Rolls-Royce Corporation

Rolls-Royce Canada Limited

Christopher Freeman

Ivor J Day

Thomas Scarinci

Adam MacGregor Bagnall

Kenneth Franklin Udall

Nicholas Howarth

David Michael Beaven

David S Knott

Martyn Richards

William B Wright

Peter G Farrar

John W Allen

Richard J Flatman

Brian Barry

Peter G G Farrar

Robert Rudolph Moritz

Ivor John Day

With highly loaded rotors and stators problems can occur with secondary flows sweeping low momentum fluid across the blades reducing efficiency. By provision of collector slots to collect the secondary air and direct that air to a return slot in a rotor hub it is possible to provide impetus to the collected secondary flow to an outlet slot such that there is dispersal of the secondary flow and therefore reduce the effects upon the overall performance of a gas turbine engine incorporating the arrangement.

Rotor arrangement

In order to achieve greater operational efficiency a propulsive arrangement is provided in which a relatively large number of small propulsive assembliesare provided in aircraft structures in particular the wing. The assembliescomprise a flow path within which a compressoris positioned to propel an air flow for propulsion. The compressorsare powered by electricity from an electricity generatorand possibly supplemented by solar cells or fuel cells or wing tip turbine generators. The assembliesare small to allow easy removal for maintenance and repair.

Propulsion arrangement

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

A fan turbine engine arrangement is provided in which a core engine is aligned such that an input of a nacelle is aligned with the upwash of a wing to which the arrangement is secured. The engine drives a fan such that bypass flows produced by the fan are guided by a duct to present a downward component for uplift. The core engine presents its output flow through a nozzle which is aligned with the fundamental axis Y&#x2014;Y of the engine. In such circumstances, the bypass flows cross the core engine flow from the nozzle unless that engine is adjustable during operation in order to maintain alignment dependent upon lift requirements.

Turbine engine arrangements

A three-stage lean burn combustion chamber () comprises a primary combustion zone (), a secondary combustion zone () and a tertiary combustion zone (). Each of the combustion zones () is supplied with premixed fuel and air by respective fuel and air mixing ducts (). The fuel and air mixing ducts () have a plurality of air injection slots () spaced apart transversely to the direction of flow through the fuel and air mixing ducts (). The air injection slots () extend in the direction of flow through the fuel and air mixing ducts () to the reduce the magnitude of the fluctuations in the fuel to air ratio of the fuel and air mixture supplied into the at least one combustion zone (). This reduces the generation of harmful vibrations in the combustion chamber ().

Combustion chamber

A three-stage lean burn combustion chamber ( ) comprises a primary combustion zone ( ), a secondary combustion zone ( ) and a tertiary combustion zone ( ). Each of the combustion zones ( ) is supplied with premixed fuel and air by respective fuel and air mixing ducts ( ). The fuel and air mixing ducts ( ) have a plurality of air injection slots ( ) spaced apart transversely to the direction of flow through the fuel and air mixing ducts ( ). The air injection slots ( ) extend in the direction of flow through the fuel and air mixing ducts ( ) to the reduce the magnitude of the fluctuations in the fuel to air ratio of the fuel and air mixture supplied into the at least one combustion zone ( ). This reduces the generation of harmful vibrations in the combustion chamber ( ).

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 variable cycle gas turbine engine ( ) includes first and second compressors ( ), combustion apparatus ( ) and first and second turbines ( ) operable to drive the first and second compressors ( ) respectively via interconnecting shafts. The capacity of one of the turbines ( ) may be varied, for example by means of guide vanes ( ) which can be adjusted to reduce or increase a throat area ( ) through which air leaves the guide vanes to impact the turbine. The capacity of the turbine may be increased at low engine speeds, to reduce the pressure ratio across the turbine and across the compressor ( ) which it drives, thereby improving the surge margin of the compressor.

Cycle gas turbine engine

A three-stage lean burn combustion chamber ( ) comprises a primary combustion zone ( ), a secondary combustion zone ( ) and a tertiary combustion zone ( ). Each of the combustion zones ( ) is supplied with premixed fuel and air by respective fuel and air mixing ducts ( ). The fuel and air mixing ducts ( ) have a plurality of air injections apertures ( ) spaced apart in the direction of flow through the fuel and air mixing ducts ( ). The apertures ( ) reduce the magnitude of the fluctuations in the fuel to air ratio of the fuel and air mixture supplied into the at least one combustion zone ( ). This reduces the generation of harmful vibrations in the combustion chamber ( ).

A gas turbine engine fan blade ( ) comprises a root portion ( ) and an aerofoil portion ( ). The aerofoil portion ( ) has a convex surface ( ), a concave surface ( ), a leading edge ( ) and a trailing edge ( ). The leading edge ( ) of the aerofoil portion ( ) is formed from a harder material than the regions of the concave surface ( ) and convex surface ( ) immediately adjacent the leading edge ( ) such that the leading edge ( ) of the aerofoil portion ( ) remains pointed. This improves the efficiency of the fan blade ( ) and hence the gas turbine engine ( ) and reduces flutter of the fan blade ( ).

Gas turbine engine blade

A ducted fan gas turbine engine is provided with a fan which is carried by a shaft. A frangible fuse ring maintains the shaft carrying the fan coaxial with the engine longitudinal axis. In the event of severe fan damage, the fuse ring (28) fractures and the fan shaft orbits about the engine longitudinal axis. A pair of threaded members (39, 41), one of which is of generally frusto-conical configuration, cooperate as a result of the orbiting motion to restore the coaxial relationship between the fan shaft and the engine longitudinal axis.

Ducted fan gas turbine engine having a frangible connection

Flow control method and means

The invention provides a method of controlling gas flow in an axial flow compressor, in which the flow at least at one chosen station in the direction of flow through the compressor is sensed at a series of circumferentially spaced positions. Flow variations above a predetermined limit are evaluated to give an actuating response if the variations indicate a disturbance above a predetermined acceptable level is present. When such a disturbance is detected, high pressure flow bled from the exit end of the compressor is injected at said station to supplement the main gas flow there. An incipient rotating stall cell will appear as a variation occurring sequentially at the circumferentially spaced positions. By responding to such a condition with a pressure injection flow, it is found possible to suppress both rotating stall and surge conditions in the compressor before the disturbance develops fully.

A rotor or stator blade for an axial flow compressor has one or both of its end portions designed to minimize the interference between the annulus boundary layer and the blade. The basis of the invention is that the aerofoil is designed so that each section of the aerofoil through the boundary layer performs a constant amount of work per unit height.

Growing community of inventors

Nottingham, United Kingdom

Christopher Freeman