An aircraft engine comprising a fan, the fan having a diameter D and including a plurality of fan blades, the fan blades having a sweep metric S.sub.tip, each fan blade having a leading edge, and a forward-most portion on the leading edge of each fan blade being in a first reference plane. The aircraft engine further comprises a nacelle, comprising an intake portion forward of the fan, a forward edge on the intake portion being in a second reference plane, wherein the intake portion has a length L measured along an axis of the aircraft engine between the first reference plane and the second reference plane, the aircraft engine having a cruise design point condition M.sub.rel, wherein M.sub.rel is between 0.4 and 0.93, L/D is between 0.2 and 0.45 and S.sub.tip is from −1 to 0.1.

An aircraft engine comprising a fan, the fan having a diameter D and including a plurality of fan blades, the fan blades having a sweep metric S.sub.tip, each fan blade having a leading edge, and a forward-most portion on the leading edge of each fan blade being in a first reference plane. The aircraft engine further comprises a nacelle, comprising an intake portion forward of the fan, a forward edge on the intake portion being in a second reference plane, wherein the intake portion has a length L measured along an axis of the aircraft engine between the first reference plane and the second reference plane, the aircraft engine having a cruise design point condition M.sub.rel, wherein M.sub.rel is between 0.4 and 0.93, L/D is between 0.2 and 0.45 and S.sub.tip is from −1 to 0.1.

A nacelle for a gas turbine engine having a longitudinal centre line. The nacelle includes an air intake disposed at an upstream end of the nacelle. The air intake includes, in flow series, an intake lip, a throat and a diffuser. The nacelle further includes a protrusion extending radially inward from the air intake downstream of the intake lip. The protrusion extends circumferentially by a protrusion angle (θ.sub.p) with respect to the longitudinal centre line of the gas turbine engine.

A gas turbine engine has a compression system radius ratio defined as the ratio of the radius of the tip of a fan blade to the radius of the tip of the most downstream compressor blade in the range of from 5 to 9. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

A gas turbine engine has a compression system radius ratio defined as the ratio of the radius of the tip of a fan blade to the radius of the tip of the most downstream compressor blade in the range of from 5 to 9. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

A variable geometry inlet system of an aircraft engine includes an inlet duct. The inlet duct includes at least first and second sections moveable between extended and retracted positions such that the inlet duct defines a variable axial length of an inlet passage for selective flight conditions. The inclusion of acoustic treatment may assist in controlling noise.

A non-metallic tailcone (202) in a tip turbine engine includes a tapered wall structure disposed (208) about a central axis. The non-metallic tailcone is fastened to a structural frame (44) in the aft portion of the tip turbine engine. The tip turbine engine produces a first temperature gas stream from a first output source and a second temperature gas stream from a second output source. The second temperature gas stream is a lower temperature than the first temperature gas stream. The second temperature gas stream is discharged at an inner diameter of the tip turbine engine over an outer surface of the tailcone. Discharging the cooler second temperature gas stream at the inner diameter allows a non-metallic to be used to form the tailcone.

A non-metallic tailcone (202) in a tip turbine engine includes a tapered wall structure disposed (208) about a central axis. The non-metallic tailcone is fastened to a structural frame (44) in the aft portion of the tip turbine engine. The tip turbine engine produces a first temperature gas stream from a first output source and a second temperature gas stream from a second output source. The second temperature gas stream is a lower temperature than the first temperature gas stream. The second temperature gas stream is discharged at an inner diameter of the tip turbine engine over an outer surface of the tailcone. Discharging the cooler second temperature gas stream at the inner diameter allows a non-metallic to be used to form the tailcone.

A disclosed fan drive gear system includes a sun gear rotatable about an axis of rotation, a plurality of intermediate gears rotatable about an intermediate gear rotation axis in meshing engagement with the sun gear and a ring gear circumscribing the intermediate gears. A bearing assembly supports at least one of the plurality of intermediate gears and includes a first beam extending in a first direction and a second beam extending from an end of the first beam in a second direction. The bearing surface supported on the second beam such that first and second beams are configured to maintain the bearing surface substantially parallel to the intermediate gear rotation axis during operation.

A disclosed fan drive gear system includes a sun gear rotatable about an axis of rotation, a plurality of intermediate gears rotatable about an intermediate gear rotation axis in meshing engagement with the sun gear and a ring gear circumscribing the intermediate gears. A bearing assembly supports at least one of the plurality of intermediate gears and includes a first beam extending in a first direction and a second beam extending from an end of the first beam in a second direction. The bearing surface supported on the second beam such that first and second beams are configured to maintain the bearing surface substantially parallel to the intermediate gear rotation axis during operation.