Gas turbine modulation valve

Abstract

A method 100 of operating a gas turbine engine of an aircraft, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, a combustor arranged to combust a fuel, and a mandrel connecting the turbine to the compressor; a fan located upstream of the engine core; a gearbox receiving an input from the mandrel and outputting a drive to the fan; an oil circuit system arranged to supply oil to the gearbox; and a heat exchange system, the heat exchange system comprising an air-oil heat exchanger through which oil flows; and a fuel-oil heat exchanger through which oil and fuel flow; and a modulation valve arranged to allow a change in the proportion of oil delivered via each heat exchanger, the method comprising: determining 102 at least one fuel characteristic of the fuel arranged to be combusted by the combustor; and controlling 104 the modulation valve based on at least one fuel characteristic in order to adjust the proportion of oil delivered via each heat exchanger under cruise conditions.

Description

[0001] Cross - reference to related applications

[0002] This specification claims the benefit of the priority of UK Patent Application No. 2319118.2, filed on December 14, 2023, the entire content of which is incorporated herein by reference. Background Art Technical Field

[0003] The present disclosure relates to aircraft propulsion systems and to methods of operating an aircraft that include managing different fluids and heat transfer between different fluids, and more particularly to managing the heat exchange system of an aircraft engine.

[0004] Description of Related Technologies

[0005] In the aviation industry, there is a trend towards using fuels different from the traditional kerosene - based jet fuels commonly used currently. These fuels may have different fuel properties relative to petroleum - based hydrocarbon fuels. Therefore, it is necessary to consider the fuel properties of these new fuels and adjust the gas turbine engines themselves and the methods of operating gas turbine engines. Summary of the Invention

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

[0007] An engine core including a turbine, a compressor, a burner arranged to burn fuel, and a spool connecting the turbine to the compressor;

[0008] A fan located upstream of the engine core;

[0009] A gearbox that receives an input from the spool and outputs drive to the fan;

[0010] An oil circuit system arranged to supply oil to the gearbox; and

[0011] A heat exchange system including:

[0012] An air - oil heat exchanger through which oil in the oil circuit system flows;

[0013] And

[0014] A fuel - oil heat exchanger through which oil and fuel in the oil circuit system flow, such that heat is transferred between the oil and the fuel; and

[0015] A modulating valve arranged to allow changing the proportion of oil delivered through each heat exchanger,

[0016] ​

[0017] The method includes:

[0018] determining at least one fuel property of fuel arranged to be burned by a burner; and

[0019] controlling a modulating valve based on the at least one fuel property so as to adjust the proportion of oil delivered via each heat exchanger under cruise conditions.

[0020] The inventors have recognized that using fuels different from traditional kerosene-based jet fuels, such as sustainable aviation fuels, can result in different fuel properties, and parameters under cruise conditions can be adjusted to take advantage of the different fuel properties. In particular, some fuels can be heated to higher temperatures in one or more fuel-oil heat exchangers than traditional fuels without significantly increasing fuel decomposition products, including increased coking and / or varnishing. This can improve the combustion efficiency of the fuel, and / or can improve the overall thermodynamic efficiency of the engine while less heat is lost to the environment (e.g., via an air-oil heat exchanger). The higher fuel temperature at the burner inlet can allow for a method of providing improved oil cooling (since the fuel can absorb more heat). For example, the thermal stability of the fuel affects how much heat the fuel can accept / how high the fuel can be raised in temperature without forming deposits within pipes, burners, and / or hydraulic machinery units or other engine components. Taking into account the thermal stability of the fuel, and depending on the thermal stability of the fuel, more or less heat is transferred from the oil to the fuel, so more effective oil cooling can be provided while avoiding coking and / or varnishing, thereby improving aircraft performance.

[0021] Using fuel to carry away more heat from the oil rather than relying on heat transfer from the oil to the environment / air (e.g., in an air-oil heat exchanger) can also provide a more thermally efficient engine. This can improve the cooling of the oil before it returns to the rest of the turbofan engine. Additionally, 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.

[0022] During transitional periods, over the life of a gas turbine engine, the available aviation fuel changes over time and depends on the geographical location of the fuel supply point as well as other variables. It is important to determine the relevant fuel properties based on the specific fuel being used and perform control of the heat exchange system. Thus, the gas turbine operation can be adjusted to make full use of the various fuels. A controllable modulating valve arranged to regulate the oil flow through each heat exchanger plays a key role in the heat exchange system control.

[0023] The fuel property can be or include calorific value, thermal stability, or the percentage of sustainable aviation fuel (SAF) in the fuel.

[0024] An air-oil heat exchanger can help remove excess heat from the oil in an oil circuit system that has not been transferred to the fuel. This can allow further cooling of the oil beyond the transfer of heat from the oil to the fuel in a fuel-oil heat exchanger.

[0025] The oil circuit system (which can also be referred to as a recirculating oil system) can branch such that a certain proportion of the oil can flow along each branch. The air-oil heat exchanger and the fuel-oil heat exchanger can be arranged in a parallel configuration on different branches of the oil circuit system. A modulating valve in such an example can be arranged to allow changing the proportion of oil delivered via each branch, and the control modulating valve can thus regulate the proportion of oil delivered via each branch under cruise conditions. In some examples, more than two heat exchangers and / or more than two branches can be provided. Two or more branches of the primary oil circuit system can rejoin after the heat exchanger system such that the oil recombines after heat has been transferred from the oil.

[0026] The modulating valve allows changing the heat transferred from the oil by the air-oil heat exchanger and the fuel-oil heat exchanger. The level of oil flow (and thus heat transfer from the oil to air or fuel) can be modified based on the temperature of the fuel leaving the fuel-oil heat exchanger or entering the burner, and can thus allow controlling the fuel temperature within defined upper and lower limits. Similarly, control of the oil flow can allow the oil temperature to be maintained within an appropriate range. This control can help ensure improved efficiency of a turbine engine (e.g., by raising the fuel temperature) by using a temperature that is too high for the durability of the fuel pump or other components downstream of the fuel-oil heat exchanger without the risk of causing unnecessary damage to the durability of said components. The modulating valve can be arranged to allow up to 100% of the oil to be delivered via the fuel-oil heat exchanger. The modulating valve can be arranged to ensure that no less than 70%, 80%, or 90% of the oil in the oil is delivered via the fuel-oil heat exchanger during cruise.

[0027] The modulating valve can be arranged to divert a fixed portion of the oil flow to each of the fuel-oil heat exchanger and the air-oil heat exchanger during engine operation, the fixed portion optionally determined at engine start or before start, or based on at least one determined fuel characteristic when reaching or before reaching cruise altitude. Alternatively, the modulating valve can be arranged to divert a variable portion of the oil flow to each of the fuel-oil heat exchanger and the air-oil heat exchanger during operation of the engine. Thus, the modulating valve can be actively controlled to change the proportion of oil delivered via each heat exchanger, particularly in embodiments where the aircraft carries multiple different fuels in different tanks, and can change the fuel (or fuel mixture) used in flight. The active control of the modulating valve can be automatic and implemented by a controller of the heat exchange system.

[0028] The gas turbine engine may further include a generator (e.g., an integrated drive generator) and a secondary oil circuit system arranged to supply oil to the generator. The oil circuit system that supplies oil to the gearbox as described above may thus be referred to as the primary oil circuit system. The heat exchange system may include an oil-oil heat exchanger arranged to transfer heat between the primary circuit system and the secondary oil circuit system. A modulating valve or an additional valve may control the oil flow from the primary oil circuit system to the oil-oil heat exchanger.

[0029] The heat exchange system may further include a secondary fuel-oil heat exchanger arranged to receive fuel and oil from the secondary oil circuit system. The method may include using the secondary fuel-oil heat exchanger to transfer heat between the oil and the fuel from the secondary oil circuit system. For simplicity, the secondary fuel-oil heat exchanger may be referred to as the secondary heat exchanger.

[0030] The fuel-oil heat exchanger through which the oil in the primary oil circuit system flows may be referred to as the primary fuel-oil heat exchanger. For simplicity, the primary fuel-oil heat exchanger may be referred to as the primary heat exchanger. The fuel may flow through the secondary fuel-oil heat exchanger before flowing through the primary fuel-oil heat exchanger, such that heat is transferred from the oil in the secondary oil circuit system to the fuel before heat is transferred from the oil in the primary oil circuit system to the fuel.

[0031] The heat exchange system may include at least one bypass tube arranged to allow fuel (or oil) to bypass one or more heat exchangers of the heat exchange system. In some embodiments, the bypass tube may effectively form an additional branch in a parallel branch oil system, and a modulating valve or another oil valve may be arranged to regulate the amount of oil conveyed through the bypass tube or each bypass tube for oil based on one or more determined fuel characteristics and optionally based on one or more temperature measurements.

[0032] The heat exchange system may further include a refrigeration cycle device, and the method may include using the refrigeration cycle device to provide a heat lift by transferring more heat from the oil to the fuel, optionally such that the fuel temperature is raised to be higher than the oil temperature. A modulating valve or another oil valve may control the amount of oil flowing through the refrigeration cycle device.

[0033] The heat exchange system may further include a branched fuel return passage and at least one valve that controls the diversion of the fuel flow. The branched passage may be arranged to return fuel from the heat exchange system to at least two different locations along a main fuel path from which fuel enters a gas turbine engine to reach a combustor. 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 embodiments, at least most of the fuel may flow through the fuel-oil heat exchanger such that the branches of the branched fuel path leaving the heat exchanger are themselves the "main" fuel path - one or more smaller branches may be split off and rejoined to the main stream at different locations (e.g., more downstream locations, such as after one or more engine components downstream of the fuel-oil heat exchanger being discussed, or actually more upstream locations (thus acting as a recirculation pipe)).

[0034] Depending on the suitability of at least one determined fuel characteristic, the method may include delivering all of the oil via the fuel-oil heat exchanger for at least one or more consecutive periods of at least 30 minutes during cruise such that no heat is lost to the environment via the air-oil heat exchanger during at least some operating periods during cruise. The method may include delivering all of the oil via the fuel-oil heat exchanger for at least 15% and optionally at least 20% of the total cruise time. Depending on the suitability of at least one determined fuel characteristic, the method may include delivering at least 95% of the oil via the fuel-oil heat exchanger for at least 90% of the time of cruise such that very little heat, if any, is lost to the environment via the air-oil heat exchanger during at least 90% of the operation during cruise. Depending on the suitability of at least one determined fuel characteristic, the method may include transferring at least 80% of the heat transferred away from the oil during cruise to the fuel. In some embodiments, all of the heat transferred away from the oil during cruise may be transferred to the fuel for at least 90% of the time of cruise.

[0035] At least one fuel characteristic may be or include thermal stability. If the fuel operates stably at a temperature above 160 °C, then at least 80% of the heat transferred away from the oil during cruise may be transferred to the fuel.

[0036] At least one fuel characteristic may be or include the aromatic content in the fuel. If the fuel has an aromatic molar percentage below 12%, then at least 80% of the heat transferred away from the oil during cruise may be transferred to the fuel.

[0037] At least one fuel characteristic may be or include the percentage of sustainable aviation fuel in the fuel - %SAF. The SAF ratio (X%) may be by volume. If the fuel has an SAF content above 50%, then at least 80% of the heat transferred away from the oil during cruise may be transferred to the fuel.

[0038] At least one fuel property may be or include the calorific value of the fuel. If the fuel has a calorific value of at least 43.5 MJ / kg, at least 80% of the heat transferred away from the oil during cruise can be transferred to the fuel.

[0039] According to a second aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising:

[0040] An engine core, the engine core including a turbine, a compressor, and a shaft connecting the turbine to the compressor;

[0041] A fan, the fan being located upstream of the engine core; and

[0042] A gearbox, the gearbox receiving an input from the shaft and outputting a drive to the fan so as to drive the fan at a rotational speed lower than that of the shaft;

[0043] An oil circuit system, the oil circuit system being arranged to supply oil to the gearbox; and

[0044] A heat exchange system, the heat exchange system including:

[0045] An air-oil heat exchanger, the oil in the oil circuit system flowing through the air-oil heat exchanger;

[0046] And

[0047] A fuel-oil heat exchanger, the oil and fuel in the oil circuit system flowing through the fuel-oil heat exchanger such that heat is transferred between the oil and the fuel; and

[0048] A modulating valve, the modulating valve being arranged to allow changing the proportion of oil delivered through each heat exchanger; and

[0049] A fuel composition determination module, the fuel composition determination module being arranged to determine at least one fuel property of the fuel arranged to be burned by a burner,

[0050] wherein the modulating valve is arranged to be controlled based on the at least one fuel property so as to adjust the proportion of oil delivered through each heat exchanger under cruise conditions.

[0051] The turbine may be a first turbine, the compressor may be a first compressor, and the shaft may be a first shaft. The engine core may further include a second turbine, a second compressor, and a second shaft connecting the second turbine to the second compressor. The second turbine, the second compressor, and the second shaft may be arranged to rotate at a rotational speed higher than that of the first shaft.

[0052] The engine of the second aspect may be arranged to perform the method of the first aspect and may have any one of the features described with respect to the first aspect.

[0053] According to a third aspect, there is provided a method of operating a gas turbine engine, the gas turbine engine comprising:

[0054] an engine core including a turbine, a compressor, a spool connecting the turbine to the compressor, and a burner arranged to combust fuel;

[0055] a fan located upstream of the engine core;

[0056] an oil system arranged to circulate oil; and

[0057] a heat exchange system including at least one fuel - oil heat exchanger arranged to transfer heat from the oil to the fuel,

[0058] wherein the method comprises:

[0059] determining at least one fuel property of fuel arranged to be supplied to the burner; and

[0060] modulating the heat exchange system so as to regulate the fuel temperature at the burner inlet during cruise to a set level, the set level being based on the at least one fuel property.

[0061] The inventors have recognized that using fuels different from conventional kerosene - based jet fuels, such as sustainable aviation fuels, can result in different fuel properties and that parameters under cruise conditions can be adjusted to take advantage of the different fuel properties. In particular, some fuels can be heated to higher temperatures in one or more fuel - oil heat exchangers than conventional fuels without significantly increasing the formation of fuel decomposition products, including coking and / or lacquering. This can improve the combustion efficiency of the fuel. A higher fuel temperature at the inlet of the burner can allow for improved oil cooling (since the fuel can absorb more heat) and / or methods for improved fuel combustion efficiency. For example, fuel viscosity affects how the fuel is delivered to the burner and how it is ignited by the burner. Viscosity can affect the droplet size from a fuel spray nozzle, which in turn can affect combustion efficiency. Therefore, taking fuel viscosity into account when delivering the fuel to the burner and appropriately controlling the fuel viscosity by varying the heat input can provide more efficient fuel combustion, thereby improving aircraft performance. During transitional periods, over the life of a gas turbine engine, the available aviation fuel changes over time and depends on the geographical location of the fuel supply point and other variables. It is important to determine the relevant fuel properties based on the specific fuel used in the engine and perform control of the heat exchange system. Thus, gas turbine operation can be adjusted to make full use of various fuels. The controllable heat exchange system plays a key role in managing fuel temperature and oil temperature.

[0062] The fuel characteristics can be or include the percentage of sustainable aviation fuel (SAF) in the fuel, the amount of sulfur in the fuel (e.g., ppm), or the thermal stability of the fuel.

[0063] The set level can be an average value of at least 140°C, 150°C, 160°C, 170°C, 180°C, 190°C, or 200°C when entering the burner under cruise conditions. The fuel temperature at the burner inlet under cruise conditions can be defined as the average value within at least 5 minutes, and optionally within 10 minutes, 20 minutes, or 30 minutes, under steady-state cruise conditions. These average temperatures do not include transient spikes in temperature, which can be defined as fluctuations in the fuel temperature during operation, typically an increase in temperature. Each fluctuation can last no more than 5 minutes. The fuel temperature at the burner inlet under the cruise conditions defined herein being at least 140°C thus requires the fuel temperature to be maintained at 140°C or above for a period of time, and a transient spike to a temperature above 140°C is not sufficient.

[0064] In some embodiments, the set level can be higher than the oil temperature. The heat exchange system can further include a refrigeration cycle device, and the method can include using the refrigeration cycle device to provide a heat boost by transferring more heat from the oil to the fuel, such that the fuel temperature is raised above the oil temperature to reach the set level.

[0065] The gas turbine engine can further include a gearbox that receives an input from the core shaft and optionally outputs a drive to the fan via a fan shaft. The oil system can be arranged to circulate oil between the gearbox and at least one fuel-oil heat exchanger, and can be described as an oil circuit system arranged to supply oil to the gearbox;

[0066] The heat exchange system can include at least one air-oil heat exchanger, and modulating the heat exchange system to adjust the fuel temperature can include modulating the amount of oil delivered via the at least one air-oil heat exchanger.

[0067] The heat exchange system can include at least one bypass tube arranged to allow oil to bypass the heat exchanger. Modulating the heat exchange system to adjust the fuel temperature to the set level can include modulating the amount of oil delivered via the bypass tube. The heat exchange system can include at least one bypass tube arranged to allow fuel to bypass the heat exchanger. Modulating the heat exchange system to adjust the fuel temperature to the set level can include modulating the amount of fuel delivered via the bypass tube.

[0068] The heat exchange system can include at least one recirculation tube arranged to allow a fluid (e.g., oil or fuel) to pass through the heat exchanger multiple times. Modulating the heat exchange system to adjust the fuel temperature to the set level can include modulating the amount of fluid delivered via the recirculation tube or each recirculation tube.

[0069] The steps of modulating the heat exchange system may include controlling the oil flow through at least one fuel-oil heat exchanger such that between 50% and 100% of the heat lost from the oil is transferred to the fuel. The steps of modulating the heat exchange system may include controlling the oil flow through at least one fuel-oil heat exchanger such that between 80% and 100% of the heat lost from the oil is transferred to the fuel.

[0070] At least one determined fuel characteristic may be or include the thermal stability of the fuel, and the method may include modulating / controlling the heat exchange system such that the fuel temperature at the burner inlet during cruise increases with an increase in thermal stability, optionally linearly.

[0071] At least one determined fuel characteristic may be or include the percentage of SAF in the fuel, and the method may include modulating the heat exchange system such that once the %SAF exceeds 60%, the fuel temperature at the burner inlet during cruise increases with an increase in %SAF.

[0072] At least one determined fuel characteristic may be or include the presence of a tracer substance in the fuel, and the method may include modulating the heat exchange system such that the fuel temperature at the burner inlet during cruise is set to a predefined level corresponding to the tracer substance.

[0073] At least one determined fuel characteristic may be or include the sulfur content of the fuel, and the method may include modulating / controlling the heat exchange system such that when the sulfur content is low, the fuel temperature at the burner inlet during cruise increases, and vice versa.

[0074] The methods of the first aspect and the third aspect may be complementary and may be performed together in various embodiments. The method of the third aspect may be performed using the engine of the second aspect.

[0075] According to a fourth aspect, there is provided a gas turbine engine for an aircraft, the engine comprising:

[0076] An engine core including a turbine, a compressor, a spool connecting the turbine to the compressor, and a burner arranged to burn fuel;

[0077] A fan located upstream of the engine core;

[0078] An oil system arranged to circulate oil;

[0079] A heat exchange system including at least one fuel-oil heat exchanger arranged to transfer heat from the oil to the fuel; and

[0080] A fuel composition determination module, which is arranged to determine at least one fuel property of a fuel arranged to be burned by a burner.

[0081] And wherein, the heat exchange system is arranged to be modulated so as to adjust the fuel temperature at the burner inlet to a set level, which set level is based on at least one fuel property.

[0082] The gas turbine engine may further include a gearbox that receives an input from a spool and outputs a drive to a fan. The oil system may be arranged to circulate oil between the gearbox and at least one fuel-oil heat exchanger, and may be described as an oil circuit system arranged to supply oil to the gearbox.

[0083] The turbine may be a first turbine, the compressor may be a first compressor, and the spool may be a first spool. The engine core may further include a second turbine, a second compressor, and a second spool connecting the second turbine to the second compressor. The second turbine, second compressor, and second spool may be arranged to rotate at a higher rotational speed than the first spool.

[0084] The heat exchange system may further include a branched fuel return passage and at least one valve that controls the diversion of the fuel flow. The branched passage may be arranged to return fuel from the heat exchange system to at least two different positions along a main fuel path, from which the fuel enters the gas turbine engine to reach the burner.

[0085] The engine of the fourth aspect may be arranged to perform the methods of the first aspect and / or the third aspect, and may have any one of the features described with respect to the first aspect or the third aspect.

[0086] In any one of the first to fourth aspects:

[0087] At least one fuel property of the fuel may include at least one of the following:

[0088] i. The percentage of sustainable aviation fuel in the fuel;

[0089] ii. The heteroatom species concentration of the fuel;

[0090] iii. The aromatic content of the fuel;

[0091] iv. The polyaromatic content of the fuel;

[0092] v. The percentage of nitrogenous substances in the fuel;

[0093] vi. The presence or percentage of tracer substances or trace elements in the fuel;

[0094] vii. The hydrogen-to-carbon ratio of the fuel;

[0095] viii. Hydrocarbon distribution of the fuel;

[0096] ix. Level of non-volatile particulate matter emissions during combustion;

[0097] x. Naphthalene content of the fuel;

[0098] xi. Sulfur content of the fuel;

[0099] xii. Naphthene content of the fuel;

[0100] xiii. Oxygen content of the fuel;

[0101] xiv. Thermal stability of the fuel;

[0102] xv. Coking level of the fuel;

[0103] xvi. Indication that the fuel is a fossil fuel;

[0104] xvii. At least one of density, viscosity, calorific value, and heat capacity.

[0105] The method may further include chemically or physically detecting one or more parameters related to the fuel in a fuel tank (the fuel tank being configured to supply fuel to a burner via a heat exchange system) after fuel replenishment. The detected parameter(s) may be a fuel property or may be used to calculate or infer a fuel property - for example, the detected parameter(s) may be shaft speed and the mass flow rate of the fuel, from which the calorific value (a fuel property) can be determined, or the detected parameter(s) may be fuel density and / or the presence of a tracer, both of which are fuel properties in themselves. Determining at least one fuel property may include obtaining stored fuel property data. Chemically and / or physically determining one or more parameters of the fuel in the fuel tank may be performed by extracting a fuel sample from the fuel tank for off-wing testing.

[0106] Determining at least one fuel property of the fuel may include obtaining at least one fuel property of any fuel already present in the fuel tank before fuel replenishment; determining at least one fuel property of the fuel added to the fuel tank during fuel replenishment; and calculating at least one fuel property of the resulting fuel in the fuel tank after fuel replenishment (based on this information).

[0107] Determining at least one fuel property may be performed based on the detection of at least one fuel property. The fuel property may be a fuel characteristic or may be used to calculate or otherwise determine (e.g., by retrieving from a look-up table) a fuel property. The detection may be performed on-wing.

[0108] Determining at least one fuel property may be performed based on received fuel composition data. The fuel composition data may be provided to the aircraft during fuel replenishment. The fuel composition data may be manually input.

[0109] During at least one of taxiing, take-off and climb of an aircraft, at least one fuel characteristic can be inferred from the performance of a gas turbine engine.

[0110] According to a fifth aspect, there is provided a method of operating a gas turbine engine, the gas turbine engine comprising:

[0111] An engine core including a turbine, a compressor, a burner arranged to combust fuel, and a shaft connecting the turbine to the compressor;

[0112] A fan located upstream of the engine core;

[0113] A gearbox that receives an input from the shaft and outputs drive to the fan so as to drive the fan at a rotational speed lower than that of the shaft;

[0114] An oil circuit system arranged to supply oil to the gearbox; and

[0115] A heat exchange system including:

[0116] An air-oil heat exchanger through which oil in the oil circuit system flows;

[0117] And

[0118] A fuel-oil heat exchanger through which oil and fuel in the oil circuit system flow such that heat is transferred between the oil and the fuel, and wherein the oil circuit system branches such that a certain proportion of the oil can flow along each branch, and the air-oil heat exchanger and the fuel-oil heat exchanger are arranged in a parallel configuration on different branches of the oil circuit system; and

[0119] A modulating valve arranged to allow changing the proportion of oil delivered via each branch,

[0120] The method includes controlling the heat exchange system such that, under cruise conditions, the heat transfer ratio:

[0121]

[0122] Is in the range of 0 to 0.67.

[0123] The present inventors have recognized that using a fuel different from traditional kerosene-based jet fuels, such as sustainable aviation fuel, can result in different fuel properties and that parameters under cruise conditions can be adjusted to take advantage of the different fuel properties. In particular, some fuels can be heated to a higher temperature than traditional fuels in one or more fuel-oil heat exchangers without significantly increasing coking or other fuel decomposition product deposition pathways (such as lacquering). This can enable a method of providing improved oil cooling (since the fuel is able to absorb more heat), and can also improve the overall thermal / thermodynamic efficiency of the engine, with less heat loss to the surrounding environment and potentially also recovering more power in the thermodynamic cycle. A controllable heat exchange system plays a key role in managing the heat transfer ratio.

[0124] While it should be understood that the ratio is dimensionless, in the examples described herein, heat transfer is measured per unit mass or volume of fuel, thus providing a heat transfer rate normalized for changes in fuel flow during cruise. For heat transfer from oil to air, the definition of "per unit mass or volume of fuel" can be equivalent to "per set time period depending on fuel flow" to similarly provide normalization of fuel flow. It should be understood that heat transfer using the heat exchanger must be completed before the fuel reaches the burner. Any additional temperature rise in the burner itself is due to combustion, not the heat exchange system. It should be understood that the heat transfer per unit mass (kg) of fuel reaching the burner is recorded in order to adjust the fuel flow and any recirculation through one or more heat exchangers or bypasses of one or more heat exchangers, as described elsewhere herein. Thus, the amount of heat transferred to the fuel can be calculated based on a comparison of the fuel temperature as it approaches or enters the burner with the fuel temperature in the aircraft's fuel tank. In most embodiments, any temperature rise of the fuel due to other engine components (as opposed to heat transfer from the oil) can be assumed to be minimal. The amount of heat transferred from the oil to the air can be determined by the temperature drop of the oil across the air-oil heat exchanger, or by comparing the heat gained by the fuel with the total heat loss of the oil and assuming that the difference is due to heat loss from the oil to the air (any other sources of loss can be neglected).

[0125] The method can include controlling the heat exchange system such that, under cruise 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.

[0126] Controlling the heat exchange system to adjust the heat transfer ratio can include reducing the amount of oil delivered via at least one air-oil heat exchanger when the heat transfer ratio is too high.

[0127] The heat exchange system may include at least one bypass pipe arranged to allow oil to bypass the heat exchanger, and controlling the heat exchange system to adjust the heat transfer ratio may include modulating the amount of oil delivered via the bypass pipe.

[0128] The heat exchange system may include at least one recirculation pipe arranged to allow fluid (oil or fuel) to pass through the heat exchanger multiple times, and controlling the heat exchange system to adjust the heat transfer ratio may include modulating the amount of fluid delivered via the recirculation pipe.

[0129] The heat exchange system may include a refrigeration cycle device. The method may include using the refrigeration cycle device to provide a heat boost by transferring more heat from the oil to the fuel, optionally raising the fuel temperature to be higher than the oil temperature. The heat transfer ratio for an example of using the refrigeration cycle device may be in the range of 0 to 0.40.

[0130] In embodiments where the refrigeration cycle device is not used / the heat exchange system is not arranged to provide a heat boost, the heat transfer ratio may be in the range of 0.38 to 0.67.

[0131] The method may include controlling the heat exchange system under cruise conditions such that if the fuel temperature at the burner inlet is at least 160 °C, the heat transfer ratio is in the range of 0 to 0.2.

[0132] The method may include controlling the heat exchange system under cruise conditions such that if the fuel temperature at the burner inlet is at least 180 °C, the heat transfer ratio is in the range of 0 to 0.1.

[0133] The method may include controlling the heat exchange system under cruise conditions such that if the fuel is at least 70% sustainable aviation fuel, the heat transfer ratio is in the range of 0 to 0.2.

[0134] The method may include controlling the heat exchange system under cruise conditions such that if the fuel is at least 80% sustainable aviation fuel, the heat transfer ratio is in the range of 0 to 0.1.

[0135] The method may include maintaining the heat transfer rate from the oil to the air within the range of 0 kJ to 240 kJ per kilogram of fuel under cruise conditions, and optionally maintaining it within 0 kJ / kg to 120 kJ / kg, where no more than 20% of the heat transferred away from the oil during cruise is transferred to the air.

[0136] The method may include maintaining the heat transfer rate from the oil to the fuel within the range of 85 kJ to 350 kJ per kilogram of fuel under cruise conditions, and optionally maintaining it within 85 kJ / kg to 170 kJ / kg, where at least 80% of the heat transferred away from the oil during cruise is transferred to the fuel.

[0137] The methods of the first, third, and fifth aspects can be complementary, and any two or more of them can be performed together in various embodiments. The method of the fifth aspect can be performed using the engine of the second or fourth aspect.

[0138] According to a sixth aspect, there is provided a gas turbine engine for an aircraft, the engine comprising:

[0139] An engine core including a turbine, a compressor, and a spool connecting the turbine to the compressor;

[0140] A fan located upstream of the engine core; and

[0141] A gearbox that receives an input from the spool and outputs a drive to the fan to drive the fan at a rotational speed lower than that of the spool;

[0142] An oil circuit system arranged to supply oil to the gearbox; and

[0143] A heat exchange system including:

[0144] An air-oil heat exchanger through which the oil in the oil circuit system flows;

[0145] A fuel-oil heat exchanger through which the oil and fuel in the oil circuit system flow, such that heat is transferred between the oil and the fuel, and wherein the oil circuit system branches such that a certain proportion of the oil can flow along each branch, and the air-oil heat exchanger and the fuel-oil heat exchanger are arranged in parallel configuration on different branches of the oil circuit system; and

[0146] A modulating valve arranged to allow changing the proportion of the oil delivered via each branch,

[0147] And wherein the heat exchange system is arranged to be controlled such that under cruise conditions, the heat transfer ratio:

[0148]

[0149] Is in the range of 0 to 0.67.

[0150] The turbine can be a first turbine, the compressor can be a first compressor, and the spool can be a first spool. The engine core may further include a second turbine, a second compressor, and a second spool connecting the second turbine to the second compressor. The second turbine, the second compressor, and the second spool may be arranged to rotate at a rotational speed higher than that of the first spool.

[0151] The heat exchange system may further include a branched fuel return passage and at least one valve that controls the diversion of the fuel flow. The branched passage may be arranged to return fuel from the heat exchange system to at least two different locations along the main fuel path, from which the fuel enters the gas turbine engine to reach the burner.

[0152] The engine of the sixth aspect may be arranged to perform the methods of the first, third, and / or fifth aspects and may have any one of the features described in any of the foregoing aspects.

[0153] According to a seventh aspect, there is provided a method of operating a gas turbine engine for an aircraft, the gas turbine engine comprising:

[0154] An engine core including a turbine, a compressor, a burner arranged to burn fuel, and a shaft connecting the turbine to the compressor;

[0155] A fan located upstream of the engine core;

[0156] A gearbox that receives an input from the shaft and outputs drive to the fan so as to drive the fan at a rotational speed lower than that of the shaft;

[0157] An oil circuit system arranged to supply oil to the gearbox; and

[0158] A heat exchange system including:

[0159] An air-oil heat exchanger through which the oil in the oil circuit system flows;

[0160] And

[0161] A fuel-oil heat exchanger through which the oil and fuel in the oil circuit system flow such that heat is transferred between the oil and the fuel, and wherein the oil circuit system branches such that a certain proportion of the oil can flow along each branch, and the air-oil heat exchanger and the fuel-oil heat exchanger are arranged in parallel configuration on different branches of the oil circuit system; and

[0162] A modulating valve arranged to allow the proportion of oil delivered via each branch to be changed,

[0163] The method includes controlling the heat exchange system such that, under idle conditions, the heat transfer ratio:

[0164]

[0165] Is in the range of 0.67 to 5.67.

[0166] The heat transfer rate is defined as discussed above with respect to the fifth aspect. The idle operation of the aircraft during ground operation may be referred to as "ground idle", and the idle operation of the aircraft during flight may be referred to as "flight idle". All options of this aspect described below may be assumed to be related to ground idle conditions. The thrust at flight idle is typically slightly higher than that at ground idle. In some embodiments, only a relatively narrow range may apply to the flight idle of a particular engine. In other embodiments, all options described below for this aspect also apply to flight idle conditions.

[0167] As discussed with respect to the fifth aspect, the inventors have recognized that using a fuel different from traditional kerosene-based jet fuels (such as sustainable aviation fuel) can result in different fuel properties, and the operating parameters can be adjusted to take advantage of the different fuel properties. In particular, some fuels can be heated to a higher temperature than traditional fuels in one or more fuel-oil heat exchangers without significantly increasing coking. This can enable a method of providing improved oil cooling (since the fuel can absorb more heat), and can also improve the overall thermal efficiency of the engine while losing less heat to the surrounding environment. A controllable heat exchange system plays a key role in managing the heat transfer ratio. Additionally, the inventors have recognized that while cruise conditions typically account for most of the aircraft engine operating time, idle operation is also important because the fuel mass flow rate at idle is much lower than that at cruise, and even though the heat load of the fuel is relatively small, it can still cause a temperature increase - thus, the use of non-traditional fuels may have a greater impact on the optimal thermal management method under idle conditions (e.g., when the aircraft starts, when it is stationary during boarding and (taxiing towards the runway or hangar, or between other ground positions) and at certain times during flight (e.g., when starting to land)). Since the operating conditions between cruise and idle are very different - in terms of altitude and the desired thrust output of the engine - different control of the heat exchange system is appropriate.

[0168] The method may include controlling the heat exchange system under idle conditions such that the heat transfer ratio is less than 5.50, and optionally less than 5.0, 4.5, or 4.0. The method may include controlling the heat exchange system such that under idle conditions, the heat transfer ratio is greater than 0.75, and optionally greater than 1.0 or 1.5.

[0169] The step of controlling the heat exchange system to adjust the heat transfer ratio may be or include controlling a modulating valve that is arranged to change the proportion of oil delivered via each branch of the oil circuit system.

[0170] The step of controlling the heat exchange system to adjust the heat transfer ratio may include reducing the amount of oil delivered via at least one air-oil heat exchanger when the heat transfer ratio is too high.

[0171] The heat exchange system may include at least one bypass pipe, and the at least one bypass pipe is arranged to allow oil or fuel to bypass one or more heat exchangers, and the step of controlling the heat exchange system to adjust the heat transfer ratio may include modulating the amount of oil delivered via the bypass pipe.

[0172] The heat exchange system may include at least one recirculation pipe, and the at least one recirculation pipe is arranged to allow a fluid (e.g., oil or fuel) to pass through one or more heat exchangers multiple times. The step of controlling the heat exchange system to adjust the heat transfer ratio may include modulating the amount of fluid delivered via the recirculation pipe.

[0173] The heat exchange system may include a refrigeration cycle device, and the method may further include using the refrigeration cycle device to provide a heat boost by transferring more heat from the oil to the fuel, optionally raising the fuel temperature to be higher than the oil temperature. In such embodiments, the heat transfer ratio is in the range of 0.67 to 4.

[0174] In other embodiments, the heat exchange system is not arranged to provide a heat boost such that the fuel temperature cannot be raised to be higher than the maximum temperature of the oil. In such embodiments, the heat transfer ratio may be in the range of 2.00 to 5.67, and optionally in the range of 3.37 to 5.67.

[0175] The method may include controlling the heat exchange system under idle conditions such that when the fuel temperature at the burner inlet is lower than 200 °C, the heat transfer ratio is in the range of 2.33 to 5.67.

[0176] The method may include controlling the heat exchange system under idle conditions such that:

[0177] (i) when the fuel temperature at the burner inlet is higher than 200 °C, the heat transfer ratio is in the range of 0.67 to 4;

[0178] (ii) when the fuel temperature at the burner inlet is higher than 250 °C, the heat transfer ratio is in the range of 0.67 to 2.67; and / or

[0179] (iii) when the fuel is higher than 280 °C, the heat transfer ratio is in the range of 0.67 to 1.22.

[0180] The method may include controlling the heat exchange system under idle conditions such that if the fuel is at least 70% sustainable aviation fuel, the heat transfer ratio is in the range of 0.67 to 3.67.

[0181] The method may include controlling the heat exchange system under idle conditions such that if the fuel is at least 80% sustainable aviation fuel, the heat transfer ratio is in the range of 0.67 to 2.67.

[0182] The methods of the first, third, fifth, and seventh aspects may be complementary, and any two or more of them may be performed together in various embodiments. The method of the seventh aspect may be performed using the engine of the second, fourth, or sixth aspect.

[0183] According to an eighth aspect, there is provided an engine core including a turbine, a compressor, and a shaft connecting the turbine to the compressor;

[0184] a fan located upstream of the engine core; and

[0185] a gearbox that receives an input from the shaft and outputs drive to the fan to drive the fan at a rotational speed lower than that of the shaft;

[0186] an oil circuit system arranged to supply oil to the gearbox; and

[0187] a heat exchange system including:

[0188] an air-oil heat exchanger through which the oil in the oil circuit system flows;

[0189] a fuel-oil heat exchanger through which the oil and fuel in the oil circuit system flow, such that heat is transferred between the oil and the fuel, and wherein the oil circuit system branches such that a certain proportion of the oil can flow along each branch, and the air-oil heat exchanger and the fuel-oil heat exchanger are arranged in a parallel configuration on different branches of the oil circuit system; and

[0190] a modulating valve arranged to allow changing the proportion of the oil delivered via each branch,

[0191] and wherein the heat exchange system is arranged to be controlled such that under idle conditions, the heat transfer ratio:

[0192]

[0193] is in the range of 0.67 to 5.67.

[0194] The turbine may be a first turbine, the compressor may be a first compressor, and the shaft may be a first shaft. The engine core may further include a second turbine, a second compressor, and a second shaft connecting the second turbine to the second compressor. The second turbine, second compressor, and second shaft may be arranged to rotate at a rotational speed higher than that of the first shaft.

[0195] The heat exchange system may further include a branched fuel return passage and at least one valve for controlling the diversion of the fuel flow. The branched passage may be arranged to return fuel from the heat exchange system to at least two different locations along the main fuel path, from which the fuel enters the gas turbine engine to reach the burner.

[0196] The engine of the eighth aspect may be arranged to perform the methods of the first, third, fifth, and / or seventh aspects, and may have any one of the features described in any of the foregoing aspects.

[0197] According to a ninth aspect, there is provided a method of operating a gas turbine engine of an aircraft, the gas turbine engine comprising:

[0198] An engine core including a turbine, a compressor, a burner arranged to burn fuel, and a shaft connecting the turbine to the compressor;

[0199] A fan located upstream of the engine core;

[0200] A gearbox that receives an input from the shaft and outputs drive to the fan so as to drive the fan at a rotational speed lower than that of the shaft;

[0201] An oil circuit system arranged to supply oil to the gearbox; and

[0202] A heat exchange system including:

[0203] An air-oil heat exchanger through which the oil in the oil circuit system flows;

[0204] A fuel-oil heat exchanger through which the oil and fuel in the oil circuit system flow such that heat is transferred between the oil and the fuel, and wherein the oil circuit system includes at least one bypass pipe arranged to allow the oil to bypass one of the air-oil heat exchanger and the fuel-oil heat exchanger; and a bypass valve arranged to allow changing the proportion of the oil conveyed via the bypass pipe.

[0205] The method includes controlling the bypass valve such that, under cruise conditions, the heat transfer ratio:

[0206]

[0207] Is in the range of 0 to 0.67.

[0208] The inventors recognize that the principles described with respect to the fifth through eighth aspects can also be applied to engines that do not have branched oil circuit passages with different heat exchangers on different branches; for example, these principles can be applied to engines with heat exchangers in a substantially linear, series arrangement, and wherein one or more bypass tubes are used as an alternative to, rather than as an optional supplement to, the branched main passage. The introduction of existing bypass tubes and / or improved control can achieve a method of providing improved oil cooling (since the fuel to be used can carry away more heat than conventional fuels), and can also improve the overall thermal efficiency of the engine while losing less heat to the surrounding environment. A controllable bypass valve plays a key role in controlling the heat transfer ratio.

[0209] The heat transfer rate used to calculate the heat transfer ratio is defined as discussed above with respect to the fifth and seventh aspects.

[0210] The method may include controlling the bypass valve such that, under cruise conditions, the heat transfer ratio is in the range from 0 to 0.60, optionally from 0 to 0.50, from 0 to 0.40, from 0 to 0.30, or from 0 to 0.20, and further optionally from 0 to 0.10.

[0211] The step of controlling the bypass valve to adjust the heat transfer ratio may include reducing the amount of oil delivered via at least one air - oil heat exchanger when the heat transfer ratio is too high.

[0212] In some embodiments, the oil circuit system may include a single bypass tube that is arranged to allow oil to bypass the air - oil heat exchanger. A bypass tube may not be provided for the fuel - oil heat exchanger. It should be understood that in most operating cases, in order to improve the engine thermal efficiency, it is desirable to transfer heat from the oil to the fuel as safely as possible, and thus bypassing the fuel - oil heat exchanger can be avoided.

[0213] The heat exchange system may include at least two oil bypass tubes, each bypass tube being arranged to allow oil to bypass one of the air - oil heat exchanger and the fuel - oil heat exchanger. The method may include modulating the amount of oil delivered via each bypass tube. The same bypass valve (which may be a three - way valve) may be used to control the flow through both bypass tubes, or different bypass valves may be provided for each bypass tube.

[0214] The heat exchange system may include at least one recirculation pipe arranged to allow fluid to pass through the heat exchanger multiple times. The method of such embodiments may also include modulating the amount of fluid conveyed via the recirculation pipe so as to adjust the heat transfer ratio. Additionally or alternatively, the heat exchange system may also include a refrigeration cycle device, and the method of such embodiments may also include using the refrigeration cycle device to provide a heat boost by transferring more heat from the oil to the fuel, optionally raising the fuel temperature to be higher than the oil temperature. Therefore, it should be understood that, in addition to the bypass valve, one or more other components may also be used to adjust the heat transfer ratio.

[0215] In embodiments having a refrigeration cycle device, the bypass valve may be controlled such that the heat transfer ratio is in the range of 0 to 0.40. In embodiments without a refrigeration cycle device, in these embodiments where the heat exchange system is not arranged to provide a heat boost, the bypass valve may be controlled such that the heat transfer ratio is in the range of 0.38 to 0.67.

[0216] The method may include controlling the bypass valve under cruise conditions such that:

[0217] (i) if the fuel temperature at the burner inlet is at least 160 °C, the heat transfer ratio is in the range of 0 to 0.2; and / or

[0218] (ii) if the fuel temperature at the burner inlet is at least 180 °C, the heat transfer ratio is in the range of 0 to 0.1.

[0219] In some embodiments, when determining how to control the bypass valve, one or more fuel characteristics may be considered. For example, the method may include controlling the bypass valve under cruise conditions such that:

[0220] (i) if the fuel is at least 70% sustainable aviation fuel, the heat transfer ratio is in the range of 0 to 0.2;

[0221] (ii) if the fuel is at least 80% sustainable aviation fuel, the heat transfer ratio is in the range of 0 to 0.1;

[0222] The method may include controlling the bypass valve under cruise conditions such that the heat transfer rate from the oil to the air under cruise conditions can be maintained in the range of 0 kJ to 240 kJ per kilogram of fuel, no more than 20% of the heat transferred away from the oil under cruise conditions is transferred to the air, and / or such that the heat transfer rate from the oil to the fuel under cruise conditions is maintained in the range of 85 kJ to 350 kJ per kilogram of fuel, and at least 80% of the heat transferred away from the oil during cruise is transferred to the fuel.

[0223] The methods of the first, third, fifth, seventh, and ninth aspects may be complementary, and any two or more of them may be performed together in various embodiments. The method of the ninth aspect may be performed using the engine of the second, fourth, sixth, or eighth aspect.

[0224] According to a tenth aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising:

[0225] An engine core including a turbine, a compressor, and a spindle connecting the turbine to the compressor;

[0226] A fan located upstream of the engine core;

[0227] A gearbox that receives an input from the spindle and outputs drive to the fan to drive the fan at a rotational speed lower than that of the spindle;

[0228] An oil circuit system arranged to supply oil to the gearbox; and

[0229] A heat exchange system including:

[0230] An air-oil heat exchanger through which the oil in the oil circuit system flows;

[0231] A fuel-oil heat exchanger through which the oil and fuel in the oil circuit system flow, such that heat is transferred between the oil and the fuel,

[0232] And wherein the oil circuit system includes at least one bypass pipe arranged to allow oil to bypass one of the air-oil heat exchanger and the fuel-oil heat exchanger; and a bypass valve arranged to allow changing the proportion of oil conveyed via the bypass pipe.

[0233] The bypass valve is arranged to be controlled such that under cruise conditions, the heat transfer ratio:

[0234]

[0235] Is in the range of 0 to 0.67.

[0236] The turbine may be a first turbine, the compressor may be a first compressor, and the spindle may be a first spindle. The engine core may further include a second turbine, a second compressor, and a second spindle connecting the second turbine to the second compressor; and the second turbine, second compressor, and second spindle may be arranged to rotate at a rotational speed higher than that of the first spindle.

[0237] The heat exchange system may also include a branched fuel return passage and at least one valve for controlling the diversion of the fuel flow, the branched passages being arranged to return fuel from the heat exchange system to at least two different locations along the main fuel path, from which the fuel enters the gas turbine engine to reach the burner.

[0238] The engine of the tenth aspect may be arranged to perform the methods of the first, third, fifth, seventh, and / or ninth aspects and may have any one of the features described in any of the foregoing aspects.

[0239] According to the eleventh aspect, there is provided a method of operating a gas turbine engine for an aircraft, the gas turbine engine comprising:

[0240] An engine core including a turbine, a compressor, a burner arranged to burn fuel, and a shaft connecting the turbine to the compressor;

[0241] A fan located upstream of the engine core;

[0242] A gearbox that receives an input from the shaft and outputs drive to the fan so as to drive the fan at a rotational speed lower than that of the shaft;

[0243] An oil circuit system arranged to supply oil to the gearbox; and

[0244] A heat exchange system including:

[0245] An air-oil heat exchanger through which the oil in the oil circuit system flows;

[0246] And

[0247] A fuel-oil heat exchanger through which the oil and fuel in the oil circuit system flow such that heat is transferred between the oil and the fuel, and wherein the oil circuit system includes at least one bypass pipe arranged to allow the oil to bypass one of the air-oil heat exchanger and the fuel-oil heat exchanger; and a bypass valve arranged to allow the proportion of the oil conveyed via the bypass pipe to be changed.

[0248] The method includes controlling the bypass valve such that under idle conditions, the heat transfer ratio:

[0249]

[0250] Is in the range of 0.67 to 5.67.

[0251] The inventors recognize that the principles described with respect to the fifth through eighth aspects can also be applied to engines that do not have branched oil circuit passages with different heat exchangers on different branches; for example, these principles can be applied to engines with a substantially linear, series arrangement of heat exchangers, and wherein one or more bypass tubes are used as an alternative to, rather than as an optional supplement to, the branched main passage. The introduction of existing bypass tubes and / or improved control can achieve a method of providing improved oil cooling, and can also improve the overall thermal efficiency of the engine while losing less heat to the surrounding environment. A controllable bypass valve plays a key role in controlling the heat transfer ratio. In addition, the inventors recognize that while cruise conditions typically account for most of the operating time of an aircraft engine, idle operation is also important because the fuel mass flow rate at idle is much lower than at cruise, and even though the thermal load of the fuel is relatively small, it can still cause temperature increases - thus, the use of non-conventional fuels may have a greater impact on the optimal thermal management method under idle conditions (e.g., when the aircraft is starting, when stationary during boarding, and (taxiing towards the runway or hangar, or between other ground positions) or during flight landing). Since the operating conditions between cruise and idle are very different in terms of altitude and the desired thrust output of the engine, different control of the bypass valve is appropriate.

[0252] The heat transfer rate is defined as discussed above with respect to the fifth aspect. Idle operation when the aircraft is on the ground can be referred to as "ground idle", and idle operation during flight can be referred to as "flight idle". All options of this aspect described below can be assumed to be related to ground idle conditions. Flight idle is typically at a slightly higher thrust, and in some embodiments, only a less restricted range may apply to flight idle for a particular engine.

[0253] The method can include controlling the bypass valve such that under idle conditions (and more specifically under ground idle in some embodiments), the heat transfer ratio is less than 5.50, and optionally less than 5.0, or less than 4.5. The method can include controlling the bypass valve such that under idle conditions, the heat transfer ratio is greater than 1.0, and optionally greater than 1.5, 2.0, 2.5, or 3.0. The method can include controlling the bypass valve such that under idle conditions, the heat transfer ratio is greater than 0.75, and optionally greater than 1.0 or 1.5.

[0254] Controlling the bypass valve to adjust the heat transfer ratio can include reducing the amount of oil delivered via at least one air - oil heat exchanger when the heat transfer ratio is too high.

[0255] The heat exchange system can include at least two oil bypass tubes, each bypass tube being arranged to allow oil to bypass a respective one of the air - oil heat exchanger and the fuel - oil heat exchanger. The method can include modulating the amount of oil delivered via each bypass tube.

[0256] The heat exchange system may include at least one recirculation pipe arranged to allow fluid to pass through the heat exchanger multiple times. The method may also include modulating the amount of fluid delivered via the recirculation pipe to adjust the heat transfer ratio.

[0257] The heat exchange system may include a refrigeration cycle device. The method may include using the refrigeration cycle device to provide a heat boost by transferring more heat from the oil to the fuel (more than the heat transferred in the heat exchanger), optionally raising the fuel temperature above the oil temperature. In such embodiments, the method may include controlling a bypass valve such that the heat transfer ratio is in the range of 0.67 to 4. In embodiments where the heat exchange system is not arranged to provide a heat boost, the method may include controlling the bypass valve such that the heat transfer ratio is in the range of 2.00 to 5.67, and optionally in the range of 3.37 to 5.67.

[0258] The method may include controlling the bypass valve under idle conditions such that the heat transfer ratio is at least one of the following:

[0259] (i) in the range of 2.33 to 5.67 when the fuel temperature at the burner inlet is below 200 °C;

[0260] (ii) in the range of 0.67 to 5.00, and optionally in the range of 0.67 to 4.00, when the fuel temperature at the inlet of the burner (16) is above 200 °C;

[0261] (iii) in the range of 0.67 to 4.00, and optionally in the range of 0.67 to 2.67, when the fuel temperature at the inlet of the burner (16) is above 250 °C; and

[0262] (iv) in the range of 0.67 to 2.33, and optionally in the range of 0.67 to 1.22, when the fuel temperature at the inlet of the burner (16) is above 280 °C.

[0263] The method may take into account the SAF content of the fuel. The method may include controlling the bypass valve such that if the fuel is at least 70% sustainable aviation fuel, the heat transfer ratio is in the range of 0.67 to 3.67, and / or if the fuel is at least 80% sustainable aviation fuel, the heat transfer ratio is in the range of 0.67 to 2.67.

[0264] The methods of the first, third, fifth, seventh, ninth, and eleventh aspects may be complementary, and any two or more of them may be performed together in various embodiments. The method of the eleventh aspect may be performed using the engine of the second, fourth, sixth, eighth, or tenth aspect.

[0265] According to a twelfth aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising:

[0266] an engine core including a turbine, a compressor and a spool connecting the turbine to the compressor;

[0267] a fan located upstream of the engine core; and

[0268] a gearbox receiving an input from the spool and driving an output to the fan so as to drive the fan at a rotational speed lower than that of the spool;

[0269] an oil circuit system arranged to supply oil to the gearbox; and

[0270] a heat exchange system including:

[0271] an air-oil heat exchanger through which the oil in the oil circuit system flows;

[0272] and

[0273] a fuel-oil heat exchanger through which the oil and fuel in the oil circuit system flow such that heat is transferred between the oil and the fuel,

[0274] and wherein the oil circuit system includes at least one bypass pipe arranged to allow oil to bypass one of the air-oil heat exchanger and the fuel-oil heat exchanger; and a bypass valve arranged to allow a change in the proportion of oil conveyed via the bypass pipe,

[0275] and wherein the bypass valve is arranged to be controlled such that under idle conditions, the heat transfer ratio:

[0276]

[0277] is in the range of 0.67 to 5.67.

[0278] The turbine may be a first turbine, the compressor may be a first compressor, and the spool may be a first spool. The engine core may further include a second turbine, a second compressor and a second spool connecting the second turbine to the second compressor. The second turbine, second compressor and second spool may be arranged to rotate at a rotational speed higher than that of the first spool.

[0279] The heat exchange system may further include a branched fuel return passage and at least one valve controlling fuel flow diversion, the branched passages being arranged to return fuel from the heat exchange system to at least two different positions along a main fuel path from which fuel enters the gas turbine engine to reach a burner.

[0280] The engine of the twelfth aspect may be arranged to perform the methods of the first, third, fifth, seventh, ninth, and / or eleventh aspects and may have any one of the features described in any of the foregoing aspects.

[0281] It should be understood that the features described for one aspect may be used in combination with any other aspect with necessary modifications.

[0282] As described elsewhere herein, the present disclosure may be applied to any relevant configuration of a gas turbine engine. Such a gas turbine engine may be, for example, a turbofan gas turbine engine, an open rotor gas turbine engine (where the propeller is not enclosed by a nacelle), a turboprop engine, or a turbojet engine. Any such engine may or may not be provided with a afterburner. Such a gas turbine engine may be configured, for example, for land - or sea - based power generation applications.

[0283] A gas turbine engine according to any aspect of the present disclosure may include an engine core that includes a turbine, a combustor, a compressor, and a shaft connecting the turbine to the compressor. Such a gas turbine engine may include a fan (with fan blades). Such a fan may be located upstream of the engine core. Alternatively, in some examples, the gas turbine engine may include a fan located downstream of the engine core, for example, in the case where the gas turbine engine is an open rotor or a turboprop engine (in which case the fan may be referred to as a propulsor).

[0284] In the case where the gas turbine engine is an open rotor or a turboprop engine, the gas turbine engine may include two contra - rotating propeller stages attached to and driven by a free power turbine via a shaft. The propulsors may rotate in opposite directions such that one propulsor rotates clockwise about the engine's axis of rotation and the other propulsor rotates counter - clockwise about the engine's axis of rotation. Alternatively, the gas turbine engine may include a propulsor stage and a stator vane stage constructed downstream of the propulsor stage. The stator vane stage may have variable pitch. Thus, the high - pressure, intermediate - pressure, and free power turbines may drive the high - pressure and intermediate - pressure compressors and the propulsors respectively via suitable interconnecting shafts. Thus, the propulsors may provide most of the propulsive thrust.

[0285] In the case where the gas turbine engine is an open rotor or a turboprop engine, one or more propulsor stages may be driven by a gearbox. The gearbox may be of the type described herein.

[0286] An engine according to the present disclosure may be a turbofan engine. Such an engine may be a direct drive turbofan engine, where the fan is directly connected to the fan drive turbine via a spindle, for example without a gearbox. In such a direct drive turbofan engine, it can be said that the fan rotates at the same rotational speed as the fan drive turbine. By way of example only, the fan drive turbine may be a first turbine, the spindle may be a first spindle, and the gas turbine engine may further include a second turbine and a second spindle connecting the second turbine to the compressor. The second turbine, compressor, and second spindle may be arranged to rotate at a higher rotational speed than the first spindle. In such an arrangement, the second turbine may be axially positioned upstream of the first turbine.

[0287] An engine according to the present disclosure may be a geared turbofan engine. In such an arrangement, the engine has a fan driven via a gearbox. Thus, such a gas turbine engine may include a gearbox that receives an input from a spindle and outputs a drive to the fan so as to drive the fan at a lower rotational speed than the spindle. The input to the gearbox may come directly from the spindle or indirectly from the spindle, for example via spur shafts and / or gears. The spindle may rigidly connect the turbine and the compressor such that the turbine and the compressor rotate at the same speed (wherein the fan rotates at a lower speed).

[0288] The gas turbine engine as described and / or claimed herein may have any suitable general architecture. For example, the gas turbine engine may have any desired number of shafts connecting the turbine and the compressor, such as one shaft, two shafts, or three shafts. By way of example only, the turbine connected to the spindle may be a first turbine, the compressor connected to the spindle may be a first compressor, and the spindle may be a first spindle. The engine core may further include a second turbine, a second compressor, and a second spindle connecting the second turbine to the second compressor. The second turbine, second compressor, and second spindle may be arranged to rotate at a higher rotational speed than the first spindle.

[0289] In such an arrangement, the second compressor may be axially positioned downstream of the first compressor. The second compressor may be arranged to receive flow (e.g., directly receive, e.g., via a generally annular duct) from the first compressor.

[0290] The gearbox may be arranged to be driven by a spindle (such as the first spindle in the above example) configured (e.g., in use) to rotate at the lowest rotational speed. For example, the gearbox may be arranged to be driven only by a spindle configured (e.g., in use) to rotate at the lowest rotational speed (e.g., in the above example, only by the first spindle and not the second spindle). Alternatively, the gearbox may be arranged to be driven by any one or more shafts, such as the first shaft and / or the second shaft in the above example.

[0291] The gearbox can be a reduction gearbox (since the output to the fan has a lower rotational speed than the input from the spool). Any type of gearbox can be used. For example, the gearbox can be a "planetary" or "stellar" gearbox, as described in more detail elsewhere herein. Such a gearbox can be single stage. Alternatively, such a gearbox can be a compound gearbox, such as a compound planetary gearbox (which can have an input on the sun gear and an output on the ring gear and is thus referred to as a "compound star" gearbox), for example having two stages of reduction.

[0292] The gearbox can have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example greater than 2.5, for example in the range of 3 to 4.2, or 3.2 to 3.8, for example, approximately or at least 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. For example, the gear ratio can be between any two values in the previous sentence. By way of example only, the gearbox can be a "stellar" gearbox having a reduction ratio in the range of 3.1 or 3.2 to 3.8. By way of a further example only, the gearbox can be a "stellar" gearbox having a reduction ratio in the range of 3.0 to 3.1. By way of a further example only, the gearbox can be a "planetary" gearbox having a reduction ratio in the range of 3.6 to 4.2. In some arrangements, the gear ratio can be outside of these ranges.

[0293] In any gas turbine engine as described and / or claimed herein, fuel of a given composition or blend is supplied to a combustor, which can be located downstream (e.g. axially downstream) of the fan and compressor with respect to the flow path. For example, in the case where a second compressor is provided, the combustor can be located directly downstream of the second compressor (e.g. at its outlet). By way of another example, in the case where a second turbine is provided, the flow at the combustor outlet can be supplied to the inlet of the second turbine. The combustor can be located upstream of one or more turbines.

[0294] The compressor or each compressor (e.g., the first and second compressors as described above) may include any number of stages, such as a plurality of stages. Each stage may include a row of rotor blades and a row of stator vanes, and the row of stator vanes may be variable stator vanes (since the angle of incidence of the row of stator vanes may be variable). The row of rotor blades and the row of stator vanes may be axially offset from each other. For example, a gas turbine engine may be a direct drive turbofan gas turbine engine including 13 or 14 compressor stages (in addition to the fan). Such an engine may include, for example, 3 stages in the first (or "low pressure") compressor and 10 or 11 stages in the second (or "high pressure") compressor. By way of a further example, a gas turbine engine may be a "geared" gas turbine engine including 11, 12 or 13 compressor stages (in addition to the fan) (where the fan is driven by a first shaft via a reduction gearbox). Such an engine may include 3 or 4 stages in the first (or "low pressure") compressor and 8 or 9 stages in the second (or "high pressure") compressor. By way of a further example, a gas turbine engine may be a "geared" gas turbine engine having 4 stages in the first (or "low pressure") compressor and 10 stages in the second (or "high pressure") compressor.

[0295] The turbine or each turbine (e.g., the first and second turbines as described above) may include any number of stages, such as a plurality of stages. Optionally, each stage may include a row of rotor blades and a row of stator vanes, and vice versa. The corresponding rows of rotor blades and stator vanes may be axially offset from each other. The second (or "high pressure") turbine may include 2 stages in any arrangement (e.g., regardless of whether it is a geared engine or a direct drive engine). A gas turbine engine may be a direct drive gas turbine engine including a first (or "low pressure") turbine having 5, 6 or 7 stages. Alternatively, a gas turbine engine may be a "geared" gas turbine engine including a first (or "low pressure") turbine having 3 or 4 stages.

[0296] Each fan blade can be defined as having a radial span that extends from a root (or hub) at a radially inner gas washing position or 0% span position to a tip at the 100% span position. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip can be less than (or approximately) any of the following: 0.4, 0.39, 0.38, 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip can be within the inclusive range defined by any two of the values in the previous sentence (i.e., these values can form an upper or lower limit), for example, in the range of 0.28 to 0.32 or 0.29 to 0.30. These ratios are generally referred to as the hub-tip ratio. The radius at the hub and the radius at the tip can both be measured at the leading edge (or axially foremost) portion of the blade. Of course, the hub-tip ratio refers to the gas washing portion of the fan blade, i.e., the portion radially outside of any platform.

[0297] The radius of the fan can be measured between the engine centerline and the tip at the leading edge of the fan blade. The fan diameter (which may simply be twice the fan radius) can be greater than (or approximately) any of the following: 140 cm, 170 cm, 180 cm, 190 cm, 200 cm, 210 cm, 220 cm, 230 cm, 240 cm, 250 cm (about 100 inches), 260 cm, 270 cm (about 105 inches), 280 cm (about 110 inches), 290 cm (about 115 inches), 300 cm (about 120 inches), 310 cm, 320 cm (about 125 inches), 330 cm (about 130 inches), 340 cm (about 135 inches), 350 cm, 360 cm (about 140 inches), 370 cm (about 145 inches), 380 cm (about 150 inches), 390 cm (about 155 inches), 400 cm, 410 cm (about 160 inches), or 420 cm (about 165 inches). The fan diameter can be within the inclusive range defined by any two of the values in the previous sentence (i.e., these values can form an upper or lower limit), for example, in the range of 210 cm to 240 cm, or 250 cm to 280 cm, or 320 cm to 380 cm. By way of non-limiting example only, the fan diameter can be in the range of 170 cm to 180 cm, 190 cm to 200 cm, 200 cm to 210 cm, 210 cm to 230 cm, 290 cm to 300 cm, or 340 cm to 360 cm.

[0298] The rotational speed of the fan can vary during use. Generally speaking, for fans with a larger diameter, the rotational speed is lower. By way of non-limiting example only, the rotational speed of the fan under cruise conditions can be less than 3500 rpm, such as less than 2600 rpm, or less than 2500 rpm, or less than 2300 rpm. By way of another non-limiting example only, for a "geared" gas turbine engine with a fan diameter in the range of 200 cm to 210 cm, the rotational speed of the fan under cruise conditions can be in the range of 2750 to 2900 rpm. By way of another non-limiting example only, for a "geared" gas turbine engine with a fan diameter in the range of 210 cm to 230 cm, the rotational speed of the fan under cruise conditions can be in the range of 2500 to 2800 rpm. By way of another non-limiting example only, for a "geared" gas turbine engine with a fan diameter in the range of 340 cm to 360 cm, the rotational speed of the fan under cruise conditions can be in the range of 1500 to 1800 rpm. By way of another non-limiting example only, for a direct drive engine with a fan diameter in the range of 190 cm to 200 cm, the rotational speed of the fan under cruise conditions can be in the range of 3600 to 3900 rpm. By way of another non-limiting example only, for a direct drive engine with a fan diameter in the range of 300 cm to 340 cm, the rotational speed of the fan under cruise conditions can be in the range of 2000 to 2800 rpm.

[0299] When using a gas turbine engine, the fan (with associated fan blades) rotates about a rotational axis. This rotation causes the tips of the fan blades to move at a speed U 尖端 . The work done by the fan blades on the flow results in an enthalpy rise dH of the flow. The fan tip loading can be defined as dH / U 尖端 2 , where dH is the enthalpy rise across the fan (e.g., 1-D average enthalpy rise), and U 尖端 is the (translational) speed 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 fan tip loading under cruise conditions can be greater than (or approximately) any of the following: 0.28, 0.29, 0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all values are dimensionless). The fan tip loading can be within the inclusive range defined by any two of the values in the previous sentence (i.e., these values can form an upper or lower limit), e.g., in the range of 0.28 to 0.31 or 0.29 to 0.3 (e.g., for a geared gas turbine engine).

[0300] A gas turbine engine according to the present disclosure may have any desired bypass ratio (BPR), where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core. In some arrangements, the bypass ratio at cruise conditions may be greater than (or approximately) any of the following: 9, 9.5, 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5, or 20. The bypass ratio at cruise conditions may be within an inclusive range defined by any two of the values in the previous sentence (i.e., these values may form an upper or lower limit), such as in the range of 12 to 16, or 13 to 15, or 13 to 14. By way of non-limiting example only, the bypass ratio at cruise conditions of a direct drive gas turbine engine according to the present disclosure may be in the range of 9:1 to 11:1. By way of another non-limiting example only, the bypass ratio at cruise conditions of a geared gas turbine engine according to the present disclosure may be in the range of 12:1 to 15:1. The bypass duct may be at least substantially annular. The bypass duct may be located radially outside of the core engine. The radially outer surface of the bypass duct may be defined by a nacelle and / or a fan casing.

[0301] 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 entering the combustor) to the stagnation pressure upstream of the fan. By way of non-limiting example only, the overall pressure ratio of a gas turbine engine as described and / or claimed herein at cruise conditions may be greater than (or approximately) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio may be within an inclusive range defined by any two of the values in the previous sentence (i.e., these values may form an upper or lower limit), such as in the range of 50 to 70. By way of non-limiting example only, the overall pressure ratio at cruise conditions of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 40 to 45. By way of non-limiting example only, the overall pressure ratio at cruise conditions of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 45 to 55. By way of non-limiting example only, the overall pressure ratio at cruise conditions of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 50 to 60. By way of non-limiting example only, the overall pressure ratio at cruise conditions of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 50 to 60.

[0302] The specific thrust of an engine can be defined as the net thrust of the engine divided by the total mass flow through the engine. In some examples, for a given thrust condition, the specific thrust can depend on the specific composition of the fuel provided to the combustor. Under cruise conditions, the specific thrust of the engines described and / or claimed herein can be less than (or approximately) any of the following: 110 N / kg -1 s, 105 N / kg -1 s, 100 N / kg -1 s, 95 N / kg -1 s, 90 N / kg -1 s, 85 N / kg -1 s or 80 N / kg -1 s. The specific thrust can be within the inclusive range defined by any two of the values in the previous sentence (i.e., these values can form an upper or lower limit), for example, between 80 N / kg -1 s and 100 N / kg -1 s, or between 85 N / kg -1 s and 95 N / kg -1 s. Compared to conventional gas turbine engines, such engines can be particularly efficient. By way of non-limiting example only, the specific thrust of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm can be between 90 N / kg -1 s and 95 N / kg -1 s. By way of non-limiting example only, the specific thrust of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm can be between 80 N / kg -1 s and 90 N / kg -1 s. By way of non-limiting example only, the specific thrust of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm can be between 70 N / kg -1 s and 90 N / kg -1 s. By way of non-limiting example only, the specific thrust of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm can be between 90 N / kg -1 s and 120 N / kg -1 s.

[0303] A gas turbine engine as described herein and / or claimed may have any desired maximum thrust. By way of non-limiting example only, a gas turbine as described herein and / or claimed may produce a maximum thrust that is at least (or approximately) any one of the following: 100 kN, 110 kN, 120 kN, 130 kN, 135 kN, 140 kN, 145 kN, 150 kN, 155 kN, 160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN, 450 kN, 500 kN, or 550 kN. The maximum thrust may be within an inclusive range defined by any two of the values in the previous sentence (i.e., these values may form an upper or lower limit). By way of non-limiting example only, a gas turbine as described herein and / or claimed may be capable of producing a maximum thrust within the range of 155 kN to 170 kN, 330 kN to 420 kN, or 350 kN to 400 kN. By way of non-limiting example only, a geared gas turbine engine having a fan diameter within the range of 200 cm to 210 cm may have a maximum thrust within the range of 140 kN to 160 kN. By way of non-limiting example only, a geared gas turbine engine having a fan diameter within the range of 210 cm to 230 cm may have a maximum thrust within the range of 150 kN to 200 kN. By way of non-limiting example only, a geared gas turbine engine having a fan diameter within the range of 340 cm to 360 cm may have a maximum thrust within the range of 370 kN to 500 kN. By way of non-limiting example only, a direct drive gas turbine engine having a fan diameter within the range of 300 cm to 340 cm may have a maximum thrust within the range of 370 kN to 500 kN. The thrust mentioned above may be the maximum net thrust at standard atmospheric conditions, at sea level, plus 15 °C (ambient pressure 101.3 kPa, temperature 30 °C), with the engine stationary.

[0304] In use, the temperature of the flow at the inlet of the high-pressure turbine can be particularly high. This temperature, which may be referred to as TET, can be measured at the outlet of the combustor, for example, just upstream of the first turbine blade which may itself be referred to as the nozzle guide vane. In some examples, for a given thrust condition, the TET can depend on the specific composition of the fuel provided to the combustor. Under cruise conditions, the TET can be at least (or approximately) any of the following: 1400K, 1450K, 1500K, 1520K, 1530K, 1540K, 1550K, 1600K or 1650K. Thus, by way of non-limiting example only, a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm can have a TET in the range of 1540K to 1600K under cruise conditions. By way of non-limiting example only, a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm can have a TET in the range of 1590K to 1650K under cruise conditions. By way of non-limiting example only, a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm can have a TET in the range of 1600K to 1660K under cruise conditions. By way of non-limiting example only, a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm can have a TET in the range of 1590K to 1650K under cruise conditions. By way of non-limiting example only, a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm can have a TET in the range of 1570K to 1630K under cruise conditions.

[0305] The TET under cruise conditions can be within an inclusive range defined by any two of the values in the previous sentence (i.e., these values can form an upper or lower limit), for example, from 1530K to 1600K. The maximum TET during engine operation can be, for example, at least (or approximately) any of the following: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K, 2000K, 2050K, or 2100K. Thus, by way of non-limiting example only, the maximum TET of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm can be in the range of 1890K to 1960K. By way of non-limiting example only, the maximum TET of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm can be in the range of 1890K to 1960K. By way of non-limiting example only, the maximum TET of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm can be in the range of 1890K to 1960K. By way of non-limiting example only, the maximum TET of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm can be in the range of 1935K to 1995K. By way of non-limiting example only, the maximum TET of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm can be in the range of 1890K to 1950K. The maximum TET can be within an inclusive range defined by any two of the values in the previous sentence (i.e., these values can form an upper or lower limit), for example, in the range of 1800K to 1950K or 1900K to 2000K. The maximum TET can occur, for example, under high thrust conditions, such as under maximum takeoff (MTO) conditions.

[0306] The fan blades and / or the airfoil portions of the fan blades described and / or claimed herein can be manufactured from any suitable material or combination of materials. For example, at least a portion of the fan blades and / or airfoils can be manufactured at least in part from a composite material, such as a metal matrix composite and / or an organic matrix composite, such as a carbon fiber composite. By way of further example, at least a portion of the fan blades and / or airfoils can be manufactured at least in part from a metal, such as a titanium-based metal or an aluminum-based material (such as an aluminum-lithium alloy) or a steel-based material. The fan blades can include at least two regions manufactured using different materials. For example, the fan blades can have a protective leading edge that can be manufactured using a material that better resists impacts (e.g., from birds, ice, or other materials) than the rest of the blade. Such leading edges can be manufactured, for example, using titanium or a titanium-based alloy. Thus, by way of example only, the fan blade can have a carbon fiber or an aluminum-based body (such as an aluminum-lithium alloy) with a titanium leading edge.

[0307] The fan as described and / or claimed herein may include a central portion from which fan blades may extend, for example, radially. The fan blades may be attached to the central portion in any desired manner. For example, each fan blade may include a fixture that may engage a corresponding slot in a hub (or disc portion). By way of example only, such a fixture may be in the form of a dovetail that may be inserted into and / or engage a corresponding slot in the hub / disc portion to secure the fan blade to the hub / disc portion. By way of a further example, the fan blades may be integrally formed with the central portion. Such an arrangement may be referred to as a bladed disc or a bladed ring. Any suitable method may be used to manufacture such a bladed disc or bladed ring. For example, at least a portion of the fan blades may be machined from a block, and / or at least portions of the fan blades may be attached to the hub / disc portion by welding (such as linear friction welding).

[0308] The gas turbine engine described and / or claimed herein may or may not be provided with a variable area nozzle (VAN). Such a variable area nozzle may allow the exit area of the bypass duct to vary in use. The general principles of the present disclosure may be applied to engines with or without a VAN.

[0309] The fan of a gas turbine as described and / or claimed herein may have any desired number of fan blades, such as 14, 16, 18, 20, 22, 24, or 26 fan blades. In the case where the fan blades have a carbon fiber composite body, there may be 16 or 18 fan blades. In the case where the fan blades have a metal body (such as an aluminum-lithium or titanium alloy), there may be 18, 20, or 22 fan blades.

[0310] As used herein, the terms idle, taxi, takeoff, climb, cruise, descent, approach, and landing (or one or more portions thereof) have their conventional meanings and will be readily understood by a person skilled in the art. Thus, for a given gas turbine engine for an aircraft, a person skilled in the art will immediately recognize that each term is used to refer to the overall or one or more portions of the operating phases of the engine of the aircraft to which the gas turbine engine is designed to be attached within a given mission.

[0311] At this point, ground idle may refer to an engine operating phase in which the aircraft is stationary and in contact with the ground, but in which there are requirements for the engine to be operated. During idle, the engine may produce between 3% and 9% of the available thrust of the engine. In additional non-limiting examples, the engine may produce between 5% and 8% of the available thrust. In additional non-limiting examples, the engine may produce between 6% and 7% of the available thrust. Taxiing may refer to an engine operating phase in which the aircraft is propelled along the ground by the thrust produced by the engine. During taxiing, the engine may produce between 5% and 15% of the available thrust. In additional non-limiting examples, the engine may produce between 6% and 12% of the available thrust. In additional non-limiting examples, the engine may produce between 7% and 10% of the available thrust. Takeoff may refer to an engine operating phase in which the aircraft is propelled by the thrust produced by the engine. During the initial phase of the takeoff phase, the aircraft may be propelled while in contact with the ground. During a later phase of the takeoff phase, the aircraft may be propelled while not in contact with the ground. During takeoff, the engine may produce between 90% and 100% of the available thrust. In additional non-limiting examples, the engine may produce between 95% and 100% of the available thrust. In additional non-limiting examples, the engine may produce 100% of the available thrust.

[0312] Climb may refer to an engine operating phase in which the aircraft is propelled by the thrust produced by the engine. During climb, the engine may produce between 75% and 100% of the available thrust. In additional non-limiting examples, the engine may produce between 80% and 95% of the available thrust. In additional non-limiting examples, the engine may produce between 85% and 90% of the available thrust. At this point, climb may refer to the operating phase between takeoff and reaching cruise conditions in the aircraft flight cycle, reaching cruise conditions thus defining the start of the cruise phase or a part thereof of the aircraft flight. Additionally or alternatively, climb may refer to one or more nominal periods at a nominal point or during the aircraft flight cycle between takeoff and landing, in which a relative increase in altitude is required, which may require additional thrust requirements of the engine.

[0313] As used herein, the cruise conditions that may define the cruise phase (or a portion thereof) of an aircraft flight have their ordinary meaning and will be readily understood by a person skilled in the art. In some examples, for a given gas turbine engine of an aircraft, the cruise conditions may refer to the operating point at which the engine cruises during a given mission (referred to in the industry as an "economic mission"), and the gas turbine engine is designed to be attached to the aircraft. In this regard, the intermediate cruise can be regarded as a key point in the aircraft flight cycle, at which point 50% of the total fuel burned between the highest point of ascent and the start of descent has been burned (which may approximate the midpoint between the highest point of ascent and the start of descent in terms of time and / or distance). Thus, considering the number of engines provided for the aircraft, the cruise conditions can define the operating point, phase, or a portion thereof of the flight that provides the thrust to ensure the steady-state operation (i.e., maintaining a constant altitude and / or a constant Mach number) or at least substantially steady-state operation (i.e., maintaining at least substantially a constant altitude and / or at least substantially a constant Mach number) of the aircraft. For example, if the engine is designed to be attached to an aircraft with two engines of the same type, then under the cruise conditions, the engine can provide half of the total thrust required for the steady-state operation or at least substantially steady-state operation of the aircraft during the intermediate cruise.

[0314] In other words, for a given gas turbine engine of an aircraft, the cruise conditions can be defined as the operating point of the engine that provides a specified thrust under the intermediate cruise atmospheric conditions (defined by the International Standard Atmosphere according to ISO 2533 at the intermediate cruise altitude), which is required to provide, in combination with any other engines on the aircraft, the steady-state operation, or at least substantially steady-state operation, of the aircraft at a given intermediate cruise Mach number. For any given gas turbine engine of an aircraft, the intermediate cruise thrust, atmospheric conditions, and Mach number are known, and thus the operating point of the engine can be clearly defined under the cruise conditions.

[0315] By way of example only, the forward speed under the cruise conditions can be any point within the range from 0.7 Mach to 0.9 Mach, such as from 0.75 to 0.85, such as from 0.76 to 0.84, such as from 0.77 to 0.83, such as from 0.78 to 0.82, such as from 0.79 to 0.81, such as approximately 0.8 Mach, approximately 0.85 Mach, or within the range from 0.8 to 0.85. Any single speed within these ranges can be part of the cruise conditions. For some aircraft, the cruise conditions may be outside these ranges, such as below 0.7 Mach or above 0.9 Mach.

[0316] By way of example only, the cruise conditions may correspond to standard atmospheric conditions (according to the International Standard Atmosphere ISA) at altitudes within the following ranges: 10000m to 15000m, for example within the range of 10000m to 12000m, for example within the range of 10400m to 11600m (about 38000 feet), for example within the range of 10500m to 11500m, for example within the range of 10600m to 11400m, for example within the range of 10700m (about 35000 feet) to 11300m, for example within the range of 10800m to 11200m, for example within the range of 10900m to 11100m, for example approximately 11000m. The cruise conditions may correspond to the standard atmospheric conditions at any given altitude within these ranges.

[0317] By way of example only, the cruise conditions may correspond to a forward Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35000ft (10668m). Under such cruise conditions, the engine may provide a known required net thrust level. Of course, the known required net thrust level depends on the engine and its intended application, and may be a value within the range of, for example, 20kN to 40kN.

[0318] As a further example only, the cruise conditions may correspond to a forward Mach number of 0.85 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38000ft (11582m). Under such cruise conditions, the engine may provide a known required net thrust level. The known required net thrust level of course depends on the engine and its intended application, and may be a value within the range of, for example, 35kN to 65kN.

[0319] In use, the gas turbine engine described and / or claimed herein may operate under the cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions of an aircraft (e.g., intermediate cruise conditions) on which at least one (e.g., 2 or 4) gas turbine engines may be installed to provide propulsion thrust.

[0320] In addition, those skilled in the art will immediately recognize that either or both of landing and approach refer to the operational phases within the flight cycle of an aircraft between cruise and landing, where approach in particular forms part of the landing and take-off (LTO) phase. During either or both of landing and approach, the engine may produce between 0% and 50% of the available thrust. In additional non-limiting examples, the engine may produce between 25% and 40% of the available thrust. In additional non-limiting examples, the engine may produce between 30% and 35% of the available thrust. Additionally or alternatively, landing may refer to a nominal point within the flight cycle of an aircraft between take-off and landing where a relative reduction in altitude is required, and this may require a reduced thrust requirement from the engine.

[0321] According to one aspect, there is provided an aircraft that includes a gas turbine engine as described herein and / or claimed. The aircraft according to this aspect is the aircraft to which the gas turbine engine has been designed to be attached. Thus, the cruise conditions according to this aspect may correspond to the operating point, phase, or a part thereof of the aircraft flight, as defined elsewhere herein.

[0322] According to one aspect, there is provided a method of operating a gas turbine engine as described herein and / or claimed. The operation can be carried out under any cruise conditions (e.g., in terms of thrust, atmospheric conditions, and Mach number) as may be defined elsewhere herein.

[0323] According to one aspect, there is provided a method of operating an aircraft that includes a gas turbine engine as described herein and / or claimed. The operation according to this aspect may include (or may be) an operation under any suitable conditions (e.g., at an intermediate cruise of the aircraft), as defined elsewhere herein.

[0324] The skilled person will understand that features or parameters described in relation to any one of the above aspects may be applied to any other aspect unless mutually exclusive. Furthermore, any feature or parameter included or described herein may be applied to any aspect and / or combined with any other feature or parameter included or described herein unless mutually exclusive.

[0325] Unless mutually exclusive, any parameter or value included or described herein may be applied to and / or combined with any one or more additional parameters and / or values included or described herein. For example, a first parameter or value (e.g., parameter A) included or described herein may be applied to and / or combined with any one or more additional parameters and / or values (e.g., parameter B; parameter C; and parameter D, etc.) included or described herein to express the product of their relationships. For example, the skilled person in the art will understand that in the case where parameter A and parameter B are disclosed separately, the product of their relationships may be expressed, as required, for example, as A / B, B / A, B*A, or any such other application, combination, or function of parameter A with respect to parameter B. BRIEF DESCRIPTION OF THE DRAWINGS

[0326] Embodiments will now be described, by way of example only, with reference to the drawings, in which:

[0327] Figure 1 is a cross-sectional side view of a gas turbine engine;

[0328] Figure 2 is a close-up cross-sectional side view of an upstream portion of a geared gas turbine engine;

[0329] Figure 3 is a partial cross-sectional view of a gearbox for a gas turbine engine;

[0330] Figure 4 is a schematic view of an aircraft having a propulsion system including two gas turbine engines;

[0331] Figure 5 is a schematic view of an exemplary fuel system;

[0332] Figure 6 is a schematic view of an alternative exemplary fuel system;

[0333] Figure 7 is a schematic view of a part of an exemplary recirculation oil system (primary oil circuit system);

[0334] Figure 8 is a schematic view of another part of an exemplary recirculation oil system (secondary oil circuit system);

[0335] Figure 9 is a schematic view of a part of an alternative exemplary recirculation oil system;

[0336] Figure 10 is Figure 5 of the exemplary fuel system and Figure 7 and Figure 8 a part of the exemplary recirculation oil system;

[0337] Figure 11 is a schematic view of another alternative exemplary recirculation oil system, showing all the main components of the heat exchange system;

[0338] Figure 12 shows an exemplary method of operating a gas turbine engine;

[0339] Figure 13 shows another exemplary method of operating a gas turbine engine;

[0340] Figure 14 shows another exemplary method of operating a gas turbine engine;

[0341] Figure 15 shows another exemplary method of operating a gas turbine engine;

[0342] Figure 16 shows another example of a recirculation oil system for a gas turbine engine;

[0343] Figure 17 shows another exemplary method of operating a gas turbine engine;

[0344] Figure 18Another exemplary method of operating a gas turbine engine is shown; and

[0345] Figure 19 A graph showing the heat transfer range of a geared gas turbine engine during cruise is shown. DETAILED DESCRIPTION

[0346] Figure 1 A gas turbine engine 10 having a main rotational axis 9 is shown. The engine 10 includes an air intake 12 and a propulsive fan 23 which generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 includes a core 11 which receives the core airflow A. The engine core 11 includes, in axial flow series, a low pressure compressor 14, a high pressure compressor 15, combustion equipment 16, a high pressure turbine 17, a low pressure turbine 19 and a core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass duct 22. The fan 23 is attached to and driven by the low pressure turbine 19 via a shaft 26 and a epicyclic gearbox 30.

[0347] In use, the core airflow A is accelerated and compressed by the low pressure compressor 14 and is directed into the high pressure compressor 15 for further compression. The compressed air discharged from the high pressure compressor 15 is directed into the combustion equipment 16 where the compressed air is mixed with fuel F and the mixture is combusted. The combustion equipment 16 may be referred to as a burner 16, where the terms "combustion equipment 16" and "burner 16" may be used interchangeably herein. The resulting hot combustion products then expand through the high pressure turbine and the low pressure turbines 17, 19 before being discharged through the nozzle 20, thereby driving the high pressure turbine and the low pressure turbines to provide some propulsive thrust. The high pressure turbine 17 drives the high pressure compressor 15 via a suitable interconnecting shaft 27. The fan 23 is generally used to apply an increased pressure to the bypass airflow B flowing through the bypass duct 22 such that the bypass airflow B is discharged through the bypass exhaust nozzle 18 to generally provide most of the propulsive thrust. The epicyclic gearbox 30 is a reduction gearbox.

[0348] Figure 2 An exemplary arrangement of a geared fan gas turbine engine 10 is shown. The low pressure turbine 19 (see Figure 1)The drive shaft 26 is coupled to the sun gear or sunwheel 28 of the epicyclic gear arrangement 30. Radially outward of the sunwheel 28 and meshing therewith are a plurality of planet gears 32 which are connected together by a planet carrier 34. The planet carrier 34 constrains the planet gears 32 to precess synchronously about the sunwheel 28 while each planet gear 32 rotates about its own axis. The planet carrier 34 is coupled via a link 36 to the fan 23 so as to drive the fan to rotate about the engine axis 9. Radially outward of the planet gears 32 and meshing therewith is a ring gear or annulus 38 which is coupled via a link 40 to the fixed support structure 24.

[0349] Note that the terms "low pressure turbine" and "low pressure compressor" as used herein may respectively denote the lowest pressure turbine stage and the lowest pressure compressor stage (i.e., excluding the fan 23), and / or the turbine stage and the compressor stage connected together by the interconnecting shaft 26 having the lowest rotational speed in the engine (i.e., excluding the gearbox output shaft driving the fan 23). In some literature, the "low pressure turbine" and "low pressure compressor" referred to herein may alternatively be termed "intermediate pressure turbine" and "intermediate pressure compressor". In the case of using such alternative nomenclature, the fan 23 may be referred to as the first or lowest pressure compression stage.

[0350] In Figure 3 the epicyclic gearbox 30 is shown in more detail by way of example. Each of the sunwheel 28, the planet gears 32 and the annulus 38 includes teeth around its periphery for meshing with other gears. However, for clarity, Figure 3 only an exemplary portion of the teeth is shown. Four planet gears 32 are shown, but it will be apparent to those skilled in the art that more or fewer planet gears 32 may be provided within the scope of the claimed invention. Practical applications of planetary epicyclic gearboxes 30 typically include at least three planet gears 32.

[0351] In Figure 2 and Figure 3 the epicyclic gearbox 30 shown by way of example is planetary, where the planet carrier 34 is coupled via a link 36 to the output shaft and the annulus 38 is fixed. However, any other suitable type of epicyclic gearbox 30 may be used. By way of a further example, the epicyclic gearbox 30 may be a stellar arrangement where the planet carrier 34 remains fixed allowing the annulus (or ring gear) 38 to rotate. In such an arrangement, the fan 23 is driven by the annulus 38. By way of another alternative example, the gearbox 30 may be a differential gearbox where both the annulus 38 and the planet carrier 34 are allowed to rotate.

[0352] It should be understood that Figure 2 and Figure 3The arrangement shown is merely exemplary, and various alternatives are within the scope of the present disclosure. By way of example only, any suitable arrangement may be used to locate the gearbox 30 in the engine 10 and / or to connect the gearbox 30 to the engine 10. By way of another example, the connecting members (such as Figure 2 the linkages 36, 40 in the example) between the gearbox 30 and other components of the engine (such as the input shaft 26, the output shaft, and the fixed structure 24) may have any desired degree of stiffness or flexibility. By way of a further example, any suitable arrangement of bearings between the rotating and fixed components of the engine (e.g., between the input and output shafts from the gearbox and a fixed structure such as the gearbox housing) may be used, and the present disclosure is not limited to Figure 2 the exemplary arrangement shown. For example, in the case where the gearbox 30 has a stellar arrangement (as described above), those skilled in the art will readily appreciate that the arrangement of the output linkages and the support linkages and the bearing positions will generally be different from Figure 2 the arrangement shown by way of example in

[0353] Accordingly, the present disclosure extends to gas turbine engines having any arrangement in the type of gearbox (e.g., stellar or planetary gears), support structure, input and output shaft arrangement, and bearing positions.

[0354] Optionally, the gearbox may drive additional and / or alternative components (e.g., an intermediate-pressure compressor and / or a booster compressor).

[0355] Other gas turbine engines to which the present disclosure may apply may have alternative configurations. For example, such engines may have an alternative number of compressors and / or turbines and / or an alternative number of interconnected shafts. By way of a further example, Figure 1 the gas turbine engine shown in

[0356] has split nozzles 18, 20, which means that the flow through the bypass duct 22 has its own nozzle 18, which is separate from and radially outside the core engine nozzle 20. However, this is not restrictive, and any aspect of the present disclosure may also apply to an engine in which the flow through the bypass duct 22 and the flow through the core 11 mix or combine before (or upstream of) a single nozzle that may be referred to as a mixed-flow nozzle. One or both nozzles (whether mixed or split) may have a fixed or variable area.

[0357] Other gas turbine engines to which the present disclosure may apply may have alternative configurations. By way of example, such engines may have an alternative number of interconnected shafts (e.g., two) and / or an alternative number of compressors and / or turbines. Additionally, the engine may include a gearbox disposed in the drive train from the turbine to the compressor and / or fan.

[0358] Although the example described relates to a turbofan engine, the present disclosure may apply to, for example, any type of gas turbine engine, such as an open rotor (where the fan stage is not surrounded by a nacelle) or, for example, a turboprop engine. In some arrangements, the gas turbine engine 10 may not include a gearbox 30.

[0359] The geometry of the gas turbine engine 10 and its components is defined by a conventional shaft system, including an axial direction (aligned with the rotational axis 9), a radial direction (the direction from bottom to top in Figure 1 ), and a circumferential direction (perpendicular to the Figure 1 plane of the view). The axial, radial, and circumferential directions are perpendicular to each other.

[0360] The fuel F supplied to the combustion equipment 16 may include a fossil-based hydrocarbon fuel, such as kerosene. Thus, the fuel F may include molecules from one or more of the chemical families of normal alkanes, isoalkanes, cycloalkanes, and aromatics. Additionally or alternatively, when blended with, mixed with, or replaced by an alternative fuel, the fuel F may contain renewable hydrocarbons produced from biological or non-biological resources, also known as sustainable aviation fuel (SAF). In each provided embodiment, the fuel F may include one or more trace elements, including, for example, sulfur, nitrogen, oxygen, inorganics, and metals.

[0361] Those skilled in the art understand SAF to refer to, for example, biofuels, renewable aviation fuels, renewable jet fuels, alternative fuels, or biojet fuels produced from biological or non-biological resources. SAF is understood to generally be synthesized from carbon-containing gases extracted from the atmosphere and / or captured from industrial processes; or from a wide range of sustainable feedstocks, such as waste oils and fats; municipal solid waste; cellulosic waste (such as corn stover); cover crops, such as camelina, carinata, and pennycress; abiogenic alternative fuels; jatropha; halophytes, and algae, rather than from fossil-based hydrocarbons. SAF is understood not to include fossil fuels.

[0362] The functional performance of a given fuel composition or fuel F blend for a given mission can be defined at least in part by the ability of the Brayton cycle of a fuel-serviced gas turbine engine 10. Parameters that define functional performance can include, for example, specific energy; energy density; thermal stability; and emissions including gases and / or particulate matter. In this regard, particulate emissions can include soot particles produced by the combustion of the fuel F, also known as non-volatile particulate matter (nvPM). Any reference to soot or smoke in this document also applies to other types of particulate emissions known in the art. Gas emissions can include one or more of the following: nitrogen oxides (NOx); carbon monoxide (CO); carbon dioxide (CO2); unburned hydrocarbons (UHC); sulfur oxides (SO), including, for example, sulfur dioxide (SO2) and / or sulfur trioxide (SO3); and volatile organic compounds (VOC) produced by the combustion of the fuel F. Any reference to gas emissions in this document also applies to other types of gas emissions known in the art.

[0363] A relatively high specific energy, expressed in MJ / kg (i.e., energy per unit mass), can at least in part reduce the takeoff weight and thus potentially provide a relative improvement in fuel efficiency. A relatively high energy density, expressed in MJ / L (i.e., energy per unit volume), can at least in part reduce the takeoff fuel volume, which can be particularly important for volume-constrained missions or military operations involving fuel replenishment. A relatively high thermal stability (i.e., inhibiting fuel degradation or coking under thermal stress) can allow the fuel to maintain elevated temperatures in the engine and fuel injectors and thus potentially provide a relative improvement in combustion efficiency. Reduced emissions, including particulate matter, can allow for reduced contrail formation while reducing the environmental impact of a given mission. Other properties of the fuel can also be critical to functional performance. For example, a relatively low freezing point (°C) can allow for long-duration missions with optimized flight profiles; a minimum aromatic concentration (%) can ensure sufficient swelling of certain materials used to construct O-rings and seals that have previously been exposed to fuels with high aromatic content; and a maximum surface tension (mN / m) can ensure sufficient spray breakup and atomization of the fuel.

[0364] The ratio of the number of hydrogen atoms to the number of carbon atoms in a molecule can affect the specific energy of a given composition or fuel blend. Fuels with a higher ratio of hydrogen atoms to carbon atoms can have a higher specific energy in the absence of bond strain. For example, fossil-based hydrocarbon fuels can include molecules having from about 7 to 18 carbons, with a significant portion of a given composition being derived from molecules having 9 to 15 carbons, with an average of 12 carbons.

[0365] Multiple sustainable aviation fuel blends have been approved for use. For example, some approved blends contain blend ratios of up to 10% sustainable aviation fuel, while other approved blends contain blend ratios of 10% to 50% sustainable aviation fuel (the remainder containing one or more fossil-based hydrocarbon fuels such as kerosene), with additional compositions awaiting approval. However, sustainable aviation fuel blends that are expected to include up to (and including) 100% sustainable aviation fuel (SAF) in the aviation industry will ultimately be approved for use.

[0366] Sustainable aviation fuel can include one or more of n-alkanes, isoalkanes, cycloalkanes, and aromatics, and can be produced, for example, from one or more of: syngas; lipids (e.g., fats, oils, and greases); sugars; and alcohols. Thus, sustainable aviation fuel can include either or both of lower aromatic and sulfur content (relative to fossil-based hydrocarbon fuels). Additionally or alternatively, sustainable aviation fuel can include either or both of higher isoalkane and cycloalkane content (relative to fossil-based hydrocarbon fuels). In some examples, sustainable aviation fuel can include either or both of a density between 90% and 98% of the density of kerosene and a calorific value between 101% and 105% of the calorific value of kerosene.

[0367] In some examples, the aromatic and / or other non-alkane content of the sustainable aviation fuel, or blend, provided to the combustion equipment 16 is relatively lower than that of kerosene. The sustainable aviation fuel can include an aromatic content, 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%. Depending on one or more of preference, fuel feedstock or supplier, and compositional variations therein, the aromatic content of the sustainable aviation fuel can be a numerical value or range that is defined or bounded by any two of the values in the previous sentence, including the end values (i.e., these values can form upper or lower limits), for example, 13.5%, 8.5%, 2.5%, 0.35%, 0.15%, 0.05%, 0.005% or 0%; or 0% to 0.75%, 0% to 0.5%, or 0.1% to 0.25%; or 0.15% to 0.65%, 0.35% to 0.55%, or 0.035% to 0.055%.

[0368] Due at least in part to the molecular structure of sustainable aviation fuel, sustainable aviation fuel can provide benefits including, for example, one or more of the following: higher specific energy (although, in some examples, lower energy density); higher specific heat capacity; higher thermal stability; higher lubricity; lower viscosity; lower surface tension; lower freezing point; lower soot emissions; lower NOx; and lower CO2 emissions, relative to fossil-based hydrocarbon fuels (e.g., when burned in combustion equipment 16). Thus, relative to fossil-based hydrocarbon fuels such as kerosene, sustainable aviation fuel can result in either or both a relative reduction in fuel consumption and a relative reduction in maintenance costs.

[0369] As Figure 4 depicted, the aircraft 1 can include a plurality of fuel tanks 50, 53; for example, a larger primary fuel tank 50 located in the aircraft fuselage and smaller fuel tanks 53a, 53b located in each wing. In other examples, the aircraft 1 can have only a single fuel tank 50, and / or the wing fuel tanks 53 can be larger than the central fuel tank 50, or the central fuel tank can be omitted (with all fuel alternatively stored in the wings of the aircraft) - it should be understood that many different tank layouts are envisioned, and the illustrated examples are provided for ease of description and are not intended to be limiting.

[0370] Figure 4 An aircraft 1 is shown having a propulsion system 2 including two gas turbine engines 10. Fuel is supplied from a fuel supply system on the aircraft 1 to the gas turbine engines 10. The fuel supply system of the illustrated example includes a single fuel source. For the purposes of this application, the term "fuel source" means: 1) a single fuel tank; or 2) a plurality of fluidly interconnected fuel tanks. Each fuel source is arranged to provide a separate fuel source, i.e., the first fuel source can contain a first fuel having one or more characteristics different from the second fuel contained in the second fuel source. Thus, the first fuel source and the second fuel source are not fluidly coupled to each other to separate different fuels (at least under normal operating conditions). The use of multiple fuel sources allows the aircraft 1 to carry multiple different fuels and change the fuel used during operation and optionally even during cruise or when changing between different operating phases in flight.

[0371] In this example, the first (and only in these examples) fuel source includes a central fuel tank 50 located primarily in the fuselage of the aircraft 1 and a plurality of wing fuel tanks 53a, 53b, with at least one wing fuel tank located in the left wing (port wing) and at least one wing fuel tank located in the right wing (starboard wing) for balance. In the illustrated example, all tanks 50, 53 are fluidly interconnected, thereby forming a single fuel source. Each of the central fuel tank 50 and the wing fuel tanks 53 can include a plurality of fluidly interconnected fuel tanks.

[0372] In another example, the wing fuel tanks 53a, 53b may not be fluidly connected to the central fuel tank 50, thereby forming a separate second fuel source. For balance purposes, one or more fuel tanks in the left wing may be fluidly connected to one or more fuel tanks in the right wing. This may be via the central fuel tank (if the tank does not form part of another fuel source), or bypass the central fuel tank, or both (for maximum flexibility and safety). In another example, the first fuel source includes the wing fuel tanks 53 and the central fuel tank 50, while the second fuel source includes another separate central fuel tank. A fluid interconnection between the wing fuel tanks and the central fuel tank of the first fuel source may be provided for the balance of the aircraft 1. In the aircraft 1 having multiple fuel sources, two or more of the fuel sources may thus contain fuels that are different from each other, such that the aircraft 1 can change fuels during flight. Thus, being able to determine which fuel is being supplied to the burner 16 may be more complex than simply recording the single identity of the fuel on the aircraft 1 or checking it once at startup.

[0373] In some examples, the distribution of the fuel tanks 50, 53 available on the aircraft 1 may be constrained such that the first fuel source and the second fuel source are each substantially symmetric with respect to the aircraft centerline. In cases where asymmetric fuel tank distribution is allowed, suitable fuel transfer mechanisms are typically provided between the fuel tanks of the first fuel source and / or between the fuel tanks of the second fuel source such that the center of mass position of the aircraft can be maintained within acceptable lateral limits throughout the flight.

[0374] An aircraft is typically refueled at multiple different airports, such as at the start and end of a long - haul flight. Despite the standards that all aviation fuels must meet, as described above, different aviation fuels have different compositions, which depend, for example, on their source (e.g., different petroleum sources, biofuels or other synthetic aviation fuels (commonly referred to as sustainable aviation fuel SAF), and / or mixtures of petroleum - based fuels, and other fuels) as well as any additives contained (e.g., such as antioxidants and metal deactivators, biocides, static eliminators, icing inhibitors, corrosion inhibitors) and any impurities. Even for a given airport or fuel supplier, the fuel composition of the available aviation fuel may vary between batches. In addition, the fuel tanks 50, 53 of the aircraft 1 are typically not emptied before being filled for a subsequent flight, thus creating a mixture of different fuels within the tanks - in effect, fuels with different compositions resulting from the mixture.

[0375] The inventors recognize that since different fuels can have different properties while still meeting the standards, knowledge of the fuel available to the aircraft 1 can allow for more effective and customized control of the aircraft 1, and more specifically of the propulsion system 2 of the aircraft (i.e., one or more gas turbine engines 10 of the aircraft 1 and associated controls and components). Thus, knowledge of the fuel can be used as a tool to improve aircraft performance, and therefore determining or monitoring the fuel composition can provide benefits. In particular, determining one or more fuel characteristics of the fuel to be supplied to the burner 16 (whether fuel from a single fuel source or a mixture of one or more fuels from different fuel sources) is thus important in determining engine operation. A key feature of engine operation is thermal management - engine thermal management is mainly carried out by utilizing the heat transfer of oil and fuel into and out of the engine, and thus control of the heat exchange system 3000 - for example by controlling the oil flow, controlling the fuel flow, and / or otherwise controlling one or more heat exchangers 1004, 1006, 2020, 2030 or other heat exchange components - can allow the performance of the engine 10 to be optimized for a fuel with specific fuel characteristics.

[0376] As used herein, the term "fuel characteristic" refers to an inherent or intrinsic fuel property (such as fuel composition), rather than a variable property (such as volume or temperature). One or more fuel characteristics can be determined, and this data is used to adjust the control of the engine 10, particularly the control of the thermal management system 3000. Examples of fuel characteristics include one or more of the following:

[0377] i. The percentage of sustainable aviation fuel (SAF) in the fuel, or an indication that the fuel is a fossil fuel (such as fossil kerosene), or an indication that the fuel is a pure SAF fuel;

[0378] ii. Parameters of the hydrocarbon distribution of the fuel, such as:

[0379] · The aromatic content of the fuel, and optionally / alternatively the polyaromatic content of the fuel;

[0380] · The hydrogen-to-carbon ratio (H / C) of the fuel;

[0381] · Percentage information on the composition of some or all of the hydrocarbons present;

[0382] iii. The presence or percentage of specific elements or species, such as:

[0383] · The percentage of nitrogen-containing substances in the fuel;

[0384] · The presence or percentage of tracer substances or trace elements in the fuel;

[0385] · The naphthalene content of the fuel;

[0386] · Sulfur content of the fuel;

[0387] · Naphthene content of the fuel;

[0388] · Oxygen content of the fuel;

[0389] iv. One or more properties of the fuel used in the gas turbine engine 10, such as:

[0390] · Level of non-volatile particulate matter (nvPM) emissions or CO2 emissions during combustion;

[0391] · Coking level of the fuel;

[0392] v. One or more properties of the fuel itself independent of the use or combustion of the engine 10, such as:

[0393] · Thermal stability of the fuel (e.g., thermal breakdown temperature; a numerical value can be assigned to the thermal stability at any temperature by taking the reciprocal of the deposition rate of the fuel decomposition products at a given temperature); and

[0394] · One or more physical properties, such as density, viscosity, calorific value, freezing temperature, and /

[0395] or heat capacity.

[0396] The fuel characteristics to be determined can be selected based on which properties of the fuel are most relevant to the changes that can be made to the thermal management system 3000. Determining the fuel characteristics can include obtaining the fuel characteristics of any fuel already present in the fuel tanks 50, 53 and an indication of the amount of remaining fuel before fuel replenishment, and then combining this information with information about the new fuel added to the fuel tanks 50, 53 during fuel replenishment.

[0397] Obtaining the fuel characteristics of any fuel already present in the fuel tanks 50, 53 before fuel replenishment, and / or obtaining the fuel characteristics of the fuel provided during fuel replenishment, can include one or more of the following:

[0398] (i) Physically and / or chemically detecting one or more characteristics or parameters of the fuel composition (this can allow direct detection of the fuel characteristics, and / or can allow the use of the detection results to determine the fuel characteristics), and / or detecting one or more tracer elements or compounds added to the fuel for ease of its identification (e.g., dyes);

[0399] (ii) Retrieving fuel characteristic information from on-vehicle memory / data repository; and / or

[0400] (iii) Receiving data, for example, from an input provided at the user interface, or receiving data transmitted to the aircraft 1.

[0401] In some examples, one or more fuel properties may be determined during operation of the gas turbine engine 10, such as by inferring fuel properties of fuel supplied to the combustor 16 from engine performance metrics or by performing in-wing detection.

[0402] In some examples, a variety of different methods may be performed to obtain fuel properties. For example, different methods may be used for different properties, and / or different methods may be used for the same property as a check. For example, stored or otherwise provided fuel property data may be compared with the results of chemically or physically detecting one or more parameters of the fuel. If there is a mismatch between the stored fuel property and the corresponding detected parameter, an alert may be provided.

[0403] Fuel properties may be determined by physically and / or chemically detecting one or more characteristics of the fuel composition (e.g., in a test unit outside the wing, or when the fuel is transported to a fuel tank on the wing, or actually when used in the gas turbine engine 10), thus allowing direct detection of fuel properties or providing data from which fuel properties can be determined as described above, and / or detecting one or more tracer elements or compounds added to the fuel for ease of its identification (e.g., dyes); or by receiving data, for example, from an input provided by a user interface, or by receiving data transmitted to the aircraft, for example, by scanning a barcode associated with fuel delivery.

[0404] When using physical and / or chemical determination, fuel properties may be detected in various ways, including direct ways (e.g., from sensor data corresponding to the fuel property in question) and indirect ways (e.g., by inference or calculation from other properties or measurements, or by reference to data of specific detected tracers in the fuel). These properties may be determined as relative values compared to another fuel or as absolute values. For example, one or more of the following detection methods may be used:

[0405] ● The aromatic or naphthene content in the fuel may be determined based on the measurement of the expansion of a sensor component made of a sealing material such as a nitrile sealing material.

[0406] ● Trace substances or species naturally present in the fuel or added as tracers may be used to determine fuel properties such as the percentage of sustainable aviation fuel in the fuel or whether the fuel is kerosene.

[0407] ● Measurement of the vibration mode of a piezoelectric crystal exposed to the fuel may be used as a basis for determining various fuel properties (including the aromatic content of the fuel, the oxygen content of the fuel, and the thermal stability or coking level of the fuel), for example, by measuring the accumulation of deposits on the upper surface of the piezoelectric crystal, which will cause a change in the vibration mode.

[0408] ● Each fuel characteristic can be determined by collecting performance parameters of the gas turbine engine 10 during a first operation (e.g., during takeoff) and optionally during a second operation (e.g., during cruise), and if a fuel of known properties is used, comparing these collected parameters with expected values.

[0409] ● Each fuel characteristic including the aromatic content of the fuel can be determined based on sensor measurements of the presence, absence, or degree of formation of a locus during the operation of the gas turbine 10.

[0410] ● The fuel characteristic including the aromatic content can be determined based on ultraviolet-visible spectroscopic measurements performed on the fuel.

[0411] ● Each fuel characteristic including sulfur content, naphthalene content, aromatic hydrogen content, and hydrogen-carbon ratio can be determined by measuring substances present in the exhaust gas emitted by the gas turbine engine 10 during its use.

[0412] ● The calorific value of the fuel can be determined during the operation of the aircraft 1 based on measurements performed during fuel combustion, for example using fuel flow rate and shaft speed or temperature variations on the burner 16.

[0413] ● Each fuel characteristic can be determined by making an operating change that is arranged to affect the operation of the gas turbine engine 10, sensing the response to the operating change; and determining one or more fuel characteristics of the fuel based on the response to the operating change.

[0414] ● By changing the fuel supplied to the gas turbine engine 10 from a first fuel to a second fuel and determining one or more fuel characteristics of the second fuel based on a change in the relationship between T30 and one of T40 and T41 (the relationship indicating a temperature rise on the burner 16), each fuel characteristic related to the fuel characteristics of the first fuel can be determined. These characteristics can be determined as relative values compared to the first fuel or as absolute values, for example by referring to known values of the first fuel.

[0415] As used herein, T30, T40, and T41 and any other numbered pressures and temperatures are defined using the station numbers listed in the standard SAE AS755, specifically:

[0416] T30 = high pressure compressor (HPC) outlet total temperature;

[0417] T40 = combustion outlet total temperature;

[0418] T41 = high pressure turbine (HPT) rotor inlet total temperature.

[0419] Any suitable method known in the art can be used, and the determination of fuel characteristics is not discussed further herein.

[0420] Accordingly, the aircraft 1 may include a fuel composition determination module 57 that is arranged to determine at least one fuel characteristic of the fuel. This determination may be performed by obtaining values from a data repository (e.g., data provided to the aircraft 1 during fuel replenishment) 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 Figure 4 the example shown, a fuel composition determination module 57 is provided as part of each engine 10 where fuel enters the engine 10. In other embodiments, the module 57 may be located differently within the engine 10 or elsewhere on the aircraft 1, such as in the fuel tanks 50, 53, on or near the fuel tanks. In embodiments having a fuel composition determination module 57, the module 57 may be arranged to provide an output to an electronic engine controller (EEC), or the module 57 may be provided as part of the EEC. Accordingly, one or more fuel characteristics determined by the fuel composition determination module 57 may be used to affect the control of the engine 10. A dedicated controller 58 may be provided to process the output of the fuel composition determination module 57 and provide control instructions to controllable engine components based on the fuel characteristic information. In other embodiments, the control function 58 may be provided by the EEC, and a separate unit or module may not be provided.

[0421] Figure 5 An exemplary fuel system 1000 for a geared gas turbine engine 10 is schematically shown, which fuel system 1000 includes a fuel flow path from a fuel tank 50 to a burner 16 of the gas turbine engine 10 of the aircraft 1. The fuel system 1000 includes both a fuel supply system 50, 1002 of the aircraft 1 that supplies fuel to the engine 10 and a fuel management system 1500 that operates within the engine 10. The fuel management system 1500 has a role in managing the fuel temperature as well as the fuel flow rate and directing the fuel through one or more heat exchangers 1004, 1006 of the engine's heat exchange system 3000. The heat exchange system 3000 includes components of the fuel management system 1500 as well as recirculating oil systems 2000, 2000' (which will be described in more detail below). The heat exchange system 3000 is a general term for the systems and components for transferring heat between fluids (specifically oil and fuel) within the engine 10 and includes heat exchangers, valves, connecting pipes, and associated components such as pumps, refrigeration devices, etc.

[0422] Returning to the fuel system 1000, fuel is pumped from the fuel tank 50 to the gas turbine engine 10 by the low-pressure fuel supply pump 1002. The fuel then flows through the secondary fuel-oil heat exchanger 1004 and the primary fuel-oil heat exchanger 1006. The primary fuel-oil heat exchanger 1006 may be described as the main fuel-oil heat exchanger because the oil flowing through it can be used to cool and lubricate the main gearbox 30 of the engine 10. The secondary fuel-oil heat exchanger 1004 may be described as the integrated drive generator fuel-oil heat exchanger because the oil flowing through it can be used to cool and / or lubricate one or more components of the integrated drive generator (IDG) of the engine 10. In other embodiments, different types of generators may be used instead of the IDG, such as a variable frequency generator (VFG) or a variable frequency starting generator (VFSG). The system 1000 of such embodiments may be equivalent in other respects. Thus, the engine 10 of the described example includes two fuel-oil heat exchangers 1004, 1006. In other embodiments, more or fewer fuel-oil heat exchangers may be provided. The illustrated fuel management system 1500 is arranged such that the fuel reaches the secondary fuel-oil heat exchanger 1004 before reaching the primary fuel-oil heat exchanger 1006. After leaving the primary fuel-oil heat exchanger 1006, the fuel then passes through the engine fuel pump 1003 and travels to the burner 16. The engine fuel pump 1003 may be described as the main fuel pump.

[0423] The primary fuel-oil heat exchanger 1006 and the secondary fuel-oil heat exchanger 1004 are configured such that, in addition to the fuel flow passing through them, an oil flow is also conveyed through each heat exchanger. 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 the fuel flowing through them. In the standard operation of the engine 10, such as under cruise 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 way, the primary fuel-oil heat exchanger 1006 and the secondary fuel-oil heat exchanger 1004 are each configured to transfer thermal energy from the oil flow to the fuel flow passing through them during operation.

[0424] The two oil flows (the oil flow through the primary heat exchanger and the oil flow through the secondary heat exchanger) can be separate - physically separate, and optionally also chemically different oils and / or having different flow rates. Thus, the oil flowing through the primary fuel-oil heat exchanger 1006 may be different from the oil flowing through the secondary fuel-oil heat exchanger 1004.

[0425] Generally speaking, at least most of the fuel passing through the secondary fuel-oil heat exchanger 1004 also passes through the primary fuel-oil heat exchanger 1006. The two heat exchangers 1004, 1006 can thus be described as being in series with each other with respect to the fuel flow, and along the main fuel flow path from the tank 50 to the burner 16. However, either or each of the heat exchangers 1004, 1006 can be provided with a bypass to allow some of the fuel to avoid passing through the corresponding heat exchanger, for example in the form of a bypass pipe 1005 as shown in Figure 5 . A valve (not shown) can determine what proportion of the fuel passes through the heat exchanger 1004 and what proportion of the fuel passes through the bypass pipe 1005. In various embodiments, bypass pipes can be provided for each of the heat exchangers 1004, 1006, allowing a portion of the fuel to avoid either or both of the heat exchangers. Additionally or alternatively, one or more bypass pipes 2005, 2005' for the oil can be provided for either or both of the fuel-oil heat exchangers 1004, 1006, allowing a portion of the oil to avoid one or more of the heat exchangers. A valve arranged to control the fluid flow through the bypass pipe 1005 can be referred to as a bypass valve.

[0426] Thus, the secondary fuel-oil heat exchanger 1004 and the primary fuel-oil heat exchanger 1006 are configured such that, in addition to the fuel flow, an oil flow is also conveyed through each fuel-oil heat exchanger - in the described embodiment, the oil flowing through one fuel-oil heat exchanger is different from the oil flowing through the other fuel-oil heat exchanger, but it should be understood that in other embodiments, the same oil can flow through one fuel-oil heat exchanger and then through the other fuel-oil heat exchanger.

[0427] Thus, in the described embodiment, the two heat exchangers 1004, 1006 are in separate closed-loop systems 2000, 2000' ( Figure 7 , Figure 8 ) with respect to the oil flow, i.e., the oil flowing through the primary fuel-oil heat exchanger and the secondary fuel-oil heat exchanger is fluidly separated and can be chemically different from each other. The two oil loop systems 2000, 2000' are used to circulate the oil through their respective fuel-oil heat exchangers 1006, 1004, and optionally also through one or more additional heat exchangers, such as air-oil or oil-oil heat exchangers, as described below. The two oil loop systems 2000, 2000' together can be described as providing a recirculation oil system for the engine 10.

[0428] Figure 6FIG. 6000 shows an alternative exemplary fuel system 6000 including a fuel supply system and a fuel management system 6500, the alternative exemplary fuel system including a fuel flow path from a fuel tank 50 to a burner 16 of a gas turbine engine 10 of an aircraft 1. As with fuel system 1000, fuel is pumped from fuel tank 50 by a low pressure fuel supply pump 1002. The fuel then flows through a secondary fuel - oil heat exchanger 1004 before reaching an engine fuel pump 1003, and then through a primary fuel - oil heat exchanger 1006, the engine fuel pump pumping the fuel along its flow path to the burner 16. Fuel system 6000 differs from fuel system 1000 in that fuel system 6000 includes a recirculation valve 6010, which is located downstream of the primary fuel - oil heat exchanger 1006 and is arranged to recirculate at least a portion of the fuel that has left the primary fuel - oil heat exchanger back to an inlet 1006a of the primary fuel - oil heat exchanger 1006, thereby allowing further heat transfer between the oil and the fuel in the primary loop system. The recirculation valve 6010 can determine the proportion of fuel recirculated via a recirculation pipe 6011 and the proportion that continues more directly to the burner 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. In some embodiments where the recirculation valve 6010 is positioned upstream of the engine fuel pump 1003, an additional recirculation pump (not shown) may be provided to provide a positive pressure gradient. In some embodiments, there may be a fuel return tank (FRTT) pump, and it also contributes to recirculation. The recirculation valve 6010 is arranged to allow a controlled amount of fuel to return to the inlet 1006a of the primary heat exchanger 1006, and thus flow through the primary heat exchanger 1006 multiple times before reaching the pump 1003 and the burner 16. It is contemplated that in an alternative embodiment, the recirculation valve may be positioned downstream of the engine fuel pump 1003, such as as Figure 10 shown. In such embodiments, the recirculation valve 6010 will be arranged to allow a controlled amount of fuel to return to the inlet 1006a of the primary heat exchanger 1006, and thus flow through the primary heat exchanger 1006 and the pump 1003 multiple times before reaching the burner 16. This recirculation provides a mechanism for controlling the fuel flow within the fuel management system 6500 and within the thermal management system 3000 without changing the fuel flow from the fuel tank 50 to the engine 10. Recirculation can be used to control the heat transfer ratio, as described in more detail below.

[0429] A pipe 6011 can thus be provided, which can be referred to as a recirculation pipe as it transports fuel from a point along the main flow path through the engine 10 to an earlier point along that flow path, such that the fuel must again pass through portions between the flow paths. The pipe leads from a recirculation valve to a point on the flow path upstream of the inlet 1006a of the primary heat exchanger 1006. In some embodiments, a recirculation pipe 6011 and a bypass pipe 1005 can be provided for any given heat exchanger 1004, 1006. In some embodiments, the same pipes 1005, 6011 can be used as both recirculation pipes and bypass pipes, and one or more valves can be used to control the direction of fluid flow therethrough.

[0430] In Figure 10 the illustrative example of Figure 6 , different from the illustrative example of

[0431] the recirculation valve 6010 is located downstream of the pump 1003 such that the recirculated fuel has passed through both the primary heat exchanger 1006 and the pump 1003, rather than just through the primary heat exchanger 1006. The recirculation pipe 6011 returns the recirculated fuel to a point on the flow path upstream of both the main pump 1003 and the primary heat exchanger 1006 such that the recirculated fuel passes through both components again. For a given shaft speed of the engine 10, recirculating fuel through the pump can allow for more adjustable control of the fuel flow to the burner 16, noting that the pump speed (or a limited set of pump speed options) is typically set by the shaft speed.

[0432] The gas turbine engine 10 of the described aircraft 1 includes a recirculation oil system that is arranged to supply oil to lubricate a plurality of components and remove heat from a plurality of components. In the described embodiments, the recirculation oil system includes a primary oil circuit system 2000 and a secondary oil circuit system 2000', each of which is a closed - loop oil system. In Figure 7FIG. schematically shows an example of a primary closed-loop oil system 2000 - for a heat exchanger, this oil circuit system is referred to as "primary" because it is responsible for lubricating and cooling the main gearbox 30 and typically for the main / primary cooling load of the engine 10. The primary closed-loop oil system 2000 includes an oil tank 2002 adapted to contain a volume of oil. In some embodiments, gas is removed from the oil within the oil tank 2002 by a degasser. The feed pump 2004 is configured to pump oil from the oil tank 2000 to the main fuel-oil heat exchanger 1006. Under cruise conditions, the average temperature of the oil entering the main fuel-oil heat exchanger 1006 is higher than the average temperature of the fuel entering the main fuel-oil heat exchanger 1006. In the main fuel-oil heat exchanger 1006, heat energy is transferred from the oil flow to the fuel flow. In this way, the average temperature of the oil flow leaving the main fuel-oil heat exchanger 1006 is lower than the average temperature of the oil flow entering the main fuel-oil heat exchanger 1006, so the oil is cooled before being reused as a lubricant and / or coolant, allowing the cooled oil to remove more heat from the system to be lubricated and / or cooled. Also in this way, the average temperature of the fuel leaving the main fuel-oil heat exchanger 1006 is higher than the average temperature of the fuel entering the main fuel-oil heat exchanger 1006.

[0433] In a standard heat exchanger 1006, the basic limiting factor for the level of cooling provided is the temperature of the coolant (in this case, the fuel) - even if an effectively infinite flow rate of fuel or oil is used, the temperature of the oil leaving the heat exchanger 1006 cannot be lower than the temperature of the fuel entering the heat exchanger 1006 (and the temperature of the fuel cannot be raised to be higher than the temperature of the oil entering the heat exchanger). By definition of the second law of thermodynamics, temperature equilibrium is the limit. In some embodiments, as Figure 7As shown, a refrigeration cycle device 1007 is provided. The refrigeration cycle device 1007 is arranged to provide a heat boost by transferring more heat from the oil to the fuel, such that the fuel temperature rises more than when simply passing through the heat exchanger, and in some cases rises above the oil temperature. The refrigeration cycle device 1007 can take the form of an additional closed loop with a refrigerant fluid, having 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 device 1007 may additionally include pumps, compressors, and / or expansion valves / instruments, and can be made in any suitable design known in the art. An oil valve can control the amount of oil flowing through the refrigeration cycle device 1007, and a fuel valve can control the amount of fuel flowing through the refrigeration cycle device 1007. The refrigeration cycle device 1007 (and more specifically, its pumps and / or compressors, and optionally expansion valves or instruments) is powered, typically electrically or mechanically (e.g., via a linkage to a rotating shaft), to provide this further forced heat transfer. The refrigeration cycle device 1007 is shown in the illustrated embodiment as being associated with the main (primary) fuel-oil heat exchanger 1006, thus providing a boost to the fuel temperature after the fuel has passed through the fuel-oil heat exchangers 1004, 1006, and further cooling the oil of the primary oil system 2000 before returning to the gearbox 30. In various embodiments, bypass tubes or recirculation tubes for the oil and / or fuel can be provided around the refrigeration cycle device 1007.

[0434] The oil flow in the primary oil circuit system 2000 is then delivered to the power gearbox 30, which can also be described as the main gearbox 30 of the gas turbine engine 10. The power gearbox 30 is arranged to receive an input from the spool 26 and deliver a drive output to the fan 23 via the fan shaft 42, and includes gears 28, 32, 38 and bearings (e.g., journal bearings) that can be lubricated and cooled by oil. The engine 10 may also include one or more additional bearings to support the shafts 26, 42, and these bearings can be journal bearings. The oil can additionally be used to lubricate and / or cool the journal bearings and, when used under cruise conditions, typically experiences a significant temperature rise, thus helping to cool the bearings and the gearbox 30 as the oil flow carries heat away from the bearings and the gearbox 30. The oil can also be used to lubricate one or more other engine components 33, such as an accessory gearbox (AGB) and / or one or more bearing housings. The AGB 33 (also known as the accessory drive when present) is a gearbox that forms part of the gas turbine engine 10 but is not part of the engine core 11 and does not drive the fan 23. Instead, the AGB drives engine accessories (e.g., fuel pumps) and typically handles large loads. As a result, a relatively large amount of heat can be discharged from the AGB into the oil. One or more bearing housings can be lubricated by the same oil and can similarly discharge heat into the oil. In many embodiments, the AGB and the bearing housings can add more heat to the oil per unit of oil flowing through them than the main gearbox 30. The oil flow can be split into two or more parallel flows, such as one flow through the main gearbox 30 and one flow through the other engine components, or multiple parallel flows (e.g., via different components of the gearbox) through the main gearbox 30 and separate flows through the AGB and the bearing housing 33 or each bearing housing.

[0435] Oil is collected in the oil sump 2008 from the power gearbox 30 (and / or other engine components 33, such as the AGB when present). The scavenge pump 2010 is configured to pump the oil from the oil sump 2008 back to the oil tank 2002 for reuse.

[0436] Figure 8Shows a secondary oil circuit system 2000', which is another closed-loop oil system 2000'. The secondary closed-loop oil system 2000' includes a secondary oil tank 2002' adapted to hold a volume of oil. In some embodiments, gas is removed from the oil within the tank 2002' by a degasser. The secondary feed pump 2004' is configured to pump oil from the secondary oil tank 2002' to the secondary fuel-oil heat exchanger 1004, which in the illustrated embodiment is the IDG fuel-oil heat exchanger 1004. Under cruise conditions, the average temperature of the oil entering the IDG fuel-oil heat exchanger 1004 is higher than the average temperature of the fuel entering the IDG fuel-oil heat exchanger 1004. In the IDG fuel-oil heat exchanger 1004, heat energy is transferred from the oil flow to the fuel flow. In this way, the average temperature of the oil flow leaving the IDG fuel-oil heat exchanger 1004 is lower than the average temperature of the oil flow entering the IDG fuel-oil heat exchanger 1004. Also in this way, the average temperature of the fuel leaving the IDG fuel-oil heat exchanger 1004 is higher than the average temperature of the fuel entering the IDG fuel-oil heat exchanger 1004. Then, the oil flow is delivered to / returns to the integrated drive generator 2006, where the oil flow lubricates and / or cools moving components and is heated in the process. In some embodiments, the oil can be used primarily as a coolant for the IDG 2006 and can provide minimal or no lubrication. The oil is collected from the integrated drive generator 2006 in the secondary oil sump 2008'. The secondary return oil pump 2010' is configured to pump the oil from the secondary oil sump 2008' back into the secondary oil tank 2002' for reuse. In some embodiments, a refrigeration cycle device can also be provided on the secondary oil circuit system 2000'.

[0437] Figure 7 and Figure 8 Each shows a series flow path of the oil, with all oil flows passing sequentially through each component (although it should be understood that one or more bypass pipes or recirculation pipes not shown may be provided for the oil). In other embodiments, the oil flow can be split into two or more parallel flows, for example, one flow through the main fuel-oil heat exchanger 1006 and another flow through the air-oil heat exchanger 2020 (described below).

[0438] Figure 9 Schematically shows Figure 7An alternative exemplary segment of the primary closed-loop oil system 2000 shown. In this segment, the oil flow is pumped by a feed pump 2004 through a valve 2016. The valve 2016 is operable to divert the oil flow between a main fuel-oil heat exchanger 1006 and a first air-oil heat exchanger 2020, where the first air-oil heat exchanger 2020 is arranged in parallel with the main fuel-oil heat exchanger 1006. The oil flow path can be described as branched, with the main fuel-oil heat exchanger 1006 on one branch and the first air-oil heat exchanger 2020 on the other branch, and these branches are in a parallel configuration such that the oil can flow via one branch or the other, but the same portion of the oil cannot cross both branches in the same cycle - flow diversion. The valve 2016 modulates the flow through the two heat exchangers 1006, 2020 and can thus be described as a modulating valve 2016. The oil flow is then recombined and delivered to a power gearbox 30 and / or other engine components 33. Any suitable percentage of the oil can flow through each of the first air-oil heat exchanger 2020 and the main fuel-oil heat exchanger 1006. In some examples, the valve 2016 is operable to change the oil flow to the main fuel-oil heat exchanger 1006 and the first air-oil heat exchanger 2020 as needed. In various examples, an oil-oil heat exchanger 2030 ( Figure 7 not shown in, but present in Figure 11 ) can be arranged, for example, in series with the first air-oil heat exchanger 2020 on this branch of the parallel diversion. The oil-oil heat exchanger 2030 can allow heat exchange between the primary closed-loop oil system 2000 and a secondary closed-loop oil system 2000'.

[0439] Figure 10 Schematically shows an exemplary arrangement and interaction of a first closed-loop oil system 2000, a second closed-loop oil system 2000', and a fuel system 1000, where the fuel flow is shown as a thick black line and the oil flow is shown as a thin black line. The striped thick black line represents a recirculation path 6011 taken by only a portion of the fuel. The combination of the fuel system 1000 and the oil systems 2000, 2000' together forms a heat exchange system 3000. The primary closed-loop oil system 2000 of this exemplary arrangement is arranged as Figure 7 shown. The secondary closed-loop oil system 2000' of this exemplary arrangement is arranged as Figure 8 shown. The fuel system 1000 of this exemplary arrangement is arranged as Figure 5 shown, but with the additional recirculation valve 6010 and tube 6011 as described above. There may also be a bypass tube 1005, but it is not shown for clarity.

[0440] In use, fuel is pumped from fuel tank 50 by low pressure fuel pump 1002. The fuel then flows through IDG fuel-oil heat exchanger 1004. The secondary closed loop oil system 2000’ is configured such that its recirculating oil flow also flows through IDG fuel-oil heat exchanger 1004. In standard operation of the engine 10, at cruise conditions and at idle, the average temperature of the oil flow entering IDG fuel-oil heat exchanger 1004 is higher than the average temperature of the fuel flow entering IDG fuel-oil heat exchanger 1004. IDG fuel-oil heat exchanger 1004 is configured such that heat is transferred from the oil flow to the fuel flow. In this way, the average temperature of the oil flow at the outlet of IDG fuel-oil heat exchanger 1004 is lower than the average temperature of the oil flow at the inlet of IDG fuel-oil heat exchanger 1004. In the same way, the average temperature of the fuel flow at the outlet of IDG fuel-oil heat exchanger 1004 is higher than the average temperature of the fuel flow at the inlet of IDG fuel-oil heat exchanger 1004.

[0441] The fuel then flows through main fuel-oil heat exchanger 1006 and additionally through refrigeration cycle device 1007. The primary closed loop oil system 2000 is configured such that its recirculating oil flow also flows through main fuel-oil heat exchanger 1006 and refrigeration cycle device 1007. In standard operation of the engine 10, at cruise conditions and at idle, the average temperature of the oil flow entering main fuel-oil heat exchanger 1006 is higher than the average temperature of the fuel flow entering main fuel-oil heat exchanger 1006. Main fuel-oil heat exchanger 1006 is configured such that heat is transferred from the oil flow to the fuel flow. In this way, the average temperature of the oil flow at the outlet of main fuel-oil heat exchanger 1006 is lower than the average temperature of the oil flow at the inlet of main fuel-oil heat exchanger 1006. In the same way, the average temperature of the fuel flow at the outlet of main fuel-oil heat exchanger 1006 is higher than the average temperature at the inlet of main fuel-oil heat exchanger 1006. Refrigeration cycle device 1007, when active (i.e., when powered on / opened and used to actively transfer heat from the oil to the fuel), can further increase the fuel temperature / decrease the oil temperature, optionally raising the fuel to a temperature higher than the oil temperature. After flowing through main fuel-oil heat exchanger 1006, the fuel flows to engine fuel pump 1003 which, in the example shown, is located downstream of primary fuel-oil heat exchanger 1006 and secondary fuel-oil heat exchanger 1004 and is arranged to deliver fuel to burner 16 of gas turbine engine 10.

[0442] Under cruise conditions, the average temperature of the oil flow through the IDG fuel-oil heat exchanger 1004 can be lower than the average temperature of the oil flow through the main fuel-oil heat exchanger 1006. In this way, the fuel first passes through the heat exchanger 1004 with a lower average oil flow temperature before passing through the heat exchanger 1006 with a higher average oil flow temperature. In addition to the branched oil flows 2000, 2000', the heat exchange system 3000 may also include a branched fuel return passage such that the fuel returns to the main fuel path, and the fuel enters the gas turbine engine 10 to the burner 16 from the main fuel path at at least two different locations (e.g., upstream or downstream of the main engine pump 1003), as Figure 6 shown, the branched path 6020 branches off the main fuel flow path from the outlet 1006b of the primary heat exchanger 1006 and rejoins the main fuel flow path downstream of the pump 1003. At least one valve (not shown) may be provided to control the diversion of the fuel flow returning from the heat exchanger 1006 through the engine 10 to the main fuel path. The valve may be controlled based on the fuel temperature. For example, if the fuel temperature is relatively high and is more likely to deteriorate the pump seal or other components, less fuel / more fuel is delivered to a location downstream of the pump via the pump 1003. The control of the fuel flow through the branched fuel return passage may be based on fuel temperature measurements (e.g., using temperature sensors at locations downstream of the heat exchanger and the fuel, and possibly also using temperature measurements upstream of the fuel-oil heat exchanger). The return of the recirculated fuel upstream of the fuel-oil heat exchanger may allow for a reduction in the heat transfer from the oil to the fuel, thereby suppressing the transient overshoot that may occur at the start of the landing phase. For example, at the start of the landing phase, the same amount of heat is generated within the oil system, but the fuel flow decreases, thus typically resulting in a temperature peak. By recirculating the fuel that has previously been heated to heat the system, the regulation of the fuel flow through these branched passages can also be used to reduce icing. Therefore, the branched fuel return passage can be used in multiple ways to improve engine thermal management. Figure 11 An exemplary configuration of the primary closed-loop oil system 2000 and the secondary closed-loop oil system 2000' is schematically shown, in which two separate recirculating oil flows form a heat exchange relationship through the oil-oil heat exchanger 2030. In this example, both of the oil circuit systems 2000, 2000' have a branched arrangement of parallel tubes / heat exchangers.

[0443] At Figure 11In the example shown, the primary closed-loop oil system 2000 is configured such that the recirculating oil flow is pumped by a feed pump 2004 through a valve 2016, which may be referred to as a modulating valve. The valve 2016 is operable to divert the oil flow such that a portion of the oil flow goes to each of the main fuel-oil heat exchanger 1006 and the first air-oil heat exchanger 2020. In the illustrated embodiment, the first air-oil heat exchanger 2020 is in series with an oil-oil heat exchanger 2030, and the air-oil heat exchanger 2020 and the oil-oil heat exchanger 2030 are arranged in parallel with the main fuel-oil heat exchanger 1006. The modulating valve 2016 determines what proportion of the oil travels through each branch of the parallel arrangement. In various embodiments, any suitable portion of the oil flow may be diverted between the main fuel-oil heat exchanger 1006 and the first air-oil heat exchanger 2020. In the example, the valve 2016 is operable to divert a fixed portion of the oil flow to each of the main fuel-oil heat exchanger 1006 and the first air-oil heat exchanger 2020. In other examples, the valve 2016 is operable to divert a variable portion of the oil flow to each of the main fuel-oil heat exchanger 1006 and the first air-oil heat exchanger 2020, such as using feedback from a temperature sensor to control the variable portion, and / or based on one or more fuel characteristics, as described below. The valve 2016 may be adjustable between a discrete number of set positions, or continuously adjustable. In some embodiments, for example, when the fuel temperature at the inlet of the burner 16 is relatively low compared to the maximum operating temperature of the fuel (e.g., based on knowledge of the fuel type or thermal stability), no oil may be delivered to the air-oil heat exchanger 2020, and all of the oil may be delivered via the fuel-oil heat exchanger 1006.

[0444] After flowing through the heat exchangers 1006, 2020, 2030, the oil flow in the primary closed-loop system 2000 is then recombined and delivered to the power gearbox 30 (and / or other engine components 33 such as the AGB) and then to the oil sump 2008. Then, a scavenge pump 2010 pumps the oil from the oil sump 2008 to the fuel tank 2002 for reuse.

[0445] The oil flow within the secondary closed-loop oil system 2000' is arranged to form a heat exchange relationship with a separate oil flow within the primary closed-loop oil system 2000 through an oil-oil heat exchanger 2030. In the oil-oil heat exchanger 2030, the oil flow within the primary closed-loop oil system 2000 does not mix with the oil flow within the secondary closed-loop oil system 2000'. The oil-oil heat exchanger 2030 is configured such that heat transfer can occur between the two separate oil flows. In this way, heat from the hotter oil flow can be transferred to the cooler oil flow within the oil-oil heat exchanger 2030. An air-oil heat exchanger is not shown in the illustrated secondary closed-loop oil system 2000', but in other examples, an air-oil heat exchanger may be provided, such as in series with the oil-oil heat exchanger 2030 or on a third parallel branch.

[0446] In Figure 11 the illustrated embodiment, the secondary closed-loop oil system 2000' is configured such that a recirculating oil flow is pumped by a secondary feed pump 2004' through a valve 2016', which may be referred to as a secondary modulating valve. The valve 2016' is operable to divert at least a portion of the oil flow between an IDG fuel-oil heat exchanger 1004 and the oil-oil heat exchanger 2030, where the oil-oil heat exchanger 2030 is arranged in parallel with the IDG fuel-oil heat exchanger 1004. In other embodiments, the secondary closed-loop oil system 2000' may not have a branched configuration. For example, the IDG fuel-oil heat exchanger 1004 and the oil-oil heat exchanger 2030 may be arranged in series such that at least a majority of the oil passing through one heat exchanger in a given cycle also passes through the other heat exchanger.

[0447] In other embodiments, an air-oil heat exchanger 2020 may be present on each closed-loop system, or only on the secondary closed-loop system, and / or the oil-oil heat exchanger may not be present.

[0448] In an example, any suitable portion of the oil flow can be diverted between the IDG fuel-oil heat exchanger 1004 and the oil-oil heat exchanger 2030. In an example, the valve 2016' is operable to divert a fixed portion of the oil flow to each of the IDG fuel-oil heat exchanger 1004 and the oil-oil heat exchanger 2030. In an example, the valve 2016' is operable to divert a variable portion of the oil flow to each of the IDG fuel-oil heat exchanger 1004 and the oil-oil heat exchanger 2030. The valve 2016' may be adjustable between a discrete number of set positions, or continuously adjustable. After flowing through the heat exchangers 1004, 2030, the oil flow is then delivered to an integrated drive generator 2006 and then to Figure 8The secondary oil sump 2008’ in the illustrated example. Then, the secondary oil return pump 2010’ pumps oil from the secondary oil sump 2008’ to the secondary oil tank 2002’ for reuse. It should be understood that in other embodiments, the oil circuit components may be arranged differently and there may be additional or alternative components of the oil circuit.

[0449] One or more temperature sensors may be provided, for example, arranged to sense the fuel temperature at the inlet of the fuel pump 1003 or the burner 16. The controller 58 of the heat exchange system 3000 may use data from one or more sensors, optionally in combination with other data (such as the fuel characteristics described above), to manage the fuel and / or oil flow through and around the heat exchangers 1004, 1006, 2020, 2030. The controller 58 may be part of or provided by the EEC, or may be a separate unit. The control may be automatic, for example, by the EEC.

[0450] The inventors recognize that in various embodiments, knowledge of one or more fuel characteristics can be used to customize the operation of the heat exchange system 3000 in order to improve the performance of the engine 10 by taking advantage of the varying properties between aviation fuels.

[0451] Figure 12 A first method 100 is shown, including using knowledge of fuel characteristics to determine control of the heat exchange system 3000. The method 100 includes determining 102 at least one fuel characteristic of the fuel arranged to be burned by the burner 16. This determination 102 may be performed by any of the methods outlined above, optionally using the fuel composition determination module 57.

[0452] The engine 10 for implementing the method 100 includes an air - oil heat exchanger 2020 and a fuel - oil heat exchanger 1006, as well as a modulating valve 2016 arranged to allow changing the proportion of oil delivered via each heat exchanger, and the method 100 includes controlling 104 the modulating valve based on at least one fuel characteristic so as to adjust the proportion of oil delivered via each heat exchanger under cruise conditions.

[0453] More specifically, for the above reference Figures 5 to 11In the described embodiments, method 100 includes controlling 104 a modulation valve 2016 of a primary oil system 2000 under cruise conditions based on at least one fuel characteristic so as to regulate the proportion of oil delivered via each branch of the primary oil circuit system 2000, thereby controlling the amount of oil passing through the main fuel - oil heat exchanger 1006 and the amount of oil passing through the air - oil heat exchanger 2020. In embodiments with different arrangements of the heat exchangers, such as a series arrangement of the heat exchangers instead of a parallel split, the modulation valve 2016 can effectively be a bypass valve, allowing one of the heat exchangers 1006, 2020 to be bypassed by a controllable proportion of the flow, thus allowing the proportion of oil delivered via each heat exchanger to be changed in this way.

[0454] As Figure 12 shown by the dashed line in, the method 100 can optionally be repeated. In some embodiments, such as embodiments where there is only one type of fuel on the aircraft 1, the method 100 can be performed only once during a flight cycle, such as during fuel replenishment. However, in other embodiments, such as embodiments with multiple fuel sources, the fuel supplied to the burner 16 may vary over time during flight. Therefore, determination 102 can be performed more than once - for example:

[0455] (i) In embodiments with multiple fuel sources, at any given time only one or the other fuel source is used (and knowledge of which fuel source the fuel is drawn from can prompt selection of appropriate stored fuel characteristics), and the determination is performed once for each fuel source 50, 53 at the time of fuel replenishment / at the start of the flight; or

[0456] (ii) The determination is performed frequently during flight, such as in response to a change in which (which of) the fuel source(s) the fuel is drawn from (note that in some embodiments, the fuel supplied to the burner 16 can be a mixture of fuels from different sources), or is performed periodically.

[0457] Accordingly, the modulation valve 2016 can be arranged to divert a fixed portion of the oil flow to each of the fuel-oil heat exchanger 1006 and the air-oil heat exchanger 2020 during operation of the engine 10, the fixed portion being determined based on at least one determined fuel characteristic at engine 10 startup (option (i) above). Alternatively, the modulation valve 2016 can be arranged to divert a variable portion of the oil flow to each of the fuel-oil heat exchanger 1006 and the air-oil heat exchanger 2020 during operation of the engine 10 during a single flight (option (ii) above). The modulation valve 2016 can thus be actively controlled to vary the proportion of oil delivered via each heat exchanger 1006, 2020, particularly in embodiments where the aircraft 1 carries multiple different fuels in different tanks, and can vary the fuel (or fuel mixture) used during flight. The active control of the modulation valve can be automatic and implemented by the controller 58 of the heat exchange system 3000, which can be a dedicated controller or part of a more general EEC. In various embodiments, based on the availability of online measurements, the control of the modulation valve 2016 can be closed-loop or open-loop - in particular, when feedback data (e.g., from an oil flow rate sensor) is available, a closed-loop control method may be preferred.

[0458] The oil circuit system 2000 (which can also be referred to as a recirculating oil system) can be branched such that a certain proportion of the oil can flow along each branch, and the air-oil heat exchanger 2020 and the fuel-oil heat exchanger 1006 can be arranged in a parallel configuration on different branches of the oil circuit system, as Figure 9 and Figure 11 shown. The modulation valve 2016 in such examples can be arranged to allow varying the proportion of oil delivered via each branch, and the control 104 of the modulation valve 2016 can thus regulate the proportion of oil delivered via each branch under cruise conditions. In some examples, more than two heat exchangers and / or more than two branches can be provided. The modulation valve 2016 thus allows varying the heat transferred from the oil by the air-oil heat exchanger 2020 and the fuel-oil heat exchanger 1006. The inventors have recognized that fuels with certain fuel characteristics, for example due to having a higher heat capacity and / or higher thermal stability, can accept more heat without decomposing than other fuels, or can benefit from being at a higher temperature (e.g., for fuel combustion efficiency). Accordingly, adaptively controlling the heat exchange system 3000 based on determined fuel characteristics enables more efficient use of different fuels as well as more efficient engine operation and oil cooling.

[0459] Fuel characteristics may not be the only data considered when adjusting the modulating valve 2016 - for example, the temperature of the fuel leaving the fuel - oil heat exchanger 1006 or entering the burner 16 may also be considered, as well as the heat resistance of the fuel system components downstream of the heat exchanger 1006 (and optionally also the refrigeration cycle device 1007, if present). In the presence of the refrigeration cycle device 1007, the modulating valve 2016 or another oil valve can control the amount of oil flowing through the refrigeration cycle device 1007. In some embodiments, all of the oil passing through the main fuel - oil heat exchanger 1006, and optionally all of the fuel passing through the heat exchanger 1006, may also pass through the refrigeration cycle device 1007 - however, the refrigeration cycle device 1007 may be powered off / inoperable for some period of time and thus not provide any temperature boost.

[0460] Depending on the suitability of at least one determined fuel characteristic, method 100 may include using the modulating valve 2016 to perform one or more of the following:

[0461] ● During one or more periods of at least 30 minutes of cruise, transport all of the oil via the fuel - oil heat exchanger 1006 (without via the air - oil heat exchanger 2020), such that during at least some operating periods of the cruise, no heat is lost to the environment via the air - oil heat exchanger;

[0462] ● During at least 15% and optionally at least 20% of the total cruise time, transport all of the oil via the fuel - oil heat exchanger 1006 (without via the air - oil heat exchanger 2020);

[0463] ● During at least 90% of the cruise time, transport at least 95% of the oil via the fuel - oil heat exchanger 1006 (not more than 5% of the oil transported via the air - oil heat exchanger 2020), such that during at least 90% of the cruise operations, very little heat (if any)

[0464] is lost to the environment via the air - oil heat exchanger;

[0465] ● Control the oil flow such that at least 80% of the heat transferred from the oil during cruise is transferred to the fuel; and / or

[0466] ● Control the oil flow such that all of the heat transferred from the oil is transferred to the fuel during at least 90% of the cruise time; and / or

[0467] ● Control the oil flow such that under cruise conditions, 200 kJ to 600 kJ and optionally 200 kJ to 500 kJ of heat is transferred from the oil to the fuel per kilogram of fuel in the heat exchange system 3000. In some embodiments, the heat transferred may be in the range of 350 kJ / kg to 450 kJ / kg of fuel.

[0468] The applicability of at least one determined fuel characteristic can be determined by comparing one of a plurality of fuel characteristics with a threshold value set for the corresponding fuel characteristic. For example, fuels with a SAF content (% SAF) exceeding 60%, 65%, 70%, 75% or 80% can be considered suitable for reducing the oil flow through the air-oil heat exchanger 2020 compared to conventional aviation fuels. In some embodiments, multiple fuel characteristics can be checked simultaneously - for example, when one fuel characteristic is within a specific range, a lower threshold value of another fuel characteristic can be used. In other embodiments, the fuel characteristics can be considered independently / isolatly.

[0469] At least one fuel characteristic can be or include one or more of the thermal stability of the fuel, the aromatic content, or the percentage of sustainable aviation fuel in the fuel, and the modulation valve control decision can be made based on the knowledge of at least one fuel characteristic. For example, if the thermal stability data indicates that the fuel can operate stably at a temperature higher than 160 °C or 170 °C, the modulation valve 2016 can be controlled such that at least 80%, and optionally 90% to 100%, of the heat transferred from the oil during cruise can be transferred to the fuel. As another example, if the molar percentage of aromatics in the fuel is less than 12% and optionally less than 10% or less than 5%, at least 80% and optionally 90% to 100% of the heat transferred from the oil during cruise can be transferred to the fuel. As another example, if the SAF content of the fuel exceeds 50% and optionally exceeds 75%, the modulation valve 2016 can be controlled such that at least 80% and optionally 90% to 100% of the heat transferred from the oil during cruise can be transferred to the fuel. As another example, if the calorific value of the fuel is at least 43.5 MJ / kg, at least 80% and optionally 90% to 100% of the heat transferred from the oil during cruise can be transferred to the fuel. Generally, if the fuel calorific value exceeds the threshold, less heat can be transferred to the conventional fuel because for fuels with a higher calorific value, the flow rate is usually reduced to achieve the same level of thrust without wasting fuel - if the amount of oil passing through the heat exchanger 1006 is not reduced, the fuel in the fuel-oil heat exchanger 1006 can thus experience an increased temperature rise due to its reduced flow rate. However, in many cases, new fuels with higher thermal stability (such as SAF) also have a higher calorific value - thus, in some cases, the increase in fuel temperature rise due to the lower flow rate is acceptable, and in fact, it is also possible to further increase the heat. This demonstrates the utility of considering multiple fuel characteristics in combination rather than considering individual fuel characteristics separately in some embodiments.

[0470] Any one of the embodiments may use only one or more fuel characteristics. A processing module may be provided, optionally as part of the fuel characteristic determination module 57, and / or as part of the general engine EEC, to make decisions regarding the control of the modulation valve 2016 based on the fuel characteristic data and optionally also on other data.

[0471] The gas turbine engine 10 for an aircraft implementing the method 100 includes: a gearbox 30 that receives an input from a core shaft and outputs a drive to a fan 23; and an oil circuit system 2000 that is arranged to supply oil to the gearbox 30. The heat exchange system 300 of the engine 10 includes: an air-oil heat exchanger 2020 through which the oil in the oil circuit system 2000 flows; and a fuel-oil heat exchanger 1006 through which the oil and fuel in the oil circuit system flow, such that heat is transferred between the oil and the fuel. A modulation valve 2016 is provided, which is arranged to allow changing the proportion of oil delivered via each heat exchanger. In Figure 9 the various examples shown, the oil circuit system 2000 branches such that a certain proportion of the oil can flow along each branch, and the air-oil heat exchanger 2020 and the fuel-oil heat exchanger 1006 are arranged in a parallel configuration, with one heat exchanger arranged on each branch of the oil circuit system 2000. The modulation valve 2016 is arranged to allow changing the proportion of oil delivered via each branch. The engine 10 further includes a fuel composition determination module 57 that is arranged to determine at least one fuel characteristic of the fuel arranged to be burned by a burner 16. The modulation valve 2016 is arranged to be controlled based on at least one fuel characteristic determined by the fuel composition determination module 57 so as to adjust the proportion of oil delivered via each heat exchanger 1006, 2020 under cruise conditions. A controller 58 may be provided to make and implement decisions based on the output of the fuel composition determination module 57.

[0472] In Figure 11 the various examples shown, the heat exchange system 3000 further includes the secondary oil circuit system 2000' and the oil-oil heat exchanger 2030 as described above, which is arranged to allow heat to be exchanged between the oil in the two oil systems 2000, 2000'. The modulation valve 2016 or an additional valve may control the oil flow from the primary oil circuit system 2000 to the oil-oil heat exchanger 2030. The secondary oil circuit system 2000' may also have a valve to control the oil flow to the oil-oil heat exchanger 2030. In Figure 11 the example shown, each of the oil systems 2000, 2000' has a total of two parallel branches, and the oil-oil heat exchanger 2030 is on the same branch as the air-oil heat exchanger 2020 of the primary oil system 2000.

[0473] The various example heat exchange systems 3000 include at least one bypass tube arranged to allow fuel (or oil) to bypass one or more heat exchangers. In some embodiments, the bypass tube can effectively form an additional branch in a parallel branch oil system, and a modulation valve 2016 or another oil valve can be arranged to regulate the amount of oil delivered through the bypass tube or each bypass tube for oil based on one or more determined fuel characteristics and optionally based on one or more temperature measurements.

[0474] Figure 13 A second method 200 is shown, including using knowledge of fuel characteristics to determine control of the heat exchange system 3000. Method 200 includes determining 202 at least one fuel characteristic of the fuel arranged to be provided to and burned by the burner 16. This determination 202 can be performed by any of the methods outlined above, optionally using a fuel composition determination module 57. The details of this determination step 202 can be the same as the determination 102 described for Figure 12 the method 100 shown. Figure 13 The method further includes modulating 204 the heat exchange system 3000 so as to adjust the fuel temperature at the inlet of the burner 16 during cruise to a set level based on at least one determined fuel characteristic.

[0475] The modulation 204 of the heat exchange system 3000 can include any one or more of the following:

[0476] ● Adjusting any suitable oil or fuel valve, such as the modulation valve 2016, to change the proportion of oil or fuel delivered via a particular heat exchanger or via the refrigeration cycle device 1007 (if present);

[0477] ● Adjusting the modulation valve 2016 (if present) to change the proportion of oil delivered via one or more branches of the oil system 2000;

[0478] ● Adjusting a bypass valve (if present) so as to adjust the proportion of fluid (such as oil or fuel) bypassing a particular heat exchanger or particular group of heat exchangers;

[0479] ● Adjusting a recirculation valve 6010 (if present) so as to adjust the proportion of fluid (such as oil or fuel) that returns to an earlier point in its path and thus passes through one or more components (usually including a particular heat exchanger or particular group of heat exchangers) multiple times;

[0480] ● Activating, deactivating, or otherwise adjusting the use of the refrigeration cycle device 1007 (if present) so as to provide a controlled enhancement of heating of one fluid / cooling of another fluid;

[0481] ● Adjust the pumping speed of the oil system pumps 2010, 2010';

[0482] ● When the oil system or each oil system has a branched oil passage, close one or more oil flow branches in the oil systems 2000,

[0483] 2000'; and / or

[0484] ● Adjust the valve to control the air flow to the air-oil heat exchanger 2020 (if present), or change the air flow through the heat exchanger 2020.

[0485] As Figure 13 shown by the dashed line in, the method 200 can optionally be repeated multiple times during flight / operation of the engine 10, as described above with respect to the Figure 12 method. Thus, in some embodiments, both the determination 202 and the subsequent modulation 204 can be performed during cruise.

[0486] Fuel characteristics may not be the only data considered when modulating the heat exchange system 3000. For example, the temperature of the ambient / surrounding air available for the air-oil heat exchanger 2020 (if present) can be considered (optionally using altitude - optionally in combination with a geographical region - as a temperature reference), as can the fuel flow rate and fuel temperature.

[0487] In some embodiments, the step 204 of modulating the heat exchange system 3000 to control the fuel temperature includes controlling the oil flow through at least one fuel-oil heat exchanger 1006 such that between 50% and 100% of the heat lost from the oil is transferred to the fuel, and optionally such that between 80% and 100% of the heat lost from the oil, or between 90% and 100% of the heat lost from the oil, is transferred to the fuel.

[0488] The determined fuel characteristics can be or include the thermal stability of the fuel; in such embodiments, the heat exchange system 3000 can be modulated such that the fuel temperature at the inlet of the burner 16 during cruise increases as the thermal stability increases, optionally linearly. Thermal stability refers to the temperature at which the fuel begins to decompose and form deposits that can cause component blockages and failures - it should be understood that this is generally not an instantaneous decomposition at a specific temperature; aviation fuel includes various components that are prone to decomposition at different temperatures, and the time spent at high temperatures is also a factor in fuel decomposition. If at a given temperature, the decomposition rate of the fuel is below a certain threshold, the fuel can be classified as stable at that temperature.

[0489] In an alternative or additional embodiment, the determined fuel characteristic may be or include the presence of a tracer substance in the fuel. In such an embodiment, the heat exchange system 3000 may be modulated such that the fuel temperature at the burner inlet during cruise is set to a predefined level corresponding to the tracer substance. Optionally, the amount of tracer substance detected may indicate whether the fuel is a mixture of a labeled aviation fuel (i.e., fuel with a tracer substance added) and another fuel, or a pure labeled aviation fuel.

[0490] In an alternative or additional embodiment, the determined fuel characteristic may be or include the percentage of SAF in the fuel (%SAF). In such an embodiment, the heat exchange system 3000 may be modulated such that once the %SAF exceeds 60%, or optionally once the %SAF exceeds 70%, 75% or 80%, the fuel temperature at the inlet of the burner 16 during cruise increases as the %SAF increases. At lower %SAF, in some embodiments modulation based on the SAF content may not be performed, but modulation based on one or more other characteristics 204 may be performed - thus it will be understood that multiple different fuel characteristics may be considered in combination (in other embodiments, modulation based on the SAF content may be performed at lower %SAF). Any one embodiment may use only one or more fuel characteristics. A processing module may be provided, optionally as part of the fuel characteristic determination module 57, and / or as part of the general engine EEC, to make decisions regarding the regulation 204 of the heat exchange system based on fuel characteristic data and optionally other data.

[0491] The gas turbine engine 10 for an aircraft implementing the method 200 does not necessarily include a gearbox 30 that receives an input from a core shaft and outputs a drive to a fan 23, i.e., the method 200 can be used for a geared engine 10 and a direct drive engine. Thus, the method 200 can be applied to a wide variety of aircraft engines as long as these aircraft engines include a burner 16, an oil system 2000 arranged to circulate oil, and a heat exchange system 3000 that includes at least one fuel - oil heat exchanger arranged to transfer heat from the oil to the fuel. A fuel composition determination module 57 is also provided, which is arranged to determine at least one fuel characteristic of the fuel arranged to be burned by the burner 16, the fuel composition determination module being an independent unit or as part of the EEC or other engine control system. The fuel composition determination module 57 may include a processing module or communicate with a processing module, which is arranged to make decisions based on one or more determined fuel characteristics and optionally other data.

[0492] In some embodiments, the heat exchange system 3000 includes a branched fuel return passage and at least one valve that controls the diversion of fuel flow. As described above, the branched passage may be arranged to return fuel from the heat exchange system 3000 - particularly fuel leaving the primary fuel - oil heat exchanger 1006 or the secondary fuel - oil heat exchanger 1004 - to at least two different locations along the main fuel path, from which the fuel enters the gas turbine engine 10 to reach the burner 16. For example, the fuel flow leaving the heat exchangers 1004, 1006 may be split into two parts, one path generally carrying most of the fuel (and thus referred to as the main fuel flow passage), and the other path diverting a portion of the fuel leaving the heat exchangers 1004, 1006 such that the diverted portion of the fuel bypasses one or more main fuel flow path components (such as pumps) and rejoins the main stream downstream of those one or more main fuel flow path components to the burner 16. The valve can control and regulate the proportion delivered via each branch. In some implementations, more than two branches may be provided.

[0493] The inventors have also recognized that using fuels different from traditional kerosene - based jet fuels (such as sustainable aviation fuels) may result in different fuel properties, and the heat transfer parameters in operation can be adjusted to take advantage of the different fuel properties. In particular, some fuels can be heated to higher temperatures in one or more fuel - oil heat exchangers than traditional fuels without significantly increasing coking. This can enable a method of providing improved oil cooling (since the fuel can absorb more heat), and can also improve the overall thermal efficiency of the engine while losing less heat to the surrounding environment. The controllable heat exchange system 3000 plays a key role in managing the heat transfer ratio.

[0494] In addition, the inventors have recognized that while cruise conditions typically account for most of the flight time of an aircraft engine, idle operation is also important because the fuel mass flow rate at idle is much lower than during cruise, and even though the heat load of the fuel is relatively small, it can still cause temperature increases - thus, the use of non - traditional fuels may have a greater impact on the optimal heat management method under idle conditions. Figure 14 and Figure 15 Methods 300, 400 address these two scenarios of aircraft operation.

[0495] Figure 14 Method 300 shows a way to achieve these considerations under cruise conditions, and Figure 15 Method 400 shows a way to achieve these considerations at idle (e.g., when the aircraft is starting up, operating stationary during boarding, and (taxiing towards the runway or hangar, or between other ground - based locations), or during certain periods of flight (such as landing)).

[0496] First, consider method 300 to be performed during cruise. Method 300 is arranged to be performed in a geared gas turbine engine 10, which includes: an oil circuit system 2000 arranged to supply oil to a gearbox 30; and a heat exchange system 3000 including an air-oil heat exchanger 2020 through which the oil in the oil circuit system flows; and a fuel-oil heat exchanger 1006 through which the oil and fuel in the oil circuit system flow such that heat is transferred between the oil and the fuel. As Figure 9 and Figure 11 shown, the oil circuit system 2000 branches such that a certain proportion of the oil can flow along each branch, and the air-oil heat exchanger 2020 and the fuel-oil heat exchanger 1006 are arranged in a parallel configuration on different branches of the oil circuit system. A modulating valve 2016 is also provided to control the proportion of oil delivered through each branch of the oil circuit system 2000. Method 300 includes controlling 302 the heat exchange system 3000 such that, under cruise conditions, the heat transfer ratio:

[0497]

[0498] is in the range from 0 to 0.67, and optionally in the range from 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 such control.

[0499] It should be understood that even for a particular engine 10 operating on a set fuel, due to varying conditions, this ratio typically has a range of values during cruise. For example, the upper limit of the range may apply to cold days (ISA - 30 conditions) at low altitude (cruise: 35,000 ft) where the oil system heat generation is low, and the lower limit may apply to hot days (ISA + 40 conditions) at high altitude (cruise: 39,000 ft) where the oil system heat generation is high.

[0500] Method 300 may also include receiving data 304 to allow calculation or inference of a heat transfer ratio, such as temperature data (temperature data of oil and / or fuel at one or more points around the oil circuit system 2000, 2000' or the fuel flow path, and / or optionally fuel tank temperature or tank temperature) and fuel flow data. Such data may be received 304 by the controller 58 and used 306 to adjust the modulation 302 of the heat exchange system 3000 during cruise so as to maintain the heat transfer ratio within a desired level or desired bounds. Such checking and adjustment / correction 306 may be performed periodically or in response to a predetermined stimulus (e.g., a change in temperature or flow rate, or a change in engine operation or altitude). These steps 304, 306 may alternatively be considered part of the control 302 of the heat exchange system 3000. Method 300 may also be arranged to utilize other information when determining 306 what control action to take, such as temperature data (ambient temperature of the oil, fuel, and / or the environment around the aircraft 1), flow rate data (oil and / or fuel), and / or one or more fuel properties.

[0501] The step of controlling 302 the heat exchange system 3000 may include any one or any combination of the examples provided for the modulation step 204 of the method 200, e.g., when the heat transfer ratio is too high, by reducing the amount of oil delivered via at least one air-oil heat exchanger 2020, or by adjusting the ratio of oil and / or fuel delivered via the respective bypass tube 1005 or recirculation tube 6011. Figure 13 In embodiments where the engine 10 includes a refrigeration cycle device 1007, the step of controlling 302 the heat exchange system 3000 may include using the refrigeration cycle device 1007 to provide a heat boost by transferring more heat from the oil to the fuel than is transferred through the fuel-oil heat exchanger 1006, such that in some cases the fuel temperature is raised to be higher than the oil temperature. In such embodiments, a lower heat transfer ratio value may be obtained compared to embodiments without the refrigeration cycle device 1007, e.g., the heat transfer ratio during cruise may not exceed 0.40, and optionally not exceed 0.38, and further optionally not exceed 0.10 or 0.05. In embodiments without the refrigeration cycle device 1007, or in embodiments where such a device is present but not used, a higher heat transfer ratio value during cruise may be maintained, e.g., greater than 0.38, and optionally greater than 0.40 or 0.50.

[0502]

[0503] The fuel temperature upon reaching the burner 16 can also be considered. The method 300 can include controlling 302 the heat exchange system 3000 under cruise conditions such that if the fuel temperature at the inlet of the burner 16 is at least 160 °C, the heat transfer ratio is in the range of 0 to 0.2, or if the fuel temperature at the inlet of the burner 16 is at least 180 °C, the heat transfer ratio is in the range of 0 to 0.1. Thus, one or more temperature sensors can be used, and their outputs can be considered when setting the control 302 of the heat exchange system 3000.

[0504] The method 300 can also consider the properties of the fuel, such as whether the fuel is or includes sustainable aviation fuel (SAF). For example, if the fuel has at least 70% SAF, the heat transfer ratio during cruise can be maintained in the range of 0 to 0.2, and if the fuel has at least 80% SAF, the heat transfer ratio during cruise can be maintained in the range of 0 to 0.1. Thus, one or more fuel properties—optionally determined by the fuel property determination module 57 as described above—can also be used; this information can be considered when setting the control 302 of the heat exchange system 3000.

[0505] The heat transfer rate from the oil to the air during cruise (in kJ per kg of fuel flow reaching the burner 16) can be in the range of 0 kJ / kg to 240 kJ / kg, and optionally in the range of 0 kJ / kg to 120 kJ / kg. In some embodiments, no more than 20% of the heat transferred away from the oil during cruise can be transferred to the air, and the heat transfer rate from the oil to the air under cruise conditions can be maintained in the range of 0 kJ to 100 kJ per kg of fuel.

[0506] The heat transfer rate from the oil to the fuel during cruise (in kJ per kg of fuel flow reaching the burner 16) can be in the range of 85 kJ / kg to 350 kJ / kg, and optionally in the range of 85 kJ / kg to 170 kJ / kg. In some embodiments, at least 80% of the heat transferred away from the oil during cruise can be transferred to the fuel, and the heat transfer rate from the oil to the fuel under cruise conditions can be maintained in the range of 110 kJ to 240 kJ per kg of fuel.

[0507] In some embodiments, the heat exchange system 3000 further includes a branched fuel return passage 6020 and at least one valve that controls the diversion of fuel flow, the branched passages being arranged to return fuel from the heat exchange system 3000 to at least two different locations along the main fuel path from which fuel enters the gas turbine engine 10 to reach the combustor 16. The valve can be controlled based on feedback from one or more temperature sensors and / or based on fuel characteristics. The method 300 can include regulating the fuel flow along each branch based on the heat transfer ratio and / or the fuel temperature when leaving the heat exchangers 1004, 1006.

[0508] For a particular geared engine 10, at cruise conditions, the percentage of the total heat transferred away from the oil in the heat exchange system 3000 that is transferred to the fuel (rather than to the air / environment) is plotted as a function of the fuel temperature limit in Figure 19 . The fuel temperature limit is the maximum fuel temperature that is considered safe for aircraft operation with that fuel and that engine 10, and can depend on fuel characteristics such as thermal stability and the heat resistance of engine components. Curve 3 shows the lower limit 4 (black dashed line) and upper limit 5 (black dashed line) of the percentage of heat transferred to the fuel - the upper and lower limits illustrate different cruise fuel flows (depending on altitude, rated power, etc.), variability in engine heat generation, aircraft fuel temperature, atmospheric temperature, and other variables. Generally, for the example engine 10 for which these data are provided, for a fuel with a temperature limit of 120°C, the lower limit can be 20% of the heat transferred to the fuel, and the upper limit of all heat loss from the oil is transferred to the fuel (100%). Curve 3 also shows a line 6 (gray solid line) for typical heat transfer to the fuel, with the fuel temperature limit. For a fuel with a temperature limit of 120°C, a typical value of the percentage of heat transferred to the fuel can be 60%; for fuel temperature limits of 170°C and above, this can increase to 100%. When the fuel temperature limit is equal to or greater than 260°C, all of the heat can be transferred to the fuel during cruise without air cooling, such that the heat transfer ratio is zero.

[0509] Turning now to method 400, which will be performed at idle, method 400 is again arranged to be performed in a geared gas turbine engine 10 that includes: an oil circuit system 2000 arranged to supply oil to the gearbox 30; and a heat exchange system 3000 that includes an air - oil heat exchanger 2020 through which the oil in the oil circuit system flows; and a fuel - oil heat exchanger 1006 through which the oil and fuel in the oil circuit system flow such that heat is transferred between the oil and the fuel. The oil circuit system 2000 is as Figure 9 and Figure 11The branches are arranged such that a certain proportion of the oil can flow along each branch, and the air-oil heat exchanger 2020 and the fuel-oil heat exchanger 1006 are arranged in a parallel configuration on different branches of the oil circuit system. A modulating valve 2016 is again provided to control the proportion of the oil delivered via each branch of the oil circuit system 2000. Method 400 includes controlling 402 the heat exchange system 3000 such that when the aircraft 1 is operating under idle conditions, the heat transfer ratio is:

[0510]

[0511] In the range of 0.67 to 5.67. Thus, the ratio at idle may be higher than the ratio during cruise. The ground idle ratio is higher than the flight idle ratio. Optionally, method 400 may include maintaining the heat transfer ratio below 5.50, 5.0, 4.5, 4.0, 3.5, or 3.0 under idle conditions. Optionally, method 400 may include maintaining the heat transfer ratio at idle above 4.0, 4.5, or 5.0 under ground idle conditions.

[0512] Method 400 may also include receiving data 404 to allow calculation or inference of the heat transfer ratio, such as temperature data (temperature data of the oil and / or fuel at one or more points around the oil circuit system 2000, 2000', or the fuel flow path, and / or optionally the fuel tank temperature or the tank temperature) and fuel flow data. Such data may be received 404 by the controller 58 and used 406 to adjust the modulation 302 of the heat exchange system 3000 at idle so as to maintain the heat transfer ratio within a desired level or desired bounds. Such checking and adjusting / correcting 406 may be performed periodically or in response to a predetermined stimulus (e.g., a change in the temperature or flow rate of the fuel or oil, or a change in the aircraft operation, such as starting to taxi). These steps 404, 406 may alternatively be considered part of the control 402 of the heat exchange system 3000. Method 400 may also be arranged to utilize other information when determining 406 what control action to take, such as temperature data (ambient temperature of the oil, fuel, and / or the surroundings of the aircraft 1), flow data (oil and / or fuel), and / or one or more fuel characteristics.

[0513] The step of controlling 402 the heat exchange system 3000 may include any one or any combination of the examples provided for the modulating step 204 of Figure 13 method 200, for example, when the heat transfer ratio is too low, by increasing the amount of oil delivered via at least one air-oil heat exchanger 2020, or by adjusting the proportion of oil and / or fuel delivered via the corresponding bypass pipe 1005 or recirculation pipe 6011.

[0514] In an embodiment where the engine 10 includes the refrigeration cycle device 1007, the step of controlling 402 the heat exchange system 3000 at idle may include using the refrigeration cycle device 1007 to provide a heat boost by transferring more heat from the oil to the fuel than is transferred through the fuel-oil heat exchanger 1006, such that in some cases the fuel temperature is raised above the oil temperature. In such embodiments, a lower heat transfer ratio can be obtained than in embodiments without the refrigeration cycle device 1007. For example, the heat transfer ratio at idle may not exceed 4, and optionally may not exceed 4.0, 3.5, 3.0, 2.5, 2.0, 1.5, or 1.0, and further optionally may not exceed 0.10 or 0.05. In embodiments without the refrigeration cycle device 1007, or in embodiments where such a device is present but not used, a higher heat transfer ratio can be maintained, such as greater than 3.37, and optionally greater than 3.4, 3.5, 4.0, 4.5, 5.0, or 5.5.

[0515] The fuel temperature at arrival at the burner 16 can also be considered. The method 400 may include controlling 402 the heat exchange system 3000 during idle operation such that if the fuel temperature at the inlet of the burner 16 is below 200 °C, the heat transfer ratio is in the range of 2.33 to 5.67, or if the fuel temperature at the inlet of the burner 16 is at or above 200 °C, the heat transfer ratio is in the range of 0.67 to 4, or in the range of 0.67 to 2.67 when the fuel temperature at the inlet of the burner 16 is at or above 250 °C, or if the fuel temperature at the inlet of the burner 16 is at or above 280 °C, the heat transfer ratio is in the range of 0.67 to 1.22. Accordingly, one or more temperature sensors can be used, and their outputs can be considered when setting the control 302 of the heat exchange system 3000.

[0516] The method 400 can additionally or alternatively consider the properties of the fuel, such as whether the fuel is or includes sustainable aviation fuel (SAF). For example, if the fuel is at least 60% or 70% sustainable aviation fuel, the heat transfer ratio at idle can be maintained in the range of 0.67 to 3.67, or if the fuel is at least 75%, 80%, or 85% sustainable aviation fuel, the heat transfer ratio can be maintained in the range of 0.67 to 2.67.

[0517] Accordingly, one or more fuel properties — optionally determined by the fuel property determination module 57 as described above — can also be used; this information can be considered when setting the control 402 of the heat exchange system 3000.

[0518] The heat transfer rate from oil to air at idle (in kJ per kg of fuel flow reaching burner 16) can be in the range of 380 kJ / kg to 1430 kJ / kg, and optionally in the range of 380 kJ / kg to 1270 kJ / kg, or in the range of 500 kJ / kg to 1080 kJ / kg. In some embodiments, no more than 80% of the heat transferred away from the oil at idle can be transferred to the air, and the heat transfer rate from oil to air under idle conditions can be maintained in the range of 630 kJ to 1430 kJ per kg of fuel. At idle, a higher percentage of air cooling is typically used than under cruise and other higher power engine conditions.

[0519] The heat transfer rate from oil to fuel at idle (in kJ per kg of fuel flow reaching burner 16) can be in the range of 0 kJ / kg to 1270 kJ / kg, and optionally in the range of 190 kJ / kg to 760 kJ / kg. In some embodiments, at least 20% of the heat transferred away from the oil at idle can be transferred to the fuel, and the heat transfer rate from oil to fuel under idle conditions can be maintained in the range of 150 kJ to 360 kJ per kg of fuel.

[0520] Regarding reference Figure 14 For method 300 described, in some embodiments, the heat exchange system 3000 further includes a branched fuel return passage and at least one valve controlling the diversion of fuel flow, the branched passages being arranged to return fuel from the heat exchange system 3000 to at least two different locations along the main fuel path from which fuel enters the gas turbine engine 10 to reach the burner 16. The valve can be controlled based on feedback from one or more temperature sensors and / or based on fuel characteristics. Method 400 can include regulating the fuel flow along each branch based on the heat transfer ratio and / or the fuel temperature when leaving the heat exchangers 1004, 1006.

[0521] Figure 16 A recirculating oil system including two loops 2000, 2000' is shown, but without heat exchangers arranged on parallel branches as Figure 11 shown. Instead, each oil loop 2000, 2000' provides a main oil flow path in series through all the heat exchangers in that loop. To regulate the oil flow through the heat exchangers, one or more bypass tubes 2005, 2005', 2005a are provided in place of branched paths leading to different heat exchangers. It should be understood that although the bypass tubes technically provide parallel, alternative flow paths, the branched paths described here with reference to parallel flow differ in that each branched route in the parallel arrangement includes a heat exchanger, while the bypass tubes are merely tubes on which there are no (significant) heat exchangers or other components.

[0522] It should be understood that in some embodiments, a combination of (i) a parallel arrangement of heat exchangers and (ii) bypass pipes can be used, and Figure 11 (both loops having parallel branches and no bypass pipes) and Figure 16 (all heat exchangers being parallel and having multiple bypass pipes) can be considered to show two different ends of the design range.

[0523] In Figure 16 the example shown, the fuel flow is represented by a thick black line to provide context for how the fuel and oil systems interact.

[0524] Figure 16 The primary oil circuit system 2000 shown in [[ ]] provides a series oil flow path that starts at tank 2002, passes through oil pump 2004, forward through primary fuel - oil heat exchanger 1006, then through air - oil heat exchanger 2020, then into gearbox 30 (and optionally other components to be cooled and lubricated), then is collected in oil sump 2008, and then is pumped back to tank 2002 by oil pump 2010. Thus, the two heat exchangers 1006, 2020 are arranged in series. In an alternative embodiment, the order of the two heat exchangers 1006, 2020 can be reversed such that the fuel - oil heat exchanger 1006 is after the air - oil heat exchanger 2020.

[0525] Figure 16The primary oil circuit system 2000 shown includes two oil bypass pipes 2005, 2005a. The first bypass pipe 2005 is arranged to allow a portion of the oil to bypass the fuel-oil heat exchanger 1006 and is controlled by the first bypass valve 2007. The first bypass pipe 2005 takes oil upstream of the inlet of the primary heat exchanger 1006 and returns it to the main oil flow path downstream of the primary heat exchanger 1006 and upstream of the air-oil heat exchanger 2020. The second bypass pipe 2005a is arranged to allow a portion of the oil to bypass the air-oil heat exchanger 2020 and is controlled by the second bypass valve 2007a. The second bypass pipe 2005a takes oil upstream of the inlet of the air-oil heat exchanger 2020 and returns it to the main oil flow path downstream of the air-oil heat exchanger 2020, and then the path reaches the gearbox 30 (and optionally other components to be cooled and lubricated). In an embodiment of the primary oil circuit system 2000 having only one oil bypass pipe 2005a, the selected location can be the location of the second bypass pipe 2005a such that there is a bypass for the air-oil heat exchanger 2020 while there is no bypass for the fuel-oil heat exchanger 1006. This helps to absorb the heat in the oil into the fuel as safely as possible and maintain the high thermal efficiency of the engine by reducing the heat loss to the environment. Having the bypass pipe 2005 on the fuel-oil heat exchanger 1006 can help to quickly adjust the heat transfer ratio if there is a risk that the heat transfer ratio drops below the desired value (for example, depending on the determined fuel characteristics, a lower limit of the heat transfer ratio significantly higher than zero can be set). Depending on the characteristics of the fuel, the temperature limits set for some fuels may be very strict. Having bypass pipes 2005, 2005a on both heat exchangers 1006, 2020 can prevent the oil from becoming too cold under certain conditions to avoid the risk of excessive oil condensation - it should be understood that this may be of more concern during cold start or ground idle than during cruising.

[0526] Figure 16 The secondary oil circuit system 2000’ shown includes a single heat exchanger 1004 and a single bypass pipe 2005’, the heat exchanger being a secondary fuel-oil heat exchanger, and the bypass pipe is arranged to allow a portion of the oil to bypass the heat exchanger 1004. In other embodiments, no bypass pipe may be provided on the secondary oil circuit system 2000’, or there may be more than one heat exchanger (for example, an air-oil heat exchanger or an oil-oil heat exchanger in addition to the secondary fuel-oil heat exchanger 1004), and more than one bypass pipe may be provided, optionally one bypass pipe for each heat exchanger. The bypass valve 2007’ is again provided to control the oil flow through the bypass pipe 2005’.

[0527] The inventors recognize that using fuels different from traditional kerosene-based jet fuels, such as sustainable aviation fuels, may result in different fuel properties, and the heat transfer parameters in operation can be adjusted to take advantage of the different fuel properties. In particular, in embodiments lacking a parallel arrangement of heat exchangers (such as those used in the methods described with respect to Figure 14 and Figure 15 ), introducing one or more oil bypass tubes with controllable valves and implementing careful control of these valves can achieve a method of providing improved oil cooling (since the fuel can absorb more heat), and can also improve the overall thermal efficiency of the engine while less heat is lost to the surrounding environment. The controllable heat exchange system 3000, in particular one or more controllable bypass valves 2007, 2007', 2007a, plays a key role in managing the heat transfer ratio in such engines 10.

[0528] In addition, the inventors recognize that while cruise conditions typically account for most of the flight time of an aircraft engine, idle operation is also important because the fuel mass flow rate during idle is much lower than during cruise, and even though the heat load of the fuel is relatively small, it may cause temperature increases - thus, the use of non-traditional fuels may have a greater impact on the optimal heat management method under idle conditions. Figure 17 and Figure 18 The methods 500, 600 of

[0529] Figure 17 address these two scenarios of aircraft operation. Figure 18 shows a method 500 for achieving these considerations under cruise conditions, and

[0530] Figure 17 shows a method 600 for achieving these considerations during idle (when the aircraft 1 is on the ground (ground idle), such as when the aircraft is starting up, stationary during boarding and taxiing (towards the runway or hangar, or between other ground-based positions), or at specific times during flight (flight idle)). Figure 16 shows the method 500 performed during cruise. The method 500 is arranged to be performed in a geared gas turbine engine 10, which includes: an oil circuit system 2000 arranged to supply oil to a gearbox 30; and a heat exchange system 3000 including an air-oil heat exchanger 2020 through which the oil in the oil circuit system flows; and a fuel-oil heat exchanger 1006 through which the oil and fuel in the oil circuit system flow such that heat is transferred between the oil and the fuel. As Figure 16As shown, the oil circuit system 2000 further includes at least one oil bypass pipe 2005, 2005a such that a certain proportion of the oil can bypass at least one heat exchanger. Bypass control valves 2007, 2007a are also provided to control the proportion of oil conveyed via the bypass pipes 2005, 2005a and thus also control the amount of oil conveyed through the respective heat exchangers 1006, 2020. Method 500 includes controlling 502 one or more bypass valves 2007, 2007a such that under cruise conditions, the heat transfer ratio:

[0531]

[0532] 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 can be provided to implement such control, which controller is optionally a stand-alone unit or part of an EEC.

[0533] Method 500 may further include receiving data 504 to allow calculation or inference of the heat transfer ratio, such as temperature data (temperature data of the oil and / or fuel at one or more points around the oil circuit system 2000, 2000' or the fuel flow path, and / or optionally fuel tank temperature or tank temperature) and fuel and / or oil flow rate data. Such data can be received 504 by the controller 58 and used 506 to adjust the control 502 of the bypass valves 2007, 2007a during cruise so as to maintain the heat transfer ratio within a desired level or desired bounds. Such checking and adjustment / correction 506 can be performed periodically or in response to a predetermined stimulus (such as a change in temperature or flow rate, or a change in engine operation or altitude). These steps 504, 506 can alternatively be considered part of the control 502 of the bypass valves 2007, 2007a. Method 500 can also be arranged to utilize other information when determining 506 what control action to take, such as temperature data (ambient temperature of the oil, fuel and / or the surroundings of the aircraft 1), flow rate data (oil and / or fuel) and / or one or more fuel characteristics.

[0534] In an embodiment where the bypass pipe 2005a is arranged to pass through the air-oil heat exchanger 2020, the controlled bypass valve can be the valve 2007a for the bypass pipe 2005a passing through the air-oil heat exchanger 2020. The step 502 of controlling the bypass valve 2007a to adjust the heat transfer ratio may include reducing the amount of oil conveyed through the air-oil heat exchanger 2020 when the heat transfer ratio is too high. In some such embodiments, the bypass pipe 2007a can be the only oil bypass pipe in the primary oil circuit system 2000. A bypass pipe may not be provided for the fuel-oil heat exchanger 1006.

[0535] In such asFigure 16 In an alternative embodiment as shown, the heat exchange system 3000 includes at least two oil bypass tubes, and optionally includes three or more oil bypass tubes, each bypass tube 2005, 2005', 2005a being arranged to allow oil to bypass one of the heat exchangers 1006, 1004, 2020. The method 500 may include modulating the amount of oil delivered via each bypass tube 2005, 2005', 2005a. In an embodiment having multiple bypass tubes within the same closed-loop oil system (e.g., as Figure 16 shown, having two bypass tubes in the primary oil circuit system 2000), the same bypass valve 2007 (which may be a three-way valve) can be used to control the flow through the two bypass tubes 2005, 2005a, or different bypass valves 2007, 2007a can be provided for each bypass tube, as Figure 16 shown.

[0536] It should be understood that controlling one or more bypass valves 2007, 2007a is not the only action that can affect the heat transfer ratio - however, the use of bypass tubes can be used to provide quick correction and plays a key role in keeping the heat transfer ratio within the desired range. Other heat exchange system components, such as the recirculation valve 6010 and / or the refrigeration cycle device 1007, can also be controlled, and their control can affect the heat transfer ratio. For example, the heat exchange system 3000 may include at least one recirculation tube 6011, which is arranged to allow the fluid to pass through the heat exchanger multiple times. In such embodiments, the method 500 may further include adjusting the amount of fluid delivered via the recirculation tube 6011 by controlling the corresponding valve 6010 in order to adjust the heat transfer ratio. Additionally or alternatively, the heat exchange system 3000 may include a refrigeration cycle device 1007, and the method 500 may further include using the refrigeration cycle device 1007 to provide a heat boost by transferring more heat from the oil to the fuel.

[0537] In embodiments having a refrigeration cycle device 1007, the bypass valves 2007, 2007a can be controlled such that the heat transfer ratio is in the range of 0 to 0.40. In embodiments without a refrigeration cycle device, in these embodiments, the heat exchange system 3000 is not arranged to provide a heat boost, and the bypass valves can be controlled such that the heat transfer ratio is generally higher, optionally in the range of 0.38 to 0.67.

[0538] In some embodiments, one or more fuel temperature sensors may be provided, and data related to the fuel temperature at the inlet of the combustor 16 may be used to fine-tune control decisions. Method 500 may include controlling 502 the bypass valves 2007, 2007a under cruise conditions such that if the fuel temperature at the inlet of the combustor 16 is at least 160 °C, the heat transfer ratio is in the range of 0 to 0.2; and / or such that if the fuel temperature at the inlet of the combustor 16 is at least 180 °C, the heat transfer ratio is in the range of 0 to 0.1.

[0539] In some embodiments, when determining how to control the bypass valves 2007, 2007a, one or more fuel properties may be considered (any of the above methods may be used to determine the fuel properties). For example, method 500 may include controlling 502 the bypass valve 2007 under cruise conditions such that if the fuel is at least 70% sustainable aviation fuel, the heat transfer ratio is in the range of 0 to 0.2; and / or such that if the fuel is at least 80% sustainable aviation fuel, the heat transfer ratio is in the range of 0 to 0.1.

[0540] Method 500 may include controlling 502 the bypass valves 2007, 2007a under cruise conditions such that the heat transfer rate from oil to air can be maintained in the range of 0 kJ to 240 kJ per kilogram of fuel, no more than 20% of the heat transferred away from the oil is transferred to the air under cruise conditions, and / or such that under cruise conditions, the heat transfer rate from oil to fuel is maintained in the range of 85 kJ to 350 kJ per kilogram of fuel, and at least 80% of the heat transferred away from the oil during cruise is transferred to the fuel.

[0541] In some embodiments, the heat exchange system 3000 further includes a branched fuel return passage 6020 and at least one valve that controls the diversion of the fuel flow, as described above. The valve may be controlled based on feedback from one or more temperature sensors and / or based on fuel properties. In embodiments having multiple fuel-oil heat exchangers 1004, 1006, a branched fuel return passage 6020 may be provided for either or both of the heat exchangers 1004, 1006. Method 500 may include adjusting the fuel flow along each branch based on the heat transfer ratio and / or the fuel temperature when leaving the heat exchangers 1004, 1006.

[0542] Figure 18 Method 600, which is shown being performed at idle, is arranged to be performed in the geared gas turbine engine 10 of method 500 as described above Figure 17 —it should be understood that the same engine 10 may be used to perform Figure 17 method 500 during cruise, and to perform Figure 18 method 600 during idle.

[0543] Method 600, which is performed at idle, includes controlling 602 bypass valves 2007, 2007a such that, under idle conditions, the heat transfer ratio:

[0544]

[0545] is in the range of 0.67 to 5.67 and optionally in the range of 0.67 to 5.50, 0.67 to 5.00, 0.67 to 4.50, 0.67 to 4.00, 0.67 to 3.50, 0.67 to 3.50 or 0.67 to 2.50. Thus, the ratio at idle may be higher than the ratio at cruise. The ground idle ratio is higher than the flight idle ratio. Optionally, method 600 may include maintaining the heat transfer ratio below 5.50, 5.0, 4.5, 4.0, 3.5 or 3.0 under ground idle conditions. Optionally, method 400 may include maintaining the heat transfer ratio at idle above 4.0, 4.5 or 5.0 under ground idle conditions.

[0546] Method 600 may include controlling the bypass valves such that, under idle conditions, the heat transfer ratio is higher than 1.0 and optionally higher than 1.5 or 2.0. A controller 58 may be provided to effect such control, which is optionally a stand-alone unit or part of an EEC.

[0547] Method 600 may further include receiving 604 data to allow calculation or inference of the heat transfer ratio, such as temperature data (temperature data of oil and / or fuel at one or more points around or along the fuel flow path in the oil circuit system 2000, 2000' and / or optionally fuel tank temperature or tank temperature), and fuel and / or oil flow data. Such data may be received 604 by controller 58 and used 606 to adjust the control 602 of bypass valves 2007, 2007a at idle so as to maintain the heat transfer ratio within a desired level or desired bounds. Such checking and adjustment / correction 606 may be performed periodically or in response to a predetermined stimulus (such as a change in temperature or flow, or a change in engine operation, such as the start of a taxi). These steps 604, 606 may alternatively be considered part of the control 602 of bypass valves 2007, 2007a. Method 600 may also be arranged to utilize other information when determining 606 what control action to take, such as temperature data (ambient temperature of oil, fuel and / or the environment around the aircraft 1), flow data (oil and / or fuel) and / or one or more fuel characteristics.

[0548] The step 602 of controlling one or more bypass valves 2007, 2007a to adjust the heat transfer ratio may include reducing the amount of oil delivered via its respective bypass pipe 2005a through the air-oil heat exchanger 2020 when the heat transfer ratio is too high. Such as Figure 16The heat exchange system 3000 of some of the illustrated embodiments includes at least two oil bypass tubes 2005, 2005a, each bypass tube being arranged to allow oil to bypass a respective one of the air-oil heat exchanger 2020 and the fuel-oil heat exchanger 1006. As for the method 500 performed during cruise, the method 600 performed during idle may include modulating the amount of oil delivered via each bypass tube 2005, 2005a. Having bypass tubes 2005, 2005a on both heat exchangers 1006, 2020 of the primary oil circuit 2000 can prevent the oil from becoming too cold under certain conditions to avoid the risk of excessive oil condensation - it should be understood that this may be of more concern during cold start or ground idle than during cruise. Thus, the oil circuit system 2000 may also include one or more oil temperature sensors, and the described method 600 may include receiving inputs from the one or more oil temperature sensors and optionally also receiving inputs from one or more temperature sensors arranged to detect the ambient temperature around the aircraft 1, and adjusting the control 602 of the bypass valve based on this temperature information - for example, during cold weather, less oil may be delivered via the air-oil heat exchanger 2020 to reduce heat loss to the environment.

[0549] In some embodiments, a bypass tube 2005' may also be provided, for example, for the oil flow in the primary oil circuit system 2000' to pass through the oil-oil heat exchanger 2030 and / or through one or more heat exchangers in the secondary oil circuit system 2000', such as through the secondary fuel-oil heat exchanger 1004.

[0550] The method 600 of some embodiments includes controlling other components in addition to the one or more bypass valves 2007, 2007a to obtain a desired heat transfer ratio. For example, at least one recirculation tube 6011 may be provided that is arranged to allow fluid to pass through the heat exchanger and / or the refrigeration cycle device 1007 multiple times, and the method may also include modulating the amount of fluid delivered via the respective additional components or, where applicable, activating or deactivating the respective components to adjust the heat transfer ratio. In embodiments having a refrigeration cycle device 1007, the bypass valves 2007, 2007a may be controlled such that the heat transfer ratio is in the range of 0.67 to 4.0. In embodiments without a refrigeration cycle device, in these embodiments where the heat exchange system 3000 is not arranged to provide heat boost, the bypass valves 2007, 2007a may be controlled such that the heat transfer ratio is generally higher, optionally in the range of 2.00 to 5.67, and optionally in the range of 3.37 to 5.67.

[0551] In some embodiments, one or more temperature sensors may be provided, and data related to the fuel temperature at the inlet of the burner 16 may be used to fine-tune control decisions. Method 600 may include controlling one or more bypass valves 2007, 2007a under idle conditions such that the heat transfer ratio is:

[0552] (i) in the range of 2.33 to 5.67 when the fuel temperature at the burner inlet is below 200 °C;

[0553] (ii) in the range of 0.67 to 5.00, and optionally in the range of 0.67 to 4.00, when the fuel temperature at the inlet of the burner (16) is above 200 °C;

[0554] (iii) in the range of 0.67 to 4.00, and optionally in the range of 0.67 to 2.67, when the fuel temperature at the inlet of the burner (16) is above 250 °C; and / or

[0555] (iv) in the range of 0.67 to 2.33, and optionally in the range of 0.67 to 1.22, when the fuel temperature at the inlet of the burner (16) is above 280 °C.

[0556] One or more fuel properties may also be used - optionally determined by the fuel property determination module 57 as described above; this information may be considered when setting the control 602 of the bypass valves 2007, 2007a. For example, method 600 may consider the SAF content of the fuel. Method 600 may include controlling 602 the bypass valves 2007, 2007a such that if the fuel is at least 70% sustainable aviation fuel, the heat transfer ratio is in the range of 0.67 to 3.67, and / or if the fuel is at least 80% sustainable aviation fuel, the heat transfer ratio is in the range of 0.67 to 2.67.

[0557] Method 600 may include controlling 602 the bypass valves 2007, 2007a under idle conditions such that the heat transfer rate from oil to air under idle conditions can be maintained in the range of 350 kJ to 1270 kJ per kilogram of fuel, no more than 20% of the heat transferred from the oil is transferred to the air at ground idle, and / or such that the heat transfer rate from oil to fuel at idle is maintained in the range of 350 kJ to 1270 kJ per kilogram of fuel, and at least 80% of the heat transferred from the oil is transferred to the fuel at idle.

[0558] In some embodiments, the heat exchange system 3000 further includes a branched fuel return passage 6020 and at least one valve for controlling the diversion of the fuel flow, as described above.

[0559] It should be understood that the present invention is not limited to the above embodiments, and various modifications and improvements can be made without departing from the concepts described herein. Unless mutually exclusive, any feature can be used alone or in combination with any other feature, and the present disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.

Claims

1. A method of operating a gas turbine engine of an aircraft, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, a combustor arranged to combust a fuel, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core; a gear box receiving input from the spindle and outputting drive to the fan; an oil circuit system arranged to supply oil to the gearbox; as well as A heat exchange system, the heat exchange system comprising: an air-oil heat exchanger through which the oil in the oil circuit system flows; and a fuel-oil heat exchanger through which the oil and the fuel in the oil circuit system flow such that heat is transferred between the oil and the fuel; and a modulating valve arranged to allow the proportion of the oil delivered via each heat exchanger to be varied, The method comprises: determining at least one fuel property of the fuel arranged to be combusted by the burner; and The modulation valve is controlled based on the at least one fuel characteristic to adjust the proportion of the oil delivered through each heat exchanger under cruise conditions.

2. The method according to claim 1, wherein: The oil circuit system is branched so that a certain proportion of the oil can flow along each branch, and the air-oil heat exchanger and the fuel-oil heat exchanger are arranged in a parallel configuration on different branches of the oil circuit system; and The modulator valve is arranged to allow variation of the proportion of the oil delivered via each branch, and wherein controlling the modulator valve adjusts the proportion of the oil delivered via each branch under cruising conditions.

3. The method according to claim 1, wherein: The at least one fuel characteristic of the fuel comprises at least one of the following: i. the percentage of sustainable aviation fuel in the fuel; ii. the heteroatom species concentration of the fuel; iii. the aromatic content of the fuel; iv. the polyaromatic content of the fuel; v. the percentage of nitrogenous substances in the fuel; vi. the presence or percentage of trace substances or trace elements in the fuel; vii. the hydrogen-to-carbon ratio of the fuel; viii. hydrocarbon distribution of the fuel; ix. Levels of non-volatile particulate matter emissions during combustion; x. the naphthalene content of the fuel; xi. the sulfur content of the fuel; xii. The cycloalkane content of the fuel; xiii. The oxygen content of the fuel; xiv. Thermal stability of the fuel; xv. the coking level of the fuel; xvi. an indication that the fuel is a fossil fuel; xvii. At least one of density, viscosity, calorific value and heat capacity.

4. The method according to claim 1, wherein: The determining of at least one fuel property of the fuel comprises: obtaining at least one fuel characteristic of any fuel already present in a fuel tank prior to refueling, the fuel tank being configured to supply fuel to the burner via the heat exchange system; determining at least one fuel characteristic of fuel added to the fuel tank during refueling; and At least one fuel property of the fuel available in the fuel tank after refueling is calculated.

5. The method according to claim 1, wherein: The determining of the at least one fuel characteristic is performed based on detecting at least one fuel property.

6. The method according to claim 1, wherein: The determining of the at least one fuel characteristic is performed based on received fuel composition data.

7. The method according to claim 1, wherein: At least one fuel characteristic is inferred from performance of the gas turbine engine during at least one of taxi, takeoff, and ascent of the aircraft.

8. The method according to claim 1, wherein: The modulation valve is arranged to ensure that at cruise no less than 70% of the oil is delivered via the fuel-oil heat exchanger.

9. The method according to claim 1, wherein: The heat exchange system also includes a refrigeration cycle device, and the method includes using the refrigeration cycle device to provide a heat boost by transferring more heat from the oil to the fuel so that the fuel temperature is raised above the oil temperature.

10. The method according to claim 1, wherein: The heat exchange system also includes a branch fuel return passage and at least one valve for controlling the diversion of the fuel flow, wherein the branch passage is arranged to return fuel from the heat exchange system to at least two different locations along a main fuel path, from which the fuel enters the gas turbine engine to the combustor.

11. The method according to claim 1, wherein: Based on the suitability of at least one determined fuel property, the method includes routing all of the oil through the fuel-oil heat exchanger during at least one continuous period of at least 30 minutes while cruising such that no heat is lost to the environment via the air-oil heat exchanger during at least some operating periods while cruising.

12. The method according to claim 1, wherein: Based on the suitability of the at least one determined fuel property, the method includes routing at least 95% of the oil via the fuel-oil heat exchanger at least 90% of the time while cruising.

13. The method according to claim 1, wherein: Based on the suitability of the at least one determined fuel property, the method includes transferring at least 80% of the heat transferred from the oil during cruising to the fuel.

14. The method according to claim 1, wherein: Based on the suitability of the at least one determined fuel property, the method includes transferring all of the heat transferred from the oil while cruising to the fuel at least 90% of the time while cruising.

15. The method according to claim 1, wherein: The at least one fuel property is or includes thermal stability, and wherein at least 80% of the heat transferred from the oil at cruise is transferred to the fuel if the fuel is stably operated at a temperature above 160°C.

16. The method according to claim 1, wherein: The at least one fuel characteristic is or includes the aromatic content of the fuel, and wherein if the mole percentage of aromatics in the fuel is less than 12%, at least 80% of the heat transferred from the oil at cruise is transferred to the fuel.

17. The method according to claim 1, wherein: The at least one fuel characteristic is or includes the percentage of sustainable aviation fuel (SAF) in the fuel, and wherein, if the fuel has a SAF content above 50%, at least 80% of the heat transferred from the oil at cruise is transferred to the fuel.

18. The method according to claim 1, wherein: The at least one fuel property is or includes a heating value of the fuel, and wherein, if the fuel has a heating value of at least 43.5 MJ / kg, at least 80% of the heat transferred from the oil at cruise is transferred to the fuel.

19. A gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core, the engine core comprising a turbine, a compressor, and a spindle connecting the turbine to the compressor; a fan located upstream of the engine core; as well as a gearbox receiving input from the spindle and outputting drive to the fan so as to drive the fan at a slower rotational speed than the spindle; an oil circuit system arranged to supply oil to the gearbox; as well as A heat exchange system, the heat exchange system comprising: an air-oil heat exchanger through which the oil in the oil circuit system flows; and a fuel-oil heat exchanger through which the oil and the fuel in the oil circuit system flow such that heat is transferred between the oil and the fuel; and a modulator valve arranged to allow variation of the proportion of the oil delivered via each heat exchanger; and a fuel composition determination module arranged to determine at least one fuel property of the fuel arranged to be combusted by the burner, Wherein the modulation valve is arranged to be controlled based on the at least one fuel property so as to adjust the proportion of the oil delivered via each heat exchanger under cruising conditions.

20. The gas turbine engine of claim 19, wherein: The turbine is a first turbine, the compressor is a first compressor, and the spindle is a first spindle; The engine core also includes a second turbine, a second compressor, and a second spindle connecting the second turbine to the second compressor; and The second turbine, the second compressor and the second spindle are arranged to rotate at a higher rotational speed than the first spindle.

Images