the core shaft, and two rearward bearings, and wherein the forward most rearward bearing has a bearing stiffness defined by the radial displacement caused by the application of a radial force at the axial centerpoint of the bearing, and wherein a stiffness ratio of the bearing stiffness at the forward most rearward bearing to the minor span is in the range from 0.08 to 0.5 kN/mm2.

Abstract: An aircraft gas turbine engine has an engine core having a turbine, compressor, and core shaft connecting the turbine to the compressor, the turbine being the lowest pressure turbine and having a turbine length being the distance between the roots of the most upstream turbine blade at its leading edge and of the most downstream turbine blade trailing edge, and the compressor being the lowest pressure compressor of the engine; a fan located upstream of the engine; and a gearbox receiving an input from the core shaft and outputting drive to the fan. The engine core has three bearings to support the core shaft, the three bearings having a forward bearing and two rearward bearings, with a minor span defined as the distance between the two rearward bearings, and wherein further a minor span to turbine length ratio of: minor ? ? span tu ? ? rbine ? ? length is equal to or less than 1.05.

Abstract: An aircraft gas turbine engine has an engine core with a turbine, compressor, and core shaft connecting the turbine to the compressor, a fan upstream of the engine core; and a gearbox. The engine core has three bearings, one forward, two rearward, to support the core shaft, a minor span being the axial distance between the two rearward bearings. A first blade to bearing ratio of the minor span divided by the product of the mass, radius at mid-height, and the square of the angular velocity at cruise for a blade of the lowest pressure set may have a value in the range from 2.0×10?6 to 7.5×10?6 kg?1.rad?2.s2. A second blade to bearing ratio of the minor span divided by the product of mass and radius at mid-height for a blade of the lowest pressure set may have a value in the range from 0.8 to 6.0 kg?1.

Abstract: An engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor, the turbine being the lowest pressure turbine of the engine and the compressor being the lowest pressure compressor of the engine; a fan located upstream of the engine core; and a gearbox that receives an input from the core shaft and outputs drive to the fan. The engine core further includes three bearings arranged to support the core shaft, the three bearings including a forward bearing and two rearward bearings, and wherein the forwardmost rearward bearing has a bearing stiffness in the range of 30 kN/mm to 100 kN/mm, the bearing stiffness being defined by the radial displacement caused by the application of a radial force at the axial centrepoint of the bearing.