Abstract: A variable stator vane in a gas turbine engine has a journal provided with a circumferential recess. An aperture extends from the recess to a face of the journal directed towards the aerofoil portion of the vane. In operation of the engine, heated air, for example from a downstream compressor stage of the engine is admitted to the recess and flows through the aperture to, for example, the pressure surface of the aerofoil portion. The aerofoil portion is thereby heated, so preventing ice accretion on the vane.

Abstract: A variable stator vane in a gas turbine engine has a journal provided with a circumferential recess. An aperture extends from the recess to a face of the journal directed towards the aerofoil portion of the vane. In operation of the engine, heated air, for example from a downstream compressor stage of the engine is admitted to the recess and flows through the aperture to, for example, the pressure surface of the aerofoil portion. The aerofoil portion is thereby heated, so preventing ice accretion on the vane.

Abstract: An aircraft (10) comprises a wing (12) having an upper surface (16) and a lower surface (18) and at least one turbofan gas turbine engine (20) mounted on the wing (12). The axis (34) of the at least one turbofan gas turbine engine (20) is arranged substantially in the plane (14) of the wing (12) of the aircraft (10). The at least one turbofan gas turbine engine (20) has an intake, the intake comprising a hollow member (44) rotatably mounted coaxially with the at least one turbofan gas turbine engine (20) such that in use the hollow member (44) is rotatable between a first position, a low-speed condition, in which air flowing over and/or above the upper surface (16) of the wing (12) flows into the intake of the at least one turbofan gas turbine engine (20) and a second position, a high-speed condition, in which air flowing along and/or below the lower surface (18) of the wing flows into the intake of the at least one turbofan gas turbine engine.

Abstract: An aircraft engine (10 comprising a nacelle (24), a first wall (32) around a set of fan blades (28) and a set of fan outlet guide vanes (48). The set of fan outlet guide vanes (48) comprises a plurality of stator vanes extending between a second wall (34) and a third wall (32). The nacelle (24) defines an aerodynamic shape for a fluid and the nacelle (24) has apertures (84,86) at a region of relatively low pressure. A means to bleed fluid (40,44,54,58,54,53,60,72) from a region of relatively high pressure at one or more of the first wall (32) around the set of fan blades (28), the second wall (34), the third wall (32) or the set of stator vanes and at least one duct (46,82) to connect the means to bleed fluid (40,44,54,58,54,53,60,73) from the region of relatively high pressure to the at least one aperture (84,86) in the nacelle (24) at the region of relatively low pressure.

Abstract: The invention relates to a gearbox arrangement for driving a first multi-bladed rotor and a second coaxial multi-bladed rotor in opposite directions.The gearbox assembly comprises a first epicyclic gear arrangement and a coaxial second epicyclic gear arrangement. The first epicyclic gear arrangement is driven by a first input shaft, and the second epicyclic gear has a static second input shaft. The first epicyclic gear arrangement is drivingly connected to the first multi-bladed rotor and the second epicyclic gear arrangement is drivingly connected to the second multi-bladed rotor. The first epicyclic gear arrangement is interconnected to the second epicyclic gear arrangement so that the first and second multi-bladed rotors are driven in opposite directions.