Gas turbine engine

Abstract

Gas Turbine Engine A gas turbine engine for an aircraft includes an engine core including a turbine, a combustion chamber, a compressor, and a core shaft connecting the turbine to the compressor. The engine includes a fan located upstream of the engine core and configured 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 outward of the engine core, where a 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 delivery system configured to provide fuel for combustion in the combustion chamber, and to provide fuel to hydraulically fuel-drive at least one actuator of the plurality of actuators. The fuel delivery system is configured to be controlled to select between: causing the fuel to actuate at least one actuator; and causing the fuel to bypass the at least one actuator, depending on the sustainable aviation fuel (SAF) content of the fuel. To be reproduced with Figure 20 upon publication. Figure for abstract: Figure 20.

Description

Title of the invention: Gas turbine engine

[0001] The present disclosure relates to aircraft actuation systems, and methods of actuating a hydraulically fueled actuation system.

[0002] There is an expectation in the aviation industry regarding a trend towards the use of fuels different from the traditional kerosene-based jet fuels generally used today. The fuels may have different fuel characteristics compared to petroleum / fossil-based hydrocarbon fuels. Thus, there is a need to consider fuel properties for these new fuels, and to adjust both the gas turbine engines themselves and the methods of operating the gas turbine engines.

[0003] According to a first aspect, a gas turbine engine for an aircraft is provided, comprising: an engine core comprising a turbine, a combustion chamber, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core and configured to be driven by the core shaft, the fan comprising a plurality of fan blades; a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside 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; a plurality of actuators; and a fuel supply system, wherein the fuel supply system is configured to supply fuel for combustion in the combustion chamber, and to supply fuel to hydraulically fuel-drive at least one actuator of the plurality of actuators, and wherein the fuel system is designed to be controlled to select between: causing the fuel to actuate at least one actuator; and causing the fuel to bypass the at least one actuator, depending on the sustainable aviation fuel (SAF) content of the fuel.

[0004] The inventors were aware that the use of fuels different from traditional kerosene-based jet fuels, such as sustainable aviation fuels, may lead to different fuel properties, and that these different fuel properties may allow the actuators to be driven hydraulically by fuel. In particular, some fuels can be heated to higher temperatures and used to actuate at least one more actuator than traditional fuels, without significantly increasing the risk of fuel thermal degradation (e.g., fuel lacquer, or fuel coking) in the actuators.

[0005] The bypass ratio is greater than or equal to 4, and may be in the range of 4 to 55. The bypass ratio may be in the range of 4 to 20. The bypass ratio may be in the range of 4 to 15.

[0006] The fuel supply system may be designed to: causing the fuel to actuate the at least one actuator when the SAF content of the fuel is above a threshold; and cause the fuel to bypass this actuator when the SAF content of the fuel is below a threshold.

[0007] The fuel system may be configured to select between actuating or bypassing multiple fuel hydraulic actuators depending on the SAF content of the fuel. One or more of the fuel hydraulic actuators may have a different threshold: for example, the fuel system may be configured to cause the fuel to actuate a first actuator when the SAF content of the fuel is greater than a first threshold or to bypass that actuator otherwise, and to cause the fuel to actuate a second actuator when the SAF content of the fuel is greater than a second threshold higher than the first threshold, or to bypass that actuator otherwise.

[0008] The core shaft may directly deliver the drive to the fan, so as to drive the fan at the same rotational speed as the core shaft. Such an engine may be a direct-drive turbine engine.

[0009] The gas turbine engine may include a reduction gear that receives an input from the core shaft and provides a drive to the fan so as to drive the fan at a rotational speed lower than that of the core shaft. Such an engine may be a reduction gear turbine engine.

[0010] The engine may include a turbine case cooling, TCC, system. The plurality of actuators may include an actuator that is part of the turbine case cooling system. The fuel delivery system may be configured to actuate or bypass the actuator that is part of a turbine case cooling system based on the SAF content of the fuel. At least two of the plurality of actuators may be part of the turbine case cooling system. The fuel delivery system may be configured to provide fuel to hydraulically fuel-drive the at least two of the plurality of actuators that are part of the TCC system. The The fuel system may be configured to actuate or bypass the at least two actuators that are part of a turbine case cooling system based on the SAF content of the fuel.

[0011] The engine may include a cabin air bleed valve. The plurality of actuators may include an actuator configured to actuate the cabin air bleed valve. The fuel system may be configured to actuate or bypass the actuator configured to actuate the cabin air bleed valve based on the SAF content of the fuel.

[0012] The engine may include a bleed air management valve. The plurality of actuators may include an actuator configured to actuate the bleed air management valve. The fuel delivery system may be configured to actuate or bypass the bleed air management valve based on the SAF content of the fuel.

[0013] The engine may include an engine heat management system including a valve. The plurality of actuators may include an actuator configured to actuate the valve in the engine heat management system. The fuel delivery system may be configured to actuate or bypass the valve in the engine heat management system based on the SAF content of the fuel.

[0014] The engine heat management system may include a heat exchanger. The heat exchanger may be an air-to-oil heat exchanger. The fuel delivery system may be configured to actuate or bypass a valve on the air side of the heat exchanger based on the SAF content of the fuel. The fuel delivery system may be configured to actuate or bypass a valve on the oil side of the heat exchanger based on the SAF content of the fuel. The heat exchanger may be a fuel-to-oil heat exchanger. The fuel delivery system may be configured to actuate or bypass a valve on the fuel side of the fuel-to-oil heat exchanger based on the SAF content of the fuel.The fuel delivery system may be configured to actuate or bypass a valve on the oil side of the fuel-oil heat exchanger depending on the SAF content of the fuel.

[0015] The engine may include a generator heat management system including a valve. The plurality of actuators may include an actuator configured to actuate the valve in the generator heat management system. The fuel delivery system may be configured to actuate or bypass the valve in the generator heat management system based on the SAF content of the fuel.

[0016] The generator heat management system may include a heat exchanger. The heat exchanger may be an air-to-oil heat exchanger. The fuel delivery system may be configured to actuate or bypass a valve on the air side of the heat exchanger based on the SAF content of the fuel. The fuel delivery system may be configured to actuate or bypass a valve on the oil side of the heat exchanger based on the SAF content of the fuel. The heat exchanger may be a fuel-to-oil heat exchanger. The fuel delivery system may be configured to actuate or bypass a valve on the fuel side of the fuel-to-oil heat exchanger based on the SAF content of the fuel.The fuel delivery system may be configured to actuate or bypass a valve on the oil side of the fuel-oil heat exchanger depending on the SAF content of the fuel.

[0017] When the fuel delivery system bypasses the at least one actuator, the at least one actuator may be actuated in a fuel-non-hydraulic manner. When the actuator is actuated in a fuel-non-hydraulic manner, it may be actuated using a fuel-non-hydraulic fluid. The fuel-non-hydraulic fluid may be supplied through a check valve to prevent mixing. Fuel may be supplied to the actuator through a check valve to prevent mixing. Instead, the actuator may be electric or pneumatic.

[0018] The fuel delivery system may be configured to cause the fuel to actuate the at least one actuator when the SAF content is above a threshold. The fuel delivery system may be configured to cause the fuel to bypass the at least one actuator when the SAF content is below a threshold.

[0019] The minimum SAF content required for actuation of the at least one actuator may be at least 25%, 30%, 35%, 40%, 45%, 50%, 52%, 55%, 60%, 65%, 70% or 75% by volume.

[0020] The fuel delivery system may be configured to be controlled to select between fuel hydraulic actuation and bypass for two or more of the plurality of actuators. For example, the fuel delivery system may be controlled to select between fuel hydraulic actuation and bypass for two, three, four, or five actuators.

[0021] When the at least one actuator is fuel hydraulically driven, the maximum operating differential pressure of that actuator under takeoff conditions may be in the range of 6,900 kPa to 10,000 kPa, or may be greater than 10,000 kPa. When the at least one actuator is fuel hydraulically driven, the maximum operating differential pressure of that actuator under takeoff conditions may be in the range of 6,900 kPa to 10,000 kPa, or may be greater than 10,000 kPa. fuel hydraulics, the maximum operating differential pressure under takeoff conditions may be greater than 7,000 kPa, 8,000 kPa, 9,000 kPa, 10,000 kPa, 11,000 kPa, 12,000 kPa, 13,000 kPa, 14,000 kPa, or greater than 15,000 kPa.

[0022] When at least one actuator is fuel hydraulically driven, the maximum operating differential pressure under takeoff conditions may be greater than 2,400 kPa, greater than 2,500 kPa, greater than 3,000 kPa, greater than 3,500 kPa, greater than 3,800 kPa, or greater than 4,000 kPa.

[0023] The same actuator may see significantly lower pressure at idle, for example with a differential pressure in the range of 1000 kPa to 1250 kPa (150 to 180 psid).

[0024] According to a second aspect, a method of operating a gas turbine engine for an aircraft is provided, the engine comprising: an engine core comprising a turbine, a combustion chamber, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core and configured to be driven by the core shaft, the fan comprising a plurality of fan blades; a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside 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; a plurality of actuators; and a fuel supply system; and wherein the method comprises: the supply, using the fuel supply system, of fuel for combustion in the combustion chamber; and the choice between: supplying, using the fuel supply system, fuel to at least one actuator of the plurality of actuators to hydraulically fuel-actuate the at least one actuator; and using the fuel delivery system, causing fuel to bypass the at least one actuator, depending on the sustainable aviation fuel (SAF) content of the fuel.

[0025] The method may comprise: supplying, using the fuel supply system, fuel to at least two of the plurality of actuators and, depending on the sustainable aviation fuel (SAF) content of the fuel, the choice for each of the at least two actuators (individually / independently) between: causing the fuel to hydraulically actuate the actuator; and causing the fuel to bypass the actuator.

[0026] The method may comprise: providing, using the fuel supply system, fuel to at least two of the plurality of actuators and, depending on the sustainable aviation fuel (SAF) content of the fuel, choosing between: causing the fuel to hydraulically actuate the at least two actuators; and causing the fuel to bypass the at least two actuators.

[0027] The bypass decision may therefore be a single decision made for all relevant actuators, or may be a series of decisions made for individual actuators.

[0028] The bypass ratio may be in the range of 4 to 55. The bypass ratio may be in the range of 4 to 20. The bypass ratio may be in the range of 4 to 15.

[0029] The method of the second aspect may be performed using the engine of the first aspect.

[0030] The method may include increasing the pressure of the fuel flowing through the at least one hydraulic fuel actuator of the fuel as a function of the SAF content of the fuel. For example, the pressure may be increased by at least 350 kPa for every 5% increase in %SAF above 60% (volumetric %). One or more pumps and / or valves may be used to increase the pressure.

[0031] According to a third aspect, a gas turbine engine for an aircraft is provided, comprising: an engine core comprising a turbine, a combustion chamber, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core and configured to be driven by the core shaft, the fan comprising a plurality of fan blades; a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside 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; a plurality of actuators; and a fuel supply system, wherein the fuel supply system is configured to supply fuel for combustion in the combustion chamber, and to supply fuel to hydraulically fuel-drive at least one actuator of the plurality of actuators.

[0032] The fuel comprises at least 25% sustainable aviation fuel (SAF) by volume. The fuel system is designed such that the maximum differential pressure of the fuel across the at least one hydraulic fuel actuator under cruise conditions is at least 2,400 kPa (350 psid).

[0033] It will be borne in mind that, even under cruising conditions, the differential pressure across an actuator varies depending on the operating conditions (e.g., fuel flow). The maximum differential pressure is therefore chosen to facilitate comparison; this pressure is more specifically a maximum steady-state differential pressure; that is, short-term transient pressure spikes are not taken into account. The maximum steady-state value may therefore be a time-averaged pressure value, calculated over five, ten, fifteen, twenty, twenty-five, or thirty seconds. The peak must last at least about five seconds, which excludes abrupt transient spikes.

[0034] The maximum fuel differential pressure across at least one fuel hydraulic actuator under cruise conditions may be at least 2500 kPa, 2600 kPa, 2750 kPa, 3000 kPa, 3200 kPa, 3400 kPa, 3500 kPa, 3600 kPa, 3700 kPa, 3800 kPa, or 4000 kPa.

[0035] The maximum differential pressure of the fuel in the at least one hydraulic fuel actuator under cruising conditions may be in the range of 2400 kPa to 4500 kPa, and optionally 2500 kPa to 4000 kPA, or 2500 kPa to 3800 kPA,

[0036] The bypass ratio is greater than or equal to 4, and may be in the range of 4 to 55. The bypass ratio may be in the range of 4 to 20. The bypass ratio may be in the range of 4 to 15.

[0037] When the at least one actuator is fuel hydraulically driven, the maximum operating differential pressure under takeoff conditions may be in the range of 6,900 kPa to 10,000 kPa, or may be greater than 7,000 kPa, 8,000 kPa, 9,000 kPa, or 10,000 kPa.

[0038] The same actuator may see significantly lower pressure at idle, for example with a differential pressure in the range of 1000 kPa to 1250 kPa (150 to 180 psid).

[0039] The core shaft can directly deliver drive to the fan, so as to drive the fan at the same rotational speed as the core shaft, such that the engine is a direct-drive turbine engine.

[0040] The turbine engine may include a reduction gear that receives an input from the core shaft and provides a drive to the fan so as to drive the fan at a rotational speed lower than that of the core shaft, such that the engine is a reduction gear turbine engine.

[0041] The fuel may comprise at least 50%, 55%, 60%, 65%, 70% or 75% SAF by volume.

[0042] The at least one fuel hydraulic actuator may be a variable stator vane actuator.

[0043] The at least one fuel hydraulic actuator may be a variable inlet guide vane actuator.

[0044] The engine of the third aspect may include any or all of the features of the engine of the first aspect, and may be used to implement the method of the second aspect.

[0045] According to a fourth aspect, there is provided a method of operating a gas turbine engine for an aircraft. The engine comprises: an engine core comprising a turbine, a combustion chamber, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core and configured to be driven by the core shaft, the fan comprising a plurality of fan blades; a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside 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; a plurality of actuators; and a fuel supply system;

[0046] The method comprises: providing, using the fuel supply system, fuel for combustion in the combustion chamber; and providing, using the fuel delivery system, fuel comprising at least 25% SAF by volume to fuel-hydraulic drive at least one of the plurality of actuators, the fuel being provided such that a maximum differential pressure of the fuel across the at least one fuel-hydraulic actuator under cruise conditions is at least 2400 kPa.

[0047] Fuel of the same composition may be supplied to both the combustion chamber and the at least one actuator. One or more valves and / or pumps controllable may be provided to adjust the fuel differential pressure across the at least one fuel hydraulic actuator.

[0048] The maximum differential fuel pressure across the at least one fuel hydraulic actuator under cruise conditions may be at least 2,500 kPa.

[0049] The bypass ratio may be in the range of 4 to 55. The bypass ratio may be in the range of 4 to 20. The bypass ratio may be in the range of 4 to 15.

[0050] The method may include controlling the fuel supply to the one or more fuel hydraulic actuators so as to adjust the maximum fuel differential pressure across the at least one fuel hydraulic actuator based on the SAF content of the fuel.

[0051] The method may include providing a fuel comprising at least 50% SAF by volume, and controlling the fuel supply to the at least one actuator such that the maximum differential pressure is at least 3200 kPa.

[0052] The method may include providing a fuel comprising at least 55% SAF by volume, and controlling the fuel supply to the at least one actuator such that the maximum differential pressure is at least 3550 kPa or 3600 kPa.

[0053] The methods of the second and fourth aspects may be complementary, and may be performed together in various implementations. The method of the fourth aspect may be performed using the engine of the first aspect or the third aspect.

[0054] According to a fifth aspect, there is provided a gas turbine engine for an aircraft, comprising: an engine core comprising a turbine, a combustion chamber, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core and configured to be driven by the core shaft, the fan comprising a plurality of fan blades; a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside 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; a plurality of actuators; a fuel supply system, wherein the fuel supply system is configured to supply fuel for combustion in the combustion chamber, and to supply fuel to hydraulically fuel-drive at least one actuator of the plurality of actuators, and two fuel-oil heat exchangers adapted to have oil and fuel passing therethrough, the heat exchangers being adapted to transfer heat from the oil to the fuel and comprising a primary fuel-oil heat exchanger adapted to heat at least the majority of the fuel, and a secondary fuel-oil heat exchanger adapted to provide additional heat to the fuel to be supplied to hydraulically drive the at least one fuel hydraulic actuator.

[0055] The heat exchangers are controlled such that, under cruising conditions, a heat transfer ratio of:

[0056] [Math.l] oil-to-fuel heat transfer rate in the secondary fuel-to-oil heat exchanger (kJkg1) oil-to-fuel heat transfer rate in the primary and secondary fuel-to-oil heat exchangers (kJkg'1) has a maximum value of at least 0.35.

[0057] While it is appreciated that the ratio is dimensionless, heat transfer is measured per unit mass or volume of fuel in the examples described herein (with units of kJ per kg noted above as an example), thereby providing a normalized heat transfer ratio for variations in cruise fuel flow. It will be appreciated that heat transfer is noted per unit mass (kg) of fuel reaching the combustion chamber so as to adjust fuel flow and any recirculation through or bypass of one or more heat exchangers as described elsewhere herein. In most implementations, any fuel temperature rise due to other engine components (as opposed to heat transfer from the oil) can be assumed to be minimal.

[0058] It will be borne in mind that, even under cruising conditions, there may be short-term transient peaks in heat transfer, which should not be taken into account. The maximum value of the heat transfer ratio may therefore be a time-averaged value, averaged over one, two, five, ten, fifteen, twenty, twenty-five or thirty minutes. It will be borne in mind that the time scales for transient changes in temperature may be larger than those for pressure in the actuators (while the full movement of an actuator is likely to be completed in a few seconds, fuel and oil may spend more time in the heat exchangers).

[0059] The heat transfer ratio may have a maximum value at cruising speed greater than 0.4. The heat transfer ratio may have a value maximum at cruising speed greater than 0.45. The heat transfer ratio may have a maximum value at cruising speed greater than 0.5.

[0060] The bypass ratio is greater than or equal to 4, and may be in the range of 4 to 55. The bypass ratio may be in the range of 4 to 20. The bypass ratio may be in the range of 4 to 15.

[0061] The maximum temperature of the fuel exiting the secondary fuel-oil heat exchanger may be at least 150°C, 160°C, 170°C, or 180°C.

[0062] The primary and secondary heat exchangers may be arranged such that the heat transfer ratio is adjustable during engine operation, for example by controlling one or more pumps and / or valves to adjust the flow of fuel and / or oil through one or each heat exchanger.

[0063] The heat exchangers may be arranged such that the heat transfer ratio is adjustable depending on the sustainable aviation fuel (SAF) content of the fuel.

[0064] The engine may include at least one bypass valve configured to allow fuel or oil to bypass one of the fuel-oil heat exchangers, so as to adjust the heat transfer ratio.

[0065] For example, the engine may include one or more of: • an operational valve to allow oil to bypass the primary heat exchanger; • an operational valve to allow oil to bypass the secondary heat exchanger; • an operational valve to allow fuel to bypass the primary heat exchanger; and • an operational valve to allow fuel to bypass the secondary heat exchanger.

[0066] The engine may include at least one recirculation valve configured to allow fuel or oil to pass multiple times through one of the fuel-oil heat exchangers, so as to adjust the heat transfer ratio.

[0067] For example, the engine may include one or more of: • an operational valve to allow oil to be recirculated through the primary heat exchanger; • an operational valve to allow oil to be recirculated through the secondary heat exchanger; • an operational valve to allow fuel to be recirculated through the primary heat exchanger; and • an operational valve to allow fuel to be recirculated through the secondary heat exchanger.

[0068] For any of the recirculation valves described herein, there may be one or more associated pumps designed to return the oil / fuel to the inlet of the heat exchanger, so as to re-pass through the heat exchanger. A, or in addition, any suitable components for repressurizing the oil / fuel to enable recirculation may be used.

[0069] The core shaft can directly deliver drive to the fan, so as to drive the fan at the same rotational speed as the core shaft, such that the engine is a direct-drive turbine engine.

[0070] The motor of the fifth aspect may comprise any or all of the features of the motor of the first and / or third aspect, and may be used to implement the method of the second or fourth aspect.

[0071] According to a sixth aspect, a method of operating a gas turbine engine for an aircraft is provided.

[0072] The engine comprises: an engine core comprising a turbine, a combustion chamber, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core and configured to be driven by the core shaft, the fan comprising a plurality of fan blades; a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside 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; a plurality of actuators; a fuel supply system; and a primary fuel-oil heat exchanger and a secondary fuel-oil heat exchanger.

[0073] The method comprises: supplying, using the fuel supply system, fuel for combustion in the combustion chamber; providing, using the fuel supply system, fuel to hydraulically fuel-drive at least one actuator of the plurality of actuators; heating at least the majority of the fuel using the primary fuel-oil heat exchanger; and heating the fuel to be supplied to hydraulically drive the at least one fuel hydraulic actuator using the secondary fuel-oil heat exchanger; in which, under cruising conditions, the power supply and heating are controlled so that a heat transfer ratio of:

[0074] [Math 2] oil-to-fuel heat transfer rate in the secondary fuel-to-oil heat exchanger (kJkg1) oil-to-fuel heat transfer rate in the primary and secondary fuel-to-oil heat exchangers (klkg'j has a maximum value of at least 0.35.

[0075] The bypass ratio may be in the range of 4 to 55. The bypass ratio may be in the range of 4 to 20. The bypass ratio may be in the range of 4 to 15.

[0076] The method may further comprise adjusting the heat transfer ratio (e.g., by adjusting the flow rate of fuel and / or oil in one or each heat exchanger) based on the sustainable aviation fuel (SAF) content of the fuel.

[0077] The methods of the second, fourth, and sixth aspects may be complementary, and may be performed together, or in any combination or subcombination, in various implementations. The method of the sixth aspect may be performed using the engine of the first aspect, the third aspect, or the fifth aspect.

[0078] According to a seventh aspect, a gas turbine engine for an aircraft is provided comprising: an engine core comprising a turbine, a combustion chamber, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core and configured to be driven by the core shaft, the fan comprising a plurality of fan blades; a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside 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; a plurality of actuators including a variable compressor vane actuator; and a fuel supply system, wherein the fuel supply system is adapted to supply fuel for combustion in the combustion chamber, and to supply fuel for hydraulically driving the fuel at least one variable compressor vane actuator, wherein the fuel is thermally stable at temperatures exceeding 280°C.

[0079] The bypass ratio is greater than or equal to 4, and may be in the range of 4 to 55. The bypass ratio may be in the range of 4 to 20. The bypass ratio may be in the range of 4 to 15.

[0080] Thermal stability may refer to the ability of the fuel to meet the requirements of a jet fuel thermal oxidation test (JFTOT). Fuel that is thermally stable at a temperature above 280°C may refer to fuel that has met the requirements of a JFTOT test at temperatures above or equal to 280°C.

[0081] The fuel may be thermally stable at temperatures equal to or greater than 290°C, 295°C, 300°C, 305°C, 310°C, 315°C, 320°C, or 325°C.

[0082] Aircraft gas turbine engines generally include a plurality of compressor blades; the variable compressor blade actuator is configured to adjust the position of a compressor blade. Since the blade is used to guide the airflow, and the position of the compressor blade is variable, it can be referred to as a variable guide vane.

[0083] Variable guide vanes of various implementations may include either or both of the following: • variable inlet guide vanes (VIGVs), which are located upstream of / in front of the compressor rotor blades, and arranged to guide and control the airflow in the compressor; and • other vanes, typically variable stator vanes (VSVs) that are located behind / downstream of the compressor rotor vanes, and arranged to guide and control airflow through the compressor.

[0084] The engine may include multiple variable stator vane actuators, each of which may be fuel hydraulic. Further, the engine may include multiple compressors, each with at least one variable stator vane actuator.

[0085] The gas turbine engine may include a plurality of variable compressor blades, and multiple variable compressor blade actuators. The fuel supply system may be configured to supply fuel to hydraulically fuel-drive the multiple variable compressor blade actuators.

[0086] The gas turbine engine may include at least two compressors, each compressor having at least one variable compressor blade and at least one variable compressor blade actuator associated therewith. The fuel supply system may be configured to supply fuel to drive hydraulically by fuel at least one variable compressor vane actuator of each compressor.

[0087] The engine may include a turbine case cooling system. The plurality of actuators may include a turbine case cooling actuator. The fuel supply system may be configured to provide fuel to hydraulically fuel-drive the turbine case cooling actuator.

[0088] An aircraft gas turbine engine typically includes multiple servomotors (also referred to as servos), a servo being a rotary or linear actuator that provides precise control of angular or linear position, speed, and / or acceleration in a mechanical system. In a given engine, one, more, or all of the servos present may be hydraulically actuated by fuel. One or more of the servomotors may include a sensor configured to provide position feedback. A dedicated controller may be provided for the engine heat management system to control the one or more servomotors.

[0089] For example, the engine may include a servomotor in a hydromechanical unit. The plurality of actuators may include at least one actuator configured to actuate the servomotor in the hydromechanical unit. The fuel supply system may be configured to provide fuel to hydraulically fuel-drive the actuator configured to actuate the servomotor valve.

[0090] Aromatic compounds may represent less than 5% of the fuel volume.

[0091] The calorific value of the fuel may be at least 43.5 MJkg *. The calorific value of the fuel may be at least 44 MJkg *.

[0092] The sulfur content of the fuel may be less than 15 parts per million.

[0093] The fuel may be or comprise a HEFA fuel, i.e. a fuel made from esters and hydrotreated fatty acids.

[0094] The core shaft can directly deliver drive to the fan, so as to drive the fan at the same rotational speed as the core shaft, such that the engine is a direct-drive turbine engine.

[0095] The turbine engine may include a reduction gear that receives an input from the core shaft and provides a drive to the fan so as to drive the fan at a rotational speed lower than that of the core shaft, such that the engine is a reduction gear turbine engine.

[0096] The motor of the seventh aspect may include any or all of the features of the motor of the first, third and / or fifth aspect, and may be used to implement the method of the second, fourth or sixth aspect.

[0097] According to an eighth aspect of the invention, a method of operating a gas turbine engine for an aircraft is provided.

[0098] The engine comprises: an engine core comprising a turbine, a combustion chamber, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core and configured to be driven by the core shaft, the fan comprising a plurality of fan blades; a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside 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; a plurality of actuators including a variable compressor vane actuator; and a fuel supply system.

[0099] The bypass ratio is greater than or equal to 4 and may be in the range of 4 to 55. The bypass ratio can be in the range of 4 to 20. The bypass ratio can be in the range of 4 to 15.

[0100] The method comprises: providing, using the fuel supply system, fuel for combustion in the combustion chamber; and providing, using the fuel supply system, fuel to hydraulically fuel-drive the variable compressor vane actuator of the plurality of actuators.

[0101] The fuel is thermally stable at temperatures above 280°C.

[0102] The methods of the second, fourth, sixth, and eighth aspects may be complementary, and may be performed together, or in any combination or subcombination, in various implementations. The method of the eighth aspect may be performed using the engine of the first aspect, the third aspect, the fifth aspect, or the seventh aspect.

[0103] According to a ninth aspect, there is provided a gas turbine engine for an aircraft. The engine comprises: an engine core comprising a turbine, a combustion chamber, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core and configured to be driven by the core shaft, the fan comprising a plurality of fan blades; a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside the engine core, where a 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; a plurality of actuators; a fuel supply system, wherein the fuel supply system is configured to supply fuel for combustion in the combustion chamber, and to supply fuel to hydraulically fuel-drive at least one actuator of the plurality of actuators; and at least one fuel-oil heat exchanger adapted to have oil and fuel passing therethrough, the at least one heat exchanger being adapted to transfer heat from the oil to the fuel.

[0104] The at least one heat exchanger is arranged such that, in cruising mode, the temperature of the fuel at the inlet of the at least one actuator is at least 5°C higher than the temperature of the fuel at the inlet of the combustion chamber.

[0105] The bypass ratio is greater than or equal to 4, and may be in the range of 4 to 55. The bypass ratio may be in the range of 4 to 20. The bypass ratio may be in the range of 4 to 15.

[0106] The at least one heat exchanger may be arranged such that, in cruising mode, the temperature of the fuel at the inlet of the at least one actuator is at least 7°C, 10°C, 12°C, 15°C, 20°C, 25°C, or 30°C higher than the temperature of the fuel at the inlet of the combustion chamber.

[0107] The core shaft can directly deliver drive to the fan, so as to drive the fan at the same rotational speed as the core shaft, such that the engine is a direct-drive turbine engine.

[0108] The turbine engine may include a reduction gear that receives an input from the core shaft and provides a drive to the fan so as to drive the fan at a rotational speed lower than that of the core shaft, such that the engine is a reduction gear turbine engine.

[0109] The at least one heat exchanger may include a primary fuel-oil heat exchanger configured to heat at least the majority of the fuel, and a secondary fuel-oil heat exchanger configured to provide additional heat to the fuel to be supplied to hydraulically drive the at least one fuel hydraulic actuator.

[0110] The primary and secondary fuel-oil heat exchangers can be controlled such that, under cruising conditions, a heat transfer ratio of: [YES] [Math 3] oil-to-fuel heat transfer rate in the secondary fuel-to-oil heat exchanger (kJkg' ) oil-to-fuel heat transfer rate in the primary and secondary fuel-to-oil heat exchangers (k.Tkg") has a maximum value of at least 0.35.

[0112] The primary and secondary heat exchangers may be arranged such that the heat transfer ratio is adjustable during engine operation.

[0113] The gas turbine engine may include at least one of the following to allow adjustment of the heat transfer ratio in operation: (i) at least one controllable oil bypass valve configured to allow oil to bypass at least one of the primary and secondary fuel-oil heat exchangers; and (ii) at least one controllable oil bypass valve adapted to allow oil to be recirculated through at least one of the primary and secondary fuel-oil heat exchangers.

[0114] The fuel delivery system may include an operable valve to allow fuel to bypass, or be recirculated through, the primary heat exchanger.

[0115] The fuel delivery system may include an operable valve to allow fuel to bypass, or be recirculated through, the secondary heat exchanger.

[0116] The engine may be designed to allow the heat transfer ratio to be adjusted based on the sustainable aviation fuel (SAF) content of the fuel. A higher heat transfer ratio may be permitted for higher SAF contents.

[0117] The fuel supply system may include or be connected to a fuel source. The fuel source may be located external to the engine (e.g., a tank in the main fuselage of an aircraft or on a wing of an aircraft) and may not be part of the engine. Fuel supplied to the at least one actuator for hydraulically fuel-driving the at least one actuator may be supplied to the combustion chamber after actuation without returning to the fuel source. The fuel may therefore be recirculated around the engine, or used in multiple ways within the engine, without being returned to a fuel tank.

[0118] The fuel supplied to the combustion chamber may comprise a mixture of a fuel that has passed through at least one fuel-to-oil heat exchanger and a fuel that has bypassed the at least one fuel-to-oil heat exchanger.

[0119] The engine of the ninth aspect may include any or all of the features of the engine of the first, third, fifth, and / or seventh aspect, and may be used to implement the method of the second, fourth, sixth, and / or eighth aspect.

[0120] According to a tenth aspect, a method of operating a gas turbine engine for an aircraft is provided.

[0121] The engine comprises: an engine core comprising a turbine, a combustion chamber, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core and configured to be driven by the core shaft, the fan comprising a plurality of fan blades; a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside 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; a plurality of actuators; a fuel supply system, wherein the fuel supply system is configured to supply fuel for combustion in the combustion chamber, and to supply fuel to hydraulically fuel-drive at least one actuator of the plurality of actuators; and at least one fuel-oil heat exchanger adapted to have oil and fuel passing therethrough, the at least one heat exchanger being adapted to transfer heat from the oil to the fuel.

[0122] The method comprises controlling the at least one heat exchanger such that, in cruising mode, the temperature of the fuel at the inlet of the at least one actuator is at least 5°C higher than the temperature of the fuel at the inlet of the combustion chamber.

[0123] The bypass ratio may be in the range of 4 to 55. The bypass ratio may be in the range of 4 to 20. The bypass ratio may be in the range of 4 to 15.

[0124] The method may include determining at least one characteristic of the fuel, and controlling a temperature difference between the temperature of the fuel at the inlet of the at least one actuator and the temperature of the fuel at the combustion chamber inlet depending on at least one characteristic of the fuel.

[0125] The fuel delivery system may obtain the fuel from a fuel source such as an aircraft fuel tank (which is generally located external to the engine); and the method may be such that fuel supplied to the at least one actuator for hydraulically fuel-driving the at least one actuator is supplied to the combustion chamber after actuation without returning to the fuel source.

[0126] The gas turbine engine may include primary and secondary fuel-oil heat exchangers, and the method may include controlling the primary and secondary fuel-oil heat exchangers such that, under cruising conditions, a heat transfer ratio of:

[0127] [Math 4] oil-to-fuel heat transfer rate in the secondary fuel-to-oil heat exchanger (Ukg) oil-to-fuel heat transfer rate in the primary and secondary fuel-to-oil heat exchangers (kJkg^) has a maximum value of at least 0.35.

[0128] The method may include controlling the primary and secondary fuel-oil heat exchangers to adjust the heat transfer ratio based on the sustainable aviation fuel (SAF) content of the fuel.

[0129] The methods of the second, fourth, sixth, eighth, and tenth aspects may be complementary, and may be performed together, or in any combination or subcombination, in various implementations. The method of the tenth aspect may be performed using the engine of the first aspect, the third aspect, the fifth aspect, the seventh aspect, or the ninth aspect.

[0130] It will be borne in mind that a feature described in relation to one aspect may be used in combination with any other aspect, mutatis mutandis.

[0131] As indicated elsewhere herein, the present disclosure may apply to any relevant configuration of a gas turbine engine. Such a gas turbine engine may be, for example, a bypass gas turbine engine, an open rotor gas turbine engine (in which the propeller is not surrounded by a nacelle), a turboprop engine, or a turbojet engine. Any such engine may or may not be provided with an afterburner. Such a gas turbine engine may be, for example, designed for land-based or marine power generation applications.

[0132] A gas turbine engine in accordance with any aspect of the present disclosure may include an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor. Such a gas turbine engine may include a fan (having blades of fan). Such a fan may be located upstream of the engine core. Alternatively, in some examples, the gas turbine engine may include a fan located downstream of the engine core, for example in the case where the gas turbine engine is an open rotor or turboprop engine (in which case the fan may be referred to as a propeller).

[0133] Where the gas turbine engine is an open rotor or turboprop engine, the gas turbine engine may include two counter-rotating propeller stages attached to and driven by a free power turbine via a shaft. The propellers may rotate in opposite directions such that one rotates clockwise and the other counter-clockwise about the engine's rotational axis. Alternatively, the gas turbine engine may include a propeller stage and a guide vane stage arranged downstream of the propeller stage. The guide vane stage may be of variable pitch. Thus, the high-pressure, intermediate-pressure, and free-power turbines may respectively drive high- and intermediate-pressure propellers and compressors via suitable interconnecting shafts. Thus, the propellers may provide the majority of the propulsive thrust.

[0134] In the case where the gas turbine engine is an open rotor or turboprop engine, one or more of the propeller stages may be driven by a reduction gear. The reduction gear may be of the type described herein.

[0135] An engine according to the present disclosure may be a bypass engine. Such an engine may be a direct-drive bypass engine in which the fan is directly connected to the fan drive turbine, for example without a reduction gear, via a core shaft. In such a direct-drive bypass engine, the fan may be said to rotate at the same rotational speed as the fan drive turbine. Strictly by way of example, the fan drive turbine may be a first turbine, the core shaft may be a first core shaft, and the gas turbine engine may further comprise a second turbine and a second core shaft connecting the second turbine to the compressor. The second turbine, the compressor, and the second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft.In such an arrangement, the second turbine may be positioned axially upstream of the first turbine.

[0136] An engine according to the present disclosure may be a geared turbofan engine. In such an arrangement, the engine has a fan that is driven via a reduction gear. Accordingly, such a gas turbine engine may include a reduction gear that receives an input from the core shaft and provides a drive to the fan so as to drive the fan at a rotational speed lower than that of the core shaft. The input to the reduction gear may be directly from the core shaft, or indirectly from the core shaft, for example via a spur shaft and / or gear. The core shaft may connect the turbine and compressor, such that the turbine and compressor rotate at the same speed (with the fan rotating at a lower speed).

[0137] The gas turbine engine as described and / or claimed herein may have any suitable general architecture. For example, the gas turbine engine may have any desired number of shafts that connect turbines and compressors, e.g., one, two, or three shafts. By way of example only, the turbine connected to the core shaft may be a first turbine, the compressor connected to the core shaft may be a first compressor, and the core shaft may be a first core shaft. The engine core may further include a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, second compressor, and second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft.

[0138] In such an arrangement, the second compressor may be positioned axially downstream of the first compressor. The second compressor may be arranged to receive (e.g., directly receive, e.g., via a generally annular conduit) a flow from the first compressor.

[0139] The reducer may be arranged to be driven by the core shaft which is designed to rotate (e.g. in use) at the lowest rotational speed (e.g. the first core shaft in the example above). For example, the reducer may be arranged to be driven only by the core shaft which is designed to rotate (e.g. in use) at the lowest rotational speed (e.g. only by the first core shaft, and not the second core shaft, in the example above). Alternatively, the reducer may be arranged to be driven by any one or more shafts, e.g. the first and / or second shafts in the example above.

[0140] The reducer may be a reduction gearbox (in that the output to the fan has a lower rotational speed than the input from the core shaft). Any type of reducer may be used. For example, the reducer may be a "planetary" or "star" reducer, as described in more detail elsewhere in this document. Such a reducer may be a single stage. Alternatively, such a reducer may be a compound reducer, for example a compound planetary reducer (which may have the input on the sun gear and the output on the ring gear, and thus be referred to as a "compound star" reducer), for example with two reduction stages.

[0141] The reducer may have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example greater than 2.5, for example in the range of 3 to 4.2 or 3.2 to 3.8, for example in the range of or at least 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. The gear ratio may be, for example, between any two of the values ​​in the preceding sentence. Strictly by way of example, the reducer may be a "star" reducer having a reduction ratio in the range of 3.1 or 3.2 to 3.8. Strictly by way of further example, the reducer may be a "star" reducer having a reduction ratio in the range of 3.0 to 3.1. Strictly by way of example, the reducer may be a "planetary" reducer having a reduction ratio in the range of 3.6 to 4.2.In some arrangements the gear ratio may be outside these ranges.

[0142] In any gas turbine engine as described and / or claimed herein, fuel of a given composition or mixture is supplied to a combustor, which may be provided downstream of the fan and the compressor(s) with respect to the flow path (e.g., axially downstream). For example, the combustor may be directly downstream of (e.g., at the outlet of) the second compressor, when a second compressor is provided. As a further example, the flow at the outlet to the combustor may be supplied at the inlet of the second turbine, when a second turbine is provided. The combustor may be provided upstream of the turbine(s).

[0143] The or each compressor (e.g., the first compressor and the second compressor as described above) may comprise any number of stages, e.g., multiple stages. Each stage may comprise a row of rotor blades and a row of stator blades, which may be variable stator blades (in that their angle of incidence may be variable). The row of rotor blades and the row of stator blades may be axially offset from each other. For example, the gas turbine engine may be a direct-drive, bypass gas turbine engine comprising 13 or 14 compressor stages (in addition to the fan). Such an engine may, for example, comprise 3 stages in the first (or "low pressure") compressor and either 10 or 11 stages in the second (or "high pressure") compressor.As a further example, the gas turbine engine may be a "geared" gas turbine engine (in which the fan is driven by a first core shaft via a reduction gearbox) comprising 11, 12, or 13 compressor stages (in addition to the fan). Such an engine may comprise 3 or 4 stages in the first (or "low pressure") compressor and 8 or 9 stages in the second (or "high pressure") compressor. As a further example, the gas turbine engine may be a . “gear” gas having 4 stages in the first (or “low pressure”) compressor and 10 stages in the second (or “high pressure”) compressor.

[0144] The or each turbine (e.g., the first turbine and the second turbine as described above) may comprise any number of stages, e.g., multiple stages. Each stage may comprise a row of rotor blades and a row of stator blades, or vice versa, as required. The respective rows of rotor blades and stator blades may be axially offset from each other. The second (or "high pressure") turbine may comprise 2 stages in any arrangement (e.g., regardless of whether it is a geared or direct-drive engine). The gas turbine engine may be a direct-drive gas turbine engine comprising a first (or "low pressure") turbine having 5, 6, or 7 stages. Alternatively, the gas turbine engine may be a "geared" gas turbine engine comprising a first (or "low pressure") turbine having 3 or 4 stages.

[0145] Each fan blade may be defined as having a radial span extending from a root (or hub) at a radially inner gas-washed location, or a 0% span position, to a tip at a 100% span position. The ratio of the fan blade radius at the hub to the fan blade radius at the tip may be less than (or in the order of) any of: 0.4, 0.39, 0.38, 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. The ratio of the fan blade radius at the hub to the fan blade radius at the tip may be included in an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, in the range 0.28 to 0.32, or 0.29 to 0.30.These ratios can be commonly referred to as the hub-to-tip ratio. Both the radius at the hub and the radius at the tip can be measured at the leading edge (or axially forward-most) portion of the blade. The hub-to-tip ratio refers, of course, to the gas-washed portion of the fan blade, i.e., the portion radially outboard of any platform.

[0146] The fan radius may be measured from the centerline of the engine to the tip of a fan blade at its leading edge. The fan diameter (which may simply be twice the fan radius) may be greater than (or of the order of) any of: 140 cm, 170 cm, 180 cm, 190 cm, 200 cm, 210 cm, 220 cm, 230 cm, 240 cm, 250 cm (about 100 inches), 260 cm, 270 cm (about 105 inches), 280 cm (about 110 inches), 290 cm (about 115 inches), 300 cm (about 120 inches), 310 cm, 320 cm (about 125 inches), 330 cm (about 130 inches), 340 cm (about 135 inches), 350 cm, 360 cm (about 140 inches), 370 cm (about 145 inches), 380 cm (about 150 inches), 390 cm (about 155 inches), 400 cm, 410 cm (about 160 inches), or 420 cm (about 165 inches). The blower diameter may be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, within the range of 210 cm to 240 cm, or 250 cm to 280 cm, or 320 cm to 380 cm. Strictly by way of non-limiting example, the blower diameter may be in the range of 170 cm to 180 cm, 190 cm to 200 cm, 200 cm to 210 cm, 210 cm to 230 cm, 290 cm to 300 cm or 340 cm to 360 cm.

[0147] The rotational speed of the fan may vary during use. Generally, the rotational speed is lower for fans with a larger diameter. Strictly by way of non-limiting example, the rotational speed of the fan at cruise conditions may be less than 3500 rpm, for example less than 2600 rpm, or less than 2500 rpm, or less than 2300 rpm. Strictly by way of further non-limiting example, the rotational speed of the fan at cruise conditions for a "geared" gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 2750 to 2900 rpm.Strictly by way of further non-limiting example, the fan rotational speed at cruise conditions for a "geared" gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 2500 to 2800 rpm. Strictly by way of further non-limiting example, the fan rotational speed at cruise conditions for a "geared" gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 1500 to 1800 rpm. Strictly by way of further non-limiting example, the fan rotational speed at cruise conditions for a direct drive engine having a fan diameter in the range of 190 cm to 200 cm may be in the range of 3,600 to 3,900 rpm.Strictly by way of further non-limiting example, the fan rotational speed at cruise conditions for a direct drive engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 2000 to 2800 rpm.

[0148] During operation of the gas turbine engine, the fan (with associated fan blades) rotates about an axis of rotation. This rotation results in a displacement of the fan blade tip with a speed Utip. The work done by the fan blades on the flow results in an increase in enthalpy dH of the flow. A fan tip load can be defined as dH / Upeak2, where dH is the enthalpy increase (e.g., the 1-D average enthalpy increase) across the fan and UpPoint is the (translational) velocity of the fan tip, e.g., at the tip leading edge (which can be defined as the fan tip radius at the leading edge times the angular velocity). The fan tip load at cruise conditions can be greater than (or of the order of) any of: 0.28, 0.29, 0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39, or 0.4 (all values ​​dimensionless).The fan peak load may be in an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, in the range 0.28 to 0.31 or 0.29 to 0.3 (e.g., for a geared gas turbine engine).

[0149] Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio (BPR), where the bypass ratio is 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. In some arrangements the bypass ratio at cruise conditions may be greater than (or of the order of) any of the following: 9. 9.5, 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5, or 20. The bypass ratio at cruise conditions may be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, in the range 12 to 16, 13 to 15, or 13 to 14.Strictly by way of non-limiting example, the bypass ratio at cruise conditions of a direct drive gas turbine engine according to the present disclosure may be in the range of 9:1 to 11:1. Strictly by way of further non-limiting example, the bypass ratio at cruise conditions of a geared gas turbine engine according to the present disclosure may be in the range of 12:1 to 15:1. The bypass duct may be at least substantially annular. The bypass duct may be radially outboard of the core engine. The radially outer surface of the bypass duct may be defined by a nacelle and / or a fan case.

[0150] The overall pressure ratio (OPR) of a gas turbine engine as described and / or claimed herein may be defined as the ratio of the stagnation pressure at the outlet of the highest pressure compressor (before entry into the combustion chamber) to the stagnation pressure upstream of the fan. As a non-limiting example, the overall pressure ratio of a gas turbine engine as described and / or claimed herein at cruise conditions may be greater than (or in the order of) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio may be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, in the range 50 to 70. Strictly by way of non-limiting example, the overall pressure ratio at cruise conditions of a geared gas turbine engine having a fan diameter in the range 200 cm to 210 cm may be in the range 40 to 45. Strictly by way of non-limiting example, the overall pressure ratio at cruise conditions of a geared gas turbine engine having a fan diameter in the range 210 cm to 230 cm can be in the range of 45 to 55.Strictly by way of non-limiting example, the overall pressure ratio at cruise conditions of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 50 to 60. Strictly by way of non-limiting example, the overall pressure ratio at cruise conditions of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 50 to 60.

[0151] The specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow rate through the engine. In some examples, specific thrust may depend, for a given thrust condition, on the specific fuel composition supplied to the combustion chamber. At cruise conditions, the specific thrust of an engine described and / or claimed herein may be less than (or on the order of) any of the following: 110 Nkg 1 s, 105 Nkg 's, 100 Nkg 's, 95 Nkg 's, 90 Nkg 's, 85 Nkg 's or 80 Nkg 's. The specific thrust may be in an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, in the range of 80 Nkg 's to 100 Nkg 's, or of 85 Nkg 's to 95 Nkg 's. Such engines may be particularly efficient compared to conventional gas turbine engines.Strictly by way of non-limiting example, the specific thrust of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 90 Nkg 's to 95 Nkg 's. Strictly by way of non-limiting example, the specific thrust of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 80 Nkg 's to 90 Nkg 's. Strictly by way of non-limiting example, the specific thrust of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 70 Nkg 's to 90 . Nkg 's. Strictly by way of non-limiting example, the specific thrust of a direct-drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 90 Nkg 's to 120 Nkg 's.

[0152] A gas turbine engine as described and / or claimed herein may have any desired maximum thrust. Strictly by way of non-limiting example, a gas turbine as described and / or claimed herein may be capable of producing a maximum thrust of at least (or in the order of) any of the following: 100 kN, 110 kN, 120 kN, 130 kN, 135 kN, 140 kN, 145 kN, 150 kN, 155 kN, 160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN, 450 kN, 500 kN, or 550 kN. The maximum thrust may be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits). By way of non-limiting example only, a gas turbine as described and / or claimed herein may be capable of producing a maximum thrust in the range of 155 kN to 170 kN, 330 kN to 420 kN, or 350 kN to 400 kN.Strictly by way of non-limiting example, the maximum thrust of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 140 kN to 160 kN. Strictly by way of non-limiting example, the maximum thrust of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 150 kN to 200 kN. Strictly by way of non-limiting example, the maximum thrust of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 370 kN to 500 kN. Strictly by way of non-limiting example, the maximum thrust of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 370 kN to 500 kN.The above mentioned thrust may be the maximum net thrust under typical atmospheric conditions at sea level plus 15 degrees C (ambient pressure 101.3 kPa, temperature 30 degrees C), with the engine static.

[0153] In use, the flow temperature at the inlet of the high pressure turbine may be particularly high. This temperature, which may be referred to as TET, may be measured at the outlet of the combustion chamber, for example immediately upstream of the first turbine blade, which itself may be referred to as the nozzle guide vane. In some examples, the TET may depend, for a given thrust condition, on the specific composition of fuel supplied to the combustion chamber. At cruise conditions, the TET may be at least (or of the order of) any one of the following values: 1400 K, 1450 K, 1500 K, 1520 K, 1530 K, 1540 K, 1550 K, 1600 K or 1650 K. Thus, solely by way of non-limiting example, the TET at cruise conditions of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 1540 K to 1600 K. Strictly by way of non-limiting example, the TET at cruise conditions of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 1590 K to 1650 K. Strictly by way of non-limiting example, the TET at cruise conditions of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 1600K to 1660K.Strictly by way of non-limiting example, the TET at cruise conditions of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 1590 K to 1650 K. Strictly by way of non-limiting example, the TET at cruise conditions of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 1570 K to 1630 K.

[0154] The TET at cruise conditions may be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, 1530 K to 1600 K. The maximum TET in engine operation may be, for example, at least (or in the order of) any one of the following values: 1700 K, 1750 K, 1800 K, 1850 K, 1900 K, 1950 K, 2000 K, 2050 K, or 2100 K. Thus, solely by way of non-limiting example, the maximum TET of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 1890 K to 1960 K. Strictly by way of As a non-limiting example, the maximum TET of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 1890 K to 1960 K.Strictly by way of non-limiting example, the maximum TET of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 1890 K to 1960 K. Strictly by way of non-limiting example, the maximum TET of a direct-drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 1935 K to 1995 K. Strictly by way of non-limiting example, the maximum TET of a direct-drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 1890 K to 1950 K. The maximum TET may be in an inclusive range bounded by any two of the values ​​in the preceding sentence. (i.e., the values ​​may form upper or lower limits), for example in the range from 1800 K to 1950 K, or from 1900 K to 2000 K.The maximum TET can occur, by . example, in a high thrust condition, for example in a maximum takeoff thrust (MTTP) condition.

[0155] A fan blade and / or airfoil portion of a fan blade described and / or claimed herein may be fabricated from any suitable material or combination of materials. For example, at least a portion of the fan blade and / or airfoil may be fabricated at least in part from a composite, for example, a metal matrix composite and / or an organic matrix composite, such as a carbon fiber composite. As a further example, at least a portion of the fan blade and / or airfoil may be fabricated at least in part from a metal, such as a titanium-based metal or an aluminum-based material (such as an aluminum-lithium alloy) or a steel-based material. The fan blade may include at least two regions fabricated using different materials.For example, the fan blade may have a protective leading edge, which may be manufactured using a material that is better able to withstand impact (e.g., by birds, ice, or other material) than the rest of the blade. Such a leading edge may, for example, be manufactured using titanium or a titanium-based alloy. Thus, strictly by way of example, the fan blade may have a carbon fiber or aluminum-based (such as an aluminum-lithium alloy) body with a titanium leading edge.

[0156] A fan as described and / or claimed herein may comprise a central portion, from which the fan blades may extend, for example in a radial direction. The fan blades may be connected to the central portion in any desired manner. For example, each fan blade may comprise a fastening element that may engage a corresponding notch in the hub (or disc). Strictly by way of example, such a fastening element may be in the form of a dovetail that may notch into and / or engage a corresponding notch in the hub / disc to secure the fan blade to the hub / disc. As a further example, the fan blades may be formed integrally with a central portion. Such an arrangement may be referred to as a bladed disc or a bladed ring.Any suitable method may be used to manufacture such a bladed disc or bladed ring. For example, at least a portion of the fan blades may be machined from a block and / or at least a portion of the fan blades may be joined to the hub / disc by welding, such as a linear friction stir weld.

[0157] The gas turbine engines described and / or claimed herein may or may not be provided with a variable area nozzle (VAN). Such a VAN may allow the bypass duct output to be varied during use. The general principles of this description can be applied to engines with or without VAN.

[0158] The fan of a gas turbine as described and / or claimed herein may have any desired number of fan blades, for example, 14, 16, 18, 20, 22, 24, or 26 fan blades. When the fan blades have a carbon fiber composite body, there may be 16 or 18 fan blades. When the fan blades have a metal body (for example, aluminum-lithium or titanium alloy), there may be 18, 20, or 22 fan blades.

[0159] As used herein, the terms idle, taxi, takeoff, climb, cruise, descent, approach, and landing (or one or more parts thereof) have the conventional meaning and would be readily understood by those skilled in the art. Thus, for a given gas turbine engine for an aircraft, one skilled in the art would readily recognize that each term refers to all, or one or more parts, of an engine operating phase within a given mission of an aircraft to which the gas turbine engine is designed to be attached.

[0160] As such, ground idling may refer to a phase of engine operation where the aircraft is stationary and in contact with the ground, but where there is a need for the engine to taxi. At idle, the engine may produce between 3% and 9% of the available engine thrust. In other non-limiting examples, the engine may produce between 5% and 8% of the available thrust. In other non-limiting examples, the engine may produce between 6% and 7% of the available thrust. Taxiing may refer to a phase of engine operation where the aircraft is propelled along the ground by the thrust produced by the engine. When taxiing, the engine may produce between 5% and 15% of the available thrust. In other non-limiting examples, the engine may produce between 6% and 12% of the available thrust. In other non-limiting examples, the engine may produce between 7% and 10% of the available thrust.Takeoff may refer to a phase of engine operation where the aircraft is propelled by thrust produced by the engine. At an initial stage in the takeoff phase, the aircraft may be propelled while the aircraft is in contact with the ground. At a later stage in the takeoff phase, the aircraft may be propelled while the aircraft is not in contact with the ground. During takeoff, the engine may produce between 90% and 100% of the available thrust. In other non-limiting examples, the engine may produce between 95% and 100% of the available thrust. In other non-limiting examples, the engine may produce 100% of the available thrust.

[0161] Climb may refer to a phase of engine operation where the aircraft is propelled by the thrust produced by the engine. During climb, the engine may produce between 75% and 100% of the available thrust. In other non-limiting examples, the engine may produce between 80% and 95% of the available thrust. In other non-limiting examples, the engine may produce between 85% and 90% of the available thrust. In this regard, climb may refer to an operational phase of an aircraft's flight cycle between takeoff and arrival at cruise conditions, where arrival at cruise conditions thereby defines the start of the cruise phase, or part thereof, of the aircraft's flight. Additionally or alternatively, climb may refer to a nominal point in, or one or more nominal periods during, an aircraft's flight cycle between takeoff and landing, where a relative increase in altitude is required, which may necessitate an additional demand for engine thrust.

[0162] As used herein, cruise conditions, which may define the cruise phase (or a portion thereof) of aircraft flight, have a conventional meaning and are readily understood by those skilled in the art. In some examples, for a given gas turbine engine for an aircraft, cruise conditions may refer to the mid-cruise engine operating point of a given mission (which may be referred to in the industry as an "economic mission") of an aircraft to which the gas turbine engine is designed to be attached. In this sense, mid-cruise may be envisioned as the point in an aircraft flight cycle at which 50% of the total fuel that is burned between the end of climb and the beginning of descent has been burned (which may be approximated by the midpoint—in terms of time and / or distance—between the end of climb and the beginning of descent).Cruise conditions may therefore define an operating point, a phase or part thereof, of the flight which provides thrust to ensure steady-state operation (i.e., maintaining a constant altitude and / or a constant Mach number), or at least substantially steady-state operation (i.e., maintaining at least a substantially constant altitude and / or at least a substantially constant Mach number) of an aircraft to which it is designed to be attached, taking into account the number of engines with which that aircraft is equipped. For example, where an engine is designed to be attached to an aircraft which has two engines of the same type, at cruise conditions the engine may provide half the total thrust which would be required for steady-state operation, or at least substantially steady-state operation, of that aircraft at mid-cruise.

[0163] In other words, for a given gas turbine engine for an aircraft, cruise conditions can be defined as the operating point of the engine that provides a specified thrust (required to provide - in combination with any other engines on the aircraft - full throttle operation) steady state, or at least substantially steady state operation, of the aircraft to which it is designed to be attached at a given mid-cruise Mach number) at mid-cruise atmospheric conditions (defined by the International Standard Atmosphere according to ISO 2533 at mid-cruise altitude). For any given gas turbine engine for an aircraft, the mid-cruise thrust, atmospheric conditions and Mach number are known, and therefore the operating point of the engine at cruise conditions can be clearly defined.

[0164] Strictly by way of example, the forward speed at the cruise condition may be any point in the range of Mach 0.7 to 0.9, e.g., 0.75 to 0.85, e.g., 0.76 to 0.84, e.g., 0.77 to 0.83, e.g., 0.78 to 0.82, e.g., 0.79 to 0.81, e.g., of the order of Mach 0.8, of the order of Mach 0.85, or in the range of 0.8 to 0.85. Any single speed within these ranges may be part of the cruise condition. For a certain aircraft, the cruise conditions may be outside these ranges, e.g., below Mach 0.7 or above Mach 0.9.

[0165] Strictly by way of example, the cruising conditions may correspond to standard atmospheric conditions (according to the International Standard Atmosphere, ISA) at an altitude which is in the range of 10,000 m to 15,000 m, for example in the range of 10,000 m to 12,000 m, for example in the range of 10,400 m to 11,600 m (approximately 38,000 feet), for example in the range of 10,500 m to 11,500 m, for example in the range of 10,600 m to 11,400 m, for example in the range of 10,700 m (approximately 35,000 feet) to 11,300 m, for example in the range of 10,800 m to 11,200 m, for example in the range from 10,900 m to 11,100 m, for example of the order of 11,000 m. Cruise conditions may correspond to typical atmospheric conditions at any given altitude within these ranges.

[0166] Strictly by way of example, cruise conditions may correspond to a forward Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35,000 feet (10,668 m). At such cruise conditions, the engine may provide a known required net thrust level. The known required net thrust level is, of course, dependent on the engine and its intended application and may be, for example, a value in the range of 20 kN to 40 kN.

[0167] Strictly by way of further example, cruise conditions may correspond to a forward Mach number of 0.85 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38,000 feet (11,582 m). At such cruise conditions, the engine may provide a known required net thrust level. The known required net thrust level is well understood to depend on the engine and its intended application and can be, for example, a value ranging from 35 kN to 65 kN.

[0168] In use, a gas turbine engine described and / or claimed herein may operate at cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions (e.g., mid-cruise conditions) of an aircraft to which at least one (e.g., 2 or 4) gas turbine engines may be mounted to provide propulsive thrust.

[0169] Further, one skilled in the art would readily recognize that either or both of a descent and an approach refer to a phase of operation within an aircraft flight cycle between aircraft cruise and landing, with the approach in particular being part of the landing and takeoff (LTO) phase. During either or both of the descent and the approach, the engine may produce between 0% and 50% of the available thrust. In other non-limiting examples, the engine may produce between 25% and 40% of the available thrust. In other non-limiting examples, the engine may produce between 30% and 35% of the available thrust. Additionally or alternatively, descent may refer to a nominal point in an aircraft flight cycle between takeoff and landing where a relative decrease in altitude is required, and which may necessitate a reduced thrust demand from the engine.

[0170] According to one aspect, there is provided an aircraft comprising a gas turbine engine as described and / or claimed herein. The aircraft according to this aspect is the aircraft to which the gas turbine engine has been designed to be attached. Accordingly, the cruise conditions according to this aspect may correspond to an operating point, a phase, or a portion thereof, of the flight of the aircraft, as defined elsewhere herein.

[0171] According to one aspect, there is provided a method of operating a gas turbine engine as described and / or claimed herein. The operation may be at any suitable condition, which may be as defined elsewhere herein (e.g., in terms of thrust, atmospheric conditions, and Mach number).

[0172] According to one aspect, there is provided a method of operating an aircraft comprising a gas turbine engine as described and / or claimed herein. Operation according to this aspect may comprise (or may be) operation at any suitable condition, for example at mid-cruise of the aircraft, as defined elsewhere herein.

[0173] Those skilled in the art would understand that, unless mutually exclusive, a feature or parameter described in connection with any of the above aspects may be applied to any other aspect. Furthermore, unless mutually exclusive, any feature or parameter contained or described herein may be applied to any appearance and / or combined with any other characteristic or parameter described herein.

[0174] Furthermore, unless mutually exclusive, any parameter or value contained or described herein may be applied to and / or combined with any one or more other parameters and / or values ​​described herein. For example, a first parameter or value contained or described herein (e.g., parameter A) may be applied to and / or combined with any one or more other parameters and / or values ​​contained or described herein (e.g., one or more parameters B; one parameter C; and one parameter D, and so on) to express a product of their relationship. For example, those skilled in the art will understand that when parameter A is described separately from parameter B, a product of their relationship may be expressed, for example, as A / B, B / A, B*A, or any other application, combination, or function of parameter A with respect to parameter B, as appropriate.

[0175] Embodiments will now be described by way of example only, with reference to the Figures, in which: [Fig.l] is a sectional side view of a gas turbine engine; [Fig.2] is a close-up sectional side view of an upstream portion of a geared gas turbine engine; [Fig.3] is a partially cutaway view of a reduction gear for a gas turbine engine; [Fig.4] is a close-up sectional side view of an upstream portion of a direct-drive gas turbine engine; [Fig.5] is a representation of an aircraft with a propulsion system comprising two gas turbine engines; [Fig.6] is a schematic representation of a fuel delivery system including a hydraulic fuel actuator; [Fig.7] is a schematic representation of a fuel delivery system comprising two hydraulic fuel actuators; [Fig.8] is a schematic representation of a fuel delivery system comprising a fuel selectively hydraulic actuator; [Fig.9] is a schematic representation of another fuel delivery system comprising a fuel selectively hydraulic actuator; [Fig. 10] is a schematic representation of a fuel delivery system comprising two hydraulic fuel actuators, one of which is a selectively hydraulic fuel actuator; [Fig. 11] is a schematic representation of a fuel delivery system comprising two selectively hydraulic fuel actuators; [Fig. 12] is a schematic representation of a fuel delivery system comprising two selectively hydraulic fuel actuators; [Fig. 13] is a schematic representation of a fuel delivery system including a fuel selectively hydraulic actuator; [Fig. 14] is a schematic representation of a supplemental fuel supply system including a fuel selectively hydraulic actuator; [Fig. 15] is a schematic representation of a fuel delivery system including a fuel selectively hydraulic actuator operative to control a valve in an engine heat management system; [Fig. 16] is a schematic representation of a supplemental fuel delivery system including a fuel selectively hydraulic actuator operable to control a valve in an engine heat management system; [Fig. 17] is a schematic representation of a fuel delivery system including a fuel selectively hydraulic actuator operable to control a valve in a generator heat management system; [Fig. 18] is a schematic representation of a supplemental fuel supply system including a fuel-selectively hydraulic actuator operable to control a valve in a generator heat management system; [Fig. 19] is a schematic representation of a fuel delivery system comprising a fuel selectively hydraulic actuator and two heat exchangers; [Fig.20] is a flowchart showing an example of a gas turbine engine operating process; [Fig.21] is a flowchart showing another example of a method of operating a gas turbine engine; [Fig.22] is a flowchart showing another example of a method of operating a gas turbine engine; [Fig.23] is a flowchart showing another example of a method of operating a gas turbine engine; [Fig.24] is a flowchart showing another example of a method of operating a gas turbine engine; and [Fig.25] illustrates an oil supply system of a gas turbine engine comprising two fuel-oil heat exchangers.

[0176] [Fig. 1] illustrates a gas turbine engine 10 having a main axis of rotation 9. The engine 10 includes an air intake 12 and a propulsion fan 23 which generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 comprises a core 11 which receives the core airflow A. The engine core 11 comprises, in axial flow series, a low pressure compressor 14, a high pressure compressor 15, combustion equipment 16, a high pressure turbine 17, a low pressure turbine 19 and a core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass duct 22. The fan 23 is attached to, and driven by, the low pressure turbine 19 via a shaft 26 and an epicyclic reduction gear 30.

[0177] In use, the core air stream A is accelerated and compressed by the low pressure compressor 14 and directed into the high pressure compressor 15 where further compression takes place. The compressed air discharged from the high pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel F and the mixture is combusted. The combustion equipment 16 may be referred to as the combustor 16, with the terms "combustion equipment 16" and "combustion chamber 16" used interchangeably herein. The resulting hot combustion products then expand, and thereby drive, the high pressure and low pressure turbines 17, 19 before being discharged through the nozzle 20 to provide some propulsive thrust. The high pressure turbine 17 drives the high pressure compressor 15 by a suitable interconnecting shaft 27.The blower 23 generally acts to impart increased pressure to the bypass airflow B flowing through the bypass duct 22, such that the bypass airflow B is exhausted through the bypass exhaust nozzle 18 to generally provide the majority of the propulsion thrust. The epicyclic reduction gear 30 is a reduction box.

[0178] An exemplary arrangement for a geared fan gas turbine engine 10 is shown in [Fig. 2]. The low pressure turbine 19 (see [Fig. 1]) drives the shaft 26, which is coupled to a sun wheel, or sun gear, 28 of the epicyclic gear arrangement 30. Radially outward of and meshing with the sun gear 28 are a plurality of planet gears 32 which are coupled together by a planet carrier 34. The planet carrier 34 forces the planet gears 32 to change orientation about the sun gear 28 in synchronism while allowing each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled via links 36 to the fan 23 in order to cause it to rotate about the engine axis 9. Radially outwardly of the planet gears 32 and meshing therewith there is a ring or toothed crown 38 which is coupled, via links 40, to a stationary support structure 24.

[0179] It should be noted that the terms "low pressure turbine" and "low pressure compressor" as used herein may be taken to indicate the lowest pressure turbine stages and the lowest pressure compressor stages (i.e., not including the fan 23) respectively and / or the turbine and compressor stages that are connected together by the interconnecting shaft 26 with the lowest rotational speed in the engine (i.e., not including the reduction gear output shaft that drives the fan 23). In some literature, the "low pressure turbine" and "low pressure compressor" referred to herein may alternatively be known as an "intermediate pressure turbine" and "intermediate pressure compressor".When such alternative nomenclature is used, the blower 23 may be designated as the first compression stage or the lower pressure compression stage.

[0180] The epicyclic reduction gear 30 is shown by way of example in more detail in [Fig. 3]. Each of the sun gear 28, the planet gears 32 and the ring gear 38 includes teeth around its periphery for meshing with the other gears. However, for clarity, only exemplary portions of the teeth are illustrated in [Fig. 3]. There are four planet gears 32 illustrated, although it will be apparent to the skilled reader that more or fewer planet gears 32 may be provided. Practical applications of an epicyclic planetary reduction gear 30 generally include at least three planet gears 32.

[0181] The epicyclic reduction gear 30 illustrated by way of example in Figures 2 and 3 is of the planetary type, in that the planet carrier 34 is coupled to an output shaft via links 36, with the ring gear 38 fixed. However, any other suitable type of epicyclic reduction gear 30 may be used. As a further example, the epicyclic reduction gear 30 may be a star arrangement, in which the planet carrier 34 is held fixed, with the ring gear 38 allowed to rotate. In such an arrangement, the fan 23 is driven by the ring gear 38. As another alternative example, the reduction gear 30 may be a differential reduction gear in which both the ring gear 38 and the planet carrier 34 are allowed to rotate.

[0182] It will be appreciated that the arrangement shown in Figures 2 and 3 is by way of example only, and that various alternatives are within the scope of this disclosure. Strictly by way of example, any suitable arrangement may be used to position the reducer 30 in the motor 10 and / or to connect the reducer 30 to the motor 10. By way of further example, the connections (such as links 36, 40 in the example of [Fig. 2]) between the reducer 30 and other parts of the motor 10 (such as the input shaft 26, the output shaft and the fixed structure 24) may have any desired degree of rigidity or flexibility. As a further example, any suitable arrangement of bearings between rotating and stationary parts of the motor (e.g. between the input and output shafts from the reducer and fixed structures, such as the reducer housing) may be used, and the description is not limited to the exemplary arrangement of [Fig. 2]. For example, where the reducer 30 has a star arrangement (described above), one skilled in the art would readily understand that the arrangement of the output and support links and bearing locations would typically be different from that shown as an example in [Fig. 2].

[0183] Thus, the present disclosure extends to a gas turbine engine having any arrangement of reduction gear styles (e.g., star or planetary), support structures, input and output shaft arrangements, and bearing locations.

[0184] Optionally, the reducer may drive additional and / or alternative components (e.g. the intermediate pressure compressor and / or a booster).

[0185] Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have another number of compressors and / or turbines and / or another number of interconnecting shafts. As a further example, the gas turbine engine shown in [Fig.l] has a split-flow nozzle 18, 20 which means that the flow through the bypass duct 22 has its own nozzle 18 which is independent of, and radially outward of, the core engine nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also be applied to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed-flow nozzle.The nozzle(s) (whether mixed or split flow) may have a fixed or variable area.

[0186] As a further example, other gas turbine engines to which the present disclosure may be applied may not have a reduction gear for the main shaft(s), being instead direct drive engines. A sectional view of such an engine is shown in [Fig.4].

[0187] With reference to [Fig.4], a gas turbine engine is generally indicated at 10, having a main rotational axis 9. The engine 10 comprises, in axial flow series, an air intake 12, a propulsion fan 23, an intermediate pressure compressor 14, a high pressure compressor 15, combustion equipment 16, a high pressure turbine 17, an intermediate pressure turbine 19a, a low-pressure turbine 19 and an exhaust nozzle 20. A nacelle 21 surrounds the engine 10 and defines both the intake 12 and the exhaust nozzle 20.

[0188] In use, air entering the intake 12 is accelerated by the fan 23 to produce two air streams: a core air stream A and a bypass air stream B. The core air stream A flows into the intermediate-pressure compressor 14, and the bypass air stream B passes through a bypass duct 22 to provide propulsive thrust. The intermediate-pressure compressor 14 compresses the air stream A before delivering this air to the high-pressure compressor 15 where further compression takes place.

[0189] Compressed air discharged from the high pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel F and the mixture is combusted. The combustion equipment 16 may be referred to as the combustion chamber 16, with the terms "combustion equipment 16" and "combustion chamber 16" used interchangeably herein. The resulting hot combustion products then expand, and thereby drive, the high pressure, intermediate pressure and low pressure turbines 17, 19a, 19 before being discharged through the nozzle 20 to provide additional propulsive thrust. The high pressure 17, intermediate pressure 19a and low pressure 19 turbines respectively drive the high pressure compressor 15, the intermediate pressure compressor 14 and the fan 23, each by a suitable interconnecting shaft.

[0190] Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of interconnecting shafts (e.g., two) and / or an alternative number of compressors and / or turbines. Further, the engine may include a reduction gear provided in the drive train from a turbine to a compressor and / or a fan.

[0191] While the described example relates to a bypass engine, the description may apply, for example, to any type of gas turbine engine, such as an open rotor (in which the fan stage is not surrounded by a nacelle) or a turboprop, for example. In some arrangements, the gas turbine engine 10 may not include a reduction gear 30.

[0192] The geometry of the gas turbine engine 10, and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned with the rotation axis 9), a radial direction (in the bottom-up direction in [Fig.l]) and a circumferential direction (perpendicular to the page in the view of [Fig.l]). The axial, radial and circumferential directions are mutually perpendicular.

[0193] The fuel F supplied to the combustion equipment 16 may comprise a fossil-based hydrocarbon fuel, such as kerosene. Thus, the fuel F may comprise molecules from one or more of the chemical families of n-alkanes, isoalkanes, cycloalkanes, and aromatics. Additionally or alternatively, when blended with, mixed with, or substituted for an alternative fuel, the fuel F may comprise renewable hydrocarbons produced from biological or non-biological resources, otherwise known as sustainable aviation fuel (SAF). In each of the examples provided, the fuel F may comprise one or more trace elements including, for example, sulfur, nitrogen, oxygen, inorganics, and metals.

[0194] By SAF, the skilled person means, for example, a biofuel, a renewable aviation fuel, a renewable jet fuel, an alternative fuel or biojet fuel, produced from biological or non-biological resources. SAFs are commonly synthesized from carbon-containing gases extracted from the atmosphere and / or captured during industrial processes; or from a wide range of sustainable feedstocks such as, for example, used oils and greases; municipal solid waste; cellulosic waste (such as corn stover); cover crops such as camelina, carinata and maidenhair fern; non-biogenic alternative fuels; jatropha; halophytes and algae, rather than fossil-based hydrocarbons. SAFs do not encompass fossil fuels.

[0195] The functional performance of a given fuel composition, or fuel mixture F for use in a given mission, may be defined, at least in part, by the fuel's ability to process the Brayton cycle of the gas turbine engine 10. Parameters defining functional performance may include, for example, specific energy; energy density; thermal stability; and, emissions including gaseous and / or particulate matter. In this regard, particulate matter emissions may include soot particles created by the combustion of said fuel F, also known as non-volatile particulate matter (nvPM). Any reference to soot or smoke herein may also apply to other types of particulate matter emissions known in the art.The gaseous emissions may include one or more of nitrogen oxides (NOx); carbon monoxide (CO); carbon dioxide (CO2); unburned hydrocarbons (UHC); sulfur oxides (SO) including, for example, sulfur dioxide (SO2) and / or sulfur trioxide (SO3); and, volatile organic compounds (VOCs) created by the combustion of said fuel F. Any reference to gaseous emissions herein may also apply to other types of gaseous emissions known in the art.

[0196] A relatively higher specific energy (i.e., energy per unit mass), expressed in MJ / kg, may at least partially reduce takeoff weight, potentially providing a relative improvement in fuel efficiency. A relatively higher energy density (i.e., energy per unit volume), expressed in MJ / L, may at least partially reduce fuel takeoff volume, which may be particularly important for volume-limited missions or military operations involving refueling. A relatively higher thermal stability (i.e., inhibiting fuel degradation or coking under thermal stress) may allow the fuel to experience elevated temperatures in the engine and fuel injectors, potentially providing relative improvements in combustion efficiency.Reduced emissions, including particulate matter, can allow for reduced contrail formation, while reducing the environmental impact of a given mission. Other fuel properties may also be critical to functional performance. For example, a relatively lower freezing point (°C) may allow long-range missions to optimize flight profiles; minimum aromatic concentrations (%) may ensure sufficient swelling of certain materials used in the construction of O-rings and seals previously exposed to high aromatic fuels; and, maximum surface tension (mN / m) may ensure sufficient spray breakup and atomization of the fuel.

[0197] The ratio of the number of hydrogen atoms to the number of carbon atoms in a molecule can influence the specific energy of a given composition, or mixture of fuels. Fuels having higher ratios of hydrogen atoms to carbon atoms may have higher specific energies in the absence of bonding strain. For example, fossil-based hydrocarbon fuels may comprise molecules having about 7 to 18 carbons, with a significant portion of a given composition derived from molecules of 9 to 15 carbons, with an average of 12 carbons.

[0198] A number of sustainable aviation fuel blends have been approved for use. For example, some approved blends include blend ratios of up to 10% sustainable aviation fuel, while other approved blends include blend ratios of between 10% and 50% sustainable aviation fuel (the remainder comprising one or more fossil-based hydrocarbon fuels, such as kerosene), with additional compositions pending approval. However, there is an anticipation in the aviation industry that sustainable aviation fuel blends, comprising up to (and including) 100% sustainable aviation fuel (SAF), will eventually be approved for use.

[0199] Sustainable aviation fuels may comprise one or more of n-alkanes, iso-alkanes, cycloalkanes, and aromatics, and may be produced, for example, from one or more of synthesis gas (syngas); lipids (e.g., fats, oils, and greases); sugars; and alcohols. Thus, sustainable aviation fuels may comprise lower aromatic and / or sulfur contents, relative to fossil-based hydrocarbon fuels. Additionally or alternatively, sustainable aviation fuels may comprise one and / or the other of higher iso-alkane content and cycloalkane content, relative to fossil-based hydrocarbon fuels. In some examples, sustainable aviation fuels may include a density between 90% and 98% of that of kerosene and / or a heating value between 101% and 105% of that of kerosene.

[0200] In some examples, the sustainable aviation fuel(s), or blend(s) supplied to the combustion equipment 16 may have a relatively lower aromatics and / or other non-paraffinic content than kerosene. The sustainable aviation fuel may have an aromatics content of 30%, 20%, 15%, 10%, 8%, 5%, or less than 5%, for example; for example, 4%, 3%, 2%, 1%, or less than 1%; for example, 0.75%, 0.5%, 0.25%, or less than 0.25%; for example, 0.2%, 0.1%, or less than 0.1%; for example, 0.01%, 0.001%, or 0%.The aromatics content of the sustainable aviation fuel may be within an inclusive figure or range bounded by one or more of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, 13.5%, 8.5%, 2.5%, 0.35%, 0.15%, 0.05%, 0.005%, or 0%; or from 0% to 0.75%, from 0% to 0.5%, or from 0.1% to 0.25%; or from 0.15% to 0.65%, 0.35% to 0.55%, or 0.035% to 0.055%; depending on one or more of a preference, a fuel stock or supplier, and compositional variations therein.

[0201] Due at least in part to the molecular structure of sustainable aviation fuels, sustainable aviation fuels may provide beneficial effects including, for example, one or more of higher specific energy (despite, in some examples, lower energy density); higher specific heat capacity; higher thermal stability; higher lubricity; lower viscosity; lower surface tension; lower freezing point; lower soot emissions; lower NOx; and, lower CO2 emissions, compared to fossil-based hydrocarbon fuels (e.g., when combusted in aircraft equipment). combustion 16). Thus, compared to fossil-based hydrocarbon fuels, such as kerosene, sustainable aviation fuels can lead to either or both of a relative decrease in specific fuel consumption and a relative decrease in maintenance costs.

[0202] As shown in [Fig. 5], an aircraft 1 may include a plurality of fuel tanks 50, 53a, b; for example, a larger primary fuel tank 50 located in the aircraft fuselage, and a smaller fuel tank 53a, 53b located in each wing. In other examples, an aircraft 1 may have only one fuel tank 50, and / or the wing fuel tanks 53a, b may be larger than the center fuel tank 50, or no center fuel tank 50 may be provided (all fuel being stored instead in the wings of the aircraft) - it will be appreciated that many different tank arrangements are contemplated and that the examples that have been illustrated are provided for ease of description and not intended to be limiting.

[0203] [Fig. 5] shows an aircraft 1 with a propulsion system 2 comprising two gas turbine engines 10. The gas turbine engines 10 are supplied with fuel from a fuel supply system on board the aircraft 1. The fuel supply system of the illustrated example comprises a single fuel source. For the purposes of the present application, the term "fuel source" means either 1) a single fuel tank or 2) a plurality of fuel tanks that are fluidically interconnected. Each fuel source is configured to provide a separate source of fuel (i.e., a first fuel source may contain a first fuel having one or more characteristics different from a second fuel contained in a second fuel source).The first and second fuel sources are therefore not fluidically coupled to each other so as to separate the different fuels (at least under normal operating conditions). The use of several fuel sources allows an aircraft 1 to carry several different fuels, and to change the fuel used during operation, and possibly even during cruise or when changing from one operating stage to another during flight.

[0204] In the present examples, the first (and, in these examples, sole) fuel source comprises a center fuel tank 50, located primarily in the fuselage of the aircraft 1 and a plurality of wing fuel tanks 53a, 53b, where at least one wing fuel tank is located in the port wing and at least one wing fuel tank is located in the starboard wing for balancing. All of the tanks 50, 53a, b are fluidically connected in the example shown, thereby forming a single fuel source. Each of the center fuel tank 50 and the Wing fuel tanks 53a, b may include a plurality of fluidically interconnected fuel tanks.

[0205] In another example, the wing fuel tanks 53a, 53b may not be fluidly connected to the center tank 50, thereby forming a second, separate fuel source. For balancing purposes, one or more port wing fuel tanks may be fluidly connected to one or more starboard wing fuel tanks. This may be done either via a center fuel tank (if that tank is not part of the other fuel source), or by bypassing the center fuel tank(s), or both (for maximum flexibility and safety). In another example, the first fuel source includes wing fuel tanks 53 and a center fuel tank 50, while a second fuel source includes another separate center fuel tank.The fluidic interconnection between the wing fuel tanks and the center fuel tank of the first fuel source may be provided for balancing the aircraft 1. In aircraft 1 with multiple fuel sources, two or more of the fuel sources may therefore contain different fuels from each other, such that the aircraft 1 may change fuels during flight. It may therefore be more complex to determine which fuel is supplied to the combustion chamber 16 than simply recording a single identity of a fuel on board the aircraft 1, or performing a check at the time of start-up.

[0206] In some examples, the distribution of fuel tanks 50, 53 available on the aircraft 1 may be limited such that the first fuel source and the second fuel source are each substantially symmetrical about the centerline of the aircraft. In cases where an asymmetrical distribution of fuel tanks is permitted, a suitable means for transferring fuel is generally provided between the fuel tanks of the first fuel source and / or between the fuel tanks of the second fuel source such that the position of the center of mass of the aircraft can be maintained within acceptable lateral limits throughout the duration of the flight.

[0207] Aircraft typically refuel at several different airports, for example at the beginning and end of a long-haul flight. While there are standards that all aviation fuels must comply with, as discussed above, different aviation fuels have different compositions, for example depending on their source (e.g., different sources of petroleum, biofuels, or other synthetic aviation fuels (often described as sustainable aviation fuels, SAFs), and / or blends of petroleum-based fuels, and other fuels) and any additives included (e.g., such as antioxidants and ratio deactivators, biocides, static reducers, icing inhibitors, corrosion inhibitors) and any impurities. The composition of the available aviation fuel may not only vary from airport to airport and fuel supplier to fuel supplier, but also from batch to batch, even for a given airport or fuel supplier. Furthermore, the fuel tanks 50, 53 of the aircraft 1 are usually not emptied before being refilled for a subsequent flight, resulting in mixtures of different fuels in the tanks, with fuel of different composition actually being obtained from the mixture. One or more fuel ports 62 may be provided for refueling.

[0208] [Fig. 6] is a schematic representation of a portion of a fuel delivery system 152 of a gas turbine engine 10. The fuel delivery system 152 includes a main fuel flow path 110 from the fuel tank 50 to the combustion chamber 16 of the gas turbine engine 10, and various branch fuel paths that take or return fuel to / from this main fuel flow path 110.

[0209] The main fuel flow path 110 of the fuel system 152 serves as a fuel source for the fuel hydraulic systems. The gas turbine engine 10 includes a combustion chamber 16 and an actuator 154. The dotted arrows in [Fig. 6] represent the direction of fuel flow. One or more fuel lines 159, 160 are provided to supply fuel to an actuator 154, to hydraulically actuate it by fuel.

[0210] The fuel supply system 152 is configured to supply fuel for combustion in the combustion chamber 16.

[0211] It will be appreciated that one or more fuel valves, pumps, sensors, taps for fuel hydraulic actuators, and heat exchangers may be present along the main fuel flow path 110, among other engine components. Fuel is supplied to the combustion chamber 16 directly via the line 110a, after passing through the most downstream engine component along the main fuel flow path 110.

[0212] Fuel delivered through lines 159 and 160 is used to drive (i.e., actuate) actuator 154. As such, actuator 154 is fuel hydraulically driven. The term "fuel hydraulic" as used herein refers to the hydraulic operation of an actuator whose hydraulic fluid is a fuel. Actuators with fuel hydraulic drive may be referred to herein as "fuel hydraulic actuators."

[0213] In the example illustrated in [Fig.6], fuel is delivered to the actuator 154 from the fuel flow path 110 via the line 159. After being used for the hydraulic fuel drive of the actuator 154, the fuel is returned to the main fuel flow path 110, via the line 160, and can then be delivered to the combustion chamber 16.

[0214] The fuel return line 160 may include a valve configured to regulate the return flow of fuel to the main fuel path 110, and the valve itself may be operated using a hydraulic fuel actuator. The fuel return line 160 may direct the fuel to a location along the main fuel flow path 110 that is before or after one or more heat exchangers of the main fuel flow path 110.

[0215] The fuel return line 160 may return fuel to the fuel tank. The valve may be a fuel tank return valve in such implementations. The fuel tank return valve may be controlled using a fuel tank return actuator. The fuel delivery system may be configured to supply fuel to the fuel tank return actuator to hydraulically drive it with fuel.

[0216] Once the fuel reaches the combustion chamber 16, it is burned to provide thrust, as explained in connection with [Fig.l].

[0217] [Fig. 6] shows a single fuel hydraulic actuator 154 on the fuel flow loop 159, 160. In other embodiments, multiple fuel hydraulic actuators 154 may be located on a single fuel flow loop 159, 160, utilizing fuel taken from a single point along the main fuel flow path 110. The multiple actuators may be arranged in series with respect to the fuel flow (such that all fuel in the sample pipe 159 passes through all of the actuators, in turn) or in parallel with respect to the fuel flow (with the sample pipe 159 branching and one or more actuators on each branch).

[0218] [Fig. 7] is a schematic representation of another portion of a fuel delivery system 152 of a gas turbine engine 10. The fuel delivery system 152 includes a fuel source 50, 53, which may be a single fuel tank or multiple fuel tanks, and a fuel flow path 110 between the fuel source 50, 53 and the combustor 16. The illustrated fuel delivery system 152 includes two actuators 254a, b, each on a fuel evacuation loop separate from the main fuel flow path 110. The dotted arrows in [Fig. 7] represent the direction of fuel flow.

[0219] The two actuators 254a, b are hydraulically driven by fuel. That is, the two actuators 254a, b are hydraulically driven using fuel from the source 50, 53 as hydraulic fluid. The fuel is supplied from the main path 110 to the actuators 254a, b respectively via the lines 259a, b. Once the fuel has been used to drive the actuators 254a, b, it returns to the main fuel flow path 110 via the lines 260a, b.

[0220] Only two actuators 254a, b are shown in [Fig. 7] but, in other implementations, any plurality of actuators may be fuel hydraulically driven. In some embodiments, the gas turbine engine 10 may include a plurality of fuel hydraulically driven actuators and a plurality of actuators that are not fuel hydraulically driven.

[0221] [Fig. 7] shows a single fuel hydraulic actuator 254 on each fuel flow loop. In other embodiments, multiple fuel hydraulic actuators 254 may be located on one or more of the multiple fuel flow loops. The multiple actuators may be arranged in series with respect to the fuel flow (such that all of the fuel in the pickup pipe 159 passes through all of the actuators, in turn) or in parallel with respect to the fuel flow (with the pickup pipe 159 branching and one or more actuators located on each branch).

[0222] In various embodiments, the gas turbine engine 10 includes at least ten, eleven, twelve, thirteen, fourteen, fifteen, sixteen, seventeen, or eighteen fuel hydraulic actuators. In some embodiments, each actuator may have its own fuel bleed pipe 159, 259 from the main fuel flow path 110, such that it has its own fuel flow loop. In other arrangements, two or more of these actuators may be located on the same fuel flow loop off the main fuel flow path 110. In some arrangements, most or all of the fuel hydraulic actuators may be located on the same fuel flow loop off the main fuel flow path 110.

[0223] In various embodiments, the gas turbine engine 10 includes one or more actuated engine systems. Each actuated engine system may include one or more individual actuators. An "actuated engine system" (also referred to as an "actuated system") refers to a group of actuators used together to control a specific system, or perform a particular function, in the engine 10. An actuated system may be described as "fuel hydraulic" if at least one of its actuators is fuel hydraulic.

[0224] [Fig.8] shows a fuel supply system 152 including an actuator 354.

[0225] [Fig. 8] is a schematic representation of another portion of a fuel delivery system 152 of a gas turbine engine 10. The fuel delivery system 152 includes a fuel source 50, 53, which may be a single fuel tank or multiple fuel tanks, and a fuel flow path 110 between the fuel source 50, 53 and the combustor 16. The illustrated fuel delivery system 152 includes an actuator 354 on a fuel bleed loop from the main fuel flow path 110. The dotted arrows in [Fig. 8] represent the direction of fuel flow.

[0226] The fuel delivery system 152 of [Fig. 8] includes a valve 304. The valve 304 is configured to direct fuel through line 361 or line 360. Line 306 includes the actuator 354, so fuel delivered through line 306 hydraulically drives the actuator 354 and is then returned to the main fuel flow path 110. The fuel delivered through line 361 returns directly to the main fuel flow path 110, without driving an actuator. The valve 304 itself may be controlled using another fuel hydraulic actuator.

[0227] The valve 304 is controllable, meaning that the fuel system 152 is designed to be controlled to choose between: • causing the fuel to actuate the actuator 354 (the fuel flows through the line 360); and • causing the fuel to bypass the at least one actuator 354 (the fuel flows via the line 361). • In some implementations, valve 304 is controlled via an actuator based on the output of a processor or other control device. Valve 304 may be controlled based on the sustainable aviation fuel (SAF) content of the fuel. In some implementations, valve 304 may be replaced with any other suitable solution for regulating the flow of fuel between two lines 360, 361. • The fuel system 152 can be designed to choose between: • causing the fuel to actuate the actuator 354 (via the line 360); and • causing the fuel to bypass the at least one actuator 354 (via the line 361).

[0228] The choice may be made based on whether the sustainable aviation fuel (SAF) content of the fuel exceeds a threshold or not.

[0229] For example, the fuel delivery system 152 may be configured so that the fuel actuates the actuator 354 (via the line 360) when the SAF content of the fuel exceeds 25%, 30%, 35%, 40%, 45%, 50%, 55%, 60%, 65%, 70%, 75%, 80%, 85%, 90%, or 95% concentration by volume.

[0230] The fuel delivery system may control the fuel in a binary manner such that all fuel reaching valve 304 is directed into one of line 360 ​​and line 361. In other implementations, the fuel may be split between line 360 ​​and line 361, for example.

[0231] To determine the SAF content of the fuel, the fuel system 152 may include a sensor. The sensor may be positioned just before the valve 304. In other implementations, the sensor may be positioned upstream of the valve 304 in the main fuel flow path 110. In other arrangements, the SAF content of the fuel is measured, or calculated, or otherwise provided, when the fuel is introduced into the fuel source 50, 53. For example, information about the SAF content of one or more fuels stored in the aircraft 1 may be communicated electronically to, or manually entered into the user interface of, an onboard controller or other computer system. It will be appreciated that any suitable approach known in the art for determining the SAF content of the fuel may be used.

[0232] [Fig. 9] is a schematic representation of another portion of a fuel delivery system 152 of a gas turbine engine 10. The fuel delivery system 152 is similar to the fuel delivery system 152 of [Fig. 8], but the line 361 that returned the fuel to the main fuel flow path 110 is replaced by the line 362 that directs the fuel to the combustion chamber 16. The valve 304 directs the fuel through one of the lines 360 and 362. The line 362 directs the fuel such that it bypasses the actuator 354 and goes to the combustion chamber 16 without returning to the main fuel flow path 110. The line 362 may include one or more pumps or valves configured to regulate the pressure and flow rate of the fuel to the combustion chamber 16.

[0233] In the implementations of Figures 8 and 9, the actuator 354 may be configured such that, when the valve 304 directs fuel to bypass the actuator 354, the actuator 354 is driven in a non-hydraulic fuel manner. For example, actuator 354 may be driven using a non-hydraulic fuel fluid or using electrical signals.

[0234] [Fig. 10] is a schematic representation of another portion of a fuel delivery system 152 of a gas turbine engine 10. The fuel delivery system 152 includes a fuel source 50, 53, which may be a single fuel tank or multiple fuel tanks, and a fuel flow path 110 between the fuel source 50, 53 and the combustor 16. The illustrated fuel delivery system 152 includes two actuators 454a, b, each actuator 454a, b being positioned on a respective fuel bleed loop from the main fuel flow path 110. The dotted arrows in [Fig. 10] represent the direction of fuel flow.

[0235] The first actuator 454a is positioned downstream of a valve 404 located on its respective fuel sampling loop. The valve 404 is configured to direct fuel via line 461 to bypass the actuator (returning directly into the main fuel flow path 110) or line 460a to hydraulically fuel actuate the actuator 454a (and then return to the main fuel flow path 110). The valve 404 is arranged similarly to the valve 304 of [Fig. 8] but, in other implementations, the valve 404 may be arranged similarly to the valve 304 of [Fig. 9] such that fuel bypassing the actuator is instead sent directly to the combustion chamber 16.In still other implementations, all of the fuel taken from the main fuel flow path 110 and sent to the valve 304, 404 may be sent to the combustion chamber 16, either directly or after passing through the actuator 354, 454a. The actuator 454a is thus selectively hydraulically driven by fuel.

[0236] In still other implementations, the selectively fuel-hydraulic-driven valve 354, 454a may not be associated with multiple fuel routes; instead, the valve 304, 404 may be configured to prevent any fuel from leaving the main fuel flow path to that actuator when the actuator 354, 454a is not to be fuel-hydraulic-driven. Flow in a single loop may therefore be started or stopped by the valve 304, 404 rather than using the valve to divert fuel flow between two different branches of the main path. Similarly for implementations in which the same fuel tapping point of the main fuel flow path 110 is used to supply multiple fuel-hydraulic actuators, a controllable valve configured to allow or prevent fuel from reaching a given actuator may either (i) divert fuel intended for that actuator along a dedicated bypass route when that actuator is not to be hydraulically fuel driven, or (ii) simply block any route to that actuator, so that fuel that would otherwise have flowed to that actuator flows to one or more other actuators supplied by the same sampling point. Choosing not to hydraulically fuel drive an actuator may therefore increase the fuel flow to one or more other actuators, unless the total flow in the sampling portion of the fuel lines is also controlled.

[0237] The second actuator 454b of the example of [Fig. 10] is located on a separate fuel flow loop from the first actuator 454a and the fuel is taken from a different point from the main fuel flow path 110. The second actuator 454b is designed to be permanently hydraulically driven by fuel (it has only one mode of actuation).

[0238] Although the implementation of [Fig. 10] shows a selectively driven actuator 454a and a fuel-hydraulic-driven actuator 454b, in other implementations, the fuel delivery system 152 may include any number of actuators, and any of these actuators may be fuel-hydraulic, non-fuel-hydraulic, or selectively fuel-hydraulic. Any number of actuators may be contained within the same fuel flow loop.

[0239] The actuators described above may be any actuator used by the engine 10. Various specific actuators that may be hydraulically driven by fuel are described below. The actuation systems described herein may be any actuation system used by the aircraft. Various specific actuation systems that may be hydraulically driven by fuel are described below. It will be understood that the specific embodiments listed are described by way of example only, and the scope of protection is limited only by the claims.

[0240] An example of an actuated system is a turbine case cooling (TCC) system. Turbine engines 10 typically include a case surrounding the turbine 17, 19 and a TCC system is used to selectively cool this case. This cooling can be used to control rotor-stator tip clearances by reducing thermal expansion. A TCC system includes one or more valves operable to modulate the flow of bleed air from the engine's compressor 17, 19 across the case to provide cooling.

[0241] In engines 10 having multiple compressors 17, 19, a TCC system may be provided for one, some, or all of the turbines.

[0242] The valve(s) of the TCC system(s) may be operated via an actuator. Each valve may have a dedicated actuator, or one actuator may control multiple valves. In various implementations, one or more of the actuators of a TCC system are hydraulically fuel driven. In one or more implementations, the fuel delivery system may be arranged to selectively bypass or hydraulically fuel actuate one or more actuators of the TCC system.

[0243] [Fig. 11] is a schematic representation of a portion of a fuel system 152 of a gas turbine engine 10. The fuel system 152 includes a TCC system 554 including two actuators 554a, b, each actuator 554a, b configured to control a corresponding valve 554a', b' of the TCC system. The valves 255a', b' may provide cold air to cool the turbine housing.

[0244] The fuel delivery system 152 of [Fig. 11] is configured to supply fuel from the fuel source 50, 53 to the combustion chamber 16 and to hydraulically fuel-drive the actuators 554a, b. As described above, the fuel delivery system 152 is configured to selectively actuate or bypass the actuators 554a, b via associated valves 504a, b. In other implementations, a single valve may be used to selectively actuate or bypass the actuators 554a, b.

[0245] The actuators 554a, b are arranged in different fuel flow loops in the implementation illustrated in [Fig. 11]. In other implementations, the actuators 554a, b may be arranged on the same fuel flow loop, and may be arranged in series or in parallel on that loop.

[0246] Another example of an actuated system used in the turbine engines 10 is a bleed air system. Bleed air systems redirect bleed air (e.g., compressed air from the compressor 14, 15, or elsewhere, e.g., from an APU) for use in other systems. Bleed air is useful in various systems because of its relatively high temperature and pressure. Bleed air systems may be referred to as "engine bleed air systems" (EBAS).

[0247] Air bleed systems may be used to control the air conditioning in the cabin of an aircraft 1. In these systems, the bleed air is filtered and cooled (by means of expansion and / or heat exchange, for example, with atmospheric air) using a conditioning unit (conditioning units or pressurization air conditioning kits (PACKS) are two accepted terms). Air conditioning units are used to cool and, if necessary, dehumidify engine or APU bleed air before it is sent to the aircraft cabin. Such systems include one or more valves and one or more actuators. Bleed air is typically introduced into the unit through a one-way check valve. In various implementations, some or more of these actuators may be selectively fuel-hydraulic.

[0248] [Fig. 12] is a schematic representation of another portion of a fuel supply system 152 of a gas turbine engine 10, including an air conditioning unit 650.

[0249] The fuel system 152 of [Fig. 12] includes a cabin air bleed valve 654' and a cabin bleed actuator 654 configured to control the cabin air bleed valve 654'. The cabin air bleed valve 654' controls the flow of bleed air from the compressor 14, 15 through a conditioning pack (as described above) to provide cooled air to the cabin of the aircraft 1.

[0250] The fuel system 152 of [Fig. 12] further includes a bleed air management valve 754' and a handling bleed actuator 754 configured to control the bleed air management valve 754'.

[0251] The bleed air management valve 754' controls the output of bleed air from the compressor 14, 15 and to other aircraft systems. The bleed air management valve 754' may be used for engine start, transient bleed, and / or overpressure control.

[0252] In some arrangements, there are multiple cabin bleed valves 654' and associated cabin bleed actuators 654, all of which may be considered part of the air conditioning group 650. In some arrangements, there are multiple handling bleed valves 754' and associated handling bleed actuators 754, all of which may be considered part of the air conditioning group 650.

[0253] In other arrangements, the cabin air bleed valve 654' and the cabin bleed actuator 654 are omitted. In other arrangements, the air bleed management valve 754' and the handling bleed actuator 754 are omitted. In other arrangements, other known actuators and associated valves may be included and selectively hydraulically actuated by fuel. In some arrangements, other known bleed valves are selectively hydraulically actuated by fuel.

[0254] Similar to the actuators described above, the cabin bleed actuator 654 is selectively hydraulically actuated by fuel to control the cabin air bleed valve 654'. The valve 604 is configured to direct fuel through line 660 to actuate the cabin bleed actuator 654 or to direct fuel through line 661 to bypass the cabin bleed actuator 654.

[0255] Similar to the actuators described above, the handling bleed actuator 754 is selectively hydraulically actuated by fuel to control the air bleed management valve 754'. The valve 704 is configured to direct fuel through line 760 to actuate the handling bleed actuator 754 or to direct fuel through line 761 to bypass the handling bleed actuator 754.

[0256] An example of an actuated system is a variable guide vane (VGV) system 850. A VGV system controls the amount and / or flow path of airflow through the compressor 14, 15 (by moving one or more vanes; and typically by moving multiple stages of compressor vanes) to achieve optimal compressor performance. The VGV system 850 varies the angle of the compressor vanes to manage compressor operability and efficiency.

[0257] The variable guide vanes of various implementations may include variable inlet guide vanes (VIGVs), which are located upstream of / in front of the compressor rotor vanes, and / or other vanes (typically variable stator vanes (VSVs) which are located behind / downstream of the compressor rotor vanes). Other compressor vanes may be located elsewhere: vane locations may vary depending on the engine architecture, as well as between compressors in a given engine. For example, variable / controllable intermediate guide vanes may be located midway through the compressor 14, 15.

[0258] Generally, the variable guide vanes are designed to be rotatable about an axis to align with the angle of the incoming airflow. This movement of the vanes allows for a greater range of non-dimensional flow through the compressor 14, 15 without stall occurring; in effect, the compressor 14, 15 can operate smoothly over a greater range of engine power settings due to the VGVs. The VGVs are controlled by the engine control system in response to a range of parameters, generally including ambient conditions and the level of thrust demanded by the engine.

[0259] In engines 10 having multiple compressors 14, 15, a VGV system 850 may be provided for one, more, or all of the compressors 14, 15.

[0260] Many VGV systems 850 include two VGV actuators, which may be used to change the angle of the compressor inlet guide vanes 14, 15. Each of the two VGV actuators may be configured to control the guide vanes on a given side of the engine 10. In various implementations, one or more of the actuators of a VGV system are hydraulically driven by fuel.

[0261] [Fig. 13] is a schematic representation of another portion of a fuel delivery system 152 of a gas turbine engine 10 including a VGV system 850. The fuel delivery system 152 includes a variable stator vane actuator 854. The fuel delivery system 152 includes a valve 804 configured to selectively hydraulically fuel actuate the variable stator vane actuator 854 by routing fuel via line 860 or line 861. The valve 804 may be configured to cause fuel to actuate or bypass the variable stator vane actuator 854 depending on the S AF content of the fuel. Although the arrangement of [Fig.13] shows a single variable stator vane actuator, in other arrangements, the fuel delivery system 152 may be arranged to selectively hydraulically actuate by fuel two or more variable stator vane actuators.

[0262] [Fig. 14] shows a diagram representing another portion of the fuel system 152 of a gas turbine engine 10 including a VGV system 850, similar to that of [Fig. 13], but with a single, simple fuel bleed loop, without an equivalent of line 861. The fuel system 152 includes a variable inlet guide vane actuator 854a. The fuel system 152 includes a valve 904 configured to selectively hydraulically actuate the variable inlet guide vane actuator 854a by fueling fuel through line 960, or preventing fuel flow along line 960.Valve 904 may be configured to cause fuel to actuate variable inlet guide vane actuator 854a, or to continue along main fuel flow path 110, thereby bypassing actuator 854a, depending on the SAF content of the fuel.

[0263] The guide vanes control the flow of air into and through the compressor 14, 15. The variable guide vane actuator(s) 854, 854a control the orientation angle of the guide vanes.

[0264] Another example of an actuated system used in turbine engines 10 is a heat management system (HMS). Turbine engines 10 often include an engine HMS and a generator HMS. The two HMS systems control the oil temperature by enabling and controlling the use (or bypass) of air-to-oil (and / or fuel-to-oil) heat exchangers. Oil-to-oil heat exchangers may also be used in some implementations, as described in more detail below with respect to [Fig. 25]. For example, oil-to-oil heat exchangers may be used to enable heat transfer between the oil in a generator HMS and that in an engine HMS.

[0265] The engine HMS system controls the temperature of the oil system that lubricates the engine components (e.g., bearing chambers, reduction gear, etc.). The engine HMS system includes one or more engine HMS valves that are controlled by one or more actuators. In various implementations, one or more of these actuators are hydraulically driven by fuel.

[0266] For example, in some implementations, the engine HMS includes a single valve and a single fuel-driven hydraulically driven actuator for controlling the valve.

[0267] The engine HMS system may include a plurality of heat exchangers arranged in a parallel or series configuration with respect to the fluid flow (e.g., fuel flow or oil flow). A single valve controlled by a single fuel hydraulic actuator may allow the use (or bypass) of all of the heat exchangers (in particular, the flow of a fluid (air, oil, or fuel, if applicable) may be controlled by the valve, with the proportion of fluid not passing through the heat exchanger being directed along a bypass pipe). In some implementations, each heat exchanger is associated with a valve controlled by an associated actuator. In various implementations, one or more of these actuators are fuel hydraulically driven.Additionally, one or more recirculation pipes may be provided to allow fluid from an outlet of the heat exchanger to be returned to the inlet of the heat exchanger, thereby passing through the heat exchanger several times. The flow rate in the recirculation pipe(s) may again be adjusted using an actuator-controlled valve. For any of the recirculation pipes described herein, there may be one or more associated pumps designed to return air / oil / fuel to the heat exchanger. Alternatively, or in addition, any suitable components for repressurizing the air / oil / fuel to enable recirculation may be used.

[0268] [Fig. 15] is a schematic representation of an engine heat management system 1000. The engine heat management system 1000 includes a Air-to-oil heat exchanger 1003 configured to remove heat from the engine oil. The air-to-oil heat exchanger 1003 includes an oil line 1001 and an air line 1002. In [Fig. 15], the heat exchanger 1003 is a parallel flow heat exchanger, with the oil and air flows in the same direction. In other implementations, the heat exchanger 1003 may be a counterflow heat exchanger, meaning that the direction of flow of the air line or the oil line is reversed such that they flow through the heat exchanger 1003 in opposite directions.

[0269] In the arrangement of [Fig.15], the oil line 1001 includes a valve 1004 operable to allow oil to bypass the air-oil heat exchanger 1003. The valve 1004 is controlled via an actuator 1054. The valve 1004 may be binary (on / off) or may be designed to allow finer control of the oil flow rate through the heat exchanger 1003, with a variable portion of the oil bypassing the heat exchanger 1003.

[0270] Actuator 1054 may be selectively fuel hydraulic, as described in connection with the above arrangements. A valve 1006 is operative to allow fuel to flow through actuator 1054 to fuel hydraulically actuate actuator 1054, or to allow fuel to bypass actuator 1054. In some arrangements, when valve 1006 causes fuel to bypass actuator 1054, actuator 1054 may be actuated in a non-fuel hydraulic manner.

[0271] Valve 1006 may be operable to only permit fuel hydraulic actuation of actuator 1054 when the SAF content of the fuel is above a predetermined threshold.

[0272] [Fig. 16] shows another arrangement of a heat management system 1000 in which the valve 1004 is positioned in the air line 1002. The engine heat management system 1000 is similar to the heat management system 1000 of [Fig. 15], but the valve 1004 is operative to allow air to bypass the air-oil heat exchanger 1003, thereby controlling the air flow therethrough. Again, the fluid flows may be arranged differently in other implementations.

[0273] In other implementations, the engine HMS systems 1000 of Figures 15 and 16 may be combined such that the HMS system includes a valve for controlling both oil flow and air flow. In some implementations, the HMS system 1000 may include multiple heat exchangers and a valve configured to allow simultaneous bypass of all heat exchangers, as well as valves for bypassing individual heat exchangers.

[0274] The generator HMS may be independent of the engine HMS 1000. It provides cooling for an electrical machine used, for example, to power the airframe with electricity to operate the aircraft systems.

[0275] [Fig. 17] is a schematic representation of a generator heat management system 1100. The generator heat management system 1100 includes an air-to-oil heat exchanger 1103 configured to remove heat from the engine oil, and more specifically to remove heat from the oil used to cool (and possibly also to lubricate) the generator. The air-to-oil heat exchanger 1103 includes an oil line 1101 and an air line 1102.

[0276] In the arrangement of [Fig.17], the oil line 1101 includes a valve 1104 operable to allow a controllable amount of oil to bypass the air-oil heat exchanger 1103. The valve 1104 is controlled via an actuator 1154.

[0277] Actuator 1154 may be selectively fuel hydraulic, as described in connection with the arrangements above. A valve 1106 is operative to allow fuel to flow through actuator 1154 to fuel hydraulically actuate the actuator, or to allow fuel to bypass the actuator. In some arrangements, actuator 1154 may be configured to be fuel non-hydraulic actuated when valve 1106 causes fuel to bypass actuator 1154.

[0278] Valve 1106 may be operable to only permit fuel hydraulic actuation of actuator 1154 when the SAF content of the fuel is above a predetermined threshold.

[0279] [Fig. 18] shows another arrangement of a generator heat management system 1100 in which the valve 1104 is positioned in the air line 1102. The generator heat management system 1100 is similar to the generator heat management system 1100 of [Fig. 17], but the valve 1104 is operative to allow air to bypass the air-to-oil heat exchanger 1103. In other implementations, valves may be provided to control fluid flow on both the air side and the oil side.

[0280] [Fig. 19] is a schematic representation of a fuel delivery system 152 including two fuel-oil heat exchangers 1220, 1230. The fuel delivery system 152 includes a primary fuel-oil heat exchanger 1220 and a secondary fuel-oil heat exchanger 1230. The heat exchangers 1220, 1230 are used to move heat from the fuel away from the oil, and can be considered part of the engine heat management system. The primary fuel-oil heat exchanger 1220 is configured to heat at least the majority of the fuel, and a heat exchanger secondary fuel-oil 1230 is configured to provide supplemental heat to the fuel to be supplied to hydraulically drive at least one fuel hydraulic actuator 1254.

[0281] The fuel supply system 152 of the implementation illustrated in [Fig. 19] comprises five fuel valves 1201 to 1205. Each valve 1201 to 1205 is controlled by an actuator, which may be fuel hydraulic.

[0282] The first valve 1201 is a fuel valve configured to allow fuel to bypass the primary heat exchanger 1220. The valve 1201 may be operable to allow any percentage of fuel to bypass, or pass through, the primary heat exchanger 1220. The diverted fuel is then returned to the main fuel flow path 110.

[0283] The primary heat exchanger 1220 allows the temperature of the fuel to be increased before it enters the combustion chamber 16 while reducing the temperature of the oil.

[0284] The second valve 1202 is configured to direct fuel from the outlet of the primary heat exchanger 1220: (i) through a recirculation pipe 1206 configured to return fuel through the primary heat exchanger 1220, (ii) to the combustion chamber 16, or (iii) to the secondary heat exchanger 1230, for use in one or more fuel hydraulic systems. In other cases, multiple separate valves may be provided for recirculation or redirection / bleed-off to a fuel hydraulic system.

[0285] The third valve 1203 is part of the fuel hydraulic system and is a fuel valve designed to allow a controllable portion of the fuel to bypass the secondary heat exchanger 1230. The valve 1203 may be operable to allow any percentage of fuel to bypass, or pass through, the secondary heat exchanger 1230.

[0286] The fourth valve 1204 is part of the fuel hydraulic system and is configured to direct fuel from downstream of the secondary heat exchanger 1230: (i) to a fuel hydraulic actuator 1254, or (ii) through a recirculation pipe 1207 configured to return fuel through the secondary heat exchanger 1230, such that some or all of the fuel directed to the secondary heat exchanger 1230 may pass through that heat exchanger multiple times. A fuel recirculation valve 1204 may be provided separately from a valve configured to control flow to the fuel hydraulic actuator 1254 in some implementations.

[0287] In some arrangements, one or both of the recirculation pipes 1206, 1207 include, or are associated with, one or more pumps configured to transporting the fuel through the recirculation pipe 1206, 1207. In some arrangements, the pump(s) are positioned upstream of the valve(s) 1202, 1204. In other arrangements, the pump(s) are positioned anywhere along the recirculation pipes 1206, 1207.

[0288] The fifth valve 1205 is part of the fuel hydraulic system and is configured to direct fuel to hydraulically fuel, or bypass, the actuator 1254. In the illustrated example, the fuel that passes through or bypasses the actuator 1254 is then sent directly to the combustion chamber 16. In other implementations, it may be returned to the main fuel flow path 110.

[0289] In some arrangements, one or both of the recirculation valves 1202, 1204 may be omitted. In some arrangements, one or both of the bypass valves 1201, 1203 may be omitted.

[0290] In some arrangements, the heat exchangers 1220, 1230 are controlled such that, under cruising conditions, a heat transfer ratio of:

[0291] [Math 5] oil-to-fuel heat transfer rate in the secondary fuel-to-oil heat exchanger (kJkg1) oil-to-fuel heat transfer rate in the primary and secondary fuel-to-oil heat exchangers (kJkg1) has a maximum (i.e., peak) value of at least 0.35.

[0292] In other arrangements, the heat transfer ratio has a maximum value at cruising speed greater than 0.4, 0.45, 0.5, 0.55 or 0.6.

[0293] In some arrangements, the fuel delivery system 152 is configured to control the heat transfer ratio by controlling the percentages of fuel that bypass, or are recirculated through, each of the heat exchangers 1220, 1230.

[0294] In additional or alternative arrangements, the heat transfer ratio is controlled by controlling the percentages of oil that bypass, or are recirculated through, each of the heat exchangers 1220, 1230. Oil bypass and recirculation valves and / or pipes may be provided accordingly. One or more air-to-oil or oil-to-oil heat exchangers may also be controlled to influence the heat transfer ratio.

[0295] Such an exemplary oil supply system 153 is illustrated in [Fig. 25]. In the implementation of [Fig. 25], each fuel-oil heat exchanger 1220, 1230 is on a separate oil loop; the two oil loops are fluidically isolated from each other and brought into thermal contact by means of an oil-oil heat exchanger 1240 (although not illustrated, it will be appreciated that one or more oil bypass and / or recirculation valves can also be supplied in conjunction with the 1240 oil-oil heat exchanger).

[0296] The oil loop containing oil configured to pass through the primary fuel-oil heat exchanger 1220 is configured to cool, and optionally lubricate, a first set of engine components 10'. The oil loop containing oil configured to pass through the secondary fuel-oil heat exchanger 1230 is configured to cool, and optionally lubricate, a second set of engine components 10”.

[0297] Each fuel-oil heat exchanger 1220, 1230 is associated with an oil bypass valve 1212, 1213 configured to allow oil to bypass the respective heat exchanger 1220, 1230. The oil bypass valve 1212, 1213 is configured to allow a controllable portion of the oil to bypass the respective heat exchanger 1220, 1230 in various implementations.

[0298] Each fuel-oil heat exchanger 1220, 1230 is associated with an oil recirculation valve 1211, 1214 configured to allow the oil to pass through the respective heat exchanger 1220, 1230 several times before returning to the engine components 10', 10” that the oil is intended to cool. The oil recirculation valve 1211, 1214 is configured to allow a controllable portion of the oil to be recirculated through the respective heat exchanger 1220, 1230 in various implementations. In the implementation of [Fig. 25], the recirculated oil is shown as being introduced into the respective bypass valve 1212, 1213, however it will be appreciated that the recirculated oil may rejoin the oil flow before or after this valve in other implementations.

[0299] In some implementations, only one of an oil recirculation valve 1211, 1214 and an oil bypass valve 1212, 1213 may be provided for one or each heat exchanger 1220, 1230, or none may be provided for one or each heat exchanger 1220, 1230.

[0300] Although no air-oil heat exchanger is shown in [Fig.25] for simplicity, it will be appreciated that one or more air-oil heat exchangers may be provided on one or each oil loop in various implementations. Furthermore, in some implementations, only one oil loop may supply oil to both fuel-oil heat exchangers; no oil-oil heat exchanger may be provided in some such implementations.

[0301] In some arrangements, the engine 10 is configured to control the heat transfer ratio based on the SAF content of the fuel.

[0302] In some arrangements, the maximum temperature of the fuel exiting the secondary fuel-to-oil heat exchanger 1230 is at least 145°C, 150°C, 155°C, 160°C, 170°C, or 180°C.

[0303] For each of the hydraulic fuel actuators described above, the fuel system 152 may be configured such that the maximum fuel differential pressure across the at least one hydraulic fuel actuator under cruise conditions is at least 2500 kPa, 3000 kPa, 3500 kPa, 4000 kPa, 4500 kPa, 5000 kPa, 5500 kPa, 6000 kPa, 6500 kPa, 7000 kPa, or 7500 kPa.

[0304] For any of the hydraulic fuel actuators described above, the fuel delivery system 152 may be configured such that a maximum fuel differential pressure across the at least one hydraulic fuel actuator during takeoff is at least 9,000 kPa. For any of the hydraulic fuel actuators described above, the fuel delivery system 152 may be configured such that a maximum fuel differential pressure across the at least one hydraulic fuel actuator during takeoff may be in the range of 6,900 to 10,000 kPa, at least 10,000 kPa, 11,000 kPa, 12,000 kPa, 13,000 kPa, 14,000 kPa, or 15,000 kPa.

[0305] Actuator 1254 may be any of the actuators described herein, or any actuator that can be made fuel hydraulic, such as the variable stator vane actuator.

[0306] In some arrangements, the fuel temperature at the actuator inlet 1254 (leaving valve 1205) is at least 5°C, 7°C, 10°C, 12°C, 15°C or 20°C higher than that of the fuel at the inlet of the combustion chamber 16 (leaving valve 1202).

[0307] In some arrangements, the fuel is thermally stable at temperatures above 280°C. In this context, being "thermally stable" at a stated temperature means that the fuel passes the Jet Fuel Thermal Oxidation Test (JFTOT) at that temperature.

[0308] In some cases, the fuel is thermally stable at temperatures above 285°C, 290°C, 295°C, 300°C, 305°C, 310°C, 315°C, 320°C or 325°C

[0309] In some arrangements, aromatic compounds represent less than 5% of the volume of the fuel. In some arrangements, the calorific value of the fuel is at least 43.5 MJkg '. In some arrangements, the calorific value of the fuel is at least 44 MJkg '. In some arrangements, the sulfur content of the fuel is less than 15 parts per million. In some arrangements, the fuel is, or includes, a HEFA fuel.

[0310] Various methods of operating a gas turbine engine 10 for an aircraft 1 are described below.

[0311] [Fig. 20] is a flowchart showing a method 1300 of operating a gas turbine engine 10 according to various implementations. The method 1300 is performed for an engine 10 comprising: an engine core 11 comprising a turbine 19, a combustor 16, a compressor 14, and a core shaft 26 connecting the turbine to the compressor; a fan 23 located upstream of the engine core 11 and configured to be driven by the core shaft 26, the fan comprising a plurality of fan blades; and a nacelle 21 surrounding the fan 23 and the engine core 11 and defining a bypass duct 22 located radially outside the engine core 11; a plurality of actuators 254; and a fuel system 152.

[0312] In the described implementations, a bypass ratio, defined as the ratio of the mass flow rate of the flow through the bypass conduit 22 to the mass flow rate of the flow through the core 11 under cruising conditions, is greater than or equal to 4. Such bypass ratios may be used in any of the engines used to perform any of the methods of FIGS. 20-24.

[0313] The method 1300 comprises: supplying 1310, using the fuel supply system 152, fuel for combustion in the combustion chamber 16; and depending on the sustainable aviation fuel (SAF) content of the fuel, the choice 1320 between: providing 1320a, using the fuel delivery system 152, fuel to at least one actuator 354 of the plurality of actuators and causing the fuel to hydraulically actuate the at least one actuator 354; and causing 1320b the fuel to bypass the at least one actuator 354, again using the fuel delivery system 152.

[0314] The step 1320 of choosing between hydraulic fuel actuation of the actuator 354 and bypassing the at least one actuator 354 may include determining the action to be taken in a processor or other computing system based on a comparison of the S AF content of the fuel with one or more defined thresholds, and actuating a valve 304 based on this determination. The valve 304 may be a bi-directional valve configured to direct fuel through a first fuel line 360 ​​or through a second fuel line 361, 362, wherein the first fuel line fuel line 360 ​​drives actuator 354 and second fuel line 361, 362 bypasses actuator 354. Valve 304 may be simply designed to open a bleed in main fuel flow path 110 only when hydraulic fuel actuation is required, thereby allowing all fuel to remain in main fuel flow path 110 when hydraulic fuel actuation is not required. When hydraulic fuel actuation is not required, fuel may therefore bypass the at least one actuator 354 via a dedicated bypass line, or simply by remaining in a main fuel flow path that does not reach actuator 354.In implementations with a dedicated bypass pipe, the bypass pipe may be actuator-specific; provide a route for fuel to bypass a single actuator, or may bypass multiple actuators (e.g., some or all actuators located on the same fuel sampling path).

[0315] The step 1320 of choosing the action to be taken may consist of determining whether the SAF content of the fuel exceeds a threshold. The fuel system 152 may include one or more sensors configured to detect one or more parameters for calculating or inferring the SAF concentration of the fuel (e.g., by detecting a tracer element and referencing a lookup table of fuel tracer elements and SAF contents), and / or may include a repository for stored data on the fuel(s) in use. The fuel system 152 may include a processor or other computer system configured to receive the stored, calculated, or otherwise determined SAF content value and to compare this value to the threshold. A controller may then be used to operate the valve 304 in accordance with the determination.An electronic engine control (EEC) device may be used to perform the determination and implement the result in some implementations; in other implementations, it is possible to use a separate dedicated computing system instead of the EEC.

[0316] In implementations with multiple fuel hydraulic actuators 354, a different threshold may be set for different actuators, or the same threshold may be used for all actuators.

[0317] The fuel used to hydraulically fuel-drive the at least one actuator 354 may then be supplied to the combustion chamber 16 to be combusted; either directly or after being returned to the main fuel flow path 110. The fuel may flow along a main fuel flow path 110 to the combustion chamber 16; the main fuel flow path 110 may thus supply 1310 the fuel to the combustion chamber 16. A portion of the fuel flowing along the main fuel flow path 110 may be diverted from the main fuel flow path 110 to the at least one actuator 354; this fuel may then be returned to the main fuel flow path 110 after being used for actuation. One or more hoses may be used to take fuel in the main fuel flow path 110 to the at least one actuator 354. These hoses may form one or more fuel flow loops. One or more valves may be used to control the flow of fuel through the, or each, hydraulic fuel-to-fuel flow loop.In some implementations, one or more hydraulic fuel pumps may be provided to actively pump fuel to or from the at least one actuator 354; however, in many implementations, the pressure provided by one or more fuel pumps on the primary fuel flow path 110 may be sufficient.

[0318] The bleed-off 159 from the main fuel flow path 110 may be located at a different point in the return line 160 of the (or each) fuel flow loop, along the main fuel flow path 110, as illustrated in [Fig. 6]. The fuel may be returned to the main fuel flow path 110 upstream, or downstream, of the bleed-off. One or more valves may be used to control the location or locations along the main fuel flow path 110 where the fuel used to operate the actuators is returned to the main fuel flow path 110.For example, the fuel used for actuation may be returned to a fuel tank 50, 53, just before the combustion chamber 16, or before or after any fuel-oil heat exchanger or fuel pump along the main fuel flow path 110 or in a fuel hydraulic system, as appropriate. In general, the fuel used for actuation is returned to a point on the main fuel flow path 110 within the engine 10 (instead of being returned directly to a fuel tank 50, 53 located elsewhere in the aircraft 1, or to a connecting pipe between the two, for example).

[0319] The supply steps 1310, 1320a may therefore comprise the control of a plurality of fuel flow valves and, possibly also of a plurality of fuel pumps.

[0320] [Fig. 21] is a flowchart showing a method 1400 of operating a gas turbine engine 10 according to various implementations. The method 1400 is performed for an engine 10 comprising: an engine core 11 comprising a turbine 19, a combustion chamber 16, a compressor 14, and a core shaft 26 connecting the turbine to the compressor; a fan 23 located upstream of the engine core 11 and adapted to be driven by the core shaft 26, the fan comprising a plurality of fan blades; and a nacelle 21 surrounding the fan 23 and the engine core 11 and defining a bypass duct 22 located radially outside the engine core 11; a plurality of actuators 254; and a fuel supply system 152.

[0321] The method 1400 comprises: supplying 1410, using the fuel supply system 152, fuel for combustion in the combustion chamber 16; providing 1420, using the fuel supply system 152, fuel to fuel-hydraulic drive at least one actuator 354 of the plurality of actuators such that a maximum differential pressure of the fuel across the at least one fuel-hydraulic actuator 354 under cruise conditions is at least 2400 kPa.

[0322] The supplied fuel has an SAF content of at least 25% by volume, and may have a much higher SAF content. The maximum differential pressure may be controlled to be higher when the SAF content is higher.

[0323] As described with respect to method 1300 of [Fig. 20], one or more fuel flow loops and associated valves may be used to control the amount of fuel diverted from a main fuel flow path 110 to the actuated systems, and where along the main fuel flow path 110 that fuel is returned.

[0324] Similarly, one or more valves and / or pumps may be controlled 1420 as appropriate to control the maximum fuel differential pressure across the fuel hydraulic actuator 354. One or more pressure sensors may be provided to provide feedback on the fuel differential pressure. It will be appreciated that controlling the supply 1420 of fuel to the actuator(s) 354 will adjust the differential pressure.

[0325] [Fig. 22] is a flowchart showing a method 1500 of operating a gas turbine engine 10 according to various implementations. The method 1500 is performed for an engine 10 comprising: an engine core 11 comprising a turbine 19, a combustor 16, a compressor 14, and a core shaft 26 connecting the turbine to the compressor; a fan 23 located upstream of the engine core 11 and configured to be driven by the core shaft 26, the fan comprising a plurality of fan blades; and a nacelle 21 surrounding the fan 23 and the engine core 11 and defining a bypass duct 22 located radially outside the engine core 11; a plurality of actuators 254; a fuel system 152; a primary fuel-oil heat exchanger 1220 and a secondary fuel-oil heat exchanger 1230.

[0326] The method 1500 comprises: supplying 1510, using the fuel supply system 152, fuel for combustion in the combustion chamber 16; providing 1520, using the fuel supply system 152, fuel to hydraulically fuel-drive at least one actuator 354 of the plurality of actuators; heating 1530 at least the majority of the fuel using the primary fuel-oil heat exchanger 1220; and (further) heating 1540 the fuel to be supplied to hydraulically drive by fuel the at least one fuel hydraulic actuator using the secondary fuel-oil heat exchanger 1230.

[0327] The method comprises controlling the heating of the fuel in the heat exchangers 1220, 1230 such that, under cruising conditions, a heat transfer ratio of:

[0328] [Math 6] oil-to-fuel heat transfer rate in the secondary fuel-to-oil heat exchanger (k.Tkg'!) oil-to-fuel heat transfer rate in the primary and secondary fuel-to-oil heat exchangers ÇkJkg'1) has a maximum value of at least 0.35.

[0329] The primary fuel-oil heat exchanger 1220 may be configured to direct heat away from the engine oil 10 that is used to lubricate and cool various engine components, such as a gearbox (if applicable) and / or shaft bearings. The secondary fuel-oil heat exchanger 1230 may be configured to direct heat away from the oil used to cool, and possibly also lubricate, the generator. Different, fluidically isolated oil flows may pass through each of the primary fuel-oil heat exchangers 1220 and the secondary fuel-oil heat exchangers 1230 in some implementations, as illustrated in [Fig. 25]. In other implementations, the same oil may pass through both heat exchangers, for example, passing through the secondary fuel-oil heat exchanger 1230 and then the primary fuel-oil heat exchanger 1220 in series.

[0330] The heat transfer ratio may be controlled by opening and closing the valves to permit: recirculation of fuel through one or both of the heat exchangers 1220, 1230, fuel bypass of one or both of the heat exchangers 1220, 1230, oil recirculation through one or both of the heat exchangers 1220, 1230, or oil bypass of one or both of the heat exchangers 1220, 1230. In addition, one or both of the heat exchangers 1220, 1230 may be controlled by opening and closing the valves to permit: recirculation of fuel through one or both of the heat exchangers 1220, 1230, fuel bypass of one or both of the heat exchangers 1220, 1230, or oil bypass of one or both of the heat exchangers 1220, 1230. several additional heat exchangers may be ordered to adjust the heat transfer ratio: for example, an air-to-oil heat exchanger may be used to cool the oil before it reaches a fuel-to-oil heat exchanger 1220, 1230, and in implementations with multiple different oil flows, an oil-to-oil heat exchanger may be provided to transfer heat between the different oils. The air and / or oil flows in such heat exchangers may therefore also be adjusted to affect the heat transfer ratio defined above.

[0331] The heat transfer ratio may be controlled through control of one or more valves configured to allow a percentage of the fuel to be recirculated through one or both of the fuel-oil heat exchangers 1220, 1230 or to bypass one or both of the heat exchangers 1220, 1230. The heat transfer ratio may be controlled through control of one or more valves configured to allow a percentage of the oil to be recirculated through one or both of the heat exchangers 1220, 1230 or to bypass one or both of the heat exchangers 1220, 1230.

[0332] As described with respect to methods 1300, 1400 of Figures 20 and 21, one or more fuel flow loops and associated valves may be used to control the amount of fuel diverted from a primary fuel flow path 110 to the actuator(s) and secondary heat exchanger 1230, and where along the primary fuel flow path 110 that fuel is returned.

[0333] [Fig. 23] is a flowchart showing a method 1600 of operating a gas turbine engine 10 according to various implementations. The method 1600 is performed for an engine 10 comprising: an engine core 11 comprising a turbine 19, a combustor 16, a compressor 14, and a core shaft 26 connecting the turbine to the compressor; a fan 23 located upstream of the engine core 11 and configured to be driven by the core shaft 26, the fan comprising a plurality of fan blades; and a nacelle 21 surrounding the fan 23 and the engine core 11 and defining a bypass duct 22 located radially outside the engine core 11; a plurality of actuators including a variable stator vane actuator 854; and a fuel system 152.

[0334] The method 1600 comprises: providing 1610, using the fuel supply system 152, fuel for combustion in the combustion chamber 16; and providing 1620, using the fuel supply system 152, fuel to hydraulically fuel-drive the variable compressor vane actuator 854 of the plurality of actuators.

[0335] Providing 1620, using the fuel delivery system 152, fuel to hydraulically fuel the variable compressor vane actuator 854 and to the combustor 16 may include providing fuel that is thermally stable at 280°C, and optionally also thermally stable at higher temperatures.

[0336] As described with respect to methods 1300, 1400, 1500 of Figures 20-22, one or more fuel flow loops and associated valves may be used to control the amount of fuel diverted from a main fuel flow path 110 to the actuated systems (and in particular to the variable stator vane actuator 854), and where along the main fuel flow path 110 that fuel is returned.

[0337] [Fig. 24] is a flowchart showing a method 1700 of operating a gas turbine engine 10 according to various implementations. The method 1700 is performed for an engine 10 comprising: an engine core 11 comprising a turbine 19, a combustor 16, a compressor 14, and a core shaft 26 connecting the turbine to the compressor; a fan 23 located upstream of the engine core 11 and configured to be driven by the core shaft 26, the fan comprising a plurality of fan blades; and a nacelle 21 surrounding the fan 23 and the engine core 11 and defining a bypass duct 22 located radially outside the engine core 11; a plurality of actuators 254; a fuel system 152; and at least one fuel-oil heat exchanger 1220 configured to have the oil and the fuel passing therethrough, the at least one heat exchanger 1220 being configured to transfer heat from the oil to the fuel.

[0338] The method 1700 comprises: supplying 1710, using the fuel supply system 152, fuel for combustion in the combustion chamber 16; providing 1720 fuel to hydraulically fuel-drive at least one actuator 1254 of the plurality of actuators; and the control 1730 of the at least one heat exchanger 1220 such that, in cruising mode, the temperature of the fuel at the inlet of the at least one actuator 1254 is at least 5°C higher than the temperature of the fuel at the inlet of the combustion chamber 16 (and possibly at least 10°C hotter).

[0339] The fuel system 152 may include two heat exchangers: a primary heat exchanger 1220 and a heat exchanger secondary 1230. Control 1730 of the at least one heat exchanger 1220 may include control of the primary heat exchanger 1220 and / or the secondary heat exchanger 1230. The secondary heat exchanger 1230 may be configured to provide supplemental heat just for the fuel to be supplied to the fuel hydraulic actuator(s) 1254.

[0340] The method 1700 may include determining at least one characteristic of the fuel and controlling a temperature difference between the temperature of the fuel at the inlet of the at least one actuator 1254 and the temperature of the fuel at the inlet of the combustion chamber 16 based on the at least one characteristic of the fuel. The characteristic of the fuel may be its SAF content. The characteristic of the fuel may be determined by any approach known in the art, ranging from extracting data from an onboard data store of fuel information to determining or inferring from one or more sensed parameters.

[0341] The fuel characteristic may be or include the calorific value, thermal stability, or the percentage of sustainable aviation fuel (SAF) in the fuel.

[0342] Controlling 1730 the fuel temperature at the inlet of the at least one actuator 1254 relative to the fuel temperature at the inlet of the combustion chamber 16 includes actuating one or more valves to recirculate the fuel through one or more heat exchangers or to bypass the fuel around one or more of the heat exchangers 1220, 1230.

[0343] Controlling 1730 the fuel temperature at the inlet of the at least one actuator 1254 relative to the fuel temperature at the inlet of the combustion chamber 16 includes actuating one or more valves to recirculate the oil through the heat exchanger 1220, 1230 or to bypass the oil around the heat exchanger.

[0344] As described with respect to methods 1300, 1400, 1500, 1600 of Figures 20-23, one or more fuel flow loops and associated valves may be used to control the amount of fuel diverted from a main fuel flow path 110 to the actuated systems, and where along the main fuel flow path 110 that fuel is returned.

[0345] It will be understood that the invention is not limited to the implementations described above and that various modifications and improvements may be made without departing from the concepts described herein. Unless mutually exclusive, any feature may be employed separately or in combination with other features and the description extends to and includes all combinations and subcombinations of one or more features described herein.

Claims

Claims

1. A gas turbine engine (10) for an aircraft comprising: an engine core (11) comprising a turbine (19), a combustion chamber (16), a compressor (14), and a core shaft (26) connecting the turbine to the compressor; a fan (23) located upstream of the engine core (11) and adapted to be driven by the core shaft (26), the fan comprising a plurality of fan blades; a nacelle (21) surrounding the fan (23) and the engine core (11) and defining a bypass duct (22) located radially outside the engine core (11), wherein a 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; a plurality of actuators (254);and a fuel supply system (152), wherein the fuel supply system is configured to supply fuel for combustion in the combustion chamber (16), and to supply fuel to hydraulically fuel-drive at least one of the plurality of actuators (254), and wherein the fuel supply system (152) is configured to be controlled to select between: causing the fuel to actuate at least one actuator (254); and causing the fuel to bypass the at least one actuator (254), depending on the sustainable aviation fuel, S AF, content of the fuel.;

2. A gas turbine engine (10) according to claim 1, wherein the core shaft (26) directly provides drive to the fan (23), so as to drive the fan at the same rotational speed as the core shaft, such that the engine (10) is a direct drive turbine engine.

3. A gas turbine engine (10) according to claim 1, wherein the turbine engine (10) comprises a reduction gear (30) which receives an input from the core shaft (26) and provides a drive to the fan (23) so as to drive the fan at a speed of lower rotation than the heart shaft, so that the motor (10) is a turbine motor with reduction gear.

4. The gas turbine engine (10) of claim 1, wherein the engine (10) comprises a turbine case cooling system (554), wherein the plurality of actuators (254) comprises an actuator (554a, 554b) that is part of the turbine case cooling system, and the fuel delivery system (152) is configured to actuate or bypass the actuator (554a, 554b) that is part of a turbine case cooling system (554) depending on the SAF content of the fuel.

5. The gas turbine engine (10) of claim 1, wherein the engine (10) comprises a cabin air bleed valve (654'), the plurality of actuators (254) comprises an actuator (654) configured to actuate the cabin air bleed valve (654'), and the fuel delivery system (152) is configured to actuate or bypass the actuator (654) configured to actuate the cabin air bleed valve based on the SAF content of the fuel.

6. The gas turbine engine (10) of claim 1, wherein the engine (10) comprises a bleed air management valve (754'), the plurality of actuators (254) comprises an actuator (754) configured to actuate the bleed air management valve, and the fuel delivery system (152) is configured to actuate or bypass the bleed air management valve (754') based on the SAF content of the fuel.

7. The gas turbine engine (10) of claim 1, wherein the engine (10) comprises an engine heat management system (1000) comprising a valve (1004), wherein the plurality of actuators (254) comprises an actuator (1054) configured to actuate the valve in the engine heat management system (1000), and the fuel delivery system (152) is configured to actuate or bypass the valve (1004) in the engine heat management system depending on the SAF content of the fuel.

8. A gas turbine engine (10) according to claim 1, wherein the engine (10) comprises a generator heat management system (1100) comprising a valve (1104), wherein the plurality of actuators (254) includes an actuator (1154) configured to actuate the valve in the generator heat management system, and the fuel delivery system (152) is configured to actuate or bypass the valve (1104) in the generator heat management system based on the SAF content of the fuel.

9. The gas turbine engine (10) of claim 1, wherein when the fuel delivery system (152) is configured to bypass the at least one actuator (254), the at least one actuator (254) is actuated in a non-hydraulic fuel manner.

10. The gas turbine engine (10) of claim 1, wherein the fuel delivery system (152) is configured to cause fuel to actuate the at least one actuator (254) when the SAF content is above a threshold, and to cause fuel to bypass the at least one actuator (254) when the SAF content is below the threshold.

11. A gas turbine engine (10) according to claim 1, wherein the minimum SAF content required for actuation of the at least one actuator (254) is at least 25%; optionally wherein the minimum SAF content required for actuation of the at least one actuator (254) is at least 50%.

12. The gas turbine engine (10) of claim 1, wherein the fuel delivery system (152) is configured to be controllable to select between fuel hydraulic actuation and bypass for two or more of the plurality of actuators (254).

13. The gas turbine engine (10) of claim 1, wherein when the at least one actuator (254) is hydraulically driven by fuel, the maximum operating differential pressure during takeoff is at least 6900 kPa; optionally wherein the maximum operating differential pressure during takeoff is greater than 7000 kPa.

14. A gas turbine engine (10) according to claim 1, wherein when the at least one actuator is hydraulically driven by fuel, the maximum operating differential pressure under cruise conditions is at least 2400 kPa; optionally wherein, when the at least one actuator is hydraulically driven by fuel, the maximum operating differential pressure under cruising conditions is greater than 2,500 kPa.

15. A method (1300) of operating a gas turbine engine (10) for an aircraft, the engine comprising: an engine core (11) comprising a turbine (19), a combustion chamber (16), a compressor (14), and a core shaft (26) connecting the turbine to the compressor; a fan (23) located upstream of the engine core (11) and configured to be driven by the core shaft (26), the fan comprising a plurality of fan blades; a nacelle (21) surrounding the fan (23) and the engine core (11) and defining a bypass duct (22) located radially outside the engine core (11), wherein a 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; a plurality of actuators (254); and a fuel supply system (152);and wherein the method (1300) comprises: providing (1310), using the fuel supply system (152), fuel for combustion in the combustion chamber (16); and selecting (1320) between: providing (1320a), using the fuel supply system (152), fuel to at least one actuator of the plurality of actuators (254) to hydraulically fuel-actuate the at least one actuator (254); and using the fuel supply system (152), causing (1320b) the fuel to bypass the at least one actuator (254), depending on the sustainable aviation fuel (SAF) content of the fuel.;

16. The method (1300) of claim 15, wherein the fuel supply system (152) is controllable to supply fuel to at least two of the plurality of actuators (254), and the method (1300) comprises, depending on the sustainable aviation fuel, S AF, content of the fuel, choosing (1320) between: causing (1320a) the fuel to hydraulically actuate the at least two actuators (254); and causing (1320b) the fuel to bypass the at least two actuators.

17. The method (1300) of claim 15, wherein the fuel supply system (152) is controllable to supply fuel to at least two of the plurality of actuators (254), and the method (1300) comprises, based on the sustainable aviation fuel (SAF) content of the fuel, choosing (1320) individually for each of the at least two actuators (254) between: causing (1320a) the fuel to hydraulically fuel the actuator (254); and causing (1320b) the fuel to bypass the actuator (254).

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