Gas turbine fuel temperature

Abstract

Gas Turbine Fuel Temperature A method of operating a gas turbine engine of an aircraft, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, a combustor for burning a fuel, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core; a fan shaft; at least one bearing for supporting the fan shaft; at least one auxiliary system for utilizing a portion of the fuel; an oil loop system for supplying oil to the at least one bearing; and a heat exchange system comprising: a primary fuel-oil heat exchanger through which the oil in the oil loop system and at least substantially all of the fuel flow such that heat is transferred between the oil and the fuel;and a secondary fuel-oil heat exchanger through which the oil in the oil loop system and a portion of the fuel flow, such that heat is transferred between the oil and the fuel, the secondary fuel-oil heat exchanger being for supplying fuel to the at least one auxiliary system; and a fuel pump for pumping the fuel, wherein the fuel pump is located downstream of the primary fuel-oil heat exchanger and upstream of the secondary fuel-oil heat exchanger along a fuel flow path; wherein the method further comprises controlling the heat exchange system such that, under cruising conditions, the fuel temperature at the outlet of the second heat exchanger is in the range of 120°C to 200°C. Figure for abstract: Fig. 16;

Description

Title of the invention: Gas turbine fuel temperature

[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 fuel, and a core shaft connecting the turbine to the compressor; a blower located upstream of the engine core; a blower shaft; at least one bearing intended to support the fan shaft; an oil loop system for supplying oil to the at least one bearing; 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; a bypass line for allowing a proportion of the oil to flow past one of the air-to-oil heat exchanger and the fuel-to-oil heat exchanger; and a bypass valve for varying the proportion of oil sent through the bypass line, the method comprising: determining at least one fuel characteristic of the fuel to be burned by the combustion chamber; and controlling the bypass 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 may 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, may therefore provide more efficient oil cooling while avoiding coking or other pathways for degradation product deposition (e.g., 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 thermodynamically 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 transitional time frame, where available aviation fuels change over the lifetime of a gas turbine engine, and also depend on the geographic 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 full advantage of a wide variety of fuels. The controllable modulating valve, designed to adjust the oil flow through each heat exchanger, plays a key role 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 air-oil heat exchanger and the fuel-oil heat exchanger may be arranged in series along the path of the oil loop system. The bypass line may be arranged such that oil passes through the fuel-oil heat exchanger while flowing past the air-oil heat exchanger or vice versa.

[0010] The bypass valve allows variation in the amount of heat transferred from the oil by the air-oil and fuel-oil heat exchangers. 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 regulation of the fuel temperature within a defined upper and lower limit. Similarly, the oil flow regulation can allow the oil temperature to be maintained within an appropriate range.This regulation may help 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 bypass valve may be intended to allow 100% of the oil to be sent through the fuel-oil heat exchanger. The bypass 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 bypass valve may be adapted to divert a fixed portion of the oil through the bypass line to bypass one of the fuel-oil heat exchanger and the air-oil exchanger during engine operation, the fixed portion optionally being determined at or before engine start, or at or before cruising altitude, based on at least one determined characteristic. of fuel. Alternatively, the bypass valve may be intended to divert a variable portion of the oil flow through the bypass line to bypass one of the fuel-to-oil heat exchanger and the air-to-oil heat exchanger during engine operation. The modulating valve may therefore be actively controlled to vary the proportion of oil sent through each heat exchanger, particularly in implementations where the aircraft carries multiple different fuels in different tanks and may change the fuel (or fuel blend) used during flight. Active control of the modulating valve may be automated and implemented by a heat exchange system controller.

[0012] Alternatively, 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 heat exchange system may therefore further comprise a modulating valve for allowing 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.

[0013] The heat exchange system may further include a secondary fuel-oil heat exchanger. The secondary fuel-oil heat exchanger may be a slave fuel-oil heat exchanger. Some, but not all, of the fuel exiting the primary fuel-oil heat exchanger may be directed to the slave fuel-oil heat exchanger. The slave fuel-oil heat exchanger may further increase the temperature of the fuel, before providing fuel for use in servo mechanisms of the gas turbine engine (e.g., for fuel hydraulic actuation and / or heating). These servo mechanisms may include a nacelle anti-icing system. The servo mechanisms may include engine actuators. The servo mechanisms may include a servo turbine case cooling (TCC) valve.Only fuel that has passed through the secondary fuel-oil heat exchanger can be used in these auxiliary systems. Fuel used in these auxiliary systems can be returned to a fuel tank for further recirculation, or be recombined with another fuel. exiting the primary fuel-oil heat exchanger and entering the combustion chamber. Fuel passing through the secondary fuel-oil heat exchanger may therefore not be supplied to the combustion chamber in some implementations, but instead returned to an aircraft fuel tank, possibly after use in auxiliary systems such as fuel-operated hydraulic actuators. The slave fuel-oil heat exchanger may be structurally similar to, or identical to, the primary fuel-oil heat exchanger. The slave fuel-oil heat exchanger may be smaller than the primary fuel-oil heat exchanger. At least some of the fuel may not pass through the secondary fuel-oil heat exchanger.

[0014] The heat exchange system may further include at least one additional bypass line for allowing oil (or fuel) 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 bypass valve, or other oil valve, may be adapted to adjust the amount of oil sent through the or each oil bypass line, based on the one or more determined fuel characteristics, and possibly on one or more temperature measurements.

[0015] 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.

[0016] 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, with 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 fuel branch 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, for example further downstream, for example after one or more engine components located downstream of the fuel-oil heat exchanger in question, or even further upstream (thus acting as a recirculation line).

[0017] Subject to the suitability of the at least one determined fuel characteristic, the method may include passing all of the oil through the fuel-oil heat exchanger for at least one or more continuous periods of time of at least 30 minutes in cruise operation, such that no heat is lost to the environment through the air-oil heat exchanger during at least some periods of cruise operation. 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 in cruise operation, such that very little heat, if any, is lost to the environment through the air-oil heat exchanger during 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.

[0018] 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 140°C.

[0019] 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%.

[0020] 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%.

[0021] The at least one fuel characteristic may be or include a heating value of the fuel. At least 80% of the heat transferred from the oil in cruising speed can be transferred to the fuel, provided that the fuel has a calorific value of at least 43.5 MJ / kg.

[0022] The at least one fuel characteristic of the fuel may comprise 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 in the fuel; iv. a multi-aromatic hydrocarbon content in 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 distribution of hydrocarbons in the fuel; ix. a level of emissions of non-volatile particulate matter 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.

[0023] 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 one or more of a shaft speed, a turbofan power ratio (TPR) and / or an engine pressure ratio (EPR) 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 include obtaining stored fuel characteristic data. Chemical and / or physical determination of one or more fuel parameters in the fuel tank may be performed by extracting a sample. on-board fuel from the fuel tank for off-site testing. Extracting a sample of the on-board fuel from the fuel tank may include taking a sample of the fuel prior to refueling.

[0024] 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).

[0025] 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.

[0026] 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.

[0027] According to a second aspect, there is provided a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core; and a reducer which receives an input from the core shaft and provides a drive to the blower so as to drive the blower at a lower rotational speed than 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; a bypass line for allowing a proportion of the oil to flow past one of the air-to-oil heat exchanger and the fuel-to-oil heat exchanger; a bypass valve for varying the proportion of oil sent through the bypass line; 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 bypass 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.

[0028] 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.

[0029] The core shaft can provide drive to the fan directly, so as to drive the fan at the same rotational speed as the core shaft, such that the engine is a direct-drive turbine engine.

[0030] 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.

[0031] 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 combustion chamber for burning fuel, and a core shaft connecting the turbine to the compressor; a blower located upstream of the engine core; a blower shaft; at least one bearing intended to support the fan shaft; an oil loop system for supplying oil to the at least one bearing; 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 bypass line for allowing a proportion of the oil to flow past one of the air-to-oil heat exchanger or the fuel-to-oil heat exchanger; and a bypass valve for permitting variation of the proportion of oil sent through the bypass line, the method comprising controlling the heat exchange system such that, under cruising conditions, a heat transfer ratio of:

[0032] [Math.l] oil to air heat transfer rate (kJkg'!) oil to fuel heat transfer rate {kJkg ) is in the range from 0 to 0.67.

[0033] 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 fuel degradation product deposition (e.g., varnishing). This may allow for a process that provides improved oil cooling (because the fuel is able to absorb more heat) and may also improve the overall thermodynamic 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.

[0034] While it is understood that the ratio is dimensionless, heat transfer is measured per mass 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 volume of fuel" can be equated with "per a fixed amount of time as a function of 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 relative to a 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. as a function of a fuel temperature at the approach to or entry into the combustion chamber as compared with a fuel temperature in a fuel tank of the aircraft. For comparison purposes, the amount of heat transferred from the oil to the air may be determined from a temperature drop of the oil across the air-oil heat exchanger(s), or by comparing the heat gained by the fuel to an overall heat loss from the oil, and assuming that the difference is due to heat loss from the oil to the air (including any other source of loss or generation, such as fuel recirculation).

[0035] 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.

[0036] 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.

[0037] The at least one bypass line may be adapted to allow oil to bypass the air-to-oil heat exchanger, such that oil flowing through the bypass line flows through the fuel-to-oil heat exchanger without flowing through the air-to-oil heat exchanger. Controlling the heat exchange system to adjust the heat transfer ratio may include increasing the amount of oil flowing through the bypass line when the heat transfer ratio is too high.

[0038] 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.

[0039] The heat exchange system may further include a secondary fuel-oil heat exchanger. The secondary fuel-oil heat exchanger may be a slave fuel-oil heat exchanger. Some, but not all, of the fuel exiting the primary fuel-oil heat exchanger may be directed to the slave fuel-oil heat exchanger. The slave fuel-oil heat exchanger may further increase the temperature of the fuel, before providing fuel for use in servo mechanisms of the gas turbine engine (e.g., for fuel hydraulic actuation and / or heating). These servo mechanisms may include an anti-icing system. of nacelle. The servo mechanisms may include engine actuators. The servo mechanisms may include a servo-controlled turbine case cooling (TCC) valve. Only fuel that has passed through the secondary fuel-oil heat exchanger may be used in these auxiliary systems. The fuel used in these auxiliary systems may be returned to a fuel tank for further recirculation, or may be recombined with other fuel exiting the primary fuel-oil heat exchanger and entering the combustion chamber. The fuel passing through the secondary fuel-oil heat exchanger may therefore not be supplied to the combustion chamber in some implementations, but instead returned to an aircraft fuel tank, possibly after use in auxiliary systems such as fuel-hydraulic actuators.The slave fuel-oil heat exchanger may be structurally similar to, or identical to, the primary fuel-oil heat exchanger. The slave fuel-oil heat exchanger may be smaller than the primary fuel-oil heat exchanger. At least some of the fuel may not pass through the secondary fuel-oil heat exchanger.

[0040] 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 examples using a refrigeration cycle apparatus may be in the range of 0 to 0.40.

[0041] 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.

[0042] 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.

[0043] 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.

[0044] 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 0.45, provided that the fuel temperature at the inlet to the combustion chamber is at least 140°C.

[0045] 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.

[0046] 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.

[0047] The method may include maintaining the oil-to-air heat transfer rate at cruise in the range of 0 to 100 kJ per kilogram of fuel at cruise conditions, and optionally 0 to 35 kJ / kg, with no more than 20% of the heat transferred from the oil at cruise being transferred to the air.

[0048] The method may include maintaining the oil-to-fuel heat transfer rate at cruise in the range of 110 to 200 kJ per kilogram of fuel at cruise conditions, and optionally 150 to 200 kJ / kg, with at least 80% of the heat transferred from the oil at cruise being transferred to the fuel.

[0049] The methods of the first and third aspects may be complementary, and they may be performed together in various implementations. The method of the third aspect may be performed using the engine of the second aspect.

[0050] 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, and a core shaft connecting the turbine to the compressor; a blower located upstream of the engine core; a blower shaft; at least one bearing intended to support the fan 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; a bypass line for allowing a proportion of the oil to flow past one of the air-to-oil heat exchanger or the fuel-to-oil heat exchanger; and a bypass valve intended to allow the proportion of oil sent through the bypass line to be varied, and wherein the heat exchange system is intended to be controlled such that, under cruising conditions, a heat transfer ratio of:

[0051] [Math.2] heat transfer rate, oil to air (klkg'1) heat transfer rate, oil to fuel (k J kg*) is in the range from 0 to 0.67.

[0052] 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.

[0053] 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.

[0054] The bypass line may be adapted to allow a proportion of the oil to flow past the air-to-oil heat exchanger, such that oil flowing through the bypass line flows through the fuel-to-oil heat exchanger without flowing through the air-to-oil heat exchanger.

[0055] The core shaft can provide drive to the fan directly, so as to drive the fan at the same rotational speed as the core shaft, such that the engine is a direct-drive turbine engine.

[0056] The air-oil heat exchanger and the fuel-oil heat exchanger may be arranged in a series configuration on the oil loop system.

[0057] Alternatively, 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-oil and fuel-oil heat exchangers may be arranged in a parallel configuration on different branches of an oil loop system.

[0058] 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 any preceding aspect.

[0059] 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 blower shaft; at least one bearing intended to support the fan shaft; an oil loop system for supplying oil to the at least one bearing; 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 bypass line for allowing a proportion of the oil to flow past one of the air-to-oil heat exchanger or the fuel-to-oil heat exchanger; and a bypass valve intended to allow the proportion of oil sent through the bypass line to be varied, the method comprising controlling the heat exchange system such that, under idle conditions, a heat transfer ratio of:

[0060] [Math.3] oil to air heat transfer rate (kJkg4) oil to fuel heat transfer rate (kJkg is in the range from 0 to 1.5.

[0061] 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 under idle 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 the deposition of fuel degradation products (e.g., varnishing). This can allow for a process that provides improved oil cooling (as the fuel is able to absorb more heat) and can also improve the overall thermodynamic 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.

[0062] 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 idling conditions - for example, when starting the aircraft, operating it while stationary during boarding, taxiing (to a runway or hangar, or between other ground locations), or during "flight idle" conditions such as during descent.Since the operating conditions are very different between cruise and idle - at least in terms of the desired engine thrust output - the heat exchange system must be controlled differently.

[0063] While it is understood that the ratio is dimensionless, heat transfer is measured per unit mass of fuel in the examples described herein, thereby providing a normalized heat transfer rate for variations in fuel flow rate at idle. For heat transfer from oil to air, the definition of "per unit mass of fuel" can be equated with "per a fixed amount 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 appreciated that 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 may therefore be calculated based on a fuel temperature at the approach or entry to the combustion chamber by comparison with a fuel temperature in a fuel tank of the aircraft. For comparison purposes, the amount of heat transferred . from oil to air can be determined from a temperature drop of the oil 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 that the difference is due to heat loss from the oil to the air (after inclusion of any other sources of loss or generation being included, such as fuel recirculation or at the fuel pump).

[0064] The method may include controlling the heat exchange system such that, under idle conditions, the heat transfer ratio is below 1.0, in the range of 0 to 0.60, 0 to 0.50, 0 to 0.40, 0 to 0.30, 0 to 0.20, or 0 to 0.10.

[0065] 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.

[0066] The at least one bypass line may be adapted to allow oil to bypass the air-to-oil heat exchanger, such that oil flowing through the bypass line flows through the fuel-to-oil heat exchanger without flowing through the air-to-oil heat exchanger. Controlling the heat exchange system to adjust the heat transfer ratio may include increasing the amount of oil flowing through the bypass line when the heat transfer ratio is too high.

[0067] 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.

[0068] The heat exchange system may further include a secondary fuel-oil heat exchanger. The secondary fuel-oil heat exchanger may be a slave fuel-oil heat exchanger. Some, but not all, of the fuel exiting the primary fuel-oil heat exchanger may be directed to the slave fuel-oil heat exchanger. The slave fuel-oil heat exchanger may further increase the temperature of the fuel, before providing fuel for use in servo mechanisms of the gas turbine engine (e.g., for fuel hydraulic actuation and / or heating). These servo mechanisms may include a nacelle anti-icing system. The servo mechanisms may include engine actuators. The servo mechanisms may include a servo valve Turbine case cooling (TCC). Only fuel that has passed through the secondary fuel-oil heat exchanger can be used in these auxiliary systems. The fuel used in these auxiliary systems may be returned to a fuel tank for further recirculation, or be recombined with other fuel exiting the primary fuel-oil heat exchanger and entering the combustion chamber. The fuel passing through the secondary fuel-oil heat exchanger may therefore not be supplied to the combustion chamber in some implementations, but instead returned to an aircraft fuel tank, possibly after use in auxiliary systems such as fuel-operated hydraulic actuators. The slave fuel-oil heat exchanger may be structurally similar to, or identical to, the primary fuel-oil heat exchanger.The slave fuel-oil heat exchanger may be smaller than the primary fuel-oil heat exchanger. At least some of the fuel may not pass through the secondary fuel-oil heat exchanger.

[0069] 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.

[0070] When the heat exchange system is not intended to provide thermal elevation, the heat exchange system may be controlled such that the heat transfer ratio is in the range of 0.38 to 1.2.

[0071] The method may comprise controlling the heat exchange system under idle conditions such that the heat transfer ratio is in the range of 0.3 to 1.5, provided that the fuel temperature at the inlet to the combustion chamber is below 180°C.

[0072] The method may comprise controlling the heat exchange system under idle conditions such that the heat transfer ratio is in the range of 0 to 0.3, provided that the fuel temperature at the inlet to the combustion chamber is above 180°C.

[0073] Under idle conditions, the method may include controlling the heat exchange system such that the heat transfer ratio is in the range of 0 to 0.2, provided that the fuel is at least 70% sustainable aviation fuel.

[0074] The method may include maintaining the oil-to-air heat transfer rate at idle in the range of 0 to 180 kJ per kilogram of fuel at idling conditions, and possibly from 0 to 60 kJ / kg, with no more than 20% of the heat transferred from the oil at idle being transferred to the air.

[0075] The method may include maintaining the oil-to-fuel heat transfer rate at idle in the range of 100 to 300 kJ per kilogram of fuel at idle conditions, and optionally 200 to 300 kJ / kg, with at least 80% of the heat transferred from the oil at idle being transferred to the fuel.

[0076] The air-oil heat exchanger and the fuel-oil heat exchanger may be arranged in series in the oil loop system.

[0077] Alternatively, 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 heat exchange system may therefore further comprise a modulating valve for allowing 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.

[0078] The heat exchange system may 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.

[0079] The methods of the first, third, and fifth aspects may be complementary, and they 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.

[0080] 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 blower located upstream of the engine core; a blower shaft; at least one bearing intended to support the fan 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; a bypass line for allowing a proportion of the oil to flow past one of the air-to-oil heat exchanger or the fuel-to-oil heat exchanger; and a bypass valve intended to allow the proportion of oil sent through the bypass line to be varied, and wherein the heat exchange system is intended to be controlled such that, under idle conditions, a heat transfer ratio of:

[0081] [Math.4] oil to air heat transfer rate CkJkgj oil to fuel heat transfer rate (kJkg1) is in the range from 0 to 1.5.

[0082] 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.

[0083] 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.

[0084] The bypass line may be adapted to allow a proportion of the oil to flow past the air-to-oil heat exchanger, such that oil flowing through the bypass line flows through the fuel-to-oil heat exchanger without flowing through the air-to-oil heat exchanger.

[0085] The core shaft can provide drive to the fan directly, so as to drive the fan at the same rotational speed as the core shaft, such that the engine is a direct-drive turbine engine.

[0086] The air-oil heat exchanger and the fuel-oil heat exchanger may be arranged in a series configuration in the oil loop system.

[0087] Alternatively, 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-oil and fuel-oil heat exchangers may be arranged in a parallel configuration on different branches of an oil loop system.

[0088] 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.

[0089] According to a seventh 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 blower shaft; at least one bearing intended to support the fan shaft; at least one auxiliary system intended to use part of the fuel; an oil loop system for supplying oil to the at least one bearing; and a heat exchange system comprising: a primary fuel-oil heat exchanger through which the oil in the oil loop system and at least substantially all of the fuel flow such that heat is transferred between the oil and the fuel; and a secondary fuel-oil heat exchanger through which the oil in the oil loop system and a portion of the fuel flow, such that heat is transferred between the oil and the fuel, the secondary fuel-oil heat exchanger being for supplying fuel to the at least one auxiliary system; and a fuel pump for pumping fuel, wherein the fuel pump is located downstream of the primary fuel-oil heat exchanger and upstream of the secondary fuel-oil heat exchanger along a fuel flow path; the method further comprises controlling the heat exchange system such that, under cruising conditions, the fuel temperature at the outlet of the secondary heat exchanger is in the range of 120°C to 200°C.

[0090] 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 under idle 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 may allow for a process that provides 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 controllable heat exchange system has a key role to play in managing the heat transfer ratio.

[0091] The secondary fuel-oil heat exchanger may be a slave fuel-oil heat exchanger. Some, but not all, of the fuel exiting the primary fuel-oil heat exchanger may be directed to the slave fuel-oil heat exchanger. The slave fuel-oil heat exchanger may further increase the temperature of the fuel, before providing fuel for use in servo mechanisms of the gas turbine engine (e.g., for fuel hydraulic actuation and / or heating). These servo mechanisms may include a nacelle anti-icing system. The servo mechanisms may include engine actuators. The servo mechanisms may include a servo turbine case cooling (TCC) valve. Only fuel that has passed through the secondary fuel-oil heat exchanger may be used in these auxiliary systems.The fuel used in these auxiliary systems may be returned to a fuel tank for further recirculation, or may be recombined with other fuel exiting the primary fuel-oil heat exchanger and entering the combustion chamber. The fuel passing through the secondary fuel-oil heat exchanger may therefore not be supplied to the combustion chamber in some implementations, but instead returned to an aircraft fuel tank, possibly after use in auxiliary systems such as fuel-operated hydraulic actuators. The slave fuel-oil heat exchanger may be structurally similar to, or identical to, the primary fuel-oil heat exchanger. The slave fuel-oil heat exchanger may be smaller than the primary fuel-oil heat exchanger. At least some of the fuel may not pass through the secondary fuel-to-oil heat exchanger.

[0092] The method may include controlling the heat exchange system such that, under cruising conditions, the fuel temperature at the outlet of the secondary heat exchanger is in the range of 120°C to 180°C.

[0093] The method may include controlling the heat exchange system such that, under cruising conditions, the fuel temperature at the outlet of the secondary heat exchanger is in the range of 135°C to 200°C.

[0094] The method may include controlling the heat exchange system such that, under cruising conditions, the fuel temperature at the outlet of the secondary heat exchanger is in the range of 135°C to 180°C.

[0095] The method may include controlling the heat exchange system such that, under cruising conditions, the fuel temperature at the outlet of the secondary heat exchanger is in the range of 150°C to 200°C, 150°C to 180°C, 150°C to 170°C. The fuel temperature at the outlet of the secondary heat exchanger may be approximately 120°C, 130°C, 140°C, 150°C, 160°C, or up to 200°C.

[0096] The heat exchange system may further include a fuel bypass line for allowing a proportion of the fuel to bypass at least one of the primary fuel-oil heat exchanger and the secondary fuel-oil heat exchanger; and a fuel bypass valve for allowing the proportion of the fuel sent through the fuel bypass line to be varied. Controlling the heat exchange system may include controlling the fuel bypass valve to adjust the proportion of the fuel sent through each of the primary and secondary fuel-oil heat exchangers to cruise conditions.

[0097] The heat exchange system may further include an oil bypass line for allowing a proportion of the oil to bypass at least one of the primary fuel-oil heat exchanger and the secondary fuel-oil heat exchanger; and an oil bypass valve for allowing the proportion of the fuel sent through the oil bypass line to be varied. Controlling the heat exchange system may include controlling the oil bypass valve to adjust the proportion of the oil sent through each of the primary and secondary fuel-oil heat exchangers to cruise conditions.

[0098] Under cruising conditions, the ratio of

[0099] [Math.5] fuel flow through secondary fuel-oil heat exchanger (m3^1) fuel flow to combustion chamber (mV) can be at least 0.3.

[0100] This ratio may be referenced as the secondary heat exchanger fuel flow ratio, for brevity.

[0101] The secondary heat exchanger fuel flow ratio may be at least 0.35, 0.4, 0.45, 0.5, 0.55. The secondary heat exchanger fuel flow ratio may be approximately 0.6.

[0102] 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. An oil valve may regulate the amount of oil that flows through the refrigeration cycle apparatus.

[0103] 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 configured to return fuel from the heat exchange system to at least two different locations along a main fuel flow path. 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 line).

[0104] The primary fuel-oil heat exchanger and the secondary fuel-oil heat exchanger may be arranged in series along the oil loop system.

[0105] The heat exchange system may further comprise at least one air-oil heat exchanger through which oil from the oil loop system flows.

[0106] According to an eighth aspect of the invention, there is provided a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core; and a reducer which receives an input from the core shaft and provides a drive to the blower so as to drive the blower at a lower rotational speed than the core shaft; an oil loop system for supplying oil to the reducer; and a heat exchange system comprising: a primary fuel-oil heat exchanger through which the oil in the oil loop system and at least substantially all of the fuel flow such that heat is transferred between the oil and the fuel; and a secondary fuel-oil heat exchanger through which the oil in the oil loop system and a portion of the fuel flow such that heat is transferred between the oil and the fuel, the secondary fuel-oil heat exchanger being for supplying fuel to the at least one auxiliary system; and a fuel pump for pumping fuel, wherein the fuel pump is located downstream of the primary fuel-oil heat exchanger and upstream of the secondary fuel-oil heat exchanger along a fuel flow path; wherein the heat exchange system is intended to be controlled such that, under cruising conditions, the fuel temperature at the outlet of the second heat exchanger is in the range of 120°C to 200°C.

[0107] The heat exchange system may be intended to be controlled such that, under cruising conditions, the fuel temperature at the outlet of the second heat exchanger is in the range of 120°C to 180°C.

[0108] 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.

[0109] The core shaft can provide drive to the fan directly, so as to drive the fan at the same rotational speed as the core shaft, such that the engine is a direct-drive turbine engine.

[0110] 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.

[0111] 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 blower shaft; at least one bearing intended to support the fan shaft; an oil loop system for supplying oil to the at least one bearing; 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 at least one valve for varying at least one of an oil flow rate and an air flow rate through at least one of the fuel-oil heat exchanger and the air-oil heat exchanger; and a temperature sensor for providing an indication of fuel temperature downstream of the fuel-oil heat exchanger (optionally at the inlet to the combustion chamber), the method comprising: determining whether the fuel temperature has increased above a set threshold at cruise conditions, based on an output from the temperature sensor; and in response to determining that the fuel temperature has increased above the threshold set at cruise conditions, controlling the at least one valve so as to change the at least one flow rate through the at least one heat exchanger.

[0112] For example, the at least one valve may be or include a valve for varying a proportion of the oil sent through the fuel-oil heat exchanger, and the method may include, in response to determining that the fuel temperature has increased above a threshold set at cruise conditions, controlling the at least one valve to send less oil through the fuel-oil heat exchanger.

[0113] The inventors were aware that it is important that the oil temperature remains within a desired range as well as the fuel temperature not exceeding a limit, so in some implementations the focal point for flow adjustment may be the air-oil heat exchanger. The at least one valve may therefore be or comprise a valve for allowing a proportion of the oil sent through the air-oil heat exchanger to be varied, and the method may comprise, in response to determining that the fuel temperature has increased above a threshold set at cruising conditions, controlling the at least one valve so as to send more oil through the air-oil heat exchanger.

[0114] A lack of an oil bypass on the air-oil heat exchanger could result in overcooling of the oil (e.g., freezing) under certain conditions. A bypass line may therefore be provided for both heat exchangers. At least some of the oil may bypass both heat exchangers in some implementations.

[0115] The at least one valve may be or include a valve for regulating a flow rate of air through the air-oil heat exchanger. The method may include sending more air through the air-oil heat exchanger in response to determining that the fuel temperature has increased above the threshold set at cruise conditions.

[0116] The method may include controlling multiple valves in some implementations, for example: • regulating oil flow through each of the fuel-oil heat exchanger and the air-oil heat exchanger separately; and / or • regulation of both air flow and oil flow.

[0117] As described for the preceding aspects, 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 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 deposition of fuel degradation products, for example by coking or varnishing, allowing the fuel to absorb more heat and thereby reducing a need for air-cooling of the oil and improving engine efficiency.However, allowing inappropriate fuels to reach these higher temperatures could be detrimental to engine performance, and even cause fuel injection nozzle blockages in some cases. Checks and balances are therefore desired to ensure that engine performance is optimized for a given fuel. The method described in this seventh aspect involves monitoring fuel temperature to control any excessively high fuel temperatures, and taking steps to reduce fuel heating if necessary. Fuel temperatures downstream of the fuel-oil heat exchanger (e.g., at the combustion chamber inlet) at cruise conditions may be defined as an average over at least 1, 2, 3, 4, or 5 minutes, and optionally over ten minutes, twenty minutes, or thirty minutes, under 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.A transient spike – for example, a spike lasting only a few seconds – at a higher temperature may therefore not be sufficient to trigger a change in oil flow.

[0118] In a transitional time frame, 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 that the level of fuel heating is decided based on the specific fuel being used. Sending less oil through the fuel-oil heat exchanger results in less heat being transferred to the fuel, thus providing a lower fuel temperature as it approaches the combustion chamber. Thus, the operation of gas turbines can be adapted to take best advantage of a wide variety of fuels. The controllable oil flow valve for adjusting the oil flow through the fuel-oil heat exchanger has a key role to play in this adaptation of engine performance.

[0119] The inventors were aware that these principles can be applied both to engines with branched oil loop paths with different heat exchangers on different branches (parallel arrangements) and to engines with series, essentially linear arrangements of heat exchangers and where the use of one or more bypass lines can be made as an alternative to a branched main path.

[0120] The oil loop system may branch such that a proportion of the oil may flow along each branch, and the air-oil and fuel-oil heat exchangers may be arranged in a parallel configuration on different branches of the oil loop system. In such implementations, the at least one valve for varying the proportion of oil sent through the fuel-oil heat exchanger may be or comprise a modulation valve intended to allow the proportion of oil sent through each branch to be varied.

[0121] The oil loop system may include at least one oil bypass line for allowing a proportion of the oil to bypass at least one of the fuel-oil heat exchanger and the air-oil heat exchanger. In such implementations, the at least one valve may be or include at least one bypass valve for allowing a proportion of the oil to bypass the respective heat exchanger / for regulating the proportion of oil sent via the bypass line. In such implementations, the air-oil and fuel-oil heat exchangers may be arranged in series in the oil loop system, or in parallel.In implementations where the air-oil and fuel-oil heat exchangers are arranged in parallel, on different branches of the oil loop system, and in which there is additionally at least one bypass line, the method may comprise controlling both the bypass valve and a modulating valve to allow the proportion of oil sent through each branch to be varied.

[0122] In implementations where the method is intended to adjust the flow rate of oil through the fuel-oil heat exchanger, at least a portion of the oil diverted from the fuel-oil heat exchanger may be sent to the air-oil heat exchanger. In some implementations, the air flow through the air-oil heat exchanger may be increased when more oil is sent through the air-oil heat exchanger.

[0123] In implementations where the method is intended to adjust the flow rate of oil through the fuel-oil heat exchanger or the air-oil heat exchanger, at least a portion of the oil diverted from the respective heat exchanger may be sent via a bypass line around the respective heat exchanger.

[0124] The method may further comprise determining the set threshold based on at least one fuel characteristic of the fuel. In such implementations, the at least one fuel characteristic of the fuel may be or comprise at least one of: a thermal stability of the fuel, a nitrogen content of the fuel, a sulfur content of the fuel, and a sustainable aviation fuel - SAF - content of the fuel.

[0125] The step of determining the set threshold may comprise increasing the set threshold, possibly linearly, with increasing thermal stability of the fuel.

[0126] The step of determining the set threshold may comprise increasing the set threshold, possibly linearly, with increasing SAF content of the fuel, for fuels with SAF content above 70%.

[0127] The method may further comprise determining the at least one fuel characteristic of the fuel, optionally by any of the methods described below.

[0128] The heat exchange system may include an air valve for regulating a flow of air through the air-oil heat exchanger. The method may further include, in response to determining that the fuel temperature has increased beyond a threshold set at cruise conditions, controlling the air valve to send more air through the air-oil heat exchanger.

[0129] The heat exchange system may include a refrigeration cycle apparatus for providing thermal lift by transferring additional heat from the oil to the fuel beyond that transferred by the fuel-oil heat exchanger. The method may further include controlling the refrigeration cycle apparatus to reduce the amount of additional heat transferred to the fuel in response to determining that the fuel temperature has increased beyond a threshold set at cruise conditions. For example, the refrigeration cycle apparatus may be turned off / disabled.

[0130] The heat exchange system may include at least one bypass line, and the at least one valve may be or include a bypass valve for regulating a flow rate through that bypass line. The heat exchange system may include multiple bypass lines, each configured to allow oil to bypass a heat exchanger (e.g., the fuel-to-oil heat exchanger or air-to-oil heat exchanger described above, or a secondary fuel-to-oil heat exchanger, an oil-to-oil heat exchanger for transferring heat between two separate oil loops of the heat exchange system, or any other suitable heat exchanger).The method may include controlling at least two bypass valves - for example, a bypass valve for the air-oil heat exchanger may be adjusted to send more oil to the air-oil heat exchanger when the bypass valve for the fuel-oil heat exchanger is adjusted to send less oil to the fuel-oil heat exchanger. The same valve (e.g., a three-way valve) may adjust an oil flow rate to both heat exchangers in some implementations.

[0131] The set threshold may be in the range of 140°C to 300°C, and optionally 250°C to 300°C. The set threshold may be 140°C, 180°C, 200°C, 210°C, 220°C, 230°C, 240°C, 250°C, 260°C, 270°C, 275°C, 280°C, 290°C, or 300°C. The determining step may be performed at regular intervals during operation of the aircraft in cruise.

[0132] 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.

[0133] According to an eighth 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 blower shaft; at least one bearing intended to support the fan shaft; an oil loop system for supplying oil to the at least one bearing; 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 at least one valve for varying at least one of an oil flow rate and an air flow rate through at least one of the fuel-oil heat exchanger and the air-oil heat exchanger; a temperature sensor for providing an indication of fuel temperature downstream of the fuel-oil heat exchanger (possibly at the inlet to the combustion chamber); and a controller for receiving an output from the temperature sensor at cruise conditions, determining whether the fuel temperature has increased above a set threshold based on that output, and in response to determining that the fuel temperature has increased above a set threshold at cruise conditions, controlling the at least one valve to change the at least one flow rate through the at least one heat exchanger.

[0134] For example, the at least one valve may be intended to allow the proportion of oil sent through the fuel-oil heat exchanger to be varied; and the control device may be intended to: receive an output from the temperature sensor at cruising conditions; determine whether the fuel temperature has increased above a set threshold based on this output; and in response to determining that the fuel temperature has increased above a set threshold at cruise conditions, controlling the at least one valve to send less oil through the fuel-oil heat exchanger.

[0135] The controller, or other processing module, may be configured to determine the set threshold based on one or more fuel characteristics of the fuel.

[0136] The heat exchange system may include a refrigeration cycle apparatus for providing thermal elevation by transferring more heat from the oil to the fuel beyond that transferred by the fuel-oil heat exchanger. Optionally, the refrigeration apparatus may allow the fuel temperature to be raised above the oil temperature. The controller may be configured to deactivate the refrigeration cycle apparatus in response to determining that the fuel temperature has increased above a set threshold at cruise conditions.

[0137] 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.

[0138] 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.

[0139] The core shaft can provide drive to the fan directly, so as to drive the fan at the same rotational speed as the core shaft, such that the engine is a direct-drive turbine engine.

[0140] The air-oil and fuel-oil heat exchangers may be arranged in series in the oil loop system.

[0141] 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.

[0142] 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. Any such engine may or may not be provided with an afterburner. Such a gas turbine engine may be, for example, designed for land or marine power generation applications.

[0143] 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).

[0144] 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 rotational axis. Alternatively, the gas turbine engine may include a propeller stage and a guide vane stage arranged 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- and intermediate-pressure propellers and compressors via suitable interconnecting shafts. Thus, the propellers may provide the majority of the propulsive thrust.

[0145] 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.

[0146] 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 a such arrangement, the second turbine may be positioned axially upstream of the first turbine.

[0147] 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 provides 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).

[0148] 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 include a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, second compressor, and second core shaft may be configured to rotate at a higher rotational speed than the first core shaft.

[0149] 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.

[0150] The reducer may be intended to be driven by the core shaft which 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 arranged to be driven only by the core shaft which 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.

[0151] The reducer may be a reduction box (in that the output to the blower has a lower rotational speed than the input from the core shaft). Any what 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 hence be referred to as a "compound star" reducer), for example with two reduction stages.

[0152] 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.

[0153] 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 flow at the outlet to the combustor may be supplied at the inlet of the second turbine, when a second turbine is provided. The combustor may be provided upstream of the turbine(s).

[0154] 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. compressor (or “high pressure”). 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.

[0155] 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.

[0156] 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.

[0157] 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.

[0158] The rotational speed of the fan 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 3500 rpm, for example less than 2600 rpm, or less than 2500 rpm, or less than 2300 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 2750 to 2900 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 rotation speed at cruising conditions for a . direct drive having a fan diameter in the range of 300 cm to 340 cm can be in the range of 2000 to 2800 rpm.

[0159] 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 in 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 preceding 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).

[0160] 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 line 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 line may be substantially annular. The bypass line may be . radially outside the core engine. The radially outer surface of the bypass duct may be defined by a nacelle and / or a fan casing.

[0161] 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.

[0162] 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 of 80 Nkg 's to 100 Nkg 's, or of 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 gas turbine engine at . gears 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.

[0163] 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 of the following: 100 kN, 110 kN, 120 kN, 130 kN, 135 kN, 140 kN, 145 kN, 150 kN, 155 kN, 160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN, 450 kN, 500 kN, or 550 kN. The maximum thrust may be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits). 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 above mentioned thrust may be the maximum net thrust under typical atmospheric conditions at sea level plus 15 degrees C (ambient pressure 101.3 kPa, temperature 30 degrees C), with the engine static.

[0164] 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 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.

[0165] 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, solely by way of non-limiting example, 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 1960 K. Strictly by way of As a 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 of 1800 K to 1950 K, or 1900 K to 2000 K. The maximum TET may occur, for example, at a high thrust condition, for example, at a maximum takeoff thrust (PMD) condition.

[0166] 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 may be fabricated 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 fabricated using different materials.For example, the fan blade may have a protective leading edge, which may be manufactured using a material that is better 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 manufactured 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.

[0167] A fan as described and / or claimed herein may include 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 include 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 integrally formed a central portion. Such an arrangement may be referred to as a bladed disc or 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.

[0168] 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 outlet area to be varied in use. The general principles of the present disclosure may be applied to engines with or without a VAN.

[0169] 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.

[0170] 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.

[0171] 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 can refer to a phase of engine operation where the aircraft is propelled by the 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.

[0172] 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 a flight cycle of an aircraft between takeoff and arrival at cruise conditions, with arrival at cruise conditions thus defining 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 time periods during, an aircraft flight cycle between takeoff and landing, where a relative increase in altitude is required, which may necessitate an additional demand for engine thrust.

[0173] 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 can 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.

[0174] 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, or at least substantially 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.

[0175] 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, e.g., 0.76 to 0.84, e.g., 0.77 to 0.83, e.g., 0.78 to 0.82, e.g., 0.79 to 0.81, e.g., of the order of Mach 0.8, of 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, e.g., below Mach 0.7 or above Mach 0.9.

[0176] 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.

[0177] Strictly by way of example, cruise conditions may correspond to a forward Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35,000 feet (10,668 m). 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 in the range 20 kN to 40 kN.

[0178] Strictly by way of further example, the cruise conditions may correspond to a forward Mach number of 0.85 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38,000 feet (11,582 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 35 kN to 65 kN.

[0179] 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.

[0180] 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 an approach in particular being part of the landing and takeoff (LTO) phase. During either or both of the 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.

[0181] 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.

[0182] 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).

[0183] 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.

[0184] Those skilled in the art would understand 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.

[0185] Except where mutually exclusive, any parameter or value contained or described herein may be applied 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 and / or combined with any one or more other parameters and / or values ​​contained or described herein (e.g., any one or more of a parameter B; a parameter C; and a parameter D, and so on) to express a product of their relationship. For example, one skilled in the art would understand that when 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.

[0186] Embodiments will now be described by way of example only, with reference to the Figures, in which: [Fig.l] is a sectional side view of a gas turbine engine; [Fig.2] is a close-up sectional side view of an upstream portion of a geared gas turbine engine; [Fig.3] is a partially cutaway view of a reduction gear for a gas turbine engine; [Fig.4] is a close-up sectional side view of an upstream portion of a direct-drive gas turbine engine; [Fig.5] is a representation of an aircraft with a propulsion system comprising two gas turbine engines; [Fig.6] is a schematic representation of an exemplary fuel system; [Fig.7A] is a schematic representation of an exemplary alternative fuel system including a recirculation loop; [Fig.7B] is a schematic representation of another example of an alternative fuel system including a recirculation loop and the use of fuel for auxiliary systems; [Fig.7C] is a schematic representation of another example of an alternative fuel system similar to that shown in [Fig.7B] but where fuel passing through the secondary heat exchanger is returned to the combustion chamber possibly after being used in one or more auxiliary systems; [Fig.8A] is a schematic representation of a portion of an exemplary oil recirculation system; [Fig.8B] is a schematic representation of a portion of an exemplary oil recirculation system; [Fig.8C] is a schematic representation of a portion of an exemplary oil recirculation system; [Fig.9] is a schematic representation of part of an exemplary oil recirculation system; [Fig. 10] is a schematic representation of a portion of the exemplary fuel system of [Fig. 6] and the exemplary oil recirculation system of [Fig. 8A]; [Fig. 11] is a schematic representation of a portion of the exemplary fuel system of [Fig.7A] and the exemplary oil recirculation system of [Fig.8A]; [Fig. 12] is a schematic representation of a portion of the exemplary fuel system of [Fig.7A] and the exemplary oil recirculation system of [Fig.8C]; [Fig. 13] illustrates an example of a method of operating a gas turbine engine; [Fig. 14] illustrates another example of a method of operating a gas turbine engine; [Fig. 15] illustrates another example of a method of operating a gas turbine engine; [Fig. 16] illustrates another example of a method of operating a gas turbine engine; [Fig. 17] illustrates another example of a method of operating a gas turbine engine; [Fig. 18] illustrates a cruise heat transfer range graph for a direct-drive gas turbine engine.

[0187] The [Fig. 1] 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 duct 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.

[0188] 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.

[0189] 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 outward 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 by means of links 36 to the fan 23 in order to cause its rotation around the engine axis 9. Radially outwardly of and meshing with the planet gears 32 is a ring or crown gear 38 which is coupled, by means of links 40, to a stationary support structure 24.

[0190] 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.

[0191] 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.

[0192] 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.

[0193] 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 parts of the motor (e.g. between the input and output shafts from the reducer and the fixed 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].

[0194] 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.

[0195] Optionally, the reducer may drive additional and / or alternative components (e.g. the intermediate pressure compressor and / or a booster).

[0196] Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have another number of compressors and / or turbines and / or another 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 be applied to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed-flow nozzle.The nozzle(s) (whether mixed or split flow) may have a fixed or variable area.

[0197] 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. A cross-sectional view of such an engine is shown in Figure 3B.

[0198] Referring to Figure 3B, a gas turbine engine is generally indicated at 10, having a main axis of rotation 9. The engine 10 includes, in axial flow series, an air intake 12, a propulsion fan 23, an intermediate pressure compressor 14, a high pressure compressor 15, a combustion equipment 16, a high pressure turbine 17, an intermediate pressure turbine 19a, a low pressure turbine 19 and an exhaust nozzle 20. A nacelle 21 surrounds the engine 10 and defines both the intake 12 and the exhaust nozzle 20.

[0199] In use, air entering the intake 12 is accelerated by the blower 23 to produce two airflows: a core airflow A and a bypass airflow B. The core airflow A flows into the intermediate pressure compressor 14, and the bypass airflow B passes through a bypass line 22 to provide propulsive thrust. The intermediate pressure compressor 14 compresses the airflow A before delivering this air to the high pressure compressor 15 where further compression takes place.

[0200] 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 combustion chamber 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, intermediate pressure and low pressure turbines 17, 19a, 19 before being discharged through the nozzle 20 to provide additional propulsive thrust. The high pressure 17, intermediate pressure 19a and low pressure 19 turbines respectively drive the high pressure compressor 15, the intermediate pressure compressor 14 and the fan 23, each by a suitable interconnecting shaft.

[0201] 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.

[0202] 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.

[0203] 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 rotational 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.

[0204] 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.

[0205] 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.

[0206] 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.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.

[0207] 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 takeoff fuel 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 aromatic fuels; and, maximum surface tension (mN / m) may ensure sufficient spray breakup and atomization of the fuel.

[0208] 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.

[0209] 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 other compositions being pending approval. However, there is an expectation 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.

[0210] 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 heating value between 101% and 105% of that of kerosene.

[0211] 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 comprise 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 compositional variations.

[0212] 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 conventional hydrocarbon fuels. fossil base (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.

[0213] As shown in [Fig. 5], an aircraft 1 may include multiple 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 the fuel instead being stored 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.

[0214] [Fig. 5] 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 distinct source of fuel, i.e., a first fuel source may contain a first fuel having a different characteristic or characteristics than 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 multiple fuel sources allows an aircraft 1 to carry multiple different fuels and to change the fuel used during operation, and possibly even during cruise or when transitioning from one stage of operation to another during flight.

[0215] 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 fluidically connected in the example shown, 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.

[0216] 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. It may therefore be more complex to determine which fuel is supplied to the combustion chamber 16 than simply recording a single identity of a fuel on board the aircraft 1 or performing a check at the time of start-up.

[0217] 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 asymmetrical 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.

[0218] Aircraft typically refuel at multiple 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 - SAFs), and / or blends of petroleum-based fuels and other fuels) and any additives included (e.g., antioxidants and metal deactivators, biocides, static electricity 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.

[0219] 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 fuel from a single fuel source or a mixture of one or more fuels from different sources - is therefore important in determining the operation of the engine.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.

[0220] 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 particularly the heat exchange 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 (nvPM) 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.

[0221] 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 exchange system 3000. Determining fuel characteristics may include obtaining fuel characteristics. fuel of any fuel already present in the fuel tank 50, 53 before refueling and an indication of the amount of fuel remaining, then combining this information with information regarding new fuel added to the tank 50, 53 during refueling.

[0222] 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.

[0223] In some examples, one or more fuel characteristics may be determined during operation of the gas turbine 10, for example by inferring fuel characteristics of the fuel supplied to the combustor 16 during operation from engine performance measurements, or by performing on-site sensing.

[0224] 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.

[0225] 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 direct detection of the fuel characteristics or the provision of data from which they may be determined, as discussed above, and / or the detection of one or more elements or tracer compounds added to the fuel to facilitate its identification (e.g., a dye); or by receiving data, for example from an input provided in a user interface, or data transmitted to the aircraft, for example by scanning a barcode associated with the fuel delivery.

[0226] 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 a tracer, can be used to determine fuel characteristics such as the percentage of sustainable aviation fuel in the fuel or whether 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 time period (such as during takeoff), and possibly also during a second operating time period (e.g., during cruise), and comparing these collected parameters to the expected values ​​if using a fuel whose properties are known. • Various fuel characteristics, including the aromatic hydrocarbon content of the fuel, can be determined based on measurements of sensor of the presence, absence or degree of formation of a condensation trail by the gas turbine 10 during its 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 intended 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.

[0227] 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 rotor (HPT).

[0228] Any suitable approach known in the art may be used, and the determination of fuel characteristics will not be discussed further herein.

[0229] 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 control functionality 58 may be provided by the EEC and no separate unit or module may be provided.

[0230] An exemplary fuel system 1000 for a direct drive gas turbine engine 10, the fuel system 1000 including a fuel flow path from the fuel tank 50 to the combustion chamber 16 of the gas turbine engine 10 of the aircraft 1 is shown schematically in [Fig. 6]. The fuel system 1000 includes both the fuel supply system 50, 1002 (which supplies fuel to the engine 10) and the fuel management system 1500 (which operates within the engine 10) of the aircraft 1. The fuel management system 1500 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 (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 lines and associated components, e.g. pumps, refrigeration units, etc.

[0231] 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 primary fuel-oil heat exchanger 1006 and a secondary fuel-oil heat exchanger 1004. The primary fuel-oil heat exchanger 1006 may be referred to as the primary fuel-oil heat exchanger because the fuel flowing therethrough may be the primary fuel flow path from the fuel tank 50 to the combustion chamber 16. The secondary fuel-oil heat exchanger 1004 may be referred to as an auxiliary fuel-oil heat exchanger, or a servo fuel-oil heat exchanger because the fuel flowing therethrough may be supplied to the auxiliary systems or servo mechanisms 33 of aircraft 1.Such auxiliary systems / servo mechanisms may include, but are not limited to, fuel hydraulic actuation; a heater; a nacelle anti-ice system; engine actuators; and / or a servo-controlled turbine case cooling (TCC) valve. Only fuel that has passed through the secondary fuel-oil heat exchanger may be used in these auxiliary systems. Fuel used in these auxiliary systems may be returned to a fuel tank for further recirculation, or be recombined with other fuel leaving the primary fuel-oil heat exchanger 1006 and entering the combustion chamber.The fuel passing through the secondary fuel-oil heat exchanger 1004 may therefore not be supplied to the combustion chamber in some implementations, but instead returned to an aircraft fuel tank 50, possibly after use in auxiliary systems such as fuel hydraulic actuators. The fuel passing through the secondary fuel-oil heat exchanger 1004 may instead be returned along the primary fuel flow path between the fuel tank 50 and the engine fuel pump 1003. More or fewer fuel-oil heat exchangers may be provided in other implementations. The illustrated fuel management system 1500 is arranged so that the fuel reaches the primary fuel-oil heat exchanger 1006 before the secondary fuel-oil heat exchanger 1004.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 may be referred to as the primary fuel pump, and in a fuel management system. fuel 1500 it is positioned upstream of the secondary fuel-oil heat exchanger 1004.

[0232] 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.

[0233] In general, at least the majority of the fuel passes through the primary fuel-oil heat exchanger 1006, and at least some of the fuel that passes through the primary fuel-oil heat exchanger 1006 also passes through the secondary fuel-oil heat exchanger 1004. The two heat exchangers 1004, 1006 may therefore be described as being in series with each other, 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 fuel bypass line 1005, as illustrated in [Fig. 6].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, 2005a 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. heat. A valve for regulating the flow of a fluid through a bypass line 1005 may be called a bypass valve.

[0234] 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 same oil flows through one and then the other in the described implementation, although it will be appreciated that different oil loop systems may feed each heat exchanger in other implementations, and thus a different oil may flow through one fuel-oil heat exchanger to the other fuel-oil heat exchanger.

[0235] In the example illustrated in [Fig. 6], the fuel system 1000 further includes a temperature sensor 1009 for sensing the temperature of the fuel approaching or reaching the combustion chamber 16. The temperature sensor 1009 is shown adjacent an inlet of the combustion chamber 16 in the illustrated example, but in various examples, the sensor 1009 may be placed anywhere downstream of the primary fuel-oil heat exchanger 1006 or the secondary fuel-oil heat exchanger on the fuel side, so as to provide a fuel temperature measurement at various points along the flow path.In gas turbine engines 10 such as those described herein, fuel passes through fuel spray nozzles before entering the combustion chamber itself - these nozzles can be very sensitive to thermal degradation of the fuel, leading to blockages in the relatively narrow passages (e.g., metering slots); placement of the sensor 1009 as close as reasonably possible to the nozzle inlet is therefore advisable in some implementations. Multiple temperature sensors 1009 may be used in some implementations. The one or more temperature sensors 1009 are intended to provide an output (temperature data) to a controller 58.

[0236] [Fig.7A] 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 flows through 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, possibly via a secondary fuel-oil heat exchanger 1004. Controlling the fuel flow through the secondary fuel-oil heat exchanger 1004 may be regulated by valve 1004a. 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 for recirculating 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 oil of the primary loop system 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 where 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) line 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 combustion chamber 16. It is contemplated that the recirculation valve could be positioned downstream of the engine fuel pump 1003 in alternative implementations.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 multiple times before reaching the pump 1003. This recirculation provides a mechanism for regulating the flow of fuel within the fuel management system 6500, as well as within the heat exchange 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.

[0237] 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 line 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 control the direction of fluid flow therethrough.

[0238] Regulation of the fuel flow in the fuel system 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.

[0239] The fuel flow control through the recirculation conduit paths 6011 may be based on a fuel temperature measurement (e.g., using a temperature sensor at a location downstream of the primary fuel-oil heat exchanger 1006, possibly also using a temperature measurement upstream of the primary fuel-oil heat exchanger 1006). A return of the recirculated fuel upstream of the primary fuel-oil heat exchanger 1006 may reduce the heat transfer from the oil to the fuel, thereby dampening a transient overshoot that might otherwise occur at the beginning of the downshift phase, for example, when the same amount of heat is generated within the oil circuit, but fuel flow is reduced, generally resulting in a temperature spike.Adjusting fuel flow through a recirculation line 6011, for example, can also be useful in reducing icing, by circulating fuel that has been preheated to warm the system. The recirculation line 6011 can therefore be used in a variety of ways to improve engine thermal management.

[0240] [Fig. 7B] illustrates an exemplary alternative fuel system 7000 to that shown in [Fig. 7A]. The fuel system 7000 is similar to the fuel system 6000 shown in [Fig. 7A] but differs in that at least a portion of the fuel exiting the secondary fuel-oil heat exchanger 1004 is supplied to additional engine and / or aircraft mechanisms 7010, instead of the combustor 16. These mechanisms 7010 may include one or more of a nacelle anti-ice system, actuators, and / or turbine case cooling (TCC) systems. At least a portion of the fuel exiting the secondary fuel-oil heat exchanger 1004 is returned to the fuel supply system (i.e., to the fuel tank 50 or the fuel supply pump 1002) or downstream of the fuel system for redistribution (i.e., upstream of the primary fuel-oil heat exchanger 1006). The portion of fuel returned to / downstream of the fuel supply system may be regulated by a valve 7011. No fuel passing through the secondary fuel-oil heat exchanger 1004 is sent to the combustion chamber 16 in the example of [Fig. 7B] (at least not directly - it will be kept in mind that the fuel returned to the fuel tank 50 may re-enter the engine 10 later, and may then be passed to the combustion chamber 16).

[0241] [Fig. 7C] illustrates another alternative example of fuel system 7000a. Fuel system 7000a is similar to fuel system 7000 illustrated in [Fig. 7B] and differs in that the portion of the fuel exiting the secondary fuel-oil heat exchanger 1004 is returned to rejoin the main fuel flow as it approaches the combustor 16 after possibly being supplied to one or more additional engine and / or aircraft mechanisms 7010 (which may be as described above). Any fuel not needed for these auxiliary systems 7010 may be returned directly to the main fuel flow path / combustor 16.At least a portion of the fuel exiting the secondary fuel-oil heat exchanger 1004 may be returned to the main fuel flow path without passing through the auxiliary systems 7010 - the proportion directed to the auxiliary systems 7010 may be adjusted based on need and regulated with a valve 7011. In other implementations, the fuel used in one or more such additional aircraft mechanisms 7010 may then be returned to a tank 50 rather than sent to the combustion chamber 16.

[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, and illustrated in [Fig.8A], the oil recirculating system includes a single closed oil loop system 2000. In other implementations, the oil recirculating system may include a primary oil loop system 2000 and a secondary oil loop system (not shown), each of which is a closed loop oil system. The secondary oil loop system may supply oil to one or both of the primary fuel-oil heat exchanger 1006 and the secondary fuel-oil heat exchanger and possibly to additional heat exchangers such as an air-oil heat exchanger.

[0243] In the illustrated implementation, the 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.

[0244] A feed pump 2004 is configured to pump oil from the oil reservoir 2002 to the secondary fuel-oil heat exchanger 1006. The average temperature of the oil entering the secondary fuel-oil heat exchanger 1006, at cruising conditions, is higher than the average temperature of fuel entering the secondary fuel-oil heat exchanger 1006. In the secondary fuel-oil heat exchanger 1006, thermal energy is transferred from the oil flow to the fuel flow. In this manner, the average temperature of the oil flow leaving the secondary fuel-oil heat exchanger 1004 is lower than the average temperature of the oil flow entering the secondary fuel-oil heat exchanger 1006.In this manner also, the average temperature of the fuel leaving the secondary fuel-oil heat exchanger 1004 is higher than the average temperature of the fuel entering the secondary fuel-oil heat exchanger 1006.

[0245] From the secondary fuel-oil heat exchanger 1004, the oil then flows to the primary fuel-oil heat exchanger 1006. In some implementations, the oil may flow through an air-oil heat exchanger (not shown) between the secondary fuel-oil heat exchanger 1004 and the primary fuel-oil heat exchanger 1006.

[0246] In the primary fuel-oil heat exchanger 1006, thermal energy is transferred from the oil flow to the fuel flow. In this manner, the average temperature of the oil flow leaving the primary fuel-oil heat exchanger 1006 is lower than the average temperature of the oil flow entering the primary fuel-oil heat exchanger 1006. Also in this manner, the average temperature of the fuel leaving the primary fuel-oil heat exchanger 1006 is higher than the average temperature of the fuel entering the primary fuel-oil heat exchanger 1006.

[0247] From the primary fuel-oil heat exchanger 1006, the oil flow is then routed through the engine components 2006 that the oil is intended to cool and / or lubricate. The oil serves as a lubricant and coolant within the engine components 2006 and helps to gain heat from the engine components 2006 in the process. The engine components 2006 may include one or more bearing chambers.

[0248] From the 2006 engine components, oil is collected in the 2008 crankcase. A scavenge pump 2010 is designed to pump oil from the 2008 crankcase back into the 2002 oil reservoir.

[0249] Figures 8B and 8C illustrate alternative oil loop systems 2000b, 2000c. In the oil loop system 2000b of [Fig.8B], oil flows through an air-to-oil heat exchanger for cooling before the primary fuel-to-oil heat exchanger 1006. The secondary fuel-to-oil heat exchanger is not present in this alternative oil loop system 2000b.

[0250] [Fig.8B] also illustrates details of the air-oil heat exchanger 2020 that are not shown in other figures for clarity, including indicating the air flow with dotted arrows. The air-oil heat exchanger 2020 has an air inlet 2020a and an air outlet 2020b. The air inlet, or intake, 2020a may be intended to capture air in the bypass duct 22 of the engine, downstream of an outlet guide vane of the fan 23. The exhaust / air outlet 2020b of the heat exchanger may be intended to return air to the bypass duct 22, downstream of the inlet 2020a, or directly into the outside atmosphere (for example, if it is provided with a dedicated outlet nozzle).

[0251] A valve 2022, referred to as an air valve, is used to regulate the flow of air through the air-to-oil heat exchanger 2020. In the illustrated example, the air valve 2022 is placed at or near the inlet 2020a of the heat exchanger 2020, and may therefore be referred to as the air inlet valve 2022. In other implementations, the air valve 2022 may be placed at or near the outlet 2020b of the heat exchanger 2020, and may therefore be referred to as the air outlet valve. Any suitable position for the valve 2022 may be selected, provided that the valve 2022 can be adjusted to regulate airflow through the air-oil heat exchanger 2020. In the example being described, the air valve 2022 is continuously adjustable between a fully closed position (no airflow through the heat exchanger) and a fully open position (maximum airflow through the heat exchanger).In other examples, the air valve 2022 may be adjustable between a plurality of distinct positions instead of being continuously adjustable - e.g., six, five, four, three, or two different positions. In implementations with two valve positions for the air valve 2022, these positions may be "open" and "closed" - the valve 2022 may be repeatedly opened and closed to provide a pulsed flow of air when an intermediate level of cooling is desired in some such implementations.

[0252] [Fig.8C] shows an oil recirculation system comprising an oil loop system 2000c. A main oil flow path passes through all of the heat exchangers 1004, 1006, 2020 in this loop. To adjust an oil flow through the heat exchangers, one or more bypass lines 2005, 2005a, 2005b are provided, with bypass valves 2007, 2007a, 2007b used to vary a flow through said bypass lines. A controller 58 may be used to actively manage fuel and / or oil flows through and around the heat exchangers 1004, 1006, 2020. 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.

[0253] In other implementations, branch paths to different heat exchangers could be used instead. 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 a parallel arrangement of heat exchangers includes a heat exchanger, whereas the bypass line is simply a line without a (significant) heat exchanger or other component thereon.

[0254] It will be appreciated that a combination of heat exchanger and bypass line arrangements may be used in some implementations, and that [Fig.8C] (all heat exchangers in series, multiple bypass lines) may be considered to illustrate one end of a design spectrum. In an alternative implementation, one or more of the heat exchangers may be part of a secondary oil loop system. In an alternative implementation, the heat exchangers may be arranged in parallel, whereby the oil loop system 2000 branches into separate branches with two or more separate branches each comprising a heat exchanger.

[0255] The oil loop system 2000 illustrated in [Fig.8C] provides a series oil flow path from a reservoir 2002, through an oil pump 2004, then through the secondary fuel-oil heat exchanger 1004, the primary fuel-oil heat exchanger 1006 and then the air-oil heat exchanger 2020, before entering the engine components 2006 to be cooled and / or lubricated and then being collected in a sump 2008 and then pumped back into the reservoir 2002 by an oil pump 2010. The heat exchangers 1004, 1006, 2020 are thus in a series arrangement. The order of heat exchangers 1004, 1006, 2020 may be reversed such that, for example, primary fuel-oil heat exchanger 1006 is after air-oil heat exchanger 2020 in alternative implementations.

[0256] The oil loop system 2000c illustrated in [Fig.8C] includes three oil bypass lines 2005, 2005a, 2005b. The first bypass line 2005 is for allowing a portion of the oil to bypass the secondary fuel-oil heat exchanger 1004 and is controlled by a first bypass valve 2007. The first bypass line 2005 takes oil upstream of the inlet of the secondary heat exchanger 1004, and returns it to the primary oil flow path downstream of the secondary heat exchanger 1004 and before the primary fuel-oil heat exchanger 1006 and the air-oil heat exchanger 2020. The second bypass line 2005a is for allowing a portion of the oil to bypass the primary fuel-oil heat exchanger 1006 and is controlled by a second bypass valve 2007a.The second bypass line 2005a takes oil upstream of the inlet of the primary fuel-oil heat exchanger 2020, and returns it to the main oil flow path downstream of the primary fuel-oil heat exchanger 2020, before the path reaches the air-oil heat exchanger 2020. The third bypass line 2005b 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 engine components 2006. In implementations of the primary oil loop system 2000 with a single oil bypass line 2005, the chosen position may be that of the third bypass line 2005b, such that there is a bypass for the 2020 air-oil heat exchanger and not for the 1006 fuel-oil heat exchanger.This can facilitate the transfer of the greatest possible amount of heat safely from the oil to the fuel, and the maintenance of the thermal efficiency of the engine at a high level by reducing heat loss to the environment. The presence of a bypass line 2005a on the primary 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, 2020 on all heat exchangers 1004, 1006, 2020 may prevent the oil from becoming too cold under certain conditions, to avoid a risk of the oil freezing - it will be borne in mind that this may be more of a concern under idle conditions (e.g., at start-up or ground idle, or during descent) in cold weather than at cruising speed.

[0257] 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. .

[0258] 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. Methods addressing both of these aircraft operating scenarios are discussed in more detail in connection with Figures 14 and 15.

[0259] 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. 9], 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, with an evaporator between the oil and the refrigerant (to transfer heat from the oil to the refrigerant) and a condenser between the fuel and the refrigerant (to transfer heat from the refrigerant 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 that flows into the refrigeration cycle apparatus 1007, and a fuel valve may regulate the amount of fuel that flows 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, generally 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 both fuel-oil heat exchangers 1004, 1006, and further cooling the oil in the oil system 2000 before it is returned to the engine components 2006. A bypass or recirculation line for the oil and / or for the fuel may be provided around the refrigeration cycle apparatus 1007 in various implementations.

[0260] [Fig. 10] schematically shows an exemplary heat exchange system 3000 including the closed-loop oil system 2000 of [Fig. 8A] and the fuel system 1000 of [Fig. 6]. The two systems 1000, 2000 are shown together to illustrate the interactions between the fuel and oil flows within the gas turbine engine 10. The fuel flow is shown with thick black arrows and the oil flow with thin black arrows.

[0261] [Fig. 11] schematically shows an exemplary alternative heat exchange system 4000, including the closed-loop oil system 2000 of [Fig. 8A] and the fuel delivery system 7000a of [Fig. 7A]. The systems 7000a, 2000 are essentially the same as those illustrated in Figures 8A and 7A respectively, with the fuel flow again shown with thick black arrows and the oil flow with thin black arrows. However, the recirculation valve 6010 is located slightly differently than in [Fig. 7A], namely upstream of the fork in the fuel flow path to the secondary heat exchanger 1006, instead of being between the inlet of 1006a and the outlet of 1006b of this heat exchanger 1006. Thus in the implementation illustrated in [Fig.l 1], all the fuel leaving the pump 1003 flows through the recirculation valve 6010, unlike the embodiment illustrated in [Fig.7A]. The proportion of fuel flowing through the secondary fuel-oil heat exchanger 1004 can be regulated by the valve 1004a. The regulation methodology for the valves 1004a, 6010 can be adjusted appropriately.

[0262] It will be appreciated that in implementations such as that illustrated in [Fig.7A], where fuel flowing through the secondary fuel-oil heat exchanger 1004 is used as a working fluid in auxiliary systems and then returned to a reservoir 50, 53 instead of being passed to the combustion chamber 16, the fuel flow line from the secondary heat exchanger 1006 to the combustion chamber 16 may be replaced with a fuel flow line returned to the reservoir 50.

[0263] The inventors were aware that using fuels different from traditional kerosene-based fuels, such as sustainable aviation fuels, may lead to different fuel properties, and that engine operation may be optimized for these different fuel properties. Fuels having a higher heating value as claimed may also have greater thermal stability, allowing the fuel to absorb more heat, which provides improved oil cooling and / or improved combustion properties in the combustion chamber. Recirculating the fuel through the primary fuel-oil heat exchanger 1006 may allow the fuel to receive more heat from the oil, increasing the fuel temperature and improving oil cooling.

[0264] A recirculation valve 6010 may be located downstream of the fuel pump 1003, and may therefore provide improved fuel flow flexibility. A lower fuel flow rate to the combustion chamber 16 may therefore be provided for higher heating value fuel, without changing the pump speed, but rather by increasing the amount of recirculation.

[0265] [Fig. 12] schematically shows an exemplary alternative heat exchange system 4500, including the closed-loop oil system 2000 of [Fig. 8C] and a portion of the fuel delivery system 7000a of [Fig. 7A]. For clarity, only the fuel flow through the primary fuel-oil heat exchanger 1006 and the secondary fuel-oil heat exchanger 1004 is shown.

[0266] In the above implementations, one or more temperature sensors may be provided, for example for sensing a 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. 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.

[0267] The inventors were aware that, in various implementations, knowledge of one or more fuel characteristics can be used to tailor the operation of the heat exchange system 3000 so as to improve the performance of the engine 10, by taking advantage of properties that vary between aviation fuels.

[0268] For the sake of clarity and brevity, a single bypass valve 2007 and a corresponding bypass line 2005 are referenced in the methods described herein. As indicated by the implementation of [Fig. 12], implementations may actually include multiple bypass valves 2007, 2007a, 2007b, each having an associated bypass line 2005, 2005a, 2005b to allow oil flow to bypass one or more of the secondary fuel-oil heat exchanger 1004, the primary fuel-oil heat exchanger 1006, and / or the air-oil heat exchanger 2020.When reference is made to "bypass valve 2007" or "bypass line 2005" in the disclosure of the methods illustrated in Figures 13-17, reference is made to one or more of the plurality of bypass valves 2007, 2007a, 2007b and associated bypass lines 2005, 2005a, 2005b, depending on the combination of heat exchangers that may be addressed.

[0269] [Fig. 13] illustrates a first method 100 involving the use of knowledge of fuel characteristics to determine a control of the heat exchange system 3000. The method 100 comprises the determination 102 of at least one fuel characteristic of the fuel intended to be burned by the combustion chamber 16. This determination 102 can be carried out by any of the methods described above, possibly using a fuel composition determination module 57.

[0270] The engine 10 used to implement the method 100 includes an air-oil heat exchanger 2020 and a fuel-oil heat exchanger 1006, a bypass line 2005 for allowing a proportion of the oil to flow past one of the air-oil heat exchanger and the fuel-oil heat exchanger; and a bypass valve 2007 for allowing the proportion of the oil sent through the bypass line to be varied, and the method 100 includes controlling 104 the bypass valve 2007 based on the at least a fuel characteristic so as to adjust the proportion of oil sent through each heat exchanger to cruising conditions.

[0271] More specifically, for the implementations described with respect to Figures 5-12 above, the method 100 includes controlling 104 the bypass valve 2007 of the 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 bypass line 2005 of the oil loop system 2000, thereby regulating the amount of oil bypassing the primary fuel-oil heat exchanger 1006, and the amount bypassing the air-oil heat exchanger 2020 and optionally the amount bypassing the secondary fuel-oil heat exchanger 1004.In implementations with different heat exchanger arrangements, for example a parallel arrangement of heat exchangers instead of a series arrangement, the diverter valve 2007 may be replaced by or act as a modulating valve directing oil flow along one or more branches of the split parallel arrangement, thereby allowing the proportion of oil sent through each heat exchanger to be varied in this manner.

[0272] As indicated by the dotted line in [Fig. 13], 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.

[0273] The bypass valve 2007 may therefore be intended to divert a fixed portion of the oil flow through the bypass line 2005 during operation of the engine 10, the fixed portion being determined at start-up of the engine 10 on the basis of the at least one determined fuel characteristic (option (i) above). Alternatively, the bypass valve 2007 may be intended to divert a variable portion of the oil flow through the bypass line 2005 during operation of the engine 10 during a single flight (option (ii) above). The bypass valve 2007 may therefore be actively controlled to vary the proportion of oil sent through the bypass line 2005 (and thus through each heat exchanger 1004, 1006, 2020), particularly in implementations where the aircraft 1 carries multiple different fuels in different tanks, and may change the fuel (or fuel mixture) used during flight. The active control of the bypass valve 2007 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, control of the bypass valve 2007 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.

[0274] The inventors were aware that fuels having certain fuel characteristics may accept more heat than others without decomposing, for example, due to a 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 allows for more efficient use of different fuels, as well as more efficient engine operation and oil cooling.

[0275] The fuel characteristic(s) may not be the only data considered in adjusting the bypass valve 2007 - 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 bypass valve 2007, 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 primary 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. not be powered / functional for some of this time, and therefore not provide any temperature rise.

[0276] Depending on the suitability of the at least one determined fuel characteristic, the method 100 may comprise using the bypass valve 2007 to perform one or more of the following operations: • sending all of the oil through the primary fuel-to-oil heat exchanger 1006 (none through the air-to-oil heat exchanger 2020) for one or more periods of time of at least 30 minutes in cruise mode, such that no heat is lost to the environment through the air-to-oil heat exchanger during at least some periods of operation in cruise mode; • send all oil through the primary fuel-to-oil heat exchanger 1006 (none through the air-to-oil heat exchanger 2020) 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 primary fuel-to-oil heat exchanger 1006 (no more than 5% through the air-to-oil heat exchanger 2020) 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-to-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.

[0277] 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 whose SAF content (%SAF) is above 60%, 65%, 70%, 75% or 80% can be considered adequate for a reduction in oil flow 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. Fuel characteristics may be considered independently / in isolation in other implementations.

[0278] 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 the thermal stability data indicates that the fuel is stable in operation at temperatures above 160°C or 170°C, the bypass valve 2007 may be controlled 104 such that at least 80%, and possibly 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 to 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 bypass valve 2007 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 possibly 90 to 100%, of the heat transferred from the oil at cruising speed can be transferred to the fuel.Generally speaking, less heat can be transferred to a conventional fuel if the heating value of the fuel exceeds a threshold, because, for higher heating value fuels, the flow rate is generally reduced to achieve the same thrust level - the fuel in the primary fuel-oil heat exchanger 1006 may therefore experience an increased temperature increase 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 increase in fuel temperature 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.

[0279] 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.

[0280] A gas turbine engine 10 for an aircraft implementing this method 100 includes a bearing 2006 for supporting the fan shaft, and an oil loop system 2000 for supplying oil to the bearing 2006. The heat exchange system 300 of the engine 10 includes an air-to-oil heat exchanger 2020 through which the oil in the oil loop system 2000 flows, and a fuel-to-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 bypass valve 2007 is provided, the valve 2007 being adapted to allow the proportion of oil sent through a bypass line 2005 to be varied and thereby bypass one or more of the heat exchangers 1006, 2020.The engine 10 also includes 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 bypass valve 2007 is intended to be controlled based on the at least one fuel characteristic, as determined by the fuel composition determination module 57, so as to adjust the proportion of oil sent via the bypass line 2005 (and thus the proportion of oil flowing through each heat exchanger 1006, 2020) to cruising conditions. A controller 58 may be provided for making and implementing decisions based on the output of the fuel composition determination module 57.

[0281] 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, some fuels may be heated to higher temperatures in one or more fuel-oil heat exchangers than traditional fuels, without significantly increasing the coking. This can allow for a process that provides 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 3000 Controllable Heat Exchange System has a key role to play in managing the heat transfer ratio.

[0282] 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 200, 300 of Figures 14 and 15 address both of these aircraft operating scenarios.

[0283] [Fig. 14] illustrates a method 200 implementing these considerations at cruise conditions, and [Fig. 15] illustrates a method 300 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).

[0284] First examining the method 200 to be performed in cruise mode, the method 200 is intended to be performed in a direct drive gas turbine engine 10 comprising an oil loop system 2000 for supplying oil to the engine components 2006; 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 primary 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 bypass line 2005 is provided, the bypass line 2005 being for allowing a proportion of the oil to flow past the air-oil heat exchanger 2020 or the fuel-oil heat exchanger 1006; and a bypass valve 2007 for allowing the proportion of the oil sent through the bypass line 2005 to be varied. The method 200 includes controlling 202 the heat exchange system 3000 such that, under cruising conditions, a heat transfer ratio of: .

[0285] [Math.6] oil to air heat transfer rate (kJkg~ oil to fuel heat transfer rate (kJkg'1) is in the range of 0 to 0.67, and optionally in the range of 0 to 0.60, 0 to 0.50, 0 to 0.40, 0 to 0.30, 0 to 0.20, or 0 to 0.10. A controller 58 may be provided to implement this control.

[0286] 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.

[0287] The method 200 may further comprise receiving data 204 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 or a fuel flow path, and / or possibly a fuel tank temperature or an oil tank temperature), and fuel flow rate data. Such data may be received 204 by a controller 58 and used 206 to adjust the control 202 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 206 may be performed at regular intervals, or in response to predetermined stimuli (e.g., a change in temperature or flow rate, or a change in engine operation or altitude). These steps 204, 206 may alternatively be considered part of the control 202 of the heat exchange system 3000 (i.e., not as separate steps). The method 200 may also be intended to use other information, e.g., temperature data (of oil, fuel, and / or ambient temperature of an environment around the aircraft 1), flow rate data (of oil and / or fuel), and / or one or more fuel characteristics, to determine 206 the control actions to be taken.

[0288] The control step 202 of the heat exchange system 3000 may include any one, or any combination, of the examples provided for step control 104 of the method 100 of [Fig. 13], for example, by decreasing the amount of oil sent through the at least one air-oil heat exchanger 2020 by increasing the amount of oil sent through the bypass line 2005 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.

[0289] In implementations where the engine 10 includes a refrigeration cycle apparatus 1007, the control step 202 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 1004, 1006, in some cases such that the fuel temperature is raised above the oil temperature. In such implementations, a lower value for the heat transfer ratio may be achieved than in implementations without a 1007 refrigeration cycle apparatus - 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 further 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.

[0290] The fuel temperature when it reaches the combustion chamber 16 may also be taken into account. The method 200 may include controlling 202 the heat exchange system 2000 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, or in the range of 0 to 0.45 provided that the fuel temperature at the inlet of the combustion chamber is at least 140°C. One or more temperature sensors 1009 may therefore be used, and their output(s) may be taken into account when adjusting the control 202 of the heat exchange system 2000.

[0291] The method 200 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 of 0 to 0.2 provided the fuel is at least 70% SAF, or 0 to 0.1 provided the fuel is at least 80% SAF. One or more fuel characteristics - possibly determined by a module 57 for determining the fuel characteristics as described above - can therefore also be used; the information can be taken into account when adjusting the control 202 of the heat exchange system 2000.

[0292] In some implementations, no more than 20% of the heat transferred from the oil at cruise may be transferred to the air and the oil-to-air heat transfer rate may be maintained in the range of 0 to 100 kJ per kilogram of fuel at cruise conditions.

[0293] In some implementations, at least 80% of the heat transferred from the oil at cruise can be transferred to the fuel and the oil-to-fuel heat transfer rate can be maintained in the range of 110 to 200 kJ per kilogram of fuel at cruise conditions.

[0294] 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 direct drive engine 10 are shown as a function of the fuel temperature limit in [Fig. 18]. 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 fuel with a fuel temperature limit. A typical value of % heat transferred to 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 heat can be transferred to the fuel at cruise for the entire cruise duration, without any need for air cooling, resulting in a heat transfer ratio of zero.

[0295] Turning now to the method 300 to be performed at idle, the method 300 is further intended to be performed in a direct drive gas turbine engine 10 comprising an oil loop system 2000 for supplying oil to engine components 2006; 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. A modulating valve 2016 is further provided for regulating the proportion of the oil sent through each branch of the oil loop system 2000.A bypass line 2005 is provided, the bypass line 2005 being for allowing a proportion of the oil to flow past the air-oil heat exchanger 2020 or the fuel-oil heat exchanger 1006; and a bypass valve 2007 for allowing the proportion of the oil sent through the bypass line 2005 to be varied. The method 300 includes controlling 302 the heat exchange system 2000 such that, when the aircraft 1 is operating at idle conditions, a heat transfer ratio of: .

[0296] [Math.7] oil to air heat transfer rate (kJkgj oil to fuel heat transfer rate (kJl <g*) is in the range of 0 to 1.5. 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 300 may include maintaining the heat transfer ratio at idle conditions below 1.0, in the range of 0 to 0.60, 0 to 0.50, 0 to 0.40, 0 to 0.30, 0 to 0.20, or 0 to 0.10.

[0297] 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. Such data may be received 304 by a controller 58 and used 206 to adjust the modulation 202 of the heat exchange system 2000 at idle, so as to maintain the heat transfer ratio at a desired level or within desired limits. This checking and adjustment / correction 206 may be performed at regular intervals or in response to predetermined stimuli (e.g., a change in temperature or flow rate of fuel or oil, or a change in aircraft operation, such as the start of taxiing). These steps 304, 206 may alternatively be considered part of the control 302 of the heat exchange system 2000. The method 300 may also be intended to use other information, for example, temperature data (of oil, fuel, and / or ambient temperature of an environment around the aircraft 1), flow rate data (of oil and / or fuel), and / or one or more fuel characteristics, in order to determine 206 the control actions to be taken.

[0298] The control step 302 of the heat exchange system 2000 may comprise any one, or any combination, of the examples provided for the modulation step 204 of the method 200 of [Fig. 14], 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 low, or by adjusting the proportion of oil and / or fuel sent through a respective bypass line 1005 or recirculation line 6011.

[0299] In implementations where the engine 10 includes a refrigeration cycle apparatus 1007, the step 302 of controlling the heat exchange system 2000 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 0.40, and optionally less than or equal to 0.30, 0.20, and further 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 in the range of 0.38 to 1.2.

[0300] The fuel temperature when it reaches the combustion chamber 16 may also be taken into account. The method 300 may include controlling 302 the heat exchange system 2000 during idle operations such that the heat transfer ratio is in the range of 0.3 to 1.5, provided that the fuel temperature at the inlet of the combustion chamber 16 is less than 180°C, or in the range of 0 to 0.3, provided that the fuel temperature at the inlet of the combustion chamber 16 is equal to or greater than 180°C. One or more temperature sensors may therefore be used, and their output(s) may be taken into account when adjusting the control 202 of the heat exchange system 3000.

[0301] The method 300 may, additionally or alternatively, take into account the nature of the fuel - for example, whether the fuel is or comprises a sustainable aviation fuel (SAF). One or more fuel characteristics - possibly determined by a fuel characteristics 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.

[0302] In some implementations, no more than 20% of the heat transferred from the oil at cruise may be transferred to the air and the oil-to-air heat transfer rate may be maintained in the range of 0 to 180 kJ per kilogram of fuel at cruise conditions.

[0303] In some implementations, 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 100 to 300 kJ per kilogram of fuel at cruise conditions. At idle, a greater % of heat rejection to air is generally used than at cruise and other higher engine power conditions.

[0304] As with the method 200 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 one different location along a main fuel path from the fuel inlet in 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 one or more heat exchangers 1004, 1006.

[0305] The inventors were also aware that the use of fuels different from traditional kerosene-based fuels, such as sustainable aviation fuels, may lead to different fuel properties, and that the parameters at cruise conditions may be adjusted to utilize the different fuel properties. In particular, some fuels may be heated to higher temperatures in the fuel-oil heat exchangers 1004, 1006 than traditional fuels. This can improve the cooling of the oil before it is returned to the rest of the turbine engine, and / or improve the combustion efficiency of the fuel. Using the fuel to absorb more heat from the oil, rather than relying on heat transfer from the oil to the surroundings / air (e.g., in an oil-to-air heat exchanger), provides a more thermally efficient turbine engine. In addition, the improved oil cooling can in turn improve the cooling effect of the oil on the engine components through which it flows.

[0306] [Fig. 16] illustrates an exemplary method 400 of operating a gas turbine engine 10. The method 400 comprises the steps of: Step 402: Providing fuel to the gas turbine engine 10 via the fuel management system 1500.

[0307] As discussed with respect to Figure 6C, the fuel management system 1500 includes a primary fuel-oil heat exchanger 1006 and a secondary fuel-oil heat exchanger 1004 for transferring heat to or from the fuel, and comprises a portion of a heat exchange system 3000. An engine fuel pump 1003 is positioned downstream of the primary fuel-oil heat exchanger 1004 and upstream of the secondary fuel-oil heat exchanger 1006. The fuel management system 1500 is arranged such that fuel reaches the primary fuel-oil heat exchanger 1006 before the secondary fuel-oil heat exchanger 1004. Step 404: Control the heat exchange system 3000 so as to raise the fuel temperature to at least 120°C at the outlet of the secondary fuel-oil heat exchanger 1004 at cruising conditions.

[0308] Control 404 of heat exchange system 3000 may include regulating a flow of fuel through the heat exchangers - for example, recirculating a variable proportion of the fuel through primary fuel-oil heat exchanger 1006, and / or allowing a variable proportion of the fuel to bypass primary fuel-oil heat exchanger 1006. While [Fig.6] shows an implementation with a bypass line 1005 but no recirculation, and Figures 7A-7C show implementations without a bypass line (at least not for the primary heat exchanger - Figures 6A and 6C provide a route for some of the fuel reaching the combustion chamber 16 to bypass the secondary heat exchanger while the remainder of the fuel passes through it) but with a recirculation line 6011, it will be appreciated that one or more bypass lines and / or one or more recirculation lines may be provided together in various implementations. Furthermore, in some systems, the direction of flow within a . same conduit may be reversible such that it may be used as a bypass conduit 1005 or a recirculation conduit 6011. It will be noted that the Figures are provided for example purposes only, and are not intended to be limiting.

[0309] Control 404 of heat exchange system 3000 may include regulating fuel flow through the heat exchangers - for example, allowing a variable proportion of the fuel to bypass primary fuel-oil heat exchanger 1006 or secondary fuel-oil heat exchanger 1006. One or more controllable valves or pumps, and optionally one or more sensors, may be provided accordingly to facilitate regulating fuel flow.

[0310] Additionally or alternatively, as with fuel flow, oil flow may be controlled using one or more bypass valves 2007 configured to control flow through one or more bypass lines 2005, when present, allowing oil to bypass one or more heat exchangers 1004, 1006, 2020 instead of flowing through them. Oil may also be recirculated in some implementations, and / or oil flow may be adjusted by controlling one or more oil pumps. One or more controllable valves and / or pumps, and optionally one or more sensors, may be provided accordingly to facilitate control of oil flow.

[0311] A controller 58 may be used to actively manage fuel and / or oil flows through and around the heat exchangers 1004, 1006, 2020. 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. The controller may receive inputs from one or more temperature sensors 1009, and may control one or more valves (e.g., a recirculation valve 6010 or a bypass valve 2007), and / or the pump 1003, based on the received data. One or more oil feed pumps and / or scavenge pumps may also be controlled by the controller.

[0312] Fuel flow may be regulated using a recirculation valve 6010 as described above. Alternatively or in addition, fuel flow may be regulated using one or more bypass lines 1005 (as illustrated in [Fig.5]) to allow a proportion of the fuel to bypass passing through either or both heat exchangers, and / or by adjusting the % split of fuel that flows directly to the combustion chamber 16 from the primary fuel-oil heat exchanger 1006 by as opposed to the % of fuel flowing from the primary fuel-oil heat exchanger into the secondary fuel-oil heat exchanger 1004 (and possibly then back to join the fuel flowing into the combustion chamber 16, depending on the implementation).

[0313] Heating the fuel to higher temperatures than previously used may improve cooling of the oil before it is returned to the rest of the turbine engine, and / or may improve fuel combustion efficiency. Positioning the secondary fuel-oil heat exchanger 1004 after the engine fuel pump 1003 may facilitate achieving higher fuel temperatures without compromising the longevity of the fuel pump.

[0314] The auxiliary system 7010 powered by the secondary fuel-oil heat exchanger 1004 may benefit from higher fuel temperatures than those appropriate for use in the combustion chamber 16 or components through which the primary fuel flow path passes. As such, it may be desirable for the fuel temperature at the outlet of the secondary fuel-oil heat exchanger 1004 to be higher than the fuel temperature at the inlet of the combustion chamber 16. Further, as illustrated by Figure 8, for example, oil may flow through the secondary fuel-oil heat exchanger 1004 before flowing through the other heat exchangers 1006, 2020 of the heat exchange system 3000.As such, the oil temperature may be higher when passing through the secondary fuel-oil heat exchanger 1004 than the other heat exchangers 1006, 2020. The control 404 of the heat exchange system 3000 may include control of a fuel bypass valve so as to adjust the proportion of fuel sent through a fuel bypass line 1005 and as such regulate the flow of fuel through each of the primary fuel-oil heat exchanger 1006 and the secondary fuel-oil heat exchanger 1004 at cruising conditions.

[0315] The heat exchange system control 404 may include controlling the oil bypass valve 2007 to adjust the proportion of oil sent through an oil bypass line 2005 and as such regulate the flow of oil through each of the primary fuel-oil heat exchanger 1006 and the secondary fuel-oil heat exchanger 1004 at cruising conditions.

[0316] Control 404 of heat exchange system 3000 may include controlling heat exchange system 3000 to raise the fuel temperature such that the fuel temperature at the outlet of the secondary heat exchanger is in the range of 120°C to 180°C, in the range of 135°C to 200°C, 135°C to 180°C, 150°C to 200°C, 150°C to 180°C, 150°C to 170°C or approximately 120°C, 130°C, 140°C, 150°C, 160°C or up to 200°C.

[0317] The inventors were aware that allowing unsuitable fuels to reach these higher temperatures could be detrimental to engine performance, however, and potentially even dangerous - excessive thermal decomposition of a fuel / deposit can cause safety issues by blocking passages and causing components to malfunction (e.g., valve sticking, and blocked nozzles, ports, valves, etc.) which could result in loss of thrust control. The method 500 illustrated in [Fig. 17] is therefore provided as a safety precaution.

[0318] The method 500 illustrated in [Fig. 17] may be performed in a gas turbine engine 10 substantially as described for the preceding methods relating to oil flow, but the one or more controllable oil valves 2007, 2007a, for varying a proportion of oil bypassing (and by extension being sent through) at least one of the air-oil heat exchanger 2020, the primary fuel-oil heat exchanger 1006 and the secondary fuel-oil heat exchanger, are replaced by at least one controllable oil or air valve 2007, 2007a, 2007b, 2022 for varying at least one of an oil flow rate and an air flow rate through at least one of the air-oil heat exchanger 2020 and the fuel-oil heat exchanger 1006. Air flow control, or oil flow, or both, can therefore be provided.In implementations where a controllable air valve 2022 is provided (possibly in addition to a controllable oil valve 2007, 2007a, 2007b), the engine may require a temperature sensor 1009 to provide an indication of fuel temperature - possibly at the inlet to the combustion chamber 16. More generally, the sensor 1009 may be located anywhere downstream of the primary fuel-oil heat exchanger 1006 on the fuel side.This temperature sensor 1009 is generally located at or near an inlet to the combustion chamber 16, and more specifically near a nozzle inlet to the combustion chamber 16 so as to sense the fuel temperature at the inlet to the combustion chamber directly, although it will be appreciated that variations in fuel temperature between the outlet of the primary fuel-oil heat exchanger 1006 and the combustion chamber 16 are likely to be small if the secondary fuel-oil heat exchanger 1004 is bypassed and calculable from knowledge of the engine 10 if the fuel temperature at a location different from the sensor 1009 is desired. In some implementations where the fuel temperature at the inlet to the combustion chamber 16 is used, . the sensor 1009 can be placed differently and the fuel temperature at the combustion chamber inlet can be calculated based on the temperature output and knowledge of the engine 10.

[0319] The method 500 includes determining 502 whether the fuel temperature has increased above a threshold set at cruise conditions, based on an output from the temperature sensor 1009. For example, a direct comparison may be made between data received from the temperature sensor 1009 and one or more temperature thresholds stored in memory. This determination 502 may therefore be made automatically, possibly by a controller 58 (which may be or include a dedicated processing module, or may be provided by a more general EEC).

[0320] The method 500 further comprises, in response to determining that the fuel temperature has increased above the threshold set at cruise conditions, controlling 504 the at least one valve 2007, 2007a, 2022 so as to change the at least one flow rate (air flow or oil flow) through the at least one heat exchanger 1004, 1006, 2020 as appropriate to reduce the fuel temperature. For example, adjusting an oil valve 2007, 2007a may send a greater proportion of oil through bypass lines 2005, 2005a to pass the primary fuel-oil heat exchanger 1006 and / or the secondary fuel-oil heat exchanger 1004 and recombine it with the remaining oil that has passed through the fuel-oil heat exchanger(s) 1004, 1006 at or downstream of the oil outlet of that heat exchanger 1006, 1004, as illustrated in [Fig. 12].In a parallel configuration (not shown), reducing oil flow through the fuel-oil heat exchanger 1006 by actuation of a modulating valve may automatically increase oil flow through the air-oil heat exchanger 2020 (unless a bypass line around the air-oil heat exchanger 2020 is present and used to compensate).

[0321] As indicated by the dotted line in [Fig. 17], this process 500 may be repeated - frequent checks may be preferred for safety reasons, possibly on a regular schedule, and / or when triggered by a change in engine operation (e.g., a change in altitude or thrust demand). The oil valve 2007, 2007a, or the air valve 2022 may therefore be actively controlled 504 to vary the flow rate of oil or air through the respective heat exchanger 1006, 2020. Active control of the valve 2007, 2007a, 2022 may be automated and implemented by the controller 58. For example, active control of the bypass valve 2007 associated with the secondary fuel-oil heat exchanger 1004 may be automated and implemented by a controller 58 of the exchange system heat exchanger 3000, which may be a dedicated control device or part of a more general EEC. In various implementations, control of one or more of the valves 2007, 2007a, 2022 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.

[0322] In implementations with both an air valve 2022 and an oil valve 2007, 2007a, airflow through the air-oil heat exchanger 2020 may be increased as more oil is sent through the air-oil heat exchanger 2020; the increase may be linear with the increase in oil flow rate. In terms of % airflow, the increase may be significant—for example, from the air valve 2022 being fully closed to fully open, or from a first position with 10% of the maximum air flow rate up to 90% or 100% of the maximum air flow rate when the fuel temperature is deemed too high.The method 500 may include, in response to the determination 502 that the fuel temperature has increased above a threshold set at cruise conditions, controlling 504a the air valve 2022 to send more air through the air-to-oil heat exchanger 2020 and controlling 504b the oil valve 2016, 2007 to send less oil through the fuel-to-oil heat exchanger 1006. The controlling / adjusting 504a of the air valve 2022 may be performed simultaneously with the controlling / adjusting 504b of the oil valve 2016, 2007.

[0323] In some implementations, the airflow through the air-to-oil heat exchanger 2020 may not be actively regulated. In such implementations, the airflow is likely to change naturally at different operating conditions based on the pressure ratio between the heat exchanger air inlet 2020a and the heat exchanger air outlet / exhaust 2020b. In this configuration, air will continue to flow through the air passages of the air-to-oil heat exchanger 2020 even when the air-to-oil heat exchanger 2020 is completely bypassed on the oil side / when the oil flow rate through the heat exchanger 2020 is zero. In alternative implementations, active control of the airflow through the air-to-oil heat exchanger 2020 is provided.For example, the air-oil heat exchanger 2020 and the fuel-oil heat exchanger(s) 1004, 1006 may be in series, without a bypass on the oil side of the air-oil heat exchanger 2020, but with an air valve 2022 on the air side that controls the air flow (both an oil bypass valve 2007b and an air flow valve 2022 may be provided in other implementations). In such a series arrangement, the exchanger . air-oil heat exchanger 2020 may be upstream or downstream of the fuel-oil heat exchanger 1006 with respect to the oil flow. In implementations with a series arrangement, having the air-oil heat exchanger 2020 upstream of the fuel-oil heat exchanger(s) 1004, 1006 with respect to the oil flow (unlike in [Fig. 12]) may help to avoid overheating the fuel - the oil can be cooled as much as desired by increasing an air flow up to a predetermined threshold.It may be kept in mind that without a dedicated mechanism to actively drive airflow, such as a gas-gas ejector or a dedicated blower, there is only so much oil can be cooled in a single pass through the air-oil heat exchanger (i.e., once the air valve 2022 is fully open, no further control adjustments can be made to increase cooling in a single pass through the air-oil heat exchanger 2020). The oil can be cooled further by recirculating oil through the air-oil heat exchanger 2020 before it reaches the fuel-oil heat exchanger 1006.

[0324] In some implementations, an oil bypass line 2005b and corresponding control valve 2007b may be implemented for the air-oil heat exchanger 2020 only. Optionally, only an oil flow rate through the air-oil heat exchanger 2020 may be actively controlled in the method 500 being described.

[0325] In various implementations: - A 2005b bypass line is provided on the oil side of the 2020 air-oil heat exchanger only (no oil bypass from the fuel-oil heat exchanger and no active air flow control); - A bypass line 2005, 2005a, 2005b may be provided for both the air-oil heat exchanger 2020, the primary fuel-oil heat exchanger 1006 and the secondary fuel-oil heat exchanger 1004 (no active airflow control); or - Air flow through the 2020 air-oil heat exchanger may be controllable, and there may be no oil bypass lines.

[0326] In general, having control of at least one flow through the air-oil heat exchanger 2020 may be deemed more important than having control of either flow through the fuel-oil heat exchanger(s) 1004, 1006. One or more oil bypass lines 2005, 2005a and / or a fuel bypass line 1005 (and a corresponding control valve) may, however, be provided for the fuel-oil heat exchanger(s) 1004, 1006 in some implementations, possibly in addition to one or more other controllable valves.

[0327] A combination of air flow control and oil flow control may be implemented in both parallel and series arrangements, but it will be appreciated that although all of the various control options may be used together, the increased complexity and number of components, and therefore the increased weight of the overall heat exchange system 3000, may not be desirable. Selecting a subset of the control options may therefore be preferred in many implementations - for example, either air flow control or oil flow control may be used to adjust the heat transfer within the air-oil heat exchanger 2020, such that control of both for a given heat exchanger 2020 may be deemed unnecessary.Similarly, in implementations where oil flows through the air-to-oil heat exchanger 2020 before the fuel-to-oil heat exchanger 1006, an adjustment of oil temperature in the air-to-oil heat exchanger 2020 (by regulating air and / or oil flow in that heat exchanger) may be used as an alternative to an adjustment of oil flow through one or more of the primary fuel-to-oil heat exchanger 1006 or the secondary fuel-to-oil heat exchanger 1004 to change the amount of heat transferred to the fuel in said one or more fuel-to-oil heat exchangers 1004, 1006.

[0328] Additionally or alternatively, in implementations where the heat exchange system 3000 includes a refrigeration cycle apparatus 1007 for transferring additional heat from the oil to the fuel beyond that transferred by the primary fuel-to-oil heat exchanger 1006 and the secondary fuel-to-oil heat exchanger 1004, the method 500 may further include controlling the refrigeration cycle apparatus 1007 to reduce the amount of additional heat transferred to the fuel in response to the determination 502 that the fuel temperature has increased beyond the threshold set at cruise conditions. This control may be achieved by reducing the flow of oil through the refrigeration cycle apparatus 1007, or by reducing the power of the refrigeration cycle apparatus 1007 (e.g., by reducing the flow rate of the refrigerant), or by disabling the refrigeration cycle apparatus 1007.The threshold set for deactivation of the refrigeration cycle apparatus 1007 may be lower than the threshold set for changing an oil and / or air flow rate through a heat exchanger 1006, 2020. .

[0329] The threshold set for the control step 504 may be in the range of 140°C to 300°C, and optionally 200°C to 300°C, and further optionally 250°C to 300°C. For example, the threshold set may be 140°C, 150°C, 160°C, 170°C, 180 °C, 190 °C, 200 °C, 210 °C, 220 °C, 230 °C, 240 °C, 250 °C, 260 °C, 270 °C, 280 °C, 290 °C, or 300 °C.

[0330] The method 500 of some implementations further comprises determining 501 the set threshold, to which the temperature sensor output is compared. This determination 501 may be performed on-site, and possibly in flight. The determination 501 may be or include the calculation of a threshold value, or the identification of an appropriate threshold value from a set of threshold values ​​stored in memory.

[0331] The inventors were aware that knowledge of one or more fuel characteristics can be used in determination 501 to ensure safety while taking advantage of properties that vary between aviation fuels. Determination 501 may therefore be based on at least one fuel characteristic. The at least one fuel characteristic may be or include any of the examples discussed above, e.g., thermal stability of the fuel, nitrogen content of the fuel, sulfur content of the fuel, and / or sustainable aviation fuel - SAF - content of the fuel (%SAF). The at least one fuel characteristic may be determined by any one or more of the approaches discussed above.

[0332] The step 501 of determining the set threshold may comprise increasing the set threshold, possibly linearly, with increasing thermal stability of the fuel. The threshold used to regulate the heat modulation may therefore be increased linearly with the thermal stability of the fuel. 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 kept in mind that this is generally not an instantaneous change when a given temperature is reached; aviation fuels include various components that are susceptible to degradation at different temperatures, and time spent at high temperature is also a factor in fuel degradation.The threshold value may be set based on comparing a maximum allowable deposition rate due to fuel degradation to the thermal stability of the fuel in use. A fuel may be considered stable at a given temperature if its degradation rate is below a certain threshold at that temperature. The method 500 may also include use of a clock or timer, or a fuel flow sensor—a higher fuel temperature may be acceptable at a higher pumping rate, for example, because faster-moving fuel will spend less time at that temperature before being combusted. Alternatively, or in addition, the method 500 may include use of one or more pressure sensors. to determine a fuel flow rate. By measuring a pressure drop across a known orifice size, a flow rate can be determined.

[0333] Alternatively or in addition, the step 501 of determining the set threshold may comprise increasing the set threshold with increasing a SAF content of the fuel (possibly linearly), for fuels with a SAF content above 70%. As discussed above, it has been found that increasing %SAF blended with a fossil-based aviation fuel increases thermal stability, but in a non-linear manner. Predetermined, distinct thresholds based on %SAF may therefore be favored instead of continually changing the set threshold proportionally to a SAF content, or a more complex (non-linear) but continuous relationship may be established.

[0334] Alternatively or in addition, the step 501 of determining the set threshold may comprise decreasing the set threshold with increasing heteroatomic species content of the fuel (possibly linearly). For example, the thermal stability of a Jet-A has been found to decrease with increasing nitrogen content (nitrogen content being a measure of the amount of nitrogen-containing species present). It is known that the interaction of sulfur-containing and nitrogen-containing species in the fuel can contribute significantly to fuel decomposition rates, so consideration of the composition of the fuel with respect to the multiple heteroatomic species may be implemented to account for these interactions.

[0335] In some implementations, particularly in implementations where a direct measure of fuel thermal stability is not available, multiple fuel characteristics may be determined and used in decisions about valve control. The one or more determined fuel characteristics may in effect be translated into a measure of thermal stability.

[0336] 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 (400) 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 fan shaft; at least one bearing (2006) for supporting the fan shaft; at least one auxiliary system (7010) for using a portion of the fuel; an oil loop system (2000) for supplying oil to the at least one bearing; and a heat exchange system (3000) comprising: a primary fuel-oil heat exchanger (1006) through which the oil in the oil loop system and at least substantially all of the fuel flow such that heat is transferred between the oil and the fuel;and a secondary fuel-oil heat exchanger (1004) through which the oil in the oil loop system and a portion of the fuel flow, such that heat is transferred between the oil and the fuel, the secondary fuel-oil heat exchanger being for supplying fuel to the at least one auxiliary system; and a fuel pump (1003) for pumping the fuel, wherein the fuel pump is located downstream of the primary fuel-oil heat exchanger (1006) and upstream of the secondary fuel-oil heat exchanger (1004) along a fuel flow path; the method further comprises controlling (402) the heat exchange system (3000) such that, under cruising conditions, the fuel temperature at the outlet of the second heat exchanger is in the range of 120°C to 200°C.;

2. The method (400) of claim 1, wherein the method comprises controlling the heat exchange system such that, under cruising conditions, the temperature of fuel at the outlet of the second heat exchanger is in the range of 120°C to 180°C.

3. The method (400) of claim 1, wherein the method comprises controlling the heat exchange system such that, under cruising conditions, the fuel temperature at the outlet of the second heat exchanger is in the range of 135°C to 200°C; and optionally in the range of 135°C to 180°C.

4. The method (400) of claim 1, wherein the heat exchange system further comprises: a fuel bypass line (1005) for allowing a proportion of the fuel to bypass at least one of the primary fuel-oil heat exchanger and the secondary fuel-oil heat exchanger; and a fuel bypass valve for allowing the proportion of the fuel sent through the fuel bypass line to be varied; wherein controlling the heat exchange system comprises controlling the fuel bypass valve to adjust the proportion of the fuel sent through each of the primary and secondary fuel-oil heat exchangers to cruise conditions.

5. The method (400) of claim 1, wherein the heat exchange system further comprises: an oil bypass line (2005) for allowing a proportion of the oil to bypass at least one of the primary fuel-oil heat exchanger and the secondary fuel-oil heat exchanger; and an oil bypass valve (2007) for allowing the proportion of the fuel sent through the fuel bypass line to be varied; wherein controlling the heat exchange system comprises controlling the oil bypass valve so as to adjust the proportion of the oil sent through each of the primary and secondary fuel-oil heat exchangers to cruise conditions.

6. The method (400) of claim 1, wherein, under cruising conditions, the ratio of [Math. 1] fuel flow through the secondary fuel-oil heat exchanger (m3) is fuel flow to the combustion chamber of at least 0.3-

7. The method (400) of claim 1, wherein under cruising conditions the ratio of [Math. 2] fuel flow through the secondary fuel-oil heat exchanger (ni3) to fuel flow to the combustion chamber is at least 0.

5.

8. The method (400) of claim 1, wherein under cruising conditions the ratio of [Math. 3] fuel flow through the secondary fuel-oil heat exchanger (m3) to fuel flow to the combustion chamber is approximately 0.

6.

9. The method (400) of claim 1, wherein the heat exchange system (3000) further comprises a refrigeration cycle apparatus (1007), and the method 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.

10. The method (400) of claim 1, wherein the primary fuel-oil heat exchanger and the secondary fuel-oil heat exchanger are arranged in series along the oil loop system.

11. The method (400) of claim 1, wherein the heat exchange system further comprises at least one air-oil heat exchanger through which oil from the oil loop system flows.

12. A gas turbine engine (10) for an aircraft comprising: an engine core (11) comprising a turbine (19), a compressor (14) and a core shaft (26) connecting the turbine to the compressor; a fan (23) located upstream of the engine core; and a reduction gear (30) which receives an input from the core shaft and provides a drive to the fan so as to drive the blower at a lower rotational speed 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: a primary fuel-oil heat exchanger (1006) through which the oil in the oil loop system and at least substantially all of the fuel flow such that heat is transferred between the oil and the fuel; and a secondary fuel-oil heat exchanger (1004) through which the oil in the oil loop system and a portion of the fuel flow such that heat is transferred between the oil and the fuel, the secondary fuel-oil heat exchanger being for supplying fuel to the at least one auxiliary system;and a fuel pump (1003) for pumping fuel, wherein the fuel pump is located downstream of the primary fuel-oil heat exchanger and upstream of the secondary fuel-oil heat exchanger along a fuel flow path; wherein the heat exchange system (3000) is to be controlled such that, under cruising conditions, the fuel temperature at the outlet of the second heat exchanger is in the range of 120°C to 200°C.;

13. The gas turbine engine (10) of claim 12, wherein: the turbine (19) is a first turbine, the compressor (14) is a first compressor, and the core shaft (26) is a first core shaft; the engine core further comprises a second turbine (17), a second compressor (15), and a second core shaft (27) connecting the second turbine to the second compressor; and the second turbine, the second compressor, and the second core shaft are adapted to rotate at a higher rotational speed than the first core shaft.

14. A gas turbine engine (10) according to claim 12, wherein the heat exchange system (3000) is adapted to be controlled such that, under cruising conditions, the temperature of

15. fuel at the outlet of the second heat exchanger is in the range of 120°C to 180°C. A gas turbine engine (10) according to claim 12, wherein the core shaft (26) provides drive to the fan (23) directly, so as to drive the fan at the same rotational speed as the core shaft, such that the engine (10) is a direct drive turbine engine.

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