Gas turbine engine

Abstract

A gas turbine engine for an aircraft includes an engine core including a turbine, a combustor, a compressor, and a mandrel connecting the turbine to the compressor. The engine includes a fan located upstream of the engine core and arranged to be driven by the mandrel. The engine includes a nacelle surrounding the fan and the engine core and defining a bypass conduit radially outward of the engine core, where a bypass ratio defined as a ratio of a mass flow rate of flow through the bypass conduit to a mass flow rate of flow through the core under cruise conditions is at least 4. The engine comprises a fuel supply system and a plurality of actuators, the fuel supply system being arranged to supply fuel for combustion in the combustor and to supply fuel to hydraulically drive at least one of the plurality of actuators by the fuel. The fuel supply system is arranged to be controlled to select, based on a sustainable aviation fuel-SAF-content of the fuel, between: causing the fuel to actuate the at least one actuator; and bypassing the fuel by the at least one actuator.

Description

[0001] Cross-reference to related applications

[0002] This specification is based on and claims the benefit of priority of UK Patent Application No. 2319154.7, filed on Dec. 14, 2023, the entire content of which is incorporated herein by reference. Background of the Invention Field of the Invention

[0003] The present disclosure relates to an aircraft actuation system and to a method of actuating an actuation system in a fuel hydraulic manner.

[0004] Description of Related Technologies

[0005] In the aviation industry, there is a trend towards using fuels different from the conventional kerosene-based jet fuels commonly used at present. These fuels may have different fuel properties relative to petroleum / fossil-based hydrocarbon fuels. Therefore, it is necessary to consider the fuel properties of these new fuels and to adapt the gas turbine engine itself and the method of operating the gas turbine engine. Summary of the Invention

[0006] According to a first aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising:

[0007] An engine core including a turbine, a combustor, a compressor, and a spool connecting the turbine to the compressor;

[0008] A fan located upstream of the engine core and arranged to be driven by the spool, the fan including a plurality of fan blades;

[0009] A nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside the engine core, wherein the bypass ratio is at least 4, the bypass ratio being 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 under cruise conditions;

[0010] A plurality of actuators; and

[0011] A fuel supply system, wherein the fuel supply system is arranged to supply fuel for combustion in the combustor and to supply fuel to drive at least one of the plurality of actuators in a fuel hydraulic manner, and

[0012] wherein the fuel supply system is arranged to be controlled to select between:

[0013] Causing the fuel to actuate at least one actuator; and

[0014] Bypass the at least one actuator with the fuel.

[0015] The inventors have recognized that using a fuel different from traditional kerosene-based jet fuel, such as sustainable aviation fuel, can result in different fuel properties, and these different fuel properties can enable actuators to be driven in a fuel hydraulic manner. In particular, some fuels can be heated to higher temperatures and used to drive at least one more actuator than traditional fuels without significantly increasing the risk of fuel thermal degradation (e.g., fuel varnish or fuel coking) within the actuator.

[0016] The bypass ratio is greater than or equal to 4 and can be in the range of 4 to 55. The bypass ratio can be in the range of 4 to 20. The bypass ratio can be in the range of 4 to 15.

[0017] The fuel system can be arranged to:

[0018] actuate at least one actuator with the fuel when the SAF content of the fuel is above a threshold; and

[0019] bypass the actuator with the fuel when the SAF content of the fuel is below the threshold.

[0020] The fuel system can be arranged 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 can have different thresholds from each other - for example, the fuel system can be arranged to actuate a first actuator with the fuel when the SAF content of the fuel is above a first threshold and otherwise bypass the actuator, and actuate a second actuator with the fuel when the SAF content of the fuel is above a second threshold higher than the first threshold and otherwise bypass the actuator.

[0021] The spool can directly output drive to the fan to drive the fan at the same rotational speed as the spool. Such an engine can be a direct drive turbofan engine.

[0022] The turbofan engine can include a gearbox that receives input from the spool and outputs drive to the fan to drive the fan at a rotational speed lower than that of the spool. Such an engine can be a geared turbofan engine.

[0023] The engine may include a turbine casing cooling - TCC - system. A plurality of actuators may include actuators that are part of the turbine casing cooling system. The fuel supply system may be configured to actuate or bypass actuators that are part of the turbine casing cooling system based on the SAF content of the fuel. At least two of the plurality of actuators may be part of the turbine casing cooling system. The fuel supply system may be arranged to supply fuel to hydraulically drive at least two of the plurality of actuators that are part of the TCC system. The fuel supply system may be configured to actuate or bypass at least two actuators that are part of the turbine casing cooling system based on the SAF content of the fuel.

[0024] The engine may include an engine nacelle drain valve. A plurality of actuators may include actuators configured to actuate the engine nacelle drain valve. The fuel supply system may be configured to actuate or bypass the actuators configured to actuate the engine nacelle drain valve based on the SAF content of the fuel.

[0025] The engine may include a manual drain valve. A plurality of actuators may include actuators configured to actuate the manual drain valve. The fuel supply system may be configured to actuate or bypass the manual drain valve based on the SAF content of the fuel.

[0026] The engine may include an engine thermal management system having a valve. A plurality of actuators may include actuators 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.

[0027] The engine thermal management system may include a heat exchanger. The heat exchanger may be an air - 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 - oil heat exchanger. The fuel supply system may be configured to actuate or bypass a valve on the fuel side of the fuel - 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 - oil heat exchanger based on the SAF content of the fuel.

[0028] The engine may include a generator thermal management system having a valve. A plurality of actuators may include actuators configured to actuate the valve within the generator thermal management system. The fuel supply system may be configured to actuate or bypass the valve within the generator thermal management system based on the SAF content of the fuel.

[0029] The generator thermal management system may include a heat exchanger. The heat exchanger may be an air-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-oil heat exchanger. The fuel supply system may be configured to actuate or bypass a valve on the fuel side of the fuel-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-oil heat exchanger based on the SAF content of the fuel.

[0030] When the fuel supply system bypasses at least one actuator, the at least one actuator may be actuated in a non-fuel hydraulic manner. When the actuator is actuated in a non-fuel hydraulic manner, it may use a non-fuel hydraulic fluid to be actuated. The non-fuel hydraulic fluid may be supplied via a check valve to avoid mixing. A check valve may be used to supply fuel to the actuator to avoid mixing. The actuator may alternatively be electrically or pneumatically actuated.

[0031] The fuel supply system may be arranged to cause the fuel to actuate at least one actuator when the SAF content is above a threshold. The fuel supply system may be arranged to cause the fuel to bypass at least one actuator when the SAF content is below a threshold.

[0032] The minimum SAF content required to actuate at least one actuator may be at least 25%, 30%, 35%, 40%, 45%, 50%, 52%, 55%, 60%, 65%, 70% or 75% by volume.

[0033] The fuel supply system may be arranged to be controlled to select between fuel hydraulic actuation and bypassing two or more of a plurality of actuators. For example, the fuel supply system may be capable of being controlled to select between fuel hydraulic actuation and bypassing two, three, four or five actuators.

[0034] When at least one actuator is driven in a fuel hydraulic manner, the maximum operating pressure difference across the actuator during takeoff conditions may be in the range of 6,900 kPa to 10,000 kPa, or may be greater than 10,000 kPa. When at least one actuator is driven in a fuel hydraulic manner, the maximum operating pressure difference during takeoff conditions may be greater than 7,000 kPa, 8,000 kPa, 9,000 kPa, 10,000 kPa, 11,000 kPa, 12,000 kPa, 13,000 kPa, 14,000 kPa or greater than 15,000 kPa.

[0035] When at least one actuator is driven in a fuel-hydraulic manner, the maximum operating differential pressure during cruise conditions can 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.

[0036] The same actuator can encounter significantly lower pressures at idle, for example having a differential pressure in the range of 1000 kPa to 1250 kPa (150 psid to 180 psid).

[0037] According to a second aspect, there is provided a method of operating a gas turbine engine of an aircraft, the engine comprising:

[0038] An engine core that includes a turbine, a combustor, a compressor, and a spool that connects the turbine to the compressor;

[0039] A fan that is located upstream of the engine core and is arranged to be driven by the spool, the fan including a plurality of fan blades;

[0040] A nacelle that surrounds the fan and the engine core and defines a bypass duct that is radially outward of the engine core, wherein the bypass ratio is at least 4, the bypass ratio being 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 under cruise conditions;

[0041] A plurality of actuators; and

[0042] A fuel supply system;

[0043] And wherein the method comprises:

[0044] Using the fuel supply system to supply fuel for combustion in the combustor; and

[0045] Based on the sustainable aviation fuel - SAF - content of the fuel, selecting between:

[0046] Using the fuel supply system to supply fuel to at least one of the plurality of actuators so as to actuated the at least one actuator in a fuel-hydraulic manner; and

[0047] Using the fuel supply system to bypass the fuel around the at least one actuator.

[0048] The method may comprise:

[0049] Using the fuel supply system to supply fuel to at least two of the plurality of actuators, and

[0050] Based on the sustainable aviation fuel - SAF - content, for each of at least two actuators (individually / independently), select between:

[0051] Actuate the actuator with the fuel in a fuel - hydraulic manner; and

[0052] Bypass the actuator with the fuel.

[0053] The method may include:

[0054] Supply fuel to at least two of a plurality of actuators using a fuel supply system, and select between:

[0055] Actuate the at least two actuators with the fuel in a fuel - hydraulic manner; and

[0056] Bypass the at least two actuators with the fuel.

[0057] Thus, the bypass decision can be a single decision made for all relevant actuators, or can be a series of decisions made for each individual actuator.

[0058] The bypass ratio can be in the range of 4 to 55. The bypass ratio can be in the range of 4 to 20. The bypass ratio can be in the range of 4 to 15.

[0059] The method of the second aspect can be performed using the engine of the first aspect.

[0060] The method may include increasing the fuel pressure flowing through at least one fuel - hydraulic actuator based on the SAF content of the fuel. For example, for %SAF above 60% (vol%), for every 5% increase, the pressure can be increased by at least 350 kPa. One or more pumps and / or valves can be used to increase the pressure.

[0061] According to a third aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising:

[0062] An engine core including a turbine, a combustor, a compressor, and a spool connecting the turbine to the compressor;

[0063] A fan located upstream of the engine core and arranged to be driven by the spool, the fan including a plurality of fan blades;

[0064] A nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside the engine core, wherein the bypass ratio is at least 4, the bypass ratio being 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 under cruise conditions;

[0065] A plurality of actuators; and

[0066] A fuel supply system, wherein the fuel supply system is arranged to supply fuel for combustion in the burner and to supply fuel to hydraulically drive at least one of the plurality of actuators by fuel pressure.

[0067] The fuel comprises at least 25% sustainable aviation fuel - SAF by volume. The fuel supply system is arranged such that the peak pressure differential of the fuel on at least one fuel - hydraulic actuator during cruise conditions is at least 2400 kPa (350 psid).

[0068] It should be understood that even during cruise conditions, the pressure differential across the actuator will vary depending on operating conditions (e.g., fuel flow rate). Thus, for ease of comparison, the peak or maximum pressure differential is selected; this pressure is more specifically the steady - state peak differential pressure; that is, any short - lived, transient pressure spikes are ignored. Thus, the steady - state peak can be a time - averaged pressure value averaged over five, ten, fifteen, twenty, twenty - five, or thirty seconds. The peak should last for at least about five seconds to exclude sharp transient spikes.

[0069] During cruise conditions, the peak pressure differential of the fuel on at least one fuel - hydraulic actuator can 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.

[0070] During cruise conditions, the peak pressure differential of the fuel on at least one fuel - hydraulic actuator can be in the range of 2400 kPa to 4500 kPa, and optionally 2500 kPa to 4000 kPa, or 2500 kPa to 3800 kPa.

[0071] The bypass ratio is greater than or equal to 4 and can be in the range of 4 to 55. The bypass ratio can be in the range of 4 to 20. The bypass ratio can be in the range of 4 to 15.

[0072] When at least one actuator is hydraulically driven by fuel, the maximum operating pressure differential during take - off conditions can be in the range of 6900 kPa to 10000 kPa, or can be greater than 7000 kPa, 8000 kPa, 9000 kPa, or 10000 kPa.

[0073] The same actuator can encounter significantly lower pressures at idle, for example having a pressure differential in the range of 1000 kPa to 1250 kPa (150 psid to 180 psid).

[0074] The mandrel can directly output drive to the fan, thereby driving the fan at the same rotational speed as the mandrel, such that the engine is a direct drive turbofan engine.

[0075] The turbofan engine can include a gearbox that receives input from the mandrel and outputs drive to the fan, thereby driving the fan at a rotational speed lower than that of the mandrel, such that the engine is a geared turbofan engine.

[0076] The fuel can comprise at least 50%, 55%, 60%, 65%, 70% or 75% SAF by volume.

[0077] At least one fuel hydraulic actuator can be a variable stator vane actuator.

[0078] At least one fuel hydraulic actuator can be a variable inlet guide vane actuator.

[0079] The engine of the third aspect can include any or all of the features of the engine of the first aspect and can be used to implement the method of the second aspect.

[0080] According to a fourth aspect, there is provided a method of operating a gas turbine engine for an aircraft.

[0081] The engine comprises:

[0082] An engine core including a turbine, a combustor, a compressor and a mandrel connecting the turbine to the compressor;

[0083] A fan located upstream of the engine core and arranged to be driven by the mandrel, the fan including a plurality of fan blades;

[0084] A nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside the engine core, wherein the bypass ratio is at least 4, the bypass ratio being 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 under cruise conditions;

[0085] A plurality of actuators; and

[0086] A fuel supply system;

[0087] The method includes:

[0088] Using the fuel supply system to supply fuel for combustion in the combustor; and

[0089] Using the fuel supply system to supply fuel comprising at least 25% SAF by volume to hydraulically drive at least one of the plurality of actuators, supplying the fuel such that the peak fuel differential pressure across the at least one fuel hydraulic actuator during cruise conditions is at least 2400 kPa.

[0090] Fuel of the same composition can be supplied to the burner and the at least one actuator. One or more controllable valves and / or pumps can be provided to regulate the fuel pressure difference across the at least one fuel hydraulic actuator.

[0091] During cruise conditions, the peak fuel pressure difference across the at least one fuel hydraulic actuator can be at least 2500 kPa.

[0092] The bypass ratio can be in the range of 4 to 55. The bypass ratio can be in the range of 4 to 20. The bypass ratio can be in the range of 4 to 15.

[0093] The method can include controlling the fuel supply to one or more fuel hydraulic actuators so as to regulate the peak fuel differential pressure across the at least one fuel hydraulic actuator based on the SAF content of the fuel.

[0094] The method can include supplying fuel comprising at least 50% by volume of SAF, and controlling the fuel supply to the at least one actuator such that the peak differential pressure is at least 3200 kPa.

[0095] The method can include supplying fuel comprising at least 55% by volume of SAF, and controlling the fuel supply to the at least one actuator such that the peak differential pressure is at least 3550 kPa or 3600 kPa.

[0096] The methods of the second and fourth aspects can be complementary and can be carried out together in various embodiments. The method of the fourth aspect can be carried out using the engine of the first or third aspect.

[0097] According to a fifth aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising:

[0098] An engine core including a turbine, a burner, a compressor, and a shaft connecting the turbine to the compressor;

[0099] A fan located upstream of the engine core and arranged to be driven by the shaft, the fan including a plurality of fan blades;

[0100] A nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside the engine core, wherein the bypass ratio is at least 4, the bypass ratio being 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 under cruise conditions;

[0101] A plurality of actuators;

[0102] A fuel supply system, wherein the fuel supply system is arranged to supply fuel for combustion in the burner and to supply fuel to hydraulically drive at least one of the plurality of actuators, and

[0103] Two fuel - oil heat exchangers are arranged to have oil and fuel flowing therethrough, the heat exchangers being arranged to transfer heat from the oil to the fuel and including a primary fuel - oil heat exchanger arranged to heat at least a majority of the fuel and a secondary fuel - oil heat exchanger arranged to provide additional heat to the fuel to be supplied for hydraulically driving at least one fuel - hydraulic actuator.

[0104] Control the heat exchanger such that, under cruise conditions, the heat transfer ratio:

[0105]

[0106] Has a maximum value of at least 0.35.

[0107] Although it should be understood that the ratio is dimensionless, in the examples described herein, heat transfer is measured per unit mass or volume of fuel (by way of example, in kJ / kg as described above), thus providing a heat transfer rate normalized for changes in fuel flow rate during cruise. It should be understood that the heat transfer per unit mass (kg) of fuel reaching the burner is recorded in order to adjust the fuel flow rate and any recirculation through one or more of the heat exchangers or bypasses of one or more of the heat exchangers, as described elsewhere herein. In most embodiments, any temperature increase of the fuel due to other engine components (as opposed to heat transfer from the oil) can be assumed to be minimal.

[0108] It should be understood that even under cruise conditions, there may be short, transient spikes in heat transfer that should be ignored. Thus, the maximum value of the heat transfer ratio can be a time average averaged over one minute, two minutes, five minutes, ten minutes, fifteen minutes, twenty minutes, twenty - five minutes, or thirty minutes. It should be understood that the time scale of the transient temperature changes can be greater than the time scale of the pressure in the actuator, while the full movement of the actuator may be completed within a few seconds and the fuel and oil may spend a longer time within the heat exchanger.

[0109] The heat transfer ratio can have a maximum value greater than 0.4 during cruise. The heat transfer ratio can have a maximum value greater than 0.45 during cruise. The heat transfer ratio can have a maximum value greater than 0.5 during cruise.

[0110] The bypass ratio is greater than or equal to 4 and can be in the range of 4 to 55. The bypass ratio can be in the range of 4 to 20. The bypass ratio can be in the range of 4 to 15.

[0111] 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.

[0112] The primary heat exchanger and the secondary heat exchanger can be arranged such that the heat transfer ratio can be adjusted during engine operation, for example, by controlling one or more pumps and / or valves to regulate the fuel and / or oil flow through one or each heat exchanger.

[0113] The heat exchanger can be arranged such that the heat transfer ratio can be adjusted based on the sustainable aviation fuel - SAF - content of the fuel.

[0114] The engine can include at least one bypass valve arranged to allow fuel or oil to bypass one of the fuel - oil heat exchangers in order to adjust the heat transfer ratio.

[0115] For example, the engine can include one or more of the following:

[0116] · A valve operable to allow oil to bypass the primary heat exchanger;

[0117] · A valve operable to allow oil to bypass the secondary heat exchanger;

[0118] · A valve operable to allow fuel to bypass the primary heat exchanger; and

[0119] · A valve operable to allow fuel to bypass the secondary heat exchanger.

[0120] The engine can include at least one recirculation valve arranged to allow fuel or oil to pass through one of the fuel - oil heat exchangers multiple times in order to adjust the heat transfer ratio.

[0121] For example, the engine can include one or more of the following:

[0122] · A valve operable to allow oil to recirculate through the primary heat exchanger;

[0123] · A valve operable to allow oil to recirculate through the secondary heat exchanger;

[0124] · A valve operable to allow fuel to recirculate through the primary heat exchanger; and

[0125] · A valve operable to allow fuel to recirculate through the secondary heat exchanger.

[0126] For any recirculation valve described herein, there can be one or more associated pumps configured to deliver the oil / fuel back to the inlet of the heat exchanger to pass through the heat exchanger again. Alternatively or additionally, any suitable component for repressurizing the oil / fuel to achieve recirculation can be used.

[0127] The mandrel can directly output drive to the fan, thereby driving the fan at the same rotational speed as the mandrel, such that the engine is a direct drive turbofan engine.

[0128] The engine of the fifth aspect may include any or all features of the engines of the first and / or third aspects, and may be used to implement the methods of the second or fourth aspects.

[0129] According to a sixth aspect, there is provided a method of operating a gas turbine engine for an aircraft.

[0130] The engine comprises:

[0131] An engine core including a turbine, a combustor, a compressor and a mandrel connecting the turbine to the compressor;

[0132] A fan located upstream of the engine core and arranged to be driven by the mandrel, the fan including a plurality of fan blades;

[0133] A nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside the engine core, wherein the bypass ratio is at least 4, the bypass ratio being defined as the ratio of the mass flow rate of the flow passing through the bypass duct to the mass flow rate of the flow passing through the core under cruise conditions;

[0134] A plurality of actuators;

[0135] A fuel supply system; and

[0136] A primary fuel-oil heat exchanger and a secondary fuel-oil heat exchanger.

[0137] The method comprises:

[0138] Supplying fuel using the fuel supply system for combustion in the combustor;

[0139] Supplying fuel using the fuel supply system to hydraulically drive at least one of the plurality of actuators;

[0140] Heating at least a majority of the fuel using the primary fuel-oil heat exchanger; and

[0141] Heating the fuel to be supplied for hydraulically driving at least one fuel hydraulic actuator using the secondary fuel-oil heat exchanger;

[0142] Wherein, under cruise conditions, the supply and heating are controlled such that the heat transfer ratio:

[0143]

[0144]

[0144]

[0145]

[0146] Has a maximum value of at least 0.35.

[0147] The bypass ratio can be in the range of 4 to 55. The bypass ratio can be in the range of 4 to 20. The bypass ratio can be in the range of 4 to 15.

[0148] The method may further include adjusting the heat transfer ratio based on the sustainable aviation fuel - SAF - content of the fuel, for example by adjusting the fuel and / or oil flow rate through one or each heat exchanger.

[0149] In various embodiments, the methods of the second, fourth, and sixth aspects can be complementary and can be performed together, or in any combination or sub - combination. The method of the sixth aspect can be performed using the engine of the first, third, or fifth aspect.

[0150] According to a seventh aspect, there is provided a gas turbine engine for an aircraft, the engine comprising:

[0151] An engine core that includes a turbine, a burner, a compressor, and a spool connecting the turbine to the compressor;

[0152] A fan located upstream of the engine core and arranged to be driven by the spool, the fan including a plurality of fan blades;

[0153] A nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside the engine core, wherein the bypass ratio is at least 4, the bypass ratio being 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 under cruise conditions;

[0154] A plurality of actuators including variable compressor vane actuators; and

[0155] A fuel supply system, wherein the fuel supply system is arranged to supply fuel for combustion in the burner and to supply fuel to hydraulically drive the variable compressor vane actuators, wherein the fuel is thermally stable at temperatures above 280 °C.

[0156] The bypass ratio is greater than or equal to 4 and can be in the range of 4 to 55. The bypass ratio can be in the range of 4 to 20. The bypass ratio can be in the range of 4 to 15.

[0157] Thermally stable can refer to the ability of the fuel to meet the requirements of the Jet Fuel Thermal Oxidation Test (JFTOT). A fuel that is thermally stable at temperatures above 280 °C can refer to a fuel that meets the JFTOT test requirements at temperatures above or equal to 280 °C.

[0158] The fuel can be thermally stable at a temperature equal to or exceeding 290 °C, 295 °C, 300 °C, 305 °C, 310 °C, 315 °C, 320 °C or 325 °C.

[0159] An aircraft gas turbine engine typically includes a plurality of compressor blades - variable compressor blade actuators are arranged to adjust the position of the compressor blades. Since the blades are used to direct the airflow and since the position of the compressor blades is variable, they can be referred to as variable guide vanes.

[0160] The variable guide vanes of various embodiments can include any one or both of the following:

[0161] · Variable inlet guide vanes (VIGV), which are located upstream / ahead of the rotor blades of the compressor and are arranged to direct and control the airflow entering the compressor; and

[0162] · Other blades - typically variable stator vanes (VSV), which are located downstream / behind the rotor blades of the compressor

[0163] and are arranged to direct and control the airflow through the compressor.

[0164] The engine can include a plurality of variable stator vane actuators, each of which can be a fuel hydraulic actuator. In addition, the engine can include a plurality of compressors, each compressor having at least one variable stator vane actuator.

[0165] A gas turbine engine can include a plurality of variable compressor blades and a plurality of variable compressor blade actuators. A fuel supply system can be arranged to supply fuel to drive the variable compressor blade actuators in a fuel hydraulic manner.

[0166] A gas turbine engine can include at least two compressors, each compressor having at least one variable compressor blade and at least one variable compressor blade actuator associated therewith. A fuel supply system can be arranged to supply fuel to drive at least one variable compressor blade actuator of each compressor in a fuel hydraulic manner.

[0167] The engine can include a turbine casing cooling system. The plurality of actuators can include a turbine casing cooling actuator. A fuel supply system can be arranged to supply fuel to drive the turbine casing cooling actuator in a fuel hydraulic manner.

[0168] An aircraft gas turbine engine typically includes a plurality of servo motors (also known as servo mechanisms), which are rotary or linear actuators that allow precise control of angular or linear position, speed, and / or acceleration in a mechanical system. In a given engine, one, some, or all of the servo mechanisms present may be actuated fuel hydraulically. One or more servo mechanisms may include sensors arranged to provide position feedback. A dedicated controller may be provided for the engine thermal management system to control one or more servo mechanisms.

[0169] For example, the engine may include a servo mechanism within a hydromechanical unit. The plurality of actuators may include at least one actuator configured to actuate a servo mechanism within the hydromechanical unit. The fuel supply system may be arranged to supply fuel to drive, fuel hydraulically, an actuator configured to actuate a servo mechanism valve.

[0170] The aromatic hydrocarbons may account for less than 5% of the fuel volume.

[0171] The calorific value of the fuel may be at least 43.5 MJ / kg -1 The calorific value of the fuel may be at least 44 MJ / kg -1 .

[0172] The sulfur content of the fuel may be less than 15 parts per million.

[0173] The fuel may be or comprise HEFA fuel, i.e., a fuel made from hydrotreated esters and fatty acids.

[0174] The spool may output drive directly to the fan, thereby driving the fan at the same rotational speed as the spool, such that the engine is a direct drive turbofan engine.

[0175] The turbofan engine may include a gearbox that receives an input from the spool and outputs drive to the fan, thereby driving the fan at a rotational speed lower than that of the spool, such that the engine is a geared turbofan engine.

[0176] The engine of the seventh aspect may include any or all of the features of the engines of the first, third, and / or fifth aspects, and may be used to implement the methods of the second, fourth, or sixth aspects.

[0177] According to an eighth aspect of the present invention, there is provided a method of operating a gas turbine engine for an aircraft.

[0178] The engine includes:

[0179] An engine core that includes a turbine, a combustor, a compressor, and a spool that connects the turbine to the compressor;

[0180] A fan that is located upstream of the engine core and is arranged to be driven by the spool, the fan including a plurality of fan blades;

[0181] A nacelle that surrounds the fan and the engine core and defines a bypass duct located radially outside the engine core, wherein the bypass ratio is at least 4, and the bypass ratio is defined as the ratio of the mass flow rate of the flow passing through the bypass duct to the mass flow rate of the flow passing through the core under cruise conditions;

[0182] A plurality of actuators, the plurality of actuators including variable compressor vane actuators; and

[0183] A fuel supply system.

[0184] The bypass ratio is greater than or equal to 4 and can be in the range of 4 to 55. The bypass ratio can be in the range of 4 to 20. The bypass ratio can be in the range of 4 to 15.

[0185] The method includes:

[0186] Using the fuel supply system to supply fuel for combustion in the burner; and

[0187] Using the fuel supply system to supply fuel to hydraulically drive the variable compressor vane actuator among the plurality of actuators.

[0188] The fuel is thermally stable at temperatures above 280 °C.

[0189] In various embodiments, the methods of the second, fourth, sixth, and eighth aspects may be complementary and may be performed together, or performed in any combination or sub - combination. The method of the eighth aspect may be performed using the engine of the first, third, fifth, or seventh aspect.

[0190] According to the ninth aspect, there is provided a gas turbine engine for an aircraft. The engine includes:

[0191] An engine core that includes a turbine, a burner, a compressor, and a spool connecting the turbine to the compressor;

[0192] A fan that is located upstream of the engine core and is arranged to be driven by the spool, the fan including a plurality of fan blades;

[0193] A nacelle that surrounds the fan and the engine core and defines a bypass duct located radially outside the engine core, wherein the bypass ratio is at least 4, and the bypass ratio is defined as the ratio of the mass flow rate of the flow passing through the bypass duct to the mass flow rate of the flow passing through the core under cruise conditions;

[0194] A plurality of actuators;

[0195] A fuel supply system, wherein the fuel supply system is arranged to supply fuel for combustion in the burner and to supply fuel to hydraulically drive at least one of the plurality of actuators; and

[0196] At least one fuel - oil heat exchanger, which is arranged to allow oil and fuel to flow therethrough, and the at least one heat exchanger is arranged to transfer heat from the oil to the fuel.

[0197] The at least one heat exchanger is arranged such that during cruise, the fuel temperature when entering at least one actuator is at least 5 °C higher than the fuel temperature when entering the burner.

[0198] The bypass ratio is greater than or equal to 4 and can be in the range of 4 to 55. The bypass ratio can be in the range of 4 to 20. The bypass ratio can be in the range of 4 to 15.

[0199] The at least one heat exchanger can be arranged such that during cruise, the fuel temperature when entering 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 when entering the burner.

[0200] The spool can directly output drive to the fan, thereby driving the fan at the same rotational speed as the spool, such that the engine is a direct drive turbofan engine.

[0201] The turbofan engine can include a gearbox that receives input from the spool and outputs drive to the fan, thereby driving the fan at a rotational speed lower than that of the spool, such that the engine is a geared turbofan engine.

[0202] The at least one heat exchanger can include a primary fuel - oil heat exchanger arranged to heat at least most of the fuel and a secondary fuel - oil heat exchanger arranged to provide additional heat to the fuel to be supplied for hydraulically driving at least one fuel - hydraulic actuator.

[0203] The primary fuel - oil heat exchanger and the secondary fuel - oil heat exchanger can be controlled such that under cruise conditions, the heat transfer ratio:

[0204]

[0205] Has a maximum value of at least 0.35.

[0206] The primary heat exchanger and the secondary heat exchanger can be arranged such that the heat transfer ratio is adjustable during the operation of the engine.

[0207] A gas turbine engine can include at least one of the following to provide adjustment of the heat transfer ratio during operation:

[0208] (i) At least one controllable oil bypass valve arranged to allow oil to bypass at least one of the primary fuel-oil heat exchanger and the secondary fuel-oil heat exchanger; and

[0209] (ii) At least one controllable oil recirculation valve arranged to allow oil to recirculate through at least one of the primary fuel-oil heat exchanger and the secondary fuel-oil heat exchanger.

[0210] The fuel supply system may include a valve operable to enable fuel to bypass or recirculate through the primary heat exchanger.

[0211] The fuel supply system may include a valve operable to enable fuel to bypass or recirculate through the secondary heat exchanger.

[0212] The engine may be arranged to allow the heat transfer ratio to be adjusted based on the sustainable aviation fuel - SAF - content of the fuel. For a higher fuel SAF content, a higher heat transfer ratio may be allowed.

[0213] The fuel supply system may include or be connected to a fuel source. The fuel source may be located external to the engine, such as a tank in the aircraft fuselage or on the aircraft wing, and may not be part of the engine. Fuel supplied to at least one actuator to hydraulically drive at least one actuator may be supplied to the burner after actuation without returning to the fuel source. Thus, fuel may be recirculated around the engine or used in various ways within the engine without returning to the fuel tank.

[0214] The fuel supplied to the burner may include a mixture of fuel that has passed through at least one fuel-oil heat exchanger and fuel that has bypassed at least one fuel-oil heat exchanger.

[0215] The engine of the ninth aspect may include any or all of the features of the engines of the first, third, fifth, and / or seventh aspects and may be used to implement the methods of the second, fourth, sixth, and / or eighth aspects.

[0216] According to a tenth aspect, there is provided a method of operating a gas turbine engine for an aircraft.

[0217] The engine includes:

[0218] An engine core including a turbine, a burner, a compressor, and a shaft connecting the turbine to the compressor;

[0219] A fan located upstream of the engine core and arranged to be driven by the shaft, the fan including a plurality of fan blades;

[0220] A nacelle that surrounds the fan and the engine core and defines a bypass duct located radially outside the engine core, wherein the bypass ratio is at least 4, the bypass ratio being 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 under cruise conditions;

[0221] A plurality of actuators;

[0222] A fuel supply system, wherein the fuel supply system is arranged to supply fuel for combustion in the burner and to supply fuel to hydraulically drive at least one of the plurality of actuators; and

[0223] At least one fuel - oil heat exchanger arranged to have oil and fuel flow therethrough, the at least one heat exchanger being arranged to transfer heat from the oil to the fuel.

[0224] The method includes controlling at least one heat exchanger such that during cruise, the fuel temperature at entry to at least one actuator is at least 5 °C higher than the fuel temperature at entry to the burner.

[0225] The bypass ratio can be in the range of 4 to 55. The bypass ratio can be in the range of 4 to 20. The bypass ratio can be in the range of 4 to 15.

[0226] The method can include determining at least one fuel property of the fuel and controlling the temperature difference between the fuel temperature at entry to at least one actuator and the fuel temperature at entry to the burner based on the at least one fuel property.

[0227] The fuel supply system can obtain fuel from a fuel source, such as an aircraft fuel tank (which is typically external to the engine); and the method can cause the fuel supplied to at least one actuator to hydraulically drive at least one actuator to be supplied to the burner after actuation without returning to the fuel source.

[0228] A gas turbine engine can include a primary fuel - oil heat exchanger and a secondary fuel - oil heat exchanger, and the method can include controlling the primary fuel - oil heat exchanger and the secondary fuel - oil heat exchanger such that under cruise conditions, the heat transfer ratio:

[0229]

[0230] Has a maximum value of at least 0.35.

[0231] The method can include controlling the primary fuel - oil heat exchanger and the secondary fuel - oil heat exchanger to adjust the heat transfer ratio based on the sustainable aviation fuel - SAF - content of the fuel.

[0232] In various embodiments, the methods of the second, fourth, sixth, eighth, and tenth aspects may be complementary and may be performed together, or in any combination or sub-combination. The method of the tenth aspect may be performed using the engine of the first, third, fifth, seventh, or ninth aspect.

[0233] It should be understood that the features described with respect to one aspect may be used in combination with any other aspect, with necessary modifications.

[0234] As described elsewhere herein, the present disclosure may be applied 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 (where the propeller is not enclosed by a nacelle), a turboprop engine, or a turbojet engine. Any such engine may or may not be provided with a afterburner. Such a gas turbine engine may be configured, for example, for land-based or marine power generation applications.

[0235] A gas turbine engine according to any aspect of the present disclosure may include an engine core that includes a turbine, a combustor, a compressor, and a spool connecting the turbine to the compressor. Such a gas turbine engine may include a fan (with fan blades). Such a fan may be located upstream of the engine core. Alternatively, in some examples, the gas turbine engine may include a fan located downstream of the engine core, such as in the case where the gas turbine engine is an open rotor or turboprop engine (in which case the fan may be referred to as a propulsor).

[0236] In the case where the gas turbine engine is an open rotor or turboprop engine, the gas turbine engine may include two contra-rotating propeller stages attached to and driven by a free power turbine via a shaft. The propulsors may rotate in opposite directions such that one propulsor rotates clockwise about the axis of rotation of the engine and the other propulsor rotates counterclockwise about the axis of rotation of the engine. Alternatively, the gas turbine engine may include a propulsor stage and a stator vane stage configured downstream of the propulsor stage. The stator vane stage may have variable pitch. Thus, the high-pressure, intermediate-pressure, and free power turbines may drive the high-pressure and intermediate-pressure compressors and the propulsors respectively via suitable interconnecting shafts. Thus, the propulsors may provide most of the propulsive thrust.

[0237] In the case where the gas turbine engine is an open rotor or turboprop engine, one or more propulsor stages may be driven by a gearbox. The gearbox may be of the type described herein.

[0238] 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 directly connected to the fan drive turbine via a spool, for example without a gearbox. In such a direct drive turbofan engine, it can be said that the fan rotates at the same rotational speed as the fan drive turbine. By way of example only, the fan drive turbine may be a first turbine, the spool may be a first spool, and the gas turbine engine may further include a second turbine and a second spool connecting the second turbine to the compressor. The second turbine, compressor and second spool may be arranged to rotate at a higher rotational speed than the first spool. In such an arrangement, the second turbine may be axially positioned upstream of the first turbine.

[0239] An engine according to the present disclosure may be a geared turbofan engine. In such an arrangement, the engine has a fan driven via a gearbox. Thus, such a gas turbine engine may include a gearbox that receives an input from a spool and outputs a drive to the fan so as to drive the fan at a lower rotational speed than the spool. The input to the gearbox may come directly from the spool or indirectly from the spool, such as via a spur shaft and / or gears. The spool may rigidly connect the turbine and the compressor such that the turbine and the compressor rotate at the same speed (wherein the fan rotates at a lower speed).

[0240] The gas turbine engine as described and / or claimed herein may have any suitable general architecture. For example, the gas turbine engine may have any desired number of shafts connecting the turbine and the compressor, such as one shaft, two shafts or three shafts. By way of example only, the turbine connected to the spool may be a first turbine, the compressor connected to the spool may be a first compressor, and the spool may be a first spool. The engine core may further include a second turbine, a second compressor and a second spool connecting the second turbine to the second compressor. The second turbine, second compressor and second spool may be arranged to rotate at a higher rotational speed than the first spool.

[0241] In such an arrangement, the second compressor may be axially positioned downstream of the first compressor. The second compressor may be arranged to receive the flow (e.g., directly receive, such as via a substantially annular conduit) from the first compressor.

[0242] The gearbox may be arranged to be driven by a spool (such as the first spool in the above example) configured to rotate at the lowest rotational speed (e.g., in use). For example, the gearbox may be arranged to be driven only by a spool configured to rotate at the lowest rotational speed (e.g., in the above example, only by the first spool and not the second spool). Alternatively, the gearbox may be arranged to be driven by any one or more shafts, such as the first shaft and / or the second shaft in the above example.

[0243] The gearbox can be a reduction gearbox (since the output to the fan has a lower rotational speed than the input from the spool). Any type of gearbox can be used. For example, the gearbox can be a "planetary" or "stellar" gearbox, as described in more detail elsewhere herein. Such a gearbox can be single stage. Alternatively, such a gearbox can be a compound gearbox, such as a compound planetary gearbox (which can have an input on the sun gear and an output on the ring gear and is thus referred to as a "compound star" gearbox), for example having two stages of reduction.

[0244] The gearbox can have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example greater than 2.5, for example in the range of 3 to 4.2, or 3.2 to 3.8, for example, approximately 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. For example, the gear tooth ratio can be between any two values in the previous sentence. By way of example only, the gearbox can be a "stellar" gearbox having a reduction ratio in the range of 3.1 or 3.2 to 3.8. By way of another example only, the gearbox can be a "stellar" gearbox having a reduction ratio in the range of 3.0 to 3.1. By way of another example only, the gearbox can be a "planetary" gearbox having a reduction ratio in the range of 3.6 to 4.2. In some arrangements, the gear tooth ratio can be outside of these ranges.

[0245] In any gas turbine engine as described and / or claimed herein, fuel of a given composition or blend is provided to a combustor, which can be disposed downstream (e.g., axially downstream) of the fan and compressor with respect to the flow path. For example, in the case where a second compressor is provided, the combustor can be located directly downstream of the second compressor (e.g., at its outlet). By way of another example, in the case where a second turbine is provided, the flow at the combustor outlet can be provided to the inlet of the second turbine. The combustor can be disposed upstream of one or more turbines.

[0246] The compressor or each compressor (e.g., the first and second compressors as described above) may include any number of stages, such as a plurality of stages. Each stage may include a row of rotor blades and a row of stator vanes, and the row of stator vanes may be variable stator vanes (since the angle of incidence of the row of stator vanes may be variable). The row of rotor blades and the row of stator vanes may be axially offset from each other. For example, a gas turbine engine may be a direct drive turbofan gas turbine engine including 13 or 14 compressor stages (in addition to the fan). Such an engine may include, for example, 3 stages in the first (or "low pressure") compressor and 10 or 11 stages in the second (or "high pressure") compressor. By way of another example, a gas turbine engine may be a "geared" gas turbine engine (where the fan is driven by a first shaft via a reduction gearbox) including 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. By way of another example, a gas turbine engine may be a "geared" gas turbine engine having 4 stages in the first (or "low pressure") compressor and 10 stages in the second (or "high pressure") compressor.

[0247] The turbine or each turbine (e.g., the first and second turbines as described above) may include any number of stages, such as a plurality of stages. As required, each stage may include a row of rotor blades and a row of stator vanes, and vice versa. The corresponding rows of rotor blades and stator vanes may be axially offset from each other. The second (or "high pressure") turbine may include 2 stages in any arrangement (e.g., regardless of whether it is a geared engine or a direct drive engine). A gas turbine engine may be a direct drive gas turbine engine including a first (or "low pressure") turbine having 5, 6 or 7 stages. Alternatively, a gas turbine engine may be a "geared" gas turbine engine including a first (or "low pressure") turbine having 3 or 4 stages.

[0248] Each fan blade can be defined as having a radial span that extends from a root (or hub) at a radially inner gas washing location or 0% span location to a tip at the 100% span location. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip can be less than (or approximately) any of the following: 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 can be within the inclusive range defined by any two values in the previous sentence (i.e., these values can form an upper or lower limit), for example, in the range of 0.28 to 0.32 or 0.29 to 0.30. These ratios are generally referred to as hub-tip ratios. The radius at the hub and the radius at the tip can both be measured at the leading edge (or axially foremost) portion of the blade. Of course, the hub-tip ratio refers to the gas washing portion of the fan blade, i.e., the portion radially outside of any platform.

[0249] The radius of the fan can be measured between the engine centerline and the tip at the leading edge of the fan blade. The fan diameter (which may simply be twice the fan radius) can be greater than (or approximately) any of the following: 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 the inclusive range defined by any two values in the previous sentence (i.e., these values can form an upper or lower limit), for example, in the range of 210 cm to 240 cm, or 250 cm to 280 cm, or 320 cm to 380 cm. By way of non-limiting example only, the fan diameter can be in the range of 170 cm to 180 cm, 190 cm to 200 cm, 200 cm to 210 cm, 210 cm to 230 cm, 290 cm to 300 cm, or 340 cm to 360 cm.

[0250] The rotational speed of the fan can vary during use. Generally speaking, for a fan with a relatively large diameter, the rotational speed is lower. By way of non-limiting examples only, the rotational speed of the fan under cruise conditions can be less than 3500 rpm, such as less than 2600 rpm, or less than 2500 rpm, or less than 2300 rpm. By way of additional non-limiting examples only, for a "geared" gas turbine engine with a fan diameter in the range of 200 cm to 210 cm, the rotational speed of the fan under cruise conditions can be in the range of 2750 rpm to 2900 rpm. By way of additional non-limiting examples only, for a "geared" gas turbine engine with a fan diameter in the range of 210 cm to 230 cm, the rotational speed of the fan under cruise conditions can be in the range of 2500 rpm to 2800 rpm. By way of additional non-limiting examples only, for a "geared" gas turbine engine with a fan diameter in the range of 340 cm to 360 cm, the rotational speed of the fan under cruise conditions can be in the range of 1500 rpm to 1800 rpm. By way of additional non-limiting examples only, for a direct drive engine with a fan diameter in the range of 190 cm to 200 cm, the rotational speed of the fan under cruise conditions can be in the range of 3600 rpm to 3900 rpm. By way of additional non-limiting examples only, for a direct drive engine with a fan diameter in the range of 300 cm to 340 cm, the rotational speed of the fan under cruise conditions can be in the range of 2000 rpm to 2800 rpm.

[0251] When using a gas turbine engine, the fan (with associated fan blades) rotates about a rotational axis. This rotation causes the tips of the fan blades to move at a speed U 尖端 . The work done by the fan blades on the flow results in an enthalpy rise dH of the flow. The fan tip loading can be defined as dH / U 尖端 2 , where dH is the enthalpy rise across the fan (e.g., 1-D mean enthalpy rise), and U 尖端 is the (translational) speed of the fan tip, such as 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 approximately) any of the following: 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 are dimensionless). The fan tip loading can be within the inclusive range defined by any two of the values in the previous sentence (i.e., these values can form an upper or lower limit), such as in the range of 0.28 to 0.31 or 0.29 to 0.3 (e.g., for a geared gas turbine engine).

[0252] A gas turbine engine according to the present disclosure may have any desired bypass ratio (BPR), where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core. In some arrangements, the bypass ratio at cruise conditions may be greater than (or approximately) any of the following: 9, 9.5, 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5, or 20. The bypass ratio at cruise conditions may be within the inclusive range defined by any two of the values in the previous sentence (i.e., these values may form an upper or lower limit), such as in the range of 12 to 16, or 13 to 15, or 13 to 14. By way of non-limiting example only, the bypass ratio at cruise conditions of a direct drive gas turbine engine according to the present disclosure may be in the range of 9:1 to 11:1. By way of another non-limiting example only, the bypass ratio at cruise conditions of a geared gas turbine engine according to the present disclosure may be in the range of 12:1 to 15:1. The bypass duct may be at least substantially annular. The bypass duct may be located radially outside the core engine. The radially outer surface of the bypass duct may be defined by a nacelle and / or a fan casing.

[0253] The overall pressure ratio (OPR) of a gas turbine engine as described and / or claimed herein may be defined as the ratio of the stagnation pressure at the exit of the highest pressure compressor (before entering the combustor) to the stagnation pressure upstream of the fan. By way of non-limiting example, the overall pressure ratio of a gas turbine engine as described and / or claimed herein at cruise conditions may be greater than (or approximately) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio may be within the inclusive range defined by any two of the values in the previous sentence (i.e., these values may form an upper or lower limit), such as in the range of 50 to 70. By way of non-limiting example only, the overall pressure ratio at cruise conditions of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 40 to 45. By way of non-limiting example only, the overall pressure ratio at cruise conditions of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 45 to 55. By way of non-limiting example only, the overall pressure ratio at cruise conditions of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 50 to 60. By way of non-limiting example only, the overall pressure ratio at cruise conditions of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 50 to 60.

[0254] The specific thrust of an engine can be defined as the net thrust of the engine divided by the total mass flow through the engine. In some examples, for a given thrust condition, the specific thrust can depend on the specific composition of the fuel provided to the combustor. Under cruise conditions, the specific thrust of the engines described and / or claimed herein can be less than (or approximately) any of the following: 110 N / kg -1 s, 105 N / kg -1 s, 100 N / kg -1 s, 95 N / kg -1 s, 90 N / kg -1 s, 85 N / kg -1 s or 80 N / kg -1 s. The specific thrust can be within the inclusive range defined by any two of the values in the previous sentence (i.e., these values can form an upper or lower limit), e.g., between 80 N / kg -1 s and 100 N / kg -1 s, or between 85 N / kg -1 s and 95 N / kg -1 s. Compared to conventional gas turbine engines, such engines can be particularly efficient. By way of non-limiting example only, the specific thrust of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm can be in the range of 90 N / kg -1 s to 95 N / kg -1 s. By way of non-limiting example only, the specific thrust of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm can be in the range of 80 N / kg -1 s to 90 N / kg -1 s. By way of non-limiting example only, the specific thrust of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm can be in the range of 70 N / kg -1 s to 90 N / kg -1 s. By way of non-limiting example only, the specific thrust of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm can be in the range of 90 N / kg -1 s to 120 N / kg -1 s.

[0255] The gas turbine engines as described herein and / or claimed may have any desired maximum thrust. By way of non-limiting example only, the gas turbines as described herein and / or claimed may produce a maximum thrust of at least (or approximately) any one 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 the inclusive range defined by any two of the values in the previous sentence (i.e., these values may form an upper or lower limit). By way of non-limiting example only, the gas turbines as described herein and / or claimed may be capable of producing a maximum thrust within 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 having a fan diameter within the range of 200 cm to 210 cm may be within the range of 140 kN to 160 kN. By way of non-limiting example only, the maximum thrust of a geared gas turbine engine having a fan diameter within the range of 210 cm to 230 cm may be within the range of 150 kN to 200 kN. By way of non-limiting example only, the maximum thrust of a geared gas turbine engine having a fan diameter within the range of 340 cm to 360 cm may be within 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 having a fan diameter within the range of 300 cm to 340 cm may be within the range of 370 kN to 500 kN. The thrust mentioned above may be the maximum net thrust at standard atmospheric conditions, at sea level, plus 15 °C (ambient pressure 101.3 kPa, temperature 30 °C), with the engine stationary.

[0256] In use, the temperature of the flow at the inlet of the high-pressure turbine can be particularly high. This temperature, which may be referred to as TET, can be measured at the outlet of the combustor, for example, just upstream of the first turbine blade which may itself be referred to as the nozzle guide vane. In some examples, for a given thrust condition, the TET can depend on the particular composition of the fuel supplied to the combustor. Under cruise conditions, the TET can be at least (or approximately) any one of the following: 1400K, 1450K, 1500K, 1520K, 1530K, 1540K, 1550K, 1600K or 1650K. Thus, by way of non-limiting example only, a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm can have a TET in the range of 1540K to 1600K under cruise conditions. By way of non-limiting example only, a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm can have a TET in the range of 1590K to 1650K under cruise conditions. By way of non-limiting example only, a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm can have a TET in the range of 1600K to 1660K under cruise conditions. By way of non-limiting example only, a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm can have a TET in the range of 1590K to 1650K under cruise conditions. By way of non-limiting example only, a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm can have a TET in the range of 1570K to 1630K under cruise conditions.

[0257] The TET under cruise conditions can be within an inclusive range defined by any two of the values in the previous sentence (i.e., these values can form an upper or lower limit), for example, from 1530K to 1600K. The maximum TET during engine operation can be, for example, at least (or approximately) any of the following: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K, 2000K, 2050K, or 2100K. Thus, by way of non-limiting example only, the maximum TET of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm can be in the range of 1890K to 1960K. By way of non-limiting example only, the maximum TET of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm can be in the range of 1890K to 1960K. By way of non-limiting example only, the maximum TET of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm can be in the range of 1890K to 1960K. By way of non-limiting example only, the maximum TET of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm can be in the range of 1935K to 1995K. By way of non-limiting example only, the maximum TET of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm can be in the range of 1890K to 1950K. The maximum TET can be within an inclusive range defined by any two of the values in the previous sentence (i.e., these values can form an upper or lower limit), for example, in the range of 1800K to 1950K or 1900K to 2000K. The maximum TET can occur, for example, under high thrust conditions, such as under maximum takeoff (MTO) conditions.

[0258] The fan blades and / or the airfoil portions of the fan blades described and / or claimed herein can be manufactured from any suitable material or combination of materials. For example, at least a portion of the fan blades and / or airfoils can be manufactured at least in part from a composite material, such as a metal matrix composite and / or an organic matrix composite, such as a carbon fiber composite. By way of further example, at least a portion of the fan blades and / or airfoils can be manufactured at least in part from a metal, such as a titanium-based metal or an aluminum-based material (such as an aluminum-lithium alloy) or a steel-based material. The fan blades can include at least two regions manufactured using different materials. For example, the fan blades can have a protective leading edge that can be manufactured using a material that better resists impact (e.g., from birds, ice, or other materials) than the rest of the blade. Such leading edges can be manufactured, for example, using titanium or a titanium-based alloy. Thus, by way of example only, the fan blade can have a carbon fiber or an aluminum-based body (such as an aluminum-lithium alloy) with a titanium leading edge.

[0259] The fan as described and / or claimed herein may include a central portion from which fan blades may extend, such as radially. The fan blades may be attached to the central portion in any desired manner. For example, each fan blade may include a fixture that may engage a corresponding slot in a hub (or disc portion). By way of example only, such a fixture may be in the form of a dovetail that may be inserted into and / or engage a corresponding slot in the hub / disc portion to secure the fan blade to the hub / disc portion. By way of a further example, the fan blades may be integrally formed with the central portion. Such an arrangement may be referred to as a bladed disc or a bladed ring. Any suitable method may be used to manufacture such a bladed disc or bladed ring. For example, at least a portion of the fan blade may be machined from a block, and / or at least part of the fan blade may be attached to the hub / disc portion by welding (such as linear friction welding).

[0260] The gas turbine engine described and / or claimed herein may or may not be provided with a variable area nozzle (VAN). Such a variable area nozzle may allow the exit area of the bypass duct to vary in use. The general principles of the present disclosure may be applied to engines with or without a VAN.

[0261] The fan of a gas turbine as described and / or claimed herein may have any desired number of fan blades, such as 14, 16, 18, 20, 22, 24 or 26 fan blades. Where the fan blades have a carbon fiber composite body, there may be 16 or 18 fan blades. Where the fan blades have a metallic body (such as an aluminum-lithium or titanium alloy), there may be 18, 20 or 22 fan blades.

[0262] As used herein, the terms idle, taxi, take-off, climb, cruise, descent, approach and landing (or one or more parts thereof) have their conventional meanings and will be readily understood by the person skilled in the art. Thus, for a given gas turbine engine for an aircraft, the person skilled in the art will immediately recognize that each term refers to all or one or more parts of the operating phase of the engine within a given mission of the aircraft to which the gas turbine engine is designed to be attached.

[0263] At this point, ground idle may refer to an engine operating phase in which the aircraft is stationary and in contact with the ground, but in which there are requirements for the engine to be operated. During idle, the engine may produce available thrust between 3% and 9% of the engine. In a further non-limiting example, the engine may produce available thrust between 5% and 8% of the engine. In a further non-limiting example, the engine may produce available thrust between 6% and 7% of the engine. Taxiing may refer to an engine operating phase in which the aircraft is propelled along the ground by thrust produced by the engine. During taxiing, the engine may produce available thrust between 5% and 15% of the engine. In a further non-limiting example, the engine may produce available thrust between 6% and 12% of the engine. In a further non-limiting example, the engine may produce available thrust between 7% and 10% of the engine. Takeoff may refer to an engine operating phase in which the aircraft is propelled by thrust produced by the engine. During the initial phase of the takeoff phase, the aircraft may be propelled while in contact with the ground. During a later phase of the takeoff phase, the aircraft may be propelled while not in contact with the ground. During takeoff, the engine may produce available thrust between 90% and 100% of the engine. In a further non-limiting example, the engine may produce available thrust between 95% and 100% of the engine. In a further non-limiting example, the engine may produce 100% of the available thrust.

[0264] Climb may refer to an engine operating phase in which the aircraft is propelled by thrust produced by the engine. During climb, the engine may produce available thrust between 75% and 100% of the engine. In a further non-limiting example, the engine may produce available thrust between 80% and 95% of the engine. In a further non-limiting example, the engine may produce available thrust between 85% and 90% of the engine. At this point, climb may refer to an operating phase during the aircraft flight cycle between takeoff and reaching cruise conditions, such that reaching cruise conditions defines the start of the cruise phase or a portion thereof of the aircraft flight. Additionally or alternatively, climb may refer to a nominal point or one or more nominal time periods during the aircraft flight cycle between takeoff and landing, in which a relative increase in altitude is required, which may require additional thrust requirements of the engine.

[0265] As used herein, the cruise conditions that may define the cruise phase (or a portion thereof) of an aircraft flight have their conventional meaning and are readily understood by a person skilled in the art. In some examples, for a given gas turbine engine of an aircraft, the cruise conditions may refer to the operating point of the engine during the intermediate cruise of a given mission of the aircraft to which the gas turbine engine is designed to be attached (which may be referred to in the industry as the "economic mission"). In this regard, the intermediate cruise may be considered as the point in the aircraft flight cycle at which 50% of the total fuel burned between the highest point of climb and the start of descent has been burned (which may approximate the midpoint between the highest point of climb and the start of descent in terms of time and / or distance). Thus, the cruise conditions may define the operating point, phase or a portion thereof of the flight that, considering the number of engines provided to the aircraft to which the gas turbine engine is designed to be attached, provides the thrust that will ensure the steady-state operation (i.e., maintaining a constant altitude and / or a constant Mach number) or at least substantially steady-state operation (i.e., maintaining at least substantially a constant altitude and / or at least substantially a constant Mach number) of the aircraft. For example, if the engine is designed to be attached to an aircraft having two engines of the same type, then under the cruise conditions, the engine may provide half of the total thrust required for the steady-state operation or at least substantially steady-state operation of the aircraft during the intermediate cruise.

[0266] In other words, for a given gas turbine engine of an aircraft, the cruise conditions may be defined as the operating point of the engine that provides a specified thrust under intermediate cruise atmospheric conditions (defined by the International Standard Atmosphere according to ISO 2533 at the intermediate cruise altitude), which is required to provide the steady-state operation or at least substantially steady-state operation of the aircraft to which the gas turbine engine is designed to be attached, in combination with any other engines on the aircraft, at a given intermediate cruise Mach number. For any given gas turbine engine of an aircraft, the intermediate cruise thrust, atmospheric conditions and Mach number are known, and thus the operating point of the engine under the cruise conditions can be well-defined.

[0267] By way of example only, the forward speed under the cruise conditions may be any point within the range from 0.7 Mach to 0.9 Mach, such as from 0.75 to 0.85, such as from 0.76 to 0.84, such as from 0.77 to 0.83, such as from 0.78 to 0.82, such as from 0.79 to 0.81, such as approximately 0.8 Mach, approximately 0.85 Mach or within the range from 0.8 to 0.85. Any single speed within these ranges can be part of the cruise conditions. For some aircraft, the cruise conditions may be outside these ranges, such as below 0.7 Mach or above 0.9 Mach.

[0268] By way of example only, the cruise conditions may correspond to standard atmospheric conditions (according to the International Standard Atmosphere ISA) at altitudes within the following ranges: 10000m to 15000m, for example within the range of 10000m to 12000m, for example within the range of 10400m to 11600m (about 38000 feet), for example within the range of 10500m to 11500m, for example within the range of 10600m to 11400m, for example within the range of 10700m (about 35000 feet) to 11300m, for example within the range of 10800m to 11200m, for example within the range of 10900m to 11100m, for example approximately 11000m. The cruise conditions may correspond to the standard atmospheric conditions at any given altitude within these ranges.

[0269] By way of example only, the 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 35000ft (10668m). Under such cruise conditions, the engine may provide a known required net thrust level. The known required net thrust level of course depends on the engine and its intended application and may be a value, for example, within the range of 20kN to 40kN.

[0270] By way of further example, the 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 38000ft (11582m). Under such cruise conditions, the engine may provide a known required net thrust level. The known required net thrust level of course depends on the engine and its intended application and may be a value, for example, within the range of 35kN to 65kN.

[0271] In use, the gas turbine engine described and / or claimed herein may operate under cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions of an aircraft (e.g., intermediate cruise conditions) on which at least one (e.g., 2 or 4) gas turbine engines may be installed to provide propulsion thrust.

[0272] In addition, those skilled in the art will immediately recognize that either or both of landing and approach refer to the operational phases between the cruise and landing of an aircraft during its flight cycle, and in particular, the approach forms part of the landing and take-off (LTO) phase. During either or both of landing and approach, the engine may produce between 0% and 50% of the available thrust. In a further non-limiting example, the engine may produce between 25% and 40% of the available thrust. In a further non-limiting example, the engine may produce between 30% and 35% of the available thrust. Additionally or alternatively, landing may refer to the nominal point in the flight cycle of an aircraft between take-off and landing where a relative reduction in altitude is required and this may require a reduced thrust demand from the engine.

[0273] According to one aspect, there is provided an aircraft that includes a gas turbine engine as described herein and / or claimed. The aircraft according to this aspect is the aircraft to which the gas turbine engine has been designed to be attached. Thus, the cruise condition according to this aspect may correspond to an operating point, phase, or part thereof of the flight of the aircraft, as defined elsewhere herein.

[0274] According to one aspect, there is provided a method of operating a gas turbine engine as described herein and / or claimed. The operation may be carried out under any cruise condition (e.g., in terms of thrust, atmospheric conditions, and Mach number) that may be defined elsewhere herein.

[0275] According to one aspect, there is provided a method of operating an aircraft that includes a gas turbine engine as described herein and / or claimed. The operation according to this aspect may include (or may be) an operation under any suitable condition (e.g., at the intermediate cruise of the aircraft), as defined elsewhere herein.

[0276] The skilled person will understand that features or parameters described with respect to any one of the above aspects may be applied to any other aspect, unless mutually exclusive. Furthermore, 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, unless mutually exclusive.

[0277] Unless mutually exclusive, any parameter or value included or described herein may be applied to any one or more additional parameters and / or values included or described herein and / or combined therewith. For example, a first parameter or value (e.g., parameter A) included or described herein may be applied to any one or more additional parameters and / or values (e.g., parameter B; parameter C; and parameter D, etc., one or more of them) included or described herein and / or combined therewith to express the product of their relationships. For example, the skilled person in the art will understand that in the case where parameter A and parameter B are disclosed separately, the product of their relationship may be expressed, as required, for example, as A / B, B / A, B*A, or any such further application, combination, or function of parameter A with respect to parameter B. Description of the Drawings

[0278] Embodiments will now be described, by way of example only, with reference to the drawings, in which:

[0279] Figure 1 is a cross-sectional side view of a gas turbine engine;

[0280] Figure 2 is a close-up cross-sectional side view of the upstream portion of a geared gas turbine engine;

[0281] Figure 3 is a partial cross-sectional view of a gearbox for a gas turbine engine;

[0282] Figure 4 is a close-up cross-sectional side view of an upstream portion of a direct drive gas turbine engine;

[0283] Figure 5 is an illustration of an aircraft having a propulsion system including two gas turbine engines;

[0284] Figure 6 is a schematic diagram showing a fuel supply system including one fuel hydraulic actuator;

[0285] Figure 7 is a schematic diagram showing a fuel supply system including two fuel hydraulic actuators;

[0286] Figure 8 is a schematic diagram showing a fuel supply system including one selective fuel hydraulic actuator;

[0287] Figure 9 is a schematic diagram showing an alternative fuel supply system including one selective fuel hydraulic actuator;

[0288] Figure 10 is a schematic diagram showing another fuel supply system including two fuel hydraulic actuators, one of which is a selective fuel hydraulic actuator;

[0289] Figure 11 is a schematic diagram showing a fuel supply system including two selective fuel hydraulic actuators;

[0290] Figure 12 is a schematic diagram showing another fuel supply system including two selective fuel hydraulic actuators;

[0291] Figure 13 is a schematic diagram showing another fuel supply system including one selective fuel hydraulic actuator;

[0292] Figure 14 is a schematic diagram showing another fuel supply system including one selective fuel hydraulic actuator;

[0293] Figure 15 is a schematic diagram showing a fuel supply system including one selective fuel hydraulic actuator operable to control a valve within an engine thermal management system;

[0294] Figure 16FIG. is a schematic diagram showing another fuel supply system including a selective fuel hydraulic actuator operable to control a valve within an engine heat management system;

[0295] Figure 17 FIG. is a schematic diagram showing a fuel supply system including a selective fuel hydraulic actuator operable to control a valve within a generator heat management system;

[0296] Figure 18 FIG. is a schematic diagram showing another fuel supply system including a selective fuel hydraulic actuator operable to control a valve within a generator heat management system;

[0297] Figure 19 FIG. is a schematic diagram showing a fuel supply system including a selective fuel hydraulic actuator and two heat exchangers;

[0298] Figure 20 FIG. is a flowchart showing an example method of operating a gas turbine engine;

[0299] Figure 21 FIG. is a flowchart showing another example method of operating a gas turbine engine;

[0300] Figure 22 FIG. is a flowchart showing another example method of operating a gas turbine engine;

[0301] Figure 23 FIG. is a flowchart showing another example method of operating a gas turbine engine;

[0302] Figure 24 FIG. is a flowchart showing another example method of operating a gas turbine engine; and

[0303] Figure 25 shows a fuel supply system of a gas turbine engine including two fuel - oil heat exchangers. DETAILED DESCRIPTION

[0304] Figure 1Figure 10 shows a gas turbine engine 10 having a primary rotational axis 9. The engine 10 includes an air intake 12 and a propulsive fan 23 which generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 includes a core 11 which receives the core airflow A. The engine core 11 includes, in axial flow series, a low pressure compressor 14, a high pressure compressor 15, combustion equipment 16, a high pressure turbine 17, a low pressure turbine 19 and a core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass duct 22. The fan 23 is attached to and driven by the low pressure turbine 19 via a shaft 26 and an epicyclic gearbox 30.

[0305] In use, the core airflow A is accelerated and compressed by the low pressure compressor 14 and is directed into the high pressure compressor 15 for further compression. The compressed air discharged from the high pressure compressor 15 is directed into the combustion equipment 16 where the compressed air is mixed with fuel F and the mixture is combusted. The combustion equipment 16 may be referred to as a burner 16, where the terms "combustion equipment 16" and "burner 16" may be used interchangeably herein. The resulting hot combustion products then expand through the high pressure turbine and the low pressure turbines 17, 19 before being discharged through the nozzle 20, thereby driving the high pressure turbine and the low pressure turbines to provide some propulsive thrust. The high pressure turbine 17 drives the high pressure compressor 15 via a suitable interconnecting shaft 27. The fan 23 is generally used to apply an increased pressure to the bypass airflow B flowing through the bypass duct 22 such that the bypass airflow B is discharged through the bypass exhaust nozzle 18 to generally provide most of the propulsive thrust. The epicyclic gearbox 30 is a reduction gearbox.

[0306] Figure 2 Figure 9 shows an exemplary arrangement of a geared fan gas turbine engine 10. The low pressure turbine 19 (see Figure 1 ) drives a shaft 26 which is coupled to a sun gear or sunwheel 28 of an epicyclic gear arrangement 30. Radially outward of and meshing with the sun gear 28 are a plurality of planet gears 32 which are joined together by a planet carrier 34. The planet carrier 34 constrains the planet gears 32 to precess synchronously about the sun gear 28 while each planet gear 32 rotates about its own axis. The planet carrier 34 is coupled to the fan 23 via a link 36 so as to drive the fan to rotate about the engine axis 9. Radially outward of and meshing with the planet gears 32 is a ring gear or annulus 38 which is coupled to a fixed support structure 24 via a link 40.

[0307] Note that, as used herein, the terms "low pressure turbine" and "low pressure compressor" may refer to the lowest pressure turbine stage and the lowest pressure compressor stage, respectively (i.e., excluding the fan 23), and / or the turbine stage and the compressor stage connected together by an interconnecting shaft 26 having the lowest rotational speed in the engine (i.e., excluding the gearbox output shaft driving the fan 23). In some literature, the "low pressure turbine" and "low pressure compressor" referred to herein may alternatively be referred to as "intermediate pressure turbine" and "intermediate pressure compressor". In the case of using such alternative nomenclature, the fan 23 may be referred to as the first or lowest pressure compression stage.

[0308] The Figure 3 epicyclic gearbox 30 is shown in more detail by way of example in Figure 3 . Each of the sun gear 28, the planet gears 32 and the ring gear 38 includes teeth around its periphery for meshing with other gears. However, for clarity,

[0309] only an exemplary portion of the teeth is shown in Figure 2 and Figure 3 . Four planet gears 32 are shown, but it will be apparent to those skilled in the art that more or fewer planet gears 32 may be provided. The actual application of the planetary epicyclic gearbox 30 typically includes at least three planet gears 32.

[0310] It should be understood that Figure 2 and Figure 3 the arrangements shown in Figure 2 are by way of example only, and various alternatives are within the scope of the present disclosure. By way of example only, any suitable arrangement may be used to locate the gearbox 30 in the engine 10 and / or to connect the gearbox 30 to the engine 10. By way of another example, the connecting members (such as Figure 2 the links 36, 40 in the example) between the gearbox 30 and other components of the engine 10 (such as the input shaft 26, the output shaft and the fixed structure 24) may have any desired degree of stiffness or flexibility. By way of a further example, any suitable arrangement of bearings between the rotating and fixed components of the engine (e.g., between the input and output shafts from the gearbox and a fixed structure such as the gearbox housing) may be used, and the present disclosure is not limited toFigure 2 Exemplary arrangements. For example, in the case where the gearbox 30 has a stellar arrangement (as described above), a person skilled in the art will readily understand that the arrangements of the output link, the support link, and the bearing positions will generally be different from Figure 2 the arrangement shown by way of example in

[0311] Accordingly, the present disclosure extends to gas turbine engines having any arrangement in terms of gearbox type (e.g., stellar or planetary gears), support structure, input and output shaft arrangements, and bearing positions.

[0312] Optionally, the gearbox may drive additional and / or alternative components (e.g., an intermediate-pressure compressor and / or a booster compressor).

[0313] Other gas turbine engines to which the present disclosure may apply may have alternative configurations. For example, such engines may have alternative numbers of compressors and / or turbines and / or alternative numbers of interconnected shafts. By way of a further example, Figure 1 the gas turbine engine shown in

[0314] has split nozzles 18, 20, which means that the flow through the bypass duct 22 has its own nozzle 18, which is separate from and radially external to the core engine nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also apply to an engine in which the flow through the bypass duct 22 and the flow through the core 11 mix or combine before (or upstream of) a single nozzle that may be referred to as a mixed-flow nozzle. One or both nozzles (whether mixed or split) may have a fixed or variable area. Figure 4 A cross-sectional view of one such engine is shown in

[0315] Referring to Figure 4 , a gas turbine engine is generally designated 10 and has a main rotational axis 9. The engine 10 includes, in an axial-flow series arrangement, an air intake 12, a propulsive fan 23, an intermediate-pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, an intermediate-pressure turbine 19a, a low-pressure turbine 19, and an exhaust nozzle 20. A nacelle 21 surrounds the engine 10 and defines the air intake 12 and the exhaust nozzle 20.

[0316] In use, air entering the air intake 12 is accelerated by the fan 23 to produce two airflows: a core airflow A and a bypass airflow B. The core airflow A flows into the intermediate-pressure compressor 14, and the bypass airflow B passes through the bypass duct 22 to provide propulsive thrust. The intermediate-pressure compressor 14 compresses the airflow A before delivering the air to the high-pressure compressor 15 where further compression occurs.

[0317] The compressed air discharged from the high-pressure compressor 15 is directed to the combustion equipment 16, where the compressed air is mixed with the fuel F and the mixture is combusted. The combustion equipment 16 may be referred to as the burner 16, and the terms "combustion equipment 16" and "burner 16" may be used interchangeably herein. Then, the resulting hot combustion products expand through the high-pressure turbine 17, the medium-pressure turbine 19a, and the low-pressure turbine 19 before being discharged through the nozzle 20, thereby driving the high-pressure turbine, the medium-pressure turbine, and the low-pressure turbine to provide additional propulsion thrust. The high-pressure turbine 17, the medium-pressure turbine 19a, and the low-pressure turbine 19 each drive the high-pressure compressor 15, the medium-pressure compressor 14, and the fan 23 respectively through suitable interconnecting shafts.

[0318] Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. By way of example, such engines may have an alternative number of interconnecting shafts (e.g., two) and / or an alternative number of compressors and / or turbines. Additionally, the engine may include a gearbox disposed in the drive train from the turbine to the compressor and / or the fan.

[0319] Although the example described relates to a turbofan engine, the present disclosure may be applied to, for example, any type of gas turbine engine, such as an open rotor (where the fan stage is not surrounded by a nacelle) or, for example, a turboprop engine. In some arrangements, the gas turbine engine 10 may not include the gearbox 30.

[0320] The geometry of the gas turbine engine 10 and its components are defined by a conventional shaft system, including an axial direction (aligned with the rotational axis 9), a radial direction (the direction from bottom to top in Figure 1 ), and a circumferential direction (perpendicular to the Figure 1 view of the page). The axial direction, the radial direction, and the circumferential direction are perpendicular to each other.

[0321] The fuel F supplied to the combustion equipment 16 may include a fossil-based hydrocarbon fuel, such as kerosene. Thus, the fuel F may include molecules from one or more of the chemical families of normal paraffins, isoparaffins, naphthenes, and aromatics. Additionally or alternatively, when blended with, mixed with, or replaced by an alternative fuel, the fuel F may include renewable hydrocarbons produced from biological or non-biological resources, also referred to as sustainable aviation fuel (SAF). In each provided embodiment, the fuel F may include one or more trace elements, including, for example, sulfur, nitrogen, oxygen, inorganic substances, and metals.

[0322] SAF is understood by those skilled in the art to refer to, for example, biofuels, renewable aviation fuels, renewable jet fuels, alternative fuels, or biojet fuels produced from biological or non-biological resources. SAF is understood to generally be synthesized from carbon-containing gases withdrawn from the atmosphere and / or captured from industrial processes; or various sustainable feedstocks such as, for example, waste oils and fats; municipal solid waste; cellulosic waste (such as corn stover); cover crops such as camelina, carinata, and pennycress; alternative fuels from non-biological sources; jatropha; halophytes, and algae, rather than from fossil-based hydrocarbons. SAF is understood not to include fossil fuels.

[0323] The functional performance of a given fuel composition or blend of fuels F for a given task can be defined at least in part by the capabilities of the Brayton cycle of the fuel-serviced gas turbine engine 10. Parameters defining the functional performance can include, for example, specific energy; energy density; thermal stability; and emissions including gaseous and / or particulate matter. In this regard, particulate matter emissions can include soot particles produced by the combustion of the fuel F, also referred to as non-volatile particulate matter (nvPM). Any reference herein to soot or smoke can equally apply to other types of particulate matter emissions known in the art. Gaseous emissions can include any one or more of the following: 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 (VOC) produced by the combustion of the fuel F. Any reference herein to gaseous emissions can equally apply to other types of gaseous emissions known in the art.

[0324] A relatively high specific energy (i.e., energy per unit mass) expressed in MJ / kg can at least in part reduce the takeoff weight and thus potentially provide a relative improvement in fuel efficiency. A relatively high energy density (i.e., energy per unit volume) expressed in MJ / L can at least in part reduce the takeoff fuel volume, which can be particularly important for volume-constrained missions or military operations involving fuel replenishment. A relatively high thermal stability (i.e., inhibiting fuel degradation or coking under thermal stress) can allow the fuel to maintain elevated temperatures in the engine and fuel injectors and thus potentially provide a relative improvement in combustion efficiency. Reduced emissions (including particulate matter) can allow reduced contrail formation while reducing the environmental impact of a given mission. Other properties of the fuel can also be critical to functional performance. For example, a relatively low freezing point (°C) can allow for long-duration missions to optimize the flight profile; a minimum aromatic concentration (%) can ensure sufficient swelling of certain materials used to construct O-rings and seals that have previously been exposed to fuels with high aromatic content; and a maximum surface tension (mN / m) can ensure sufficient spray breakup and atomization of the fuel.

[0325] 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 composition or fuel blend. Fuels with a higher ratio of hydrogen atoms to carbon atoms can have a higher specific energy in the absence of bond strain. For example, fossil-based hydrocarbon fuels can include molecules having from about 7 to 18 carbons, where a significant portion of a given composition is derived from molecules having 9 to 15 carbons and having an average of 12 carbons.

[0326] A variety of sustainable aviation fuel blends have been approved for use. For example, some approved blends contain a blend ratio of up to 10% sustainable aviation fuel, while other approved blends contain a blend ratio of 10% to 50% sustainable aviation fuel (the remainder containing one or more fossil-based hydrocarbon fuels such as kerosene), where additional compositions are pending approval. However, sustainable aviation fuel blends that are expected to include up to (and including) 100% sustainable aviation fuel (SAF) in the aviation industry will ultimately be approved for use.

[0327] Sustainable aviation fuel can include one or more of n-alkanes, isoalkanes, cycloalkanes, and aromatics, and can be produced, for example, from one or more of: syngas; lipids (e.g., fats, oils, and greases); sugars; and alcohols. Thus, sustainable aviation fuel can include one or both of lower aromatic and sulfur content (relative to fossil-based hydrocarbon fuels). Additionally or alternatively, sustainable aviation fuel can include one or both of higher isoalkane and cycloalkane content (relative to fossil-based hydrocarbon fuels). In some examples, sustainable aviation fuel can include one or both of a density between 90% and 98% of the density of kerosene and a heat value between 101% and 105% of the heat value of kerosene.

[0328] In some examples, the aromatic and / or other non-alkane content of the sustainable aviation fuel or blend provided to the combustion equipment 16 may be relatively lower than that of kerosene. The sustainable aviation fuel may include an aromatic content, such as 30%, 20%, 15%, 10%, 8%, 5% or less than 5%; such as 4%, 3%, 2%, 1% or less than 1%; such as 0.75%, 0.5%, 0.25% or less than 0.25%; such as 0.2%, 0.1% or less than 0.1%; such as 0.01%, 0.001% or 0%. The aromatic content of the sustainable aviation fuel may be within an inclusive number or range defined by or within any two of the values in the previous sentence (i.e., the values may form an upper or lower limit), such as 13.5%, 8.5%, 2.5%, 0.35%, 0.15%, 0.05%, 0.005% or 0%; 0% to 0.75%, 0% to 0.5%, or 0.1% to 0.25%; 0.15% to 0.65%, 0.35% to 0.55% or 0.035% to 0.055%; determined based on one or more of preferences, fuel feedstocks or suppliers, and compositional variations therein.

[0329] At least in part due to the molecular structure of the sustainable aviation fuel, the sustainable aviation fuel may provide benefits, including, for example, one or more of the following: higher specific energy (although, in some examples, lower energy density); higher specific heat capacity; higher thermal stability; higher lubricity; lower viscosity; lower surface tension; lower freezing point; lower soot emissions; lower NOx; and lower CO2 emissions, relative to fossil-based hydrocarbon fuels (e.g., when burned in the combustion equipment 16). Thus, relative to fossil-based hydrocarbon fuels (such as kerosene), the sustainable aviation fuel may result in either or both a relative reduction in fuel consumption and a relative reduction in maintenance costs.

[0330] As Figure 5 As shown, the aircraft 1 may include a plurality of fuel tanks 50, 53a, 53b; for example, a larger primary fuel tank 50 located in the aircraft fuselage and smaller fuel tanks 53a, 53b located in each wing. In other examples, the aircraft 1 may have only a single fuel tank 50, and / or the wing fuel tanks 53a, 53b may be larger than the central fuel tank 50, or the central fuel tank 50 may not be provided (instead, all fuel is stored in the wings of the aircraft) - it should be understood that many different tank layouts are conceivable, and the illustrated examples are provided for ease of description and are not intended to be limiting.

[0331] Figure 5An aircraft 1 having a propulsion system 2 including two gas turbine engines 10 is shown. Fuel is supplied from a fuel supply system on the aircraft 1 to the gas turbine engines 10. The fuel supply system illustrated in the figure includes a single fuel source. For the purposes of the present application, the term "fuel source" means: 1) a single fuel tank; or 2) a plurality of fluidly interconnected fuel tanks. Each fuel source is arranged to provide a separate fuel source (i.e., the first fuel source may contain a first fuel having one or more characteristics different from the second fuel contained in the second fuel source). Thus, the first fuel source and the second fuel source are not fluidly coupled to each other in order to separate different fuels (at least under normal operating conditions). The use of multiple fuel sources may allow the aircraft 1 to carry a variety of different fuels and, during operation, optionally even switch between different operating phases during cruise or in flight, change the fuel being used.

[0332] In this example, the first (and only in these examples) fuel source includes a central fuel tank 50 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 left wing and at least one wing fuel tank located in the right wing for balance. In the example shown, all tanks 50, 53a, 53b are fluidly interconnected, thereby forming a single fuel source. Each of the central fuel tank 50 and the wing fuel tanks 53a, 53b may include a plurality of fluidly interconnected fuel tanks.

[0333] In another example, the wing fuel tanks 53a, 53b may not be fluidly connected to the central tank 50, thereby forming a separate second fuel source. For balance purposes, one or more fuel tanks in the left wing may be fluidly connected to one or more fuel tanks in the right wing. This may be accomplished via the central fuel tank (if the fuel tank is not part of another fuel source) or bypassing the central fuel tank or both (for maximum flexibility and safety). In another example, the first fuel source includes the wing fuel tanks 53 and the central fuel tank 50, while the second fuel source includes another separate central fuel tank. Fluid interconnectivity may be provided between the wing fuel tanks and the central fuel tank of the first fuel source for balancing the aircraft 1. In an aircraft 1 having multiple fuel sources, two or more of the fuel sources may thus contain fuels different from each other such that the aircraft 1 may change fuel during flight. Therefore, determining the fuel supplied to the burner 16 may be more complex than simply recording a single identifier of the fuel on the aircraft 1 or checking it once at startup.

[0334] In some examples, the distribution of the fuel tanks 50, 53 available on the aircraft 1 may be restricted such that the first fuel source and the second fuel source are each substantially symmetric with respect to the aircraft centerline. In cases where asymmetric fuel tank distribution is allowed, suitable fuel transfer means are typically provided between the fuel tanks of the first fuel source and / or between the fuel tanks of the second fuel source such that the position of the aircraft center of mass can be kept within acceptable lateral limits throughout the flight.

[0335] Aircraft typically refuel at a number of different airports, such as at the start and end of a long-haul flight. While there are standards that all aviation fuels must meet, as noted above, different aviation fuels have different compositions, depending, for example, on their source (e.g., different petroleum sources, biofuels or other synthetic aviation fuels (commonly described as sustainable aviation fuel - SAF) and / or mixtures of petroleum-based fuels, and other fuels) and any additives included (e.g., such as antioxidants and metal deactivators, biocides, static reducers, icing inhibitors, corrosion inhibitors) as well as any impurities. In addition to differences between airports and fuel suppliers, even for a given airport or fuel supplier, the fuel composition of the available aviation fuel may vary between batches. Furthermore, the fuel tanks 50, 53 of the aircraft 1 are not typically emptied before being filled for a subsequent flight, resulting in a mixture of different fuels within the tanks - effectively a fuel with a different composition resulting from the mixture. One or more fuel ports 62 may be provided for refueling.

[0336] Figure 6 A schematic view of a portion of a fuel supply system 152 for a gas turbine engine 10 is shown. The fuel supply system 152 includes a main fuel flow path 110 from the fuel tank 50 to the burner 16 of the gas turbine engine 10, and various branch fuel paths that obtain fuel from or return fuel to the main fuel flow path.

[0337] The main fuel flow path 110 of the fuel supply system 152 serves as a fuel source for a fuel hydraulic system. The gas turbine engine 10 includes a burner 16 and an actuator 154. Figure 6 The dashed arrows in indicate the direction of fuel flow. One or more fuel lines 159, 160 are provided to supply fuel to the actuator 154 to actuate the actuator in a fuel hydraulic manner.

[0338] The fuel supply system 152 is arranged to supply fuel for combustion in the burner 16.

[0339] It should be understood that, in addition to other engine components, one or more valves, pumps, sensors, discharge devices for fuel hydraulic actuators, and heat exchangers may be present along the main fuel flow path 110. After the fuel passes through any preceding engine component most downstream along the main fuel flow path 110, it is directly supplied to the burner 16 via the line 110a.

[0340] The fuel traveling via the lines 159 and 160 is used to drive (i.e., actuate) the actuator 154. Thus, the actuator 154 is hydraulically actuated by fuel. The term "fuel hydraulic" as used herein refers to the hydraulic operation of an actuator that uses a hydraulic fluid as fuel. An actuator hydraulically actuated by fuel may be referred to herein as a "fuel hydraulic actuator".

[0341] In Figure 6 the example shown, the fuel travels from the fuel flow path 110 via the line 159 to the actuator 154. After being used to hydraulically actuate the actuator 154, the fuel returns to the main fuel flow path 110 via the line 160 and may then be delivered to the burner 16.

[0342] The fuel return line 160 may include a valve that is configured to regulate the fuel flow returning to the main fuel path 110, and the valve itself may be operated using a fuel hydraulic actuator. The fuel return line 160 may direct the fuel to a location along the main fuel flow path 110, either before or after one or more heat exchangers in the main fuel flow path 110.

[0343] The fuel return line 160 may return the fuel to the fuel tank. In such embodiments, the valve may be a fuel return tank valve. The fuel return tank valve may be controlled using a fuel return tank actuator. The fuel supply system may be arranged to supply fuel to hydraulically actuate the fuel return tank actuator.

[0344] Once the fuel reaches the burner 16, it burns to provide thrust, as explained with respect to Figure 1 as explained.

[0345] Figure 6 A single fuel hydraulic actuator 154 on the fuel flow circuits 159, 160 is shown. In an alternative embodiment, multiple fuel hydraulic actuators 154 may be located on a single fuel flow circuit 159, 160, using fuel taken from a single point along the main fuel flow path 110. The multiple actuators may be arranged in series with respect to the fuel flow (such that all the fuel in the discharge pipe 159 passes through all the actuators in sequence) or in parallel with respect to the fuel flow (with the discharge pipe 159 branching and having one or more actuators on each branch).

[0346] Figure 7FIG. shows an alternative portion of the fuel supply system 152 of the gas turbine engine 10. The fuel supply system 152 includes fuel sources 50, 53 (which may be a single fuel tank or multiple fuel tanks) and a fuel flow path 110 between the fuel sources 50, 53 and the burner 16. The illustrated fuel supply system 152 includes two actuators 254a, 254b, each actuator being located on a fuel drain circuit separate from the main fuel flow path 110. Figure 7 The dashed arrows in FIG. indicate the direction of fuel flow.

[0347] The two actuators 254a, 254b are hydraulically actuated by fuel. That is, both actuators 254a, 254b are hydraulically actuated using fuel from the fuel sources 50, 53 as the hydraulic fluid. Fuel travels from the main fuel flow path 110 to the actuators 254a, 254b via lines 259a, 259b respectively. After the fuel is used to drive the actuators 254a, 254b, it returns to the main fuel flow path 110 via lines 260a, 260b.

[0348] Figure 7 Only two actuators 254a, 254b are shown in FIG., but in other embodiments, any number of actuators may be hydraulically actuated by fuel. In some embodiments, the gas turbine engine 10 may include multiple fuel hydraulically actuated actuators and multiple non-fuel hydraulically actuated actuators.

[0349] Figure 7 FIG. shows a single fuel hydraulically actuated actuator 254 on each fuel flow circuit. In an alternative embodiment, multiple fuel hydraulically actuated 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 (such that all the fuel in the drain pipe 159 passes through all the actuators in sequence) or in parallel with respect to the fuel flow (with the drain pipe 159 branching and having one or more actuators on each branch).

[0350] In various embodiments, the gas turbine engine 10 includes at least ten, eleven, twelve, thirteen, fourteen, fifteen, sixteen, seventeen, or eighteen fuel hydraulically actuated actuators. In some embodiments, each actuator may have its own dedicated fuel drain pipe 159, 259 on the main fuel flow path 110 such that it has 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 flow path 110. In some arrangements, most or all of the fuel hydraulically actuated actuators may be located on the same fuel return circuit away from the main fuel flow path 110.

[0351] In various embodiments, the gas turbine engine 10 includes one or more actuated engine systems. Each actuated engine system may include one or more individual actuators. An "actuated engine system" (also referred to as an "actuated system") refers to a set of actuators that are used in combination to control a particular system within the engine 10 or perform a particular function. If at least one actuator in the actuated system is fuel hydraulic, the system may be described as a "fuel hydraulic" system.

[0352] Figure 8 A fuel supply system 152 including an actuator 354 is shown.

[0353] Figure 8 A schematic illustration of an alternative portion of the fuel supply system 152 of the gas turbine engine 10 is shown. The fuel supply system 152 includes fuel sources 50, 53 (which may be a single fuel tank or multiple fuel tanks) and a fuel flow path 110 between the fuel sources 50, 53 and the burner 16. The illustrated fuel supply system 152 includes an actuator 354 in the fuel discharge return line of the main fuel flow path 110. Figure 8 The dashed arrows in [figure] indicate the direction of fuel flow.

[0354] Figure 8 The fuel supply system 152 of [figure] includes a valve 304. The valve 304 is configured to direct fuel via line 361 or line 360. Line 306 includes the actuator 354, so the fuel directed via line 306 drives the actuator 354 in a fuel hydraulic manner and then returns to the main fuel flow path 110. The fuel directed via line 361 returns directly to the main fuel flow path 110 without driving the actuator. The valve 304 itself may be controlled using additional fuel hydraulic actuators.

[0355] The valve 304 is controllable, which means the fuel supply system 152 is arranged to be controlled to select between:

[0356] · causing fuel to actuate the actuator 354 (fuel flows via line 360); and

[0357] · causing fuel to bypass at least one actuator 354 (fuel flows via line 361).

[0358] In some embodiments, the valve 304 is controlled via an actuator in accordance with an output from a processor or other controller. The valve 304 may be controlled based on the sustainable aviation fuel - SAF - content of the fuel. In some embodiments, the valve 304 may be replaced with any alternative suitable for regulating the fuel flow between the two lines 360, 361.

[0359] The fuel supply system 152 may be arranged to select between:

[0360] · Actuate the fuel actuator 354 (via line 360); and

[0361] · Bypass at least one actuator 354 with the fuel (via line 361).

[0362] This selection can be made based on whether the sustainable aviation fuel - SAF - content of the fuel exceeds a threshold value.

[0363] For example, the fuel supply system 152 can be arranged to actuate the fuel actuator 354 (via line 360) when the SAF content of the fuel exceeds a concentration of 25%, 30%, 35%, 40%, 45%, 50%, 55%, 60%, 65%, 70%, 75%, 80%, 85%, 90% or 95% by volume.

[0364] The fuel supply system can control the fuel in a binary manner such that all the fuel reaching the valve 304 is directed into one of line 360 and line 361. In other embodiments, the fuel can be split, for example, between line 360 and line 361.

[0365] To determine the SAF content of the fuel, the fuel supply system 152 can include a sensor. The sensor can be positioned immediately upstream of the valve 304. In other embodiments, the sensor can be positioned upstream of the valve 304 within the main fuel flow path 110. In an alternative arrangement, when the fuel is input into the fuel sources 50, 53, the SAF content of the fuel is measured or calculated or otherwise provided. For example, information regarding the SAF content of one or more fuels stored on the aircraft 1 can be transmitted electronically to an on - board controller or other computing system, or manually input into a user interface of the on - board controller or other computing system. It should be understood that any suitable method known in the art can be used to determine the SAF content of the fuel.

[0366] Figure 9 A schematic illustration of an alternative portion of the fuel supply system 152 of the gas turbine engine 10 is shown. The fuel supply system 152 is similar to Figure 8 the fuel supply system 152, but the line 361 that returns fuel to the main fuel flow path 110 is replaced by a line 362 that directs the fuel to the burner 16. The valve 304 directs the fuel via one of lines 360 and 362. Line 362 directs the fuel such that it bypasses the actuator 354 and travels to the burner 16 without returning to the main fuel flow path 110. Line 362 can include one or more pumps or valves configured to regulate the pressure and flow rate of the fuel reaching the burner 16.

[0367] In Figure 8 and Figure 9In an embodiment, the actuator 354 can be configured such that when the valve 304 directs fuel to bypass the actuator 354, the actuator 354 is driven in a non-fuel hydraulic manner. For example, a non-fuel hydraulic fluid or an electrical signal can be used to drive the actuator 354.

[0368] Figure 10 FIG. shows a schematic view of an alternative portion of the fuel supply system 152 of the gas turbine engine 10. The fuel supply system 152 includes fuel sources 50, 53 (which can be a single fuel tank or multiple fuel tanks) and a fuel flow path 110 between the fuel sources 50, 53 and the burner 16. The illustrated fuel supply system 152 includes two actuators 454a, 454b, each actuator 454a, 454b being positioned on a respective fuel discharge loop of the main fuel flow path 110. Figure 10 The dashed arrows in FIG. indicate the direction of fuel flow.

[0369] The first actuator 454a is positioned downstream of a valve 404 located on its respective fuel discharge loop. The valve 404 is configured to direct fuel to bypass the actuator (directly back to the main fuel flow path 110) via line 461 or to hydraulically actuate the actuator 454a with fuel (and then return to the main fuel flow path 110) via line 460a. The valve 404 is arranged to be similar to Figure 8 valve 304, but in other embodiments, the valve 404 can be arranged to be similar to Figure 9 valve 304 such that fuel bypassing the actuator is instead directed directly to the burner 16. In still other embodiments, all fuel taken from the main fuel flow path 110 and reaching the valves 304, 404 can be directed directly to the burner 16 or after passing through the actuators 354, 454a. Thus, the actuator 454a is selectively driven in a fuel hydraulic manner.

[0370] In a further embodiment, the valves 354, 454a selectively driven in a fuel hydraulic manner may not have a plurality of fuel paths associated therewith. Instead, the valves 304, 404 may be arranged to prevent any fuel from leaving the main fuel flow path and flowing towards the actuators when the actuators 354, 454a are not driven in a fuel hydraulic manner. Thus, the flow through the simple circuit can be started or stopped by the valves 304, 404, rather than using the valves to split the fuel flow between two different branches of the main path. Similarly, for embodiments that supply fuel to a plurality of fuel hydraulic actuators using the same fuel discharge point of the main fuel flow path 110, the controllable valves arranged to allow or prevent fuel from reaching a given actuator may (i) divert the fuel intended for the actuator along a dedicated bypass route when the actuator is not actuated in a fuel hydraulic manner, or (ii) simply block any route towards the actuator, so that the fuel that would otherwise flow to the actuator flows towards one or more other actuators supplied by the same discharge device. Therefore, choosing not to actuated an actuator in a fuel hydraulic manner can increase the fuel flow rate to one or more other actuators, unless the total flow through the discharge portion of the fuel flow line is also controlled.

[0371] Figure 10 A second actuator 454b of the example is located on a fuel flow return path separate from the first actuator 454a, and the fuel is taken from a different point in the main fuel flow path 110. The second actuator 454b is arranged to always be driven in a fuel hydraulic manner (it only has a single actuation mode).

[0372] Although Figure 10 the embodiment shows one actuator 454a selectively driven and one actuator 454b always driven in a fuel hydraulic manner, in other embodiments, the fuel supply system 152 may include any number of actuators, and any one 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 loop.

[0373] The actuators described above may be any actuators used by the engine 10. Various specific actuators that can be driven in a fuel hydraulic manner are described below. The actuation systems described herein may be any actuation systems used by an aircraft. Various specific actuation systems that can be driven in a fuel hydraulic manner are described below. It should be understood that the specific embodiments listed are only described as examples, and the scope of protection is only defined by the claims.

[0374] An example of an actuated system is a turbine case cooling (TCC) system. A turbine engine 10 typically includes a case surrounding turbines 17, 19, and the TCC system is used to selectively cool the case. This cooling can be used to control the rotor-stator tip clearance by reducing the amount of thermal expansion. The TCC system includes one or more valves that can be used to regulate the bleed air flow through the case from the engine's compressors 17, 19 to provide cooling.

[0375] In an engine 10 having multiple turbines 17, 19, a TCC system can be provided for one, some, or all of the turbines.

[0376] One or more valves of the TCC system can be operated via an actuator. Each valve can have a dedicated actuator, or one actuator can control multiple valves. In various embodiments, one or more of the actuators of the TCC system are fuel hydraulically actuated. In one or more embodiments, the fuel supply system can be arranged to selectively bypass or fuel hydraulically actuate one or more actuators of the TCC system.

[0377] Figure 11 A schematic diagram of a portion of a fuel supply system 152 of a gas turbine engine 10 is shown. The fuel supply system 152 includes a TCC system 554 that includes two actuators 554a, 554b, each actuator 554a, 554b being configured to control a corresponding valve 554a', 554b' of the TCC system. Valves 255a', 255b' can supply cold air to cool the turbine case.

[0378] Figure 11 The fuel supply system 152 is arranged to supply fuel from fuel sources 50, 53 to burners 16 and fuel hydraulically actuate actuators 554a, 554b. As described above, the fuel supply system 152 is arranged to selectively actuate or bypass actuators 554a, 554b via associated valves 504a, 504b. In an alternative embodiment, a single valve can be used to selectively actuate or bypass actuators 554a, 554b.

[0379] In Figure 11 the illustrated embodiment, actuators 554a, 554b are arranged in different fuel flow circuits. In other embodiments, actuators 554a, 554b can be arranged on the same fuel flow circuit and can be arranged in series or parallel on the circuit.

[0380] Another example of an actuated system used within the turbine engine 10 is the bleed air system. The bleed air system redirects bleed air (e.g., compressed air taken from compressor 14, 15 or elsewhere such as from the APU) for use in other systems. Bleed air is useful in various systems due to its relatively high temperature and pressure. The bleed air system may be referred to as an "engine bleed air system" (EBAS).

[0381] The bleed air system can be used to control the air conditioning within the cabin of the aircraft 1. In such systems, a conditioning assembly (either a conditioning assembly or a pressurized air conditioning kit (PACK) are acceptable terms) filters and cools the bleed air (via expansion and / or heat exchange with ambient air for example). The air conditioning assembly is used to cool the bleed air from the engine or APU and, where appropriate, dehumidify it, and then supply it to the aircraft cabin. Such systems include one or more valves and one or more actuators. The bleed air is typically supplied to the assembly via a non-return check valve. In various embodiments, one or more of these actuators can be selective fuel hydraulic actuators.

[0382] Figure 12 A schematic illustration of an alternative portion of the fuel supply system 152 of the gas turbine engine 10 is shown, which includes an air conditioning assembly 650.

[0383] Figure 12 The fuel supply system 152 includes a cabin bleed valve 654' and a cabin bleed actuator 654 configured to control the cabin bleed valve 654'. The cabin bleed valve 654' controls the flow of bleed air from compressors 14, 15 through the air conditioning assembly (as described above) to supply cooling air to the cabin of the aircraft 1.

[0384] Figure 12 The fuel supply system 152 also includes a manual bleed valve 754' and a manual bleed actuator 754 configured to control the manual bleed valve 754'.

[0385] The manual bleed valve 754' controls the discharge of bleed air from compressors 14, 15 and into other aircraft systems. The manual bleed valve 754' can be used for engine start, transient bleed, and / or surge control.

[0386] In some arrangements, there are multiple cabin bleed valves 654' and associated cabin bleed actuators 654, all of which can be considered part of the air conditioning assembly 650. In some arrangements, there are multiple manual bleed valves 754' and associated manual bleed actuators 754, all of which can be considered part of the air conditioning assembly 650.

[0387] In an alternative arrangement, the nacelle drain valve 654' and the nacelle drain actuator 654 are omitted. In an alternative arrangement, the controlled drain valve 754' and the controlled drain actuator 754 are omitted. In an alternative arrangement, alternative known actuators and associated valves may be included and selectively actuated in a fuel hydraulic manner. In some arrangements, other known drain valves are selectively actuated in a fuel hydraulic manner.

[0388] Similar to the above actuator, the nacelle drain actuator 654 is selectively actuated in a fuel hydraulic manner to control the nacelle drain valve 654'. The valve 604 is arranged to direct fuel via line 660 to actuate the nacelle drain actuator 654, or to bypass the nacelle drain actuator 654 by directing fuel via line 661.

[0389] Similar to the above actuator, the controlled drain actuator 754 is selectively actuated in a fuel hydraulic manner to control the controlled drain valve 754'. The valve 704 is arranged to direct fuel via line 760 to actuate the controlled drain actuator 754, or to bypass the controlled drain actuator 754 by directing fuel via line 761.

[0390] Another example of an actuated system is the variable guide vane (VGV) system 850. The VGV system controls the amount and / or flow path of the airflow through the compressors 14, 15 (by moving one or more compressor vanes; and typically by moving multi-stage compressor vanes) to provide optimal compressor performance. The VGV system 850 changes the angle of the compressor vanes to manage the operability and efficiency of the compressor.

[0391] The variable guide vanes of various embodiments may include either or both of variable inlet guide vanes (VIGV) located upstream / ahead of the compressor rotor blades and other vanes (typically variable stator vanes (VSV)) located downstream / behind the compressor rotor blades. Other compressor vanes may be located elsewhere, and the vane positions may vary depending on the engine architecture and also depending on the different compressors of a given engine. For example, variable / controllable intermediate guide vanes may be located midway through the compressors 14, 15.

[0392] Generally, the variable guide vanes are arranged to be rotatable about an axis to align with the angle of the incoming airflow. This movement of the vanes allows a wider range of dimensionless flow through the compressors 14, 15 without stalling; thanks to the VGV, the compressors 14, 15 can operate effectively over a wider range of engine power settings without surging. The VGV is controlled by the engine control system in response to a series of parameters, which typically include environmental conditions and the thrust level required by the engine.

[0393] In an engine 10 having multiple compressors 14, 15, a VGV system 850 may be provided for one, some, or all of the compressors 14, 15.

[0394] Many VGV systems 850 include two VGV actuators that may be used to vary the angle of the inlet guide vanes of the compressors 14, 15. Each of the two VGV actuators, the VSV actuator, may be configured to control the guide vanes on a given side of the engine 10. In various embodiments, one or more of the actuators of the VGV system are hydraulically actuated by fuel.

[0395] Figure 13 A schematic illustration of an alternative portion of the fuel supply system 152 of a gas turbine engine 10 is shown, which 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 that is configured to selectively hydraulically actuate the variable stator vane actuator 854 by directing fuel via line 860 or line 861. The valve 804 may be configured to actuate the fuel or bypass the variable stator vane actuator 854 based on the SAF content of the fuel. Although Figure 13 the arrangement shown depicts a single variable stator vane actuator, in other arrangements, the fuel supply system 152 may be arranged to selectively hydraulically actuate two or more variable stator vane actuators.

[0396] Figure 14 A schematic illustration is shown depicting an alternative portion of the fuel supply system 152 of a gas turbine engine 10, which includes a VGV system 850, similar to Figure 13 the VGV system of, but having only a single simple fuel drain circuit and 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 that is configured to selectively hydraulically actuate the variable inlet guide vane actuator 854a by directing fuel via line 960 to the actuator or preventing fuel from flowing along line 960. The valve 904 may be configured to actuate the fuel to actuate the variable inlet guide vane actuator 854a based on the SAF content of the fuel, or continue to flow along the main fuel flow path 110, thereby bypassing the actuator 854a.

[0397] The guide vanes control the airflow entering and passing through the compressors 14, 15. The variable guide vane actuators 854, 854a control the orientation angle of the guide vanes.

[0398] Another example of an actuation system used within a turbine engine 10 is a thermal management system (HMS). The turbine engine 10 typically includes an engine HMS and a generator HMS. Both HMS systems control the temperature of the oil by enabling and controlling the use (or bypass) of an air - oil (and / or fuel - oil) heat exchanger. An oil - oil heat exchanger may also be used in some embodiments, which will be described in more detail with respect to Figure 25 For example, heat transfer between the oil in the generator HMS and the oil in the engine HMS can be achieved using an oil - oil heat exchanger.

[0399] The engine HMS system controls the temperature of the oil system that lubricates engine components (e.g., bearing housings, gearboxes, etc.). The engine HMS system includes one or more engine HMS valves controlled by one or more actuators. In various embodiments, one or more of these actuators are fuel - hydraulically driven.

[0400] For example, in some embodiments, the engine HMS includes a single valve and a single actuator that is fuel - hydraulically driven and arranged to control the valve.

[0401] The engine HMS system may include multiple heat exchangers arranged in a parallel configuration or a series configuration with respect to a fluid flow (e.g., a fuel flow or an oil flow). A single valve controlled by a single fuel - hydraulic actuator enables the use (or bypass) of all heat exchangers. In particular, the flow rate of the fluid (air, oil, or fuel, if applicable) can be controlled by the valve, and the portion of the fluid that does not pass through the heat exchanger is directed along a bypass pipe. In some embodiments, each heat exchanger has an associated valve controlled by an associated actuator. In various embodiments, one or more of these actuators are fuel - hydraulically driven. Additionally, one or more recirculation pipes may be provided to allow the fluid to return from the heat exchanger outlet to the heat exchanger inlet, thereby passing through the heat exchanger multiple times. The flow rate through the recirculation pipe can again be regulated using a valve controlled by an actuator. For any of the recirculation pipes described herein, one or more associated pumps may be present, which are configured to transport air / oil / fuel back to the heat exchanger. Alternatively or additionally, any suitable component for repressurizing the air / oil / fuel to achieve recirculation may be used.

[0402] Figure 15 A schematic diagram representing an engine thermal management system 1000 is shown. The engine thermal management system 1000 includes an air - oil heat exchanger 1003 that is configured to transfer heat from the oil within the engine. The air - oil heat exchanger 1003 includes an oil line 1001 and an air line 1002. In Figure 15In [the system], the heat exchanger 1003 is a parallel flow heat exchanger, and the oil flow and air flow are in the same direction. In other embodiments, the heat exchanger 1003 can be a counterflow heat exchanger, which means that the flow direction of the air pipeline or the oil pipeline is reversed, such that they flow through the heat exchanger 1003 in opposite directions.

[0403] In Figure 15 the arrangement, the oil pipeline 1001 includes a valve 1004, which can be operated to enable the oil to bypass the air-oil heat exchanger 1003. The valve 1004 is controlled by an actuator 1054. The valve 1004 can be binary (open / closed), or can be arranged to allow finer control of the oil flow rate through the heat exchanger 1003, where a variable portion of the oil bypasses the heat exchanger 1003.

[0404] The actuator 1054 can be a selective fuel hydraulic actuator, as described for the above arrangement. The valve 1006 can be operated such that fuel can flow through the actuator 1054 to actuate the actuator 1054 in a fuel hydraulic manner, or such that the fuel can bypass the actuator 1054. In some arrangements, when the valve 1006 causes the fuel to bypass the actuator 1054, the actuator 1054 can be actuated in a non-fuel hydraulic manner.

[0405] The valve 1006 can be operated to enable fuel hydraulic actuation of the actuator 1054 only when the SAF content of the fuel is higher than a predetermined threshold.

[0406] Figure 16 An alternative arrangement of the thermal management system 1000 is shown, where the valve 1004 is positioned within the air pipeline 1002. The engine thermal management system 1000 is similar to Figure 15 the thermal management system 1000, but the valve 1004 can be operated such that air can bypass the air-oil heat exchanger 1003, thereby controlling the air flow rate through it. Similarly, in other embodiments, the fluid flows can be arranged differently.

[0407] In other embodiments, the Figure 15 and Figure 16 engine HMS systems 1000 can be combined such that the HMS system includes valves for controlling both the oil flow and the air flow. In some embodiments, the HMS system 1000 can include multiple heat exchangers and valves arranged to be able to bypass all the heat exchangers simultaneously, as well as valves for bypassing individual heat exchangers.

[0408] The generator HMS can be independent of the engine HMS1000. It provides cooling for an electric machine, which is used, for example, to supply power to the fuselage to operate aircraft systems.

[0409] Figure 17A schematic diagram representing a generator thermal management system 1100 is shown. The generator thermal management system 1100 includes an air-oil heat exchanger 1103 that is configured to transfer heat from the oil within the engine and more specifically from the oil used to cool (and optionally also lubricate the generator). The air-oil heat exchanger 1103 includes an oil line 1101 and an air line 1102.

[0410] In Figure 17 the arrangement, the oil line 1101 includes a valve 1104 that can be operated to enable a controlled amount of oil to bypass the air-oil heat exchanger 1103. The valve 1104 is controlled by an actuator 1154.

[0411] The actuator 1154 can be a selective fuel hydraulic actuator as described with respect to the above arrangement. A valve 1106 can be operated so that fuel can flow through the actuator 1154 to hydraulically actuate the actuator by fuel or so that the fuel can bypass the actuator. In some arrangements, the actuator 1154 can be configured to be actuated in a non-fuel hydraulic manner when the valve 1106 causes the fuel to bypass the actuator 1154.

[0412] The valve 1106 can be operated to enable fuel hydraulic actuation of the actuator 1154 only when the SAF content of the fuel is higher than a predetermined threshold.

[0413] Figure 18 An alternative arrangement of the generator thermal management system 1100 is shown, where the valve 1104 is positioned within the air line 1102. The generator thermal management system 1100 is similar to Figure 17 the generator thermal management system 1100 of

[0414] Figure 19 A schematic diagram representing a fuel supply system 152 is shown, which includes two fuel-oil heat exchangers 1220, 1230. The fuel supply system 152 includes a primary fuel-oil heat exchanger 1220 and a secondary fuel-oil heat exchanger 1230. The heat exchangers 1220, 1230 are used to transfer heat from the oil to the fuel and can be regarded as part of the engine thermal management system. The primary fuel-oil heat exchanger 1220 is arranged to heat at least most of the fuel, and the secondary fuel-oil heat exchanger 1230 is arranged to provide additional heat to the fuel to be supplied for fuel hydraulically driving at least one fuel hydraulic actuator 1254.

[0415] Figure 19The fuel supply system 152 of the illustrated embodiment includes five fuel valves 1201-1205. Each valve 1201-1205 is controlled by an actuator, which can be a fuel hydraulic actuator.

[0416] The first valve 1201 is a fuel valve arranged such that fuel can bypass the primary heat exchanger 1220. The valve 1201 is operable to allow any percentage of the fuel to bypass or pass through the primary heat exchanger 1220. The split fuel then returns to the main fuel flow path 110.

[0417] The primary heat exchanger 1220 allows the fuel temperature to increase before the fuel enters the burner 16 while reducing the temperature of the oil.

[0418] The second valve 1202 is arranged to direct fuel from the outlet of the primary heat exchanger 1220: (i) through a recirculation pipe 1206, which is configured to direct the fuel back to the primary heat exchanger 1220, (ii) to the burner 16, or (iii) towards the secondary heat exchanger 1230 for one or more fuel hydraulic systems. In other embodiments, multiple individual valves may be provided for recirculation or redirection / discharge to the fuel hydraulic system.

[0419] The third valve 1203 forms part of the fuel hydraulic system and is a fuel valve arranged such that a controllable portion of the fuel can bypass the secondary heat exchanger 1230. The valve 1203 is operable to allow any percentage of the fuel to bypass or pass through the secondary heat exchanger 1230.

[0420] The fourth valve 1204 forms part of the fuel hydraulic system and is arranged to direct fuel downstream of the secondary heat exchanger 1230 (i) to the fuel hydraulic actuator 1254, or (ii) through a recirculation pipe 1207, which is configured to direct the fuel back to the secondary heat exchanger 1230 such that some or all of the fuel directed to the secondary heat exchanger 1230 can pass through the heat exchanger multiple times. In some embodiments, the fuel recirculation valve 1204 may be provided separately from the valve arranged to control the flow to the fuel hydraulic actuator 1254.

[0421] In some arrangements, one or both of the recirculation pipes 1206, 1207 include one or more pumps or are associated with the one or more pumps, which are configured to convey fuel through the recirculation pipes 1206, 1207. In some arrangements, the pumps are located upstream of the valves 1202, 1204. In other arrangements, the pumps are located anywhere along the length of the recirculation pipes 1206, 1207.

[0422] The fifth valve 1205 forms part of the fuel hydraulic system and is arranged to direct fuel to drive the actuator 1254 by fuel hydraulic means or to bypass the actuator. In the example shown, the fuel passing through or bypassing the actuator 1254 is then sent directly to the burner 16. In other embodiments, it may return to the main fuel flow path 110.

[0423] 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.

[0424] In some arrangements, the heat exchangers 1220, 1230 are controlled such that, under cruise conditions, the heat transfer ratio:

[0425]

[0426] Has a maximum value (i.e., peak value) of at least 0.35.

[0427] In other arrangements, the heat transfer ratio has a maximum value greater than 0.4, 0.45, 0.5, 0.55 or 0.6 during cruise.

[0428] In some arrangements, the fuel supply system 152 is arranged to control the heat transfer ratio by controlling the percentage of fuel bypassing or recirculating through each of the heat exchangers 1220, 1230.

[0429] 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 pipes may be provided accordingly. One or more air-oil or oil-oil heat exchangers may also be controlled to affect the heat transfer ratio.

[0430] Figure 25 One such exemplary fuel supply system 153 is shown. In Figure 25 In the embodiment, each fuel-oil heat exchanger 1220, 1230 is on a separate oil circuit, the two oil circuits being fluidly isolated from each other and in thermal contact by means of an oil-oil heat exchanger 1240 (although not shown, it should be understood that one or more oil bypass and / or recirculation valves associated with the oil-oil heat exchanger 1240 may also be provided).

[0431] The oil circuit containing the oil arranged to pass through the primary fuel-oil heat exchanger 1220 is arranged to cool and optionally lubricate the first set of engine components 10'. The oil circuit containing the oil arranged to pass through the secondary fuel-oil heat exchanger 1230 is arranged to cool and optionally lubricate the second set of engine components 10".

[0432] Each fuel - oil heat exchanger 1220, 1230 has an oil bypass valve 1212, 1213 associated therewith, which is arranged to allow oil to bypass the respective heat exchanger 1220, 1230. In various embodiments, the oil bypass valves 1212, 1213 are arranged to allow a controllable portion of the oil to bypass the respective heat exchangers 1220, 1230.

[0433] Each fuel oil heat exchanger 1220, 1230 has an oil recirculation valve 1211, 1214 associated therewith, which is arranged to allow oil to pass through the respective heat exchanger 1220, 1230 multiple times before returning to engine components 10', 10" where the return oil needs to be cooled. In various embodiments, the oil recirculation valves 1211, 1214 are arranged to allow a controllable portion of the oil to recirculate through the respective heat exchangers 1220, 1230. In Figure 25 the embodiments, the recirculated oil is shown as being supplied into the respective bypass valves 1212, 1213, however, it should be understood that in other embodiments, the recirculated oil may be rejoined to the oil flow before or after the valve.

[0434] In some embodiments, only one of the oil recirculation valves 1211, 1214 and the oil bypass valves 1212, 1213 may be provided for one or each of the heat exchangers 1220, 1230, or neither may be provided for one or each of the heat exchangers 1220, 1230.

[0435] Although for simplicity, Figure 25 the air - oil heat exchanger is not shown in,

[0436] it should be understood that in various embodiments, one or more air - oil heat exchangers may be provided on one or each oil return line. Additionally, in some embodiments, a single oil return line may supply oil to two fuel - oil heat exchangers; in some such embodiments, an oil - oil heat exchanger may not be provided.

[0437] In some arrangements, the engine 10 is arranged to control the heat transfer ratio based on the SAF content of the fuel.

[0438] For any of the above fuel hydraulic actuators, the fuel supply system 152 may be arranged such that the peak fuel pressure differential across 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.

[0439] For any of the above fuel hydraulic actuators, the fuel supply system 152 may be arranged such that the peak fuel differential pressure across at least one fuel hydraulic actuator during takeoff is at least 9000 kPa. For any of the above fuel hydraulic actuators, the fuel supply system 152 may be arranged such that the peak fuel differential pressure across at least one fuel hydraulic actuator during takeoff may be in the range of 6,900 kPa to 10,000 kPa, at least 10000 kPa, 11,000 kPa, 12,000 kPa, 13,000 kPa, 14,000 kPa or 15,000 kPa.

[0440] Actuator 1254 may be any of the actuators described herein, or any actuator that can be made into a fuel hydraulic actuator, such as a variable stator vane actuator.

[0441] In some arrangements, the temperature of the fuel entering actuator 1254 (exiting valve 1205) is at least 5 °C, 7 °C, 10 °C, 12 °C, 15 °C or 20 °C higher than the fuel entering burner 16 (exiting valve 1202).

[0442] In some arrangements, the fuel is thermally stable at temperatures above 280 °C. Herein, being "thermally stable" at the said temperature means that the fuel passes the Jet Fuel Thermal Oxidation Test (JFTOT) at that temperature.

[0443] 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.

[0444] In some arrangements, the aromatic hydrocarbons account for less than 5% of the fuel volume. In some arrangements, the calorific value of the fuel is at least 43.5 MJ / kg -1 . In some arrangements, the calorific value of the fuel is at least 44 MJ / kg -1 . In some arrangements, the sulfur content of the fuel is less than 15 parts per million. In some arrangements, the fuel is HEFA fuel or includes HEFA fuel.

[0445] Various methods for operating the gas turbine engine 10 of the aircraft 1 are described below.

[0446] Figure 20A flowchart of a method 1300 for operating a gas turbine engine 10 according to various embodiments is shown. The method 1300 is performed on an engine 10 that includes: an engine core 11 that includes a turbine 19, a combustor 16, a compressor 14, and a shaft 26 that connects the turbine to the compressor; a fan 23 that is located upstream of the engine core 11 and is arranged to be driven by the shaft 26, the fan including a plurality of fan blades; and a nacelle 21 that surrounds the fan 23 and the engine core 11 and defines a bypass duct 22 that is located radially outward of the engine core 11; a plurality of actuators 254; and a fuel supply system 152.

[0447] In the described embodiments, the bypass ratio is greater than or equal to 4, the bypass ratio being defined as the ratio of the mass flow rate of the flow through the bypass duct 22 to the mass flow rate of the flow through the core 11 under cruise conditions. Such a bypass ratio can be used for any engine performing Figures 20 to 24 any method.

[0448] The method 1300 includes:

[0449] Supplying 1310 fuel using the fuel supply system 152 for combustion in the combustor 16; and

[0450] Selecting 1320 between:

[0451] Supplying 1320a fuel from the fuel supply system 152 to at least one of the plurality of actuators 354 and actuating at least one of the actuators 354 in a fuel hydraulic manner; and

[0452] Again using the fuel supply system 152 to bypass 1320b fuel around at least one of the actuators 354.

[0453] The step 1320 of selecting between actuating the actuator 354 in a fuel hydraulic manner and bypassing at least one actuator 354 may include determining, in a processor or other computing system, an action to take based on a comparison of the fuel SAF content with one or more set thresholds, and operating the 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 second fuel lines 361, 362, where the first fuel line 360 drives the actuator 354 and the second fuel lines 361, 362 bypass the actuator 354. The valve 304 may simply be arranged to open the discharge device on the main fuel flow path 110 only when fuel hydraulic actuation is required, thereby allowing all fuel to remain on the main fuel flow path 110 when fuel hydraulic actuation is not required. Thus, when fuel hydraulic actuation is not required, the fuel may bypass at least one actuator 354 through a dedicated bypass line, or simply remain on the main fuel flow path without reaching the actuator 354. In embodiments having a dedicated bypass line, the bypass line may be actuator-specific; providing a route for fuel to bypass a single actuator, or may bypass multiple actuators (e.g., some or all of the actuators located on the same fuel discharge path).

[0454] The step 1320 of selecting which action to take may include determining whether the SAF content of the fuel exceeds a threshold. The fuel supply system 152 may include one or more sensors configured to sense one or more parameters that permit the calculation or inference of the SAF concentration of the fuel (e.g., by detecting tracer elements and referring to a lookup table of fuel tracer elements and SAF content), and / or may include a repository for storing data regarding the fuel in use. The fuel supply system 152 may include a processor or other computing system configured to receive the stored, calculated, or otherwise determined SAF content value and compare that value to the threshold. The controller may then be used to operate the valve 304 based on that determination. In some embodiments, the determination may be made and the result implemented using an engine electronic controller (EEC); in other embodiments, a separate, dedicated computing system may be used instead of the EEC.

[0455] In embodiments having multiple fuel hydraulic actuators 354, different thresholds may be set for different actuators, or the same threshold may be used for all actuators.

[0456] Then fuel for driving at least one actuator 354 in a fuel hydraulic manner can be supplied to the burner 16 for combustion; this can be done directly or after returning to the main fuel flow path 110. The fuel can flow along the main fuel flow path 110 to the burner 16, so the main fuel flow path 110 can supply fuel 1310 to the burner 16. A portion of the fuel flowing along the main fuel flow path 110 can be diverted from the main fuel flow path 110 to at least one actuator 354, and then this fuel can return to the main fuel flow path 110 after being used for actuation. One or more tubes can be used to convey fuel from the main fuel flow path 110 to at least one actuator 354. These tubes can form one or more fuel flow circuits. One or more valves can be used to control the fuel flow rate through the said or each fuel hydraulic flow circuit. In some embodiments, one or more fuel hydraulic fuel pumps can be provided to actively pump fuel to or from at least one actuator 354. However, in many embodiments, the pressure provided by one or more fuel pumps on the main fuel flow path 110 may be sufficient.

[0457] The discharge device 159 on the main fuel flow path 110 can be located at a point along the main fuel flow path 110 different from the return pipe 160 of the (or each) fuel flow circuit, as Figure 6 shown. The fuel can return to the main fuel flow path 110 at a location upstream or downstream of this discharge device. One or more valves can be used to control the location along the main fuel flow path 110 at which the fuel for driving the actuator returns to the main fuel flow path 110. For example, the fuel for actuation can return to the fuel tanks 50, 53, return just before the burner 16, or return before or after any fuel - oil heat exchanger or fuel pump along the main fuel flow path 110 or within the fuel hydraulic system, as the case may be. Generally, the fuel for actuation returns to a point on the main fuel flow path 110 within the engine 10 (rather than, for example, directly returning to the fuel tanks 50, 53 located at other positions in the aircraft 1 or returning to the connecting pipe between them).

[0458] Therefore, the supply steps 1310, 1320a can include controlling a plurality of fuel flow valves and optionally also include controlling a plurality of fuel pumps.

[0459] Figure 21FIG. 0 shows a flow chart of a method 1400 of operating a gas turbine engine 10 in accordance with various embodiments. The method 1400 is performed on an engine 10 that includes: an engine core 11 including a turbine 19, a combustor 16, a compressor 14, and a 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 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 supply system 152.

[0460] The method 1400 includes:

[0461] supplying 1410 fuel using the fuel supply system 152 for combustion in the combustor 16;

[0462] supplying 1420 fuel using the fuel supply system 152 to hydraulically drive at least one of the plurality of actuators 354 such that a peak fuel differential pressure across at least one fuel hydraulic actuator 354 during cruise conditions is at least 2400 kPa.

[0463] The fuel supplied has an SAF content of at least 25% by volume and may have a much higher SAF content. When the SAF content is higher, the peak differential pressure can be controlled at a higher level.

[0464] As described with respect to Figure 20 method 1300, one or more fuel flow circuits and associated valves can be used to control how much fuel diverges from the main fuel flow path 110 to the actuated system and where the fuel returns along the main fuel flow path 110.

[0465] Similarly, one or more valves and / or pumps can be appropriately controlled 1420 to control the peak fuel differential pressure across the fuel hydraulic actuator 354. One or more pressure sensors can be provided to provide feedback on the fuel differential pressure. It should be understood that controlling the fuel supply 1420 to the actuator 354 allows the differential pressure to be adjusted.

[0466] Figure 22FIG. 1500 is a flow chart showing a method 1500 of operating a gas turbine engine 10 according to various embodiments. The method 1500 is performed on an engine 10 that includes: an engine core 11 that includes a turbine 19, a combustor 16, a compressor 14, and a spool 26 that connects the turbine to the compressor; a fan 23 that is located upstream of the engine core 11 and is arranged to be driven by the spool 26, the fan including a plurality of fan blades; and a nacelle 21 that surrounds the fan 23 and the engine core 11 and defines a bypass duct 22 that is located radially outward of the engine core 11; a plurality of actuators 254; a fuel supply system 152; a primary fuel-oil heat exchanger 1220 and a secondary fuel-oil heat exchanger 1230.

[0467] The method 1500 includes:

[0468] supplying 1510 fuel using the fuel supply system 152 for combustion in the combustor 16;

[0469] supplying 1520 fuel using the fuel supply system 152 to hydraulically drive at least one of the plurality of actuators 354;

[0470] heating 1530 at least a majority of the fuel using the primary fuel-oil heat exchanger 1220; and

[0471] heating 1540 the fuel to be supplied to hydraulically drive at least one fuel hydraulic actuator further using the secondary fuel-oil heat exchanger 1230.

[0472] The method includes controlling the heating of the fuel in the heat exchangers 1220, 1230 such that, under cruise conditions, the heat transfer ratio:

[0473]

[0474] has a maximum value of at least 0.35.

[0475] The primary fuel-oil heat exchanger 1220 may be configured to transfer heat from oil within the engine 10 that is used to lubricate and cool various engine components such as a gearbox (if present) and / or bearings. The secondary fuel-oil heat exchanger 1230 may be configured to transfer heat from oil that is used to cool and optionally also lubricate a generator. As Figure 25 shown, in some embodiments, different, fluid-isolated oil flows may pass through each of the primary fuel-oil heat exchanger 1220 and the secondary fuel-oil heat exchanger 1230. In other embodiments, the same oil may pass through both heat exchangers, for example through the secondary fuel-oil heat exchanger 1230 and then through the primary fuel-oil heat exchanger 1220 in series.

[0476] The heat transfer ratio can be controlled by opening and closing valves so as to: recycle fuel through one or both of the heat exchangers 1220, 1230, bypass one or both of the heat exchangers 1220, 1230 with the fuel, recycle oil through one or both of the heat exchangers 1220, 1230, or bypass one or both of the heat exchangers 1220, 1230 with the oil. Additionally, one or more additional heat exchangers can be controlled to adjust the heat transfer ratio. For example, an air-oil heat exchanger can be used to cool the oil before the oil reaches the fuel-oil heat exchangers 1220, 1230, and in embodiments having multiple different oil flows, an oil-oil heat exchanger can be provided to transfer heat between the different oils. Accordingly, the flow rate of air and / or oil entering such heat exchangers can also be adjusted so as to affect the heat transfer ratio defined above.

[0477] The heat transfer ratio can be controlled by controlling one or more valves configured to enable a certain percentage of the fuel to be recycled through one or both of the fuel-oil heat exchangers 1220, 1230 or bypass one or both of the heat exchangers 1220, 1230. The heat transfer ratio can be controlled by controlling one or more valves configured to enable a certain percentage of the oil to be recycled through one or both of the heat exchangers 1220, 1230 or bypass one or both of the heat exchangers 1220, 1230.

[0478] As described with respect to Figure 20 and Figure 21 methods 1300, 1400, one or more fuel flow circuits and associated valves can be used to control how much fuel diverges from the main fuel flow path 110 to the actuators and the secondary heat exchanger 1230, and where the fuel returns along the main fuel flow path 110.

[0479] Figure 23 A flowchart of a method 1600 of operating a gas turbine engine 10 in accordance with various embodiments is shown. Method 1600 is performed on an engine 10 that includes: an engine core 11 including a turbine 19, a combustor 16, a compressor 14, and a 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 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 supply system 152.

[0480] Method 1600 includes:

[0481] supplying 1610 fuel using the fuel supply system 152 for combustion in the combustor 16; and

[0482] The fuel supply system 152 supplies 1620 fuel to drive the variable compressor vane actuator 854 among a plurality of actuators in a fuel hydraulic manner.

[0483] Supplying 1620 fuel using the fuel supply system 152 to drive the variable compressor vane actuator 854 in a fuel hydraulic manner and supplying to the combustor 16 may include supplying fuel that is thermally stable at 280 °C and optionally also thermally stable at higher temperatures.

[0484] As described with respect to Figures 20 to 22 the methods 1300, 1400, 1500, one or more fuel flow circuits and associated valves can be used to control how much fuel diverges from the main fuel flow path 110 to the actuated systems - in particular the variable stator vane actuator 854, and where the fuel returns along the main fuel flow path 110.

[0485] Figure 24 A flowchart of a method 1700 of operating a gas turbine engine 10 in accordance with various embodiments is shown. The method 1700 is performed on the engine 10, which includes: an engine core 11 including a turbine 19, a combustor 16, a compressor 14, and a spool 26 connecting the turbine to the compressor; a fan 23 located upstream of the engine core 11 and arranged to be driven by the spool 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; a fuel supply system 152; and at least one fuel - oil heat exchanger 1220 arranged to have oil and fuel flow therethrough, the at least one heat exchanger 1220 arranged to transfer heat from the oil to the fuel.

[0486] The method 1700 includes:

[0487] Using the fuel supply system 152 to supply 1710 fuel for combustion in the combustor 16;

[0488] Supplying 1720 fuel to drive at least one actuator 1254 among the plurality of actuators in a fuel hydraulic manner; and

[0489] Controlling 1730 at least one heat exchanger 1220 such that during cruise, the fuel temperature at entry to at least one actuator 1254 is at least 5 °C (and optionally at least 10 °C) higher than the fuel temperature at entry to the combustor 16.

[0490] The fuel supply system 152 may include two heat exchangers - a primary heat exchanger 1220 and a secondary heat exchanger 1230. Controlling 1730 at least one of the heat exchangers 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 to the fuel to be supplied to the fuel hydraulic actuator 1254.

[0491] The method 1700 may include determining at least one fuel characteristic of the fuel and controlling the temperature difference between the fuel temperature when entering at least one actuator 1254 and the fuel temperature when entering the burner 16 based on the at least one fuel characteristic. The fuel characteristic may be the SAF content of the fuel. The fuel characteristic may be determined by any method known in the art, including retrieving data from an on-board fuel information data store and determining or inferring from one or more sensed parameters.

[0492] The fuel characteristic may be or include calorific value, thermal stability, or the percentage of sustainable aviation fuel (SAF) in the fuel.

[0493] Controlling 1730 the fuel temperature when entering at least one actuator 1254 relative to the fuel temperature when entering the burner 16 may include actuating one or more valves so as to recirculate the fuel through one or more of the heat exchangers or bypass one or more of the heat exchangers 1220, 1230.

[0494] Controlling 1730 the fuel temperature when entering at least one actuator 1254 relative to the fuel temperature when entering the burner 16 may include actuating one or more valves so as to recirculate the oil through the heat exchangers 1220, 1230 or bypass the heat exchangers.

[0495] As described with respect to Figures 20 to 23 the methods 1300, 1400, 1500, 1600, one or more fuel flow circuits and associated valves may be used to control how much fuel diverges from the main fuel flow path 110 to the actuated system and where the fuel returns along the main fuel flow path 110.

[0496] It should be understood that the present invention is not limited to the above-described embodiments and various modifications and improvements can be made without departing from the concepts described herein. Unless mutually exclusive, any feature may be used alone or in combination with any other feature, and the present disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.

Claims

1. A gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core, the engine core comprising a turbine, a combustor, a compressor, and a mandrel connecting the turbine to the compressor; a fan located upstream of the engine core and arranged to be driven by the spindle, the fan comprising a plurality of fan blades; a nacelle surrounding the fan and the engine core and defining a bypass duct radially outward of the engine core, wherein a bypass ratio is at least 4, the bypass ratio being defined as the ratio of the mass flow rate of flow through the bypass duct to the mass flow rate of flow through the core under cruise conditions; multiple actuators; and a fuel supply system, wherein the fuel supply system is arranged to supply fuel for combustion in the burner and to supply fuel to hydraulically drive at least one actuator of the plurality of actuators with the fuel, and wherein the fuel supply system is arranged to be controlled to select, based on the sustainable aviation fuel - SAF - content of the fuel, between: causing the fuel to actuate at least one actuator; and The fuel is bypassed around the at least one actuator.

2. A gas turbine engine according to claim 1, wherein the spindle outputs drive directly to the fan, thereby driving the fan at the same rotational speed as the spindle, making the engine a direct drive turbine engine.

3. A gas turbine engine according to claim 1, wherein the turbine engine includes a gearbox, which receives input from the spindle and outputs drive to the fan, thereby driving the fan at a lower rotational speed than the spindle, making the engine a geared turbine engine.

4. The gas turbine engine of claim 1 , wherein the engine includes a turbine casing cooling system, wherein the plurality of actuators includes an actuator that is part of the turbine casing cooling system, and wherein the fuel supply system is configured to actuate or bypass the actuator that is part of the turbine casing cooling system based on the SAF content of the fuel.

5. The gas turbine engine of claim 1 , wherein the engine comprises a cabin bleed valve, the plurality of actuators comprises an actuator configured to actuate the cabin bleed valve, and the fuel supply system is configured to actuate or bypass the actuator configured to actuate the cabin bleed valve based on the SAF content of the fuel.

6. The gas turbine engine of claim 1 , wherein the engine includes a piloted bleed valve, the plurality of actuators includes an actuator configured to actuate the piloted bleed valve, and the fuel supply system is configured to actuate or bypass the piloted bleed valve based on the SAF content of the fuel.

7. A gas turbine engine according to claim 1, wherein the engine includes an engine thermal management system, the engine thermal management system includes a valve, wherein the plurality of actuators include an actuator configured to actuate the valve within the engine thermal management system, and the fuel supply system is configured to actuate or bypass the valve within the engine thermal management system based on the SAF content of the fuel.

8. A gas turbine engine according to claim 1, wherein the engine includes a generator thermal management system, the generator thermal management system includes a valve, wherein the plurality of actuators includes an actuator configured to actuate the valve within the generator thermal management system, and the fuel supply system is configured to actuate or bypass the valve within the generator thermal management system based on the SAF content of the fuel.

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

10. The gas turbine engine of claim 1, wherein the fuel supply system is arranged to cause the fuel to actuate the at least one actuator when the SAF content is above a threshold value, and to cause the fuel to bypass the at least one actuator when the SAF content is below the threshold value.

11. The gas turbine engine of claim 1 , wherein a minimum SAF content required to actuate the at least one actuator is at least 25%.

12. The gas turbine engine of claim 1, wherein a minimum SAF content required to actuate the at least one actuator is at least 50%.

13. The gas turbine engine of claim 1, wherein the fuel supply system is arranged to be controlled to select between hydraulic actuation of the fuel and bypassing two or more of the plurality of actuators.

14. The gas turbine engine of claim 1, wherein a maximum operating pressure differential during takeoff conditions when the at least one actuator is fuel hydraulically driven is at least 6,900 kPa.

15. The gas turbine engine of claim 1, wherein a maximum operating pressure differential during takeoff conditions when the at least one actuator is fuel hydraulically driven is greater than 7,000 kPa.

16. The gas turbine engine of claim 1, wherein a maximum operating pressure differential during cruise conditions when the at least one actuator is fuel hydraulically driven is at least 2,400 kPa.

17. The gas turbine engine of claim 1, wherein a maximum operating pressure differential during cruise conditions when the at least one actuator is fuel hydraulically driven is greater than 2,500 kPa.

18. A method of operating a gas turbine engine for an aircraft, the engine comprising: an engine core, the engine core comprising a turbine, a combustor, a compressor, and a mandrel connecting the turbine to the compressor; a fan located upstream of the engine core and arranged to be driven by the spindle, the fan comprising a plurality of fan blades; a nacelle surrounding the fan and the engine core and defining a bypass duct radially outward of the engine core, wherein a bypass ratio is at least 4, the bypass ratio being defined as the ratio of the mass flow rate of flow through the bypass duct to the mass flow rate of flow through the core under cruise conditions; multiple actuators; and Fuel supply system; And wherein the method comprises: supplying fuel for combustion in the burner using the fuel supply system; as well as Based on the sustainable aviation fuel -SAF- content of the fuel, a choice is made between: supplying fuel to at least one actuator of the plurality of actuators using the fuel supply system to hydraulically actuate the at least one actuator; and Using the fuel supply system, the fuel bypasses the at least one actuator.

19. The method of claim 18, wherein the fuel supply system is controllable to supply fuel to at least two of the plurality of actuators, and the method comprises selecting based on a sustainable aviation fuel (SAF) content of the fuel between: causing the fuel to hydraulically actuate the at least two actuators; and The fuel is bypassed around the at least two actuators.

20. The method of claim 18, wherein the fuel supply system is controllable to supply fuel to at least two actuators of the plurality of actuators, and the method comprises selecting, for each of the at least two actuators, individually between: causing the fuel to hydraulically actuate the actuator; and The fuel is bypassed around the actuator.

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