DE102023135099A1 - Method for operating a gas turbine engine - Google Patents

Abstract

A method (11000) of operating a gas turbine engine (10) is disclosed, the gas turbine engine (10) comprising an engine core (11) comprising a combustion chamber (16) arranged to combust a fuel; a turbine (19), the turbine comprising a plurality of turbine blades (19'); a compressor (14) arranged to be used as a cooling air source for the turbine blades (19'); and an inductor (1901) arranged to accelerate and direct the cooling air onto the turbine blades (19') and comprising a plurality of airflow passages (1903) and a modulation valve (1902) arranged to allow or block cooling air flow into a subset of the passages; and a fuel management system (1500) arranged to provide the fuel to the combustion chamber (16), the fuel management system (1500) comprising: two fuel-oil heat exchangers (1004,1006) through which the oil and the fuel flow, the fuel-oil 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). The method (11000) includes using (11100) the modulation valve (1902) to adjust the cooling air flow based on the turbine inlet temperature; and transferring (11200) 200 - 600 kJ/m3 of heat to the fuel from the oil in the primary fuel-oil heat exchanger (1004) under cruise conditions.

Description

The present disclosure relates to aircraft propulsion systems and methods of operating aircraft that include managing different fluids.

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. These 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 processes for operating gas turbine engines.

• einen Triebwerkskern, umfassend eine Brennkammer, die angeordnet ist, um einen Kraftstoff zu verbrennen; eine Turbine, wobei die Turbine eine Vielzahl von Turbinenschaufeln umfasst; einen Verdichter, der angeordnet ist, um als eine Kühlluftquelle für die Turbinenschaufeln verwendet zu werden; und einen Induktor, der angeordnet ist, um die Kühlluft auf die Turbinenschaufeln zu beschleunigen und zu richten und eine Vielzahl von Luftstromdurchgängen und ein Modulationsventil umfasst, das angeordnet ist, um den Kühlluftstrom in eine Teilmenge der Durchgänge zu steuern; und
• ein Kraftstoffverwaltungssystem, umfassend: zwei Kraftstoff-Öl-Wärmetauscher, durch die das Öl und der Kraftstoff 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 umfasst;
• wobei das Verfahren umfasst:
  • Verwenden des Kraftstoffverwaltungssystems, um einen Kraftstoff mit einem Schwefelgehalt von weniger als 30 ppm an die Brennkammer bereitzustellen; und
  • Steuern des Modulationsventils, um den Kühlluftstrom basierend auf der Turbineneintrittstemperatur einzustellen.

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

• an engine core comprising a combustion chamber arranged to combust a fuel; a turbine, the turbine comprising a plurality of turbine blades; a compressor arranged to be used as a source of cooling air for the turbine blades; and an inductor arranged to accelerate and direct the cooling air onto the turbine blades and comprising a plurality of airflow passages and a modulation valve arranged to control the flow of cooling air into a subset of the passages; and
• a fuel management system comprising: two fuel-oil heat exchangers through which the oil and the fuel flow, the fuel-oil 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 method comprising:
  • Using the fuel management system to deliver a fuel having a sulfur content of less than 30 ppm to the combustion chamber; and
  • Controlling the modulating valve to adjust the cooling air flow based on the turbine inlet temperature.

The primary fuel-oil heat exchanger and the secondary fuel-oil heat exchanger can be referred to as the primary heat exchanger and the secondary heat exchanger, respectively, for abbreviation.

The modulating valve can allow or block airflow (maximum flow or no flow) in a given passage or subset of passages, or can control a flow rate so that the flow rate can vary in each passage or subset of passages.

The inventors have recognized that the use of fuels other than conventional 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. The sulfur content of the fuel should therefore be taken into account when considering heat transfer from the oil to the fuel, as the sulfur content is related to the thermal stability of the fuel and a less thermally stable fuel can absorb less heat before performance problems occur, while a more thermally stable fuel can absorb more heat and so can be used for more efficient engine cooling.

The fuel may additionally have a density in the range of 760-840 kg/m ³ at 15 °C. The density of the fuel may affect the fuel delivery to and spray pattern of combustion chamber fuel spray nozzles during a reignition procedure and/or during cruise operation - correct selection of a fuel may therefore improve combustion efficiency and/or reignition speed.

The method may include selecting a suitable fuel to use. Selecting the fuel may include selecting a single fuel. Selecting the fuel may include selecting a fuel blend. The selected fuel may be the only fuel on board the aircraft. Thus, selecting the fuel may be performed during refueling of the aircraft. Alternatively, the selected fuel may be one of multiple fuels on board the aircraft or a blend of one or more of the multiple fuels on board the aircraft; selecting may therefore be performed in flight and automated. Selecting may therefore be performed in flight or performed prior to departure.

The fuel may have an aromatic content of less than or equal to 10% by volume. The fuel may have an aromatic content of 5 vol.% or less. The fuel may have an aromatic content of 1 vol.% or less.

The fuel may have a sulfur content of less than 15 ppm.

The fuel may have a density in the range of 760-840 kg/m ³ at 15 °C. The fuel may additionally have a density in the range of 760-800 kg/m ³ at 15 °C.

• einen Triebwerkskern, umfassend eine Brennkammer, die angeordnet ist, um einen Kraftstoff zu verbrennen; eine Turbine, wobei die Turbine eine Vielzahl von Turbinenschaufeln umfasst; einen Verdichter, der angeordnet ist, um als eine Kühlluftquelle für die Turbinenschaufeln verwendet zu werden; und einen Induktor, der angeordnet ist, um die Kühlluft auf die Turbinenschaufeln zu beschleunigen und zu richten und eine Vielzahl von Luftstromdurchgängen und ein Modulationsventil umfasst, das angeordnet ist, um den Kühlluftstrom in eine Teilmenge der Durchgänge zu steuern, wobei das Modulationsventil konfiguriert ist, um den Kühlluftstrom basierend auf der Turbineneintrittstemperatur einzustellen; und
• ein Kraftstoffverwaltungssystem, das angeordnet ist, um den Kraftstoff der Brennkammer bereitzustellen, wobei das Kraftstoffverwaltungssystem umfasst: zwei Kraftstoff-Öl-Wärmetauscher, durch die das Öl und der Kraftstoff strömen, wobei die KraftstoffÖl-Wärmetauscher angeordnet sind, um Wärme zwischen dem Öl und dem Kraftstoff zu übertragen, und einen primären Kraftstoff-Öl-Wärmetauscher und einen sekundären KraftstoffÖl-Wärmetauscher umfasst;
• wobei der Kraftstoff einen Schwefelgehalt von weniger als 30 ppm aufweist.

According to a second aspect, there is provided a method of operating a gas turbine engine on an aircraft, the gas turbine engine comprising:

• an engine core comprising a combustion chamber arranged to combust a fuel; a turbine, the turbine comprising a plurality of turbine blades; a compressor arranged to be used as a cooling air source for the turbine blades; and an inductor arranged to accelerate and direct the cooling air to the turbine blades and comprising a plurality of airflow passages and a modulation valve arranged to control the cooling air flow into a subset of the passages, the modulation valve configured to adjust the cooling air flow based on the turbine inlet temperature; 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 the oil and the fuel flow, the fuel-oil 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;
• wherein the fuel has a sulphur content of less than 30 ppm.

The gas turbine engine can therefore be used to implement the method of the first aspect and can offer the same technical effects and advantages.

The engine core may further include a core shaft connecting the turbine to the compressor. The gas turbine engine may further include a fan disposed upstream of the engine core, the fan being driven (directly or indirectly) by the core shaft.

The gas turbine engine may be a direct drive engine.

The gas turbine engine may further include an auxiliary gearbox and a pump shaft, the auxiliary gearbox receiving drive from the core shaft and outputting drive to the fuel pump via the pump shaft. The pump speed may therefore be linked to the core shaft speed and may have a fixed speed relationship between the core shaft (or in some engine architectures, another intermediate shaft) and the fuel pump via the auxiliary gearbox. In some implementations, the auxiliary gearbox may instead be arranged to provide a fixed number of different speeds for a given core shaft speed.

The gas turbine engine may further include engine bearing chambers. 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 loop such that the oil is cooled by the fuel-oil heat exchangers before entering the engine bearing chambers, with the hotter oil exiting the engine bearing chambers then being returned to the fuel-oil heat exchangers.

It is intended that any of the features described for the first or second aspect may equally apply to the other of the first or second aspect.

• einen Triebwerkskern, umfassend eine Brennkammer, die angeordnet ist, um einen Kraftstoff zu verbrennen; eine Turbine, wobei die Turbine eine Vielzahl von Turbinenschaufeln umfasst; einen Verdichter, der angeordnet ist, um als eine Kühlluftquelle für die Turbinenschaufeln verwendet zu werden; und einen Induktor, der angeordnet ist, um die Kühlluft auf die Turbinenschaufeln zu beschleunigen und zu richten und eine Vielzahl von Luftstromdurchgängen und
• ein Modulationsventil umfasst, das angeordnet ist, um den Kühlluftstrom in eine Teilmenge der Durchgänge zu steuern; und
• ein Kraftstoffverwaltungssystem, das angeordnet ist, um den Kraftstoff der Brennkammer bereitzustellen, wobei das Kraftstoffverwaltungssystem umfasst:
  • zwei Kraftstoff-Öl-Wärmetauscher, durch die das Öl und der Kraftstoff strömen, wobei die KraftstoffÖl-Wärmetauscher angeordnet sind, um Wärme zwischen dem Öl und dem Kraftstoff zu übertragen, und einen primären Kraftstoff-Öl-Wärmetauscher und einen sekundären KraftstoffÖl-Wärmetauscher umfasst;
  • wobei das Verfahren umfasst:
    • Verwenden des Kraftstoffverwaltungssystems, um einen Kraftstoff mit einer Dichte im Bereich von 760-840 kg/m³ bei 15 °C an die Brennkammer bereitzustellen; und
    • Steuern des Modulationsventils, um den Kühlluftstrom basierend auf der Turbineneintrittstemperatur einzustellen.

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

• an engine core comprising a combustion chamber arranged to combust a fuel; a turbine, the turbine comprising a plurality of turbine blades; a compressor arranged to be used as a source of cooling air for the turbine blades; and an inductor arranged to accelerate and direct the cooling air onto the turbine blades and to define a plurality of airflow passages and
• a modulation valve arranged to control the flow of cooling air into a subset of the passages; 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 the oil and the fuel flow, the fuel-oil 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 method comprising:
    • Using the fuel management system to deliver a fuel having a density in the range of 760-840 kg/m ³ at 15 °C to the combustion chamber; and
    • Controlling the modulating valve to adjust the cooling air flow based on the turbine inlet temperature.

The inventors have recognized that the use of fuels other than conventional 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. The density of the fuel should therefore be considered when considering heat transfer from the oil to the fuel, as lower density fuel at 15°C is likely to require less heat input to achieve a desired density for combustion than a higher density fuel at 15°C. It is understood that fuel properties as they enter the combustion chamber can affect engine performance, e.g. due to droplet size and nozzle spray characteristics, affecting fuel-air mixing and combustion efficiency.

The fuel may have a sulfur content of less than 30 ppm.

The fuel may have an aromatic content of less than or equal to 10% by volume.

• einen Triebwerkskern, umfassend eine Brennkammer, die angeordnet ist, um einen Kraftstoff zu verbrennen; eine Turbine, wobei die Turbine eine Vielzahl von Turbinenschaufeln umfasst; einen Verdichter, der angeordnet ist, um als eine Kühlluftquelle für die Turbinenschaufeln verwendet zu werden; und einen Induktor, der angeordnet ist, um die Kühlluft auf die Turbinenschaufeln zu beschleunigen und zu richten und eine Vielzahl von Luftstromdurchgängen und ein Modulationsventil umfasst, das angeordnet ist, um den Kühlluftstrom in eine Teilmenge der Durchgänge zu steuern, wobei das Modulationsventil konfiguriert ist, um den Kühlluftstrom basierend auf der Turbineneintrittstemperatur einzustellen; und
• ein Kraftstoffverwaltungssystem, das angeordnet ist, um den Kraftstoff der Brennkammer bereitzustellen, wobei das Kraftstoffverwaltungssystem umfasst:
  • zwei Kraftstoff-Öl-Wärmetauscher, durch die das Öl und der Kraftstoff 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 umfasst;
  • wobei der Kraftstoff eine Dichte im Bereich von 760-840 kg/m³ bei 15 °C aufweist.

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

• an engine core comprising a combustion chamber arranged to combust a fuel; a turbine, the turbine comprising a plurality of turbine blades; a compressor arranged to be used as a cooling air source for the turbine blades; and an inductor arranged to accelerate and direct the cooling air to the turbine blades and comprising a plurality of airflow passages and a modulation valve arranged to control the cooling air flow into a subset of the passages, the modulation valve configured to adjust the cooling air flow based on the turbine inlet temperature; 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 the 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;
  • the fuel has a density in the range of 760-840 kg/m ³ at 15 °C.

The fuel may have a sulfur content of less than 30 ppm.

• einen Triebwerkskern, umfassend eine Brennkammer, die angeordnet ist, um einen Kraftstoff zu verbrennen; eine Turbine, wobei die Turbine eine Vielzahl von Turbinenschaufeln umfasst; einen Verdichter, der angeordnet ist, um als eine Kühlluftquelle für die Turbinenschaufeln verwendet zu werden; und einen Induktor, der angeordnet ist, um die Kühlluft auf die Turbinenschaufeln zu beschleunigen und zu richten, und eine Vielzahl von Luftstromdurchgängen umfasst, und ein Modulationsventil, das angeordnet ist, um einen Kühlluftstrom in eine Teilmenge der Durchgänge zu ermöglichen oder zu blockieren; und
• ein Kraftstoffverwaltungssystem, das angeordnet ist, um den Kraftstoff der Brennkammer bereitzustellen, wobei das Kraftstoffverwaltungssystem umfasst:
  • zwei Kraftstoff-Öl-Wärmetauscher, durch die das Öl und der Kraftstoff 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 umfasst;
  • wobei das Verfahren umfasst:
    • Verwenden des Modulationsventils, um den Kühlluftstrom basierend auf der Turbineneintrittstemperatur einzustellen; und
    • Steuern des Kraftstoffverwaltungssystems, um zwischen 200 und 600 kJ/m³ von Wärme an den Kraftstoff aus dem Öl im primären Kraftstoff-Öl-Wärmetauscher unter Reiseflugbedingungen zu übertragen.

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

• an engine core comprising a combustion chamber arranged to combust a fuel; a turbine, the turbine comprising a plurality of turbine blades; a compressor arranged to be used as a source of cooling air for the turbine blades; and an inductor arranged to accelerate and direct the cooling air onto the turbine blades and comprising a plurality of airflow passages and a modulation valve arranged to allow or block cooling airflow into a subset of the passages; 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 the oil and the fuel flow, the fuel-oil 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 method comprising:
    • Using the modulating valve to adjust the cooling air flow based on the turbine inlet temperature; and
    • Controlling the fuel management system to transfer between 200 and 600 kJ/ ^m3 of heat to the fuel from the oil in the primary fuel-oil heat exchanger under cruise conditions.

Heat transfer is measured per cubic meter of fuel as it enters the combustion chamber. The amount of heat transferred to the fuel can therefore be calculated based on a temperature of fuel as it approaches or enters the combustion chamber compared to a temperature of fuel in an aircraft fuel tank. Any recirculation through or bypass from a heat exchanger can therefore be taken into account. Since heat transfer is measured per unit volume of fuel, this can be viewed as a heat transfer rate that is normalized for fuel flow variations in cruise flight

The inventors have recognized that using fuels other than traditional kerosene-based jet fuels, such as sustainable aviation fuels, may allow more heat to be transferred from the oil into the fuel per unit volume of fuel through the heat exchange system. Higher fuel temperatures upon entering the combustion chamber may allow for improved fuel combustion efficiency as well as improved oil cooling (since the fuel can absorb more heat). For example, a higher temperature capable fuel may be less viscous, which may result in a different nozzle spray pattern and droplet size within the combustion chamber, altering the fuel-air mixture and improving combustion efficiency. Heat transfer may be optimized to tailor fuel spray characteristics within the combustion chamber. In addition, the cooler oil may allow more heat to be dissipated from engine components, such as bearings, for the same oil flow rate, or a lower oil flow rate may be used for the same level of cooling.

Controlling the fuel management system may include controlling the flow of fuel through the heat exchangers - for example, to recirculate a variable portion of the fuel through the primary fuel-oil heat exchanger, and/or to allow a variable portion of the fuel to bypass the primary fuel-oil heat exchanger.

Controlling the fuel management system may include controlling the flow of oil through the heat exchangers - for example, by allowing a variable portion of the oil to bypass the primary fuel-oil heat exchanger or the secondary fuel-oil heat exchanger and/or by allowing a variable portion of the oil to recirculate through the primary and/or secondary fuel-oil heat exchangers.

One or more valves, pumps and/or recirculation or bypass lines may be provided accordingly to facilitate this control.

The method may include transferring 300 - 500 kJ/m ³ of heat to the fuel in the primary fuel-oil heat exchanger under cruise conditions. The method may include transferring 350 - 450 kJ/m ³ of heat to the fuel in the primary fuel-oil heat exchanger under cruise conditions. The method may include transferring 400 kJ/m ³ of heat to the fuel in the primary fuel-oil heat exchanger under cruise conditions.

• einen Triebwerkskern, umfassend eine Brennkammer, die angeordnet ist, um einen Kraftstoff zu verbrennen; eine Turbine, wobei die Turbine eine Vielzahl von Turbinenschaufeln umfasst; einen Verdichter, der angeordnet ist, um als eine Kühlluftquelle für die Turbinenschaufeln verwendet zu werden; und einen Induktor, der angeordnet ist, um die Kühlluft auf die Turbinenschaufeln zu beschleunigen und zu richten und eine Vielzahl von Luftstromdurchgängen und
• ein Modulationsventil umfasst, das angeordnet ist, um einen Kühlluftstrom in eine Teilmenge der Durchgänge zu ermöglichen oder zu blockieren, wobei das Modulationsventil konfiguriert ist, um den Kühlluftstrom basierend auf der Turbineneintrittstemperatur einzustellen; und
• ein Kraftstoffverwaltungssystem, das angeordnet ist, um den Kraftstoff der Brennkammer bereitzustellen, wobei das Kraftstoffverwaltungssystem umfasst:
  • zwei Kraftstoff-Öl-Wärmetauscher, durch die das Öl und der Kraftstoff 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 umfasst;
  • und wobei das Kraftstoffverwaltungssystem angeordnet ist, um 200 - 600 kJ/m³ von Wärme an den Kraftstoff im primären KraftstoffÖl-Wärmetauscher unter Reisebedingungen zu übertragen.

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

• an engine core comprising a combustion chamber arranged to combust a fuel; a turbine, the turbine comprising a plurality of turbine blades; a compressor arranged to be used as a source of cooling air for the turbine blades; and an inductor arranged to accelerate and direct the cooling air onto the turbine blades and to define a plurality of airflow passages and
• a modulation valve arranged to allow or block cooling air flow into a subset of the passages, the modulation valve configured to adjust the cooling air flow based on the turbine inlet temperature; 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 the 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;
  • and wherein the fuel management system is arranged to transfer 200 - 600 kJ/m ³ of heat to the fuel in the primary fuel-oil heat exchanger under cruising conditions.

The inventors have recognized that, particularly when using fuels other than conventional kerosene fuels, such as sustainable aviation fuels, fuel properties under cruise conditions can be controlled through careful engine management to improve performance. For example, the fuel management system may recirculate fuel through one or more fuel-oil heat exchangers (e.g., using a recirculation valve) and/or route fuel past one or more fuel-oil heat exchangers (e.g., using a bypass valve) or change oil or fuel flow rates to adjust the amount of heat transferred per unit volume of fuel. As discussed in relation to the third aspect, this can be used to optimize combustion conditions.

The fuel management system may further include one or more valves, pumps, and/or recirculation or bypass lines that may be used to control oil and/or fuel through the two fuel-oil heat exchangers to adjust the amount of heat transferred.

The engine core may further include a core shaft connecting the turbine to the compressor. The gas turbine engine may further include a fan disposed 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 gearbox and a pump shaft, the auxiliary gearbox receiving drive from the core shaft and outputting drive to the fuel pump via the pump shaft. The pump speed may therefore be linked to the core shaft speed and may have a fixed speed relationship between the core shaft (or in some engine architectures, another intermediate shaft) and the fuel pump via the auxiliary gearbox. In some implementations, the auxiliary gearbox may instead be arranged to provide a fixed number of different speeds for a given core shaft speed.

The gas turbine engine may further include engine bearing chambers. These engine bearing chambers may be cooled using the oil flow before the oil flows through one or both of 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 loop such that the oil is cooled by the fuel-oil heat exchanger before entering the engine bearing chambers, then the hotter oil exiting the engine bearing chambers is returned to the fuel-oil heat exchanger.

It is intended that any feature described for the third or fourth aspect may equally apply to the other of the third or fourth aspect and that the third or fourth aspect may be used in conjunction with the first or second aspect or with any of the other aspects discussed in relation to the first and second aspects.

In any or all of the above aspects, one or more of the features described below may apply.

The oil may enter the fuel-oil heat exchanger at a higher temperature than the fuel under cruising conditions, so the fuel-oil heat exchanger is arranged to transfer heat from the oil to the fuel.

The fuel may flow through the primary fuel-oil heat exchanger before passing through the secondary fuel-oil 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. A portion, but not all, of the fuel exiting the main fuel-oil heat exchanger may be directed to the servo fuel-oil heat exchanger. The servo fuel-oil heat exchanger may further increase the temperature of the fuel, thereby providing additional oil cooling before the fuel is provided for use in servomechanisms of the gas turbine engine (e.g., for fuel hydraulic actuation and/or heating). These servomechanisms may include a nacelle anti-icing system. The servomechanisms may include engine actuators. tuators. The servomechanisms may include a turbine housing and/or a turbine blade cooling servo valve. In these auxiliary systems, only fuel that has passed through the secondary fuel-oil heat exchanger may be used. The fuel used in these systems may be returned to a fuel tank for later recirculation or may be combined with other fuel leaving the main fuel-oil heat exchanger and entering the combustion chamber.

Fuel passing through the secondary fuel-oil heat exchanger cannot be supplied to the combustion chamber, so any heat transferred from the oil to fuel that reaches the combustion chamber is transferred to the primary fuel-oil heat exchanger. It will be understood that fuel returned to the fuel tank from the auxiliary systems after passing through the secondary heat exchanger can later be supplied to the combustion chamber, but any residual heat transferred from the secondary fuel-oil heat exchanger during the fuel's previous passage through the engine is likely to be negligible.

The fuel management system may include an engine fuel pump arranged to pump the fuel towards the combustion chamber and/or the secondary fuel-oil heat exchanger. The primary fuel-oil heat exchanger may be upstream of the pump so that the fuel passes through the pump after leaving the primary fuel-oil heat exchanger. The secondary fuel-oil heat exchanger may be downstream of the pump so that the fuel leaving the secondary fuel-oil heat exchanger does not then pass through the pump. The pump may therefore be shielded from the higher temperatures that may be reached for fuel in the secondary fuel-oil heat exchanger.

Therefore, in implementations where fuel passing through the secondary fuel-oil heat exchanger is not supplied to the combustion chamber, any heat transferred from the oil to fuel entering the combustion chamber may be transferred to the primary fuel-oil heat exchanger. The secondary fuel-oil heat exchanger therefore cannot affect the temperature of the fuel entering the combustion chamber.

At least some of the fuel cannot pass through the secondary fuel-oil heat exchanger. At least some 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 some of the fuel to bypass that heat exchanger.

The fuel management system may further include a recirculation valve disposed downstream of the primary heat exchanger and between the primary heat exchanger and the combustion chamber. The recirculation valve may be arranged to allow a controlled amount of fuel that has passed through the primary heat exchanger to be returned to the inlet of the primary heat exchanger.

The method may therefore include controlling the flow of fuel through the heat exchangers by controlling the amount of fuel returned to the inlet of the primary heat exchanger. The recirculated fuel may therefore pass through the primary heat exchanger at least twice, gaining more heat. A portion of the fuel may be passed through 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 may be delivered to the secondary fuel-oil heat exchanger via the primary fuel-oil heat exchanger (i.e., the fuel passed through or segmented through the primary 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, between 10% and 20%, or between 13% and 17%, of the fuel may be delivered to the secondary fuel-oil heat exchanger. Thus, more fuel generally flows from the primary heat exchanger into the combustion chamber without passing through the secondary heat exchanger than the fuel that 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 the path to flow into the secondary fuel-oil heat exchanger and optionally some or all of that portion being returned to the main fuel flow path prior to the combustion chamber.

Approximately 0-10%, 20-30%, 20-40%, or 30-40% of the fuel may 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 fuel can pass through the secondary fuel-oil heat exchanger. A variable proportion of fuel can pass through the secondary fuel-oil heat exchanger; the variation may be controlled by an automated feedback system (e.g. implemented by a fuel management system controller), optionally based on data from one or more temperature sensors. The one or more temperature sensors may be arranged to sense fuel or oil temperature.

A ratio of heat transfer from the oil to the fuel for the primary and secondary fuel-oil heat exchangers for fuel passing through both heat exchangers may be approximately between 70:30 and 90:10. For example, for fuel passing through both heat exchangers, 70% of the heat recovered by that fuel may be provided to the fuel in the primary heat exchanger and the remainder by the secondary heat exchanger, for a ratio of 70:30.

The ratio of heat transfer from oil to fuel for the primary and secondary fuel-oil heat exchangers can be approximately 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 may flow through the secondary fuel-oil heat exchanger before passing through the primary fuel-oil heat exchanger. The oil may not flow through any components that would increase its temperature between leaving the secondary fuel-oil heat exchanger and entering the primary fuel-oil heat exchanger. Thus, the temperature of the oil flowing through the secondary fuel-oil heat exchanger may be higher than the temperature of the oil flowing through the primary fuel-oil heat exchanger under cruise conditions. The fuel may flow through the primary fuel-oil heat exchanger before passing through the secondary fuel-oil heat exchanger. It will be understood that the fuel exiting the secondary fuel-oil heat exchanger will therefore generally be hotter than the fuel exiting the primary fuel-oil heat exchanger. The fuel may flow through the primary fuel-oil heat exchanger before passing through the secondary fuel-oil heat exchanger, while the oil may flow through the secondary fuel-oil heat exchanger before passing through the primary fuel-oil heat exchanger - the oil and fuel flow directions may therefore be opposite.

In some implementations, the fuel exiting the secondary fuel-oil heat exchanger is delivered to the combustion chamber rather than being routed to a pump and then recirculated, for example. This reduces the flow path length and time spent with the fuel at its highest temperature during the fuel delivery process, thereby reducing temperature losses to the environment. Furthermore, the shorter fuel path can reduce the number of components interacting with the fuel at its highest temperature, which can improve component life (e.g., protecting the pump).

In other implementations, the fuel exiting the secondary fuel-oil heat exchanger is instead delivered to a fuel tank, optionally through use in one or more auxiliary systems as mentioned above (e.g., fuel draulics). This flow path may in turn avoid the engine fuel pump, thus protecting that pump. Mixing with other fuel in the fuel tank may lower the average fuel temperature of the fuel just passed through the fuel-oil heat exchanger (it will generally raise the average fuel temperature of the entire body of fuel in the tank, which may provide a secondary benefit, particularly in cold conditions).

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 can be increased as it leaves the primary heat exchanger and increased even further as it leaves 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 temperatures of the fuel stream and therefore reduces the damage associated with these high temperatures.

The fuel flow may be controlled by using a recirculation valve arranged downstream of the primary heat exchanger, the recirculation valve arranged to allow a controlled amount of fuel not passed through the secondary heat exchanger to be returned to the inlet of the primary heat exchanger.

Alternatively or additionally, the fuel flow may be controlled by using a bypass pipe located across one or each fuel-oil heat exchanger. The bypass pipe may be arranged to allow a controlled amount of fuel completely bypasses the respective heat exchanger.

Alternatively or additionally, the oil flow may be controlled by using a bypass pipe located across one or each fuel-oil heat exchanger. The bypass pipe may be arranged to allow a controlled amount of oil to bypass the respective heat exchanger entirely.

The gas turbine engine may be part of an aircraft. The aircraft may include a fuel tank and a fuel supply pump configured to supply a flow of fluid 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 therefore upstream of the primary fuel-oil heat exchanger.

A fuel pump of the fuel management system may be described herein as a 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 may be located downstream of the primary fuel-oil heat exchanger. The main fuel pump may be located upstream of the secondary fuel-oil heat exchanger. One or more auxiliary fuel pumps may be located at any suitable position along the fuel flow path.

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

Alternatively or additionally, at least a portion of the fuel flowing through the secondary fuel-oil heat exchanger may be returned to any location along the fuel flow path; for example, back to the inlet of the primary or secondary fuel-oil heat exchanger, or to re-enter the flow from the primary fuel-oil heat exchanger on its way to 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. The closed loop oil system may be configured to supply a circulating 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 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 is heated in the process of lubricating and/or cooling other system components.

The closed loop system may include at least a first oil pump configured to pump an oil flow around at least a portion of the recirculating oil system. The first oil pump may be located at any suitable location around the recirculating oil system. The recirculating oil system may be configured so that the oil flow flows through engine components (e.g., the engine bearing chamber) and may then be collected in a pump sump. The first oil pump may be configured to pump oil from the sump to a first oil tank. Thus, the first oil pump may be described as an interceptor pump.

The oil tank may be adapted to hold a volume of oil. The oil tank may be configured to contain any suitable 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 disposed between the first oil tank and the primary or secondary fuel-oil heat exchanger. The second oil pump may be described as a feed pump. The second oil pump may be configured to pump oil from the first oil tank into 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 it follows the opposite flow path of the fuel.

The gas turbine engine may further include one or more air-to-oil heat exchangers. The one or more air-to-oil heat exchangers may be part of the recirculating oil system. The one or more air-to-oil heat exchangers may be arranged in a 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 after exiting the primary or secondary fuel-to-oil heat exchanger. 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 oil cooler) and then the primary fuel-to-oil heat exchanger.

In the examples described above, the engine architecture may be a direct drive engine, with no gearbox separating the core shaft from a fan shaft.

In other engine architectures, the engine may be designed differently, for example a geared motor instead of a direct drive motor.

• einen Triebwerkskern, der eine Turbine, einen Verdichter und eine Kernwelle, die die Turbine mit dem Verdichter verbindet, umfasst;
• einen Fan, der stromaufwärts des Triebwerkskerns angeordnet ist;
• eine Fanwelle;
• ein Getriebe (das als Hauptgetriebe bezeichnet werden kann), das einen Antrieb von der Kernwelle empfängt und über die Fanwelle einen Antrieb an den Fan ausgibt, wobei das Getriebe Zahnräder und Zapfenlager umfasst, die ihnen zugeordnet sind;
• ein Umwälzschmiersystem, das angeordnet ist, um Öl zu versorgen, um die Zahnräder und Zapfenlager zu schmieren, wobei das Schmiersystem einen ersten Öltank umfasst, der angeordnet ist, um Öl an die Zahnräder und Zapfenlager zuzuführen, und einen zweiten Öltank, der angeordnet ist, um Öl nur an die Zapfenlager zuzuführen; und
• ein Wärmeaustauschsystem, das angeordnet ist, um Wärme zwischen dem Öl und dem Kraftstoff zu übertragen.

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

• an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor;
• a fan located upstream of the engine core;
• a wave of fans;
• a gearbox (which may be referred to as a main gearbox) receiving drive from the core shaft and outputting drive to the fan via the fan shaft, the gearbox comprising gears and journal bearings associated therewith;
• a recirculating lubrication system arranged to supply oil to lubricate the gears and journal bearings, the lubrication system comprising a first oil tank arranged to supply oil to the gears and journal bearings and a second oil tank arranged to supply oil only to the journal bearings; and
• a heat exchange system arranged to transfer heat between the oil and the fuel.

• Steuern des Wärmetauschsystems, um 200 - 600 kJ/m³ von Wärme an den Kraftstoff aus dem Öl zu übertragen und
• Bereitstellen von kühlerem Öl an die Zapfenlager als an die Zahnräder.

The procedure includes under cruise flight conditions:

• Controlling the heat exchange system to transfer 200 - 600 kJ/m ³ of heat to the fuel from the oil and
• Providing cooler oil to the journal bearings than to the gears.

Heat transfer is measured per cubic meter of fuel as it enters the combustion chamber. Since heat transfer is measured per unit volume of fuel, this can be viewed as a heat transfer rate normalized for fuel flow variations in cruise flight. This transfer of heat from the oil to the fuel can be used to control the oil temperature as it enters the gearbox, and in particular the journal bearings and gears of the gearbox.

The inventors have recognized that the use of fuels other than traditional kerosene-based jet fuels, such as sustainable aviation fuels, may allow more heat to be transferred from the oil into the fuel per unit volume of fuel through the heat exchange system. Higher fuel temperatures upon entering the combustion chamber may enable improved fuel combustion efficiency as well as improved oil cooling.

Controlled transfer of heat from the oil to the fuel through the heat exchange system under cruising conditions plays a role in controlling oil temperature, and different fuels can provide improved oil cooling so that oil temperatures can be optimized.

The journal bearings may take on a greater load and/or heat up more than the gears, and it may therefore be more important to provide cool oil to the journal bearings than to the gears to improve lubrication and/or remove more heat from the oil. A lower oil temperature may be achieved if a smaller volume of oil is cooled, for the same temperature of the cooling medium (here probably the fuel). Deciding that a proportion of the oil leaving the oil tank is not cooled may therefore allow the cooled portion of the oil to reach a lower temperature. The inventors have realised that for improved cooling and lubrication efficiency, the recirculating lubrication system may therefore be arranged to provide oil at different temperatures to different components based on their needs, rather than the entire volume of oil. in the system in the same way. Having portions of the oil at different temperatures and controlled mixing of these oils at different temperatures can allow for improved oil temperature matching for a given use/component.

Furthermore, journal bearings are generally less tolerant of variations in lubricant film thickness than gears. "Touching down" with metal-to-metal contact in journal bearings can cause a much faster failure progression than with gear mesh. The journal bearings rely on a constant oil supply under all operating conditions, even in the event of a failure in the oil supply system (e.g. oil pump failure or pipe leakage). There is therefore also a positive impact of using a lower oil flow to the journal bearings. Since cooler oil allows a reduced oil flow to be used for the same cooling effect, the redundancy for failures required in the oil system can be more easily managed and the components in the oil system can be of a smaller size. For example, in current systems, oil can be supplied to the journal bearings from two different oil supplies and the supply is twice the required minimum oil flow, so that if one supply fails, the other supply can still provide the required flow to the bearings. At a reduced oil temperature for the journal bearings, twice the required flow may still be required, but the required flow is lower than with warmer oil. This has a positive impact in terms of component size (e.g. oil pump size) and improves power transmission efficiency and hence the specific fuel consumption (SFC) of the engine. Focusing cooling on the portion of the total oil flow destined for the journal bearings can therefore be used to increase the amount of heat that can be absorbed by the journal bearings (if the flow rate is kept constant, or reduced by an amount too small to fully offset the benefit provided by the cooler) and/or to allow a reduced flow rate with the same heat dissipation.

For example, in some implementations, some oil from a main oil tank is cooled (by passing through one or more heat exchangers), while other oil from the main oil tank is supplied to the transmission without such cooling. In various implementations, no uncooled oil from the main oil tank is supplied to the journal bearings; i.e., only cooled oil from the main oil tank can reach the journal bearings of the transmission. All uncooled oil, mixed with a percentage of the cooled oil, can reach the gears of the transmission.

The oil provided to the journal bearings may be at least 10e cooler than the oil supplied to the gears under cruising conditions. The oil provided to the journal bearings may be at least 15°C, 20°C, 25°C, 30°C, 35°C, 40°C or 45°C cooler than the oil supplied to the gearbox. The oil provided to the journal bearings may be 30°C cooler than the oil supplied to the gearbox. The oil provided to the journal bearings may be between 20°C and 40°C cooler than the oil supplied to the gearbox.

The oil supplied to the journal bearings may have a temperature between 80 °C and 120 °C under cruising conditions and optionally 100 °C. The oil supplied to the gears may have a temperature between 100 °C and 140 °C under cruising conditions and optionally 130 °C.

The method may include transferring 300 - 500 kJ/m ³ of heat to the fuel from the oil through the heat exchange system under cruising conditions. The method may include transferring 350 - 450 kJ/m ³ of heat to the fuel from the oil through the heat exchange system under cruising conditions. The method may include transferring 400 kJ/m ³ of heat to the fuel from the oil through the heat exchange system under cruising conditions. This heat transfer may be used to control the oil temperature entering the transmission.

The oil may have an average temperature of up to 160 °C, 180 °C, 200 °C or 220 °C when entering the heat exchange system under cruising conditions.

In addition to extracting heat from the main gearbox, the oil may pass through and cool one or more other engine components, optionally including an auxiliary gearbox (AGB) and/or one or more bearing chambers. These engine components may add more heat to the oil to raise its temperature to a temperature higher than if it were extracted from the main gearbox alone.

The oil in the recirculating lubrication system may therefore flow through an auxiliary gearbox and one or more of the engine journal bearings as well as the main gearbox, and the oil may have an average temperature of up to 220 °C when entering the heat exchange system under cruise conditions (even if oil leaving the main gearbox is considerably cooler).

The gas turbine engine may include an integrated propulsion generator.

The recirculating lubrication system may include first and second closed loop systems, the second closed loop system providing oil to the transmission and the first closed loop system providing oil to the integrated drive generator. The method may include transferring heat from oil in the second closed loop system to the fuel. The method may include transferring heat from oil in the first closed loop system to the fuel.

Heat can be transferred from the oil in the first closed loop to the fuel before heat is transferred from the oil to the fuel in the second closed loop system.

More heat can be dissipated into oil in the second closed control loop by the gearbox than is dissipated into oil in the first closed control loop by the integrated drive generator (IDG).

It will be appreciated that while the gearbox generally dissipates more heat in terms of absolute heat output, the heat output from the main gearbox, which may also be referred to as the power gearbox (PGB), will typically result in low quality with relatively high oil flow rates, i.e. the oil flow rate may be kept high so that the oil leaving the gearbox does not reach the temperature that would be achievable if the oil flow rate were equal to that through the IDG. The oil leaving the PGB will still generally be hotter than the oil leaving the IDG, although it should be noted that this may vary between implementations.

For example, a PGB may output about 75 kW of heat at cruise. The volumetric flow of oil to the PGB at the same conditions may be about 0.002 m3/s. In contrast, an IDG may only output about 18.4 kW of heat at cruise conditions, only about 25% of the heat output of the PGB - so the PGB can therefore output about four times as much heat as the IDG. However, the volumetric flow of oil to the IDG may only be about 0.00062 m3/s. The oil leaving the PGB may therefore only have about 1.2 times the heat transferred to it per unit volume compared to oil leaving the IDG, even though the heat output of the PGB is four times as high.

The PGB oil flow rate at cruise can be between 100 liters per minute and 150 liters per minute and can optionally be around or equal to 126 l/min. The IDG oil flow rate at cruise can be between 30 liters per minute and 45 liters per minute and can optionally be around or equal to 37 l/min.

PGB heat output in different implementations can range from 50 to 100 kW. IDG heat output in different implementations can range from 5 kW to 25 kW.

1. (i) einen Luft-Öl-Wärmetauscher, und
2. (ii) einen Öl-Öl-Wärmetauscher, der von einem Ölstrom von einer anderen Quelle durchströmt wird (z. B. zur Wärmeübertragung zwischen dem ersten und dem zweiten geschlossenen Regelkreis).

A plurality of heat exchangers may be provided to cool the oil before it re-enters the transmission (and/or other engine components to be lubricated and/or cooled), the plurality of heat exchangers including a fuel-oil heat exchanger and at least one of the following:

1. (i) an air-oil heat exchanger, and
2. (ii) an oil-to-oil heat exchanger through which an oil stream from another source flows (e.g. for heat transfer between the first and second closed loops).

In such implementations, the plurality of heat exchangers may be arranged in a parallel configuration and the method may include supplying a portion of the oil through each branch of the parallel configuration and adjusting that portion to vary the amount of oil flowing through the fuel-oil heat exchanger and flowing through a heat exchanger in the other branch. The parallel configuration may include a first branch with the fuel-oil heat exchanger and a second branch with an air-oil heat exchanger and/or an oil-oil heat exchanger.

The method may comprise transferring at least 40% of the heat lost from the oil to the fuel, with the remainder of the heat being transferred from the oil to air or to another oil. The method may comprise transferring at least 50%, 60%, or 70% of the heat lost from the oil to the fuel, with the remainder of the heat being transferred from the oil to air or to another oil.

A closed loop oil system may split or branch so that oil flows in two (or more) different branches which are then rejoined so that parallel oil streams flow around a portion of the loop. The plurality of heat exchangers may therefore be arranged in a parallel configuration and the method may include feeding a proportion of the oil through each branch of the parallel configuration and adjusting that proportion to achieve the To vary the amount of oil through the fuel-oil heat exchanger and the amount of oil through at least one heat exchanger on the other branch.

The parallel configuration may include a first branch with the fuel-oil heat exchanger and a second branch with one or more of an air-oil heat exchanger and an oil-oil heat exchanger.

The first closed loop system and the second closed loop system may be configured to interact via at least one oil-to-oil heat exchanger so that heat can be transferred from one oil stream to the other.

• einen Triebwerkskern, der eine Turbine, einen Verdichter und eine Kernwelle, die die Turbine mit dem Verdichter verbindet, umfasst;
• einen Fan, der stromaufwärts des Triebwerkskerns angeordnet ist;
• eine Fanwelle;
• ein Getriebe, das einen Antrieb von der Kernwelle aufnimmt und über die Fanwelle einen Antrieb an den Fan ausgibt, wobei das Getriebe Zahnräder und Zapfenlager umfasst;
• ein Umwälzschmiersystem, das angeordnet ist, um Öl zu versorgen, um die Zahnräder und Zapfenlager zu schmieren, wobei das Schmiersystem einen ersten Öltank umfasst, der angeordnet ist, um Öl an die Zahnräder und Zapfenlager zuzuführen, und einen zweiten Öltank, der angeordnet ist, um Öl nur an die Zapfenlager zuzuführen; und
• ein Wärmeaustauschsystem, das angeordnet ist, um Wärme zwischen dem Öl und dem Kraftstoff zu übertragen,
• wobei das Umwälzschmiersystem angeordnet ist, um dem Zapfenlager kühleres Öl bereitzustellen als an den Zahnrädern unter Reiseflugbedingungen; und
• wobei das Wärmetauschsystem angeordnet ist, um 200 - 600 kJ/m³ von Wärme an den Kraftstoff aus dem Öl unter Reisebedingungen zu übertragen.

According to a sixth aspect there is provided a gas turbine engine for an aircraft comprising:

• an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor;
• a fan located upstream of the engine core;
• a wave of fans;
• a gearbox receiving a drive from the core shaft and outputting a drive to the fan via the fan shaft, the gearbox comprising gears and journal bearings;
• a recirculating lubrication system arranged to supply oil to lubricate the gears and journal bearings, the lubrication system comprising a first oil tank arranged to supply oil to the gears and journal bearings and a second oil tank arranged to supply oil only to the journal bearings; and
• a heat exchange system arranged to transfer heat between the oil and the fuel,
• wherein the recirculating lubrication system is arranged to provide cooler oil to the journal bearing than to the gears under cruise conditions; and
• wherein the heat exchange system is arranged to transfer 200 - 600 kJ/m ³ of heat to the fuel from the oil under cruising conditions.

The transfer of heat to the fuel from the oil under cruising conditions can be used to control the oil temperature entering the transmission.

The gas turbine engine may further comprise an auxiliary gearbox. The oil in the recirculating lubrication system may be arranged to cool the auxiliary gearbox, thereby increasing the temperature.

The gas turbine engine may further include one or more bearing chambers. The oil in the recirculating lubrication system may be arranged to cool the one or more bearing chambers, thereby increasing the temperature.

Heat transfer is measured per cubic meter of fuel entering the combustion chamber. The amount of heat transferred to the fuel can therefore be calculated based on a temperature of fuel as it approaches or enters the combustion chamber compared to a temperature of fuel in an aircraft fuel tank. Any recirculation through or bypass from a heat exchanger can therefore be taken into account.

Since heat transfer is measured per unit volume of fuel, this can be viewed as a heat transfer rate normalized for fuel flow variations during cruise flight.

The gas turbine engine can therefore be used to implement the method of the fifth aspect and can offer the same technical effects and advantages.

It is intended that any of the features described for the fifth or sixth aspect may equally apply to the other of the fifth or sixth aspect.

Under cruising 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 stream of fuel entering the same fuel-oil heat exchanger. In this way, thermal energy may be transferred from the oil stream to the stream of fuel flowing through the one or more fuel-oil heat exchangers under cruising conditions. As such, oil exiting each heat exchanger may have a lower temperature than the oil entering that heat exchanger under cruising conditions.

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 enclosed in a nacelle), a turboprop engine or a turbojet engine. Each such engine may or may not be equipped with an afterburner. Such a gas turbine engine may, for example, be configured for power generation applications on land or at sea.

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).

Where the gas turbine engine is an open rotor or a turboprop engine, the gas turbine engine may comprise two counter-rotating propeller stages connected to and driven by a free power turbine by a shaft. 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 drive high and intermediate pressure compressors and propellers respectively 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 directly connected 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 configured to rotate (for example in use) at the lowest speed (for example from 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 (for example, in use) at the lowest speed (for example, only by the first core shaft and not by the second core shaft in the above example). Alternatively, the gearbox may be arranged to be driven by any one or more shafts, for example, by the first and/or second shaft in the above example.

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 equal to 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 having 5, 6 or 7 stages. Alternatively, the gas turbine engine may be a "geared" gas turbine engine having a first (or "low pressure") turbine having 3 or 4 stages.

Each fan blade may be defined to have 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 can 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, that is, 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 210 cm to 240 cm, or 250 cm to 280 cm, or 320 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 2600 rpm, or less than 2500 rpm, or less than 2300 rpm. As a purely 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, may be in the range of 2750 to 2900 rpm. As a purely 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, may be in the range of 2500 to 2800 rpm. 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), where the bypass ratio is defined as the ratio of the mass flow rate of flow through the bypass channel to the mass flow rate of 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, 9.5, 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 defined 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. As a non-limiting example only, 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 only, 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. By way of non-limiting example only, 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, 135 kN, 140 kN, 145 kN, 150 kN, 155 kN, 160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN, 450 kN, 500 kN, or 550 kN. The maximum thrust may be 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, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust in the range of 155 kN to 170 kN, 330 kN to 420 kN, or 350 kN to 400 kN. 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 with a fan diameter in the range 210 cm to 230 cm may be in the range 150 kN to 200 kN. Purely as a non-limiting example, the maximum thrust of a geared gas turbine engine with a fan diameter in the range 340 cm to 360 cm may be in the range 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 running 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.

Purely as a non-limiting example, the maximum TET of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 1935 K to 1995 K. Purely as a non-limiting example, the maximum TET of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 1890 K to 1950 K. The maximum TET may be in an inclusive range bounded by any two of the values in the preceding sentence (i.e., the values may form upper or lower limits), for example in the range of 1800 K to 1950 K or 1900 K to 2000 K. The maximum TET may, 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 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 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 that is better able to withstand impact (for example, from birds, ice, or other material) than the rest of the running edge. fel. Such a leading edge can be manufactured using titanium or a titanium-based alloy, for example. Thus, purely by way of example, the fan blade can have 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 during which the aircraft is propelled by the thrust produced by the engine. During climb, the engine may produce between 75% and 100% of the available thrust. In further non-limiting examples, the engine may produce between 80% and 95% of the available thrust. In further non-limiting examples, the engine may produce between 85% and 90% of the available thrust. In this regard, climb may refer to a phase of operation within an aircraft flight cycle between takeoff and arrival under 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.

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 a forward Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35,000 feet (10,668 m). Under such cruise conditions, the engine may provide a known required net thrust level. The known required net thrust level will of course depend on the engine and its intended application and may, for example, be a value in the range of 20 kN to 40 kN.

Further purely by way of example, the cruise conditions may correspond to a forward Mach number of 0.85 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38,000 feet (11,582 m). Under such cruise conditions, the engine may provide a known required net thrust level. The known required net thrust level will of course depend on the engine and its intended application and may, for example, be a value in the range of 35 kN to 65 kN.

In use, a gas turbine engine as described and/or claimed herein may operate under cruise conditions experienced at other Such cruise conditions may be determined by the cruise conditions (e.g., 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 occur 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 seitliche Schnittansicht eines Gasturbinentriebwerks ist, wobei die Direktantriebstriebwerk kein Getriebe aufweist, wie in 3A gezeigt ist;
• 4 ein beispielhaftes Flugzeug, das zwei Gasturbinentriebwerke umfasst, die vom Getriebe- oder Direktantrieb-Typ sein können, ist;
• 5 eine schematische Darstellung eines beispielhaften Kraftstoffsystems ist;
• 6a eine schematische Darstellung eines alternativen beispielhaften Kraftstoffsystems, einschließlich einer Kraftstoffumwälzschleife, ist;
• 6b eine schematische Darstellung eines weiteren alternativen beispielhaften Kraftstoffsystems ist, ähnlich wie in 6a gezeigt, wobei jedoch Kraftstoff, der den sekundären Wärmetauscher passiert, nicht an die Brennkammer zugeführt wird
• 6c eine schematische Darstellung eines weiteren alternativen beispielhaften Kraftstoffsystems ist, das dem in 6b gezeigten ähnlich ist, aber in dem Kraftstoff, der durch den sekundären Wärmetauscher strömt, optional nach dem Verwenden in einem oder mehreren Hilfssystemen in die Brennkammer zurückgeführt wird;
• 7 eine schematische Darstellung eines beispielhaften Umwälzölsystems ist;
• 8 eine schematische Darstellung einer Kombination des beispielhaften Kraftstoffsystems von 5 und des beispielhaften Umwälzölsystems aus 7 ist;
• 9 eine schematische Darstellung einer Kombination des beispielhaften Kraftstoffsystems von 6c und des beispielhaften Umwälzölsystems aus 7 mit kleineren Modifikationen ist;
• 10 ein beispielhaftes Verfahren zum Betreiben eines Gasturbinentriebwerks veranschaulicht, das ein Direktantriebstriebwerk sein kann; und
• 11 ein weiteres beispielhaftes Verfahren zum Betreiben eines Gasturbinentriebwerks ist, das ein Direktantriebstriebwerk sein kann.
• 12 eine schematische Darstellung eines beispielhaften Kraftstoffsystems ist, das für ein Getriebetriebwerk sein kann;
• 13 eine schematische Darstellung eines Teils einer ersten geschlossenen Schleife eines beispielhaften Umwälzschmiersystems ist;
• 14 eine schematische Darstellung eines Teils einer zweiten geschlossenen Schleife eines beispielhaften Umwälzschmiermittelsystems ist, das in Verbindung mit der in 13 gezeigten ersten geschlossenen Schleife verwendet werden kann;
• 15 eine schematische Darstellung eines Teils einer alternativen zweiten geschlossenen Schleife eines beispielhaften Umwälzschmiermittelsystems ist.
• 16 eine schematische Darstellung eines beispielhaften Kraftstoffsystems ist, das mit einem beispielhaften Umwälzschmiersystem interagiert
• 17 eine schematische Darstellung eines Teils eines beispielhaften Umwälzschmiermittelsystems ist;
• 18 ein beispielhaftes Verfahren zum Betreiben eines Gasturbinentriebwerks ist, das ein Getriebegasturbinentriebwerk sein kann;
• 19 eine schematische Darstellung einer Turbinenkühlvorrichtung ist, die Kraftstoff-hydraulisch betätigt werden kann; und
• 20 ist eine schematische Darstellung eines weiteren alternativen Teils eines beispielhaften Umwälzschmiermittelsystems.

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 side sectional view of a gas turbine engine, the direct drive engine having no gearbox, as in 3A is shown;
• 4 an exemplary aircraft comprising two gas turbine engines, which may be of the geared or direct drive type;
• 5 is a schematic representation of an example fuel system;
• 6a is a schematic representation of an alternative example fuel system including a fuel recirculation loop;
• 6b is a schematic representation of another alternative exemplary fuel system, similar to 6a shown, however, fuel passing through the secondary heat exchanger is not supplied to the combustion chamber
• 6c is a schematic representation of another alternative exemplary fuel system similar to that shown in 6b similar to that shown, but in which fuel flowing 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 an exemplary recirculating oil system;
• 8th a schematic representation of a combination of the exemplary fuel system of 5 and the exemplary circulating oil system from 7 is;
• 9 a schematic representation of a combination of the exemplary fuel system of 6c and the exemplary circulating oil system from 7 with minor modifications;
• 10 illustrates an exemplary method of operating a gas turbine engine, which may be a direct drive engine; and
• 11 is another exemplary method of operating a gas turbine engine, which may be a direct drive engine.
• 12 is a schematic representation of an exemplary fuel system that may be for a geared engine;
• 13 is a schematic representation of a portion of a first closed loop of an exemplary recirculating lubrication system;
• 14 is a schematic representation of a portion of a second closed loop of an exemplary circulating lubricant system used in conjunction with the 13 shown first closed loop;
• 15 is a schematic representation of a portion of an alternative second closed loop of an exemplary recirculating lubricant system.
• 16 is a schematic representation of an example fuel system interacting with an example recirculating lubrication system
• 17 is a schematic representation of a portion of an exemplary recirculating lubricant system;
• 18 is an exemplary method of operating a gas turbine engine, which may be a geared gas turbine engine;
• 19 is a schematic representation of a turbine cooling device that can be actuated by fuel hydraulics; and
• 20 is a schematic representation of another alternative portion of an exemplary recirculating lubricant system.

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 air streams: a core air stream A and a bypass air stream B. The gas turbine engine 10 includes a core 11 which receives the core air stream A. The engine core 11 comprises, in axial flow sequence, a low-pressure compressor 14, a high-pressure compressor 15, a combustor 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 air flow B flows through the bypass duct 22. The fan 23 is attached to the low-pressure turbine 19 via a shaft 26 and an epicyclic gear 30 and is driven thereby.

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 propulsion 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.

In implementations such as those described above with respect to the fifth and sixth aspects mentioned, one or more of the gears of the transmission may be mounted on journal bearings 44 (as in 20 shown). In particular, each planetary gear 32 may be rotatably mounted on a journal bearing 44 for rotation about an axis. These journal bearings may be described as journal bearings of the transmission 30 and as associated with the respective gears.

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 periphery for engaging the other gears. However, for clarity, only exemplary portions of the teeth are shown in 3A 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 the 2 and 3 is of the planetary type in that the planet carrier 34 is coupled to an output shaft via linkages 36 with the ring gear 38 being fixed. However, any other suitable type of epicyclic gear 30 may be used. As another example, the epicyclic gear 30 may be a star arrangement in which the planet carrier 34 is held fixed and the ring gear (or ring gear) 38 is allowed to rotate. In such an arrangement, the fan 23 is driven by the ring gear 38. As another alternative example, the gear 30 may be a differential gear in which both the ring gear 38 and the 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) can have a fixed or variable area.

As another example, other gas turbine engines with which the present disclosure may be applied may not include 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 1 a gas turbine engine is indicated generally at 10 with a main axis of rotation 9. The engine 10 includes, in axial flow order, an air inlet 12, a drive fan 23, an intermediate pressure compressor 14, a high pressure compressor 15, a combustor 16, a high pressure turbine 17, an intermediate pressure turbine 18, 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 combustor 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, 18, 19 before being discharged through the nozzle 20 to provide additional propulsive thrust. The high pressure turbine 17, the intermediate pressure turbine 18 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 combustor 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 fuels 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 missions or military missions with one refueling. A relatively higher thermal Stability (i.e., inhibition of fuel degradation or coking under thermal stress) can enable the fuel to maintain elevated temperatures in the engine and fuel injectors, potentially providing relative improvements in combustion efficiency. Reduced emissions, including particulate matter, can 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) can enable long-distance missions to optimize flight profiles; minimum aromatic concentrations (%) can ensure sufficient swelling of certain materials used in the design of O-rings and seals that have been previously exposed to fuels with high aromatic contents; and maximum surface tension (mN/m) can 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 can comprise molecules of about 7 to 18 carbons, with a significant portion of a given composition coming from molecules of 9 to 15 carbons, with an average of 12 carbons.

A number of sustainable aviation fuel blends have been approved for use. For example, some approved blends include blend ratios of up to 10% sustainable aviation fuel, while other approved blends include blend ratios of between 10% and 50% sustainable aviation fuel (with the remainder comprising one or more fossil-based hydrocarbon fuels, such as kerosene), with further 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 CO ₂ emissions relative to fossil-based hydrocarbon fuels (e.g., when burned in the combustor 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 of the aircraft 1. The fuel supply system of the examples shown 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) multiple 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 for balancing. All tanks 50, 53 are fluidly connected to each other in the example shown, thereby forming a single fuel source. Each of the central fuel tank 50 and the wing fuel tanks 53 may include a plurality of fluidly connected fuel tanks. It should be understood that this tank arrangement is provided by way of example only and is not limited to the scope of this disclosure. For example, the wing tanks 53a, 53b may be fluidly connected to each other but fluidly isolated from the central fuel tank 50, thereby providing two separate fuel sources that may contain chemically different fuels.

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 schematically shown in 5 shown.

The fuel system 1000 includes both the fuel supply system (which supplies the engine) and the fuel management system 1500 (which operates within the engine) of the aircraft.

Fuel is pumped from the fuel tank 50 to the gas turbine engine 10 by a low pressure fuel supply pump 1002. The fuel then flows through a primary fuel-oil heat exchanger 1004 before being pumped 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 passing through the secondary fuel-oil heat exchanger 1006. In the 5 In the example shown, the portion of the fuel flowing through the secondary fuel-oil heat exchanger 1006 then flows to the combustion chamber 16 of the gas turbine engine 10. In other examples, the portion of the fuel flowing through the secondary fuel-oil heat exchanger 1006 may not be supplied to the combustion chamber, but may instead be returned to a fuel tank 50, optionally after use in one or more auxiliary systems.

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.

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 the fuel flowing therethrough. Under cruising conditions, the average temperature of the oil flow entering the primary fuel-oil heat exchanger 1004 and the secondary fuel-oil heat exchanger 1006 is higher than the average temperature of the fuel entering the primary fuel-oil heat exchanger 1004 and the secondary fuel-oil heat exchanger 1006, respectively. In this manner, the primary fuel-oil heat exchanger 1004 and the secondary fuel-oil heat exchanger 1006 are configured to transfer thermal energy from an oil stream to an air stream flowing therethrough at flight conditions.

In the arrangement of the fuel management system 1500, the oil flows through the secondary fuel-oil heat exchanger 1006 before passing through the primary fuel-oil heat exchanger 1004 and does not leak through any components that would cause them to absorb heat between the secondary fuel-oil heat exchanger 1006 and the primary fuel-oil heat exchanger 1004. 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, however, flows through the primary fuel-oil heat exchanger 1004 before passing through the secondary fuel-oil heat exchanger 1006. In this arrangement, the fuel temperature exiting the secondary fuel-oil heat exchanger 1006 is higher than the fuel temperature exiting the primary fuel-oil heat exchanger 1004. The engine fuel pump 1003 is located upstream of the secondary fuel-oil heat exchanger 1006, meaning that it is not exposed to further elevated fuel temperatures.

The oil may flow through an air-oil heat exchanger (not shown) as well as through the primary and secondary fuel-oil heat exchangers 1004,1006 if further heat loss from the oil is desired to be removed from the fuel. For example, the oil may flow through the air-oil heat exchanger after passing through the secondary fuel-oil heat exchanger 1006 and before entering the primary fuel-oil heat exchanger 1004. The air-oil heat exchanger may be the only component (apart from tubes) through which the oil flows between the secondary fuel-oil heat exchanger 1006 and the primary fuel-oil heat exchanger 1004. In other examples, no air-oil heat exchanger may be provided. The oil may be fed to the bearings and/or other engine components used to lubricate and/or cool them after exiting the primary heat exchanger 1004. The oil may be returned to an oil tank 2002 as described below after exiting the primary heat exchanger 1004 or optionally after exiting the bearings and/or other engine components to lubricate and/or cool them before then being returned to the secondary heat exchanger 1006.

6a shows an alternative exemplary fuel system 6000 that includes a fuel delivery system and a fuel management system 6500 that includes a fuel flow path from the fuel tank 50 to the combustion chamber 16 of the gas turbine engine 10 of the aircraft 1. As with the fuel system 1000, fuel is pumped from the fuel tank 50 by the low pressure fuel supply pump 1002. The fuel then flows through a primary fuel-oil heat exchanger 1004 before being pumped by an engine fuel pump 1003. The engine fuel pump 1003 may be described as a main fuel pump.

The fuel system 6000 differs from the fuel system 1000 in that the fuel system 6000 includes a recirculation valve 6010 located downstream of the primary fuel-oil heat exchanger 1004 and upstream of an outlet 1006b of the secondary fuel-oil heat exchanger 1006. In the illustrated example, the recirculation valve 6010 is located downstream of the pump 1003.

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 the 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. The recirculation valve 6010 is arranged to return a controlled amount of fuel that has already passed through the primary heat exchanger 1004, but has not passed through the secondary heat exchanger 1006, to the inlet 1004a of the primary heat exchanger 1004. This provides a mechanism for controlling the fuel flow of the fuel system 6000. Controlling fuel flow in fuel system 6000 may operate to provide desired characteristics of the fuel (such as temperature or amount of heat dissipated by the oil thereon) upon entering combustion chamber 16, and also to provide desired oil temperatures.

Therefore, a tube 6011 may be provided that may be referred to as a recirculation tube because it transports fuel from a point along the main flow path through the engine to an earlier point along that flow path such that it must flow along a portion of the flow path multiple times to flow from the recirculation valve to a point on the flow path prior to the inlet to the primary heat exchanger 1004. In the illustrated example, the bypass tube 6011 delivers 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 flows through those components at least twice.

6b illustrates another alternative exemplary fuel system 7000. Fuel system 7000 is similar to fuel system 6000 and differs in that at least a portion of the fuel exiting secondary fuel-oil heat exchanger 1006 is supplied to additional aircraft mechanisms 7010 and then returned to a tank 50 rather than flowing to the combustor 16. Any fuel not required for these auxiliary systems 7010 may be returned directly to tank 50. In other implementations, the fuel may be returned to a point on the fuel flow path immediately upstream of the pump rather than to the tank itself. These mechanisms 7010 may include one or more of a nacelle anti-icing system, actuators and/or turbine shrouds and/or blade cooling servo valves (as described below). At least a portion of the fuel exiting the secondary fuel-oil heat exchanger 1006 is returned to the fuel supply system (e.g., to the fuel tank 50 or the fuel supply pump 1002) for redistribution. In other implementations, fuel used in one or more such additional aircraft mechanisms 7010 may then be routed to the combustion chamber 16 rather than being returned to a tank 50.

6c illustrates another alternative exemplary fuel system 7000a. Fuel system 7000a is similar to fuel system 7000 and differs in that the portion of fuel exiting secondary fuel-oil heat exchanger 1006 is directed to flow into the main fuel stream at its approach to combustion chamber 16 after optionally being directed to one or more auxiliary aircraft mechanisms 7010 (which may be as described above). Any fuel not required for these auxiliary systems 7010 may be directed directly into 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 if desired. In other implementations, fuel used in one or more such auxiliary aircraft mechanisms 7010 may then be returned to a tank 50 rather than flowing into the combustion chamber 16.

The aircraft 1 includes a recirculating oil system arranged to supply oil for lubrication and to remove heat from several components. An example of a closed oil-oil circulation system 2000 is shown schematically in 7 The recirculating oil system as set out in claim 7 comprises a closed loop oil system. In the example described, a single closed loop oil system provides the entirety of the recirculating oil system, in alternative implementations, multiple fluidly separated closed loops may be used, optionally with chemically different oils and/or different oil flow rates.

The closed oil-to-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 feed pump 2004 is configured to pump oil from the oil tank 2002 to 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 cruising conditions. In the secondary fuel-oil heat exchanger 1006, heat 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. Also in this way, 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 (optionally via an air-oil heat exchanger, in some implementations). 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. Also 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. The oil flow is then conveyed from the primary fuel-oil heat exchanger 1004 through the engine components 2006 for cooling and/or lubrication, such as shaft bearings. The oil acts as a lubricant and coolant within the engine components 2006 and receives heat from the engine components 2006 during the process.

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

8th shows schematically an exemplary heat exchange system that the closed oil-oil system 2000 of 7 and the fuel system 1000 from 5 The two systems 1000,2000 are shown together to illustrate the interactions between the fuel and oil flow within the gas turbine engine 10. Thick black lines indicate fuel flow and thin black lines indicate oil flow.

9 shows schematically an alternative exemplary heat exchange system 9000, which the closed oil system 2000 of 7 and the 7000a fuel system from 6c including the recirculation loop 6010, 6011, but the position of the recirculation valve 6010 differs slightly from that shown in 6c In particular, the recirculation valve 6010 is located upstream of the branch in the fuel flow path to the secondary heat exchanger 1006, rather than between the inlet to 1006a and the outlet of 1006b of the heat exchanger 1006. Thus, in the 9 shown implementation, all the fuel exiting the pump 1003 in contrast to the 6c shown implementation by the recirculation valve 6010. The control methodology for the valve 6010 can be adapted accordingly. Again, thick black lines indicate fuel flow and thin black lines indicate oil flow. 9 additionally includes the auxiliary systems 7010 through which the fuel can pass after leaving the secondary fuel-oil heat exchanger 1006, indicating that they can use the fuel before it is directed into the combustion chamber 16 in some implementations.

The auxiliary systems 7010 include one or more turbine blade cooling servo valves 1902 in some implementations - these valves 1902 may be fueldraulic; using fuel to actuate them, as described in more detail below with respect to 19 discussed. In other implementations, these valves may be actuated differently (e.g., electronically or with a different working fluid) rather than being fueldraulic. These valves 1902 may be used to modulate airflow to cool the turbine casing and/or blades, and thus may be described as modulation valves 1902.

The modulation valves 1902 may be used to control the flow of cooling air to a turbine 17,19 of the gas turbine engine 10 using air from a compressor 14,15 of the gas turbine engine.

19 illustrates a turbine cooling device 1900 comprising at least one modulation valve 1902. The turbine cooling device 1900 comprises an inductor 1901 arranged to accelerate and direct the cooling air from the compressor 14,15 onto the turbine blades 17', 19' of the turbine 17,19.

The inductor 1901 may be an annular static inductor and may accelerate the air flow in a substantially circumferential direction.

The inductor 1901 includes a plurality of airflow passages 1903 arranged to direct airflow to specific locations within the turbine 17,19 to be cooled. The one or more modulating valves 1902 may be used to control the flow of cooling air into one or more of the passages 1903. For example, multiple modulating valves 1902 may be provided and each modulating valve 1902 may control air access to and flow through a subset of the passages. Some passages 1903' are not valved so they cannot be closed and air always flows during operation.

Each flow passage 1903 includes an inlet arranged to receive air from a compressor 14,15 and an outlet arranged to direct the air onto the turbine 17,19, and in particular to a desired location on the turbine 17,19 (e.g., a particular portion of the blades 17', 19' or across each entire blade). A single, larger input flow from the compressor 14,15 can therefore be converted into multiple smaller flows onto the turbine 17,19. Narrowing the flow area - from a less confined space into the passages 1903 - can accelerate the flow.

The inductor 1901 of the described implementation has an annular series of identical first and second flow passages, optionally alternating around a circumference of the inductor, thereby forming two subsets of passages. The cooling device 1900 includes a flow guide for directing a first portion of the air flow from the compressor 14,15 to the inlets of the first subset of passages 1903 and a second portion of the air flow from the compressor 14,15 to the inlets of the second subset of passages 1903'.

The modulation valves 1902 selectively modulate the first part of the air flow.

A controller 1904 may be used to control the or each valve 1902 such that flow through the first subset of passages 1903 may be controlled - for example, allowed or blocked (binary control: on or off) or controlled over a range of different flow rate options. The control may be automatic, for example based on feedback from one or more temperature sensors to the controller 1904. In particular, the air temperature at the inlet of the turbine 17,19 to be cooled may be sensed (directly or indirectly by calculation from other sensor data), and control of the modulation valves 1902 may be performed based on this temperature - for example, by increasing the air flow rate when the turbine inlet temperature rises above a setpoint.

It will be understood that the core air stream A enters the turbine 19 after leaving the combustion chamber 16 (and also after passing through another turbine 17 in some implementations). Upon leaving the combustion chamber 16, this core air stream A includes combustion products and potentially traces of unburned fuel as well as air, but is also referred to as an air stream to avoid possible confusion with fuel and oil streams. An average temperature of this air stream upon entering the turbine 19 - i.e., the turbine inlet temperature for that turbine - therefore provides an indication of how hot the turbine blades 19' will be without cooling. and so can be used to guide the amount of cooling provided to the blades 19' of that turbine 19. Likewise, an inlet temperature for the high pressure turbine 17 can be used to guide an amount of cooling provided to the blades 17' of that turbine 17. A turbine inlet temperature can be used to guide the amount of cooling provided to the blades 19', 17' of multiple turbines in some implementations. The turbine inlet temperature of the highest pressure turbine can be used in such implementations.

The same cooling device 1900 may be used to cool multiple turbines 17,19 in some implementations. In other implementations, each turbine 17,19 may be provided with its own cooling device 1900, or only some turbines 17,19 may be cooled. Similarly, if there are multiple compressors 14,15, air from one or both may be used for turbine cooling.

In the example described above, a grouped subset of passes may be controlled, but the passes 1903 may not be individually controlled. In other implementations, individual pass control may be provided and/or more or fewer subsets of passes may be defined.

In engines with multiple compressors and turbines, multiple such cooling devices 1900 may be provided, optionally one for each turbine.

The inventors recognized that the use of fuels different from the traditional kerosene-based jet fuels, such as sustainable aviation fuels, can result in different fuel properties, and that the operation of the engine can be optimized for these different fuel properties. The sulfur content of the fuel should therefore be taken into account when considering heat transfer from the oil to the fuel, as the sulfur content is linked to the thermal stability of the fuel, and a less thermally stable fuel can absorb less heat before performance problems occur, while a more thermally stable fuel can absorb more heat and so can be used for more efficient engine cooling.

• Schritt 10100: Erhalten eines Kraftstoffs,wobei der Kraftstoff eineSchwefelgehalt von weniger als 30 ppm aufweist, zur Bereitstellung an die Brennkammer 16 des Gasturbinentriebwerks 10, und optional auch an einem oder mehreren Hilfstriebwerkssystemen, im Reiseflug.

10 illustrates an exemplary method 10000 of operating a gas turbine engine 10, which may be a direct drive gas turbine engine 10. The method 10000 includes the steps of:

• Step 10100: Obtaining a fuel, the fuel having a sulfur content of less than 30 ppm, for supply to the combustion chamber 16 of the gas turbine engine 10, and optionally also to one or more auxiliary power systems, during cruise flight.

The fuel may have a density in the range of 760-840 kg/m ³ at 15 °C.

The fuel provided to the gas turbine engine 10 may therefore be provided to the combustor 16, or to one or more auxiliary power systems 7010, or in some cases, to the one or more auxiliary power systems 7010 and then subsequently passed to the combustor 16.

The selected fuel may have an aromatic content of less than 10% by volume.

Obtaining the fuel may include selecting a single fuel. Obtaining the fuel may alternatively include selecting a fuel blend. 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 1. The selection 10100 may therefore be performed prior to departure. Alternatively, the aircraft 1 may have multiple fuel sources and the selected fuel may be one of multiple fuels on board the aircraft 1 or a blend of multiple fuels on board the aircraft. Thus, the selection 10100 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 blend fuel(s) stored in one or more fuel tanks 50,53 to provide a fuel with the desired properties. The controller may store the fuel properties and refueling stations for the fuels on board and calculate the ratio of fuel blending (which may not include mixing of fuels) to provide a fuel with the desired properties.

• Schritt 10200: Verwenden des Modulationsventils, um den Kühlluftstrom basierend auf einer Turbineneintrittstemperatur einzustellen.

The gas turbine engine 10 includes a turbine cooling device 1900 which includes one or more modulation valves 1902 arranged to control the cooling air flow of the turbine 17,19.

• Step 10200: Use the modulating valve to adjust the cooling air flow based on a turbine inlet temperature.

The one or more modulation valves 1902 may be fueldraulic valves - that is, the fuel provided to the gas turbine engine 10 may be used to actuate the one or more modulation valves 1902, as well as to direct a portion (or all) thereof to the combustion chamber 16.

The fuel density may be considered and selected as suitable for use in fueldraulic systems and/or the fueldraulic systems may be adapted to operate with a fuel having the density selected based on combustion efficiency.

The inventors also recognize that the use of fuels other than traditional kerosene-based jet fuels, such as sustainable aviation fuels, may allow more heat to be transferred from the oil to the fuel per unit volume of fuel through the heat exchange system. Higher fuel temperatures upon entering the combustor may enable improved fuel combustion efficiency and/or improved oil cooling, as described above.

11 illustrates an example method 11000 for operating a gas turbine engine 10 including primary and secondary fuel-oil heat exchangers 1004,1006. The gas turbine engine may be a direct drive gas turbine engine 10.

• Schritt 11100: Steuern des Kraftstoffverwaltungssystems, um zwischen 200 und 600 kJ/m³ von Wärme an den Kraftstoff aus dem Öl im primären Kraftstoff-Öl-Wärmetauscher unter Reisebedingungen zu übertragen.

The method 11000 comprises the steps:

• Step 11100: Control the fuel management system to transfer between 200 and 600 kJ/ ^m3 of heat to the fuel from the oil in the primary fuel-oil heat exchanger under cruise conditions.

The fuel management system 1500 then provides some or all of the heated fuel to the combustion chamber 16.

A total of 200 - 600 kJ/m ³ of heat can be transferred from the oil to the fuel as the fuel passes from a fuel tank to the combustion chamber under cruising conditions.

As with regard to 5 discussed, the fuel management system 1500 includes the primary fuel-oil heat exchanger 1004 and the secondary fuel-oil heat exchanger 1006 arranged to transfer heat to or from the fuel. An engine fuel pump 1003 is positioned downstream of the primary fuel-oil heat exchanger 1004 and upstream of the secondary fuel-oil heat exchanger 1006. The fuel management system 1000 is arranged so that the fuel reaches the primary fuel-oil heat exchanger 1004 before the secondary fuel-oil heat exchanger 1006.

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 downstream of the engine fuel pump may facilitate obtaining higher fuel temperatures without compromising fuel pump life since the high temperature fuel does not have to pass through the pump 1003. The reduced maximum temperature to which the pump 1003 is exposed can reduce creep, reduce the deposition of thermal decomposition products of the fuel within the pump (which can cause clogging, resulting in degradation of flow over the life of the pump), reduce thermal damage to components such as seals, and/or reduce cycle damage resulting from differential thermal expansion of different pump components, thus providing an extension of pump life/improvement of pump performance for a given pump age. In addition, the bearing film thickness in the pump 1003 can be improved if temperatures of fuel in the pump are kept lower by placing the pump in front of another heat exchanger. Fuel is the lubricant for the pump bearings and fuel viscosity generally decreases with increasing temperature, thereby damagingly reducing film thickness. Reducing the fuel temperature can therefore result in an increase in the viscosity of the fuel, which generally increases the performance of the bearing surfaces within the pump 1003, thereby reducing wear and thus reducing the deterioration of the flow output over time. In addition, lower fuel temperatures generally reduce the formation of fuel degradation products, which also have an adverse effect on the life and reliability of the pump 1003. The lower fuel temperature can therefore also increase reliability. The lower temperature can also reduce damage to journal bearings and thrust bearings. Other relevant factors include variation in cavitation (cooler fuel is denser and has a lower vapor pressure) and lubrication; A cooler fuel is advantageous from a volumetric pump output perspective, allowing a pump 1003 to last longer/ and when operated with cool Using a different fuel will also result in a longer operating time.

Controlling the fuel management system may include controlling the oil flow and/or the fuel flow through the primary and secondary fuel-oil heat exchangers 1004,1006, for example using one or more bypass lines or recirculation loops 6010,6011 as described above.

In various implementations, as discussed above, controlling 11100 the fuel management system may include controlling the flow of fuel through the primary and secondary fuel-oil heat exchangers 1004, 1006.

Fuel flow may be controlled using a recirculation valve 6010, where provided. An actively controlled amount of fuel exiting the primary heat exchanger 1004 may be recirculated into the primary heat exchanger 1004 rather than flowing directly into the combustion chamber 16. This recirculation may also transfer fuel that has already passed through the engine fuel pump 1003 to a position upstream of the engine fuel pump 1003.

Alternatively or additionally, the fuel flow may be controlled by using one or more bypass pipes arranged to allow a portion of the fuel to flow through either one or both fuel-oil heat exchangers.

The speed of pump 1003 can also be adjusted to either increase the fuel flow rate (so that the heat transfer per unit volume through the heat exchangers is reduced) or decrease the fuel flow rate (so that the heat transfer per unit volume through the heat exchangers is increased).

Controlling 11100 the fuel management system may include controlling the flow of oil through the primary and secondary fuel-oil heat exchangers 1004,1006 and/or through one or more other heat exchangers - e.g., an oil-oil heat exchanger 2030 between separate closed loop oil systems, in engine architectures with multiple closed oil-oil systems, or an air-oil heat exchanger 2020.

Additionally or alternatively, as for fuel flow, oil flow may be controlled using one or more bypass tubes where present to allow oil to bypass one or more heat exchangers 1004,1006 rather than flowing therethrough. The oil may also be recirculated in some implementations and/or the oil flow rate may be adjusted by controlling one or more oil pumps.

The heat exchange system may include a controller arranged to implement this control. The controller may receive inputs from one or more temperature sensors and may control one or more valves and/or the pump 1003 based on the received data.

• • Kernwellendrehzahl und Triebwerksschubanforderung;
• • Aktuelle Kraftstofftemperatur und/oder Öltemperatur an einem oder mehreren Positionen;
• • Kraftstoffkheizwert;
• • Kraftstoffviskosität;
• • Fan-Drehzahl;
• • Kraftstoffdurchflussrate an die Brennkammer (allgemein als WFE-Gewicht des Hauptmotors-Kraftstoffstroms bezeichnet); und
• • Haupt-/Triebwerk-Kraftstoffpumpendrehzahl oder Drehzahloptionen.

Active control of fuel and/or oil flow rates can be performed based on one or more parameters, such as:

• • Core shaft speed and engine thrust requirement;
• • Current fuel temperature and/or oil temperature at one or more locations;
• • Fuel calorific value;
• • Fuel viscosity;
• • 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.

In alternative examples, controlling the amount of fuel exiting the primary heat exchanger 1004 to be recirculated may not be an active process step - instead, a set, fixed, portion of the fuel may be recirculated. Alternatively, no fuel may be recirculated, and in some implementations, no recirculation path may be available.

Step 11200: Use the modulating valve to adjust the cooling air flow based on turbine inlet temperature.

As for the method 10000 described above, the one or more modulation valves 1902 may be fueldraulic valves - i.e., the fuel provided to the gas turbine engine 10 may be used to actuate the one or more modulation valves 1902, as well as some (or all) of them being passed to the combustor 16.

The fuel density may be considered and selected as suitable for use in fueldraulic systems and/or the fueldraulic systems may be adapted to operate with a fuel having the density selected based on combustion efficiency.

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

The methods 10000,11000 described above have been described primarily with respect to a direct current gas turbine engine 10, but it is understood that they could be implemented for a geared gas turbine engine 10 with appropriate adjustments to the oiling system (e.g., adding another closed loop to lubricate and cool a gearbox 30).

In other engine architectures, the engine can be designed differently, for example a geared motor instead of a direct drive motor. 12 to 17 and 20 and the description thereof illustrate exemplary arrangements of a geared motor, although it will be appreciated that many of the features and options could equally be applied to a direct drive motor, with the exception of features relating to the main gearbox 30. The use of the same reference numerals indicates equivalent features.

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 schematically shown in 12 . The fuel system 1000 includes both the fuel supply system arranged to supply fuel from the fuel source 50,53 (as described above) to each engine 10 and a fuel management system within the engine arranged to supply the supplied fuel to the combustion chamber 16. The fuel management system manages the fuel temperature as well as the fuel flow, directing the fuel across one or more heat exchangers of the engine's heat exchange system.

A low pressure fuel pump 1002 pumps fuel from the fuel tank 50 toward the gas turbine engine 10. The fuel flows from the fuel tank 50 through an integrated drive generator fuel-oil heat exchanger (IDG fuel-oil heat exchanger) 1005 before passing through a primary fuel-oil heat exchanger 1007. The two heat exchangers 1005, 1007 form part of the engine's heat exchange system.

The first heat exchanger reached by the fuel may have oil used to lubricate and/or cool one or more components of an integrated drive generator 2006 passing therethrough and thus may be described as an integrated drive generator oil heat exchanger 1005.

The second heat exchanger reached by the fuel may include oil used to lubricate and/or cool a main gearbox 30 of the gas turbine engine 10, described as a main fuel-oil heat exchanger.

From the main fuel-oil heat exchanger 1007, the fuel is pumped by the engine fuel pump 1003 into the combustion chamber 16 of the gas turbine engine 10, where it is burned to drive the gas turbine engine 10.

In the described implementation, the primary fuel-oil heat exchanger 1007 and the IDG fuel-oil heat exchanger 1005 are configured to promote a flow of fuel therethrough. Generally, at least the majority of the fuel supplied to the engine 10 passes through both heat exchangers 1005,1007, although each heat exchanger 1005,1007 may be bypassed to allow a portion of the fuel to flow through the respective heat exchanger.

Generally, at least the majority of the fuel passing through the IDG fuel-oil heat exchanger 1005 also flows along a single flow path through the primary fuel-oil heat exchanger 1007. The two heat exchangers 1005,1007 can therefore be described as being in series with each other in terms of fuel flow. The IDG fuel-oil heat exchanger 1005 and the primary fuel-oil heat exchanger 1007 are configured to also pass an oil stream through each - the oil flowing through one is different from the oil flowing through the other in the implementation described, although it will be appreciated that the same oil may flow through one fuel-oil heat exchanger and then through another fuel-oil heat exchanger in other implementations. The two heat exchangers 1005,1007 are therefore in the described implementation in separate closed circuit systems 2000,2000' with respect to oil flow.

The IDG fuel-oil heat exchanger 1005 and the primary fuel-oil heat exchanger 1007 are each configured to allow heat to be transferred between the oil flowing therethrough and the fuel flowing therethrough. Under cruising conditions, the average temperature of the oil flow entering the primary fuel-oil heat exchanger 1007 and the oil entering the IDG fuel-oil heat exchanger 1005 is higher than the average Temperature of the fuel entering the primary fuel-oil heat exchanger 1007 and the IDG fuel-oil heat exchanger 1005, respectively. In this manner, the primary fuel-oil heat exchanger 1007 and the IDG fuel-oil heat exchanger 1005 are configured to transfer thermal energy from an oil stream to an air stream flowing through the flight conditions.

The aircraft 1 includes a circulating lubrication and/or cooling system arranged to lubricate oil from and/or remove heat from a plurality of components. It will be appreciated that the relative importance of lubrication and cooling may vary for different aircraft components and for different phases of flight and that the circulating lubrication and/or cooling system may simply be referred to as a circulating lubrication system for brevity. The circulating lubrication system in some examples includes a closed loop oil system or two separate closed oil-oil systems as mentioned above.

An example of a first closed oil-oil system 2000 is shown schematically in 13 It is understood that such a first closed oil-oil system 2000 may additionally comprise other components which are 13 are not shown or described herein, and that only the key components are described and shown with respect to the features described. Those skilled in the art will therefore recognize that the various diagrams of the closed systems shown and described herein are not intended to be limited to the presence of other components in that circuit.

The closed oil-to-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 feed pump 2004 is configured to pump oil from the oil tank 2002 into the IDG fuel-oil heat exchanger 1005. The average temperature of the oil entering the IDG fuel-oil heat exchanger 1005 is higher than the average temperature of the fuel entering the IDG fuel-oil heat exchanger 1005 under cruising conditions. In the IDG fuel-oil heat exchanger 1005, heat energy is transferred from the oil stream to the fuel stream. In this way, the average temperature of the oil stream leaving the IDG fuel-oil heat exchanger 1005 is lower than the average temperature of the oil stream entering the IDG fuel-oil heat exchanger 1005. Also in this way, the average temperature of the fuel leaving the IDG fuel-oil heat exchanger 1005 is higher than the average temperature of the fuel entering the IDG fuel-oil heat exchanger 1005.

The oil stream is then directed to an integrated drive generator 2006 where it lubricates moving components and is heated in the process. In some implementations, the oil may be used primarily as a coolant for the IDG 2006 and may perform minimal or no lubrication.

From the integrated drive generator 2006, oil collects in a sump 2008. A scavenging pump 2010 is configured to pump oil from the sump 2008 back to the oil tank 2002, which is ready for reuse.

A second closed oil-oil system 2000 as in 14 shown includes a further oil tank 2002' adapted to hold a volume of oil. In some implementations, gases are removed from oil within the oil tank 2002' of the second closed loop system by a vent.

A second feed pump 2004' is configured to pump oil from the oil tank 2002' of the second closed oil-to-oil system to the primary fuel-to-oil heat exchanger 1007. The average temperature of the oil entering the primary fuel-to-oil heat exchanger 1007 under cruise conditions is higher than the average temperature of the fuel entering the primary fuel-to-oil heat exchanger 1007. In the primary fuel-to-oil heat exchanger 1007, thermal energy is transferred from the oil stream to the fuel stream. In this way, the average temperature of the oil leaving the primary fuel-oil heat exchanger 1007 is lower than the average temperature of the oil entering the primary fuel-oil heat exchanger 1007, so it is cooled before being reused as a lubricant, whereby the cooled oil removes more heat from the system being lubricated. Also in this way, the average temperature of the fuel leaving the primary fuel-oil heat exchanger 1007 is higher than the average temperature of the fuel entering the primary fuel-oil heat exchanger 1007.

The oil flow is then directed to a power transmission 30, which may also be described as the main transmission 30 of the gas turbine engine 10, and generally to other engine components 33, including an auxiliary transmission (AGB) and one or more bearing chambers.

The oil flow may be divided into two or more parallel flows, for example one flow through the main gearbox 30 and one flow through the other engine components 33, or multiple parallel flows through the main gearbox 30 (e.g., via different components of the gearbox) and separate flows through the AGB 33 and the or each bearing chamber 33.

The power gearbox 30 is arranged to receive drive from the core shaft and output the drive to the fan via the fan shaft 42, and thus may include or be associated with one or more bearings to support the shafts, which may be journal bearings. In addition, the gearbox 30 has gears therein and may have one or more journal bearings associated with the gears therein, which may also require lubrication and/or cooling. The oil may be used to lubricate the journal bearings and/or gears, and its temperature generally increases significantly when used under cruising conditions, thereby assisting in cooling the bearings and/or gears as the oil flow carries heat away from them.

It will be appreciated that the main transmission 30 may include several different components to be lubricated - e.g. gears (and e.g. in particular gear teeth that mesh with each other) and also one or more journal bearings within the transmission 30. A more complex implementation of this second closed oil-oil system 2000' arranged to allow different oil flows to different transmission components is described below with reference to 20 described.

The engine components 33 that are cooled and optionally also lubricated generally include an AGB 33. The AGB, also known as an auxiliary drive, is a gearbox that forms part of the gas turbine engine 10, although it is not part of the engine core 11 and does not drive the fan 23. The AGB instead drives the engine accessories, e.g., fuel pumps, and generally handles large loads. A relatively large amount of heat can therefore be discharged from the AGB into the oil. One or more bearing chambers may be lubricated by the same oil and may similarly dissipate heat into the oil. Per unit of oil passing through, the AGB and bearing chambers 33 can transfer more heat to the oil than the main gearbox 30 in most implementations.

For example, in various implementations under cruise conditions, the exit temperature of oil from the power gearbox 30 may have a maximum of 160°C and optionally range from 100°C to 160°C. In contrast, oil exiting the AGB and/or the various bearing chambers 33 may have a temperature in the range from 160°C to 220°C. In implementations where the flow is not split, the oil may flow through the main gearbox 30 before entering the AGB 33. One or more valves may be provided to control the oil flow distribution.

Back to 14 , the oil from the power transmission 30 and other engine components 33 collects in a second sump 2008'. The pump sump 2008' may be a single pump sump or may consist of a plurality of separate pumps, for example as described below with respect to 20 described in more detail. A second interceptor pump 2010' is configured to pump oil from the second sump 2008' back to the oil tank 2002' of the second closed oil-to-oil system ready for reuse. Similar to the pump sump(s), multiple interceptor pumps may be used in some implementations.

15 shows schematically an alternative exemplary section of the 14 shown second closed oil system 2000'. In this portion, an oil flow is pumped by the feed pump 2004' through a valve 2016. The valve 2016 is operable to split the oil flow between both the primary fuel-oil heat exchanger 1007 and a first air-oil heat exchanger 2020, where the first air-oil heat exchanger 2020 is disposed in parallel with the primary fuel-oil heat exchanger 1007. The branched oil flow may then be recombined before the oil is supplied to the transmission 30 and other engine components 33 to be cooled and/or lubricated. The recombined oil flow is then delivered to the power transmission 30 and other engine components 33 to be cooled and/or lubricated. In various examples, an oil-to-oil heat exchanger (not shown) may be arranged in series with/on the same branch of the parallel arrangement as the first air-to-oil heat exchanger 2020. The oil-to-oil heat exchanger may enable heat exchange between the first and second closed oil-to-oil systems 2000,2000'.

Any suitable proportion of oil can flow through each of the first air-to-oil heat exchanger 2020 and the primary fuel-to-oil heat exchanger 1007. In some examples, the valve 2016 is operable to vary the flow of oil into the primary fuel-to-oil heat exchanger 1007 and the first air-to-oil heat exchanger 2020.

It is also understood that in various implementations, each of the fuel-oil heat exchangers 1005,1007 may be arranged in series or in parallel with one or more air-oil heat exchangers.

16 schematically shows an exemplary arrangement and interaction of the first closed oil system 2000, the second closed oil system 2000' and the fuel system 1000, wherein the fuel flow is shown in a thick black line and the oil flow in a thinner black line.

The first closed oil system 2000 of this exemplary arrangement is as in 13 The second closed oil system 2000' of this exemplary arrangement is as shown in 14 The fuel system 1000 of this exemplary arrangement is as shown in 12 arranged as shown.

In use, fuel is pumped from the fuel tank 50 through the low pressure fuel pump 1002. The fuel then flows through the IDG fuel-oil heat exchanger 1005. The first closed oil system 2000 is configured so that the recirculating flow of oil also flows through the IDG fuel-oil heat exchanger 1005.

Under cruise conditions, the average temperature of the oil stream entering the IDG fuel-oil heat exchanger 1005 is higher than the average temperature of the fuel stream entering the IDG fuel-oil heat exchanger 1005. The IDG fuel-oil heat exchanger 1005 is configured to transfer heat from the oil stream to the fuel stream. In this way, the average temperature of the oil stream exiting the IDG fuel-oil heat exchanger 1005 is lower than the average temperature of the oil stream entering the IDG fuel-oil heat exchanger 1005. Likewise, the average temperature of the fuel stream exiting the IDG fuel-oil heat exchanger 1005 is higher than the average temperature of the fuel stream entering the IDG fuel-oil heat exchanger 1005.

The fuel then flows through the primary fuel-oil heat exchanger 1007. The second closed oil system 2000' is configured so that the recirculating flow of oil also flows through the primary fuel-oil heat exchanger 1007.

Under cruising conditions, the average temperature of the oil stream entering the primary fuel-oil heat exchanger 1007 is higher than the average temperature of the fuel entering the primary fuel-oil heat exchanger 1007. The primary fuel-oil heat exchanger 1007 is configured to transfer heat from the oil stream to the fuel stream. In this way, the average temperature of the oil stream exiting the primary fuel-oil heat exchanger 1007 is lower than the average temperature of the oil stream entering the primary fuel-oil heat exchanger 1007. Likewise, the average temperature of the fuel exiting the primary fuel-oil heat exchanger 1007 is higher than the average temperature of the fuel stream entering the primary fuel-oil heat exchanger 1007.

During cruising conditions, the average temperature of the oil stream flowing through the IDG fuel-oil heat exchanger 1005 may be lower than the average temperature of the oil stream flowing through the primary fuel-oil heat exchanger 1007. In this way, the fuel first flows through the heat exchanger with a lower average oil stream temperature before passing through the heat exchanger with a higher average oil stream temperature.

After flowing through the primary fuel-oil heat exchanger 1007, the fuel flows to the combustion chamber 16 of the gas turbine engine 10, which is pumped by the engine fuel pump 1003.

In some examples, heat transferred from the oil to the fuel before it reaches the combustion chamber 16 may increase the fuel temperature to an average of at least 135°C, 140°C, 150°C, 160°C, 170°C, 180°C, 190°C, or 200°C at the entrance to the combustion chamber 16 under cruise conditions.

In some examples, heat transferred from the oil to the fuel may increase the fuel temperature to an average between 135-150°C, 135-160°C, 135-170°C, 135-180°C, 135-190°C, or 135-200°C at the entrance to the combustion chamber 16 under cruise conditions.

The fuel temperatures entering the combustion chamber under cruise conditions may be defined as an average over at least 5 minutes and optionally over ten minutes under steady cruise conditions. The use of average temperatures avoids the temperature measurement being unduly distorted by any transient spike in temperature, which may be defined as temperature fluctuations of the fuel during operation, often an increase in temperature. Each fluctuation shall not last more than 5 minutes. A fuel A minimum combustion chamber entry temperature of 135 °C under cruising conditions therefore requires that the fuel temperature remain at or above 135 °C for a period of time and a short transient spike to a temperature above 135 °C is not sufficient.

Optionally, one or more additional heat sources can be used to heat the fuel in addition to the oil-fuel heat transfer to achieve the desired temperature.

In some examples, the heat transferred to the fuel from the oil before the fuel reaches the combustion chamber 16 may be 200 - 600 kJ/m ³ under cruise conditions.

17 schematically shows an exemplary configuration of the first closed oil system 2000 and the second oil system 2000', wherein the two independent circulating streams of oil are brought into heat exchange relationship by an oil-oil heat exchanger 2030.

In the 17 In the example shown, the first closed oil system 2000 is configured such that the recirculating oil flow is pumped by the feed pump 2004 through a valve 2016. The valve 2016 is operable to divide the oil flow between the IDG fuel-oil heat exchanger 1005 and the oil-oil heat exchanger 2030, wherein the oil-oil heat exchanger is arranged in parallel with the IDG fuel-oil heat exchanger 1005.

In various examples, any suitable portion of the oil flow may be diverted between the IDG fuel-oil heat exchanger 1005 and the oil-oil heat exchanger 2030. In such examples, the valve 2016 may be operable to divert a fixed portion of the oil flow to each of the IDG fuel-oil heat exchanger 1005 and the oil-oil heat exchanger 2030. Alternatively, the valve 2016 may be operable to divert a variable portion of the oil flow to each of the IDG fuel-oil heat exchanger 1005 and the oil-oil heat exchanger 2030.

In alternative implementations, the IDG fuel-to-oil heat exchanger 1005 and the oil-to-oil heat exchanger 2030 may be in series, and the first closed oil-to-oil system 2000 may not branch.

After passing through the heat exchangers, the oil flow is then fed to the integrated drive generator 2006 and then to the sump 2008. The interceptor pump 2010 then pumps the oil from the sump 2008 into the oil tank 2002 for reuse.

The oil stream within the first closed oil-oil system 2000 is arranged to be brought into heat exchange relationship with the separate oil stream within the second closed oil-oil system 2000' by the oil-oil heat exchanger 2030. In the oil-oil heat exchanger 2030, the oil stream within the first closed oil-oil system 2000 does not mix with the oil stream within the second closed oil-oil system 2000'. The oil-oil heat exchanger 2030 is configured so that heat transfer can occur between the two separate oil streams. In this way, heat can be transferred from a hotter oil stream to the cooler oil stream within the oil-oil heat exchanger 2030.

The second closed oil-to-oil system 2000' is configured such that the circulating oil flow is pumped by the second feed pump 2004' through a second valve 2016' where the oil flow branches off. The second valve 2016' is operable to split the oil flow between the primary fuel-to-oil heat exchanger 1007 and a first air-to-oil heat exchanger 2020, where the first air-to-oil heat exchanger 2020 is in series on the same branch with the oil-to-oil heat exchanger 2030 such that the air-to-oil heat exchanger 2020 and the oil-to-oil heat exchanger 2030 are arranged in parallel with the primary fuel-to-oil heat exchanger 1007. Additional air-to-oil heat exchangers may be present in some implementations.

In various examples, any suitable portion of the oil flow may be diverted between the primary fuel-oil heat exchanger 1007 and the first air-oil heat exchanger 2020. In such examples, the second valve 2016' may be operable to divert a fixed portion of the oil flow to each of the primary fuel-oil heat exchanger 1007 and the first air-oil heat exchanger 2020. Alternatively, the second valve 2016' may be operable to divert a variable portion of the oil flow to each of the primary fuel-oil heat exchanger 1007 and the first air-oil heat exchanger 2020.

After passing through the heat exchangers, the oil stream is then delivered to the power transmission 30 and other engine components 33 (generally including an AGB) and then to the second sump 2008'. The second scavenge pump 2010' then pumps the oil from the second pump sump 2008' to the oil tank 2002' of the second closed oil-to-oil system 2000' for reuse.

20 illustrates a more complex implementation of the second closed oil-to-oil system 2000', wherein the oil used to cool the main transmission 30 is controlled to specifically target different transmission components.

In particular, the transmission comprises gears 28, 32,38 and also journal bearings 44 which are associated with one or more of the gears.

The circulating lubrication (and/or cooling) system 2000a` of the examples described herein is arranged to provide cooler oil to the journal bearing 44 than to the gears 32.

In the implementation shown, the feed pump 2004' pumps oil from the main oil tank 2002' in two different directions, with some of the oil passing through the parallel arrangement of heat exchangers 2020,1007 described above and the remainder instead passing through a bypass pipe 2005, thus eliminating the heat exchangers 2020,1007 (it will be understood that one or more additional or alternative heat exchangers not shown may be present in this heat exchange arrangement, for example an IDG fuel-oil heat exchanger 1005 as described above). The oil passing through the bypass pipe 2005 is therefore hotter than the oil flowing through the parallel heat exchange arrangement 2020,1007, as it has not had the opportunity to lose heat to the fuel and/or air. A valve 2005a may be provided to directly control the bypass flow in some implementations. The valve 2005a may be actively controlled in conjunction with the valve 2016', which is operable to divide the oil flow between the primary fuel-oil heat exchanger 1007 and the first air-oil heat exchanger 2020 based on control rules and the output from one or more temperature sensors. Alternatively, the valve 2005a may be a passive valve (e.g., a thermostatic valve). In other implementations, no separate valve 2005a may be provided, and instead the valve 2016' may be a more complex valve that divides the oil flow between the heat exchangers 1007,2020 and the bypass 2005. For example, a single valve controlling the split between the three routes (two parallel heat exchangers and a bypass route) may have a different fixed percentage of opening of a port that allows the oil flow to flow through the bypass tube 2005 when split between the first air-oil heat exchanger 2020 and the primary fuel-oil heat exchanger 1007, respectively. However, the two separately controllable valves 2005a, 2016' may provide more flexibility.

A valve 2005a operable to vary the proportion of fuel flow directed along the bypass tube 2005 versus the proportion of fuel passing into the one or more heat exchangers 1007,2020 may therefore be provided in various implementations to control cooling. In some implementations, a one-way valve may additionally be provided where the bypass tube reconnects the oil flow to the gears 28,32,38 to prevent backflow. Since the bypass flow may generally be at a higher pressure than the flow that passed through the heat exchangers 2020,1007, the one-way valve may serve to prevent the bypassing flow from traveling in the wrong direction along the tubes and inadvertently reaching the journal bearings 44.

The 2000a'' recirculating lubrication system is generally similar to that in 15 However, the second closed circuit system 2000' shown additionally comprises a ring gear catcher pump 38a arranged to pump the used oil from the ring gear 38 of the transmission 30, and a second (auxiliary) oil tank 2008b' arranged to receive the oil received from the ring gear 38 through the ring gear catcher pump 38a. The oil from the second (auxiliary) oil tank 2008b' is supplied to enter the journal bearings 44 directly, bypassing the heat exchangers 1005,1007 and the main tank 2002'. The second oil tank 2008b' is therefore arranged only to supply oil to the journal bearings.

The sump 2008a'' collects oil from the other components of the transmission 30 and acts as the sump 2008' of the 15 example shown. From this pump sump 2008a', a portion of the oil is returned to the main oil tank 2002', but the remainder of the oil is directed directly to the journal bearings 44 in the implementation shown, bypassing the heat exchangers 1005,1007 and the main tank 2002'. A dedicated pump 2011' may be provided to pump oil from the pump sump 2008a'' to the journal bearings 44. In alternative implementations, the pump 38a may be arranged differently and may be fed from the ring gear 38 and/or from the pump sump 2008a, such that oil returns to the journal bearings 44 from both the ring gear 38 and the pump sump 2008a.

In still other implementations, all of the oil in the pump sump 2008' may be returned to the main tank 2002'.

The main tank 2002' and also the pump sump 2008a (indirectly via its connection to the main tank 2002) therefore supplies oil to both the gearbox gear wheels 28,32, 38 as well as the gear journal bearings 44, while the auxiliary tank 2008b only supplies oil to the gear journal bearing 44.

All of the uncooled oil from the main tank 2002' - i.e., the oil that passes through the bypass tube 2005 rather than passing through the heat exchange assembly and is therefore warmer - is supplied to the gears 28,32,38 mixed with a percentage of the cooled oil. The percentage of cooled oil supplied to the gears 28,32,38 may be controllable, optionally automatically based on a temperature feedback loop using sensor data from one or more temperature sensors. One or more additional valves and/or pumps may be provided to facilitate this control. Generally, none of the uncooled oil from the main tank 2002' goes to the power transmission journal bearings 44 in the described implementations, at least not under cruising conditions. Instead, a percentage of the cooled oil is supplied to the journal bearings 44 along with some oil from the pump sump 2008a'' and some from the auxiliary oil tank 2008b' in the implementation shown. In alternative implementations, only cooled oil from the parallel heat exchange assemblies may be supplied to the journal bearings 44 and all of the oil from the pump sump 2008a'' and from the auxiliary oil tank 2008b' may be returned to the main tank rather than recirculating around the transmission 30.

The second closed loop 2000a` is therefore arranged to provide cooler oil to the journal bearings 44 than to the gears 28,32,38. This may allow for more efficient cooling overall by cooling a smaller amount of oil to the lowest desired temperature and using warmer oil sufficient for the intended purpose/component(s), as discussed above.

In various implementations, additional features and branches may be present in the oil system. For example, the oil supply pipe 2007 to the gears 28,32,38 of the main power transmission 30 may branch and supply oil to an auxiliary gearbox (AGB) and/or to one or more shaft bearings 33. An additional auxiliary oil tank or sump 2008c may collect oil leaving these engine components, and an additional pump 2012, which may be an interceptor pump, may return the used oil to the main oil tank 2002'. Additionally or alternatively, a portion of the oil from the main tank 2002' - optionally the uncooled oil before it reaches the heat exchange assembly or bypass pipe 2005 - may be taken to cool one or more other engine components - e.g., a vibration dampening device for a shaft. In such implementations, the auxiliary transmission cooling, shaft bearing cooling, and/or vibration dampening device cooling may reach the hottest oil temperatures - for example, the oil temperature exiting the auxiliary transmission cooling and/or vibration dampening device may be about 180°C or more, while oil exiting the power transmission 30 may be about 130°C.

After mixing in the main tank 2002', the uncooled oil from the main tank 2002' may have a temperature of about 140°C. As such, oil exiting the main transmission 30 or being recirculated around the main transmission 30 without returning to the main tank 2002' may be cooler (e.g., by about 10°C) than oil exiting the main tank 2002'. The oil flowing through the heat exchangers may have a temperature drop to about 80°C.

To supply oil to the transmission journal bearings 44, a portion of this cooled oil may be mixed with the uncooled oil from the sump 2008a' and/or the auxiliary tank 2008b' to provide oil at a temperature of approximately 100°C to the journal bearings 44.

To supply oil to the gears 28,32,38, the remainder of this cooled oil may be mixed with the uncooled oil from the main tank 2002' to provide oil at a temperature of about 120°C to the gears.

It is understood that these temperatures are provided only as example values in travel operation and may vary between implementations.

In some implementations, the hottest oil (e.g., from the engine components 33 including the AGB) may be fed directly to a fuel-oil heat exchanger 1007 rather than being mixed with cooler oil in the main tank 2002' to increase a temperature rise of the fuel.

The oil system may further include one or more valves arranged to control the mixing of oil flows - for example, a Journal Shut-Off Valve (JSOV) and a Variable Oil Recirculation Valve (VORV) may be provided to control the oil flow to the journal bearings 44. The JSOV may be a passive valve based on oil pressure - when the oil pressure is low, the valve may automatically create a closed loop circuit that feeds oil from the power transmission (PGB) sump 2008a` back to the PGB journal bearings 44. The VORV may control the proportions of oil from the pump sump 2008a', the auxiliary tank 2008b' and the heat exchange system 2020,1007 in the oil supplied to the PGB journal bearings 44, for example based on feedback from one or more temperature sensors.

The inventors have recognized that the use of fuels other than conventional kerosene-based jet fuels, such as sustainable aviation fuels, may allow more heat to be transferred from the oil into the fuel per unit volume of fuel through the heat exchange system. Higher fuel temperatures upon entering the combustion chamber may enable improved fuel combustion efficiency and/or improved oil cooling, as described above.

• Schritt 18100: Steuern des Wärmetauschsystems, um 200 - 600 kJ/m³ von Wärme an den Kraftstoff aus dem Öl unter Reisebedingungen zu übertragen.

18 illustrates an exemplary method 18000 of operating a geared gas turbine engine 10 having a transmission 30 including gears 28, 32, 38 and journal bearings 44. The method 18000 includes the steps of:

• Step 18100: Control the heat exchange system to transfer 200 - 600 kJ/m ³ of heat to the fuel from the oil under cruise conditions.

This heat transfer from the oil to the fuel can be used to control the oil temperature entering the transmission.

Controlling the heat exchange system may include controlling the oil flow and/or fuel flow through the main and IDG fuel-oil heat exchangers 1005,1007, for example using one or more bypass lines or recirculation loops 6010,6011 as described above.

In various implementations, as discussed above, controlling 18100 the heat exchange system may include controlling the flow of fuel through the main and IDG fuel-oil heat exchangers 1005,1007.

The fuel flow may be controlled using a recirculation valve 6010 where provided. An actively controlled amount of fuel exiting the primary fuel-oil heat exchanger 1007 may be recirculated into the inlet of the primary fuel-oil heat exchanger 1007 rather than flowing directly into the combustion chamber 16. This recirculation may also pick up fuel that has already passed through the engine fuel pump 1003 to a position upstream of the engine fuel pump 1003, depending on the location of the recirculation loop.

Alternatively or additionally, the fuel flow may be controlled by using one or more bypass pipes arranged to allow a portion of the fuel to flow through either one or both fuel-oil heat exchangers.

The speed of pump 1003 can also be adjusted to either increase the fuel flow rate (so that the heat transfer per unit volume through the heat exchangers is reduced) or decrease the fuel flow rate (so that the heat transfer per unit volume through the heat exchangers is increased).

Controlling 18100 the heat exchange system may include controlling the flow of oil through the main and/or IDG fuel-oil heat exchangers 1005,1007 and/or through one or more other heat exchangers - e.g., an oil-oil heat exchanger 2030 between separate closed oil-oil systems, in engine architectures with multiple closed oil-oil systems, or an air-oil heat exchanger 2020.

Additionally or alternatively, as for fuel flow, oil flow may be controlled using one or more bypass pipes, where provided, to allow oil to bypass one or more heat exchangers 1005,1007 rather than flowing through them. The oil may also be recirculated in some implementations and/or the oil flow rate may be adjusted by controlling one or more oil pumps.

The heat exchange system may include a controller arranged to implement this control. The controller may receive inputs from one or more temperature sensors and may control one or more valves and/or the pump 1003 based on the data received. The heat exchange system may include one or more valves, pumps, and/or recirculation or bypass lines, in addition to those explicitly shown and described, to facilitate this control.

• • Kernwellendrehzahl und Triebwerksschubanforderung;
• • Aktuelle Kraftstofftemperatur und/oder Öltemperatur an einem oder mehreren Positionen;
• • Kraftstoffkheizwert;
• • Kraftstoffviskosität;
• • Kraftstoffdurchflussrate an die Brennkammer (allgemein als WFE-Gewicht des Hauptmotors-Kraftstoffstroms bezeichnet);
• • Fan-Drehzahl; und
• • Haupt-/Triebwerk-Kraftstoffpumpendrehzahl oder Drehzahloptionen.

Active control of fuel and/or oil flow rates can be performed based on one or more parameters, such as:

• • Core shaft speed and engine thrust requirement;
• • Current fuel temperature and/or oil temperature at one or more locations;
• • Fuel calorific value;
• • Fuel viscosity;
• • Fuel flow rate to the combustion chamber (commonly referred to as WFE – weight of main engine fuel flow);
• • Fan speed; and
• • Main/engine fuel pump speed or speed options.

In alternative examples, controlling the amount of fuel leaving the main fuel-oil heat exchanger 1007 to be recirculated may not be an active process step - instead, a specific, fixed proportion of the fuel may be recirculated. Alternatively, no fuel may be recirculated, and in some implementations, no recirculation path may be available.

Step 18200: Supply cooler oil to the journal bearings 44 of the gearbox 30 than to the gears 28,32, 38.

The lubrication (and/or cooling) system 2000a'' includes a first oil tank 2002 arranged to supply oil to the gears and journal bearings and a second oil tank 2008b' arranged to supply oil only to the journal bearings, thus allowing for different oil flows and individual oil temperature control. The lubrication system 2000a'' may also include a bypass pipe 2005 that allows a portion of the oil from the main tank 2002' to bypass one or more heat exchangers - this uncooled oil may be supplied to the gears 28,32,38 of the power transmission 30, but not to the journal bearings 44 of the power transmission 30, so that the oil temperature of oil provided to the gears is again increased compared to the oil provided to the journal bearings.

It is to be understood that the invention is not limited to the implementations 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 (15)

A method (11000) of operating a gas turbine engine (10), the gas turbine engine (10) comprising: an engine core (11) comprising a combustion chamber (16) arranged to combust a fuel; a turbine (19), the turbine comprising a plurality of turbine blades (19'); a compressor (14) arranged to be used as a cooling air source for the turbine blades (19'); and an inductor (1901) arranged to accelerate and direct the cooling air onto the turbine blades (19') and comprising a plurality of airflow passages (1903) and a modulation valve (1902) arranged to allow or block cooling air flow into a subset of the passages; and a fuel management system (1500) arranged to provide the fuel to the combustion chamber (16), the fuel management system comprising: two fuel-oil heat exchangers (1004,1006) through which the oil and the fuel flow, the fuel-oil 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); the method (11000) comprising: using (11100) the modulation valve (1902) to adjust the cooling air flow based on the turbine inlet temperature; and controlling (11200) the fuel management system (1500) to transfer between 200 and 600 kJ/m ³ of heat to the fuel from the oil in the primary fuel-oil heat exchanger (1004) under cruise conditions. Procedure (11000) according to Claim 1 , the method comprising transferring 300 - 500 kJ/m ³ of heat to the fuel in the primary fuel-oil heat exchanger (1004) under cruising conditions. Procedure (11000) according to Claim 1 wherein the secondary fuel-oil heat exchanger (1006) is a servo fuel-oil heat exchanger. Procedure (11000) according to Claim 1 , wherein between 10% and 30% of the fuel is fed to the secondary fuel-oil heat exchanger (1006). Procedure (11000) according to Claim 1 , wherein for fuel passing through both fuel-oil heat exchangers (1004,1006), a ratio of heat transfer from the oil to the fuel for the primary (1004) and secondary (1006) fuel-oil heat exchangers is between 70:30 and 90:10. Procedure (11000) according to Claim 1 , wherein for fuel passing through both fuel-oil heat exchangers (1004,1006), a ratio of heat transfer from the oil to the fuel for the primary (1004) and secondary (1006) fuel-oil heat exchangers is approximately 80:20. Procedure (11000) 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). Procedure (11000) according to Claim 1 wherein: the fuel management system (1500) further comprises: a recirculation valve (6010) disposed downstream of the primary fuel-oil heat exchanger (1004) and between the primary fuel-oil heat exchanger and the combustion chamber (16), the recirculation valve (6010) disposed to allow a controlled amount of fuel that has passed through the primary fuel-oil heat exchanger (1004) to be returned to the inlet (1004a) of the primary fuel-oil heat exchanger; and wherein the flow of fuel through the fuel-oil heat exchanger (1004,1006) is controlled at least in part by controlling the amount of fuel to be returned to the inlet of the primary fuel-oil heat exchanger (1004). Procedure (11000) according to Claim 1 wherein fuel passing through the secondary fuel-oil heat exchanger (1006) is not provided to the combustion chamber (16) so that all heat transferred from the oil to the fuel reaching the combustion chamber (16) is transferred in the primary fuel-oil heat exchanger (1004). A gas turbine engine (10) for an aircraft (1), the gas turbine engine (10) comprising: an engine core (11) comprising a combustion chamber (16) arranged to combust a fuel; a turbine (19), the turbine comprising a plurality of turbine blades (19'); a compressor (14) arranged to be used as a cooling air source for the turbine blades (19'); and an inductor (1901) arranged to accelerate and direct the cooling air onto the turbine blades (19') and comprising a plurality of airflow passages (1903) and a modulation valve (1902) arranged to allow or block cooling air flow into a subset of the passages, the modulation valve (1902) configured to adjust the cooling air flow based on the turbine inlet temperature; and a fuel management system (1500) arranged to provide the fuel to the combustion chamber (16), the fuel management system comprising: two fuel-oil heat exchangers (1004,1006) through which the oil and the fuel flow, the fuel-oil 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 the fuel management system (1500) arranged to transfer 200 - 600 kJ/m ³ of heat to the fuel in the primary fuel-oil heat exchanger (1004) under cruising conditions. Gas turbine engine (10) to Claim 10 wherein the fuel management system (1500) is arranged to transfer 300 - 500 kJ/m ³ of heat to the fuel in the primary fuel-oil heat exchanger (1004) under cruising conditions. Gas turbine engine (10) to Claim 10 wherein the secondary fuel-oil heat exchanger (1006) is a servo fuel-oil heat exchanger. Gas turbine engine (10) to Claim 10 , wherein between 10% and 30% of the fuel is fed to the secondary fuel-oil heat exchanger (1006). Gas turbine engine (10) to Claim 10 , wherein a ratio of heat transfer from the oil to the fuel for the primary (1004) and secondary (1006) fuel-oil heat exchangers is between 70:30 and 90:10. Gas turbine engine (10) to Claim 10 , 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).

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