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 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. The gas turbine engine comprises a rotatable shaft defining a rotational axis extending between rearward and forward ends of the gas turbine engine and a compressor drum surrounding, and coupled to, the shaft so as to define an annular gap therebetween. The gas turbine engine further comprises an intermediate case arranged axially rearward of the compressor drum and comprising a bearing support element extending into the annular gap, and a forward bearing mounted between the bearing support element and the shaft proximate a forward end of the compressor drum.

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: A gas turbine engine includes an engine core including: a compressor system including first, lower pressure compressor, and second, higher pressure compressor; and an outer core casing. The engine includes a front mount arranged for connection to a pylon; and a fan located upstream of the engine core. The outer core casing includes a first flange connection that: is arranged to allow separation of the outer core casing at an axial position thereof, and is the first flange connection downstream of an axial position defined by the axial midpoint between the mid-span axial location on trailing edge of the most downstream aerofoil of first compressor and mid-span axial location on leading edge of the most upstream aerofoil of the second compressor. A front mount position ratio of: axial ? ? distance ? ? between ? ? the ? ? first ? ? flange ? ? connection and ? ? the ? ? front ? ? mount first ? ? flange ? ? radius is equal to or less than 1.18.

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: A gas turbine engine includes an engine core including: a compressor system including first, lower pressure compressor, and second, higher pressure compressor; and inner and outer core casings that define a core working gas flow path (A) therebetween, which has an outer radius that defines a gas path radius. The outer core casing includes a first flange connection that: has a first flange radius, is arranged to allow separation of the outer core casing at an axial position thereof, and is the first flange connection that is downstream of an axial position defined by the axial midpoint between the mid-span axial location on trailing edge of the most downstream aerofoil of first compressor and mid-span axial location on leading edge of the most upstream aerofoil of the second compressor. A gas path ratio of: first ? ? flange ? ? radius gas ? ? path ? ? radius is equal to or greater than 1.10.

Abstract: A coupling arrangement for a gas turbine engine. The arrangement comprises first, second and third members. The first member has a first threaded mating surface extending in a first direction (X) and a flange extending in a direction generally normal to the first direction (X). The second member has a second threaded mating surface extending in the first direction (X) and a flange extending in a direction generally normal to the first direction (X), the flanges of the first and second members engaging against one another. The third member has a third threaded mating surface configured to engage against the first threaded mating surface, and a fourth threaded mating surface configured to engage against the second threaded mating surface.

Abstract: A gas turbine engine for an aircraft includes: an engine core with a compressor system including a first, lower pressure, compressor, and a second, higher pressure, compressor; an inner core casing provided radially inwardly of the compressor blades of the compressor system; and an outer core casing surrounding the compressor system, the inner core casing and the outer core casing defining a core working gas flow path therebetween. The outer core casing includes: 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 having a first flange radius, A gas path radius is defined as the outer radius of the core gas flow path at the axial position of the first flange connection, and a gas path ratio of: first ? ? flange ? ? radius gas ? ? path ? ? radius is equal to or greater than 1.10.

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 including: a compressor system including first, lower pressure compressor, and second, higher pressure compressor; and an outer core casing. The engine includes a front mount arranged for connection to a pylon; and a fan located upstream of the engine core. The outer core casing includes a first flange connection that: is arranged to allow separation of the outer core casing at an axial position thereof, and is the first flange connection downstream of an axial position defined by the axial midpoint between the mid-span axial location on trailing edge of the most downstream aerofoil of first compressor and mid-span axial location on leading edge of the most upstream aerofoil of the second compressor. A front mount position ratio of: axial ? ? distance ? ? between ? ? the ? ? first ? ? flange ? ? connection and ? ? the ? ? front ? ? mount first ? ? flange ? ? radius is equal to or less than 1.18.

Abstract: A gas turbine engine for an aircraft and an engine core including: a compressor system, a first lower pressure compressor, a second, higher pressure compressor; an outer core casing surrounding the compressor system. The outer core casing includes 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 having a first flange radius, wherein the first flange connection is the 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, and a front mount arranged to be connected to a pylon.

Abstract: A gas turbine engine for an aircraft includes: an engine core with: a compressor system including a first, lower pressure, compressor, and a second, higher pressure, compressor; and an outer core casing surrounding the compressor system. The gas turbine engine further includes a fan located upstream of the engine core with a plurality of fan blades and having a fan diameter. The outer core casing includes a first flange connection arranged to allow separation of the outer core casing at an axial position of the first flange connection which has a first flange radius, wherein the first flange connection 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.

Abstract: A gas turbine engine includes an engine core including: a compressor system including first, lower pressure compressor, and second, higher pressure compressor; and inner and outer core casings that define a core working gas flow path (A) therebetween, which has an outer radius that defines a gas path radius. The outer core casing includes a first flange connection that: has a first flange radius, is arranged to allow separation of the outer core casing at an axial position thereof, and is the first flange connection that is downstream of an axial position defined by the axial midpoint between the mid-span axial location on trailing edge of the most downstream aerofoil of first compressor and mid-span axial location on leading edge of the most upstream aerofoil of the second compressor. A gas path ratio of: first ? ? flange ? ? radius gas ? ? path ? ? radius is equal to or greater than 1.10.

Abstract: A gas turbine engine for an aircraft. The gas turbine engine comprises a rotatable shaft defining a rotational axis extending between rearward and forward ends of the gas turbine engine and a compressor drum surrounding, and coupled to, the shaft so as to define an annular gap therebetween. The gas turbine engine further comprises an intermediate case arranged axially rearward of the compressor drum and comprising a bearing support element extending into the annular gap, and a forward bearing mounted between the bearing support element and the shaft proximate a forward end of the compressor drum.

Abstract: There is disclosed an auxiliary rotation device 28 for a gas turbine engine 10. The auxiliary rotation device 28 comprises: an electric machine 30 configured to be coupled to a rotor of the gas turbine engine 10; and a dedicated electrical storage device 32 coupled to the electric machine 30. During a period of powered operation of the gas turbine engine 10 the electric machine 30 is configured to act as a generator so as to charge the dedicated electrical storage device 32. Following the period of powered operation of the gas turbine engine 10 the electric machine 30 is configured to act as a motor by discharging the dedicated electrical storage device 32 so as to cause the rotor to rotate for a period of time to provide even cooling of the rotor around its circumference. There is also disclosed a method of cooling a rotor of a gas turbine engine 10 using an auxiliary rotation device 28.

Abstract: A gas turbine engine (10) for an aircraft comprises an engine core (11) comprising: a compressor system comprising a first, lower pressure, compressor (14), and a second, higher pressure, compressor (15); and an outer core casing (70) surrounding the compressor system. The gas turbine engine further comprises a fan (23) located upstream of the engine core (11), the fan comprising a plurality of fan blades.

Abstract: A gas turbine engine includes an engine core; a fan located upstream of engine core and having a fan diameter; a nacelle surrounding engine core and defining a bypass duct between engine core and nacelle; and a fan OGV extending radially across bypass duct, fan OGV having a radially inner and outer edge, wherein the radially inner edge's axial midpoint is the fan OGV root centrepoint. The engine core includes a compressor system encompassing a 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 outer core casing at an axial position of first flange connection, the first flange connection having a first flange radius, wherein the first flange connection is downstream of an axial position.

Abstract: A thrust balancing mechanism for balancing axial loads on a rotor thrust bearing 3 is described. The mechanism comprises a piston arrangement 6 axially mounted on a stationary structure 2, about a center axis arranged, in use, in coaxial alignment with a rotating shaft 1 carrying the rotor thrust bearing 3. A hydrodynamic thrust bearing 8 is mounted, in use, between the piston 6 and the rotor thrust bearing 3. The piston 6 is pressurized so as to impart to the rotor thrust bearing 3, via the hydrodynamic thrust bearing 8, an axial load which counters an axial load imparted to the rotor thrust bearing 3 by the rotating shaft 1.