HEAT EXCHANGE BYPASS

Abstract

A method of operating a gas turbine engine, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, a combustion chamber arranged to combust a fuel, and a core shaft connecting the turbine to the compressor; a fan arranged upstream of the engine core; a fan shaft; at least one bearing arranged to support the fan shaft; an oil circulation system arranged to supply oil to the at least one bearing; and a heat exchange system comprising: an air-to-oil heat exchanger through which the oil in the oil circulation system flows; and a fuel-to-oil heat exchanger through which the oil in the oil circulation system and the fuel flow, such that heat is transferred between the oil and the fuel; a bypass pipe arranged to allow a portion of the oil to flow past one of the air-to-oil heat exchanger or the fuel-to-oil heat exchanger.and a bypass valve arranged to vary the proportion of oil passed through the bypass pipe; the method comprising controlling the heat exchange system such that under idle conditions a heat transfer ratio of: Heat transfer rate from oil to air ( kJkg − 1 ) Heat transfer rate from oil to fuel ( kJkg − 1 ) in the range of 0 to 1.5.

Description

The present disclosure relates to aircraft propulsion systems and methods of operating aircraft that include management of various fluids and heat exchange therebetween, and more particularly to management of an aircraft engine heat exchange system.In the aeronautical industry, development toward the use of fuels different from the currently widely used conventional kerosine-based jet fuels is expected. The fuels may have different fuel properties compared to petroleum-based hydrocarbon fuels. Thus, there is a need to account for fuel properties for these new fuels and adapt both the gas turbine engines themselves and the methods of operating gas turbine engines.According to a first aspect, there is provided a method of operating a gas turbine engine of an aircraft, the gas turbine engine comprising:an engine core comprising a turbine, a compressor, a combustor arranged to combust a fuel, and a core shaft connecting the turbine to the compressor;a fan disposed upstream of the engine core;a fan shaft;at least one bearing arranged for supporting the fan shaft;an oil circuit system arranged to supply the at least one bearing with oil; anda heat exchange system comprising:an air-oil heat exchanger through which the oil in the oil circulation system flows; anda fuel-oil heat exchanger through which the oil in the oil circulation system and the fuel flow so that heat is transferred between the oil and the fuel;a bypass pipe arranged such that a portion of the oil may bypass one of the air-oil heat exchanger and the fuel-oil heat exchanger; anda bypass valve configured to vary the amount of oil directed through the bypass tube,the method comprising:determining at least one fuel property of the fuel to be burned by the combustion chamber; andcontrolling the bypass valve based on the at least one fuel characteristic to adjust the amount of oil directed across each heat exchanger during cruise conditions.The inventors have recognized that the use of fuels different from conventional kerosine-based jet fuels, such as persistent flying fuels, may result in different fuel properties and that parameters under cruise conditions may be adjusted to take advantage of the different fuel properties. In particular, some fuels may be heated to higher temperatures than conventional fuels in one or more fuel-oil heat exchangers without significant increase in degradation products of the fuel, including coking and / or paint. This may improve the combustion efficiency of the fuel and / or increase the overall thermodynamic efficiency of the engine, with less heat being lost to the environment (e.g., via the air-oil heat exchanger). Higher fuel temperatures at the entrance to the combustion chamber may also enable a method that provides improved oil cooling (as the fuel may absorb more heat). The thermal stability of the fuel has an effect, for example, on how much heat the fuel can absorb or to which temperature the fuel can be brought without deposits forming in lines, burners and / or a hydromechanical unit or another engine component. Thus, considering the thermal stability of the fuel and transferring more or less heat from the oil to the fuel depending on the thermal stability of the fuel may provide more efficient oil cooling while avoiding coking or other deposition pathways of fuel degradation products (e.g., paint), thereby improving the performance of the aircraft.Using the fuel to remove more heat from the oil, rather than rely on heat transfer from the oil to the environment / air (e.g., in the air-to-oil heat exchanger), may also provide a thermodynamically more efficient engine.This may improve cooling of the oil prior to recirculation to the remainder of the turbine engine. Additionally, the improved cooling of the oil may in turn improve the cooling effect of the oil on the components of the engine through which it flows, such that, for example, a lower flow rate of the oil may provide the same cooling effect.In a transition time, in which the available flying fuels change over time during the lifetime of a gas turbine engine and depend, inter alia, on the geographical position of a refueling point, it is important that a relevant fuel property is determined and the control of the heat exchange system takes place on the basis of the respectively used fuel. Thus, the operation of gas turbines may be adjusted to retrieve the best from a variety of fuels. The controllable modulation valve arranged to adjust the oil flow through each heat exchanger plays a key role in the control of the heat exchange system.The fuel properties may be or include the heating value, thermal stability, or percentage of persistent flying fuel (SAF) in the fuel.The air-to-oil heat exchanger may help remove excess heat from the oil of the oil circulation system that is not transferred to the fuel. This may allow further cooling of the oil beyond the transfer of heat from the oil to the fuel in the fuel-oil heat exchanger.The air-oil heat exchanger and the fuel-oil heat exchanger may be arranged in series along the path of the oil circulation system. The bypass pipe may be arranged such that the oil flows through the fuel-oil heat exchanger as it passes the air-oil heat exchanger, or vice versa.The bypass valve allows the amount of heat transferred from the oil through the air-oil and fuel-oil heat exchangers to be varied. The amount of oil flow (and thus the heat transfer from the oil to the air or fuel) may be varied based on the temperature of the fuel leaving the fuel-oil heat exchanger or entering the combustion chamber, thus enabling control of the fuel temperature within a defined upper and lower limit. Similarly, by controlling the flow of oil, maintenance of the oil temperature in an appropriate range can be made possible. This control may help ensure improved turbine engine efficiency (e.g., by increasing fuel temperature) without risking unnecessary damage to a fuel pump or other components downstream of the fuel-oil heat exchanger by using a temperature that is too high for the durability of these components. The bypass valve may be configured to direct up to 100% of the oil via the fuel-oil heat exchanger. The bypass valve may be configured to ensure that not less than 70%, 80% or 90% of the oil is directed across the fuel-oil heat exchanger during cruise.The bypass valve may be configured to bypass a fixed portion of the oil through the bypass pipe to bypass one of the fuel-oil heat exchanger and the air-oil heat exchanger during operation of the engine, optionally wherein the fixed portion is determined based on the at least one determined fuel property at or prior to starting the engine or at or prior to reaching the cruise altitude. Alternatively, the bypass valve may be configured to bypass a variable portion of the oil flow through the bypass tube to bypass one of the fuel-oil heat exchanger and the air-oil heat exchanger during operation of the engine. The modulation valve may therefore be actively controlled to vary the proportion of oil routed across each heat exchanger, particularly in implementations where the aircraft carries multiple different fuels in different tanks and to change which fuel (or fuel mixture) is used in flight. The active control of the modulation valve may be automated and implemented by a controller of the heat exchange system.Alternatively, the oil circulation system, which may also be referred to as an oil circulation system, may branch such that a portion of the oil may flow along each branch. The air-oil and fuel-oil heat exchangers may be arranged in a parallel configuration on different branches of an oil circulation system. The heat exchange system may further include a modulation valve configured to vary the proportion of oil directed across each branch, and controlling the modulation valve may therefore adjust the proportion of oil directed across each branch during cruise conditions. In some examples, more than two heat exchangers and / or more than two branches may be provided. The two or more branches of the primary oil circulation system may re-join after the heat exchange system such that the oil re-joins after heat transfer from the oil.The heat exchange system may further comprise a secondary fuel-oil heat exchanger. The secondary fuel-oil heat exchanger may be a servo fuel-oil heat exchanger. A portion, but not all, of the fuel exiting the main fuel-oil heat exchanger may be directed to the servo fuel-oil heat exchanger. The power fuel-oil heat exchanger may further increase the temperature of the fuel before the fuel is provided for use in gas turbine engine power assist mechanisms (e.g., for fuel hydraulic actuation and / or heating). These servo mechanisms may include a nacelle anti-icing system. The servomechanisms may include engine actuators. The servomechanisms may include a turbine case cooling (TCC) servovalve. In these auxiliary systems, only fuel that has passed through the secondary fuel-oil heat exchanger may be used. The fuel used in these auxiliary systems may be recirculated to a fuel tank for later recirculation or may be recombined with other fuel exiting the main fuel-oil heat exchanger and entering the combustion chamber. The fuel passing through the secondary fuel-oil heat exchanger may therefore, in some implementations, not be provided to the combustion chamber, but instead be returned to an aircraft fuel tank, optionally after use in auxiliary systems such as fuel hydraulic actuators. The servo fuel-oil heat exchanger may be structurally similar or identical to the primary fuel-oil heat exchanger. The servo fuel-oil heat exchanger may be smaller than the primary fuel-oil heat exchanger. At least a portion of the fuel may not flow through the secondary fuel-oil heat exchanger.The heat exchange system may further include an additional bypass pipe arranged such that oil (or fuel) may bypass one or more heat exchangers of the heat exchange system. A bypass pipe may, in some implementations, effectively form an additional branch in a parallel branched oil system - the bypass valve or other oil valve may be configured to adjust the amount of oil directed through the or each bypass pipe for oil based on the one or more determined fuel properties and optionally based on one or more temperature measurements.The heat exchange system may further include a refrigeration cycle device, and the method may include using the refrigeration cycle device to provide thermal buoyancy by transferring further heat from the oil to the fuel, optionally such that the fuel temperature is raised above the oil temperature. The modulation valve or another oil valve may control how much oil flows through the refrigeration cycle device.The heat exchange system may further comprise branching fuel return paths and at least one valve that controls a division of the fuel flow. The branching paths may be arranged to return fuel from the heat exchange system at at least two different locations along a main fuel path leading to the combustor from where fuel enters the gas turbine engine. For example, fuel exiting a fuel-oil heat exchanger may be split into two or more branches, each branch opening back into the main fuel flow path at a different point. In some implementations, at least most of the fuel may flow through the fuel-oil heat exchanger such that a branch of the branching fuel path exiting that heat exchanger is itself the "main" fuel path - one or more smaller branches may be branched and re-connect to the main flow at another location, e.g., further downstream, e.g., after one or more components of the engine located downstream of the respective fuel-oil heat exchanger, or even further upstream (and thus functioning as a recirculation pipe).Subject to the suitability of the at least one determined fuel property, the method may comprise directing all of the oil over the fuel-oil heat exchanger for at least one or more uninterrupted periods of at least 30 minutes in cruise so that no heat is lost to the environment via the air-oil heat exchanger during at least some periods of time in cruise. Subject to the suitability of the at least one determined fuel property, the method may comprise passing at least 95% of the oil over the fuel-oil heat exchanger during at least 90% of the time spent in cruise, such that only very little, if any, heat is lost to the environment via the air-oil heat exchanger during at least 90% of cruise operation. Subject to the suitability of the at least one ascertained fuel property, the method can comprise at least 80% of the heat dissipated by the oil during cruise flight being transferred to the fuel. In some implementations, all of the heat dissipated from the oil during cruise may be transferred to the fuel over at least 90% of the time spent during cruise.The at least one fuel property may be or comprise thermal stability. At least 80% of the heat dissipated by the oil during cruise can be transferred to the fuel provided the fuel is operationally stable at temperatures above 140° C.The at least one fuel property may be or include the content of aromatic hydrocarbons in the fuel. At least 80% of the heat dissipated by the oil during cruise can be transferred to the fuel provided the fuel has a molar proportion of aromatic hydrocarbons of less than 12%.The at least one fuel property may be or may comprise the percentage of sustainable flying fuel - % SAF - in the fuel. The SAF fraction (X %) may be volumetric. At least 80% of the heat dissipated by the oil during cruise can be transferred to the fuel provided the fuel has a SAF content of more than 50%.The at least one fuel property may be or include the heating value of the fuel. At least 80% of the heat dissipated by the oil during cruise flight can be transferred to the fuel provided the fuel has a heating value of at least 43.5 MJ / kg.The at least one fuel property of the fuel may include at least one of the following features: i. percent of residual flying fuel in the fuel; ii. concentration of the heteroatomic species of the fuel; iii. content of aromatic hydrocarbons in the fuel; iv. content of multiaromatic hydrocarbons in the fuel; v. percent of nitrogenic species in the fuel; vi. presence or percent of a tracer species or trace element in the fuel; vii. hydrogen-carbon ratio of the fuel; viii. hydrocarbon distribution of the fuel; ix. level of non-volatile particle emissions during combustion; x. naphthalene content of the fuel; xi. sulfur content of the fuel; xii. cycloparaffin content of the fuel; xiii. oxygen content of the fuel; xiv. thermal stability of the fuel; xv. the degree of coking of the fuel; xvi. an indication that the fuel is a fossil fuel; xvii. at least one of density, viscosity, heating value, and heat capacity.The method may further include chemically or physically sensing one or more parameters relevant to fuel in a fuel tank after refueling (the fuel tank configured to supply fuel to the combustion chamber via the heat exchange system). The detected parameters may be fuel characteristics or may be used to calculate or derive fuel characteristics. For example, the detected parameters may be one or more of shaft speed, turbo fan power ratio (TPR), and / or engine pressure ratio (EPR), and mass flow rate of the fuel from which the heating value (a fuel property) may be determined, or the detected parameters may be fuel density and / or the presence of a tracer, both of which are themselves fuel properties. Determining at least one fuel property may include retrieving stored data about fuel properties. Chemically and / or physically determining one or more parameters of the fuel in the fuel tank may be performed by sampling the entrained fuel from the fuel tank for off-wing testing. Obtaining a sample of the entrained fuel from the fuel tank may include obtaining a fuel sample prior to refueling.Determining the at least one fuel property of the fuel may include: obtaining the at least one fuel property of a fuel already present in the fuel tank before refueling; determining at least one fuel property of a fuel supplied to the fuel tank during refueling; and calculating (based on this information) at least one fuel property of the resulting fuel in the fuel tank after refueling.The determining of the at least one fuel property may be performed based on the detection of at least one fuel feature. The fuel characteristic may be the fuel property or may be used to calculate or otherwise determine (e.g., by retrieving from a look-up table) the fuel property. The detection can be performed on the wing.The determination of the at least one fuel property may be performed based on the received data regarding the composition of the fuel. The fuel composition data may be provided to the aircraft during refueling. The fuel composition data may be manually input.According to a second aspect, there is provided a gas turbine engine for an aircraft comprising:an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor;a fan disposed upstream of the engine core; anda gearbox receiving input from the core shaft and outputting drive to the fan to drive the fan at a lower speed than the core shaft;an oil circuit system arranged to supply the transmission with oil; anda heat exchange system comprising:an air-oil heat exchanger through which the oil in the oil circulation system flows; anda fuel-oil heat exchanger through which the oil in the oil circulation system and the fuel flow so that heat is transferred between the oil and the fuel;a bypass pipe arranged such that a portion of the oil may bypass one of the air-oil heat exchanger and the fuel-oil heat exchanger;a bypass valve configured to vary the proportion of oil directed through the bypass tube; and a fuel composition determination module configured to determine at least one fuel property of fuel to be burned by the combustion chamber,wherein the bypass valve is configured to be controlled based on the at least one fuel property to adjust the proportion of oil directed across each heat exchanger during cruise conditions.The turbine may be a first turbine, the compressor may be a first compressor, and the core shaft may be a first core shaft. The engine core may further include a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, the second compressor, and the second core shaft may be arranged to rotate at a higher speed than the first core shaft.The core shaft may output the drive directly to the fan to drive the fan at the same speed as the core shaft, such that the engine is a direct drive turbine engine.The engine of the second aspect may be configured to perform the method of the first aspect and may include any of the features described with respect to the first aspect.According to a third aspect, there is provided a method of operating a gas turbine engine comprising:an engine core comprising a turbine, a compressor, a combustor arranged to combust a fuel, and a core shaft connecting the turbine to the compressor;a fan disposed upstream of the engine core;a fan shaft;at least one bearing arranged for supporting the fan shaft;an oil circuit system arranged to supply the at least one bearing with oil; anda heat exchange system comprising:an air-oil heat exchanger through which the oil in the oil circulation system flows; anda fuel-oil heat exchanger through which the oil in the oil circulation system and the fuel flow so that heat is transferred between the oil and the fuel; anda bypass pipe arranged such that a portion of the oil may bypass one of the air-oil heat exchanger or the fuel-oil heat exchanger; anda bypass valve configured to vary the amount of oil directed through the bypass tube,the method comprising controlling the heat exchange system such that a heat transfer ratio of:The range is from 0 to 0.67.The inventors have recognized that the use of fuels different from conventional kerosine-based jet fuels, such as persistent flying fuels, may result in different fuel properties and that parameters under cruise conditions may be adjusted to take advantage of the different fuel properties. In particular, some fuels may be heated to higher temperatures than conventional fuels in one or more fuel-oil heat exchangers without substantially increasing coking or other deposition pathways for fuel degradation products (e.g., fouling). This may enable a method that provides improved oil cooling (as the fuel may draw more heat) and may also improve overall thermodynamic efficiency of the engine as less heat is lost to the environment. The controllable heat exchange system is key to controlling the heat transfer ratio.While it should be appreciated that the ratio is dimensionless, in the examples described herein, heat transfer is measured from the unit mass of fuel to provide a heat transfer rate normalized to cruise fuel flow rate variations. For oil-to-air heat transfer, the definition "per unit volume of fuel" may be set equal to "per fixed time period as a function of fuel flow rate" to also provide normalization of fuel flow rate. It should be appreciated that heat transfer through the heat exchanger must be completed before the fuel reaches the combustion chamber. Any additional temperature rise in the combustion chamber itself is due to combustion, not to the heat exchange system. It will be appreciated that heat transfer is indicated in terms of the mass (kg) of fuel reaching the combustion chamber to account for the fuel flow rate and any recirculation through one or more heat exchangers or bypass of one or more heat exchangers as described elsewhere herein. The amount of heat transferred to the fuel may therefore be calculated based on a temperature of fuel as it approaches or enters the combustion chamber compared to a temperature of fuel in a fuel tank of the aircraft. For comparison purposes, the amount of heat transferred from the oil to the air may be determined from the drop in temperature of the oil in the air-to-oil heat exchanger(s) or by comparing the heat recovered from the fuel with the total heat loss of the oil, assuming that the difference is due to the heat loss of the oil to the air (considering all other loss or generation sources such as fuel recirculation).The method may comprise controlling the heat exchange system such that the heat transfer ratio under cruise conditions is in the range of from 0 to 0.60, from 0 to 0.50, from 0 to 0.40, from 0 to 0.30, from 0 to 0.20, or from 0 to 0.10.Controlling the heat exchange system to adjust the heat transfer ratio may include decreasing the amount of oil directed via the at least one air-oil heat exchanger when the heat transfer ratio is too high.The at least one bypass pipe may be arranged such that oil may bypass the air-oil heat exchanger such that oil flowing through the bypass pipe flows through the fuel-oil heat exchanger without flowing through the air-oil heat exchanger. Controlling the heat exchange system to adjust the heat transfer ratio may include increasing the amount of oil flowing through the bypass pipe when the heat transfer ratio is too high.The heat exchange system may include at least one return pipe arranged such that a liquid (oil or fuel) may flow through a heat exchanger multiple times, and controlling the heat exchange system to adjust the heat transfer ratio may include modulating the amount of liquid directed through the return pipe.The heat exchange system may further comprise a secondary fuel-oil heat exchanger. The secondary fuel-oil heat exchanger may be a servo fuel-oil heat exchanger. A portion, but not all, of the fuel exiting the main fuel-oil heat exchanger may be directed to the servo fuel-oil heat exchanger. The power fuel-oil heat exchanger may further increase the temperature of the fuel before the fuel is provided for use in gas turbine engine power assist mechanisms (e.g., for fuel hydraulic actuation and / or heating). These servo mechanisms may include a nacelle anti-icing system. The servomechanisms may include engine actuators. The servomechanisms may include a turbine case cooling (TCC) servovalve. In these auxiliary systems, only fuel that has passed through the secondary fuel-oil heat exchanger may be used. The fuel used in these auxiliary systems may be recirculated to a fuel tank for later recirculation or may be recombined with other fuel exiting the main fuel-oil heat exchanger and entering the combustion chamber. The fuel passing through the secondary fuel-oil heat exchanger may therefore, in some implementations, not be provided to the combustion chamber, but instead be returned to an aircraft fuel tank, optionally after use in auxiliary systems such as fuel hydraulic actuators. The servo fuel-oil heat exchanger may be structurally similar or identical to the primary fuel-oil heat exchanger. The servo fuel-oil heat exchanger may be smaller than the primary fuel-oil heat exchanger. At least a portion of the fuel may not flow through the secondary fuel-oil heat exchanger.The heat exchange system may include a refrigeration cycle device. The method may include using the refrigeration cycle device to provide thermal lift by transferring further heat from the oil to the fuel, optionally such that the fuel temperature is raised above the oil temperature. The heat transfer ratio of examples using a refrigeration cycle device may be in the range of 0 to 0.40.In implementations where no refrigeration cycle device is used or where the heat exchange system is not configured to provide thermal buoyancy, the heat transfer ratio may be in the range of 0.38 to 0.67.The method may include controlling the heat exchange system under cruise conditions such that the heat transfer ratio is in the range of 0 to 0.2 if the temperature of the fuel entering the combustion chamber is at least 160° C.The method may include controlling the heat exchange system under cruise conditions such that the heat transfer ratio is in the range of 0 to 0.1 if the temperature of the fuel entering the combustion chamber is at least 180° C.The method may include controlling the heat exchange system under cruise conditions such that the heat transfer ratio is in the range of 0 to 0.45 if the temperature of the fuel entering the combustion chamber is at least 140° C.The method may include controlling the heat exchange system under cruise conditions such that the heat transfer ratio is in the range of 0 to 0.2 if the fuel is at least 70% sustainable flight fuel.The method may include controlling the heat exchange system under cruise conditions such that the heat transfer ratio is in the range of 0 to 0.1 if the fuel is at least 80% sustainable flight fuel.The method may comprise maintaining the oil to air heat transfer rate in the range of 0 to 100 kJ per kilogram of fuel at cruise conditions and optionally 0 to 35 kJ / kg, wherein no more than 20% of the heat dissipated by the oil in cruise is transferred to the air.The method may comprise maintaining the oil to fuel heat transfer rate in the range of 110 to 200 kJ per kilogram of fuel at cruise conditions and optionally 150 to 200 kJ / kg, wherein at least 80% of the heat dissipated from the oil during cruise is transferred to the fuel.The methods of the first and third aspects may supplement each other and may be performed together in various implementations. The method of the third aspect may be performed using the engine of the second aspect.According to a fourth aspect, there is provided a gas turbine engine for an aircraft, the engine comprising:an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor;a fan disposed upstream of the engine core;a fan shaft;at least one bearing arranged for supporting the fan shaft;an oil circuit system arranged to supply the transmission with oil; anda heat exchange system comprising:an air-oil heat exchanger through which the oil in the oil circulation system flows;a fuel-oil heat exchanger through which the oil in the oil circulation system and the fuel flow so that heat is transferred between the oil and the fuel;a bypass pipe arranged such that a portion of the oil may bypass one of the air-oil heat exchanger or the fuel-oil heat exchanger; anda bypass valve configured to vary the amount of oil directed through the bypass tube,and wherein the heat exchange system is adapted to be controlled so that a heat transfer ratio of:The range is from 0 to 0.67.The turbine may be a first turbine, the compressor may be a first compressor, and the core shaft may be a first core shaft. The engine core may further include a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, the second compressor, and the second core shaft may be arranged to rotate at a higher speed than the first core shaft.The heat exchange system may further comprise branching fuel return paths and at least one valve that controls a division of the fuel flow. The branching paths may be arranged to return fuel from the heat exchange system at at least two different locations along a main fuel path leading to the combustor from where fuel enters the gas turbine engine.The bypass pipe may be arranged such that a portion of the oil is bypassed by the air-oil heat exchanger such that the oil flowing through the bypass pipe flows through the fuel-oil heat exchanger without flowing through the air-oil heat exchanger.The core shaft may output the drive directly to the fan to drive the fan at the same speed as the core shaft, such that the engine is a direct drive turbine engine.The air-oil heat exchanger and the fuel-oil heat exchanger may be arranged in a series configuration on the oil circulation system.Alternatively, the oil circulation system, which may also be referred to as an oil circulation system, may branch such that a portion of the oil may flow along each branch. The air-oil and fuel-oil heat exchangers may be arranged in a parallel configuration on different branches of an oil circulation system.The engine of the fourth aspect may be configured to perform the method of the first and / or third aspect and may include any of the features described with respect to any of the preceding aspects.According to a fifth aspect, there is provided a method of operating a gas turbine engine comprising:an engine core comprising a turbine, a compressor, a combustor arranged to combust a fuel, and a core shaft connecting the turbine to the compressor;a fan disposed upstream of the engine core;a fan shaft;at least one bearing arranged for supporting the fan shaft;an oil circuit system arranged to supply the at least one bearing with oil; anda heat exchange system comprising:an air-oil heat exchanger through which the oil in the oil circulation system flows; anda fuel-oil heat exchanger through which the oil in the oil circulation system and the fuel flow so that heat is transferred between the oil and the fuel; anda bypass pipe arranged such that a portion of the oil may bypass one of the air-oil heat exchanger or the fuel-oil heat exchanger; anda bypass valve configured to vary the amount of oil directed through the bypass tube,the method comprising controlling the heat exchange system such that a heat transfer ratio of:The range is from 0 to 1.5.The inventors have recognized that the use of fuels different from conventional kerosine-based jet fuels, such as persistent flying fuels, may result in different fuel properties and that parameters under idle conditions may be adjusted to take advantage of the different fuel properties. In particular, some fuels may be heated to higher temperatures than conventional fuels in one or more fuel-oil heat exchangers without substantially increasing coking or other deposition pathways for fuel degradation products (e.g., fouling). This may enable a method that provides improved oil cooling (as the fuel may draw more heat) and may also improve overall thermodynamic efficiency of the engine as less heat is lost to the environment. The controllable heat exchange system is key to controlling the heat transfer ratio.Furthermore, cruise conditions generally account for a much greater proportion of engine operating time, but the inventors also recognized the meaning of idle operation. Since the fuel flow rate is much lower at idle than at cruise, even a relatively low heat load on the fuel may result in a high temperature rise. The use of non-conventional fuels may therefore have an even greater impact on optimal approaches to thermal management under idle conditions, e.g., during aircraft launch, stand-by running during boarding, rolling (to a runway or hangar, or between other ground locations), or "flight idle conditions" such as descent. Since the operating conditions are very different between cruise and idle, at least with respect to the desired thrust of the engine, a different control of the heat exchange system is appropriate.Although the ratio is dimensionless, in the examples described herein, heat transfer is measured per unit mass of fuel, providing a heat transfer rate normalized for idle fuel flow rate variations. For oil-to-air heat transfer, the definition "per unit mass of fuel" may be set equal to "per fixed time period as a function of fuel flow rate" to also provide normalization of fuel flow rate. It should be appreciated that heat transfer through the heat exchanger must be completed before the fuel reaches the combustion chamber. Any additional temperature rise in the combustion chamber itself is due to combustion, not to the heat exchange system. It should be appreciated that heat transfer per unit mass (kg) of fuel reaching the combustion chamber is indicated to account for fuel flow rate and any recirculation through one or more heat exchangers or bypass of one or more heat exchangers as described elsewhere herein. The amount of heat transferred to the fuel may therefore be calculated based on a temperature of fuel as it approaches or enters the combustion chamber compared to a temperature of fuel in a fuel tank of the aircraft. For comparison purposes, the amount of heat transferred from the oil to the air may be determined from a drop in temperature of the oil across the air-to-oil heat exchanger / exchangers or may be determined by comparing the heat recovered from the fuel with the total heat loss of the oil, assuming that the difference is due to the heat loss of the oil to the air (after considering other loss or generation sources such as the fuel recirculation or fuel pump).The method may comprise controlling the heat exchange system such that the heat transfer ratio is below 1.0 under idle conditions, in the range of 0 to 0.60, 0 to 0.50, 0 to 0.40, 0 to 0.30, 0 to 0.20, or 0 to 0.10.Controlling the heat exchange system to adjust the heat transfer ratio may include decreasing the amount of oil directed via the at least one air-oil heat exchanger when the heat transfer ratio is too high.The at least one bypass pipe may be arranged such that oil may bypass the air-oil heat exchanger such that oil flowing through the bypass pipe flows through the fuel-oil heat exchanger without flowing through the air-oil heat exchanger. Controlling the heat exchange system to adjust the heat transfer ratio may include increasing the amount of oil flowing through the bypass pipe when the heat transfer ratio is too high.The heat exchange system may include at least one return pipe arranged such that a liquid (oil or fuel) may flow through a heat exchanger multiple times, and controlling the heat exchange system to adjust the heat transfer ratio may include modulating the amount of liquid directed through the return pipe.The heat exchange system may further comprise a secondary fuel-oil heat exchanger. The secondary fuel-oil heat exchanger may be a servo fuel-oil heat exchanger. A portion, but not all, of the fuel exiting the main fuel-oil heat exchanger may be directed to the servo fuel-oil heat exchanger. The power fuel-oil heat exchanger may further increase the temperature of the fuel before the fuel is provided for use in gas turbine engine power assist mechanisms (e.g., for fuel hydraulic actuation and / or heating). These servo mechanisms may include a nacelle anti-icing system. The servomechanisms may include engine actuators. The servomechanisms may include a turbine case cooling (TCC) servovalve. In these auxiliary systems, only fuel that has passed through the secondary fuel-oil heat exchanger may be used. The fuel used in these auxiliary systems may be recirculated to a fuel tank for later recirculation or may be recombined with other fuel exiting the main fuel-oil heat exchanger and entering the combustion chamber. The fuel passing through the secondary fuel-oil heat exchanger may therefore, in some implementations, not be provided to the combustion chamber, but instead be returned to an aircraft fuel tank, optionally after use in auxiliary systems such as fuel hydraulic actuators. The servo fuel-oil heat exchanger may be structurally similar or identical to the primary fuel-oil heat exchanger. The servo fuel-oil heat exchanger may be smaller than the primary fuel-oil heat exchanger. At least a portion of the fuel may not flow through the secondary fuel-oil heat exchanger.The heat exchange system may include a refrigeration cycle device. The method may include using the refrigeration cycle device to provide thermal lift by transferring further heat from the oil to the fuel, optionally such that the fuel temperature is raised above the oil temperature. The heat transfer ratio of examples using a refrigeration cycle device may be in the range of 0 to 0.40.If the heat exchange system is not arranged to provide thermal buoyancy, the heat exchange system may be controlled such that the heat transfer ratio is in the range of 0.38 to 1.2.The method may include controlling the heat exchange system under idle conditions such that the heat transfer ratio is in the range of 0.3 to 1.5 if the temperature of the fuel entering the combustion chamber is below 180° C.The method may include controlling the heat exchange system under idle conditions such that the heat transfer ratio is in the range of 0 to 0.3 if the temperature of the fuel entering the combustion chamber is above 180° C.Under idle conditions, the method may include controlling the heat exchange system such that the heat transfer ratio is in the range of 0 to 0.2 if the fuel is at least 70% sustainable aviation fuel.The method may comprise maintaining the heat transfer rate of oil to idle air in the range of 0 to 180 kJ per kilogram of fuel at idle conditions and optionally 0 to 60 kJ / kg, wherein no more than 20% of the heat dissipated by the idle oil is transferred to the air.The method may comprise maintaining the heat transfer rate of oil to fuel at idle in the range of 100 to 300 kJ per kilogram of fuel at idle conditions and optionally 200 to 300 kJ / kg, wherein at least 80% of the heat dissipated from the oil at idle is transferred to the fuel.The air-oil heat exchanger and the fuel-oil heat exchanger may be arranged in series in the oil circulation system.Alternatively, the oil circulation system, which may also be referred to as an oil circulation system, may branch such that a portion of the oil may flow along each branch. The air-oil and fuel-oil heat exchangers may be arranged in a parallel configuration on different branches of an oil circulation system. The heat exchange system may further include a modulation valve configured to vary the proportion of oil directed across each branch, and controlling the modulation valve may therefore adjust the proportion of oil directed across each branch during cruise conditions. In some examples, more than two heat exchangers and / or more than two branches may be provided. The two or more branches of the primary oil circulation system may re-join after the heat exchange system such that the oil re-joins after heat transfer from the oil.Further, the heat exchange system may comprise branching fuel return paths and at least one valve controlling a division of the fuel flow, the branching paths being arranged to return fuel from the heat exchange system at least two different locations along a main fuel path leading to the combustion chamber from where the fuel enters the gas turbine engine.The methods of the first, third, and fifth aspects may complement each other and may be performed together in various implementations. The method of the fifth aspect may be performed using the engine of the second or fourth aspect.According to a sixth aspect, there is provided a gas turbine engine for an aircraft, the engine comprising:an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor;a fan disposed upstream of the engine core;a fan shaft;at least one bearing arranged for supporting the fan shaft;an oil circuit system arranged to supply the transmission with oil; anda heat exchange system comprising:an air-oil heat exchanger through which the oil in the oil circulation system flows;a fuel-oil heat exchanger through which the oil in the oil circulation system and the fuel flow so that heat is transferred between the oil and the fuel;a bypass pipe arranged such that a portion of the oil may bypass one of the air-oil heat exchanger or the fuel-oil heat exchanger; anda bypass valve configured to vary the amount of oil directed through the bypass tube,and wherein the heat exchange system is arranged to be controlled such that under idle conditions a heat transfer ratio of:The range is from 0 to 1.5.The turbine may be a first turbine, the compressor may be a first compressor, and the core shaft may be a first core shaft. The engine core may further include a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, the second compressor, and the second core shaft may be arranged to rotate at a higher speed than the first core shaft.The heat exchange system may further comprise branching fuel return paths and at least one valve that controls a division of the fuel flow. The branching paths may be arranged to return fuel from the heat exchange system at at least two different locations along a main fuel path leading to the combustor from where fuel enters the gas turbine engine.The bypass pipe may be arranged such that a portion of the oil is bypassed by the air-oil heat exchanger such that the oil flowing through the bypass pipe flows through the fuel-oil heat exchanger without flowing through the air-oil heat exchanger.The core shaft may output the drive directly to the fan to drive the fan at the same speed as the core shaft, such that the engine is a direct drive turbine engine.The air-oil heat exchanger and a fuel-oil heat exchanger may be arranged in a series configuration in the oil circulation system.Alternatively, the oil circulation system, which may also be referred to as an oil circulation system, may branch such that a portion of the oil may flow along each branch. The air-oil and fuel-oil heat exchangers may be arranged in a parallel configuration on different branches of an oil circulation system.The engine of the sixth aspect may be configured to perform the method of the first, third and / or fifth aspect and may include any of the features described with respect to any of the preceding aspects.According to a seventh aspect, there is provided a method of operating a gas turbine engine of an aircraft, the gas turbine engine comprising:an engine core comprising a turbine, a compressor, a combustor arranged to combust a fuel, and a core shaft connecting the turbine to the compressor;a fan disposed upstream of the engine core;a fan shaft;at least one bearing arranged for supporting the fan shaft;at least one assist system arranged to consume a portion of the fuel;an oil circuit system arranged to supply the at least one bearing with oil; anda heat exchange system comprising:a primary fuel-oil heat exchanger through which the oil in the oil circulation system and at least substantially all of the fuel flow flow so that heat is transferred between the oil and the fuel; anda secondary fuel-oil heat exchanger through which the oil in the oil circulation system and a portion of the fuel flow so as to transfer heat between the oil and the fuel, the secondary fuel-oil heat exchanger being arranged to supply fuel to the at least one auxiliary system; anda fuel pump arranged to pump the fuel, the fuel pump being arranged along a fuel flow path downstream of the primary fuel-oil heat exchanger and upstream of the secondary fuel-oil heat exchanger;wherein the method further comprises controlling the heat exchange system such that, under cruise conditions, the fuel temperature upon exiting the secondary heat exchanger is in the range of 120° C. to 200° C.The inventors have recognized that the use of fuels different from conventional kerosine-based jet fuels, such as persistent flying fuels, may result in different fuel properties and that parameters under idle conditions may be adjusted to take advantage of the different fuel properties. In particular, some fuels may be heated to higher temperatures than conventional fuels in one or more fuel-oil heat exchangers without substantially increasing coking or other deposition pathways for fuel degradation products (e.g., fouling). This may enable a method that provides improved oil cooling (as the fuel may draw more heat) and may also improve overall thermal efficiency of the engine as less heat is lost to the environment. The controllable heat exchange system is key to controlling the heat transfer ratio.The secondary fuel-oil heat exchanger may be a servo fuel-oil heat exchanger. A portion, but not all, of the fuel exiting the main fuel-oil heat exchanger may be directed to the servo fuel-oil heat exchanger. The power fuel-oil heat exchanger may further increase the temperature of the fuel before the fuel is provided for use in gas turbine engine power assist mechanisms (e.g., for fuel hydraulic actuation and / or heating). These servo mechanisms may include a nacelle anti-icing system. The servomechanisms may include engine actuators. The servomechanisms may include a turbine case cooling (TCC) servovalve. In these auxiliary systems, only fuel that has passed through the secondary fuel-oil heat exchanger may be used. The fuel used in these auxiliary systems may be recirculated to a fuel tank for later recirculation or may be recombined with other fuel exiting the main fuel-oil heat exchanger and entering the combustion chamber. The fuel passing through the secondary fuel-oil heat exchanger may therefore, in some implementations, not be provided to the combustion chamber, but instead be returned to an aircraft fuel tank, optionally after use in auxiliary systems such as fuel hydraulic actuators. The servo fuel-oil heat exchanger may be structurally similar or identical to the primary fuel-oil heat exchanger. The servo fuel-oil heat exchanger may be smaller than the primary fuel-oil heat exchanger. At least a portion of the fuel may not flow through the secondary fuel-oil heat exchanger.The method may include controlling the heat exchange system such that, during cruise conditions, the fuel temperature upon exiting the secondary heat exchanger is in the range of 120° C. to 180° C.The method may include controlling the heat exchange system such that, during cruise conditions, the fuel temperature upon exiting the secondary heat exchanger is in the range of 135° C. to 200° C.The method may include controlling the heat exchange system such that, during cruise conditions, the fuel temperature upon exiting the secondary heat exchanger is in the range of 135° C. to 180° C.The method may include controlling the heat exchange system such that, under cruise conditions, the fuel temperature upon exiting the secondary heat exchanger is in the range of 150° C. to 200° C., 150° C. to 180° C., or 150° C. to 170° C. The fuel temperature upon exiting the secondary heat exchanger may be about 120° C., 130° C., 140° C., 150° C., 160° C., or up to 200° C.The heat exchange system may further include a fuel bypass pipe arranged such that a portion of the fuel may bypass at least one of the primary fuel-oil heat exchanger and the secondary fuel-oil heat exchanger; and a fuel bypass valve arranged such that the portion of the fuel directed through the fuel bypass pipe may be varied. Controlling the heat exchange system may include controlling the fuel bypass valve to adjust the proportion of fuel directed over each of the primary and secondary fuel-oil heat exchangers during cruise conditions.The heat exchange system may further include an oil bypass pipe arranged such that a portion of the oil may bypass at least one of the primary fuel-oil heat exchanger and the secondary fuel-oil heat exchanger; and an oil bypass valve arranged such that the portion of the fuel directed through the oil bypass pipe may be varied. Controlling the heat exchange system may include controlling the oil bypass valve to adjust the amount of oil directed over each of the primary and secondary fuel-oil heat exchangers during cruise conditions.Under cruise conditions, the ratio may be at least 0.3.This ratio may be referred to as the secondary heat exchanger fuel flow ratio for brevity.The fuel flow ratio of the secondary heat exchanger may be at least 0.35, 0.4, 0.45, 0.5, or 0.55. The fuel flow ratio of the secondary heat exchanger may be about 0.6.The heat exchange system may further include a refrigeration cycle device, and the method may include using the refrigeration cycle device to provide thermal buoyancy by transferring further heat from the oil to the fuel, optionally such that the fuel temperature is raised above the oil temperature. An oil valve may control how much oil flows through the refrigeration cycle device.The heat exchange system may further comprise branching fuel return paths and at least one valve controlling a division of the fuel flow, the branching paths being arranged to return fuel from the heat exchange system at at least two different locations along a main fuel flow path. For example, fuel exiting a fuel-oil heat exchanger may be split into two or more branches, each branch opening back into the main fuel flow path at a different point. In some implementations, at least most of the fuel may flow through the fuel-oil heat exchanger such that a branch of the branching fuel path exiting that heat exchanger is itself the "main" fuel path - one or more smaller branches may be branched and re-connect to the main flow at another location, e.g., further downstream, e.g., after one or more components of the engine located downstream of the respective fuel-oil heat exchanger, or even further upstream (and thus functioning as a recirculation pipe).The primary fuel-oil heat exchanger and the secondary fuel-oil heat exchanger may be arranged in series along the oil circulation system.The heat exchange system may further comprise at least one air-oil heat exchanger through which the oil of the oil circulation system flows.According to an eighth aspect of the invention there is provided a gas turbine engine for an aircraft comprising:an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor;a fan disposed upstream of the engine core; anda gearbox receiving input from the core shaft and outputting drive to the fan to drive the fan at a lower speed than the core shaft;an oil circuit system arranged to supply the transmission with oil; anda heat exchange system comprising:a primary fuel-oil heat exchanger through which the oil in the oil circulation system and at least substantially all of the fuel flow flow so that heat is transferred between the oil and the fuel; anda secondary fuel-oil heat exchanger through which the oil in the oil circulation system and a portion of the fuel flow so as to transfer heat between the oil and the fuel, the secondary fuel-oil heat exchanger being arranged to supply fuel to the at least one auxiliary system; anda fuel pump arranged to pump the fuel, the fuel pump being arranged along a fuel flow path downstream of the primary fuel-oil heat exchanger and upstream of the secondary fuel-oil heat exchanger;wherein the heat exchange system is configured to be controlled such that, under cruise conditions, the fuel temperature upon exiting the second heat exchanger is in the range of 120° C. to 200° C.The heat exchange system may be configured to be controlled such that, under cruise conditions, the fuel temperature upon exiting the second heat exchanger is in the range of 120° C. to 180° C.The turbine may be a first turbine, the compressor may be a first compressor, and the core shaft may be a first core shaft. The engine core may further include a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor; and the second turbine, the second compressor, and the second core shaft may be arranged to rotate at a higher speed than the first core shaft.The core shaft may output the drive directly to the fan to drive the fan at the same speed as the core shaft, such that the engine is a direct drive turbine engine.The engine of the eighth aspect may be configured to perform the method of the first, third, fifth and / or seventh aspect, and may include any of the features described with respect to any of the preceding aspects.According to a ninth aspect, there is provided a method of operating a gas turbine engine of an aircraft, the gas turbine engine comprising:an engine core comprising a turbine, a compressor, a combustor arranged to combust a fuel, and a core shaft connecting the turbine to the compressor;a fan disposed upstream of the engine core;a fan shaft;at least one bearing arranged for supporting the fan shaft;an oil circuit system arranged for supplying the at least one bearing with oil;a heat exchange system comprising:an air-oil heat exchanger through which the oil in the oil circulation system flows;a fuel-oil heat exchanger through which the oil in the oil circulation system and the fuel flow so that heat is transferred between the oil and the fuel; andat least one valve configured such that at least one of the oil and air flow rates can be varied by at least one of the fuel-oil heat exchanger and the air-oil heat exchanger; anda temperature sensor arranged to provide an indication of fuel temperature downstream of the fuel-oil heat exchanger (optionally at the inlet to the combustion chamber),the method comprising:determining whether the fuel temperature has increased above a set threshold during cruise conditions based on an output of the temperature sensor; andresponsive to determining that the fuel temperature has increased above the set threshold during cruise conditions, controlling the at least one valve to change the at least one flow rate through the at least one heat exchanger.For example, the at least one valve may be or may comprise a valve configured to vary a proportion of oil directed via the fuel-oil heat exchanger, and the method may comprise, responsive to determining that the fuel temperature has increased above a predetermined threshold during cruise conditions, controlling the at least one valve to direct less oil through the fuel-oil heat exchanger.The inventors have recognized that it is important that the oil temperature remain within a desired range and that the fuel temperature does not exceed a threshold. Thus, in some implementations, the air-oil heat exchanger may be at the center of adjusting the flow rate. The at least one valve may therefore be or comprise a valve configured to vary a proportion of oil directed via the air-to-oil heat exchanger, and the method may comprise, responsive to determining that the fuel temperature has increased above a predetermined threshold during cruise conditions, controlling the at least one valve to direct more oil through the air-to-oil heat exchanger.The lack of an oil bypass at the air-oil heat exchanger may result in overcooling of the oil (e.g., solidification) under certain conditions. Therefore, a bypass pipe can be provided for both heat exchangers. In some implementations, at least a portion of the oil may be directed past both heat exchangers.The at least one valve may be or may comprise a valve arranged to control the air flow rate through the air-oil heat exchanger. The method may include directing more air through the fuel-oil heat exchanger when it is determined that the fuel temperature has increased above the set threshold during cruise conditions.The method may include controlling multiple valves, for example: in some implementations.• controlling the flow of oil through each of the fuel-oil heat exchangers and the air-oil heat exchangers separately; and / or• Controlling the air flow and the oil flow.As already described from the foregoing aspects, the inventors have recognized that the use of fuels different from the conventional kerosine-based jet fuels, such as persistent flying fuels, may result in different fuel properties, and that the parameters may be adjusted to utilize the different fuel properties. In particular, some fuels may be heated to higher temperatures than conventional fuels in one or more fuel-oil heat exchangers without significantly increasing the deposition of fuel degradation products, e.g., by coking or paint. This allows the fuel to absorb more heat, reducing the need for air cooling oil and improving engine efficiency. However, when improper fuels reach these higher temperatures, this may impair engine performance and in some cases even result in injector blockages. Thus, controls and compensations are desired to ensure that the performance of the engine is optimized for a particular fuel. The method of this seventh aspect comprises monitoring the fuel temperature to check whether the fuel temperature is too high and taking measures to reduce the heating of the fuel, if necessary. Fuel temperatures downstream of the fuel-oil heat exchanger (e.g., upon entering the combustion chamber) at cruise conditions may be defined as an average over at least 1, 2, 3, 4, or 5 minutes, and optionally over ten, twenty, or thirty minutes under steady state cruise conditions. These average temperatures do not include transient temperature spikes, which may be defined as fluctuations in the temperature of the fuel during operation, often increasing the temperature. A transient rise - for example a rise lasting only a few seconds - to a higher temperature may therefore not be sufficient to trigger a change in the oil flow.In a transient period, where available flying fuels change over time during the life of a gas turbine engine, and also depend on the geographical location of a refueling point (among other variables), it is important that the amount of heating of the fuel be determined based on the particular fuel used. As less oil is directed through the fuel-oil heat exchanger, less heat is transferred to the fuel, which in turn results in the fuel having a lower temperature upon reaching the combustion chamber. Thus, the operation of gas turbines may be adjusted to retrieve the best from a variety of fuels. The controllable oil flow valve, which regulates the flow rate of oil through the fuel-oil heat exchanger, plays a decisive role in adapting the performance of the engine.The inventors have realized that these principles can be applied to engines having branching oil circulation paths with different heat exchangers on different branches (parallel arrangements) as well as to engines having substantially linear, serial arrangements of heat exchangers in which, as an alternative to a branched main circuit, one or more bypass pipes can be used.The oil circuit system may branch such that a portion of the oil may flow along each branch, and the air-oil and fuel-oil heat exchangers may be arranged in a parallel configuration on different branches of the oil circuit system. In such implementations, the at least one valve configured such that the proportion of oil routed via the fuel-oil heat exchanger may be varied may be or may comprise a modulation valve arranged such that the proportion of oil routed via each branch may be varied.The oil circulation system may include at least one bypass pipe arranged such that a portion of the oil may bypass at least one of the fuel-oil heat exchanger and the air-oil heat exchanger. In such implementations, the at least one valve may be or comprise at least one bypass valve configured such that a portion of the oil bypasses the respective heat exchanger / controls the portion of the oil routed via the bypass pipe. In such implementations, the air-oil and fuel-oil heat exchangers may be arranged in the oil circulation system in series or in parallel. In implementations where the air-oil and fuel-oil heat exchangers are arranged in parallel on different branches of the oil circulation system and where there is additionally at least one bypass pipe, the method may comprise controlling both the bypass valve and a modulation valve which are arranged such that the proportion of oil conducted via each branch can be varied.In implementations where the method is configured to adjust the flow rate of the oil through the fuel-oil heat exchanger, at least a portion of the oil diverted from the fuel-oil heat exchanger may be directed to the air-oil heat exchanger. In some of these implementations, the air flow through the air-to-oil heat exchanger may be increased as more oil is directed across the air-to-oil heat exchanger.In implementations where the method is configured to adjust the flow rate of the oil through the fuel-oil heat exchanger or the air-oil heat exchanger, at least a portion of the oil directed away from the respective heat exchanger may be directed around the respective heat exchanger via a bypass pipe.The method may further comprise determining the set threshold based on at least one fuel property of the fuel. In such implementations, the at least one fuel property of the fuel may be or include at least one of: thermal stability of the fuel, nitrogen content of the fuel, sulfur content of the fuel, and sustainable flying fuel, SAF, content of the fuel.The step of determining the predetermined threshold may comprise increasing the predetermined threshold, optionally linearly, with increasing thermal stability of the fuel.The step of determining the fixed threshold may comprise increasing the fixed threshold, optionally linearly, as the SAF content of the fuel increases for fuels having a SAF content of over 70%.The method may further comprise determining the at least one fuel property of the fuel, optionally by any of the methods described below.The heat exchange system may include an air valve arranged to control the air flow rate through the air-oil heat exchanger. The method may further comprise controlling the air valve to direct more air through the air-to-oil heat exchanger when it is determined that the fuel temperature has increased above a set threshold during cruise conditions.The heat exchange system may include a refrigeration cycle device configured to provide thermal buoyancy by transferring additional heat from the oil to the fuel that exceeds the heat transferred from the fuel-oil heat exchanger. The method may further include controlling the refrigeration cycle device to reduce the amount of additional heat transferred to the fuel when it is determined that the fuel temperature has increased above a set threshold during cruise conditions. For example, the refrigeration cycle device may be turned off / deactivated.The heat exchange system may comprise at least one bypass pipe and the at least one valve may be or comprise a bypass valve arranged to control the flow rate through that bypass pipe. The heat exchange system may include a plurality of bypass tubes, each arranged such that oil may bypass a heat exchanger (e.g., the fuel-oil heat exchanger or air-oil heat exchanger described above, or a secondary fuel-oil heat exchanger, an oil-oil heat exchanger arranged to transfer heat between two separate oil circuits of the heat exchange system, or another suitable heat exchanger). The method may comprise controlling at least two bypass valves - for example, a bypass valve for the air-oil heat exchanger may be adapted to direct more oil to the air-oil heat exchanger when the bypass valve for the fuel-oil heat exchanger is adapted to direct less oil to the fuel-oil heat exchanger. In some implementations, the same valve (e.g., a three-way valve) may adjust the oil flow to both heat exchangers.The predetermined threshold may be in the range of 140° C. to 300° C., and optionally 250° C. to 300° C. The predetermined threshold may be 140° C., 180° C., 200° C., 210° C., 220° C., 230° C., 240° C., 250° C., 260° C., 270° C., 275° C., 280° C., 290° C., or 300° C. The step of determining may be performed at regular intervals during aircraft operation in cruise flight.The methods of the first, third, fifth, and seventh aspects may complement each other, and two or more of them may be performed together in various implementations. The method of the seventh aspect may be performed using the engine of the second, fourth or sixth aspect.According to an eighth aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising:an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor;a fan disposed upstream of the engine core;a fan shaft;at least one bearing arranged for supporting the fan shaft;an oil circuit system arranged for supplying the at least one bearing with oil;a heat exchange system comprising:an air-oil heat exchanger through which the oil in the oil circulation system flows;a fuel-oil heat exchanger through which the oil in the oil circulation system and the fuel flow so that heat is transferred between the oil and the fuel; andat least one valve configured such that at least one of the oil and air flow rates can be varied by at least one of the fuel-oil heat exchanger and the air-oil heat exchanger;a temperature sensor arranged to provide an indication of fuel temperature downstream of the fuel-oil heat exchanger (optionally at the inlet to the combustion chamber); anda controller configured to receive an output from the temperature sensor at cruise conditions, determine whether the fuel temperature has increased above a predetermined threshold based on this output, and in response to determining that the fuel temperature has increased above a predetermined threshold at cruise conditions, control the at least one valve to change the at least one flow rate through the at least one heat exchanger.For example, the at least one valve may be configured such that a portion of the oil directed via the fuel-oil heat exchanger may be varied; and the controller may be configured to:receiving an output from the temperature sensor during cruise conditions; based on this output, determining whether the temperature of the fuel has increased above a predetermined threshold; andresponsive to determining that the fuel temperature has increased above a predetermined threshold during cruise conditions, controlling the at least one valve to change the at least one flow rate through the at least one heat exchanger.The controller or other processing module may be configured to determine the set threshold based on one or more fuel characteristics of the fuel.The heat exchange system may include a refrigeration cycle device configured to provide thermal buoyancy by transferring further heat from the oil to the fuel that exceeds the heat transferred from the fuel-oil heat exchanger. Optionally, the chiller may raise the fuel temperature above the oil temperature. The controller may be configured to deactivate the refrigeration cycle device when it determines that the fuel temperature has increased above a set threshold during cruise conditions.The turbine may be a first turbine, the compressor may be a first compressor, and the core shaft may be a first core shaft. The engine core may further include a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, the second compressor, and the second core shaft may be arranged to rotate at a higher speed than the first core shaft.The heat exchange system may further comprise branching fuel return paths and at least one valve controlling a division of the fuel flow, the branching paths arranged to return fuel from the heat exchange system to at least two different locations along a main fuel path leading to the combustion chamber from where the fuel enters the gas turbine engine.The core shaft may output the drive directly to the fan to drive the fan at the same speed as the core shaft, such that the engine is a direct drive turbine engine.The air-oil and fuel-oil heat exchangers may be arranged in series in the oil circulation system.The engine of the tenth aspect may be configured to perform the method of the first, third, fifth, seventh and / or ninth aspect, and may include any of the features described with respect to any of the preceding aspects.As noted elsewhere herein, the present disclosure may be applicable to a relevant configuration of a gas turbine engine. Such a gas turbine engine may be, for example, a turbofan gas turbine engine, an open rotor gas turbine engine (in which the propeller is not surrounded by a nacelle), a turboprop engine, or a turbojet engine. Each such engine may or may not be equipped with an afterburner. Such a gas turbine engine may be configured for, for example, land or sea power generation applications.A gas turbine engine according to the present disclosure may include an engine core including a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor. Such a gas turbine engine may comprise 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 disposed downstream of the engine core, for example, when the gas turbine engine has an open rotor or is a turboprop engine (in which case the fan may be referred to as a propeller).If the gas turbine engine has an open rotor or is a turboprop engine, the gas turbine engine may comprise two contrarotating propeller stages which are connected via a shaft to a free power turbine and driven by it. The propellers may rotate in the opposite direction so that one rotates clockwise and the other rotates counter-clockwise about the axis of rotation of the engine. Alternatively, the gas turbine engine may include a propeller stage and a vane stage configured downstream of the propeller stage. The guide vane stage may have a variable pitch.Accordingly, high pressure, intermediate pressure and free turbines may drive high and intermediate pressure compressors and propellers, respectively, through suitable connecting shafts. The propellers can thus provide the majority of the drive thrust.When the gas turbine engine has an open rotor or is a turboprop engine, one or more of the propeller stages may be driven by a gearbox. The transmission may be of the type described herein.An engine according to the present disclosure may be a turbo fan engine. Such an engine may be a direct drive turbo fan engine in which the fan is directly connected to the fan drive turbine via a core shaft, for example without a gearbox. In such a direct drive turbo fan engine, the fan may be arranged to rotate at the same speed as the fan drive turbine. By way of example only, the fan power turbine may be a first turbine, the core shaft may be a first core shaft, and the gas turbine engine may further include a second turbine and a second core shaft connecting the second turbine to the compressor. The second turbine, compressor, and second core shaft may be arranged to rotate at a higher speed than the first core shaft. In such an arrangement, the second turbine may be positioned axially upstream of the first turbine.An engine according to the present disclosure may be a gear fan engine. In such an arrangement, the engine includes a fan driven via a gearbox. Accordingly, such a gas turbine engine may include a gearbox that receives input from the core shaft and outputs output to the fan such that the fan is driven at a lower speed than the core shaft. The input to the transmission can be effected directly from the core shaft or indirectly from the core shaft, for example via a spur gear shaft and / or a gearwheel. The core shaft may rigidly connect the turbine and the compressor such that the turbine and compressor rotate at the same speed (with the fan rotating at a lower speed).The gas turbine engine as described and / or claimed herein may have any suitable general architecture. For example, the gas turbine engine may include any number of shafts connecting turbines and compressors, for example, one, two, or three shafts. For example only, the turbine connected to the core shaft may be a first turbine, the compressor connected to the core shaft may be a first compressor, and the core shaft may be a first core shaft. The engine core may further include a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, the second compressor, and the second core shaft may be arranged to rotate at a higher speed than the first core shaft.In such an arrangement, the second compressor may be positioned axially downstream of the first compressor. The second compressor may be arranged to receive (e.g., directly receive) a flow from the first compressor, e.g., via a generally annular channel).The transmission may be arranged to be driven by the core shaft configured to rotate (e.g., in use) at the lowest speed (e.g., by the first core shaft in the above example). For example, the transmission may be arranged to be driven only by the core shaft configured to rotate (e.g., in use) at the lowest speed (e.g., only by the first core shaft and not by the second core shaft in the above example). Alternatively, the transmission may be arranged to be driven by any one or more shafts, for example the first and / or second shaft in the above example.The gearbox may be a reduction gearbox (in that the output to the fan has a lower speed than the input from the core shaft). Any type of transmission may be used. For example, the transmission may be a "planetary gear" or a "star gear" as described in more detail elsewhere herein. Such a transmission can be a single-stage transmission. Alternatively, such a transmission may be a compound transmission, for example a compound planetary transmission (which may have the input at the sun gear and the output at the ring gear and thus may be referred to as a "compound star" transmission) having, for example, two reduction stages.The transmission may have any desired reduction ratio (defined as the speed of the input shaft divided by the speed of the output shaft), for example greater than 2.5, for example in the range of 3 to 4.2 or 3.2 to 3.8, for example on the order of or at least 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. the transmission ratio may be between any two of the values mentioned in the previous sentence, for example. By way of example only, the transmission may be a "star" transmission having a gear ratio in the range of 3.1 or 3.2 to 3.8. Further, by way of example only, the transmission may be a "star" transmission having a gear ratio in the range of 3.0 to 3.1. Further, by way of example only, the transmission may be a "planetary gear" having a gear ratio in the range of 3.6 to 4.2. In some arrangements, the gear ratio may be outside of these ranges.In any gas turbine engine described and / or claimed herein, fuel of a given composition or mixture is provided to a combustor, which may be provided downstream of the fan and compressor(s) with respect to the flowpath (e.g., axially downstream). For example, the combustor may be located directly downstream of (e.g., at the exit of) the second compressor where a second compressor is provided. As another example, the flow at the exit to the combustor may be provided to the inlet of the second turbine where a second turbine is provided. The combustor may be provided upstream of the turbine(s).The or each compressor (e.g., the first compressor and the second compressor as described above) may comprise any number of stages, e.g., multiple stages. Each stage may include a row of rotor blades and a row of stator blades, which may be variable stator blades (as their angle of incidence may be variable). The row of rotor blades and the row of stator blades may be axially offset from each other. For example, the gas turbine engine may be a direct drive turbo fan engine that includes 13 or 14 compressor stages (in addition to the fan). Such an engine may, for example, comprise 3 stages in the first (or "low pressure") compressor and either 10 or 11 stages in the second (or "high pressure") compressor. As another example, the gas turbine engine may be a "gearbox" gas turbine engine (in which the fan is driven by a first core shaft via a reduction gearbox) that includes 11, 12 or 13 compressor stages (in addition to the fan). Such an engine may comprise 3 or 4 stages in the first (or "low pressure") compressor and 8 or 9 stages in the second (or "high pressure") compressor. As another example, the gas turbine engine may be a "gearbox" gas turbine engine having 4 stages in the first (or "low pressure") compressor and 10 stages in the second (or "high pressure") compressor.The or each turbine (e.g., the first turbine and the second turbine as described above) may comprise any number of stages, e.g., multiple stages. Each stage may comprise a row of rotor blades and a row of stator blades or vice versa, as required. The respective rows of rotor blades and stator blades may be axially offset from each other. The second (or "high pressure") turbine may include 2 stages in any arrangement (e.g., whether it be a transmission or direct drive engine). The gas turbine engine may be a direct drive turbo fan engine that includes a first (or "low pressure") 5, 6 or 7 stage turbine. Alternatively, the gas turbine engine may be a "gearbox" gas turbine engine having a first (or "low pressure") 3 or 4 stage turbine.Each fan blade may be defined as having a radial span extending from a root (or hub) at a radially inner gas washed location or 0% span position to a tip at a 100% span position. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be less than (or on the order of) 0.4, 0.39, 0.38, 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25 The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be in an inclusive range bounded by any two of the values in the previous set (i.e., the values may form upper or lower limits), for example, in the range of 0.28 to 0.32, or 0.29 to 0.30 These ratios may be referred to generally as the hub-to-tip ratio. The radius at the hub and the radius at the tip may both be measured at the leading edge (or the axially foremost part) of the blade. The hub-to-tip ratio refers, of course, to the gas washed portion of the fan blade, i.e., the portion radially outward of any platform.The radius of the fan may be measured between the centerline of the engine and the tip of a fan blade at its leading edge. The fan diameter (which may be simply twice the fan radius) may be greater than (or on the order of) 140 cm, 170 cm, 180 cm, 190 cm, 200 cm, 210 cm, 220 cm, 230 cm, 240 cm, 250 cm (about 100 inches), 260 cm, 270 cm (about 105 inches), 280 cm (about 110 inches), 290 cm (about 115 inches), 300 cm (about 120 inches), 310 cm, 320 cm (about 125 inches), 330 cm (about 130 inches), 340 cm (about 135 inches), 350 cm, 360 cm (about 140 inches), 370 cm (about 145 inches), 380 cm (about 150 inches), 390 cm (about 155 inches), 400 cm, 410 cm (about 160 inches), or 420 cm (about 165 inches). The fan diameter may be in an inclusive range bounded by any two of the values in the previous set (i.e., the values may form upper or lower bounds), 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 may range from 170 cm to 180 cm, 190 cm to 200 cm, 200 cm to 210 cm, 210 cm to 230 cm, 290 cm to 300 cm, or 340 cm to 360 cm.The speed of the fan may vary in use. Generally, the speed is lower for larger diameter fans. Purely as a non-limiting example, the speed of the fan may be less than 3500 U / min, for example less than 2600 U / min, or less than 2500 U / min, or less than 2300 U / min, under cruise conditions. By way of further non-limiting example, the speed of the fan may be in the range of 2750 to 2900 U / min under cruise conditions for a "gearbox" gas turbine engine having a fan diameter in the range of 200 cm to 210 cm. By way of further non-limiting example, the speed of the fan may be in the range of 2500 to 2800 U / min under cruise conditions for a "gearbox" gas turbine engine having a fan diameter in the range of 210 cm to 230 cm. By way of further non-limiting example, the speed of the fan may be in the range of 1500 to 1800 U / min under cruise conditions for a "gearbox" gas turbine engine having a fan diameter in the range of 340 cm to 360 cm. By way of further non-limiting example, the speed of the fan may be in the range of 3600 to 3900 U / min under cruise conditions for a direct drive engine having a fan diameter in the range of 190 cm to 200 cm. By way of further non-limiting example, the speed of the fan may be in the range of 2000 to 2800 U / min under cruise conditions for a direct drive engine having a fan diameter in the range of 300 cm to 340 cm.In use of the gas turbine engine, the fan (with associated fan blades) rotates about an axis of rotation. This rotation causes the tip of the fan blade to move Spitze at a speed U. The work performed by the fan blades in the flow results in an enthalpy increase dH of the flow. A fan tip load may be defined as dH / U Spitze2 where dH is the enthalpy increase (e.g., the average 1-D enthalpy increase) across the fan and U Spitze is the (translatory) speed of the fan tip, e.g., at the leading edge of the tip (which may be defined as the fan tip radius at the leading edge multiplied by the angular speed). The fan peak load under cruise conditions may be greater than (or on the order of) 0.28, 0.29, 0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39, or 0.4 (all values being dimensionless). The fan tip load may be in an inclusive range bounded by any two of the values in the previous set (i.e., the values may form upper or lower limits), for example, in the range of 0.28 to 0.31 or 0.29 to 0.3 (for example, for a transmission gas turbine engine).Gas turbine engines according to the present disclosure may have a desired bypass ratio (BPR), where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core. In some arrangements, the bypass ratio under cruise conditions may be greater than (or on the order of) one of: 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. By way of further non-limiting example, the bypass ratio of a gas turbine engine according to the present disclosure may be in the range of 12:1 to 15:1 under cruise conditions. The bypass channel may be at least substantially annular. The bypass duct may be located radially outside of the core engine. The radially outer surface of the bypass duct may be defined by a nacelle and / or a fan casing.The total pressure ratio (OPR) of a gas turbine engine as described and / or claimed herein may be defined as the ratio of the ram pressure at the exit of the super compressor (prior to entry into the combustor) to the ram pressure upstream of the fan. By way of non-limiting example, the total pressure ratio of a gas turbine engine as described and / or claimed herein may be greater than (or on the order of) one of the following under cruise conditions): 35, 40, 45, 50, 55, 60, 65, 70, 75. The total pressure ratio may be in an inclusive range bounded by any two of the values in the preceding set (i.e., the values may form upper or lower limits), for example, in the range of 50 to 70. Purely by way of non-limiting example, the total pressure ratio under cruise conditions of a transmission 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 total pressure ratio may be in the range of 45-55 during cruise conditions of a gas turbine engine having a fan diameter in the range of 210 cm to 230 cm. By way of non-limiting example only, the total pressure ratio may be in the range of 50 to 60 under cruise conditions of a gas turbine engine having a fan diameter in the range of 340 cm to 360 cm. By way of non-limiting example only, the total pressure ratio may be in the range of 50 to 60 under cruise conditions of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm.The specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow through the engine. In some examples, the specific thrust for a given state of thrust may depend on the specific composition of fuel provided to the combustion chamber. Under cruise conditions, the specific thrust of an engine described and / or claimed herein may be less than (or on the order of) 110 Nkg -1 s, 105 Nkg -1 s, 100 Nkg -1 s, 95 Nkg -1 s, 90 Nkg -1 s, 85 Nkg -1 s, or 80 Nkg -1 s. The specific thrust may be in an inclusive range bounded by any two of the values in the previous set (i.e., the values may form upper or lower bounds), for example, in the range of 80 Nkg -1 s to 100 Nkg -1 s or 85 Nkg -1 s to 95 Nkg -1 s. Such engines may be particularly efficient compared to conventional gas turbine engines. By way of non-limiting example only, the specific thrust of a gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 90 Nkg -1 s to 95 Nkg -1 s. By way of non-limiting example only, the specific thrust of a gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 80 Nkg -1 s to 90 Nkg -1 s. By way of non-limiting example only, the specific thrust of a gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 70 Nkg -1 s to 90 Nkg -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 may be in the range of 90 Nkg -1 s to 120 Nkg -1 s.A gas turbine engine as described and / or claimed herein may have any desired maximum thrust. By way of non-limiting example only, a gas turbine as described and / or claimed herein may be capable of generating a maximum thrust of at least one of (or on the order of) 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 in an inclusive range bounded by any two of the values in the previous set (i.e., the values may form upper or lower bounds). By way of non-limiting example only, a gas turbine as described and / or claimed herein may be capable of generating a maximum thrust in the range of 155 kN to 170 kN, 330 kN to 420 kN, or 350 kN to 400 kN. By way of non-limiting example only, the maximum thrust of a gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 140 kN to 160 kN. By way of non-limiting example only, the maximum thrust of a gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 150 kN to 200 kN. By way of non-limiting example only, the maximum thrust of a gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 370 kN to 500 kN. By way of non-limiting example only, the maximum thrust of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 370 kN to 500 kN. The above thrust may be the maximum net thrust at normal sea level atmospheric conditions at +15 degrees C. (ambient pressure 101.3 kPa, temperature 30 degrees C.) for a static engine.In use, the temperature of the stream at the inlet to the high pressure turbine may be particularly high. This temperature, which may be referred to as TET, may be measured at the exit to the combustor, for example immediately upstream of the first turbine vane, which may itself be referred to as a nozzle vane. In some examples, for a given state of thrust, the TET may depend on the specific composition of fuel provided to the combustion chamber. Under cruise conditions, the TET may be at least one of (or on the order of): 1400 K, 1450 K, 1500 K, 1520 K, 1530 K, 1540 K, 1550 K, 1600 K, or 1650 K. Thus, purely by way of non-limiting example, under cruise conditions of a gas turbine engine having a fan diameter in the range of 200 cm to 210 cm, the TET may be in the range of 1540 K to 1600 K. By way of non-limiting example only, under cruise conditions of a gas turbine engine having a fan diameter in the range of 210 cm to 230 cm, the TET may be in the range of 1590 K to 1650 K. By way of non-limiting example only, under cruise conditions of a gas turbine engine having a fan diameter in the range of 340 cm to 360 cm, the TET may be in the range of 1600 K to 1660 K. By way of non-limiting example only, the TET of a direct drive gas turbine engine having a fan diameter under cruise conditions may be in the range of 300 cm to 340 cm in the range of 1590 K to 1650 K. By way of non-limiting example only, under cruise conditions of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm, the TET may be in the range of 1570 K to 1630 K.The TET under cruise conditions may be in an inclusive range bounded by any two of the values in the previous set (i.e., the values may form upper or lower limits), for example 1530 K to 1600 K. The maximum TET in use of the engine may be, for example, at least one of (or on the order of): 1700 K, 1750 K, 1800 K, 1850 K, 1900 K, 1950 K, 2000 K, 2050 K, or 2100 K. Thus, purely by way of non-limiting example, the maximum TET of a gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 1890 K to 1960 K. By way of non-limiting example only, the maximum TET of a gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 1890 K to 1960 K. By way of non-limiting example only, the maximum TET of a gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 1890 K to 1960 K. 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 may be in the range of 1935 K to 1995 K. By way of non-limiting example only, the maximum TET of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 1890 K to 1950 K. The maximum TET may be in an inclusive range bounded by any two of the values in the previous set (i.e., the values may form upper or lower limits), for example, in the range of 1800 K to 1950 K or 1900 K to 2000 K. The maximum TET may occur, for example, at a high thrust condition, for example, at a maximum lift-off (MTO) condition.A fan blade and / or airfoil portion of a fan blade as described and / or claimed herein may be made of any suitable material or combination of materials. For example, at least a portion of the fan blade and / or the air baffle may be at least partially made of a composite material, for example a metal matrix composite material and / or an organic matrix composite material, such as a carbon fiber composite material. As another example, at least a portion of the fan blade and / or air baffle may be made at least partially from a metal, such as a titanium-based metal or an aluminum-based material (such as an aluminum-lithium alloy) or a steel-based material. The fan blade may include at least two regions made using different materials. For example, the fan blade may have a protective leading edge that may be made using a material that is better able to resist impact (e.g., birds, ice, or other material) than the remainder of the blade. Such a leading edge may be made using titanium or a titanium-based alloy, for example. Thus, for example only, the fan blade may include a carbon fiber or aluminum-based body (such as an aluminum-lithium alloy) having a leading edge of titanium.A fan as described and / or claimed herein may comprise a central portion from which the fan blades may extend, for example in a radial direction. The fan blades may be attached to the central portion in any desired manner. For example, each fan blade may include a fastener that may engage a corresponding slot in the hub (or disc). Such a fastening may be in the form of a dovetail, purely by way of example, which may be inserted and / or snap-fitted into a corresponding slot in the hub / disk in order to fasten the fan blade to the hub / disk. As another example, the fan blades may be integrally formed with a central portion. Such an arrangement may be referred to as a blade disk or blade ring. Any suitable method may be used to produce such a blade disc or ring. For example, at least a portion of the fan blades may be made from a block and / or at least a portion of the fan blades may be attached to the hub / disk by welding, such as linear friction welding.The gas turbine engines described and / or claimed herein may or may not be provided with a variable area nozzle (VAN). Such a variable area nozzle may allow for varying the exit area of the bypass passage in use. The general principles of the present disclosure may be applied to engines with or without VAN.The fan of a gas turbine as described and / or claimed herein may include any number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26 fan blades. If the fan blades comprise a carbon fibre composite body, 16 or 18 fan blades may be present. If the fan blades comprise a metallic body (for example an aluminium-lithium or titanium alloy), 18, 20 or 22 fan blades may be present.As used herein, the terms coasting, rolling, takeoff, climbing, cruise, descent, approach, and landing (or one or more portions thereof) have the usual meaning and are readily understood by those skilled in the art. Thus, for a particular gas turbine engine for an aircraft, those skilled in the art would readily appreciate that each term refers to all or part of an operating phase of the engine within a given deployment of an aircraft for which the gas turbine engine was designed for mounting thereto.In this regard, ground idle may refer to an operating phase of the engine in which the aircraft is stationary and is in contact with the ground, but there is a requirement that the engine be running. During idle, the engine may generate between 3% and 9% of the available thrust of the engine. In further non-limiting examples, the engine may generate between 5% and 8% of available thrust. In further non-limiting examples, the engine may generate between 6% and 7% of available thrust. Coasting may refer to an operating phase of the engine in which the aircraft is propelled along the ground by the thrust generated by the engine. During coasting, the engine may generate between 5% and 15% of the available thrust. In further non-limiting examples, the engine may generate between 6% and 12% of the available thrust. In further non-limiting examples, the engine may generate between 7% and 10% of the available thrust. Takeoff may refer to an operating phase of the engine in which the aircraft is propelled by the thrust generated by the engine. At an initial stage within the launch phase, the aircraft may be propelled while the aircraft is in contact with the ground. At a later stage within the launch phase, the aircraft may be propelled while the aircraft is not in contact with the ground. During launch, the engine may generate between 90% and 100% of the available thrust. In further non-limiting examples, the engine may generate between 95% and 100% of the available thrust. In further non-limiting examples, the engine may generate 100% of the available thrust.Climbing flight may refer to an operating phase of the engine in which the aircraft is driven by the thrust generated by the engine. During climbing flight, the engine may generate between 75% and 100% of the available thrust. In further non-limiting examples, the engine may generate between 80% and 95% of the available thrust. In further non-limiting examples, the engine may generate between 85% and 90% of the available thrust. In this context, the ascent may refer to an operating phase within the flight cycle of an aircraft between the start and the attainment of the cruise conditions, wherein the attainment of the cruise conditions defines the beginning of the cruise phase or a portion thereof of the aircraft flight. Additionally or alternatively, the term climbing flight may refer to a nominal point or time periods during the flight cycle of an aircraft between takeoff and landing where a relative elevation is required that may require additional propulsion demands of the engine.As used herein, cruise conditions that may define the cruise phase (or a portion thereof) of aircraft flight are of conventional significance and will be readily understood by those skilled in the art. In some examples, the cruise conditions for a particular gas turbine engine for an aircraft may refer to the operating point of the engine at cruise of a given mission (which may be referred to in the industry as "economic mission") of an aircraft for which the gas turbine engine has been designed for attachment thereto. In this context, cruise may be considered to be the point in the flight cycle of an aircraft where 50% of the total fuel burned between the highest point of climb and the beginning of descent has burned (which may be approximated by the midpoint - in terms of time and / or distance - between the highest point of climb and the beginning of descent). The cruise conditions may thus define an operating point, phase, or portion thereof of the flight that provides thrust that provides stable operation (i.e., maintaining a constant altitude and / or Mach number) or at least substantially stable operation (i.e., maintaining an at least substantially constant altitude and / or Mach number) of an aircraft for which it was designed for attachment thereto, taking into account the number of engines provided for that aircraft. For example, if an engine is configured to be mounted on an aircraft having two engines of the same type, the engine may provide, under cruise conditions, half of the total thrust that would be required for stable operation or at least substantially stable operation of the aircraft during cruise.In other words, the cruise conditions for a given gas turbine engine for an aircraft may be defined as the operating point of the engine that provides a particular thrust (required to provide, in combination with any other engines on the aircraft, stable operation or at least substantially stable operation of the aircraft for which it has been designed to be mounted thereto at a given Mach number during cruise) under cruise atmospheric conditions (defined by the international standard atmosphere according to ISO 2533 at altitude during cruise). For any given gas turbine engine for an aircraft, the thrust at cruise, atmospheric conditions and Mach number are known and thus the operating point of the engine under cruise conditions can be clearly defined.Purely by way of example, the forward speed under cruise conditions may be any point in the range of Mach 0.7 to 0.9, for example 0.75 to 0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example 0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach 0.8, on the order of Mach 0.85, or in the range of 0.8 to 0.85. Any single speed within these ranges may be part of the cruise conditions. In some aircraft, cruise conditions may be outside these ranges, for example below Mach 0.7 or above Mach 0.9.By way of example only, the cruise conditions may correspond to standard atmospheric conditions (according to the International Standard Atmosphere, ISA) at a height that is in the range of 10,000 m to 15,000 m, for example in the range of 10,000 m to 12,000 m, for example in the range of 10,400 m to 11,600 m (about 38,000 feet), for example in the range of 10,500 m to 11,500 m, for example in the range of 10,600 m to 11,400 m, for example in the range of 10,700 m (about 35,000 feet) to 11,300 m, for example in the range of 10,800 m to 11,200 m, for example in the range of 10,900 m to 11,100 m, for example in the order of 11,000 m. The cruise conditions may correspond to standard atmospheric conditions at any given altitude in these ranges.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 a level of 35,000 feet (10,668 m). Under such cruise conditions, the engine may provide a known net thrust level required. The known net thrust level required is, of course, dependent on the engine and its intended use and may be, for example, a value in the range of 20 kN to 40 kN.By way of example only, cruise conditions may correspond to a forward Mach number of 0.85 and standard atmospheric conditions (according to the International Standard Atmosphere) at a altitude of 38,000 feet (11,582 m). Under such cruise conditions, the engine may provide a known net thrust level required. The known net thrust level required is, of course, dependent on the engine and its intended use and may be, for example, a value in the range of 35 kN to 65 kN.In use, a gas turbine engine as described and / or claimed herein may operate under the cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions (e.g., the conditions during cruise) of an aircraft to which at least one (e.g., 2 or 4) gas turbine engine may be mounted to provide propulsion thrust.Furthermore, the skilled person would immediately recognize that descent and approach, or both, relate to an operating phase within a flight cycle between cruise and landing of the aircraft, wherein in particular the approach forms part of the landing and take-off phase (LTO phase). During descent or landing approach, or both, the engine may generate between 0% and 50% of the available thrust. In further non-limiting examples, the engine may generate between 25% and 40% of the available thrust. In further non-limiting examples, the engine may generate between 30% and 35% of the available thrust. Additionally or alternatively, descent may refer to a nominal point in an aircraft flight cycle between takeoff and landing where a relative decrease in altitude is required and which may require a reduced propulsion demand of the engine.According to one aspect, there is provided an aircraft comprising a gas turbine engine as described and / or claimed herein. The aircraft of this aspect is the aircraft for which the gas turbine engine has been designed for mounting thereto. Accordingly, in this aspect, cruise conditions may correspond to an operating point, phase, or portion thereof of aircraft flight, as defined elsewhere herein.According to one aspect, there is provided a method of operating a gas turbine engine as described and / or claimed herein. Operation may be under suitable conditions as defined elsewhere herein (e.g., in terms of thrust, atmospheric conditions, and Mach number).According to one aspect, there is provided a method of operating an aircraft comprising a gas turbine engine as described and / or claimed herein. Operation in accordance with this aspect may include (or be) operation under suitable conditions, for example, at the center of flight of the aircraft as defined elsewhere herein.The skilled person will understand that, except in the case of mutual exclusion, a feature or parameter described with respect to one of the above-mentioned aspects can be applied to any other aspect. Moreover, except in the case of mutual exclusion, any feature or parameter included or described herein may be applied to any aspect and / or combined with any other feature or parameter included or described herein.Unless mutually exclusive, any parameter or value included or described herein may be applied to and / or combined with one or more other parameters and / or values included or described herein. For example, a first parameter or value (e.g., parameter A) included or described herein may be applied to and / or combined with one or more other parameters and / or values included or described herein (e.g., one or more parameters B; parameter C; and parameter D, etc.) to express a product of their relationship. For example, those skilled in the art would understand that in cases where parameter A is disclosed separately from parameter B, a product of their relationship may be expressed as, for example, A / B, B / A, B*A, or any other such application, combination, or function of parameter A with respect to parameter B, as desired.Embodiments will now be described, by way of example only, with reference to the figures, in which: FIG. 1 is a cross-sectional side view of a gas turbine engine; FIG. 2 is a cross-sectional side view of an upstream portion of a close-up gas turbine engine; FIG. 3 is a partial cross-sectional view of a transmission for a gas turbine engine; FIG. 4 is a cross-sectional side view of an upstream portion of a close-up direct drive gas turbine engine; FIG. 5 is an illustration of an aircraft having a propulsion system including two gas turbine engines; FIG. 6 is a schematic illustration of an example fuel system; FIG. 7A is a schematic illustration of an alternative exemplary fuel system including a recirculation circuit; FIG. 7B is a schematic illustration of another alternative exemplary fuel system that includes a recirculation loop and the use of fuel for auxiliary systems; FIG. 7C is a schematic illustration of another alternative exemplary fuel system similar to that shown in FIG. 7B ,However, in this case, the fuel flowing through the secondary heat exchanger,optionally being recirculated to the combustion chamber after use in one or more auxiliary systems; FIG. 8A is a schematic illustration of a portion of an example oil circulation system; FIG. 8B is a schematic illustration of a portion of an example oil circulation system; FIG. 8C is a schematic illustration of a portion of an example oil circulation system; FIG. 9 is a schematic illustration of a portion of an example oil circulation system; FIG. 10 is a schematic illustration of a portion of the example fuel system of FIG. 6 and the example oil circulation system of FIG. 8A ; FIG. 11 is a schematic illustration of a portion of the example fuel system of FIG. 7A and the example oil recirculation system of FIG. 8A ; FIG. 12 is a schematic illustration of a portion of the example fuel system of FIG. 7A and the example oil recirculation system of FIG. 8C ; FIG. 13 illustrates an example method of operating a gas turbine engine; FIG. 14 illustrates another example method of operating a gas turbine engine; FIG. 15 illustrates another example method of operating a gas turbine engine; FIG. 16 illustrates another example method of operating a gas turbine engine; FIG. 17 illustrates another example method of operating a gas turbine engine; FIG. 18 illustrates a plot of cruise heat transfer areas for a direct drive gas turbine engine.FIG. 1 illustrates a gas turbine engine 10 having a main axis of rotation 9. the engine 10 includes an air inlet 12 and a propulsion fan 23 that produces two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 includes a core 11 that receives the core airflow A. The engine core 11 comprises, in axial flow sequence, a low pressure compressor 14, a high pressure compressor 15, combustion equipment 16, a high pressure turbine 17, a low pressure turbine 19 and a core exhaust nozzle 20, a nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18, the bypass airflow B flows through the bypass duct 22, the fan 23 is mounted on and driven by the low pressure turbine 19 via a shaft 26 and an epicyclic gearing 30.In use, the core air stream A is accelerated and compressed by the low pressure compressor 14 and directed into the high pressure compressor 15 where it is further compressed. The compressed air discharged from the high pressure compressor 15 is introduced into the combustion equipment 16 where it is mixed with fuel F and the mixture is burned. Combustion equipment 16 may be referred to as combustor 16, with the terms "combustion equipment 16" and "combustor 16" being used interchangeably herein. The resulting hot combustion products then expand through and thereby drive the high pressure and low pressure turbines 17, 19 before being discharged through the nozzle 20 to provide some propulsion thrust. The high-pressure turbine 17 drives the high-pressure compressor 15 via a suitable connecting shaft 27. The fan 23 generally acts to pressurize the bypass airflow B flowing through the bypass duct 22 such that the bypass airflow B is expelled through the bypass outlet nozzle 18 to generally provide the majority of the motive thrust. The planetary gear 30 is a reduction gear.An exemplary arrangement for a gear fan gas turbine engine 10 is shown in FIG. 2. The low pressure turbine 19 (see FIG. 1 ) drives the shaft 26 which is coupled to a central pinion or sun gear 28 of the epicyclic gear assembly 30. Radially outwardly of and engaging sun gear 28 is a plurality of planet gears 32 which are coupled together by a planet carrier 34. The planet carrier 34 restricts the planet gears 32 to orbit synchronously about the sun gear 28 while enabling each planet gear 32 to rotate about its own axis. Planet carrier 34 is coupled via linkages 36 to fan 23 to drive rotation thereof about engine axis 9. Radially outwardly of and engaging the planet gears 32 is a ring gear 38 which is coupled via linkages 40 to a stationary support structure 24.It should be noted that the terms "low pressure turbine" and "low pressure compressor" as used herein mean the lowest pressure turbine stages and the lowest pressure compressor stages, respectively (i.e., not including the fan 23) and / or the turbine and compressor stages connected by the lowest speed in the engine connecting shaft 26 (i.e., not including the transmission output shaft driving the fan 23). In some references, the "low pressure turbine" and "low pressure compressor" referred to herein may alternatively be known as the "intermediate pressure turbine" and "intermediate pressure compressor.". Where such alternative nomenclature is used, the fan 23 may be referred to as a first or lowest compression stage.The planetary gear 30 is shown in more detail in FIG. 3 by way of example. Each of the sun gear 28, the planet gears 32, and the ring gear 38 includes teeth on its periphery to engage the other gears. However, for clarity, only exemplary portions of the teeth are illustrated in FIG. 3. Four planet gears 32 are illustrated, although it will be apparent to those skilled in the art that more or fewer planet gears 32 may be provided within the scope of the claimed invention. Practical applications of planetary epicyclic gearing 30 generally include at least three planetary gears 32.The planetary gear 30 exemplarily illustrated in FIGS. 2 and 3 is of the planetary type because the planetary carrier 34 is coupled to an output shaft via linkages 36 with the ring gear 38 fixed. However, any other suitable type of epicyclic gear train 30 may be used. As another example, the epicyclic gear 30 may be a star arrangement in which the planet carrier 34 is held fixed and the ring gear (or ring gear) 38 may rotate. In such an arrangement, the fan 23 is driven by the ring gear 38. As another alternative example, the transmission 30 may be a differential transmission in which both the ring gear 38 and the planet carrier 34 may rotate.It should be understood that the arrangement shown in FIGS. 2 and 3 is exemplary only and that various alternatives are within the scope of the present disclosure. By way of example only, any suitable arrangement may be used for housing the transmission 30 in the engine 10 and / or for connecting the transmission 30 to the engine 10. As another example, the connections (such as linkage 36, 40 in the example of FIG. 2 ) between transmission 30 and other parts of engine 10 (such as driveshaft 26, output shaft, and fixed structure 24) may have any desired degree of stiffness or flexibility. As another example, any suitable arrangement of bearings between rotating and stationary parts of the engine (e.g., between the drive and output shafts of the transmission and the stationary structures such as the transmission case) may be used, and the disclosure is not limited to the example arrangement of FIG. 2. For example, if the transmission 30 has a star configuration (described above), one skilled in the art would readily understand that the configuration of the takeoff and support linkages and bearing positions would typically vary from the configuration shown in FIG. 2 by way of example.Accordingly, the present disclosure extends to a gas turbine engine having any arrangement of transmission types (e.g., star or planetary gear), support structures, drive and output shaft arrangements, and bearing positions.Optionally, the transmission may drive additional and / or alternative components (e.g., the intermediate pressure compressor and / or a booster compressor).Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may include an alternative number of compressors and / or turbines and / or an alternative number of connection shafts. As another example, the gas turbine engine shown in FIG. 1 has a split flow nozzle 18, 20, meaning that the flow through the bypass duct 22 has its own nozzle 18 separate from and located radially outside the core engine nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also be applied to engines in which the flow through the bypass passage 22 and the flow through the core 11 are mixed or combined before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area.As another example, other gas turbine engines to which the present disclosure may be applied may not include a gearbox for the mainshaft(s) and instead may be direct drive engines. A cross-sectional view of such an engine is shown in FIG. 3B.Referring to FIG. 3B, a gas turbine engine is generally indicated at 10 with a main axis of rotation 9. The engine 10 includes, in axial flow order, an air inlet 12, a drive 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.In use, air entering the inlet 12 is accelerated by the fan 23 to produce two air streams: a core air stream A and a bypass air stream B. The core air stream A flows into the intermediate pressure compressor 14, and the bypass air stream B passes through a bypass duct 22 to provide motive thrust. The intermediate pressure compressor 14 compresses the air stream A before delivering the air to the high pressure compressor 15 where further compression takes place.The compressed air discharged from the high pressure compressor 15 is introduced into the combustion equipment 16 where it is mixed with fuel F and the mixture is burned. Combustion equipment 16 may be referred to as combustor 16, with the terms "combustion equipment 16" and "combustor 16" being used interchangeably herein. The resulting hot combustion products then expand through and thereby drive the high pressure, intermediate pressure and low pressure turbines 17, 19a, 19 before being discharged through the nozzle 20 to provide additional drive thrust. The high-pressure turbine 17, the intermediate-pressure turbine 19 aand the low-pressure turbine 19 each drive the high-pressure compressor 15, the intermediate-pressure compressor 14 and the fan 23 by a suitable connecting shaft.Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. As an example, such engines may include an alternative number of connection shafts (e.g., two) and / or an alternative number of compressors and / or turbines. Further, the engine may include a gearbox provided in the powertrain from a turbine to a compressor and / or a fan.For example, while the described example relates to a turbo fan engine, the disclosure may be applied to any type of gas turbine engine, such as an open rotor (where the fan stage is not surrounded by a nacelle) or a turboprop engine. In some arrangements, the gas turbine engine 10 may not include a gearbox 30.The geometry of the gas turbine engine 10 and components thereof is defined by a conventional axis system that includes an axial direction (aligned with the axis of rotation 9), a radial direction (in the bottom-up direction in FIG. 1 ), and a circumferential direction (perpendicular to the side in the view of FIG. 1 ). The axial, radial and circumferential directions are perpendicular to each other.The fuel F provided to the combustion equipment 16 may include a fossil-based hydrocarbon fuel such as kerosine. Thus, fuel F may comprise molecules from one or more of the chemical families of n-alkanes, isoalkanes, cycloalkanes, and aromatics. Additionally or alternatively, the fuel F, when mixed with or replaced by an alternative fuel, may comprise renewable hydrocarbons made from biological or non-biological resources, also known as sustainable aviation fuel (SAF). In any of the examples provided, the fuel F may include one or more trace elements including, for example, sulfur, nitrogen, oxygen, inorganics, and metals.For example, SAF is understood by those skilled in the art to mean a bio-fuel, a renewable flying fuel, a renewable nozzle fuel, an alternative fuel or a biojet fuel made from biological or non-biological resources. SAF is typically synthesized from carbonaceous gases taken from the atmosphere and / or separated in industrial processes; or from a wide range of sustainable raw materials such as waste oil and fats; solid boiling wastes; cellulosic wastes (such as corn stalks); crop crops such as flax yolk, carinate and light herb; non-biogenic alternative fuels; jatropha; halophytes and algae instead of fossil hydrocarbons. SAF is to be understood as not comprising fossil fuels.The functional performance of a particular fuel composition or mixture F for use in a particular mission may be defined, at least in part, by the fuel's ability to operate the Brayton cycle of the gas turbine engine 10. The parameters defining the functional performance may include, for example, specific energy; energy density; thermal stability; and emissions including gaseous and / or particulate matter. In this context, the particulate emissions may include soot particles that arise during combustion of the fuel F and are also referred to as non-volatile particulates (nvPM). All references herein to soot or smoke apply equally to other types of particulate emissions known in the art. Gaseous emissions may include one or more of the following: oxides of nitrogen (NOx), carbon monoxide (CO), carbon dioxide (CO 2), unburned hydrocarbons (UHC), oxides of sulfur (SO) including, for example, sulfur dioxide (SO 2) and / or sulfur trioxide (SO 3), and volatile organic compounds (VOC) produced during combustion of fuel F. All references herein to gaseous emissions apply equally to other types of gaseous emissions known in the art.A relatively higher specific energy (i.e., energy per unit mass) expressed as MJ / kg may at least partially reduce the starting weight, thereby potentially providing a relative improvement in fuel efficiency. A relatively higher energy density (i.e., energy per unit volume) expressed as MJ / 1 may at least partially reduce the starting fuel volume, which may be particularly important for volume limited missions or military missions with refueling. Relatively higher thermal stability (i.e., inhibiting degradation or coking of the fuel under thermal load) may allow the fuel to maintain elevated temperatures in the engine and fuel injectors, thereby potentially providing relative improvements in combustion efficiency. Reduced emissions, including particulate matter, may allow for reduced condensate streaking while reducing the environmental impact of a given mission. Other properties of the fuel may also be central to functional performance. For example, a relatively lower freezing point (°C) may allow remote deployments to optimize flight profiles; minimum aromatic concentrations (%) may ensure sufficient swelling of certain materials used in the construction of O-rings and gaskets previously exposed to fuels with high aromatic ingredients; and maximum surface tension (mN / m) may ensure sufficient spray distribution and atomization of the fuel.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 mixture of fuel. Fuels having higher hydrogen to carbon ratios may have higher specific energies in the absence of bonding stress. For example, fossil hydrocarbon fuels may comprise molecules having about 7 to 18 carbons, with a substantial portion of a particular composition derived from molecules having 9 to 15 carbons, with an average of 12 carbons.A number of mixtures of sustainable aviation fuels have been approved for use. For example, some permitted mixtures comprise mixing ratios of up to 10% sustainable aviation fuel, while other permitted mixtures comprise mixing ratios of between 10% and 50% sustainable aviation fuel (the balance comprising one or more fossil-based hydrocarbon fuels, such as kerosine), wherein permission is pending for further compositions. However, in the aviation industry, mixtures of sustainable flight fuels comprising up to (and including) 100% sustainable flight fuel (SAF) are expected to be eventually approved for use.Sustainable aviation fuels may include one or more of n-alkanes, isoalkanes, cycloalkanes, and aromatics, and may be made, for example, from one or more of syngas (syngas); lipids (e.g., fats, oils, and greases); sugars; and alcohols. Thus, persistent flying fuels may comprise lower aromatics and sulfur or both as compared to fossil hydrocarbon fuels. Additionally or alternatively, sustainable aviation fuels may comprise a higher content of isoalkanes and cycloalkanes or both compared to fossil hydrocarbon fuels. Thus, in some examples, persistent flying fuels may include a density between 90% and 98% of the density of kerosine and a heating value between 101 % and 105% of the density of kerosine, or both.In some examples, the sustainable fuel(s) or mixture(s) provided to the combustion equipment 16 may have a relatively lower content of aromatic and / or other non-paraffinic constituents than kerosine. The sustainable flying fuel may comprise an aromatic content of, for example, 30%, 20%, 15%, 10%, 8%, 5%, or less than 5%; e.g., 4%, 3%, 2%, 1%, or less than 1%; e.g., 0.75%, 0.5%, 0.25%, or less than 0.25%; e.g., 0.2%, 0.1%, or less than 0.1%; e.g., 0.01%, 0.001%, or 0%. The aromatic content of the sustainable flying fuel may be in an inclusive value or range that is limited by or within any of two of the values mentioned in the previous sentence (i.e. the values may form upper or lower limits), e.g. 13.5%, 8.5%, 2.5%, 0.35%, 0.15%, 0.05%, 0.005 % or 0%; or 0% to 0.75%, 0% to 0.5% or 0.1% to 0.25%; or 0.15% to 0.65%, 0.35% to 0.55% or 0.035% to 0.055%; depending on one or more of preference, fuel supply or supplier and variations of the composition thereof.Due, at least in part, to the molecular structure of sustainable flying fuels, sustainable flying fuels may provide advantages including, for example, one or more of a higher specific energy (despite, in some examples, a lower energy density); a higher specific heat capacity; a higher thermal stability; a higher lubricity; a lower viscosity; a lower surface tension; a lower freezing point; lower soot emissions; less NOx; and lower CO 2- emissions compared to fossil-based hydrocarbon fuels (e.g., when burned in the combustion equipment 16). Accordingly, persistent flying fuels may result in one or both of a relative decrease in specific fuel consumption and a relative decrease in maintenance costs, as compared to fossil hydrocarbon fuels such as kerosine.As shown in FIG. 5, an aircraft 1 may include a plurality of fuel tanks 50, 53; for example, a larger primary fuel tank 50 located in the aircraft fuselage and a smaller fuel tank 53 a, 53 b located in each wing. In other examples, an aircraft 1 may have only a single fuel tank 50 and / or the wing fuel tanks 53 may be larger than the central fuel tank 50, or a central fuel tank may not be provided (with all fuel instead stored in the aircraft's wings) - it should be appreciated that many different tank layouts are included and that the depicted examples are provided for ease of description and are not intended to be limiting.FIG. 5 shows an aircraft 1 with a drive system 2 comprising two gas turbine engines 10. The gas turbine engines 10 are supplied with fuel from a fuel supply system on board the aircraft 1. The fuel delivery system of the illustrated example includes a single fuel source. For purposes of the present application, the term "fuel source" means either 1) a single fuel tank or 2) a plurality of fuel tanks that are fluidly connected together. Each fuel source is arranged to provide a separate fuel source, i.e., a first fuel source may include a first fuel having a different property or properties than a second fuel included in a second fuel source. Thus, the first and second fuel sources are not fluidly connected to each other to separate the different fuels (at least under normal operating conditions). The use of several fuel sources allows an aircraft 1 to carry several different fuels and to change the fuel used during operation, optionally even in cruise flight or when changing between different operating stages in flight.In the present example, the first (and only in these examples) fuel source comprises a central fuel tank 50 located primarily in the fuselage of the aircraft 1, and a plurality of wing fuel tanks 53 a, 53 b, wherein for balancing at least one wing fuel tank is located in the port wing and at least one wing fuel tank is located in the starboard wing. All tanks 50, 53 are fluidly connected to each other in the example shown, thereby forming a single fuel source. Each of the middle fuel tank 50 and the wing fuel tank 53 may include a plurality of fuel tanks fluidly connected to each other.In another example, the wing fuel tanks 53 a, 53 bmay not be fluidly connected to the middle tank 50, thus forming a separate, second fuel source. For purposes of balancing, one or more fuel tanks in the port wing may be fluidly connected to one or more fuel tanks in the starboard wing. This may be done either via a middle fuel tank (unless this tank is part of the other fuel source), or bypassing the middle fuel tank / tanks, or both (for maximum flexibility and safety). In another example, the first fuel source includes wing fuel tanks 53 and a middle fuel tank 50, while a second fuel source includes another separate middle fuel tank. For balancing the aircraft 1, a liquid connection may be provided between the wing fuel tanks and the central fuel tank of the first fuel source. Thus, in multi-fuel source aircraft 1, two or more fuel sources may contain fuels that are different from each other so that aircraft 1 may change fuel during flight. Being able to determine which fuel is provided to the combustion chamber 16 may therefore be more complex than simply detecting a single identity of a fuel on board the aircraft 1 or checking once at start.In some examples, the allocation of the fuel tanks 50, 53 available in aircraft 1 may be limited such that the first fuel source and the second fuel source are each substantially symmetrical with respect to the centerline of the aircraft. In cases where asymmetric fuel tank allocation is permitted, a suitable means of fuel transfer is typically provided between the fuel tanks of the first fuel source and / or between the fuel tanks of the second fuel source so that the position of the center of gravity of the aircraft can be maintained within acceptable lateral limits throughout flight.Aircraft typically stop at several different airports, for example at the beginning and at the end of a long-distance flight. As mentioned above, while there are standards to which all of the aviation fuels must meet, different aviation fuels have different compositions, for example, depending on the source (e.g., different petroleum sources, bio-fuels, or other synthetic aviation fuels (often referred to as sustainable aviation fuels - SAFs - ) and / or mixtures of petroleum-based fuels and other fuels) and the additives included (e.g., antioxidants and metal deactivators, biocides, anti-static agents, anti-icing agents, anti-corrosion agents) and impurities. The composition of the available aviation fuel may vary not only from airport to airport and from fuel supplier to fuel supplier, but even from batch to batch of the same airport or fuel supplier. Moreover, the fuel tanks 50, 53 of the aircraft 1 are usually not emptied before being refilled for a further flight, which results in mixtures of different fuels being present in the tanks - and a fuel having a different composition being effectively formed from the mixture.The inventors have realized that by allowing different fuels to have different characteristics without conflicting with the standards, knowing the fuel / fuels available / available to an aircraft 1 may enable more efficient, individual control of the aircraft 1, and in particular the propulsion system 2 of the aircraft (i.e., the one or more gas turbine engines 10 of the aircraft 1 and the associated controllers and components). Knowledge of the fuel may therefore be used as an aid to improving aircraft performance, thus determining or monitoring the composition of the fuel may provide advantages. In particular, determining one or more fuel properties of the fuel to be supplied to the combustion chamber 16, whether it be a fuel from a single fuel source or a mixture of one or more fuels from different fuel sources, is therefore important for determining the operation of the engine. An essential feature of the operation of the engine is thermal management-thermal management of the engine is primarily via heat transfer to and from oil and fuel in the engine, such that control of a heat exchange system 3000-for example, by controlling the oil flow, controlling the fuel flow, and / or otherwise controlling one or more heat exchangers 1004, 1006, 2020, 2030, or other heat exchange components-may allow the performance of the engine 10 to be optimized for a fuel having certain fuel properties.As used herein, the term "fuel properties" refers to intrinsic or inherent fuel properties such as the composition of the fuel, and not to variable properties such as volume or temperature. One or more fuel characteristics may be determined and this data may be used to adjust the control of the engine 10, and more particularly the heat exchange system 3000. Examples of fuel properties include one or more of the following: i. the percentage of sustainable flying fuel (SAF) in the fuel or the indication that the fuel is a fossil fuel, e.g. fossil kerosine, or that the fuel is a pure SAF fuel; ii. parameters of a hydrocarbon distribution of the fuel, such as:• the content of aromatic hydrocarbons of the fuel and optionally also / alternatively the content of multiaromatic hydrocarbons of the fuel;• the hydrogen-carbon ratio (H / C) of the fuel;• Information on the % composition of some or all hydrocarbons present;iii. the presence or percentage of a particular element or species, such as:• the percentage of nitrogenous species in the fuel;• the presence or percentage of a tracer species or trace element in the fuel;• the naphthalene content of the fuel;• the sulfur content of the fuel;• the cycloparaffin content of the fuel;• the oxygen content of the fuel;iv. one or more properties of the fuel used in a gas turbine engine 10, such as:• Level of non-volatile particulate (nvPM) or CO 2- emissions during combustion;• Degree of coking of fuel;v. one or more properties of the fuel itself, regardless of use in an engine 10 or combustion, such as:• Thermal stability of the fuel (e.g., thermal decay temperature); the thermal stability may be assigned a numerical value at each temperature by forming the inverse of the deposition rate of fuel degradation products at the given temperature); and• one or more physical properties such as density, viscosity, heating value, freezing temperature and / or heat capacity.The fuel property(s) to be determined may be selected based on those fuel properties most relevant to possible changes to the heat exchange system 3000. The determination of the fuel properties may include obtaining the fuel properties of the fuel that was already in the fuel tank 50, 53 before refueling and an indication of the amount of fuel remaining, and then combining this information with the information about the new fuel that is filled into the fuel tank 50, 53 during refueling.Obtaining the fuel properties of a fuel already in the fuel tank 50, 53 prior to refueling and / or obtaining the fuel properties of a fuel provided during refueling may comprise one or more of the following elements: (i) physical and / or chemical detection of one or more features or parameters of the composition of the fuel (this may allow direct detection of the fuel properties and / or allow determination of the fuel properties based on the detection results) and / or detecting one or more tracer elements or compounds added to the fuel to facilitate its identification (e.g., a colorant); (ii) retrieving information about fuel properties from an integrated memory / data store; and / or (iii) receiving data, for example from an input provided at a user interface or from data transmitted to the aircraft 1.In some examples, one or more fuel characteristics may be determined during operation of the gas turbine engine 10, for example, by deriving the fuel characteristics of the fuel provided to the combustor 16 during operation from the performance metrics of the engine or by making a detection on the airfoil.In some examples, multiple different methods may be performed to obtain the fuel properties - for example, different methods may be used for different properties and / or different methods may be used for testing for the same property. For example, stored or otherwise provided fuel property data may be compared to the results of chemically or physically detecting one or more parameters of the fuel. If the stored fuel property and the corresponding detected parameter do not match, a warning message may be provided.The fuel properties may be determined by physically and / or chemically detecting one or more characteristics of the composition of the fuel (e.g., in a test unit external to the wing or during transport of the fuel to a fuel tank on the wing or even during use in the gas turbine engine 10) such that the fuel properties may be directly detected or data may be provided from which they may be determined as mentioned above, and / or by detecting one or more tracer elements or compounds added to the fuel to facilitate its identification (e.g., a colorant); or by receiving data, for example via input provided at a user interface, or by data transmitted to the aircraft, e.g., by scanning a barcode associated with the fuel delivery.In the physical and / or chemical determination, the fuel properties may be determined in various ways, both directly (e.g., based on sensor data corresponding to the fuel property in question) and indirectly (e.g., by derivation or calculation from other properties or measurements, or by reference to data for a particular tracer detected in the fuel). The properties may be determined as relative values compared to another fuel or as absolute values. For example, one or more of the following detection methods may be used:• The aromatic or cycloparaffin content of the fuel can be determined from measurements of swelling of a sensor component made of a sealing material, such as a nitrile sealing material.• Trace materials or species either naturally present in the fuel or added as tracer may be used to determine fuel properties, for example the percentage of sustainable flying fuel in the fuel or whether the fuel is kerosine.• Measurements of the vibration mode of a piezoelectric crystal exposed to the fuel can be used as a basis for determining various fuel properties, including the aromatic content of the fuel, the oxygen content of the fuel and the thermal stability or degree of coking of the fuel - for example, by measuring the formation of surface deposits on the piezoelectric crystal which lead to a change of the vibration mode.• Various fuel properties may be determined by collecting performance parameters of the gas turbine engine 10 during a first operating period (e.g., during takeoff), and optionally also during a second operating period (e.g., during cruise flight), and comparing these collected parameters to expected values using fuel with known properties.• Various fuel properties, including the content of aromatic hydrocarbons in the fuel, may be determined based on sensor measurements regarding the presence, absence or degree of formation of a condensate strip by the gas turbine 10 during its operation.• The fuel properties, including the content of aromatic hydrocarbons, can be determined by means of a UV-Vis spectroscopy measurement performed on the fuel.• Various fuel properties, including sulfur content, naphthalene content, aromatic hydrogen content, and hydrogen-carbon ratio, can be determined by measuring substances contained in the exhaust gases exhausted from the gas turbine engine 10 during operation.• The heating value of the fuel may be determined during operation of the aircraft 1 from measurements made during combustion of the fuel - for example from the fuel flow rate and the shaft speed or the temperature change in the combustion chamber 16.• Various fuel characteristics may be determined by making an operational change that affects operation of the gas turbine engine 10, detecting a response to the operational change; and determining the one or more fuel characteristics of the fuel based on the response to the operational change.• Different fuel characteristics may be determined with respect to the fuel characteristics of a first fuel by changing a fuel supplied to the gas turbine engine 10 from the first fuel to a second fuel and determining the one or more fuel characteristics of the second fuel based on a change in the ratio between T 30 and one of the values T 40 and T 41 (the ratio indicating the temperature rise in the combustion chamber 16). The properties may be determined as relative values to the first fuel or as absolute values, e.g., by reference to known values for the first fuel.As used herein, T30, T40, and T41, as well as any other numbered pressure readings and temperatures, are defined based on the station numbering set forth in the SAE AS755 standard, particularly:• T30 = total temperature at the outlet of the high pressure compressor (HPC);• T40 = total combustion outlet temperature;• T41 = Total temperature at the rotor inlet of a high pressure turbine (HPT).Any suitable approach known in the art may be used and the determination of fuel properties will not be discussed further herein.The aircraft 1 may therefore comprise a fuel composition determination module 57 configured to determine at least one fuel property of the fuel. The determination may be made by retrieving a value from a data store (e.g., from data provided to the aircraft 1 during refueling) or by performing one or more calculations based on data provided by one or more sensors or other engine components using any of the methods described above. In the example shown in FIG. 4, a fuel composition determination module 57 is provided as part of each engine 10 where fuel enters the engine 10. In other implementations, module 57 may be located elsewhere in engine 10 or elsewhere on aircraft 1, e.g., in, on, or adjacent to a fuel tank 50, 53. In implementations with a fuel composition determination module 57, module 57 may be configured to provide output to an electronic engine control unit (EEC), or module 57 may be provided as part of an EEC. The one or more fuel characteristics determined by the fuel composition determination module 57 may therefore be used to affect control of the engine 10. A special controller 58 may be provided to process the output of a fuel composition determination module 57 and provide control instructions to controllable components of the engine based on the fuel property information. In other implementations, the control functionality 58 may be provided by the EEC and no separate unit or module may be provided.An example fuel system 1000 for a direct drive gas turbine engine 10, the fuel system 1000 including a fuel flow path from the fuel tank 50 to the combustor 16 of the gas turbine engine 10 of the aircraft 1, is schematically illustrated in FIG. 6. The fuel system 1000 includes both the fuel supply system 50, 1002 (which supplies fuel to the engine 10) and the fuel management system 1500 (which operates within the engine 10) of the aircraft 1. the fuel management system 1500 is operable to control fuel temperature as well as fuel flow by directing fuel via one or more heat exchangers 1004, 1006 of the engine's heat exchange system 3000. The heat exchange system 3000 includes portions of the fuel management system 1500 as well as the oil circulation system 2000 (described in more detail below). Heat exchange system 3000 is the general term for the systems and components used to transfer heat between fluids (particularly oil and fuel) within engine 10, and includes heat exchangers, valves, connection pipes, and associated components, e.g., pumps, refrigerators, etc.Returning to fuel system 1000, fuel is pumped from fuel tank 50 to gas turbine engine 10 by low pressure fuel supply pump 1002. The fuel then flows through a primary fuel-oil heat exchanger 1006 and a secondary fuel-oil heat exchanger 1004. The primary fuel-oil heat exchanger 1006 may be referred to as a main fuel-oil heat exchanger because the fuel flowing through it may be the main fuel flow path from the fuel tank 50 to the combustion chamber 16. The secondary fuel-oil heat exchanger 1004 may be referred to as an auxiliary fuel-oil heat exchanger or a power-fuel-oil heat exchanger because fuel flowing therethrough may be supplied to the auxiliary systems or servomechanisms 33 of the aircraft 1. Such auxiliary systems / servomechanisms may include, but are not limited to, fuel hydraulic actuation; heating; a nacelle anti-icing system; engine actuators; and / or a turbine case cooling (TCC) servovalve. In these auxiliary systems, only fuel that has passed through the secondary fuel-oil heat exchanger may be used. The fuel used in these auxiliary systems may be recirculated to a fuel tank for later recirculation or may be recombined with other fuel exiting the main fuel-oil heat exchanger 1006 and entering the combustion chamber. The fuel passing through the secondary fuel-oil heat exchanger 1004 may therefore, in some implementations, not be provided to the combustion chamber, but instead be returned to an aircraft fuel tank 50, optionally after use in auxiliary systems such as fuel hydraulic actuators. The fuel flowing through the secondary fuel-oil heat exchanger 1004 may instead be recirculated along the main fuel flow path between the fuel tank 50 and the engine fuel pump 1003. In other implementations, more or fewer fuel-oil heat exchangers may be provided. Depicted fuel management system 1500 is configured such that fuel reaches primary fuel-oil heat exchanger 1006 prior to secondary fuel-oil heat exchanger 1004. After exiting the fuel-oil heat exchanger 1006, the fuel passes through an engine fuel pump 1003 and then enters the combustion chamber 16.The primary fuel-oil heat exchanger 1006 and the secondary fuel-oil heat exchanger 1004 are configured to also carry a stream of oil therethrough in addition to the stream of fuel. The primary fuel-oil heat exchanger 1006 and the secondary fuel-oil heat exchanger 1004 are configured such that heat may be transferred between the oil and the fuel flowing through them. During normal operation of the engine 10, for example, cruise conditions, the average temperature of the oil stream entering the primary fuel-oil heat exchanger 1006 is higher than the average temperature of the fuel entering the primary fuel-oil heat exchanger 1006, and the average temperature of the oil stream entering the secondary fuel-oil heat exchanger 1004 is higher than the average temperature of the fuel entering the secondary fuel-oil heat exchanger 1004. In this manner, the primary fuel-oil heat exchanger 1006 and the secondary fuel-oil heat exchanger 1004 are each configured to transfer thermal energy from an oil stream to a fuel stream flowing therethrough during operation.Generally, at least a majority of the fuel passes through the primary fuel-oil heat exchanger 1006, and at least a portion of the fuel passing through the primary fuel-oil heat exchanger 1006 also passes through the secondary fuel-oil heat exchanger 1004. The two heat exchangers 1004, 1006 may therefore be described as being in series with each other and along a main fuel flow path from the tank 50 to the combustion chamber 16 with respect to fuel flow. However, one or each of the heat exchangers 1004, 1006 may be bypassed so that a portion of the fuel does not flow through the respective heat exchanger, for example in the form of a bypass pipe for fuel 1005, as shown in FIG. 6. A valve (not shown) may determine which portion of the fuel is directed through the heat exchanger 1004 and which portion is directed through the bypass tube 1005. In various implementations, a bypass tube may be provided for each heat exchanger 1004, 1006 so that a portion of the fuel may bypass one or both heat exchangers. Additionally or alternatively, one or more bypass pipes 2005, 2005 afor the oil may be provided for one or both fuel-oil heat exchangers 1004, 1006, such that a portion of the oil may bypass one or more heat exchangers. A valve arranged to control the flow of a liquid through a bypass pipe 1005 may be referred to as a bypass valve.The secondary fuel-oil heat exchanger 1004 and the main fuel-oil heat exchanger 1006 are therefore configured such that, in addition to the fuel stream, an oil stream is also passed through each of them - in the described implementation, the same oil flows through the one and then the other, although it is understood that in other implementations, the different oil circuit systems may feed each heat exchanger and thus different oil may flow through a fuel-oil heat exchanger to the other fuel-oil heat exchanger.In the example shown in FIG. 6, fuel system 1000 further includes a temperature sensor 1009 arranged to sense the temperature of fuel approaching or reaching combustion chamber 16. The temperature sensor 1009 is shown adjacent an inlet of the combustion chamber 16 in the depicted example, but in various examples, the sensor 1009 may be disposed at any location downstream of the primary fuel-oil heat exchanger 1006 or the secondary fuel-oil heat exchanger on the fuel side to provide a measurement of fuel temperature at various points along the flow path. In gas turbine engines 10 such as those described herein, the fuel flows through fuel spray nozzles before entering the combustor itself-these nozzles may be very susceptible to thermal decomposition of the fuel, resulting in blockages in the relatively narrow passages (e.g., measurement slots); therefore, in some implementations, it is recommended to place the sensor 1009 as close as possible to the inlet to the nozzles. In some implementations, multiple temperature sensors 1009 may be used. The one or more temperature sensors 1009 are arranged to provide an output (temperature data) to a controller 58.FIG. 7A shows an alternative example fuel system 6000 including a fuel supply system and a fuel management system 6500 including a fuel flow path from the fuel tank 50 to the combustor 16 of the gas turbine engine 10 of the aircraft 1. As with fuel system 1000, fuel is pumped from fuel tank 50 by low pressure fuel supply pump 1002. The fuel then flows through a primary fuel-oil heat exchanger 1006 before reaching an engine fuel pump 1003 that pumps the fuel along its flow path to the combustion chamber 16, optionally via a secondary fuel-oil heat exchanger 1004. Control of fuel flow through the secondary fuel-oil heat exchanger 1004 may be with the valve 1004 a. Fuel system 6000 differs from fuel system 1000 in that fuel system 6000 includes a recirculation valve 6010 located downstream of primary fuel-oil heat exchanger 1006 and configured to recirculate at least a portion of the fuel that has left the primary fuel-oil heat exchanger back to inlet 1006 aof primary fuel-oil heat exchanger 1006, thereby allowing further heat transfer between the primary circulation system oil and the fuel. Recirculation valve 6010 may determine which portion of fuel is recirculated via recirculation pipe 6011 and which portion is directed more directly to combustion chamber 16. In the illustrated example, the recirculation valve 6010 is downstream of the primary fuel-oil heat exchanger 1006. In the illustrated example, the recirculation valve 6010 is positioned downstream of the engine fuel pump 1003. In some implementations where the recirculation valve 6010 is positioned upstream of the engine fuel pump 1003, an additional recirculation pump (not shown) may be provided to provide a positive pressure gradient. A conduit for recirculating fuel to the tank (FRTT) may be present and may also aid in recirculation in some implementations. The recirculation valve 6010 is arranged to allow a controlled amount of fuel to be recirculated to the inlet 1006 aof the primary heat exchanger 1006 and thus flow through the primary heat exchanger 1006 a plurality of times before reaching the combustion chamber 16. It is contemplated that the recirculation valve could be positioned upstream of the engine fuel pump 1003 in alternative implementations. In such implementations, the recirculation valve 6010 would be arranged such that a controlled amount of fuel can be recirculated to the inlet 1006 aof the primary heat exchanger 1006, thus flowing multiple times through the primary heat exchanger 1006 before reaching the pump 1003. This recirculation provides a mechanism for controlling fuel flow within fuel management system 6500, and also within heat exchange system 3000, without altering fuel flow from tank 50 to engine 10. The feedback may be used to control the heat transfer ratio.Thus, a tube 6011 may be provided, which may be referred to as a return tube, as it carries fuel from a point on the main flow path through the engine 10 to an earlier point on that flow path, so that the fuel must pass through the intermediate portion of the flow path again, and which leads from the return valve to a point on the flow path upstream of the inlet 1006a to the primary heat exchanger 1006. In some implementations, a return pipe 6011 and a bypass pipe 1005 may be provided for any given heat exchanger 1004, 1006. In some implementations, the same pipe 1005, 6011may serve as both a return pipe and a bypass pipe. One or more valves may be used to control the direction of liquid flow through the tube.Controlling fuel flow in fuel system 6000 by recirculating through and / or bypassing one or more heat exchangers may assist in thermal management (e.g., affecting the temperature of the fuel as it enters combustion chamber 16 or pump 1003, or the amount of heat transferred to the fuel). Thus, a portion of the control of the heat exchange system 3000 may be the control of one or more valves that control the flow of oil and / or fuel through one or more return and / or bypass pipes 6011, 1005.Control of fuel flow through the paths of the recirculation pipe 6011 may be based on the measurement of fuel temperature (e.g., using a temperature sensor at a location downstream of the primary fuel-oil heat exchanger 1006, possibly also using a temperature measurement upstream of the primary fuel-oil heat exchanger 1006). Recirculation of the recirculated fuel upstream of the primary fuel-oil heat exchanger 1006 may allow for reduced heat transfer from the oil to the fuel and thus dampen transient overshoot that might occur, e.g., at the beginning of the descent phase when the same amount of heat is generated in the oil system, but fuel flow is reduced, which typically results in a temperature increase. Adjusting the flow of fuel through the recirculation pipe 6011 may also be useful, for example, to reduce icing by circulating fuel that has been previously heated to heat the system. The recirculation pipe 6011 can therefore be used in various ways to improve engine thermal management.FIG. 7B illustrates an alternative example fuel system 7000 to that shown in FIG. 7A. Fuel system 7000 is similar to fuel system 6000 shown in FIG. 7A, but differs in that at least a portion of the fuel exiting secondary fuel-oil heat exchanger 1004 is supplied to additional engine and / or aircraft mechanisms 7010, rather than combustor 16. These mechanisms 7010 may be one or more of a nacelle anti-icing system, actuators, and / or turbine housing cooling (TCC) systems. At least a portion of the fuel exiting the secondary fuel-oil heat exchanger 1004 is returned to the fuel supply system (i.e., the fuel tank 50 or the fuel supply pump 1002) or downstream of the redistribution supply system (i.e., upstream of the primary fuel-oil heat exchanger 1006). The portion of fuel returned to / downstream of the fuel supply system may be controlled by the valve 7011. No fuel passing through the secondary fuel-oil heat exchanger 1004 is delivered to the combustor 16 (at least not directly - it should be appreciated that fuel returned to the fuel tank 50 may later re-enter the engine 10 and then be delivered to the combustor 16) in the example of FIG. 7B.FIG. 7C illustrates another alternative example fuel system 7000 a. Fuel system 7000a is similar to fuel system 7000 shown in FIG. 7B and differs in that the portion of fuel exiting secondary fuel-oil heat exchanger 1004 is recirculated to join the main fuel stream as it approaches combustion chamber 16 after it is optionally supplied to one or more additional engine and / or aircraft mechanisms 7010 (which may be as described above). Any fuel not needed for these auxiliary systems 7010 may be returned to the main fuel flow path / combustion chamber 16 without bypassing. At least a portion of the fuel exiting the secondary fuel-oil heat exchanger 1004 may be returned to the main fuel flow path without flowing through the auxiliary systems 7010-the portion directed to the auxiliary systems 7010 may be adjusted as needed and controlled with the valve 7011. In other implementations, fuel used in one or more such additional aircraft mechanisms 7010 may then be returned to a tank 50 instead of being sent to the combustion chamber 16.The described gas turbine engine 10 of the aircraft 1 includes an oil circulation system arranged to provide oil for lubricating and heat removing a plurality of components. In the implementation described and illustrated in FIG. 8A, the oil circulation system comprises a single closed oil circulation system 2000. In other implementations, the oil circulation system may include a primary oil circulation system 2000 and a secondary oil circulation system (not shown), each of which is closed oil circulation systems. The secondary oil circulation system may supply oil to the primary fuel-oil heat exchanger 1006 and / or the secondary fuel-oil heat exchanger, and possibly additional heat exchangers such as an air-oil heat exchanger.In the illustrated implementation, the closed loop oil system 2000 includes an oil tank 2002 suitable for receiving a particular volume of oil. In some implementations, gases are removed from the oil in oil tank 2002 by a breather.A feed pump 2004 is configured to pump oil from the oil tank 2002 to the secondary fuel-oil heat exchanger 1006. The average temperature of the oil entering the secondary fuel-oil heat exchanger 1006 is higher than the average temperature of the fuel entering the secondary fuel-oil heat exchanger 1006 during cruise conditions. In the secondary fuel-oil heat exchanger 1006, thermal energy is transferred from the oil stream to the fuel stream. In this way, the average temperature of the oil stream exiting the secondary fuel-oil heat exchanger 1004 is lower than the average temperature of the oil stream entering the secondary fuel-oil heat exchanger 1006. Also in this way, the average temperature of the fuel exiting the secondary fuel-oil heat exchanger 1004 is higher than the average temperature of the fuel entering the secondary fuel-oil heat exchanger 1006.From the secondary fuel-oil heat exchanger 1004, the oil then flows to the primary fuel-oil heat exchanger 1006. In some implementations, the oil may flow through an air-to-oil heat exchanger (not shown) between the secondary fuel-to-oil heat exchanger 1004 and the primary fuel-to-oil heat exchanger 1006.In the primary fuel-oil heat exchanger 1006, thermal energy is transferred from the oil stream to the fuel stream. In this way, the average temperature of the oil stream exiting the primary fuel-oil heat exchanger 1006 is lower than the average temperature of the oil stream entering the primary fuel-oil heat exchanger 1006. Also in this way, the average temperature of the fuel exiting the primary fuel-oil heat exchanger 1006 is higher than the average temperature of the fuel entering the primary fuel-oil heat exchanger 1006.From the primary fuel-oil heat exchanger 1006, the oil stream is then directed through the engine components 2006 that are to cool and / or lubricate the oil. The oil acts as a lubricant and a coolant within the engine components 2006 and absorbs heat from the engine components 2006 during the process. The engine components 2006 may include one or more bearing chambers.Coming from the engine components 2006, the oil collects in the oil pan 2008. A scavenge pump 2010 is configured to pump oil from the oil pan 2008 back into the oil tank 2002.FIGS. 8B and 8C illustrate alternative oil circuit systems 2000 b, 2000 c. In the oil circulation system 2000 bof FIG. 8B, the oil flows through an air-oil heat exchanger for cooling before entering the primary fuel-oil heat exchanger 1006. The secondary fuel-oil heat exchanger is not present in this alternative oil circulation system 2000 b.FIG. 8B also illustrates details of the air-oil heat exchanger 2020 not shown in other figures for clarity, and particularly indicates the air flow with dashed arrows. The air-oil heat exchanger 2020 has an air inlet 2020a and an air outlet 2020b. The air inlet or intake passage 2020a may be arranged to receive air in the bypass duct 22 of the engine downstream of an outlet guide vane of the fan 23. The heat exchanger exhaust / air outlet 2020b may be arranged to return the air downstream of the inlet 2020a into the bypass duct 22 or directly into the outside atmosphere (e.g., when a particular outlet nozzle is provided).A valve 2022, referred to as an air valve, is used to control the air flow rate through the air-oil heat exchanger 2020. In the depicted example, air valve 2022 is provided at or near inlet 2020a to heat exchanger 2020 and may therefore be referred to as air inlet valve 2022. In other implementations, the air valve 2022 may be provided at or near the outlet 2020 bfrom the heat exchanger 2020 and may therefore be referred to as an air outlet valve. Any suitable position may be selected for the valve 2022, provided that the valve 2022 can be adjusted so that the flow of air through the air-oil heat exchanger 2020 can be controlled. In the described example, the air valve 2022 may be continuously adjusted between a fully closed position (no air flow through the heat exchanger) and a fully open position (maximum air flow through the heat exchanger). In other examples, the air valve 2022 may be adjustable between a plurality of discrete positions, rather than being continuously adjustable - for example, between six, five, four, three, or two different positions. In implementations with only two valve positions for the air valve 2022, these positions may be "open" and "closed"-the valve 2022 may be repeatedly opened and closed to provide pulsating air flow when a medium cooling level is desired in some of these implementations.FIG. 8C shows an oil circulation system including an oil circulation system 2000 c. A main oil flow path passes through all heat exchangers 1004, 1006, 2020 in this circuit. To adjust the flow of oil through the heat exchangers, one or more bypass pipes 2005, 2005a, 2005b are provided, wherein bypass valves 2007, 2007a, 2007b are used to vary the flow through the bypass pipes. A controller 58 may be used to actively control the fuel and / or oil flows through and around the heat exchangers 1004, 1006, 2020. Controller 58 may be part of or provided by an EEC, or it may be a separate unit. The control can be automated, for example by an EEC.In other implementations, branching paths to different heat exchangers could be used instead. It will be appreciated that although a bypass pipe provides a parallel, alternative flow path from a technical point of view, a branching path as described herein with respect to parallel flows differs in that each branching route of a parallel arrangement of heat exchangers comprises a heat exchanger, whereas the bypass pipe is merely a pipe without (appreciable) heat exchange or other components thereon.It should be appreciated that in some implementations, a combination of arrangements of heat exchangers and bypass tubes may be used and that FIG. 8C (all heat exchangers in series, multiple bypass tubes) may be viewed as representing an end of a design spectrum. In an alternative implementation, one or more of the heat exchangers may be part of a secondary oil circulation system. In an alternative implementation, the heat exchangers may be arranged in parallel, with the oil circulation system 2000 branching into separate branches, with two or more separate branches each comprising a heat exchanger.The oil circulation system 2000 illustrated in FIG. 8C provides a series oil flow path leading from a tank 2002 through an oil pump 2004, further through the secondary fuel-oil heat exchanger 1004, the primary fuel-oil heat exchanger 1006, and then the air-oil heat exchanger 2020, before entering the engine components 2006 for cooling and / or lubrication thereat, and then being collected in an oil pan 2008, and then being pumped back into the tank 2002 with the oil pump 2010. The heat exchangers 1004, 1006, 2020 are thus in a series arrangement. The order of the heat exchangers 1004, 1006, 2020 may be reversed such that, for example, the primary fuel-oil heat exchanger 1006 is disposed after the air-oil heat exchanger 2020 in alternative implementations.The oil circulation system 2000 cillustrated in FIG. 8C includes three bypass pipes 2005, 2005 aand 2005 b. The first bypass pipe 2005 is disposed such that a part of the oil can bypass the secondary fuel-oil heat exchanger 1004, and is controlled by a first bypass valve 2007. The first bypass pipe 2005 directs oil from upstream of the inlet to the secondary heat exchanger 1004 and returns it to the main fuel flow path downstream of the secondary heat exchanger 1004 and before the primary fuel-oil heat exchanger 1006 and the air-oil heat exchanger 2020. The second bypass pipe 2005 ais disposed so that a part of the oil can bypass primary fuel-oil heat exchangers 1006, and is controlled by a second bypass valve 2007 a. The second bypass pipe 2005 aroutes oil from upstream of the inlet to the primary fuel-oil heat exchanger 2020 and returns it to the main fuel flow path downstream of the primary fuel-oil heat exchanger 2020 before the path reaches the air-oil heat exchanger 2020. The third bypass pipe 2005 broutes oil from upstream of the inlet to the air-oil heat exchanger 2020 and returns it to the main oil flow path downstream of the air-oil heat exchanger 2020 before the path reaches the engine components 2006. In implementations of the primary oil circulation system 2000 with only one oil bypass pipe 2005, the selected position may be that of the third bypass pipe 2005 b, such that there is a bypass for the air-oil heat exchanger 2020 and not for the fuel-oil heat exchanger 1006. As a result, as much heat as possible can be reliably dissipated from the oil into the fuel, and by reducing the heat loss to the environment, the thermal efficiency of the engine can be kept high. The presence of a bypass pipe 2005 aat the primary fuel-oil heat exchanger 1006 may promote a quick adjustment of the heat transfer ratio when there is a risk that this ratio falls below the desired value (for example, depending on the fuel properties determined, a lower limit may be set for the ratio that is significantly above zero). Severe fixed temperature limits may apply to some fuels depending on their properties. The presence of a bypass pipe 2005, 2005a, 2020 at all heat exchangers 1004, 1006, 2020 can protect the oil from cooling too much under certain conditions in order to avoid the risk of oil clotting-it is understood that this can represent a greater problem on a cold day under idling conditions (e.g. during take-off or under ground idling or in descent) than during cruise.The inventors have recognized that the use of fuels different from conventional kerosine-based jet fuels, such as persistent flying fuels, may result in different fuel properties and that the heat transfer parameters may be adjusted in operation to take advantage of the different fuel properties. Particularly in implementations where there are no parallel arrangements of heat exchangers (as used in the methods described with reference to FIGS. 14 and 15 ), the introduction of one or more bypass pipes with controllable valves and the implementation of careful control of these valves may enable a method that provides improved oil cooling (as the fuel may possibly absorb more heat) and may also improve overall thermal efficiency of the engine as less heat is lost to the environment. The controllable heat exchange system 3000 and in particular one or more controllable bypass valves 2007, 2007', 2007a is assigned a key role in the regulation of the heat transfer ratio in such engines 10.Although cruise conditions generally account for a much greater proportion of engine operating time, the inventors have recognized that idle operation is also important. Since the fuel flow rate is much lower at idle than at cruise, even a relatively low heat load on the fuel can result in a high temperature rise. The use of non-conventional fuels may therefore have an even greater impact on optimal approaches to thermal management under idle conditions. The methods relating to these two scenarios of aircraft operation will be discussed in more detail with reference to FIGS. 14 and 15.In a standard heat exchanger 1006, the temperature of the coolant (in this case, the fuel) is the basic limiting factor for the cooling power provided - the temperature of the oil as it exits the heat exchanger 1006 cannot be lowered below the temperature of the fuel entering the heat exchanger 1006 (and also the temperature of the fuel cannot be raised higher than that of the oil entering the heat exchanger), even if a substantially infinite flow rate of fuel or oil were used. Temperature compensation is the threshold as defined in the second principle set of thermodynamics. Therefore, in some implementations, as illustrated in FIG. 9, a refrigeration cycle device 1007 is provided. The refrigeration cycle device 1007 is arranged to provide thermal buoyancy by transferring further heat from the oil to the fuel so that the temperature of the fuel is raised more than would be the case with a simple pass through the heat exchanger, and in some cases even beyond the oil temperature. The refrigeration cycle device 1007 may take the form of an additional closed circuit including a refrigerant liquid, an evaporator between oil and refrigerant (to transfer heat from the oil to the refrigerant), and a condenser between fuel and refrigerant (to transfer heat from the refrigerant to the fuel. The refrigeration cycle device 1007 may additionally include a pump, compressor, and / or expansion valve / meter, and may be embodied in any suitable embodiment known in the art. An oil valve may control how much oil flows through the refrigeration cycle device 1007, and a fuel valve may control how much fuel flows through the refrigeration cycle device 1007. The refrigeration cycle device 1007 (and particularly its pump and / or compressor, and optionally also the expansion valve or meter) is generally electrically or mechanically driven (e.g., via a connection to a rotating shaft) to provide this further forced heat transfer. The refrigeration cycle device 1007 is shown in connection with the main fuel-oil heat exchanger (primary fuel-oil heat exchanger) 1006 in the depicted implementation, thus providing an increase in fuel temperature after fuel passes both fuel-oil heat exchangers 1004, 1006, and further cooling oil of the oil system 2000 before being recirculated to the engine components 2006. In various implementations, a bypass or return pipe for the oil and / or fuel may be provided around the refrigeration cycle device 1007.FIG. 10 schematically illustrates an example heat exchange system 3000 including the closed oil circuit system 2000 of FIG. 8A and the fuel system 1000 of FIG. 6. The two systems 1000, 2000 are shown together to illustrate the interactions between the fuel and the oil stream within the gas turbine engine 10. The fuel flow is shown with thick black arrows and the oil flow with thin black arrows.FIG. 11 schematically illustrates an alternative example heat exchange system 4000 including the closed oil circuit system 2000 of FIG. 8A and the fuel delivery system 7000 aof FIG. 7A. Systems 7000a and 2000 are substantially the same as those illustrated in FIGS. 8A and 7A, with the fuel flow again represented by thick black arrows and the oil flow by thin black arrows. However, the recirculation valve 6010 is located somewhat different from FIG. 7A, upstream of the bifurcation in the fuel flow path to the secondary heat exchanger 1006, rather than being between the inlet to 1006 aand the outlet of 1006 bof the heat exchanger 1006. Thus, in the implementation shown in FIG. 11, all of the fuel exiting pump 1003 flows through recirculation valve 6010, unlike the embodiment shown in FIG. 7A. The proportion of fuel flowing through the secondary fuel-oil heat exchanger 1004 may be controlled with the valve 1004 a. The control methodology of valves 1004a, 6010 may be adjusted as desired.It should be appreciated that in implementations such as that shown in FIG. 7A, where fuel flowing through the secondary fuel-oil heat exchanger 1004 is used as working fluid in auxiliary systems and then returned to a tank 50, 53 instead of being directed to the combustion chamber 16, the fuel flow line from the secondary heat exchanger 1006 to the combustion chamber 16 may be replaced with a fuel flow line back to the tank 50.The inventors have recognized that the use of fuels different from conventional kerosine-based jet fuels, such as sustainable flying fuels, may result in different fuel properties and that the operation of the engine may be optimized for these different fuel properties. Fuels with a higher heating value than claimed may also have a higher thermal stability, whereby the fuel may absorb more heat, whereby improved oil cooling and / or combustion properties may be provided in the combustion chamber. Recycling the fuel through the primary fuel-oil heat exchanger 1006 may allow the fuel to draw more heat from the oil, thereby increasing fuel temperature and improving oil cooling.A recirculation valve 6010 may be located downstream of the fuel pump 1003 and may therefore provide improved flexibility in fuel flow. Therefore, for a higher heating fuel, a lower fuel flow rate may be provided to the combustion chamber 16 without changing pump speed by instead increasing the amount of recirculation.FIG. 12 schematically illustrates an alternative example heat exchange system 4500 including the closed oil circuit system 2000 of FIG. 8C and a portion of the fuel delivery system 7000 aof FIG. 7A. For clarity, only the fuel passing through the primary fuel-oil heat exchanger 1006 and the secondary fuel-oil heat exchanger 1004 is shown.In the above implementations, one or more temperature sensors may be provided, for example, arranged to sense fuel temperature as the fuel pump 1003 or the combustion chamber 16 approaches. The data from one or more sensors, optionally in combination with other data (e.g., fuel characteristics as described above), may be used by a controller 58 of the heat exchange system 3000 to control the flow of fuel and / or oil through and around the heat exchangers 1004, 1006, 2020. Controller 58 may be part of or provided by an EEC, or it may be a separate unit. The control can be automated, for example by an EEC.The inventors have recognized that in various implementations, knowledge of one or more fuel characteristics may be utilized to adjust the operation of the heat exchange system 3000 to improve the performance of the engine 10 by utilizing the different fuel characteristics of the individual flying fuels.For the sake of clarity and brevity, reference is made in the methods described herein to a single bypass valve 2007 and a corresponding bypass pipe 2005. As indicated by the implementation of FIG. 12, implementations may actually include multiple bypass valves 2007, 2007 a, 2007 b, each of which has an associated bypass pipe 2005, 2005 a, 2005 bto allow the oil stream to bypass one or more of the secondary fuel-oil heat exchangers 1004, the primary fuel-oil heat exchanger 1006, and / or the air-oil heat exchanger 2020. When reference is made to the "bypass valve 2007" or the "bypass pipe 2005" in the discussion of the methods illustrated in FIGS. 13 to 17, it is intended to refer to one or more of the plurality of bypass valves 2007, 2007 a, 2007 band the associated bypass pipes 2005, 2005 a, 2005 b, depending on which combination of heat exchangers is just discussed.FIG. 13 illustrates a first method 100 in which knowledge of the fuel properties is used to determine control of the heat exchange system 3000. The method 100 includes determining 102 at least one fuel property of the fuel to be burned by the combustion chamber 16. This determination 102 can be carried out according to one of the methods described above, wherein optionally a module for determining the fuel composition 57 can be used.The engine 10 employed to implement the method 100 includes an air-to-oil heat exchanger 2020 and a fuel-to-oil heat exchanger 1006, a bypass pipe 2005 arranged such that a portion of the oil may bypass one of the air-to-oil heat exchanger and the fuel-to-oil heat exchanger; and a bypass valve 2007 arranged such that the portion of the oil directed through the bypass pipe may be varied, and the method 100 includes controlling 104 the bypass valve 2007 based on the at least one fuel characteristic to adjust the portion of the oil directed over each heat exchanger under cruise conditions.More specifically, the method 100 for the implementations described with reference to FIGS. 5-12 above includes controlling 104 the bypass valve 2007 of the oil system 2000 under cruise conditions based on the at least one fuel property to adjust the amount of oil directed across each bypass pipe 2005 of the oil circuit system 2000 so as to control how much oil bypasses the primary fuel-oil heat exchanger 1006 and how much the air-oil heat exchanger 2020 bypasses, and optionally how much the secondary fuel-oil heat exchanger 1004. In implementations with different arrangements of heat exchangers, e.g. a parallel arrangement of heat exchangers instead of a serial arrangement, the bypass valve 2007 may be replaced by or function as a modulation valve that directs the oil flow along one or more branches of the parallel split arrangement, so that the proportion of oil directed across each heat exchanger can be varied in this way.As indicated by the dashed line in FIG. 13, this method 100 can optionally be repeated. In some implementations, for example, implementations with only one fuel on board the aircraft 1, the method 100 may be performed only once during a flight cycle, e.g., during refueling. In other implementations, such as implementations with multiple fuel sources, the fuel supplied to the combustion chamber 16 may vary over a flight. The determination 102 may therefore be performed more than once - e.g.: (i) once for each fuel source 50, 53 at fueling / commencement of a flight in implementations with multiple fuel sources, only one or the other of which is used at any particular time (and knowing from which fuel source fuel is drawn may cause selection of the appropriate stored fuel property); or (ii) multiple times during a flight, e.g., in response to a change in fuel source(s) from which fuel is drawn (note that the fuel provided to the combustion chamber 16 may be a mixture of fuels from different sources in some implementations), or at regular intervals.The bypass valve 2007 may therefore be configured to bypass a fixed portion of the oil flow through the bypass pipe 2005 during operation of the engine 10, wherein the fixed portion is determined at the start of the engine 10 based on the at least one determined fuel property (option (i) above). Alternatively, bypass valve 2007 may be configured such that during operation of engine 10 during a single flight, a variable portion of the oil flow is diverted through bypass tube 2005 (option (ii) above). The bypass valve 2007 may therefore be actively controlled to vary the proportion of oil directed through the bypass pipe 2005 (and thus through each heat exchanger 1004, 1006, 2020), particularly in implementations where the aircraft 1 carries a plurality of different fuels in different tanks and may change the fuel (or fuel mixture) used during flight. The active control of the bypass valve 2007 can be automated and implemented by a control unit 58 of the heat exchange system 3000, which can be a special control unit or a part of a more general EEC. In various implementations, the control of the bypass valve 2007 may be a closed or open loop depending on the availability of inline measurements-in particular, a closed loop approach may be preferable when feedback data (e.g., from oil flow rate sensors) is available.The inventors have recognized that fuels with certain fuel properties may absorb more heat than others without decomposing, e.g., because they have a higher heat capacity and / or thermal stability, or otherwise benefit from a higher temperature (e.g., in terms of fuel combustion efficiency). Adaptively controlling the heat exchange system 3000 based on the determined fuel characteristics therefore allows for more efficient use of various fuels, as well as more efficient engine operation and oil cooling.The fuel property(s) may not be the only data taken into account in adjusting the bypass valve 2007. For example, the temperature of the fuel exiting or entering the combustion chamber 16 from the fuel-oil heat exchanger 1006 may also be considered, as may the thermal compatibility of the components of the fuel system downstream of the heat exchanger 1006 (and optionally also the refrigeration circuit device 1007, if present). When a refrigeration cycle device 1007 is present, the bypass valve 2007 or other oil valve may control how much oil flows through the refrigeration cycle device 1007. In some implementations, all of the oil passing through the main fuel-oil heat exchanger 1006 and optionally all of the fuel passing through this heat exchanger 1006 may also pass through the refrigeration cycle device 1007. However, the refrigeration cycle device 1007 cannot be energized / operated for a portion of this time, so that it does not provide a temperature rise.Subject to the suitability of the at least one determined fuel property, method 100 may include using bypass valve 2007 to perform one or more of the following actions:• directing all of the oil via the primary fuel-oil heat exchanger 1006 (no oil via the air-oil heat exchanger 2020) for one or more periods of at least 30 minutes in cruise so that, for at least some periods of time of operation in cruise, no heat is lost to the environment via the air-oil heat exchanger;• directing all oil over the primary fuel-oil heat exchanger 1006 (no oil over the air-oil heat exchanger 2020) for at least 15% of total time spent in cruise and optionally for at least 20% of total time spent in cruise;• Passing at least 95% of the oil over the primary fuel-oil heat exchanger 1006 (no more than 5% over the air-oil heat exchanger 2020) during at least 90% of the time spent in cruise so that during at least 90% of cruise operation, if any, very little heat is lost to the environment via the air-oil heat exchanger;• controlling the flow of oil such that at least 80% of the heat dissipated from the oil during cruise is transferred to the fuel; and / or• controlling the flow of oil so that all heat dissipated by the oil during cruise is transferred to the fuel over at least 90% of the time spent during cruise; and / or• Controlling the flow of oil so that in the heat exchange system 3000, heat from the oil is transferred to the fuel under cruise conditions per kilogram of fuel between 200 and 600 kJ and optionally 200-500 kJ. In some implementations, the transferred heat may be in a range of 350 to 450 kJ / kg fuel.The suitability of the at least one ascertained fuel property can be determined by comparing one or more fuel properties with a threshold value established for the respective fuel properties. For example, a fuel having a SAF content (% SAF) of over 60%, 65%. For a reduction in oil flow rate through the air-oil heat exchanger 2020 as compared to conventional flying fuels, 70%, 75% or 80% may be considered suitable. In some implementations, multiple fuel characteristics may be checked together - for example, by using a lower threshold for one fuel characteristic when another fuel characteristic is within a certain range. The fuel properties may be considered independent / isolated in other implementations.The at least one fuel property may be or may comprise one or more of thermal stability, aromatic hydrocarbon content in the fuel, or percentage of persistent flying fuel in the fuel, and modulation valve control decisions may be made based on knowledge of this at least one fuel property. If, for example, the data on thermal stability indicate that the fuel is stable in operation at temperatures above 160° C. or 170° C., the bypass valve 2007 can be controlled 104 such that at least 80% and optionally 90-100% of the heat dissipated from the oil during cruise flight is transferred to the fuel. As a further example: If the molar fraction of the aromatic hydrocarbons in the fuel is below 12% and optionally below 10% or below 5%, at least 80% and optionally 90-100% of the heat dissipated from the oil during cruise flight can be transferred to the fuel. As another example, if the SAF content of the fuel is above 50% and optionally above 75%, the bypass valve 2007 may be controlled such that at least 80% and optionally 90-100% of the heat dissipated by the oil during cruise flight is transferred to the fuel. As another example, if the heating value of the fuel is at least 43.5 MJ / kg, at least 80%, and optionally 90-100%, of the heat dissipated by the oil during cruise flight can be transferred to the fuel. As a rule, less heat can be transferred to a conventional fuel if the heating value of the fuel exceeds a threshold value, because in fuels with a higher heating value, the flow rate is generally reduced in order to achieve the same thrust force. The fuel in the primary fuel-oil heat exchanger 1006 may therefore experience an increased temperature rise due to its reduced flow rate if the amount of oil passing through the heat exchanger 1006 is not reduced. However, newer fuels (e.g., SAF) with higher thermal stability also have higher heating values in many cases. The increased temperature rise of the fuel due to the lower flow rate may therefore be acceptable in some scenarios, and indeed, a fuel addition beyond this may also be possible. This shows how useful to take into account multiple fuel characteristics together in some implementations, rather than just a single fuel characteristic.Each implementation may use only one or more fuel characteristics. A processing module may be provided, optionally as part of the fuel property determination module 57 and / or as part of a general engine EEC, to make decisions about control of the modulation valve 2016 based on the fuel property data and optionally other data as well.A gas turbine engine 10 for an aircraft implementing this method 100 includes a bearing 2006 arranged to support the fan shaft and an oil circulation system 2000 arranged to supply oil to the bearing 2006. The heat exchange system 300 of the engine 10 includes an air-oil heat exchanger 2020 through which the oil in the oil circulation system 2000 flows, and a fuel-oil heat exchanger 1006 through which the oil in the oil circulation system and the fuel flow so that heat is transferred between the oil and the fuel. A bypass valve 2007 is provided, wherein the valve 2007 is configured such that the proportion of the oil which is conducted via a bypass pipe 2005 and thus bypasses one or more of the heat exchangers 1006, 2020 can be varied. The engine 10 also includes a fuel composition determination module 57 configured to determine at least one fuel property of the fuel to be burned by the combustion chamber 16. The bypass valve 2007 is configured to be controlled based on the at least one fuel property determined by the fuel composition determination module 57 to adjust the amount of oil directed via the bypass pipe 2005 (and thus the amount of oil flowing through each heat exchanger 1006, 2020) under cruise conditions. A controller 58 may be provided to make and implement decisions based on the output of the fuel composition determination module 57.The inventors have also realized that the use of fuels different from conventional kerosine-based jet fuels, such as sustainable flying fuels, may result in different fuel properties and that the heat transfer parameters may be adjusted in operation to take advantage of the different fuel properties. In particular, some fuels may be heated to higher temperatures than conventional fuels in one or more fuel-oil heat exchangers without substantially increasing coking. This may enable a method that provides improved oil cooling (as the fuel may draw more heat) and may also improve overall thermal efficiency of the engine as less heat is lost to the environment. The controllable heat exchange system 3000 is key to controlling the heat transfer ratio.Although cruise conditions generally account for a much greater proportion of engine operating time, the inventors have recognized that idle operation is also important. Since the fuel flow rate is much lower at idle than at cruise, even a relatively low heat load on the fuel can result in a high temperature rise. The use of non-conventional fuels may therefore have an even greater impact on optimal approaches to thermal management under idle conditions. The methods 200, 300 of FIGS. 14 and 15 are concerned with these two scenarios of aircraft operation.FIG. 14 illustrates a method 200 implementing these considerations under cruise conditions and FIG. 15 illustrates a method 300 implementing these considerations at idle, e.g., during aircraft launch, stand-by running during boarding and rolling (to a runway or hangar, or between other ground locations) or at certain time periods during flight (e.g., in descent).Considering first the method 200 to be performed in cruise flight, the method 200 is configured to be performed in a direct drive gas turbine engine 10 that includes an oil cycle system 2000 configured to supply oil to the engine components 2006; and a heat exchange system 3000 that includes an air-to-oil heat exchanger 2020 through which the oil in the oil cycle system flows; and a primary fuel-to-oil heat exchanger 1006 through which the oil in the oil cycle system and the fuel flow such that heat is transferred between the oil and the fuel. A bypass pipe 2005 is provided, the bypass pipe 2005 being arranged such that a portion of the oil can bypass one of the air-oil heat exchanger 2020 or the fuel-oil heat exchanger 1006; and a bypass valve 2007 is arranged such that the portion of the oil that is guided through the bypass pipe 2005 can be varied. The method 200 includes controlling 202 the heat exchange system 3000 such that, under cruise conditions, a heat transfer ratio of: in the range of 0 to 0.67, and optionally m is in the range of 0 to 0.60, 0 to 0.50, 0 to 0.40, 0 to 0.30, 0 to 0.20, or 0 to 0.10. A control device 58 can be provided for implementing this control.It will be appreciated that even for a particular engine 10 operating on a particular fuel, a range of values for that ratio is common in cruise flight due to different conditions. For example, an upper limit of the range may apply on a cold low altitude day (ISA-30 conditions) (cruise: 35,000 ft) with low heat buildup of the oil system, and a lower limit may apply on a hot high altitude day (ISA+40 conditions) (cruise: 39,000 ft) with high heat buildup of the oil system.The method 200 may further include receiving data 204 to enable calculation or derivation of the heat transfer ratio, e.g., temperature data (of the oil and / or fuel at one or more points around an oil circulation system 2000 or fuel flow path and / or optionally fuel tank temperature or oil tank temperature) and fuel flow rate data. Such data may be received 204 and used 206 by a controller 58 to adjust the controller 202 of the cruise heat exchange system 3000 to maintain the heat transfer ratio at a desired level or within the desired limits. This checking and adjustment / correction 206 may be performed at regular intervals or in response to predetermined stimuli (e.g., a change in temperature or flow rate or a change in engine operation or altitude). These steps 204, 206 may alternatively also be considered part of controlling 202 the heat exchange system 3000 (i.e., not separate steps). The method 200 may also be configured to use other information, e.g., temperature data (of the oil, fuel, and / or ambient temperature of an environment around the aircraft 1), flow rate data (of the oil and / or fuel), and / or one or more fuel characteristics, in determining 206 which control actions to take.The step of controlling 202 the heat exchange system 3000 may comprise one or any combination of the examples provided for the control step 104 of the method 100 of FIG. 13, for example by reducing the amount of oil routed via the at least one air-oil heat exchanger 2020, by increasing the amount of oil routed via the bypass pipe 2005 if the heat transfer ratio is too high, or by adjusting the proportion of oil and / or fuel routed via a corresponding bypass pipe 1005 or return pipe 6011.In implementations where the engine 10 includes a refrigeration cycle device 1007, the step of controlling 202 the heat exchange system 3000 may include using the refrigeration cycle device 1007 to provide thermal lift by transferring further heat from the oil to the fuel that exceeds the heat transferred as it passes through the fuel-oil heat exchanger / exchangers 1004, 1006, in some cases such that the fuel temperature is raised above the oil temperature. In such implementations, a lower value for the heat transfer ratio may be achieved than in implementations without the refrigeration cycle device 1007. For example, the cruise heat transfer ratio may be no more than 0.40, and optionally no more than 0.38, and also optionally no more than 0.10, or 0.05. In implementations where there is no refrigeration cycle device 1007 or where such a device is present but not in use, a higher value of cruise heat transfer ratio may be maintained, for example, greater than 0.38 and optionally greater than 0.40 or 0.50.The fuel temperature when reaching the combustion chamber 16 may also be taken into account. The method 200 may include controlling 202 the heat exchange system 2000 under cruise conditions such that the heat transfer ratio is in the range of 0 to 0.2 if the fuel temperature entering the combustion chamber 16 is at least 160° C., or in the range of 0 to 0.1 if the fuel temperature entering the combustion chamber 16 is at least 180° C., or in the range of 0 to 0.45 if the fuel temperature entering the combustion chamber is at least 140° C. Therefore, one or more temperature sensors 1009 may be used, the output(s) of which may be considered in adjusting the controller 202 of the heat exchange system 2000.The method 200 may also consider the type of fuel-e.g., whether the fuel is or comprises a persistent flight fuel (SAF). For example, the cruise heat transfer ratio may be maintained in a range of from 0 to 0.2 if the fuel is at least 70% SAF or from 0 to 0.1 if the fuel is at least 80% SAF. One or more fuel properties - optionally determined by a fuel property determination module 57 as described above - may therefore also be used; the information may be taken into account in adapting the controller 202 of the heat exchange system 2000.In some implementations, no more than 20% of the oil-removed heat may be dissipated to the air during cruise, and the oil-to-air heat transfer rate may be maintained in the range of 0 to 100 kJ per kilogram of fuel during cruise conditions.In some implementations, at least 80% of the oil-dissipated heat during cruise may be transferred to the fuel, and the oil-to-fuel heat transfer rate may be maintained in the range of 110-200 kJ per kilogram of fuel during cruise conditions.The percentages of total heat dissipated from the oil in heat exchange system 3000 that is transferred to the fuel (rather than being transferred to the air / environment) under cruise conditions for a particular direct drive engine 10 are plotted in FIG. 18 as a function of the fuel temperature limit. The fuel temperature limit is the highest fuel temperature that is considered safe for the operation of the aircraft for this fuel and engine 10, and may depend on fuel properties such as thermal stability and also on the thermal compatibility of the engine components. Diagram 3 shows a lower limit 4 (black dotted line) and an upper limit 5 (black dotted line) for the heat fraction transferred to the fuel. The upper and lower limits 4, 5 take into account different cruise fuel flows (which depend, inter alia, on flight altitude, rating, etc.), variations in engine heat generation, aircraft fuel temperature, atmospheric temperature, and other variables. Generally speaking, for the example engine 10 to which this data is provided, for a fuel having a temperature limit of 120° C., the lower limit may be 20% of the heat transferred to the fuel, with an upper limit being the total heat lost from the oil transferred to the fuel (100%). Graph 3 also shows a line 6 (gray solid line) for typical heat transfer to fuel with a fuel temperature limit. A typical value for the % heat transferred to the fuel for a fuel with a temperature limit of 120° C. may be 60%; for a fuel temperature limit of 170° C. and above, this value may increase to 100%. When the fuel temperature limit is equal to or higher than 260 °C, all the heat in cruise may be transferred to the fuel during cruise without requiring air cooling, resulting in a zero heat transfer ratio.Turning now to method 300 to be performed at idle, method 300 is again configured to be performed in a direct drive gas turbine engine 10 that includes an oil cycle system 2000 configured to supply oil to engine components 2006; and a heat exchange system 3000 that includes an air-to-oil heat exchanger 2020 through which the oil in the oil cycle system flows; and a fuel-to-oil heat exchanger 1006 through which the oil in the oil cycle system and the fuel flow such that heat is transferred between the oil and the fuel. A modulation valve 2016 is in turn provided to control the portion of the oil that is routed via each branch of the oil circulation system 2000. A bypass pipe 2005 is provided, the bypass pipe 2005 being arranged such that a portion of the oil can bypass one of the air-oil heat exchanger 2020 or the fuel-oil heat exchanger 1006; and a bypass valve 2007 is arranged such that the portion of the oil that is guided through the bypass pipe 2005 can be varied. The method 300 includes controlling 302 the heat exchange system 2000 such that when the aircraft 1 is operated under idle conditions, a heat transfer ratio of: is in the range of 0 to 1.5. Therefore, the ratio may be higher at idle than at cruise. The ratio can be higher in ground idling than in flight idling. Optionally, the method 300 may include maintaining the heat transfer ratio under idle conditions under idle conditions below 1.0, in the range of 0 to 0.60, 0 to 0.50, 0 to 0.40, 0 to 0.30, 0 to 0.20, or 0 to 0.10.The method 300 may further include receiving data 304 to enable calculation or derivation of the heat transfer ratio, e.g., temperature data (of the oil and / or fuel at one or more points around an oil circulation system 2000, 2000' or fuel flow path and / or optionally fuel tank temperature or oil tank temperature) and fuel flow rate data. Such data may be received 304 and used 206 by a controller 58 to adjust the modulation 202 of the idle heat exchange system 2000 to maintain the heat transfer ratio at a desired level or within the desired limits. This checking and adjustment / correction 206 may be performed at regular intervals or in response to predetermined stimuli (e.g., a change in the temperature or flow rate of the fuel or oil, or a change in aircraft operation such as the start of coasting). These steps 304, 206 may alternatively also be considered part of controlling 302 the heat exchange system 2000. The method 300 may also be configured to use other information, e.g., temperature data (of the oil, fuel, and / or ambient temperature of an environment around the aircraft 1), flow rate data (of the oil and / or fuel), and / or one or more fuel characteristics, in determining 206 which control actions to take.The step of controlling 302 the heat exchange system 2000 may comprise any or any combination of the examples provided for the modulation step 204 of the method 200 of FIG. 14, e.g. by increasing the amount of oil directed via the at least one air-oil heat exchanger 2020 when the heat transfer ratio is too low, or by adjusting the proportion of oil and / or fuel directed via a corresponding bypass pipe 1005 or return pipe 6011.In implementations where the engine 10 includes a refrigeration cycle device 1007, the step of controlling 302 the idle heat exchange system 2000 may include using the refrigeration cycle device 1007 to provide thermal lift by transferring further heat from the oil to the fuel that exceeds the heat transferred as it passes through the fuel-oil heat exchanger / exchangers 1006, in some cases such that the fuel temperature is raised above the oil temperature. In such implementations, a lower value for the heat transfer ratio may be achieved than in implementations without the refrigeration cycle device 1007 - for example, the idle heat transfer ratio may be no more than 0.40, and optionally no more than 0.30, 0.20, and further optionally no more than 0.10, or 0.05. In implementations where no refrigeration cycle device 1007 is present or where such a device is present but not in use, a higher value of the heat transfer ratio may be maintained, for example, in the range of 0.38 to 1.2.The fuel temperature when reaching the combustion chamber 16 may also be taken into account. The method 300 may include controlling 302 the heat exchange system 2000 during idle operations such that the heat transfer ratio is in the range of 0.3 to 1.5 if the fuel temperature entering the combustion chamber 16 is below 180° C., or in the range of 0 to 0.3 if the fuel temperature entering the combustion chamber 16 is at or above 180° C. Therefore, one or more temperature sensors may be used, the output(s) of which may be taken into account in adjusting the controller 202 of the heat exchange system 3000.The method 300 may additionally or alternatively consider the type of fuel-e.g., whether the fuel is or includes a persistent flight fuel (SAF). One or more fuel properties - optionally determined by a fuel property determination module 57 as described above - may therefore also be used; the information may be taken into account in adapting the controller 302 of the heat exchange system 3000.In some implementations, no more than 20% of the oil-removed heat may be dissipated to the air during cruise, and the oil-to-air heat transfer rate may be maintained in the range of 0 to 180 kJ per kilogram of fuel during cruise conditions.In some implementations, at least 80% of the oil-dissipated heat during cruise may be transferred to the fuel, and the oil-to-fuel heat transfer rate may be maintained in the range of 100-300 kJ per kilogram of fuel during cruise conditions. At idle, a higher percentage of heat release to the air is generally used than at cruise and other higher power engine conditions.As with the method 200 described with respect to FIG. 14, in some implementations, the heat exchange system 3000 further includes branching fuel return paths and at least one valve that controls a split of the fuel flow, the branching paths arranged to return fuel from the heat exchange system 3000 to at least one different location along a main fuel path leading from where the fuel enters the gas turbine engine 10 to the combustor 16. The valve may be controlled based on feedback from one or more temperature sensors and / or based on fuel characteristics. The method 300 may include adjusting the flow of fuel along each branch based on the heat transfer ratio and / or the temperature of the fuel as it exits the heat exchanger(s) 1004, 1006.The inventors have also realized that the use of fuels different from conventional kerosine-based jet fuels, such as persistent flying fuels, can result in different fuel properties and that parameters under cruise conditions can be adjusted to take advantage of the different fuel properties. In particular, some fuels in the fuel-oil heat exchangers 1004, 1006 may be heated to higher temperatures than conventional fuels. This may improve cooling of the oil before it is recirculated to the remainder of the turbine engine and / or may improve combustion efficiency of the fuel. Using the fuel to remove more heat from the oil instead of relying on heat transfer from the oil to the environment / air (e.g., in an oil-to-air heat exchanger) provides a more heat efficient turbine engine. Additionally, the improved cooling of the oil may in turn improve the cooling effect of the oil on the components of the engine through which it flows.FIG. 16 illustrates an example method 400 for operating a gas turbine engine 10.Step 402: Providing fuel to the gas turbine engine 10 via the fuel management system 1500.As discussed with respect to FIG. 6C, the fuel management system 1500 includes a primary fuel-oil heat exchanger 1006 and a secondary fuel-oil heat exchanger 1004 configured to transfer heat to or from the fuel and includes a portion of the heat exchange system 3000. Positioned downstream of the primary fuel-oil heat exchanger 1004 and upstream of the secondary fuel-oil heat exchanger 1006 is an engine fuel pump 1003. The fuel management system 1500 is configured such that the fuel reaches the primary fuel-oil heat exchanger 1006 prior to the secondary fuel-oil heat exchanger 1004.Step 404: Control the heat exchange system 3000 to increase the fuel temperature to at least 120° C. upon exiting the secondary fuel-oil heat exchanger 1004 during cruise conditions.Controlling 404 the heat exchange system 3000 may include controlling fuel flow through the heat exchangers - for example, recirculating a variable portion of fuel through the primary fuel-oil heat exchanger 1006 and / or allowing a variable portion of fuel to bypass the primary fuel-oil heat exchanger 1006. While FIG. 6 shows an implementation with a bypass pipe 1005, but without recirculation, and FIGS. 7A-7C show implementations without bypass pipe (at least not for the primary heat exchanger-FIGS. 6A and 6C provide a path through which a portion of the fuel reaching the combustion chamber 16 may bypass the secondary heat exchanger while the remainder of the fuel passes through it), but with a recirculation pipe 6011, it should be appreciated that one or more bypass pipes and / or one or more recirculation pipes may be provided together in various implementations. Further, in some systems, the flow direction within the same pipe may be reversible so that it may be used as bypass pipe 1005 or return pipe 6011. It is understood that the figures are provided by way of example only and are not intended to be limiting.Controlling 404 the heat exchange system 3000 may include controlling the flow of fuel through the heat exchangers such that, for example, a variable fraction of the fuel may bypass the primary fuel-oil heat exchanger 1006 or the secondary fuel-oil heat exchanger 1006. One or more controllable valves or pumps and optionally one or more sensors may be provided accordingly to enable control of the fuel flow.Additionally or alternatively, the oil flow, as well as the fuel flow, may be controlled using one or more bypass valves 2007 configured to control the flow through one or more bypass tubes 2005, if present, such that the oil may bypass, rather than flow through, one or more heat exchangers 1004, 1006, 2020. The oil may also be recirculated in some implementations and / or the oil flow rate may be adjusted by controlling one or more oil pumps. One or more controllable valves and / or pumps and optionally one or more sensors may be provided accordingly to enable control of the oil flow.A controller 58 may be used to actively control the fuel and / or oil flows through and around the heat exchangers 1004, 1006, 2020. Controller 58 may be part of or provided by an EEC, or it may be a separate unit. The control can be automated, for example by an EEC. The controller may receive inputs from one or more temperature sensors 1009, and may control one or more valves (e.g., recirculation valve 6010 or bypass valve 2007) and / or pump 1003 based on the received data. One or more oil supply pumps and / or scavenge pumps may also be controlled via the controller.Fuel flow may be controlled using a recirculation valve 6010 as described above. Alternatively or additionally, the fuel flow may be controlled using one or more bypass pipes 1005 (as shown in FIG. 5 ) arranged to allow a portion of the fuel not to flow through one or both heat exchangers and / or by adjusting the % split of fuel flowing from the primary fuel-oil heat exchanger 1006 directly to the combustion chamber 16 versus the % of fuel flowing from the primary fuel-oil heat exchanger into the secondary fuel-oil heat exchanger 1004 (and optionally then back to flow together with the fuel flowing into the combustion chamber 16, as implemented).By heating the fuel to higher temperatures than previously used, the oil may be better cooled before being recirculated to the remaining turbine engine and / or the combustion efficiency of the fuel may be improved. Positioning the secondary fuel-oil heat exchanger 1004 downstream of the engine fuel pump 1003 may help achieve higher fuel temperatures without compromising fuel pump longevity.The auxiliary system 7010 supplied by the secondary fuel-oil heat exchanger 1004 may benefit from higher fuel temperatures than are suitable for use in the combustion chamber 16 or in components through which the main fuel flow path passes. Therefore, it may be desirable that the temperature of the fuel upon exiting the secondary fuel-oil heat exchanger 1004 be higher than the temperature of the fuel upon entering the combustion chamber 16. Therefore, the oil temperature may be higher when passing through the secondary fuel-oil heat exchanger 1004 than the other heat exchangers 1006, 2020. Controlling 404 the heat exchange system 3000 may include controlling a bypass valve for fuel to adjust the proportion of fuel directed via a bypass pipe 1005 to control the flow of fuel through each of the primary fuel-oil heat exchangers 1006 and secondary fuel-oil heat exchangers 1004 during cruise conditions.Controlling 404 the heat exchange system may include controlling the bypass valve 2007 to adjust the amount of oil directed through a bypass pipe 2005 to control the flow of oil through each of the primary fuel-oil heat exchangers 1006 and secondary fuel-oil heat exchangers 1004 under cruise conditions.Controlling 404 the heat exchange system 3000 may include controlling the heat exchange system 3000 to increase the temperature of the fuel such that the fuel temperature upon exiting the secondary heat exchanger is in the range of 120° C. to 180° C., in the range of 135° C. to 200° C., 135° C. to 180° C., 150° C. to 200° C., 150° C. to 180° C., 150° C. to 170° C., or about 120° C., 130° C., 140° C., 150° C., 160° C., or up to 200° C.The inventors have recognized that it may adversely affect engine performance and even become dangerous when improper fuels reach these higher temperatures. Excessive thermal decomposition of a fuel / deposit can lead to safety issues by blocking passageways and causing malfunctions of components (e.g., jamming valves and clogged nozzles, orifices, valves, etc.), which can result in loss of thrust control. The method 500 illustrated in FIG. 17 is therefore provided as a safety measure.The method 500 illustrated in FIG. 17 may be performed in a gas turbine engine 10 substantially as described for the foregoing methods with respect to oil flow, however, one or more controllable oil valve(s) 2007, 2007 aare arranged such that a portion of the oil may bypass (and later be directed across) at least one of the air-oil heat exchangers 2020, the primary fuel-oil heat exchanger 1006 and the secondary fuel-oil heat exchanger to be varied are replaced by at least one controllable oil or air valve 2007, 2007a, 2007b, 2022 arranged such that at least one of the oil flow rate and air flow rate may be varied by at least one of the air-oil heat exchangers 2020 and the fuel-oil heat exchanger 1006. Therefore, air flow control or oil flow control or both may be provided. In implementations where a controllable air valve 2022 is provided (optionally in addition to a controllable oil valve 2007, 2007a, 2007b), the engine may require a temperature sensor 1009 arranged to provide an indication of fuel temperature - optionally upon entry into the combustion chamber 16. This temperature sensor 1009 is generally located at or near an inlet to the combustion chamber 16, and more particularly, near a nozzle inlet of the combustion chamber 16, to directly sense the fuel temperature as it enters the combustion chamber, although it will be appreciated that variations in fuel temperature between the outlet of the primary fuel-oil heat exchanger 1006 and the combustion chamber 16 are likely small when bypassing the secondary fuel-oil heat exchanger 1004, and are calculable from knowledge about the engine 10 when the fuel temperature is desired elsewhere than the sensor 1009. In some implementations where the fuel temperature is used upon entering the combustion chamber 16, the sensor 1009 may be arranged differently and the fuel temperature upon entering the combustion chamber may be calculated based on the temperature output and knowledge about the engine 10.The method 500 includes determining 502 whether the fuel temperature has increased above a set threshold during cruise conditions based on an output of the temperature sensor 1009. For example, a direct comparison may be made between the data received from temperature sensor 1009 and one or more temperature thresholds stored in memory. This determination 502 may therefore be performed automatically, optionally by a controller 58 (which may be or include a dedicated processing module or may be provided by a more general EEC).The method 500 further comprises, responsive to determining that the fuel temperature has increased above the set threshold during cruise conditions, controlling 504 the at least one valve 2007, 2007 a, 2022 to change the at least one flow rate (air or oil flow rate) through the at least one heat exchanger 1004, 1006, 2020 to decrease the fuel temperature. For example, if an oil valve 2007, 2007 ais adjusted, a greater proportion of the oil can be directed via bypass pipes 2005, 2005 ato bypass the primary fuel-oil heat exchanger 1006 and / or the secondary fuel-oil heat exchanger 1004 and re-unite with the remaining oil that has flowed through the fuel-oil heat exchanger / s 1004, 1006 at or downstream of the oil outlet from this heat exchanger 1006, 1004, as illustrated in FIG. 12. In a parallel configuration (not shown), reducing the oil flow through the fuel-oil heat exchanger 1006 by controlling a modulation valve may automatically increase the oil flow through the air-oil heat exchanger 2020 (unless a bypass pipe is present around the air-oil heat exchanger 2020 and is used for balancing).As indicated by the dashed line in FIG. 17, this method 500 may be repeated-for safety reasons, frequent checks may be preferable, optionally at regular intervals and / or when triggered by a change in engine operation (e.g., a change in flight altitude or thrust demand). The oil valve 2007, 2007 aor the air valve 2022 can therefore be actively controlled 504 in order to vary the oil or air flow rate through the respective heat exchanger 1006, 2020. The active control of the valve 2007, 2007 a, 2022 can be automated and implemented by the control unit 58. For example, the active control of the bypass valve 2007 associated with the secondary fuel-oil heat exchanger 1004 may be automated and implemented by a controller 58 of the heat exchange system 3000, which may be a dedicated controller or part of a more general EEC. In various implementations, the control of one or more of the valves 2007, 2007a, 2022may be a closed or open loop depending on the availability of inline measurements-in particular, a closed loop approach may be preferable for control when feedback data (e.g., from oil flow rate sensors) is available.In implementations with an air valve 2022 and an oil valve 2007, 2007 a, the air flow through the air-to-oil heat exchanger 2020 can be increased as more oil is directed over the air-to-oil heat exchanger 2020; the increase can be made linearly with increasing oil flow rate. With respect to the percentage of airflow, the increase may be large--for example, when transitioning from a fully closed to a fully open air valve 2022, or from a first position at 10% of the maximum air flow rate to 90% or 100% of the maximum air flow rate when the fuel temperature is deemed to be too high. The method 500 may include, responsive to determining 502 that the fuel temperature has increased above a set threshold during cruise conditions, controlling 504 aof the air valve 2022 to direct more air through the air-to-oil heat exchanger 2020, and controlling 504 bof the oil valve 2016, 2007 to direct less oil through the fuel-to-oil heat exchanger 1006. The control / adjustment 504 aof the air valve 2022 may be performed simultaneously with the control / adjustment 504 bof the oil valve 2016, 2007.In some implementations, the flow of air through the air-to-oil heat exchanger 2020 may not be actively controlled. In such implementations, the air flow is likely to naturally change under different operating conditions based on the pressure ratio between the heat exchanger air inlet 2020 aand the heat exchanger exhaust / air outlet 2020 b. In this configuration, even if the oil-side air-oil heat exchanger 2020 is completely bypassed / the oil flow rate through the heat exchanger 2020 is zero, the air continues to flow through the air passages of the air-oil heat exchanger 2020. In alternative implementations, active control of the air flow through the air-oil heat exchanger 2020 is provided. For example, the air-to-oil heat exchanger 2020 and the fuel-to-oil heat exchanger / exchangers 1004, 1006 may be connected in series without bypass on the oil side of the air-to-oil heat exchanger 2020, but with an air valve 2022 on the air side that controls the air flow (both an oil bypass valve 2007 band an air flow valve 2022 may be provided in other implementations). In such an in-line arrangement, the air-to-oil heat exchanger 2020 may be disposed upstream or downstream of the fuel-to-oil heat exchanger 1006 with respect to the oil stream. In series arrangement implementations, it may be helpful to place the air-to-oil heat exchanger 2020 upstream of the fuel-to-oil heat exchanger(s) 1004, 1006 (as opposed to FIG. 12 ) with respect to the oil flow in order to avoid overheating of the fuel - the oil may be cooled to any desired extent by increasing the air flow to a predetermined threshold value. It should be appreciated that without a particular mechanism for actively driving the airflow, such as a gas-gas ejector or a particular fan, cooling of the oil is limited on a single pass through the air-oil heat exchanger (i.e., once the air valve 2022 is fully open, no further control adjustments may be made to increase cooling on a single pass through the air-oil heat exchanger 2020). The oil may be further cooled by being recirculated through the air-oil heat exchanger 2020 before reaching the fuel-oil heat exchanger 1006.In some implementations, a bypass pipe 2005 band a corresponding control valve 2007 bmay be implemented only for the air-oil heat exchanger 2020. Optionally, in the described method 500, only the oil flow rate through the air-oil heat exchanger 2020 may be actively controlled.In various implementations:A bypass pipe 2005 bis provided only on the oil side of the air-oil heat exchanger 2020 (no oil bypass of the fuel-oil heat exchanger and no active air flow control);A bypass pipe 2005, 2005 a, 2005 bmay be provided for each of the air-oil heat exchanger 2020, the primary fuel-oil heat exchanger 1006, and the secondary fuel-oil heat exchanger 1004 (no active air flow control); orThe air flow rate through the air-oil heat exchanger 2020 may be controllable and no bypass pipes for oil may be present.Generally, control over at least one flow through the air-oil heat exchanger 2020 may be considered more important than control over one of the flows through the fuel-oil heat exchanger / s 1004, 1006. However, one or more oil bypass pipes 2005, 2005 aand / or a fuel bypass pipe 1005 (and corresponding control valve) may be provided to the fuel-oil heat exchanger / s 1004, 1006 in some implementations, optionally in addition to one or more further controllable valves.A combination of air flow control and oil flow control may be implemented in both parallel and series arrangements. However, it should be appreciated that while all of the various control options may be used together, increasing the complexity and number of components, and thus the increased weight of the overall heat exchange system 3000, may not be desirable. Thus, in many implementations, selection of a subset of the control options may be preferred - for example, either air flow control or oil flow control may be used to adjust heat transfer within the air-oil heat exchanger 2020, such that control of both may be deemed superfluous for a particular heat exchanger 2020. Similarly, in an implementation where the oil flows through the air-oil heat exchanger 2020 before passing through the fuel-oil heat exchanger 1006, the oil temperature adjustment in the air-oil heat exchanger 2020 (by controlling the air and / or oil flow in that heat exchanger) may be used as an alternative to oil flow adjustment by one or more of the primary fuel-oil heat exchangers 1006 or secondary fuel-oil heat exchangers 1004 to change the amount of heat transferred to the fuel in the fuel-oil heat exchanger(s) 1004, 1006.Additionally or alternatively, in implementations where the heat exchange system 3000 includes a refrigeration cycle device 1007 configured to transfer additional heat from the oil to the fuel that exceeds that transferred from the primary fuel-oil heat exchanger 1006 and secondary fuel-oil heat exchanger 1004, the method 500 may further include controlling the refrigeration cycle device 1007 to reduce the amount of additional heat transferred to the fuel when it is determined 502 that the fuel temperature has increased above the set threshold during cruise conditions. This control may be performed by decreasing the flow of oil through the refrigeration cycle device 1007, or by decreasing the performance of the refrigeration cycle device 1007 (e.g., by decreasing the flow rate of the refrigerant), or by deactivating the refrigeration cycle device 1007. The set threshold for deactivation of the refrigeration cycle device 1007 may be lower than the set threshold for changing a flow rate of oil and / or air through a heat exchanger 1006, 2020.The threshold established for the step of controlling 504 may be in the range of 140° C. to 300° C., and optionally 200° C. to 300° C., and further optionally 250° C. to 300° C. The predetermined threshold may be, for example, 140° C., 150° C., 160° C., 170° C., 180° C., 190° C., 200° C., 210° C., 220° C., 230° C., 240° C., 250° C., 260° C., 270° C., 280° C., 290° C., or 300° C.The method 500 of some implementations further comprises determining 501 the set threshold value to which the output of the temperature sensor is compared. This determination 501 can be carried out on the wing and optionally in flight. The determination 501 may include calculating a threshold or identifying an appropriate threshold from a prestored set of thresholds stored in the memory.The inventors have recognized that knowledge of one or more fuel properties may be utilized in the determination 501 to ensure safety while still utilizing the advantages of the different properties of flying fuels. The determination 501 may therefore be based on at least one fuel property of the fuel. The at least one fuel property may be or include one of the examples discussed above, e.g., the thermal stability of the fuel, the nitrogen content of the fuel, the sulfur content of the fuel, and / or the Sustainable Flight Fuel, SAF, content of the fuel (% SAF). The at least one fuel property may be determined by one or more of the approaches discussed above.The step of determining 501 the established threshold value may comprise increasing the established threshold value, optionally linearly, with increasing thermal stability of the fuel. The fixed threshold used to control the thermal modulation may therefore be increased linearly with the thermal stability of the fuel. Thermal stability is the temperature at which the fuel begins to decompose and form deposits which can lead to blockages and malfunctions of components. It will be appreciated that this is generally not an immediate change in reaching a particular temperature; aviation fuels include various components that can decompose at different temperatures and residence time at high temperatures is also a factor for fuel degradation. The threshold may be set based on a comparison of the maximum allowable deposition rate from fuel split with the thermal stability of the fuel used. A fuel may be deemed stable at a particular temperature if its decomposition rate at that temperature is below a particular threshold. Method 500 may also include using a clock or timer or fuel flow rate sensor - a higher fuel temperature may be acceptable, for example, at a higher pump rate because the faster moving fuel spends less time at that temperature before being burned. Alternatively or additionally, method 500 may include using one or more pressure sensors to determine the fuel flow rate. By measuring a pressure drop through a known orifice size, the flow rate can be determined.Alternatively or additionally, the step of determining 501 the fixed threshold may comprise increasing the fixed threshold as the SAF content of the fuel increases (optionally linearly) for fuels having a SAF content of over 70%. As discussed above, it has been found that increasing SAF content in fossil aviation fuel increases thermal stability, but not linearly. Therefore, predetermined discrete SAF content-based thresholds may be preferable to continuously change the predetermined threshold in proportion to the SAF content, or a more complex (non-linear) but continuous relationship may be established.Alternatively or additionally, the step of determining the fixed threshold 501 may comprise decreasing the fixed threshold as the amount of heteroatomic species of the fuel increases (optionally linearly). For example, it has been found that the thermal stability of Jet-A decreases with increasing nitrogen content (the nitrogen content being a measure of the amount of nitrogenous species present). It is known that the interaction of sulphur- and nitrogen-containing species in the fuel can make an important contribution to the decomposition rate of the fuel. Thus, consideration of the composition of the fuel with respect to multiple hetero-atomic species may be implemented to account for these interactions.In some implementations, particularly implementations where a direct measure of the thermal stability of the fuel is not available, multiple fuel properties may be determined and used in valve control decisions. The one or more determined fuel properties may actually be converted to a measure of thermal stability.It should be understood that the invention is not limited to the above-described embodiments, and that various modifications and improvements may be made without departing from the concepts described herein. Except in the case of mutual exclusion, each of the features may be employed separately or in combination with any other features, and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.

Claims

A method (300) of operating a gas turbine engine (10), the gas turbine engine (10) comprising: an engine core (11) comprising a turbine (19), a compressor (14), a combustor (16) arranged to combust a fuel, and a core shaft (26) connecting the turbine to the compressor; a fan (23) arranged upstream of the engine core (11); a fan shaft; at least one bearing (2006) arranged to support the fan shaft; an oil circulation system (2000) arranged to supply the at least one bearing with oil; and a heat exchange system (3000) comprising: an air-to-oil heat exchanger (2020) through which the oil in the oil circulation system flows; a fuel-oil heat exchanger (1006) through which the oil in the oil circulation system and the fuel flow so as to transfer heat between the oil and the fuel; a bypass pipe (2005) disposed so that a portion of the oil can bypass one of the air-oil heat exchanger or the fuel-oil heat exchanger; and a bypass valve (2007) configured such that the proportion of the oil that is passed through the bypass pipe (2005) can be varied, the method (300) comprising controlling (302) the heat exchange system such that, under idle conditions, a heat transfer ratio of: W a%0020̈ rme u%0020̈̈ transfer rate of O%0020̈ l to air (kJkg-1) W a%0020̈rme transfer rate of O%0020̈ l to fuel (kJkg-1) is in the range of 0 to 1.5.The method (300) of claim 1, comprising controlling (302) the heat exchange system (3000) such that the heat transfer ratio is below 1.0 under idle conditions.The method (300) of claim 1, wherein the bypass pipe (2005) is arranged such that a portion of the oil can bypass the air-oil heat exchanger (2020) such that the oil flowing through the bypass pipe (2005) flows through the fuel-oil heat exchanger without flowing through the air-oil heat exchanger (2020), and wherein controlling (302) the heat exchange system to adjust the heat transfer ratio comprises increasing the amount of oil flowing through the bypass pipe (2005) when the heat transfer ratio is too high.The method (300) of claim 1, wherein the heat exchange system (3000) comprises at least one return pipe (6011) arranged such that a liquid can pass multiple times through a heat exchanger (1006, 2020), and wherein controlling (302) the heat exchange system to adjust the heat transfer ratio comprises modulating the amount of liquid routed via the return pipe (6011).The method (300) of claim 1, wherein the heat exchange system (3000) further comprises a refrigeration cycle device (1007), and controlling (302) the heat exchange system comprises using the refrigeration cycle device to provide thermal buoyancy by transferring further heat from the oil to the fuel such that the fuel temperature is raised above the oil temperature; optionally wherein the heat exchange system (3000) is controlled such that the heat transfer ratio is in the range of 0 to 0.4.The method (300) of claim 1, wherein the heat exchange system (3000) is not configured to provide thermal buoyancy, and wherein the heat exchange system (3000) is controlled such that the heat transfer ratio is in the range of 0.38 to 1.2.The method (300) of claim 1, wherein the method (300) comprises: controlling (302) the heat exchange system (3000) under idle conditions such that the heat transfer ratio is in the range of 0.3 to 1.5 if the fuel temperature entering the combustion chamber (16) is below 180°C; or controlling (302) the heat exchange system (3000) under idle conditions such that the heat transfer ratio is in the range of 0 to 0.13 if the fuel temperature entering the combustion chamber (16) is above 180°C.The method (300) of claim 1, wherein under idle conditions, the method (300) comprises: controlling (302) the heat exchange system (3000) such that the heat transfer ratio is in the range of 0 to 0.2 if the fuel is at least 70% sustainable aviation fuel.The method (300) of claim 1, the method comprising: controlling (302) the heat exchange system (3000) such that, at idle conditions, the heat transfer rate of oil to air is maintained in the range of 0 to 180 kJ per kilogram of fuel, wherein no more than 20% of the heat dissipated from the oil at idle conditions is transferred to the air; or controlling (302) the heat exchange system (3000) such that, at idle conditions, the heat transfer rate of oil to fuel is maintained in the range of 100 to 300 kJ per kilogram of fuel, wherein at least 80% of the heat dissipated from the oil at idle is transferred to the fuel.A gas turbine engine (10) for an aircraft, comprising: an engine core (11) comprising a turbine (19), a compressor (14), and a core shaft (26) connecting the turbine to the compressor; a fan (23) disposed upstream of the engine core; a fan shaft; at least one bearing disposed to support the fan shaft; an oil circulation system (2000) disposed to supply the at least one bearing with oil; and a heat exchange system (3000) comprising: an air-oil heat exchanger (2020) through which the oil in the oil circulation system flows; a fuel-oil heat exchanger (1006) through which the oil in the oil circulation system and the fuel flow such that heat is transferred between the oil and the fuel; a bypass pipe (2005) arranged such that a portion of the oil can bypass one of the air-oil heat exchanger or the fuel-oil heat exchanger; and a bypass valve (2007) arranged such that the portion of the oil that is passed through the bypass pipe (2005) can be varied, and wherein the heat exchange system (3000) is arranged such that, under idling conditions, a heat transfer ratio of: W a%0020̈ rme u%0020̈̈ transfer rate of O%0020̈ l to air (kJkg-1) W a%0020̈ rmetransfer rate of O%0020̈ l to fuel (kJkg-1) is in the range of 0 to 1.5.The gas turbine engine (10) of claim 10, wherein: the turbine (19) is a first turbine, the compressor (14) is a first compressor, and the core shaft (26) is a first core shaft; the engine core further comprises a second turbine (17), a second compressor (15), and a second core shaft (27) connecting the second turbine to the second compressor; and the second turbine, the second compressor, and the second core shaft are arranged to rotate at a higher speed than the first core shaft.The gas turbine engine (10) of claim 10, wherein the air-oil heat exchanger (2020) and the fuel-oil heat exchanger (1006) are arranged in series in the oil circulation system (2000).The gas turbine engine (10) of claim 10, wherein the heat exchange system (3000) is configured to be controlled such that the heat transfer ratio is below 1.0 under idle conditions.The gas turbine engine (10) of claim 10, wherein: the bypass pipe (2005) is arranged such that a portion of the oil can bypass the air-oil heat exchanger (2020) such that the oil flowing through the bypass pipe (2005) flows through the fuel-oil heat exchanger without flowing through the air-oil heat exchanger (2020).The gas turbine engine (10) of claim 10, wherein the core shaft (26) outputs drive directly to the fan (23) to drive the fan at the same speed as the core shaft such that the engine (10) is a direct drive turbine engine.

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