Rolls-royce Corporation

Nicholas Grech

Jacopo Tacconi

Richard Geoffrey Stretton

Michael Christopher Willmot

Chloe J Palmer

Matthew John Willshee

John Wells

Christopher Mosley

An aircraft propulsion system includes a gas turbine engine configured to burn hydrogen fuel supplied from a cooled hydrogen supply via a fuel conduit. A fuel offtake is configured and arranged to divert a portion of hydrogen fuel from the fuel conduit. A burner is configured to burn the portion of hydrogen fuel diverted from the main fuel conduit with air bled from the core compressor via an offtake to produce exhaust gases. A heat exchanger is configured to transfer heat from the exhaust gasses produced by the burner to hydrogen fuel in the main fuel conduit.

Gas turbine engine

A gas turbine engine fuel system includes a fuel offtake configured to divert a portion of hydrogen fuel from a main fuel conduit, a burner configured to burn the portion of hydrogen fuel diverted from the main fuel conduit, a heat exchanger configured to transfer heat from exhaust gasses produced by the burner to hydrogen fuel in the main fuel conduit and a power recovery device configured to extract power from exhaust gasses of the burner downstream of the heat-exchanger.

Gas turbine engine fuel system

A combined gas turbine engine and hydrogen fuel cell system includes a hydrogen fuelled gas turbine engine, a cryogenic liquid hydrogen fuel tank, a first fuel offtake configured and arranged to divert a portion of hydrogen fuel from a main fuel conduit, a burner configured and arranged to burn the portion of hydrogen fuel diverted from the main fuel conduit, a heat exchanger configured and arranged to transfer heat from exhaust gasses produced by the burner to hydrogen fuel in the main fuel conduit, a second fuel offtake arranged to divert a portion of hydrogen fuel from the main fuel conduit downstream of the heat exchanger, and a hydrogen fuel cell configured and arranged to produce electric power using hydrogen fuel diverted from the second fuel offtake.

Combined gas turbine engine and fuel cell

A propulsive aircraft gas turbine engine comprises a turbine disposed in a gas turbine engine core flow and a recuperator heat exchanger disposed downstream of the turbine in gas turbine engine core flow, the recuperator heat exchanger being configured to transfer heat from the gas turbine engine core flow to gas turbine engine fuel. The recuperator heat exchanger is configured to accommodate a portion of the gas turbine engine core flow therethrough, the remainder being bypassed around the recuperator heat exchanger.

A gas turbine engine () for an aircraft comprises an engine core () comprising a turbine (), a compressor (), a core shaft (), and a core exhaust nozzle (), the core exhaust nozzle () having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit (); a fan () comprising a plurality of fan blades; and a nacelle () surrounding the fan () and the engine core () and defining a bypass duct (), the bypass duct () comprising a bypass exhaust nozzle (), the bypass exhaust nozzle () having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit; wherein a bypass to core ratio of: is configured to be in the range from 1.1 to 1.4 under aircraft cruise conditions.

Turbine engine core and bypass flows having a defined fan-turbine radial distance

A gas turbine engine includes a core engine having a compressor, a combustor and a turbine in sequential air flow series. The engine further includes a fuel offtake configured and arranged to divert a portion of hydrogen fuel from a main fuel conduit, a burner configured and arranged to burn the portion of hydrogen fuel diverted from the main fuel conduit and a heat exchanger configured and arranged to transfer heat from exhaust gasses produced by the burner to hydrogen fuel in the main fuel conduit. At least first and second compressor bleed offtakes are at different pressure stages of the compressor, each being configured to bleed a portion of air from the compressor. At least first and second compressor bleed offtake valves are configured to control flow through the first and second bleed offtakes respectively. The burner is configured to receive bleed air from the compressor bleed offtakes.

A gas turbine engine for an aircraft includes an engine core including a turbine, a compressor, a core shaft, and a core exhaust nozzle, the core exhaust nozzle having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit; a fan including a plurality of fan blades; and a nacelle surrounding the fan and the engine core and defining a bypass duct, the bypass duct including a bypass exhaust nozzle, the bypass exhaust nozzle having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit; wherein a bypass to core ratio of: is configured to be in the range from 1.1 to 2.0 under aircraft cruise conditions.

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Derby, United Kingdom

Nicholas Grech