Rolls-royce Corporation

Rolls-royce Deutschland Ltd & Co. Kg

Andrea Minelli

Craig W Bemment

Benjamin J Keeler

Martin Kenneth Yates

Alastair G Hobday

Kevin R McNally

Peter Swann

Christopher P Madden

Paul William Ferra

Andrew T Smith

David Michael Beaven

Ivo Martin

Björn Petersen

A gas turbine engine for an aircraft includes an engine core including a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor. The engine includes a fan upstream of the engine core and driven by the core shaft. The engine includes a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside of the engine core, where the bypass ratio, defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at cruise conditions, is greater than or equal to 4. The engine includes a plurality of actuators and a fuel supply system. The fuel supply system is arranged to supply fuel for combustion in the combustor, and to supply fuel to fueldraulically drive at least ten of the plurality of actuators.

Fueldraulic actuation

A method of operating a gas turbine engine is disclosed, the gas turbine engine comprising a combustor arranged to combust a fuel; and a fuel management system arranged to provide the fuel to the combustor. The fuel management system comprises two fuel-oil heat exchangers through which oil and the fuel flow, the heat exchangers arranged to transfer heat to the fuel and comprising a primary fuel-oil heat exchanger and a secondary fuel-oil heat exchanger; and a fuel pump arranged to deliver the fuel to the combustor, wherein the fuel pump is located between the two heat exchangers. The method comprises controlling the fuel management system so as to transfer between 200 and 600 kJ/mof heat to the fuel from the oil in the primary fuel-oil heat exchanger at cruise conditions.

Gas turbine performance

A gas turbine engine for an aircraft, comprising: a combustor; a fuel-oil heat exchanger arranged to receive fuel from a fuel tank onboard the aircraft and transfer heat from the oil to the fuel; a fuel return line arranged to return at least some fuel that has passed through the heat exchanger to the fuel tank; and a modulator valve arranged to modulate the flow of fuel along the fuel return line, wherein the modulator valve is arranged to initiate return of fuel to the fuel tank when fuel having passed through the heat exchanger is at a temperature of at least 120° C. A method of operating a gas turbine engine is also disclosed.

Gas turbine fuel temperature management

A gas turbine engine for an aircraft includes an engine core including a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor. The engine includes a fan located upstream of the engine core and arranged to be driven by the core shaft. The engine includes a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside of the engine core, where the bypass ratio, defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at cruise conditions, is at least 4. The engine includes a plurality of actuators and a fuel supply system arranged to supply fuel to fueldraulically drive at least one actuator of the plurality of actuators. The engine includes two fuel-oil heat exchangers.

Fuel supply system

A gas turbine engine includes engine core, combustor, compressor, and core shaft. The engine includes fan located upstream of engine core and arranged to be driven by core shaft. The engine includes a nacelle surrounding fan and engine core and defining a bypass duct located radially outside of engine core, where a bypass ratio, defined as ratio of mass flow rate of flow through bypass duct to mass flow rate of flow through core at cruise conditions, is at least 4. The engine includes plurality of actuators and fuel supply system arranged to supply fuel for combustion in combustor, and to supply fuel to fueldraulically drive at least one actuator of plurality of actuators. The fuel supply system is arranged to select between: causing fuel to actuate at least one actuator; and causing fuel to bypass the at least one actuator, based on sustainable aviation fuel—SAF—content of fuel.

Gas turbine engine

A method of operating a gas turbine engine having an engine core having a turbine, a compressor, a combustor arranged to combust a fuel, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core; a fan shaft; a gearbox that receives an input from the core shaft and outputs drive to the fan via the fan shaft; a heat exchange system having at least one fuel-oil heat exchanger arranged to transfer heat to the fuel; and a fuel pump arranged to deliver the fuel to the combustor, wherein the fuel pump is located downstream of the at least one fuel-oil heat exchanger. The method comprises controlling the heat exchange system so as to raise the fuel temperature to at least 135° C. on entry to the combustor at cruise conditions.

Heat transfer

A gas turbine engine includes an engine core with a combustor; a turbine including turbine blades; a compressor as a source of cooling air for the turbine blades; and an inducer to accelerate and direct the cooling air onto the turbine blades, and a modulating valve to allow or block cooling air flow into a subset of airflow passageways of the inducer; and a fuel management system to provide fuel to the combustor. The fuel management system includes a primary fuel-oil heat exchanger and a secondary fuel-oil heat exchanger through which oil and the fuel flow, to transfer heat between the oil and the fuel. A method of operating the engine includes using the modulating valve to adjust the cooling air flow based on turbine inlet temperature; and transferring 200-600 kJ/mof heat to the fuel from the oil in the primary fuel-oil heat exchanger at cruise conditions.

Aircraft heat management

A combination of a gas turbine engine and a power electronics for powering aircraft and/or engine systems. The engine includes an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor, and a fuel circuit for supplying a fuel flow to the combustor. The power electronics is configured to transfer heat produced by the power electronics to a cooling flow formed by a portion of the fuel flow. The fuel circuit is configured to circulate the cooling flow in a loop during selected engine conditions such that the cooling flow transfers heat from the power electronics to a phase change material located on the loop. The phase change material has a phase change temperature at a predetermined limiting temperature whereby the phase change material stores heat from the cooling flow to prevent the power electronics exceeding the limiting temperature.

Combination of a gas turbine engine and a power electronics

A gas turbine engine for an aircraft includes an engine core with a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor. The engine includes a fan upstream of the engine core driven by the core shaft. The engine includes a nacelle surrounding the fan and engine core and defining a bypass duct located radially outside of the engine core, where the bypass ratio of the mass flow rate through the bypass duct to the mass flow rate through the core at cruise conditions, is at least 4. The engine includes an engine heat management system and an actuator configured to actuate at least one valve within that system. The engine includes a fuel supply system to supply fuel to fueldraulically drive the actuator. The actuator is configured to actuate said valve to enable non-binary position adjustment between an open and closed valve positions.

Fueldraulic heat management system

A gas turbine engine for an aircraft, comprising: a combustor; a fuel-oil heat exchanger arranged to receive fuel from a fuel tank on board the aircraft and transfer heat from the oil to the fuel; a fuel return line arranged to return at least some fuel that has passed through the heat exchanger to the fuel tank; and a modulator valve arranged to modulate the flow of fuel along the fuel return line such that a ratio of a temperature, in Kelvin, of fuel in the fuel tank to a temperature, in Kelvin, of fuel being delivered to the combustor is less than 0.56. Methods of operating a gas turbine engine are also disclosed.

Gas turbine fuel return line

A method of operating an aircraft gas turbine engine including: a core including a turbine, a compressor, a combustor, and a core shaft connecting the turbine to the compressor; a fan upstream of the core; a fan shaft; at least one bearing supporting the fan shaft; at least one auxiliary system using some of the fuel; an oil loop system to supply oil to the bearing; and a heat exchange system including: a primary fuel-oil heat exchanger through which oil in the oil loop system and substantially all the fuel flow to transfer heat between the oil and the fuel; and a secondary fuel-oil heat exchanger through which oil in the oil loop system and a portion of the fuel flow, to transfer heat between the oil and the fuel; and a fuel pump; wherein, under cruise conditions, fuel temperature on leaving the second heat exchanger is 120° C. to 200° C.

Gas turbine fuel temperature

A method of operating a gas turbine engine including an engine core including a turbine, compressor, combustor to combust a fuel, and core shaft connecting the turbine and compressor; a fan upstream of the engine core; a fan shaft; a gearbox that receives an input from the core shaft and outputs drive to the fan via the fan shaft; a primary oil loop system to supply oil to the gearbox; and a heat exchange system. The method includes controlling the heat exchange system to adjust fuel viscosity to be lower than or equal to 0.58 mm/s on entry to the combustor at cruise conditions.

Fuel viscosity

An aircraft gas turbine engine includes: an engine core including a compressor, combustor, turbine, and a core shaft connecting the turbine to the compressor; a fan including fan blades and arranged upstream of the engine core; turbomachinery bearings; a power gearbox to drive the fan at a lower rotation speed than the turbine; and a heat management system providing lubrication and cooling to the power gearbox and turbomachinery bearings, and including a pipe assembly adapted to provide a lubricant flow to the power gearbox and turbomachinery bearings to remove the heat generated by the power gearbox and turbomachinery bearings, at least one air-lubricant heat exchanger to dissipate a first amount of heat to a first heat sink, and at least one fuel-lubricant heat exchanger to dissipate a second amount of heat to a second heat sink, wherein the first heat sink is air and the second heat sink is fuel.

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Derby, United Kingdom

Andrea Minelli