VARIETY OF ACTUATORS

Abstract

A gas turbine engine for an aircraft comprises an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor. The engine includes a fan located upstream of the engine core and arranged to be driven by the core. The engine includes a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outward of the engine core, wherein the bypass ratio, defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at cruise conditions, is at least 4. The engine includes a plurality of actuators and a fuel supply system configured to supply fuel for combustion in the combustor and to supply fuel for fuel-hydraulic drive of at least one actuator.The fuel comprises at least 25 vol% SAF and the fuel supply system is designed so that a peak differential pressure of the fuel across the at least one fuel hydraulic actuator during cruise flight conditions is at least 2400 kPa.

Description

[0001] The present disclosure relates to aircraft actuation systems and to methods for fuel-hydraulic actuation of actuation systems.

[0002] The aviation industry is expected to evolve toward the use of fuels that differ from the currently commonly used conventional kerosene-based jet fuels. These fuels may exhibit different fuel properties compared to petroleum / fossil-based hydrocarbon fuels. Thus, there is a need to consider fuel properties for these new fuels and adapt both the gas turbine engines themselves and the processes for operating them.

[0003] According to a first 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 arranged 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 outward of the engine core, 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, being at least 4; a variety 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 fuel-hydraulically drive at least one actuator of the plurality of actuators, and wherein the fuel supply system is arranged 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, based on the Sustainable Aviation Fuel (SAF) content of the fuel.

[0004] The inventors recognized that the use of fuels other than conventional kerosene-based jet fuels, such as sustainable aviation fuels, can result in different fuel properties, and that these different fuel properties can enable actuators to be driven by fuel hydraulics. In particular, some fuels can be heated to higher temperatures and used to drive at least one additional actuator than conventional fuels without significantly increasing the risk of thermal fuel degradation (e.g., fuel varnishing or fuel coking) in the actuators.

[0005] The bypass ratio is greater than or equal to 4 and can range from 4 to 55. The bypass ratio can range from 4 to 20. The bypass ratio can range from 4 to 15.

[0006] The fuel system may be configured 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 based on the SAF content of the fuel. One or more of the fuel hydraulic actuators may have a different threshold from one another—for example, the fuel system may be configured to cause the fuel to actuate a first actuator if the SAF content of the fuel is above a first threshold, or otherwise bypass that actuator, and to cause the fuel to actuate a second actuator if the SAF content of the fuel is above a second threshold that is higher than the first threshold, or otherwise bypass that actuator.

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

[0009] The turbine engine may include a gearbox that receives input from the core shaft and outputs drive to the fan to drive the fan at a lower speed than the core shaft. Such an engine may be a geared 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 supply 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 supply system may be configured to supply fuel to fuel-hydraulically drive the at least two of the plurality of actuators that are part of the TTC system. The fuel supply 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 vent valve. The plurality of actuators may include an actuator configured to actuate the cabin vent valve. The fuel supply system may be configured to actuate or override the actuator configured to actuate the cabin vent valve based on the SAF content of the fuel.

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

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

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

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

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

[0017] If the fuel supply system bypasses the at least one actuator, the at least one actuator may be actuated non-fuel hydraulically. If the actuator is actuated non-fuel hydraulically, it may be actuated using a non-fuel-based hydraulic fluid. To avoid mixing, the non-fuel-based hydraulic fluid may be supplied via a check valve. To avoid mixing, the fuel may be supplied to the actuator using a check valve. The actuator may instead be actuated electrically or pneumatically.

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

[0019] The minimum SAF content required to actuate the at least one actuator may be at least 25 vol%, 30 vol%, 35 vol%, 40 vol%, 45 vol%, 50 vol%, 52 vol%, 55 vol%, 60 vol%, 65 vol%, 70 vol% or 75 vol%.

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

[0021] If the at least one actuator is fuel-hydraulically driven, the maximum differential operating pressure across that actuator during start-up conditions may be in the range of 6,900 kPa to 10,000 kPa or greater than 10,000 kPa. If the at least one actuator is fuel-hydraulically driven, the maximum differential operating pressure during start-up 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] If the at least one actuator is fuel-hydraulically driven, the maximum differential operating pressure during cruise conditions may be greater than or equal to 2400 kPa, greater than 2500 kPa, greater than 3000 kPa, greater than 3500 kPa, greater than 3800 kPa, or greater than 4000 kPa.

[0023] The same actuator can be exposed to significantly lower pressure when idling - for example, a differential pressure in the range of 1000 kPa to 1250 kPa (150-180 psid).

[0024] According to a second aspect, there is provided a method of operating a gas turbine engine for an aircraft, the gas turbine 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 arranged 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 outward of the engine core, 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, being at least 4; a variety of actuators; and a fuel supply system; and wherein the method comprises: Supplying fuel using the fuel supply system for combustion in the combustion chamber; and Choose between: Supplying fuel using the fuel supply system to at least one actuator of the plurality of Actuators for actuating the at least one actuator using fuel hydraulics; and Using the fuel supply system to cause the fuel to bypass the at least one actuator, based on the content of sustainable aviation fuel - SAF - in the fuel.

[0025] The procedure may include: Supplying fuel using the fuel supply system to at least two of the plurality of actuators and, based on the content of Sustainable Aviation Fuel (SAF) in the fuel, selecting for each of the at least two actuators (individually / independently) between: Causing the fuel to actuate the actuator hydraulically; and Cause the fuel to bypass the actuator.

[0026] The procedure may include: Supplying fuel using the fuel supply system to at least two of the plurality of actuators and selecting, based on the Sustainable Aviation Fuel (SAF) content of the fuel, between: Causing the fuel to actuate the at least two actuators in a fuel-hydraulic manner; and Cause the fuel to bypass at least two actuators.

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

[0028] The bypass ratio can range from 4 to 55. The bypass ratio can range from 4 to 20. The bypass ratio can range from 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 fuel pressure flowing through the at least one fuel hydraulic actuator based on the SAF content of the fuel. For example, the pressure may be increased by at least 350 kPa for each 5% increase in the SAF content above 60% (volume percent). One or more pumps and / or valves may be used to increase the pressure.

[0031] According to a third 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 arranged 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 outward of the engine core, 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, being at least 4; a variety of actuators; and a fuel supply system, the fuel supply system being configured to supply fuel for combustion in the combustion chamber and to supply fuel to fuel-hydraulically drive at least one actuator of the plurality of actuators.

[0032] The fuel consists of at least 25% by volume of sustainable aviation fuel (SAF). The fuel supply system is designed to ensure a peak fuel differential pressure across at least one fuel hydraulic actuator during cruise flight conditions of at least 2400 kPa (350 psid).

[0033] It is understood that the differential pressure across an actuator changes depending on operating conditions (e.g., fuel flow rate), even at cruise conditions. The peak or maximum differential pressure is therefore selected for ease of comparison; more accurately, this pressure is a steady-state peak differential pressure; that is, any short-term, transient pressure spikes are not taken into account. The steady-state peak value can therefore be a time-averaged pressure value, averaged over five, ten, fifteen, twenty, twenty-five, or thirty seconds. The peak duration should be at least approximately five seconds, so that abrupt transient spikes are excluded.

[0034] The peak differential pressure of the fuel across the at least one fuel hydraulic actuator 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 during cruise flight conditions.

[0035] The peak differential pressure of the fuel across the at least one fuel hydraulic actuator during cruise conditions may be in the range of 2400 kPa to 4500 kPa and optionally from 2500 kPa to 4000 kPa or from 2500 kPa to 3800 kPa.

[0036] The bypass ratio is greater than or equal to 4 and can range from 4 to 55. The bypass ratio can range from 4 to 20. The bypass ratio can range from 4 to 15.

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

[0038] The same actuator can be exposed to significantly lower pressure when idling - for example, a differential pressure in the range of 1000 kPa to 1250 kPa (150-180 psid).

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

[0040] The turbine engine may include a gearbox that receives input from a core shaft and outputs drive to the fan to drive the fan at a lower speed than the core shaft, so that the engine is a geared turbine engine.

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

[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 an actuator for variable inlet guide vanes.

[0044] The engine of the third aspect may include some 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, a method of operating a gas turbine engine for an aircraft is provided. 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 arranged 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 outward of the engine core, 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, being at least 4; a variety of actuators; and a fuel supply system;

[0046] The procedure includes: Supplying fuel using the fuel supply system for combustion in the combustion chamber; and Supplying fuel comprising at least 25 vol% SAF using the fuel supply system to fuel-hydraulically drive at least one actuator of the plurality of actuators, wherein the fuel is supplied such that a peak differential pressure of the fuel across the at least one fuel-hydraulic actuator during cruise conditions is at least 2400 kPa.

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

[0048] The peak differential pressure of the fuel across the at least one fuel hydraulic actuator may be at least 2500 kPa during cruise flight conditions.

[0049] The bypass ratio can range from 4 to 55. The bypass ratio can range from 4 to 20. The bypass ratio can range from 4 to 15.

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

[0051] The method may include supplying fuel comprising at least 50 vol% SAF and controlling the fuel supply to at least one actuator such that the peak differential pressure is at least 3200 kPa.

[0052] The method may include supplying fuel comprising at least 55 vol% SAF and controlling the fuel supply to at least one actuator such that the peak differential pressure is at least 3550 kPa or 3600 kPa.

[0053] The methods of the second and fourth aspects may be complementary and 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 arranged 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 outward of the engine core, 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, being at least 4; a variety 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 fuel-hydraulically drive at least one actuator of the plurality of actuators, and two fuel-to-oil heat exchangers arranged to allow oil and fuel to flow therethrough, the heat exchangers being arranged to transfer heat from the oil to the fuel and comprising a primary fuel-to-oil heat exchanger arranged to heat at least a majority of the fuel and a secondary fuel-to-oil heat exchanger arranged to add additional heat to the fuel to be supplied to fuel-hydraulicly drive the at least one fuel-hydraulic actuator.

[0055] The heat exchangers are controlled to achieve a heat transfer ratio of: Heat transfer rate from oil to fuel in the secondary fuel-oil heat exchanger (kJkg−1)Heat transfer rate from oil to fuel in the primary and secondary fuel-oil heat exchanger (kJkg−1) has a maximum value of at least 0.35.

[0056] While it is understood that the ratio is dimensionless, in the examples described herein, heat transfer is measured per unit mass or volume of fuel (with the units kJ per kg exemplified above), thus providing a heat transfer rate normalized for variations in cruise fuel flow rate. It is understood that heat transfer is reported per unit mass (kg) of fuel reaching the combustion chamber to account for fuel flow rate and any recirculation through one or more of the heat exchangers or bypass of one or more of the heat exchangers, as described elsewhere herein. In most implementations, any fuel temperature increase due to other engine components (rather than heat transfer from the oil) can be assumed to be minimal.

[0057] It is understood that even under cruise conditions, short-term, transient peaks in heat transfer can occur that should not be considered. The maximum value for 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 is understood that the timescales for transient temperature changes may be larger than those for pressure in actuators. While the full movement of an actuator is likely completed within a few seconds, fuel and oil may remain in the heat exchangers for longer.

[0058] The heat transfer ratio can have a maximum value of over 0.4 during cruise flight. The heat transfer ratio can have a maximum value of over 0.45 during cruise flight. The heat transfer ratio can have a maximum value of over 0.5 during cruise flight.

[0059] The bypass ratio is greater than or equal to 4 and can range from 4 to 55. The bypass ratio can range from 4 to 20. The bypass ratio can range from 4 to 15.

[0060] The maximum temperature of the fuel leaving the secondary fuel-oil heat exchanger can be at least 150 °C, 160 °C, 170 °C or 180 °C.

[0061] The primary and secondary heat exchangers may be arranged so that the heat transfer ratio is adjustable during operation of the engine, for example by controlling one or more pumps and / or valves to adjust the fuel and / or oil flow through one or each heat exchanger.

[0062] The heat exchangers can be arranged so that the heat transfer ratio is adjustable based on the content of sustainable aviation fuel - SAF - in the fuel.

[0063] The engine may include at least one bypass valve arranged to allow fuel or oil to bypass one of the fuel-oil heat exchangers to adjust the heat transfer ratio.

[0064] For example, the engine may comprise one or more of the following elements: • a valve operable to allow oil to bypass the primary heat exchanger; • a valve operable to allow the oil to bypass the secondary heat exchanger; • a valve operable to allow the fuel to bypass the primary heat exchanger; and • a valve operable to allow fuel to bypass the secondary heat exchanger.

[0065] The engine may include at least one recirculation valve arranged to allow fuel or oil to be passed through one of the fuel-oil heat exchangers multiple times to adjust the heat transfer ratio.

[0066] For example, the engine may comprise one or more of the following elements: • a valve operable to allow the recirculation of oil through the primary heat exchanger; • a valve operable to allow the recirculation of oil through the secondary heat exchanger; • a valve operable to allow the recirculation of fuel through the primary heat exchanger; and • a valve operable to allow the recirculation of fuel through the secondary heat exchanger.

[0067] Each of the recirculation valves described herein may have one or more associated pumps configured to return the oil / fuel to the heat exchanger inlet for recirculation through the heat exchanger. Alternatively or additionally, any suitable components may be used to repressurize the oil / fuel to enable recirculation.

[0068] The core shaft can output the drive directly to the fan to drive the fan at the same speed as the core shaft, so that the engine is a direct drive turbine engine.

[0069] The engine of the fifth aspect may include some or all of the features of the engine of the first and / or third aspect and may be used to implement the method of the second or fourth aspect.

[0070] According to a sixth aspect, a method of operating a gas turbine engine for an aircraft is provided.

[0071] The engine includes: 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 arranged 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 outward of the engine core, 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, being at least 4; a variety of actuators; a fuel supply system; and a primary fuel-oil heat exchanger and a secondary fuel-oil heat exchanger.

[0072] The procedure includes: Supplying fuel using the fuel supply system for combustion in the combustion chamber; Supplying fuel using the fuel supply system to fuel-hydraulically 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 for fuel-hydraulic driving of the at least one fuel-hydraulic actuator using the secondary fuel-oil heat exchanger; whereby under cruising conditions the supply and heating are controlled so that a heat transfer ratio of: Heat transfer rate from oil to fuel in the secondary fuel-oil heat exchanger (kJkg−1)Heat transfer rate from oil to fuel in the primary and secondary fuel-oil heat exchanger (kJkg−1) has a maximum value of at least 0.35.

[0073] The bypass ratio can range from 4 to 55. The bypass ratio can range from 4 to 20. The bypass ratio can range from 4 to 15.

[0074] The method may further comprise adjusting the heat transfer ratio - for example, by adjusting the fuel and / or oil flow rate through one or each heat exchanger - based on the Sustainable Aviation Fuel (SAF) content in the fuel.

[0075] 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.

[0076] According to a seventh 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 arranged 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 outward of the engine core, 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, being at least 4; a plurality of actuators, including a variable compressor vane actuator; and a fuel supply system, the fuel supply system being configured to supply fuel for combustion in the combustion chamber and supply fuel to fuel-hydraulically drive the variable compressor guide vane actuator, the fuel being thermally stable at temperatures above 280°C.

[0077] The bypass ratio is greater than or equal to 4 and can range from 4 to 55. The bypass ratio can range from 4 to 20. The bypass ratio can range from 4 to 15.

[0078] "Thermally stable" may refer to the ability of the fuel to meet the requirements of a jet fuel thermal oxidation test (JFTOT). A fuel that is thermally stable at temperatures above 280°C may be considered to be a fuel that has met the requirements of a JFTOT test at temperatures above or equal to 280°C.

[0079] The fuel may be thermally stable at temperatures of 290 °C, 295 °C, 300 °C, 305 °C, 310 °C, 315 °C, 320 °C or 325 °C or above.

[0080] Aircraft gas turbine engines generally incorporate a plurality of compressor vanes. The variable compressor vane actuator is designed to adjust the position of a compressor vane. Since the vane serves to direct the airflow and the position of the compressor vane is variable, it can be referred to as a variable vane.

[0081] The variable vanes of different implementations may include one or both of the following elements: • Variable Inlet Guide Vanes (VIGV), located upstream of / in front of the compressor rotor blades and arranged to guide and control the airflow into the compressor; and • other guide vanes - generally variable stator vanes (VSV), which are located behind / downstream of the compressor rotor blades and are arranged to direct and control the airflow through the compressor.

[0082] The engine may include multiple variable stator vane actuators, each of which may be fuel-hydraulic. Furthermore, the engine may include multiple compressors, each of which has at least one variable stator vane actuator.

[0083] The gas turbine engine may include a plurality of variable compressor vanes and a plurality of variable compressor vane actuators. The fuel supply system may be configured to supply fuel to fuel-hydraulically drive the plurality of variable compressor vane actuators.

[0084] The gas turbine engine may include at least two compressors, each compressor having at least one variable compressor vane and at least one variable compressor vane actuator connected thereto. The fuel supply system may be configured to supply fuel to fuel-hydraulically drive at least one variable compressor vane actuator of each compressor.

[0085] The engine may include a turbine casing cooling system. The plurality of actuators may include a turbine casing cooling actuator. The fuel supply system may be configured to supply fuel to fuel-hydraulically drive the turbine casing cooling actuator.

[0086] An aircraft gas turbine engine typically includes multiple servomotors (also called servos), where a servo is a rotary or linear actuator that enables precise control of angular or linear position, velocity, and / or acceleration in a mechanical system. In a given engine, one, some, or all of the servos may be fuel-hydraulically actuated. One or more servos may include a sensor configured to provide position feedback. A dedicated control unit may be provided for the engine thermal management system to control the one or more servos.

[0087] For example, the engine may include a servo within a hydromechanical unit. The plurality of actuators may include at least one actuator configured to actuate the servo within the hydromechanical unit. The fuel supply system may be configured to supply fuel to fuel-hydraulically drive the actuator configured to actuate the servo valve.

[0088] Aromatics can make up less than 5% of the fuel volume.

[0089] The calorific value of the fuel can be at least 43.5 MJkg -1 The calorific value of the fuel can be at least 44 MJkg -1 be.

[0090] The sulfur content of the fuel can be less than 15 parts per million.

[0091] The fuel may be or comprise a HEFA fuel, i.e. a fuel made from hydrotreated esters and fatty acids.

[0092] The core shaft can output the drive directly to the fan to drive the fan at the same speed as the core shaft, so that the engine is a direct drive turbine engine.

[0093] The turbine engine may include a gearbox that receives input from a core shaft and outputs drive to the fan to drive the fan at a lower speed than the core shaft, so that the engine is a geared turbine engine.

[0094] The engine of the seventh aspect may include some or all of the features of the engine of the first, third and / or fifth aspect and may be used to implement the method of the second, fourth or sixth aspect.

[0095] According to an eighth aspect of the invention, a method of operating a gas turbine engine for an aircraft is provided.

[0096] The engine includes: 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 arranged 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 outward of the engine core, 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, being at least 4; a plurality of actuators, including a variable compressor vane actuator; and a fuel supply system.

[0097] The bypass ratio is greater than or equal to 4 and can range from 4 to 55. The bypass ratio can range from 4 to 20. The bypass ratio can range from 4 to 15.

[0098] The procedure includes: Supplying fuel using the fuel supply system for combustion in the combustion chamber; and Supplying fuel using the fuel supply system to fuel-hydraulically drive the variable compressor vane actuator of the plurality of actuators.

[0099] The fuel is thermally stable at temperatures above 280 °C.

[0100] 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.

[0101] According to a ninth aspect, a gas turbine engine for an aircraft is provided. 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 arranged 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 outward of the engine core, wherein a bypass ratio, defined as the ratio of the mass flow rate through the bypass duct to the mass flow rate through the core at cruise conditions, is at least 4; a variety of actuators; a fuel supply system, the fuel supply system being configured to supply fuel for combustion in the combustion chamber and to supply fuel to fuel-hydraulically drive at least one actuator of the plurality of actuators; and at least one fuel-oil heat exchanger arranged to allow oil and fuel to flow therethrough, the at least one heat exchanger being arranged to transfer heat from the oil to the fuel.

[0102] The at least one heat exchanger is arranged such that, during cruise flight, the fuel temperature entering the at least one actuator is at least 5 °C higher than the fuel temperature entering the combustion chamber.

[0103] The bypass ratio is greater than or equal to 4 and can range from 4 to 55. The bypass ratio can range from 4 to 20. The bypass ratio can range from 4 to 15.

[0104] The at least one heat exchanger can be arranged such that, during cruise flight, the fuel temperature upon entry into 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 fuel temperature upon entry into the combustion chamber.

[0105] The core shaft can output the drive directly to the fan to drive the fan at the same speed as the core shaft, so that the engine is a direct drive turbine engine.

[0106] The turbine engine may include a gearbox that receives input from a core shaft and outputs drive to the fan to drive the fan at a lower speed than the core shaft, so that the engine is a geared turbine engine.

[0107] The at least one heat exchanger may comprise a primary fuel-oil heat exchanger arranged to heat at least the majority of the fuel and a secondary fuel-oil heat exchanger arranged to add additional heat to the fuel to be supplied to fuel-hydraulic drive the at least one fuel-hydraulic actuator.

[0108] The primary and secondary fuel-oil heat exchangers can be controlled to achieve a heat transfer ratio of: Heat transfer rate from oil to fuel in the secondary fuel-oil heat exchanger (kJkg−1)Heat transfer rate from oil to fuel in the primary and secondary fuel-oil heat exchanger (kJkg−1) has a maximum value of at least 0.35.

[0109] The primary and secondary heat exchangers can be arranged so that the heat transfer ratio can be adjusted during engine operation.

[0110] To provide adjustment of the heat transfer ratio during operation, the gas turbine engine may comprise at least one of the following elements: (i) at least one controllable oil bypass valve arranged to allow oil to bypass at least one of the primary and secondary fuel-oil heat exchangers; and (ii) at least one controllable oil return valve arranged to allow return of oil through at least one of the primary and secondary fuel-oil heat exchangers.

[0111] The fuel supply system may include a valve operable to bypass or return fuel through the primary heat exchanger.

[0112] The fuel supply system may include a valve operable to bypass or return fuel through the secondary heat exchanger.

[0113] The engine may be configured 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 at higher SAF contents of the fuel.

[0114] 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., in a tank in the main fuselage of an aircraft or on an aircraft wing—and need not necessarily be part of the engine. The fuel supplied to the at least one actuator for fuel-hydraulic drive of the at least one actuator may be supplied after actuation of the combustion chamber without returning to the fuel source. The fuel may therefore be recirculated around the engine or used in various ways within the engine without having to be returned to a fuel tank.

[0115] The fuel supplied to the combustion chamber may comprise a mixture of fuel that has passed through the at least one fuel-oil heat exchanger and fuel that has bypassed the at least one fuel-oil heat exchanger.

[0116] The engine of the ninth aspect may include some 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.

[0117] According to a tenth aspect, a method of operating a gas turbine engine for an aircraft is provided.

[0118] The engine includes: 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 arranged 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 outward of the engine core, 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, being at least 4; a variety of actuators; a fuel supply system, the fuel supply system being configured to supply fuel for combustion in the combustion chamber and to supply fuel to fuel-hydraulically drive at least one actuator of the plurality of actuators; and at least one fuel-oil heat exchanger arranged to allow oil and fuel to flow therethrough, the at least one heat exchanger being arranged to transfer heat from the oil to the fuel.

[0119] The method comprises controlling the at least one heat exchanger such that, during cruise flight, the fuel temperature upon entry into the at least one actuator is at least 5 °C higher than the fuel temperature upon entry into the combustion chamber.

[0120] The bypass ratio can range from 4 to 55. The bypass ratio can range from 4 to 20. The bypass ratio can range from 4 to 15.

[0121] The method may include determining at least one fuel property of the fuel and controlling a temperature difference between the fuel temperature entering the at least one actuator and the fuel temperature entering the combustion chamber based on the at least one fuel property.

[0122] The fuel supply system may draw the fuel from a fuel source, for example an aircraft fuel tank (generally located external to the engine); and the method may be such that the fuel supplied to the at least one actuator for fuel-hydraulic drive of the at least one actuator is supplied to the combustion chamber after actuation without returning to the fuel source.

[0123] 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 so that, under cruise conditions, a heat transfer ratio of: Heat transfer rate from oil to fuel in the secondary fuel-oil heat exchanger (kJkg−1)Heat transfer rate from oil to fuel in the primary and secondary fuel-oil heat exchanger (kJkg−1) has a maximum value of at least 0.35.

[0124] 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 in the fuel.

[0125] 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.

[0126] It is understood that a feature described in relation to one aspect may, mutatis mutandis, also be used in combination with any other aspect.

[0127] As indicated elsewhere herein, the present disclosure may be applicable to any relevant configuration of a gas turbine engine. Such a gas turbine engine may be, for example, a turbofan gas turbine engine, an open-rotor gas turbine engine (in which the propeller is not enclosed by a nacelle), a turboprop engine, or a turbojet engine. Each such engine may or may not be provided with an afterburner. Such a gas turbine engine may be configured, for example, for land-based or marine power generation applications.

[0128] A gas turbine engine according to the present disclosure may include an engine core including 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 (with fan blades). Such a fan may be disposed upstream of the engine core. Alternatively, in some examples, the gas turbine engine may include a fan disposed downstream of the engine core, for example, if the gas turbine engine has an open rotor or is a turboprop engine (in which case the fan may be referred to as a propeller).

[0129] If the gas turbine engine has an open rotor or is a turboprop engine, the gas turbine engine may comprise two counter-rotating propeller stages connected by a shaft to and driven by a free power turbine. The propellers may rotate in opposite directions, so that one rotates clockwise and the other counterclockwise about the engine's axis of rotation. Alternatively, the gas turbine engine may comprise a propeller stage and a guide vane stage configured downstream of the propeller stage. The guide vane stage may have a variable pitch. Accordingly, high-pressure, intermediate-pressure, and free power turbines may each drive high- and intermediate-pressure compressors and propellers through suitable connecting shafts. The propellers may thus provide the majority of the propulsive thrust.

[0130] If the gas turbine engine has an open rotor or is a turboprop engine, one or more of the propeller stages may be driven by a gearbox. The gearbox may be of the type described herein.

[0131] An engine according to the present disclosure may be a turbofan engine. Such an engine may be a direct-drive turbofan engine in which the fan is connected directly to the fan drive turbine via a core shaft, for example, without a gearbox. In such a direct-drive turbofan engine, the fan may be arranged to rotate at the same speed as the fan drive turbine. For example only, 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 speed than the first core shaft. In such an arrangement, the second turbine may be positioned axially upstream of the first turbine.

[0132] An engine according to the present disclosure may be a geared turbofan engine. In such an arrangement, the engine includes a fan driven by a gearbox. Accordingly, such a gas turbine engine may include a gearbox that receives input from the core shaft and outputs drive to the fan to drive the fan at a lower speed than the core shaft. The input to the gearbox may be directly from the core shaft or indirectly from the core shaft, for example, via a spur gear shaft and / or a gear. The core shaft may rigidly connect the turbine and compressor so that the turbine and compressor rotate at the same speed (with the fan rotating at a lower speed).

[0133] 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 number of shafts connecting turbines and compressors, for example, one, two, or three shafts. For 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, the second compressor, and the second core shaft may be arranged to rotate at a higher speed than the first core shaft.

[0134] In such an arrangement, the second compressor may be positioned axially downstream of the first compressor. The second compressor may be arranged to receive flow from the first compressor (e.g., directly, e.g., via a generally annular channel).

[0135] The gearbox may be arranged to be driven by the core shaft configured to rotate (e.g., in use) at the lowest speed (e.g., the first core shaft in the example above). For example, the gearbox may be arranged to be driven only by the core shaft configured to rotate (e.g., in use) at the lowest speed (e.g., only by the first core shaft and not by the second core shaft in the example above). Alternatively, the gearbox may be arranged to be driven by any one or more shafts, for example, the first and / or second shafts in the example above.

[0136] The gearbox may be a reduction gearbox (in that the output to the fan has a lower speed than the input from the core shaft). Any type of gearbox may be used. For example, the gearbox may be a "planetary gearbox" or a "star gearbox," as described in more detail elsewhere herein. Such a gearbox may be single-stage. Alternatively, such a gearbox may be a compound gearbox, for example, a compound planetary gearbox (where the input is on the sun gear and the output is on the ring gear, and which is therefore also referred to as a "compound star" gearbox), e.g., with two reduction stages.

[0137] The transmission may have any desired reduction ratio (defined as the speed of the input shaft divided by the 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 order 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, for example, be between any two of the values ​​mentioned in the preceding sentence. By way of example only, the transmission may be a "star gear" with a gear ratio in the range of 3.1 or 3.2 to 3.8. Further by way of example only, the transmission may be a "star gear" with a gear ratio in the range of 3.0 to 3.1. Further, purely by way of example, the transmission may be a "planetary gear" with a gear ratio in the range of 3.6 to 4.2. In some arrangements, the gear ratio may be outside these ranges.

[0138] In any gas turbine engine described and / or claimed herein, fuel of a given composition or mixture is provided to a combustor, which may be provided downstream of the fan and compressor(s) with respect to the flowpath (e.g., axially downstream). For example, the combustor may be located directly downstream of the second compressor (e.g., at its exit), where a second compressor is provided. As another example, flow at the exit to the combustor may be provided to the inlet of the second turbine, where a second turbine is provided. The combustor may be provided upstream of the turbine(s).

[0139] The or each compressor (for example, the first compressor and the second compressor, as described above) may comprise any number of stages, for example multiple stages. Each stage may comprise a row of rotor blades and a row of stator blades, which may be variable stator blades (since 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 turbofan 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 another example, the gas turbine engine may be a "geared" gas turbine engine (where the fan is driven by a first core shaft via a reduction gear) that includes 11, 12, or 13 compressor stages (in addition to the fan). Such an engine may include 3 or 4 stages in the first (or "low-pressure") compressor and 8 or 9 stages in the second (or "high-pressure") compressor. As another example, the gas turbine engine may be a "geared" gas turbine engine with 4 stages in the first (or "low-pressure") compressor and 10 stages in the second (or "high-pressure") compressor.

[0140] The or each turbine (for example, the first turbine and the second turbine as described above) may comprise any number of stages, for example, multiple stages. Each stage may comprise one row of rotor blades and one row of stator blades, or vice versa, as required. The respective rows of rotor blades and stator blades may be axially offset from one another. The second (or "high pressure") turbine may comprise two stages in any arrangement (for example, regardless of whether it is a geared or direct drive engine). The gas turbine engine may be a direct drive turbofan engine comprising a first (or "low pressure") turbine with five, six, or seven stages. Alternatively, the gas turbine engine may be a "geared" gas turbine engine having a first (or "low pressure") turbine with three or four stages.

[0141] Each fan blade may be defined as having a radial span extending from a root (or hub) at a radially inner gas-swept or 0% span location to a tip at a 100% span location. The ratio of the fan blade radius at the hub to the fan blade radius at the tip may be less than (or on the order 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 radius of the fan blade at the hub to the radius of the fan blade at the tip 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 0.28 to 0.32 or 0.29 to 0.30. These ratios may be generally referred to as the hub-to-tip ratio.The hub radius and the tip radius can both be measured at the leading edge (or axially forwardmost part) of the blade. The hub-to-tip ratio, of course, refers to the gas-swept section of the fan blade, i.e., the section radially outward from any platform.

[0142] The radius of the fan can be measured between the centerline of the engine and the tip of a fan blade at its leading edge. The fan diameter (which can be simply twice the fan radius) can be larger than (or on the order 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 fan diameter can be within an inclusive range limited by any two of the values ​​in the previous sentence (i.e., the values ​​can be upper or lower limits), for example, in the range from 210 cm to 240 cm, or 250 cm to 280 cm, or 320 cm to 380 cm. Purely as a non-limiting example, the fan diameter can be in the range from 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.

[0143] The speed of the fan may vary during use. Generally, the speed is lower for fans with a larger diameter. Purely as a non-limiting example, the fan speed under cruise conditions may be less than 3500 rpm, for example, less than 2600 rpm, or less than 2500 rpm, or less than 2300 rpm. Purely as a further non-limiting example, the fan speed under cruise conditions for a geared gas turbine engine with a fan diameter in the range of 200 cm to 210 cm may be in the range of 2750 to 2900 rpm. Purely as a further non-limiting example, the fan speed under cruise conditions for a geared gas turbine engine with a fan diameter in the range of 210 cm to 230 cm may be in the range of 2500 to 2800 rpm.By way of further non-limiting example, the fan speed under cruise conditions for a geared gas turbine engine with a fan diameter in the range of 340 cm to 360 cm may be in the range of 1500 to 1800 rpm. By way of further non-limiting example, the fan speed under cruise conditions for a direct drive engine with a fan diameter in the range of 190 cm to 200 cm may be in the range of 3600 to 3900 rpm. By way of further non-limiting example, the fan speed under cruise conditions for a direct drive engine with a fan diameter in the range of 300 cm to 340 cm may be in the range of 2000 to 2800 rpm.

[0144] During operation of the gas turbine engine, the fan (with associated fan blades) rotates around a rotational axis. This rotation causes the tip of the fan blade to rotate at a speed U SpitzeThe work performed by the fan blades in the flow leads to an enthalpy increase dH of the flow. A fan peak load can be expressed as dH / U Spitze 2 where dH is the enthalpy rise (for example, the mean 1-D enthalpy rise) across the fan and U Spitzeis the (translational) fan tip velocity, for example, at the leading edge of the tip (which can be defined as the fan tip radius at the leading edge multiplied by the angular velocity). The fan tip loading under cruise conditions can be greater than (or on the order 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 ​​being dimensionless). The fan peak loading may lie within an inclusive range bounded by any two of the values ​​in the previous 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 (for example, for a geared gas turbine engine).

[0145] Gas turbine engines according to the present disclosure may have a desired bypass ratio (BPR), where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass channel to the mass flow rate of the flow through the core. In some arrangements, the bypass ratio under cruise conditions may be greater than (or of the order of) one 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 under cruise conditions 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 12 to 16, 13 to 15, or 13 to 14.By way of non-limiting example only, the bypass ratio under cruise conditions of a direct drive gas turbine engine according to the present disclosure may range from 9:1 to 11:1. By way of further non-limiting example only, the bypass ratio of a geared gas turbine engine according to the present disclosure may range from 12:1 to 15:1 under cruise conditions. The bypass duct may be at least substantially annular. The bypass duct may be located radially outward from the core engine. The radially outer surface of the bypass duct may be defined by a nacelle and / or a fan casing.

[0146] The total pressure ratio (OPR) of a gas turbine engine, as described and / or claimed herein, may be defined as the ratio of the ram pressure at the exit of the maximum pressure compressor (before entering the combustor) to the ram pressure upstream of the fan. As a non-limiting example, the total pressure ratio of a gas turbine engine, as described and / or claimed herein, under cruise conditions may be greater than (or on the order of) one of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The total pressure ratio 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 50 to 70.By way of non-limiting example only, the overall pressure ratio under cruise conditions of a geared gas turbine engine with a fan diameter in the range of 200 cm to 210 cm may be in the range of 40 to 45. By way of non-limiting example only, the overall pressure ratio under cruise conditions of a geared gas turbine engine with a fan diameter in the range of 210 cm to 230 cm may be in the range of 45 to 55. By way of non-limiting example only, the overall pressure ratio under cruise conditions of a geared gas turbine engine with a fan diameter in the range of 340 cm to 360 cm may be in the range of 50 to 60. By way of non-limiting example only, the overall pressure ratio under cruise conditions of a direct drive gas turbine engine with a fan diameter in the range of 300 cm to 340 cm may be in the range of 50 to 60.

[0147] The specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow through the engine. In some examples, the specific thrust for a given thrust condition may depend on the specific composition of fuel provided to the combustion chamber. Under cruise conditions, the specific thrust of an engine described and / or claimed herein may be less than (or on the order of) one of the following: 110 Nkg -1 s, 105 Nkg -1 s, 100 Nkg -1 s, 95 Nkg -1 s, 90 Nkg -1 s, 85 Nkg -1 s or 80 Nkg -1 s. The specific thrust may lie within an inclusive range limited by any two of the values ​​in the previous sentence (ie the values ​​may form upper or lower limits), for example in the range of 80 Nkg -1 s up to 100 Nkg -1 s or 85 Nkg -1 s up to 95 Nkg-1 s. Such engines can be particularly efficient compared to conventional gas turbine engines. Purely as a non-limiting example, the specific thrust of a geared gas turbine engine with a fan diameter in the range of 200 cm to 210 cm can be in the range of 90 Nkg -1 s up to 95 Nkg -1 s. Purely as a non-limiting example, the specific thrust of a geared gas turbine engine with a fan diameter in the range of 210 cm to 230 cm can be in the range of 80 Nkg -1 s up to 90 Nkg -1 s. Purely as a non-limiting example, the specific thrust of a geared gas turbine engine with a fan diameter in the range of 340 cm to 360 cm can be in the range of 70 Nkg -1 s up to 90 Nkg -1s. Purely as a non-limiting example, the specific thrust of a direct drive gas turbine engine with a fan diameter in the range of 300 cm to 340 cm can be in the range of 90 Nkg -1 s up to 120 Nkg -1 s lie.

[0148] A gas turbine engine as described and / or claimed herein may have any desired maximum thrust. Purely as a non-limiting example, a gas turbine as described and / or claimed herein may be capable of producing a maximum thrust of at least one of (or on the order 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). Purely as a non-limiting example, 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.By way of non-limiting example only, the maximum thrust of a geared gas turbine engine with a fan diameter in the range of 200 cm to 210 cm may be in the range of 140 kN to 160 kN. By way of non-limiting example only, the maximum thrust of a geared gas turbine engine with a fan diameter in the range of 210 cm to 230 cm may be in the range of 150 kN to 200 kN. By way of non-limiting example only, the maximum thrust of a geared gas turbine engine with a fan diameter in the range of 340 cm to 360 cm may be in the range of 370 kN to 500 kN. By way of non-limiting example only, the maximum thrust of a direct drive gas turbine engine with 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 thrust may be the maximum net thrust at normal atmospheric conditions at sea level at +15 degrees C (ambient pressure 101.3 kPa, temperature 30 degrees C) with static engine.

[0149] In use, the flow temperature at the inlet to the high-pressure turbine may be particularly high. This temperature, which may be referred to as the TET, may be measured at the exit to the combustor, for example, immediately upstream of the first turbine vane, which may itself be referred to as the nozzle vane. In some examples, the TET for a given thrust condition may depend on the specific composition of fuel supplied to the combustor. Under cruise conditions, the TET may be at least one of (or on the order of) the following: 1400 K, 1450 K, 1500 K, 1520 K, 1530 K, 1540 K, 1550 K, 1600 K, or 1650 K. Thus, purely as a non-limiting example, under cruise conditions of a geared gas turbine engine with a fan diameter in the range of 200 cm to 210 cm, the TET may be in the range of 1540 K to 1600 K.By way of non-limiting example only, the TET under cruise conditions of a geared gas turbine engine with a fan diameter in the range of 210 cm to 230 cm may be in the range of 1590 K to 1650 K. By way of non-limiting example only, the TET under cruise conditions of a geared gas turbine engine with a fan diameter in the range of 340 cm to 360 cm may be in the range of 1600 K to 1660 K. By way of non-limiting example only, the TET under cruise conditions of a direct drive gas turbine engine with a fan diameter in the range of 300 cm to 340 cm may be in the range of 1590 K to 1650 K. Purely as a non-limiting example, the TET under cruise flight conditions of a direct drive gas turbine engine with a fan diameter in the range of 300 cm to 340 cm may be in the range of 1570 K to 1630 K.

[0150] The TET under cruise conditions 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 1530 K to 1600 K. The maximum TET in use of the engine may, for example, be at least one of (or be of the order of magnitude of): 1700 K, 1750 K, 1800 K, 1850 K, 1900 K, 1950 K, 2000 K, 2050 K, or 2100 K. Thus, purely as a non-limiting example, the maximum TET of a geared gas turbine engine with a fan diameter in the range 200 cm to 210 cm may be in the range 1890 K to 1960 K. Purely as a non-limiting example, the maximum TET of a geared gas turbine engine with a fan diameter in the range of 210 cm to 230 cm may be in the range of 1890 K to 1960 K.By way of non-limiting example only, the maximum TET of a geared gas turbine engine with a fan diameter in the range of 340 cm to 360 cm may be in the range of 1890 K to 1960 K. By way of non-limiting example only, the maximum TET of a direct drive gas turbine engine with a fan diameter in the range of 300 cm to 340 cm may be in the range of 1935 K to 1995 K. By way of non-limiting example only, the maximum TET of a direct drive gas turbine engine with 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 previous sentence (i.e., the values ​​may form upper or lower limits), for example, in the range 1800 K to 1950 K or 1900 K to 2000 K.The maximum TET can occur, for example, in a high thrust condition, for example, in a maximum take-off (MTO) condition.

[0151] A fan blade and / or a vane portion of a fan blade as described and / or claimed herein may be made from any suitable material or combination of materials. For example, at least a portion of the fan blade and / or vane may be made at least partially from a composite material, for example, a metal matrix composite and / or an organic matrix composite, such as a carbon fiber composite. As another example, at least a portion of the fan blade and / or vane may be made at least partially 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 made using different materials.For example, the fan blade may include a protective leading edge, which may be fabricated using a material better able to withstand impact (e.g., from birds, ice, or other material) than the rest of the blade. Such a leading edge may, for example, be fabricated using titanium or a titanium-based alloy. Thus, purely by way of example, the fan blade may include a carbon fiber or aluminum-based body (such as an aluminum-lithium alloy) with a titanium leading edge.

[0152] 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 attached to the central portion in any desired manner. For example, each fan blade may include a fastener engageable with a corresponding slot in the hub (or disc). Such a fastener may, by way of example only, be in the form of a dovetail that is insertable and / or snap-fittable into a corresponding slot 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 blade disc or blade ring. Any suitable method may be used to manufacture such a blade disc or blade 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 attached to the hub / disk by welding, such as linear friction welding.

[0153] The gas turbine engines described and / or claimed herein may or may not be equipped with a variable area nozzle (VAN). Such a variable area nozzle may enable the bypass duct exit area to be varied during use. The general principles of the present disclosure may be applied to engines with or without a VAN.

[0154] The fan of a gas turbine as described and / or claimed herein may have any number of fan blades, for example, 14, 16, 18, 20, 22, 24, or 26 fan blades. If the fan blades have a carbon fiber composite body, there may be 16 or 18 fan blades. If the fan blades have a metallic body (for example, an aluminum-lithium or titanium alloy), there may be 18, 20, or 22 fan blades.

[0155] As used herein, the terms idle, taxi, takeoff, climb, cruise, descent, approach, and landing (or one or more parts thereof) have their usual meaning and are readily understood by one skilled in the art. Thus, for a particular gas turbine engine for an aircraft, one skilled in the art would immediately recognize that each term refers to all or one or more parts of a phase of engine operation within a given application of an aircraft to which the gas turbine engine is designed to be attached.

[0156] In this regard, ground idle may refer to a phase of engine operation in which the aircraft is stationary and in contact with the ground, but with a requirement that the engine be running. During idle, the engine may produce between 3% and 9% of the engine's available thrust. In further non-limiting examples, the engine may produce between 5% and 8% of the available thrust. In further non-limiting examples, the engine may produce between 6% and 7% of the available thrust. Taxi may refer to a phase of engine operation in which the aircraft is propelled along the ground by the thrust generated by the engine. During taxi, the engine may produce between 5% and 15% of the available thrust. In further non-limiting examples, the engine may produce between 6% and 12% of the available thrust.In further non-limiting examples, the engine may produce between 7% and 10% of the available thrust. Takeoff may refer to a phase of engine operation in which the aircraft is propelled by the thrust generated by the engine. In an initial stage within the takeoff phase, the aircraft may be propelled while the aircraft is in contact with the ground. In a later stage within 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 further non-limiting examples, the engine may produce between 95% and 100% of the available thrust. In further non-limiting examples, the engine may produce 100% of the available thrust.

[0157] Climb may refer to a phase of engine operation during which the aircraft is propelled by the thrust generated by the engine. During climb, the engine may produce between 75% and 100% of the available thrust. In further non-limiting examples, the engine may produce between 80% and 95% of the available thrust. In further non-limiting examples, the engine may produce between 85% and 90% of the available thrust. In this context, climb may refer to a phase of operation within an aircraft's flight cycle between takeoff and the achievement of cruise conditions, where the achievement of cruise conditions defines the beginning of the cruise phase or portion thereof of the aircraft's flight.Additionally or alternatively, the term climb may refer to a nominal point or one or more nominal periods during an aircraft's flight cycle between takeoff and landing during which a relative increase in altitude is required, which may require additional engine thrust.

[0158] As used herein, cruise conditions, which may define the cruise phase (or portion thereof) of aircraft flight, have the conventional meaning and are readily understood by those skilled in the art. In some examples, the cruise conditions for a particular gas turbine engine for an aircraft may refer to the engine's cruise 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 context, cruise can be considered the point in an aircraft's flight cycle at which 50% of the total fuel burned between the peak of climb and the start of descent has been burned (which can be approximated by the midpoint, in terms of time and / or distance, between the peak of climb and the start of descent). Cruise conditions can thus define an operating point, phase, or part of flight that provides thrust sufficient to support 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 an at least substantially constant altitude and / or an at least substantially constant Mach number) of an aircraft to which it is designed for attachment, taking into account the number of engines provided for that aircraft. For example, if an engine is designed to be attached to an aircraft having two engines of the same type, the engine may provide, under cruise conditions, half the total thrust that would be required for stable operation, or at least substantially stable operation, of the aircraft at cruise.

[0159] In other words, the cruise flight conditions for a given aircraft gas turbine engine can be defined as the engine operating point that provides a given thrust (required to provide, in combination with any other engines on the aircraft, stable operation, or at least substantially stable operation, of the aircraft to which it is designed for attachment at a given cruise Mach number) under cruise atmospheric conditions (defined by the international standard atmosphere according to ISO 2533 at a cruise altitude). For any given aircraft gas turbine engine, the cruise thrust, atmospheric conditions, and Mach number are known, and thus the engine operating point under cruise conditions can be clearly defined.

[0160] By way of example only, forward speed under cruise conditions may be any point in the range of Mach 0.7 to 0.9, for example, 0.75 to 0.85, for example, 0.76 to 0.84, for example, 0.77 to 0.83, for example, 0.78 to 0.82, for example, 0.79 to 0.81, for example, on the order of Mach 0.8, on the order of Mach 0.85, or in the range of 0.8 to 0.85. Any individual speed within these ranges may be part of the cruise conditions. For some aircraft, cruise conditions may be outside these ranges, for example, below Mach 0.7 or above Mach 0.9.

[0161] By way of example only, the cruise flight conditions may correspond to standard atmosphere 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 (about 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 (about 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 of 10,900 m to 11,100 m, for example in the order of 11,000 m. Cruise conditions can be equivalent to standard atmospheric conditions at any given altitude in these areas.

[0162] For 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). Under such cruise conditions, the engine can provide a known required net thrust level. The known required net thrust level naturally depends on the engine and its intended use and may, for example, be a value in the range of 20 kN to 40 kN.

[0163] Further, purely by way of 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). Under such cruise conditions, the engine may provide a known required net thrust level. The known required net thrust level naturally depends on the engine and its intended use and may, for example, be a value in the range of 35 kN to 65 kN.

[0164] In use, a gas turbine engine as described and / or claimed herein may operate under the cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions (e.g., the conditions at cruise) of an aircraft on which at least one (e.g., two or four) gas turbine engines may be mounted to provide propulsive thrust.

[0165] Furthermore, one skilled in the art would readily recognize that descent and approach, or both, refer to an operational phase within a flight cycle between cruise and landing of the aircraft, with the approach, in particular, forming part of the landing and takeoff (LTO) phase. During one or both of descent and approach, the engine may produce between 0% and 50% of available thrust. In further non-limiting examples, the engine may produce between 25% and 40% of available thrust. In further non-limiting examples, the engine may produce between 30% and 35% of available thrust. Additionally or alternatively, the term descent may refer to a nominal point in an aircraft's flight cycle between takeoff and landing at which a relative decrease in altitude is required and which may require a reduced thrust requirement from the engine.

[0166] According to one aspect, an aircraft comprising a gas turbine engine as described and / or claimed herein is provided. The aircraft according to this aspect is the aircraft for which the gas turbine engine is designed to be mounted. Accordingly, the cruise flight conditions according to this aspect may correspond to an operating point, phase, or portion thereof of the aircraft's flight, as defined elsewhere herein.

[0167] According to one aspect, a method of operating a gas turbine engine as described and / or claimed herein is provided. Operation may occur under any suitable conditions, as may be defined elsewhere herein (e.g., with respect to thrust, atmospheric conditions, and Mach number).

[0168] According to one aspect, a method of operating an aircraft comprising a gas turbine engine as described and / or claimed herein is provided. Operation according to this aspect may include (or be) operation under any suitable condition, for example, during cruise of the aircraft, as defined elsewhere herein.

[0169] Those skilled in the art will understand that, except in cases of mutual exclusion, a feature or parameter described with respect to one of the above aspects may be applied to any other aspect. Furthermore, except in cases of mutual exclusion, any feature or parameter included or described herein may be applied to any aspect and / or combined with any other feature or parameter included or described herein.

[0170] Unless mutually exclusive, each parameter or value included or described herein may be applied to and / or combined with one or more further parameters and / or values ​​included or described herein. For example, a first parameter or value included or described herein (e.g., Parameter A) may be applied to and / or combined with one or more further parameters and / or values ​​included or described herein (e.g., one or more Parameter B; Parameter C; and Parameter D, etc.) to express a product of their relationship. For example, one of skill in the art would understand that where Parameter A is disclosed separately from Parameter B, a product of their relationship may be expressed, for example, as A / B, B / A, B*A, or any other such application, combination, or function of Parameter A with respect to Parameter B, as appropriate.

[0171] Embodiments will now be described by way of example only with reference to the figures in which: Fig. Figure 1 is a cross-sectional side view of a gas turbine engine; Fig. Figure 2 is a close-up cross-sectional side view of an upstream portion of a geared gas turbine engine; Fig. Figure 3 is a partial sectional view of a gearbox for a gas turbine engine; Fig. Figure 4 is a close-up cross-sectional side view of an upstream portion of a direct drive gas turbine engine; Fig. Figure 5 is an illustration of an aircraft having a propulsion system including two gas turbine engines; Fig. 6 is a schematic diagram of a fuel supply system including a fuel hydraulic actuator; Fig. Figure 7 is a schematic diagram of a fuel supply system including two fuel hydraulic actuators; Fig. 8 is a schematic representation of a fuel supply system including a selective fuel hydraulic actuator; Fig. 9 is a schematic illustration of an alternative fuel supply system including a selective fuel hydraulic actuator; Fig. 10 is a schematic representation of another fuel supply system comprising two fuel hydraulic actuators, one of which is a selective fuel hydraulic actuator; Fig. 11 is a schematic diagram of a fuel supply system including two selective fuel hydraulic actuators; Fig. 12 is a schematic representation of another fuel supply system including two selective fuel hydraulic actuators; Fig. 13 is a schematic representation of another fuel supply system including a selective fuel hydraulic actuator; Fig. 14 is a schematic representation of another fuel supply system including a selective fuel hydraulic actuator; Fig. 15 is a schematic representation of a fuel supply system including a selective fuel hydraulic actuator operable to control a valve within an engine thermal management system; Fig. 16 is a schematic illustration of another fuel supply system including a selective fuel hydraulic actuator operable to control a valve within an engine thermal management system; Fig. 17 is a schematic illustration of a fuel supply system including a selective fuel hydraulic actuator operable to control a valve within a generator thermal management system; Fig. 18 is a schematic illustration of another fuel supply system including a selective fuel hydraulic actuator operable to control a valve within a generator thermal management system; Fig. 19 is a schematic diagram of a fuel supply system including a selective fuel hydraulic actuator and two heat exchangers; Fig. 20 is a flowchart illustrating an exemplary method for operating a gas turbine engine; Fig. 21 is a flowchart illustrating another exemplary method for operating a gas turbine engine; Fig. 22 is a flowchart illustrating another exemplary method for operating a gas turbine engine; Fig. 23 is a flowchart illustrating another exemplary method for operating a gas turbine engine; Fig. 24 is a flowchart illustrating another exemplary method for operating a gas turbine engine; and Fig. 25 illustrates an oil supply system for a gas turbine engine comprising two fuel-oil heat exchangers.

[0172] Fig. 1 illustrates a gas turbine engine 10 having a main axis of rotation 9. The engine 10 includes an air inlet 12 and a propulsion fan 23 that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 includes a core 11 that receives the core airflow A. The engine core 11 comprises, in axial flow sequence, 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 the low-pressure turbine 19 via a shaft 26 and an epicyclic gear train 30 and is driven thereby.

[0173] In use, the core air stream A is accelerated and compressed by the low-pressure compressor 14 and passed into the high-pressure compressor 15 where it is further compressed. The compressed air discharged from the high-pressure compressor 15 is passed 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 a combustor 16, with the terms "combustion equipment 16" and "combustor 16" being used interchangeably herein. The resulting hot combustion products then expand through, 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 via a suitable connecting shaft 27.Fan 23 generally acts to apply increased pressure to the bypass airflow B flowing through bypass duct 22, so that bypass airflow B is expelled through bypass exhaust nozzle 18 to generally provide the majority of the propulsive thrust. Epicyclic gearing 30 is a reduction gear.

[0174] An exemplary arrangement for a geared turbofan 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 central pinion or sun gear 28 of the epicyclic gear assembly 30. Radially outward from and meshing with the sun gear 28 are a plurality of planet gears 32 coupled together by a planet carrier 34. The planet carrier 34 constrains the planet gears 32 to orbit synchronously about the sun gear 28 while enabling each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled to the fan 23 via linkages 36 to drive its rotation about the engine axis 9. Radially outward from and meshing with the planet gears 32 is a ring gear or ring gear 38 coupled to a stationary support structure 24 via linkages 40.

[0175] It should be noted that the terms "low pressure turbine" and "low pressure compressor," as used herein, mean the lowest pressure turbine stages and the lowest pressure compressor stages, respectively (i.e., not including the fan 23) and / or the turbine and compressor stages connected by the lowest speed connecting shaft 26 in the engine (i.e., not including the gearbox output shaft that drives the fan 23). In some references, the "low pressure turbine" and "low pressure compressor" referred to herein may alternatively be known as the "intermediate pressure turbine" and "intermediate pressure compressor." Where such alternative nomenclature is used, the fan 23 may be referred to as a first or lowest pressure compression stage.

[0176] The epicyclic gear 30 is exemplary in Fig. 3 in more detail. Each of the sun gear 28, the planet gears 32 and the ring gear 38 includes teeth on its circumference to mesh with the other gears. For clarity, Fig. However, only exemplary portions of the teeth are shown in Figure 3. Four planetary gears 32 are illustrated, although it will be apparent to one skilled in the art that more or fewer planetary gears 32 may be provided. Practical applications of a planetary gear train 30 generally include at least three planetary gears 32.

[0177] This is exemplified in the Fig. 2 and Fig. The epicyclic gear train 30 illustrated in Figure 3 is of the planetary type, as the planet carrier 34 is coupled to an output shaft via linkages 36, with the ring gear 38 being fixed. However, any other suitable type of epicyclic gear train 30 may be used. As another example, the epicyclic gear train 30 may be a star arrangement, in which the planet carrier 34 is held fixed and the ring gear (or ring gear) 38 is allowed to rotate. In such an arrangement, the fan 23 is driven by the ring gear 38. As another alternative example, the train 30 may be a differential gear train, in which both the ring gear 38 and the planet carrier 34 are allowed to rotate.

[0178] It is understood that the Fig. 2 and Fig. 3 is merely exemplary and that various alternatives are within the scope of the present disclosure. By way of example only, any suitable arrangement for housing the transmission 30 in the engine 10 and / or for connecting the transmission 30 to the engine 10 may be used. As another example, the connections (such as the linkages 36, 40 in the example of Fig. 2) between the transmission 30 and other parts of the engine 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 engine (for example, between the input and output shafts of the transmission and the fixed structures, such as the transmission housing) may be used, and the disclosure is not limited to the exemplary arrangement of Fig. 2. For example, if the gearbox 30 has a star arrangement (described above), one skilled in the art would readily understand that the arrangement of output and support rods and bearing positions would typically be different from that shown in Fig. 2 would deviate from the arrangement shown as an example.

[0179] Accordingly, the present disclosure extends to a gas turbine engine having any arrangement of gear types (for example, star or planetary gears), support structures, input and output shaft arrangement, and bearing positions.

[0180] Optionally, the gearbox can drive additional and / or alternative components (e.g. the intermediate pressure compressor and / or a booster compressor).

[0181] 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 compressors and / or turbines and / or an alternative number of connecting shafts. As a further example, the Fig. 1 has a split-flow nozzle 18, 20, meaning that the flow through the bypass duct 22 has its own nozzle 18 separate from and located radially outward from 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. One or both nozzles (whether mixed or split flow) may have a fixed or variable area.

[0182] As a further example, other gas turbine engines to which the present disclosure may be applied may not have a gearbox for the main shaft(s) and may instead be direct drive engines. A cross-sectional view of such an engine is shown in Fig. 4 shown.

[0183] With reference to Fig. 4 is a gas turbine engine generally designated 10 and having a main axis of rotation 9. The engine 10 comprises, in axial flow series, an air inlet 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 core 10 and defines both the inlet 12 and the exhaust nozzle 20.

[0184] In use, air entering inlet 12 is accelerated by fan 23 to create two airflows: a core airflow A and a bypass airflow B. Core airflow A flows into intermediate pressure compressor 14, and bypass airflow B passes through a bypass duct 22 to provide motive thrust. Intermediate pressure compressor 14 compresses airflow A before discharging the air to high pressure compressor 15, where further compression occurs.

[0185] 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 a combustor 16, with the terms "combustion equipment 16" and "combustor 16" being used interchangeably herein. The resulting hot combustion products then expand through, 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 turbine 17, the intermediate-pressure turbine 19a, and the low-pressure turbine 19 each drive the high-pressure compressor 15, the intermediate-pressure compressor 14, and the fan 23, respectively, through a suitable connecting shaft.

[0186] 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 connecting shafts (e.g., two) and / or an alternative number of compressors and / or turbines. Further, the engine may include a gearbox provided in the driveline from a turbine to a compressor and / or a fan.

[0187] While the described example refers to a turbofan engine, the disclosure may be applied to any type of gas turbine engine, such as an open rotor (where the fan stage is not surrounded by a nacelle) or a turboprop engine. In some arrangements, the gas turbine engine 10 may not include a gearbox 30.

[0188] The geometry of the gas turbine engine 10 and components thereof is defined by a conventional axis system comprising an axial direction (aligned with the rotational axis 9), a radial direction (in the direction from bottom to top in Fig. 1) and a circumferential direction (perpendicular to the side in the view of Fig. 1). The axial, radial, and circumferential directions are perpendicular to each other.

[0189] The fuel F provided 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, the fuel F, when blended with or replaced by an alternative fuel, may comprise renewable hydrocarbons produced from biological or non-biological resources, also known as sustainable aviation fuel (SAF). In each of the provided examples, the fuel F may include one or more trace elements, including, for example, sulfur, nitrogen, oxygen, inorganics, and metals.

[0190] For example, SAF is understood by those skilled in the art to mean a biofuel, a renewable aviation fuel, a renewable jet fuel, an alternative fuel, or a biojet fuel, which are produced from biological or non-biological resources. SAF is typically 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 waste oils and fats; municipal solid waste; cellulosic waste (such as corn stalks); cover crops such as camelina, carinata, and speedwell; non-biogenic alternative fuels; jatropha; halophytes, and algae, rather than from fossil hydrocarbons. SAF is understood to encompass no fossil fuels.

[0191] The functional performance of a particular fuel composition or fuel blend F for use in a particular application may be defined, at least in part, by the ability of the fuel to serve the Brayton cycle of the gas turbine engine 10. Parameters that define functional performance may include, for example, specific energy; energy density; thermal stability; and emissions, including gaseous and / or particulate matter. In this context, particulate emissions may include soot particles produced during the combustion of the fuel F, also referred to as non-volatile particulate matter (nvPM). Any references herein to soot or smoke apply equally to other types of particulate emissions known in the art.Gaseous emissions may include one or more of the following elements: 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) produced during the combustion of the fuel F. All references herein to gaseous emissions apply equally to other types of gaseous emissions known in the art.

[0192] A relatively higher specific energy (i.e., energy per unit mass), expressed as MJ / kg, can 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 as MJ / l, can at least partially reduce takeoff fuel volume, which may be particularly important for volume-limited or single-refueling military operations. A relatively higher thermal stability (i.e., inhibition of fuel degradation or coking under thermal stress) can allow the fuel to maintain elevated temperatures in the engine and fuel injectors, potentially providing relative improvements in combustion efficiency.Reduced emissions, including particulate matter, can enable reduced contrail formation while reducing the environmental impact of a given mission. Other fuel properties can also be central to functional performance. For example, a relatively lower freezing point (°C) can enable long-range missions to optimize flight profiles; minimum aromatic concentrations (%) can ensure sufficient swelling of certain materials used in the construction of O-rings and seals previously exposed to fuels with high aromatic content; and maximum surface tension (mN / m) can ensure sufficient spray dispersion and atomization of the fuel.

[0193] The ratio of the number of hydrogen atoms to the number of carbon atoms in a molecule can affect the specific energy of a given fuel composition or blend. Fuels with higher ratios of hydrogen atoms to carbon atoms can have higher specific energies in the absence of bond strain. For example, fossil-based hydrocarbon fuels may comprise molecules of approximately 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.

[0194] 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 between 10% and 50% sustainable aviation fuel (with the remainder comprising one or more fossil-based hydrocarbon fuels, such as kerosene), with additional blends pending approval. However, within the aviation industry, sustainable aviation fuel blends comprising up to (and including) 100% sustainable aviation fuel (SAF) are expected to eventually be approved for use.

[0195] Sustainable aviation fuels may comprise one or more of n-alkanes, isoalkanes, cycloalkanes, and aromatics, and may, for example, be made 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 levels of aromatics and sulfur, or both, compared to fossil hydrocarbon fuels. Additionally or alternatively, sustainable aviation fuels may comprise higher levels of isoalkanes and cycloalkanes, or both, compared to fossil hydrocarbon fuels. In some examples, sustainable aviation fuels may comprise one or both of a density of between 90% and 98% of that of kerosene and a calorific value of between 101% and 105% of that of kerosene.

[0196] In some examples, the sustainable aviation fuel(s) or blend(s) provided to combustion equipment 16 may have a relatively lower content of aromatic and / or other non-paraffinic components than kerosene. The sustainable aviation fuel may include an aromatic content of, for example, 30%, 20%, 15%, 10%, 8%, 5%, or less than 5%; e.g., 4%, 3%, 2%, 1%, or less than 1%; e.g., 0.75%, 0.5%, 0.25%, or less than 0.25%; e.g., 0.2%, 0.1%, or less than 0.1%; e.g., 0.01%, 0.001%, or 0%. The aromatics content of the sustainable aviation fuel may be within an inclusive value or range limited by or within any two of the values ​​mentioned in the previous sentence (i.e. the values ​​may form upper or lower limits), e.g.13.5%, 8.5%, 2.5%, 0.35%, 0.15%, 0.05%, 0.005% or 0%; or 0% to 0.75%, 0% to 0.5% or 0.1% to 0.25%; or 0.15% to 0.65%, 0.35% to 0.55% or 0.035% to 0.055%; depending on one or more of preference, fuel stock or supplier and variations in the composition thereof.

[0197] Due at least in part to the molecular structure of sustainable aviation fuels, sustainable aviation fuels may provide benefits 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; less NOx; and lower CO2 emissions compared to fossil hydrocarbon fuels (e.g., when burned in combustion equipment 16). Accordingly, sustainable aviation fuels may result in one or both of a relative decrease in specific fuel consumption and a relative decrease in maintenance costs compared to fossil hydrocarbon fuels such as kerosene.

[0198] As in Fig. 5, an aircraft 1 may include multiple 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 a single fuel tank 50 and / or the wing fuel tanks 53a, b may be larger than the central fuel tank 50, or there may be no central fuel tank 50 (with all fuel instead being stored in the wings of the aircraft) - it is understood that many different tank layouts are envisaged and that the illustrated examples are provided for ease of description and are not intended to be limiting.

[0199] Fig. 5 shows an aircraft 1 having 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 fluidly connected to one another. Each fuel source is arranged to provide a separate fuel source (i.e., a first fuel source may contain a first fuel that has a different property or properties than a second fuel contained in a second fuel source).The first and second fuel sources are therefore not fluidly connected to separate the different fuels (at least under normal operating conditions). The use of multiple fuel sources allows an aircraft 1 to carry several different fuels and to switch the fuel used during operation, and optionally even during cruise or when changing between different operating modes in flight.

[0200] In the present example, the first (and in these examples, only) fuel source comprises a center fuel tank 50, which is located primarily in the fuselage of the aircraft 1, and a plurality of wing fuel tanks 53a, 53b, with at least one wing fuel tank located in the port wing and at least one wing fuel tank located in the starboard wing for balancing purposes. In the example shown, all tanks 50, 53a, b are fluidly connected to one another, thereby forming a single fuel source. Each of the center fuel tanks 50 and the wing tanks 53a, b may comprise a plurality of fluidly connected fuel tanks.

[0201] In another example, the wing fuel tanks 53a, 53b may not be fluidly connected to the center tank 50 and thus form a separate, second fuel source. For balancing purposes, one or more fuel tanks in the port wing may be fluidly connected to one or more fuel tanks in the starboard wing. This may be done either via a center fuel tank (provided that tank is not part of the other fuel source), bypassing the center fuel tank(s), or both (for maximum flexibility and safety). In another example, the first fuel source comprises wing fuel tanks 53 and a center fuel tank 50, while a second fuel source comprises another separate center fuel tank.To balance the aircraft 1, a fluid connection may be provided between the wing fuel tanks and the center fuel tank of the first fuel source. For aircraft 1 with multiple fuel sources, two or more fuel sources may therefore contain different fuels, allowing the aircraft 1 to switch fuels during flight. Being able to determine which fuel is being supplied to the combustion chamber 16 may therefore be more complex than simply recording a single fuel identity on board the aircraft 1 or checking it once upon takeoff.

[0202] In some examples, the allocation of the fuel tanks 50, 53 available in the aircraft 1 may be restricted such that the first fuel source and the second fuel source are each substantially symmetrical with respect to the aircraft's centerline. In cases where asymmetric fuel tank allocation is permitted, a suitable means of fuel transfer is typically provided between the fuel tanks of the first fuel source and / or between the fuel tanks of the second fuel source so that the position of the aircraft's center of gravity can be maintained within acceptable lateral limits throughout the flight.

[0203] Aircraft typically refuel at several different airports, for example, at the beginning and end of a long-haul flight. As mentioned above, while there are standards that all aviation fuels must comply with, different aviation fuels have different compositions, for example, depending on their source (e.g., various petroleum sources, biofuels or other synthetic aviation fuels (often referred to as sustainable aviation fuels - SAFs) and / or blends of petroleum-based fuels and other fuels) and the additives included (e.g., antioxidants and metal deactivators, biocides, antistatic agents, anti-icing agents, corrosion inhibitors) and contaminants. The composition of available aviation fuel can vary not only from airport to airport and from fuel supplier to fuel supplier, but even from batch to batch from the same airport or fuel supplier.Furthermore, the fuel tanks 50, 53 of the aircraft 1 are generally not emptied before being refilled for another flight, resulting in mixtures of different fuels in the tanks—and the mixture effectively resulting in a fuel with a different composition. One or more fuel connections 62 may be provided for refueling.

[0204] Fig. 6 shows a schematic diagram illustrating a portion of a fuel supply system 152 of a gas turbine engine 10. The fuel supply 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 branching fuel paths that withdraw fuel from or return fuel to this main fuel flow path 110.

[0205] The main fuel flow path 110 of the fuel supply system 152 serves as a fuel source for fuel hydraulic systems. The gas turbine engine 10 includes a combustion chamber 16 and an actuator 154. The dashed arrows in Fig. 6 represent the direction of the fuel flow. One or more fuel lines 159, 160 are provided to supply fuel to an actuator 154 to actuate it hydraulically.

[0206] The fuel supply system 152 is configured to supply fuel for combustion in the combustion chamber 16.

[0207] It is understood that, among other engine components, one or more valves, pumps, sensors, vents for fuel hydraulic actuators, and heat exchangers may be present along the main fuel flow path 110. The fuel is supplied directly to the combustion chamber 16 via line 110a after passing through the engine component located downstream on the main fuel flow path 110.

[0208] The fuel flowing through lines 159 and 160 is used to drive (i.e., actuate) actuator 154. As such, actuator 154 is fuel-hydraulic driven. As used herein, the term "fuel-hydraulic" refers to the hydraulic operation of an actuator where the hydraulic fluid is fuel. Actuators that are fuel-hydraulic driven may be referred to herein as "fuel-hydraulic actuators."

[0209] In the Fig. 6, the fuel flows from the fuel flow path 110 to the actuator 154 via the line 159. After the fuel has been used to drive the actuator 154 hydraulically, it is returned to the main fuel flow path 110 via the line 160 and can then be directed to the combustion chamber 16.

[0210] The fuel return line 160 may include a valve configured to regulate the flow of fuel back to the main fuel path 110, and the valve itself may be operated using a fuel hydraulic actuator.

[0211] The fuel return line 160 may direct fuel to a location along the main fuel flow path 110 that is upstream or downstream of one or more heat exchangers of the main fuel flow path 110.

[0212] The fuel return line 160 may return fuel to the fuel tank. In such implementations, the valve may be a fuel-to-tank return valve. The fuel-to-tank return valve may be controlled using a fuel-to-tank return actuator. The fuel supply system may be configured to supply fuel to fuel-hydraulically drive the fuel-to-tank return actuator.

[0213] Once the fuel reaches the combustion chamber 16, it is burned to provide thrust, as described with respect to Fig. 1 explained.

[0214] Fig. 6 shows a single fuel hydraulic actuator 154 in the fuel flow circuit 159, 160. In alternative implementations, multiple fuel hydraulic actuators 154 may be located in a single fuel flow circuit 159, 160 and utilize fuel drawn 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 (so that all of the fuel in the bleeder tube 159 flows sequentially through all actuators) or arranged in parallel with respect to the fuel flow (with a branching bleeder tube 159 and one or more actuators on each branch).

[0215] Fig. Figure 7 shows a schematic diagram illustrating an alternative portion of a fuel supply system 152 of a gas turbine engine 10. The fuel supply 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 combustion chamber 16. The fuel supply system 152 shown includes two actuators 254a, b, each located on a separate fuel extraction circuit of the main fuel flow path 110. The dashed arrows in Fig. 7 represent the direction of fuel flow.

[0216] Both actuators 254a,b are fuel-hydraulically driven. That is, both actuators 254a,b are hydraulically actuated, using fuel from the fuel source 50, 53 as the hydraulic fluid. The fuel flows from the main fuel flow path 110 via lines 259a,b to the actuators 254a,b, respectively. After the fuel has been used to drive the actuators 254a,b, it returns to the main fuel flow path 110 via lines 260a,b.

[0217] In Fig. 7, only two actuators 254a,b are shown, but in other implementations, any plurality of actuators may be fuel-hydraulically driven. In some implementations, the gas turbine engine 10 may include a plurality of fuel-hydraulically driven actuators and a plurality of non-fuel-hydraulically driven actuators.

[0218] Fig. 7 shows a single fuel hydraulic actuator 254 on each fuel flow circuit. In alternative implementations, multiple fuel hydraulic actuators 254 may be located on one or more of the multiple fuel flow circuits. The multiple actuators may be arranged in series with respect to the fuel flow (so that all of the fuel in the bleeder tube 159 flows sequentially through all actuators) or arranged in parallel with respect to the fuel flow (with a branching bleeder tube 159 and one or more actuators on each branch).

[0219] In various implementations, the gas turbine engine 10 includes at least ten, eleven, twelve, thirteen, fourteen, fifteen, sixteen, seventeen, or eighteen fuel hydraulic actuators. In some implementations, each actuator may have its own dedicated fuel bleed pipe 159, 259 from the main fuel flowpath 110, thus having its own fuel flow circuit. In other arrangements, two or more of these actuators may be located on the same fuel flow circuit outside the main fuel flowpath 110. In some arrangements, most or all of the fuel hydraulic actuators may be located on the same fuel circuit outside the main fuel flowpath 110.

[0220] In various implementations, 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 in conjunction with one another to control a particular system or perform a particular function within the engine 10. An actuated system may be described as "fuel hydraulic" if at least one of its actuators is fuel hydraulic.

[0221] Fig. 8 shows a fuel supply system 152 that includes an actuator 354.

[0222] Fig. Figure 8 shows a schematic diagram illustrating an alternative portion of a fuel supply system 152 of a gas turbine engine 10. The fuel supply 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 combustion chamber 16. The fuel supply system 152 shown includes an actuator 354 on a fuel withdrawal circuit from the main fuel flow path 110. The dashed arrows in Fig. 8 represent the direction of fuel flow.

[0223] The fuel supply system 152 of Fig. 8 includes a valve 304. Valve 304 is configured to direct fuel via line 361 or line 360. Line 306 includes actuator 354, such that fuel directed via line 306 fuel-hydraulically drives actuator 354 and is then returned to the main fuel flowpath 110. Fuel directed via line 361 returns directly to the main fuel flowpath 110 without driving an actuator. Valve 304 itself may be controlled using another fuel-hydraulic actuator.

[0224] The valve 304 is controllable, which means that the fuel supply system 152 is configured to be controlled to select between: • Causing the fuel to actuate actuator 354 (fuel flows via line 360); and • Causing fuel to bypass the at least one actuator 354 (fuel flows via line 361).

[0225] In some implementations, valve 304 is controlled via an actuator according to the output from a processor or other control device. Valve 304 may be controlled based on the Sustainable Aviation Fuel (SAF) content in the fuel. In some implementations, valve 304 may be replaced by any suitable alternative for regulating fuel flow between two lines 360, 361.

[0226] The fuel supply system 152 may be configured to select between: • Causing the fuel to actuate actuator 354 (via line 360); and • Causing the fuel to bypass the at least one actuator 354 (via line 361).

[0227] The selection may be based on whether the content of sustainable aviation fuel (SAF) in the fuel exceeds a threshold or not.

[0228] For example, the fuel supply system 152 may be configured to cause the fuel to actuate the actuator 354 (via line 360) when the SAF content of the fuel exceeds a concentration of 25 vol%, 30 vol%, 35 vol%, 40 vol%, 45 vol%, 50 vol%, 55 vol%, 60 vol%, 65 vol%, 70 vol%, 75 vol%, 80 vol%, 85 vol%, 90 vol%, or 95 vol%.

[0229] The fuel supply system may control the fuel in a binary manner so that all fuel reaching valve 304 is directed into line 360 ​​or line 361. In other implementations, the fuel may be split between line 360 ​​and line 361, for example.

[0230] To determine the SAF content of the fuel, the fuel supply system 152 may include a sensor. The sensor may be positioned immediately upstream of the valve 304. In other implementations, the sensor may be positioned upstream of the valve 304 within the main fuel flowpath 110. In alternative arrangements, the SAF content of the fuel is measured, calculated, or otherwise provided as the fuel is introduced into the fuel source 50, 53. For example, information regarding the SAF content of one or more fuels stored in the aircraft 1 may be communicated electronically to an onboard control unit or other computer system, or manually entered into a user interface of that system. It is understood that any suitable approach known in the art for determining the SAF content of a fuel may be used.

[0231] Fig. Figure 9 shows a schematic diagram illustrating an alternative portion of a fuel supply system 152 of a gas turbine engine 10. The fuel supply system 152 is similar to the fuel supply system 152 of Fig. 8, but line 361, which returned the fuel to the main fuel flow path 110, is replaced by line 362, which directs the fuel to the combustion chamber 16. The valve 304 directs the fuel via one of lines 360 and 362. Line 362 directs the fuel to bypass the actuator 354 and reach the combustion chamber 16 without returning to the main fuel flow path 110. 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.

[0232] In the implementations of the Fig. 8 and Fig. 9, the actuator 354 may be configured such that the actuator 354 is driven non-fuel hydraulically when the valve 304 directs fuel to bypass the actuator 354. For example, the actuator 354 may be driven using a non-fuel-based hydraulic fluid or using electrical signals.

[0233] Fig. 10 shows a schematic diagram illustrating an alternative portion of a fuel supply system 152 of a gas turbine engine 10. The fuel supply 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 combustion chamber 16. The fuel supply system 152 shown includes two actuators 454a, b, each actuator 454a, b positioned on a respective fuel withdrawal circuit from the main fuel flow path 110. The dashed arrows in Fig. 10 represent the direction of fuel flow.

[0234] The first actuator 454a is positioned downstream of a valve 404 located in its corresponding fuel withdrawal circuit. The valve 404 is configured to direct fuel via line 461 to bypass the actuator (and return directly to the main fuel flow path 110), or via line 460a to fuel-hydraulically actuate the actuator 454a (and then return to the main fuel flow path 110). The valve 404 is similar to the valve 304 of Fig. 8, but in other implementations, the valve 404 may be similar to the valve 304 of Fig. 9, so that the fuel bypassing the actuator is instead directed directly to the combustion chamber 16. In still further implementations, all of the fuel taken from the main fuel flow path 110 and supplied to the valve 304, 404 may be directed to the combustion chamber 16, either directly or after passing through the actuator 354, 454a. The actuator 454a is therefore selectively driven by fuel hydraulics.

[0235] In still further implementations, the selectively fuel-hydraulically driven valve 354, 454a may not have multiple fuel routes associated with it—instead, the valve 304, 404 may be arranged to prevent any fuel from leaving the main fuel flow path and flowing toward that actuator when the actuator 354, 454a is to be non-fuel-hydraulically driven. Flow through a simple circuit 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, in implementations where the same fuel takeoff point from the main fuel flow path 110 is used to feed multiple fuel hydraulic actuators, a controllable valve arranged to allow or prevent fuel from reaching a particular actuator can either (i) divert the fuel intended for that actuator along a dedicated bypass route when that actuator is to be actuated non-fuel hydraulically, or (ii) simply block any route to that actuator, thereby causing the fuel that would otherwise have flowed to that actuator to flow to one or more other actuators fed from the same takeoff point.The decision to operate an actuator non-fuel hydraulically may therefore increase the fuel flow rate to one or more other actuators unless the total flow through the drain portion of the fuel flow lines is also controlled.

[0236] The second actuator 454b of the example from Fig. 10 is in a separate fuel flow circuit to the first actuator 454a, and the fuel is drawn from a different point in the main fuel flow path 110. The second actuator 454b is arranged to be continuously fuel-hydraulically driven (it has only a single actuation mode).

[0237] Although the implementation of Fig. While FIG. 10 shows an actuator 454a that is selectively driven and an actuator 454b that is always fuel-hydraulic driven, in other implementations, the fuel supply system 152 may include any number of actuators, and each of these actuators may be fuel-hydraulic, non-fuel-hydraulic, or selectively fuel-hydraulic. Any number of actuators may be included in the same fuel flow circuit.

[0238] The actuators described above may be any actuators used by the engine 10. Various specific actuators that may be fuel-hydraulically driven are described below. The actuation systems described herein may be any actuation system used by the aircraft. Various specific actuation systems that may be fuel-hydraulically driven are described below. It should be understood that the specific implementations listed are described only by way of example, and that the scope of protection is limited only by the claims.

[0239] An example of an actuated system is a turbine casing cooling (TCC) system. Turbine engines 10 typically include a casing surrounding the turbine 17, 19, and a TCC system is used to selectively cool this casing. This cooling can be used to control the rotor-stator tip clearance by reducing thermal expansion. A TCC system includes one or more valves operable to modulate the bleed air flow from the engine's compressor 17, 19 through the casing to provide cooling.

[0240] For engines 10 with multiple turbines 17, 19, a TCC system may be provided for one, some or all of the turbines.

[0241] The one or more valves of the TCC system are operable 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 fuel-hydraulically driven. In one or more implementations, the fuel supply system may be configured to selectively bypass or fuel-hydraulically actuate one or more actuators of the TCC system.

[0242] Fig. Figure 11 shows a schematic diagram illustrating a portion of a fuel supply system 152 of a gas turbine engine 10. The fuel supply system 152 includes a TCC system 554 comprising two actuators 554a, b, each actuator 554a, b configured to control a corresponding valve 554a', b' of the TCC system. The valves 554a', b' can supply cool air to cool the turbine casing.

[0243] The fuel supply system 152 of Fig. 11 is configured to supply fuel from the fuel source 50, 53 to the combustion chamber 16 and to fuel-hydraulically drive the actuators 554a, b. As described above, the fuel supply system 152 is configured so that the actuators 554a, b can be selectively actuated or bypassed via associated valves 504a, b. In alternative implementations, a single valve may be used to selectively actuate or bypass the actuators 554a, b.

[0244] In the Fig. In the implementation illustrated in Figure 11, the actuators 554a, b are arranged in different fuel flow circuits. In other implementations, the actuators 554a, b may be arranged on the same fuel flow circuit and may be arranged in series or parallel within that circuit.

[0245] Another example of an actuated system used in turbine engines 10 is a bleed air system. Bleed air systems divert bleed air (e.g., compressed air taken from the compressor 14, 15 or elsewhere, e.g., from an APU) for use in other systems. Due to its relatively high temperature and pressure, bleed air is useful in various systems. Bleed air systems may be referred to as "engine bleed air systems" (EBAS).

[0246] Bleed air systems can be used to control the air conditioning within the cabin of an aircraft 1. In such systems, the bleed air is filtered and cooled (by expansion and / or heat exchange, e.g., with atmospheric air) using a climate pack (climate packs or pressurization air conditioning kits (PACKS) are both accepted terms). Climate packs serve to cool and, if necessary, dehumidify the bleed air from the engine or APU before it is fed into the aircraft cabin. Such systems comprise one or more valves and one or more actuators. The bleed air is generally supplied to the package via a one-way check valve. In various implementations, one or more of these actuators may be selectively fuel-hydraulic.

[0247] Fig. 12 shows a schematic diagram illustrating an alternative portion of a fuel supply system 152 of a gas turbine engine 10 that includes an air conditioning pack 650.

[0248] The fuel supply system 152 of Fig. 12 includes a cabin vent valve 654' and a cabin vent actuator 654 configured to control the cabin vent valve 654'. The cabin vent valve 654' controls the bleed air flow from the compressor 14, 15 through an air conditioning pack (as described above) to deliver cooled air into the cabin of the aircraft 1.

[0249] The fuel supply system 152 of Fig. 12 further includes a handling vent valve 754' and a handling vent actuator 754 configured to control the handling vent valve 754'.

[0250] The handling vent valve 754' controls the exhaust of bleed air from the compressor 14, 15 and into other aircraft systems. The handling vent valve 754' can be used for engine starting, temporary venting, and / or surge control.

[0251] In some arrangements, there are multiple cabin vent valves 654' and associated cabin vent actuators 654, all of which may be considered part of the climate pack 650. In some arrangements, there are multiple handling vent valves 754' and associated handling vent actuators 754, all of which may be considered part of the climate pack 650.

[0252] In alternative arrangements, the cabin vent valve 654' and the cabin vent actuator 654 are omitted. In alternative arrangements, the handling vent valve 754' and the handling vent actuator 754 are omitted. In alternative arrangements, alternative known actuators and associated valves may be included and selectively fuel-hydraulically actuated. In some arrangements, other known vent valves are selectively fuel-hydraulically actuated.

[0253] Similar to the actuators described above, the cabin vent actuator 654 is selectively fuel-hydraulically actuated to control the cabin vent valve 654'. The valve 654 is arranged to direct fuel via line 660 to actuate the cabin vent actuator 654 or to direct fuel via line 661 to bypass the cabin vent actuator 654.

[0254] Similar to the actuators described above, the handling vent actuator 754 is selectively fuel-hydraulically actuated to control the handling vent valve 754'. The valve 754 is arranged to direct fuel via line 760 to actuate the handling vent actuator 754 or to direct fuel via line 761 to bypass the handling vent actuator 754.

[0255] Another 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 compressor guide vanes; and generally by moving multiple stages of compressor guide vanes) to provide optimal compressor performance. The VGV system 850 varies the angle of the compressor's guide vanes to control compressor performance and efficiency.

[0256] The variable guide vanes of various implementations may include either variable inlet guide vanes (VIGVs) located upstream of / in front of the compressor rotor blades, or other guide vanes—generally variable stator vanes (VSVs) located behind / downstream of the compressor rotor blades—or both. Other compressor guide vanes may be located elsewhere—guide vane position may vary depending on the engine architecture and also between compressors of a given engine. For example, intermediate guide vanes, which are variable / controllable, may be located midway through the compressor 14, 15.

[0257] Generally, variable-angle guide vanes are arranged to rotate about an axis to adapt to the angle of the incoming airflow. This movement of the guide vanes allows a wider range of non-dimensional flow through the compressor 14, 15 without causing stall; because of the VGVs, the compressor 14, 15 can operate effectively over a wider range of engine settings without stalling. VGVs are controlled by the engine control system in response to a number of parameters, generally including the ambient conditions and the thrust required by the engine.

[0258] For engines 10 having multiple compressors 14, 15, a VGV system 850 may be provided for one, some, or all of the compressors 14, 15.

[0259] Many VGV systems 850 include two VGV actuators operable to change the angle of the inlet guide vanes of the compressor 14, 15. Each of the two VGV actuators can be configured to control the guide vanes on a specific side of the engine 10. In various implementations, one or more of the actuators of a VGV system are fuel-hydraulically driven.

[0260] Fig. 13 shows a schematic diagram illustrating an alternative portion of a fuel supply system 152 of a gas turbine engine 10 that includes a VGV system 850. The fuel supply system 152 includes a variable stator vane actuator 854. The fuel supply system 152 includes a valve 804 configured to selectively fuel-hydraulically actuate the variable stator vane actuator 854 by directing fuel via line 860 or line 861. The valve 804 may be configured to cause the variable stator vane actuator 854 to be actuated or bypassed based on the SAF content of the fuel. Although the arrangement of Fig. 13 shows a single variable stator vane actuator, in other arrangements, the fuel supply system 152 may be configured to selectively fuel-hydraulically actuate two or more variable stator vane actuators.

[0261] Fig. Figure 14 is a schematic diagram illustrating an alternative portion of a fuel supply system 152 of a gas turbine engine 10 that includes a VGV system 850 similar to that of Fig. 13, but having only a single, simple fuel bleed circuit with no equivalent of line 861. The fuel supply system 152 includes a variable inlet guide vane actuator 854a. The fuel supply system 152 includes a valve 904 configured to selectively fuel-hydraulically actuate the variable inlet guide vane actuator 854a by directing fuel to the actuator via line 960 or preventing fuel flow along line 960. The valve 904 may be configured to cause the fuel to actuate the variable inlet guide vane actuator 854a or to continue flowing along the main fuel flowpath 110, thus bypassing the actuator 854a, based on the SAF content of the fuel.

[0262] The guide vanes control the airflow into and through the compressor 14, 15. The variable guide vane actuator(s) 854, 854a control the orientation angle of the guide vanes.

[0263] 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. Both HMS systems control oil temperature by enabling and controlling the use (or bypassing) of air-to-oil heat exchangers (and / or fuel-to-oil heat exchangers). In some implementations, oil-to-oil heat exchangers may also be used, as discussed further below with respect to Fig. 25. For example, oil-to-oil heat exchangers can be used to enable heat transfer between oil in a generator HMS and an engine HMS.

[0264] The engine HMS system controls the temperature of the oil system that lubricates the engine components (e.g., bearing chambers, gearboxes, etc.). The engine HMS system includes one or more engine HMS valves controlled by one or more actuators. In various implementations, one or more of these actuators are fuel-hydraulically driven.

[0265] For example, in some implementations, the engine HMS includes a single valve and a single actuator that is fuel hydraulically driven and arranged to control the valve.

[0266] The engine HMS system may include a plurality of heat exchangers arranged in a parallel or series configuration with respect to fluid flow (e.g., fuel flow or oil flow). A single valve controlled by a single fuel-hydraulic actuator may enable the use (or bypass) of all heat exchangers—in particular, the flow rate of a fluid (air, oil, or fuel, as needed) through the valve may be controlled, with the portion of the fluid not flowing through the heat exchanger being diverted via a bypass tube. In some implementations, each heat exchanger has an associated valve controlled by an associated actuator. In various implementations, one or more of these actuators are fuel-hydraulic driven.Additionally, one or more recirculation tubes may be provided to return the fluid from the heat exchanger outlet to the heat exchanger inlet, thereby passing through the heat exchanger multiple times. The flow through the recirculation tube(s) may again be adjusted using an actuator-controlled valve. Each of the recirculation tubes described herein may have one or more associated pumps configured to recirculate the air / oil / fuel through the heat exchanger. Alternatively or additionally, any suitable components may be used to repressurize the air / oil / fuel to enable recirculation.

[0267] Fig. Figure 15 shows a schematic diagram illustrating an engine thermal management system 1000. The engine thermal management system 1000 includes an air-to-oil heat exchanger 1003 configured to remove heat from the oil in the engine. The air-to-oil heat exchanger 1003 includes an oil line 1001 and an air line 1002. Fig. 15, the heat exchanger 1003 is a cocurrent heat exchanger, where the oil and air flows in the same direction. In other implementations, the heat exchanger 1003 may be a countercurrent heat exchanger, meaning that the flow direction of the air line or the oil line is opposite, so they flow in opposite directions through the heat exchanger 1003.

[0268] In the arrangement of Fig. 15, the oil line 1001 includes a valve 1004 operable to bypass oil past the air-to-oil heat exchanger 1003. The valve 1004 is controlled by an actuator 1054. The valve 1004 may be binary (on / off) or configured 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.

[0269] The actuator 1054 may be selectively fuel-hydraulic, as described with respect to the above arrangements. A valve 1006 is operable to allow fuel to flow through the actuator 1054 to fuel-hydraulically actuate the actuator 1054, or to bypass the fuel flow past the actuator 1054. In some arrangements, when the valve 1006 causes the fuel to bypass the actuator 1054, the actuator 1054 may be non-fuel-hydraulically actuated.

[0270] The valve 1006 may be operable to enable fuel hydraulic actuation of the actuator 1054 only when the SAF content of the fuel is above a predetermined threshold.

[0271] Fig. 16 shows an alternative arrangement of a thermal management system 1000 in which the valve 1004 is positioned within the air line 1002. The engine thermal management system 1000 is similar to the thermal management system 1000 of Fig. 15, but the valve 1004 is operable to allow air to bypass the air-to-oil heat exchanger 1003, thereby controlling the air flow rate therethrough. Again, the fluid flows could be arranged differently in other implementations.

[0272] In other implementations, the engine HMS systems can support 1000 of the Fig. 15 and Fig. 16 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 arranged to allow simultaneous bypassing of all heat exchangers, as well as valves for bypassing individual heat exchangers.

[0273] The generator HMS can be independent of the engine HMS 1000. It provides cooling for an electrical machine that, for example, supplies the airframe with electrical energy to operate the aircraft systems.

[0274] Fig. Figure 17 shows a schematic diagram illustrating a generator thermal management system 1100. The generator thermal management system 1100 includes an air-to-oil heat exchanger 1103 configured to remove heat from the oil in the engine, and in particular, to remove heat from the oil used to cool—and optionally also lubricate—the generator. The air-to-oil heat exchanger 1103 includes an oil line 1101 and an air line 1102.

[0275] 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-to-oil heat exchanger 1103. The valve 1104 is controlled by an actuator 1154.

[0276] The actuator 1154 may be selectively fuel-hydraulic, as described with respect to the above arrangements. A valve 1106 is operable to allow fuel to flow through the actuator 1154 to fuel-hydraulically actuate the actuator, or to bypass fuel. In some arrangements, the actuator 1154 may be configured to be non-fuel-hydraulically actuated when the valve 1106 causes fuel to bypass the actuator 1154.

[0277] The valve 1106 may be operable to enable fuel hydraulic actuation of the actuator 1154 only when the SAF content of the fuel is above a predetermined threshold.

[0278] Fig. 18 shows an alternative arrangement of a generator thermal management system 1100 in which the valve 1104 is positioned within the air line 1102. The generator thermal management system 1100 is similar to the generator thermal management system 1100 of Fig. 17, but valve 1104 is operable to allow air to bypass 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.

[0279] Fig. Figure 19 shows a schematic diagram illustrating a fuel supply system 152 including two fuel-to-oil heat exchangers 1220, 1230. The fuel supply system 152 includes a primary fuel-to-oil heat exchanger 1220 and a secondary fuel-to-oil heat exchanger 1230. The heat exchangers 1220, 1230 are used to transfer heat from the oil to the fuel and can be considered part of the engine thermal management system. The primary fuel-to-oil heat exchanger 1220 is arranged to heat at least the majority of the fuel, and the secondary fuel-to-oil heat exchanger 1230 is arranged to add additional heat to the fuel to be supplied to fuel-hydraulicly drive the at least one fuel-hydraulic actuator 1254.

[0280] The fuel supply system 152 of the Fig. The implementation shown in Figure 19 includes five fuel valves 1201-1205. Each valve 1201-1205 is controlled by an actuator, which may be fuel hydraulic.

[0281] The first valve 1201 is a fuel valve arranged to allow fuel to bypass the primary heat exchanger 1220. The valve 1201 may be operable to allow any percentage of the fuel to bypass or flow through the primary heat exchanger 1220. The bypassed fuel is then returned to the main fuel flow path 110.

[0282] The primary heat exchanger 1220 allows the fuel temperature to be increased before entering the combustion chamber 16 while the temperature of the oil is reduced.

[0283] The second valve 1202 is arranged to direct the fuel from the outlet of the primary heat exchanger 1220 either: (i) through a return pipe 1206 configured to return the 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 implementations, multiple separate valves may be provided for return or bypass / drain for a fuel system.

[0284] The third valve 1203 forms part of the fuel hydraulic system and is a fuel valve arranged 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 the fuel to bypass or flow through the secondary heat exchanger 1230.

[0285] The fourth valve 1204 forms part of the fuel hydraulic system and is arranged to direct the fuel from downstream of the secondary heat exchanger 1230 either: (i) to a fuel hydraulic actuator 1254 or (ii) through a return pipe 1207 configured to return the 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. In some implementations, a fuel return valve 1204 may be provided separately from a valve arranged to control flow to the fuel hydraulic actuator 1254.

[0286] In some arrangements, one or both of the return tubes 1206, 1207 include or are connected to one or more pumps configured to pump the fuel through the return tube 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 length of the return tubes 1206, 1207.

[0287] The fifth valve 1205 forms part of the fuel hydraulic system and is arranged to direct fuel to hydraulically drive or bypass the actuator 1254. In the example shown, the fuel flowing through or bypassing the actuator 1254 is then directed directly to the combustion chamber 16. In other implementations, it may be recirculated into the main fuel flow path 110.

[0288] 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.

[0289] In some arrangements, the heat exchangers 1220, 1230 are controlled to provide a heat transfer ratio of: Heat transfer rate from oil to fuel in the secondary fuel-oil heat exchanger (kJkg−1)Heat transfer rate from oil to fuel in the primary and secondary fuel-oil heat exchanger (kJkg−1) has a maximum value (i.e. peak value) of at least 0.35.

[0290] In other arrangements, the heat transfer ratio in cruise flight has a maximum value of more than 0.4, 0.45, 0.5, 0.55 or 0.6.

[0291] In some arrangements, the fuel supply system 152 is configured to control the heat transfer ratio by controlling the percentage of fuel bypassed or recirculated through each of the heat exchangers 1220, 1230.

[0292] In additional or alternative arrangements, the heat transfer ratio is controlled by controlling the percentage of oil bypassing or recirculating through each of the heat exchangers 1220, 1230. Oil bypass and recirculation valves and / or tubes may be provided accordingly. To influence the heat transfer ratio, one or more air-to-oil or oil-to-oil heat exchangers may additionally be controlled.

[0293] An example of such an oil supply system 153 is shown in Fig. 25. In the implementation of Fig. 25, each fuel-oil heat exchanger 1220, 1230 is located in a separate oil circuit - the two oil circuits are fluidly isolated from each other and are brought into thermal contact by means of an oil-oil heat exchanger 1240 (although not shown, it is understood that one or more oil bypass and / or return valves may also be provided in conjunction with the oil-oil heat exchanger 1240).

[0294] The oil circuit, containing oil arranged to flow through the primary fuel-oil heat exchanger 1220, is arranged to cool and optionally lubricate a first set of engine components 10'. The oil circuit, containing oil arranged to flow through the secondary fuel-oil heat exchanger 1230, is arranged to cool and optionally lubricate a second set of engine components 10".

[0295] Each fuel-oil heat exchanger 1220, 1230 is associated with an oil bypass valve 1212, 1213 arranged to allow the oil to bypass the respective heat exchanger 1220, 1230. The oil bypass valve 1212, 1213 is arranged so that a controllable portion of the oil can bypass the respective heat exchanger 1220, 1230 in various implementations.

[0296] Each fuel-oil heat exchanger 1220, 1230 is assigned an oil return valve 1211, 1214, which is arranged so that the oil can flow several times through the respective heat exchanger 1220, 1230 before being returned to the engine components 10', 10", where the oil is to be cooled. The oil return valve 1211, 1214 is arranged so that a controllable portion of the oil can be returned through the respective heat exchanger 1220, 1230 in various implementations. In the implementation of Fig. 25, the recirculated oil is shown as being fed into the respective bypass valve 1212, 1213, but it is understood that in other implementations the recirculated oil may re-enter the oil stream before or after this valve.

[0297] In some implementations, only one of an oil return valve 1211, 1214 and an oil bypass valve 1212, 1213 may be provided for one or each heat exchanger 1220, 1230, or neither may be provided for one or each heat exchanger 1220, 1230.

[0298] Although in Fig. 25, for simplicity, no air-to-oil heat exchangers are shown, it should be understood that in various implementations, one or more air-to-oil heat exchangers may be provided on one or each oil circuit. Furthermore, in some implementations, a single oil circuit may supply oil to both fuel-to-oil heat exchangers; in some such implementations, no oil-to-oil heat exchanger may be provided.

[0299] In some arrangements, the engine 10 is configured to control the heat transfer ratio based on the SAF content of the fuel.

[0300] In some arrangements, the maximum temperature of the fuel exiting the secondary fuel-oil heat exchanger 1230 is at least 145°C, 150°C, 155°C, 160°C, 170°C, or 180°C.

[0301] For each of the fuel hydraulic actuators described above, the fuel supply system 152 may be configured such that a peak differential pressure of the fuel across the at least one fuel hydraulic actuator during 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.

[0302] For each of the fuel-hydraulic actuators described above, the fuel supply system 152 may be configured such that a peak differential pressure of the fuel across the at least one fuel-hydraulic actuator during startup is at least 9000 kPa. For each of the fuel-hydraulic actuators described above, the fuel supply system 152 may be configured such that a peak differential pressure of the fuel across the at least one fuel-hydraulic actuator during startup may be in the range of 6900 to 10 000 kPa, but at least 10 000 kPa, 11 000 kPa, 12 000 kPa, 13 000 kPa, 14 000 kPa, or 15 000 kPa.

[0303] The actuator 1254 may be any of the actuators described herein or any actuator that can be implemented using fuel hydraulics, such as the variable stator vane actuator.

[0304] In some arrangements, the temperature of the fuel entering the actuator 1254 (exiting valve 1205) is at least 5°C, 7°C, 10°C, 12°C, 15°C, or 20°C higher than that of the fuel entering the combustion chamber 16 (exiting valve 1202).

[0305] In some configurations, the fuel is thermally stable at temperatures above 280°C. In this context, "thermally stable" at a specified temperature means that the fuel passes a Jet Fuel Thermal Oxidation Test (JFTOT) at that temperature.

[0306] In some arrangements, 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.

[0307] In some configurations, aromatics make up less than 5% of the fuel volume. In some configurations, the fuel's calorific value is at least 43.5 MJkg -1 In some arrangements, the calorific value of the fuel is at least 44 MJkg -1 In some arrangements, the sulfur content of the fuel is below 15 parts per million. In some arrangements, the fuel is or includes a HEFA fuel.

[0308] Various methods for operating a gas turbine engine 10 for an aircraft 1 are described below.

[0309] Fig. 20 shows a flowchart illustrating a method 1300 for operating a gas turbine engine 10 according to various implementations. The method 1300 is performed for an engine 10 comprising: an engine core 11 including 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 arranged to be driven by the core shaft 26, the fan including 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 outward of the engine core 11; a plurality of actuators 254; and a fuel delivery system 152.

[0310] In the described implementations, a bypass ratio, defined as the ratio of the mass flow rate of the flow through the bypass channel 22 to the mass flow rate of the flow through the engine core 11 under cruise 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 Fig. 20 to 24 can be used.

[0311] The method 1300 includes: Supplying 1310 fuel using the fuel supply system 152 for combustion in the combustion chamber 16; and based on the Sustainable Aviation Fuel (SAF) content of the fuel, select 1320 between: Supplying 1320a fuel using the fuel supply system 152 to at least one actuator 354 of the plurality of actuators and causing the fuel to fuel-hydraulically actuate at least one actuator 354; and Causing 1320b the fuel to bypass the at least one actuator 354, again using the fuel supply system 152.

[0312] The step of selecting 1320 between fuel-hydraulic actuation of the actuator 354 and bypassing the at least one actuator 354 may include making a determination, in a processor or other computer system, of which action to perform based on a comparison of the SAF content of the fuel to one or more predetermined thresholds, and actuating a valve 304 based on that determination. The valve 304 may be a two-way valve configured to direct fuel through a first fuel line 360 ​​or through a second fuel line 361, 362, with the first fuel line 360 ​​driving the actuator 354 and the second fuel line 361, 362 bypassing the actuator 354.The valve 304 may simply be arranged to open a vent from a main fuel flowpath 110 only when fuel hydraulic actuation is required, so that all fuel remains in the main fuel flowpath 110 when fuel hydraulic actuation is not required. Therefore, when fuel hydraulic actuation is not required, the fuel may bypass the at least one actuator 354 through a dedicated bypass tube or simply remain in a main fuel flowpath that does not reach the actuator 354. In implementations with a dedicated bypass tube, the bypass tube may be actuator-specific; providing a route for the fuel to bypass a single actuator or to bypass multiple actuators (e.g., some or all actuators located on the same fuel vent path).

[0313] The step of selecting 1320 the action to be taken may include determining whether or not the SAF content of the fuel exceeds a threshold. The fuel delivery system 152 may include one or more sensors configured to sense one or more parameters that enable calculation or derivation of the SAF concentration of the fuel (e.g., by detecting a tracer element and referring to a lookup table of fuel tracer elements and SAF levels) and / or may include a storage location for stored data about the fuel(s) in use. The fuel delivery system 152 may include a processor or other computing system configured to receive the stored, calculated, or otherwise determined SAF content and compare that value to the threshold.A controller may then be used to actuate valve 304 according to the determination. In some implementations, an engine electronic controller (EEC) may be used to make the determination and implement the result; in other implementations, a separate, dedicated computer system may be used instead of the EEC.

[0314] 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.

[0315] The fuel used to fuel-hydraulically 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 flowpath 110. The fuel may flow along a main fuel flowpath 110 to the combustion chamber 16—the main fuel flowpath 110 may therefore supply 1310 fuel to the combustion chamber 16. A portion of the fuel flowing along the main fuel flowpath 110 may be diverted from the main fuel flowpath 110 to the at least one actuator 354—this fuel may then be returned to the main fuel flowpath 110 after its use in actuation. One or more tubes may be used to direct fuel from the main fuel flowpath 110 to the at least one actuator 354. These pipes can form one or more fuel flow circuits.One or more valves may be used to control the fuel flow rate through the or each fuel-hydraulic fuel flow circuit. In some implementations, one or more fuel-hydraulic fuel pumps may be provided to actively pump the 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 main fuel flow path 110 may be sufficient.

[0316] The outlet 159 from the main fuel flow path 110 may be located at a different point along the main fuel flow path 110 than the return pipe 160 of the (or each) fuel flow circuit, as in Fig. 6. The fuel may be returned to the main fuel flowpath 110 at a location upstream or downstream of the location of the vent. One or more valves may be used to control the location(s) along the main fuel flowpath 110 at which the fuel used to drive the actuators is returned to the main fuel flowpath 110. For example, the fuel used in actuation may be returned to a fuel tank 50, 53, to just upstream of the combustion chamber 16, or to upstream of or downstream of a fuel-oil heat exchanger or fuel pump along the main fuel flowpath 110, or within a fuel hydraulic system, as appropriate.Generally, the fuel used in actuation is returned to a point on the main fuel flow path 110 within the engine 10 (as opposed to, for example, a direct return to a fuel tank 50, 53 located elsewhere in the aircraft 1, or, for example, to an intermediate connecting pipe).

[0317] The supply steps 1310, 1320a may therefore include controlling a plurality of fuel flow valves and optionally also a plurality of fuel pumps.

[0318] Fig. 21 shows a flowchart illustrating a method 1400 for operating a gas turbine engine 10 according to various implementations. The method 1400 is performed for an engine 10 comprising: an engine core 11 including 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 arranged to be driven by the core shaft 26, the fan including 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 outward of the engine core 11; a plurality of actuators 254; and a fuel delivery system 152.

[0319] The method 1400 includes: Supplying 1410 fuel using the fuel supply system 152 for combustion in the combustion chamber 16; Supplying 1420 fuel using the fuel supply system 152 to fuel-hydraulically drive at least one actuator 354 of the plurality of actuators such that a peak differential pressure of the fuel across the at least one fuel-hydraulic actuator 354 is at least 2400 kPa during cruise conditions.

[0320] The supplied fuel has a SAF content of at least 25 vol%, but can also have a significantly higher content. The peak differential pressure can be controlled to be higher when the SAF content is higher.

[0321] As with respect to Procedure 1300 of Fig. 20, one or more fuel flow circuits and associated valves may be used to control how much fuel is 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.

[0322] Similarly, one or more valves and / or pumps may be controlled 1420 as needed to control the peak differential fuel pressure across the fuel hydraulic actuator 354. One or more pressure sensors may be provided to provide feedback on the differential fuel pressure. It should be understood that controlling 1420 the supply of fuel to the actuator(s) 354 enables adjustment of the differential pressure.

[0323] Fig. 22 shows a flowchart illustrating a method 1500 for 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 arranged 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 outward of the engine core 11; a plurality of actuators 254; a fuel delivery system 152; a primary fuel-oil heat exchanger 1220 and a secondary fuel-oil heat exchanger 1230.

[0324] The 1500 procedure includes: Supplying 1510 fuel using the fuel supply system 152 for combustion in the combustion chamber 16; Supplying 1520 fuel using the fuel supply system 152 to fuel-hydraulically 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 for fuel-hydraulic driving of the at least one fuel-hydraulic actuator using the secondary fuel-oil heat exchanger 1230.

[0325] The method includes controlling the heating of the fuel in the heat exchangers 1220, 1230 so that under cruise conditions a heat transfer ratio of: Heat transfer rate from oil to fuel in the secondary fuel-oil heat exchanger (kJkg−1)Heat transfer rate from oil to fuel in the primary and secondary fuel-oil heat exchanger (kJkg−1) has a maximum value of at least 0.35.

[0326] The primary fuel-to-oil heat exchanger 1220 may be configured to remove heat from the oil in the engine 10 used to lubricate and cool various engine components, such as a gearbox (if present) and / or shaft bearings. The secondary fuel-to-oil heat exchanger 1230 may be configured to remove heat from the oil used to cool—and optionally also lubricate—the generator. In some implementations, different, fluidly isolated oil streams may flow through the primary fuel-to-oil heat exchanger 1220 and the secondary fuel-to-oil heat exchanger 1230, as shown in Fig. 25. In other implementations, the same oil may flow through both heat exchangers—for example, sequentially through the secondary fuel-oil heat exchanger 1230 and then through the primary fuel-oil heat exchanger 1220.

[0327] The heat transfer ratio may be controlled by opening and closing valves to allow for fuel return through one or both heat exchangers 1220, 1230, fuel bypass of one or both heat exchangers 1220, 1230, oil return through one or both heat exchangers 1220, 1230, or oil bypass of one or both heat exchangers 1220, 1230. In addition, one or more other heat exchangers may be controlled to adjust the heat transfer ratio - for example, an air-to-oil heat exchanger may be used to cool oil before it reaches a fuel-to-oil heat exchanger 1220, 1230, and in implementations with multiple different oil streams, an oil-to-oil heat exchanger may be provided to transfer heat between the different oils. Flows of air and / or oil into such heat exchangers can therefore also be adjusted to influence the heat transfer ratio defined above.

[0328] The heat transfer ratio may be controlled by controlling 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 by controlling 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.

[0329] As in relation to procedures 1300, 1400 of the Fig. 20 and Fig. 21, one or more fuel flow circuits and associated valves may be used to control how much fuel is diverted from a main fuel flow path 110 to the actuator(s) and the secondary heat exchanger 1230 and where along the main fuel flow path 110 that fuel is returned.

[0330] Fig. 23 shows a flowchart illustrating a method 1600 for operating a gas turbine engine 10 according to various implementations. The method 1600 is performed for an engine 10 comprising: an engine core 11 including 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 arranged to be driven by the core shaft 26, the fan including 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 outward of the engine core 11; a plurality of actuators including a variable stator vane actuator 854; and a fuel delivery system 152.

[0331] The 1600 method includes: Supplying 1610 fuel using the fuel supply system 152 for combustion in the combustion chamber 16; and Supplying 1620 fuel using the fuel supply system 152 to fuel-hydraulically drive the variable compressor vane actuator 854 of the plurality of actuators.

[0332] Supplying 1620 fuel using the fuel supply system 152 to fuel-hydraulically drive the variable compressor vane actuator 854 and to the combustor 16 may include supplying fuel that is thermally stable at 280°C and optionally also thermally stable at higher temperatures.

[0333] As with procedures 1300, 1400, 1500 of the Fig. 20-22, one or more fuel flow circuits and associated valves may be used to control how much fuel is 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.

[0334] Fig. 24 shows a flowchart illustrating a method 1700 for 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 arranged 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 outward of the engine core 11; a plurality of actuators 254; a fuel delivery system 152; and at least one fuel-oil heat exchanger 1220 arranged to have oil and fuel flowing therethrough, the at least one heat exchanger 1220 arranged to transfer heat from the oil to the fuel.

[0335] The 1700 method includes: Supplying 1710 fuel using the fuel supply system 152 for combustion in the combustion chamber 16; Supplying 1720 fuel to fuel-hydraulically drive at least one actuator 1254 of the plurality of actuators; and Controlling 1730 the at least one heat exchanger 1220 such that, under cruise conditions, the fuel temperature entering the at least one actuator 1254 is at least 5°C higher than the fuel temperature entering the combustion chamber 16 (and optionally at least 10°C warmer).

[0336] The fuel supply system 152 may include two heat exchangers—a primary heat exchanger 1220 and a secondary heat exchanger 1230. Controlling 1730 the at least one heat exchanger 1220 may include controlling the primary heat exchanger 1220 and / or the secondary heat exchanger 1230. The secondary heat exchanger 1230 may be arranged to provide additional heat only for the fuel to be supplied to the fuel hydraulic actuator(s) 1254.

[0337] Method 1700 may include determining at least one fuel property of the fuel and controlling a temperature difference between the fuel temperature entering the at least one actuator 1254 and the fuel temperature entering the combustion chamber 16 based on the at least one fuel property. The fuel property may be the SAF content of the fuel. The fuel property may be determined by any approach known in the art, from retrieving data from an onboard fuel information data store to determining or inferring it from one or more sensed parameters.

[0338] The fuel properties may be or include the calorific value, thermal stability, or the percentage of sustainable aviation fuel (SAF) in the fuel.

[0339] Controlling 1730 the fuel temperature entering the at least one actuator 1254 relative to the fuel temperature entering the combustion chamber 16 may include actuating one or more valves to recirculate fuel through one or more of the heat exchangers or to bypass fuel around one or more of the heat exchangers 1220, 1230.

[0340] Controlling 1730 the fuel temperature entering the at least one actuator 1254 relative to the fuel temperature entering the combustion chamber 16 may include actuating one or more valves to recirculate oil through the heat exchanger 1220, 1230 or to bypass oil around the heat exchanger.

[0341] As with respect to procedures 1300, 1400, 1500, 1600 of the Fig.20-23, one or more fuel flow circuits and associated valves may be used to control how much fuel is 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.

[0342] It should 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. Except in cases of mutual exclusion, each of the features may be used separately or in combination with any other features, and the disclosure extends to and includes all combinations and subcombinations of one or more features described herein.

Claims

[1] 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 arranged 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 outward of 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), the fuel supply system being configured to supply fuel for combustion in the combustion chamber (16) and to supply fuel to fuel-hydraulically drive at least one actuator of the plurality of actuators (254), and wherein the fuel comprises at least 25 vol% sustainable aviation fuel - SAF - and wherein the fuel supply system (152) is configured such that a peak differential pressure of the fuel across the at least one fuel hydraulic actuator (254) during cruise conditions is at least 2400 kPa. [2] The gas turbine engine (10) of claim 1, wherein the fuel supply system (152) is configured such that the peak differential pressure of the fuel across the at least one fuel hydraulic actuator (254) during cruise conditions is at least 2500 kPa. [3] Gas turbine engine (10) according to claim 1, wherein the fuel supply system (152) is arranged such that the peak differential pressure of the fuel across the at least one fuel hydraulic actuator (254) during cruise conditions is at least 2800 kPa, optionally at least 3000 kPa [4] The gas turbine engine (10) of claim 1, wherein the fuel supply system (152) is configured such that the peak differential pressure of the fuel across the at least one fuel hydraulic actuator (254) during start-up is at least 6900 kPa. [5] The gas turbine engine (10) of claim 1, wherein the fuel supply system (152) is configured such that the peak differential pressure of the fuel across the at least one fuel hydraulic actuator (254) during start-up is at least 7000 kPa, optionally at least 8000 kPa. [6] The gas turbine engine (10) of claim 1, wherein the fuel supply system (152) is configured such that the peak differential pressure of the fuel across the at least one fuel hydraulic actuator (254) at idle is in the range of 1000 kPa to 1250 kPa. [7] A gas turbine engine (10) according to claim 1, wherein the core shaft (26) outputs drive directly to the fan (23) to drive the fan at the same speed as the core shaft, such that the engine (10) is a direct drive turbine engine. [8] A gas turbine engine (10) according to claim 1, wherein the turbine engine (10) includes a gearbox (30) receiving an input from the core shaft (26) and outputting a drive to the fan (23) to drive the fan at a lower speed than the core shaft, such that the engine (10) is a geared turbine engine. [9] A gas turbine engine (10) according to claim 1, wherein the fuel comprises at least 50 vol% SAF. [10] Gas turbine engine (10) according to claim 1, wherein the fuel comprises at least 55 vol% SAF, optionally at least 60 vol% SAF. [11] The gas turbine engine (10) of claim 1, wherein the at least one fuel hydraulic actuator (254) is a variable stator vane actuator (854). [12] The gas turbine engine (10) of claim 1, wherein the at least one fuel hydraulic actuator (254) is a variable inlet guide vane actuator (854a). [13] A method (1400) for 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 arranged 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 outward of 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 (1400) comprises: Supplying (1410) fuel comprising at least 25 vol% SAF using the fuel supply system (152) for combustion in the combustion chamber (16); and Supplying (1420) fuel comprising at least 25 vol% SAF using the fuel supply system (152) to fuel-hydraulically drive at least one actuator of the plurality of actuators such that a peak differential pressure of the fuel across the at least one fuel-hydraulic actuator during cruise conditions is at least 2400 kPa. [14] The method (1400) of claim 13, wherein the method comprises supplying (1410, 1420) fuel comprising at least 50 vol% SAF and controlling (1420) the fuel supply such that the peak differential pressure is at least 3200 kPa. [15] The method (1400) of claim 13, wherein the method comprises supplying (1410, 1420) fuel comprising at least 55 vol% SAF and controlling (1420) the fuel supply such that the peak differential pressure is at least 3600 kPa.

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