Abstract: A gas turbine engine for an aircraft is provided. The gas turbine engine comprises an engine core comprising a compressor, a combustor, a turbine, and a core shaft connecting the turbine to the compressor. The gas turbine engine further comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades, the fan generating a core airflow which enters the engine core and a bypass airflow which flows through a bypass duct surrounding the engine core. The gas turbine engine further comprises a circumferential row of outer guide vanes located in the bypass duct rearwards of the fan, the outer guide vanes extending radially outwardly from an inner ring which defines a radially inner surface of the bypass duct.

Abstract: An oil tank filling system for filling the oil tank of a gas turbine engine that includes an aft core cowl. The system includes an oil tank that has an oil tank top and an oil tank bottom and is located within a core of the engine, an oil access port located on the aft core cowl, and an oil tank filling pipe that leads from the oil access port to a tank filling port located at or adjacent the oil tank bottom. The system is configured so that oil that is supplied to the oil access port flows to and into the oil tank using gravitational force. A method of filling the oil tank of a gas turbine engine and a gas turbine engine that includes the oil tank filling system.

Abstract: The disclosure relates to a gas turbine engine having an accessory gearbox assembly having an accessory gearbox. The accessory gearbox assembly is mounted adjacent a core of the gas turbine engine and has a support formation for supporting a conduit. The conduit extends between a first location on the engine and a second location on the engine. The first location and the second location may be spaced from the accessory gearbox such that the conduit does not operatively communicate with the accessory gearbox in use. The accessory gearbox may be axially mounted and may provide a bridge for supporting a plurality of conduits.

Abstract: An oil tank filling system for filling the oil tank of a gas turbine engine that includes an aft core cowl. The system includes: an oil tank that is located within a core of the engine; an oil access port located on the aft core cowl; an oil tank filling pipe that connects the oil access port to a tank filling port on the oil tank; and a hydraulic lock that prevents oil entering the oil tank once the oil tank contains a predetermined volume of oil. A method of filling the oil tank of a gas turbine engine and a gas turbine engine that includes the oil tank filling system are also disclosed.

Abstract: The disclosure relates to a gas turbine engine (10) having an accessory gearbox assembly having an accessory gearbox (42). The accessory gearbox assembly is mounted adjacent a core (11) of the gas turbine engine (10) and has a support formation for supporting a conduit (50). The conduit (50) extends between a first location (52) on the engine (10) and a second location (54) on the engine (10). The first location (52) and the second location (54) may be spaced from the accessory gearbox (42) such that the conduit (50) does not operatively communicate with the accessory gearbox (42) in use. The accessory gearbox (42) may be axially mounted and may provide a bridge for supporting a plurality of conduits.

Abstract: A turbine-tip clearance control system offtake for a gas turbine engine, comprising: an air input duct in fluid communication with a compressor bleed air outlet, the air input duct comprising an exit aperture, the exit aperture defining an exit flow path for at least a portion of the air flowing in the air input duct during use of the gas turbine engine. The direction of airflow through the exit aperture, during use of the gas turbine engine, has at least a component that is anti-parallel to the direction of airflow in the air input duct in a region of the air input duct adjacent to the exit aperture. A turbine-tip clearance control system, a gas turbine engine for an aircraft and a method of supplying airflow to a turbine-tip clearance control system in a gas turbine engine are also disclosed.

Abstract: An oil tank filling system for filling the oil tank of a gas turbine engine that includes an aft core cowl. The system includes an oil tank that has an oil tank top and an oil tank bottom and is located within a core of the engine, an oil access port located on the aft core cowl; and an oil tank filling pipe that leads from the oil access port to a tank filling port located at or adjacent the oil tank bottom. The system is configured so that oil that is supplied to the oil access port flows to and into the oil tank using gravitational force. A method of filling the oil tank of a gas turbine engine and a gas turbine engine that includes the oil tank filling system are also disclosed.

Abstract: An oil tank filling system for filling the oil tank of a gas turbine engine that includes an aft core cowl. The system includes: an oil tank that is located within a core of the engine; an oil access port located on the aft core cowl; an oil tank filling pipe that connects the oil access port to a tank filling port on the oil tank; and a hydraulic lock that prevents oil entering the oil tank once the oil tank contains a predetermined volume of oil. A method of filling the oil tank of a gas turbine engine and a gas turbine engine that includes the oil tank filling system are also disclosed.

Abstract: A gas turbine engine for an aircraft is provided. The gas turbine engine comprises an engine core comprising a compressor, a combustor, a turbine, and a core shaft connecting the turbine to the compressor. The gas turbine engine further comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades, the fan generating a core airflow which enters the engine core and a bypass airflow which flows through a bypass duct surrounding the engine core. The gas turbine engine further comprises a circumferential row of outer guide vanes located in the bypass duct rearwards of the fan, the outer guide vanes extending radially outwardly from an inner ring which defines a radially inner surface of the bypass duct.

Abstract: A gas turbine engine (10) comprising: an engine core (11), comprising a compressor (14, 15); an outer casing (25A) separating the engine core (11) from a bypass airflow; a compressor bleed valve (50) in communication with the compressor (14, 15) and configured to release bleed air from the compressor (14, 15); a bleed air duct (51) connected to the compressor bleed valve (50) and configured to eject the bleed air released by the compressor bleed valve (50) into an airflow at a location radially inward of the outer casing (25A).

Abstract: A bleed valve arrangement for a gas turbine engine comprises a bleed valve housing defining a bleed duct through which a bleed flow may pass in use and a valve within the bleed duct operable between a closed position and an open position. The valve position is continuously variable between the closed position and the open position. The arrangement further comprises an ejection duct downstream of the bleed duct. The arrangement is characterised in that the ejection duct has a greater cross-sectional area than the bleed duct immediately downstream of the bleed duct so that the bleed flow experiences a sudden expansion on passing from the bleed duct into the ejection duct.

Abstract: A cowl cavity is defined between a core engine cowl and an engine inner casing. A core fairing is provided at the forward end of the engine inner casing and has a radially outer surface which, with the cowl, provides an airwashed surface of a bypass duct. A plenum is provided within the cowl cavity and receives air from bypass flow through openings in the surface. Conduits extend from a partition defining the plenum to distribute air to locations and components within the engine.