HEAT EXCHANGE SYSTEM

Abstract

A method 600 for operating a gas turbine engine comprising an engine core including a turbine, a compressor, a combustor arranged to combust a fuel, and a core shaft connecting the turbine to the compressor; a fan arranged upstream of the engine core; and a gearbox receiving input from the core shaft and outputting drive to the fan; an oil circulation system arranged to supply oil to the gearbox; and a heat exchange system comprising: an air-to-oil heat exchanger through which the oil flows; and a fuel-to-oil heat exchanger through which the oil and fuel flow, and wherein the oil circulation system includes a bypass pipe arranged to allow oil to bypass a heat exchanger;and a bypass valve arranged to vary the proportion of oil passed through the bypass tube, wherein the method 600 comprises controlling 602 the bypass valve so that under idle conditions a heat transfer ratio ; Heat transfer rate from oil to air Heat transfer rate from oil to fuel in the range of 0.67 to 5.67.

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;and a transmission receiving input from the core shaft and outputting drive to the fan;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 circuit system is filtered, andthe fuel flows to transfer heat between the oil and the fuel; anda modulation valve configured such that the proportion of oil directed across each heat exchanger can be varied,the method comprising:determining at least one fuel property of the fuel to be burned by the combustion chamber; andcontrolling the modulation valve based on the at least one fuel property to adjust the proportion of oil directed across each heat exchanger under 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 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 cooling of the oil while avoiding coking and / or paint and thus 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 more heat 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 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 modulation valve of such examples may be configured such that the proportion of oil directed via each branch may be varied, and the controller of the modulation valve may therefore adjust the proportion of oil directed via 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 modulation 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 modulation valve may be configured such that up to 100% of the oil may be directed via the fuel-oil heat exchanger. The modulation valve may be arranged to ensure that no less than 70%, 80% or 90% of the oil is directed across the fuel-oil heat exchanger during cruise.The modulation valve may be configured to redirect a fixed portion of the oil flow to both the fuel-oil heat exchanger and the air-oil heat exchanger during operation of the engine, the fixed portion optionally being determined at or prior to starting the engine or at or prior to reaching the cruise altitude based on the at least one particular fuel characteristic. Alternatively, the modulation valve may be arranged to divert a variable portion of the oil flow to the fuel-oil heat exchanger and the air-oil heat exchanger, respectively, 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.The gas turbine engine may further include a generator (e.g., an integrated propulsion generator) and a secondary oil circulation system configured to provide oil to the generator. The above-described oil circuit system that supplies the transmission with oil may therefore be referred to as a primary oil circuit system. The heat exchange system may comprise an oil-to-oil heat exchanger configured to transfer heat between the primary and secondary oil circuit systems. The modulation valve or an additional valve may control the flow of oil from the primary oil circulation system to the oil-to-oil heat exchanger.The heat exchange system may further include a secondary fuel-oil heat exchanger configured to receive the fuel and oil from the secondary oil circuit system. The method may include transferring heat between the oil from the secondary oil circuit system and the fuel using the secondary fuel-oil heat exchanger. The secondary fuel-oil heat exchanger may also be referred to as a secondary heat exchanger for simplicity.The fuel-oil heat exchanger through which the oil of the primary oil circulation system flows may be referred to as a primary fuel-oil heat exchanger. The primary fuel-oil heat exchanger may also be referred to as a primary heat exchanger for simplicity. The fuel may flow through the secondary fuel-oil heat exchanger before flowing through the primary fuel-oil heat exchanger such that heat is transferred from the oil in the secondary oil circuit system to the fuel before heat is transferred from the oil in the primary oil circuit system to the fuel.The heat exchange system may include at least one bypass pipe arranged to allow fuel (or oil) to 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 modulation 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. The method may include directing all oil over the fuel-oil heat exchanger for at least 15%, and optionally at least 20%, of the total time spent in cruise flight. 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 from the oil during cruise can be transferred to the fuel provided the fuel is operationally stable at temperatures above 160° 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.According to a second 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; 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 circuit system is filtered, andthe fuel flows to transfer heat between the oil and the fuel; anda modulation valve configured to vary the proportion of oil directed across each heat exchanger; 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 modulation 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 under 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 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, the gas turbine engine comprising:an engine core comprising a turbine, a compressor, a core shaft connecting the turbine to the compressor, and a combustor configured to combust a fuel;a fan disposed upstream of the engine core;an oil system configured to circulate oil; anda heat exchange system comprising at least one fuel-oil heat exchanger,which is configured to transfer heat from the oil to the fuel,wherein the method comprises:determining at least one fuel property of the fuel to be provided to the combustion chamber; andmodulating the heat exchange system to adjust the temperature of the fuel at entry into the combustion chamber in cruise, the set value based on the at least one fuel property.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 significantly increasing the formation of degradation products of the fuel, including coking and / or fouling. This can improve the combustion efficiency of the fuel. Higher fuel temperatures upon entering the combustion chamber may enable a method that provides improved oil cooling (when the fuel may draw more heat) and / or fuel combustion efficiency. For example, fuel viscosity affects how fuel is delivered to and ignited by the combustion chamber. Viscosity may affect droplet size from fuel spray nozzles, which in turn may affect combustion efficiency. Therefore, considering the fuel viscosity when supplying fuel to the combustion chamber and appropriately controlling it by varying the heat input may provide more efficient fuel combustion, thereby improving aircraft performance. In a transient period, in which the available flying fuels change over time during the lifetime of a gas turbine engine and depend, among other things, 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 is carried out on the basis of the specific fuel used in the engine. Thus, the operation of gas turbines may be adjusted to retrieve the best from a variety of fuels. The controllable heat exchange system is key to the control of fuel and oil temperatures.The fuel property may be or include the percentage of persistent flying fuel (SAF) in the fuel, the amount (e.g., ppm) of sulfur in the fuel, or the thermal stability of the fuel.The set value may be an average of at least 140° C., 150° C., 160° C., 170° C., 180° C., 190° C., or 200° C. upon entering the combustor under cruise conditions. The fuel temperatures entering the combustion chamber under cruise conditions may be defined as an average over at least 5 minutes and optionally over ten, twenty or thirty minutes under stable 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 increase in temperature. Each variation must not take more than 5 minutes. A fuel temperature of at least 140° C. upon entry into the combustion chamber under the cruise conditions defined herein therefore requires that the fuel temperature remain at or above 140° C. for a period of time and a transient temperature increase to above 140° C. is not sufficient.The set value may be above oil temperature in some implementations. 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 such that the fuel temperature is raised above the oil temperature to reach the set value.The gas turbine engine may further include a gearbox that receives input from a core shaft and outputs propulsion to the fan, optionally via a fan shaft. The oil system may be configured to circulate oil between the transmission and the at least one fuel-oil heat exchanger, and may be described as an oil circulation system configured to supply oil to the transmission.The heat exchange system may include at least one air-oil heat exchanger, and modulating the heat exchange system to adjust the fuel temperature may include modulating the amount of oil directed across the at least one air-oil heat exchanger.The heat exchange system may comprise at least one bypass pipe arranged to bypass oil past a heat exchanger. Modulating the heat exchange system to adjust the fuel temperature to the set value may include modulating the amount of oil supplied via the bypass pipe. The heat exchange system may comprise at least one bypass pipe arranged to bypass fuel past a heat exchanger. Modulating the heat exchange system to adjust the fuel temperature to the set value may include modulating the amount of fuel supplied via the bypass pipe.The heat exchange system may include at least one return pipe arranged such that a liquid (e.g., the oil or fuel) may pass through a heat exchanger multiple times. Modulating the heat exchange system to adjust the fuel temperature to the set value may comprise modulating the amount of liquid directed via the or each return tube.The step of modulating the heat exchange system may comprise controlling the flow of oil through the at least one fuel-oil heat exchanger such that between 50% and 100% of the heat lost from the oil is transferred to the fuel. The step of modulating the heat exchange system may comprise controlling the flow of oil through the at least one fuel-oil heat exchanger such that between 80% and 100% of the heat lost from the oil is transferred to the fuel.The at least one determined fuel property may be or include the thermal stability of the fuel, and the method may include modulating / controlling the heat exchange system such that the temperature of the fuel as it enters the combustion chamber during cruise flight is increased with increasing thermal stability, optionally linearly.The at least one determined fuel property may be or include the percentage of SAF in the fuel, and the method may include modulating the heat exchange system such that the fuel temperature when entering the combustion chamber during cruise is increased as % SAF increases as % SAF exceeds 60%.The at least one determined fuel property may be or include the presence of a tracer species in the fuel, and the method may include modulating the heat exchange system such that the fuel temperature upon entering the combustion chamber during cruise flight is adjusted to a predefined value corresponding to the tracer species.The at least one determined fuel property may be or include the sulfur content of the fuel, and the method may include modulating / controlling the heat exchange system such that the fuel temperature is increased when entering the combustion chamber in cruise flight, when the sulfur content is lower, and vice versa.The methods of the first and third aspects may complement 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, a core shaft connecting the turbine to the compressor, and a combustor configured to combust a fuel;a fan disposed upstream of the engine core;an oil system configured to circulate oil;a heat exchange system comprising at least one fuel-oil heat exchanger configured to transfer heat from the oil to the fuel; anda fuel composition determination module configured to determine at least one fuel property of the fuel to be burned by the combustion chamber,and wherein the heat exchange system is configured to be modulated such that the temperature of the fuel upon entering the combustion chamber is adjusted to a predetermined value, the predetermined value based on the at least one fuel property.The gas turbine engine may further include a gearbox that receives input from the core shaft and outputs propulsion to the fan. The oil system may be configured to circulate oil between the transmission and the at least one fuel-oil heat exchanger, and may be described as an oil circulation system configured to supply oil to the transmission.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 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 the first or third aspect.In each of the first to fourth aspects:At least one fuel property of the fuel may comprise 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 nitrogenous 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 the shaft speed and the mass flow rate of the fuel from which the heating value (a fuel property) may be determined, or the detected parameters may be the 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 fuel from the fuel tank for off-wing testing.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.From the performance of the gas turbine engine during coasting, takeoff, and climb of the aircraft, at least one fuel property may be inferred.According to a fifth aspect, there is provided a method of operating a gas turbine engine, 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 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 circuit system and the fuel flow so that heat is transferred between the oil and the fuel, and wherein the oil circuit system branches so that a portion of the oil can flow along each branch, and the air-oil and fuel-oil heat exchangers are arranged in a parallel configuration on different branches of the oil circuit system; anda modulation valve configured such that the proportion of oil directed via each branch can be varied,the method comprising controlling the heat exchange system such that a heat transfer ratio is established during cruise conditionsThe 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 process that provides improved oil cooling (as the fuel may draw more heat) and may also improve overall thermal efficiency / thermodynamic efficiency of the engine as less heat is lost to the environment and possibly also more energy is recovered in the thermodynamic cycle. The controllable heat exchange system is key to controlling the heat transfer ratio.While it should be understood that the ratio is dimensionless, in the examples described herein, heat transfer is measured per unit mass or volume of fuel, thus providing a heat transfer rate normalized to cruise fuel flow rate variations. For oil-to-air heat transfer, the definition "per unit mass or volume of fuel" may be set equal to "per fixed time period depending on the fuel flow rate" to also provide normalization of the 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. In most implementations, it may be assumed that any temperature rise of the fuel due to other engine components (and not due to heat transfer from the oil) is minimal. 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 gained 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 (all other sources of loss being negligible).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 heat exchange system may include at least one bypass pipe arranged to bypass oil past a heat exchanger, and controlling the heat exchange system to adjust the heat transfer ratio may include modulating the amount of oil directed through the bypass pipe.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 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 0.2 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 240 kJ per kilogram of fuel at cruise conditions and optionally 0 to 120 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 85-350 kJ per kilogram of fuel at cruise conditions and optionally 85-170 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, third, and fifth aspects may complement each other, and two or more of them 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; and a 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 circulation system disposed to supply oil to the gearbox; and 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 circuit system and the fuel flow so that heat is transferred between the oil and the fuel, and wherein the oil circuit system branches so that a portion of the oil can flow along each branch, and the air-oil and fuel-oil heat exchangers are arranged in a parallel configuration on different branches of the oil circuit system; anda modulation valve configured such that the proportion of oil directed via each branch can be varied, and wherein the heat exchange system is adapted to be controlled such that a heat transfer ratio is in the range of 0 to 0.67 under 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 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 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 for 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 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 circuit system and the fuel flow so that heat is transferred between the oil and the fuel, and wherein the oil circuit system branches so that a portion of the oil can flow along each branch, and the air-oil and fuel-oil heat exchangers are arranged in a parallel configuration on different branches of the oil circuit system; anda modulation valve configured such that the proportion of oil directed via each branch can be varied,the method comprising controlling the heat exchange system such that a heat transfer ratio is maintained under idle conditionsThe range is from 0.67 to 5.67.The heat transfer rates are defined as described above with respect to the fifth aspect. Idle operation while the aircraft is operating on the ground may be referred to as "ground idle" and idle operation while the aircraft is in air may be referred to as "flight idle.". All options described below for this aspect may be considered relevant to ground idle conditions. In flight idle, the thrust is generally somewhat higher than in ground idle. In some implementations, only the less restrictive areas may apply to the idle flight of a particular engine. In other implementations, all options described below for this aspect may also apply to the flight idle conditions.As discussed in the fifth aspect, 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 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 is key to controlling the heat transfer ratio. Moreover, the inventors have recognized that although cruise conditions generally account for a much greater portion of the operating time of an engine, idle operation is also important. Since the fuel mass flow rate is much lower at idle than at cruise, even a relatively low heat load on the fuel can result in a large 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, stationary running during boarding and rolling (to a runway or slope or between other locations on the ground) and also at certain times in flight (e.g. at the beginning of a descent). Since the operating conditions differ greatly between cruise and idle - both in terms of altitude and desired engine thrust - a different control of the heat exchange system is appropriate.The method may include controlling the heat exchange system under idle conditions such that the heat transfer ratio is below 5.50, and optionally below 5.0, 4.5, or 4.0. The method may comprise controlling the heat exchange system under idle conditions such that the heat transfer ratio is above 0.75 and optionally above 1.0 or 1.5.The step of controlling the heat exchange system to adjust the heat transfer ratio may be or comprise controlling the modulation valve arranged to vary the proportion of oil directed across each branch of the oil circulation system.The step of controlling the heat exchange system to adjust the heat transfer ratio may include reducing the amount of oil directed across the at least one air-oil heat exchanger when the heat transfer ratio is too high.The heat exchange system may comprise at least one bypass pipe arranged such that oil or fuel may be directed past one or more heat exchangers, and the step of controlling the heat exchange system to adjust the heat transfer ratio may comprise modulating the amount of oil directed across the bypass pipe.The heat exchange system may include at least one return pipe arranged such that a liquid (e.g., the oil or fuel) may pass through one or more heat exchangers a plurality of times. The step of controlling the heat exchange system to adjust the heat transfer ratio may comprise modulating the amount of liquid directed through the return tube.The heat exchange system may include a refrigeration cycle device, and the method may further 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. In such implementations, the heat transfer ratio is in the range of 0.67 to 4.In other implementations, the heat exchange system is not configured to provide thermal lift so that the fuel temperature cannot be raised above a maximum temperature of the oil. In such implementations, the heat transfer ratio may be in the range of 2.00 to 5.67, and optionally 3.37 to 5.67.The method may include controlling the heat exchange system under idle conditions such that the heat transfer ratio is in the range of 2.33 to 5.67 when the fuel temperature entering the combustion chamber is below 200° C.The method may comprise controlling the heat exchange system under idle conditions such that: (i) the heat transfer ratio is in the range of 0.67 to 4 when the fuel temperature at entry into the combustion chamber is above 200° C.; (ii) the heat transfer ratio is in the range of 0.67 to 2.67 when the fuel temperature at entry into the combustion chamber is above 250° C.; and / or (iii) the heat transfer ratio is in the range of 0.67 to 1.22 when the fuel temperature is above 280° C.The method may comprise controlling the heat exchange system under idle conditions such that the heat transfer ratio is in the range of 0.67 to 3.67 if the fuel is at least 70% sustainable aviation fuel.The method may comprise controlling the heat exchange system under idle conditions such that the heat transfer ratio is in the range of 0.67 to 2.67 if the fuel is at least 80% sustainable aviation fuel.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 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; and a transmission 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 circulation system disposed to supply the transmission with oil; and 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 circuit system and the fuel flow so that heat is transferred between the oil and the fuel, and wherein the oil circuit system branches so that a portion of the oil can flow along each branch, and the air-oil and fuel-oil heat exchangers are arranged in a parallel configuration on different branches of the oil circuit system; anda modulation valve configured such that the proportion of oil directed via each branch can be varied, and wherein the heat exchange system is adapted to be controlled such that a heat transfer ratio is in the range of 0.67 to 5.67 under idling 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 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 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 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;a fuel-oil heat exchanger through which the oil in the oil circuit system and the fuel flow so that heat is transferred between the oil and the fuel, and wherein the oil circuit system comprises at least one bypass pipe, the bypass pipe being arranged to allow oil to be bypassed by one of the air-oil heat exchanger and the fuel-oil heat exchanger; and a bypass valve arranged so that the proportion of the oil that is routed via the bypass pipe can be varied,The method includes controlling the bypass valve such that a heat transfer ratio is in the range of 0 to 0.67 during travel conditions.The inventors have realised that the principles described with respect to the fifth to eighth aspects can also be applied to engines which do not have branching oil circulation paths with different heat exchangers on different branches; this applies, for example, to engines with a substantially linear, serial arrangement of heat exchangers in which the use of one or more bypass pipes takes place as an alternative to a branched main path and not as an optional supplement thereto. The introduction of bypass pipes and / or improved control of existing bypass pipes may enable a method that provides better oil cooling (as the fuel to be used may absorb more heat than conventional fuels) and may also improve overall thermal efficiency of the engine as less heat is lost to the environment. The controllable bypass valve is key to controlling the heat transfer ratio.The heat transfer rates used to calculate the heat transfer ratio are defined as discussed above with respect to the fifth and seventh aspects.The method may comprise controlling the bypass valve such that the heat transfer ratio is in the range of 0 to 0.60, optionally 0 to 0.50, 0 to 0.40, 0 to 0.30, or 0 to 0.20, and further optionally 0 to 0.10, under cruise conditions.The step of controlling the bypass valve to adjust the heat transfer ratio may include reducing the amount of oil directed across the at least one air-oil heat exchanger when the heat transfer ratio is too high.In some implementations, the oil circulation system may include a single bypass pipe, the single bypass pipe arranged such that oil may be routed past the air-oil heat exchanger. No bypass pipe may be provided for the fuel-oil heat exchanger. It will be appreciated that in most operating scenarios and to improve engine thermal efficiency, as much heat is to be transferred from the oil to the fuel as safely / as possible, and that therefore, bypassing the fuel-oil heat exchanger may be avoided.The heat exchange system may include at least two bypass pipes for oil, each bypass pipe arranged to allow oil to be bypassed by one of the air-oil heat exchanger and the fuel-oil heat exchanger; the method may include modulating the amount of oil routed via each bypass pipe. For controlling the flow through both bypass pipes, the same bypass valve - which may be a three-way valve - may be used, or a different bypass valve may be provided for each.The heat exchange system may comprise at least one return pipe arranged such that a liquid may pass multiple times through a heat exchanger. The method of such implementations may further include modulating the amount of liquid directed through the return tube to adjust the heat transfer ratio. Additionally or alternatively, the heat exchange system may further comprise a refrigeration cycle device, and the method for such implementations may further comprise 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. It will therefore be appreciated that in addition to the bypass valve, one or more other components may be used to adjust the heat transfer ratio.In implementations with a refrigeration cycle device, the bypass valve may be controlled such that the heat transfer ratio is in the range of 0 to 0.40.In implementations without a refrigeration cycle device in which the heat exchange system is not configured to provide thermal lift, the bypass valve may be controlled such that the heat transfer ratio is in the range of 0.38 to 0.67.The method may include controlling the bypass valve under cruise conditions such that: (i) the heat transfer ratio is in the range of 0 to 0.2 when the fuel temperature enters the combustion chamber is at least 160° C.; and / or (ii) the heat transfer ratio is in the range of 0 to 0.1 when the fuel temperature enters the combustion chamber is at least 180° C.In some implementations, one or more fuel characteristics may be considered when it is concerned with determining how to control the bypass valve. For example, the method may comprise controlling the bypass valve under cruise conditions such that: (i) the heat transfer ratio is in the range of 0 to 0.2 if the fuel is at least 70% sustainable flight fuel; (ii) 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 controlling the bypass valve under cruise conditions such that the oil-to-air heat transfer rate may be maintained in the range of 0 to 240 kJ per kilogram of fuel at cruise conditions, wherein no more than 20% of the heat dissipated from the oil at cruise conditions is transferred to the air, and / or such that the oil-to-fuel heat transfer rate is maintained in the range of 85 to 350 kJ per kilogram of fuel at cruise conditions, wherein at least 80% of the heat dissipated from the oil at cruise conditions is transferred to the fuel.The methods of the first, third, fifth, seventh, and ninth aspects may complement each other, and two or more of them may be performed together in various implementations. The method of the ninth aspect may be performed using the engine of the second, fourth, sixth, or eighth aspect.According to a tenth 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 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;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,and wherein the oil circulation system comprises at least one bypass pipe, wherein the bypass pipe is arranged to allow oil to be bypassed by one of the air-oil heat exchanger and the fuel-oil heat exchanger, and wherein a bypass valve is arranged such that the proportion of oil routed via the bypass pipe can be varied.The bypass valve is configured to be controllable to have a heat transfer ratio in the range of 0 to 0.67 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; 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 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 being arranged to return fuel from the heat exchange system at least two different locations along a main fuel path leading to the combustor from where the fuel enters the gas turbine engine.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.According to an eleventh aspect, there is provided a method of operating a gas turbine engine for an aircraft, the gas turbine engine comprising:an engine core comprising a turbine, a 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 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 circuit system and the fuel flow so that heat is transferred between the oil and the fuel, and wherein the oil circuit system comprises at least one bypass pipe, the bypass pipe being arranged to allow oil to be bypassed by one of the air-oil heat exchanger and the fuel-oil heat exchanger; and a bypass valve arranged so that the proportion of the oil that is routed via the bypass pipe can be varied,The method includes controlling the bypass valve such that a heat transfer ratio is in the range of 0.67 to 5.67 under idle conditions.The inventors have realised that the principles described with respect to the fifth to eighth aspects can also be applied to engines which do not have branching oil circulation paths with different heat exchangers on different branches; this applies, for example, to engines with a substantially linear, serial arrangement of heat exchangers in which the use of one or more bypass pipes takes place as an alternative to a branched main path and not as an optional supplement thereto. The introduction of bypass tubes and / or improved control of existing bypass tubes may enable a method that provides better oil cooling and may also improve overall thermal efficiency of the engine as less heat is dissipated to the environment. The controllable bypass valve is key to controlling the heat transfer ratio. Moreover, the inventors have recognized that although cruise conditions generally account for a much greater proportion of aircraft engine operating time, idle operation is also important. Since the fuel mass flow rate is much lower at idle than at cruise, even a relatively low heat load on the fuel can result in a large 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, stationary running during boarding and rolling (to a runway or slope or between other locations on the ground) or during descent in flight. Since the operating conditions differ greatly between cruise and idle, both in terms of altitude and in terms of the desired thrust of the engine, a different control of the bypass valve is appropriate.The heat transfer rates are defined as described above with respect to the fifth aspect. Idle operation while the aircraft is operating on the ground may be referred to as "ground idle" and idle operation in flight may be referred to as "flight idle.". All options described below for this aspect may be considered relevant to ground idle conditions. In flight idle, a somewhat higher thrust is generally achieved, and in some implementations, only the less restrictive areas may apply to the flight idle of a particular engine.The method may include controlling the bypass valve such that under idle conditions (and in some implementations specifically at ground idle), the heat transfer ratio is below 5.50, and optionally below 5.0, or below 4.5. The method may include controlling the bypass valve such that the heat transfer ratio is above 1.0, and optionally above 1.5, 2.0, 2.5, or 3.0 under idle conditions. The method may include controlling the bypass valve such that the heat transfer ratio is above 0.75 and optionally above 1.0, 1.5 under idle conditions.Controlling the bypass valve 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 heat exchange system may include at least two bypass pipes for oil, each bypass pipe arranged to allow oil to be bypassed by a respective one of the air-oil heat exchanger and the fuel-oil heat exchanger; the method may include modulating the amount of oil routed across each bypass pipe.The heat exchange system may comprise at least one return pipe arranged such that a liquid may pass multiple times through a heat exchanger. The method may further comprise modulating the amount of liquid directed through the recirculation pipe to adjust the heat transfer ratio.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 (beyond what would be transferred in the heat exchanger), optionally such that the fuel temperature is raised above the oil temperature. In such implementations, the method may include controlling the bypass valve such that the heat transfer ratio is in the range of 0.67 to 4. In implementations where the heat exchange system is not configured to provide thermal lift, the method may include controlling the bypass valve such that the heat transfer ratio is in the range of 2.00 to 5.67, and optionally 3.37 to 5.67.The method may include controlling the bypass valve under idle conditions such that the heat transfer ratio is at least one of: (i) in the range of 2.33 to 5.67 when the fuel temperature entering the combustion chamber is below 200° C.; (ii) in the range of 0.67 to 5.00 and optionally 0.67 to 4.00 when the fuel temperature entering the combustion chamber (16) is above 200° C.; (iii) in the range of 0.67 to 4.00 and optionally 0.67 to 2.67 when the fuel temperature entering the combustion chamber (16) is above 250° C.; and (iv) in the range of 0.67 to 2.33 and optionally 0.67 to 1.22 when the fuel temperature entering the combustion chamber (16) is above 280° C.The method may take into account the SAF content of the fuel. The method may include controlling the bypass valve such that the heat transfer ratio is in the range of 0.67 to 3.67 if the fuel is at least 70% sustainable flight fuel and / or in the range of 0.67 to 2.67 if the fuel is at least 80% sustainable flight fuel.The methods of the first, third, fifth, seventh, ninth, and eleventh aspects may complement each other, and two or more of them may be performed together in various implementations. The method of the eleventh aspect may be performed using the engine of the second, fourth, sixth, eighth, or tenth aspect.According to a twelfth 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; 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 circuit system is filtered, andthe fuel flows so that heat is transferred between the oil and the fuel,and wherein the oil circulation system comprises at least one bypass pipe, the bypass pipe being arranged to allow the oil to bypass one of the air-oil heat exchanger and the fuel-oil heat exchanger; and a bypass valve arranged to vary the proportion of the oil directed across the bypass pipe,and wherein the bypass valve is configured to be controlled such that a heat transfer ratio is established under idle conditionsThe range is from 0.67 to 5.67.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 being arranged to return fuel from the heat exchange system at least two different locations along a main fuel path leading to the combustor from where the fuel enters the gas turbine engine.The engine of the twelfth aspect may be configured to perform the method of the first, third, fifth, seventh, ninth and / or eleventh aspect, and may include any of the features described with respect to any of the preceding aspects.It is understood that a feature described with respect to one aspect may be used, mutatismutantly, in combination with any other aspect.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 (in which the input is on the sun gear and the output is on the ring gear and which is therefore also referred to as a "compound star" transmission), e.g. with 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 be in the range of 170 cm to 180 cm, 190 cm to 200 cm, 200 cm to 210 cm, 210 cm to 230 cm, 290 cm to 300 cm, or 340 cm to 360 cm.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 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, 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 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, the term descent may refer to a nominal point in the flight cycle of an aircraft between takeoff and landing where a relative decrease in altitude is required and which may require a reduced engine thrust demand.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 any suitable condition 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 any suitable condition, for example, in cruise 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 an illustration of an aircraft having a propulsion system including two gas turbine engines; FIG. 5 is an illustration of an example fuel system; FIG. 6 is an illustration of an alternative exemplary fuel system; FIG. 7 is a diagram of a portion of an example oil circulation system (a primary oil circulation system); FIG. 8 is a diagram of another part of an exemplary oil circulation system (a secondary oil circulation system); FIG. 9 is a diagram of a portion of an alternative exemplary oil circulation system; FIG. 10 is an illustration of a portion of the example fuel system of FIG. 5 and the example oil circulation system of FIGS. 7 and 8 ; FIG. 11 is a diagram of another alternative exemplary oil circulation system illustrating all major components of a heat exchange system; FIG. 12 illustrates an example method of operating a gas turbine engine; FIG. 13 illustrates another 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 a recirculating oil system for a gas turbine engine; FIG. 17 illustrates another example method of operating a gas turbine engine; FIG. 18 illustrates another example method of operating a gas turbine engine; and FIG. 19 illustrates a plot of cruise heat transfer ranges for a transmission 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.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 / L may at least partially reduce the starting fuel volume, which may be particularly important for volume limited or military uses 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 the 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-based hydrocarbon fuels may comprise molecules of about 7 to 18 carbons, with a significant portion of a given composition being derived from molecules of 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 hydrocarbon fuels such as kerosine), with the approval being pending further compositions. However, in the aeronautical industry, it is believed that persistent flight fuel mixtures comprising (and including) up to 100% persistent flight fuel (SAF) are permitted for use at some time.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. 4, 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 no central fuel tank may be provided (with all fuel instead stored in the aircraft's wings) - it should be appreciated that many different tank layouts are provided and that the depicted examples are provided for ease of description and are not intended to be limiting.FIG. 4 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 onboard 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 center fuel tank 50 and the wing fuel tanks 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 control of the oil flow, control of the fuel flow, and / or control of one or more heat exchangers 1004, 1006, 2020, 2030 or other heat exchange components in some other way-may allow the performance of the engine 10 to be optimized for a fuel having particular 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 control of the engine 10, and more particularly the thermal management 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 / may be selected based on those fuel properties most relevant to possible changes to the thermal management 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, may be determined by measuring substances contained in the exhaust gases that gas turbine engine 10 discharges during operation thereof. • The heating value of the fuel may be determined during operation of aircraft 1 from measurements made during combustion of the fuel - for example, from the fuel flow rate and shaft speed or temperature change in 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 transmission 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. 5. The fuel system 1000 includes both the fuel delivery 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. Heat exchange system 3000 includes portions of fuel management system 1500 as well as oil circulation system 2000, 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 secondary fuel-oil heat exchanger 1004 and a primary fuel-oil heat exchanger 1006. The primary fuel-oil heat exchanger 1006 may be described as a main fuel-oil heat exchanger because the oil flowing through it may be used to cool and lubricate the main transmission 30 of the engine 10. The secondary fuel-oil heat exchanger 1004 may be described as an integrated propulsion generator fuel-oil heat exchanger because the oil flowing through it may be used to cool and / or lubricate one or more integrated propulsion generator (IDG) components of the engine 10. In other implementations, a different type of generator may be used in place of an IDG - e.g., a variable frequency generator (VFG) or a variable frequency starter generator (VFSG). The system 1000 of such implementations may be otherwise equivalent. The engine 10 of the described example therefore comprises two fuel-oil heat exchangers 1004, 1006. In other implementations, more or fewer fuel-oil heat exchangers may be provided. Depicted fuel management system 1500 is configured such that fuel reaches secondary fuel-oil heat exchanger 1004 prior to primary fuel-oil heat exchanger 1006. After exiting the fuel-oil heat exchanger 1006, the fuel passes through an engine fuel pump 1003 and then enters the combustor 16.The primary fuel-oil heat exchanger 1006 and the secondary fuel-oil heat exchanger 1004 are configured to also pass an oil stream through each of them in addition to the fuel stream. 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.The two oil streams (oil stream through the primary heat exchanger and oil stream through the secondary heat exchanger) may be separate-physically separate and optionally also chemically different oils and / or have a different flow rate. Therefore, oil may flow through the primary fuel-oil heat exchanger 1006 other than through the secondary fuel-oil heat exchanger 1004.Generally, at least most of the fuel passing through the secondary fuel-oil heat exchanger 1004 also passes through the main fuel-oil heat exchanger 1006. 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 provided with a bypass so that a portion of the fuel does not flow through the respective heat exchanger, for example in the form of a bypass pipe 1005, as shown in FIG. 5. A valve (not shown) may determine which portion of the fuel flows through the heat exchanger 1004 and which portion flows 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' for 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 to also pass an oil stream through them in addition to the fuel stream - wherein in the described implementation the oil flowing through the one heat exchanger is different from the oil flowing through the other heat exchanger, although it is understood that in other implementations the same oil may flow through one fuel-oil heat exchanger and then through the other fuel-oil heat exchanger.The two heat exchangers 1004, 1006 are therefore located in separate closed-loop systems 2000, 2000' (FIGS. 7, 8 ) in the described implementation with respect to the oil flow, i.e. the oils, the oils flowing through the primary and secondary fuel-oil heat exchangers are fluidically separated and may be chemically different from one another. The two oil circulation systems 2000, 2000' serve to circulate oil through their respective fuel-oil heat exchangers 1006, 1004, and optionally also through one or more further heat exchangers, e.g. air-oil or oil-oil heat exchangers, as described below. The two oil circulation systems 2000, 2000' may be provided together as an oil circulation system for the engine 10.FIG. 6 shows an alternative example fuel system 6000 including a fuel delivery 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 secondary fuel-oil heat exchanger 1004 and then a primary fuel-oil heat exchanger 1006 before reaching an engine fuel pump 1003 that pumps the fuel along its flowpath to the combustor 16. 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 upstream 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 pump for returning fuel to the tank (FRTT) may be present and in some implementations may also aid in the recirculation. 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 pump 1003 and the combustion chamber 16. It is contemplated that the recirculation valve could be positioned downstream of the engine fuel pump 1003, for example as shown in FIG. 10, 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 and thus multiple times flow through the primary heat exchanger 1006 and the pump 1003 before reaching the combustion chamber 16. This recirculation provides a mechanism for controlling fuel flow within fuel management system 6500, and also within thermal management system 3000, without altering fuel flow from tank 50 to engine 10. The feedback may be used to control the heat transfer ratio, as described in more detail below.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 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.In the example shown in FIG. 10, unlike FIG. 6, the recirculation valve 6010 is downstream of the pump 1003, so that the recirculated fuel has already passed through both the primary heat exchanger 1006 and the pump 1003, and not just the primary heat exchanger 1006. The return pipe 6011 returns the returned fuel to a point on the flow path that is upstream of both the main pump 1003 and the primary heat exchanger 1006, so that the returned fuel passes through both of these components once more. Recirculation of fuel through the pump may allow more adjustable control of fuel flow rate to the combustor 16 for a particular shaft speed of the engine 10, it being appreciated that pump speed (or a limited set of pump speed options) is often determined by shaft speed.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.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 described implementation, the oil circulation system comprises a primary oil circulation system 2000 and a secondary oil circulation system 2000', each of which is a closed oil circulation system. An example of a primary closed loop oil system 2000 is schematically shown in FIG. 7 --as with the heat exchangers, this oil loop system is also referred to as "primary" because it is responsible for lubricating and cooling the main transmission 30 and generally for the main / primary cooling load of the engine 10. The primary closed-loop oil system 2000 includes an oil tank 2002 suitable for receiving a volume of oil. In some implementations, gases are removed from the oil in the oil tank 2002 by a breather. A feed pump 2004 is configured to pump oil from the oil tank 2000 into the main fuel-oil heat exchanger 1006. The average temperature of the oil entering the main fuel-oil heat exchanger 1006 is higher than the average temperature of the fuel entering the main fuel-oil heat exchanger 1006 during cruise conditions. In the main 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 main fuel-oil heat exchanger 1006 is lower than the average temperature of the oil stream entering the main fuel-oil heat exchanger 1006, such that it is cooled prior to reuse as lubricant and / or coolant, thereby allowing the cooled oil to draw more heat from the system to be lubricated and / or cooled. Also in this manner, the average temperature of the fuel exiting the main fuel-oil heat exchanger 1006 is higher than the average temperature of the fuel entering the main fuel-oil heat exchanger 1006.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. 7, 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 cools the oil of the primary oil system 2000 before being recirculated to the transmission 30. In various implementations, a bypass or return pipe for the oil and / or fuel may be provided around the refrigeration cycle device 1007.The oil stream in the primary oil circulation system 2000 is then conveyed to a power transmission 30, which may also be referred to as the main transmission 30 of the gas turbine engine 10. The power transmission 30 is arranged to receive input from the core shaft 26 and output to the fan 23 via the fan shaft 42 and includes the gears 28, 32, 38 and bearings (e.g., journal bearings) that can be lubricated and / or cooled by the oil. The engine 10 may also include one or more additional bearings for supporting the shafts 26, 42, which may be journal bearings. The oil can additionally be used to lubricate and / or cool the journal bearings and generally significantly increases temperature when used under cruise conditions, thereby assisting in cooling the bearings and transmission 30, as the oil stream carries heat away from the bearings and transmission 30. The oil may also be used to lubricate one or more other engine components 33, e.g., an auxiliary gear box (AGB) and / or one or more bearing chambers. The AGB 33, also known as a power assist, if present, is a gearbox that forms part of the gas turbine engine 10, although it is not part of the engine core 11 and does not drive the fan 23. The AGB instead drives engine accessories, e.g., fuel pumps, and generally handles large loads. A relatively large amount of heat can therefore be dissipated from the AGB into the oil. One or more bearing chambers may be lubricated by the same oil and may also provide heat to the oil. Per unit of oil flowing through, the AGB and bearing chambers may add more heat to the oil than the main transmission 30 in many implementations. The oil flow may be split into two or more parallel flows, for example, one flow through the main transmission 30 and one flow through the other engine components, or multiple parallel flows through the main transmission 30 (e.g., via different components of the transmission), and separate flows through the AGB and the or each bearing chamber 33.The oil from the power transmission 30 (and / or other engine components 33, such as the AGB, if present) collects in an oil pan 2008. A return pump 2010 is configured to pump oil from the oil pan 2008 back into the oil tank 2002 where it is ready for reuse.FIG. 8 illustrates the secondary oil circulation system 2000', which is another closed oil circulation system 2000'. The secondary closed oil circuit system 2000' includes a secondary oil tank 2002' adapted to receive a volume of oil. In some implementations, gases are removed from the oil in oil tank 2002' by a breather. A secondary feed pump 2004' is configured to pump oil from the secondary oil tank 2002' to the secondary fuel-oil heat exchanger 1004, which in the illustrated implementation is the IDG fuel-oil heat exchanger 1004. The average temperature of the oil entering the IDG fuel-oil heat exchanger 1004 is higher than the average temperature of the fuel entering the IDG fuel-oil heat exchanger 1004 during cruise conditions. In the IDG fuel-oil heat exchanger 1004, thermal energy is transferred from the oil stream to the fuel stream. In this way, the average temperature of the oil stream exiting the IDG fuel-oil heat exchanger 1004 is lower than the average temperature of the oil stream entering the IDG fuel-oil heat exchanger 1004. Also in this way, the average temperature of the fuel exiting the IDG fuel-oil heat exchanger 1004 is higher than the average temperature of the fuel entering the IDG fuel-oil heat exchanger 1004. The oil stream is then conveyed back to an integrated propulsion generator 2006, where it lubricates and / or cools moving components and is warmed while doing so. In some implementations, the oil may be used primarily as a coolant for IDG 2006 and serves only minimally or not at all for lubrication. The oil from the integrated propulsion generator 2006 collects in a secondary oil pan 2008'.A secondary return pump 2010' is configured to pump oil from the secondary oil pan 2008' back into the secondary oil tank 2002' where it is ready for reuse. In some implementations, a refrigeration cycle device may also be provided in the secondary oil cycle system 2000'.FIGS. 7 and 8 each show a serial flow path of the oil, wherein the entire oil stream flows successively through the individual components (wherein, of course, one or more bypass pipes or return pipes, not shown, can also be provided for the oil). In other implementations, the oil stream may be split into two or more parallel streams, e.g., a stream through the main fuel-oil heat exchanger 1006 and a stream through an air-oil heat exchanger 2020 (as described below).FIG. 9 schematically illustrates an alternative example portion of the primary closed oil circuit system 2000 shown in FIG. 7. In this section, an oil flow is pumped by the feed pump 2004 through a valve 2016. The valve 2016 may be operated to split the oil flow between the main fuel oil heat exchanger 1006 and a first air oil heat exchanger 2020, the first air oil heat exchanger 2020 being arranged in parallel with the main fuel oil heat exchanger 1006. The oil flow path may be described as branching, with the main fuel-oil heat exchanger 1006 on one branch and the first air-oil heat exchanger 2020 on the other branch being arranged in a parallel configuration such that oil may flow over one or the other branch, but the same portion of oil cannot flow through both in the same cycle - the flow splits. The valve 2016 modulates the flow through the two heat exchangers 1006, 2020 and can therefore be described as a modulation valve 2016. The oil streams are then merged again and conveyed to the power transmission 30 and / or other engine components 33. Any suitable percentage of oil may flow through each of the first air-oil heat exchanger 2020 and the main fuel-oil heat exchanger 1006. In some examples, the valve 2016 is operable to vary the flow of oil into the main fuel-oil heat exchanger 1006 and the first air-oil heat exchanger 2020 as desired. In various examples, an oil-to-oil heat exchanger 2030 (not shown in FIG. 7, but present in FIG. 11 ) may be provided, for example, arranged in series with the first air-to-oil heat exchanger 2020 on this branch of the parallel split. The oil-to-oil heat exchanger 2030 may allow heat exchange between the primary and secondary closed oil loop systems 2000, 2000'.FIG. 10 schematically shows an exemplary arrangement and the interaction of the first closed oil circuit system 2000, the second closed oil circuit system 2000' and the fuel system 1000, wherein the fuel flow is illustrated in thick black lines and the oil flow is illustrated in thinner black lines. The hatched thick black line indicates a return path 6011 through which only a part of the fuel passes. The combination of fuel systems 1000 and oil systems 2000, 2000' together form heat exchange system 3000. The primary closed oil circuit system 2000 of this exemplary arrangement is arranged as shown in FIG. 7. The secondary closed oil circuit system 2000' of this exemplary arrangement is arranged as shown in FIG. 8. The fuel system 1000 of this exemplary arrangement is arranged as shown in FIG. 5, but with an additional recirculation valve 6010 and pipe 6011 as described above. A bypass tube 1005 may also be present, but is not shown for clarity.In use, fuel is pumped from the fuel tank 50 by the low pressure fuel pump 1002. The fuel then flows through the IDG fuel-oil heat exchanger 1004. The secondary closed oil loop system 2000' is configured such that its oil return flow also flows through the IDG fuel-oil heat exchanger 1004. During normal operation of the engine 10, both in cruise conditions and idle, the average temperature of the oil flow entering the IDG fuel-oil heat exchanger 1004 is higher than the average temperature of the fuel flow entering the IDG fuel-oil heat exchanger 1004. The IDG fuel-oil heat exchanger 1004 is configured to transfer heat from the oil stream to the fuel stream. In this way, the average temperature of the oil stream exiting the IDG fuel-oil heat exchanger 1004 is lower than the average temperature of the oil stream entering the IDG fuel-oil heat exchanger 1004. Similarly, the average temperature of the fuel stream exiting the IDG fuel-oil heat exchanger 1004 is higher than the average temperature of the fuel stream entering the IDG fuel-oil heat exchanger 1004.The fuel then flows through the main fuel-oil heat exchanger 1006 and, in addition, through the refrigeration cycle device 1007. The primary closed oil cycle system 2000 is configured such that its oil return flow also flows through the main fuel-oil heat exchanger 1006 and the refrigeration cycle device 1007. During normal operation of the engine 10, both in cruise conditions and idle, the average temperature of the oil flow entering the main fuel-oil heat exchanger 1006 is higher than the average temperature of the fuel flow entering the main fuel-oil heat exchanger 1006. The main fuel-oil heat exchanger 1006 is configured to transfer heat from the oil stream to the fuel stream. In this way, the average temperature of the oil stream exiting the main fuel-oil heat exchanger 1006 is lower than the average temperature of the oil stream entering the main fuel-oil heat exchanger 1006. Similarly, the average temperature of the fuel stream exiting the main fuel-oil heat exchanger 1006 is higher than the average temperature entering the main fuel-oil heat exchanger 1006. The refrigeration cycle device 1007 may, when active (i.e., when energized / energized and used to actively transfer heat from the oil to the fuel), provide a further increase in fuel temperature and a decrease in oil temperature, respectively, where the fuel is optionally raised above the oil temperature. After passing through the main fuel-oil heat exchanger 1006, the fuel flows to the engine fuel pump 1003, which in the examples shown is downstream of the primary and secondary fuel-oil heat exchangers 1006, 1004 and is arranged to supply fuel to the combustor 16 of the gas turbine engine 10.During cruise conditions, the average temperature of the oil stream flowing through the IDG fuel-oil heat exchanger 1004 may be lower than the average temperature of the oil stream flowing through the main fuel-oil heat exchanger 1006. In this manner, the fuel first passes through the heat exchanger 1004 having a lower average oil flow temperature before passing through the heat exchanger 1006 having a higher average oil flow temperature. In addition to the diverging oil streams 2000, 2000', the heat exchange system 3000 may further include diverging fuel return paths such that fuel from where fuel enters the gas turbine engine 10 to the combustor 16 is returned to the main fuel path at at least two different locations - for example, upstream or downstream of the main engine pump 1003 as shown in FIG. 6, with a diverging path 6020 branching the main fuel flow path at the exit 1006b of the primary heat exchanger 1006 and opening back into the main fuel flow path downstream of the pump 1003. At least one valve (not shown) may be provided to control a division of the fuel flow from the heat exchanger 1006 back to the main fuel path through the engine 10. The valve may be controlled based on fuel temperature - for example, to direct less fuel via pump 1003 / more fuel to a position downstream of the pump when fuel temperature is relatively high and the likelihood of damage to the pump seals or other components is higher. Control of fuel flow through the diverging fuel return paths may be based on measurement of fuel temperature (e.g., using a temperature sensor at a location downstream of the heat exchanger and fuel, possibly also using a temperature measurement upstream of the fuel-oil heat exchanger). Recirculation of the recirculated fuel upstream of the fuel-oil heat exchanger may allow for reduced heat transfer from the oil to the fuel and thus dampen transient overshoot, which may occur, for example, at the beginning of the descent phase when the same amount of heat is generated in the oil system, but the fuel flow is reduced, which typically results in a temperature increase. Adjusting the flow of fuel through these diverging paths may also be useful for reducing icing by recirculating previously heated fuel to warm the system. The branching fuel return paths can therefore be used in a variety of ways to improve engine thermal management. FIG. 11 schematically shows an exemplary configuration of the primary closed oil circuit system 2000 and the secondary closed oil circuit system 2000' where the two independent oil return streams are brought into a heat exchange relationship by an oil-oil heat exchanger 2030. Both oil circulation systems 2000, 2000' have in this example a branched arrangement of parallel pipes / heat exchangers.In the example shown in FIG. 11, the primary closed loop oil system 2000 is configured such that the oil return flow from the feed pump 2004 is pumped through a valve 2016, which may be referred to as a modulation valve. The valve 2016 may be operated to split the oil stream such that a portion of the oil stream flows to the main fuel-oil heat exchanger 1006 and a first air-oil heat exchanger 2020, respectively. In the illustrated implementation, the first air-oil heat exchanger 2020 is connected in series with the oil-oil heat exchanger 2030, and the arrangement of the air-oil heat exchanger 2020 and the oil-oil heat exchanger 2030 is connected in parallel with the main fuel-oil heat exchanger 1006. The modulation valve 2016 determines which portion of the oil flows through each branch of the parallel arrangement. In various implementations, any suitable portion of the oil stream may be diverted between the main fuel-oil heat exchanger 1006 and the first air-oil heat exchanger 2020. In examples, the valve 2016 is operable to redirect a fixed portion of the oil stream to the main fuel-oil heat exchanger 1006 and the first air-oil heat exchanger 2020, respectively. In other examples, the valve 2016 is operable to direct a variable portion of the oil stream to the main fuel-oil heat exchanger 1006 and the first air-oil heat exchanger 2020, respectively, for example, using temperature sensor data as feedback to control the variable portion and / or based on one or more fuel characteristics, as described below. The valve 2016 may be adjustable between a discrete number of fixed positions or continuously. In some implementations, for example, when the temperature of the fuel entering the combustion chamber 16 is relatively low compared to the maximum operating temperature of the fuel (e.g., based on knowledge of the fuel type or thermal stability), no oil may be directed to the air-oil heat exchanger 2020 and all of the oil may be directed via the fuel-oil heat exchanger 1006.After passing through the heat exchangers 1006, 2020, 2030, the oil stream in the primary closed loop system 2000 is then merged again and conveyed to the power transmission 30 (and / or other engine components 33 such as the AGB) and then to the oil pan 2008. Subsequently, the return pump 2010 pumps the oil from the oil pan 2008 into the oil tank 2002 for reuse.The oil stream in the secondary closed oil circuit system 2000' is arranged to be brought into a heat exchange relationship with the separate oil stream in the primary closed oil circuit system 2000 via the oil-to-oil heat exchanger 2030. In the oil-to-oil heat exchanger 2030, the oil flow in the primary closed oil circuit system 2000 does not mix with the oil flow in the secondary closed oil circuit system 2000'. The oil-to-oil heat exchanger 2030 is configured such that heat transfer can take place between the two separate oil streams. In this way, the heat of a hotter oil stream may be transferred to the cooler oil stream in the oil-to-oil heat exchanger 2030. No air-to-oil heat exchanger is shown in the depicted secondary closed oil circuit system 2000', however, in other examples, an air-to-oil heat exchanger may be provided - e.g., in series with the oil-to-oil heat exchanger 2030 or on a third parallel branch.In the implementation shown in FIG. 11, the secondary closed loop oil system 2000' is configured such that the oil return flow from the secondary feed pump 2004' is pumped through a valve 2016', which may be referred to as a secondary modulation valve. The valve 2016' is operable to redirect at least a portion of the oil flow between the IDG fuel-oil heat exchanger 1004 and the oil-oil heat exchanger 2030, the oil-oil heat exchanger 2030 being disposed parallel to the IDG fuel-oil heat exchanger 1004. In other implementations, the secondary closed oil loop system 2000' may not have a branched configuration. For example, the IDG fuel-oil heat exchanger 1004 and the oil-oil heat exchanger 2030 may be arranged in series such that in a particular cycle, at least a majority of the oil flowing through the one heat exchanger also flows through the other heat exchanger.In other implementations, an air-to-oil heat exchanger 2020 may be present in each closed loop system or only in the secondary closed loop system and / or no oil-to-oil heat exchanger may be present.In examples, any suitable portion of the oil stream may be diverted between the IDG fuel-oil heat exchanger 1004 and the oil-oil heat exchanger 2030. In examples, valve 2016' is operable to redirect a fixed portion of the oil stream to IDG fuel-oil heat exchanger 1004 and oil-oil heat exchanger 2030, respectively. In examples, valve 2016' is operable to redirect a variable portion of the oil stream to fuel-oil heat exchanger 1004 and oil-oil heat exchanger 2030, respectively. The valve 2016' may be adjustable between a discrete number of fixed positions or continuously. After the oil stream has flowed through the heat exchangers 1004, 2030, in the example shown in Figure 8, it is then conveyed to the integrated propulsion generator 2006 and then into the secondary oil pan 2008'. The secondary return pump 2010' then pumps the oil from the secondary oil pan 2008' into the secondary oil tank 2002' for reuse. It will be appreciated that in other implementations, the components of the oil circuit may be arranged differently and that additional or alternative components of the oil circuit may be present.One or more temperature sensors may be provided, which are designed, for example, to detect the fuel temperature when the fuel pump 1003 or the combustion chamber 16 is approached. 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, 2030. 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.FIG. 12 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 may be performed by any of the methods described above, optionally using a fuel composition determination module 57.The engine 10 used to implement the method 100 includes an air-to-oil heat exchanger 2020 and a fuel-to-oil heat exchanger 1006 and a modulation valve 2016 arranged such that the proportion of oil directed across each heat exchanger can be varied, and the method 100 includes controlling 104 the modulation valve based on the at least one fuel property to adjust the proportion of oil directed across each heat exchanger under cruise conditions.More specifically, the method 100 for the implementations described above with reference to FIGS. 5-11 includes controlling 104 the modulation valve 2016 of the primary oil system 2000 under travel conditions based on the at least one fuel property to adjust the amount of oil directed across each branch of the primary oil circuit system 2000 and control how much oil flows through the main fuel-oil heat exchanger 1006 and how much flows through the air-oil heat exchanger 2020. In implementations with different arrangements of heat exchangers, for example a series arrangement of heat exchangers instead of a parallel split, the modulation valve 2016 may effectively be a bypass valve that allows one of the heat exchangers 1006, 2020 to bypass a controllable portion of the flow, whereby the portion of oil routed across each heat exchanger may be varied in this way.As indicated by the dashed line in FIG. 12, 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 modulation valve 2016 may therefore be configured to redirect a fixed portion of the oil flow to each of the fuel-oil heat exchanger 1006 and the air-oil heat exchanger 2020 during operation of the engine 10, the fixed portion being determined at start-up of the engine 10 based on the at least one determined fuel characteristic (option (i) above). Alternatively, the modulation valve 2016 may be configured to redirect a variable portion of the oil flow to the fuel-oil heat exchanger 1006 and the air-oil heat exchanger 2020, respectively, during operation of the engine 10 during a single flight (option (ii) above). The modulation valve 2016 may therefore be actively controlled to vary the proportion of oil routed across each heat exchanger 1006, 2020, particularly in implementations where the aircraft 1 carries multiple different fuels in different tanks, and may change which fuel (or fuel mixture) is used during flight. The active control of the modulation valve may be automated and implemented by a controller 58 of the heat exchange system 3000, which may be a dedicated controller or a portion of a more general EEC. In various implementations, the control of the modulation valve 2016 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 oil circulation system 2000, 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, and the air-to-oil heat exchangers 2020 and fuel-to-oil heat exchangers 1006 may be arranged in a parallel configuration on different branches of the oil circulation system, as shown in FIGS. 9 and 11. The modulation valve 2016 of such examples may be configured such that the proportion of oil directed via each branch may be varied, and the controller 104 of the modulation valve 2016 may therefore adjust the proportion of oil directed via each branch under cruise conditions. In some examples, more than two heat exchangers and / or more than two branches may be provided. The modulation valve 2016 therefore enables the amount of heat transferred from the oil through the air-oil heat exchangers 2020 and fuel-oil heat exchangers 1006 to be varied. 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 the adjustment of the modulation valve 2016. 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 modulation valve 2016 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, the method 100 may include using the modulation valve 2016 to perform one or more of the following actions:• directing all of the oil via the 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 via the fuel-oil heat exchanger 1006 (no oil via 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 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. For example, if the thermal stability data indicates that the fuel is stable in operation at temperatures above 160° C. or 170° C., the modulation valve 2016 may be controlled ( 104) such that at least 80% and optionally 90-100% of the heat dissipated from the oil during cruise 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 modulation valve 2016 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 for fuels with a higher heating value, the flow rate is generally reduced in order to achieve the same thrust force without wasting fuel. The fuel in the 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 gearbox 30 that receives input from the core shaft and outputs propulsion to the fan 23, and an oil circulation system 2000 configured to supply oil to the gearbox 30. 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 modulation valve 2016 is provided, wherein the valve 2016 is configured such that the proportion of the oil conducted via each heat exchanger can be varied. In various examples, including that shown in FIG. 9, the oil circulation system 2000 branches such that a portion of the oil may flow along each branch, and the air-oil and fuel-oil heat exchangers 1006, 2020 are arranged in a parallel configuration, one on each branch of the oil circulation system 2000. The modulation valve 2016 is configured so that the proportion of the oil that is conducted via each branch 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 modulation valve 2016 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 that is directed across each heat exchanger 1006, 2020 during travel conditions. A controller 58 may be provided to make and implement decisions based on the output of the fuel composition determination module 57.In various examples, including that shown in FIG. 11, the heat exchange system 3000 also includes a secondary oil circulation system 2000' as described above and an oil-to-oil heat exchanger 2030 configured to allow heat exchange between the oil in the two oil systems 2000, 2000'. The modulation valve 2016 or an additional valve may control the flow of oil from the primary oil circulation system 2000 to the oil-to-oil heat exchanger 2030. The secondary oil circulation system 2000' may also include a valve to control the flow of oil to the oil-to-oil heat exchanger 2030. In the example shown in FIG. 11, each oil system 2000, 2000' has a total of two parallel branches, and the oil-oil heat exchanger 2030 is located on the same branch as the air-oil heat exchanger 2020 of the primary oil system 2000.The heat exchange systems 3000 of various examples include at least one bypass pipe arranged such that fuel (or oil) may be routed past one or more heat exchangers. In some implementations, a bypass pipe may effectively form an additional branch in a parallel branched oil system. The modulation valve 2016 or other oil valve may be configured to adjust the amount of oil directed through the or each oil bypass pipe based on the one or more determined fuel characteristics and optionally based on one or more temperature measurements.FIG. 13 illustrates a second method 200 in which knowledge of the fuel properties is used to determine control of the heat exchange system 3000. The method 200 includes determining 202 at least one fuel property of the fuel to be provided to and burned by the combustion chamber 16. This determination 202 may be performed by any of the methods described above, optionally using a fuel composition determination module 57. the details of this determination step 202 may be as described in the determination 102 for the method 100 shown in FIG. 12. The method of FIG. 13 further includes modulating 204 the heat exchange system 3000 to adjust the temperature of the fuel at entry into the combustion chamber 16 during cruise, the set value based on the at least one particular fuel property.The modulation 204 of the heat exchange system 3000 may include one or more of the following elements:• adjusting an appropriate oil or fuel valve, e.g., the modulation valve 2016, to change the proportion of the oil or fuel that is directed over a particular heat exchanger or, if present, over a refrigeration circuit device 1007;• adjusting a modulation valve 2016, if present, to change the proportion of oil routed via one or more branches of an oil system 2000;• adjusting a bypass valve, if any, to adjust the proportion of a liquid (e.g., oil or fuel) that bypasses a particular heat exchanger or set of heat exchangers;• adjusting a recirculation valve 6010, if present, to adjust the proportion of a liquid (e.g., oil or fuel) that is recirculated to an earlier point on its path and thus multiple-pass one or more components (generally including a particular heat exchanger or set of heat exchangers);• Activate, deactivate, or otherwise adjust the use of a refrigeration cycle device 1007, if present, to provide a controlled increase in heating of one liquid / cooling of another liquid;• adjusting a pumping speed of an oil system oil pump 2010, 2010';• shutting off one or more branches of the oil flow in an oil system 2000, 2000' if the or each oil system has branching oil passages; and / or• adjusting a valve to control the flow of air to an air-to-oil heat exchanger 2020, if present, or otherwise changing the flow of air through this heat exchanger 2020.As indicated by the dashed line in FIG. 13, this method 200 may optionally be repeated multiple times during flight / operation of the engine 10, as described above with respect to the method of FIG. 12. Thus, both the determination 202 and subsequent modulation 204 may be performed during cruise flight in some implementations.The fuel property(s) may not be the only data taken into account in modulating the heat exchange system 3000. For example, the temperature of the environment / ambient air available for use in the air-to-oil heat exchanger 2020 (if present) may be considered (optionally using altitude - optionally in conjunction with the geographic region - as a reference for temperature), as well as fuel flow rates and fuel temperatures.In some implementations, the step of modulating 204 the heat exchange system 3000 to control fuel temperature includes controlling oil flow through the at least one fuel-oil heat exchanger 1006 such that between 50% and 100% of heat lost from the oil is transferred to the fuel, and optionally such that between 80% and 100% or between 90% and 100% of heat lost from the oil is transferred to the fuel.The fuel property determined may be or include the thermal stability of the fuel; in such implementations, the heat exchange system 3000 may be modulated such that the temperature of the fuel as it enters the combustion chamber 16 during cruise is increased with increasing thermal stability, optionally linearly. 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 degradation upon reaching a particular temperature; aviation fuels include various components that can degrade at different temperatures and residence time at high temperatures is also a factor for fuel degradation. A fuel may be deemed stable at a particular temperature if its decomposition rate at that temperature is below a particular threshold.In alternative or additional implementations, the determined fuel property may be or include the presence of a tracer species in the fuel. In such implementations, the heat exchange system 3000 may be modulated such that the temperature of the fuel as it enters the combustion chamber during cruise flight is set to a predefined value corresponding to the tracer species. Optionally, the amount of tracer species detected may indicate whether the fuel is a mixture of a labeled flying fuel (i.e., a fuel with added tracer species) with another fuel or the pure labeled flying fuel.In alternative or additional implementations, the determined fuel property may be or include the percentage of SAF in fuel (% SAF). In such implementations, the heat exchange system 3000 may be modulated such that the temperature of the fuel as it enters the combustion chamber 16 in cruise flight is increased as % SAF increases, once % SAF exceeds 60%, or optionally, once % SAF exceeds 70%, 75%, or 80%. At lower % SAF values, in some implementations, no modulation based on SAF content may be performed, but modulation 204 based on one or more other properties may be performed. It is therefore understood that multiple different fuel properties may be taken into account together (in other embodiments, modulation based on SAF content may be performed at lower & SAF values). 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 the modulation of the heat exchange system 204 based on the fuel property data and optionally other data as well.A gas turbine engine 10 for an aircraft implementing this method 200 does not necessarily include a gearbox 30 that receives input from the core shaft and outputs propulsion to the fan 23 - i.e., the method 200 may be applied to both transmission-driven engines 10 and direct-drive engines. The method 200 may therefore be applied to a wide range of aircraft engines provided they include a combustion chamber 16, an oil system 2000 arranged to circulate oil, and a heat exchange system 3000 comprising at least one fuel-oil heat exchanger arranged to transfer heat from the oil to the fuel. 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 is also provided - either as a stand-alone unit or as part of an EEC or other engine control system. The fuel composition determination module 57 may include or be in communication with a processing module configured to make decisions based on the one or more determined fuel characteristics and optionally other data as well.In some implementations, heat exchange system 3000 includes branching fuel return paths and at least one valve that controls a division of the fuel flow. The branching paths may be arranged to recirculate fuel from the heat exchange system 3000--and in particular fuel exiting the primary 1006 or secondary 1004 fuel-oil heat exchanger--at at least two different locations along a main fuel path leading from the location where fuel enters the gas turbine engine 10 to the combustor 16 as described above. For example, the fuel flow exiting the heat exchanger 1004, 1006 may be split into two paths, one path generally receiving the majority of the fuel (and is thus referred to as a main fuel flow path) and the other path diverting a portion of the fuel exiting the heat exchanger 1004, 1006 such that the diverted portion of the fuel bypasses one or more components of the main fuel flow path (e.g., a pump) and connects back to the main flow to the combustion chamber 16 downstream of these one or more components of the main fuel flow path. The valve may control and adjust the proportion routed through each branch. In some implementations, more than two branches may be provided.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 300, 400 of FIGS. 14 and 15 are concerned with these two scenarios of aircraft operation.FIG. 14 illustrates a method 300 implementing these considerations under cruise conditions and FIG. 15 illustrates a method 400 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 300 to be performed in cruise flight, the method 300 is configured to be performed in a transmission gas turbine engine 10 comprising an oil cycle system 2000 configured to supply oil to the transmission 30; and a heat exchange system 3000 comprising an air-oil heat exchanger 2020 through which the oil in the oil cycle system flows; and a fuel-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. The oil circulation system 2000 branches as shown in FIGS. 9 and 11 such that a portion of the oil may flow along each branch, with the air-oil and fuel-oil heat exchangers 1006, 2020 arranged in a parallel configuration on different branches of the oil circulation system. A modulation valve 2016 is also provided to control the portion of the oil that is routed via each branch of the oil circulation system 2000. The method 300 includes controlling 302 the heat exchange system 3000 such that, under cruise conditions, a heat transfer ratio is in the range of from 0 to 0.67, and optionally 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. 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 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 from a controller 58 and used 306 to adjust the modulation 302 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 306 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 304, 306 may alternatively be considered part of controlling 302 the heat exchange system 3000. 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 surrounding the aircraft 1), flow rate data (of the oil and / or fuel), and / or one or more fuel characteristics, in determining 306 which control actions to take.The step of controlling 302 the heat exchange system 3000 may comprise any or any combination of the examples provided for the modulation step 204 of the method 200 of FIG. 13, e.g., by reducing the amount of oil directed via the at least one air-oil heat exchanger 2020 when the heat transfer ratio is too high, 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 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 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 300 may include controlling 302 the heat exchange system 3000 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. Therefore, one or more temperature sensors may be used, the output(s) of which may be taken into account in adjusting the controller 302 of the heat exchange system 3000.The method 300 may also consider the type of fuel-e.g., whether the fuel is or includes 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 adjusting the controller 302 of the heat exchange system 3000.The heat transfer rate of oil to cruise air (measured in kJ per kilogram of fuel flow reaching the combustion chamber 16) may be in the range of 0 to 240 kJ / kg, and optionally in the range of 0 to 120 kJ / kg. 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.The heat transfer rate of oil to cruise fuel (measured in kJ per kilogram of fuel flow reaching the combustion chamber 16) may be in the range of 85 to 350 kJ / kg, and optionally in the range of 85 to 170 kJ / kg. 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 110-240 kJ per kilogram of fuel during cruise conditions.In some implementations, the heat exchange system 3000 further includes branching fuel return paths 6020 and at least one valve that controls a division of the fuel flow, the branching paths arranged to return fuel from the heat exchange system 3000 at least two different locations along a main fuel path leading from the location 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 1004, 1006.The percentages of total heat dissipated from the oil in heat exchange system 3000 that is transferred to the fuel (as opposed to transfer to the air / environment) under cruise conditions for a particular transmission engine 10 are plotted as a function of the fuel temperature limit in FIG. 19. 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 400 to be performed at idle, method 400 is again configured to be executed in a transmission gas turbine engine 10 comprising an oil cycle system 2000 configured to supply oil to transmission 30; and a heat exchange system 3000 comprising 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. The oil circulation system 2000 branches as shown in FIGS. 9 and 11 such that a portion of the oil may flow along each branch, with the air-oil and fuel-oil heat exchangers 1006, 2020 arranged in a parallel configuration on different branches of the oil circulation system. 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. The method 400 includes controlling 402 the heat exchange system 3000 such that when the aircraft 1 is operated under idle conditions, a heat transfer ratio is in the range of 0.67 to 5.67. Therefore, the ratio may be higher at idle than at cruise. The ratio can be higher in ground idling than in flight idling.Optionally, method 400 may comprise maintaining the heat transfer ratio under idle conditions at below 5.50, 5.0, 4.5, 4.0, 3.5, or 3.0. Optionally, method 400 may include maintaining the heat transfer ratio at above 4.0, 4.5 under idle conditions or at 5.0 under ground idle conditions.The method 400 may further include receiving data 404 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 404 and used 406 from a controller 58 to adjust the modulation 302 of the idle heat exchange system 3000 to maintain the heat transfer ratio at a desired level or within the desired limits. This checking and adjustment / correction 406 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 404, 406 may alternatively also be considered part of controlling 402 the heat exchange system 3000. The method 400 may also be configured to use other information, e.g., temperature data (of the oil, fuel, and / or ambient temperature of an environment surrounding the aircraft 1), flow rate data (of the oil and / or fuel), and / or one or more fuel characteristics, in determining 406 which control actions to take.The step of controlling 402 the heat exchange system 3000 may comprise any or any combination of the examples provided for the modulation step 204 of the method 200 of FIG. 13, 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 402 the idle 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 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 heat transfer ratio at idle may be not more than 4, and optionally not more than 4.0, 3.5, 3.0, 2.5, 2.0, 1.5, or 1.0, and also optionally not 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 heat transfer ratio may be maintained, for example, greater than 3.37 and optionally greater than 3.4, 3.5, 4.0, 4.5, 5.0 or 5.5.The fuel temperature when reaching the combustion chamber 16 may also be taken into account. The method 400 may include controlling 402 the heat exchange system 3000 during idle operation such that the heat transfer ratio is in the range of 2.33 to 5.67 if the fuel temperature entering the combustion chamber 16 is below 200° C., or in the range of 0.67 to 4 if the fuel temperature entering the combustion chamber 16 is at or above 200° C., or in the range of 0.67 to 2.67 if the fuel temperature entering the combustion chamber 16 is at or above 250° C., or in the range of 0.67 to 1.22 if the fuel temperature entering the combustion chamber 16 is at or above 280° C. Therefore, one or more temperature sensors may be used, the output(s) of which may be taken into account in adjusting the controller 302 of the heat exchange system 3000.The method 400 may additionally or alternatively consider the type of fuel-e.g., whether the fuel is or includes a persistent flight fuel (SAF). For example, the idle heat transfer ratio may be maintained in the range of 0.67 to 3.67 if the fuel is at least 60%, or 70% of sustainable flight fuel, or 0.67 to 2.67 if the fuel is at least 75%, 80%, or 85% of sustainable flight fuel.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 adjusting the controller 402 of the heat exchange system 3000.The heat transfer rate of oil to air at idle (measured in kJ per kilogram of fuel flow reaching the combustion chamber 16) may be in the range of 380 to 1430 kJ / kg, and optionally in the range of 380 to 1270 kJ / kg, or in the range of 500 to 1080 kJ / kg. In some implementations, no more than 80% of the idle heat dissipated from the oil may be dissipated to the air, and the oil-to-air heat transfer rate may be maintained in the range of 630 to 1430 kJ per kilogram of fuel under idle conditions. At idle, a higher percentage of heat release to the air is generally used than at cruise and other higher power engine conditions.The heat transfer rate of oil to fuel at idle (measured in kJ per kilogram of fuel flow reaching the combustion chamber 16) may be in the range of 0 to 1270 kJ / kg, and optionally in the range of 190 to 760 kJ / kg. In some implementations, at least 20% of the heat dissipated from the oil at idle may be transferred to the fuel, and the heat transfer rate of oil to fuel may be maintained in the range of 150 to 360 kJ per kilogram of fuel under idle conditions.As with the method 300 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 at at least two different locations along a main fuel path leading from the location 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 400 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 1004, 1006.Figure 16 shows an oil circulation system comprising two circuits 2000, 2000' but without heat exchangers arranged on parallel branches as shown in Figure 11. Instead, each oil circuit 2000, 2000' provides a series connected main oil flow path through all heat exchangers in that circuit. In order to adjust the oil flow through the heat exchangers, instead of branching paths to different heat exchangers, one or more bypass pipes 2005, 2005', 2005a are provided. It should 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 the parallel arrangement 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 (i) parallel arrangements of heat exchangers and (ii) bypass tubes may be used and that FIGS. 11 (both circuits have parallel branches, no bypass tubes) and 16 (all heat exchangers parallel, multiple bypass tubes) may be considered to show two different ends of a design spectrum.In the example shown in FIG. 16, the fuel flow is shown with a thick black line to provide a context for the interaction between the fuel and oil systems.The primary oil circulation system 2000 illustrated in FIG. 16 provides a series-connected oil flow path from a tank 2002 via an oil pump 2004, further through the primary fuel-oil heat exchanger 1006, and then the air-oil heat exchanger 2020, before entering the transmission 30 (and optionally also into other components to be cooled and lubricated) and then being collected in an oil pan 2008 and then pumped back into the tank 2002 by the oil pump 2010. The two heat exchangers 1006, 2020 are thus present in a series circuit. The order of the two heat exchangers 1006, 2020 may be reversed such that the fuel-oil heat exchanger 1006 is after the air-oil heat exchanger 2020 in alternative implementations.The primary oil circulation system 2000 illustrated in FIG. 16 comprises two bypass pipes for oil 2005, 2005 a. The first bypass pipe 2005 is disposed such that a part of the oil can bypass the fuel-oil heat exchanger 1006, and is controlled by a first bypass valve 2007. The first bypass pipe 2005 guides the oil from upstream of the inlet to the primary heat exchanger 1006 and returns it to the main oil flow path downstream of the primary heat exchanger 1006 and upstream of the air-oil heat exchanger 2020. The second bypass pipe 2005 ais disposed such that a part of the oil can bypass the air-oil heat exchanger 2020, and is controlled by a second bypass valve 2007 a. The second bypass pipe 2005 guides the 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 transmission 30 (and optionally also other components to be cooled and lubricated). In implementations of the primary oil circulation system 2000 with only one bypass pipe for oil 2005 a, the selected position may be that of the second bypass pipe 2005 a, 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 on the fuel-oil heat exchanger 1006 may facilitate a quick adjustment of the heat transfer ratio when there is a risk of it falling below the desired value (for example, depending on the fuel properties determined, a lower limit for the ratio may be set significantly above zero). Severe fixed temperature limits may apply to some fuels depending on their properties. The presence of a bypass pipe 2005, 2005 aon both heat exchangers 1006, 2020 can prevent the oil from becoming too cold under certain circumstances, and thus avoid the risk of excessive oil clotting. It will be appreciated that this may pose a greater problem at take-off or ground idle on a cold day than at cruise.The secondary oil circulation system 2000' shown in FIG. 16 includes a single heat exchanger 1004, which is the secondary fuel-oil heat exchanger, and a single bypass pipe 2005' arranged so that a part of the oil can bypass this heat exchanger 1004. In other implementations, no bypass pipe may be provided on the secondary oil circulation system 2000', or more than one heat exchanger may be present (e.g., an air-oil or oil-oil heat exchanger in addition to the secondary fuel-oil heat exchanger 1004), and more than one bypass pipe may be provided, optionally one for each heat exchanger. A bypass valve 2007' is again provided to control the flow of oil through the bypass pipe 2005'.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 500, 600 of FIGS. 17 and 18 are concerned with these two scenarios of aircraft operation.FIG. 17 illustrates a method 500 implementing these considerations under cruise conditions and FIG. 18 illustrates a method 600 implementing these considerations idle when the aircraft 1 is on the ground (ground idle), e.g., while the aircraft is being launched, while stationary running during boarding and while rolling (to a runway or hangar, or between other locations on the ground), or at certain times during flight (flight idle).FIG. 17 illustrates the method 500 to be performed in cruise flight. The method 500 is configured to be performed in a transmission gas turbine engine 10 comprising an oil cycle system 2000 configured to supply oil to the transmission 30; and a heat exchange system 3000 comprising 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. The oil circulation system 2000 further includes at least one bypass pipe 2005, 2005a as shown in FIG. 16, so that a part of the oil can be led past at least one heat exchanger. In addition, a bypass control valve 2007, 2007 ais provided in order to control the proportion of the oil conducted via the bypass pipe 2005, 2005 aand thus also the amount of oil conducted through the respective heat exchanger 1006, 2020. The method 500 includes controlling 502 the one or more bypass valves 2007, 2007 asuch that under cruise conditions a heat transfer ratio is in the range of 0 to 0.67, and optionally 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 controller 58 - optionally either a stand-alone unit or a portion of an EEC - may be provided to implement this control.The method 500 may further include receiving data 504 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 and / or oil flow rate data. Such data may be received 504 and used 506 from a controller 58 to adjust the controller 502 of the cruise bypass valve 2007, 2007 ato maintain the heat transfer ratio at a desired level or within desired limits. This checking and adjustment / correction 506 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 504, 506 can alternatively also be considered part of the control 502 of the bypass valve 2007, 2007 a. The method 500 may also be configured to use other information, e.g., temperature data (of the oil, fuel, and / or ambient temperature of an environment surrounding the aircraft 1), flow rate data (of the oil and / or fuel), and / or one or more fuel characteristics, in determining 506 which control actions to take.In implementations in which a bypass pipe 2005 ais provided via the air-oil heat exchanger 2020, the controlled bypass valve may be the valve 2007 afor the bypass pipe 2005 avia the air-oil heat exchanger 2020. The step 502 of controlling the bypass valve 2007 ato adjust the heat transfer ratio may include reducing the amount of oil routed via the air-oil heat exchanger 2020 when the heat transfer ratio is too high. In some such implementations, this bypass pipe 2007 acan be the only bypass pipe for oil in the primary oil circulation system 2000. No bypass pipe may be provided to the fuel-oil heat exchanger 1006.In alternative implementations, such as that shown in FIG. 16, the heat exchange system 3000 includes at least two bypass pipes for oil and optionally three or more, each bypass pipe 2005, 2005', 2005a being arranged such that oil can bypass one of the heat exchangers 1006, 1004, 2020. Method 500 may include modulating the amount of oil directed across each bypass pipe 2005, 2005', 2005a. In embodiments with multiple bypass pipes within the same closed oil circuit system (e.g., with two bypass pipes in the primary oil circuit system 2000 as shown in FIG. 16 ), the same bypass valve 2007-which may be a three-way valve-may be used to control the flow through both bypass pipes 2005, 2005 a, or a different bypass valve 2007, 2007 amay be provided for each as shown in FIG. 16.It will be appreciated that control of the one or more bypass valves 2007, 2007a is not the only measure that can influence the heat transfer ratio - however, the use of bypass tubes can be used to provide quick corrections and plays a key role in maintaining the ratio within the desired limits. Other components of the heat exchange system, e.g., a recirculation valve 6010 and / or a refrigeration cycle device 1007, may also be controlled, and their control may affect the heat transfer ratio. For example, the heat exchange system 3000 may include at least one return pipe 6011 arranged such that a liquid may pass through a heat exchanger multiple times. The method 500 may further include, in such implementations, modulating the amount of liquid routed through the return tube 6011 by controlling the corresponding valve 6010 to adjust the heat transfer ratio. Additionally or alternatively, the heat exchange system 3000 may include a refrigeration cycle device 1007, and the method 500 may further include using the refrigeration cycle device 1007 to provide thermal buoyancy by transferring further heat from the oil to the fuel.In implementations with a refrigeration cycle device 1007, the bypass valve 2007, 2007 acan be controlled such that the heat transfer ratio is in the range of 0 to 0.40. In implementations without a refrigeration cycle device, in which implementations the heat exchange system 3000 is not configured to provide thermal lift, the bypass valve may be controlled such that the heat transfer ratio is generally higher and may be in the range of 0.38 to 0.67.In some implementations, one or more fuel temperature sensors may be provided, and data regarding fuel temperature entering the combustion chamber 16 may be used for fine tuning control decisions. The method 500 may include controlling 502 the bypass valve 2007, 2007 aunder cruise conditions such that the heat transfer ratio is in the range of 0 to 0.2 if the fuel temperature enters the combustion chamber 16 is at least 160° C.; and / or such that the heat transfer ratio is in the range of 0 to 0.1 if the fuel temperature enters the combustion chamber 16 is at least 180° C.In some implementations, one or more fuel characteristics may be considered when determining how to control the bypass valve 2007, 2007 a(the fuel characteristics may be determined using any of the approaches described above). For example, the method 500 may include controlling 502 the bypass valve 2007 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; and / or 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 500 may include controlling 502 the bypass valve 2007, 2007 aunder cruise conditions such that the oil-to-air heat transfer rate may be maintained in the range of 0 to 240 kJ per kilogram of fuel under cruise conditions, wherein no more than 20% of the heat dissipated from the oil under cruise conditions is transferred to the air, and / or such that the oil-to-fuel heat transfer rate is maintained in the range of 85 to 350 kJ per kilogram of fuel under cruise conditions, wherein at least 80% of the heat dissipated from the oil under cruise conditions is transferred to the fuel.In some implementations, heat exchange system 3000 further includes branching fuel return paths 6020 and at least one valve that controls a split of the fuel flow as described above. The valve may be controlled based on feedback from one or more temperature sensors and / or based on fuel characteristics. In implementations with multiple fuel-oil heat exchangers 1004, 1006, branching fuel return paths 6020 may be provided for one or both of the heat exchangers 1004, 1006. The method 500 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 1004, 1006.FIG. 18 illustrates method 600 to be performed at idle. Method 600 is configured to be performed in a gas turbine engine 10, as described above for method 500 of FIG. 17. It should be appreciated that the same engine 10 may be used to implement the cruise method 500 of FIG. 17 and the idle method 600 of FIG. 18.The idle method 600 includes controlling 602 the bypass valve 2007, 2007 ato have a heat transfer ratio in the range of from 0.67 to 5.67, and optionally from 0.67 to 5.50, from 0.67 to 5.00, from 0.67 to 4.50, from 0.67 to 4.00, from 0.67 to 3.50, from 0.67 to 3.50, or from 0.67 to 2.50, under idle conditions. Therefore, the ratio may be higher at idle than at cruise. The ratio can be higher in ground idling than in flight idling. Optionally, method 600 may include maintaining the heat transfer ratio under ground idle conditions below 5.50, 5.0, 4.5, 4.0, 3.5, or 3.0. Optionally, method 400 may include maintaining the heat transfer ratio at above 4.0, 4.5 under idle conditions or at 5.0 under ground idle conditions.Method 600 may include controlling the bypass valve such that the heat transfer ratio is above 1.0, and optionally above 1.5, 2.0, under idle conditions. A controller 58 - optionally either a stand-alone unit or a portion of an EEC - may be provided to implement this control.The method 600 may further include receiving 604 data 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 along a fuel flow path and / or optionally a fuel tank temperature or oil tank temperature) and fuel and / or oil flow rate data. Such data may be received 604 and used 606 by a controller 58 to adjust the controller 602 of the idle bypass valve 2007, 2007 ato maintain the heat transfer ratio at a desired level or within desired limits. This checking and adjustment / correction 606 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, such as the start of coasting). These steps 604, 606 can alternatively be considered part of the control 602 of the bypass valve 2007, 2007 a. The method 600 may also be configured to use other information, e.g., temperature data (of the oil, fuel, and / or ambient temperature of an environment surrounding the aircraft 1), flow rate data (of the oil and / or fuel), and / or one or more fuel characteristics, in determining 606 which control actions to take.The step 602 of controlling one or more bypass valves 2007, 2007 afor adjusting the heat transfer ratio may include reducing the amount of oil directed through the air-oil heat exchanger 2020 via its respective bypass pipe 2005 awhen the heat transfer ratio is too high. The heat exchange system 3000 of some implementations, such as that shown in FIG. 16, includes at least two bypass pipes for oil 2005, 2005a, each bypass pipe being arranged such that oil can be passed by the air-oil heat exchanger 2020 and the fuel-oil heat exchanger 1006, respectively. As with the cruise method 500, the idle method 600 may include modulating the amount of oil directed across each bypass pipe 2005, 2005a. The presence of a bypass pipe 2005, 2005 aat both heat exchangers 1006, 2020 of the primary oil circuit 2000 can prevent the oil from becoming too cold under certain circumstances, thus avoiding the risk of excessive oil clotting. It will be appreciated that this may pose a greater problem at take-off or ground idle on a cold day than at cruise. The oil circuit system 2000 may therefore also include one or more oil temperature sensors, and the method 600 described may include receiving inputs from the one or more oil temperature sensors and optionally also from one or more temperature sensors arranged to sense an ambient temperature around the aircraft 1, and adjusting the bypass valve or valves controller 602 based on this temperature information. For example, on a cold day, less oil may be directed across the air-to-oil heat exchanger 2020 to reduce heat loss to the environment.In some implementations, further bypass pipes 2005' may be provided, for example for the flow of oil in the primary oil circuit system 2000' via an oil-to-oil heat exchanger 2030 and / or via one or more heat exchangers in a secondary oil circuit system 2000', e.g. via a secondary fuel-to-oil heat exchanger 1004.The method 600 of some implementations includes controlling other components in addition to the one or more bypass valves 2007, 2007 ato achieve the desired heat transfer ratio. For example, at least one return pipe 6011 arranged such that a liquid can pass multiple times through a heat exchanger and / or a refrigeration cycle device 1007 may be provided, and the method may further comprise modulating the amount of liquid routed via the respective additional component(s) or, as needed, activating or deactivating the respective component(s) to adjust the heat transfer ratio. In implementations with a refrigeration cycle device 1007, the bypass valve 2007, 2007 acan be controlled such that the heat transfer ratio is in the range of 0.67 to 4.0. In non-refrigeration cycle device implementations in which the heat exchange system 3000 is not configured to provide thermal lift, the bypass valve 2007, 2007 acan be controlled such that the heat transfer ratio is generally higher, optionally in the range of 2.00 to 5.67, and optionally 3.37 to 5.67.In some implementations, one or more temperature sensors may be provided, and data regarding fuel temperature entering the combustion chamber 16 may be used for fine tuning control decisions. The method 600 may include controlling the one or more bypass valves 2007, 2007a under idle conditions such that the heat transfer ratio is one of: (i) in the range of 2.33 to 5.67 when the fuel temperature at entry into the combustion chamber is below 200° C.; (ii) in the range of 0.67 to 5.00 and optionally 0.67 to 4.00 when the fuel temperature at entry into the combustion chamber (16) is above 200° C.; (iii) in the range of 0.67 to 4.00 and optionally 0.67 to 2.67 when the fuel temperature at entry into the combustion chamber (16) is above 250° C.; and / or (iv) in the range from 0.67 to 2.33 and optionally from 0.67 to 1.22 if the fuel temperature on entry into the combustion chamber (16) is above 280° C.One or more fuel properties - optionally determined by a fuel property determination module 57 as described above - may also be used; the information may be taken into account in adjusting the controller 602 of the bypass valve 2007, 2007a. For example, method 600 may take into account the SAF content of the fuel. The method 600 may include controlling 602 the bypass valve 2007, 2007 asuch that the heat transfer ratio is in the range of 0.67 to 3.67 if the fuel is at least 70% persistent flight fuel and / or in the range of 0.67 to 2.67 if the fuel is at least 80% persistent flight fuel. The method 600 may include controlling 602 the bypass valve 2007, 2007 aunder idle conditions such that the oil-to-air heat transfer rate at idle conditions may be maintained in the range of 350 to 1270 kJ per kilogram of fuel, wherein no more than 20% of the oil-removed heat at ground idle is transferred to the air, and / or such that the oil-to-fuel heat transfer rate at idle is maintained in the range of 350 to 1270 kJ per kilogram of fuel, wherein at least 80% of the oil-removed heat at idle is transferred to the fuel.In some implementations, heat exchange system 3000 further includes branching fuel return paths 6020 and at least one valve that controls a split of the fuel flow as described above.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.References included in the specificationThis list of documents cited by the applicant has been produced in an automated manner and is only included for the better information of the reader. The list is not part of the German patent application or utility model application. The DPMA does not take any adhesion for any faults or omissions.Cited Non-Patent LiteratureSAE AS755

[0210]

Claims

A method (600) of operating a gas turbine engine (10) for an aircraft, 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 gearbox (30) receiving input from the core shaft (26) and outputting drive to the fan (23) to drive the fan at a lower speed than the core shaft, an oil circulation system (2000) arranged to supply the gearbox (30) 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; and a fuel-oil heat exchanger (1006) through which the oil in the oil circuit system and the fuel flow so that heat is transferred between the oil and the fuel, and wherein the oil circuit system (2000) comprises at least one bypass pipe (2005, 2005a), the bypass pipe being arranged to allow oil to be bypassed by one of the air-oil heat exchanger and the fuel-oil heat exchanger; and a bypass valve (2007, 2007a) configured such that the proportion of the oil conducted via the bypass pipe (2005, 2005a) can be varied, the method (600) comprising controlling (602) the bypass valve (2007, 2007a) such that, under idling conditions, a heat transfer ratio of: W a%0020̈ rme u%0020̈ the transfer rate of O%0020̈ l to air (kJkg-1) W a%0020̈ rme u%0020̈ the transfer rate of O%0020̈ l to fuel (kJkg-1) is in the range of 0.67 to 5.67.The method (600) of claim 1, comprising controlling (602) the bypass valve (2007, 2007a) such that under idle conditions the heat transfer ratio is below 5.50.The method (600) of claim 1, wherein controlling (602) the bypass valve (2007, 2007a) to adjust the heat transfer ratio comprises decreasing the amount of oil directed across the at least one air-oil heat exchanger (2020) when the heat transfer ratio is too high.The method (600) of claim 1, wherein the heat exchange system (3000) comprises at least two bypass pipes for oil (2005, 2005a), each bypass pipe arranged to allow oil to bypass one of the air-oil heat exchanger (2020) and the fuel-oil heat exchanger (1006), and wherein the method comprises modulating the amount of oil directed across each bypass pipe (2005, 2005a).The method (600) 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 the method (600) further comprises modulating the amount of liquid routed via the return pipe (6011) to adjust the heat transfer ratio.The method (600) of claim 1, wherein the heat exchange system (3000) further comprises a refrigeration cycle device (1007), and the method (800) 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 increased above the oil temperature.The method (600) of claim 6, comprising controlling (602) the bypass valve (2007, 2007a) such that the heat transfer ratio is in the range of 0.67 to 4.The method (600) of claim 1, wherein the heat exchange system (3000) is not configured to provide thermal lift, and wherein the method comprises controlling (602) the bypass valve (2007, 2007a) such that the heat transfer ratio is in the range of 2.00 or 3.37 to 5.67.The method (600) of claim 1, wherein the method (600) comprises controlling (602) the bypass valve (2007, 2007a) under idle conditions such that the heat transfer ratio is in the range of 2.33 to 5.67 when the temperature of the fuel entering the combustion chamber (16) is below 200°C.The method (600) of claim 1, wherein the method (600) comprises controlling (602) the bypass valve (2007, 2007a) under idle conditions such that the heat transfer ratio is in the range of 0.67 to 4.00 when the temperature of the fuel entering the combustion chamber (16) is above 200°C.The method (600) of claim 1, wherein the method (600) comprises controlling (602) the bypass valve (2007, 2007a) such that the heat transfer ratio is in the range of 0.67 to 3.67 when the fuel is at least 70% sustainable flight fuel under idle conditions.The method (600) of claim 1, comprising controlling (602) the bypass valve (2007, 2007a) such that under idle conditions the heat transfer ratio is above 0.75.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; and a gearbox (30) 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 circulation system (2000) disposed to supply the gearbox (30) 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; and a fuel-oil heat exchanger (1006) through which the oil in the oil circuit system and the fuel flow so that heat is transferred between the oil and the fuel, and wherein the oil circuit system (2000) comprises at least one bypass pipe (2005, 2005a), the bypass pipe being arranged to allow oil to be bypassed by one of the air-oil heat exchanger and the fuel-oil heat exchanger; and at least one bypass valve (2007, 2007a) configured such that the proportion of the oil which is conducted via the bypass pipe (2005, 2005a) can be varied, and wherein the bypass valve (2007, 2007a) is configured such that it can be controlled such that, under idling conditions, a heat transfer ratio W a%0020̈ rme u%0020̈̈ transfer rate of O%0020̈ l on air (kJkg-1) W a%0020̈ rme u%0020̈ transfer rate of O%0020̈ l on fuel (kJkg-1) is in the range from 0.67 to 5.67.The gas turbine engine (10) of claim 13, 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 13, wherein the heat exchange system (3000) further comprises branching fuel return paths and at least one valve that controls a split 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 path leading from the location where the fuel enters the gas turbine engine (10) to the combustor (16).

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