Abstract: A gas turbine engine includes a gearbox that receives an input from a core shaft and drives a fan at a lower speed than the core shaft; an oil loop system to supply oil to the gearbox; and a heat exchange system with an air-oil heat exchanger; and a fuel-oil heat exchanger. The oil loop system branches such that a proportion of the oil can flow along each branch and the air-oil and fuel-oil heat exchangers are in parallel on different branches of the oil loop system. A modulation valve allows the proportion of the oil sent via each branch to be varied. A method includes controlling the heat exchange system such that, under idle conditions, a heat transfer ratio of rate of heat transfer from oil to air/rate of heat transfer from oil to fuel is from 0.67 to 5.67.

Abstract: There is provided a method of operating a gas turbine engine including a combustor. The combustor includes a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the chamber. The nozzles include a first and second subset of fuel spray nozzles. The combustor is operable in a condition wherein the first subset is supplied with more fuel than the second. A ratio of the number of nozzles in the first subset to the second is 1:2 to 1:5. The method includes operating the engine so a reduction of 20-80% in an average of particles/kg of nvPM in the exhaust when the engine is operating at 85% available thrust for given operating conditions and when the engine is operating at 30% is obtained when fuel provided to the nozzles is a sustainable aviation fuel instead of a fossil-based hydrocarbon fuel. Also provided is a gas turbine engine for an aircraft.

Abstract: 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.

Abstract: 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.

Abstract: 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.

Abstract: A method of operating a gas turbine engine includes an engine core with a turbine, compressor, fuel combustor, and core shaft; a fan upstream of the engine core; a gearbox that receives an input from the core shaft and outputs drive to the fan; an oil loop system supplying oil to the gearbox; and a heat exchange system including: air-oil and fuel-oil heat exchangers, and wherein the oil loop system branches such that part of the oil flows along each branch and the air-oil and fuel-oil heat exchangers are parallel on different branches; and a modulation valve allows the oil sent via each branch to be varied, the method including controlling the heat exchange system wherein, under cruise conditions, a heat transfer ratio of: rate ? of ? heat ? transfer ? from ? oil ? to ? air rate ? of ? heat ? transfer ? from ? oil ? to ? fuel is in the range from 0 to 0.67.

Abstract: A gas turbine engine for an aircraft comprises an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor. The engine comprises a fan located upstream of the engine core and arranged to be driven by the core shaft. The engine comprises 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 comprises 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.

Abstract: 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 the engine core and defining a bypass duct located radially outside 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 greater than or equal to 4. The engine includes a plurality of actuated engine systems, including a heat management system and a turbine case cooling system. The engine includes a fuel supply system arranged to supply fuel for combustion in the combustor, and to supply fuel to fueldraulically drive at least three of the plurality of actuated engine systems.

Abstract: 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.

Abstract: A method of operating a gas turbine engine comprising an engine core comprising a turbine, a compressor, a core shaft connecting the turbine to the compressor, and a combustor arranged to combust a fuel; a fan located upstream of the engine core; an oil system arranged to circulate oil; and a heat exchange system comprising at least one fuel-oil heat exchanger arranged to transfer heat from the oil to the fuel, wherein the method comprises: determining at least one fuel characteristic of the fuel arranged to be provided to the combustor; and modulating the heat exchange system so as to adjust the fuel temperature on entry to the combustor at cruise to a set level, the set level being based on the at least one fuel characteristic.

Abstract: A method of operating a gas turbine engine includes an engine core with a turbine, compressor, fuel combustor, and core shaft connecting the turbine and compressor; a fan located upstream of the engine core; a gearbox that receives an input from the core shaft and outputs drive to the fan; an oil loop system supplying oil to the gearbox; and a heat exchange system including: air-oil and fuel-oil heat exchangers, and wherein the oil loop system includes a bypass pipe allowing oil to bypass a heat exchanger; and a bypass valve allowing the oil sent via the bypass pipe to be varied, the method including controlling the bypass valve wherein, under cruise conditions, a heat transfer ratio of: rate ? of ? heat ? transfer ? from ? oil ? to ? air rate ? of ? heat ? transfer ? from ? oil ? to ? fuel is in the range from 0 to 0.67.

Abstract: A method of operating a gas turbine engine of an aircraft with an engine core including a turbine, compressor, combustor, and core shaft connecting the turbine to the compressor; a fan upstream of the engine core; a gearbox that receives an input from the core shaft and outputs drive to the fan; an oil loop system to supply oil to the gearbox; and a heat exchange system with an air-oil heat exchanger through which the oil flows; and a fuel-oil heat exchanger through which the oil and fuel flow; and a modulation valve to allow the oil sent via each heat exchanger to be varied, the method including determining a fuel characteristic of the fuel to be combusted by the combustor; and controlling the modulation valve based on the at least one fuel characteristic to adjust the proportion of the oil sent via each heat exchanger at cruise conditions.

Abstract: A gas turbine engine for an aircraft includes an engine core including a turbine, combustor, compressor, and core shaft connecting the turbine to the compressor. The engine includes a fan located upstream of engine core and arranged to be driven by core shaft. The engine includes a nacelle surrounding the fan and engine core and defining bypass duct located radially outside of engine core, where bypass ratio, defined as the ratio of the mass flow rate of the flow through bypass duct to the mass flow rate of the flow through the core at cruise conditions, is at least 4. The engine includes plurality of actuators and fuel supply system arranged to supply fuel to fueldraulically drive at least one actuator of the plurality of actuators. At cruise, fuel temperature on entry into the at least one actuator is at least 5° C. greater than fuel temperature on entry to combustor.

Abstract: 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 mm2/s on entry to the combustor at cruise conditions.

Abstract: A method of operating a gas turbine engine includes an engine core with a turbine, compressor, fuel combustor, and core shaft; a fan upstream of the engine core; a gearbox that receives an input from the core shaft and outputs drive to the fan; an oil loop system supplying oil to the gearbox; and a heat exchange system including: air-oil and fuel-oil heat exchangers, and wherein the oil loop system branches such that part of the oil flows along each branch and the air-oil and fuel-oil heat exchangers are parallel on different branches; and a modulation valve allows the oil sent via each branch to be varied, the method including controlling the heat exchange system wherein, under cruise conditions, a heat transfer ratio of: rate ? of ? heat ? transfer ? from ? oil ? to ? air rate ? of ? heat ? transfer ? from ? oil ? to ? fuel is in the range from 0 to 0.67.

Abstract: A method of operating a gas turbine engine; the gas turbine engine includes a combustor. The combustor includes a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber. The plurality of fuel spray nozzles includes a first subset of fuel spray nozzles and a second subset of fuel spray nozzles. The combustor is operable in a condition in which the first subset of fuel spray nozzles are supplied with more fuel than the second subset of fuel spray nozzles. A ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5. The method includes: providing fuel to the one or more fuel-oil heat exchangers. Also provided is a gas turbine engine for an aircraft.

Abstract: A gas turbine engine has a combustor with a combustion chamber and a plurality of fuel spray nozzles including a first subset of nozzles and a second subset of nozzles. The first subset of nozzles are supplied with more fuel than the second subset of nozzles. A ratio of the number of nozzles in the first subset to the number of nozzles in the second subset is in the range of 1:2 to 1:5. A method includes providing fuel to the one or more fuel-oil heat exchangers, transferring heat from oil to the fuel, and providing the fuel from the one or more fuel-oil heat exchangers to the fuel spray nozzles. Heat is transferred from the oil to the fuel to lower a viscosity of the fuel to 0.58 mm2/s or lower on injection of the fuel into the combustion chamber at cruise conditions.

Abstract: There is provided a method of operating a gas turbine engine including a combustor. The combustor includes a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the chamber. The nozzles include a first and second subset of fuel spray nozzles. The combustor is operable in a condition wherein the first subset is supplied with more fuel than the second. A ratio of the number of nozzles in the first subset to the second is 1:2 to 1:5. The method includes operating the engine so a reduction of 20-80% in an average of particles/kg of nvPM in the exhaust when the engine is operating at 85% available thrust for given operating conditions and when the engine is operating at 30% is obtained when fuel provided to the nozzles is a sustainable aviation fuel instead of a fossil-based hydrocarbon fuel. Also provided is a gas turbine engine for an aircraft.

Abstract: A method of operating a gas turbine engine having a combustor having a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber. The fuel spray nozzles have a first subset of fuel spray nozzles and a second subset of fuel spray nozzles. The combustor is operable in a condition in which the first subset of fuel spray nozzles are supplied with more fuel than the second subset of fuel spray nozzles. A ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5. The method includes providing a fuel to the plurality of fuel spray nozzles having a calorific value of at least 43.5 MJ/kg. A gas turbine engine can be for an aircraft.

Abstract: A method of operating a gas turbine engine including: a combustor arranged to combust a fuel; and a fuel management system arranged to provide the fuel to the combustor, wherein the fuel management system includes two fuel-oil heat exchangers through which oil and fuel flow, which are arranged to transfer heat between the oil and fuel and include primary and secondary fuel-oil heat exchangers; a fuel pump arranged to deliver the fuel to the combustor, wherein the fuel pump is located between the heat exchangers; and a recirculation valve located downstream of the primary heat exchanger, the recirculation valve arranged to allow a controlled amount of fuel which has passed through the primary heat exchanger to be returned to the inlet. The method includes selecting one or more fuels such that the calorific value of the fuel provided to the gas turbine engine is at least 43.5 MJ/kg.