DE102023135089A1 - Fuel management - Google Patents

Abstract

A method (12000) of operating a gas turbine engine (10) is disclosed, the gas turbine engine (10) comprising a combustion chamber (16) arranged to combust a fuel; and a fuel management system arranged to provide the fuel to the combustion chamber (16). The fuel management system comprises two fuel-oil heat exchangers through which oil and the fuel flow, the heat exchangers arranged to transfer heat to the fuel, a primary fuel-oil heat exchanger (1004) and a secondary fuel-oil heat exchanger (1006); and a fuel pump (1003) arranged to deliver the fuel to the combustion chamber (16), the fuel pump (1003) located between the two heat exchangers. The method includes controlling (12200) the fuel management system to adjust the fuel viscosity to less than 0.58 mm2/s upon entering the combustion chamber (16) under cruise conditions.

Description

The present disclosure relates to aircraft propulsion systems and methods of operating aircraft involving the management of different fluids and the heat transfer therebetween, and more particularly, the management of fuel properties as they enter the combustion chamber.

A trend is expected in the aviation industry towards the use of fuels that are different from the conventional kerosene-based jet fuels currently in general use. The fuels may have different fuel properties compared to petroleum-based hydrocarbon fuels.

There is therefore a need to consider fuel properties for these new fuels and to adapt the procedures for operating gas turbine engines accordingly.

• eine Brennkammer, die angeordnet ist, um einen Kraftstoff zu verbrennen; und
• ein Kraftstoffmanagementsystem, das angeordnet ist, um der Brennkammer den Kraftstoff bereitzustellen, wobei das Kraftstoffmanagementsystem umfasst:
  • zwei Kraftstoff-Öl-Wärmetauscher, durch die Öl und der Kraftstoff hindurch strömen, wobei die Wärmetauscher angeordnet sind, um Wärme zwischen dem Öl und dem Kraftstoff zu übertragen, und einen primären Kraftstoff-Öl-Wärmetauscher und einen sekundären Kraftstoff-Öl-Wärmetauscher umfassen;
  • eine Kraftstoffpumpe, die angeordnet ist, um den Kraftstoff an die Brennkammer abzugeben, wobei sich die Kraftstoffpumpe zwischen den zwei Kraftstoff-Öl-Wärmetauschern befindet; und
  • ein Rezirkulationsventil, das sich stromabwärts des primären Kraftstoff-Öl-Wärmetauschers befindet, wobei das Rezirkulationsventil angeordnet ist, um zu ermöglichen, dass eine kontrollierte Menge an Kraftstoff, die durch den primären Kraftstoff-Öl-Wärmetauscher hindurchgetreten ist, an einen Einlass des primären Kraftstoff-Öl-Wärmetauschers zurückgeführt wird;
• wobei das Verfahren das Auswählen von Kraftstoff dahingehend umfasst, dass der Brennwert des Kraftstoffs, der dem Gasturbinentriebwerk unter Reiseflugbedingungen bereitgestellt wird, mindestens 43,5 MJ/kg beträgt.

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

• a combustion chamber arranged to burn a fuel; and
• a fuel management system arranged to provide the fuel to the combustion chamber, the fuel management system comprising:
  • two fuel-oil heat exchangers through which oil and the fuel flow, 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;
  • a fuel pump arranged to deliver the fuel to the combustion chamber, the fuel pump being 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 amount 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 method comprising selecting fuel such that the calorific value of the fuel provided 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 primary heat exchanger and secondary heat exchanger for brevity.

The inventors have recognized that the use of fuels other than traditional kerosene-based jet 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 calorific value may also have greater thermal stability, allowing the fuel to absorb more heat, thereby providing improved oil cooling and/or improved combustion characteristics in the combustion chamber. Recirculation of the fuel through the primary fuel-oil heat exchanger may allow the fuel to absorb more heat from the oil, thereby increasing the fuel temperature and improving oil cooling.

It is understood that herein, when the fuel pump is described as being located "between" the two fuel-oil heat exchangers, this is intended to mean 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 leaving the primary fuel-oil heat exchanger and entering the secondary fuel-oil heat exchanger (or vice versa), regardless of whether a straight line would pass through the pump between the primary fuel-oil heat exchanger and the secondary fuel-oil heat exchanger. It is understood 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 redirect a portion of the flow leaving the pump to a point in the flow path upstream of the pump inlet. This may provide additional flexibility in the fuel flow rate 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 required fuel burn rate for a given shaft speed under cruise conditions. In particular, for engine architectures where the fuel pump speed is directly coupled to shaft speed (e.g., where a shaft speed to pump speed ratio is fixed or has only a set of specific discrete values, at least under cruise conditions), fuel may be wasted if engine operation is not adjusted. The inventors have recognized that more recirculation of fuel can therefore be performed, returning fuel that has left the pump to the pump inlet, such that a fuel flow rate to the combustion chamber can be lowered while keeping a fuel flow rate through the pump constant. Therefore, for fuels with a higher calorific value, more recirculation can be performed/a higher percentage of the fuel leaving the pump can be returned to the pump inlet. The arrangement of the above fuel management system, in which recirculation of fuel occurs before the fuel enters the secondary heat exchanger (located downstream of the pump), can therefore provide improved flexibility of fuel flow without passing hotter fuel through the pump, which could damage the pump. Additionally, or alternatively, in examples where the recirculation valve is located downstream of the primary heat exchanger but upstream of the pump, recirculating a portion of the fuel through the primary heat exchanger may improve and/or provide more flexibility for heat transfer arrangements and oil cooling.

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

For a given fuel and engine, the % recirculated may depend on the engine operating condition (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 deliver less current than a new one and thus provide reduced % recirculation). Fuel temperature may also affect the % recirculated. The values given above assume a reasonably new/well functioning pump and standard fuel temperatures for the given engine under cruise conditions.

Even under cruise conditions, power requirements may vary (e.g. depending on altitude, cruise level, and mass of fuel on board). Recirculation % may therefore vary based on current power requirements and hence required engine fuel burn rates.

The fuel management system may further comprise a bypass pipe arranged to allow a portion 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 bypassing a heat exchanger may depend on the engine operating condition (e.g., cruise, idle, descent, or maximum takeoff) and also, for example, on the fuel and/or oil temperature. 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 controlled and a relationship between the valve position and the fuel flow may be known so that an engine electronic controller (EEC) can determine or "meter" the fuel flow based on the valve position.

The FMV and the recirculation valve may cooperate to deliver a required fuel flow to the combustion chamber. The FMV position may be actively controlled in a closed loop, with excess fuel being returned via the recirculation valve to an earlier part of the fuel flow path, e.g., before an engine pump. The method may therefore include using the FMV and a recirculation tube 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 arranged to control the fuel flow rate therethrough and to provide information about the fuel flow therethrough (e.g. based on the valve position).

Obtaining/selecting the fuel may include selecting a single fuel. Selecting the fuel may include selecting a fuel mixture. The selected fuel may be the only fuel on board the aircraft. Thus, selecting the fuel during the fueling of the flight Alternatively, the fuel selected may be one of several fuels on board the aircraft or a mixture of several fuels on board the aircraft; the selection can therefore be made in flight.

The recirculation valve may be located downstream of an inlet of the secondary heat exchanger. The recirculation valve may therefore not affect what proportion of fuel is passed through the secondary heat exchanger in such examples (in the current pass) - this split may be fixed at a set 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 affect (in the current pass) a maximum amount of fuel that can 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 cannot be recirculated and therefore the pump is never exposed to the higher temperature of fuel leaving the secondary heat exchanger (since the pump is located upstream of the secondary heat exchanger). The pump may therefore be shielded from the higher temperatures of fuel leaving 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.

• eine Brennkammer, die angeordnet ist, um einen Kraftstoff zu verbrennen; und
• ein Kraftstoffmanagementsystem, das angeordnet ist, um der Brennkammer den Kraftstoff bereitzustellen, wobei das Kraftstoffmanagementsystem umfasst:
  • zwei Kraftstoff-Öl-Wärmetauscher, die dazu angeordnet sind, dass Öl und der Kraftstoff durch sie hindurch strömen, wobei die Kraftstoff-Öl-Wärmetauscher angeordnet sind, um Wärme zwischen dem Öl und dem Kraftstoff zu übertragen,
  • und einen primären Kraftstoff-Öl-Wärmetauscher und einen sekundären Kraftstoff-Öl-Wärmetauscher umfassen;
  • eine Kraftstoffpumpe, die angeordnet ist, um den Kraftstoff an die Brennkammer abzugeben, wobei sich die Kraftstoffpumpe zwischen den zwei Kraftstoff-Öl-Wärmetauschern befindet; und
  • ein Rezirkulationsventil, das sich stromabwärts des primären Kraftstoff-Öl-Wärmetauschers befindet, wobei das Rezirkulationsventil angeordnet ist, um zu ermöglichen, dass eine kontrollierte Menge an Kraftstoff, die durch den primären Kraftstoff-Öl-Wärmetauscher hindurchgetreten ist, an einen Einlass des primären Kraftstoff-Öl-Wärmetauschers zurückgeführt wird.

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

• a combustion chamber arranged to burn a fuel; and
• a fuel management system arranged to provide the fuel to the combustion chamber, the fuel management system comprising:
  • two fuel-oil heat exchangers arranged to allow oil and the fuel to flow through them, the fuel-oil heat exchangers being arranged to transfer heat between the oil and the fuel,
  • and comprise a primary fuel-oil heat exchanger and a secondary fuel-oil heat exchanger;
  • a fuel pump arranged to deliver the fuel to the combustion chamber, the fuel pump being 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 amount 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, in addition to the combustor, a turbine, a compressor, and a core shaft connecting the turbine to the compressor. The gas turbine engine may further include a fan located upstream of the engine core, the fan being driven by the core shaft.

The gas turbine engine may be a direct drive engine.

The gas turbine engine may further include an auxiliary power transmission and a pump shaft, the auxiliary power transmission receiving input from the core shaft and outputting output to the fuel pump via the pump shaft. The pump speed may therefore be coupled to the core shaft speed and may, by means of the transmission, have a fixed speed relationship between the core shaft (or in some engine architectures, another intermediate shaft) and the fuel pump. In some implementations, the auxiliary power transmission may instead be arranged to provide a fixed number of different speeds for a given core shaft speed. Recirculation of fuel may allow for a wider variety of fuel flow rates into the combustion chamber for the given number of pump speeds available at a given shaft speed. Recirculation of fuel may allow for multiple different fuel flow rates in implementations where there is a single fixed speed relationship between the core shaft and the fuel pump. rates into the combustion chamber at a given core shaft speed.

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

It is intended that all the features described for the first aspect can equally apply to the second aspect.

• eine Brennkammer, die angeordnet ist, um einen Kraftstoff zu verbrennen; und
• ein Kraftstoffmanagementsystem, das angeordnet ist, um der Brennkammer den Kraftstoff bereitzustellen, wobei das Kraftstoffmanagementsystem umfasst:
  • zwei Kraftstoff-Öl-Wärmetauscher, durch die Öl und der Kraftstoff hindurch strömen, wobei die Wärmetauscher angeordnet sind, um Wärme zwischen dem Öl und dem Kraftstoff zu übertragen, und einen primären Kraftstoff-Öl-Wärmetauscher und einen sekundären Kraftstoff-Öl-Wärmetauscher umfassen; und
  • eine Kraftstoffpumpe, die angeordnet ist, um den Kraftstoff an die Brennkammer abzugeben, wobei sich die Kraftstoffpumpe zwischen den zwei Wärmetauschern befindet;
  • wobei das Verfahren das Steuern des Kraftstoffmanagementsystems derart, dass die Kraftstofftemperatur auf mindestens 135 °C bei Eintritt in die Brennkammer unter Reiseflugbedingungen erhöht wird, umfasst.

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

• a combustion chamber arranged to burn a fuel; and
• a fuel management system arranged to provide the fuel to the combustion chamber, the fuel management system comprising:
  • two fuel-oil heat exchangers through which oil and the 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 the fuel to the combustion chamber, the fuel pump being located between the two heat exchangers;
  • the method comprising controlling the fuel management system to increase the fuel temperature to at least 135°C upon entering the combustion chamber under cruise conditions.

The inventors have recognized that using fuels other than traditional kerosene-based jet fuels, such as sustainable aviation fuels, can result in different fuel properties and that parameters can be adjusted under cruise conditions to take advantage of the different fuel properties. In particular, some fuels can be heated to higher temperatures in the fuel-oil heat exchangers than conventional fuels. This can improve cooling of the oil before it is returned to the rest of the turbine engine and/or can improve the combustion efficiency of the fuel, for example by influencing droplet size and distribution from fuel injectors. Using the fuel to remove more heat from the oil rather than relying on heat transfer from the oil to the environment/air (e.g. in an oil-air heat exchanger) provides a more thermally efficient turbine engine. Additionally, the improved cooling of the oil can in turn improve the cooling effect of the oil on the components of the engine through which it flows.

The fuel management system can be controlled to increase the fuel temperature to between 135 °C and 170 °C upon entering the combustion chamber under cruise conditions.

The fuel management system can be controlled to increase the fuel temperature to between 150 °C and 170 °C upon entering the combustion chamber under cruise conditions.

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

Fuel temperatures entering the combustion chamber under cruise conditions may be defined as an average over at least 5 minutes, 10 minutes or 30 minutes under steady cruise conditions. These average temperatures do not include transient temperature spikes, which may be defined as fluctuations in the temperature of the fuel during operation, often an increase in temperature. No fluctuation shall last more than 5 minutes.

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 passing through the primary heat exchanger passed through is returned to an inlet of the primary heat exchanger. The recirculation valve can also be located downstream of the fuel pump.

The step of controlling the fuel management system may include controlling the proportion of fuel returned to the inlet of the primary heat exchanger 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 controlled and a relationship between the valve position and the fuel flow may be known so that an engine electronic controller (EEC) can determine or "meter" the fuel flow based on the valve position.

The FMV and the recirculation valve may cooperate to deliver a required fuel flow to the combustion chamber. The FMV position may be actively controlled in a closed loop, with excess fuel being returned via the recirculation valve to an earlier part of the fuel flow path, e.g., before an engine pump. The method may therefore include using the FMV and a recirculation tube 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 arranged to control the fuel flow rate therethrough and to provide information about the fuel flow therethrough (e.g., based on valve position). The step of controlling the fuel management system may include controlling the fuel metering valve and the recirculation valve based on the 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 comprise a bypass pipe arranged to allow a portion 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 enable this control.

The step of controlling the fuel management system may include controlling the proportion of fuel that passes through the bypass pipe rather than the primary heat exchanger.

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

The step of controlling the fuel management system may comprise controlling the proportion of oil passing through the oil bypass pipe rather than through the at least one heat exchanger, for example by adjusting a controllable valve.

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

• eine Brennkammer, die angeordnet ist, um einen Kraftstoff zu verbrennen; und
• ein Kraftstoffmanagementsystem, das angeordnet ist, um der Brennkammer den Kraftstoff bereitzustellen, wobei das Kraftstoffmanagementsystem umfasst:
  • zwei Kraftstoff-Öl-Wärmetauscher, die dazu angeordnet sind, dass Öl und der Kraftstoff durch sie hindurch strömen, wobei die Wärmetauscher angeordnet sind, um Wärme zwischen dem Öl und dem Kraftstoff zu übertragen, und einen primären Kraftstoff-Öl-Wärmetauscher und einen sekundären Kraftstoff-Öl-Wärmetauscher umfassen; und
  • eine Kraftstoffpumpe, die angeordnet ist, um den Kraftstoff an die Brennkammer abzugeben, wobei sich die Kraftstoffpumpe zwischen den zwei Kraftstoff-Öl-Wärmetauschern befindet;
• wobei das Kraftstoffmanagementsystem angeordnet ist, um die Kraftstofftemperatur auf mindestens 135 °C bei Eintritt in die Brennkammer unter Reiseflugbedingungen zu erhöhen.

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

• a combustion chamber arranged to burn a fuel; and
• a fuel management system arranged to provide the fuel to the combustion chamber, the fuel management system comprising:
  • two fuel-oil heat exchangers arranged to have oil and the 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; and
  • a fuel pump arranged to deliver the fuel to the combustion chamber, the fuel pump being located between the two fuel-oil heat exchangers;
• wherein the fuel management system is arranged to increase the fuel temperature to at least 135 °C upon entering the combustion chamber under cruise conditions.

It is intended that all the features described for the third aspect can equally apply to the fourth aspect.

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

• eine Brennkammer, die angeordnet ist, um einen Kraftstoff zu verbrennen; und
• ein Kraftstoffmanagementsystem, das angeordnet ist, um der Brennkammer den Kraftstoff bereitzustellen, wobei das Kraftstoffmanagementsystem umfasst:
  • zwei Kraftstoff-Öl-Wärmetauscher, durch die Öl und der Kraftstoff hindurch strömen, wobei die Wärmetauscher angeordnet sind, um Wärme zwischen dem Öl und dem Kraftstoff zu übertragen, und einen primären Kraftstoff-Öl-Wärmetauscher und einen sekundären Kraftstoff-Öl-Wärmetauscher umfassen; und
  • eine Kraftstoffpumpe, die angeordnet ist, um den Kraftstoff an die Brennkammer abzugeben, wobei sich die Kraftstoffpumpe zwischen den zwei Wärmetauschern befindet;
• wobei das Verfahren das Steuern des Kraftstoffmanagementsystems derart umfasst, dass die Kraftstoffviskosität auf weniger als 0,58 mm²/s (d. h. weniger als 0,58 cSt) bei Eintritt in die Brennkammer unter Reiseflugbedingungen angepasst wird.

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

• a combustion chamber arranged to burn a fuel; and
• a fuel management system arranged to provide the fuel to the combustion chamber, the fuel management system comprising:
  • two fuel-oil heat exchangers through which oil and the 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 the fuel to the combustion chamber, the fuel pump being located between the two heat exchangers;
• the method comprising controlling the fuel management system to adjust the fuel viscosity to less than 0.58 mm ² /s (ie, less than 0.58 cSt) upon entering the combustion chamber under cruise conditions.

The inventors have recognized that, particularly when using fuels other than traditional kerosene-based jet fuels, such as sustainable aviation fuels, fuel properties under cruise conditions can be controlled through careful engine management to improve performance. In this case, a lower viscosity can be obtained, which can affect combustion efficiency, particularly with respect to fuel nozzle spray performance within the combustion chamber. Fuel nozzle spray performance impacts fuel combustion efficiency. Lower fuel viscosity under cruise conditions can conducive to a more efficient engine. Fuel flow can be optimized to improve turbine engine efficiency, which the lower viscosity of the fuel can contribute to. For example, fuel viscosity affects how fuel is delivered to and ignited by the combustion chamber. Viscosity can affect droplet size and spray distribution from fuel spray nozzles, which in turn can affect combustion efficiency.

Therefore, taking fuel viscosity into account when supplying fuel to the combustion chamber and controlling it appropriately by varying the heat input can provide more efficient fuel combustion, thereby improving aircraft performance.

The fuel may pass 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 minor portion of the fuel may pass through the secondary heat exchanger. Fuel passing through the secondary heat exchanger may, optionally after use in one or more auxiliary systems, be returned to a fuel tank instead of going to the combustion chamber.

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

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

The fuel management system may further comprise a recirculation valve located downstream of the primary heat exchanger, the recirculation valve being 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 fuel viscosity, 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 the fuel metering valve (and optionally based on other information, e.g. temperature sensor data).

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

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

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

The step of controlling 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 tube rather than through the at least one heat exchanger.

The method may include controlling the fuel management system to increase the fuel temperature to between 150°C and 170°C upon entering the combustion chamber under cruise conditions.

According to a sixth aspect, there is provided a gas turbine engine for an aircraft. The gas turbine engine comprises a combustion chamber arranged to combust a fuel and a fuel management system arranged to provide the fuel to the combustion chamber. The fuel management system comprises two fuel-oil heat exchangers arranged to have oil and the fuel flowing 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 deliver 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 under cruise conditions.

It is intended that all the features described for the fifth aspect can equally apply to the sixth aspect.

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

• eine Brennkammer, die angeordnet ist, um einen Kraftstoff zu verbrennen; und
• ein Kraftstoffmanagementsystem, das angeordnet ist, um der Brennkammer den Kraftstoff bereitzustellen, wobei das Kraftstoffmanagementsystem umfasst:
  • zwei Kraftstoff-Öl-Wärmetauscher, durch die Öl und der Kraftstoff hindurch strömen, wobei die Wärmetauscher angeordnet sind, um Wärme zwischen dem Öl und dem Kraftstoff zu übertragen, und einen primären Kraftstoff-Öl-Wärmetauscher und einen sekundären Kraftstoff-Öl-Wärmetauscher umfassen; und
  • eine Kraftstoffpumpe, die angeordnet ist, um den Kraftstoff an die Brennkammer abzugeben, wobei sich die Kraftstoffpumpe zwischen den zwei Wärmetauschern befindet;
• wobei das Verfahren das Steuern des Kraftstoffmanagementsystems derart, dass zwischen 200 und 600 kJ/m³ an Wärme aus dem Öl in dem primären Kraftstoff-Öl-Wärmetauscher auf den Kraftstoff unter Reiseflugbedingungen übertragen werden, umfasst.

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

• a combustion chamber arranged to burn a fuel; and
• a fuel management system arranged to provide the fuel to the combustion chamber, the fuel management system comprising:
  • two fuel-oil heat exchangers through which oil and the 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 the fuel to the combustion chamber, the fuel pump being located between the two heat exchangers;
• the method comprising controlling the fuel management system such that between 200 and 600 kJ/m ³ of heat is transferred from the oil in the primary fuel-oil heat exchanger to the fuel under cruise conditions.

It is understood that the heat transfer per unit volume (m ³ ) of fuel reaching the combustion chamber is specified to compensate for 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 temperature of fuel as it approaches or enters the combustion chamber compared to a temperature of fuel in a fuel tank of the aircraft, taking into account heat transferred from other sources to fuel reaching the combustion chamber (e.g., heat stored in the secondary fuel-oil heat exchanger). exchanger from which oil is transferred to the fuel, in implementations where this fuel is routed to the combustion chamber rather than back to a fuel tank for later recirculation). Heat absorbed by the fuel from other engine components - e.g. the pump - is expected to be negligible compared to the heat absorbed 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 heat transfer rate normalized for fuel flow variations at cruise.

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

The inventors have recognized that the use of fuels other than traditional kerosene-based jet fuels, such as sustainable aviation fuels, may result in different fuel properties and that parameters may be adjusted under cruise conditions to take advantage of the different fuel properties. For example, some fuels may have higher thermal stability and/or higher heat capacity, allowing for increased heat transfer to the fuel and/or higher operating temperature. In particular, in some implementations, a greater amount of heat may be transferred from the oil to the fuel in the fuel-oil heat exchangers. This may improve cooling of the oil before returning it to the rest of the turbine engine. This, in turn, improves the cooling effect of the oil on the components of the engine through which it flows. Increased fuel temperature may also improve combustion efficiency in the combustor, as described above for other aspects.

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

The fuel management system can be controlled to transfer between 300 and 500 kJ/m ³ of heat from the oil to the fuel under cruise conditions.

The fuel management system can be controlled to transfer between 350 and 450 kJ/m ³ of heat from the oil to the fuel under cruise conditions.

The fuel management system can be controlled to transfer between 250-550 kJ/m ³ , 250-450 kJ/m ³ , 300-450 kJ/m ³ , 300-400 kJ/m ³ , 350-400 kJ/m ³ , 400-450 kJ/m ³ or 400-500 kJ/m ³ of heat from the oil to the fuel 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 include controlling the proportion of fuel returned to the inlet of the primary heat exchanger via the recirculation valve.

The gas turbine engine may include a fuel metering valve (as described above), the fuel metering valve arranged to control the fuel flow rate therethrough and to provide information about the fuel flow therethrough (e.g. based on valve position). The step of controlling the fuel management system may include controlling the fuel metering valve and the recirculation valve based on the 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 pipe arranged to allow a portion 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 portion of fuel that passes through the bypass tube rather than the primary heat exchanger.

The fuel management system may further comprise an oil bypass pipe arranged to allow a portion of the oil to bypass at least one of the primary heat exchanger and the secondary heat exchanger, and may optionally include two oil bypass pipes, 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 tube rather than through the at least one heat exchanger.

According to an eighth aspect, there is provided a gas turbine engine for an aircraft. The gas turbine engine comprises a combustion chamber arranged to combust a fuel and a fuel management system arranged to provide the fuel to the combustion chamber. The fuel management system comprises two fuel-oil heat exchangers arranged to have oil and the fuel flowing 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 deliver 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 from the oil to the fuel in the primary fuel-oil heat exchanger under cruise conditions.

It is intended that all the features described for the seventh aspect can equally apply to the eighth aspect.

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

The following characteristics may apply to any or all of the above aspects.

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

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

The primary fuel-oil heat exchanger may be referred to as the main 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 main fuel-oil heat exchanger.

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

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

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

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

At least a portion of the fuel cannot pass through the secondary fuel-oil heat exchanger. At least a portion of the fuel cannot pass 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 tank rather than provided to the combustion chamber, the portion of fuel exiting the primary fuel-oil heat exchanger bypassing the secondary fuel-oil heat exchanger may be greater than the portion passing through the secondary fuel-oil heat exchanger, optionally by a factor of at least three. A bypass pipe 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 of the fuel used for the 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 between 10% and 30% of the fuel downstream of the primary fuel-oil heat exchanger (i.e., the 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 between 10% and 30% of the fuel in the main fuel path upstream of the secondary fuel-oil heat exchanger. For example, approximately between 10% and 20% of the fuel may be delivered to the secondary fuel-oil heat exchanger. Approximately between 13% and 17% of the fuel may be delivered to the secondary fuel-oil heat exchanger. Thus, more of the fuel generally flows from the primary heat exchanger into 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 portion of the fuel being diverted from this path to flow through the secondary fuel-oil heat exchanger and, optionally, some or all of this portion being returned to the main fuel flow path.

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

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

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

A ratio of heat transfer from the oil to the fuel for the primary and secondary fuel-oil heat exchangers may be approximately between 70:30 and 90:10. The primary fuel-oil heat exchanger may therefore be responsible for 70 to 90% of the heat transfer - it may be called "primary" because it is responsible for the majority of the heat transfer, i.e. it is the primary heat source for heating the fuel before it enters the combustion chamber, and because it is the first fuel-oil heat exchanger reached by the fuel.

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

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

The oil can flow through the secondary fuel-oil heat exchanger before it flows through the primary fuel-oil heat exchanger. The oil cannot pass through components that add heat to it before it leaves the secondary heat exchanger and enters the primary heat exchanger. Thus, under cruise conditions, the temperature of the oil flowing through the secondary fuel-oil heat exchanger can be higher than the temperature of the oil flowing through the primary fuel-oil heat exchanger. The fuel can flow through the primary fuel-oil heat exchanger before it flows through the secondary fuel-oil heat exchanger.

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

In light of the foregoing, it is understood 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 (optionally via fuel hydraulic components or other fuel utilizing systems) rather than, for example, being conveyed to a pump and then optionally recirculated. Further, the shorter fuel path of fuel exiting the secondary heat exchanger compared to that exiting the primary heat exchanger may reduce the number of components with which the fuel interacts at its highest temperature, which may improve component longevity.

Furthermore, placing the fuel pump between the primary and secondary fuel-oil heat exchangers can help preserve the life of the fuel pump. The temperature of the fuel may be elevated as it exits the primary heat exchanger and even further elevated as it exits the secondary heat exchanger. Placing the fuel pump between the two, and especially before the secondary heat exchanger, reduces the fuel pump's exposure to the highest temperature fuel stream and therefore reduces the damage associated with 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 flow of fuel may include adjusting the rate of fuel flow through one or more of the heat exchangers.

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

Controlling the fuel management system may include changing parameters of the 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 that air-to-oil heat exchanger. A dedicated controller may be used to control the fuel management system, or the EEC may be used.

The fuel flow may be controlled for any of the third to 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 form part of an aircraft. The aircraft may include a fuel tank and a fuel supply pump configured to supply a fuel flow to the fuel management system of the engine itself. The fuel supply 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 delivery pump, it is a part of the engine itself. The main fuel pump is located downstream of the primary fuel-oil heat exchanger. The primary fuel pump is located upstream of the outlet of the secondary fuel-oil heat exchanger and generally also upstream of the inlet to the secondary fuel-oil heat exchanger. One or more auxiliary fuel pumps may be located at any suitable location along the fuel flow path.

As discussed above, at least a portion of the fuel flowing through the secondary fuel-oil heat exchanger may be provided to aircraft servo mechanisms. At least a portion of this fuel may then be returned to an aircraft fuel tank or to 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 earlier point on the fuel flow path; for example, back to the inlet of the primary or secondary fuel-oil heat exchanger, or may be recombined with other fuel approaching the combustion chamber.

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

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

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

The oil tank may be configured to contain a volume of oil. The oil tank may be configured to contain any suitable volume of oil. The oil tank may be arranged to remove gases from the oil in the first oil tank. Oil exiting the oil tank may pass through a filter, strainer or the like.

A second oil pump may be located between the first oil tank and the primary or secondary fuel-oil heat exchanger. The second oil pump may be described as a transfer pump. The second oil pump may be configured to pump oil from the first oil tank 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 above, the oil may first flow through the secondary fuel-oil heat exchanger and then through the primary fuel-oil heat exchanger, such that the opposite flow path to that of the fuel is present.

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 recirculation 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 the 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 portion of the oil to bypass the or each air-to-oil heat exchanger. As part of the fuel management system control, the bypass percentage can be adjusted to decrease or increase the amount of heat in the oil available to the fuel.

Under cruise conditions, and for all aspects described herein, the oil stream entering one of the fuel-oil heat exchangers may have a higher average temperature than the fuel stream entering the same fuel-oil heat exchanger under cruise conditions. In this way, thermal energy may be transferred from the oil stream to the fuel stream flowing through the one or more fuel-oil heat exchangers under cruise conditions. As such, the oil exiting the respective heat exchanger may have a lower temperature than the oil entering that heat exchanger under cruise conditions.

It is understood that in any of the examples described herein, an intermediate heat transfer fluid could be used instead of direct oil-to-fuel heat transfer, and that the term "fuel-to-oil heat exchanger" may include heat exchangers that include such an intermediate fluid. Similarly, an oil-to-oil, fuel-to-air, or oil-to-air heat exchanger may utilize an intermediate heat transfer fluid in some implementations.

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

A gas turbine engine according to the present disclosure may include an engine core including 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 disposed upstream of the engine core. Alternatively, in some examples, the gas turbine engine may include a fan disposed downstream of the engine core, for example, when the gas turbine engine has an open rotor or is a turboprop engine (in which case the fan may be referred to as a propeller).

If the gas turbine engine has an open rotor or is a turboprop engine, the gas turbine engine may comprise two counter-rotating propeller stages connected by a shaft to and driven by a free power turbine. The propellers may rotate in opposite directions so that one rotates clockwise and the other anti-clockwise about the axis of rotation of the engine. Alternatively, the gas turbine engine may comprise a propeller stage and a nozzle stage configured downstream of the propeller stage. The nozzle stage may have a variable pitch. Accordingly, high pressure, intermediate pressure and free power turbines may respectively drive high and intermediate pressure compressors and propellers through suitable connecting shafts. The propellers may thus provide the majority of the propulsion thrust.

If the gas turbine engine has an open rotor or is a turboprop engine, one or more of the propeller stages may be driven by a gearbox. The gearbox 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 connected directly to the fan drive turbine via a core shaft, for example, without a gearbox. In such a direct drive turbofan engine, the fan may be arranged to rotate at the same speed as the fan drive turbine. For example only, 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 include a second turbine and a second core shaft connecting the second turbine to the compressor. The second turbine, compressor, and second core shaft may be arranged to rotate at a higher 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 engine. In such an arrangement, the engine includes a fan driven via a gearbox. Accordingly, such a gas turbine engine may include a gearbox that receives input from the core shaft and outputs output to the fan to drive the fan at a lower speed than the core shaft. The drive to the gearbox may be directly from the core shaft or indirectly from the core shaft, for example via a spur shaft and/or a gear. The core shaft may rigidly connect the turbine and compressor so 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 number of shafts connecting turbines and compressors, for example, one, two, or three shafts. For 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, the second compressor, and the second core shaft may be arranged to rotate at a higher 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 flow from the first compressor (e.g., directly receive it, for example via a generally annular channel).

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

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

The transmission may have any desired gear ratio (defined as the speed of the input shaft divided by the 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 on the order 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, for example, be between any two of the values recited in the preceding sentence. By way of example only, the transmission may be a "star gear" with a gear ratio in the range of 3.1 or 3.2 to 3.8. Further by way of example only, the transmission may be a "star gear" with a gear ratio in the range of 3.0 to 3.1. Further purely by way of example, the transmission may be a "planetary transmission" with a gear ratio in the range of 3.6 to 4.2. In some arrangements, the gear ratio may be outside these ranges.

In any gas turbine engine described and/or claimed herein, fuel of a given composition or mixture is provided to a combustor, which may be provided downstream of the fan and compressor(s) with respect to the flow path (e.g., axially downstream). For example, the combustor may be located directly downstream of the second compressor (e.g., at the exit thereof) where a second compressor is provided. As another example, flow at the exit to the combustor may be provided to 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 (for example the first compressor and the second compressor as described above) may comprise any number of stages, for example multiple stages. Each stage may comprise a row of rotor blades and a row of stator blades, which may be variable stator blades (since their angle of incidence may be variable). The row of rotor blades and the row of stator blades may be axially offset from one another. For example, the gas turbine engine may be a direct drive turbofan 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 another example, the gas turbine engine may be a "geared" gas turbine engine (in which the fan is driven by a first core shaft via a reduction gear) that includes 11, 12, or 13 compressor stages (in addition to the fan). Such an engine may include 3 or 4 stages in the first (or "low pressure") compressor and 8 or 9 stages in the second (or "high pressure") compressor. As another example, the gas turbine engine may be a "geared" gas turbine engine with 4 stages in the first (or "low pressure") compressor and 10 stages in the second (or "high pressure") compressor.

The or each turbine (for example the first turbine and the second turbine as described above) may comprise any number of stages, for example 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 one another. The second (or "high pressure") turbine may comprise 2 stages in any arrangement (for example, regardless of whether it is a geared or direct drive engine). The gas turbine engine may be a direct drive turbofan engine comprising a first (or "low pressure") turbine with 5, 6 or 7 stages Alternatively, the gas turbine engine may be a "geared" gas turbine engine having a first (or "low pressure") turbine with 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-swept location or 0% span position to a tip at a 100% span position. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be less than (or on the order 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 radius of the fan blade at the hub to the radius of the fan blade at the tip 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 0.28 to 0.32 or 0.29 to 0.30. These ratios may be generally referred to as the hub-to-tip ratio. The radius at the hub and the radius at the tip may both be measured at the leading edge (or axially forwardmost part) of the blade. The hub-to-tip ratio, of course, refers to the gas-swept section of the fan blade, i.e., the section radially outward from any platform.

The radius of the fan can be measured between the centerline of the engine and the tip of a fan blade at its leading edge. The fan diameter (which can be simply twice the fan radius) can be larger than (or on the order 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 fan diameter can be in an inclusive range bounded by any two of the values in the preceding sentence (i.e., the values can form upper or lower limits), for example, in the range of 240 cm to 280 cm or 330 cm to 380 cm. Purely as a non-limiting example, the fan diameter can be in the range of 170 cm to 180 cm, 190 cm to 200 cm, 200 cm to 210 cm, 210 cm to 230 cm, 290 cm to 300 cm, or 340 cm to 360 cm.

The speed of the fan may vary with use. Generally, the speed is lower for larger diameter fans. As a purely non-limiting example, the speed of the fan under cruise conditions may be less than 3500 rpm, for example less than 2500 rpm, for example less than 2300 rpm. As a purely further non-limiting example, the speed of the fan under cruise conditions for a "geared" gas turbine engine having a fan diameter in the range of 200 cm to 210 cm, the speed of the fan under cruise conditions may be in the range of 2750 to 2900 rpm. As a purely further non-limiting example, the speed of the fan under cruise conditions for a "geared" gas turbine engine having a fan diameter in the range of 210 cm to 230 cm. As a further non-limiting example, the speed of the fan under 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. As a further non-limiting example, the speed of the fan under 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. As a further non-limiting example, the speed of the fan under 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.

In use of the gas turbine engine, the fan (with associated fan blades) rotates about an axis of rotation. This rotation causes the tip of the fan blade to move with a velocity U _tip . The work done by the fan blades on the flow results in an enthalpy rise dH of the flow. A fan tip loading can be defined as dH/U _tip ² , where dH is the enthalpy rise (e.g. the mean 1-D enthalpy rise) across the fan and U _tip is the (translational) velocity of the fan tip, for example at the leading edge of the tip (which can be defined as the fan tip radius at the leading edge multiplied by the angular velocity). The fan tip loading under cruise conditions may be greater than (or of the order 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 fan tip loading may be in an inclusive range bounded by any two of the values in the preceding sentence (ie the values may form upper or lower limits), for example in the range 0.28 to 0.31 or 0.29 to 0.3 (for example for a geared gas turbine engine).

Gas turbine engines according to the present disclosure may have a desired bypass ratio (BPR), wherein the bypass ratio ratio is defined as the ratio of the mass flow rate of the flow through the bypass channel to the mass flow rate of the flow through the core. In some arrangements, the bypass ratio under cruise conditions may be greater than (or on the order of) one of the following: 9, 9, 9.5, 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5, or 20. The bypass ratio under cruise conditions 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 12 to 16, 13 to 15, or 13 to 14. Purely as a non-limiting example, the bypass ratio under cruise conditions of a direct drive gas turbine engine according to the present disclosure may be in the range of 9:1 to 11:1. As a further non-limiting example, the bypass ratio under 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 located radially outward of the core engine. The radially outer surface of the bypass duct may be defined by a nacelle and/or a fan casing.

The overall pressure ratio (OPR) of a gas turbine engine as described and/or claimed herein may be defined as the ratio of the ram pressure at the exit of the ultrahigh pressure compressor (before entering the combustor) to the ram pressure upstream of the fan. As a non-limiting example, the total pressure ratio of a gas turbine engine as described and/or claimed herein under cruise conditions may be greater than (or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The total pressure ratio 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 50 to 70. Purely as a non-limiting example, the total pressure ratio under cruise conditions of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 40 to 45. By way of non-limiting example only, the overall pressure ratio under cruise conditions of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 45 to 55. By way of non-limiting example only, the overall pressure ratio under cruise conditions of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 50 to 60. By way of non-limiting example only, the overall pressure ratio under 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 through the engine. In some examples, the specific thrust for a given thrust condition may depend on the specific composition of fuel provided to the combustion chamber. Under cruise conditions, the specific thrust of an engine described and/or claimed herein may be less than (or of the order of magnitude 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 (ie, the values may form upper or lower limits), for example in the range 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 compared to conventional gas turbine engines. By way of non-limiting example only, the specific thrust of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 90 Nkg ^-1 s to 95 Nkg ^-1 s. By way of non-limiting example only, the specific thrust of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 80 Nkg ^-1 s to 90 Nkg ^-1 s. By way of non-limiting example only, the specific thrust of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 70 Nkg ^-1 s to 90 Nkg ^-1 s. Purely as a non-limiting example, the specific thrust of a direct drive gas turbine engine with 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. Purely as a non-limiting example, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust of at least one of (or on the order of) 100 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 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). Purely as a non-limiting example For example, 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. Purely as a 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. Purely as a 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. Purely as a 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. Purely as a non-limiting example, the maximum thrust of a direct drive gas turbine engine with a fan diameter in the range 300 cm to 340 cm may be in the range 370 kN to 500 kN. The above thrust may be the maximum net thrust at standard atmospheric conditions at sea level at +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 inlet to the high pressure turbine may be particularly high. This temperature, which may be referred to as the TET, may be measured at the exit to the combustor, for example immediately upstream of the first turbine vane, which may itself be referred to as the nozzle vane. In some examples, the TET for a given thrust condition may depend on the specific composition of fuel provided to the combustor. Under cruise conditions, the TET may be at least one of (or on the order of) the following: 1400 K, 1450 K, 1500 K, 1520 K, 1530 K, 1540 K, 1550 K, 1600 K, or 1650 K. Thus, purely as a non-limiting example, under cruise conditions of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm, the TET may be in the range of 1540 K to 1600 K. Purely as a non-limiting example, under cruise conditions of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm, the TET may be in the range of 1590 K to 1650 K. By way of non-limiting example only, the TET under 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. By way of non-limiting example only, the TET under 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. By way of non-limiting example only, the TET under 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 under cruise conditions 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 1530 K to 1600 K. The maximum TET in use of the engine may, for example, be at least one of (or on the order of): 1700 K, 1750 K, 1800 K, 1850 K, 1900 K, 1950 K, 2000 K, 2050 K, or 2100 K. Thus, purely as a non-limiting example, the maximum TET of a geared gas turbine engine with a fan diameter in the range of 200 cm to 210 cm may be in the range of 1890 K to 1960 K. By way of non-limiting example only, the maximum TET of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 1890 K to 1960 K. By way of non-limiting example only, the maximum TET of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 1890 K to 1960 K. By way of non-limiting example only, the maximum TET of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 1935 K to 1995 K. By way of non-limiting example only, the maximum TET of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm 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 previous sentence (i.e., the values may form upper or lower limits), for example in the range 1800 K to 1950 K or 1900 K to 2000 K. The maximum TET may, for example, occur at a high thrust condition, for example at a maximum lift-off (MTO) condition.

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

A fan as described and/or claimed herein may include a central portion from which the fan blades may extend, for example in a radial direction. The fan blades may be attached to the central portion in any desired manner. For example, each fan blade may include a fastener engageable with a corresponding slot in the hub (or disc). Such a fastener may, by way of example only, be in the form of a dovetail that is insertable and/or snappable into a corresponding slot in the hub/disc to secure the fan blade to the hub/disc. As a further example, the fan blades may be formed integrally with a central portion. Such an arrangement may be referred to as a blade disc or blade ring. Any suitable method may be used to manufacture such a blade disc 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 attached to the hub/disk 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 enable the exit area of the bypass duct to be varied in 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 include any number of fan blades, for example, 14, 16, 18, 20, 22, 24, or 26 fan blades. If the fan blades include a carbon fiber composite body, there may be 16 or 18 fan blades. If the fan blades include a metallic body (for example, an aluminum-lithium or titanium alloy), there may be 18, 20, or 22 fan blades.

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

In this regard, ground idle may refer to a phase of engine operation in which the aircraft is stationary and in contact with the ground, but with a requirement that the engine be running. During idle, the engine may produce between 3% and 9% of the engine's available thrust. In further non-limiting examples, the engine may produce between 5% and 8% of the engine's available thrust. In further non-limiting examples, the engine may produce between 6% and 7% of the engine's available thrust. Taxiing may refer to a phase of engine operation in which the aircraft is propelled along the ground by thrust generated by the engine. During taxiing, the engine may produce between 5% and 15% of the available thrust. In further non-limiting examples, the engine may produce between 6% and 12% of the engine's available thrust. In further non-limiting examples, the engine may produce between 7% and 10% of the engine's available thrust. Takeoff may refer to a phase of engine operation in which the aircraft is propelled by thrust generated by the engine. At an initial stage within the takeoff phase, the aircraft may be propelled while the aircraft is in contact with the ground. At a later stage within the takeoff phase, the aircraft may be propelled while the aircraft is not in contact with the ground. During takeoff, the engine may produce between 90% and 100% of the available thrust. In further non-limiting examples, the engine may produce between 95% and 100% of the available thrust. In further non-limiting examples, the engine may produce 100% of the available thrust.

Climb may refer to a phase of engine operation in which the aircraft is propelled by the thrust generated by the engine. During climb, the engine may engine may produce between 75% and 100% of available thrust. In further non-limiting examples, the engine may produce between 80% and 95% of available thrust. In further non-limiting examples, the engine may produce between 85% and 90% of available thrust. In this regard, climb may refer to an operational phase 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 at which a relative increase in altitude is required, which may require additional engine thrust demand.

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

In other words, the cruise conditions for a given aircraft gas turbine engine are defined as the operating point of the engine which 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 to be attached at a given mid-cruise Mach number) under mid-cruise atmospheric conditions (defined by the International Standard Atmosphere according to ISO 2533 at cruise altitude). For any given aircraft gas turbine engine, the mid-cruise thrust, atmospheric conditions and Mach number are known and thus the operating point of the engine under cruise conditions is clearly defined.

For example only, forward speed under cruise conditions 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 on the order of Mach 0.8, on 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 conditions. For some aircraft, the cruise conditions may be outside these ranges, for example below Mach 0.7 or above Mach 0.9.

By way of example only, the cruise conditions may correspond to standard atmosphere 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 (about 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 (about 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 in the order of 11,000 m. Cruise conditions may correspond to standard atmosphere conditions at any given altitude in these areas.

For example only, the cruise conditions may correspond to an engine operating point that provides a known required thrust level (for example, a value in the range of 30 kN to 35 kN) at a forward Mach number of 0.8 and standard atmosphere conditions (according to the International Standard Atmosphere) at an altitude of 38,000 feet (11,582 m). For another example only, the cruise conditions may correspond to an engine operating point that provides a known required thrust level (for example, a value in the range of 50 kN to 65 kN) at a forward Mach number of 0.85 and standard atmosphere conditions (according to the International Standard Atmosphere) at an altitude of 35,000 feet (10,668 m).

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

Furthermore, one of ordinary skill in the art would readily recognize that either or both of descent and 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 further non-limiting examples, the engine may produce between 25% and 40% of available thrust. In further 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 at which a relative decrease in altitude is required, and which may require a reduced thrust requirement of the engine.

According to one aspect, there is provided an aircraft comprising a gas turbine engine as described and/or claimed herein. The aircraft according to this aspect is the aircraft to which the gas turbine engine is designed to be attached. Accordingly, the cruise conditions according to this aspect correspond to the mid-range flight of the aircraft as defined elsewhere herein.

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

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

Those skilled in the art will understand that, except in the case of mutual exclusion, a feature or parameter described with respect to any of the above aspects may be applied to any other aspect. Furthermore, except in the case of mutual exclusion, any feature or parameter described herein may be applied to any aspect and/or combined with any other feature or parameter described herein.

• 1 eine Querschnittsseitenansicht eines Gasturbinentriebwerks ist;
• 2 eine Querschnittsseitenansicht eines stromaufwärtigen Abschnitts eines Gasturbinentriebwerks aus der Nähe ist;
• 3A eine teilweise weggeschnittene Ansicht eines Getriebes für ein Gasturbinentriebwerk ist;
• 3B eine Querschnittsseitenansicht eines Direktantriebsgasturbinentriebwerks ist, wobei das Direktantriebstriebwerk kein Getriebe wie in 3A gezeigt aufweist;
• 4 ein beispielhaftes Flugzeug zeigt, das zwei Gasturbinentriebwerke umfasst;
• 5 eine schematische Darstellung eines beispielhaften Kraftstoffsystems ist;
• 6A eine schematische Darstellung eines alternativen beispielhaften Kraftstoffsystems, das einen Rezirkulationsregelkreis umfasst, ist;
• 6B eine schematische Darstellung eines weiteren alternativen beispielhaften Kraftstoffsystems, das einen Rezirkulationsregelkreis und die Verwendung von Kraftstoff für Hilfssysteme umfasst, ist;
• 6C eine schematische Darstellung eines weiteren alternativen beispielhaften Kraftstoffsystems ist, das dem in 6b gezeigten ähnlich ist, in welchem jedoch Kraftstoff, der durch den sekundären Wärmetauscher hindurchtritt, optional nach Verwendung in einem oder mehreren Hilfssystemen an die Brennkammer zurückgeführt wird;
• 7 eine schematische Darstellung eines Abschnitts eines beispielhaften Rezirkulationsölsystems ist;
• 8 eine schematische Darstellung eines Abschnitts des beispielhaften Kraftstoffsystems von 5 und des beispielhaften Rezirkulationsölsystems von 7 ist;
• 9 eine schematische Darstellung eines Abschnitts des beispielhaften Kraftstoffsystems von 6C und des beispielhaften Rezirkulationsölsystems von 7 ist;
• 10 ein beispielhaftes Verfahren zum Betreiben eines Gasturbinentriebwerks veranschaulicht;
• 11 ein weiteres beispielhaftes Verfahren zum Betreiben eines Gasturbinentriebwerks veranschaulicht;
• 12 ein weiteres beispielhaftes Verfahren zum Betreiben eines Gasturbinentriebwerks veranschaulicht; und
• 13 ein weiteres beispielhaftes Verfahren zum Betreiben eines Gasturbinentriebwerks veranschaulicht.

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

• 1 is a cross-sectional side view of a gas turbine engine;
• 2 is a close up cross-sectional side view of an upstream section of a gas turbine engine;
• 3A is a partially cutaway view of a gearbox for a gas turbine engine;
• 3B is a cross-sectional side view of a direct drive gas turbine engine, wherein the direct drive engine does not have a gearbox as in 3A shown;
• 4 shows an exemplary aircraft incorporating two gas turbine engines;
• 5 is a schematic representation of an example fuel system;
• 6A is a schematic representation of an alternative example fuel system including a recirculation control loop;
• 6B is a schematic representation of another alternative exemplary fuel system including a recirculation control loop and use of fuel for auxiliary systems;
• 6C is a schematic representation of another alternative exemplary fuel system similar to that shown in 6b shown, but in which fuel passing through the secondary heat exchanger is optionally returned to the combustion chamber after use in one or more auxiliary systems;
• 7 is a schematic representation of a portion of an exemplary recirculation oil system;
• 8th a schematic representation of a portion of the exemplary fuel system of 5 and the exemplary recirculation oil system of 7 is;
• 9 a schematic representation of a portion of the exemplary fuel system from 6C and the exemplary recirculation oil system of 7 is;
• 10 illustrates an exemplary method for operating a gas turbine engine;
• 11 illustrates another exemplary method for operating a gas turbine engine;
• 12 illustrates another exemplary method for operating a gas turbine engine; and
• 13 illustrates another exemplary method for operating a gas turbine engine.

1 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 which generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 includes a core 11 which receives the core airflow A. The engine core 11 includes, in axial flow sequence, 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 gear 30.

In use, the core air stream A is accelerated and compressed by the low pressure compressor 14 and passed into the high pressure compressor 15 where it is further compressed. The compressed air discharged from the high pressure compressor 15 is passed 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 a combustor 16, the terms "combustion equipment 16" and "combustor 16" being used interchangeably herein. The resulting hot combustion products then expand through and thereby drive the high pressure and low pressure turbines 17, 19 before being discharged through the nozzle 20 to provide some propulsive thrust. The high pressure turbine 17 drives the high pressure compressor 15 via a suitable connecting shaft 27. The fan 23 generally acts to apply increased pressure to the bypass air stream B flowing through the bypass duct 22 so that the bypass air stream B is expelled through the bypass exhaust nozzle 18 to generally provide the majority of the propulsive thrust. The epicyclic gear train 30 is a reduction gear train.

An exemplary arrangement for a geared turbofan gas turbine engine 10 is shown in 2 The low-pressure turbine 19 (see 1 ) drives shaft 26 which is coupled to a central pinion or sun gear 28 of epicyclic gear assembly 30. Radially outward from and engaging sun gear 28 are a plurality of planet gears 32 coupled together by a planet carrier 34. Planet carrier 34 constrains planet gears 32 to synchronously orbit sun gear 28 while allowing each planet gear 32 to rotate about its own axis. Planet carrier 34 is coupled to fan 23 by linkages 36 to drive its rotation about engine axis 9. Radially outward from and engaging planet gears 32 is a ring gear or ring gear 38 coupled to a stationary support structure 24 by linkages 40.

It should be noted that the terms "low pressure turbine" and "low pressure compressor" as used herein mean the lowest pressure turbine stages and the lowest pressure compressor stages, respectively (i.e., not including the fan 23) and/or the turbine and compressor stages connected by the lowest speed connecting shaft 26 in the engine (i.e., not including the transmission output shaft that drives the fan 23). In some references, the "low pressure turbine" and "low pressure compressor" referred to herein may alternatively be known as the "intermediate pressure turbine" and "intermediate pressure compressor." Where such alternative nomenclature is used, the fan 23 may be referred to as a first or lowest pressure compression stage.

The epicyclic gear 30 is shown in greater detail in 3A Each of the sun gear 28, the planet gears 32 and the ring gear 38 includes teeth on its circumference for engaging the other gears. For clarity, 3 however, only exemplary portions of the teeth are illustrated. Four planetary gears 32 are illustrated, although it will be apparent to those skilled in the art that more or fewer planetary gears 32 may be provided within the scope of the claimed invention. Practical applications of a planetary gear train 30 generally include at least three planetary gears 32.

This is exemplified in 2 and 3A The illustrated epicyclic gear 30 is of the planetary type, since the planet carrier 34 is coupled to an output shaft via linkage 36, whereby the ring gear 38 is fixed. However, any other suitable type of epicyclic gear 30 may be used. As another example, epicyclic gear 30 may be a star arrangement in which planet carrier 34 is held fixed and ring gear (or ring gear) 38 is allowed to rotate. In such an arrangement, fan 23 is driven by ring gear 38. As another alternative example, gear 30 may be a differential gear in which both ring gear 38 and planet carrier 34 are allowed to rotate.

It is understood that the 2 and 3A is merely exemplary and that various alternatives are within the scope of the present disclosure. For example only, any suitable arrangement for housing the transmission 30 in the engine 10 and/or for connecting the transmission 30 to the engine 10 may be used. As another example, the connections (such as the linkages 36, 40 in the example of 2 ) between the transmission 30 and other parts of the engine 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 another example, any suitable arrangement of the bearings between rotating and stationary parts of the engine (for example, between the input and output shafts of the transmission and the fixed structures, such as the transmission housing) may be used, and the disclosure is not limited to the exemplary arrangement of 2 For example, if the gearbox 30 has a star arrangement (described above), one skilled in the art would readily understand that the arrangement of output and support rods and bearing positions would typically be different from that shown in 2 would deviate from the arrangement shown as an example.

Accordingly, the present disclosure extends to a gas turbine engine having any arrangement of gear types (e.g., star or planetary gears), support structures, input and output shaft arrangement, and bearing locations.

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

Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of compressors and/or turbines and/or an alternative number of connecting shafts. As another example, the gas turbine engine disclosed in 1 has a split flow nozzle 18, 20, meaning that the flow through the bypass duct 22 has its own nozzle 18 separate from and located radially outward from the core engine nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also be applied to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed or combined before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area.

As another example, other gas turbine engines to which the present disclosure may be applied may not have a gearbox for the main shaft(s), but may instead be direct drive engines. A cross-sectional view of such an engine is shown in 3B shown.

With reference to 3B a gas turbine engine is indicated generally at 10 with 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 core 10 and defines both the inlet 12 and the exhaust nozzle 20.

In use, air entering the inlet 12 is accelerated by the fan 23 to create 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 motive thrust. The intermediate pressure compressor 14 compresses the air stream A before discharging the air to the high pressure compressor 15 where further compression takes place.

The compressed air discharged from the high pressure compressor 15 is passed 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 a combustor 16, the terms "combustion equipment 16" and "combustor 16" being used interchangeably herein. The resulting hot combustion products then expand through and thereby drive the high pressure, intermediate pressure and low pressure turbines 17, 19a, 19 before being discharged through the nozzle 20 to provide additional propulsive thrust. The high pressure turbine 17, the intermediate pressure turbine 19a and the low pressure turbine 19 each drive the high pressure compressor 15, the intermediate pressure compressor 14 and the fan 23 through a suitable connecting shaft.

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

For example, while the example described refers to a turbofan engine, the disclosure may be applied to any type of gas turbine engine, such as an open rotor (where the fan stage is not surrounded by a nacelle) or a turboprop engine. In some arrangements, the gas turbine engine 10 may not include a gearbox 30.

The geometry of the gas turbine engine 10 and components thereof is defined by a conventional axis system having an axial direction (aligned with the axis of rotation 9), a radial direction (in the direction from bottom to top in 1 ) and a circumferential direction (perpendicular to the side in the view of 1 ). The axial, radial and circumferential directions are perpendicular to each other.

The fuel F provided 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, also known as sustainable aviation fuel (SAF). In any of the examples provided, the fuel F may include one or more trace elements, including, for example, sulfur, nitrogen, oxygen, inorganics, and metals.

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

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

A range 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 (with the remainder being one or more fossil-based hydrocarbon fuels, such as kerosene, ), with additional compositions pending approval. However, it is expected 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 made from, for example, one or more of synthesis gas (syngas); lipids (e.g., fats, oils, and greases); sugars; and alcohols. Thus, sustainable aviation fuels may include one or both of a lower aromatic and sulfur content relative to fossil-based hydrocarbon fuels. Additionally or alternatively, sustainable aviation fuels may include one or both of a higher isoalkane and cycloalkane content relative to fossil-based hydrocarbon fuels. Thus, in some examples, sustainable aviation fuels may include one or both of a density of between 90% and 98% of that of kerosene and a calorific value of between 101% and 105% of that of kerosene.

Due at least in part to the molecular structure of sustainable aircraft fuels, sustainable aircraft fuels may provide benefits 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 CO2 emissions relative to fossil-based hydrocarbon fuels (e.g., when burned in combustion equipment 16). Accordingly, sustainable aircraft fuels may result in one or both of a relative decrease in specific fuel consumption and a relative decrease in maintenance costs relative to fossil-based hydrocarbon fuels such as kerosene.

4 shows an aircraft 1 with a propulsion system 2 comprising two gas turbine engines 10. The gas turbine engines 10 are supplied with fuel from a fuel supply system on board the aircraft 1. The fuel supply system of the illustrated 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 fluidly connected to one another.

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, with at least one wing fuel tank located in the port wing and at least one wing fuel tank located in the starboard wing to provide balance. All of the tanks 50, 53 are fluidly connected to one another in the example shown, 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 connected fuel tanks. It is understood that this tank arrangement is provided by way of example only and does not limit the scope of this disclosure.

In alternative examples, the wing fuel tanks 53a, 53b may be fluidly connected to each other, but not to the central fuel tank 50, such that two separate fuel sources are present, each of which may be used to contain a different fuel. In other examples, three or more fluidly isolated fuel sources may be provided. A fuel management system, as described below, may receive fuel from one or more of the fuel sources 50, 53 to supply the engine 10.

An exemplary fuel system 1000 comprising a fuel flow path from the fuel tank 50 to the combustion chamber 16 of the gas turbine engine 10 of the aircraft 1 is shown schematically in 5 The fuel system 1000 comprises 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 implementation described, each engine 10 has its own fuel management system 1500. In other implementations, a single fuel management system 1500 may manage the fuel supply to multiple engines and may, for example, duplicate several of the fuel management systems described in 5 elements shown for the other engine(s).

Fuel is pumped by a low pressure fuel supply pump 1002 from the fuel tank 50 (exemplarily selected from the various fuel tanks 50, 53) to the gas turbine engine 10. The fuel then flows through a primary fuel-oil heat exchanger 1004 before it flows through an engine fuel pump 1003. The engine fuel pump 1003 may be described as a main 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 5 From the secondary fuel-oil heat exchanger 1006, the fuel that has flowed therethrough then flows to the combustion chamber 16 of the gas turbine engine 10, where it rejoins fuel that has flowed only through the primary fuel-oil heat exchanger 1004. In alternatives, such as the example shown in 6B (described below), the portion of the fuel that flows through the secondary fuel-oil heat exchanger 1006 may be returned to a tank 50, 53 rather than being delivered to the combustion chamber 16, optionally being used in one or more auxiliary systems 7010 between leaving the secondary fuel-oil heat exchanger 1006 and reaching the tank 50.

The primary fuel-oil heat exchanger 1004 may be described as a main fuel-oil heat exchanger. The secondary fuel-oil heat exchanger 1006 may be described as a servo fuel-oil heat exchanger. The fuel management system 1500 is arranged so that the fuel reaches the primary fuel-oil heat exchanger 1004 before the secondary fuel-oil heat exchanger 1006. In this example, a bypass pipe 1005 is provided to allow a controlled amount of fuel to pass through this pipe 1005 instead of through the primary fuel-oil heat exchanger 1004. Such a bypass pipe 1005 may be provided for one of the heat exchangers 1004, 1006, or both, in various implementations. One or more equivalent bypass tubes may be provided for oil in various implementations so that a controlled amount of the oil passes through this tube 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 configured to also pass an oil flow therethrough. The primary fuel-oil heat exchanger 1004 and the secondary fuel-oil heat exchanger 1006 are configured to allow heat to be transferred between the oil and fuel flowing therethrough. During cruise conditions, the average temperature of the oil flow entering the primary fuel-oil heat exchanger 1004 and the secondary fuel-oil heat exchanger 1006, respectively, 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 stream to a fuel stream flowing therethrough during cruise conditions.

In various configurations of the fuel management system 1500, the oil flows through the secondary fuel-oil heat exchanger 1006 before flowing through the primary fuel-oil heat exchanger 1004, and does not flow through engine components that would increase its temperature therebetween. The oil is therefore hotter when entering the secondary fuel-oil heat exchanger 1006 than when entering the primary fuel-oil heat exchanger 1004. The fuel, on the other hand, flows through the primary fuel-oil heat exchanger 1004 before flowing through the secondary fuel-oil heat exchanger 1006. In this configuration, the temperature of the fuel exiting the secondary fuel-oil heat exchanger 1006 is higher than the temperature of the fuel exiting the primary fuel-oil heat exchanger 1004. The engine fuel pump 1003 is located upstream of the secondary fuel-oil heat exchanger 1006, meaning that it is not exposed to these further elevated fuel temperatures.

The oil may flow through an air-to-oil heat exchanger (not shown) as well as through the primary and secondary fuel-to-oil heat exchangers 1004, 1006 if further heat loss from the oil beyond that absorbed by the fuel is desired. For example, the oil may pass through the air-to-oil heat exchanger after passing through the secondary fuel-to-oil heat exchanger 1006 and before entering the primary fuel-to-oil heat exchanger 1004. The air-to-oil heat exchanger may be the only component (apart from tubing) through which the oil passes between the secondary fuel-to-oil heat exchanger 1006 and the primary fuel-to-oil heat exchanger 1004. In other examples, no air-to-oil heat exchanger may be provided, or an air-to-oil heat exchanger that is provided may be located elsewhere. The oil may be returned to the bearings and/or other engine components it is used to lubricate and/or cool after leaving the primary heat exchanger 1004. The oil may be returned to an oil tank 2002 as next after leaving the primary heat exchanger 1004 or optionally after leaving the bearings and/or other engine components it is used to lubricate and/or cool. described above, before it is then returned to the secondary heat exchanger 1006.

6A shows an alternative exemplary fuel system 6000 as described above with respect to 5 described fuel system 1000, but with a fuel recirculation control loop instead of a fuel bypass pipe 1005 (in alternative implementations, both a recirculation control loop and a bypass pipe may be provided, optionally for the same heat exchanger). This alternative exemplary fuel system 6000 includes a fuel delivery 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 delivery 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 can be described as the main fuel pump.

The fuel system 6000 from 6A differs from the fuel system 1000 of 5 by the fact that the fuel system 6000 of 6A a recirculation valve 6010 located downstream of the primary fuel-oil heat exchanger 1004 but upstream of the combustion chamber 16. The recirculation valve may be located upstream of an outlet 1006b of the secondary fuel-oil heat exchanger 1006, with fuel from this heat exchanger rejoining the main fuel flow path. In the illustrated example, the recirculation valve 6010 is located downstream of the pump 1003. 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 in other examples, the recirculation valve 6010 could be located upstream of the inlet 1006a of the secondary heat exchanger 1006 (e.g., as described below with respect to 6C and 9 In still further implementations, there may be no outlet of the secondary heat exchanger 1006 back to the main fuel flow path to the combustion chamber 16, as described below with respect to 6B described.

The recirculation valve 6010 is arranged to allow a controlled amount of fuel that has passed through the primary heat exchanger 1004 but 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, whereby fuel that has passed through the primary heat exchanger 1004 (and optionally also the pump 1003) is returned to a position in the flow path upstream of the primary heat exchanger 1004 (and optionally also the pump 1003).

Controlling the flow of fuel in the fuel system 6000 may result in desired characteristics of the fuel (such as temperature and flow rate) being provided as it enters the combustion chamber 16. In particular, if the recirculation valve 6010 is located downstream of the pump 1003 and the proportion recirculated is varied, the fuel flow rate to the combustion chamber 16 may be adjusted without changing the speed of the pump 1003 - this may be particularly useful in implementations where the pump speed is directly coupled to the shaft speed of the engine, thereby allowing a variation in the flow rate to the combustion chamber 16 for a given pump speed. In implementations where fuel flowing through the secondary heat exchanger 1006 is not returned to the combustion chamber 16, changing a proportion of fuel delivered to this heat exchanger 1006 may also be used to control the fuel flow rate into the combustion chamber 16.

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

6B illustrates an alternative exemplary fuel system 7000 compared to that in 6A The 7000 fuel system is similar to that shown in 6A , but differs in that at least a portion of the fuel exiting the secondary fuel-oil heat exchanger 1006 is delivered to additional aircraft mechanisms 7010 rather than 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., the fuel tank 50 or the fuel supply pump 1002) for redistribution. No fuel passing through the secondary fuel-oil heat exchanger 1006 is recycled in the example of 6B to the combustion chamber 16 (at least not directly - it is understood that the fuel returned to the fuel tank 50 can later re-enter the engine 10 and then be conveyed to the combustion chamber 16).

6c illustrates another alternative exemplary fuel system 7000a. The fuel system 7000a is similar to the 6B and differs in that the portion of fuel exiting the secondary fuel-oil heat exchanger 1006 is recirculated to merge with the main fuel stream as it approaches the combustion chamber 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 portion directed to the auxiliary systems 7010 may be adjusted as needed. In other implementations, fuel used in one or more such additional aircraft mechanisms 7010 may then be returned to a tank 50 rather than being delivered to the combustion chamber 16.

In the various examples described above, the aircraft 1 includes a recirculating oil system 2000 arranged to supply oil to a plurality of components for lubrication and/or for heat removal. The recirculating oil system includes a closed loop oil system and may, in some implementations, consist of a single closed loop. An example of a closed loop oil system 2000 is schematically shown in 7 shown.

The closed loop oil system 2000 includes an oil tank 2002 adapted to hold a volume of oil. In some implementations, gases are removed from oil within the oil tank 2002 by a vent.

A transfer pump 2004 is configured to pump oil from the oil tank 2002 into the secondary fuel-oil heat exchanger 1006. The average temperature of the oil entering the secondary fuel-oil heat exchanger 1006 is higher than the average temperature of the fuel entering the secondary fuel-oil heat exchanger 1006 under cruise conditions. In the secondary fuel-oil heat exchanger 1006, thermal energy is transferred from the oil stream to the fuel stream. In this way, the average temperature of the oil stream leaving the secondary fuel-oil heat exchanger 1006 is lower than the average temperature of the oil stream entering the secondary fuel-oil heat exchanger 1006. Likewise, in this manner, the average temperature of the fuel leaving 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, heat energy is transferred from the oil stream to the fuel stream. In this way, the average temperature of the oil stream leaving the primary fuel-oil heat exchanger 1004 is lower than the average temperature of the oil stream entering the primary fuel-oil heat exchanger 1004. Likewise, in this way, the average temperature of the fuel leaving 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 conveyed through the engine components 2006, which the oil is intended to cool and/or lubricate. The oil acts as a lubricant and as a coolant within the engine engine components 2006 and absorbs heat from the engine components 2006 during the process. The engine components 2006 may include one or more storage chambers.

Oil from the engine components 2006 collects in the sump 2008. A return pump 2010 is configured to pump oil from the sump 2008 back to the oil tank 2002.

8th 12 schematically shows an exemplary heat exchange system including 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 flow within the gas turbine engine 10. The fuel flow is shown with thick black arrows and the oil flow with thin black arrows.

9 schematically shows an alternative exemplary heat exchange system 9000 including the closed loop oil system 2000 and the fuel supply system 7000a. The systems 7000a, 2000 are substantially similar to those shown in the 6C or 7 where the fuel flow is again shown with thick black arrows and the oil flow with thin black arrows. However, the recirculation valve 6010 is placed slightly differently than in 6C , upstream of the fork in the fuel flow path to the secondary heat exchanger 1006, rather than between the inlet to 1006a and the outlet of 1006b of this heat exchanger 1006. Thus, in the 9 shown implementation, all fuel exiting the pump 1003 passes through the recirculation valve 6010, unlike in the 6B shown embodiment. The control methodology for the valve 6010 can be adapted accordingly.

It is understood that in implementations such as the one in 6B where the fuel flowing through the secondary fuel-oil heat exchanger 1006 is used as a working fluid in auxiliary systems and then returned to a tank 50, 53 rather than being delivered to the combustion chamber 16, the fuel flow line from the secondary heat exchanger 1006 to the combustion chamber 16 can be replaced by a fuel flow line back to the tank 50.

The inventors have recognized that the use of fuels other than traditional kerosene-based jet 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 calorific value, as claimed, may also have greater thermal stability, allowing the fuel to absorb more heat, thereby providing improved oil cooling and/or improved combustion characteristics in the combustion chamber. Recirculation of the fuel through the primary fuel-oil heat exchanger 1004 may allow the fuel to absorb more heat from the oil, thereby increasing the fuel temperature and improving oil cooling.

A recirculation valve 6010 may be located downstream of the fuel pump 1003 and may therefore provide improved flexibility of the fuel flow.

Therefore, a lower flow rate of fuel to the combustion chamber 16 can be provided for a higher calorific value fuel without changing the pump speed by instead increasing the amount of recirculation.

• Schritt 10100: Erhalten/Auswählen eines Kraftstoffs, der dem Gasturbinentriebwerk 10 bereitgestellt werden soll, wobei der Brennwert des Kraftstoffs mindestens 43,5 MJ/kg und optional zwischen 43,5 und 44 MJ/kg beträgt.

10 illustrates an exemplary method 10000 for operating a gas turbine engine 10. The method 10000 includes the steps:

• Step 10100: Obtaining/selecting a fuel to be provided 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 the fuel may include selecting a single fuel. Alternatively, obtaining the fuel may include selecting a fuel mixture.

• Schritt 10200: Bereitstellen des Kraftstoffs an die Brennkammer 16 über das Kraftstoffmanagementsystem 6500.

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

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

As with regard to 6A and 6B discussed, 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 control loop 6010, 6011 arranged to allow a portion of the fuel to pass through the primary heat exchanger 1004 multiple times. An engine fuel pump 1003 is located downstream of the primary fuel-oil heat exchanger 1004 and upstream of the secondary fuel-oil heat exchanger 1006. The fuel management system 6500 is arranged so that the fuel reaches the primary fuel-oil heat exchanger 1004 before the secondary fuel-oil heat exchanger 1006.

• Schritt 10300: Aktives Steuern einer Menge an Kraftstoff, die den primären Wärmetauscher 1004 verlässt, um zu einem Einlass 1004a des primären Wärmetauschers 1004 rezirkuliert zu werden, anstatt zu dem sekundären Kraftstoff-Öl-Wärmetauscher 1006 oder direkt zu der Brennkammer 16 zu strömen, unter Verwendung eines Rezirkulationsventils 6010. Diese Rezirkulation kann auch Kraftstoff, der bereits durch die Triebwerk-Kraftstoffpumpe 1003 hindurchgeströmt ist, zu einer Position stromaufwärts der Triebwerk-Kraftstoffpumpe 1003 zurückführen.

Optionally, the method may further comprise:

• Step 10300: Actively control 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 return fuel that has already passed through the engine fuel pump 1003 to a location upstream of the engine fuel pump 1003.

• • Bedarf an Kernwellendrehzahl und Triebwerksschub;
• • Aktuelle Kraftstofftemperatur und/oder Öltemperatur;
• • Kraftstoffbrennwert;
• • Fan-Drehzahl;
• • Kraftstoffströmungsrate an die Brennkammer (allgemein als WFE bezeichnet - Weight of Main Engine Fuel Flow, Gewicht des Haupttriebwerkskraftstoffstroms); und
• • Drehzahl oder Drehzahloptionen der Haupt-/Triebwerk-Kraftstoffpumpe.

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

• • Core shaft speed and engine thrust requirements;
• • Current fuel temperature and/or oil temperature;
• • Fuel calorific value;
• • Fan speed;
• • Fuel flow rate to the combustion chamber (commonly referred to as WFE - Weight of Main Engine Fuel Flow); and
• • Main/engine fuel pump speed or speed options.

Therefore, for fuels with a higher calorific value, for example, a higher percentage of the fuel leaving the pump can be returned to the pump inlet

In alternative examples, controlling the amount of fuel leaving the primary heat exchanger 1004 for recirculation may not be an active process step - instead, a set, fixed proportion of the fuel may be recirculated. For example, 70%, 75%, 80%, 85%, or 90% of the fuel may be recirculated under cruise conditions. 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 where fuel pump speed is directly coupled to shaft speed (e.g., where a shaft speed to pump speed ratio is fixed or where only certain shaft speed to pump speed ratios are possible, at least under cruise conditions), fuel may be wasted if engine operation is not ceased. Therefore, recirculating more of the fuel may allow a lower fuel rate to reach the combustion chamber while maintaining a constant fuel flow rate through pump 1003.

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

The inventors also recognized that using fuels other than traditional kerosene-based jet fuels, such as sustainable aviation fuels, can result in different fuel properties and that parameters can be adjusted under cruise conditions to take advantage of the different fuel properties. In particular, some fuels can be heated to higher temperatures in the fuel-oil heat exchangers 1004, 1006 than conventional fuels. This can improve cooling of the oil before it is returned to the rest of the turbine engine and/or can improve the combustion efficiency of the fuel. Using the fuel to remove 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) can improve the combustion efficiency of the fuel. cher) provides a more thermally efficient turbine engine. In addition, the improved cooling of the oil can in turn improve the cooling effect of the oil on the engine components through which it flows.

• Schritt 11100: Bereitstellen eines Kraftstoffs an die Brennkammer 16 des Gasturbinentriebwerks 10 über das Kraftstoffmanagementsystem 1500.

11 illustrates an exemplary method 11000 for operating a gas turbine engine 10. The method 11000 includes the steps:

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

• Schritt 11200: Steuern des Kraftstoffmanagementsystems 1500 derart, dass die Kraftstofftemperatur auf mindestens 135 °C bei Eintritt in die Brennkammer 16 unter Reiseflugbedingungen erhöht wird.

As with regard to 5 discussed, 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 located downstream of the primary fuel-oil heat exchanger 1004 and upstream of the secondary fuel-oil heat exchanger 1006. The fuel management system 1500 is arranged so that the fuel reaches the primary fuel-oil heat exchanger 1004 before the secondary fuel-oil heat exchanger 1006.

• Step 11200: Controlling the fuel management system 1500 to increase the fuel temperature to at least 135 °C upon entering the combustion chamber 16 under cruise conditions.

Controlling 11200 the fuel management system 1500 may include controlling the flow of fuel through the heat exchangers - for example, recirculating a variable portion of the fuel through the primary fuel-oil heat exchanger 1004 and/or allowing a variable portion of the fuel to bypass the primary fuel-oil heat exchanger 1004. While 5 shows an implementation with a bypass pipe 1005, but without recirculation and the 6A-6C Implementations without bypass pipe (at least not for the primary heat exchanger - 6A and 6C provide a path 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 tube 6011, it is understood that one or more bypass tubes and/or one or more recirculation tubes may be provided together in various implementations. Further, in some systems, the flow direction within the same tube may be reversible so that it may be used as a bypass tube 1005 or as a recirculation tube 6011. It is understood that the figures are provided by way of example only and are not intended to be limiting.

Controlling 11200 the fuel management system 1500 may include controlling the flow of fuel through the heat exchangers - for example, allowing a variable portion of the fuel to bypass the primary fuel-oil heat exchanger 1004 or the secondary fuel-oil heat exchanger 1006. One or more controllable valves or pumps, and optionally one or more sensors, may be provided as appropriate to enable control of the flow of fuel.

Additionally or alternatively, as for fuel flow, oil flow may be controlled using one or more bypass pipes, if present, allowing oil to bypass one or more heat exchangers 1004, 1006 rather than flowing therethrough. Oil may also be recirculated in some implementations and/or the oil flow rate 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 enable control of oil flow.

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

The fuel flow may be controlled using a recirculation valve 6010 as described above. Alternatively or additionally, the fuel flow may be controlled by using one or more bypass tubes 1005 (as in 5 shown) arranged to allow a portion of the fuel to not pass through one or both heat exchangers, and/or by adjusting the split of % of fuel flowing from the primary fuel-oil heat exchanger 1004 directly to the combustion chamber 16 versus the % of fuel flowing from the primary fuel-oil heat exchanger into the secondary fuel-oil heat exchanger 1006 (and optionally then back to merge with 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 rest of the turbine engine and/or may improve combustion efficiency of the fuel. Positioning the secondary fuel-oil heat exchanger 1006 after the engine fuel pump 1003 may allow higher fuel temperatures to be obtained without compromising the longevity of the fuel pump.

The procedure of 11 can be used in conjunction with the procedure of 10 be used.

The inventors have recognized that, particularly when using fuels other than traditional kerosene-based jet fuels, such as sustainable aviation fuels, fuel properties under cruise conditions can be controlled through careful engine management to improve performance. In this case, a lower viscosity can be obtained, which can affect combustion efficiency, particularly with respect to fuel nozzle spray performance within the combustion chamber 16. Fuel nozzle spray performance affects the combustion efficiency of the fuel. A lower viscosity of the fuel under cruise conditions can conducive to a more efficient engine. Fuel flow can be optimized to improve turbine engine efficiency, which the lower viscosity of the fuel can contribute to.

• Schritt 12100: Bereitstellen eines Kraftstoffs an die Brennkammer 16 des Gasturbinentriebwerks 10 über das Kraftstoffmanagementsystem 1500.

12 illustrates an exemplary method 12000 for operating a gas turbine engine 10. The method 12000 includes the steps:

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

• Schritt 12200: Steuern des Kraftstoffmanagementsystems 1500 derart, dass die Kraftstoffviskosität auf weniger als 0,58 mm2/s bei Eintritt in die Brennkammer 16 unter Reiseflugbedingungen angepasst wird.

As with regard to 5 discussed, 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 located downstream of the primary fuel-oil heat exchanger 1004 and upstream of the secondary fuel-oil heat exchanger 1006. The fuel management system 1500 is arranged so that the fuel reaches the primary fuel-oil heat exchanger 1004 before the secondary fuel-oil heat exchanger 1006.

• Step 12200: Control the fuel management system 1500 to adjust the fuel viscosity to less than 0.58 mm2/s entering the combustion chamber 16 under cruise conditions.

By carefully managing the engine to achieve a lower viscosity of fuel entering the combustion chamber 16, combustion efficiency can be improved, particularly with respect to fuel nozzle spray performance within the combustion chamber. Fuel flow, oil flow, and/or one or more other heat exchange characteristics can be adjusted to provide this viscosity difference.

Controlling 12200 the fuel management system 1500 may therefore include controlling the flow of fuel and/or oil 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 or oil to pass through the primary and/or secondary fuel-oil heat exchangers 1004, 1006 as described above with respect to 11 described, bypasses.

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

The procedure of 12 can be used in conjunction with the procedure of 10 and/or the procedure of 11 be used.

The inventors have recognized that the use of fuels other than traditional kerosene-based jet fuels, such as sustainable aviation fuels, may result in different fuel properties and that parameters may be adjusted under cruise conditions to take advantage of the different fuel properties. For example, some fuels may have higher thermal stability and/or higher heat capacity, allowing for increased heat transfer to the fuel and/or higher operating temperature. In particular, in some implementations, a greater amount of heat may be transferred from the oil to the fuel in the fuel-oil heat exchangers 1004, 1006. This may improve cooling of the oil before returning it to the rest of the turbine engine. This, in turn, improves the cooling effect of the oil on the components of the engine through which it flows. Increased fuel temperature ture can also improve the combustion efficiency in the combustion chamber 16.

• Schritt 13100: Bereitstellen eines Kraftstoffs an die Brennkammer 16 des Gasturbinentriebwerks 10 über das Kraftstoffmanagementsystem 1500.

13 illustrates an exemplary method 13000 for operating a gas turbine engine 10. The method 13000 includes the steps:

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

As with regard to 5 discussed, 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 located downstream of the primary fuel-oil heat exchanger 1004 and upstream of the secondary fuel-oil heat exchanger 1006. The fuel management system 1500 is arranged so that the fuel reaches the primary fuel-oil heat exchanger 1004 before the secondary fuel-oil heat exchanger 1006.

Step 13200: Control the fuel management system 1500 such that 200 to 600 kJ/m3 of heat is transferred from the oil to the fuel in the primary fuel-oil heat exchanger under cruise conditions (where the heat transfer is 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 fuel path - i.e. downstream of the primary heat exchanger but upstream of the secondary heat exchanger.

Controlling 13200 the fuel management system 1500 may therefore include controlling the flow of fuel and/or oil 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 or oil to pass through the primary and/or secondary fuel-oil heat exchangers 1004, 1006, as described above with respect to 11 described. The fuel management system 1500 may therefore include one or more additional components not shown in the figures - such as bypass pipes, recirculation pipes, controllable pumps, and/or controllable valves to enable or facilitate this control. One or more sensors - e.g., temperature and flow rate sensors - may be used to provide feedback and direct fuel and oil management. 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 passed through the secondary fuel-oil heat exchanger 1004.

For example, a fuel metering valve (FMV) 6014, as shown in 6B illustrated, may be provided to control the fuel delivered to the combustion chamber 16. The position of the valve 6014 (i.e. how far it is open/closed) may be controlled and may also be used to determine the fuel flow rate since a controller of (or communication with) the fuel management system has access to a record of the relationship between valve position and flow rate. The FMV 6014 and the recirculation valve 6010 may cooperate to deliver a required flow of fuel to the combustion chamber 16. The FMV position may be actively controlled in a closed loop, with the excess fuel being returned to the pump 1003 via the recirculation valve and pipe 6010, 6011. The FMV 6014 may, in some implementations, be located at an outlet 1006b of the secondary fuel-oil heat exchanger 1006, downstream of the recirculation valve 6010 on the main fuel flow path through the engine 10, as in 6C shown.

By adjusting heat exchange parameters under cruise conditions, more efficient engine performance can be provided.

The procedure of 13 may be used in connection with the procedure by any or all of the 10 to 12 Similarly, an FMV 6014 may be used in any or all of the disclosed methods.

It is to 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 in the case of mutual exclusion, each of the features may be used separately or in combination with any other features, and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.

Claims (18)

Method (12000) for operating a gas turbine engine (10), wherein the gas turbine engine plant (10) comprising: a combustion chamber (16) arranged to combust a fuel; and a fuel management system arranged to provide the fuel to the combustion chamber (16), the fuel management system comprising: two fuel-oil heat exchangers through which oil and the 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 the fuel to the combustion chamber (16), the fuel pump (1003) being located between the two heat exchangers; the method comprising controlling (12200) the fuel management system such that the fuel viscosity is adjusted to less than 0.58 mm ² /s upon entering the combustion chamber (16) under cruise conditions. Procedure (12000) according to Claim 1 wherein the oil enters the fuel-oil heat exchangers (1004, 1006) at a higher temperature than the fuel, such that the fuel-oil heat exchangers are arranged to transfer heat from the oil to the fuel. Procedure (12000) according to Claim 1 the method comprising controlling (12200) the fuel management system such that the fuel viscosity is adjusted to between 0.35 mm ² /s and 0.53 mm ² /s upon entering the combustion chamber (16) under cruise conditions. Procedure (12000) according to Claim 1 wherein: the fuel management system further comprises: 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 (12200) 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). Procedure (12000) according to Claim 1 wherein: the fuel management system further comprises: a bypass pipe (1005) arranged to allow a portion of the fuel to bypass the primary fuel-oil heat exchanger (1004); and wherein controlling (12200) the fuel management system comprises controlling the portion of the fuel that passes through the bypass pipe rather than the primary fuel-oil heat exchanger (1004). Procedure (12000) according to Claim 1 wherein: the fuel management system further comprises an oil bypass pipe arranged to allow a portion of the oil to bypass at least one of the primary fuel-oil heat exchanger (1004) and the secondary fuel-oil heat exchanger (1006); and wherein controlling (12200) the fuel management system comprises controlling the portion of the oil that passes through the oil bypass pipe rather than through the at least one heat exchanger. Procedure (12000) according to Claim 4 wherein the gas turbine engine (10) comprises a fuel metering valve (6014) downstream of the recirculation valve (6010) arranged to control the fuel flow rate therethrough and to provide information about the fuel flow therethrough, and wherein controlling (13200) the fuel management system comprises controlling the fuel metering valve (6014) and the recirculation valve (6010) based on the information provided by the fuel metering valve (6014). Procedure (12000) according to Claim 1 the method comprising controlling (12200) the fuel management system such that the fuel temperature is increased to between 150 °C and 170 °C upon entering the combustion chamber (16) under cruise conditions. Procedure (12000) according to Claim 1 , with between 10% and 20% of the fuel being delivered to the secondary fuel-oil heat exchanger (1006). Procedure (12000) according to Claim 1 , wherein a ratio of heat transfer from the oil to the fuel for the primary and secondary fuel-oil heat exchangers (1004, 1006) is between 70:30 and 90:10. Procedure (12000) according to 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). Gas turbine engine (10) for an aircraft (1), the gas turbine engine (10) comprising: a combustion chamber (16) arranged to combust a fuel; and a fuel management system arranged to provide the fuel to the combustion chamber (16), the fuel management system comprising: two fuel-oil heat exchangers through which oil and the fuel flow, 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 the fuel to the combustion chamber (16), the fuel pump (1003) located between the two fuel-oil heat exchangers; the fuel management system arranged to adjust the fuel viscosity to less than 0.58 mm ² /s upon entering the combustion chamber (16) under cruise conditions. Gas turbine engine (10) to Claim 12 wherein the fuel management system is arranged to adjust the fuel viscosity to between 0.35 mm ² /s and 0.53 mm ² /s upon entering the combustion chamber (16) under cruise conditions. Gas turbine engine (10) to Claim 12 , the fuel management system further comprising 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). Gas turbine engine (10) to Claim 12 wherein the fuel management system is arranged to control the fuel flow through the heat exchangers to increase the fuel temperature to between 150 °C and 170 °C upon entering the combustion chamber (16) under cruise conditions. Gas turbine engine (10) to Claim 12 wherein the secondary fuel-oil heat exchanger (1006) is a servo fuel-oil heat exchanger. Gas turbine engine (10) to Claim 12 , wherein a ratio of heat transfer from the oil to the fuel for the primary and secondary fuel-oil heat exchangers (1004, 1006) is between 70:30 and 90:10. Gas turbine engine (10) to Claim 12 , wherein the fuel management system is arranged such that 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).

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