FR3144225A1 - Fueling system - Google Patents

Abstract

A method (11000) of operating a gas turbine engine (10) is disclosed, the gas turbine engine (10) comprising a combustion chamber (16) arranged to burn a fuel; and a fuel management system arranged to supply fuel to the combustion chamber (16). The fuel management system includes two fuel-oil heat exchangers through which oil and fuel flow, the heat exchangers arranged to transfer heat to the fuel and including a primary fuel-oil heat exchanger (1004) and a secondary fuel-oil heat exchanger (1006); and a fuel pump (1003) arranged to distribute fuel to the combustion chamber (16), wherein the fuel pump (1003) is located between the two heat exchangers. The method includes controlling (11200) the fuel management system to raise the fuel temperature to at least 135°C at the inlet of the combustion chamber (16) at cruise conditions. Figure for abstract: Fig. 6A

Description

Fueling system

The present description relates to aircraft propulsion systems, as well as methods of operating aircraft involving the management of different fluids and heat transfer between them, and in particular the management of fuel properties at the entrance to the combustion chamber.

There is an expectation in the aviation industry that there will be a trend towards the use of fuels that are different from the traditional kerosene-based fuels commonly used today. The fuels may have different fuel characteristics compared to petroleum-based hydrocarbon fuels.

Thus, there is a need to consider fuel properties for these new fuels and to adjust the operating processes of gas turbine engines.

According to a first aspect, a method of operating a gas turbine engine is provided, the gas turbine engine comprising:

a combustion chamber arranged to burn a fuel; and

a fuel management system arranged to supply fuel to the combustion chamber, wherein the fuel management system comprises:

two fuel-oil heat exchangers through which oil and fuel flow, the heat exchangers being arranged to transfer heat between the oil and the fuel and comprising a primary fuel-oil heat exchanger and a secondary fuel-oil heat exchanger;

a fuel pump arranged to deliver fuel to the combustion chamber, wherein the fuel pump is located between the two fuel-oil heat exchangers; and

a recirculation valve located downstream of the primary fuel-oil heat exchanger, the recirculation valve arranged to allow a controlled quantity of fuel that has passed through the primary fuel-oil heat exchanger to be returned to an inlet of the primary fuel-oil heat exchanger;

wherein the method comprises selecting fuel such that the calorific value of the fuel supplied to the gas turbine engine under cruise conditions is at least 43.5 MJ/kg.

The primary and secondary fuel-oil heat exchangers may be referred to herein as the primary heat exchanger and the secondary heat exchanger for the sake of brevity.

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

It will be appreciated that, herein, when the fuel pump is described as being located "between" the two fuel-oil heat exchangers, it is meant that the fuel pump is located in the fuel path between the primary fuel-oil heat exchanger and the secondary fuel-oil heat exchanger, such that fuel flows through the pump between the outlet of the primary fuel-oil heat exchanger and the inlet to the secondary fuel-oil heat exchanger (or vice versa ), regardless of whether a straight line between the primary fuel-oil heat exchanger and the secondary fuel-oil heat exchanger would pass through the pump. It will be appreciated that exact relative physical positions of the pump and the heat exchangers are not significant.

The recirculation valve may be located downstream of the fuel pump, and may direct a proportion of the flow exiting the pump back to a point in the flow path before the pump inlet. This may provide additional flexibility in fuel flow to the combustion chamber (for a given pumping rate), in addition to the heat exchange flexibility provided by the recirculation valve.

A higher calorific value of a fuel also reduces the fuel burn rate required for a given shaft speed at cruise conditions. Particularly for engine architectures in which the fuel pump speed is directly related to the shaft speed (e.g., in which a ratio of shaft speed to pump speed is fixed, or has only a set of specific discrete values, at least at cruise conditions), fuel can be wasted if the engine operation is not adjusted. The inventors were aware that more fuel recirculation can therefore be achieved, by taking fuel that has left the pump back to the pump inlet such that a fuel flow rate to the combustion chamber can be lowered while maintaining a constant fuel flow rate through the pump. More recirculation can therefore be achieved/a higher percentage of the fuel leaving the pump can be directed back to the pump inlet for higher calorific value fuels. The above fuel management system arrangement, with fuel recirculation occurring before the fuel enters the secondary heat exchanger (which is downstream of the pump), can therefore provide improved flexibility in fuel flow without introducing warmer fuel through the pump, which could damage the pump. Additionally, or alternatively in examples where the recirculation valve is downstream of the primary heat exchanger but upstream of the pump, recirculation of a portion of the fuel through the primary heat exchanger can improve, and/or provide more flexibility in, heat transfer arrangements and oil cooling.

At least 75%, 80%, or 82% of the fuel flow can be recirculated through the recirculation valve at cruise conditions.

For a given fuel and engine, the % recirculated can depend on the operating condition of the engine (e.g. cruise, idle, descent, or maximum takeoff) and also on the life of the pump (i.e. an older, deteriorated fuel pump would generally provide less flow than a new one and therefore have a reduced % recirculated). Fuel temperature can also affect the % recirculated. The values given above assume a reasonably new/performing pump and standard fuel temperatures for the given engine at cruise conditions.

Even at cruise conditions, power requirements may vary (e.g. depending on altitude, cruise phase, and fuel mass on board). The recirculation percentage may therefore be varied depending on the current power requirements, and hence the fuel burn rate requirements, of the engine.

The fuel management system may further include a bypass line arranged to allow a proportion of the fuel to bypass at least one of the primary heat exchanger and the secondary heat exchanger. A controllable bypass valve may be used to control the percentage of fuel passing therethrough. For a given fuel and engine, the percentage of fuel that bypasses a heat exchanger may depend on the operating condition of the engine (e.g., cruise, idle, descent, or maximum takeoff) and also on the fuel and/or oil temperature, for example. One or more temperature sensors may be provided accordingly.

A fuel metering valve (FMV) may also be provided to control the fuel delivered to the combustion chamber. The position of the valve may be monitored, and a relationship between the valve position and the fuel flow rate may be known such that an electronic engine controller (EEC) can determine, or “meter,” the fuel flow rate based on the valve position.

The FMV and the recirculation valve may work together to deliver a required fuel flow to the combustion chamber. The position of the FMV may be actively controlled in a closed loop, with excess fuel returned to an earlier portion of the fuel flow path, for example before an engine pump, via the recirculation valve. The method may therefore include using the FMV and a recirculation line to control the fuel flow to the combustion chamber.

The gas turbine engine may therefore include a fuel metering valve (optionally downstream of a recirculation valve, along the main fuel flow path through the engine to the combustion chamber), the fuel metering valve being arranged to control the fuel flow therethrough and to provide information about the fuel flow therethrough (e.g., as a function of valve position).

Obtaining/selecting fuel may include selecting a single fuel. Selecting fuel may include selecting a fuel blend. The selected fuel may be the only fuel on board the aircraft. As such, the fuel selection may be made while the aircraft is being refueled. Alternatively, the selected fuel may be one of multiple fuels on board the aircraft, or a blend of multiple fuels on board the aircraft; thus, the selection may be made in flight.

The recirculation valve may be located downstream of an inlet to the secondary heat exchanger. The recirculation valve may therefore have no influence on the proportion of fuel that is passed through the secondary heat exchanger (on the current pass) in such examples – this division may be fixed at a defined level, or may be controlled separately. Alternatively, the recirculation valve may be located upstream of the inlet to the secondary heat exchanger. The recirculation valve may therefore influence a maximum quantity of fuel (on the current pass) that may flow to the secondary heat exchanger.

The recirculation valve may be located upstream of the secondary heat exchanger, so that fuel that has passed through the secondary heat exchanger may not be recirculated, and thus the pump is never exposed to the higher fuel temperature exiting the secondary heat exchanger (because the pump is upstream of the secondary heat exchanger). The pump may therefore be protected from the higher fuel temperatures exiting the secondary heat exchanger.

The calorific value of the fuel supplied to the gas turbine engine may be between 43.5 MJ/kg and 44 MJ/kg. The calorific value of the fuel supplied to the gas turbine engine may be between 43.8 MJ/kg and 44 MJ/kg.

The calorific value of the fuel supplied to the gas turbine engine may be approximately 43.5 MJ/kg, 43.6 MJ/kg, 43.7 MJ/kg, 43.8 MJ/kg, 43.9 MJ/kg, or 44 MJ/kg.

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

a combustion chamber arranged to burn a fuel; and

a fuel management system arranged to supply fuel to the combustion chamber, wherein the fuel management system comprises:

two fuel-oil heat exchangers arranged for oil and fuel to flow therethrough, the fuel-oil heat exchangers being arranged to transfer heat between the oil and the fuel and comprising a primary fuel-oil heat exchanger and a secondary fuel-oil heat exchanger;

a fuel pump arranged to deliver fuel to the combustion chamber, wherein the fuel pump is located between the two fuel-oil heat exchangers; and

a recirculation valve located downstream of the primary fuel-oil heat exchanger, the recirculation valve arranged to allow a controlled quantity of fuel that has passed through the primary fuel-oil heat exchanger to be returned to an inlet of the primary fuel-oil heat exchanger.

The calorific value of the fuel supplied to the gas turbine engine is at least 43.5 MJ/kg.

The gas turbine engine may further include an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor, in addition to the combustor. The gas turbine engine may further include a fan located upstream of the engine core, wherein the fan is driven by the core shaft.

The gas turbine engine can be a direct drive engine.

The gas turbine engine may further include an accessory gearbox and a pump shaft, wherein the accessory gearbox receives an input from the core shaft and outputs a drive to the fuel pump via the pump shaft. The pump speed may thus be related to the rotational speed of the core shaft, and may have, by means of the gearbox, a fixed speed relationship between the core shaft (or in some engine architectures, another intermediate shaft) and the fuel pump. In some implementations, the accessory gearbox may instead be arranged to provide a fixed number of different speeds for a given core shaft speed. Fuel recirculation may allow for a greater variety of fuel flow rates into the combustion chamber for the given number of pump speeds available at a given shaft speed. Fuel recirculation can allow for multiple different fuel flow rates into the combustion chamber at a given core shaft speed in implementations where there is a single fixed speed relationship between the core shaft and the fuel pump.

The gas turbine engine may further include engine bearing chambers (e.g., shaft bearings or other rotating components). These engine bearing chambers may be cooled using the oil flow prior to the oil flowing through the fuel-oil heat exchangers. The oil exiting the engine bearing chambers may be warmer than the oil entering the engine bearing chambers. The oil flow path may be a loop, such that the oil is cooled by the fuel-oil heat exchangers before entering the engine bearing chambers, and then the warmer oil exiting the engine bearing chambers is recirculated to the fuel-oil heat exchangers.

It is envisaged that all the features described for the first aspect may also apply to the second aspect.

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

a combustion chamber arranged to burn a fuel; and

a fuel management system arranged to supply fuel to the combustion chamber, wherein the fuel management system comprises:

two fuel-oil heat exchangers through which oil and fuel flow, the heat exchangers being arranged to transfer heat between the oil and the fuel and comprising a primary fuel-oil heat exchanger and a secondary fuel-oil heat exchanger; and

a fuel pump arranged to deliver fuel to the combustion chamber, wherein the fuel pump is located between the two heat exchangers;

wherein the method comprises controlling the fuel management system so as to raise the fuel temperature to at least 135°C at the combustion chamber inlet at cruise conditions.

The inventors were aware that the use of fuels other than traditional kerosene-based fuels, such as sustainable aviation fuels, can lead to different fuel properties, and that the cruise parameters can be adjusted to utilize the different fuel properties. In particular, some fuels can be heated to higher temperatures in fuel-oil heat exchangers than traditional fuels. This can improve the cooling of the oil before it is returned to the rest of the turbine engine, and/or improve the efficiency of fuel combustion, for example by affecting the size and distribution of droplets from fuel injection nozzles. 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, improved oil cooling can in turn improve the cooling effect of the oil on the engine components through which it flows.

The fuel management system can be controlled to raise the fuel temperature to between 135°C and 170°C at the combustion chamber entrance at cruise conditions.

The fuel management system can be controlled to raise the fuel temperature to between 150°C and 170°C at the combustion chamber entrance at cruising conditions.

The fuel management system may be controlled to raise the fuel temperature to between 135 and 150°C, 135 and 160°C, 135 and 170°C, 135 and 180°C, 135 and 190°C, or 135 and 200°C at the inlet to the combustion chamber at cruise conditions. The fuel management system may be controlled to raise the fuel temperature to an average of at least 140°C, 150°C, 160°C, 170°C, 180°C, 190°C, or 200°C at the inlet to the combustion chamber at cruise conditions.

Combustion chamber inlet fuel temperatures at cruise conditions may be defined as an average over at least 5 minutes, 10 minutes, or 30 minutes of steady-state cruise conditions. These average temperatures do not include transient temperature spikes, which may be defined as fluctuations in operating fuel temperature, often a temperature rise. Each fluctuation may last no longer than 5 minutes.

The fuel management system may further include a recirculation valve located downstream of the primary heat exchanger, the recirculation valve arranged to allow a controlled amount of fuel that has passed through the primary heat exchanger to be returned to an inlet of the primary heat exchanger. The recirculation valve may also be located downstream of the fuel pump.

The control step of the fuel management system may include controlling the proportion of fuel returned to the primary heat exchanger inlet via the recirculation valve.

A fuel metering valve (FMV) may also be provided to control the fuel delivered to the combustion chamber. The position of the valve may be monitored, and a relationship between the valve position and the fuel flow rate may be known such that an electronic engine controller (EEC) can determine, or “meter,” the fuel flow rate based on the valve position.

The FMV and the recirculation valve may work together to deliver a required fuel flow to the combustion chamber. The position of the FMV may be actively controlled in a closed loop, with excess fuel returned to an earlier portion of the fuel flow path, for example before an engine pump, via the recirculation valve. The method may therefore include using the FMV and a recirculation line to control the fuel flow to the combustion chamber.

The gas turbine engine may include a fuel metering valve (optionally located downstream of a recirculation valve, along the main fuel flow path through the engine to the combustion chamber), the fuel metering valve being arranged to control the fuel flow therethrough and to provide information about the fuel flow therethrough (e.g., as a function of the valve position). The control step of the fuel management system may include controlling the fuel metering valve and the recirculation valve as a function of information provided by the fuel metering valve (and optionally also as a function of other information, e.g. temperature sensor data).

The fuel management system may further include a bypass line arranged to allow a proportion of the fuel to bypass at least one of the primary heat exchanger and the secondary heat exchanger. One or more valves may be provided to facilitate this control.

The control step of the fuel management system may include controlling the proportion of fuel that passes through the bypass line instead of through the primary heat exchanger.

The fuel management system may further comprise an oil bypass line arranged to allow a proportion of the oil to bypass at least one of the primary heat exchanger and the secondary heat exchanger, and optionally may comprise an oil bypass line for each heat exchanger. One or more valves may be provided to facilitate control of a proportion of oil that bypasses the or each heat exchanger.

The control step of the fuel management system may include controlling the proportion of oil that passes through the oil bypass line instead of through the at least one heat exchanger, for example by adjusting a controllable valve.

The control step of the fuel management system may include controlling the proportion of fuel that is recirculated to an earlier point in the fuel flow path, for example by adjusting the position of a fuel metering valve as described above.

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

a combustion chamber arranged to burn a fuel; and

a fuel management system arranged to supply fuel to the combustion chamber, wherein the fuel management system comprises:

two fuel-oil heat exchangers arranged for oil and fuel to flow therethrough, the heat exchangers arranged to transfer heat between the oil and the fuel and comprising a primary fuel-oil heat exchanger and a secondary fuel-oil heat exchanger; and

a fuel pump arranged to deliver fuel to the combustion chamber, wherein the fuel pump is located between the two fuel-oil heat exchangers;

wherein the fuel management system is arranged to raise the fuel temperature to at least 135°C upon entering the combustion chamber under cruising conditions.

It is envisaged that all the features described for the third aspect may also apply to the fourth aspect.

Similarly, the third or fourth aspect can be implemented in conjunction with the first or second aspect.

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

a combustion chamber arranged to burn a fuel; and

a fuel management system arranged to supply fuel to the combustion chamber, wherein the fuel management system comprises:

two fuel-oil heat exchangers through which oil and fuel flow, the heat exchangers being arranged to transfer heat between the oil and the fuel and comprising a primary fuel-oil heat exchanger and a secondary fuel-oil heat exchanger; and

a fuel pump arranged to deliver fuel to the combustion chamber, wherein the fuel pump is located between the two heat exchangers;

wherein the method comprises controlling the fuel management system so as to adjust the fuel viscosity to less than 0.58 mm ² /s (ie less than 0.58 cSt) at the combustion chamber inlet at cruise conditions.

The inventors were aware that, particularly with the use of fuels other than traditional kerosene-based fuels, such as sustainable aviation fuels, fuel properties at cruise conditions can be controlled by careful engine management to improve performance. In this case, a lower viscosity can be achieved, which can affect combustion efficiency, particularly with respect to the spray performance of the fuel nozzle within the combustion chamber. The fuel nozzle spray performance affects the combustion efficiency of the fuel. A lower viscosity of the fuel at cruise conditions can be more conducive to a more efficient engine. Fuel flow can be optimized to improve turbine engine efficiency, to which the lower viscosity of said fuel can contribute. For example, fuel viscosity affects how the fuel is delivered into and ignited by the combustion chamber. Viscosity can affect the droplet size and spray distribution from fuel spray nozzles, which in turn can impact combustion efficiency. Taking into account fuel viscosity when delivering fuel to the combustion chamber, and controlling it appropriately by varying the heat input, can therefore provide more efficient fuel combustion, improving aircraft performance.

Fuel may flow through the primary heat exchanger before reaching the secondary heat exchanger. At least substantially all of the fuel may pass through the primary heat exchanger. Only a smaller proportion of the fuel may pass through the secondary heat exchanger. Fuel that passes through the secondary heat exchanger may be returned to a fuel tank, possibly after use in one or more auxiliary systems, instead of passing to the combustion chamber.

The fuel management system can be controlled to adjust the fuel viscosity between 0.35 mm ² /s and 0.53 mm ² /s at the combustion chamber entrance at cruise conditions.

The fuel management system can be controlled to adjust the fuel viscosity between 0.4 mm ² /s and 0.48 mm ² /s at the combustion chamber entrance at cruise conditions.

The fuel management system may be controlled to adjust the fuel viscosity to less than or equal to 0.50 mm ² /s, 0.48 mm ² /s, 0.46 mm ² /s, 0.44 mm ² /s, or 0.42 mm ² /s at the combustion chamber inlet at cruise conditions. The fuel management system may be controlled to adjust the fuel viscosity to less than 0.50 mm ² /s, 0.48 mm ² /s, 0.46 mm ² /s, 0.44 mm ² /s, or 0.42 mm ² /s at the combustion chamber inlet at cruise conditions.

The fuel management system may further comprise a recirculation valve located downstream of the primary heat exchanger, the recirculation valve arranged to allow a controlled amount of fuel that has passed through the primary heat exchanger to be returned to an inlet of the primary heat exchanger. The step of controlling the fuel management system, which may be performed to adjust the viscosity of the fuel, may comprise controlling the proportion of fuel returned to the inlet of the primary heat exchanger via the recirculation valve. The step of controlling the fuel management system may comprise controlling a fuel metering valve (as described above) and the recirculation valve based on information provided by the fuel metering valve (and optionally also based on other information, e.g. temperature sensor data).

The fuel management system may further include a bypass line arranged to allow a proportion of the fuel to bypass at least one of the primary heat exchanger and the secondary heat exchanger.

The control step of the fuel management system, which may be performed to adjust fuel viscosity, may include controlling the proportion of fuel that passes through the bypass line instead of through the primary heat exchanger.

The fuel management system may further include an oil bypass line arranged to allow a proportion of the oil to bypass at least one of the primary heat exchanger and the secondary heat exchanger, and optionally an oil bypass line for each heat exchanger.

The control step of the fuel management system, which may be performed to adjust fuel viscosity, may include controlling the proportion of oil that passes through the oil bypass line instead of through the at least one heat exchanger.

The method may include controlling the fuel management system to raise the fuel temperature to between 150°C and 170°C at the inlet into the combustion chamber under cruise conditions.

According to a sixth aspect, a gas turbine engine for an aircraft is provided. The gas turbine engine comprises a combustion chamber arranged to burn a fuel, and a fuel management system arranged to supply the fuel to the combustion chamber. The fuel management system comprises two fuel-oil heat exchangers arranged to have oil and fuel flow therethrough, the heat exchangers arranged to transfer heat between the oil and the fuel and comprising a primary fuel-oil heat exchanger and a secondary fuel-oil heat exchanger. The fuel management system further comprises a fuel pump arranged to supply the fuel to the combustion chamber. The fuel pump is located between the two fuel-oil heat exchangers. The fuel management system is arranged to adjust the fuel viscosity to less than 0.58 mm ² /s upon entering the combustion chamber at cruise conditions.

It is envisaged that all the characteristics described for the fifth aspect may also apply to the sixth aspect.

Similarly, the sixth or seventh aspect can be implemented in conjunction with the first or second aspect and/or the third or fourth aspect.

According to a seventh aspect, a method of operating a gas turbine engine is provided, the gas turbine engine comprising:

a combustion chamber arranged to burn a fuel; and

a fuel management system arranged to supply fuel to the combustion chamber, wherein the fuel management system comprises:

two fuel-oil heat exchangers through which oil and fuel flow, the heat exchangers being arranged to transfer heat between the oil and the fuel and comprising a primary fuel-oil heat exchanger and a secondary fuel-oil heat exchanger; and

a fuel pump arranged to deliver fuel to the combustion chamber, wherein the fuel pump is located between the two heat exchangers;

wherein the method comprises controlling the fuel management system so as to transfer between 200 and 600 kJ/m ³ of heat to the fuel from the oil in the primary fuel-oil heat exchanger under cruise conditions.

It will be appreciated that heat transfer is denoted per unit volume (m ³ ) of fuel reaching the combustion chamber so as to adjust the fuel flow rate and any recirculation through the primary fuel-oil heat exchanger or bypass of the primary fuel-oil heat exchanger as described below. The amount of heat transferred to the fuel may therefore be calculated based on a fuel temperature at the approach or entry to the combustion chamber as compared to a fuel temperature in an aircraft fuel tank, taking into account heat transferred from other sources to the fuel reaching the combustion chamber (e.g., heat transferred from the oil to the fuel in the secondary fuel-oil heat exchanger, in implementations in which the fuel is sent to the combustion chamber instead of being returned to a fuel tank for subsequent recirculation). The heat gained by the fuel from other engine components, e.g. the pump, should be negligible compared to the heat gain in the heat exchangers and can therefore be ignored in most implementations.

Since heat transfer is measured per unit volume of fuel, this can be considered a normalized heat transfer rate for fuel flow variations under cruise conditions.

The fuel pump is located between the two heat exchangers along the flow path – i.e. downstream of the primary heat exchanger but upstream of the secondary heat exchanger.

The inventors were aware that the use of fuels other than traditional kerosene-based fuels, such as sustainable aviation fuels, may result in different fuel properties, and that the parameters at cruise conditions may be adjusted to utilize the different fuel properties. For example, some fuels may have superior thermal stability and/or higher heat capacity, which allows for increased heat transfer to the fuel and/or a higher operating temperature. In particular, more heat may be transferred from the oil to the fuel in the fuel-oil heat exchangers in some implementations. This may improve the cooling of the oil before it is returned to the rest of the turbine engine. This in turn improves the cooling effect of the oil on the engine components through which it flows. Increased fuel temperature may also improve combustion efficiency in the combustor, as previously described for other aspects.

Placing the pump before the secondary heat exchanger may allow the fuel temperature to be increased more than otherwise due to the sensitivity of the pump components to high temperatures.

The fuel management system can be controlled to transfer between 300 and 500 kJ/ ^m3 of heat to the fuel from the oil at cruise conditions.

The fuel management system can be controlled to transfer between 350 and 450 kJ/ ^m3 of heat to the fuel from the oil at cruise conditions.

The fuel management system can be controlled to transfer between 250 and 550 kJ/ ^m3 , 250 and 450 kJ/ ^m3 , 300 and 450 kJ/ ^m3 , 300 and 400 kJ/ ^m3 , 350 and 400 kJ/ ^m3 , 400 and 450 kJ/ ^m3 or 400 and 500 kJ/ ^m3 of heat to the fuel from the oil under cruise conditions.

The fuel management system may further include a recirculation valve located downstream of the primary heat exchanger, the recirculation valve arranged to allow a controlled amount of fuel that has passed through the primary heat exchanger to be returned to an inlet of the primary heat exchanger.

Controlling the fuel management system to control the amount of heat transferred may involve controlling the proportion of fuel returned to the primary heat exchanger inlet via the recirculation valve.

The gas turbine engine may include a fuel metering valve (as described above), the fuel metering valve being arranged to control the flow of fuel therethrough and to provide information about the flow of fuel therethrough (e.g., as a function of the valve position). The controlling step of the fuel management system may include controlling the fuel metering valve and the recirculation valve as a function of the information provided by the fuel metering valve (and possibly also as a function of other information, e.g. temperature sensor data).

The fuel management system may further include a bypass line arranged to allow a proportion of the fuel to bypass at least one of the primary heat exchanger and the secondary heat exchanger.

Controlling the fuel management system to control the amount of heat transferred may include controlling the proportion of fuel that passes through the bypass line instead of through the primary heat exchanger.

The fuel management system may further comprise an oil bypass line arranged to allow a proportion of the oil to bypass at least one of the primary heat exchanger and the secondary heat exchanger, and optionally may comprise two oil bypass lines, one for each heat exchanger.

Controlling the fuel management system to control the amount of heat transferred may include controlling the proportion of oil that passes through the (or each) oil bypass line instead of through the at least one heat exchanger.

According to an eighth aspect, a gas turbine engine for an aircraft is provided. The gas turbine engine comprises a combustion chamber arranged to burn a fuel, and a fuel management system arranged to supply the fuel to the combustion chamber. The fuel management system comprises two fuel-oil heat exchangers arranged to have oil and fuel flow therethrough, the heat exchangers arranged to transfer heat between the oil and the fuel and comprising a primary fuel-oil heat exchanger and a secondary fuel-oil heat exchanger. The fuel management system further comprises a fuel pump arranged to supply the fuel to the combustion chamber. The fuel pump is located between the two fuel-oil heat exchangers. The fuel management system is arranged to transfer between 200 and 600 kJ/m ³ of heat to the fuel from the oil in the primary fuel-oil heat exchanger under cruise conditions.

It is envisaged that all the characteristics described for the seventh aspect may also apply to the eighth aspect.

Similarly, the seventh or eighth aspect can be implemented in conjunction with any or all of the previous aspects.

The characteristics below may apply in all or part of the above aspects.

In the above aspects, the oil may enter the fuel-oil heat exchangers at a higher temperature than the fuel enters the respective heat exchanger under cruising conditions, such that the fuel-oil heat exchangers are arranged to transfer heat from the oil to the fuel.

Fuel may flow through the primary fuel-oil heat exchanger before flowing through the secondary fuel-oil heat exchanger. Fuel may flow through the fuel pump after exiting the primary heat exchanger and before entering the secondary heat exchanger.

The primary fuel-oil heat exchanger may be referred to as the primary fuel-oil heat exchanger. The majority of the heat transfer between the oil and the fuel may occur in the primary fuel-oil heat exchanger. The primary function of the primary fuel-oil heat exchanger may be to heat the fuel before it is supplied to the combustion chamber. At least substantially all of the fuel may pass through the primary fuel-oil heat exchanger.

The secondary fuel-oil heat exchanger may be a slave fuel-oil heat exchanger. Some, but not all, of the fuel exiting the primary fuel-oil heat exchanger may be directed to the slave fuel-oil heat exchanger. The slave heat exchanger is typically smaller than the primary heat exchanger due to the reduced fuel flow rate compared to that through the primary heat exchanger. The slave fuel-oil heat exchanger may further increase the fuel temperature, prior to providing fuel for use in servo mechanisms of the gas turbine engine (e.g. for fuel hydraulic actuation and/or heating). These servo mechanisms may include a nacelle anti-icing system. The servo mechanisms may include engine actuators. The servo mechanisms may include a servo turbine case cooling (TCC) valve. Only fuel that has passed through the secondary fuel-oil heat exchanger can be used in these auxiliary systems.

Fuel used in these auxiliary systems can be returned to a fuel tank for later recirculation, or recombined with other fuel exiting the main fuel-oil heat exchanger and entering the combustion chamber.

Fuel passing through the secondary fuel-oil heat exchanger may therefore not be supplied to the combustion chamber in some implementations, but instead returned to an aircraft fuel tank, possibly after use in auxiliary systems such as fuel hydraulic actuators.

The slave fuel-oil heat exchanger may be structurally similar to, or identical to, the primary fuel-oil heat exchanger. The slave fuel-oil heat exchanger may be smaller than the primary fuel-oil heat exchanger.

At least a portion of the fuel may be bypassed through the secondary fuel-oil heat exchanger. At least a portion of the fuel may be bypassed through the primary fuel-oil heat exchanger. A bypass path may be provided for one or each heat exchanger to allow a portion of the fuel to bypass that heat exchanger.

In particular, in examples where fuel passing through the secondary fuel-oil heat exchanger is returned to a reservoir rather than supplied to the combustion chamber, the proportion of fuel exiting the primary fuel-oil heat exchanger that bypasses the secondary fuel-oil heat exchanger may be greater than the proportion flowing through the secondary fuel-oil heat exchanger, optionally by a factor of at least three. A bypass line for the secondary fuel-oil heat exchanger may be provided in addition to an alternate fuel route from the primary heat exchanger to the combustion chamber that bypasses the secondary heat exchanger, in some implementations. Some fuel used for auxiliary systems may therefore have bypassed the secondary heat exchanger rather than passing through it.

A portion of the fuel may be directed to the secondary fuel-oil heat exchanger by any suitable means such as one or more valves or the like.

Approximately 10% to 30% of the fuel beyond the primary fuel-oil heat exchanger (i.e., fuel that has passed through or bypassed the primary fuel-oil heat exchanger) may be delivered to the secondary fuel-oil heat exchanger. This may be described as 10% to 30% of the fuel in the primary fuel path upstream of the secondary fuel-oil heat exchanger. For example, approximately 10% to 20% of the fuel may be delivered to the secondary fuel-oil heat exchanger. Approximately 13% to 17% of the fuel may be delivered to the secondary fuel-oil heat exchanger. As such, more fuel typically flows from the primary heat exchanger to the combustion chamber without flowing through the secondary heat exchanger than flows through the secondary heat exchanger. The flow path from the primary heat exchanger to the combustion chamber can therefore be described as a main fuel flow path, with a proportion of the fuel being diverted from this path to flow through the secondary fuel-oil heat exchanger, and possibly some or all of this proportion being returned to the main fuel flow path before the combustion chamber.

Approximately 0 to 10%, 20 to 30%, 20 to 40%, or 30 to 40% of the fuel is allowed to pass through the secondary fuel-to-oil heat exchanger.

In other examples, a higher proportion of fuel may pass through the secondary fuel-oil heat exchanger.

A fixed portion of fuel may pass through the secondary fuel-oil heat exchanger. Alternatively, a variable portion of fuel may pass through the secondary fuel-oil heat exchanger – control of the fuel management system may include controlling the proportion of fuel that is sent to the secondary heat exchanger.

A ratio of oil to fuel heat transfer for the primary and secondary fuel-oil heat exchangers can be approximately between 70:30 and 90:10. The primary fuel-oil heat exchanger can therefore be responsible for 70 to 90% of the heat transfer – it can be called “primary” because it is responsible for most of the heat transfer, i.e. being the primary source of heat to heat the fuel before entering the combustion chamber, as well as being the first fuel-oil heat exchanger to be reached by the fuel.

A ratio of oil to fuel heat transfer for primary and secondary fuel-oil heat exchangers can be approximately 80:20.

In other examples, the oil to fuel heat transfer ratio for the secondary fuel-to-oil heat exchanger may be higher.

Oil may flow through the secondary fuel-oil heat exchanger before flowing through the primary fuel-oil heat exchanger. The oil may not pass through components that add heat to it between the outlet of the secondary heat exchanger and the inlet to the primary heat exchanger. Thus, the temperature of the oil flowing through the secondary fuel-oil heat exchanger may be higher than the temperature of the oil flowing through the primary fuel-oil heat exchanger under cruising conditions. Fuel may flow through the primary fuel-oil heat exchanger before flowing through the secondary fuel-oil heat exchanger.

Indeed, the fuel management system may be arranged such that fuel flows through the primary fuel-oil heat exchanger before flowing through the secondary fuel-oil heat exchanger while oil flows through the secondary fuel-oil heat exchanger before flowing through the primary fuel-oil heat exchanger.

In view of the above, it will be appreciated that the fuel exiting the secondary fuel-oil heat exchanger is generally hotter than the fuel exiting the primary fuel-oil heat exchanger. The fuel exiting the secondary fuel-oil heat exchanger may be delivered to the combustion chamber, or may be returned to a fuel tank (possibly via fuel hydraulics, or other fuel-using systems), rather than being passed to a pump and then possibly recirculated, for example. In addition, the shorter fuel path of the fuel exiting the secondary heat exchanger compared to that exiting the primary heat exchanger may reduce the number of components that have interacted with the fuel at its highest temperature, which may improve the longevity of the component.

Additionally, locating the fuel pump between the primary and secondary fuel-oil heat exchangers can help preserve fuel pump life. Fuel temperatures can be high at the outlet of the primary heat exchanger, and even higher at the outlet of the secondary heat exchanger. Locating the fuel pump between the two, and specifically before the secondary heat exchanger, reduces the fuel pump’s exposure to the higher temperature fuel flow, and therefore reduces damage related to these high temperatures.

Controlling the fuel management system may include controlling the flow of fuel through one or more of the heat exchangers. Controlling the fuel flow may include adjusting the flow of fuel through one or more of the heat exchangers.

Controlling the fuel management system may include controlling the flow of oil through one or more of the heat exchangers. Controlling the oil flow may include adjusting the flow of oil through one or more of the heat exchangers.

Controlling the fuel management system may include changing parameters of one or more components of the system, including one or more of the heat exchangers or the fuel pump. Controlling the fuel management system may further include using additional components, such as an air-to-oil heat exchanger and one or more valves arranged to divert a controlled portion of the oil to the air-to-oil heat exchanger. A dedicated controller may be used to control the fuel management system, or the EEC may be used.

The flow of fuel may be controlled for each of the third through eighth aspects by using a recirculation valve located downstream of the primary heat exchanger, the recirculation valve arranged to allow a controlled amount of fuel that has not passed through the secondary heat exchanger to be returned to the inlet of the primary heat exchanger.

The gas turbine engine may be part of an aircraft. The aircraft may include a fuel tank, and a fuel feed pump configured to supply a flow of fuel to the fuel management system of the engine itself. The fuel feed pump may be described as a fuel tank pump or a low pressure pump, and is located upstream of the engine, and thus upstream of the primary fuel-oil heat exchanger.

The fuel pump of the fuel management system may be described herein as the main fuel pump or engine fuel pump, because, unlike the fuel feed pump, it is a part of the engine itself. The main fuel pump is located downstream of the primary fuel-oil heat exchanger. The main fuel pump is located upstream of the outlet of the secondary fuel-oil heat exchanger, and typically also upstream of the inlet of the secondary fuel-oil heat exchanger. One or more auxiliary fuel pumps may be located at any suitable position along the fuel flow path.

As discussed above, at least a portion of the fuel passing through the secondary fuel-oil heat exchanger may be supplied to servo mechanisms of the aircraft. At least a portion of this fuel may then be returned to an aircraft fuel tank or the low pressure pump for recirculation.

Alternatively or additionally, at least a portion of the fuel flowing through the secondary fuel-oil heat exchanger may be returned to any prior point in the fuel flow path; for example, back to the inlet of the primary or secondary fuel-oil heat exchanger, or may be recombined with another fuel approaching the combustion chamber.

In some exemplary implementations, a third, fourth, or any suitable number of other fuel-oil heat exchangers may be present, optionally with a fuel valve operable to control a flow of fuel therethrough.

The oil flow within the gas turbine engine may follow a closed loop. The closed loop oil system may be configured to supply a recirculating flow of oil within the engine and may be described as a recirculating lubrication and/or cooling system, or as a recirculating oil system. At least one of the primary and secondary fuel-to-oil heat exchangers may be described as part of the closed loop system. The recirculating lubrication and/or cooling system may be described as an oil heat management system, since heat is removed from the oil after it has been heated in the process of lubricating and/or cooling other system components.

The closed loop system may include at least a first oil pump configured to pump an oil flow around at least a portion of the recirculating oil system. The first oil pump may be located at any suitable position around the recirculating oil system. The recirculating oil system may be configured such that the oil flow flows through engine components to be lubricated and/or cooled (e.g., the engine bearing chamber) and may then be collected in a crankcase. The first oil pump may be configured to pump oil from the crankcase to a first oil reservoir. As such, the first oil pump may be described as a scavenge pump.

The oil reservoir may be suitable for holding a volume of oil. The oil reservoir may be designed to hold any suitable volume of oil. The oil reservoir may be arranged to remove gases from the oil in the first oil reservoir. Oil exiting the oil reservoir may pass through a filter, strainer or the like.

A second oil pump may be located between the first oil reservoir and the primary or secondary fuel-oil heat exchanger. The second oil pump may be described as a feed pump. The second oil pump may be configured to pump oil from the first oil reservoir to the primary or secondary fuel-oil heat exchanger.

In some implementations, at least a portion of the oil flows through the primary fuel-oil heat exchanger and at least a portion of the oil flows through the secondary fuel-oil heat exchanger. All of the oil may flow through both the primary and secondary fuel-oil heat exchangers. As discussed previously, the oil may first flow through the secondary fuel-oil heat exchanger and then through the primary fuel-oil heat exchanger, thereby having the opposite flow path to that of the fuel.

The gas turbine engine may further include one or more air-to-oil heat exchangers. The one or more air-to-oil heat exchangers may be part of the recirculating oil system. The one or more air-to-oil heat exchangers may be arranged in flow series with the primary and secondary fuel-to-oil heat exchangers such that at least a portion of the oil flow flows through an air-to-oil heat exchanger before entering the primary or secondary fuel-to-oil heat exchanger, or vice versa. In one example, the oil may flow through a secondary fuel-to-oil heat exchanger before flowing through an air-to-oil heat exchanger (e.g., a surface air-cooled oil cooler) and then the primary fuel-to-oil heat exchanger. An oil bypass pipe may be provided to allow a variable proportion of the oil to bypass the or each air-to-oil heat exchanger. As part of the fuel management system control, the bypass ratio can be adjusted to decrease or increase the amount of heat in the oil available to the fuel.

At cruise conditions, and for all aspects described herein, the oil flow entering any of the fuel-oil heat exchangers may have a higher average temperature than the fuel flow entering the same fuel-oil heat exchanger at cruise conditions. In this manner, thermal energy may be transferred from the oil flow to the fuel flow flowing through the fuel-oil heat exchanger(s) at cruise conditions. Thus, the oil exiting each heat exchanger may have a lower temperature than the oil entering that heat exchanger at cruise conditions.

It will be appreciated that in any of the examples described herein, an intermediate heat transfer fluid could be used rather than direct oil-to-fuel heat transfer, and that the term "fuel-to-oil heat exchanger" may include heat exchangers using such an intermediate fluid. Similarly, an oil-to-oil, fuel-to-air, or oil-to-air heat exchanger may use an intermediate heat transfer fluid in certain implementations.

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

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

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

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

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 the fan drive turbine, for example without a reduction gear, via a core shaft. In such a direct-drive turbofan engine, the fan may be said to rotate at the same rotational speed as the fan drive turbine. Strictly by way of example, the fan drive turbine may be a first turbine, the core shaft may be a first core shaft, and the gas turbine engine may further comprise a second turbine and a second core shaft connecting the second turbine to the compressor. The second turbine, the compressor, and the second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft. In such an arrangement, the second turbine may be positioned axially upstream of the first turbine.

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

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

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

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

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

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

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

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

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

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

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

The rotational speed of the fan may vary during use. Typically, the rotational speed is lower for fans with a larger diameter. Strictly by way of non-limiting example, the rotational speed of the fan at cruise conditions may be less than 3500 rpm, for example less than 2500 rpm, for example less than 2300 rpm. Strictly by way of further non-limiting example, the rotational speed of the fan at cruise conditions for a "geared" gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 2750 to 2900 rpm. Strictly by way of further non-limiting example, the fan rotational speed at cruise conditions for a "geared" gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 2500 to 2800 rpm. Strictly by way of further non-limiting example, the fan rotational speed at cruise conditions for a "geared" gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 1500 to 1800 rpm. Strictly by way of further non-limiting example, the fan rotational speed at cruise conditions for a direct drive engine having a fan diameter in the range of 190 cm to 200 cm may be in the range of 3600 to 3900 rpm. Strictly by way of further non-limiting example, the fan rotation speed at cruise conditions for a direct drive engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 2000 to 2800 rpm.

During operation of the gas turbine engine, the fan (with associated fan blades) rotates about an axis of rotation. This rotation results in a displacement of the fan blade tip with a velocity U _tip . The work done by the fan blades on the flow results in an enthalpy increase dH of the flow. A fan tip load can be defined as dH/U _tip ² , where dH is the enthalpy increase (e.g. the 1-D average enthalpy increase) across the fan and U _tip is the (translational) velocity of the fan tip, e.g. at the leading edge of the fan tip (which can be defined as the fan tip radius at the leading edge multiplied by the angular velocity). The peak fan load at cruise conditions may be greater than (or of the order of) any of: 0.28, 0.29, 0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all values being dimensionless). The peak fan load may be within an inclusive range bounded by any two of the values in the preceding sentence (i.e. the values may form upper or lower limits), for example in the range from 0.28 to 0.31 or from 0.29 to 0.3 (e.g. for a geared gas turbine engine).

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

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

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

A gas turbine engine as described and/or claimed herein may have any desired maximum thrust. Strictly by way of non-limiting example, a gas turbine engine as described and/or claimed herein may be capable of producing a maximum thrust of at least (or on the order of) any of the following: 110 kN, 110 kN, 120 kN, 130 kN, 140 kN, 150 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 an inclusive range bounded by any two of the values in the preceding sentence (i.e., the values may form upper or lower limits). By way of non-limiting example only, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust in the range of 330 kN to 420 kN, for example 350 kN to 400 kN. Strictly by way of non-limiting example, the maximum thrust of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 140 kN to 160 kN. Strictly by way of non-limiting example, the maximum thrust of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 150 kN to 200 kN. Strictly by way of non-limiting example, the maximum thrust of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 370 kN to 500 kN. Strictly by way of non-limiting example, the maximum thrust of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 370 kN to 500 kN. The above mentioned thrust may be the maximum net thrust under standard atmospheric conditions at sea level plus 15 degrees C (ambient pressure 101.3 kPa, temperature 30 degrees C), with the engine static.

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

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

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

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

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

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

As used herein, the terms rest, taxi, takeoff, climb, cruise, descent, approach and landing have the conventional meaning and would be readily understood by those skilled in the art. Thus, for a given gas turbine engine for an aircraft, those skilled in the art would readily recognize that each term refers to a phase of engine operation within a given mission of an aircraft to which the gas turbine engine is designed to be attached.

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

Climb may refer to a phase of engine operation where the aircraft is propelled by thrust produced by the engine. During climb, the engine may produce between 75% and 100% of available thrust. In other non-limiting examples, the engine may produce between 80% and 95% of available thrust. In other non-limiting examples, the engine may produce between 85% and 90% of available thrust. As such, climb may refer to a phase of operation within an aircraft flight cycle between takeoff and arrival at cruise conditions. Additionally or alternatively, climb may refer to a nominal point in an aircraft flight cycle between takeoff and landing where a relative increase in altitude is required, which may require an additional demand for engine thrust.

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

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

Strictly by way of example, the forward speed at the cruise condition may be any point in the range of Mach 0.7 to 0.9, for example 0.75 to 0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example 0.78 to 0.82, for example 0.79 to 0.81, for example of the order of Mach 0.8, of the order of Mach 0.85 or in the range of 0.8 to 0.85. Any single speed within these ranges may be part of the cruise condition. For a given aircraft, the cruise conditions may be outside these ranges, for example below Mach 0.7 or above Mach 0.9.

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

Strictly by way of example, cruise conditions may correspond to an engine operating point that provides a known required thrust level (e.g., a value in the range of 30 kN to 35 kN) at a forward Mach number of 0.8 and typical atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38,000 feet (11,582 m). Strictly by way of further example, cruise conditions may correspond to an engine operating point that provides a known required thrust level (e.g., a value in the range of 50 kN to 65 kN) at a forward Mach number of 0.85 and typical atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35,000 feet (10,668 m).

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

Further, one skilled in the art would readily recognize that either or both of a descent and an approach refer to a phase of operation within an aircraft flight cycle between cruise and landing of the aircraft. During either or both of descent and approach, the engine may produce between 20% and 50% of available thrust. In other non-limiting examples, the engine may produce between 25% and 40% of available thrust. In other non-limiting examples, the engine may produce between 30% and 35% of available thrust. Additionally or alternatively, descent may refer to a nominal point in an aircraft flight cycle between takeoff and landing where a relative decrease in altitude is required, and which may require a reduced thrust demand from the engine.

In one aspect, there is provided an aircraft comprising a gas turbine engine as described and/or claimed herein. The aircraft in this aspect is the aircraft to which the gas turbine engine has been designed to be attached. Thus, the cruise conditions in this aspect correspond to the mid-cruise of the aircraft, as defined elsewhere herein.

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

In one aspect, there is provided a method of operating an aircraft comprising a gas turbine engine as described and/or claimed herein. Operation in accordance with this aspect may include (or may be) operation at any suitable condition, for example at mid-cruise of the aircraft, as defined elsewhere herein.

One skilled in the art would understand that, unless mutually exclusive, any feature or parameter described in connection with any of the above aspects may be applied to any other aspect. Furthermore, unless mutually exclusive, any feature or parameter described herein may be applied to any aspect and/or in combination with any other feature or parameter described herein.

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

is a side sectional view of a gas turbine engine;

is a close-up sectional side view of an upstream portion of a gas turbine engine;

is a partially cutaway view of a gearbox for a gas turbine engine;

is a sectional side view of a direct drive gas turbine engine, the direct drive engine having no reduction gear as shown in FIG.

shows an example of an aircraft incorporating two gas turbine engines;

is a schematic representation of an example fuel system;

is a schematic representation of an exemplary alternative fuel system including a recirculation loop;

is a schematic representation of another example of an alternative fuel system including a recirculation loop and use of fuel for auxiliary systems;

[Fig. 6C] is a schematic representation of another example of an alternative fuel system similar to that shown in Figure 6b but in which fuel passing through the secondary heat exchanger is returned to the combustion chamber optionally after being used in one or more auxiliary systems;

is a schematic representation of part of an exemplary recirculating oil system;

is a schematic representation of part of the example fuel system of the and the example of the recirculating oil system of the ;

is a schematic representation of a portion of the exemplary fuel system of [Fig. 6C] and the exemplary recirculating oil system of the ;

illustrates an example of a method of operating a gas turbine engine;

illustrates another example of a method of operating a gas turbine engine;

illustrates another example of a method of operating a gas turbine engine; and

illustrates another example of a gas turbine engine operating method.

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

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

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

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

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

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

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

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

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

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

As a further example, other gas turbine engines to which the present disclosure may be applied may not have a reduction gear for the main shaft(s), being instead direct drive engines. A cross-sectional view of such an engine is shown in FIG. .

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

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

Compressed air exhausted from the high pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel F and the mixture is combusted. The combustion equipment 16 may be referred to as the combustion chamber 16, with the terms “combustion equipment 16” and “combustion chamber 16” used interchangeably herein. The resulting hot combustion products then expand, thereby driving the high pressure, intermediate pressure and low pressure turbines 17, 19a, 19 before being exhausted through the nozzle 20 to provide additional propulsive thrust. The high pressure 17, intermediate pressure 19a and low pressure 19 turbines respectively drive the high pressure compressor 15, the intermediate pressure compressor 14 and the fan 23, each by a suitable interconnecting shaft.

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

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

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

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

The functional performance of a given composition, or fuel mixture for use in a given mission, may be defined, at least in part, by the fuel's ability to handle the Brayton cycle of the gas turbine engine 10. Parameters defining functional performance may include, for example, specific energy; energy density; thermal stability; and particulate matter emissions. A relatively higher specific energy (i.e., energy per unit mass), expressed in MJ/kg, may at least partially reduce takeoff weight, potentially providing a relative improvement in fuel efficiency. A relatively higher energy density (i.e., energy per unit volume), expressed in MJ/L, may at least partially reduce fuel volume at takeoff, which may be particularly important for volume-limited missions or military operations involving refueling. Relatively higher thermal stability (i.e., inhibiting fuel degradation or coking under thermal stress) can allow the fuel to withstand elevated engine and fuel injector temperatures, potentially providing relative improvements in combustion efficiency. Reduced emissions, including particulate matter, can allow for reduced contrail formation, while reducing the environmental impact of a given mission. Other fuel properties can also be critical to operational performance. For example, a relatively lower freezing point (°C) can allow long-range missions to optimize flight profiles; minimum aromatic concentrations (%) can ensure sufficient swelling of certain materials used in the construction of O-rings and gaskets previously exposed to high aromatic fuels; and, maximum surface tension (mN/m) can ensure sufficient spray breakup and atomization of the fuel.

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

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

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

Due at least in part to the molecular structure of sustainable aviation fuels, sustainable aviation fuels may provide beneficial effects including, for example, one or more of higher specific energy (despite, in some examples, lower energy density); higher specific heat capacity; higher thermal stability; higher lubricity; lower viscosity; lower surface tension; lower freezing point; lower soot emissions; and, lower CO ₂ emissions, relative to fossil-based hydrocarbon fuels (eg, when burned in combustion equipment 16). Thus, relative to fossil-based hydrocarbon fuels, such as kerosene, sustainable aviation fuels may result in either or both of a relative decrease in specific fuel consumption, and a relative decrease in maintenance costs.

There shows an aircraft 1 with a propulsion system 2 comprising two gas turbine engines 10. The gas turbine engines 10 are supplied with fuel from a fuel supply system on board the aircraft 1. The fuel supply system of the illustrated examples comprises a single fuel source 50, 53.

For the purposes of this application, the term "fuel source" means either 1) a single fuel tank or 2) a plurality of fuel tanks that are fluidly interconnected.

In the present examples, the first (and, in these examples, only) fuel source comprises a central fuel tank 50, located primarily in the fuselage of the aircraft 1 and a plurality of wing fuel tanks 53a, 53b, where at least one wing fuel tank is located in the port wing and at least one wing fuel tank is located in the starboard wing for balancing. All of the tanks 50, 53 are fluidly connected in the illustrated example, thereby forming a single fuel source. Each of the central fuel tank 50 and the wing fuel tanks 53 may comprise a plurality of fluidly interconnected fuel tanks. It will be appreciated that this tank arrangement is provided by way of example only, and is not limiting of the scope of this disclosure.

In alternative examples, the wing fuel tanks 53a, 53b may be fluidically interconnected with each other, but not with the center fuel tank 50, such that there are two separate fuel sources, each of which may be used to hold a different fuel. Three or more fluidically isolated fuel sources may be provided in other examples. A fuel management system as described below may take fuel from one or more of the fuel sources 50, 53 to power the engine 10.

An exemplary fuel system 1000 including a fuel flow path from the fuel tank 50 to the combustion chamber 16 of the gas turbine engine 10 of the aircraft 1 is schematically shown in FIG. . The fuel system 1000 includes both the fuel supply system 50, 53, 1002 (which supplies fuel to the engine) and the fuel management system 1500 (which operates within the engine) of the aircraft 1. In the described implementation, each engine 10 has its own fuel management system 1500. In other implementations, a single fuel management system 1500 may manage the supply of fuel to multiple engines, and may for example include a duplicate of various of the elements shown in the for the other engine(s).

Fuel is pumped from the fuel tank 50 (exemplarily captured from the various fuel tanks 50, 53) to the gas turbine engine 10 by a low pressure fuel feed pump 1002. The fuel then flows through a primary fuel-oil heat exchanger 1004 before flowing through an engine fuel pump 1003. The engine fuel pump 1003 may be described as a primary fuel pump. At least a portion of the fuel then flows through a secondary fuel-oil heat exchanger 1006 and at least a portion of the fuel flows to the combustion chamber 16 without flowing through the secondary fuel-oil heat exchanger 1006.

In the example shown in the , from the secondary fuel-oil heat exchanger 1006 the portion of fuel that has flowed therethrough then flows to the combustion chamber 16 of the gas turbine engine 10, joining the fuel that has flowed through the primary fuel-oil heat exchanger 1004 only. In alternatives, such as that shown in FIG. (described below), the portion of fuel that flows through the secondary fuel-oil heat exchanger 1006 may be returned to a reservoir 50, 53 rather than being delivered to the combustion chamber 16, possibly used in one or more auxiliary systems 7010 between the outlet of the secondary fuel-oil heat exchanger 1006 and before reaching the reservoir 50.

The primary fuel-oil heat exchanger 1004 may be referred to as a primary fuel-oil heat exchanger. The secondary fuel-oil heat exchanger 1006 may be referred to as a slave fuel-oil heat exchanger. The fuel management system 1500 is arranged such that fuel reaches the primary fuel-oil heat exchanger 1004 before the secondary fuel-oil heat exchanger 1006. In this example, a bypass line 1005 is provided to allow a controlled amount of fuel to pass through this line 1005 instead of through the primary fuel-oil heat exchanger 1004. Such a bypass line 1005 may be provided for one or both of the heat exchangers 1004, 1006 in different implementations. One or more equivalent bypass lines may be provided for oil in various implementations, thereby allowing a controlled amount of the oil to pass through that line instead of through the respective heat exchanger 1004, 1006.

The primary fuel-oil heat exchanger 1004 and the secondary fuel-oil heat exchanger 1006 are designed such that an oil flow is also carried therethrough. The primary fuel-oil heat exchanger 1004 and the secondary fuel-oil heat exchanger 1006 are designed such that heat can be transferred between the oil and fuel flowing therethrough. At cruising conditions, the average temperature of the oil flow entering the primary fuel-oil heat exchanger 1004 and the secondary fuel-oil heat exchanger 1006 is higher than the average temperature of the fuel entering the primary fuel-oil heat exchanger 1004 and the secondary fuel-oil heat exchanger 1006 respectively. In this manner, the primary fuel-oil heat exchanger 1004 and the secondary fuel-oil heat exchanger 1006 are configured to transfer thermal energy from an oil flow to a fuel flow flowing therethrough at cruise conditions.

In various arrangements of the fuel management system 1500, oil flows through the secondary fuel-oil heat exchanger 1006 before passing through the primary fuel-oil heat exchanger 1004, and does not flow through engine components which would increase its temperature therebetween. The oil is therefore hotter at the inlet of the secondary fuel-oil heat exchanger 1006 than at the inlet of the primary fuel-oil heat exchanger 1004. In contrast, the fuel flows through the primary fuel-oil heat exchanger 1004 before flowing through the secondary fuel-oil heat exchanger 1006. In this arrangement, the fuel temperature exiting the secondary fuel-oil heat exchanger 1006 is higher than the fuel temperature exiting the primary fuel-oil heat exchanger 1004. The engine fuel pump 1003 is positioned upstream of the secondary fuel-oil heat exchanger 1006, i.e. it is not exposed to these higher fuel temperatures.

The oil may flow through an air-oil heat exchanger (not shown) as well as through the primary and secondary fuel-oil heat exchangers 1004, 1006 if additional heat loss from the oil is desired, beyond that to be absorbed by the fuel. For example, the oil may pass through the air-oil heat exchanger after passing through the secondary fuel-oil heat exchanger 1006 and before entering the primary fuel-oil heat exchanger 1004. The air-oil heat exchanger may be the only component (other than the pipes) through which the oil passes between the secondary fuel-oil heat exchanger 1006 and the primary fuel-oil heat exchanger 1004. No air-oil heat exchanger may be provided in other examples, or a provided air-oil heat exchanger may be differently located. The oil may be returned to the bearings and/or other engine components for lubrication and/or cooling after exiting the primary heat exchanger 1004. The oil may be returned to an oil reservoir 2002, as described below, after exiting the primary heat exchanger 1004, or optionally after exiting the bearings and/or other engine components for lubrication and/or cooling, before then being returned to the secondary heat exchanger 1006.

There illustrates an example of an alternative fuel system 6000 like that of the fuel system 1000 described above with respect to the , but with a fuel recirculation loop in place of a fuel bypass line 1005 (in alternative implementations, both a recirculation loop and a bypass line may be provided, optionally for the same heat exchanger). This exemplary alternative fuel system 6000 includes a fuel supply system 50, 1002 as described above and a fuel management system 6500 including a fuel flow path from the fuel tank 50 to the combustion chamber 16 of the gas turbine engine 10 of the aircraft 1. As with the fuel system 1000, fuel is pumped from the fuel tank 50 by the low pressure fuel supply pump 1002. The fuel then flows through a primary fuel-oil heat exchanger 1004 before flowing through an engine fuel pump 1003. The engine fuel pump 1003 may be described as a primary fuel pump.

The 6000 fuel system of the differs from the 1000 fuel system of the in that the 6000 fuel system of the includes a recirculation valve 6010 located downstream of the primary fuel-oil heat exchanger 1004, but before the combustion chamber 16. The recirculation valve may be located upstream of an outlet 1006b of the secondary fuel-oil heat exchanger 1006 where fuel from that heat exchanger rejoins the primary fuel flow path. In the illustrated example, the recirculation valve 6010 is located downstream of the pump 1003. The recirculated fuel therefore passes through the primary fuel-oil heat exchanger 1004 and the pump 1003 more than once.

In the illustrated example, the recirculation valve 6010 is located downstream of the inlet 1006a of the secondary fuel-oil heat exchanger 1006, but upstream of its outlet 1006b. It is contemplated that the recirculation valve 6010 may be located upstream of the inlet 1006a of the secondary heat exchanger 1006 in other examples (e.g. as described below with respect to [Fig. 6C] and [Fig. 6D]). ). In yet other implementations, there may be no outlet from the secondary heat exchanger 1006 back to the primary fuel flow path to the combustion chamber 16, as described below with respect to the .

The recirculation valve 6010 is arranged to allow a controlled amount of fuel that has passed through the primary heat exchanger 1004 but has not passed through the secondary heat exchanger 1006 to be returned to the inlet 1004a of the primary heat exchanger 1004. This provides a mechanism for controlling the fuel flow of the fuel system 6000, by returning fuel that has passed through the primary heat exchanger 1004 (and possibly also the pump 1003) to a position in the flow path before the primary heat exchanger 1004 (and possibly also before the pump 1003).

Control of fuel flow in the fuel system 6000 may operate to provide desired fuel characteristics (such as temperature and flow rate) at the inlet to the combustion chamber 16. In particular, the fuel flow rate to the combustion chamber 16 may be adjusted without changing the speed of the pump 1003 if the recirculation valve 6010 is located after the pump 1003 and the proportion recirculated is varied – this may be of particular utility in implementations in which the pump speed is directly related to the engine shaft speed, allowing for variation in flow rate to the combustion chamber 16 at a given pump speed. In implementations in which fuel flowing through the secondary heat exchanger 1006 is not returned to the combustion chamber 16, changing a proportion of fuel sent to that heat exchanger 1006 may also be used to control the flow rate of fuel into the combustion chamber 16.

A line 6011, which may be referred to as a recirculation line because 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 recirculated fuel must pass through a section of the flow path more than once, may thus be provided, leading from the recirculation valve 6010 to a point in the flow path upstream of the inlet of the primary heat exchanger 1004. In the illustrated example, the recirculation line 6011 returns the recirculated fuel to a point in the flow path upstream of both the main pump 1003 and the primary heat exchanger 1004, such that the recirculated fuel passes through both of these components twice.

There illustrates an example of an alternative fuel system 7000 to that shown in FIG. . The 7000 fuel system is similar to the 6000 fuel system shown in the but differs in that at least a portion of the fuel exiting the secondary fuel-oil heat exchanger 1006 is supplied to additional aircraft mechanisms 7010, rather than to the combustor 16. These mechanisms 7010 may include one or more of a nacelle anti-icing system, actuators, and/or turbine case cooling (TCC) systems. At least a portion of the fuel exiting the secondary fuel-oil heat exchanger 1006 is returned to the fuel supply system (i.e., fuel tank 50 or fuel feed pump 1002) for redistribution. No fuel passing through the secondary fuel-oil heat exchanger 1006 is sent to the combustor 16 in the example of the (at least not directly – it will be kept in mind that fuel returned to the fuel tank 50 may re-enter the engine 10 later, and may then be passed to the combustion chamber 16).

There illustrates another alternative example of fuel system 7000a. Fuel system 7000a is similar to fuel system 7000 illustrated in FIG. and differs in that the portion of the fuel exiting the secondary fuel-oil heat exchanger 1006 is returned to join the main fuel flow as it approaches the combustor 16 after optionally being supplied to one or more additional aircraft mechanisms 7010 (which may be as described above). Any fuel not required for these auxiliary systems 7010 may be returned directly to the main fuel flow path/combustion chamber 16. At least a portion of the fuel exiting the secondary fuel-oil heat exchanger 1006 may be returned to the main fuel flow path without passing through the auxiliary systems 7010 – the proportion directed to the auxiliary systems 7010 may be adjusted as needed. In other implementations, the fuel used in one or more such additional aircraft mechanisms 7010 may then be returned to a tank 50 rather than being sent to the combustor 16.

In the various examples described above, the aircraft 1 includes a recirculating oil system 2000 arranged to supply oil to lubricate and/or remove heat from a plurality of components. The recirculating oil system includes a closed-loop oil system, and may be comprised of a single closed loop in some implementations. An example of a closed-loop oil system 2000 is schematically shown in FIG. .

The closed loop oil system 2000 includes an oil reservoir 2002 suitable for holding a volume of oil. In some implementations, gases are removed from the oil within the oil reservoir 2002 by a de-aerator.

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

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

In the primary fuel-oil heat exchanger 1004, thermal energy is transferred from the oil flow to the fuel flow. In this manner, the average temperature of the oil flow exiting the primary fuel-oil heat exchanger 1004 is lower than the average temperature of the oil flow entering the primary fuel-oil heat exchanger 1004. Also in this manner, the average temperature of the fuel exiting the primary fuel-oil heat exchanger 1004 is higher than the average temperature of the fuel entering the primary fuel-oil heat exchanger 1004.

From the primary fuel-oil heat exchanger 1004, the oil flow is then routed through the engine components 2006, which the oil is intended to cool and/or lubricate. The oil serves as a lubricant and coolant within the engine components 2006 and helps to gain heat from the engine components 2006 in the process. The engine components 2006 may include one or more bearing chambers.

Since the 2006 engine components, oil has been collected in the 2008 crankcase. A 2010 scavenge pump is designed to pump oil from the 2008 crankcase back into the 2002 oil reservoir.

There schematically shows an exemplary heat exchange system comprising the closed loop oil system 2000 and the fuel system 1000. The two systems 1000, 2000 are shown together to illustrate the interactions between the fuel and oil flows within the gas turbine engine 10. The fuel flow is shown with thick black arrows and the oil flow with thin black arrows.

There schematically depicts an exemplary alternative heat exchange system 9000, including the closed-loop oil system 2000 and the fuel delivery system 7000a. The systems 7000a, 2000 are substantially the same as those illustrated in FIGS. 6C and 7 respectively, with the fuel flow again shown with thick black arrows and the oil flow with thin black arrows. However, the recirculation valve 6010 is located slightly differently than in [FIG. 6C], namely upstream of the fork in the fuel flow path to the secondary heat exchanger 1006, rather than being between the inlet of 1006a and the outlet of 1006b of that heat exchanger 1006. Thus in the implementation illustrated in FIG. , all of the fuel exiting the pump 1003 flows through the recirculation valve 6010, unlike the embodiment illustrated in the . The control methodology for the 6010 valve can be adjusted as appropriate.

It will be kept in mind that in implementations such as that illustrated in the , wherein fuel flowing through the secondary fuel-oil heat exchanger 1006 is used as a working fluid in auxiliary systems and then returned to a reservoir 50, 53 instead of passing to the combustion chamber 16, the fuel flow line from the secondary heat exchanger 1006 to the combustion chamber 16 may be replaced by a fuel flow line returned to the reservoir 50.

The inventors were aware that using fuels different from traditional kerosene-based fuels, such as sustainable aviation fuels, can result in different fuel properties, and that engine operation can be optimized for these different fuel properties. Fuels having a higher heating value as claimed can also have greater thermal stability, allowing the fuel to absorb more heat, which provides improved oil cooling and/or improved combustion properties in the combustion chamber. Recirculating the fuel through the primary fuel-oil heat exchanger 1004 can allow the fuel to receive more heat from the oil, increasing the fuel temperature and improving oil cooling.

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

There illustrates an exemplary method 10000 of a gas turbine engine 10. The method 10000 comprises the steps of:

Step 10100: Obtaining/selecting a fuel to be supplied to the gas turbine engine 10, wherein the calorific value of the fuel is at least 43.5 MJ/kg, and optionally between 43.5 and 44 MJ/kg.

Obtaining fuel may include selecting a single fuel. Obtaining fuel may alternatively include selecting a fuel blend.

The selected fuel may be the only fuel on board the aircraft 1. As such, the fuel selection may be performed while the aircraft is being refueled. Alternatively, the selected fuel may be one of multiple fuels on board the aircraft 1, or a blend of multiple fuels on board the aircraft. Thus, the fuel selection may be performed during flight. The selection process may be automated. For example, an onboard controller may be configured to select and/or blend fuel(s) stored in one or more fuel tanks 50, 53 to provide a fuel of the desired calorific value. The controller may store the calorific values and tank locations for the onboard fuels and calculate the fuel blend ratio (which may not include a blend of fuels) to provide a fuel of the desired calorific value.

Step 10200: Delivering fuel to the combustion chamber 16 via the fuel management system 6500.

As discussed in relation to the and to the , the fuel management system 6500 includes a primary fuel-oil heat exchanger 1004 and a secondary fuel-oil heat exchanger 1006 arranged to transfer heat between the oil and the fuel, and a recirculation loop 6010, 6011 arranged to allow a proportion of the fuel to pass through the primary heat exchanger 1004 multiple times. An engine fuel pump 1003 is positioned downstream of the primary fuel-oil heat exchanger 1004 and upstream of the secondary fuel-oil heat exchanger 1006. The fuel management system 6500 is arranged such that the fuel reaches the primary fuel-oil heat exchanger 1004 before the secondary fuel-oil heat exchanger 1006.

Optionally, the method may further comprise:

Step 10300: Actively controlling an amount of fuel exiting the primary heat exchanger 1004 to be recirculated to an inlet 1004a of the primary heat exchanger 1004 rather than flowing to the secondary fuel-oil heat exchanger 1006 or directly to the combustion chamber 16, using a recirculation valve 6010. This recirculation may also bring fuel that has already passed through the engine fuel pump 1003 back to a position upstream of the engine fuel pump 1003, as described above.

This active control can be performed based on one or more parameters such as:

• Heart shaft speed and engine thrust demand;

• Current fuel temperature and/or oil temperature;

• Calorific value of fuel;

• Blower rotation speed;

• • Fuel flow to the combustion chamber (commonly referred to as WFE – weight of main engine fuel flow); and

• Main fuel pump/engine speed, or speed options.

For example, a higher percentage of the fuel exiting the pump can be directed back to the pump inlet for fuels with a higher calorific value.

In alternative examples, controlling the amount of fuel exiting the primary heat exchanger 1004 to be recirculated may not be an active process step—instead, a defined, fixed proportion of the fuel may be recirculated. For example, 70%, 75%, 80%, 85%, or 90% of the fuel may be recirculated at cruise. The proportion recirculated may be higher for a newer pump, for a given fuel and engine. The proportion recirculated may be higher for a higher calorific value fuel, for a given pump condition and engine.

Particularly for engine architectures in which the fuel pump speed is directly related to the shaft speed (e.g., a ratio of shaft speed to pump speed is fixed, or in which only certain ratios of shaft speed to pump speed are possible, at least under cruising conditions), fuel can be wasted if the engine operation is not adjusted. Recirculating more fuel can therefore allow a lower velocity of fuel reaching the combustion chamber while maintaining a constant fuel flow rate through the pump 1003.

The arrangement of the above fuel management system 6500, with fuel recirculation occurring, may therefore provide greater flexibility in fuel flow. One or more additional or alternative controllable valves, pumps and/or sensors may be provided in various implementations to facilitate fuel flow control.

The inventors were also aware that using fuels other than traditional kerosene-based fuels, such as sustainable aviation fuels, can result in different fuel properties, and that the cruise parameters can be adjusted to utilize the different fuel properties. In particular, certain 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 it is returned to the rest of the turbine engine, and/or improve the combustion efficiency of the fuel. Using the fuel to absorb more heat from the oil, rather than relying on heat transfer from the oil to the environment/air (e.g., in an oil-air heat exchanger) provides a more thermally efficient turbine engine. Furthermore, the improved oil cooling can in turn improve the cooling effect of the oil on the engine components through which it flows.

There illustrates an exemplary method 11000 of a gas turbine engine 10. The method 11000 comprises the steps of:

Step 11100: Providing fuel to the combustion chamber 16 of the gas turbine engine 10 via a fuel management system 1500.

As discussed in relation to the , the fuel management system 1500 includes a primary fuel-oil heat exchanger 1004 and a secondary fuel-oil heat exchanger 1006 arranged to transfer heat to or from the fuel. An engine fuel pump 1003 is positioned downstream of the primary fuel-oil heat exchanger 1004 and upstream of the secondary fuel-oil heat exchanger 1006. The fuel management system 1500 is arranged such that fuel reaches the primary fuel-oil heat exchanger 1004 before the secondary fuel-oil heat exchanger 1006.

Step 11200: Control the fuel management system 1500 so as to raise the fuel temperature to at least 135°C at the inlet of the combustion chamber 16 at cruise conditions.

Control 11200 of fuel management system 1500 may include controlling fuel flow through the heat exchangers—for example, recirculating a variable proportion of the fuel through primary fuel-oil heat exchanger 1004, and/or allowing a variable proportion of the fuel to bypass primary fuel-oil heat exchanger 1004. While the shows an implementation with a bypass line 1005 but no recirculation, and Figures 6A-6C show implementations without a bypass line (at least not for the primary heat exchanger - Figures 6A and 6C provide a route for a portion of the fuel reaching the combustion chamber 16 to bypass the secondary heat exchanger while the remainder of the fuel passes through it) but with a recirculation line 6011, it will be appreciated that one or more bypass lines and/or one or more recirculation lines may be provided together in various implementations. Furthermore, in some systems, the flow direction within a single line may be reversible such that it can be used as either a bypass line 1005 or a recirculation line 6011. It will be appreciated that the figures are exemplary only and are not intended to be limiting.

Control 11200 of fuel management system 1500 may include controlling fuel flow through the heat exchangers – for example allowing a variable proportion of the fuel to bypass primary fuel-oil heat exchanger 1004 or secondary fuel-oil heat exchanger 1006. One or more controllable valves or pumps, and optionally one or more sensors, may be provided accordingly to facilitate controlling fuel flow.

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

The heat exchange system may include a controller arranged to implement this control (of the fuel flow and/or the oil flow). The controller may receive inputs from one or more temperature sensors, and may control one or more valves (e.g. the recirculation valve 6010), and/or the pump 1003, based on the data received. One or more oil feed pumps and/or scavenge pumps may also be controlled by the controller.

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

Heating the fuel to higher temperatures than previously used may improve cooling of the oil before it is returned to the remainder of the turbine engine, and/or may improve fuel combustion efficiency. Positioning the secondary fuel-oil heat exchanger 1006 after the engine fuel pump 1003 may facilitate achieving higher fuel temperatures without compromising fuel pump longevity.

The process of the can be used in conjunction with the method of the .

The inventors were aware that, particularly with the use of fuels other than traditional kerosene-based fuels, such as sustainable aviation fuels, fuel properties at cruise conditions can be controlled by careful engine management to improve performance. In this case, a lower viscosity can be achieved, which can affect combustion efficiency, particularly with respect to the atomization performance of the fuel nozzle within the combustion chamber 16. The fuel nozzle atomization performance affects the combustion efficiency of the fuel. A lower viscosity of the fuel at cruise conditions can be more conducive to a more efficient engine. The fuel flow rate can be optimized to improve the efficiency of the turbine engine, to which the lower viscosity of said fuel can contribute.

There illustrates an exemplary method 12000 of a gas turbine engine 10. The method 12000 comprises the steps of:

Step 12100: Providing fuel to the combustion chamber 16 of the gas turbine engine 10 via a fuel management system 1500.

As discussed in relation to the , the fuel management system 1500 includes a primary fuel-oil heat exchanger 1004 and a secondary fuel-oil heat exchanger 1006 arranged to transfer heat to the fuel. An engine fuel pump 1003 is positioned downstream of the primary fuel-oil heat exchanger 1004 and upstream of the secondary fuel-oil heat exchanger 1006. The fuel management system 1500 is arranged such that the fuel reaches the primary fuel-oil heat exchanger 1004 before the secondary fuel-oil heat exchanger 1006.

Step 12200: Controlling the fuel management system 1500 so as to adjust the fuel viscosity to less than 0.58 mm ² /s at the inlet of the combustion chamber 16 under cruise conditions.

Careful engine management to achieve a lower fuel viscosity at the inlet to the combustion chamber 16 can improve combustion efficiency, particularly relative to fuel nozzle spray performance within the combustion chamber. Fuel flow properties, oil flow properties, and/or one or more other heat exchange properties can be adjusted to provide this viscosity difference.

Control 12200 of the fuel management system 1500 may therefore comprise controlling fuel flow and/or oil flow through the heat exchangers as described above – for example recirculating a variable proportion of the fuel through the primary fuel-oil heat exchanger 1004, and/or allowing a variable proportion of the fuel to bypass the primary and/or secondary fuel-oil heat exchanger 1004, 1006, as described above with respect to the .

For example, fuel flow may be controlled using the recirculation valve 6010, similar to the fuel management system 6000, and/or by adjusting the proportion of fuel sent through the secondary fuel-oil heat exchanger 1004 or through a bypass line 1005.

The process of the can be used in conjunction with the method of the and/or the process of the .

The inventors were aware that the use of fuels other than traditional kerosene-based fuels, such as sustainable aviation fuels, may result in different fuel properties, and that the parameters at cruise conditions may be adjusted to utilize the different fuel properties. For example, some fuels may have greater thermal stability and/or heat capacity, which allows for increased heat transfer to the fuel and/or a higher operating temperature. In particular, more heat may be transferred from the oil to the fuel in the fuel-oil heat exchangers 1004, 1006 in some implementations. This may improve the cooling of the oil before it is returned to the remainder of the turbine engine. This in turn improves the cooling effect of the oil on the engine components through which it flows. Increased fuel temperature may also improve combustion efficiency in the combustor 16.

There illustrates an exemplary method 13000 of a gas turbine engine 10. The method 13000 comprises the steps of:

Step 13100: Providing fuel to the combustion chamber 16 of the gas turbine engine 10 via a fuel management system 1500.

As discussed in relation to the , the fuel management system 1500 includes a primary fuel-oil heat exchanger 1004 and a secondary fuel-oil heat exchanger 1006 arranged to transfer heat to or from the fuel. An engine fuel pump 1003 is positioned downstream of the primary fuel-oil heat exchanger 1004 and upstream of the secondary fuel-oil heat exchanger 1006. The fuel management system 1500 is arranged such that fuel reaches the primary fuel-oil heat exchanger 1004 before the secondary fuel-oil heat exchanger 1006.

Step 13200: Controlling the fuel management system 1500 so as to transfer 200 to 600 kJ/m ³ of heat to the fuel from the oil in the primary fuel-oil heat exchanger under cruise conditions (heat transfer being measured per unit volume of fuel reaching the combustion chamber 16).

The fuel pump 1003 is located between the two heat exchangers 1004, 1006 along the flow path – i.e. downstream of the primary heat exchanger but upstream of the secondary heat exchanger.

Control 13200 of the fuel management system 1500 may therefore include controlling fuel flow and/or oil flow through the heat exchangers – for example, recirculating a variable proportion of the fuel through the primary fuel-oil heat exchanger 1004, and/or allowing a variable proportion of the fuel to bypass the primary and/or secondary fuel-oil heat exchanger 1004, 1006, as described above with respect to the . The fuel management system 1500 may therefore include one or more additional components not shown in the Figures—such as bypass lines, recirculation lines, controllable pumps, and/or controllable valves—to enable or facilitate such control. One or more sensors—for example, temperature and flow sensors—may be used to provide fuel and oil management feedback and guidance. For example, fuel flow may be controlled using the recirculation valve 6010, similar to the fuel management system 6000, and/or by adjusting the proportion of fuel sent through the secondary fuel-oil heat exchanger 1004.

For example, a 6014 fuel metering valve (FMV) as shown in the may be provided to control the fuel delivered to the combustion chamber 16. The position of the valve 6014 (ie its degree of opening/closing) may be monitored, and may also be used to determine the fuel flow rate due to a controller of (or in communication with) the fuel management system having access to a record of the relationship between the valve position and the flow rate. The FMV 6014 and the recirculation valve 6010 may work together to deliver a required fuel flow to the combustion chamber 16. The position of the FMV may be actively controlled in a closed loop, with excess fuel being returned to the pump 1003 via the recirculation valve and line 6010, 6011. The FMV 6014 may be located at an outlet 1006b of the secondary fuel-oil heat exchanger 1006 in some implementations, downstream of the recirculation valve 6010 on the primary fuel flow path through the engine 10, as illustrated in [Fig. 6C].

Adjusting heat exchange parameters under cruising conditions can provide more efficient engine performance.

The process of the may be used in conjunction with the method of any or all of Figures 10 through 12. Similarly, an FMV 6014 may be used in any or all of the methods described.

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

Claims (15)

1 A method (11000) of operating a gas turbine engine (10), the gas turbine engine (10) comprising:
a combustion chamber (16) arranged to burn a fuel; and
a fuel management system arranged to supply fuel to the combustion chamber (16), wherein the fuel management system comprises:
two fuel-oil heat exchangers through which oil and fuel flow, the heat exchangers arranged to transfer heat to the fuel and comprising a primary fuel-oil heat exchanger (1004) and a secondary fuel-oil heat exchanger (1006); and
a fuel pump (1003) arranged to deliver fuel to the combustion chamber (16), wherein the fuel pump (1003) is located between the two heat exchangers;
wherein the method comprises controlling (11200) the fuel management system so as to raise the fuel temperature to at least 135°C at the inlet to the combustion chamber (16) at cruise conditions. 2 A method (11000) according to claim 1, wherein the method comprises controlling (11200) the fuel management system so as to raise the fuel temperature to between 135°C and 170°C at the inlet into the combustion chamber (16) under cruising conditions. 3 Method (11000) according to claim 1, in which:
The fuel management system further includes:
a recirculation valve (6010) located downstream of the primary fuel-oil heat exchanger (1004), the recirculation valve (6010) arranged to allow a controlled amount of fuel that has passed through the primary fuel-oil heat exchanger (1004) to be returned to an inlet of the primary fuel-oil heat exchanger (1004);
and wherein controlling the fuel management system comprises controlling the proportion of fuel returned to the inlet of the primary fuel-oil heat exchanger (1004) via the recirculation valve (6010). 4 Method (11000) according to claim 1, in which:
The fuel management system further includes:
a bypass line (1005) arranged to allow a proportion of the fuel to bypass the primary fuel-oil heat exchanger (1004);
and wherein controlling the fuel management system includes controlling the proportion of fuel that passes through the bypass line instead of through the primary fuel-oil heat exchanger (1004). 5 Method (11000) according to claim 1, wherein:
the fuel management system further comprises an oil bypass line arranged to allow a proportion of the oil to bypass at least one of the primary fuel-oil heat exchanger (1004) and the secondary fuel heat exchanger (1006);
and wherein controlling the fuel management system includes controlling the proportion of oil that passes through the oil bypass line instead of through the at least one heat exchanger. 6. A method (11000) according to claim 3, wherein the gas turbine engine (10) comprises a fuel metering valve (6014) downstream of the recirculation valve (6010) arranged to control the flow of fuel therethrough and to provide information about the flow of fuel therethrough, and wherein controlling (13200) the fuel management system comprises controlling the fuel metering valve (6014) and the recirculation valve (6010) based on information provided by the fuel metering valve (6014). 7 The method (11000) of claim 1, wherein between 10% and 30% of the fuel is delivered to the secondary fuel-oil heat exchanger (1006). 8 The method (11000) of claim 1, wherein a ratio of heat transfer from oil to fuel for the primary and secondary fuel-to-oil heat exchangers (1004, 1006) is between 70:30 and 90:10. 9 The method (11000) of claim 1, wherein the fuel flows through the primary fuel-oil heat exchanger (1004) before flowing through the secondary fuel-oil heat exchanger (1006) while the oil flows through the secondary fuel-oil heat exchanger (1006) before flowing through the primary fuel-oil heat exchanger (1004). 10 A gas turbine engine (10) for an aircraft (1), the gas turbine engine (10) comprising:
a combustion chamber (16) arranged to burn a fuel; and
a fuel management system arranged to supply fuel to the combustion chamber (16), wherein the fuel management system comprises:
two fuel-oil heat exchangers arranged for oil and fuel to flow therethrough, the heat exchangers arranged to transfer heat between the oil and the fuel and comprising a primary fuel-oil heat exchanger (1004) and a secondary fuel-oil heat exchanger (1006); and
a fuel pump (1003) arranged to deliver fuel to the combustion chamber (16), wherein the fuel pump (1003) is located between the two fuel-oil heat exchangers;
wherein the fuel management system is arranged to raise the fuel temperature to at least 135°C at the inlet into the combustion chamber (16) under cruising conditions. 11 A gas turbine engine (10) according to claim 10, wherein the fuel management system is arranged to raise the fuel temperature to between 135°C and 170°C upon entering the combustion chamber (16) under cruise conditions. 12 A gas turbine engine (10) according to claim 10, wherein:
The fuel management system further includes:
a recirculation valve (6010) located downstream of the primary fuel-oil heat exchanger (1004), the recirculation valve (6010) arranged to allow a controlled amount of fuel that has passed through the primary fuel-oil heat exchanger (1004) to be returned to an inlet of the primary fuel-oil heat exchanger (1004). 13 A gas turbine engine (10) according to claim 10, wherein the secondary fuel-oil heat exchanger (1006) is a slave fuel-oil heat exchanger. 14 A gas turbine engine (10) according to claim 10, wherein a heat transfer ratio of oil to fuel for the primary and secondary fuel-to-oil heat exchangers (1004, 1006) is between 70:30 and 90:10. 15 A gas turbine engine (10) according to claim 10, wherein the fuel management system is arranged such that fuel flows through the primary fuel-oil heat exchanger (1004) before flowing through the secondary fuel-oil heat exchanger (1006) while oil flows through the secondary fuel-oil heat exchanger (1006) before flowing through the primary fuel-oil heat exchanger (1004).