Rolls-royce Corporation

Rolls-royce Deutschland Ltd & Co. Kg

David Peter Scothern

David Reginald Trainer

Gareth Edward Moore

Stephen Mark Husband

Paul R Davies

Luke George

Craig W Bemment

David S Knott

David Edwards

Stephan Uhkötter

Brian Scott Maners

Paul Simms

Edward J Spalton

Steven P Culwick

Mark Boden

William Williamson

Neil Robert Fomison

Douglas M M Herbert

Don Klemen

Shane Perera

Douglas Herbert

A gas turbine engine for an aircraft comprises, in axial flow sequence, a compressor module, a combustor module, and a turbine module, with a first electric machine being rotationally connected to the turbine module. The first electrical machine is configured to generate a maximum electrical power P(W), and the gas turbine engine is configured to generate a maximum dry thrust T (N); and a ratio S of: is in a range of between 2.0 and 10.0.

Gas turbine engine

Multi-engine aircraft power and propulsion systems and methods of restarting an engine of a multi-engine aircraft during fight are provided. One such method comprises: determining a condition to the effect that a flame in the combustion equipment of the second gas turbine engine has been extinguished; responsive to the determination, supplying electrical power from the electrical energy storage system to one or more of the electric machines of the second gas turbine engine and operating said one or more electric machines as motors to limit a reduction in a speed of the one or more spools of the second gas turbine engine following extinguishment of the flame in its combustion equipment; and restarting the second gas turbine engine by relighting the combustion equipment of the second gas turbine engine.

Restarting a gas turbine engine

An aircraft gas turbine engine includes a heat exchanger module, and a core engine. The core engine includes an intermediate-pressure compressor, high-pressure compressor, and high and low-pressure turbines. The high-pressure compressor rotationally connects to the high-pressure turbine by a first shaft, and the intermediate-pressure compressor rotationally connects to the low-pressure turbine by a second shaft. The heat exchanger module fluidly communicates with the core engine by an inlet duct. The heat exchanger module includes a central hub and multiple heat transfer elements extending radially from the hub and spaced in a circumferential array, for heat energy transfer from a first fluid within the elements to an inlet airflow passing over a surface of the elements prior to airflow entry into an inlet to the core engine. The gas turbine engine further includes a first electric machine rotationally connected to the first shaft, and positioned downstream of the heat exchanger module.

A hybrid aircraft propulsion system. The system comprises a gas turbine engine comprising a compressor, a combustor, one or more turbines, a shaft coupled to one of the turbines, and a bypass fan mechanically driven by the shaft. The system further comprises an electrical generator mechanically coupled to the shaft, and an auxiliary propulsor mechanically coupled to an electric motor and electrically coupled to the electric generator. At maximum power, the gas turbine engine is configured to produce a turbine entry temperature at maximum power between 1800 Kelvin and 2000 Kelvin, the engine comprises a fan bypass ratio of between 4:1 and 13:1, and the generator is configured to absorb between 10% and 60% of the mechanical power generated by the turbine.

Hybrid aircraft propulsion system

An electric starter-generator is described. The generator may comprise a rotor, a housing, two bearings, a stator, and a cooling-fluid flowpath. The cooling-fluid flowpath may comprise a cooling-fluid entrance and exit, and a cooling-fluid channel in fluid communication with the entrance and exit. At least a portion of the channel may be defined by a fluid-tight coupling of the housing and a sleeve radially surrounding the outer race of either the bearings. The portion may form an annulus about the axis. The portion may comprise a radially inner surface defined by the sleeve, a radially outer surface define by said housing, and two axial surfaces. The two axial surfaces may extend a distance from the radially inner to outer surfaces that is less than a distance from one of the two axial surfaces to the other of the two axial surfaces.

Fluid cooling of grease-packed bearings

A gas turbine engine comprising: an engine core comprising a compressor; a compressor bleed valve in communication with the compressor and configured to release bleed air from the compressor; at least one component provided at the inlet of the engine core having a de-icing conduit, configured to receive the bleed air; and a flow controller, configured to provide bleed air to the de-icing conduit of the at least one component in response to either or both of a requirement to de-ice the component and a requirement to release bleed air from the compressor to optimise operation of the core.

An auxiliary oil supply apparatus for a rotating device (), the rotating device comprising a primary oil supply (), a rotating component () and a static component () situated radially outwardly of a centre of rotation of the rotating component () and arranged to collect oil () held radially outwardly from the rotating component () as the rotating component rotates. The auxiliary oil apparatus comprises a scoop () associated with the rotating component () and is responsive to a change in a known parameter to move between a first position in the static component () to a second position between the static component () and the centre of rotation.

Auxiliary oil supply apparatus for a rotating component

A method of operating a gas turbine engine compressor. The engine comprises a compressor having an environmental control system bleed port having an outlet in fluid communication with an aircraft environmental control system air duct, and an air turbine starter configured to rotate a compressor shaft of the gas turbine engine. The air turbine starter has an inlet in fluid communication with the environmental control system air duct via an air turbine valve. The method comprises determining a surge margin of the compressor, and where the surge margin of the compressor is determined to be below a predetermined minimum surge margin, opening the air turbine valve to supply air to the air turbine.

Gas turbine engine compressor control method

An electrical power generator system has: an input shaft for receiving a rotary input drive; plural output shafts connected by respective gear shifting arrangements to the input shaft thereby providing different gear ratios between the input shaft and each output shaft; and plural electrical generators powered by rotation of the respective output shafts, electrical power outputs of the generators being combined to supply a total power to a load. The input shaft operates over a range of rotation frequencies, and the gear shifting arrangements are configured to shift the gear ratios between the input shaft and the output shafts such that the output shafts operate over a narrower range of rotation frequencies.

Electrical power generator system

A fluid transfer coupling comprises a first shaft assembly a second shaft assembly, and a seal assembly. The first shaft assembly comprises a first shaft and an annular fin attached to the first shaft, the annular fin having one or more internal passages extending therethrough. Each internal passage connects a centre portion of the first shaft to a radially outwardly facing side of the annular fin. The second shaft assembly comprises a second shaft and an annular trough extending radially outwardly of the second shaft.

Fluid transfer coupling

The present disclosure concerns the detection of ice within a system. More specifically, but not exclusively, the disclosure concerns the detection of ice accretion within a gas turbine engine. The apparatus and method relies on heating a first region () of a component () and comparing the measured temperature of the first region () with a second temperature value, possibly measured at a distinct second region () of the component ().

Ice detection

A valve monitoring apparatus has a system to determine a state of a valve. The system determines a fluid flow first pressure at a first location within a gas turbine engine, and a second pressure of a compressed fluid at a second location within the engine when the valve is in the first position; and compares the first and second pressures to determine the valve state. The system is arranged to command the valve to move from the first position towards a second position; determine the second pressure of the compressed fluid at the second location; compare the pressure at the second location when the valve is in the first position to the pressure at the second location when the valve has been commanded to move towards the second position; and, determine whether the valve has moved from the first position towards the second position when commanded to do so.

System to determine a state of a valve

A bearing arrangement comprising a shaft with a recess in its end, a bearing housing, an annular first bush comprising an inner first portion and an inner second portion and a second bush that engages the recess of the shaft. The first bush bears against the bearing housing. The shaft engages the inner first portion of the first bush. The second bush engages the inner second portion of the first bush.

Bearing assembly

A fluid transfer coupling includes a first and second shaft assembly. The first shaft assembly includes a shaft and annular fin attached to shaft, annular fin having one or more internal passages extending therethrough. The or each internal passage connects a center portion of shaft to a radially outwardly facing side of annular fin. The second shaft assembly includes a shaft and annular trough extending radially outwardly of shaft. The first shaft is concentrically accommodated within second shaft with annular fin being accommodated within annular trough. In use, first shaft assembly rotates at a first rotational speed and second shaft assembly rotates at a second rotational speed, with second rotational speed being greater than first rotational speed. The difference between second and first rotational speeds cause a fluid contained within annular trough to be driven radially inwardly through the or each internal passage to center portion of first shaft.

Coupling and a method of using the same

Stator vane assembly includes plurality of vanes arranged in circumferential array, radially proximal mounting portion, radially distal mounting portion, and first latching mechanism. Each of plurality of vanes extends radially between radially proximal mounting portion and radially distal mounting portion, with each of plurality of vanes being pivotally mounted between radially proximal mounting portion and radially distal mounting portion. Plurality of vanes is movable between first arrangement allowing gas to flow across vanes, and second arrangement blocking gas flow across vanes. Pivot axis of each of plurality of vanes is offset from the line of aerodynamic center of pressure of vane towards trailing edge of vane, and first latching mechanism holds plurality of vanes in first arrangement. Upon receipt of signal, first latching mechanism releases plurality of vanes so that gas flow through vane assembly causes plurality of vanes to move from first arrangement to second arrangement.

Growing community of inventors

Derby, United Kingdom

David Peter Scothern