Gas turbine modulating valve

Abstract

A method 100 of operating a gas turbine engine of an aircraft comprising an engine core comprising a turbine, a compressor, a combustion chamber for burning fuel, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core; a reduction gear which receives an input from the core shaft and provides an output drive to the fan; an oil loop system for supplying oil to the reduction gear; and a heat exchange system comprising an air-oil heat exchanger through which oil flows; and a fuel-oil heat exchanger through which oil and fuel flow; and a modulating valve for allowing the proportion of oil sent through each heat exchanger to be varied, the method comprising: determining 102 at least one fuel characteristic of the fuel to be burned by the combustion chamber;and controlling 104 the modulation valve based on the at least one fuel characteristic so as to adjust the proportion of oil sent through each heat exchanger to cruising conditions. To be reproduced with Figure 12 upon publication Figure for abstract: Fig. 12;

Description

Title of the invention: Gas turbine modulation valve

[0001] The present description relates to aircraft propulsion systems, as well as to methods of operating aircraft involving the management of different fluids and the transfer of heat between them, and in particular to the management of a heat exchange system of an aircraft engine.

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

[0003] According to a first aspect, a method of operating a gas turbine engine of an aircraft is provided, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, a combustion chamber for burning a fuel, and a core shaft connecting the turbine to the compressor; a blower located upstream of the engine core; a reducer which receives an input from the heart shaft and delivers an output to the blower; an oil loop system for supplying oil to the reducer; and a heat exchange system comprising: an air-oil heat exchanger through which the oil in the oil loop system flows; and a fuel-oil heat exchanger through which the oil in the oil loop system and the fuel flow such that heat is transferred between the oil and the fuel; and a modulating valve for varying the proportion of oil sent through each heat exchanger, the method comprising: determining at least one fuel characteristic of the fuel to be burned by the combustion chamber; and controlling the modulation valve based on the at least one fuel characteristic so as to adjust the proportion of oil sent through each heat exchanger to cruising conditions.

[0004] The inventors were aware that the use of fuels different from traditional kerosene-based jet fuels, such as sustainable aviation fuels, may lead to different fuel properties, and that parameters at cruise conditions may be adjusted to utilize the different fuel properties. In particular, certain fuels may be heated to higher temperatures in one or more fuel-oil heat exchangers than traditional fuels, without significantly increasing fuel degradation products, including without increasing coking and / or varnishing. This may improve fuel combustion efficiency and / or improve the overall thermodynamic efficiency of the engine, with less heat lost to the environment (e.g., through the air-oil heat exchanger).Higher fuel temperatures at the combustion chamber inlet can enable a process that provides improved oil cooling (because the fuel is able to absorb more heat). For example, fuel thermal stability affects the amount of heat the fuel can accept / the temperature to which the fuel can be heated without forming deposits in lines, burners and / or a hydromechanical unit or other engine component. Taking fuel thermal stability into account and transferring more or less heat from the oil to the fuel depending on the fuel thermal stability can therefore allow for more efficient oil cooling while avoiding coking and / or varnishing, thereby improving aircraft performance.

[0005] Using the fuel to absorb more heat from the oil, rather than relying on heat transfer from the oil to the environment / air (e.g., in an air-to-oil heat exchanger) can also provide a more thermally efficient engine. This can improve the cooling of the oil before it is returned to the rest of the turbine engine. In addition, the improved cooling of the oil can in turn improve the cooling effect of the oil on the engine components through which it flows, for example allowing a lower oil flow rate to provide the same cooling effect.

[0006] In a transition period, where available aviation fuels change over the lifetime of a gas turbine engine, and also depend on the geographical location of a refueling point, among other variables, it is important to determine a relevant fuel characteristic and control the heat exchange system based on the specific fuel used. Thus, the operation of gas turbines can be adapted to take best advantage of a wide variety of fuels. The controllable modulation valve, intended to adjust The flow of oil through each heat exchanger has a key role to play in controlling the heat exchange system.

[0007] The fuel characteristic may be or include a calorific value, thermal stability, or the percentage of sustainable aviation fuel (SAF) in the fuel.

[0008] The air-oil heat exchanger may help remove excess heat from the oil in the oil loop system that is not transferred to the fuel. This may allow for further cooling of the oil beyond the heat transfer from the oil to the fuel in the fuel-oil heat exchanger.

[0009] The oil loop system, which may also be referred to as an oil recirculation system, may branch such that a proportion of the oil may flow along each branch. The air-to-oil and fuel-to-oil heat exchangers may be arranged in a parallel configuration on different branches of an oil loop system. The modulating valve of such examples may be arranged to allow the proportion of oil sent through each branch to be varied, and control of the modulating valve may therefore adjust the proportion of oil sent through each branch to suit cruising 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 loop system may join after the heat exchanger system, so that the oil recombines after heat transfer from the oil.

[0010] The modulating valve allows the amount of heat transferred from the oil by the air-oil and fuel-oil heat exchangers to be varied. The level of oil flow (and thus the heat transfer from the oil to the air or fuel) can be varied based on the temperature of the fuel exiting the fuel-oil heat exchanger, or entering the combustion chamber, and can thus allow the fuel temperature to be controlled within a defined upper and lower limit. Similarly, the oil flow control can allow the oil temperature to be maintained within an appropriate range.This regulation may help to ensure improved turbine engine efficiency (e.g., by raising the fuel temperature) without risking unnecessary damage to a fuel pump or other components downstream of the fuel-oil heat exchanger by using a temperature too high for the durability of said component. The modulation valve may be intended to allow 100% of the oil to be sent through the fuel-oil heat exchanger. The modulation valve may be intended to ensure that not less than 70%, 80%, or 90% of the oil is sent through the fuel-oil heat exchanger at cruise.

[0011] The modulation valve may be adapted to divert a fixed portion of the oil flow to each of the fuel-oil heat exchanger and the air-oil heat exchanger during engine operation, the fixed portion optionally being determined at or before engine start, or upon reaching cruising altitude or before, based on at least one determined fuel characteristic. Alternatively, the modulation valve may be adapted to divert a variable portion of the oil flow to each of the fuel-oil heat exchanger and the air-oil heat exchanger during engine operation.The modulation valve can therefore be actively controlled to vary the proportion of oil sent through each heat exchanger, particularly in implementations where the aircraft carries several different fuels in different tanks and may change the fuel (or fuel mixture) used during flight. Active control of the modulation valve can be automated and implemented by a controller of the heat exchange system.

[0012] The gas turbine engine may further comprise a generator (e.g., an integrated drive generator) and a secondary oil loop system for supplying oil to the generator. The oil loop system described above, which supplies oil to the reduction gear, may therefore be referred to as a primary oil loop system. The heat exchange system may comprise an air-oil heat exchanger for transferring heat between the primary and secondary oil loop systems. The modulating valve, or an additional valve, may control an oil flow from the primary oil loop system to the oil-oil heat exchanger.

[0013] The heat exchange system may further include a secondary fuel-oil heat exchanger for receiving fuel and oil from the secondary oil loop system. The method may include transferring heat between oil from the secondary oil loop system and fuel using the secondary fuel-oil heat exchanger. The secondary fuel-oil heat exchanger may be referred to as a secondary heat exchanger for brevity.

[0014] The fuel-oil heat exchanger through which oil of the primary oil loop system flows may be referred to as the primary fuel-oil heat exchanger. The primary fuel-oil heat exchanger may be referred to as the primary heat exchanger for brevity. Fuel may flow through the secondary fuel-oil heat exchanger before flow through the primary fuel-oil heat exchanger, such that heat is transferred from the oil in the secondary oil loop system. to the fuel before heat is transferred from the oil in the primary oil loop system to the fuel.

[0015] The heat exchange system may include at least one bypass line for allowing fuel (or oil) to bypass a heat exchanger, or multiple heat exchangers, of the heat exchange system. A bypass line may effectively form an additional branch in a parallel branch oil system, in some implementations - the modulating valve, or other oil valve, may be for adjusting the amount of oil sent through the or each bypass line for oil, based on the one or more determined fuel characteristics, and possibly on one or more temperature measurements.

[0016] The heat exchange system may further include a refrigeration cycle apparatus, and the method may include using the refrigeration cycle apparatus to provide a thermal lift by transferring more heat from the oil to the fuel, possibly such that the fuel temperature is raised above the oil temperature. The modulating valve, or another oil valve, may regulate the amount of oil that flows through the refrigeration cycle apparatus.

[0017] The heat exchange system may further include branched fuel return paths and at least one valve controlling a fuel flow split. The branched paths may be configured to return fuel from the heat exchange system to at least two different locations along a main fuel path from where the fuel enters the gas turbine engine to the combustion chamber. For example, fuel leaving a fuel-oil heat exchanger may be split into two or more branches, each branch rejoining the main fuel flow path at a different point.In some implementations, at least the majority of the fuel may flow through the fuel-oil heat exchanger, such that a branch of the branched fuel path leaving that heat exchanger is itself the "main" fuel path - one or more smaller branches may be split and rejoin the main flow at a different location, e.g., further downstream, e.g., past one or more engine components located downstream of the fuel-oil heat exchanger in question, or even further upstream (thus acting as a recirculation pipe).

[0018] Subject to the suitability of the at least one determined fuel characteristic, the method may comprise passing all of the oil through the fuel-oil heat exchanger for at least one or more periods continuous for at least 30 minutes at cruise, such that no heat is lost to the environment through the air-oil heat exchanger during at least some periods of cruise operation. The method may include passing all of the oil through the fuel-oil heat exchanger for at least 15% of the total time spent at cruise, and optionally at least 20%. Subject to the suitability of the at least one determined fuel characteristic, the method may include passing at least 95% of the oil through the fuel-oil heat exchanger for at least 90% of the time spent at cruise, such that very little, if any, heat is lost to the environment through the air-oil heat exchanger for at least 90% of the cruise operation.Subject to the suitability of the at least one determined fuel characteristic, the method may include transferring at least 80% of the heat transferred from the oil at cruise to the fuel. All of the heat transferred from the oil at cruise may be transferred to the fuel for at least 90% of the time spent at cruise in some implementations.

[0019] The at least one fuel characteristic may be or include thermal stability. At least 80% of the heat transferred from the oil at cruise may be transferred to the fuel, provided that the fuel is stable in operation at temperatures above 160°C.

[0020] The at least one fuel characteristic may be or include an aromatic hydrocarbon content of the fuel. At least 80% of the heat transferred from the oil at cruise may be transferred to the fuel provided that the fuel has a mole percentage of aromatic hydrocarbons below 12%.

[0021] The at least one fuel characteristic may be or include a percentage of sustainable aviation fuel - %SAF - in the fuel. The proportion of SAF (X%) may be volumetric. At least 80% of the heat transferred from the oil at cruise may be transferred to the fuel, provided that the fuel has a SAF content above 50%.

[0022] The at least one fuel characteristic may be or include a calorific value of the fuel. At least 80% of the heat transferred from the oil at cruising speed may be transferred to the fuel, provided that the fuel has a calorific value of at least 43.5 MJ / kg.

[0023] According to a second aspect, a gas turbine engine for an aircraft is provided, 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 located upstream of the engine core; and a reducer which receives an input from the core shaft and outputs a drive to the blower so as to drive the blower at a lower rotational speed than that of the core shaft; an oil loop system for supplying oil to the reducer; and a heat exchange system comprising: an air-oil heat exchanger through which the oil in the oil loop system flows; and a fuel-oil heat exchanger through which the oil in the oil loop system and the fuel flow such that heat is transferred between the oil and the fuel; and a modulation valve intended to allow the proportion of oil sent through each heat exchanger to be varied; and a fuel composition determination module, intended to determine at least one fuel characteristic of the fuel intended to be burned by the combustion chamber, wherein the modulation valve is adapted to be controlled based on the at least one fuel characteristic so as to adjust the proportion of oil sent through each heat exchanger to cruising conditions.

[0024] 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 comprise 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 configured to rotate at a higher rotational speed than the first core shaft.

[0025] The motor of the second aspect may be for carrying out the method of the first aspect, and may have any of the characteristics described in relation to the first aspect.

[0026] According to a third aspect, a method of operating a gas turbine engine is provided, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, a core shaft connecting the turbine to the compressor, and a combustion chamber for burning a fuel; a blower located upstream of the engine core; an oil system for circulating oil; and a heat exchange system comprising at least one fuel-oil heat exchanger for transferring heat from the oil to the fuel, wherein the method comprises: determining at least one fuel characteristic of the fuel to be supplied to the combustion chamber; and modulating the heat exchange system so as to adjust the fuel temperature at the combustion chamber inlet at cruising speed to a set level, the set level being based on the at least one fuel characteristic.

[0027] The inventors were aware that the use of fuels different from traditional kerosene-based jet fuels, such as sustainable aviation fuels, may lead to different fuel properties, and that parameters at cruise conditions may be adjusted to utilize the different fuel properties. In particular, certain fuels may be heated to higher temperatures in one or more fuel-oil heat exchangers than traditional fuels, without significantly increasing the formation of fuel degradation products, including without increasing coking and / or varnishing. This may improve the combustion efficiency of the fuel.Higher fuel temperatures at the combustion chamber inlet can enable a process that provides improved oil cooling (because the fuel is able to absorb more heat) and / or improved fuel combustion efficiency. For example, fuel viscosity affects how fuel is delivered into and ignited by the combustion chamber. Viscosity can affect the droplet size from fuel spray nozzles, which in turn can impact combustion efficiency. Taking fuel viscosity into account when delivering fuel to the combustion chamber, and controlling it appropriately by varying the heat input, can therefore provide more efficient fuel combustion, improving aircraft performance.In a transition period, where available aviation fuels change over the lifetime of a gas turbine engine, and also depend on the geographical location of a refueling point, among other variables, it is important to determine a relevant fuel characteristic and control the heat exchange system based on the specific fuel used in the engine. Thus, the operation of gas turbines can be adapted to take best advantage of a wide variety of fuels. The controllable heat exchange system has a key role to play in managing fuel and oil temperatures.

[0028] The fuel characteristic may be or include the percentage of sustainable aviation fuel (SAF) in the fuel, the amount (e.g., ppm) of sulfur in the fuel, or the thermal stability of the fuel.

[0029] The set level may be an average of at least 140°C, 150°C, 160°C, 170°C, 180°C, 190°C, or 200°C at the combustion chamber inlet at cruise conditions. The fuel temperatures at the combustion chamber inlet at cruise conditions may be defined as an average for at least 5 minutes, and optionally for ten minutes, twenty minutes, or thirty minutes, at steady-state cruise conditions. These average temperatures do not include transient temperature spikes, which may be defined as fluctuations in fuel temperature during operation, often a rise in temperature. Each fluctuation may last no more than 5 minutes.A fuel temperature of at least 140°C at the combustion chamber inlet at cruise conditions as defined herein therefore requires that the fuel temperature remain at 140°C or above for a period of time, and a transient spike to a temperature above 140°C is not sufficient.

[0030] The set level may be higher than the oil temperature in some implementations. The heat exchange system may further include a refrigeration cycle apparatus, and the method may include using the refrigeration cycle apparatus to provide a thermal lift by transferring more heat from the oil to the fuel such that the fuel temperature is raised above the oil temperature to reach the set level.

[0031] The gas turbine engine may further include a reduction gear that receives an input from the core shaft and outputs a drive to the fan, possibly via a fan shaft. The oil system may be for circulating oil between the reduction gear and the at least one fuel-oil heat exchanger, and may be described as an oil loop system for supplying oil to the reduction gear.

[0032] The heat exchange system may include at least one air-to-oil heat exchanger, and modulating the heat exchange system to adjust the fuel temperature may include modulating the amount of oil sent through the at least one air-to-oil heat exchanger.

[0033] The heat exchange system may include at least one bypass line for allowing oil to bypass a heat exchanger. Modulating the heat exchange system to adjust the fuel temperature to the set level may include modulating the amount of oil sent through the bypass line. The heat exchange system may include at least one bypass line for allowing fuel to bypass a heat exchanger. Modulating the heat exchange system to adjust the fuel temperature to the set level may include modulating the amount of fuel sent through the bypass line.

[0034] The heat exchange system may include at least one recirculation line for allowing a fluid (e.g., oil or fuel) to pass multiple times through a heat exchanger. Modulating the heat exchange system to adjust the fuel temperature to the set level may include modulating the amount of fluid sent through the or each recirculation line.

[0035] The step of modulating the heat exchange system may comprise regulating oil flow through the at least one fuel-oil heat exchanger such that between 50% and 100% of the waste heat from the oil is transferred to the fuel. The step of modulating the heat exchange system may comprise regulating oil flow through the at least one fuel-oil heat exchanger such that between 80% and 100% of the waste heat from the oil is transferred to the fuel.

[0036] The at least one determined fuel characteristic may be or include thermal stability of the fuel, and the method may include modulating / controlling the heat exchange system such that the fuel temperature at the inlet of the combustion chamber at cruising speed increases with increasing thermal stability, possibly linearly.

[0037] The at least one determined fuel characteristic may be or include a percentage of SAF in the fuel, and the method may include modulating the heat exchange system such that the fuel temperature at the inlet of the combustion chamber at cruise increases with increasing %SAF once the %SAF exceeds 60%.

[0038] The at least one determined fuel characteristic 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 at the inlet to the combustion chamber at cruising speed is set at a predefined level corresponding to the tracer species.

[0039] The at least one determined characteristic of the fuel 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 at the inlet to the combustion chamber at cruise is increased when the sulfur content is lower, and vice versa.

[0040] The methods of the first and third aspects may be complementary, and may be performed together in various implementations. The method of the third aspect may be performed using the engine of the second aspect.

[0041] According to a fourth aspect, a gas turbine engine for an aircraft is provided, the engine comprising: an engine core comprising a turbine, a compressor, a core shaft connecting the turbine to the compressor, and a combustion chamber for burning a fuel; a blower located upstream of the engine core; an oil system for circulating oil; a heat exchange system comprising at least one fuel-oil heat exchanger for transferring heat from the oil to the fuel; and a fuel composition determination module for determining at least one fuel characteristic of the fuel to be burned by the combustion chamber, and wherein the heat exchange system is adapted to be modulated so as to adjust the fuel temperature at the inlet of the combustion chamber to a set level, the set level being based on the at least one fuel characteristic.

[0042] The gas turbine engine may further include a reduction gear that receives an input from the core shaft and outputs a drive to the fan. The oil system may be for circulating oil between the reduction gear and the at least one fuel-oil heat exchanger, and may be described as an oil loop system for supplying oil to the reduction gear.

[0043] 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 comprise 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 configured to rotate at a higher rotational speed than the first core shaft.

[0044] The heat exchange system may further include branched fuel return paths and at least one valve controlling a fuel flow split. The branched paths may be configured to return fuel from the heat exchange system to at least two different locations along a main fuel path from where the fuel enters the gas turbine engine to the combustion chamber.

[0045] The motor of the fourth aspect may be for carrying out the method of the first and / or third aspect, and may have any of the characteristics described in relation to the first or third aspect.

[0046] In any one of the first to fourth aspects: The at least one fuel characteristic of the fuel may include at least one of: i. a percentage of sustainable aviation fuel in the fuel; ii. a concentration of heteroatomic species in the fuel; iii. an aromatic hydrocarbon content of the fuel; iv. a multi-aromatic hydrocarbon content of the fuel; v. a percentage of nitrogen-containing species in the fuel; vi. a presence or percentage of a tracer species or trace element in the fuel; vii. a hydrogen / carbon ratio of the fuel; viii. a hydrocarbon distribution of the fuel; ix. a level of non-volatile particulate matter emissions during combustion; x. a naphthalene content of the fuel; xi. a sulfur content of the fuel; xii. a cycloparaffin content of the fuel; xiii. an oxygen content of the fuel; xiv. a thermal stability of the fuel; xv. a coking level of the fuel; xvi. an indication that the fuel is a fossil fuel; xvii. at least one of a density, a viscosity, a calorific value and a heat capacity.

[0047] The method may further comprise chemically or physically detecting one or more fuel-related parameters in a fuel tank after refueling (the fuel tank being 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 a shaft speed and a fuel mass flow rate, from which a heating value (a fuel characteristic) may be determined, or the detected parameters may be a fuel density and / or the presence of a tracer, both of which are fuel characteristics. Determining at least one fuel characteristic may comprise obtaining stored data on the fuel characteristics.The chemical and / or physical determination of one or more fuel parameters in the fuel tank may be accomplished by extracting a sample of the fuel from the fuel tank for off-site testing.

[0048] Determining the at least one fuel characteristic may include obtaining at least one fuel characteristic of any fuel already present in the fuel tank prior to refueling; determining at least one fuel characteristic of a fuel added to the fuel tank during refueling; and calculating at least one fuel characteristic of the resulting fuel in the fuel tank after refueling (based on this information).

[0049] The determination of at least one fuel characteristic may be performed based on a detection of at least one fuel property. The fuel property may be the fuel characteristic, or may be used to calculate or otherwise determine (e.g., by retrieving from a lookup table) the fuel characteristic. The detection may be performed on-site.

[0050] The determination of at least one fuel characteristic may be performed based on received fuel composition data. The fuel composition data may be provided to the aircraft during refueling. The fuel composition data may be entered manually.

[0051] At least one fuel characteristic may be derived from the performance of the gas turbine engine during at least one of taxi, takeoff, and climb of the aircraft.

[0052] According to a fifth aspect, a method of operating a gas turbine engine is provided, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, a combustion chamber for burning fuel, and a core shaft connecting the turbine to the compressor; a blower located upstream of the engine core; a reducer which receives an input from the core shaft and outputs a drive to the blower so as to drive the blower at a lower rotational speed than that of the core shaft; an oil loop system for supplying oil to the reducer; and a heat exchange system comprising: an air-oil heat exchanger through which the oil in the oil loop system flows; and a fuel-oil heat exchanger through which the oil in the oil loop system and the fuel flow such that heat is transferred between the oil and the fuel, and wherein the oil loop system branches such that a proportion 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 loop system; and a modulation valve intended to allow the proportion of oil sent through each branch to be varied, the method comprising controlling the heat exchange system such that, under cruising conditions, a heat transfer ratio of: oil to air heat transfer rate (kJkgk oil to fuel heat transfer rate (kJkg1) is in the range from 0 to 0.67.

[0053] The inventors were aware that the use of fuels different from traditional kerosene-based jet fuels, such as sustainable aviation fuels, may lead to different fuel properties, and that parameters at cruise conditions may be adjusted to utilize the different fuel properties. In particular, certain fuels may be heated to higher temperatures in one or more fuel-oil heat exchangers than traditional fuels, without significantly increasing coking or other pathways for deposition of fuel degradation products (e.g., varnishing).This can enable a process that provides improved oil cooling (as the fuel is able to absorb more heat) and can also improve the overall thermal / thermodynamic efficiency of the engine, with less heat lost to the environment and potentially also more energy recovered in the thermodynamic cycle. The controllable heat exchange system has a key role to play in managing the heat transfer ratio.

[0054] While it is understood that the ratio is dimensionless, heat transfer is measured per unit mass or volume of fuel in the examples described herein, thereby providing a normalized heat transfer rate for variations in fuel flow rate at cruise. For heat transfer from oil to air, the definition of "per unit mass or volume of fuel" can be equated with "per a fixed period of time relative to the fuel flow rate" to similarly provide a normalization for fuel flow rate. It will be understood that heat transfer using the heat exchanger must be completed before the fuel reaches the combustion chamber. Any additional temperature increase in the combustion chamber itself is due to combustion, not the heat exchange system.It will be borne in mind that the heat transfer is denoted per unit mass (kg) of fuel reaching the combustion chamber so as to adjust the fuel flow rate and any recirculation through one or more heat exchangers or bypass of one or more heat exchangers as described elsewhere herein. The amount of heat transferred to the fuel can therefore be calculated based on a fuel temperature approaching or entering the combustion chamber by comparison. with a fuel temperature in an aircraft fuel tank. In most implementations, any fuel temperature rise due to other engine components (as opposed to heat transfer from the oil) can be assumed to be minimal. The amount of heat transferred from the oil to the air can be determined from an oil temperature drop across the air-to-oil heat exchanger(s), or by comparing the heat gained by the fuel to an overall heat loss from the oil, and assuming the difference is due to heat loss from the oil to the air (all other sources of loss being negligible).

[0055] The method may include controlling the heat exchange system such that, under cruising conditions, the heat transfer ratio is in the range of 0 to 0.60, 0 to 0.50, 0 to 0.40, 0 to 0.30, 0 to 0.20, or 0 to 0.10.

[0056] Controlling the heat exchange system to adjust a heat transfer ratio may include decreasing the amount of oil sent through the at least one air-oil heat exchanger when the heat transfer ratio is too high.

[0057] The heat exchange system may include at least one bypass line for allowing oil to bypass a heat exchanger, and controlling the heat exchange system to adjust the heat transfer ratio may include modulating the amount of oil sent through the bypass line.

[0058] The heat exchange system may include at least one recirculation line for allowing a fluid (oil or fuel) to pass multiple times through a heat exchanger, and controlling the heat exchange system to adjust the heat transfer ratio may include modulating the amount of fluid sent through the recirculation line.

[0059] The heat exchange system may include a refrigeration cycle apparatus. The method may include using the refrigeration cycle apparatus to provide a thermal lift by transferring more heat from the oil to the fuel, possibly such that the fuel temperature is raised above the oil temperature. The heat transfer ratio of the examples using a refrigeration cycle apparatus may be in the range of 0 to 0.40.

[0060] In implementations where no refrigeration cycle apparatus is used / where the heat exchange system is not intended to provide thermal lift, the heat transfer ratio may be in the range of 0.38 to 0.67.

[0061] The method may include controlling the heat exchange system under cruising conditions such that the heat transfer ratio is in the range of 0 to 0.2, provided that the fuel temperature at the inlet to the combustion chamber is at least 160°C.

[0062] The method may include controlling the heat exchange system under cruising conditions such that the heat transfer ratio is in the range of 0 to 0.1, provided that the fuel temperature at the inlet to the combustion chamber is at least 180°C.

[0063] 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, provided that the fuel is at least 70% sustainable aviation fuel.

[0064] 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, provided that the fuel is at least 80% sustainable aviation fuel.

[0065] The method may include maintaining the oil-to-air heat transfer rate at cruise conditions in the range of 0 to 240 kJ per kilogram of fuel at cruise conditions, and optionally 0 to 120 kJ / kg, with no more than 20% of the heat transferred from the oil at cruise being transferred to the air.

[0066] The method may include maintaining the oil-to-fuel heat transfer rate at cruise in the range of 85 to 350 kJ per kilogram of fuel at cruise conditions, and optionally 85 to 170 kJ / kg, with at least 80% of the heat transferred from the oil at cruise being transferred to the fuel.

[0067] The methods of the first, third, and fifth aspects may be complementary, and any two or more thereof 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.

[0068] According to a sixth aspect, a gas turbine engine for an aircraft is provided, the engine comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core; and a reducer which receives an input from the core shaft and outputs a drive to the blower so as to drive the blower at a lower rotational speed than that of the core shaft; an oil loop system for supplying oil to the reducer; and a heat exchange system comprising: an air-oil heat exchanger through which the oil in the oil loop system flows; a fuel-oil heat exchanger through which the oil in the oil loop system and the fuel flow such that heat is transferred between the oil and the fuel, and wherein the oil loop system branches such that a proportion 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 loop system; and a modulating valve for allowing the proportion of the oil sent through each branch to be varied, and wherein the heat exchange system is intended to be controlled such that, under cruising conditions, a heat transfer ratio of: heat transfer rate, oil to air (kJkg1) heat transfer rate from oil to fuel (kJkg'J) is in the range from 0 to 0.67.

[0069] 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 comprise 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 configured to rotate at a higher rotational speed than the first core shaft.

[0070] The heat exchange system may further include branched fuel return paths and at least one valve controlling a fuel flow split. The branched paths may be configured to return fuel from the heat exchange system to at least two different locations along a main fuel path from where the fuel enters the gas turbine engine to the combustion chamber.

[0071] The motor of the sixth aspect may be for carrying out the method of the first, third, and / or fifth aspect, and may have any of the characteristics described in relation to any preceding aspect.

[0072] According to a seventh aspect, a method of operating a gas turbine engine for an aircraft is provided, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, a combustion chamber for burning fuel, and a core shaft connecting the turbine to the compressor; a blower located upstream of the engine core; a reducer which receives an input from the core shaft and outputs a drive to the blower so as to drive the blower at a lower rotational speed than that of the core shaft; an oil loop system for supplying oil to the reducer; and a heat exchange system comprising: an air-oil heat exchanger through which the oil in the oil loop system flows; and a fuel-oil heat exchanger through which the oil in the oil loop system and the fuel flow such that heat is transferred between the oil and the fuel, and wherein the oil loop system branches such that a proportion 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 loop system; and a modulating valve for allowing the proportion of the oil sent through each branch to be varied, the method comprising controlling the heat exchange system such that, under idle conditions, a heat transfer ratio of: oil to air heat transfer rate (kJ kg') oil to fuel heat transfer rate (kJkg') is in the range of 0.67 to 5.67.

[0073] The heat transfer rates are defined as mentioned above with respect to the fifth aspect. Idling operation when the aircraft is in ground operation may be referred to as "ground idle", and idling operation when the aircraft is in flight may be referred to as "flight idle". All of the options described below for this aspect may be considered relevant for ground idle conditions. Flight idle is generally at slightly higher thrust than ground idle. In some implementations, only the least restrictive ranges may apply to flight idle for a given engine. In other implementations, all of the options described below for this aspect may also apply to flight idle conditions.

[0074] As mentioned for the fifth aspect, the inventors were aware that the use of fuels different from traditional kerosene-based jet fuels, such as sustainable aviation fuels, may lead to different fuel properties, and that operating parameters may be adjusted to utilize the different fuel properties. In particular, certain fuels may be heated to higher temperatures in one or more fuel-oil heat exchangers than traditional fuels, without significantly increasing coking. This may allow a process providing a Improved oil cooling (as the fuel is able to absorb more heat) and can also improve the overall thermal efficiency of the engine, with less heat lost to the environment. The controllable heat exchange system has a key role to play in managing the heat transfer ratio.Furthermore, the inventors were aware that, although cruise conditions typically represent a much larger proportion of an aircraft engine's operating time, idling 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 increase - the use of non-traditional fuels may therefore have an even greater effect on optimal approaches to heat management under idling conditions - for example, during aircraft start-up, operation from idle during boarding and taxiing (to a runway or hangar, or between other ground locations), as well as at certain times during flight (for example, at the beginning of a descent).Since the operating conditions are very different between cruise and idle - both in terms of altitude and in terms of the desired engine thrust output - the heat exchange system must be controlled differently.

[0075] 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 include controlling the heat exchange system such that under idle conditions the heat transfer ratio is greater than 0.75, and optionally greater than 1.0 or 1.5.

[0076] The step of controlling the heat exchange system to adjust the heat transfer ratio may be or include controlling the modulating valve arranged to vary the proportion of oil sent through each branch of the oil loop system.

[0077] The step of controlling the heat exchange system to adjust the heat transfer ratio may include decreasing the amount of oil sent through at least one air-oil heat exchanger when the heat transfer ratio is too high.

[0078] The heat exchange system may include at least one bypass line for allowing oil or fuel to bypass one or more heat exchangers, and the step of controlling the heat exchange system to adjust the heat transfer ratio may include modulating the amount of oil sent through the bypass line.

[0079] The heat exchange system may include at least one recirculation line for allowing a fluid (e.g., oil or fuel) to pass multiple times through one or more heat exchangers. The step of controlling the heat exchange system to adjust the heat transfer ratio may include modulating the amount of fluid sent through the recirculation line.

[0080] The heat exchange system may include a refrigeration cycle apparatus, and the method may further include using the refrigeration cycle apparatus to provide a thermal lift by transferring more heat from the oil to the fuel, possibly 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.

[0081] In other implementations, the heat exchange system is not intended to provide thermal elevation, such that the fuel temperature cannot exceed a maximum oil temperature. In such implementations, the heat transfer ratio may be in the range of 2.00 to 5.67, and possibly 3.37 to 5.67.

[0082] 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 at the inlet to the combustion chamber is below 200°C.

[0083] 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 the combustion chamber inlet is above 200°C; (ii) the heat transfer ratio is in the range of 0.67 to 2.67 when the fuel temperature at the combustion chamber inlet is above 250°C; and / or (iii) the heat transfer ratio is in the range of 0.67 to 1.22 when the fuel is above 280°C.

[0084] The method may include controlling the heat exchange system under idle conditions such that the heat transfer ratio is in the range of 0.67 to 3.67, provided that the fuel is at least 70% sustainable aviation fuel.

[0085] The method may include controlling the heat exchange system under idle conditions such that the heat transfer ratio is within the range from 0.67 to 2.67, provided that the fuel is at least 80% sustainable aviation fuel.

[0086] The methods of the first, third, fifth, and seventh aspects may be complementary, and any two or more thereof 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.

[0087] 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 located upstream of the engine core; and a reducer which receives an input from the core shaft and outputs a drive to the blower so as to drive the blower at a lower rotational speed than that of the core shaft; an oil loop system for supplying oil to the reducer; and a heat exchange system comprising: an air-oil heat exchanger through which the oil in the oil loop system flows; a fuel-oil heat exchanger through which the oil in the oil loop system and the fuel flow such that heat is transferred between the oil and the fuel, and wherein the oil loop system branches such that a proportion 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 loop system; and a modulating valve for allowing the proportion of the oil sent through each branch to be varied, and wherein the heat exchange system is intended to be controlled such that, under idle conditions, a heat transfer ratio of: oil-to-air heat transfer rate Ik.lkjLj heat transfer rate from oil to fuel (klkg'1) is in the range from 0.67 to 5.67.

[0088] 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 comprise 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 configured to rotate at a higher rotational speed than the first core shaft.

[0089] The heat exchange system may further include branched fuel return paths and at least one valve controlling a fuel flow division. The branched paths may be configured to return fuel from the heat exchange system to at least two different locations along a main fuel path, from where the fuel enters the gas turbine engine to the combustion chamber.

[0090] The motor of the eighth aspect may be for carrying out the method of the first, third, fifth and / or seventh aspect, and may have any of the characteristics described in relation to any preceding aspect.

[0091] According to a ninth aspect, a method of operating a gas turbine engine of an aircraft is provided, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, a combustion chamber for burning fuel, and a core shaft connecting the turbine to the compressor; a blower located upstream of the engine core; a reducer which receives an input from the core shaft and outputs a drive to the blower so as to drive the blower at a lower rotational speed than that of the core shaft; an oil loop system for supplying oil to the reducer; and a heat exchange system comprising: an air-oil heat exchanger through which the oil in the oil loop system flows; a fuel-oil heat exchanger through which oil in the oil loop system and fuel flow such that heat is transferred between the oil and the fuel, and wherein the oil loop system comprises at least one bypass line, the bypass line being for allowing oil to bypass one of the air-oil heat exchanger and the fuel-oil heat exchanger; and a bypass valve for allowing the proportion of oil sent through the bypass line to be varied.

[0092] The method comprises controlling the bypass valve such that, under cruising conditions, a heat transfer ratio of: oil to air heat transfer rate (kJkgk heat transfer rate from oil to fuel (kJkg') is in the range from 0 to 0.67.

[0093] The inventors were aware that the principles described in relation to the fifth to eighth aspects may also be applied to engines which do not have branched oil loop paths with different heat exchangers on different branches; for example being applicable to engines with series, substantially linear arrangements of heat exchangers and in which the use of one or more bypass lines is made as an alternative to a branched main path, rather than as an optional addition. to it. Introducing and / or improving the control of existing bypass lines can enable a process that provides improved oil cooling (as the fuel used may be able to absorb more heat than traditional fuels) and can also improve the overall thermal efficiency of the engine, with less heat lost to the environment. The controllable bypass valve has a key role to play in managing the heat transfer ratio.

[0094] The heat transfer rates used to calculate the heat transfer ratio are defined as mentioned above with respect to the fifth and seventh aspects.

[0095] The method may include controlling the bypass valve such that, under cruising conditions, 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.

[0096] The step of controlling the bypass valve to adjust the heat transfer ratio may include decreasing the amount of oil sent through the at least one air-oil heat exchanger when the heat transfer ratio is too high.

[0097] In some implementations, the oil loop system may include a single bypass line, the single bypass line being for allowing oil to bypass the air-to-oil heat exchanger. No bypass line may be provided for the fuel-to-oil heat exchanger. It will be kept in mind that in most operating scenarios and to improve the thermal efficiency of the engine, there is a desire to ensure as much heat transfer as is safe / practicable from the oil to the fuel, and thus to be able to avoid bypassing the fuel-to-oil heat exchanger.

[0098] The heat exchange system may include at least two oil bypass lines, each bypass line being adapted to allow oil to bypass one of the air-to-oil heat exchanger and the fuel-to-oil heat exchanger. The method may include modulating the amount of oil sent through each bypass line. The same bypass valve—which may be a three-way valve—may be used to regulate the flow through both bypass lines, or a different bypass valve may be provided for each.

[0099] The heat exchange system may include at least one recirculation line for allowing a fluid to pass multiple times through a heat exchanger. The method of such implementations may further include modulating the amount of fluid sent through the recirculation pipe so as to adjust the heat transfer ratio. Additionally or alternatively, the heat exchange system may include a refrigeration cycle apparatus, and the method of such implementations may further include using the refrigeration cycle apparatus to provide thermal elevation by transferring more heat from the oil to the fuel, possibly such that the fuel temperature is raised above the oil temperature. It will therefore be appreciated that one or more other components in addition to the bypass valve may be used to adjust the heat transfer ratio.

[0100] In implementations with a refrigeration cycle apparatus, 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 apparatus, in which the heat exchange system is not intended 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.

[0101] The method may comprise controlling the bypass valve under cruise conditions such that: (i) the heat transfer ratio is in the range 0 to 0.2, provided that the fuel temperature at the combustion chamber inlet is at least 160°C; and / or (ii) the heat transfer ratio is in the range from 0 to 0.1, provided that the fuel temperature at the combustion chamber inlet is at least 180°C.

[0102] In some implementations, one or more fuel characteristics may be taken into account when determining how to control the bypass valve. For example, the method may include controlling the bypass valve under cruise conditions such that: (i) the heat transfer ratio is in the range 0 to 0.2, provided that the fuel is at least 70% sustainable aviation fuel; (ii) the heat transfer ratio is in the range of 0 to 0.1, provided that the fuel is at least 80% sustainable aviation fuel;

[0103] The method may include controlling the bypass valve under cruise conditions such that the heat transfer rate from the oil to the air can be maintained in the range of 0 to 240 kJ per kilogram of fuel at cruise conditions, with no more than 20% of the heat transferred from the oil at cruise being transferred to the air, and / or such that the heat transfer rate from the oil to the fuel is maintained in the range of 85 to 350 kJ per kilogram of fuel at cruise conditions, with at least 80% of the heat transferred from the oil at cruise being transferred to the fuel.

[0104] The methods of the first, third, fifth, seventh, and ninth aspects may be complementary, and any two or more thereof 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.

[0105] According to a tenth aspect, a gas turbine engine for an aircraft is provided, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a blower located upstream of the engine core; a reducer which receives an input from the core shaft and outputs a drive to the blower so as to drive the blower at a lower rotational speed than that of the core shaft; an oil loop system for supplying oil to the reducer; and a heat exchange system comprising: an air-oil heat exchanger through which the oil in the oil loop system flows; a fuel-oil heat exchanger through which the oil in the oil loop system and the fuel flow such that heat is transferred between the oil and the fuel, and wherein the oil loop system comprises at least one bypass line, the bypass line being adapted to allow oil to bypass one of the air-to-oil heat exchanger and the fuel-to-oil heat exchanger, and a bypass valve adapted to allow the proportion of oil sent through the bypass line to be varied.

[0106] The bypass valve is intended to be controlled such that, under cruising conditions, a heat transfer ratio of: oil to air heat transfer rate Ik.lkjLj oil to fuel heat transfer rate (klkg'1) is in the range from 0 to 0.67.

[0107] 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 comprise 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 configured to rotate at a higher rotational speed than the first core shaft.

[0108] The heat exchange system may further comprise branched fuel return paths and at least one valve controlling a fuel flow division, the branched paths being for returning fuel fuel from the heat exchange system at at least two different locations along a main fuel path, from the fuel inlet to the gas turbine engine to the combustion chamber.

[0109] The motor of the tenth aspect may be for carrying out the method of the first, third, fifth, seventh and / or ninth aspect, and may have any of the characteristics described in relation to any preceding aspect.

[0110] According to an eleventh aspect, a method of operating a gas turbine engine for an aircraft is provided, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, a combustion chamber for burning fuel, and a core shaft connecting the turbine to the compressor; a blower located upstream of the engine core; a reducer which receives an input from the core shaft and outputs a drive to the blower so as to drive the blower at a lower rotational speed than that of the core shaft; an oil loop system for supplying oil to the reducer; and a heat exchange system comprising: an air-oil heat exchanger through which the oil in the oil loop system flows; and a fuel-oil heat exchanger through which oil in the oil loop system and fuel flow such that heat is transferred between the oil and the fuel, and wherein the oil loop system comprises at least one bypass line, the bypass line being for allowing oil to bypass one of the air-oil heat exchanger and the fuel-oil heat exchanger; and a bypass valve for allowing the proportion of oil sent through the bypass line to be varied.

[0111] The method comprises controlling the bypass valve such that, under idle conditions, a heat transfer ratio of: oil to air heat transfer rate (kJkgk oil to fuel heat transfer rate (kJkg') is in the range of 0.67 to 5.67.

[0112] The inventors were aware that the principles described in relation to the fifth to eighth aspects may also be applied to engines which do not have branched oil loop paths with different heat exchangers on different branches; for example being applicable to engines with series, substantially linear arrangements of heat exchangers and in which the use of one or more bypass lines is made as an alternative to a branched main path, rather than as an optional addition. to it. Introducing and / or improving the control of existing bypass lines can enable a process that provides oil cooling and can also improve the overall thermal efficiency of the engine, with less heat lost to the environment. The controllable bypass valve has a key role to play in managing the heat transfer ratio.Furthermore, the inventors were aware that, although cruise conditions typically represent a much larger proportion of an aircraft engine's operating time, idling 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 increase - the use of non-traditional fuels may therefore have an even greater effect on optimal approaches to heat management under idling conditions - for example, during aircraft start-up, operation from standstill during boarding and taxiing (to a runway or hangar, or between other ground locations), or during descent in flight.Since the operating conditions are very different between cruise and idle - both in terms of altitude and in terms of the desired engine thrust output - the bypass valve must be controlled differently.

[0113] The heat transfer rates are defined as mentioned above with respect to the fifth aspect. Idling operation when the aircraft is in ground operation may be referred to as "ground idle" and idling operation in flight may be referred to as "in-flight idle". All of the options described below for this aspect may be considered relevant for ground idle conditions. In-flight idle is typically at slightly higher thrust and in some implementations only the least restrictive ranges may apply to in-flight idle for a given engine.

[0114] The method may include controlling the bypass valve such that, under idle conditions (and in some implementations, more 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 heat exchange system such that, under idle conditions, the heat transfer ratio is greater than 1.0, and optionally greater than 1.5, 2.0, 2.5, or 3.0. The method may include controlling the bypass valve such that, under idle conditions, the heat transfer ratio is greater than 0.75, and optionally greater than 1.0 or 1.5.

[0115] Controlling the bypass valve to adjust the heat transfer ratio may include decreasing the amount of oil sent through at least one air-oil heat exchanger when the heat transfer ratio is too high.

[0116] The heat exchange system may include at least two oil bypass lines, each bypass line being configured to allow oil to bypass a respective heat exchanger of the air-to-oil heat exchanger and the fuel-to-oil heat exchanger. The method may include modulating the amount of oil sent through each bypass line.

[0117] The heat exchange system may include at least one recirculation line for allowing a fluid to pass multiple times through a heat exchanger. The method may further include modulating the amount of fluid sent through the recirculation pipe to adjust the heat transfer ratio.

[0118] The heat exchange system may include a refrigeration cycle apparatus. The method may include using the refrigeration cycle apparatus to provide thermal lift by transferring more heat from the oil to the fuel (beyond what would be transferred in the heat exchanger), possibly 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 in which the heat exchange system is not intended 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 possibly 3.37 to 5.67.

[0119] 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 at the combustion chamber inlet is below 200°C; (ii) is in the range of 0.67 to 5.00, and optionally 0.67 to 4.00, when the fuel temperature at the inlet of the combustion chamber (16) is above 200°C; (iii) is in the range of 0.67 to 4.00, and optionally 0.67 to 2.67, when the fuel temperature at the inlet of the combustion chamber (16) is above 250°C; and (iv) in the range from 0.67 to 2.33, and optionally from 0.67 to 1.22, when the fuel temperature at the inlet of the combustion chamber (16) is above 280°C.

[0120] The method may take into account the S AF 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, provided that the fuel is at least 70% sustainable aviation fuel, and / or in the range of 0.67 to 2.67, provided that the fuel is at least 80% sustainable aviation fuel.

[0121] The methods of the first, third, fifth, seventh, ninth, and eleventh aspects may be complementary, and any two or more thereof may be performed together in various implementations. The method of the eleventh aspect may be performed using the motor of the second, fourth, sixth, eighth, or tenth aspect.

[0122] According to a twelfth aspect, a gas turbine engine for an aircraft is provided, 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 located upstream of the engine core; and a reducer which receives an input from the core shaft and outputs a drive to the blower so as to drive the blower at a lower rotational speed than that of the core shaft; an oil loop system for supplying oil to the reducer; and a heat exchange system comprising: an air-oil heat exchanger through which the oil in the oil loop system flows; and a fuel-oil heat exchanger through which the oil in the oil loop system and the fuel flow such that heat is transferred between the oil and the fuel, and wherein the oil loop system comprises at least one bypass line, the bypass line being adapted to allow oil to bypass one of the air-to-oil heat exchanger and the fuel-to-oil heat exchanger; and a bypass valve adapted to allow the proportion of oil sent through the bypass line to be varied, and wherein the bypass valve is intended to be controlled such that, under idle conditions, a heat transfer ratio of: heat transfer rate from oil to air (kJkg'!) heat transfer rate from oil to fuel (kJkg'!) is in the range from 0.67 to 5.67.

[0123] 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 comprise 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 intended to rotate at a higher rotational speed than that of the first core shaft.

[0124] The heat exchange system may further include branched fuel return paths and at least one valve controlling a fuel flow split, the branched paths being for returning fuel from the heat exchange system to at least two different locations along a main fuel path from the fuel inlet in the gas turbine engine to the combustion chamber.

[0125] The motor of the twelfth aspect may be for carrying out the method of the first, third, fifth, seventh, ninth, and / or eleventh aspect, and may have any of the characteristics described in relation to any preceding aspect.

[0126] It will be noted that a feature described in relation to one aspect may be used in combination with any aspect, mutatis mutandis.

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

[0128] A gas turbine engine in accordance with any aspect of the present disclosure may include an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor. Such a gas turbine engine may include a fan (having fan blades). Such a fan may be located upstream of the engine core. Alternatively, in some examples, the gas turbine engine may include a fan located downstream of the engine core, for example in the case where the gas turbine engine is an open rotor or turboprop engine (in which case the fan may be referred to as a propeller).

[0129] Where the gas turbine engine is an open rotor or turboprop engine, the gas turbine engine may include two counter-rotating propeller stages attached to and driven by a free power turbine via a shaft. The propellers may rotate in opposite directions such that one rotates clockwise and the other counter-clockwise about the engine's axis of rotation. Alternatively, the gas turbine engine may include a propeller stage and a guide vane stage designed downstream of the propeller stage. The guide vane stage may be of variable pitch. Thus, the high-pressure, intermediate-pressure, and free-power turbines may respectively drive high-pressure and intermediate-pressure propellers and compressors through suitable interconnecting shafts. Thus, the propellers may provide the majority of the propulsion thrust.

[0130] In the case where the gas turbine engine is an open rotor or turboprop engine, one or more of the propeller stages may be driven by a reduction gear. The reduction gear may be of the type described herein.

[0131] An engine according to the present disclosure may be a bypass engine. Such an engine may be a direct-drive bypass engine in which the fan is directly connected to the fan drive turbine, for example without a reduction gear, via a core shaft. In such a direct-drive bypass engine, the fan may be said to rotate at the same rotational speed as the fan drive turbine. Strictly by way of example, the fan drive turbine may be a first turbine, the core shaft may be a first core shaft, and the gas turbine engine may further comprise a second turbine and a second core shaft connecting the second turbine to the compressor. The second turbine, the compressor, and the second core shaft may be intended to rotate at a higher rotational speed than the first core shaft.In such an arrangement, the second turbine may be positioned axially upstream of the first turbine.

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

[0133] The gas turbine engine as described and / or claimed herein may have any suitable general architecture. For example, the gas turbine engine may have any desired number of shafts that connect turbines and compressors, e.g., one, two, or three shafts. By way of example only, the turbine connected to the core shaft may be a first turbine, the compressor connected to the core shaft may be a first compressor, and the core shaft may be a first core shaft. The engine core may further comprise 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 configured to rotate at a higher rotational speed than the first core shaft.

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

[0135] The reducer may be intended to be driven by the core shaft that is designed to rotate (e.g., in use) at the lowest rotational speed (e.g., the first core shaft in the example above). For example, the reducer may be intended to be driven only by the core shaft that is designed to rotate (e.g., in use) at the lowest rotational speed (e.g., only by the first core shaft, and not the second core shaft, in the example above). Alternatively, the reducer may be intended to be driven by any one or more shafts, e.g., the first and / or second shafts in the example above.

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

[0137] The reducer may have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example greater than 2.5, for example in the range of 3 to 4.2, or 3.2 to 3.8, for example in the range of or at least 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. The gear ratio may be, for example, between any two of the values ​​in the preceding sentence. Strictly by way of example, the reducer may be a "star" reducer having a reduction ratio in the range of 3.1 or 3.2 to 3.8. Strictly by way of further example, the reducer may be a "star" reducer having a reduction ratio in the range of 3.0 to 3.1. Strictly by way of example, the reducer may be a "planetary" reducer having a reduction ratio in the range of 3.6 to 4.2. In some arrangements, the gear ratio may be outside these ranges.

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

[0139] The or each compressor (e.g., the first compressor and the second compressor as described above) may comprise any number of stages, e.g., multiple stages. Each stage may comprise a row of rotor blades and a row of stator blades, which may be variable stator blades (in that their angle of incidence may be variable). The row of rotor blades and the row of stator blades may be axially offset from each other. For example, the gas turbine engine may be a direct-drive, bypass gas turbine engine comprising 13 or 14 compressor stages (in addition to the fan). Such an engine may, for example, comprise 3 stages in the first (or "low pressure") compressor and either 10 or 11 stages in the second (or "high pressure") compressor.As a further example, the gas turbine engine may be a "geared" gas turbine engine (in which the fan is driven by a first core shaft through a reduction gearbox) having 11, 12, or 13 compressor stages (in addition to the fan). Such an engine may have 3 or 4 stages in the first (or "low pressure") compressor and 8 or 9 stages in the second (or "high pressure") compressor. As a further example, the gas turbine engine may be a "geared" gas turbine engine having 4 stages in the first (or "low pressure") compressor and 10 stages in the second (or "high pressure") compressor.

[0140] The or each turbine (e.g., the first turbine and the second turbine as described above) may comprise any number of stages, e.g., multiple stages. Each stage may comprise a row of rotor blades and a row of stator blades, or vice versa, as required. The respective rows of rotor blades and stator blades may be axially offset from each other. The second (or "high pressure") turbine may comprise 2 stages in any arrangement (e.g., regardless of whether it is a geared or direct-drive engine). The gas turbine engine may be a direct-drive gas turbine engine comprising a first (or "low pressure") turbine having 5, 6 or 7 stages. Alternatively, the gas turbine engine may be a "geared" gas turbine engine comprising a first (or "low pressure") turbine having 3 or 4 stages.

[0141] Each fan blade may be defined as having a radial throw extending from a root (or hub) at a radially inner gas-washed location, or a 0% throw position, to a tip at a 100% throw position. The ratio of the fan blade radius at the hub to the fan blade radius at the tip may be less than (or in the order of) any of: 0.4, 0.39, 0.38, 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. The ratio of the fan blade radius at the hub to the fan blade radius at the tip may be included in an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, in the range 0.28 to 0.32, or 0.29 to 0.30.These ratios can be commonly referred to as the hub-to-tip ratio. Both the radius at the hub and the radius at the tip can be measured at the leading edge (or axially forward-most) portion of the blade. The hub-to-tip ratio refers, of course, to the gas-washed portion of the fan blade, i.e., the portion radially outboard of any platform.

[0142] The fan radius can be measured between the engine centerline and the tip of a fan blade at its leading edge. The blower diameter (which may simply be twice the blower radius) may be greater than (or of the order of) any of: 140 cm, 170 cm, 180 cm, 190 cm, 200 cm, 210 cm, 220 cm, 230 cm, 240 cm, 250 cm (about 100 inches), 260 cm, 270 cm (about 105 inches), 280 cm (about 110 inches), 290 cm (about 115 inches), 300 cm (about 120 inches), 310 cm, 320 cm (about 125 inches), 330 cm (about 130 inches), 340 cm (about 135 inches), 350 cm, 360 cm (about 140 inches), 370 cm (about 145 inches), 380 cm (approximately 150 inches), 390 cm (approximately 155 inches), 400 cm, 410 cm (approximately 160 inches) or 420 cm (approximately 165 inches).The blower diameter may be in an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, in the range of 210 cm to 240 cm, or 250 cm to 280 cm, or 320 cm to 380 cm. Strictly by way of non-limiting example, the blower diameter may be in the range of 170 cm to 180 cm, 190 cm to 200 cm, 200 cm to 210 cm, 210 cm to 230 cm, 290 cm to 300 cm, or 340 cm to 360 cm.

[0143] The rotation speed of the blower may vary during use. Generally, the rotational speed is lower for fans with a larger diameter. Strictly by way of non-limiting example, the rotational speed of the fan at cruise conditions may be less than 3,500 rpm, for example less than 2,600 rpm, or less than 2,500 rpm, or less than 2,300 rpm. Strictly by way of further non-limiting example, the rotational speed of the fan at cruise conditions for a "geared" gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 2,750 to 2,900 rpm. Strictly by way of further non-limiting example, the fan rotational speed at cruise conditions for a "geared" gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 2,500 to 2,800 rpm.Strictly by way of further non-limiting example, the fan rotational speed at cruise conditions for a "geared" gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 1,500 to 1,800 rpm. Strictly by way of further non-limiting example, the fan rotational speed at cruise conditions for a direct drive engine having a fan diameter in the range of 190 cm to 200 cm may be in the range of 3,600 to 3,900 rpm. Strictly by way of further non-limiting example, the fan rotational speed at cruising conditions for a direct drive engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 2,000 to 2,800 rpm.

[0144] During operation of the gas turbine engine, the fan (with associated fan blades) rotates about an axis of rotation. This rotation results in a displacement of the fan blade tip with a velocity Utip . The work done by the fan blades on the flow results in an enthalpy increase dH of the flow. A fan tip load can be defined as dH / UpOint2, where dH is the enthalpy increase (e.g., the 1-D average enthalpy increase) across the fan and Utip is the (translational) velocity of the fan tip, e.g., at the leading edge of the tip (which can be defined as the fan tip radius at the leading edge multiplied by the angular velocity).The peak fan load at cruise conditions may be greater than (or of the order of) any one 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 peak fan load may be within an inclusive range bounded by any two of the values ​​in the . previous sentence (i.e. the values ​​may form upper or lower limits), for example in the range 0.28 to 0.31 or 0.29 to 0.3 (e.g. for a geared gas turbine engine).

[0145] Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio (BPR), where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core. In some arrangements the bypass ratio at cruise conditions may be greater than (or of the order of) any of the following: 9. 9.5, 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5, or 20. The bypass ratio at cruise conditions may be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, in the range 12 to 16, 13 to 15, or 13 to 14.Strictly by way of non-limiting example, the bypass ratio at cruise conditions of a direct drive gas turbine engine according to the present disclosure may be in the range of 9:1 to 11:1. Strictly by way of further non-limiting example, the bypass ratio at cruise conditions of a geared gas turbine engine according to the present disclosure may be in the range of 12:1 to 15:1. The bypass duct may be substantially annular. The bypass duct may be radially outboard of the core engine. The radially outer surface of the bypass duct may be defined by a nacelle and / or a fan case.

[0146] The overall pressure ratio (OPR) of a gas turbine engine as described and / or claimed herein may be defined as the ratio of the stagnation pressure at the outlet of the highest pressure compressor (before an inlet into the combustor) to the stagnation pressure upstream of the fan. By way of non-limiting example, the overall pressure ratio of a gas turbine engine as described and / or claimed herein at cruise conditions may be greater than (or in the order of) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio may be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, in the range of 50 to 70.Strictly by way of non-limiting example, the overall pressure ratio at cruise conditions of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 40 to 45. Strictly by way of non-limiting example, the overall pressure ratio at cruise conditions of a geared gas turbine engine having a fan diameter . in the range of 210 cm to 230 cm may be in the range of 45 to 55. Strictly by way of non-limiting example, the overall pressure ratio at cruise conditions of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 50 to 60. Strictly by way of non-limiting example, the overall pressure ratio at cruise conditions of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 50 to 60.

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

[0148] A gas turbine engine as described and / or claimed herein may have any desired maximum thrust. Strictly by way of non-limiting example, a gas turbine as described and / or claimed herein may be capable of producing a maximum thrust of at least (or in the order of) any one of the following: 100 kN, HOkN, 120 kN, 130 kN, 135 kN, 140 kN, 145 kN, 150 kN, 155 kN, 160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN, 450 kN, 500 kN, or 550 kN. The maximum thrust may be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits). By way of non-limiting example only, a gas turbine as described and / or claimed herein may be capable of producing a maximum thrust in the range of 155 kN to 170 kN, 330 kN to 420 kN, or 350 kN to 400 kN. Strictly by way of non-limiting example, the maximum thrust of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 140 kN to 160 kN. Strictly by way of non-limiting example, the maximum thrust of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 150 kN to 200 kN.Strictly by way of non-limiting example, the maximum thrust of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 370 kN to 500 kN. Strictly by way of non-limiting example, the maximum thrust of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 370 kN to 500 kN. The thrust mentioned above may be the maximum net thrust under typical atmospheric conditions at sea level plus 15 degrees C (ambient pressure of 101.3 kPa, temperature of 30 degrees C), with the engine static.

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

[0150] The TET at cruise conditions may be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, 1530 K to 1600 K. The maximum TET in engine operation may be, for example, at least (or in the order of) any one of the following: 1700 K, 1750 K, 1800 K, 1850 K, 1900 K, 1950 K, 2000 K, 2050 K, or 2100 K. Thus, by way of non-limiting example only, the maximum TET of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 1890 K to 1 960 K. Strictly by way of non-limiting example, the maximum TET of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 1890 K to 1960 K.Strictly by way of non-limiting example, the maximum TET of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 1890 K to 1960 K. Strictly by way of non-limiting example, the maximum TET of a direct-drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 1935 K to 1995 K. Strictly by way of non-limiting example, the maximum TET of a direct-drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 1890 K to 1950 K. The maximum TET may be in an inclusive range bounded by any two of the values ​​in the preceding sentence. (i.e., the values ​​may form upper or lower limits), for example in the range from 1800 K to 1950 K, or from 1900 K to 2000 K.Maximum TET may occur, for example, under a high thrust condition, such as a maximum takeoff thrust (MTTP) condition.

[0151] A fan blade and / or airfoil portion of a fan blade described and / or claimed herein may be fabricated from any suitable material or combination of materials. For example, at least a portion of the fan blade and / or airfoil may be fabricated at least in part from a composite, for example, a metal matrix composite and / or an organic matrix composite, such as a carbon fiber composite. As a further example, at least a portion of the fan blade and / or airfoil aerodynamic may be made at least in part from a metal, such as a titanium-based metal or an aluminum-based material (such as an aluminum-lithium alloy) or a steel-based material. The fan blade may include at least two regions made using different materials. For example, the fan blade may have a protective leading edge, which may be made using a material that is more able to withstand impact (e.g., by birds, ice, or other material) than the rest of the blade. Such a leading edge may, for example, be made using titanium or a titanium-based alloy. Thus, strictly by way of example, the fan blade may have a carbon fiber or aluminum-based (such as an aluminum-lithium alloy) body with a titanium leading edge.

[0152] A fan as described and / or claimed herein may comprise a central portion, from which the fan blades may extend, for example in a radial direction. The fan blades may be connected to the central portion in any desired manner. For example, each fan blade may comprise a fastening element that may engage a corresponding notch in the hub (or disc). Strictly by way of example, such a fastening element may be in the form of a dovetail that may notch into and / or engage a corresponding notch in the hub / disc to secure the fan blade to the hub / disc. As a further example, the fan blades may be formed integrally with a central portion. Such an arrangement may be referred to as a bladed disc or a bladed ring.Any suitable method may be used to manufacture such a bladed disc or bladed ring. For example, at least a portion of the fan blades may be machined from a block and / or at least a portion of the fan blades may be joined to the hub / disc by welding, such as a linear friction stir weld.

[0153] 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 the bypass duct exit area to be varied during use. The general principles of the present disclosure may be applied to engines with or without a VAN.

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

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

[0156] As such, ground idling may refer to a phase of engine operation where the aircraft is stationary and in contact with the ground, but where there is a need for the engine to taxi. At idle, the engine may produce between 3% and 9% of the available engine thrust. In other non-limiting examples, the engine may produce between 5% and 8% of the available thrust. In other non-limiting examples, the engine may produce between 6% and 7% of the available thrust. Taxiing may refer to a phase of engine operation where the aircraft is propelled along the ground by the thrust produced by the engine. When taxiing, the engine may produce between 5% and 15% of the available thrust. In other non-limiting examples, the engine may produce between 6% and 12% of the available thrust. In other non-limiting examples, the engine may produce between 7% and 10% of the available thrust.Takeoff may refer to a phase of engine operation where the aircraft is propelled by thrust produced by the engine. At an initial stage in the takeoff phase, the aircraft may be propelled while the aircraft is in contact with the ground. At a later stage in the takeoff phase, the aircraft may be propelled while the aircraft is not in contact with the ground. During takeoff, the engine may produce between 90% and 100% of the available thrust. In other non-limiting examples, the engine may produce between 95% and 100% of the available thrust. In other non-limiting examples, the engine may produce 100% of the available thrust.

[0157] Climb may refer to a phase of engine operation where the aircraft is propelled by thrust produced by the engine. During climb, the engine may produce between 75% and 100% of the available thrust. In other non-limiting examples, the engine may produce between 80% and 95% of the available thrust. In other non-limiting examples, the engine may produce between 85% and 90% of the available thrust. In this regard, climb may refer to an operational phase within an aircraft flight cycle between takeoff and arrival at cruise conditions, where arrival at cruise conditions thus defines the start of the cruise phase, or a portion thereof, of the aircraft flight. Additionally or alternatively, climb may refer to a nominal point in, or one or more nominal periods during, an aircraft flight cycle between takeoff and landing, where a relative increase in altitude is required, which may require an additional demand for engine thrust.

[0158] As used herein, cruise conditions, which may define the cruise phase (or a portion thereof) of aircraft flight, have a conventional meaning and will be readily understood by those skilled in the art. In some examples, for a given gas turbine engine for an aircraft, cruise conditions may refer to an engine operating point at mid-cruise of a given mission (which may be referred to in the industry as an "economic mission") of an aircraft to which the gas turbine engine is designed to be attached. In this sense, mid-cruise may be considered the point in an aircraft flight cycle at which 50% of the total fuel that is burned between the end of climb and the beginning of descent has been burned (which may be approximated by the midpoint—in terms of time and / or distance—between the end of climb and the beginning of descent).Cruise conditions may therefore define an operating point, phase, or part thereof, of flight that provides thrust that would ensure steady-state operation (i.e., maintaining a constant altitude and / or a constant Mach number), or at least substantially steady-state operation (i.e., maintaining at least a substantially constant altitude and / or at least a substantially constant Mach number) of an aircraft to which it is designed to be attached, taking into account the number of engines provided on that aircraft. For example, where an engine is designed to be attached to an aircraft that has two engines of the same type, at cruise conditions the engine may provide half the total thrust that would be required for steady-state operation, or at least substantially steady-state operation, of that aircraft at mid-cruise.

[0159] In other words, for a given gas turbine engine for an aircraft, the cruise conditions can be defined as the operating point of the engine that provides a specified thrust (required to provide - in combination with any other engines on the aircraft - steady-state operation of the aircraft to which it is designed to be attached at a given mid-cruise Mach number) at the mid-cruise atmospheric conditions (defined by the International Standard Atmosphere according to ISO 2533 at the mid-cruise altitude). For any given gas turbine engine for an aircraft, the mid-cruise thrust, atmospheric conditions and Mach number are known, and thus the operating point of the engine at cruise conditions can be clearly defined.

[0160] Strictly by way of example, the forward speed at the cruise condition may be any point in the range of Mach 0.7 to 0.9, e.g., 0.75 to 0.85, for example. 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 in the order of Mach 0.8, in the order of Mach 0.85 or in the range of 0.8 to 0.85. Any single speed within these ranges may be part of the cruise condition. For a certain aircraft, the cruise conditions may be outside these ranges, for example below Mach 0.7 or above Mach 0.9.

[0161] Strictly by way of example, the cruising conditions may correspond to standard atmospheric conditions (according to the International Standard Atmosphere, ISA) at an altitude which is in the range of 10,000 m to 15,000 m, for example in the range of 10,000 m to 12,000 m, for example in the range of 10,400 m to 11,600 m (approximately 38,000 feet), for example in the range of 10,500 m to 11,500 m, for example in the range of 10,600 m to 11,400 m, for example in the range of 10,700 m (approximately 35,000 feet) to 11,300 m, for example in the range of 10,800 m to 11,200 m, for example in the range from 10,900 m to 11,100 m, for example of the order of 11,000 m. Cruise conditions may correspond to typical atmospheric conditions at any given altitude within these ranges.

[0162] Strictly by way of example, the cruise conditions may correspond to a forward Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35,000 feet (10,668 m). At such cruise conditions, the engine may provide a known required net thrust level. The known required net thrust level is of course dependent on the engine and its intended application and may be, for example, a value ranging from 20 kN to 40 kN.

[0163] Strictly by way of further example, the cruising conditions may correspond to a forward Mach number of 0.85 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38,000 feet (11,582 m). At such cruising conditions, the engine can provide a known required net thrust level. The known required net thrust level is of course dependent on the engine and its intended application and may be, for example, a value ranging from 35 kN to 65 kN.

[0164] In use, a gas turbine engine described and / or claimed herein may operate at cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions (e.g., mid-cruise conditions) of an aircraft to which at least one (e.g., 2 or 4) gas turbine engines may be mounted to provide propulsive thrust.

[0165] Further, one skilled in the art would readily recognize that either or both of a descent and an approach refer to a phase of operation within an aircraft flight cycle between aircraft cruise and landing, with a particular approach being part of the landing and takeoff (LTO) phase. During either or both of descent and approach, the engine may produce between 0% and 50% of available thrust. In other non-limiting examples, the engine may produce between 25% and 40% of available thrust. In other non-limiting examples, the engine may produce between 30% and 35% of available thrust. Additionally or alternatively, descent may refer to a nominal point in an aircraft flight cycle between takeoff and landing, where a relative decrease in altitude is required, and which may necessitate a reduced thrust demand from the engine.

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

[0167] According to one aspect, there is provided a method of operating a gas turbine engine as described and / or claimed herein. The operation may be at any suitable condition, which may be as defined elsewhere herein (e.g., in terms of thrust, atmospheric conditions, and Mach number).

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

[0169] Those skilled in the art will appreciate that, unless mutually exclusive, any feature or parameter described in connection with any of the above aspects may be applied to any other aspect. Furthermore, unless mutually exclusive, any feature or parameter contained or described herein may be applied to any aspect and / or in association with any other feature or parameter contained or described herein.

[0170] Except where mutually exclusive, any parameter or value contained or described herein may be applied to and / or combined with any one or more parameters and / or values ​​contained or described herein. For example, a first parameter or value contained or described herein (e.g., parameter A) may be applied to and / or combined with any one or more other parameters and / or values ​​contained or described herein (e.g., any one or more of parameter B; parameter C; and parameter D, and so on) to express a product of their relationship. For example, the The trade would understand that 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 application, combination, or function of parameter A with respect to parameter B, as appropriate.

[0171] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0172] [Fig.l] is a sectional side view of a gas turbine engine;

[0173] [Fig.2] is a close-up sectional side view of an upstream portion of a motor geared gas turbine;

[0174] [Fig.3] is a partially cutaway view of a reduction gear for a gas turbine engine;

[0175] [Fig.4] is a representation of an aircraft with a propulsion system comprising two gas turbine engines;

[0176] [Fig.5] is a representation of an exemplary fuel system;

[0177] [Fig.6] is a representation of an alternative example of a fuel system;

[0178] [Fig.7] is a representation of a part of an example of a system of oil recirculation (a primary oil loop system);

[0179] [Fig.8] is a representation of another portion of an exemplary recirculating oil system (a secondary oil loop system);

[0180] [Fig.9] is a representation of a portion of an exemplary reciprocating recirculating oil system;

[0181] [Fig.10] is a representation of a portion of the exemplary fuel system of [Fig.5] and the exemplary recirculating oil system of Figures 7 and 8;

[0182] [Fig.l 1] is a representation of another example of an oil recirculation system, illustrating all the main components of a heat exchange system;

[0183] [Fig.12] illustrates an example of a method of operating a gas turbine engine;

[0184] [Fig. 13] illustrates another example of a method of operating a gas turbine engine;

[0185] [Fig.14] illustrates another example of a method of operating a gas turbine engine;

[0186] [Fig. 15] illustrates another example of a method of operating a gas turbine engine;

[0187] [Fig. 16] illustrates another example of an oil recirculation system of a gas turbine engine;

[0188] [Fig.17] illustrates another example of a method of operating a gas turbine engine;

[0189] [Fig. 18] illustrates another example of a method of operating a gas turbine engine; and

[0190] [Fig. 19] illustrates a cruise heat transfer range graph for a geared gas turbine engine.

[0191] [Fig.l] illustrates a gas turbine engine 10 having a main rotational axis 9. The engine 10 includes an air intake 12 and a propulsion fan 23 that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 includes a core 11 that receives the core airflow A. The engine core 11 includes, in axial flow series, a low-pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, a low-pressure turbine 19, and a core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass line 22. The blower 23 is attached to, and driven by, the low pressure turbine 19 by means of a shaft 26 and an epicyclic reduction gear 30.

[0192] In use, the core airflow A is accelerated and compressed by the low pressure compressor 14 and directed into the high pressure compressor 15 where further compression takes place. The compressed air discharged from the high pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel F and the mixture is combusted. The combustion equipment 16 may be referred to as the combustor 16, with the terms "combustion equipment 16" and "combustion chamber 16" used interchangeably herein. The resulting hot combustion products then expand, and thereby drive, the high pressure and low pressure turbines 17, 19 before being discharged through the nozzle 20 to provide some propulsive thrust. The high pressure turbine 17 drives the high pressure compressor 15 by a suitable interconnecting shaft 27.The blower 23 generally acts to impart increased pressure to the bypass airflow B flowing through the bypass duct 22, such that the bypass airflow B is exhausted through the bypass exhaust nozzle 18 to generally provide the majority of the propulsion thrust. The epicyclic reduction gear 30 is a reduction gearbox.

[0193] An exemplary arrangement for a geared fan gas turbine engine 10 is shown in [Fig. 2]. The low-pressure turbine 19 (see [Fig. 1]) drives the shaft 26, which is coupled to a sun wheel, or sun gear, 28 of the epicyclic gear arrangement 30. Radially outwardly of and meshing with the sun gear 28 are a plurality of planet gears 32 which are coupled together by a planet carrier 34. The planet carrier 34 forces the planet gears 32 to change orientation about the sun gear 28 in synchronism while allowing each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled via links 36 to the fan 23 to cause it to rotate about the engine axis 9. Radially outwardly of the planet gears 32 and meshing therewith is a ring or crown gear 38 which is coupled, via links 40, to a stationary support structure 24.

[0194] It should be noted that the terms "low pressure turbine" and "low pressure compressor" as used herein may be taken to indicate the lowest pressure turbine stages and the lowest pressure compressor stages (i.e., not including the fan 23) respectively and / or the turbine and compressor stages that are connected together by the interconnecting shaft 26 with the lowest rotational speed in the engine (i.e., not including the reduction gear output shaft that drives the fan 23). In some literature, the "low pressure turbine" and "low pressure compressor" referred to herein may alternatively be known as an "intermediate pressure turbine" and "intermediate pressure compressor".When such alternative nomenclature is used, the blower 23 may be designated as the first compression stage or the lower pressure compression stage.

[0195] The epicyclic reduction gear 30 is shown by way of example in more detail in [Fig. 3]. Each of the sun gear 28, the planet gears 32 and the ring gear 38 includes teeth around its periphery for meshing with the other gears. However, for clarity, only exemplary portions of the teeth are illustrated in [Fig. 3]. There are four planet gears 32 illustrated, although it will be apparent to the skilled reader that more or fewer planet gears 32 may be provided within the scope of the claimed invention. Practical applications of an epicyclic planetary reduction gear 30 generally include at least three planet gears 32.

[0196] The epicyclic reduction gear 30 illustrated by way of example in Figures 2 and 3 is of the planetary type, in that the planet carrier 34 is coupled to an output shaft via links 36, with the ring gear 38 fixed. However, any other suitable type of epicyclic reduction gear 30 may be used. As a further example, the epicyclic reduction gear 30 may be a star arrangement, in which the planet carrier 34 is held fixed, with the ring gear 38 allowed to rotate. In such an arrangement, the fan 23 is driven by the ring gear 38. As another alternative example, the reduction gear 30 may be a differential reduction gear in which both the ring gear 38 and the planet carrier 34 are allowed to rotate.

[0197] It will be appreciated that the arrangement shown in Figures 2 and 3 is by way of example only, and that various alternatives are within the scope of this disclosure. Strictly by way of example, any suitable arrangement may be used to position the reducer 30 in the motor 10 and / or to connect the reducer 30 to the motor 10. As a further example, the connections (such as the links 36, 40 in the example of [Fig. 2]) between the reducer 30 and other parts of the motor 10 (such as the input shaft 26, the output shaft, and the fixed structure 24) may have any desired degree of rigidity or flexibility.As a further example, any suitable arrangement of bearings between rotating and stationary portions of the motor (e.g., between input and output shafts from the reducer and stationary structures, such as the reducer housing) may be used, and the description is not limited to the exemplary arrangement of [Fig. 2]. For example, where the reducer 30 has a star arrangement (described above), one skilled in the art would readily understand that the arrangement of the output and support links and bearing locations would typically be different from that shown as an example in [Fig. 2].

[0198] Thus, the present disclosure extends to a gas turbine engine having any arrangement of reduction gear styles (e.g., star or planetary), support structures, input and output shaft arrangements, and bearing locations.

[0199] Optionally, the reducer may drive additional and / or alternative components (e.g. the intermediate pressure compressor and / or a booster).

[0200] Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have a different number of compressors and / or turbines and / or a different number of interconnecting shafts. As a further example, the gas turbine engine shown in [Fig.l] has a split-flow nozzle 18, 20 which means that the flow through the bypass duct 22 has its own nozzle 18 which is independent of, and radially outward of, the core engine nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed-flow nozzle.The nozzle(s) (whether mixed or split flow) may have a fixed or variable area.

[0201] As a further example, other gas turbine engines to which the present disclosure may be applied may not have a reduction gear for the main shaft(s), being instead direct drive engines.

[0202] Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of interconnecting shafts (e.g., two) and / or an alternative number of compressors and / or turbines. Further, the engine may include a reduction gear provided in the drive train from a turbine to a compressor and / or a fan.

[0203] While the described example relates to a bypass engine, the description may apply, for example, to any type of gas turbine engine, such as an open rotor (in which the fan stage is not surrounded by a nacelle) or a turboprop, for example. In some arrangements, the gas turbine engine 10 may not include a reduction gear 30.

[0204] The geometry of the gas turbine engine 10, and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned with the rotation axis 9), a radial direction (in the bottom-up direction in [Fig.l]) and a circumferential direction (perpendicular to the page in the view of [Fig.l]). The axial, radial and circumferential directions are mutually perpendicular.

[0205] The fuel F supplied to the combustion equipment 16 may comprise a fossil-based hydrocarbon fuel, such as kerosene. Thus, the fuel F may comprise molecules from one or more of the chemical families of n-alkanes, isoalkanes, cycloalkanes, and aromatics. In addition to or alternatively, when cut with, or blended with, or replaced by an alternative fuel, the fuel F may comprise renewable hydrocarbons produced from biological or non-biological resources, otherwise known as sustainable aviation fuel (SAF). In each of the examples provided, the fuel F may comprise one or more trace elements including, for example, sulfur, nitrogen, oxygen, inorganics, and metals.

[0206] SAF is understood by those skilled in the art to refer to, for example, a biofuel, a renewable aviation fuel, a renewable jet fuel, an alternative fuel or a biofuel, produced from biological or non-biological resources. SAF is understood to be generally synthesized from carbon-containing gases extracted from the atmosphere and / or captured during industrial processes; or from a wide range of sustainable feedstocks such as, for example, used oils and greases; municipal solid waste; cellulosic waste (such as corn stover); cover crops such as camelina, carinata and stinkweed; non-biogenic alternative fuels; jatropha; halophytes and algae, rather than fossil-based hydrocarbons. SAF is not understood to encompass fossil fuels.

[0207] The functional performance of a given fuel composition, or fuel mixture F for use in a given mission, may be defined, at least in part, by the fuel's ability to process the Brayton cycle of the gas turbine engine 10. Parameters defining functional performance may include, for example, specific energy; energy density; thermal stability; and emissions including gaseous and / or particulate matter emissions. In this regard, particulate matter emissions may include soot particles created by the combustion of said fuel F, also known as non-volatile particulate matter (nvPM). Any reference made herein to soot or smoke may also apply to other types of particulate matter emissions known in the art.The gaseous emissions may include any one or more of nitrogen oxides (NOx); carbon monoxide (CO); carbon dioxide (CO2); unburned hydrocarbons (UHC); sulfur oxides (SO), including, for example, sulfur dioxide (SO2) and / or sulfur trioxide (SO3); and volatile organic compounds (VOCs) created by the combustion of said fuel F. Any reference made herein to gaseous emissions may also apply to other types of gaseous emissions known in the art.

[0208] A relatively higher specific energy (i.e., energy per unit mass), expressed in MJ / kg, may at least partially reduce takeoff weight, potentially providing a relative improvement in fuel efficiency. A relatively higher energy density (i.e., energy per unit volume), expressed in MJ / L, may at least partially reduce fuel takeoff volume, which may be particularly important for volume-limited missions or military operations involving refueling. A relatively higher thermal stability (i.e., inhibiting fuel degradation or coking under thermal stress) may allow the fuel to experience elevated temperatures in the engine and fuel injectors, potentially providing relative improvements in combustion efficiency.Reduced emissions, including particulate matter, can allow for reduced contrail formation, while reducing the environmental impact of a given mission. Other fuel properties may also be critical to functional performance. For example, a relatively lower freezing point (°C) may allow long-range missions to optimize flight profiles; minimum aromatic concentrations (%) may ensure sufficient swelling of certain materials used in the construction of O-rings and seals previously exposed to high-temperature fuels. aromatic content; and, a maximum surface tension (mN / m) can ensure sufficient spray breakage and atomization of the fuel.

[0209] The ratio of the number of hydrogen atoms to the number of carbon atoms in a molecule can influence the specific energy of a given composition, or fuel mixture. Fuels having higher ratios of hydrogen atoms to carbon atoms may have higher specific energies in the absence of bonding strain. For example, fossil-based hydrocarbon fuels may comprise molecules having about 7 to 18 carbons, with a significant portion of a given composition derived from molecules of 9 to 15 carbons, with an average of 12 carbons.

[0210] A number of sustainable aviation fuel blends have been approved for use. For example, some approved blends include blend ratios of up to 10% sustainable aviation fuel, while other approved blends include blend ratios of between 10% and 50% sustainable aviation fuel (the remainder comprising one or more fossil-based hydrocarbon fuels, such as kerosene), with further compositions pending approval. However, there is an anticipation in the aviation industry that sustainable aviation fuel blends comprising up to (and including) 100% sustainable aviation fuel (SAF) will eventually be approved for use.

[0211] Sustainable aviation fuels may comprise one or more of n-alkanes, iso-alkanes, cycloalkanes, and aromatics, and may be produced, for example, from one or more of synthesis gas (syngas); lipids (e.g., fats, oils, and greases); sugars; and alcohols. Thus, sustainable aviation fuels may comprise lower aromatic and / or sulfur contents, relative to fossil-based hydrocarbon fuels. Additionally or alternatively, sustainable aviation fuels may comprise one and / or the other of higher iso-alkane content and cycloalkane content, relative to fossil-based hydrocarbon fuels. In some examples, sustainable aviation fuels may include a density between 90% and 98% of that of kerosene and / or a calorific value between 101% and 105% of that of kerosene.

[0212] In some examples, the sustainable aviation fuel(s), or blend(s) supplied to the combustion equipment 16 may have a relatively lower aromatics and / or other non-paraffinic content than kerosene. The sustainable aviation fuel may include an aromatics content of, for example, 30%, 20%, 15%, 10%, 8%, 5%, or less than 5%; for example, 4%, 3%, 2%, 1%, or less than 1%; for example, 0.75%, 0.5%, 0.25%, or less than 0.25 %; for example, 0.2%, 0.1% or less than 0.1%; for example, 0.01%, 0.001% or 0%. The aromatics content of the sustainable aviation fuel may be within an inclusive figure or range bounded by any two or more of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits), for example, 13.5%, 8.5%, 2.5%, 0.35%, 0.15%, 0.05%, 0.005% or 0%; or from 0% to 0.75%, from 0% to 0.5%, or from 0.1% to 0.25%; or from 0.15% to 0.65%, from 0.35% to 0.55%, or from 0.035% to 0.055%; depending on one or more preferences, fuel stock or supplier, and composition variations.

[0213] Due at least in part to the molecular structure of sustainable aviation fuels, sustainable aviation fuels may provide beneficial effects including, for example, one or more of higher specific energy (despite, in some examples, lower energy density); higher specific heat capacity; higher thermal stability; higher lubricity; lower viscosity; lower surface tension; lower freezing point; lower soot emissions; lower NOx; and, lower CO2 emissions, compared to fossil-based hydrocarbon fuels (e.g., when burned in combustion equipment 16).Thus, compared to fossil-based hydrocarbon fuels, such as kerosene, sustainable aviation fuels can lead to either or both a relative decrease in specific fuel consumption, and a relative decrease in maintenance costs.

[0214] 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 53a, 53b located in each wing. In other examples, an aircraft 1 may have only one fuel tank 50, and / or the wing fuel tanks 53 may be larger than the center fuel tank 50, or no center fuel tank may be provided (with all fuel being stored instead in the wings of the aircraft) - it will be understood that many different tank arrangements are contemplated and that the examples that have been illustrated are provided for ease of description and not intended to be limiting.

[0215] [Fig. 4] shows an aircraft 1 with a propulsion system 2 comprising two gas turbine engines 10. The gas turbine engines 10 are supplied with fuel from a fuel supply system on board the aircraft 1. The fuel supply system of the illustrated example comprises a single fuel source. For the purposes of the present application, the term "fuel source" means either 1) a single fuel tank or 2) a plurality of fuel tanks that are fluidically interconnected. Each fuel source is intended to provide a separate fuel source, i.e., a first fuel source may contain a first fuel having a characteristic or characteristics different from those of a second fuel contained in a second fuel source. The first and second fuel sources are therefore not fluidically coupled to each other so as to separate the different fuels (at least under normal operating conditions). The use of several fuel sources allows an aircraft 1 to carry several different fuels and to change the fuel used during operation, and possibly even during cruise or when changing from one stage of operation to another during flight.

[0216] In the present example, the first (and, in these examples, the only) fuel source comprises a center fuel tank 50, located primarily in the fuselage of the aircraft 1 and a plurality of wing fuel tanks 53a, 53b, where at least one wing fuel tank is located in the port wing and at least one wing fuel tank is located in the starboard wing for balancing. All of the tanks 50, 53 are fluidly connected in the illustrated example, thereby forming a single fuel source. Each of the center fuel tank 50 and the wing fuel tanks 53 may comprise a plurality of fluidly interconnected fuel tanks.

[0217] In another example, the wing fuel tanks 53a, 53b may not be fluidly connected to the center tank 50, thereby forming a second, separate fuel source. For balancing purposes, one or more port wing fuel tanks may be fluidly connected to one or more starboard wing fuel tanks. This may be done either through a center fuel tank (if that tank is not part of the other fuel source), or by bypassing the center fuel tank(s), or both (for maximum flexibility and safety). In another example, the first fuel source includes wing fuel tanks 53 and a center fuel tank 50, while a second fuel source includes another, separate center fuel tank.Fluid interconnection between the wing fuel tanks and the center fuel tank of the first fuel source may be provided for balancing the aircraft 1. In the aircraft 1 with multiple fuel sources, two or more of the fuel sources may therefore contain different fuels from each other, so that the aircraft 1 may change fuels during flight. Determining which fuel is supplied to the combustion chamber 16 may therefore be more complex than simply recording a single identity of a fuel on board the aircraft 1, or performing a check at the time of start-up.

[0218] In some examples, the distribution of fuel tanks 50, 53 available on the aircraft 1 may be limited such that the first fuel source and the second fuel source are each substantially symmetrical about the aircraft centerline. In cases where an asymmetric distribution of fuel tanks is permitted, a suitable means for transferring fuel is generally provided between the fuel tanks of the first fuel source and / or between the fuel tanks of the second fuel source such that the position of the center of mass of the aircraft can be maintained within acceptable lateral limits throughout the duration of the flight.

[0219] Aircraft typically refuel at several different airports, for example at the beginning and end of a long-haul flight. While there are standards that all aviation fuels must comply with, as mentioned above, different aviation fuels have different compositions, for example depending on their source (e.g., different sources of petroleum, biofuels, or other synthetic aviation fuels (often described as sustainable aviation fuels - SAF), and / or blends of petroleum-based fuels and other fuels) and any included additives (e.g., antioxidants and metal deactivators, biocides, static reducers, icing inhibitors, corrosion inhibitors) and any impurities.The composition of available aviation fuel may not only vary from airport to airport and fuel supplier to fuel supplier, but also from batch to batch, even for a given airport or fuel supplier. Furthermore, the fuel tanks 50, 53 of the aircraft 1 are generally not emptied before being refilled for a subsequent flight, resulting in mixtures of different fuels in the tanks - fuel with a different composition effectively resulting from the mixture.

[0220] The inventors were aware that, as different fuels may have different properties, while still remaining compliant with standards, knowledge of the fuel(s) available for an aircraft 1 may allow for more efficient and tailored control of the aircraft 1, and more specifically of the aircraft's propulsion system 2 (i.e., the gas turbine engine(s) 10 of the aircraft 1, and associated controls and components). Knowledge of the fuel may therefore be used as a tool to improve aircraft performance, such that determining or monitoring the fuel composition may have advantages.In particular, determining one or more characteristics of the fuel to be supplied to the combustion chamber 16 - whether it is a fuel from a single fuel source or a mixture of one or more fuels from different sources - is therefore important in determining . engine operation. A key element of engine operation is heat management - engine heat management is primarily accomplished using heat transfer to and from the oil and fuel in the engine, such that controlling a heat exchange system 3000 - e.g., by regulating oil flow, regulating fuel flow, and / or controlling one or more heat exchangers 1004, 1006, 2020, 2030 or other heat exchange components in other ways - may optimize the performance of the engine 10 for a fuel with particular fuel characteristics.

[0221] As used herein, the term "fuel characteristics" refers to intrinsic or inherent properties of the fuel, such as fuel composition, and not to variable properties such as volume or temperature. One or more fuel characteristics may be determined, and this data used to adjust control of the engine 10, and in particular the heat management system 3000. Examples of fuel characteristics include one or more of: i. the percentage of sustainable aviation fuel (SAF) in the fuel, or an indication that the fuel is a fossil fuel, e.g. fossil kerosene, or that the fuel is pure SAF fuel; ii. parameters of a distribution of hydrocarbons in the fuel, such as: • the aromatic hydrocarbon content of the fuel, and possibly also / alternatively the multi-aromatic hydrocarbon content of the fuel; • the hydrogen / carbon (H / C) ratio of the fuel; • % composition information for some or all of the hydrocarbons present; iii. the presence or percentage of a particular element or species, such as: • the percentage of nitrogen-containing species in the fuel; • the presence or percentage of a tracer species or trace element in fuel; • a naphthalene content of the fuel; • a sulfur content of the fuel; • a cycloparaffin content of the fuel; • oxygen content of the fuel; iv. one or more properties of the fuel used in a gas turbine engine 10, such as: • a level of emissions of non-volatile particulate matter or CO2 during combustion; • a level of fuel coking; v. one or more properties of the fuel itself, independent of its use in an engine 10 or of combustion, such as: • a thermal stability of the fuel (e.g., thermal degradation temperature); a thermal stability may be assigned a numerical value at any temperature by taking the inverse of a fuel degradation product deposition rate at the given temperature); and • one or more physical properties such as density, viscosity, calorific value, freezing temperature and / or heat capacity.

[0222] The fuel characteristic(s) to be determined may be chosen based on the fuel properties most relevant to modifications that may be made to the heat management system 3000. Determining fuel characteristics may include obtaining fuel characteristics of any fuel already present in the fuel tank 50, 53 prior to refueling and an indication of the amount of fuel remaining, and then combining this information with information regarding new fuel added to the tank 50, 53 upon refueling.

[0223] Obtaining fuel characteristics of any fuel already present in the fuel tank 50, 53 prior to refueling, and / or obtaining fuel characteristics of a fuel supplied at refueling, may include one or more of: (i) the physical and / or chemical detection of one or more features or parameters of the fuel composition (which may allow the direct detection of fuel characteristics and / or may allow the determination of fuel characteristics using the detection results), and / or the detection of one or more tracer elements or compounds added to the fuel to facilitate its identification (e.g., a dye); (ii) retrieving fuel characteristic information from an on-board memory or data store; and / or (iii) receiving data, for example from an input provided in a user interface, or data transmitted to the aircraft 1.

[0224] In some examples, one or more fuel characteristics may be determined during operation of the gas turbine 10, for example by deriving fuel characteristics of the fuel supplied to the combustion chamber. combustion 16 in operation from engine performance measurements, or by performing on-site detection.

[0225] In some examples, multiple different methods may be performed to obtain the fuel characteristics - e.g., different methods may be used for different characteristics, and / or different methods may be used for the same characteristics as a control. For example, stored or otherwise provided data on the fuel characteristics may be compared to the results of chemical or physical detection of one or more parameters of the fuel. If there is a mismatch between the stored fuel characteristic and the corresponding detected parameter, an alert may be provided.

[0226] Fuel characteristics may be determined by physically and / or chemically detecting one or more features of the fuel composition (e.g., in an off-site test unit, or when the fuel is transported to an on-site fuel tank, or when used in the gas turbine engine 10), thereby enabling the fuel characteristics to be directly detected or data to be provided from which they may be determined, as discussed above, and / or detecting one or more tracer elements or compounds added to the fuel to aid its identification (e.g., a dye); or by receiving data, e.g., from an input provided in a user interface, or data transmitted to the aircraft, e.g., by scanning a barcode associated with the fuel delivery.

[0227] When a physical and / or chemical determination is used, the fuel characteristics may be detected in various ways, both directly (e.g., from sensor data corresponding to the characteristic of the fuel in question) and indirectly (e.g., by deduction or calculation from other characteristics or measurements, or by reference to data of a specific tracer detected in the fuel). The characteristics may be determined as relative values ​​by comparison with 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 based on measurements of the swelling of a sensor component made from a sealing material such as a nitrile sealing material. • Trace substances or species, either naturally occurring in the fuel or added as tracers, can be used to determine fuel characteristics such as the percentage of sustainable aviation fuel in the fuel or the fact that the fuel is kerosene. Measurements of the vibrational mode of a piezoelectric crystal exposed to fuel can be used as a basis for determining various fuel characteristics, including the aromatic content of the fuel, the oxygen content of the fuel, and the thermal stability or coking level of the fuel - for example, by measuring the buildup of surface deposits on the piezoelectric crystal, which will cause a change in vibrational mode. Various fuel characteristics may be determined by collecting performance parameters of the gas turbine engine 10 during a first operating period (e.g., during takeoff), and possibly also during a second operating period (e.g., during cruise), and comparing these collected parameters to expected values ​​using a fuel of known properties. Various fuel characteristics, including the aromatic hydrocarbon content of the fuel, may be determined based on sensor measurements of the presence, absence, or degree of condensation trail formation by the gas turbine 10 during operation. Fuel characteristics, including aromatic hydrocarbon content, can be determined based on a UV-Vis spectroscopy measurement performed on the fuel. Various fuel characteristics, including sulfur content, naphthalene content, aromatic hydrogen content, and hydrogen-to-carbon ratio, may be determined by measuring substances present in the exhaust gases emitted by the gas turbine engine 10 during operation. The calorific value of the fuel may be determined during operation of the aircraft 1 based on measurements taken as the fuel is burned - for example using the fuel flow rate and shaft speed or the temperature change in the combustion chamber 16. Various fuel characteristics may be determined by making an operational change to affect the operation of the gas turbine engine 10, and detecting a response to the operational change; and determining the one or more fuel characteristics based on the response to the operational change. • Various fuel characteristics may be determined relative to the fuel characteristics of a first fuel by replacing the first fuel supplied to the gas turbine engine 10 with a second fuel, and determining the one or more fuel characteristics of the second fuel based on a change in a relationship between T30 and one of T40 and T41 (the relationship indicating the temperature increase in the combustion chamber 16). The characteristics may be determined as relative values ​​by comparison with the first fuel, or as absolute values, for example by reference to known values ​​for the first fuel.

[0228] As used herein, T30, T40, and T41, and any other numbered pressures and temperatures, are defined using the station numbering listed in the SAE AS755 standard, specifically: • T30 = Total outlet temperature of the high pressure compressor (HPC); • T40 = Total temperature at the combustion outlet; • T41 = Total temperature at the inlet of the high pressure turbine (HPT) rotor.

[0229] Any suitable approach known in the art may be used, and the determination of fuel characteristics will not be discussed further herein.

[0230] The aircraft 1 may therefore comprise a fuel composition determination module 57 for determining at least one fuel characteristic of the fuel. The determination may be made by obtaining 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 of [Fig. 4], a fuel composition determination module 57 is integrated into each engine 10, where the fuel enters the engine 10. In other implementations, the module 57 may be located differently within the engine 10 or located elsewhere on the aircraft 1, for example in, on, or adjacent to a fuel tank 50, 53.In implementations with a fuel composition determination module 57, the module 57 may be intended to provide an output to an electronic engine control (EEC), or the module 57 may be provided as part of an EEC. The fuel characteristic(s) determined by the fuel composition determination module 57 may therefore be used to influence control of the engine 10. A specialized controller 58 may be provided to process the output of the fuel composition determination module 57 and provide control instructions to the controllable engine components based on the fuel characteristic information. In other implementations, the . 58 control functionality may be provided by the EEC and no separate unit or module may be provided.

[0231] An exemplary fuel system 1000 for a geared gas turbine engine 10, the fuel system 1000 including a fuel flow path from the fuel tank 50 to the combustion chamber 16 of the gas turbine engine 10 of the aircraft 1 is shown schematically in [Fig. 5]. The fuel system 1000 includes both the fuel supply system 50, 1002 (which supplies fuel to the engine 10) and the fuel management system 1500 (which operates within the engine 10) of the aircraft 1. The fuel management system 1500 has a role in fuel temperature management as well as fuel flow, directing the fuel through one or more heat exchangers 1004, 1006 of the engine heat exchange system 3000.The heat exchange system 3000 includes portions of the fuel management system 1500, as well as the oil recirculation system 2000, 2000' (which is described in more detail below). The heat exchange system 3000 is the general term used for the systems and components used to transfer heat between fluids (particularly, oil and fuel) within the engine 10, and includes heat exchangers, valves, connecting pipes, and associated components, e.g., pumps, refrigeration devices, etc.

[0232] Returning to the fuel system 1000, fuel is pumped from the fuel tank 50 to the gas turbine engine 10 by a low-pressure fuel feed 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 referred to as a primary fuel-oil heat exchanger because oil flowing therethrough may be used to cool and lubricate the primary reduction gear 30 of the engine 10. The secondary fuel-oil heat exchanger 1004 may be referred to as an integrated drive generator fuel-oil heat exchanger because oil flowing therethrough may be used to cool and / or lubricate one or more components of an integrated drive generator (IDG) of the engine 10.In other implementations, a different type of generator may be used instead of an IDG—for example, 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 example being described thus includes two fuel-oil heat exchangers 1004, 1006. More or fewer fuel-oil heat exchangers may be provided in other implementations. The illustrated fuel management system 1500 is arranged so that fuel reaches the secondary fuel-oil heat exchanger 1004 before . the primary fuel-oil heat exchanger 1006. After leaving the primary fuel-oil heat exchanger 1006, the fuel then passes through an engine fuel pump 1003 and moves to the combustion chamber 16. The engine fuel pump 1003 can be described as a primary fuel pump.

[0233] The primary fuel-oil heat exchanger 1006 and the secondary fuel-oil heat exchanger 1004 are configured such that an oil flow is routed through each in addition to the fuel flow therethrough. The primary fuel-oil heat exchanger 1006 and the secondary fuel-oil heat exchanger 1004 are configured such that heat can be transferred between the oil and fuel flowing therethrough.In standard operation of the engine 10, for example at cruising conditions, the average temperature of the oil flow 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 flow 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 flow to a fuel flow flowing therethrough in operation.

[0234] The two oil flows (oil flow through the primary heat exchanger and oil flow through the secondary heat exchanger) may be separate - physically separate and possibly also chemically distinct oils, and / or having a different flow rate. A different oil may therefore flow through the primary fuel-oil heat exchanger 1006 than that flowing through the secondary fuel-oil heat exchanger 1004.

[0235] In general, at least the majority of the fuel that passes through the secondary fuel-oil heat exchanger 1004 also passes through the primary fuel-oil heat exchanger 1006. The two heat exchangers 1004, 1006 may therefore be described as being in series with each other, with respect to the fuel flow, and as being along a primary fuel flow path from the tank 50 to the combustion chamber 16. However, either or each heat exchanger 1004, 1006 may be provided with a bypass to allow some of the fuel to avoid passing through the respective heat exchanger, for example in the form of a bypass line 1005, as illustrated in [Fig. 5]. A valve (not shown) may determine what proportion of the fuel passes through the heat exchanger 1004 and what proportion passes through the bypass line 1005. In various implementations, a bypass line may be provided for each heat exchanger 1004, 1006, allowing a portion of the fuel to bypass one or both of the heat exchangers. In addition or alternatively, one or more bypass lines 2005, 2005' may be provided for the oil for one or both of the fuel-oil heat exchangers 1004, 1006, thereby allowing a portion of the oil to bypass one or more of the heat exchangers. A valve for regulating the flow of a fluid through a bypass line 1005 may be referred to as a bypass valve.

[0236] The secondary fuel-oil heat exchanger 1004 and the primary fuel-oil heat exchanger 1006 are thus designed such that an oil flow is routed through each in addition to the fuel flow - the oil flowing through one is different from the oil flowing through the other in the implementation being described, although it will be appreciated that the same oil may flow through one fuel-oil heat exchanger and then through another fuel-oil heat exchanger in other implementations.

[0237] The two heat exchangers 1004, 1006 are thus in separate closed loop systems 2000, 2000', (Figures 7, 8) with respect to the oil flow, in the implementation being described, namely the oils flowing through the primary and secondary fuel-oil heat exchangers are fluidically separated, and may be chemically distinct from each other. The two oil loop systems 2000, 2000' serve to circulate oil through their respective fuel-oil heat exchangers 1006, 1004, and optionally also through one or more other heat exchangers, for example air-oil or oil-oil heat exchangers as described below. The two oil loop systems 2000, 2000' together can be described as providing an oil recirculation system for the engine 10.

[0238] [Fig. 6] shows an exemplary alternative 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 combustion chamber 16 of the gas turbine engine 10 of the aircraft 1. As with the fuel system 1000, fuel is pumped from the fuel tank 50 by the low-pressure fuel feed pump 1002. The fuel then passes through a secondary fuel-oil heat exchanger 1004 and then a primary fuel-oil heat exchanger 1006 before reaching an engine fuel pump 1003, which pumps the fuel along its flow path to the combustion chamber 16. The fuel system 6000 differs from the fuel system 1000 in that the fuel system 6000 includes a recirculation valve 6010 located downstream of the primary fuel-oil heat exchanger 1006 and adapted to recirculate at least a portion of the fuel that has left the primary fuel-oil heat exchanger back to the inlet 1006a of the primary fuel-oil heat exchanger 1006, allowing additional heat transfer between the primary loop system oil and the fuel. The recirculation valve 6010 can determine what proportion of the fuel is recirculated, via the recirculation line 6011 and what proportion continues more directly to the combustion chamber 16. In the illustrated example, the recirculation valve 6010 is located 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.An additional recirculation pump (not shown) may be provided in some implementations in which the recirculation valve 6010 is positioned upstream of the engine fuel pump 1003, to provide a positive pressure gradient. A fuel return to tank (FRTT) pump may be present and also contribute to recirculation in some implementations. The recirculation valve 6010 is arranged to allow a controlled amount of fuel to be returned to the inlet 1006a of the primary heat exchanger 1006, thereby flowing through the primary heat exchanger 1006 multiple times before reaching the pump 1003 and the combustion chamber 16. It is contemplated that the recirculation valve may be positioned downstream of the engine fuel pump 1003 in alternative implementations, for example as illustrated in [Fig. 10].In such implementations, the recirculation valve 6010 would be intended to allow a regulated amount of fuel to be returned to the inlet 1006a of the primary heat exchanger 1006, thereby flowing through the primary heat exchanger 1006 and the pump 1003 multiple times before reaching the combustion chamber 16. This recirculation provides a mechanism for regulating the flow of fuel within the fuel management system 6500, as well as within the heat management system 3000, without altering the flow of fuel from the tank 50 to the engine 10. The recirculation may be used to regulate the heat transfer ratio, as described in more detail below.

[0239] A line 6011, which may be referred to as a recirculation line because it transports fuel from a point along the primary flow path through the engine 10 to an earlier point along that flow path such that the fuel must pass through the portion of the flow path between them again, may therefore be provided, leading from the recirculation valve to a point on the flow path upstream of the inlet 1006a to the primary heat exchanger 1006. In some implementations, a recirculation line 6011 and a recirculation line 6011 may be provided. Bypass lines 1005 may be provided for any given heat exchanger 1004, 1006. In some implementations, the same line 1005, 6011 may serve as both a recirculation line and a bypass line - one or more valves may be used to regulate the direction of fluid flow therethrough.

[0240] In the example illustrated in [Fig. 10], unlike that in [Fig. 6], the recirculation valve 6010 is located downstream of the pump 1003, such that the recirculated fuel has already passed through both the primary heat exchanger 1006 and the pump 1003, rather than just the primary heat exchanger 1006. The recirculation line 6011 returns the recirculated fuel to a point in the flow path upstream of both the primary pump 1003 and the primary heat exchanger 1006, such that the recirculated fuel passes through both of these components once more. The recirculated fuel through the pump may allow for more adjustable regulation of fuel flow to the combustion chamber 16 for a given shaft speed of the engine 10, noting that the pump speed (or a limited group of pump speed options) is often fixed by the shaft speed.

[0241] Regulation of the fuel flow in the fuel circuit 6000 - by recirculation and / or bypass of one or more heat exchangers - may contribute to heat management (for example, by influencing the fuel temperature at the inlet of the combustion chamber 16 or the pump 1003, or the amount of heat transferred to the fuel). Part of the control of the heat exchange system 3000 may therefore be the control of one or more valves regulating the flow of oil and / or fuel through one or more recirculation and / or bypass lines 6011, 1005.

[0242] The gas turbine engine 10 of the aircraft 1 being described also includes a recirculating oil system for supplying oil to lubricate and remove heat from a plurality of components. In the implementation being described herein, the oil recirculating system includes a primary oil loop system 2000 and a secondary oil loop system 2000', each of which is a closed loop oil system. An example of a closed loop primary oil system 2000 is schematically illustrated in [Fig.7] - as with the heat exchangers, this oil loop system is referred to as "primary" because it is responsible for oiling and cooling the main reduction gear 30 and, generally, is responsible for the main / primary cooling load of the engine 10. The primary closed loop oil system 2000 includes an oil reservoir 2002 adapted to contain a volume of oil.In some implementations, gases are removed from the oil within the oil reservoir 2002 by a degasser. A feed pump 2004 is provided. to pump oil from the oil reservoir 2000 to the main fuel-oil heat exchanger 1006. The average temperature of the oil entering the main fuel-oil heat exchanger 1006, at cruising conditions, is higher than the average temperature of the fuel entering the main fuel-oil heat exchanger 1006. In the main fuel-oil heat exchanger 1006, thermal energy is transferred from the oil flow to the fuel flow. In this way, the average temperature of the oil flow leaving the main fuel-oil heat exchanger 1006 is lower than the average temperature of the oil flow entering the main fuel-oil heat exchanger 1006, so that it is cooled before reuse as a lubricant and / or coolant, allowing the cooled oil to remove more heat from the system to be lubricated and / or cooled.In this way also, 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.

[0243] In a standard heat exchanger 1006, the fundamental limiting factor on the level of cooling provided is the temperature of the coolant (in this case, the fuel) - the temperature of the oil at the outlet of the heat exchanger 1006 cannot be lowered below the temperature of the fuel entering the heat exchanger 1006 (nor can the temperature of the fuel be raised above that of the oil entering the heat exchanger), even if an effectively infinite flow of fuel or oil were used. Temperature equilibrium is the limit, as defined in the second law of thermodynamics. In some implementations, as illustrated in [Fig. 7], a refrigeration cycle apparatus 1007 is therefore provided.The refrigeration cycle apparatus 1007 is intended to provide thermal elevation by transferring more heat from the oil to the fuel, such that the fuel temperature is increased more than it would be by simply passing through the heat exchanger, and in some cases, it is above the oil temperature. The refrigeration cycle apparatus 1007 may take the form of an additional closed-loop circuit with a refrigerant fluid, with an evaporator between the oil and the refrigerant liquid (to transfer heat from the oil to the refrigerant liquid) and a condenser between the fuel and the refrigerant liquid (to transfer heat from the refrigerant liquid to the fuel). The refrigeration cycle apparatus 1007 may additionally include a pump, a compressor, and / or an expansion valve / meter, and may be made according to any suitable design known in the art.An oil valve may regulate the amount of oil flowing into the refrigeration cycle apparatus 1007, and a fuel valve may regulate the amount of . fuel flowing through the refrigeration cycle apparatus 1007. The refrigeration cycle apparatus 1007 (and more particularly its pump and / or compressor, and possibly also the expansion valve or meter) is powered, usually electrically or mechanically (e.g., via a connection to a rotating shaft), to provide this additional, forced heat transfer. The refrigeration cycle apparatus 1007 is shown in association with the main (primary) fuel-oil heat exchanger 1006 in the illustrated implementation, thereby providing a fuel temperature boost after the fuel has passed through the two fuel-oil heat exchangers 1004, 1006, and further cooling the oil in the primary oil system 2000 before it is returned to the reducer 30.A bypass or recirculation line for oil and / or fuel may be provided around the refrigeration cycle apparatus 1007 in various implementations.

[0244] The oil flow in the primary oil loop system 2000 is then routed to a power reducer 30, which may also be described as a main reducer 30 of the gas turbine engine 10. The power reducer 30 is to receive an input from the core shaft 26 and to output a drive to the fan 23 via the fan shaft 42 and includes gears 28, 32, 38 and bearings (e.g., journal bearings) that may be oil lubricated and cooled. The engine 10 may also include one or more additional bearings to support the shafts 26, 42, which may be journal bearings.The oil may further be used to lubricate and / or cool the journal bearings, and generally increases in temperature significantly in use at cruising conditions, so as to help cool the bearings and reduction gear 30 as the oil flow carries heat away from the bearings and reduction gear 30. The oil may also be used to lubricate one or more other engine components 33, e.g., an auxiliary reduction gear (AGB) and / or one or more bearing chambers. The AGB 33, also known as an accessory drive, when present, is a reduction gear that is 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 the engine accessories, e.g., the fuel pumps, and generally handles large loads. A relatively large amount of heat may therefore be dumped into the oil from the AGB.One or more bearing chambers may be lubricated by the same oil, and may similarly dump heat into the oil. Per unit of oil flowing through them, the AGB and bearing chambers may add more heat to the oil than the main gear 30 in most implementations. The oil flow may be split into two or more parallel flows, for example a flow at . through the main reducer 30 and one flow through the other engine components, or several parallel flows through the main reducer 30 (e.g. through different components of the reducer) and separate flows through the AGB and the or each bearing chamber 33.

[0245] From the power reducer 30 (and / or other engine components 33 such as the AGB, if applicable), oil is collected in a sump 2008. A scavenge pump 2010 is adapted to pump oil from the sump 2008 back into the oil reservoir 2002, ready for reuse.

[0246] [Fig.8] illustrates the secondary oil loop system 2000', which is another closed loop oil system 2000'. The secondary closed loop oil system 2000' includes a secondary oil reservoir 2002' adapted to contain a volume of oil. In some implementations, gases are removed from the oil within the oil reservoir 2002' by a degasser. A secondary feed pump 2004' is configured to pump oil from the secondary oil reservoir 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, at cruise conditions, is higher than the average temperature of fuel entering the IDG fuel-oil heat exchanger 1004.In the IDG fuel-oil heat exchanger 1004, thermal energy is transferred from the oil flow to the fuel flow. In this way, the average temperature of the oil flow leaving the IDG fuel-oil heat exchanger 1004 is lower than the average temperature of the oil flow entering the IDG fuel-oil heat exchanger 1004. In this way, the average temperature of the fuel leaving 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 flow is then routed to / back to an integrated drive generator 2006, where it lubricates and / or cools moving components and is heated in the process.In some implementations, the oil may be used primarily as a coolant for the IDG 2006, and may perform minimal or no lubrication. From the integrated drive generator 2006, the oil is collected in a secondary sump 2008'. A secondary scavenge pump 2010' is configured to pump oil from the secondary sump 2008' back into the secondary oil reservoir 2002' ready for reuse. In some implementations, a refrigeration cycle apparatus may also be provided on the secondary oil loop system 2000'.

[0247] Figures 7 and 8 each illustrate a serial oil flow path, with the entire oil flow passing sequentially through each component (although it will be appreciated that one or more bypass or recirculation lines not shown may be provided for the oil). In other implementations, the oil flow may be split into two or more parallel flows, for example, a flow through the main fuel-to-oil heat exchanger 1006 and a flow through an air-to-oil heat exchanger 2020 (as described below).

[0248] [Fig. 9] schematically represents an exemplary alternative section of the primary closed-loop oil 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' is operable to divide the oil flow between the primary 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 primary fuel-oil heat exchanger 1006.The oil flow path can be referred to as a branch, with the main fuel-oil heat exchanger 1006 on one branch and the first air-oil heat exchanger 2020 on the other branch, in a parallel configuration such that oil can flow through one branch or the other, but the same portion of oil cannot pass through both on the same cycle - the flow splits. Valve 2016 modulates the flow through both heat exchangers 1006, 2020, and thus may be described as a modulating valve 2016. The oil flows are then recombined and routed to the power reducer 30 and other engine components 33. Any appropriate percentage of oil may flow through each of the first air-to-oil heat exchanger 2020 and the main fuel-to-oil heat exchanger 1006.In some examples, the valve 2016 is operable to vary the flow of oil to the primary fuel-oil heat exchanger 1006 and the first air-oil heat exchanger 2020 upon request. In various examples, an oil-oil heat exchanger 2030 (not shown in [Fig. 7] but present in [Fig. 11]) may be provided, for example being arranged in series with the first air-oil heat exchanger 2020 on this branch of the parallel split. The oil-oil heat exchanger 2030 may provide heat exchange between the primary and secondary closed-loop oil systems 2000, 2000'.

[0249] [Fig. 10] schematically shows an exemplary arrangement and interaction of the first closed-loop oil system 2000, the second closed-loop oil system 2000' and the fuel system 1000, with fuel flow shown in thick black lines and oil flow shown in lines thinner black lines. The thick black striped line indicates a recirculation path 6011 taken by only a portion of the fuel. The combination of the fuel 1000 and oil systems 2000, 2000' together forms the heat exchange system 3000. The primary closed-loop oil system 2000 of this exemplary arrangement is arranged as shown in [Fig. 7]. The secondary closed-loop oil 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 line 6011 as described above. A bypass line 1005 may also be present, but is not shown for clarity.

[0250] 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-loop oil system 2000' is designed such that its oil recirculation flow also flows through the IDG fuel-oil heat exchanger 1004. In standard operation of the engine 10, at both cruising and idling conditions, 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 designed such that heat is transferred from the oil flow to the fuel flow.In this way, the average temperature of the oil flow at the outlet of the fuel-oil heat exchanger of IDG 1004 is lower than the average temperature of the oil flow at the inlet of the fuel-oil heat exchanger of IDG 1004. In the same way, the average temperature of the fuel flow at the outlet of the fuel-oil heat exchanger of IDG 1004 is higher than the average temperature of the fuel flow at the inlet of the fuel-oil heat exchanger of IDG 1004.

[0251] The fuel then flows through the primary fuel-oil heat exchanger 1006, and additionally through the refrigeration cycle apparatus 1007. The primary closed-loop oil system 2000 is designed such that its oil recirculation flow also flows through the primary fuel-oil heat exchanger 1006 and the refrigeration cycle apparatus 1007. In standard operation of the engine 10, at both cruising and idling conditions, the average temperature of the oil flow entering the primary fuel-oil heat exchanger 1006 is higher than the average temperature of the fuel flow entering the primary fuel-oil heat exchanger 1006. The primary fuel-oil heat exchanger 1006 is designed to such that heat is transferred from the oil flow to the fuel flow. In this manner, the average temperature of the oil flow at the outlet of the main fuel-oil heat exchanger 1006 is lower than the average temperature of the oil flow at the inlet of the main fuel-oil heat exchanger 1006. Similarly, the average temperature of the fuel flow at the outlet of the main fuel-oil heat exchanger 1006 is higher than the average temperature at the inlet of the main fuel-oil heat exchanger 1006. The refrigeration cycle apparatus 1007, when active (i.e., when energized / turned on and used to actively move heat from the oil to the fuel), may provide an additional fuel temperature boost / oil temperature drop, possibly by raising the fuel above the oil temperature.After flowing through the primary fuel-oil heat exchanger 1006, fuel flows to the engine fuel pump 1003, which is located downstream of the primary and secondary fuel-oil heat exchangers 1006, 1004 in the examples shown and is intended to deliver fuel to the combustion chamber 16 of the gas turbine engine 10.

[0252] At cruising conditions, the average temperature of the oil flow flowing through the IDG fuel-oil heat exchanger 1004 may be lower than the average temperature of the oil flow flowing through the main fuel-oil heat exchanger 1006. In this manner, the fuel passes through the heat exchanger 1004 having a lower average oil flow temperature first, before passing through the heat exchanger 1006 having a higher average oil flow temperature.In addition to the branched oil flows 2000, 2000', the heat exchange system 3000 may further include branched fuel return paths, such that fuel is returned to the main fuel path from where fuel enters the gas turbine engine 10 to the combustion chamber 16 in at least two different locations - for example upstream of, or downstream of, the main engine pump 1003 as shown in [Fig. 6] with a branched path 6020 branching off from the main fuel flow path at the outlet 1006b from the primary heat exchanger 1006 and rejoining the main fuel flow path downstream of the pump 1003. At least one valve (not shown) may be provided to regulate a split of the fuel flow from the heat exchanger 1006 back to the main fuel path through the engine 10.The valve can be controlled based on fuel temperature - for example to send less fuel through pump 1003 / more fuel to a position downstream of the pump if the temperature of . fuel is relatively high and more likely to degrade pump seals or other components. Fuel flow control in branched return paths may be based on fuel temperature measurement (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). A return of recirculated fuel upstream of the fuel-oil heat exchanger may reduce heat transfer from the oil to the fuel, thereby dampening a transient overshoot that might otherwise occur at the beginning of the downshift, for example, when the same amount of heat is generated within the oil circuit, but fuel flow is reduced, typically resulting in a temperature spike.Adjusting the fuel flow through these branched paths can also be useful in reducing icing, by recirculating fuel that has been preheated to warm the system. The branched fuel return paths can therefore be used in a variety of ways to improve engine thermal management. [Fig. 11] schematically shows an example configuration of the primary closed-loop oil system 2000 and the secondary closed-loop oil system 2000' where the two independent oil recirculation flows are brought into a heat exchange relationship through an oil-to-oil heat exchanger 2030. Both oil loop systems 2000, 2000' have a branched arrangement of parallel pipes / heat exchangers in this example.

[0253] In the example illustrated in [Fig. 11], the primary closed-loop oil system 2000 is configured such that the oil recirculation flow is pumped by the feed pump 2004 through a valve 2016, which may be referred to as a modulating valve. The valve 2016 is operable to split the oil flow such that a portion of the oil flow passes to each of the main fuel-oil heat exchanger 1006 and a first air-oil heat exchanger 2020. In the illustrated embodiment, the first air-oil heat exchanger 2020 is in series with the oil-oil heat exchanger 2030, and the arrangement of the air-oil 2020 and oil-oil 2030 heat exchanger is in parallel with the main fuel-oil heat exchanger 1006. The modulating valve 2016 determines what proportion of the oil passes through each branch of the parallel arrangement.In various implementations, any suitable portion of the oil flow may be diverted between the primary fuel-oil heat exchanger 1006 and the first air-oil heat exchanger 2020. In examples, the valve 2016 is operable to divert a fixed portion of the oil flow to each of the primary fuel-oil heat exchanger 1006 and the first. air-oil heat exchanger 2020. In other examples, the valve 2016 may operate to divert a variable portion of the oil flow to each of the primary fuel-oil heat exchanger 1006 and the first air-oil heat exchanger 2020, for example using data from a temperature sensor for feedback to regulate the variable portion, and / or based on one or more fuel characteristics, as described above. The valve 2016 may be adjustable between a distinct number of fixed positions or continuously adjustable.In some implementations, for example when the fuel temperature at the inlet to the combustion chamber 16 is relatively low compared to the maximum operational fuel temperature (e.g., based on knowledge of the fuel type or thermal stability), no oil may be sent to the air-to-oil heat exchanger 2020, and all of the oil may be sent through the fuel-to-oil heat exchanger 1006.

[0254] After flowing through the heat exchangers 1006, 2020, 2030, the oil flow in the primary closed loop system 2000 is then recombined and routed to the power reducer 30 (and / or other engine components 33 such as the AGB) and then to the crankcase 2008. The scavenge pump 2010 then pumps the oil from the crankcase 2008 to the oil reservoir 2002, for reuse.

[0255] The oil flow within the secondary closed-loop oil system 2000' is intended to be brought into a heat exchange relationship with the separate oil flow within the primary closed-loop oil system 2000 through the oil-oil heat exchanger 2030. In the oil-oil heat exchanger 2030 the oil flow within the primary closed-loop oil system 2000 does not mix with the oil flow within the secondary closed-loop oil system 2000'. The oil-oil heat exchanger 2030 is designed such that heat transfer can take place between the two separate oil flows. In this way, heat from a warmer oil flow can be transferred to the cooler oil flow within the oil-oil heat exchanger 2030.No air-oil heat exchanger is shown in the illustrated 2000' closed-loop secondary system, but an air-oil heat exchanger may be provided—for example, in series with the 2030 oil-oil heat exchanger or on a third parallel branch—in other examples.

[0256] In the implementation shown in [Fig. 11], the secondary closed-loop oil system 2000' is configured such that the oil recirculation flow is pumped by the secondary feed pump 2004' through a valve 2016', which may be referred to as a secondary modulating valve. The valve 2016' is operable to divert at least a portion of the oil flow between the IDG fuel-to-oil heat exchanger 1004 and the oil-to-oil heat exchanger 2030, where the oil-to-oil heat exchanger 2030 is arranged in parallel with the IDG fuel-to-oil heat exchanger 1004. In other implementations, the secondary closed-loop oil system 2000' may not have a branched configuration. For example, the IDG fuel-to-oil heat exchanger 1004 and the oil-to-oil heat exchanger 2030 may be arranged in series such that at least a majority of the oil passing through one also passes through the other in a given cycle.

[0257] In other implementations, an air-to-oil heat exchanger 2020 may be present on each closed-loop system, or only on the secondary closed-loop system, and / or no oil-to-oil heat exchanger may be present.

[0258] In examples, any suitable portion of the oil flow may be diverted between the IDG fuel-oil heat exchanger 1004 and the oil-oil heat exchanger 2030. In examples, the valve 2016' is operable to divert a fixed portion of the oil flow to each of the IDG fuel-oil heat exchanger 1004 and the oil-oil heat exchanger 2030. In examples, the valve 2016' is operable to divert a variable portion of the oil flow to each of the IDG fuel-oil heat exchanger 1004 and the oil-oil heat exchanger 2030. The valve 2016' may be adjustable between a distinct number of fixed positions, or continuously adjustable. After flowing through the heat exchangers 1004, 2030, the oil flow is then routed to the integrated drive generator 2006 and then to the secondary housing 2008' in the example shown in [Fig.8].The secondary scavenge pump 2010' then pumps oil from the secondary sump 2008' to the secondary oil reservoir 2002', for reuse. It will be appreciated that the oil loop components may be arranged differently and that additional or alternative oil loop components may be present in other implementations.

[0259] One or more temperature sensors may be provided, for example, for detecting fuel temperature as it approaches the fuel pump 1003 or the combustion chamber 16. Data from the one or more sensors, possibly 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 manage fuel and / or oil flows through and around the heat exchangers 1004, 1006, 2020, 2030. The controller 58 may be part of, or provided by, an EEC, or may be a separate unit. The control may be automated, for example, by an EEC.

[0260] The inventors were aware that, in various implementations, knowledge of one or more fuel characteristics can be used to adapt the operation of the heat exchange system 3000 so as to improve the engine performance 10, taking advantage of properties that vary between aviation fuels.

[0261] [Fig. 12] illustrates a first method 100 of using knowledge of fuel characteristics to determine control of the heat exchange system 3000. The method 100 comprises determining 102 at least one fuel characteristic of the fuel to be burned by the combustion chamber 16. This determination 102 may be performed by any of the methods described above, possibly using a fuel composition determination module 57.

[0262] The engine 10 used to implement the method 100 comprises an air-oil heat exchanger 2020 and a fuel-oil heat exchanger 1006, and a modulation valve 2016 for varying the proportion of oil sent through each heat exchanger, and the method 100 comprises controlling 104 the modulation valve based on the at least one fuel characteristic so as to adjust the proportion of oil sent through each heat exchanger to cruise conditions.

[0263] More specifically, for the implementations described with respect to Figures 5-11 above, the method 100 includes controlling 104 the modulation valve 2016 of the primary oil system 2000 at cruise conditions based on the at least one fuel characteristic so as to adjust the proportion of oil sent through each branch of the primary oil loop system 2000, thereby regulating the amount of oil passing through the main fuel-oil heat exchanger 1006, and the amount passing through the air-oil heat exchanger 2020.In implementations with different heat exchanger arrangements, for example a series arrangement of heat exchangers instead of a parallel split, the modulating valve 2016 may effectively be a bypass valve allowing one of the heat exchangers 1006, 2020 to be bypassed by a controllable proportion of the flow, thereby allowing the proportion of oil sent through each heat exchanger to be varied in this manner.

[0264] As indicated by the dotted line in [Fig. 12], this method 100 may optionally be repeated. In some implementations, for example in implementations with a single fuel on board the aircraft 1, the method 100 may only be performed once during a flight cycle - for example during refueling. In other implementations, however, for example in implementations with multiple fuel sources, the fuel supplied to the combustion chamber 16 may vary over time during a flight. The determination 102 may therefore be performed more than once - for example: (i) once for each fuel source 50, 53 when refueling / starting a flight in implementations with multiple fuel sources of which only one or the other is used at any given time (and knowledge of where the fuel source is drawn from may result in the selection of the appropriate stored fuel characteristic); or (ii) frequently during a flight, for example in response to a change in the fuel source(s) (noting that the fuel supplied to the combustion chamber 16 may be a mixture of fuels from different sources, in some implementations), or at regular intervals.

[0265] The modulation valve 2016 may therefore be intended to divert 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 intended to divert a variable 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 during a single flight (option (ii) above).The modulation valve 2016 may therefore be actively controlled to vary the proportion of oil sent through each heat exchanger 1006, 2020, particularly in implementations where the aircraft 1 carries several different fuels in different tanks and may change the fuel (or fuel blend) used mid-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 part of a more general EEC. In various implementations, the control of the modulation valve 2016 may be closed-loop or open-loop, depending on the availability of online measurements—in particular, a closed-loop approach to control may be preferred when feedback data (e.g., from oil flow sensors) is available.

[0266] The oil loop system 2000, which may also be referred to as an oil recirculation system, may branch such that a proportion of the oil may flow along each branch, and the air-to-oil 2020 and fuel-to-oil 1006 heat exchangers may be arranged in a parallel configuration on different branches of the oil loop system, as shown in Figures 9 and 11. The modulating valve 2016 of such examples may be intended to allow the proportion of the oil sent through each branch to be varied, and the control 104 of the modulating valve 2016 may therefore adjust the proportion of the oil sent through each branch at cruising conditions. In some examples, more than two heat exchangers and / or more than two branches may be provided. Modulating valve 2016 therefore allows the amount of heat transferred from the oil by the air-to-oil 2020 and fuel-to-oil 1006 heat exchangers to be varied. The inventors were aware that fuels having certain fuel characteristics may accept more heat than others without breaking down, for example, due to higher heat capacity and / or higher thermal stability, or may otherwise benefit from a higher temperature (e.g., for fuel combustion efficiency).Adaptive control of the heat exchange system 3000 based on the determined fuel characteristic(s) therefore enables more efficient use of different fuels, as well as more efficient engine operation and oil cooling.

[0267] The fuel characteristic(s) may not be the only data considered in adjusting the modulation valve 2016 - for example, the temperature of the fuel leaving the fuel-oil heat exchanger 1006 or entering the combustion chamber 16 may also be considered, as well as the heat tolerance of fuel system components downstream of the heat exchanger 1006 (and possibly also the refrigeration cycle apparatus 1007, when present). In the presence of a refrigeration cycle apparatus 1007, the modulation valve 2016, or another oil valve, may regulate the amount of oil that flows through the refrigeration cycle apparatus 1007.In some implementations, all of the oil passing through the main fuel-oil heat exchanger 1006, and possibly all of the fuel passing through that heat exchanger 1006, may also pass through the refrigeration cycle apparatus 1007 - the refrigeration cycle apparatus 1007, however, may not be powered / operational for some of that time, and thus not provide any temperature rise.

[0268] Depending on the suitability of the at least one determined fuel characteristic, the method 100 may comprise using the modulation valve 2016 to perform one or more of the following operations: • sending all of the oil through the fuel-to-oil heat exchanger 1006 (none through the air-to-oil heat exchanger 2020) for one or more periods of at least 30 minutes at cruise, such that no heat is lost to the environment through the air-to-oil heat exchanger during at least some periods of cruise operation; • send all oil through the 1006 fuel-oil heat exchanger (none through the 2020 air-oil heat exchanger) for at least 15% of the total time spent in cruise mode, and optionally for at least 20% of the total time spent in cruise mode; • send at least 95% of the oil through the 1006 fuel-oil heat exchanger (no more than 5% through the 2020 air-oil heat exchanger) for at least 90% of the time spent in cruise mode, such that very little, if any, heat is lost to the environment through the air-oil heat exchanger for at least 90% of the cruise operation; • regulate the oil flow so that at least 80% of the heat transferred from the oil at cruising speed is transferred to the fuel; and / or • regulate the oil flow so that all heat transferred from the oil at cruise is transferred to the fuel for at least 90% of the time spent at cruise; and / or • regulate the oil flow such that between 200 and 600 kJ, and possibly between 200 and 500 kJ, of heat from the oil is transferred to the fuel in the heat exchange system 3000 at cruising conditions, per kilogram of fuel. In some implementations, the heat transferred may be in the range of 350 to 450 kJ / kg of fuel.

[0269] The suitability of the at least one determined fuel characteristic may be determined by comparing one or more fuel characteristics to a threshold value set for the respective fuel characteristics. For example, a fuel with a SAF content (%SAF) above 60%, 65%, 70%, 75% or 80% may be considered suitable for an oil flow reduction through the 2020 air-oil heat exchanger compared to traditional aviation fuels. Multiple fuel characteristics may be considered together in some implementations - for example with a lower threshold for one fuel characteristic when another fuel characteristic falls within a specific range. The fuel characteristics may be considered independently / in isolation in other implementations.

[0270] The at least one fuel characteristic may be or include one or more of thermal stability, aromatic hydrocarbon content of the fuel, or percentage of sustainable aviation fuel in the fuel, and decisions regarding control of the modulation valve may be made based on knowledge of the at least one fuel characteristic. For example, provided that 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 transferred from the oil at cruise may be transferred to the fuel. As a further example, provided that the mole percentage of 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 transferred from the oil at cruise may be transferred to the fuel.As a further example, provided that the S AF 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 to 100% of the heat transferred from the oil at cruise may be transferred to the fuel. As a further example, provided that the calorific value of the fuel is at least 43.5 MJ / kg, at least 80%, and optionally 90 to 100%, of the heat transferred from the oil at cruise may be transferred to the fuel.Generally speaking, less heat can be transferred to a conventional fuel if the fuel's heating value exceeds a threshold, because, for higher heating value fuels, the flow rate is usually reduced to achieve the same thrust level 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., FAS), with higher thermal stabilities, also have higher heating values ​​in many cases - the fuel temperature rise due to a lower flow rate may therefore be acceptable in some scenarios, and even additional heat addition beyond this may also be possible.This demonstrates the usefulness of considering multiple fuel characteristics jointly in some implementations, instead of a single fuel characteristic.

[0271] Any implementation may use a single or multiple fuel characteristics. A processing module may be provided, possibly as part of the fuel characteristics determination module 57, and / or as part of a general EEC engine, to make decisions about the control of the modulation valve 2016 based on the fuel characteristics data, and possibly also on other data.

[0272] A gas turbine engine 10 for an aircraft implementing this method 100 comprises a reduction gear 30 which receives an input from the core shaft and delivers as output a blower drive 23, and an oil loop system 2000 for supplying oil to the reduction gear 30. The heat exchange system 300 of the engine 10 includes an air-oil heat exchanger 2020 through which the oil in the oil loop system 2000 flows, and a fuel-oil heat exchanger 1006 through which the oil in the oil loop system and the fuel flow, such that heat is transferred between the oil and the fuel. A modulating valve 2016 is provided, the valve 2016 being for allowing the proportion of oil sent through each heat exchanger to be varied. In various examples, including that illustrated in [Fig.9], the oil loop system 2000 branches such that a proportion of the oil can 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 loop system 2000. The modulation valve 2016 is intended to allow the proportion of the oil sent through each branch to be varied. The engine 10 also comprises a fuel composition determination module 57, arranged to determine at least one characteristic of fuel arranged to be combusted by the combustion chamber 16.The modulation valve 2016 is to be controlled based on at least one fuel characteristic, as determined by the fuel composition determination module 57, so as to adjust the proportion of oil sent through each heat exchanger 1006, 2020 to cruise conditions. A controller 58 may be provided to make and implement decisions based on the output of the fuel composition determination module 57.

[0273] In various examples, including that illustrated in [Fig.l 1], the heat exchange system 3000 also includes a secondary oil loop system 2000' as described above, and an oil-to-oil heat exchanger 2030 for allowing heat to be exchanged from one oil to the other in the two oil systems 2000, 2000'. Modulating valve 2016, or an additional valve, may control oil flow from primary oil loop system 2000 to oil-oil heat exchanger 2030. Secondary oil loop system 2000' may also have a valve to regulate oil flow to oil-oil heat exchanger 2030. In the example shown in [Fig.l 1], each oil system 2000, 2000' has a total of two parallel branches, and oil-oil heat exchanger 2030 is on the same branch as air-oil heat exchanger 2020 of primary oil system 2000.

[0274] The heat exchange systems 3000 of various examples include at least one bypass line for allowing fuel (or oil) to bypass one or more heat exchangers. A bypass line may effectively form an additional branch in a parallel branch oil system, in some implementations - the modulating valve 2016, or another oil valve, may be intended to adjust the amount of oil sent through the or each bypass line for oil, based on one or more determined fuel characteristics, and possibly on one or more temperature measurements.

[0275] [Fig. 13] illustrates a second method 200 of using knowledge of fuel characteristics to determine control of the heat exchange system 3000. The method 200 comprises determining 202 at least one characteristic of fuel to be supplied to, and combusted by, the combustion chamber 16. This determination 202 may be performed by any of the methods described above, possibly using a fuel composition determination module 57. The details of this determination step 202 may be identical to those of the determination 102 described for the method 100 illustrated in [Fig. 12]. The method of [Fig. 13] further comprises modulating 204 the heat exchange system 3000 so as to adjust the fuel temperature at the inlet of the combustion chamber 16 in cruising mode to a fixed level, this level being based on the at least one determined fuel characteristic.

[0276] Modulation 204 of heat exchange system 3000 may include any one or more of: • adjusting any suitable oil or fuel valve, for example modulating valve 2016, to alter the proportion of oil or fuel sent through a particular heat exchanger, or through a refrigeration cycle apparatus 1007, where present; • adjusting a modulation valve 2016 - when present - to modify the proportion of oil sent through one or more branches of an oil system 2000; • adjusting a bypass valve - where provided - so as to adjust the proportion of a fluid (e.g., oil or fuel) that bypasses a particular heat exchanger or set of heat exchangers; • adjusting a recirculation valve 6010 - when present - so as to adjust the proportion of a fluid (e.g., oil or fuel) that is returned to an earlier point in its path, thereby passing multiple times through one or more components (typically including a particular heat exchanger, or set of heat exchangers); • activating, deactivating or otherwise adjusting the use of a 1007 refrigeration cycle apparatus - where present - so as to provide a regulated boost to the heating of one fluid / the cooling of another fluid; • adjusting a pumping speed of an oil pump of the 2010, 2010' oil system; • closing one or more branches of the oil flow in an oil system 2000, 2000', when the or each oil system has branched oil passages; and / or • adjusting a valve to regulate airflow to an air-oil heat exchanger 2020, when present, or otherwise modifying airflow through that heat exchanger 2020.

[0277] As indicated by the dotted line in [Fig. 13], this method 200 may optionally be repeated multiple times during a flight / during operation of the engine 10, as described above with respect to the method of [Fig. 12]. Both the determination 202 and the subsequent modulation 204 may therefore be performed during cruise in some implementations.

[0278] The fuel characteristic(s) may not be the only data considered 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 (when present) may be considered (possibly using altitude - possibly in conjunction with geographic region - as a reference for temperature), as may fuel flow rates and fuel temperatures.

[0279] In some implementations, the step of modulating 204 the heat exchange system 3000 to regulate the fuel temperature comprises regulating the flow of oil through the at least one fuel-oil heat exchanger 1006 such that between 50% and 100% of the heat lost by the oil is transferred to the fuel, and optionally such that between 80% and 100%, or between 90% and 100%, of the heat lost by the oil is transferred to the fuel.

[0280] The determined fuel characteristic may be or include thermal stability of the fuel; in such implementations, the heat exchange system 3000 may be modulated such that the fuel temperature at the inlet of the combustion chamber 16 at cruising speed increases, possibly linearly, with increasing thermal stability. Thermal stability is the temperature at which the fuel begins to degrade and form deposits that can cause component blockages and malfunctions - it will be appreciated that this is generally not a degradation instantaneous when a given temperature is reached; aviation fuels include various components that are subject to degradation at different temperatures, and time spent at high temperature is also a factor in fuel degradation. A fuel can be considered stable at a given temperature if its degradation rate is below a certain threshold at that temperature.

[0281] In alternative or additional implementations, the determined fuel characteristic 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 fuel temperature at the inlet to the combustor during cruise operation is set at a predefined level corresponding to the tracer species. Optionally, the amount of the tracer species detected may indicate whether the fuel is a mixture of a marked aviation fuel (i.e., a fuel to which a tracer species has been added) with another fuel, or whether it is solely a marked aviation fuel.

[0282] In alternative or additional implementations, the determined fuel characteristic may be or include the percentage of SAF in the fuel (%SAF). In such implementations, the heat exchange system 3000 may be modulated such that the fuel temperature at the inlet of the combustion chamber 16 at cruise increases with increasing %SAF once the %SAF exceeds 60%, or possibly once the %SAF exceeds 70%, 75%, or 80%. When the %SAF is lower, no modulation based on SAF content may be performed in some implementations, but modulation 204 based on one or more other characteristics may be—so it will be kept in mind that multiple different fuel characteristics may be considered together (in other implementations, modulation based on SAF content may be performed when the %SAF is lower).Any implementation may use a single or multiple fuel characteristics. A processing module may be provided, possibly as part of the fuel characteristics determination module 57, and / or as part of a general EEC engine, to make decisions about modulating the heat exchange system 204 based on the fuel characteristics data, and possibly other data as well.

[0283] A gas turbine engine 10 for an aircraft implementing this method 200 does not necessarily include a reduction gear 30 that receives an input from the core shaft and outputs a drive to the fan 23 - i.e., the method 200 can be used with both geared engines 10 and direct drive engines. The method 200 can therefore be applied to a wide range of aircraft engines, provided that they comprise a combustion chamber 16, an oil system 2000 for circulating oil and a heat exchange system 3000 comprising at least one fuel-oil heat exchanger for transferring heat from the oil to the fuel. A fuel composition determination module 57, for determining at least one fuel characteristic of the fuel to be combusted 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 comprise, or be in communication with, a processing module for making decisions based on the one or more determined fuel characteristics, and possibly also other data.

[0284] In some implementations, the heat exchange system 3000 includes branched fuel return paths and at least one valve controlling a split of the fuel flow. The branched paths may be for returning fuel from the heat exchange system 3000—and particularly fuel leaving the primary 1006 or secondary 1004 fuel-oil heat exchanger—to at least two different locations along a primary fuel path, from where the fuel enters the gas turbine 10 to the combustor 16, as described above.For example, the fuel flow leaving the heat exchanger 1004, 1006 may be split in two, with one path generally taking the majority of the fuel (and thus being referred to as the main fuel flow path) and the other path diverting a portion of the fuel leaving the heat exchanger 1004, 1006 such that the diverted portion of fuel bypasses one or more main fuel flow path components (e.g., a pump) and rejoins the main flow to the combustion chamber 16 downstream of those one or more main fuel flow path components. The valve may control and adjust the proportion sent through each branch. More than two branches may be provided in some implementations.

[0285] The inventors were also aware that the use of fuels different from traditional kerosene-based jet fuels, such as sustainable aviation fuels, may lead to different fuel properties, and that operating heat transfer parameters may be adjusted to utilize the different fuel properties. In particular, certain fuels may be heated to higher temperatures in one or more fuel-oil heat exchangers than traditional fuels, without significantly increasing coking. This may allow a process providing improved oil cooling (because the fuel is able to absorb more heat) and may also improve the overall thermal efficiency of the engine, with less heat lost to the environment. The 3000 Controllable Heat Exchange System plays a key role in managing the heat transfer ratio.

[0286] Furthermore, although cruise conditions generally represent a much larger proportion of an aircraft engine's operating time, the inventors were aware that idling 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 increase - therefore, the use of non-traditional fuels may have an even greater effect on optimal approaches to heat management under idle conditions. The methods 300, 400 of Figures 14 and 15 address both of these aircraft operating scenarios.

[0287] [Fig. 14] illustrates a method 300 implementing these considerations at cruise conditions, and [Fig. 15] illustrates a method 400 implementing these considerations at idle, for example, when starting the aircraft, operating it while stationary during boarding, and taxiing (to a runway or hangar, or between other ground locations), or at certain periods of flight (for example, descent).

[0288] First examining the method 300 to be performed in cruise mode, the method 300 is intended to be performed in a geared gas turbine engine 10 comprising an oil loop system 2000 for supplying oil to the reduction gear 30; and a heat exchange system 3000 comprising an air-oil heat exchanger 2020 through which the oil in the oil loop system flows; and a fuel-oil heat exchanger 1006 through which the oil in the oil loop system and the fuel flow such that heat is transferred between the oil and the fuel. The oil loop system 2000 branches, as shown in Figures 9 and 11, such that a proportion of the oil can flow along each branch, with the air-to-oil and fuel-to-oil heat exchangers 1006, 2020 arranged in a parallel configuration on different branches of the oil loop system.A modulating valve 2016 is also provided to regulate the proportion of oil sent through each branch of the oil loop system 2000. The method 300 includes controlling 302 the heat exchange system 3000 such that, under cruising conditions, a heat transfer ratio of: . oil to air heat transfer rate (kJkg'1) heat transfer rate from oil to fuel (kJkg'l) 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 may be provided to implement this control.

[0289] It will be borne in mind that even for a specific engine 10 operating with a fixed fuel, it is usual to have a range of values ​​for this ratio at cruise due to varying conditions - for example, an upper limit of the range may apply in cold weather (ISA-30 conditions) at low altitude (cruise: 35,000 feet), with low heat generation in the oil system, and a lower limit may apply in hot weather (ISA+40 conditions) at high altitude (cruise: 39,000 feet), with high heat generation in the oil system.

[0290] The method 300 may further comprise receiving data 304 to enable the heat transfer ratio to be calculated or derived, for example, temperature data (of the oil and / or fuel, at one or more points around an oil loop system 2000, 2000' or a fuel flow path, and / or possibly a fuel tank temperature or an oil tank temperature), and fuel flow rate data. This data may be received 304 by a controller 58 and used 306 to adjust the modulation 302 of the heat exchange system 3000 in cruise mode, so as to maintain the heat transfer ratio at a desired level or within 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 as part of the control 302 of the heat exchange system 3000. The method 300 may also be intended to use other information, e.g., temperature data (of the oil, fuel, and / or ambient temperature of an environment around the aircraft 1), flow rate data (of the oil and / or fuel), and / or one or more fuel characteristics, to determine 306 the control actions to be taken.

[0291] The control step 302 of the heat exchange system 3000 may comprise any one, or any combination, of the examples provided for the modulation step 204 of the method 200 of [Fig. 13], for example, by decreasing the amount of oil sent through 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 sent through a respective bypass line 1005 or recirculation line 6011.

[0292] In implementations where the engine 10 includes a refrigeration cycle apparatus 1007, the control step 302 of the heat exchange system 3000 may include using the refrigeration cycle apparatus 1007 to provide thermal lift by transferring more heat from the oil to the fuel beyond that transferred when passing through the one or more fuel-oil heat 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 apparatus 1007—for example, the cruise heat transfer ratio may be less than or equal to 0.40 and possibly less than or equal to 0.38, and possibly less than or equal to 0.10 or 0.05.In implementations where there is no refrigeration cycle apparatus 1007, or where such apparatus is present but not used, a higher value of the cruise heat transfer ratio may be maintained, e.g., greater than 0.38, and possibly greater than 0.40 or 0.50.

[0293] The fuel temperature when it reaches the combustion chamber 16 may also be taken into account. The method 300 may comprise controlling 302 the heat exchange system 3000 under cruising conditions such that the heat transfer ratio is in the range of 0 to 0.2, provided that the fuel temperature at the inlet of the combustion chamber 16 is at least 160°C, or in the range of 0 to 0.1 provided that the fuel temperature at the inlet of the combustion chamber 16 is at least 180°C. One or more temperature sensors may therefore be used, and their output(s) may be taken into account when adjusting the control 302 of the heat exchange system 3000.

[0294] The method 300 may also take into account the nature of the fuel - for example, whether the fuel is or comprises sustainable aviation fuel (SAF). For example, the cruise heat transfer ratio may be maintained in the range from 0 to 0.2 provided that the fuel is at least 70% SAF, or from 0 to 0.1 provided that the fuel is at least 80% SAF. One or more fuel characteristics - possibly determined by a fuel characteristic determination module 57 as described above - may therefore also be used; the information may be taken into account when adjusting the control 302 of the heat exchange system 3000.

[0295] The oil-to-air heat transfer rate at cruise (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 possibly 0 to 120 kJ / kg. In some implementations, no more than 20% of the heat transferred from the oil in cruise regime can be transferred to air and the oil-to-air heat transfer rate can be maintained in the range of 0 to 100 kJ per kilogram of fuel at cruise conditions.

[0296] The oil-to-fuel heat transfer rate at cruise (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 possibly 85 to 170 kJ / kg. In some cases, at least 80% of the heat transferred from the oil at cruise may be transferred to the fuel and the oil-to-fuel heat transfer rate may be maintained in the range of 110 to 240 kJ per kilogram of fuel at cruise conditions.

[0297] In some implementations, the heat exchange system 3000 further includes branched fuel return paths 6020 and at least one valve controlling a fuel flow split, the branched paths being for returning fuel from the heat exchange system 3000 to at least two different locations along a main fuel path, from the fuel inlet to 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 fuel flow along each branch based on the heat transfer ratio and / or the fuel temperature at the outlet of the heat exchanger 1004, 1006.

[0298] The percentages of the total heat transferred from the oil in the heat exchange system 3000 that are transferred to the fuel (as opposed to being transferred to the air / environment) at cruise conditions for a particular geared engine 10 are shown as a function of the fuel temperature limit in [Fig. 19]. The fuel temperature limit is the highest temperature deemed safe for aircraft operation for that fuel and engine 10, and may depend on fuel characteristics, such as thermal stability, as well as the heat tolerance of engine components.Graph 3 represents a lower limit 4 (black dotted line) and an upper limit 5 (black dotted line) for the % heat transferred to the fuel - the upper and lower limits 4, 5 take into account different fuel flows in cruise (which depend on, among other things, altitude, rated power, etc.), variability in engine heat generation, aircraft fuel temperature, atmospheric temperature and other variables. Generally speaking, for the example engine 10 for which this data is provided, for a fuel with a temperature limit of 120 °C, the lower limit might be 20% of the heat transferred to the fuel, with an upper limit of all heat. lost by the oil being transferred to the fuel (100%). Graph 3 also represents a line 6 (solid gray line) for a typical heat transfer to the fuel with a fuel temperature limit. A typical value of % heat transferred to the fuel for a fuel with a temperature limit of 120 °C might be 60%; this value can be increased to 100% for fuel temperature limits of 170 °C and above. If the fuel temperature limit is 260 °C or higher, all the heat can be transferred to the fuel at cruise for the entire duration of the cruise, without any need for air cooling, resulting in a heat transfer ratio of zero.

[0299] Turning now to the method 400 to be carried out at idle, the method 400 is again intended to be carried out in a geared gas turbine engine 10 comprising an oil loop system 2000 for supplying oil to the reduction gear 30; and a heat exchange system 3000 comprising an air-oil heat exchanger 2020 through which the oil in the oil loop system flows; and a fuel-oil heat exchanger 1006 through which the oil in the oil loop system and the fuel flow such that heat is transferred between the oil and the fuel. The oil loop system 2000 branches, as shown in Figures 9 and 11, such that a proportion of the oil can flow along each branch, with the air-to-oil and fuel-to-oil heat exchangers 1006, 2020 arranged in a parallel configuration on different branches of the oil loop system.A modulating valve 2016 is further provided to regulate the proportion of oil sent through each branch of the oil loop system 2000. The method 400 includes controlling 402 the heat exchange system 3000 such that, when the aircraft 1 is operating at idle conditions, a heat transfer ratio of: oil to air heat transfer rate (kJ kg') oil to fuel heat transfer rate (kJkg') is in the range of 0.67 to 5.67. The ratio may therefore be higher at idle than at cruise. The ratio may be higher at ground idle than at flight idle. Optionally, the method 400 may include maintaining the heat transfer ratio at 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 under idle conditions above 4.0, 4.5, or 5.0 under ground idle conditions.

[0300] The method 400 may further comprise receiving data 404 to enable the heat transfer ratio to be calculated or derived, for example, temperature data (of the oil and / or fuel, at one or more points around an oil loop system 2000, 2000' or a flow path of fuel, and / or possibly a fuel tank temperature or an oil tank temperature), and fuel flow rate data. This data may be received 404 by a controller 58 and used 406 to adjust the modulation 302 of the heat exchange system 3000 at idle, so as to maintain the heat transfer ratio at a desired level or within 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 fuel or oil temperature or flow rate, or a change in aircraft operation, such as the start of taxiing). These steps 404, 406 may alternatively be considered part of the control 402 of the heat exchange system 3000.The method 400 may also be intended to use other information, for example temperature data (of the oil, fuel and / or ambient temperature of an environment around the aircraft 1), flow rate data (of the oil and / or fuel) and / or one or more fuel characteristics, in order to determine 406 the control actions to be taken.

[0301] The control step 402 of the heat exchange system 3000 may comprise any one, or any combination, of the examples provided for the modulation step 204 of the method 200 of [Fig. 13], for example, by increasing the amount of oil sent through 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 sent through a respective bypass line 1005 or recirculation line 6011.

[0302] In implementations where the engine 10 includes a refrigeration cycle apparatus 1007, the step 402 of controlling the heat exchange system 3000 at idle may include using the refrigeration cycle apparatus 1007 to provide thermal lift by transferring more heat from the oil to the fuel beyond that transferred when passing through the one or more fuel-oil heat 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 obtained than in implementations without a 1007 refrigeration cycle apparatus - for example, the idle heat transfer ratio may be less than or equal to 4 and optionally less than or equal to 4.0, 3.5, 3.0, 2.5, 2.0, 1.5 or 1.0, and optionally less than or equal to 0.10 or 0.05.In implementations where there is no refrigeration cycle apparatus 1007, or where such apparatus is present but not used, a higher value of the heat transfer ratio may be maintained, for example greater than 3.37, and possibly greater than 3.4, 3.5, 4.0, 4.5, 5.0 or 5.5.

[0303] The fuel temperature when it reaches the combustion chamber 16 can also be taken into account. The method 400 may include controlling 402 the heat exchange system 3000 during idle operations such that the heat transfer ratio is in the range of 2.33 to 5.67 provided that the fuel temperature at the inlet of the combustion chamber 16 is less than 200°C, or in the range of 0.67 to 4, provided that the fuel temperature at the inlet of the combustion chamber 16 is equal to or greater than 200°C, or in the range of 0.67 to 2.67 when the fuel temperature at the inlet of the combustion chamber 16 is equal to or greater than 250°C, or 0.67 to 1.22 when the fuel temperature at the inlet of the combustion chamber 16 is equal to or greater than 280°C.One or more temperature sensors may therefore be used, and their output(s) may be taken into account when adjusting the control 302 of the heat exchange system 3000.

[0304] The method 400 may, additionally or alternatively, take into account the nature of the fuel - for example, whether the fuel is or comprises sustainable aviation fuel (SAF). For example, the idle heat transfer ratio may be maintained in the range of 0.67 to 3.67, provided that the fuel is at least 60% or 70% sustainable aviation fuel, or 0.67 to 2.67, provided that the fuel is at least 75%, 80% or 85% sustainable aviation fuel.

[0305] One or more fuel characteristics - possibly determined by a fuel characteristics determination module 57 as described above - can therefore also be used; the information can be taken into account when adjusting the control 402 of the heat exchange system 3000.

[0306] The oil-to-air heat transfer rate at idle (measured in kJ per kilogram of fuel flow reaching the combustion chamber 16) may be in the range of 380 to 1430 / kg, and possibly in the range of 380 to 1270 kJ / kg, or 500 to 1080 kJ / kg. In some implementations, no more than 80% of the heat transferred from the oil at idle may be transferred 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 at idle conditions. At idle, a greater % heat rejection to the air is generally used than at cruise and other higher engine power conditions.

[0307] The oil-to-fuel heat transfer rate 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 possibly 190 to 760 kJ / kg. In some cases, at least 20% of the heat transferred from the oil at idle may be transferred to the fuel and the oil-to-fuel heat transfer rate may be maintained in the range of 150 to 360 kJ per kilogram of fuel at idle conditions.

[0308] As with the method 300 described in [Fig. 14], in some implementations, the heat exchange system 3000 further includes branched fuel return paths and at least one valve controlling a fuel flow split, the branched paths being for returning fuel from the heat exchange system 3000 to at least two different locations along a main fuel path, from the fuel inlet to 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 fuel flow along each branch based on the heat transfer ratio and / or the fuel temperature at the outlet of the heat exchanger 1004, 1006.

[0309] [Fig. 16] illustrates an oil recirculation system comprising two loops 2000, 2000' but without heat exchangers arranged on parallel branches, as illustrated in [Fig. 11]. Instead, each oil loop 2000, 2000' provides a series of main oil flow paths through all the heat exchangers in that loop. To adjust the oil flow through the heat exchangers, one or more bypass lines 2005, 2005', 2005a are provided instead of the bypass paths to different heat exchangers.It will be appreciated that while a bypass line technically provides an alternative parallel flow path, a branch path as described herein with reference to parallel flows is different in that each branch route of the parallel arrangement includes a heat exchanger, whereas the bypass line is simply a line without a (significant) heat exchanger or other component thereon.

[0310] It will be appreciated that a combination of (i) parallel arrangements of heat exchangers and (ii) bypass lines may be used in some implementations, and that Figures 11 (both loops have parallel branches, no bypass lines) and 16 (all heat exchangers in parallel, multiple bypass lines) may be considered to illustrate the two different ends of a design spectrum.

[0311] In the example illustrated in [Fig. 16], the fuel flow is represented by a thick black line to provide context for the interaction between the fuel and oil systems.

[0312] The primary oil loop system 2000 illustrated in [Fig. 16] provides series oil flow paths from a reservoir 2002, through a pump oil 2004, then through the primary fuel-oil heat exchanger 1006 and then the air-oil heat exchanger 2020, before entering the reduction gear 30 (and possibly other components to be cooled and lubricated), then being collected in a sump 2008 and then pumped back into the reservoir 2002 by the oil pump 2010. The two heat exchangers 1006, 2020 are therefore in a series arrangement. The order of the two heat exchangers 1006, 2020 can be reversed such that the fuel-oil heat exchanger 1006 is after the air-oil heat exchanger 2020 in alternative implementations.

[0313] The primary oil loop system 2000 illustrated in [Fig. 16] includes two oil bypass lines 2005, 2005a. The first bypass line 2005 is for allowing a portion of the oil to bypass the fuel-oil heat exchanger 1006 and is controlled by a first bypass valve 2007. The first bypass line 2005 takes oil upstream of the inlet of the primary heat exchanger 1006 and returns it to the main oil flow path downstream of the primary heat exchanger 1006 and before the air-oil heat exchanger 2020. The second bypass line 2005a is for allowing a portion of the oil to bypass the air-oil heat exchanger 2020 and is controlled by a second bypass valve 2007a.The second bypass line 2005a takes oil upstream of the inlet of 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 reducer 30 (and possibly other components to be cooled and lubricated). In implementations of the primary oil loop system 2000 with a single oil bypass line 2005a, the position chosen may be that of the second bypass line 2005a, such that there is a bypass for the air-to-oil heat exchanger 2020 and not for the fuel-to-oil heat exchanger 1006. This may facilitate transferring the greatest amount of heat safely from the oil to the fuel, and maintaining high engine thermal efficiency by reducing heat loss to the environment.The presence of a bypass line 2005 on the fuel-oil heat exchanger 1006 can facilitate rapid adjustment of the heat transfer ratio if there is a risk of it falling lower than desired (for example, depending on determined fuel characteristics, a lower limit for the ratio significantly above zero can be set). The temperature limits set can be strict for certain fuels, depending on their characteristics. The presence of a bypass line 2005, 2005a on both heat exchangers 1006, 2020 can prevent the oil from becoming too cold under certain conditions, to avoid a risk of . Excessive oil freezing - keep in mind that this may be more of a concern when starting or idling on the ground in cold weather than when cruising.

[0314] The secondary oil loop system 2000' illustrated in [Fig. 16] includes a single heat exchanger 1004, which is the secondary fuel-oil heat exchanger, and a single bypass line 2005' for allowing a portion of the oil to bypass that heat exchanger 1004. In other implementations, no bypass line may be provided on the secondary oil loop 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 line may be provided, possibly one for each heat exchanger. A bypass valve 2007' is still provided to regulate the flow of oil through the bypass line 2005'.

[0315] The inventors were aware that the use of fuels different from traditional kerosene-based jet fuels, such as sustainable aviation fuels, may lead to different fuel properties, and that operating heat transfer parameters may be adjusted to utilize the different fuel properties. In particular, in implementations that lack parallel arrangements of heat exchangers (such as those used in the methods described with respect to Figures 14 and 15), the introduction of one or more oil bypass lines with controllable valves, and the implementation of careful control of these valves, may allow a method providing improved oil cooling (as the fuel may be able to absorb more heat) and may also improve the overall thermal efficiency of the engine, with less heat being lost to the environment.The controllable heat exchange system 3000, and in particular one or more controllable bypass valves 2007, 2007', 2007a, have a key role to play in managing the heat transfer ratio in such engines 10. .

[0316] Furthermore, although cruise conditions generally represent a much larger proportion of an aircraft engine's operating time, the inventors were aware that idling 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 increase - therefore, the use of non-traditional fuels may have an even greater effect on optimal approaches to heat management under idle conditions. The methods 500, 600 of Figures 17 and 18 address both of these aircraft operating scenarios.

[0317] [Fig. 17] illustrates a method 500 implementing these considerations at cruise conditions, and [Fig. 18] illustrates a method 600 implementing these considerations at idle, when the aircraft 1 is on the ground (ground idle), for example, when starting the aircraft, operating it while stationary during boarding, and taxiing (to a runway or hangar, or between other ground locations), or at certain times during flight (in-flight idle).

[0318] [Fig. 17] illustrates the method 500 to be performed in cruise mode. The method 500 is intended to be performed in a geared gas turbine engine 10 comprising an oil loop system 2000 for supplying oil to the reduction gear 30; and a heat exchange system 3000 comprising an air-oil heat exchanger 2020 through which the oil in the oil loop system flows; and a fuel-oil heat exchanger 1006 through which the oil in the oil loop system and the fuel flow such that heat is transferred between the oil and the fuel. The oil loop system 2000 further comprises at least one oil bypass line 2005, 2005a, as shown in [Fig. 16], such that a proportion of the oil can bypass at least one heat exchanger.A bypass control valve 2007, 2007a is also provided for controlling the proportion of oil sent through the bypass line 2005, 2005a, and thus also the amount of oil sent through the respective heat exchanger 1006, 2020. The method 500 comprises controlling 502 the one or more bypass valves 2007, 2007a such that, under cruising conditions, a heat transfer ratio of: . heat transfer rate, oil to air (kJkg1) heat transfer rate from oil to fuel (kJkg'J) 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 control device 58 - which may be either a stand-alone unit or part of an EEC - may be provided to implement this control.

[0319] The method 500 may further comprise receiving data 504 to enable the heat transfer ratio to be calculated or derived, for example, temperature data (of the oil and / or fuel, at one or more points around an oil loop system 2000, 2000' or a fuel flow path, and / or possibly a fuel tank temperature or an oil tank temperature), and fuel and / or oil flow rate data. Such data may be received 504 by a controller 58 and used 506 to adjust the control 502 of the bypass valve 2007, 2007a in cruise mode, so as to maintain the heat transfer ratio at a desired level or within desired limits. This checking and adjustment / this 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 may also be considered as part of the control 502 of the bypass valve 2007, 2007a. The method 500 may also be intended to use other information, for example, temperature data (of the oil, fuel, and / or ambient temperature of an environment around the aircraft 1), flow rate data (of the oil and / or fuel), and / or one or more fuel characteristics, to determine 506 the control actions to be taken.

[0320] In implementations where a bypass line 2005a is provided through the air-oil heat exchanger 2020, the controlled bypass valve may be the valve 2007a for the bypass line 2005a through the air-oil heat exchanger 2020. The step 502 of controlling the bypass valve 2007a to adjust the heat transfer ratio may include decreasing the amount of oil sent through the air-oil heat exchanger 2020 when the heat transfer ratio is too high. In some implementations, this bypass line 2007a may be the only oil bypass line in the primary oil loop system 2000. No bypass line may be provided for the fuel-oil heat exchanger 1006.

[0321] In alternative implementations such as that illustrated in [Fig. 16], the heat exchange system 3000 includes at least two oil bypass lines, and possibly three or more, each bypass line 2005, 2005', 2005a being to allow oil to bypass one of the heat exchangers 1006, 1004, 2020. The method 500 may include modulating the amount of oil sent through each bypass line 2005, 2005', 2005a. In embodiments with multiple bypass lines in the same closed-loop oil system (e.g., with two bypass lines in the primary oil loop system 2000, as illustrated in [Fig.16]), the same bypass valve 2007 - which may be a three-way valve - may be used to regulate the flow in both bypass lines 2005, 2005a, or a different bypass valve 2007, 2007a may be provided for each of them, as shown in [Fig.16].

[0322] It will be borne in mind that the control of the one or more bypass valves 2007, 2007a is not the only action that can affect the heat transfer ratio - the use of bypass lines can, however, be used to provide rapid corrections and has a key role to play in maintaining the ratio within the desired limits. Other components of the heat exchange system - for example, a recirculation valve 6010 and / or a refrigeration cycle apparatus 1007 - may also be controlled, and their control may influence the heat transfer ratio. For example, the heat exchange system 3000 may include at least one recirculation line 6011 for allowing a fluid to pass multiple times through a heat exchanger. The method 500 in such implementations may further include modulating the amount of the fluid sent through the recirculation line 6011, by controlling the respective valve 6010, so as to adjust the heat transfer ratio. Additionally or alternatively, the heat exchange system 3000 may include a refrigeration cycle apparatus 1007, and the method 500 may further include using the refrigeration cycle apparatus 1007 to provide thermal elevation by transferring more heat from the oil to the fuel.

[0323] In implementations with a refrigeration cycle apparatus 1007, the bypass valve 2007, 2007a may be controlled such that the heat transfer ratio is in the range of 0 to 0.40. In implementations without a refrigeration cycle apparatus, in which the heat exchange system 3000 is not intended to provide thermal lift, the bypass valve may be controlled such that the heat transfer ratio is generally higher, possibly in the range of 0.38 to 0.67.

[0324] In some implementations, one or more fuel temperature sensors may be provided, and data relating to the fuel temperature at the inlet of the combustion chamber 16 may be used to refine control decisions. The method 500 may include controlling 502 the bypass valve 2007, 2007a under cruise conditions such that the heat transfer ratio is in the range of 0 to 0.2, provided that the fuel temperature at the inlet of 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, provided that the fuel temperature at the inlet of the combustion chamber 16 is at least 180°C.

[0325] In some implementations, one or more fuel characteristics may be considered when determining how to control the bypass valve 2007, 2007a (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, provided that the fuel is at least 70% sustainable aviation fuel; and / or such that the heat transfer ratio is in the range of 0 to 0.1, provided that the fuel is at least 80% sustainable aviation fuel.

[0326] The method 500 may include controlling 502 the bypass valve 2007, 2007a under cruise conditions such that the heat transfer rate from the oil to the air may be maintained in the range of 0 to 240 kJ per kilogram of fuel at cruise conditions, with no more than 20% of the heat transferred from the oil at cruise being transferred to the air, and / or such that the heat transfer rate from the oil to the fuel is maintained in the range of 85 to 350 kJ per kilogram of fuel at cruise conditions, with at least 80% of the heat transferred from the oil at cruise being transferred to the fuel.

[0327] In some implementations, the heat exchange system 3000 further includes branched fuel return paths 6020 and at least one valve controlling 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, branched fuel return paths 6020 may be provided for either or both of the heat exchangers 1004, 1006. The method 500 may include adjusting the fuel flow along each branch based on the heat transfer ratio and / or the fuel temperature at the outlet of the heat exchanger 1004, 1006.

[0328] [Fig. 18] illustrates method 600 to be performed at idle. Method 600 is intended to be performed in a geared gas turbine engine 10, as described above for method 500 of [Fig. 17] - it will be appreciated that the same engine 10 can be used to implement method 500 of [Fig. 17] at cruise and method 600 of [Fig. 18] at idle.

[0329] The method 600 carried out at idle includes the control 602 of the bypass valve 2007, 2007a such that, under idle conditions, a heat transfer ratio of: oil to air heat transfer rate (kJkgk oil to fuel heat transfer rate (kJkg') is in the range of 0.67 to 5.67, and optionally 0.67 to 5.50, 0.67 to 5.00, 0.67 to 4.50, 0.67 to 4.00, 0.67 to 3.50, 0.67 to 3.50, or 0.67 to 2.50. The ratio may therefore be higher at idle than at cruise. The ratio may be higher at ground idle than at flight idle. Optionally, the method 600 may include maintaining the heat transfer ratio at ground idle conditions below 5.50, 5.0, 4.5, 4.0, 3.5, or 3.0. Optionally, the method 400 may include maintaining the heat transfer ratio at idle conditions above 4.0, 4.5 or 5.0 at ground idle conditions.

[0330] The method 600 may include controlling the bypass valve such that, under idle conditions, the heat transfer ratio is greater than 1.0, and possibly greater than 1.5 or 2.0. A controller 58—which may be either a stand-alone unit or part of an EEC—may be provided to implement this control.

[0331] The method 600 may further comprise receiving 604 data to enable the heat transfer ratio to be calculated or derived, for example, temperature data (of the oil and / or fuel, at one or more points around an oil loop system 2000, 2000' or along a fuel flow path, and / or possibly a fuel tank temperature or an oil tank temperature), and fuel and / or oil flow rate data. Such data may be received 604 by a controller 58 and used 606 to adjust control 602 of the bypass valve 2007, 2007a at idle, so as to 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 taxiing). These steps 604, 606 may also be considered part of the control 602 of the bypass valve 2007, 2007a. The method 600 may also be intended to use other information, e.g., temperature data (of the oil, fuel, and / or ambient temperature of an environment around the aircraft 1), flow rate data (of the oil and / or fuel), and / or one or more fuel characteristics, to determine 606 the control actions to be taken.

[0332] Step 602 of controlling the one or more bypass valves 2007, 2007a to adjust the heat transfer ratio may include decreasing the amount of oil sent through the air-oil heat exchanger 2020 through its respective bypass line 2005a when the heat transfer ratio is too high. The heat exchange system 3000 of some implementations, such as that illustrated in [Fig. 16], includes at least two oil bypass lines 2005, 2005a, each bypass line being for allowing oil to bypass the air-to-oil heat exchanger 2020 and the fuel-to-oil heat exchanger 1006. As with the process 500 performed at cruise, the process 600 performed at idle may include modulating the amount of oil sent through each bypass line 2005, 2005a.The presence of a bypass line 2005, 2005a on the two heat exchangers 1006, 2020 of the primary oil loop 2000 can prevent the oil from becoming too cold under certain conditions, to avoid a risk of excessive freezing of . oil - it will be borne in mind that this may be of greater concern at startup or ground idle in cold weather than at cruise. The oil loop system 2000 may therefore also include one or more oil temperature sensors, and the method 600 being described may include receiving inputs from the one or more oil temperature sensors, and possibly also from one or more temperature sensors for sensing an ambient temperature around the aircraft 1, and adjusting the control 602 of the one or more bypass valves based on this temperature information - for example, in cold weather, less oil may be sent through the air-oil heat exchanger 2020 to reduce heat loss to the environment.

[0333] Other bypass lines 2005' may still be provided in some implementations, for example for the flow of oil in the primary oil loop system 2000' through an oil-to-oil heat exchanger 2030, and / or through one or more heat exchangers in a secondary oil loop system 2000', for example through a secondary fuel-to-oil heat exchanger 1004.

[0334] The method 600 of some implementations includes controlling other components in addition to the one or more bypass valves 2007, 2007a so as to achieve the desired heat transfer ratio. For example, at least one recirculation line 6011 for allowing a fluid to pass multiple times through a heat exchanger and / or a refrigeration cycle apparatus 1007 may be provided, and the method may further include modulating the amount of the fluid sent through the respective additional component(s), or activating or deactivating the respective component(s) as applicable, to adjust the heat transfer ratio. In implementations with a refrigeration cycle apparatus 1007, the bypass valve 2007, 2007a may be controlled such that the heat transfer ratio is in the range of 0.67 to 4.0.in implementations without a refrigeration cycle apparatus, in which implementations the heat exchange system 3000 is not intended to provide thermal lift, the bypass valve 2007, 2007a may be controlled such that the heat transfer ratio is generally higher, possibly in the range of 2.00 to 5.67, and possibly 3.37 to 5.67.

[0335] In some implementations, one or more temperature sensors may be provided, and data relating to the fuel temperature at the inlet to the combustion chamber 16 may be used to refine 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: (i) in the range of 2.33 to 5.67 when the fuel temperature at the combustion chamber inlet is below 200°C; (ii) is in the range of 0.67 to 5.00, and optionally 0.67 to 4.00, when the fuel temperature at the inlet of the combustion chamber (16) is above 200°C; (iii) is in the range of 0.67 to 4.00, and optionally 0.67 to 2.67, when the fuel temperature at the inlet of 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, when the fuel temperature at the inlet of the combustion chamber (16) is above 280°C.

[0336] One or more fuel characteristics—optionally determined by a fuel characteristic determination module 57 as described above—may also be used; the information may be taken into account when adjusting the control 602 of the bypass valve 2007, 2007a. For example, the method 600 may take into account the SAF content of the fuel. The method 600 may include controlling 602 of the bypass valve 2007, 2007a such that the heat transfer ratio is in the range of 0.67 to 3.67, provided that the fuel is at least 70% sustainable aviation fuel, and / or in the range of 0.67 to 2.67, provided that the fuel is at least 80% sustainable aviation fuel.

[0337] The method 600 may include controlling 602 the bypass valve 2007, 2007a under idle conditions such that the heat transfer rate from the oil to the air may be maintained in the range of 350 to 1270 kJ per kilogram of fuel under idle conditions, with no more than 20% of the heat transferred from the oil at ground idle being transferred to the air, and / or such that the heat transfer rate from the oil to the fuel is maintained in the range of 350 to 1270 kJ per kilogram of fuel at idle, with at least 80% of the heat transferred from the oil at idle being transferred to the fuel.

[0338] In some implementations, the heat exchange system 3000 further includes branched fuel return paths 6020 and at least one valve controlling a split of the fuel flow, as described above.

[0339] It will be understood that the invention is not limited to the embodiments described above and that various modifications and improvements may be made without departing from the concepts described herein. Except where mutually exclusive, any feature may be employed separately or in combination with other features and the description extends to and includes all combinations and subcombinations of one or more features described herein.

Claims

Claims

1. A method (100) of operating a gas turbine engine (10) of an aircraft, the gas turbine engine (10) comprising: an engine core (11) comprising a turbine (19), a compressor (14), a combustion chamber (16) for burning fuel, and a core shaft (26) connecting the turbine to the compressor; a fan (23) located upstream of the engine core (11); a reduction gear (30) which receives an input from the core shaft (26) and outputs a drive to the fan (23); an oil loop system (2000) for supplying oil to the reduction gear (30); and a heat exchange system (3000) comprising: an air-oil heat exchanger (2020) through which oil in the oil loop system flows; and a fuel-oil heat exchanger (1006) through which the oil in the oil loop system and the fuel flow such that heat is transferred between the oil and the fuel;and a modulation valve (2016) for varying the proportion of oil sent through each heat exchanger, the method (100) comprising: determining (102) at least one fuel characteristic of the fuel to be combusted by the combustion chamber (16); and controlling (104) the modulation valve (2016) based on the at least one fuel characteristic so as to adjust the proportion of oil sent through each heat exchanger (1006, 2020) to cruising conditions.;

2. The method (100) of claim 1, wherein: the oil loop system (2000) branches such that a proportion of the oil can flow along each branch and the air-oil and fuel-oil heat exchangers (1006, 2020) are arranged in a parallel configuration on different branches of the oil loop system; and the modulating valve (2016) is adapted to allow the proportion of the oil sent through each branch to be varied, and wherein the modulating valve control (2016) adjusts the proportion of oil sent through each branch to cruising conditions.

3. The method (100) of claim 1, wherein the at least one characteristic of the fuel comprises at least one of: i. a percentage of sustainable aviation fuel in the fuel; ii. a concentration of heteroatomic species of the fuel; iii. an aromatic hydrocarbon content of the fuel; iv. a multi-aromatic hydrocarbon content of the fuel; v. a percentage of nitrogen-containing species in the fuel; vi. a presence or percentage of a tracer species or trace element in the fuel; vii. a hydrogen-to-carbon ratio of the fuel; viii. a hydrocarbon distribution of the fuel; ix. a level of non-volatile particulate matter emissions upon combustion; x. a naphthalene content of the fuel; xi. a sulfur content of the fuel; xii. a cycloparaffin content of the fuel; xiii. an oxygen content of the fuel; xiv. a thermal stability of the fuel; xv.a fuel coking level; xvi. an indication that the fuel is a fossil fuel; xvii. at least one of a density, a viscosity, a calorific value and a heat capacity.

4. The method (100) of claim 1, wherein determining (102) at least one fuel characteristic of the fuel comprises: obtaining at least one fuel characteristic of any fuel already present in a fuel tank (50) before refueling, the fuel tank (50) being configured to supply fuel to the combustion chamber (16) via the heat exchange system (3000); determining at least one fuel characteristic of a fuel added to the fuel tank (50) upon refueling; and calculating at least one fuel characteristic of the resulting fuel in the tank (50) after refueling.

5. The method (100) of claim 1, wherein the determination (102) of the at least one fuel characteristic is performed based on a detection of at least one fuel property, or based on received fuel composition data.

6. The method (100) of claim 1, wherein at least one fuel characteristic is derived from performance of the gas turbine engine (10) during at least one of taxiing, takeoff, and climb of the aircraft.

7. The method (100) of claim 1, wherein the heat exchange system (3000) further comprises a refrigeration cycle apparatus (1007), and the method (100) comprises using the refrigeration cycle apparatus to provide thermal elevation by transferring more heat from the oil to the fuel such that the fuel temperature is raised above the oil temperature.

8. The method (100) of claim 1, wherein the heat exchange system (3000) further comprises branched fuel return paths and at least one valve controlling a fuel flow split, the branched paths being for returning fuel from the heat exchange system to at least two different locations along a main fuel path from where the fuel enters the gas turbine engine (10) to the combustion chamber (16).

9. The method (100) of claim 1, wherein, subject to the suitability of the at least one determined fuel characteristic, the method (100) comprises passing all of the oil through the fuel-to-oil heat exchanger for at least one continuous period of at least 30 minutes at cruise, such that no heat is lost to the environment through the air-to-oil heat exchanger (2020) during at least some periods of cruise operation.

10. The method (100) of claim 1, wherein, subject to the suitability of the at least one determined fuel characteristic, the method (100) comprises transferring at least 80% of the heat transferred from the oil at cruise to the fuel.

11. The method (100) of claim 1, wherein the at least one fuel characteristic is or comprises thermal stability, and in which at least 80% of the heat transferred from the oil at cruising speed is transferred to the fuel provided that the fuel is stable in operation at temperatures above 160°C.

12. The method (100) of claim 1, wherein the at least one fuel characteristic is or comprises an aromatic hydrocarbon content of the fuel, and wherein at least 80% of the heat transferred from the oil at cruise is transferred to the fuel provided that the fuel has a mole percentage of aromatic hydrocarbons below 12%.

13. / V. The method (100) of claim 1, wherein the at least one fuel characteristic is or comprises a percentage of sustainable aviation fuel - S AF - in the fuel, and wherein at least 80% of the heat transferred from the oil at cruise is transferred to the fuel provided the fuel has a SAF content above 50%.

14. The method (100) of claim 1, wherein the at least one fuel characteristic is or comprises a calorific value of the fuel, and wherein at least 80% of the heat transferred from the oil at cruise is transferred to the fuel provided that the fuel has a calorific value of at least 43.5 MJ / kg.

15. A gas turbine engine (10) for an aircraft comprising: an engine core (11) including a turbine (19), a compressor (14), and a core shaft (26) connecting the turbine to the compressor; a fan (23) located upstream of the engine core; and a reduction gear (30) which receives an input from the core shaft and outputs a drive to the fan so as to drive the fan at a rotational speed lower than that of the core shaft; an oil loop system (2000) for supplying oil to the reduction gear (30); and a heat exchange system (3000) comprising: an air-oil heat exchanger (2020) through which the oil in the oil loop system flows; and a fuel-oil heat exchanger (1006) through which the oil in the oil loop system and the fuel flow such that heat is transferred between the oil and the fuel; and a modulation valve (2016) intended to allow the proportion of oil sent through each heat exchanger to be varied; and a fuel composition determining module (57) for determining at least one fuel characteristic of the fuel to be burned by the combustion chamber (16), wherein the modulation valve (2016) is to be controlled based on the at least one fuel characteristic so as to adjust the proportion of oil sent through each heat exchanger (1006, 2020) to cruising conditions.

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