Rolls-royce Corporation

Siemens Aktiengesellschaft

Cambridge Enterprise Limited

Siemens Energy Global Gmbh & Co. Kg

Robert John Miller

James Vincent Taylor

Christopher R Hall

Malcolm L Hillel

Bryce D Conduit

Anthony M Dickens

Mark David Taylor

Samuel Alexander Mason

Florin Udrea

Anthony G Razzell

Chloe J Palmer

Martin N Goodhand

Jacopo Tacconi

Howard P Hodson

Anthony J Rae

Hang Wai Lung

Torbjorn O Lindquist

John McGill

Jonathan J H Heffer

Roger Wells

Nathan Clark

Kiran S Auchoybur

Ibraheem Haneef

Chloe Jo Palmer

Martin Goodhand

Hang Lung

Michael Arthur Howard

Martin Neil Goodhand

A combined cycle engine is formed of a gas turbine cycle coupled with a fuel expansion cycle in which fuel for the gas turbine cycle is the working fluid of the fuel expansion cycle. The fuel expansion cycle comprises, in flow series, a fuel pump for pumping a cryogenic fuel for the gas turbine cycle, a first heat exchanger for heating the fuel, an expander to recover power by expansion of the fuel and to thereby drive a load, and a second heat exchanger to reheat the fuel. The gas turbine cycle comprises, in flow series, a compressor, a combustor for combustion of reheated fuel delivered from the second heat exchanger, and a turbine configured to drive the compressor. The first heat exchanger and the second heat exchanger are stationed downstream of the combustor and configured to transfer heat from combustion products to the fuel in the fuel expansion cycle.

Combined cycles

A computer-implemented method comprising: controlling input of data quantifying damage received by a compressor of a gas turbine engine into a first machine learning algorithm; receiving data quantifying a first operating parameter of the compressor as an output of the first machine learning algorithm; and determining operability of the compressor by comparing the received data quantifying the first operating parameter of the compressor with a threshold.

Computer-implemented methods for determining compressor operability

A computer-implemented method controls input of at least a portion of a first training data set into a first machine learning algorithm. The first training data set includes data quantifying damage to a first compressor and data quantifying a first operating parameter of the first compressor. The first machine learning algorithm is executed, and data quantifying the first operating parameter is received as an output of the first machine learning algorithm. The first machine learning algorithm is trained using the received data output from the first machine learning algorithm and data quantifying the first operating parameter of the first compressor. The trained first machine learning algorithm is configured to enable determination of operability of a second compressor of a gas turbine engine.

Computer-implemented methods for training a machine learning algorithm

An aerofoil member that extends between radially inner and radially outer endwalls which define a gas annulus of the compressor. Aerofoil member has a leading edge, a trailing edge, a pressure surface and a suction surface such that successive cross-sections through the aerofoil member transverse to the radial direction provide respective aerofoil sections. The external shape of the aerofoil member is defined by the stacking of the aerofoil sections on a stacking axis which passes through a reference point common to each aerofoil section. The aerofoil member has lean produced by the projection of the stacking axis onto a plane normal to the engine axis intersecting a first one of the endwalls at an angle of from 5&#xb0; to 25&#xb0; to the circumferential direction direction such that the pressure surface faces the first endwall.

Compressor aerofoil member

A multistage compressor for a gas turbine engine is disclosed. The compressor includes a plurality of compressor stages, each stage including in axial flow series a row of rotor blades and a row of stator vanes; a hub to which a radially inner end of each rotor blade of the rows of rotor blades is connected; and a casing that circumscribes the rows of rotor blades and the rows of stator vanes. An aerofoil portion of each of the rotor blades and/or the stator vanes of one or more of the plurality of compressor stages is profiled in a spanwise direction such that the aerofoil portion has a trailing edge that is angled more towards an axial direction of the multistage compressor in regions of the aerofoil portion proximal to the casing and the hub than in a mid-region of the aerofoil portion.

Multistage compressor with aerofoil portion profiled in a spanwise direction

A compressor aerofoil for a turbine engine or an axial process compressor, the compressor aerofoil includes a suction surface wall having a suction surface and a pressure surface wall having a pressure surface, the suction surface wall and the pressure surface wall meet at a leading edge and a trailing edge and define a tip having a tip surface, the aerofoil has a maximum thickness T. A mean camber line is defined as passing through the leading edge and the trailing edge. The compressor aerofoil further includes a winglet at the tip and which extends from the suction surface, the winglet has an overhang W that has a perpendicular extent from the suction surface in the range 0.1Tto 1.5T. The winglet has a maximum overhang Wthat occurs within 50% of the length of the mean camber line from the leading edge.

Compressor aerofoil

An aerofoil having a leading edge point within a leading edge region and a pressure surface with a profile wherein within the leading edge region the pressure surface profile has a local minimum. The local minimum reduces the loss which may be caused by high negative incidence on to the blade.

Aerofoils

This invention relates to hot film shear stress sensors and their fabrication. We describe a hot film shear stress sensor comprising a silicon substrate supporting a membrane having a cavity underneath, said membrane bearing a film of metal and having electrical contacts for heating said film, and wherein said membrane comprises a silicon oxide membrane, where in said metal comprises aluminium or tungsten, and wherein said membrane has a protective layer of a silicon-based material over said film of metal. In preferred embodiments the sensor is fabricated by a CMOS process and the metal comprises aluminium or tungsten.

Shear stress sensors

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 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

Growing community of inventors

Cambridge, United Kingdom

Robert John Miller