Rolls-royce Corporation

Adam MacGregor Bagnall

Christopher Freeman

Clive Lockwood

Richard Geoffrey Stretton

Peter Beecroft

Angus Roy Smith

Gareth L Jones

Stephane Michel Marcel Baralon

Philip Geoffrey Woodrow

Kevin P Vyze

Ross D Maxwell

A gas turbine engine is disclosed including a bifurcation fairing located in a bypass duct of the gas turbine engine and traversing the radial extent of the bypass duct. The bifurcation fairing has a delivery conduit inlet leading to a delivery conduit extending inside the bifurcation fairing, the delivery conduit being arranged in use for delivery of bypass air to one or more components of the gas turbine engine. A diverter conduit has a diverter conduit inlet from the delivery conduit upstream of the delivery conduit reaching the one or more components of the gas turbine engine. The diverter conduit has an outlet to a location other than the one or more components of the gas turbine engine.

Fluid system

A ducted fan gas turbine engine has a propulsive fan, a fan case surrounding the fan, a core engine, and a plurality of support structures which transmit loads from the core engine to the fan case. Each support structure has two support struts which extend from the core engine to a joint radially outwardly of the fan case. The support struts are spaced apart at the core engine but converge to meet at the joint. Each support structure further has two structural elements which extend from the joint to respective fixing positions on the fan case at opposite sides of the joint.

Gas turbines engine support structures

A fan casing for fitment around an array of radially extending fan blades of a gas turbine engine. The fan casing includes an annular casing element; and an annular fan track liner positioned radially inward of the annular casing element. The fan track liner includes an abradable layer and an abrasive layer, the abrasive layer being positioned radially inward of the abradable layer and proximal, in use, to the fan blades.

Method of manufacturing a gas turbine engine having a fan track liner with an abradable layer

A method of reducing the fluid flow effects of one or more flow modifiers () in a passage (), the method comprising: providing a plurality of aerofoil structures () in the passage (), wherein the geometry of each aerofoil structure () is initially substantially the same; and shortening the trailing edge () of one or more selected aerofoil structures () in a chordwise direction and over at least a spanwise portion of the one or more selected aerofoil structures such that, when in use, the direction of the fluid flow () in the vicinity of the one or more selected aerofoil structures is altered and the fluid flow effects of the one or more flow modifiers () in the passage are reduced.

Method of reducing asymmetric fluid flow effects in a passage

A seal assembly arranged adjacent to a primary flow region, the seal assembly including: first and second components; a seal arranged between the first and second components to seal a secondary flow region from the primary flow region; and a first recess portion provided in the first component and arranged adjacent to the seal and the primary flow region, the first recess portion further being arranged to receive flow from at least one of the primary flow region and the secondary flow region and shed flow to the primary flow region, wherein the first recess portion is configured to promote a first flow feature within the first recess portion, the first flow feature flowing with a portion of the first flow feature adjacent to the primary flow region and the portion of the first flow feature being shed to the primary flow region in substantially the same direction as the flow in the primary flow region.

Seal assembly

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.

Variable stator vane

A nose cone assembly for a gas turbine engine comprises a spinner and a plurality of circumferentially spaced mounting members for mounting the spinner on a fan hub of the gas turbine engine. Each of the mounting members is rigid in a generally axial direction and flexible in a generally radial direction.

Gas turbine engine having a nose cone assembly rigidly attached in an axial direction but flexible in a radial direction

An apparatus for controlling the temperature of a component, which is situated in use in a gas stream, provides a nozzle to create a jet of air at an angle to the gas stream, the jet being directed into the region of the stagnation point of the component so as to control the temperature of the component. The invention is particularly suited to preventing or reducing the formation of ice on vanes of gas turbine engines, but may also be applied to other components, and may equally be used in situations where a component is to be cooled rather than heated.

Temperature controlling apparatus

An aircraft engine comprising a nacelle, a first wall and fan outlet guide vanes. The fan outlet guide vanes comprises a plurality of stator vanes extending between a second wall and a third wall. The nacelle having apertures at a region of low pressure. A means to bleed fluid from a region of high pressure at one or more of the first wall, the second wall, the third wall or the stator vanes and at least one duct to connect the means to bleed fluid from the region of high pressure to the at least one aperture in the nacelle. The static pressure at the high pressure region is greater than that in the low pressure region such that at least some of the boundary layer of the fluid flows through the at least one duct to the at least one aperture in the nacelle.

Arrangement for bleeding the boundary layer from an aircraft engine

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

Variable position intake for an aircraft mounted gas turbine engine

A fan ( ) for a turbofan gas turbine engine ( ) comprises a fan rotor ( ) carrying a first set of circumferentially spaced radially extending fan blades ( ) and a second set of circumferentially spaced radially extending fan blades ( ). The second set of fan blades ( ) is arranged downstream of the first set of fan blades ( ). The hub to tip ratio (R /R ) of the first set of fan blades ( ) is substantially the same as the hub to tip ratio (R /R ) of the second set of fan blades ( ) and the radius (R ) of the radially outer ends of the first set of fan blades ( ) is less than the radius (R ) of the radially outer ends of the second set of fan blades ( ). This increases the flow area of the fan ( ) by about 7% compared to a conventional fan of the same radius and thus increases the mass flow by about 7% and/or increases the pressure ratio.

Fan for a turbofan gas turbine engine

A nose cone assembly ( ) for a gas turbine engine ( ) comprising a spinner ( ) having a generally conical forward portion ( ) and a base portion ( ). The base portion ( ) of the spinner ( ) is disposed radially inside an extended profile of the forward portion ( ). Upon the base ( ) there is a radially outwardly extending flange ( ) through which mounting fasteners ( ) are arranged to attach the spinner ( ) to a fan hub ( ) of the gas turbine engine ( ). A fairing ( ) is arranged to surround and cover the base portion ( ) of the spinner ( ) when the nose cone assembly ( ) is fitted to the fan hub ( ). The fairing ( ) having a generally circumferentially and axially smooth continuous outer surface, the profile of which continues that of the forward portion of the spinner . Fasteners ( ) preferably extend from the fairing ( ) in a substantially radially inwardly directed direction to attach the fairing ( ). At least three mounting brackets ( ) are disposed radially inside the fairing ( ) to attach the fairing ( ).

Gas turbine nose cone assembly

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. It is preferred that the first and second epicyclic gear arrangements are dimensioned and configured to drive the first and second multi-bladed rotors with equal and opposite torques, so that the static carrying structure can be of reduced strength and weight.

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Derby, United Kingdom

Adam MacGregor Bagnall