Gas turbine fuel temperature

Abstract

A method of operating a gas turbine engine of an aircraft, the gas turbine engine comprising: an engine core, a fan, a fan shaft, at least one bearing, at least one auxiliary system, an oil circuit system, and a heat exchange system comprising: a primary fuel-oil heat exchanger, a secondary fuel-oil heat exchanger, a fuel pump; wherein the method further comprises controlling the heat exchange system such that, under cruise conditions, the fuel temperature when exiting the second heat exchanger is in the range of 120 DEG C to 200 DEG C.

Description

[0001] Cross - reference to related applications

[0002] This specification claims the benefit of priority based on and claiming priority from UK Patent Application No. 2319128.1, 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, which include managing different fluids and heat transfer between different fluids, and particularly to managing the heat exchange system of an aircraft engine.

[0004] Description of Related Art

[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 engine itself and the methods of operating the gas turbine engine. 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 fan shaft;

[0010] At least one bearing arranged to support the fan shaft;

[0011] An oil circuit system arranged to supply oil to the at least one bearing; and

[0012] A heat exchange system including:

[0013] An air - oil heat exchanger through which oil in the oil circuit system flows; 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;

[0015] A bypass duct arranged to allow a proportion of the oil to flow through at least one of the air - oil heat exchanger and the fuel - oil heat exchanger;

[0016] A bypass valve, which is arranged to allow a proportion of the oil conveyed through the bypass pipe to be changed,

[0017] The method includes:

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

[0019] Controlling the bypass valve based on the at least one fuel property so as to adjust the proportion of the oil conveyed via each heat exchanger under cruise conditions.

[0020] The present 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 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 than conventional fuels in one or more fuel-oil heat exchangers 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 the air-oil heat exchanger). A 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 / what temperature the fuel can be raised to without forming deposits within the pipes, burner, and / or hydraulic machinery unit 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 or other fuel decomposition product deposition pathways (such as varnishing), thereby improving aircraft performance.

[0021] Using the 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 the air-oil heat exchanger) can also provide a more thermodynamically efficient engine. This can improve the cooling of the oil before it returns to the rest of the turbomachine. 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 a transition period, over the life of a gas turbine engine, the available aviation fuel changes over time, and depending 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 being used and to 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 control of the heat exchange system.

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

[0024] The air-oil heat exchanger can help remove excess heat from the oil in the oil circuit system that has not been transferred to the fuel. This can allow further cooling of the oil by transferring heat from the oil to outside the fuel in the fuel-oil heat exchanger.

[0025] The air-oil heat exchanger and the fuel-oil heat exchanger can be arranged in series along the path of the oil circuit system. A bypass pipe can be arranged such that the oil passes through the fuel-oil heat exchanger while flowing through the air-oil heat exchanger and vice versa.

[0026] The bypass 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 the heat transfer from the oil to the air or fuel) can be modified based on the temperature of the fuel leaving the fuel-oil heat exchanger or entering the burner, and thus can allow the fuel temperature to be controlled within defined upper and lower limits. Similarly, control of the oil flow can allow the oil temperature to be kept within an appropriate range. Such control can help ensure improved efficiency of the turbofan engine (e.g., by raising the fuel temperature) by using temperatures that are 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 bypass valve can be arranged to allow up to 100% of the oil to be delivered via the fuel-oil heat exchanger. The bypass valve can be arranged to ensure that at least 70%, 80%, or 90% of the oil in the oil is delivered via the fuel-oil heat exchanger during cruise.

[0027] The bypass valve can be arranged to divert a fixed portion of the oil to the bypass pipe during engine operation to bypass one of the fuel-oil heat exchanger and the air-oil heat exchanger, the fixed portion being 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 bypass valve can be arranged to divert a variable portion of the oil flow through the bypass pipe during engine operation to bypass one of the fuel-oil heat exchanger and the air-oil heat exchanger. Thus, the modulating valve can be actively controlled to change the proportion of oil delivered via each heat exchanger, especially in embodiments where the aircraft carries multiple different fuels in different tanks, and can change the fuel (or fuel mixture) used during flight. The active control of the modulating valve can be automatic and implemented by a controller of the heat exchange system.

[0028] Alternatively, the oil circuit system (which may 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. 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. Accordingly, the heat exchange system can further include a modulating valve that is arranged to allow the proportion of the oil delivered via each branch to vary, and controlling the modulating valve can thus adjust the proportion of the 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.

[0029] The heat exchange system can further include a secondary fuel-oil heat exchanger. The secondary fuel-oil heat exchanger can be a servo fuel-oil heat exchanger. A portion but not all of the fuel leaving the main fuel-oil heat exchanger can be directed to the servo fuel-oil heat exchanger. The servo fuel-oil heat exchanger can further increase the temperature of the fuel before the fuel is provided for the servo mechanisms of the gas turbine engine (such as for fuel hydraulic actuation and / or heating). These servo mechanisms can include a nacelle anti-icing system. The servo mechanisms can include engine actuators. The servo mechanisms can include a turbine case cooling (TCC) servo valve. Only the fuel passing through the secondary fuel-oil heat exchanger is used in these auxiliary systems. The fuel used in these auxiliary systems can be returned to the fuel tank for later recirculation, or can be recombined with other fuel leaving the main fuel-oil heat exchanger and enter the burner. Thus, in some embodiments, the fuel passing through the secondary fuel-oil heat exchanger may not be provided to the burner, but instead optionally returned to the aircraft fuel tank after being used for an auxiliary system (such as a fuel hydraulic actuator). The servo fuel-oil heat exchanger can be structurally similar or identical to the primary fuel-oil heat exchanger. The servo fuel-oil heat exchanger can be smaller than the primary fuel-oil heat exchanger. At least a portion of the fuel can bypass the secondary fuel-oil heat exchanger.

[0030] The heat exchange system can further include an additional bypass duct that is arranged to allow oil (or fuel) to bypass one or more heat exchangers of the heat exchange system. In some embodiments, the bypass duct can effectively form an additional branch in a parallel branch oil system, and a bypass valve or another oil valve can be arranged to regulate the amount of oil delivered through the bypass duct for oil or each bypass duct based on one or more determined fuel characteristics and optionally based on one or more temperature measurements.

[0031] 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 boost by transferring more heat from the oil to the fuel, optionally raising the fuel temperature 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.

[0032] The heat exchange system may further include a branched fuel return passage and at least one valve that controls the diversion of 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 the fuel enters a gas turbine engine to reach a combustor. For example, the fuel leaving the 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 (such as a more downstream location, for example after one or more engine components downstream of the fuel-oil heat exchanger being discussed, or actually a more upstream location (thus acting as a recirculation duct)).

[0033] 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. 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 during 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 operations 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 during cruise.

[0034] At least one fuel characteristic may be or include thermal stability. If the fuel operates stably at a temperature higher than 140 °C, then at least 80% of the heat transferred away from the oil during cruise may be transferred to the fuel.

[0035] At least one fuel characteristic may be or include the aromatic content in the fuel. If the fuel has an aromatic molar percentage lower than 12%, 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 property may be or include the percentage of sustainable aviation fuel in the fuel - % SAF. The SAF proportion (X%) may be by volume. If the fuel has an SAF content higher than 50%, at least 80% of the heat transferred away from the oil during cruise can be transferred to the fuel.

[0037] 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.

[0038] At least one fuel property of the fuel may include at least one of the following:

[0039] i. The percentage of sustainable aviation fuel in the fuel;

[0040] ii. The heteroatom species concentration of the fuel;

[0041] iii. The aromatic content of the fuel;

[0042] iv. The polyaromatic content of the fuel;

[0043] v. The percentage of nitrogenous substances in the fuel;

[0044] vi. The presence or percentage of tracer substances or trace elements in the fuel;

[0045] vii. The hydrogen - to - carbon ratio of the fuel;

[0046] viii. The hydrocarbon distribution of the fuel;

[0047] ix. The level of non - volatile particulate matter emissions during combustion;

[0048] x. The naphthalene content of the fuel;

[0049] xi. The sulfur content of the fuel;

[0050] xii. The cycloalkane content of the fuel;

[0051] xiii. The oxygen content of the fuel;

[0052] xiv. The thermal stability of the fuel;

[0053] xv. The coking level of the fuel;

[0054] xvi. An indication that the fuel is a fossil fuel;

[0055] xvii. At least one of density, viscosity, calorific value, and heat capacity.

[0056] The method may also include chemically or physically detecting one or more parameters related to the fuel in the 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 one or more of shaft speed, turbine fan power ratio (TPR), and / or engine pressure ratio (EPR) and fuel mass flow rate, from which the calorific value (a fuel property) may 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 loaded fuel sample from the fuel tank for out-wing testing. Extracting a sample of the loaded fuel from the fuel tank may include extracting a fuel sample before fuel replenishment.

[0057] 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).

[0058] 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.

[0059] 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.

[0060] According to a second aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising:

[0061] An engine core including a turbine, a compressor, and a spool connecting the turbine to the compressor;

[0062] A fan located upstream of the engine core; and

[0063] A gearbox receiving an input from the spool and outputting a driving force to the fan so as to drive the fan at a rotational speed lower than that of the spool;

[0064] An oil circuit system arranged to supply oil to the gearbox; and

[0065] A heat exchange system including:

[0066] An air - oil heat exchanger through which oil in an oil circuit system flows; and

[0067] A fuel - oil heat exchanger through which oil and fuel in an oil circuit system flow such that heat is transferred between the oil and the fuel;

[0068] A bypass pipe arranged to allow a proportion of the oil to flow through at least one of the air - oil heat exchanger and the fuel - oil heat exchanger;

[0069] A bypass valve arranged to allow the proportion of the oil conveyed through the bypass pipe to be changed; and

[0070] A fuel composition determination module arranged to determine at least one fuel characteristic of fuel arranged to be burned by a burner,

[0071] wherein the bypass valve is arranged to be controlled based on the at least one fuel characteristic so as to adjust the proportion of the oil conveyed through each heat exchanger under cruise conditions.

[0072] 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, the second compressor, and the second spool may be arranged to rotate at a higher rotational speed than the first spool.

[0073] The spool may directly output a driving force to the fan so as to drive the fan at the same rotational speed as the spool, such that the engine is a direct - drive turbofan engine.

[0074] An engine according to a 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.

[0075] According to a third aspect, there is provided a method of operating a gas turbine engine, the gas turbine engine comprising:

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

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

[0078] A fan shaft;

[0079] At least one bearing arranged to support the fan shaft;

[0080] An oil circuit system arranged to supply oil to the at least one bearing; and

[0081] A heat exchange system, the heat exchange system comprising:

[0082] An air - oil heat exchanger through which oil in an oil circuit system flows; and

[0083] A fuel - oil heat exchanger through which oil and fuel in an oil circuit system flow such that heat is transferred between the oil and the fuel; and

[0084] A bypass pipe arranged to allow a proportion of the oil to flow through at least one of the air - oil heat exchanger or the fuel - oil heat exchanger; and

[0085] A bypass valve arranged to allow the proportion of oil flowing through the bypass pipe to be changed,

[0086] The method includes controlling the heat exchange system such that, under cruise conditions, the heat transfer ratio:

[0087]

[0088] Is in the range from 0 to 0.67.

[0089] 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 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 in one or more fuel - oil heat exchangers than traditional fuels without significantly increasing coking or other fuel decomposition product deposition pathways (such as varnishing). This can enable a method of providing improved oil cooling (since the fuel can absorb more heat), and can also improve the overall thermodynamic efficiency of the engine while less heat is lost to the surrounding environment. A controllable heat exchange system plays a key role in managing the heat transfer ratio.

[0090] Although it should be understood that this ratio is dimensionless, in the examples described herein, heat transfer is measured by fuel mass, 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 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 heat transfer is recorded relative to the mass (kg) of fuel reaching the burner in order to regulate 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 when approaching or entering the burner with the fuel temperature in the aircraft's fuel tank. For comparison purposes, the amount of heat transferred from oil to air can be determined by the temperature drop of the oil through 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 oil to air (taking into account any other sources of loss or generation such as fuel recirculation).

[0091] The method may 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.

[0092] 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.

[0093] At least one bypass duct may be arranged to allow oil to bypass the air-oil heat exchanger such that the oil flowing through the bypass duct flows through the fuel-oil heat exchanger and not through the air-oil heat exchanger. Controlling the heat exchange system to adjust the heat transfer ratio may include increasing the amount of oil flowing through the bypass duct when the heat transfer ratio is too high.

[0094] The heat exchange system may include at least one recirculation duct 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 duct.

[0095] The heat exchange system may further include a secondary fuel - oil heat exchanger. The secondary fuel - oil heat exchanger may be a servo fuel - oil heat exchanger. A portion but not all of the fuel leaving the primary fuel - oil heat exchanger may be directed to the servo fuel - oil heat exchanger. The servo fuel - oil heat exchanger may further increase the temperature of the fuel before providing fuel for the servo mechanisms of the gas turbine engine (such as for fuel hydraulic actuation and / or heating). These servo mechanisms may include nacelle anti - ice systems. The servo mechanisms may include engine actuators. The servo mechanisms may include turbine case cooling (TCC) servo valves. Only the fuel passing through the secondary fuel - oil heat exchanger is used in these auxiliary systems. The fuel used in these auxiliary systems may be returned to the fuel tank for later recirculation, or may be recombined with other fuel leaving the primary fuel - oil heat exchanger and enter the burner. Thus, in some embodiments, the fuel passing through the secondary fuel - oil heat exchanger may not be provided to the burner, but instead may optionally be returned to the aircraft fuel tank after being used in an auxiliary system (such as a fuel hydraulic actuator). The servo fuel - oil heat exchanger may be structurally similar or identical to the primary fuel - oil heat exchanger. The servo fuel - oil heat exchanger may be smaller than the primary fuel - oil heat exchanger. At least a portion of the fuel may not pass through the secondary fuel - oil heat exchanger.

[0096] The heat exchange system may include a refrigeration cycle device. 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 raising the fuel temperature to be higher than the oil temperature. An example heat transfer ratio using the refrigeration cycle device may be in the range of 0 to 0.40.

[0097] In embodiments where no refrigeration cycle device is used / no heat exchange system is arranged to provide a heat lift, the heat transfer ratio may be in the range of 0.38 to 0.67.

[0098] 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.

[0099] 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.

[0100] 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 140 °C, the heat transfer ratio is in the range of 0 to 0.45.

[0101] 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.

[0102] The method may include controlling a 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.

[0103] The method may include maintaining the heat transfer rate from oil to air under cruise conditions in the range of 0 kJ to 100 kJ per kilogram of fuel, and optionally in the range of 0 kJ / kg to 35 kJ / kg, wherein no more than 20% of the heat transferred away from the oil during cruise is transferred to the air.

[0104] The method may include maintaining the heat transfer rate from oil to fuel during cruise in the range of 110 kJ to 200 kJ per kilogram of fuel, and optionally in the range of 150 kJ / kg to 200 kJ / kg, wherein at least 80% of the heat transferred away from the oil during cruise is transferred to the fuel.

[0105] The methods of the first aspect and the third aspect may be complementary and may be carried out together in various embodiments. The method of the third aspect may be carried out using the engine of the second aspect.

[0106] According to a fourth aspect, there is provided a gas turbine engine for an aircraft, the engine comprising:

[0107] An engine core including a turbine, a compressor, and a spool connecting the turbine to the compressor;

[0108] A fan located upstream of the engine core;

[0109] A fan shaft;

[0110] At least one bearing arranged to support the fan shaft;

[0111] An oil circuit system arranged to supply oil to a gearbox; and

[0112] A heat exchange system including:

[0113] An air-oil heat exchanger through which the oil in the oil circuit system flows;

[0114] 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;

[0115] A bypass duct arranged to allow a proportion of the oil to flow through at least one of the air-oil heat exchanger or the fuel-oil heat exchanger; and

[0116] A bypass valve arranged to allow variation of the proportion of the oil conveyed through the bypass duct

[0117] And wherein, the heat exchange system is arranged to be controlled such that, under cruise conditions, the heat transfer ratio:

[0118]

[0119] is in the range of 0 to 0.67.

[0120] 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.

[0121] The heat exchange system may further include a branched fuel return passage and at least one valve for controlling the diversion of fuel flow. The branched passage may be arranged to return fuel from the heat exchange system to at least two different positions along the main fuel path, from which fuel enters the gas turbine engine to reach the burner.

[0122] The bypass duct may be arranged to allow a proportion of the oil flow to pass through the air-oil heat exchanger such that the oil flowing through the bypass duct passes through the fuel-oil heat exchanger without passing through the air-oil heat exchanger.

[0123] The spool may directly output driving force to the fan so as to drive the fan at the same rotational speed as the spool, such that the engine is a direct drive turbofan engine.

[0124] The air-oil heat exchanger and the fuel-oil heat exchanger may be arranged in a series configuration on the oil circuit system.

[0125] Alternatively, the oil circuit system (which may also be referred to as a recirculating oil system) may branch such that a proportion of the oil can flow along each branch. The air-oil heat exchanger and the fuel-oil heat exchanger may be arranged in a parallel configuration on different branches of the oil circuit system.

[0126] 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 in any of the foregoing aspects.

[0127] According to a fifth aspect, there is provided a method of operating a gas turbine engine, the gas turbine engine comprising:

[0128] An engine core including a turbine, a compressor, a burner arranged to burn fuel, and a spool connecting the turbine to the compressor;

[0129] A fan located upstream of the engine core;

[0130] Fan shaft;

[0131] At least one bearing arranged to support the fan shaft;

[0132] An oil circuit system arranged to supply oil to the at least one bearing; and

[0133] A heat exchange system, the heat exchange system comprising:

[0134] An air - oil heat exchanger through which the oil in the oil circuit system flows; and

[0135] 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

[0136] A bypass pipe arranged to allow a certain proportion of the oil to flow through at least one of the air - oil heat exchanger or the fuel - oil heat exchanger; and

[0137] A bypass valve arranged to allow changing the proportion of the oil conveyed through the bypass pipe,

[0138] The method includes controlling the heat exchange system such that under idle conditions, the heat transfer ratio:

[0139]

[0140] Is in the range of 0 to 1.5.

[0141] 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 idle 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 coking or other fuel decomposition product deposition pathways (such as varnishing). This can enable a method of providing improved oil cooling (since the fuel can absorb more heat), and can also improve the overall thermodynamic efficiency of the engine while less heat is lost to the surrounding environment. The controllable heat exchange system plays a key role in managing the heat transfer ratio.

[0142] In addition, while cruise conditions typically account for the majority of an aircraft engine's operating time, the inventors have recognized that idle operation is also important because the fuel mass flow rate during idle is much lower than during cruise, and even relatively small heat loads on the fuel can cause temperature increases - thus, the use of non-conventional fuels may have a greater impact on the optimal thermal management methods during idle conditions (e.g., when the aircraft is starting up, operating stationary during boarding, taxiing (towards the runway or hangar, or between other ground positions), or during "flight idle" conditions such as during descent). Since the operating conditions between cruise and idle are very different - at least in terms of the desired thrust output of the engine - different controls of the heat exchange system are appropriate.

[0143] While it should be understood that this ratio is dimensionless, in the examples described herein, heat transfer is measured per unit mass of fuel, thus providing a heat transfer rate normalized for changes in fuel flow rate during idle. For heat transfer from oil to air, the definition of "per unit mass of fuel" can be equivalent to "per set time period depending on fuel flow rate" to similarly provide normalization of fuel flow rate. It should be understood that heat transfer using a heat exchanger must be completed before the fuel reaches the burner. Any additional temperature increase 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 regulate the fuel flow rate 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. For comparison purposes, the heat transferred from the oil to the air can be determined by the temperature drop of the oil as it passes through 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 (after accounting for any other sources of loss or generation such as fuel recirculation or fuel pumps).

[0144] The method may include controlling the heat exchange system such that, during idle conditions, the heat transfer ratio is less than 1.0, within 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.

[0145] Controlling the heat exchange system to regulate 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.

[0146] At least one bypass duct may be arranged to allow oil to bypass the air-oil heat exchanger such that the oil flowing through the bypass duct flows through the fuel-oil heat exchanger and not through the air-oil heat exchanger. Controlling the heat exchange system to adjust the heat transfer ratio may include increasing the amount of oil flowing through the bypass duct when the heat transfer ratio is too high.

[0147] The heat exchange system may include at least one recirculation duct arranged to allow a 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 conveyed via the recirculation duct.

[0148] The heat exchange system may further include a secondary fuel-oil heat exchanger. The secondary fuel-oil heat exchanger may be a servo fuel-oil heat exchanger. A portion but not all of the fuel leaving the primary fuel-oil heat exchanger may be directed to the servo fuel-oil heat exchanger. The servo fuel-oil heat exchanger may further increase the temperature of the fuel before providing fuel to servo mechanisms of a gas turbine engine (such as for fuel hydraulic actuation and / or heating). These servo mechanisms may include nacelle anti-icing systems. The servo mechanisms may include engine actuators. The servo mechanisms may include turbine case cooling (TCC) servo valves. Only the fuel passing through the secondary fuel-oil heat exchanger is used in these auxiliary systems. The fuel used in these auxiliary systems may be returned to the fuel tank for later recirculation or may be recombined with other fuel leaving the primary fuel-oil heat exchanger and enter the combustor. Thus, in some embodiments, the fuel passing through the secondary fuel-oil heat exchanger may not be provided to the combustor but instead may optionally be returned to the aircraft fuel tank after being used for an auxiliary system (such as a fuel hydraulic actuator). The servo fuel-oil heat exchanger may be structurally similar or identical to the primary fuel-oil heat exchanger. The servo fuel-oil heat exchanger may be smaller than the primary fuel-oil heat exchanger. At least a portion of the fuel may not pass through the secondary fuel-oil heat exchanger.

[0149] The heat exchange system may include a refrigeration cycle device. 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 raising the fuel temperature to be higher than the oil temperature. An example heat transfer ratio using the refrigeration cycle device may be in the range of 0 to 0.40.

[0150] In cases where the heat exchange system is not arranged to provide a heat lift, the heat exchange system may be controlled such that the heat transfer ratio is in the range of 0.38 to 1.2.

[0151] The method may include controlling the heat exchange system under idle conditions such that if the fuel temperature at the combustor inlet is below 180 °C, the heat transfer ratio is in the range of 0.3 to 1.5.

[0152] The method may include controlling a heat exchange system under idle conditions such that if the fuel temperature at the burner inlet is higher than 180 °C, the heat transfer ratio is in the range of 0 to 0.3.

[0153] Under idle conditions, the method may include controlling the heat exchange system 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.

[0154] The method may include maintaining the heat transfer rate from oil to air in the range of 0 kJ to 180 kJ per kilogram of fuel under idle conditions, and optionally in the range of 0 kJ / kg to 60 kJ / kg, wherein no more than 20% of the heat removed from the oil at idle is transferred to the air.

[0155] The method may include maintaining the heat transfer rate from oil to fuel in the range of 100 kJ to 300 kJ per kilogram of fuel under idle conditions, and optionally in the range of 200 kJ / kg to 300 kJ / kg, wherein at least 80% of the heat removed from the oil at idle is transferred to the fuel.

[0156] The air - oil heat exchanger and the fuel - oil heat exchanger may be arranged in series in the oil circuit system.

[0157] Alternatively, the oil circuit system (which may also be referred to as a recirculating oil system) may 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 may be arranged in a parallel configuration on different branches of the oil circuit system. Thus, the heat exchange system may further include a modulating valve arranged to allow variation in the proportion of oil delivered via each branch, and controlling the modulating valve can thus adjust 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 may be provided. Two or more branches of the primary oil circuit system may re - join after the heat exchanger system such that the oil re - combines after heat is transferred from the oil.

[0158] The heat exchange system may further include branched fuel return passages and at least one valve controlling fuel flow diversion, these branched passages being arranged to return fuel from the heat exchange system to at least two different positions along the main fuel path from which fuel enters the gas turbine engine to reach the burner.

[0159] The methods of the first, third, and fifth aspects may be complementary and they may be performed together in various embodiments. The method of the fifth aspect may be performed using the engine of the second or fourth aspect.

[0160] According to a sixth aspect, there is provided a gas turbine engine for an aircraft, the engine comprising:

[0161] An engine core, the engine core including a turbine, a compressor, and a spool connecting the turbine to the compressor;

[0162] A fan, the fan being located upstream of the engine core;

[0163] A fan shaft;

[0164] At least one bearing, the at least one bearing being arranged to support the fan shaft;

[0165] An oil circuit system, the oil circuit system being arranged to supply oil to a gearbox; and

[0166] A heat exchange system, the heat exchange system including:

[0167] An air-oil heat exchanger, oil in the oil circuit system flowing through the air-oil heat exchanger;

[0168] A fuel-oil heat exchanger, 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;

[0169] A bypass pipe, the bypass pipe being arranged to allow a certain proportion of the oil to flow through at least one of the air-oil heat exchanger or the fuel-oil heat exchanger; and

[0170] A bypass valve, the bypass valve being arranged to allow the proportion of the oil conveyed through the bypass pipe to be changed,

[0171] And wherein, the heat exchange system is arranged to be controlled such that under idle conditions, the heat transfer ratio:

[0172]

[0173] Is in the range of 0 to 1.5.

[0174] 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, the second compressor, and the second spool may be arranged to rotate at a higher rotational speed than the first spool.

[0175] The heat exchange system may further include a branched fuel return passage and at least one valve controlling the fuel flow diversion. The branched passage may be arranged to return fuel from the heat exchange system to at least two different positions along the main fuel path, from which the fuel enters the gas turbine engine to reach the burner.

[0176] The bypass pipe may be arranged to allow a certain proportion of the oil to flow through the air-oil heat exchanger such that the oil flowing through the bypass pipe flows through the fuel-oil heat exchanger without flowing through the air-oil heat exchanger.

[0177] The mandrel can directly output a driving force to the fan so as to drive the fan at the same rotational speed as the mandrel, such that the engine is a direct drive turbofan engine.

[0178] The air-oil heat exchanger and the fuel-oil heat exchanger can be arranged in a series configuration in the oil circuit system.

[0179] Alternatively, the oil circuit system (which may 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.

[0180] The engine of the sixth aspect can be arranged to perform the methods of the first, third and / or fifth aspects and can have any one of the features described in any of the foregoing aspects.

[0181] According to a seventh aspect, there is provided a method of operating a gas turbine engine of an aircraft, the gas turbine engine comprising:

[0182] An engine core including a turbine, a compressor, a burner arranged to burn fuel, and a mandrel connecting the turbine to the compressor;

[0183] A fan located upstream of the engine core;

[0184] A fan shaft;

[0185] At least one bearing arranged to support the fan shaft;

[0186] At least one auxiliary system arranged to use some of the fuel;

[0187] An oil circuit system arranged to supply oil to at least one bearing; and

[0188] A heat exchange system including:

[0189] A primary fuel-oil heat exchanger through which oil in the oil circuit system and at least substantially all of the fuel flow such that heat is transferred between the oil and the fuel; and

[0190] A secondary fuel-oil heat exchanger through which oil in the oil circuit system and a portion of the fuel flow such that heat is transferred between the oil and the fuel, the secondary fuel-oil heat exchanger being arranged to supply fuel to at least one auxiliary system; and

[0191] A fuel pump, which is arranged to pump fuel, wherein the fuel pump is located downstream of a primary fuel-oil heat exchanger and upstream of a secondary fuel-oil heat exchanger along a fuel flow path;

[0192] The method further includes controlling the heat exchange system such that, under cruise conditions, the fuel temperature leaving the secondary heat exchanger is in the range of 120°C to 200°C.

[0193] The present 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 idle conditions can be adjusted to take advantage of the different fuel properties. In particular, some fuels can be heated to higher temperatures than traditional fuels in one or more fuel-oil heat exchangers without significantly increasing coking or other fuel decomposition product deposition pathways (such as varnishing). 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 less heat is lost to the surrounding environment. A controllable heat exchange system plays a key role in managing the heat transfer ratio.

[0194] The secondary fuel-oil heat exchanger can be a servo fuel-oil heat exchanger. A portion but not all of the fuel leaving the primary fuel-oil heat exchanger can be directed to the servo fuel-oil heat exchanger. The servo fuel-oil heat exchanger can further increase the temperature of the fuel before providing fuel to the servo mechanisms of a gas turbine engine (such as for fuel hydraulic actuation and / or heating). These servo mechanisms can include nacelle anti-icing systems. The servo mechanisms can include engine actuators. The servo mechanisms can include a turbine case cooling (TCC) servo valve. Only the fuel passing through the secondary fuel-oil heat exchanger is used in these auxiliary systems. The fuel used in these auxiliary systems can be returned to the fuel tank for later recirculation, or can be recombined with other fuel leaving the primary fuel-oil heat exchanger and enter the burner. Thus, in some embodiments, the fuel passing through the secondary fuel-oil heat exchanger may not be provided to the burner, but instead optionally returned to the aircraft fuel tank after being used in an auxiliary system (such as a fuel hydraulic actuator). The servo fuel-oil heat exchanger can be structurally similar or identical to the primary fuel-oil heat exchanger. The servo fuel-oil heat exchanger can be smaller than the primary fuel-oil heat exchanger. At least a portion of the fuel can not pass through the secondary fuel-oil heat exchanger.

[0195] The method can include controlling the heat exchange system such that, under cruise conditions, the fuel temperature leaving the secondary heat exchanger is in the range of 120°C to 180°C.

[0196] The method can include controlling the heat exchange system such that, under cruise conditions, the fuel temperature leaving the secondary heat exchanger is in the range of 135°C to 200°C.

[0197] The method may include controlling a heat exchange system such that, under cruise conditions, the fuel temperature leaving the secondary heat exchanger is in the range of 135°C to 180°C.

[0198] The method may include controlling a heat exchange system such that, under cruise conditions, the fuel temperature leaving the secondary heat exchanger is in the range of 150°C to 200°C, 150°C to 180°C, 150°C to 170°C. The fuel temperature leaving the secondary heat exchanger may be about 120°C, 130°C, 140°C, 150°C, 160°C or up to 200°C.

[0199] The heat exchange system may further include a fuel bypass line arranged to allow a portion of the fuel to bypass at least one of the primary fuel - oil heat exchanger and the secondary fuel - oil heat exchanger; and a fuel bypass valve arranged to allow changing the proportion of the fuel conveyed via the fuel bypass line. Controlling the heat exchange system may include controlling the fuel bypass valve to adjust the proportion of the fuel conveyed through each of the primary fuel - oil heat exchanger and the secondary fuel - oil heat exchanger under cruise conditions.

[0200] The heat exchange system may further include an oil bypass line arranged to allow a portion of the oil to bypass at least one of the primary fuel - oil heat exchanger and the secondary fuel - oil heat exchanger; and an oil bypass valve arranged to allow changing the proportion of the oil conveyed via the oil bypass line. Controlling the heat exchange system may include controlling the oil bypass valve to adjust the proportion of the oil conveyed through each of the primary fuel - oil heat exchanger and the secondary fuel - oil heat exchanger under cruise conditions.

[0201] Under cruise conditions, the ratio of the following formula

[0202]

[0203] may be at least 0.3.

[0204] For simplicity, this ratio may be referred to as the secondary heat exchanger fuel flow ratio.

[0205] The secondary heat exchanger fuel flow ratio may be at least 0.35, 0.4, 0.45, 0.5, 0.55. The secondary heat exchanger fuel flow ratio may be about 0.6.

[0206] 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 boost by transferring more heat from the oil to the fuel, optionally such that the fuel temperature is raised above the oil temperature. An oil valve may control the amount of oil flowing through the refrigeration cycle device.

[0207] The heat exchange system may also include a branched fuel return passage and at least one valve controlling the diversion of the fuel flow, the branched passage being arranged to return fuel from the heat exchange system to at least two different locations along the main fuel flow path. For example, the fuel leaving the 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., a more downstream location, such as after one or more engine components downstream of the fuel - oil heat exchanger in question, or indeed a more upstream location (thus acting as a recirculation duct)).

[0208] The primary fuel - oil heat exchanger and the secondary fuel - oil heat exchanger may be arranged in series along the oil circuit system.

[0209] The heat exchange system may also include at least one air - oil heat exchanger through which the oil of the oil circuit system flows.

[0210] According to an eighth aspect of the present invention, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising:

[0211] An engine core including a turbine, a compressor, and a spool connecting the turbine to the compressor;

[0212] A fan located upstream of the engine core; and

[0213] A gearbox receiving an input from the spool and outputting a driving force to the fan so as to drive the fan at a rotational speed lower than that of the spool;

[0214] An oil circuit system arranged to supply oil to the gearbox; and

[0215] A heat exchange system including:

[0216] A primary fuel - oil heat exchanger through which the oil in the oil circuit system and at least substantially all of the fuel flow such that heat is transferred between the oil and the fuel; and

[0217] A secondary fuel - oil heat exchanger through which the oil in the oil circuit system and a portion of the fuel flow such that heat is transferred between the oil and the fuel, the secondary fuel - oil heat exchanger being arranged to supply fuel to at least one auxiliary system; and

[0218] A fuel pump, which is arranged to pump fuel, wherein the fuel pump is located downstream of a primary fuel-oil heat exchanger and upstream of a secondary fuel-oil heat exchanger along a fuel flow path;

[0219] Wherein, the heat exchange system is arranged to be controlled such that, under cruise conditions, the fuel temperature when leaving the second heat exchanger is in the range of 120°C to 200°C.

[0220] The heat exchange system can be arranged to be controlled such that, under cruise conditions, the fuel temperature when leaving the second heat exchanger is in the range of 120°C to 180°C.

[0221] 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 can further include a second turbine, a second compressor, and a second spool connecting the second turbine to the second compressor; and the second turbine, the second compressor, and the second spool can be arranged to rotate at a higher rotational speed than the first spool.

[0222] The spool can directly output driving force to the fan so as to drive the fan at the same rotational speed as the spool, such that the engine is a direct drive turbofan engine.

[0223] The engine of the eighth aspect can be arranged to perform the methods of the first aspect, the third aspect, the fifth aspect, and / or the seventh aspect, and can have any one of the features described with respect to any of the foregoing aspects.

[0224] According to a ninth aspect, there is provided a method of operating a gas turbine engine of an aircraft, the gas turbine engine comprising:

[0225] An engine core, the engine core including a turbine, a compressor, a burner arranged to burn fuel, and a spool connecting the turbine to the compressor;

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

[0227] A fan shaft;

[0228] At least one bearing, the at least one bearing being arranged to support the fan shaft;

[0229] An oil circuit system, the oil circuit system being arranged to supply oil to the at least one bearing;

[0230] A heat exchange system, the heat exchange system including:

[0231] An air-oil heat exchanger, through which the oil in the oil circuit system flows;

[0232] A fuel - oil heat exchanger through which oil and fuel in an oil circuit system flow such that heat is transferred between the oil and the fuel; and

[0233] At least one valve arranged to allow changing at least one of the oil flow rate and the air flow rate through at least one of the fuel - oil heat exchanger and the air - oil heat exchanger; and

[0234] A temperature sensor arranged to provide an indication of the fuel temperature downstream of the fuel - oil heat exchanger (optionally at the burner inlet),

[0235] The method includes:

[0236] Based on the output from the temperature sensor, determining whether the fuel temperature has increased above a set threshold under cruise conditions; and

[0237] In response to determining that the fuel temperature has increased above the set threshold under cruise conditions, controlling the at least one valve to change at least one of the flow rates through at least one of the heat exchangers.

[0238] For example, the at least one valve may be or include a valve arranged to allow changing the proportion of oil delivered via the fuel - oil heat exchanger, and the method may include, in response to determining that the fuel temperature has increased above the set threshold under cruise conditions, controlling the at least one valve to deliver less oil through the fuel - oil heat exchanger.

[0239] The inventors recognize that it is important for the oil temperature to be maintained within a desired range and for the fuel temperature not to exceed a limit. Thus, in some embodiments, the focus of flow regulation may be on the air - oil heat exchanger. Accordingly, the at least one valve may be or include a valve arranged to allow changing the proportion of oil delivered via the air - oil heat exchanger, and the method may include, in response to determining that the fuel temperature has increased above the set threshold under cruise conditions, controlling the at least one valve to deliver more oil through the air - oil heat exchanger.

[0240] In some cases, the lack of an oil bypass on the air - oil heat exchanger can cause the oil to become too cold (e.g., condensate). Thus, bypass pipes may be provided for both heat exchangers. In some embodiments, at least a portion of the oil may bypass both heat exchangers.

[0241] The at least one valve may be or include a valve arranged to control the air flow rate through the air - oil heat exchanger. The method may include delivering more air through the air - oil heat exchanger in response to determining that the fuel temperature has increased above the set threshold under cruise conditions.

[0242] In some embodiments, the method may include controlling a plurality of valves, for example:

[0243] ●Control the oil flow through each of the fuel - oil heat exchanger and the air - oil heat exchanger separately; and / or

[0244] ●Control the air flow and the oil flow.

[0245] As described in the foregoing aspect, 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 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 the deposition of fuel decomposition products, such as by coking or varnishing, thereby allowing the fuel to absorb more heat and thus reducing the need for air cooling of the oil and improving engine efficiency. However, allowing inappropriate fuels to reach these higher temperatures can be harmful to engine performance and can even cause fuel nozzle blockage in some cases. Therefore, checks and balances are needed to ensure that engine performance is optimal for a given fuel. The method of this seventh aspect includes detecting the fuel temperature to check for any excessive fuel temperature and taking measures to reduce the heating of the fuel when appropriate. Under cruise conditions, the fuel temperature downstream of the fuel - oil heat exchanger (e.g., at the inlet of the burner) can be defined as the average value over at least 1 minute, 2 minutes, 3 minutes, 4 minutes, or 5 minutes, and optionally over 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 fuel temperature during operation, typically an increase in temperature. Thus, an instantaneous peak reaching a higher temperature (e.g., a peak lasting only a few seconds) may not be sufficient to trigger a change in the oil flow.

[0246] During a transition period, 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 (among other variables). It is important to determine the heating level of the fuel based on the specific fuel being used. The less oil is passed through the fuel - oil heat exchanger, the less heat is transferred to the fuel, and thus the lower the fuel temperature provided as it approaches the burner. Therefore, gas turbine operation can be adjusted to make the most of various fuels. A controllable oil flow valve arranged to regulate the oil flow through the fuel - oil heat exchanger plays a key role in the adjustment of engine performance.

[0247] The inventors have realized that these principles can be applied to engines with branched oil circuit paths that have different heat exchangers on different branches (arranged in parallel), and these principles can be applied to engines with heat exchangers in a substantially linear series arrangement, and where one or more bypass pipes can be used as an alternative to the branched main path.

[0248] The oil circuit system can be branched 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 can be arranged in a parallel configuration on different branches of the oil circuit system. In such embodiments, at least one valve arranged to allow changing the proportion of the oil conveyed via the fuel-oil heat exchanger can be or include a modulating valve arranged to allow changing the proportion of the oil conveyed via each branch.

[0249] The oil circuit system can include at least one bypass conduit arranged to allow a certain proportion of the oil to bypass at least one of the fuel-oil heat exchanger and the air-oil heat exchanger. In such embodiments, the at least one valve can be or include at least one bypass valve arranged to allow a certain proportion of the oil to bypass the corresponding heat exchanger / control the proportion of the oil conveyed via the bypass conduit. In such embodiments, the air-oil heat exchanger and the fuel-oil heat exchanger can be arranged in series or in parallel in the oil circuit system. In embodiments where the air-oil heat exchanger and the fuel-oil heat exchanger are arranged in parallel on different branches of the oil circuit system and there is additionally at least one bypass conduit, the method can include controlling both the bypass valve and the modulating valve, the modulating valve being arranged to allow changing the proportion of the oil conveyed via each branch.

[0250] In embodiments where the method is arranged to regulate the flow of oil through the fuel-oil heat exchanger, at least some of the oil transferred from the fuel-oil heat exchanger can be conveyed to the air-oil heat exchanger. In some such embodiments, when more oil is conveyed via the air-oil heat exchanger, the air flow through the air-oil heat exchanger can be increased.

[0251] In embodiments where the method is arranged to regulate the flow of oil through the fuel-oil heat exchanger or the air-oil heat exchanger, at least some of the oil transferred from the respective heat exchanger can be conveyed via a bypass conduit around the respective heat exchanger.

[0252] The method can also include determining a set threshold based on at least one fuel property of the fuel. In such embodiments, the at least one fuel property of the fuel can be or include at least one of the following: the thermal stability of the fuel, the nitrogen content of the fuel, the sulfur content of the fuel, and the sustainable aviation fuel (SAF) content of the fuel.

[0253] The step of determining the set threshold can include optionally linearly increasing the set threshold as the fuel thermal stability increases.

[0254] For fuels with a SAF content higher than 70%, the step of determining the set threshold can include optionally linearly increasing the set threshold as the SAF content of the fuel increases.

[0255] The method may further include determining at least one fuel property of the fuel, optionally by any one of the following methods.

[0256] The heat exchange system may include an air valve that is arranged to control the air flow through the air-oil heat exchanger. The method may further include, in response to determining that the fuel temperature has increased above a set threshold under cruise conditions, controlling the air valve to deliver more air through the air-oil heat exchanger.

[0257] The heat exchange system may include a refrigeration cycle device that is arranged to provide a heat boost by transferring additional heat from the oil to the fuel in addition to the heat transferred by the fuel-oil heat exchanger. The method may further include controlling the refrigeration cycle device to reduce the additional heat transferred to the fuel in response to determining that the fuel temperature has increased above a set threshold under cruise conditions. For example, the refrigeration cycle device may be turned off / deactivated.

[0258] The heat exchange system may include at least one bypass duct, and at least one valve may be or include a bypass valve that is arranged to control the flow through the bypass duct. The heat exchange system may include a plurality of bypass ducts, each bypass duct being arranged to allow oil to bypass a heat exchanger (e.g., the fuel-oil heat exchanger or air-oil heat exchanger described above, or a secondary fuel-oil heat exchanger, an oil-oil heat exchanger arranged to transfer heat between two separate oil circuits of the heat exchange system, or any other suitable heat exchanger). The method may include controlling at least two bypass valves - for example, when the bypass valve of the fuel-oil heat exchanger is adjusted to deliver less oil to the fuel-oil heat exchanger, the bypass valve of the air-oil heat exchanger may be adjusted to deliver more oil to the air-oil heat exchanger. In some embodiments, the same valve (e.g., a three-way valve) may regulate the oil flow to two heat exchangers.

[0259] The set threshold may be in the range of 140 °C to 300 °C, and optionally in the range of 250 °C to 300 °C. The set threshold may be 140 °C, 180 °C, 200 °C, 210 °C, 220 °C, 230 °C, 240 °C, 250 °C, 260 °C, 270 °C, 275 °C, 280 °C, 290 °C or 300 °C. The determining step may be performed periodically during the cruise operation of the aircraft.

[0260] 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.

[0261] According to an eighth aspect, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising:

[0262] an engine core including a turbine, a compressor and a spool connecting the turbine to the compressor;

[0263] a fan located upstream of the engine core;

[0264] a fan shaft;

[0265] at least one bearing arranged to support the fan shaft;

[0266] an oil circuit system arranged to supply oil to the at least one bearing;

[0267] a heat exchange system including:

[0268] an air-oil heat exchanger through which the oil in the oil circuit system flows;

[0269] 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

[0270] at least one valve arranged to allow changing at least one of the oil flow rate and the air flow rate through at least one of the fuel-oil heat exchanger and the air-oil heat exchanger;

[0271] a temperature sensor arranged to provide an indication of the fuel temperature downstream of the fuel-oil heat exchanger (optionally at the inlet of the combustor); and

[0272] a controller arranged to receive the output from the temperature sensor under cruise conditions, determine based on the output whether the fuel temperature has risen above a set threshold, and in response to determining that the fuel temperature has risen above the set threshold under cruise conditions, control the at least one valve to change at least one of the flow rates through at least one of the heat exchangers.

[0273] For example, the at least one valve may be arranged to allow changing the proportion of oil delivered via the fuel-oil heat exchanger; and the controller may be arranged to:

[0274] receive the output from the temperature sensor under cruise conditions;

[0275] determine based on the output whether the fuel temperature has increased above a set threshold; and

[0276] in response to determining that the fuel temperature has increased above the set threshold under cruise conditions, control the at least one valve so as to deliver less oil through the fuel-oil heat exchanger.

[0277] A controller or another processing module may be arranged to determine a set threshold based on one or more fuel properties of the fuel.

[0278] The heat exchange system may include a refrigeration cycle device arranged to provide a heat boost by transferring more heat from the oil to the fuel than the heat transferred by the fuel - oil heat exchanger. Optionally, the refrigeration device may allow the fuel temperature to increase to be higher than the oil temperature. The controller may be arranged to deactivate the refrigeration cycle device in response to determining that the fuel temperature has increased to be higher than the set threshold under cruise conditions.

[0279] 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.

[0280] The heat exchange system may further include a branched fuel return passage and at least one valve 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.

[0281] The spool may directly output a driving force to the fan so as to drive the fan at the same rotational speed as the spool, such that the engine is a direct - drive turbine engine.

[0282] The air - oil heat exchanger and the fuel - oil heat exchanger may be arranged in series in the oil circuit system.

[0283] 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.

[0284] As described elsewhere herein, the present disclosure may be applied to any relevant configuration of a gas turbine engine. Such gas turbine engines may be, for example, turbofan gas turbine engines, open rotor gas turbine engines (where the propellers are not enclosed by nacelles), turboprop engines, or turbojet engines. Any such engine may or may not be provided with an afterburner. Such gas turbine engines may be configured, for example, for land - or sea - based power generation applications.

[0285] A gas turbine engine according to any aspect of the present disclosure may include an engine core including a turbine, a combustor, a compressor, and a spool 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, a gas turbine engine may include a fan located downstream of the engine core, such as in the case where the gas turbine engine is an open rotor or turboprop engine (in which case the fan may be referred to as a propulsor).

[0286] In the case where the gas turbine engine is an open rotor or 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 counterclockwise 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 propulsor via suitable interconnecting shafts respectively. Thus, the propulsor may provide most of the propulsive thrust.

[0287] In the case where the gas turbine engine is an open rotor or turboprop engine, one or more propulsor stages may be driven by a gearbox. The gearbox may be of the type described herein.

[0288] An engine according to the present disclosure may be a turbofan engine. Such an engine may be a direct drive turbofan engine in which the fan is directly connected to a fan drive turbine via a spool, 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 spool may be a first spool, and the gas turbine engine may further include a second turbine and a second spool connecting the second turbine to the compressor. The second turbine, compressor, and second spool may be arranged to rotate at a higher rotational speed than the first spool. In such an arrangement, the second turbine may be axially located upstream of the first turbine.

[0289] 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 the spool and outputs a driving force to the fan in order to drive the fan at a lower rotational speed than the spool. The input to the gearbox may come directly from the spool or indirectly from the spool, such as via spur shafts and / or gears. The spool 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).

[0290] A 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 turbines and compressors, such as one shaft, two shafts or three shafts. By way of example only, the turbine connected to the spool may be a first turbine, the compressor connected to the spool may be a first compressor, and the spool may be a first spool. The engine core may also 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.

[0291] In such an arrangement, the second compressor may be axially located downstream of the first compressor. The second compressor may be arranged to receive flow (e.g. directly, e.g. via a substantially annular duct) from the first compressor.

[0292] The gearbox may be arranged to be driven by a spool (such as the first spool 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 spool configured (e.g. in use) to rotate at the lowest rotational speed (e.g., in the above example, only by the first spool and not the second spool). 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.

[0293] The gearbox may be a reduction gearbox (since the output to the fan has a lower rotational rate than the input from the spool). Any type of gearbox may be used. For example, the gearbox may be a "planetary" or "stellar" gearbox, as described in more detail elsewhere herein. Such a gearbox may be single stage. Alternatively, such a gearbox may be a compound gearbox, such as a compound planetary gearbox (which may have an input on the sun gear and an output on the ring gear and is thus referred to as a "compound star" gearbox), such as having two stages of reduction.

[0294] 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), such as greater than 2.5, such as in the range of 3 to 4.2, or 3.2 to 3.8, such as, 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 another 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 yet another 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.

[0295] In any gas turbine engine as described and / or claimed herein, fuel of a given composition or blend is provided to a combustor that can be disposed downstream (e.g., axially downstream) of a fan and a 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 provided to the inlet of the second turbine. The combustor can be disposed upstream of one or more turbines.

[0296] The compressor or each compressor (e.g., the first compressor and the second compressor as described above) can include any number of stages, such as a plurality of stages. Each stage can include a row of rotor blades and a row of stator vanes, and the row of stator vanes can be variable stator vanes (since the angle of incidence of the row of stator vanes can be variable). The row of rotor blades and the row of stator vanes can be axially offset from each other. For example, the gas turbine engine can be a direct drive turbofan gas turbine engine including 13 or 14 compressor stages (in addition to the fan). Such an engine can 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 another example, the gas turbine engine can 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 can 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 yet another example, the gas turbine engine can 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.

[0297] 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). The 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, the gas turbine engine may be a “geared” gas turbine engine including a first (or “low pressure”) turbine having 3 or 4 stages.

[0298] Each fan blade may 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 a 100% span position. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be less than (or 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 may be within the inclusive range defined by any two of the values in the previous sentence (i.e., these values may form an upper or lower limit), e.g., in the range from 0.28 to 0.32 or from 0.29 to 0.30. These ratios may generally be referred to as hub-tip ratios. The radius at the hub and the radius at the tip may both be measured at the leading (or axially foremost) part 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 any platform.

[0299] 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), such as 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.

[0300] 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 examples 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 additional non-limiting examples 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 additional non-limiting examples 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 additional non-limiting examples 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 additional non-limiting examples 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 additional non-limiting examples 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.

[0301] When using a gas turbine engine, the fan (with associated fan blades) rotates about an axis of rotation. 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).

[0302] 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 the inclusive range defined by any two of the values in the previous sentence (i.e., these values may form an upper or lower limit), for example, 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 the core engine. The radially outer surface of the bypass duct may be defined by a nacelle and / or a fan casing.

[0303] 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 exit of the highest pressure compressor (before entering the combustor) to the stagnation pressure upstream of the fan. By way of non-limiting example, the overall pressure ratio of a gas turbine engine as described and / or claimed herein at cruise conditions may be greater than (or approximately) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio may be within the inclusive range defined by any two of the values in the previous sentence (i.e., these values may form an upper or lower limit), for example, 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.

[0304] The specific thrust of an engine can be defined as the net thrust of the engine divided by the total mass flow rate 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 traditional 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.

[0305] 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 of at least (or approximately) any of the following: 100 kN, 110 kN, 120 kN, 130 kN, 135 kN, 140 kN, 145 kN, 150 kN, 155 kN, 160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN, 450 kN, 500 kN, or 550 kN. The maximum thrust may be within an inclusive range 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.

[0306] 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 itself may be referred to as the nozzle guide vane. In some examples, for a given thrust condition, the TET can depend on the particular composition of the fuel supplied 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.

[0307] 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.

[0308] The fan blades and / or the airfoil portions of the fan blades described and / or claimed herein can be made of any suitable material or combination of materials. For example, at least a portion of the fan blades and / or airfoils can be made at least in part of a composite material, such as a metal matrix composite and / or an organic matrix composite, such as a carbon fiber composite. By way of additional example, at least a portion of the fan blades and / or airfoils can be made at least in part of 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 made of different materials. For example, the fan blades can have a protective leading edge that can be made of 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 made, for example, of 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.

[0309] 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 blade may be machined from a block, and / or at least a portion of the fan blade may be attached to the hub / disc portion by welding (such as linear friction welding).

[0310] 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.

[0311] 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.

[0312] As used herein, the terms idle, taxi, takeoff, climb, cruise, descent, approach, and landing (or one or more parts 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 parts of the operating phase of the engine of the aircraft to which the gas turbine engine is designed to be attached within a given mission.

[0313] At this point, ground idle can refer to an engine operating phase where the aircraft is stationary and in contact with the ground, but where there are requirements for the engine to be operated. During idle, the engine can produce available thrust between 3% and 9% of the engine. In another non-limiting example, the engine can produce available thrust between 5% and 8% of the engine. In another non-limiting example, the engine can produce available thrust between 6% and 7% of the engine. Taxiing can refer to an engine operating phase where the aircraft is propelled along the ground by the thrust produced by the engine. During taxiing, the engine can produce available thrust between 5% and 15% of the engine. In another non-limiting example, the engine can produce available thrust between 6% and 12% of the engine. In another non-limiting example, the engine can produce available thrust between 7% and 10% of the engine. Takeoff can refer to an engine operating phase where the aircraft is propelled by the thrust produced by the engine. During the initial phase of the takeoff phase, the aircraft can be propelled while in contact with the ground. During a later phase of the takeoff phase, the aircraft can be propelled while not in contact with the ground. During takeoff, the engine can produce available thrust between 90% and 100% of the engine. In another non-limiting example, the engine can produce available thrust between 95% and 100% of the engine. In another non-limiting example, the engine can produce 100% of the available thrust.

[0314] Climb can refer to an engine operating phase where the aircraft is propelled by the thrust produced by the engine. During climb, the engine can produce available thrust between 75% and 100% of the engine. In another non-limiting example, the engine can produce available thrust between 80% and 95% of the engine. In another non-limiting example, the engine can produce available thrust between 85% and 90% of the engine. At this point, climb can refer to the operating phase between takeoff and reaching cruise conditions within 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 can refer to one or more nominal periods at a nominal point or during the aircraft flight cycle between takeoff and landing, where a relative increase in altitude is required, which can require additional thrust requirements of the engine.

[0315] As used herein, the cruise conditions that may define the cruise phase (or a part thereof) of an aircraft flight have their conventional 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, taking into account the number of engines provided for the aircraft, the cruise conditions can define the operating point, phase, or a part thereof of the flight that provides the thrust to ensure the steady-state operation of the aircraft (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). For example, if the engine is designed to be attached to an aircraft with two engines of the same type, then under 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 intermediate cruise.

[0316] 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 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 under cruise conditions can be clearly defined.

[0317] By way of example only, the forward speed under 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.

[0318] 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 about 11000m. The cruise conditions may correspond to the standard atmospheric conditions at any given altitude within these ranges.

[0319] 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.

[0320] By way of 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.

[0321] In use, the gas turbine engine described and / or claimed herein may operate under cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions of an aircraft on which at least one (e.g., 2 or 4) gas turbine engines may be mounted to provide propulsion thrust.

[0322] In addition, those skilled in the art will immediately recognize that either or both of landing and approach refer to 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 demand from the engine.

[0323] According to one aspect, there is provided an aircraft comprising 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 flight of the aircraft, as defined elsewhere herein.

[0324] According to one aspect, there is provided a method of operating a gas turbine engine as described herein and / or claimed. The operation may be carried out under any cruise conditions (e.g., in terms of thrust, atmospheric conditions and Mach number) as may be defined elsewhere herein.

[0325] According to one aspect, there is provided a method of operating an aircraft comprising a gas turbine engine as described herein and / or claimed. The operation according to this aspect may comprise (or may be) the operation under any suitable conditions (e.g., at the intermediate cruise of the aircraft), as defined elsewhere herein.

[0326] 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.

[0327] 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, for example, A / B, B / A, B*A, or any such other application, combination or function of parameter A with respect to parameter B as required. BRIEF DESCRIPTION OF THE DRAWINGS

[0328] Embodiments will now be described, by way of example only, with reference to the accompanying drawings, in which:

[0329] Figure 1 is a cross-sectional side view of a gas turbine engine;

[0330] Figure 2 is a close-up cross-sectional side view of an upstream portion of a geared gas turbine engine;

[0331] Figure 3 is a partial cross-sectional view of a gearbox for a gas turbine engine;

[0332] Figure 4 is a close-up cross-sectional side view of the upstream portion of a directly-driven gas turbine engine;

[0333] Figure 5 is a schematic view of an aircraft having a propulsion system including two gas turbine engines;

[0334] Figure 6 is a schematic view of an exemplary fuel system;

[0335] Figure 7A is a schematic view of an alternative exemplary fuel system including a recirculation loop;

[0336] Figure 7B is a schematic view of another alternative exemplary fuel system including a recirculation loop and fuel for an auxiliary system;

[0337] Figure 7C is similar to Figure 7B a schematic view of another alternative exemplary fuel system as shown, but wherein the fuel passing through the secondary heat exchanger is optionally returned to the burner after being used in one or more auxiliary systems;

[0338] Figure 8A is a schematic view of a part of an exemplary recirculation oil system;

[0339] Figure 8B is a schematic view of a part of an exemplary recirculation oil system;

[0340] Figure 8C is a schematic view of a part of an exemplary recirculation oil system;

[0341] Figure 9 is a schematic view of a part of an exemplary recirculation oil system;

[0342] Figure 10 is Figure 6 a schematic view of a part of an exemplary fuel system of Figure 8A and an exemplary recirculation oil system of

[0343] Figure 11 is Figure 7A a schematic view of a part of an exemplary fuel system of Figure 8A and an exemplary recirculation oil system of

[0344] Figure 12 is Figure 7A a schematic view of a part of an exemplary fuel system of Figure 8C and an exemplary recirculation oil system of

[0345] Figure 13 Illustrates an exemplary method of operating a gas turbine engine;

[0346] Figure 14 Illustrates another exemplary method of operating a gas turbine engine;

[0347] Figure 15 Illustrates another exemplary method of operating a gas turbine engine;

[0348] Figure 16 Illustrates another exemplary method of operating a gas turbine engine;

[0349] Figure 17 Illustrates another exemplary method of operating a gas turbine engine;

[0350] Figure 18 Illustrates a graph of the heat transfer range of a direct drive gas turbine engine during cruise. Detailed description

[0351] Figure 1 Illustrates a gas turbine engine 10 having a main rotational axis 9. The engine 10 includes an air inlet 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 an epicyclic gearbox 30.

[0352] In use, the core airflow A is accelerated and compressed by the low pressure compressor 14 and 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 are then expanded 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.

[0353] Figure 2 An exemplary arrangement of a geared fan gas turbine engine 10 is shown. The low pressure turbine 19 (see Figure 1 ) drives a shaft 26 which is coupled to a sun or sun gear 28 of an epicyclic gear arrangement 30. Radially outward of and meshing with the sun gear 28 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 sun gear 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 and meshing with the planet gears 32 is a ring or annulus gear 38 which is coupled via a link 40 to a fixed support structure 24.

[0354] 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., not including the fan 23), and / or turbine and compressor stages connected together by an interconnecting shaft 26 having the lowest rotational speed in the engine (i.e., not including 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.

[0355] In Figure 3 the epicyclic gearbox 30 is shown in more detail by way of example. Each of the sun gear 28, the planet gears 32 and the annulus gear 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 an epicyclic gearbox 30 typically include at least three planet gears 32.

[0356] In Figure 2 and Figure 3The epicyclic gearbox 30 shown by way of example is planetary, where the planet carrier 34 is connected to the output shaft via a link 36 and the annulus gear 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 gear (or ring gear) 38 to rotate. In such an arrangement, the fan 23 is driven by the annulus gear 38. By way of another alternative example, the gearbox 30 may be a differential gearbox, where both the annulus gear 38 and the planet carrier 34 are allowed to rotate.

[0357] It should be understood that Figure 2 and Figure 3 the arrangements shown are 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 within the engine 10 and / or for connecting the gearbox 30 to the engine 10. By way of another example, the connecting members (such as Figure 2 the links 36, 40 in the example) between the gearbox 30 and other components of the engine 10 (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 (for example, 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 arrangements 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 arrangements of the output link and the support link and the bearing positions will generally be different from Figure 2 the arrangements shown by way of example.

[0358] Accordingly, the present disclosure extends to gas turbine engines having any arrangement in terms of gearbox type (such as stellar or planetary gears), support structure, input and output shaft arrangements and bearing positions.

[0359] Optionally, the gearbox may drive additional and / or alternative components (for example, an intermediate pressure compressor and / or a supercharger compressor).

[0360] Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of compressors and / or turbines and / or an alternative number of interconnected shafts. By way of a further example, Figure 1The gas turbine engine shown in [description] has split nozzles 18, 20, which means that the flow through the bypass duct 22 has its own nozzle 18 that is separate from and radially external to the core engine nozzle 20. However, this is not restrictive, and any aspect of the present disclosure can also be applied to an engine in which the flow through the bypass duct 22 and the flow through the core 11 are mixed or combined before (or upstream of) a single nozzle that can be referred to as a mixed-flow nozzle. One or both nozzles (whether mixed or split) can have a fixed or variable area.

[0361] As a further example, other gas turbine engines to which the present disclosure can be applied can be direct drive engines without a gearbox for the main shaft. Figure 4 A cross-sectional view of one such engine is shown in [description].

[0362] Referring Figure 4 , the gas turbine engine is generally designated by 10 and has a main rotational axis 9. The engine 10 includes, in an axial flow series arrangement, an air intake 12, a propulsive fan 23, an intermediate pressure compressor 14, a high pressure compressor 15, combustion equipment 16, a high pressure turbine 17, an intermediate pressure turbine 19a, a low pressure turbine 19, and an exhaust nozzle 20. A nacelle 21 surrounds the engine 10 and defines the air intake 12 and the exhaust nozzle 20.

[0363] In use, air entering the air intake 12 is accelerated by the fan 23 to produce two airflows: a core airflow A and a bypass airflow B. The core airflow A flows into the intermediate pressure compressor 14, and the bypass airflow B passes through the bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 14 compresses the airflow A before delivering the air to the high pressure compressor 15 where further compression occurs.

[0364] 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 can be referred to as a burner 16, where the terms "combustion equipment 16" and "burner 16" can be used interchangeably herein. The resulting hot combustion products then expand through the high pressure turbine 17, the intermediate pressure turbine 19a, and the low pressure turbine 19 before being discharged through the nozzle 20, thereby driving the high pressure turbine, the intermediate pressure turbine, and the low pressure turbine to provide additional propulsive thrust. The high pressure turbine 17, the intermediate pressure turbine 19a, and the low pressure turbine 19 each drive the high pressure compressor 15, the intermediate pressure compressor 14, and the fan 23 respectively via suitable interconnecting shafts.

[0365] Other gas turbine engines to which the present disclosure may apply may have alternative configurations. By way of example, such engines may have alternative numbers of interconnected shafts (e.g., two) and / or alternative numbers 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.

[0366] Although the examples described relate to turbofan engines, 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.

[0367] The geometry of the gas turbine engine 10 and its components are 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.

[0368] The fuel F supplied to the combustion equipment 16 may include fossil-based hydrocarbon fuels 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, inorganic substances, and metals.

[0369] 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; alternative fuels from abiogenic sources; jatropha; halophytes, and algae, rather than from fossil-based hydrocarbons. SAF is understood not to include fossil fuels.

[0370] 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 defining 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 herein to soot or smoke equally applies to other types of particulate emissions known in the art. Gaseous 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 herein to gaseous emissions equally applies to other types of gaseous emissions known in the art.

[0371] A relatively high specific energy (i.e., energy per unit mass), expressed in MJ / kg, can at least in part reduce the takeoff weight and thus potentially provide a relative improvement in fuel efficiency. A relatively high energy density (i.e., energy per unit volume), expressed in MJ / L, 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 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 to optimize the flight profile; 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.

[0372] 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, where a significant portion of a given composition is derived from molecules having 9 to 15 carbons, with an average of 12 carbons.

[0373] 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 pending approval. However, in the aviation industry, sustainable aviation fuel blends that are expected to include up to (and including) 100% sustainable aviation fuel (SAF) will ultimately be approved for use.

[0374] 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 aromatics 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.

[0375] 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 variations in its composition, the aromatic content of the sustainable aviation fuel can be within a numerical value or range that is defined by or defined within any two of the values in the previous sentence, including the end values (i.e., these values can form an upper or lower limit), 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%.

[0376] At least in part due 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.

[0377] As Figure 5 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.

[0378] Figure 5 An aircraft 1 with a propulsion system 2 including two gas turbine engines 10 is shown. 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 that has 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.

[0379] 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.

[0380] In another example, the wing fuel tanks 53a, 53b may not be fluidly connected to the central fuel tank 50, thus 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 bypassing 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 different from each other, such that the aircraft 1 can change fuels during flight. Therefore, 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.

[0381] 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.

[0382] Aircraft are typically refueled at multiple different airports, for example at the start and end of a long - haul flight. Although there are 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) and 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.

[0383] The inventors recognize that since different fuels can have different properties while still meeting standards, knowledge of the fuel available for the aircraft 1 can allow for more effective and customized control of the aircraft 1, and more particularly 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.

[0384] As used herein, the term "fuel characteristic" refers to an intrinsic or inherent 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 heat exchange system 3000. Examples of fuel characteristics include one or more of the following:

[0385] i. The percentage of sustainable aviation fuel (SAF) in the fuel, or an indication that the fuel is a fossil fuel (e.g., fossil kerosene), or an indication that the fuel is a pure SAF fuel;

[0386] ii. Parameters of the hydrocarbon distribution of the fuel, such as:

[0387] · The aromatic content of the fuel, and optionally / alternatively the polyaromatic content of the fuel;

[0388] · The hydrogen-to-carbon ratio (H / C) of the fuel;

[0389] · Percentage information on the composition of some or all of the hydrocarbons present;

[0390] iii. The presence or percentage of specific elements or species, such as:

[0391] · The percentage of nitrogenous substances in the fuel;

[0392] · The presence or percentage of tracer substances or trace elements in the fuel;

[0393] · The naphthalene content of the fuel;

[0394] · Sulfur content of the fuel;

[0395] · Naphthene content of the fuel;

[0396] · Oxygen content of the fuel;

[0397] iv. One or more properties of the fuel used in the gas turbine engine 10, such as:

[0398] · The level of non-volatile particulate matter (nvPM) emissions or CO2 emissions during combustion;

[0399] · The coking level of the fuel;

[0400] v. One or more properties of the fuel itself independent of the use or combustion of the engine 10, such as:

[0401] · The thermal stability of the fuel (e.g., the thermal breakdown temperature; a numerical value for thermal stability at any temperature can be assigned by taking the reciprocal of the deposition rate of the fuel decomposition products at a given temperature); and

[0402] · One or more physical properties, such as density, viscosity, calorific value, freezing temperature, and / or heat capacity.

[0403] 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 heat exchange system 3000. The determination of 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.

[0404] 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:

[0405] (i) Physically and / or chemically detecting one or more characteristics or parameters of the fuel composition (which 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);

[0406] (ii) Retrieving fuel characteristic information from on-board memory / data repositories; and / or

[0407] (iii) Receiving data, for example, from an input provided at a user interface, or receiving data transmitted to the aircraft 1.

[0408] In some examples, one or more fuel properties may be determined during operation of the gas turbine engine 10, such as by inferring the fuel properties of fuel supplied to the combustor 16 from engine performance metrics or by performing in-wing detection.

[0409] 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 properties and the corresponding detected parameters, an alert may be provided.

[0410] 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 from an input provided by a user interface, or by receiving data transmitted to the aircraft by scanning a barcode associated with fuel delivery, for example.

[0411] 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 on 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:

[0412] ● The aromatic or naphthene content in the fuel may be determined based on measurements of the expansion of a sensor component made of a sealing material such as a nitrile sealing material.

[0413] ● 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.

[0414] ● Measurements 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.

[0415] ● 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.

[0416] ● 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 trajectory during the operation of the gas turbine 10.

[0417] ● The fuel characteristic including the aromatic content can be determined based on ultraviolet-visible spectroscopic measurements performed on the fuel.

[0418] ● 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.

[0419] ● The calorific value of the fuel can be determined during the operation of the aircraft 1 based on measurements made during fuel combustion, for example using fuel flow and shaft speed or temperature changes at the burner 16.

[0420] ● 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.

[0421] ● 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 (which relationship indicates a temperature rise at 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 reference to known values of the first fuel.

[0422] 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:

[0423] · T30 = High-pressure compressor (HPC) outlet total temperature;

[0424] · T40 = Combustion outlet total temperature;

[0425] · T41 = High-pressure turbine (HPT) rotor inlet total temperature.

[0426] Any suitable method known in the art can be used, and the determination of fuel characteristics is not discussed further herein.

[0427] Accordingly, the aircraft 1 may include a fuel composition determination module 57 that is arranged to determine at least one fuel property 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 properties 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 property information. In other embodiments, the control function 58 may be provided by the EEC, and a separate unit or module may not be provided.

[0428] Figure 6 An exemplary fuel system 1000 for directly driving a gas turbine engine 10 is schematically shown, the fuel system 1000 including 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, 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 portions of the fuel management system 1500 as well as a recirculating oil system 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.

[0429] 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 primary fuel-oil heat exchanger 1006 and the secondary fuel-oil heat exchanger 1004. The primary fuel-oil heat exchanger 1006 may be described as the main fuel-oil heat exchanger because the fuel flowing through it can be the main fuel flow path from the fuel tank 50 to the burner 16. The secondary fuel-oil heat exchanger 1004 may be described as an auxiliary fuel-oil heat exchanger or a servo fuel-oil heat exchanger because the fuel flowing through it can be supplied to the auxiliary systems or servo mechanisms 33 of the aircraft 1. Such auxiliary systems / servo mechanisms may include, but are not limited to, fuel hydraulic actuation; heating; nacelle anti-icing systems; engine actuators; and / or turbine case cooling (TCC) servo valves. Only the fuel passing through the secondary fuel-oil heat exchanger is used in these auxiliary systems. The fuel used in these auxiliary systems can be returned to the fuel tank for later recirculation, or it can be recombined with other fuel leaving the primary fuel-oil heat exchanger 1006 and enter the burner. Thus, in some embodiments, the fuel passing through the secondary fuel-oil heat exchanger 1004 may not be supplied to the burner, but instead may optionally be returned to the aircraft fuel tank 50 after being used for an auxiliary system (such as a fuel hydraulic actuator). The fuel passing through the secondary fuel-oil heat exchanger 1004 may instead return along the main fuel flow path between the fuel tank 50 and the engine fuel pump 1003. 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 primary fuel-oil heat exchanger 1006 before the secondary fuel-oil heat exchanger 1004. After leaving the primary fuel-oil heat exchanger 1006, the fuel then passes through the engine fuel pump 1003 and travels to the burner 16. The engine fuel pump 1003 may be described as the main fuel pump and is positioned upstream of the secondary fuel-oil heat exchanger 1004 in the fuel management system 1500.

[0430] The primary fuel - oil heat exchanger 1006 and the secondary fuel - oil heat exchanger 1004 are configured such that an oil flow is conveyed through each heat exchanger in addition to the fuel flow passing therethrough. The primary fuel - oil heat exchanger 1006 and the secondary fuel - oil heat exchanger 1004 are configured such that heat can be transferred between the oil and the fuel passing therethrough. During 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 therethrough during operation.

[0431] Typically, at least most of the fuel passes through the primary fuel - oil heat exchanger 1006, and at least a portion of the fuel passing through the primary fuel - oil heat exchanger 1006 also passes through the secondary fuel - oil heat exchanger 1004. The two heat exchangers 1004, 1006 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 respective heat exchanger, for example in the form of a fuel bypass conduit 1005 as Figure 6 shown. 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 conduit 1005. In various embodiments, bypass conduits can be provided for each of the heat exchangers 1004, 1006, thereby allowing a portion of the fuel to bypass either or both of the heat exchangers. Additionally or alternatively, one or more bypass conduits 2005, 2005a for the oil can be provided for either or both of the fuel - oil heat exchangers 1004, 1006, thereby allowing a portion of the oil to bypass one or more of the heat exchangers. A valve arranged to control the fluid flow through the bypass conduit 1005 can be referred to as a bypass valve.

[0432] 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 heat exchanger - in the described embodiment, the same oil passes through one heat exchanger and then through the other heat exchanger, but it should be understood that in other embodiments, different oil circuit systems can supply each heat exchanger and thus different oils can flow through one fuel - oil heat exchanger to reach the other fuel - oil heat exchanger.

[0433] During Figure 6In the example shown, the fuel system 1000 further includes a temperature sensor 1009 that is arranged to sense the temperature of the fuel proximate to or reaching the burner 16. In the illustrated example, the temperature sensor 1009 is shown adjacent to the inlet of the burner 16, but in various examples, the sensor 1009 can be located anywhere downstream of the primary fuel - oil heat exchanger 1006 or the secondary fuel - oil heat exchanger on the fuel side so as to provide a measure of the fuel temperature at different points along the flow path. In the gas turbine engine 10 as described herein, the fuel passes through fuel nozzles before entering the combustion chamber itself, and these nozzles are very susceptible to fuel thermal decomposition, which can cause blockage of relatively narrow passages (e.g., metering slots); thus, in some embodiments, it is advisable to place the sensor 1009 as reasonably close as possible to the inlet of the nozzle. In some embodiments, multiple temperature sensors 1009 can be used. One or more temperature sensors 1009 are arranged to provide an output (temperature data) to the controller 58.

[0434] Figure 7AAn alternative exemplary fuel system 6000 is shown that includes 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 the fuel system 1000, fuel is pumped from the fuel tank 50 by a low-pressure fuel supply pump 1002. The fuel then flows through a primary fuel-oil heat exchanger 1006 before reaching an engine fuel pump 1003 that pumps the fuel along its flow path to the burner 16, optionally via a secondary fuel-oil heat exchanger 1004. The fuel flow through the secondary fuel-oil heat exchanger 1004 can be controlled with a valve 1004a. The fuel system 6000 differs from the fuel system 1000 in that the fuel system 6000 includes a recirculation valve 6010 that 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 line 6011 and the proportion that continues more directly to the burner 16. In the illustrated example, the recirculation valve 6010 is positioned downstream of the primary fuel-oil heat exchanger 1006. In the illustrated example, the recirculation valve 6010 is positioned downstream of the engine fuel pump 1003. In some 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) line 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 burner 16. It is contemplated that in an alternative embodiment, the recirculation valve may be positioned upstream of the engine fuel pump 1003. In such embodiments, the recirculation valve 6010 would 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 multiple times before reaching the pump 1003. This recirculation provides a mechanism for controlling the fuel flow within the fuel management system 6500 and within the heat exchange 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.

[0435] A recirculation line 6011 can thus be provided, which can be referred to as a recirculation line 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, and the line 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 line 6011 and a bypass line 1005 can be provided for any given heat exchanger 1004, 1006. In some embodiments, the same lines 1005, 6011 can be used as recirculation lines and bypass lines, and one or more valves can be used to control the direction of fluid flow therethrough.

[0436] Controlling the fuel flow in the fuel system 6000 - by recirculating and / or bypassing one or more heat exchangers - can contribute to thermal management (e.g., affecting the fuel temperature at the burner 16 or the inlet of the pump 1003, or the heat transferred to the fuel). Thus, part of the control of the heat exchange system 3000 can be the control of one or more valves that control the oil and / or fuel flow through one or more recirculation and / or bypass lines 6011, 1005.

[0437] The control of the fuel flow through the recirculation line 6011 can be based on fuel temperature measurements (e.g., using a temperature sensor at a location downstream of the primary fuel - oil heat exchanger 1006, and possibly also using temperature measurements upstream of the primary fuel - oil heat exchanger 1006). The return of recirculated fuel upstream of the primary fuel - oil heat exchanger 1006 can allow for a reduction in heat transfer from the oil to the fuel, thereby suppressing an instantaneous overshoot that can 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 is reduced, thus typically resulting in a temperature peak. Also, by heating the system by circulating pre - heated fuel, regulating the fuel flow through the recirculation line 6011 can also be used to reduce icing. Thus, the recirculation line 6011 can be used in multiple ways to improve engine thermal management.

[0438] Figure 7B is shown Figure 7A An alternative exemplary fuel system 7000 is shown. The fuel system 7000 is similar to Figure 7AThe fuel system 6000 shown, but different in that at least a portion of the fuel leaving the secondary fuel-oil heat exchanger 1004 is supplied to additional engine and / or aircraft mechanisms 7010 rather than to the burner 16. These mechanisms 7010 may include one or more of a nacelle anti-ice system, an actuator, and / or a turbine case cooling (TCC) system. At least a portion of the fuel leaving the secondary fuel-oil heat exchanger 1004 returns to the fuel supply system (i.e., the fuel tank 50 or the fuel supply pump 1002) or downstream of the supply system for redistribution (i.e., upstream of the primary fuel-oil heat exchanger 1006). The portion of the fuel returning to the fuel supply system / downstream of the fuel supply system may be controlled by a valve 7011. In Figure 7B the example of, fuel that does not pass through the secondary fuel-oil heat exchanger 1004 is delivered to the burner 16 (at least not directly to the burner), it should be understood that fuel returning to the fuel tank 50 may later re-enter the engine 10 and may then be sent to the burner 16.

[0439] Figure 7C Another alternative exemplary fuel system 7000a is shown. The fuel system 7000a is similar to the fuel system 7000 shown in Figure 7B except that the portion of the fuel leaving the secondary fuel-oil heat exchanger 1004 returns to join the main fuel flow in its path to the burner 16 after optionally being supplied to one or more additional engine and / or aircraft mechanisms 7010 (as may be described above). Any fuel not needed by these auxiliary systems 7010 may be returned directly to the main fuel flow path / burner 16. At least a portion of the fuel leaving the secondary fuel-oil heat exchanger 1004 may return to the main fuel flow path without passing through the auxiliary system 7010 - the proportion directed to the auxiliary system 7010 may be adjusted as needed and is controlled by a valve 7011. In other embodiments, the fuel used in one or more such additional aircraft mechanisms 7010 may then return to the tank 50 rather than being delivered to the burner 16.

[0440] The gas turbine engine 10 of the described aircraft 1 includes a recirculating 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 and Figure 8AIn the illustrated embodiment, the recirculating oil system includes a single closed oil circuit system 2000. In other embodiments, the recirculating oil system may include a primary oil circuit system 2000 and a secondary oil circuit system (not shown), each of the primary oil circuit system and the secondary oil circuit system being a closed circuit oil system. The secondary oil circuit system may supply oil to one or both of the primary fuel-oil heat exchanger 1006 and the secondary fuel-oil heat exchanger, and possibly additional heat exchangers (such as an air-oil heat exchanger).

[0441] In the illustrated embodiment, the closed circuit oil system 2000 includes a fuel tank 2002 adapted to hold a large quantity of oil. In some embodiments, gas is removed from the oil within the fuel tank 2002 by a degasser.

[0442] The feed pump 2004 is configured to pump oil from the fuel tank 2002 to the secondary fuel-oil heat exchanger 1006. Under cruise conditions, the average temperature of the oil entering the secondary fuel-oil heat exchanger 1006 is higher than the average temperature of the fuel entering the secondary fuel-oil heat exchanger 1006. In the secondary fuel-oil heat exchanger 1006, thermal energy is transferred from the oil flow to the fuel flow. In this way, the average temperature of the oil flow leaving the secondary fuel-oil heat exchanger 1004 is lower than the average temperature of the oil flow entering the secondary fuel-oil heat exchanger 1006. Also in this way, the average temperature of the fuel leaving the secondary fuel-oil heat exchanger 1004 is higher than the average temperature of the fuel entering the secondary fuel-oil heat exchanger 1006.

[0443] The oil then flows from the secondary fuel-oil heat exchanger 1004 to the primary fuel-oil heat exchanger 1006. In some embodiments, the oil may flow through an air-oil heat exchanger (not shown) between the secondary fuel-oil heat exchanger 1004 and the primary fuel-oil heat exchanger 1006.

[0444] In the primary fuel-oil heat exchanger 1006, thermal energy is transferred from the oil flow to the fuel flow. In this way, the average temperature of the oil flow leaving the primary fuel-oil heat exchanger 1006 is lower than the average temperature of the oil flow entering the primary fuel-oil heat exchanger 1006. Also in this way, the average temperature of the fuel leaving the primary fuel-oil heat exchanger 1006 is higher than the average temperature of the fuel entering the primary fuel-oil heat exchanger 1006.

[0445] From the primary fuel-oil heat exchanger 1006, the oil flow is then conveyed through the engine components 2006, which oil flow is intended to cool and / or lubricate these engine components. The oil acts as a lubricant and coolant within the engine components 2006 and acquires heat from the engine components 2006 during this process. The engine components 2006 may include one or more bearing chambers.

[0446] Oil from the engine component 2006 is collected in the oil sump 2008. The scavenge pump 2010 is configured to pump the oil from the oil sump 2008 back into the fuel tank 2002.

[0447] Figure 8B and Figure 8C Alternative oil circuit systems 2000b, 2000c are shown. In Figure 8B the oil circuit system 2000b, the oil flows through an air - oil heat exchanger for cooling before the primary fuel - oil heat exchanger 1006. The secondary fuel - oil heat exchanger does not exist in this alternative oil circuit system 2000b.

[0448] Figure 8B Details of the air - oil heat exchanger 2020 are also shown, which are not shown in other figures for clarity, particularly with dashed arrows indicating the air flow. The air - oil heat exchanger 2020 has an air inlet 2020a and an air outlet 2020b. The air inlet or intake 2020a may be arranged to capture the air in the engine bypass duct 22 downstream of the outlet guide vanes of the fan 23. The heat exchanger exhaust / air outlet 2020b may be arranged to output the air back into the bypass duct 22 downstream of the inlet 2020a, or directly to the external atmosphere (e.g., if a dedicated outlet nozzle is provided).

[0449] A valve 2022, referred to as an air valve, is used to control the flow of air through the air - oil heat exchanger 2020. In the example shown, the air valve 2022 is provided at or near the inlet 2020a of the heat exchanger 2020 and can thus be referred to as an air inlet valve 2022. In other embodiments, the air valve 2022 may be provided at or near the outlet 2020b of the heat exchanger 2020 and can thus be referred to as an air outlet valve. If the valve 2022 can be adjusted to control the air flow through the air - oil heat exchanger 2020, any suitable position of the valve 2022 can be selected. In the example described, the air valve 2022 can be continuously adjusted between a fully closed position (no air flow through the heat exchanger) and a fully open position (maximum air flow through the heat exchanger). In other examples, the air valve 2022 can be adjusted between a plurality of discrete positions rather than continuously, e.g., six, five, four, three, or two different positions. In embodiments where the air valve 2022 has only two valve positions, these positions can be "open" and "closed" - in some such embodiments, when a medium level of cooling is required, the valve 2022 can be repeatedly opened and closed to provide a pulsed air flow.

[0450] Figure 8CA recirculation oil system including an oil circuit system 2000c is shown. The main oil flow path passes through all of the heat exchangers 1004, 1006, 2020 in the circuit. To regulate the oil flow through the heat exchangers, one or more bypass pipes 2005, 2005a, 2005b are provided, and bypass valves 2007, 2007a, 2007b are used to vary the flow through the bypass pipes. The controller 58 can be used to actively manage the fuel flow and / or oil flow through and around the heat exchangers 1004, 1006, 2020. The controller 58 can be part of or provided by the EEC, or can be a separate unit. The control can be automatic, for example, by the EEC.

[0451] In other embodiments, branch paths leading to different heat exchangers can be used instead. It should be understood that although the bypass pipes technically provide parallel, alternative flow paths, the branch paths described here with reference to parallel flow differ in that each branch route of the parallel arrangement of heat exchangers includes a heat exchanger, while the bypass pipes are merely pipes on which there are no (significant) heat exchangers or other components.

[0452] It should be understood that in some embodiments, a combined arrangement of heat exchangers and bypass pipes can be used, and Figure 8C (all heat exchangers in series, multiple bypass pipes) can be considered to show one end of the design range. In an alternative embodiment, one or more of the heat exchangers can be part of a secondary oil circuit system. In an alternative embodiment, the heat exchangers can be arranged in parallel, whereby the oil circuit system 2000 branches into separate branches, where two or more separate branches each include a heat exchanger.

[0453] Figure 8C The oil circuit system 2000 shown provides a series oil flow path that starts from the tank 2002, passes through the oil pump 2004, forward through the secondary fuel - oil heat exchanger 1004, the primary fuel - oil heat exchanger 1006, then through the air - oil heat exchanger 2020, then into the engine components 2006 to be cooled and / or lubricated, then is collected in the oil sump 2008, and then is pumped back to the tank 2002 by the oil pump 2010. Thus, the heat exchangers 1004, 1006, 2020 are arranged in series. The order of the heat exchangers 1004, 1006, 2020 can be reversed, for example, such that in an alternative embodiment, the primary fuel - oil heat exchanger 1006 is after the air - oil heat exchanger 2020.

[0454] Figure 8CThe oil circuit system 2000c shown includes three oil bypass pipes 2005, 2005a, 2005b. The first bypass pipe 2005 is arranged to allow a portion of the oil to bypass the secondary fuel - oil heat exchanger 1004 and is controlled by a first bypass valve 2007. The first bypass pipe 2005 takes oil upstream of the inlet of the secondary heat exchanger 1004 and returns it to the main oil flow path downstream of the secondary heat exchanger 1004 and before the primary fuel - oil heat exchanger 1006 and the air - oil heat exchanger 2020. The second bypass pipe 2005a is arranged to allow a portion of the oil to bypass the primary fuel - oil heat exchanger 1006 and is controlled by a second bypass valve 2007a. The second bypass pipe 2005a takes oil upstream of the inlet of the primary fuel - oil heat exchanger 2020 and returns it to the main oil flow path downstream of the primary fuel - oil heat exchanger 2020 before the path reaches the air - oil heat exchanger 2020. The third bypass pipe 2005b takes oil upstream of the inlet of the air - oil heat exchanger 2020 and returns it to the main oil flow path downstream of the air - oil heat exchanger 2020 before the path reaches the engine component 2006. In an embodiment of the primary oil circuit system 2000 having only one oil bypass pipe 2005, the selected location can be the location of the third bypass pipe 2005b such that there is a bypass for the air - oil heat exchanger 2020 and 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 2005a on the primary fuel - oil heat exchanger 1006 can help in the rapid adjustment of the heat transfer ratio if there is a risk that the heat transfer ratio drops below the desired value (e.g., depending on the determined fuel properties, a lower limit of the heat transfer ratio significantly above zero can be set). Depending on the fuel properties, the temperature limits set for some fuels can be very strict. Having the bypass pipes 2005, 2005a, 2020 on all the heat exchangers 1004, 1006, 2020 can prevent the oil from getting too cold under certain conditions to avoid the risk of oil condensation - it should be understood that this may be of more concern during cold - day idle conditions (e.g., at start - up or ground idle, or during descent) than during cruise.

[0455] The inventors recognize that using fuels different from traditional kerosene - based jet fuels (such as sustainable aviation fuels) can result in different fuel properties and that 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 (as in relation to Figure 14 and Figure 15In the described method (used in), introducing one or more oil bypass pipes 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, especially one or more controllable bypass valves 2007, 2007', 2007a, plays a key role in managing the heat transfer ratio in such an engine 10.

[0456] In addition, the inventors recognized that while cruise conditions typically account for most of the flight vehicle engine operation time, 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 can still cause a temperature increase - thus, the use of non-conventional fuels may have a greater impact on the optimal heat management method under idle conditions. Regarding Figure 14 and Figure 15 methods for addressing these two flight vehicle operating scenarios are discussed in more detail.

[0457] 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 fuel or oil with an effectively infinite flow rate 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 rise above the temperature of the oil entering the heat exchanger). By the definition of the second law of thermodynamics, temperature equilibrium is the limit. In some embodiments, such as Figure 9As 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 is raised more than when simply passing through the heat exchanger, and in some cases raised 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 in fuel temperature after the fuel has passed through the fuel-oil heat exchangers 1004, 1006, and further cooling the oil of the oil system 2000 before returning to the engine component 2006. In various embodiments, bypass pipes or recirculation pipes for the oil and / or fuel can be provided around the refrigeration cycle device 1007.

[0458] Figure 10 An exemplary heat exchange system 3000 is schematically shown, which includes Figure 8A a closed-loop oil system 2000 and Figure 6 a fuel system 1000. The two systems 1000, 2000 are depicted together to show the interaction between the fuel and oil flows within the gas turbine engine 10. The fuel flow is shown by thick black arrows, and the oil flow is shown by thin black arrows.

[0459] Figure 11 An alternative exemplary heat exchange system 4000 is schematically shown, including Figure 8A a closed-loop oil system 2000 and Figure 7A a fuel delivery system 7000a. The systems 7000a, 2000 are substantially the same as those shown in Figure 8A and Figure 7A respectively, the fuel flow is again shown by thick black arrows, and the oil flow is shown by thin black arrows. However, the position of the recirculation valve 6010 is slightly different from that in Figure 7A , i.e., upstream of the bifurcation in the fuel flow path leading to the secondary heat exchanger 1006, rather than between the inlet 1006a and the outlet 1006b of the heat exchanger 1006. Thus, compared withFigure 7A Unlike the embodiments shown, in Figure 11 the embodiments shown, all of the fuel leaving the pump 1003 flows through the recirculation valve 6010. The proportion of fuel flowing through the secondary fuel-oil heat exchanger 1004 can be controlled by the valve 1004a. The control methods of the valves 1004a and 6010 can be adjusted appropriately.

[0460] It should be understood that in embodiments such as Figure 7A the embodiments shown, where the fuel flowing through the secondary fuel-oil heat exchanger 1004 is used as the working fluid in the auxiliary system and then returned to the tanks 50, 53 instead of being delivered to the burner 16, the fuel flow line from the secondary heat exchanger 1006 to the burner 16 can be replaced with a fuel flow line returning to the tank 50.

[0461] 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 engine operation can be optimized for these different fuel properties. The claimed fuel with a higher calorific value may also have greater thermal stability, allowing the fuel to absorb more heat, thereby providing improved oil cooling and / or improved combustion properties in the burner. Recirculating the fuel through the primary fuel-oil heat exchanger 1006 allows the fuel to receive more heat from the oil, thereby increasing the fuel temperature and improving oil cooling.

[0462] The recirculation valve 6010 can be located downstream of the fuel pump 1003 and can thus provide improved flexibility of the fuel flow. Therefore, without changing the pump speed, by instead increasing the recirculation amount, a lower flow rate of fuel can be provided to the burner 16 for a higher calorific value fuel.

[0463] Figure 12 An alternative exemplary heat exchange system 4500 is schematically shown, including Figure 8C a closed-loop oil system 2000 and Figure 7A a part of the fuel delivery system 7000a. For clarity, only the fuel flow through the primary fuel-oil heat exchanger 1006 and the secondary fuel-oil heat exchanger 1004 is shown.

[0464] In the above-described embodiments, 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 the 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. 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.

[0465] The inventors have recognized 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 among aviation fuels.

[0466] For clarity and conciseness, a single bypass valve 2007 and a corresponding bypass duct 2005 are referred to in the methods described herein. As Figure 12 shown in the embodiments of, the embodiments may actually include multiple bypass valves 2007, 2007a, 2007b, each bypass valve having an associated bypass duct 2005, 2005a, 2005b, such that the oil flow can bypass one or more of the secondary fuel-oil heat exchanger 1004, the primary fuel-oil heat exchanger 1006, and / or the air-oil heat exchanger 2020. When referring to the "bypass valve 2007" or the "bypass duct 2005" in the discussion of the Figures 13 to 17 methods shown, it refers to one or more of the multiple bypass valves 2007, 2007a, 2007b and the associated bypass ducts 2005, 2005a, 2005b, depending on which combination of heat exchangers is being discussed.

[0467] Figure 13 A first method 100 is shown, including using knowledge of fuel characteristics to determine control of the heat exchange system 3000. 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.

[0468] The engine 10 for implementing method 100 includes an air-oil heat exchanger 2020 and a fuel-oil heat exchanger 1006, a bypass duct 2005 arranged to allow a proportion of the oil flow to pass through one of the air-oil heat exchanger and the fuel-oil heat exchanger; and a bypass valve 2007 arranged to allow changing the proportion of the oil conveyed through the bypass duct, and method 100 includes controlling 104 the bypass valve 2007 based on at least one fuel characteristic so as to adjust the proportion of the oil conveyed through each heat exchanger under cruise conditions.

[0469] More specifically, for the embodiments described above with reference to Figures 5 to 12 the method 100 includes controlling 104 a bypass valve 2007 of an oil system 2000 under cruise conditions based on at least one fuel characteristic so as to regulate the proportion of oil delivered via each bypass conduit 2005 of the oil circuit system 2000, thereby controlling the amount of oil bypassing a primary fuel-oil heat exchanger 1006, the amount of oil bypassing an air-oil heat exchanger 2020, and optionally the amount of oil bypassing a secondary fuel-oil heat exchanger 1004. In embodiments having heat exchangers with different arrangements (e.g., heat exchangers arranged in parallel rather than in series), the bypass valve 2007 may be replaced by or used as a modulating valve that directs the oil flow along one or more branches of a parallel shunt arrangement, thereby allowing the proportion of oil delivered via each heat exchanger to be varied in this way.

[0470] As Figure 13 shown by the dashed line in

[0471] (i) In embodiments having multiple fuel sources, where only one or the other fuel source is used at any given time (and knowledge of which fuel source the fuel is drawn from may prompt selection of appropriate stored fuel characteristics), the determination is performed once for each fuel source 50, 53 at fuel replenishment / at the start of flight; or

[0472] (ii) The determination is performed frequently during flight, e.g., in response to a change in which (which) fuel source the fuel is drawn from (note that in some embodiments, the fuel supplied to the burner 16 may be a mixture of fuels from different sources), or periodically.

[0473] Accordingly, the bypass valve 2007 can be arranged to divert a fixed portion of the oil flow through the bypass conduit 2005 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 bypass valve 2007 can be arranged to divert a variable portion of the oil flow through the bypass conduit 2005 during operation of the engine 10 during a single flight (option (ii) above). Accordingly, the bypass valve 2007 can be actively controlled to vary the proportion of oil delivered through the bypass conduit 2005 (and thus through each of the heat exchangers 1004, 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 bypass valve 2007 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 bypass valve 2007 can be closed-loop or open-loop - in particular, when feedback data (e.g., from an oil flow sensor) is available, a closed-loop control method may be preferred.

[0474] The inventors have recognized that fuels having certain fuel characteristics, such as 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.

[0475] Fuel characteristics may not be the only data considered when adjusting the bypass valve 2007 - for example, the temperature of the fuel leaving the fuel-oil heat exchanger 1006 or entering the burner 16, as well as the heat resistance of fuel system components downstream of the heat exchanger 1006 (and optionally also the refrigeration cycle device 1007, if present), can also be considered. In the presence of the refrigeration cycle device 1007, the bypass valve 2007 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, can also pass through the refrigeration cycle device 1007 - however, the refrigeration cycle device 1007 may be powered off / inoperable for a period of time and thus not provide any temperature boost.

[0476] According to the suitability of at least one determined fuel characteristic, the method 100 can include using the bypass valve 2007 to perform one or more of the following:

[0477] ● During one or more periods of at least 30 minutes of cruise, transport all of the oil via the primary fuel - oil heat exchanger 1006 (without passing through the air - oil heat exchanger 2020), such that during at least some operating periods of cruise, no heat is lost to the environment via the air - oil heat exchanger;

[0478] ● During at least 15% and optionally at least 20% of the total cruise time, transport all of the oil via the primary fuel - oil heat exchanger 1006 (without passing through the air - oil heat exchanger 2020);

[0479] ● During at least 90% of the cruise time, transport at least 95% of the oil via the primary 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) is lost to the environment via the air - oil heat exchanger;

[0480] ● 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

[0481] ● 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

[0482] ● 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 in the heat exchange system 3000 per kilogram of fuel. In some embodiments, the heat transferred can be in the range of 350 kJ / kg to 450 kJ / kg of fuel.

[0483] The suitability of at least one determined fuel property can be determined by comparing one of the multiple fuel properties with a threshold set for the corresponding fuel property. 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 properties can be checked simultaneously - for example, when one fuel property is within a specific range, a lower threshold of another fuel property is used. In other embodiments, the fuel properties can be considered independently / isolatly.

[0484] 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 bypass valve 2007 can be controlled 104 such that at least 80%, and optionally 90% to 100%, of the heat transferred away 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%, then at least 80%, and optionally 90% to 100%, of the heat transferred away 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%, then the bypass valve 2007 can be controlled such that at least 80%, and optionally 90% to 100%, of the heat transferred away 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, then at least 80%, and optionally 90% to 100%, of the heat transferred away 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 higher calorific values, the flow rate is typically reduced to achieve the same level of thrust - if the amount of oil passing through the heat exchanger 1006 is not reduced, the fuel in the primary 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 higher calorific values - thus, in some cases, the increased fuel temperature rise due to the lower flow rate is acceptable, and in fact, it may also be possible to further increase the heat. This demonstrates the utility of considering multiple fuel characteristics in combination rather than considering individual fuel characteristics alone in some embodiments.

[0485] Any one of the embodiments can use only one or more fuel characteristics. A processing module can be provided, optionally as part of the fuel characteristic determination module 57, and / or as part of the general engine EEC, to make decisions on the control of the modulation valve 2016 based on the fuel characteristic data and optionally also based on other data.

[0486] The gas turbine engine 10 for an aircraft implementing the method 100 includes a bearing 2006 arranged to support a fan shaft, and an oil circuit system 2000 arranged to supply oil to the bearing 2006. The heat exchange system 300 of the engine 10 includes: an air-oil heat exchanger 2020 through which oil in the oil circuit system 2000 flows; and a fuel-oil heat exchanger 1006 through which oil and fuel in the oil circuit system flow, such that heat is transferred between the oil and the fuel. A bypass valve 2007 is provided, which is arranged to allow changing the proportion of oil conveyed via a bypass duct 2005 and thus bypassing one or more heat exchangers 1006, 2020. The engine 10 further includes a fuel composition determination module 57 arranged to determine at least one fuel characteristic of the fuel arranged to be burned by a burner 16. The bypass valve 2007 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 conveyed via the bypass duct 2005 under cruise conditions (and thus the proportion of oil flowing through each heat exchanger 1006, 2020). A controller 58 may be provided to make and implement decisions based on the output of the fuel composition determination module 57.

[0487] The inventors also 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, some fuels can be heated to higher temperatures 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 less heat is lost to the surrounding environment. The controllable heat exchange system 3000 plays a key role in managing the heat transfer ratio.

[0488] 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 during idle is much lower than during cruise, and even though the heat load of the fuel is relatively small, it may cause a temperature increase - 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 The methods 200, 300 address these two scenarios of aircraft operation.

[0489] Figure 14 The method 200 shows how to achieve these considerations under cruise conditions, and Figure 15Method 300 for achieving these considerations at idle (e.g., when the aircraft is starting, stationary during boarding, and (taxiing towards the runway or hangar, or between other ground-based positions), or during certain periods of flight (e.g., landing)) is shown.

[0490] Looking first at method 200 to be performed during cruise, method 200 is arranged to be performed in a direct drive gas turbine engine 10, which includes: an oil circuit system 2000 arranged to supply oil to engine components 2006; and a heat exchange system 3000 including an air-oil heat exchanger 2020 through which oil in the oil circuit system flows; and a primary fuel-oil heat exchanger 1006 through which oil and fuel in the oil circuit system flow such that heat is transferred between the oil and the fuel. A bypass duct 2005 is provided and is arranged to allow a certain proportion of the oil to flow through one of the air-oil heat exchanger 2020 or the fuel-oil heat exchanger 1006; and a bypass valve 2007 is arranged to allow the proportion of oil conveyed through the bypass duct 2005 to be changed. Method 200 includes controlling 202 the heat exchange system in the range of 0 to 0.30, 0 to 0.20, or 0 to 0.10. A controller 58 may be provided to achieve such control.

[0491] It should be understood that even for a particular engine 10 operating on a set fuel, due to varying conditions, there is typically a range of values for this ratio 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.

[0492] Method 200 may also include receiving data 204 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 or the fuel flow path, and / or optionally fuel tank temperature or tank temperature) and fuel flow rate data. Such data may be received 204 by the controller 58 and used 206 to adjust the control 202 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 206 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 204, 206 may alternatively be considered part of the control 202 of the heat exchange system 3000 (i.e., not separate steps). Method 200 may also be arranged to utilize other information when determining 206 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 characteristics.

[0493] The step of controlling 202 the heat exchange system 3000 may include any one or any combination of the examples provided for the control step 104 of Figure 13 method 100, for example, when the heat transfer ratio is too high, reducing the amount of oil conveyed via at least one air-oil heat exchanger 2020 by increasing the amount of oil conveyed via the bypass pipe 2005 or by adjusting the ratio of oil and / or fuel conveyed via the respective bypass pipes 1005 or the recirculation pipe 6011.

[0494] In an embodiment where the engine 10 includes a refrigeration cycle device 1007, the step of controlling 202 the heat exchange system 3000 may include using the refrigeration cycle device 1007 to provide a heat boost by transferring more heat from the oil to the fuel than is transferred through the fuel-oil heat exchangers 1004, 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. For example, 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 an embodiment without the refrigeration cycle device 1007, or in an embodiment where such a device is present but not used, a higher heat transfer ratio value may be maintained during cruise, such as greater than 0.38, and optionally greater than 0.40 or 0.50.

[0495] The fuel temperature upon reaching the burner 16 can also be considered. The method 200 can include controlling 202 the heat exchange system 2000 under cruise conditions such that the heat transfer ratio is in the range of 0 to 0.2, 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, or if the fuel temperature at the inlet of the burner is at least 140 °C, the heat transfer ratio is in the range of 0 to 0.45. Accordingly, one or more temperature sensors 1009 can be used and their outputs can be considered when setting the control 202 of the heat exchange system 2000.

[0496] The method 200 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. 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 202 of the heat exchange system 2000.

[0497] 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 kilogram of fuel.

[0498] 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 200 kJ per kilogram of fuel.

[0499] For a particular direct drive engine 10, under 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 (as opposed to being transferred to the air / environment) is plotted as a function of the fuel temperature limit in Figure 18Among them, the fuel temperature limit is the maximum fuel temperature considered safe for the operation of the aircraft with this fuel and this engine 10, and it 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 the upper limit 5 (black dashed line) of the percentage of heat transferred to the fuel - the upper limit 4 and the lower limit 5 illustrate different cruise fuel flows (depending on altitude, rated power, etc.), the variability of engine heat generation, the aircraft fuel temperature, the atmospheric temperature, and other variables. Generally speaking, for the example engine 10 that provides these data, 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 losses from the oil is transferred to the fuel (100%). Curve 3 also shows line 6 (gray solid line) for typical heat transfer to the fuel, with a fuel temperature limit. For a fuel with a temperature limit of 120 °C, the typical value of the percentage of heat transferred to the fuel can be 60%; for fuel temperature limits of 170 °C and above, this may increase to 100%. When the fuel temperature limit is equal to or greater than 260 °C, throughout the cruise, all the heat can be transferred to the fuel during cruise without air cooling, so the heat transfer ratio is zero.

[0500] Now turning to the method 300 to be executed at idle, the method 300 is again arranged to be executed in a direct - drive gas turbine engine 10, which includes: an oil circuit system 2000 arranged to supply oil to engine components 2006; 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. A modulating valve 2016 is again provided to control the proportion of oil delivered through each branch of the oil circuit system 2000. A bypass pipe 2005 is arranged to allow a certain proportion of the oil to flow through one of the air - oil heat exchanger 2020 or the fuel - oil heat exchanger 1006; and a bypass valve 2007 is arranged to allow the proportion of oil flowing through the bypass pipe 2005 to be changed. The method 300 includes controlling 302 the heat - exchange system 2000 such that when the aircraft 1 is operating under idle conditions, the heat transfer ratio:

[0501]

[0502] Within the range of 0 to 1.5. 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 300 may include maintaining the heat transfer ratio under idle conditions below 1.0, within 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.

[0503] Method 300 may also include receiving data 304 to allow the 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 the oil circuit system 2000, 2000', or fuel flow path, and / or optionally fuel tank temperature or tank temperature) and fuel flow data. Such data may be received 304 by controller 58 and used 206 to adjust the modulation 202 of heat exchange system 2000 at idle so as to maintain the heat transfer ratio within a desired level or desired bounds. Such checking and adjustment / correction 206 may be performed periodically or in response to a predetermined stimulus (e.g., a change in the temperature or flow rate of fuel or oil, or a change in the aircraft operation, such as starting to taxi). These steps 304, 206 may alternatively be considered part of the control 302 of heat exchange system 2000. Method 300 may also be arranged to utilize other information when determining 206 what control action to take, such as temperature data (ambient temperature of oil, fuel, and / or the environment around aircraft 1), flow rate data (oil and / or fuel), and / or one or more fuel characteristics.

[0504] The step of controlling 302 heat exchange system 2000 may include any one or any combination of the examples provided for Figure 14 the modulation step 204 of method 200, e.g., 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 respective bypass pipe 1005 or recirculation pipe 6011.

[0505] In an embodiment where the engine 10 includes a refrigeration cycle device 1007, the step of controlling 302 the heat exchange system 2000 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 0.40, and optionally not exceed 0.30, 0.20, 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 can be maintained, for example in the range of 0.38 to 1.2.

[0506] The fuel temperature upon reaching the burner 16 can also be considered. The method 300 may include controlling 302 the heat exchange system 2000 during idle operation such that if the fuel temperature at the inlet of the burner 16 is below 180 °C, the heat transfer ratio is in the range of 0.3 to 1.5, or if the fuel temperature at the inlet of the burner 16 is at or above 180 °C, the heat transfer ratio is in the range of 0 to 0.3. Accordingly, one or more temperature sensors can be used, and their outputs can be considered when setting the control 202 of the heat exchange system 3000.

[0507] The method 300 can additionally or alternatively consider the properties of the fuel, such as whether the fuel is or includes sustainable aviation fuel (SAF). Accordingly, one or more fuel characteristics - optionally determined by the fuel characteristics 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.

[0508] In some embodiments, no more than 20% of the heat transferred 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 180 kJ per kilogram of fuel.

[0509] In some embodiments, at least 80% of the heat transferred 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 100 kJ to 300 kJ per kilogram of fuel. At idle, a higher percentage of heat dissipation to air is typically used than during cruise and other higher power engine conditions.

[0510] Regarding the reference Figure 14The described method 200, in some embodiments, the heat exchange system 3000 further includes a branched fuel return passage and at least one valve that controls the diversion of fuel flow, and these branched passages are arranged to return fuel from the heat exchange system 3000 to at least one different location along the main fuel path, from which the 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. 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.

[0511] The present inventors have also recognized that using fuels different from traditional kerosene-based jet fuels (such as sustainable aviation fuels) can result in different fuel properties, and the parameters under cruise conditions can be adjusted to utilize the different fuel properties. In particular, some fuels can be heated to higher temperatures in the fuel-oil heat exchangers 1004, 1006 than traditional fuels. This can improve the cooling of the oil before the oil returns to the rest of the turbine engine, and / or can improve the combustion efficiency of the fuel. Using the fuel to absorb more heat from the oil rather than relying on heat transfer from the oil to the environment / air (e.g., in an oil-air heat exchanger) provides a more thermally efficient turbine engine. Additionally, the increased cooling of the oil can in turn improve the cooling effect of the oil on the engine components through which it flows.

[0512] Figure 16 An exemplary method 400 for operating a gas turbine engine 10 is shown. Method 400 includes the following steps:

[0513] Step 402: Supply fuel to the gas turbine engine 10 via the fuel management system 1500.

[0514] As discussed with reference to Figure 6 The fuel management system 1500 includes a primary fuel-oil heat exchanger 1006 and a secondary fuel-oil heat exchanger 1004 that are arranged to transfer heat to or from the fuel, and is part of the heat exchange system 3000. The engine fuel pump 1003 is located downstream of the primary fuel-oil heat exchanger 1004 and upstream of the secondary fuel-oil heat exchanger 1006. The fuel management system 1500 is arranged such that the fuel reaches the primary fuel-oil heat exchanger 1006 before reaching the secondary fuel-oil heat exchanger 1004.

[0515] Step 404: Control the heat exchange system 3000 to raise the fuel temperature when leaving the secondary fuel-oil heat exchanger 1004 to at least 120 °C under cruise conditions.

[0516] The control 404 of the heat exchange system 3000 may include controlling the fuel flow through the heat exchanger - for example, recirculating a variable proportion of the fuel through the primary fuel - oil heat exchanger 1006, and / or allowing a variable proportion of the fuel to bypass the primary fuel - oil heat exchanger 1006. Although Figure 6 an embodiment is shown having a bypass duct 1005 but no recirculation, and Figures 7A to 7C an embodiment is shown having no bypass duct (at least none for the primary heat exchanger - Figure 7A and Figure 7C which does provide some route for the fuel reaching the burner 16 to bypass the secondary heat exchanger while the remaining fuel passes through the secondary heat exchanger) but having a recirculation duct 6011, it should be understood that one or more bypass ducts and / or one or more recirculation ducts may be provided together in various embodiments. Additionally, in some systems, the flow direction within the same duct may be reversible such that it can be used as the bypass duct 1005 or the recirculation duct 6011. It should be understood that the drawings are provided by way of example only and are not intended to be limiting.

[0517] The control 404 of the heat exchange system 3000 may include controlling the fuel flow through the heat exchanger - for example, allowing a variable proportion of the fuel to bypass the primary fuel - oil heat exchanger 1006 or the secondary fuel - oil heat exchanger 1006. Accordingly, one or more controllable valves or pumps, and optionally one or more sensors, may be provided to facilitate fuel flow control.

[0518] Additionally or alternatively, for the fuel flow, one or more bypass valves 2007 may be used to control the oil flow, which are configured to control the flow through one or more bypass ducts 2005 (if present), thereby allowing the oil to bypass one or more of the heat exchangers 1004, 1006, 2020 rather than flowing through them. In some embodiments, the oil may also be recirculated and / or the oil flow may be regulated by controlling one or more oil pumps. Accordingly, one or more controllable valves and / or pumps, and optionally one or more sensors, may be provided to facilitate oil flow control.

[0519] The controller 58 may be used to actively manage the fuel flow and / or oil through and around the heat exchangers 1004, 1006, 2020. 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. The controller may receive inputs from one or more temperature sensors 1009 and may control one or more valves (such as the recirculation valve 6010 or the bypass valve 2007) and / or pumps 1003 based on the received data. One or more fuel supply pumps and / or fuel return pumps may also be controlled by the controller.

[0520] As described above, the recirculation valve 6010 can be used to control the fuel flow. Alternatively or additionally, the fuel flow can be controlled by using one or more bypass conduits 1005 arranged to allow a proportion of the fuel to avoid passing through either or both of the heat exchangers, as Figure 5 shown, and / or by adjusting the percentage of fuel flowing directly from the primary fuel-oil heat exchanger 1006 to the burner 16 as compared to the percentage of fuel flowing from the primary fuel-oil heat exchanger into the secondary fuel-oil heat exchanger 1004 (and then optionally returning to rejoin the fuel flowing into the burner 16, depending on the implementation) to control the fuel flow rate.

[0521] Heating the fuel to a higher temperature than previously used can improve the cooling of the oil before it returns to the rest of the turbine engine, and / or can improve the combustion efficiency of the fuel. Positioning the secondary fuel-oil heat exchanger 1004 after the engine fuel pump 1003 can help achieve a higher fuel temperature without compromising the life of the fuel pump.

[0522] The auxiliary system 7010 supplied by the secondary fuel-oil heat exchanger 1004 can benefit from a fuel temperature higher than that applicable to the fuel in the burner 16 or components through which the main fuel flow path passes. Thus, it may be desirable for the temperature of the fuel leaving the secondary fuel-oil heat exchanger 1004 to be higher than the fuel temperature at the burner 16 inlet. Additionally, as shown in FIG. 8, for example, the oil can flow through the secondary fuel-oil heat exchanger 1004 before flowing through the other heat exchangers 1006, 2020 of the heat exchange system 3000. Thus, the oil temperature may be higher when passing through the secondary fuel-oil heat exchanger 1004 than when passing through the other heat exchangers 1006, 2020. The control 404 of the heat exchange system 3000 can include controlling a fuel bypass valve to adjust the proportion of fuel delivered via the fuel bypass conduit 1005 and also controlling the fuel flow through each of the primary fuel-oil heat exchanger 1006 and the secondary fuel-oil heat exchanger 1004 under cruise conditions.

[0523] The control 404 of the heat exchange system can include controlling an oil bypass valve 2007 to adjust the proportion of oil delivered via the oil bypass conduit 2005 and also controlling the oil flow through each of the primary fuel-oil heat exchanger 1006 and the secondary fuel-oil heat exchanger 1004 under cruise conditions.

[0524] The control 404 of the heat exchange system 3000 may include controlling the heat exchange system 3000 to raise the temperature of the fuel such that the fuel temperature upon leaving the secondary heat exchanger is in the range of 120 °C to 180 °C, in the range of 135 °C to 200 °C, 135 °C to 180 °C, 150 °C to 200 °C, 150 °C to 180 °C, 150 °C to 170 °C, or is about 120 °C, 130 °C, 140 °C, 150 °C, 160 °C or up to 200 °C.

[0525] However, the inventors have recognized that allowing inappropriate fuel to reach these higher temperatures can be detrimental to engine performance and even potentially dangerous - excessive thermal decomposition of the fuel / sediments can cause safety issues by blocking passages and causing component failures (e.g., valve sticking and blocked nozzles, orifices, valves, etc.), which can lead to thrust runaway. Therefore, Figure 17 the method 500 shown is provided as a safety precaution.

[0526] Figure 17The method 500 shown in [description] can be implemented in a gas turbine engine 10, substantially as described in the foregoing method regarding oil flow, but one or more controllable oil valves 2007, 2007a (arranged to allow changing the proportion of oil bypassing at least one of the air-oil heat exchanger 2020, the primary fuel-oil heat exchanger 1006, and the secondary fuel-oil heat exchanger (and conveyed therethrough by extension)) are replaced by at least one controllable oil or air valve 2007, 2007a, 2007b, 2022, which is arranged to allow changing at least one of the oil flow rate and the air flow rate through at least one of the air-oil heat exchanger 2020 and the fuel-oil heat exchanger 1006. Thus, air flow control or oil flow control, or both, can be provided. In embodiments providing a controllable air valve 2022 (optionally also controllable oil valves 2007, 2007a, 2007b), the engine may require a temperature sensor 1009, which is arranged to optionally provide an indication of the fuel temperature at the inlet of the burner 16. More generally, the sensor 1009 can be located anywhere downstream of the primary fuel-oil heat exchanger 1006 on the fuel side. The temperature sensor 1009 is typically located at or near the inlet of the burner 16, and more specifically near the nozzle inlet of the burner 16, so as to directly sense the fuel temperature at the burner inlet. However, it should be understood that if the secondary fuel-oil heat exchanger 1004 is bypassed, the change in fuel temperature between the outlet of the primary fuel-oil heat exchanger 1006 and the burner 16 may be small, and if the fuel temperature at a location different from the sensor 1009 is required, the fuel temperature change can be calculated based on the knowledge of the engine 10. In some embodiments using the fuel temperature at the inlet of the burner 16, the sensor 1009 can be located at a different location, and the fuel temperature at the burner inlet can be calculated based on the temperature output and the knowledge of the engine 10.

[0527] The method 500 includes determining 502 whether the fuel temperature has increased above a set threshold under cruise conditions based on the output of the temperature sensor 1009. For example, a direct comparison can be made between the data received from the temperature sensor 1009 and one or more temperature thresholds stored in the memory. Thus, the determination 502 can be performed automatically, optionally by a controller 58 (which can be or include a dedicated processing module, or can be provided by a more general EEC).

[0528] Method 500 further includes, in response to determining that the fuel temperature has increased above a set threshold under cruise conditions, controlling 504 at least one of valves 2007, 2007a, 2022 so as to appropriately vary at least one flow rate (air or oil flow rate) through at least one of heat exchangers 1004, 1006, 2020 to reduce the fuel temperature. For example, adjusting oil valves 2007, 2007a can deliver a greater proportion of the oil via bypass pipes 2005, 2005a to pass through primary fuel-oil heat exchanger 1006 and / or secondary fuel-oil heat exchanger 1004, and recombine with the remaining oil passing through fuel-oil heat exchangers 1004, 1006 at or downstream of the oil outlets of heat exchangers 1006, 1004, as Figure 12 shown. In a parallel configuration (not shown), reducing the oil flow through fuel-oil heat exchanger 1006 by controlling the modulating valve can automatically increase the oil flow through air-oil heat exchanger 2020 (unless a bypass pipe around air-oil heat exchanger 2020 is present and used to compensate).

[0529] As Figure 17 shown by the dashed lines in, the method 500 can be repeated - for safety reasons, optionally according to a regular schedule and / or when triggered by changes in engine operation (e.g., changes in altitude or thrust demand), and frequent checks may be beneficial. Thus, the oil valves 2007, 2007a or the air valve 2022 can be actively controlled 504 to vary the oil or air flow rate through the respective heat exchangers 1006, 2020. The active control of valves 2007, 2007a, 2022 can be automatic and implemented by controller 58. For example, the active control of bypass valve 2007 associated with secondary fuel-oil heat exchanger 1004 can be automatic and implemented by controller 58 of 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 one or more of valves 2007, 2007a, 2022 can be closed-loop or open-loop - in particular, when feedback data (e.g., from an oil flow sensor) is available, a closed-loop control method may be preferred.

[0530] In an embodiment having an air valve 2022 and oil valves 2007, 2007a, when more oil is conveyed via the air-oil heat exchanger 2020, the air flow through the air-oil heat exchanger 2020 can be increased; with the increase in oil flow rate, this increase can be linear. In terms of the percentage of air flow, the increase can be significant - for example, from the air valve 2022 being fully closed to fully open, or when the fuel temperature is considered too high, from a first position of 10% of the maximum air flow rate to 90% or 100% of the maximum air flow rate. The method 500 can include, in response to determining 502 that the fuel temperature has increased above a set threshold under cruise conditions, controlling 504a the air valve 2022 to convey more air through the air-oil heat exchanger 2020, and controlling 504b the oil valves 2016, 2007 to convey less oil via the fuel-oil heat exchanger 1006. The control / regulation 504a of the air valve 2022 can be performed simultaneously with the control / regulation 504b of the oil valves 2016, 2007.

[0531] In some embodiments, the air flow through the air-oil heat exchanger 2020 may not be actively controlled. In such embodiments, based on the pressure ratio between the heat exchanger inlet 2020a and the heat exchanger exhaust / outlet 2020b, the air flow may vary naturally under different operating conditions. In this configuration, even when the air-oil heat exchanger 2020 is completely bypassed on the oil side / when the oil flow rate through the heat exchanger 2020 is zero, air will still flow through the air channels of the air-oil heat exchanger 2020. In an alternative embodiment, active control of the air flow through the air-oil heat exchanger 2020 is provided. For example, the air-oil heat exchanger 2020 and the fuel-oil heat exchangers 1004, 1006 can be in series, with no bypass on the oil side of the air-oil heat exchanger 2020, but an air valve 2022 for controlling the air flow on the air side (in other embodiments, an oil bypass valve 2007b and an air flow valve 2022 can be provided). In such a series arrangement structure, the air-oil heat exchanger 2020 can be upstream or downstream of the fuel-oil heat exchanger 1006 relative to the oil flow. In an embodiment of the series arrangement, the air-oil heat exchanger 2020 is located upstream of the fuel-oil heat exchangers 1004, 1006 relative to the oil flow (different from Figure 12) can help prevent the fuel from overheating - the oil can be cooled as much as possible by increasing the air flow to a predetermined threshold. It will be appreciated that in the absence of a dedicated mechanism for actively driving the air flow (such as a gas-gas injector or a dedicated fan), there are limitations on how much oil can be cooled in a single pass through the air-oil heat exchanger (i.e., once the air valve 2022 is fully open, no further control adjustment can be made to increase the cooling in a single pass through the air-oil heat exchanger 2020). The oil can be further cooled by recirculating the oil through the air-oil heat exchanger 2020 before it reaches the fuel-oil heat exchanger 1006.

[0532] In some embodiments, the oil bypass conduit 2005b and the corresponding control valve 2007b can be implemented only for the air-oil heat exchanger 2020. Optionally, in the described method 500, only the oil flow through the air-oil heat exchanger 2020 can be actively controlled.

[0533] In various embodiments:

[0534] - The bypass conduit 2005b is provided only on the oil side of the air-oil heat exchanger 2020 (no bypass of the oil of the fuel-oil heat exchanger and no active air flow control);

[0535] - Bypass conduits 2005, 2005a, 2005b (no active air flow control) can be provided for the air-oil heat exchanger 2020, the primary fuel-oil heat exchanger 1006, and the secondary fuel-oil heat exchanger 1004; or

[0536] - The air flow through the air-oil heat exchanger 2020 can be controllable and there can be no oil bypass conduit.

[0537] Generally speaking, controlling at least one of the flows through the air-oil heat exchanger 2020 can be considered more important than controlling either of the flows through the fuel-oil heat exchangers 1004, 1006. However, in some embodiments, one or more oil bypass conduits 2005, 2005a and / or fuel bypass conduits 1005 (and corresponding control valves) can be provided for the fuel-oil heat exchangers 1004, 1006, optionally along with one or more additional controllable valves.

[0538] A combination of air flow control and oil flow control can be achieved in both parallel and series arrangements. However, it should be understood that while all the various control options can be used together, an increase in complexity and the number of components and thus an increase in the weight of the entire heat exchange system 3000 may be undesirable. Thus, in many embodiments, it may be preferable to select a subset of control options - for example, controlling the air flow or controlling the oil flow can be used to regulate heat transfer within the air - oil heat exchanger 2020, and thus using both controls for a given heat exchanger 2020 may be considered redundant. Similarly, in embodiments where the oil flows through the air - oil heat exchanger 2020 before the fuel - oil heat exchanger 1006, the regulation of the oil temperature in the air - oil heat exchanger 2020 (by controlling the air and / or oil flow in that heat exchanger) can be used as an alternative to regulating the oil flow through one or more of the primary fuel - oil heat exchanger 1006 or the secondary fuel - oil heat exchanger 1004 to vary the heat transferred to the fuel in the fuel - oil heat exchangers 1004, 1006.

[0539] Additionally or alternatively, in embodiments in which the heat exchange system 3000 includes a refrigeration cycle device 1007, the refrigeration cycle device is arranged to transfer additional heat from the oil to the fuel in addition to the heat transferred by the primary fuel - oil heat exchanger 1006 and the secondary fuel - oil heat exchanger 1004. The method 500 may further include, in response to determining 502 that the fuel temperature has increased above a set threshold under cruise conditions, controlling the refrigeration cycle device 1007 to reduce the additional heat transferred to the fuel. This control can be performed by reducing the oil flow through the refrigeration cycle device 1007, or by reducing the power of the refrigeration cycle device 1007 (e.g., by reducing the flow rate of the refrigerant fluid), or by deactivating the refrigeration cycle device 1007. The set threshold for deactivation of the refrigeration cycle device 1007 can be lower than the set threshold for changing the flow rate of the oil and / or air through the heat exchangers 1006, 2020.

[0540] The set threshold for controlling step 504 can be in the range of 140°C to 300°C, and optionally in the range of 200°C to 300°C, and further optionally in the range of 250°C to 300°C. For example, the set threshold can be 140°C, 150°C, 160°C, 170°C, 180°C, 190°C, 200°C, 210°C, 220°C, 230°C, 240°C, 250°C, 260°C, 270°C, 280°C, 290°C or 300°C.

[0541] The method 500 of some embodiments further includes determining 501 a set threshold, to which the output of the temperature sensor is compared. This determination 501 can be performed on-wing and optionally in-flight. The determination 501 can be or include calculating the threshold or identifying a suitable threshold from a pre-stored set of thresholds stored in a memory.

[0542] The inventors have recognized that knowledge of one or more fuel properties can be used in the determination 501 to ensure safety while still taking advantage of the varying nature between aviation fuels. Thus, the determination 501 can be based on at least one fuel property of the fuel. The at least one fuel property can be or include any of the above examples, such as the thermal stability of the fuel, the nitrogen content of the fuel, the sulfur content of the fuel, and / or the sustainable aviation fuel (SAF) content (% SAF) of the fuel. The at least one fuel property can be determined by any one or more of the above methods.

[0543] The step 501 of determining the set threshold can include optionally linearly increasing the set threshold as the fuel thermal stability increases. Thus, the set threshold for controlling the thermal modulation can increase linearly with the fuel thermal stability. 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 change at a specific temperature; aviation fuels include various components that are prone to decomposition at different temperatures, and the time spent at high temperatures is also a factor in fuel decomposition. The threshold can be set based on a comparison of the maximum allowable deposition rate of fuel decomposition with the thermal stability of the fuel in use. 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. The method 500 can also include using a clock or timer, or a fuel flow sensor—for example, at higher pumping rates, a higher fuel temperature is acceptable because faster-moving fuel spends less time at that temperature before combustion. Alternatively or additionally, the method 500 can include using one or more pressure sensors to determine the fuel flow. The flow rate can be determined by measuring the pressure drop across a known orifice.

[0544] Alternatively or additionally, for fuels with a SAF content higher than 70%, the step of determining 501 the set threshold can include increasing the set threshold as the SAF content of the fuel increases (optionally linearly). As mentioned above, it has been found that an increase in % SAF blended with fossil-based aviation fuel increases thermal stability, but in a non-linear manner. Thus, pre-set, discrete thresholds based on % SAF may be advantageous rather than continuously changing the set threshold proportionally to the SAF content, or a more complex (non-linear) but continuous relationship can be established.

[0545] Alternatively or additionally, the step of determining the set threshold 501 may include decreasing the set threshold as the content of heteroatom species in the fuel increases (optionally linearly). For example, it has been found that the thermal stability of jet A decreases as the nitrogen content increases (the nitrogen content is a measure of the amount of nitrogenous substances present). It is well known that the interaction between sulfurous and nitrogenous substances in the fuel can be an important contributor to the fuel decomposition rate, and thus consideration of the fuel composition with respect to multiple heteroatom species can be implemented to account for these interactions.

[0546] In some embodiments, particularly those in which a direct measure of the thermal stability of the fuel is not available, multiple fuel properties may be determined and used in the decision-making for valve control. One or more of the determined fuel properties can be effectively translated into a measure of thermal stability.

[0547] It should be understood that the present invention is not limited to the above-described 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 of the 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; Fan shaft; at least one bearing arranged to support the fan shaft; at least one auxiliary system arranged to use some of said fuel; an oil circuit system arranged to supply oil to the at least one bearing; as well as A heat exchange system, the heat exchange system comprising: a primary fuel-oil heat exchanger through which the oil and at least substantially all of the fuel in the oil circuit system flow such that heat is transferred between the oil and the fuel; and a secondary fuel-oil heat exchanger through which a portion of the oil and the fuel in the oil circuit system flows so that heat is transferred between the oil and the fuel, the secondary fuel-oil heat exchanger being arranged to supply fuel to the at least one auxiliary system; as well as a fuel pump arranged to pump the fuel, wherein the fuel pump is located downstream of the primary fuel-oil heat exchanger and upstream of the secondary fuel-oil heat exchanger along a fuel flow path; The method further comprises controlling the heat exchange system such that under cruising conditions, the temperature of the fuel exiting the second heat exchanger is in the range of 120°C to 200°C.

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

3. The method according to claim 1, wherein: The method comprises controlling the heat exchange system such that under cruising conditions the temperature of the fuel exiting the second heat exchanger is in the range of 135°C to 200°C.

4. The method according to claim 1, wherein: The method comprises controlling the heat exchange system such that under cruising conditions the temperature of the fuel exiting the second heat exchanger is in the range of 135°C to 180°C.

5. The method according to claim 1, wherein: The heat exchange system also includes: a fuel bypass conduit arranged to allow a portion of the fuel to bypass at least one of the primary fuel-oil heat exchanger and the secondary fuel-oil heat exchanger; and a fuel bypass valve arranged to allow a proportion of the fuel delivered via the fuel bypass conduit to be varied; Wherein controlling the heat exchange system includes controlling the fuel bypass valve to adjust a proportion of the fuel delivered through each of the primary fuel-oil heat exchanger and the secondary fuel-oil heat exchanger under cruise conditions.

6. The method according to claim 1, wherein: The heat exchange system also includes: an oil bypass conduit arranged to allow a portion of the oil to bypass at least one of the primary fuel-oil heat exchanger and the secondary fuel-oil heat exchanger; and an oil bypass valve arranged to allow a proportion of the fuel delivered via the fuel bypass conduit to be varied; Wherein controlling the heat exchange system includes controlling the oil bypass valve to adjust a proportion of the oil delivered through each of the primary fuel-oil heat exchanger and the secondary fuel-oil heat exchanger under cruise conditions.

7. The method according to claim 1, wherein: Under cruising conditions, the following The ratio is at least 0.

3.

8. The method according to claim 1, wherein: Under cruising conditions, the following The ratio is at least 0.

5.

9. The method according to claim 1, wherein: Under cruising conditions, the following The ratio is about 0.

6.

10. 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.

11. The method according to claim 1, wherein: The primary fuel-oil heat exchanger and the secondary fuel-oil heat exchanger are arranged in series along the oil circuit system.

12. The method according to claim 1, wherein: The heat exchange system further comprises at least one air-oil heat exchanger, through which the oil of the oil circuit system flows.

13. 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 gear box receiving input from the spindle and outputting a driving force to the fan so as to drive the fan at a lower 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: a primary fuel-oil heat exchanger through which the oil and at least substantially all of the fuel in the oil circuit system flow such that heat is transferred between the oil and the fuel; and a secondary fuel-oil heat exchanger through which a portion of the oil and the fuel in the oil circuit system flows so that heat is transferred between the oil and the fuel, the secondary fuel-oil heat exchanger being arranged to supply fuel to the at least one auxiliary system; and a fuel pump arranged to pump the fuel, wherein the fuel pump is located downstream of the primary fuel-oil heat exchanger and upstream of the secondary fuel-oil heat exchanger along a fuel flow path; Wherein the heat exchange system is arranged to be controlled such that under cruising conditions the temperature of the fuel when leaving the second heat exchanger is in the range of 120°C to 200°C.

14. The gas turbine engine according to claim 13, 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.

15. The gas turbine engine of claim 13, wherein: The heat exchange system is arranged to be controlled such that under cruising conditions the fuel temperature upon exiting the second heat exchanger is in the range of 120°C to 180°C.

16. The gas turbine engine of claim 13, wherein: The spindle outputs driving force directly to the fan so as to drive the fan at the same rotational speed as the spindle, so that the engine is a direct drive turbine engine.

Images