Rolls-Royce Corporation

Rolls-Royce Deutschland Ltd & Co. Kg

Mark David Taylor

Neil Wiliam Harvey

Nicholas Howarth

Rory Douglas Stieger

Robert John Miller

Samuel Alexander Mason

Frank G Noppel

John Richard Webster

Anthony G Razzell

Arthur Laurence Rowe

Shahrokh Shahpar

Martin G Rose

Anthony Brian Parry

Erik Janke

Daniel Swain

Ritindar Singh

Riti Singh

Torbjorn O Lindquist

Alasdair Gardner

Jonathan J H Heffer

Frank Gunter Noppel

An auxiliary power unit () for an aircraft () comprises a gas turbine engine comprising an engine core (). The engine core comprises a core compressor () and core turbine () coupled by a core shaft (). The auxiliary power unit further comprises a load spool comprising a load compressor (), a load turbine () and a permanent magnet electric machine (), each being coupled by a load shaft (). The load shaft () and core shaft () are configured to rotate independently of one another.

Aircraft auxiliary power unit

A turboshaft gas turbine engine comprises, in fluid flow series, a gas-generator compressor, a combustor, a gas-generator turbine, and a free power turbine. The expansion ratios of the turbines are dependent upon an outlet temperature of the combustor, and an expansion relationship between the turbines is defined as the ratio of expansion ratios of the turbines over a running range of non-dimensional power outputs of the gas turbine engine. A second derivative of the expansion relationship is from 0.6 to 1.0.

Turboshaft gas turbine engine and expansion ratio relationship

An engine that has, in axial flow series, booster compressor, core compressor, combustion equipment, core turbine, and low-pressure turbine. Core turbine drives core compressor via an interconnecting high-pressure shaft. Low-pressure turbine drives booster compressor via an interconnecting low-pressure shaft. Low-pressure turbine also drives external load having a defined speed characteristic that dictates speed of the low-pressure turbine and booster compressor. Booster compressor has one or more rows of variable stator vanes. The method includes: scheduling variation in the angle of variable stator vanes as a function of speed of the booster compressor wherein the vanes open as booster compressor speed increases; measuring or setting one or more operational parameters which are determinative of temperature at entry to core turbine; and biasing scheduling of angle variation of variable stator vanes as a function of operational parameter(s) to reduce variation in temperature at entry to core turbine.

Control of a gas turbine engine

A gas turbine engine () having a pressure-rise combustor () and the pressure-rise combustor () is positioned upstream of a stage of turbine nozzle guide vanes (). The vanes () of the stage of turbine nozzle guide vanes () forms an ejector and each of the vanes () has an upstream portion () and a downstream portion (). The upstream portions () of the vanes () have leading edges () and the upstream portions () are arranged substantially straight and parallel to define constant area mixing passages () for a flow of gases there-through. The downstream portions () of the vanes () are arranged at an angle to the upstream portions () of the vanes () to turn the flow of gases there-through.

Gas turbine engine

A method of operating an aircraft, and the aircraft itself, comprising a gas turbine engine that exhausts a plume of gases in use, the aircraft is characterized by comprising an electromagnetic radiation generator and a waveguide to direct electromagnetic radiation at the exhaust plume to avoid the formation of contrails.

Method and apparatus for suppressing aeroengine contrails

A method of operating a contra-rotating propeller engine that preferably comprises a 12 bladed front and a 9 bladed rear propeller. As is conventional, the engine is operated during at least a take-off phase, a cruise phase and an approach phase; during the cruise phase the engine operates with a generally constant propeller tip speed. The method is characterized by the step of operating the engine such that the tip speed of either or both of the propellers, during at least one of take-off, climb or approach, at least 10% greater than cruise tip speed. With a specific front to rear propeller spacing, increasing the tip speed reduces overall noise generated by the propellers.

Aeroengine

Environmental concerns have resulted in consideration of the effects of condensation trails from gas turbine engines. By provision of a condensation stage in a heat exchanger arrangement for a gas turbine engine the level of condensation within a final exhaust gas flow G from an engine is reduced. Furthermore, by cooling of the exhaust gas flows for mixing with by-pass air flows a water partial pressure at an exhaust gas exit temperature can be provided which is below the eutectic liquid to vapor phase transition and therefore avoid condensation (contrail) formation.

Gas turbine engine having a heat exchanger arrangement for exhaust gas flows

A rotor and a stator assembly, wherein the rotor () comprises at least one stage of rotor blades () and the stator () comprises at least one stage of turbine stator vanes (). The rotor blades () have aerofoils and platforms () and the turbine stator vanes () having aerofoils and platforms (). A seal () is defined between the rotor blade () platforms () and the stator vane () platforms () wherein a portion (B) of the rotor blade platforms () and/or a portion (B) of the stator vane () platforms () comprise a shape memory alloy member or a bimetallic member. The shape memory alloy member, or bimetallic member, controls the flow of cooling air through the seals () allowing a greater cooling flow at higher temperatures than at lower temperatures.

Seal between rotor blade platforms and stator vane platforms, a rotor blade and a stator vane

A gas turbine engine turbine comprises an annular array of nozzle guide vanes and an annular array of turbine blades mounted within its annular casing. An array of radially extending protrusions are positioned axially upstream of said array of said nozzle guide vanes and protruding inwardly from the inner casing wall so as to mix the tangential momentum component of the overtip leakage fluid flow before it reaches the array of nozzle guide vanes.

Turbine

A pressure rise combustor is provided with fuel provided at intermittent periods. The fuel is pulsed at timings such that the phase lag between the addition of the fuel and a resultant pressure rise is minimized. The fuel is pulsed such that the unsteady addition of heat reinforces the amplitude of an unsteady pressure fluctuation.

Pulse jet engine having pressure sensor means for controlling fuel delivery into a combustion chamber

A turbine engine arrangement is provided in which contra rotating shaftsare respectively secured to a fan and a gearboxwhich is also coupled to the shaft. In such circumstances the relative rotational speed ratio between the shaftscan be determined with a first low pressure turbinesecured to the first shaftarranged to rotate at a lower speed but provide high work whilst a second low pressure turbinesecured to the second shaftrotates at a higher speed governed by the gearbox. By such an arrangement a smaller gearboxmay be used as less power is transferred through that gearboxthan with previous arrangements. By contra rotation of the turbinesa lower flow deflection guide vane assemblymay be used and a further stator/guide vane assembly is not required between the turbines

Counter-rotating turbine engine including a gearbox

A guide vane is adapted to be mounted in an annular fluid-conveying duct of a gas turbine engine, and has an aerofoil part which extends radially across the duct when the guide vane is thus-mounted. On a transverse cross-section relative to the direction of intended fluid flow across the aerofoil part, at least a portion of the aerofoil part is S-shaped.

Guide vane for a gas turbine engine

An aerofoil ( ), preferably of a high lift, highly loaded design, for an axial flow turbo machine ( ). The aerofoil having a span, a leading edge (LE), a trailing edge (TE) and a cambered sectional profile comprising a pressure surface ( ) and a suction surface ( ) extending between the leading edge (LE) and trailing edge (TE). The aerofoil ( ) having at least one aerofoil cross bleed passage ( ) defined in the aerofoil ( ) which extends from the pressure surface ( ) through the aerofoil ( ) to the suction surface ( ). The at least one passage ( ) preferably disposed generally at a location on the suction surface ( ) at which boundary layer separation from the suction surface ( ) would normally occur. The passage ( ) arranged to provide a bleed from the pressure surface ( ) to the suction surface ( ) with the passage ( ) preferably angled towards the trailing edge (TE) at a shallow angle relative to the suction surface ( ). The aerofoil ( ) may be an aerofoil of a vane or blade of for example a gas turbine engine compressor or turbine.

Growing community of inventors

Derby, United Kingdom

Mark David Taylor