Abstract: A gas turbine engine exhaust cone (40) comprising one or more acoustic absorption chambers (48), each chamber (48) comprising two side walls (50) and an acoustic attenuation wall (62). The side walls (50) define the circumferential extent of the chamber (48). The acoustic attenuation wall (62) has one or more acoustic attenuators (64) and is positioned circumferentially intermediate the side walls (50). The side walls (50) extend axially along the exhaust cone (40) at least one of them with a circumferential progression. The acoustic attenuation wall (62) extends axially along the exhaust cone (40) with or without a circumferential progression.

Abstract: A pylon for attachment of a gas turbine engine to a wing of an aircraft has a trailing edge which is rearward of the trailing edge of the core fairing and the trailing edge of the fan nacelle. The pylon has two laterally-spaced side faces which extend in the rearward direction of the engine to end at the trailing edge of the pylon. Each side face has a bottom edge which extends in a rearward direction of the engine from the core fairing to the bottom end of the trailing edge of the pylon. The bottom edges merge such that the bottom edges form a single bottom edge. The pylon is intended to exert control on the bypass flow of a gas turbine engine. Other pylons are also provided which can also exert such control.

Abstract: A gas turbine engine has an annular bypass duct defined between a fan nacelle and a core fairing through which fan air is discharged. One or more bifurcator members are positioned inside the exit of the bypass duct to bifurcate the air flow in the bypass duct around the bifurcator members prior to exiting the bypass duct. The bifurcator members modify the flow field of the air in the bypass duct radially and/or tangentially with respect to the axis of the annular bypass duct.

Abstract: A gas turbine engine exhaust cone (40) comprising one or more acoustic absorption chambers (48), each chamber (48) comprising two side walls (50) and an acoustic attenuation wall (62). The side walls (50) define the circumferential extent of the chamber (48). The acoustic attenuation wall (62) has one or more acoustic attenuators (64) and is positioned circumferentially intermediate the side walls (50). The side walls (50) extend axially along the exhaust cone (40) at least one of them with a circumferential progression. The acoustic attenuation wall (62) extends axially along the exhaust cone (40) with or without a circumferential progression.

Abstract: Noise generated by a gas turbine engine 6 supported by a pylon 8 on a wing 2 of an aircraft is reduced by influencing a shear layer generated between the free-stream flowfield and flow from the engine 6. The shear layer is influenced by means of one or more winglets 18, 20, 22 which interact with the free-stream flowfield to deflect the shear layer 14 downwardly, to avoid interaction with a flap 4 on the wing 2, or to reduce the strength of the shear layer 14.

Abstract: An aircraft wing has an engine attachment position at which a gas turbine engine is attached beneath the wing, in use the engine ejecting a propulsive gas jet with a jet shear layer being formed between the gas jet and the surrounding air. The aircraft wing further has one or more elongate, flow-modifying formations protruding from the underside of the wing. The length direction of the or each formation is along the fore and aft direction of the wing with the trailing edge of the formation being rearward of the trailing edge of the wing. The flow-modifying formations can be arranged such that they interact with the jet shear layer to reduce noise generated by the interaction of the jet shear layer and the wing. They can also be arranged to block, attenuate and/or to scatter the noise reflected by the wing.

Abstract: Noise generated by a gas turbine engine 6 supported by a pylon 8 on a wing 2 of an aircraft is reduced by influencing a shear layer generated between the free-stream flowfield and flow from the engine 6. The shear layer is influenced by means of one or more winglets 18, 20, 22 which interact with the free-stream flowfield to deflect the shear layer 14 downwardly, to avoid interaction with a flap 4 on the wing 2, or to reduce the strength of the shear layer 14.

Abstract: A pylon is provided for attachment of a gas turbine engine to a wing of an aircraft. The gas turbine engine comprises a core fairing surrounding a core generator and defining a nozzle for discharging a core gas flow, and a fan nacelle surrounding the core fairing to define an annular bypass duct therebetween for discharging fan air. The pylon has a trailing edge which is rearward of the trailing edge of the core fairing and the trailing edge of the fan nacelle. The pylon further has having two laterally-spaced side faces which span the bypass duct between the fan nacelle and the core fairing and extend in the rearward direction of the engine to end at the trailing edge of the pylon. Each side face has a bottom edge which extends in a rearward direction of the engine from the core fairing to the bottom end of the trailing edge of the pylon.

Abstract: A gas turbine engine has an annular bypass duct defined between a fan nacelle and a core fairing through which fan air is discharged. One or more bifurcator members are positioned inside the exit of the bypass duct to bifurcate the air flow in the bypass duct around the bifurcator members prior to exiting the bypass duct. The bifurcator members modify the flow field of the air in the bypass duct radially and/or tangentially with respect to the axis of the annular bypass duct.