A nacelle for a gas turbine engine having a longitudinal centre line includes an intake lip disposed at an upstream end of the nacelle. The intake lip includes a crown and a keel. The crown includes a crown leading edge and the keel includes a keel leading edge. The crown leading edge and the keel leading edge define a scarf line therebetween. The scarf line forms a scarf angle (θ.sub.scarf) relative to a reference line perpendicular to the longitudinal centre line. A fan casing is disposed downstream of the intake lip and includes a casing leading edge. The casing leading edge defines a droop line normal to the casing leading edge. The droop line forms a droop angle (θ.sub.droop) relative to the longitudinal centre line. A relationship between the droop angle (θ.sub.droop) and the scarf angle (θ.sub.scarf) is given by: θ.sub.droop=θ.sub.scarf/1.5±1 degree.

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 turbofan gas turbine engine comprises, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, and a turbine module. The fan assembly comprises a plurality of fan blades defining a fan diameter (D). The heat exchanger module comprises a plurality of heat transfer elements. The heat exchanger module is in fluid communication with the fan assembly by an inlet duct. The inlet duct has a fluid path length along a central axis of the inlet duct between a downstream-most face of the heat transfer elements and an upstream-most face of the fan assembly. The fluid path length is less than 10.0*D.

A compact inlet design including a single bulkhead and/or an acoustic panel extending into nacelle lip region for noise reduction.

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 diffuser further includes a diffuser angle (θ.sub.diff), indicating a degree of divergence of the diffuser relative to the longitudinal centre line. The diffuser angle (θ.sub.diff) is from about 0 degrees to about 12 degrees.

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 coupled to an aircraft includes an engine core arranged axially along an axis, a bypass duct arranged circumferentially around the engine core to define a bypass channel, and a heat exchanger system. The bypass channel is arranged to conduct bypass air around the engine core to provide thrust for the gas turbine engine. The heat exchanger system is configured to provide cooling for the engine core.

An acoustic attenuation panel for a propulsion unit including a nacelle and a turbojet engine includes a cellular core disposed between an inner skin and an outer skin, called acoustic skin, the acoustic skin including a plurality of acoustic apertures, the acoustic apertures being inclined, at a non-zero inclination angle (β) relative to the direction normal to the acoustic skin, upstream with respect to the flow direction of the air or gas flow to which the panel is intended to be subjected under normal operating conditions, that is to say upstream of the propulsion unit when the panel is mounted in such a unit.

An anti-icing system is disclosed. In various embodiments, the anti-icing system includes an inner duct having a first end configured to deliver heated gas to a plenum and a second end spaced from the first end; an outer duct circumferentially encompassing at least a portion of the inner duct; and a seal system disposed proximate the second end, the seal system including a first annular seal having a radially inner end positioned proximate a flange disposed on the inner duct.