FR3144223A1 - Method of operating a gas turbine engine - Google Patents

Abstract

A method (1100) of operating a gas turbine engine (10) is disclosed. The gas turbine engine (10) includes a staged combustion chamber (16) having an arrangement of fuel spray nozzles (403, 404) in which a flow of fuel is directed to a subset of the nozzles (403, 404). 404) adjacent to one or more igniters (405) during a relighting procedure. The method (1100) includes providing fuel (1101) to the combustion chamber (16) having an aromatic content of 10% or less by volume. The invention also relates to a gas turbine engine (10). Figure for abstract: Figure 9

Description

Method of operating a gas turbine engine

The present description relates to a method of operating a gas turbine engine using fuels other than conventional kerosene-based jet fuels.

There is an expectation in the aviation industry for a trend toward the use of fuels different from the traditional kerosene-based jet fuels typically used today.

According to a first aspect, a method of operating a gas turbine engine is provided. The gas turbine engine includes a staged combustion chamber in which fuel is burned. The combustion chamber includes a fuel spray nozzle arrangement in which fuel flow is directed to a subset of the fuel spray nozzles adjacent one or more igniters during a relighting procedure. The method includes supplying fuel to the combustion chamber having a calorific value of at least 43.5 MJ/kg.

The inventors were aware that the calorific value of the fuel has an effect on the ability of the gas turbine engine to re-ignite (e.g., allowing easier re-ignition). Fuels with a higher calorific value may also have greater thermal stability, allowing the fuel to absorb more heat and therefore improve combustion of the fuel in the combustion chamber and/or improve cooling of the combustion chamber. oil (for example, if heat is transferred to the fuel from the oil through a fuel-oil heat exchanger). The calorific value of the fuel must therefore be taken into account when distributing fuel to the fuel spray nozzles of the combustion chamber during a relighting procedure.

The method may include providing fuel to the combustion chamber having a calorific value of between 43.5 MJ/kg and 44 MJ/kg.

The method may include providing fuel to the combustion chamber having a calorific value of at least 43.8 MJ/kg.

The method may include providing fuel to the combustion chamber having a calorific value of between 43.8 MJ/kg and 44 MJ/kg.

The fuel supplied to the combustion chamber may have an aromatic content of 10% or less by volume, preferably 5% or less, and more preferably 1% or less.

The gas turbine engine may include a fuel-oil heat exchanger. The method may include transferring heat from the oil to the fuel before the fuel enters the combustion chamber so as to lower the viscosity of the fuel to 0.58 mm ² /s or less upon entry. in the combustion chamber at cruising conditions, preferably 0.48 mm ² /s or less, preferably between 0.40 mm ² /s and 0.48 mm ² /s and preferably between 0.42 mm ² / s and 0.44 mm ² /s.

In a second aspect, a gas turbine engine is provided for an aircraft. The gas turbine engine includes a staged combustion chamber in which fuel is burned. The combustion chamber includes an arrangement of fuel spray nozzles and one or more igniters. The gas turbine engine includes a controller configured to direct a flow of fuel to a subset of the fuel spray nozzles adjacent the at least one igniter during a re-ignition procedure, the fuel having a calorific value of 'at least 43.5 MJ/kg.

The fuel can have a calorific value between 43.5 MJ/kg and 44 MJ/kg.

The fuel may have a calorific value of at least 43.8 MJ/kg.

The fuel can have a calorific value between 43.8 MJ/kg and 44 MJ/kg.

The fuel supplied to the combustion chamber may have an aromatic content of 10% or less by volume, preferably 5% or less, and more preferably 1% or less.

The gas turbine engine may include a fuel-oil heat exchanger. The controller may be configured to transfer heat from the oil to the fuel within the fuel-oil heat exchanger before the fuel enters the combustion chamber so as to lower the viscosity of the oil. fuel at 0.58 mm ² /s or less when entering the combustion chamber at cruise conditions, preferably 0.48 mm ² /s or less, preferably between 0.40 mm ² /s and 0 .48 mm ² /s and preferably between 0.42 mm ² /s and 0.44 mm ² /s.

According to a third aspect, a method of operating a gas turbine engine is provided. The gas turbine engine includes a staged combustion chamber in which fuel is burned. The combustion chamber includes a fuel spray nozzle arrangement in which fuel flow is directed to a subset of the fuel spray nozzles adjacent one or more igniters during a relighting procedure. The method includes providing fuel to the combustion chamber having an aromatic content of 10% or less by volume.

The inventors were aware that the aromatic content of the fuel has an effect on the manner in which the fuel is distributed to and ignited in the combustion chamber during a relighting procedure (e.g., droplet size from the spray nozzles of fuel, which can impact how the fuel is ignited). The aromatic content of the fuel must therefore be taken into account when distributing fuel to the fuel spray nozzles of the combustion chamber during a relighting procedure.

The method may include providing fuel to the combustion chamber having an aromatic content of 5% or less by volume.

The method may include providing fuel to the combustion chamber having an aromatic content of 1% or less by volume.

The fuel supplied to the combustion chamber may have a calorific value of at least 43.5 MJ/kg, preferably between 43.5 MJ/kg and 44 MJ/kg, or at least 43.8 MJ/kg and preferably between 43.8 MJ/kg and 44 MJ/kg.

The gas turbine engine may include a fuel-oil heat exchanger. The method may include transferring heat from the oil to the fuel before the fuel enters the combustion chamber so as to lower the viscosity of the fuel to 0.58 mm ² /s or less upon entry. in the combustion chamber at cruising conditions, preferably 0.48 mm ² /s or less, preferably between 0.40 mm ² /s and 0.48 mm ² /s and preferably between 0.42 mm2/s and 0.44 mm2/s.

In a fourth aspect, a gas turbine engine is provided for an aircraft. The gas turbine engine includes a staged combustion chamber in which fuel is burned. The combustion chamber includes an arrangement of fuel spray nozzles and one or more igniters. The gas turbine engine includes a controller configured to direct a flow of fuel to a subset of the fuel spray nozzles adjacent the one or more igniters during a re-ignition procedure, the fuel having a content of compounds aromatics of 10% or less by volume.

The fuel may have an aromatic content of 5% or less by volume.

The fuel may have an aromatic content of 1% or less by volume.

The fuel supplied to the combustion chamber may have a calorific value of at least 43.5 MJ/kg, preferably between 43.5 MJ/kg and 44 MJ/kg, or at least 43.8 MJ/kg and preferably between 43.8 MJ/kg and 44 MJ/kg.

The gas turbine engine may include a fuel-oil heat exchanger. The controller may be arranged to transfer heat from the oil to the fuel before the fuel enters the combustion chamber so as to lower the fuel viscosity to 0.58 mm ² /s or less when entering the combustion chamber at cruise conditions, preferably 0.48 mm ² /s or less, preferably between 0.40 mm ² /s and 0.48 mm ² /s and preferably between 0 .42 mm ² /s and 0.44 mm ² /s.

According to a fifth aspect, a method of operating a gas turbine engine is provided. The gas turbine engine includes a staged combustion chamber in which fuel is burned. The combustion chamber includes a fuel spray nozzle arrangement in which fuel flow is directed to a subset of the fuel spray nozzles adjacent one or more igniters during a relighting procedure. The method includes transferring heat from the oil to the fuel before the fuel enters the combustion chamber so as to lower the fuel viscosity to 0.58 mm2/s or less at the entrance to the chamber combustion at cruise conditions.

The inventors were aware that fuel viscosity has an effect on how fuel is distributed to and ignited in the combustion chamber (e.g., droplet size from fuel spray nozzles, which can impact on atomization and combustion efficiency) in cruise and/or during a relighting procedure (which can occur most frequently in cruise). Taking into account the viscosity of the fuel when distributing the fuel to the combustion chamber, and regulating it according to the situation by varying the thermal input to the fuel, can therefore enable more fuel combustion. efficient, improving the performance of the aircraft. Lower cruise fuel viscosity can lend itself to a more efficient engine and easier relighting during a relight procedure. In addition to more efficient fuel combustion, lower fuel viscosity can improve pump performance and can also improve pump longevity. Fuel viscosity must therefore be taken into account when delivering fuel to the combustion chamber fuel spray nozzles at cruise and/or during a relighting procedure.

The method may include transferring heat from the oil to the fuel before the fuel enters the combustion chamber so as to lower the fuel viscosity to between 0.58 mm ² /s and 0.30 mm ² / s upon entry into the combustion chamber at cruise conditions.

The method may include transferring heat from the oil to the fuel before the fuel enters the combustion chamber so as to lower the fuel viscosity to 0.48 mm ² /s or less at the fuel inlet. the combustion chamber at cruising conditions.

The method may include transferring heat from the oil to the fuel before the fuel enters the combustion chamber so as to lower the fuel viscosity to a value between 0.50 mm ² /s and 0.35 mm ² /s, or between 0.48 mm ² /s and 0.40 mm ² /s, or between 0.44 mm ² /s and 0.42 mm ² /s at the entrance to the combustion chamber at cruising conditions.

The method may include transferring heat from the oil to the fuel before the fuel enters the combustion chamber so as to lower the fuel viscosity to 0.57, 0.56, 0.55, 0, 54, 0.53, 0.52, 0.51, 0.50, 0.49, 0.48, 0.47, 0.46, 0.45, 0.44, 0.43, 0.42, 0.41, 0.40, 0.39, 0.38, 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31 or 0.30 mm ² / s or less when entering the combustion chamber at cruise conditions. The method may include transferring heat from the oil to the fuel before the fuel enters the combustion chamber so as to lower the viscosity of the fuel so that it is within a defined range between any two of the values of the previous sentence.

The gas turbine engine may include a fuel-oil heat exchanger.

The fuel supplied to the combustion chamber may have a calorific value of at least 43.5 MJ/kg, preferably between 43.5 MJ/kg and 44 MJ/kg, or at least 43.8 MJ/kg and preferably between 43.8 MJ/kg and 44 MJ/kg.

The fuel supplied to the combustion chamber may have an aromatic content of 10% or less by volume, preferably 5% or less, and more preferably 1% or less.

In a sixth aspect, a gas turbine engine is provided for an aircraft. The gas turbine engine includes a staged combustion chamber in which fuel is burned. The combustion chamber includes an arrangement of fuel spray nozzles and one or more igniters. The gas turbine engine includes a fuel-oil heat exchanger. The gas turbine engine includes a controller. The controller is configured to direct fuel flow to a subset of the fuel spray nozzles adjacent the igniter or igniters during a relighting procedure. The controller is configured to control operation of the fuel-oil heat exchanger to transfer heat from the oil to the fuel so as to lower the fuel viscosity to 0.58 mm ² /s or less when entering the combustion chamber at cruise conditions.

The controller may be configured to control operation of the fuel-oil heat exchanger to transfer heat from the oil to the fuel before the fuel enters the combustion chamber so as to lower the fuel viscosity between 0.58 mm ² /s and 0.30 mm ² /s when entering the combustion chamber at cruise conditions.

The controller may be configured to control operation of the fuel-oil heat exchanger to transfer heat from the oil to the fuel before the fuel enters the combustion chamber so as to lower the fuel viscosity at 0.48 mm ² /s or less when entering the combustion chamber at cruise conditions.

The controller may be configured to control operation of the fuel-oil heat exchanger to transfer heat from the oil to the fuel before the fuel enters the combustion chamber so as to lower the fuel viscosity at a value between 0.50 mm ² /s and 0.35 mm ² /s, or between 0.48 mm ² /s and 0.40 mm ² /s, or between 0.44 mm ² /s and 0.42 mm ² /s when entering the combustion chamber at cruise conditions.

The controller may be configured to control operation of the fuel-oil heat exchanger to transfer heat from the oil to the fuel before the fuel enters the combustion chamber so as to lower the fuel viscosity at 0.57, 0.56, 0.55, 0.54, 0.53, 0.52, 0.51, 0.50, 0.49, 0.48, 0.47, 0, 46, 0.45, 0.44, 0.43, 0.42, 0.41, 0.40, 0.39, 0.38, 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31 or 0.30 mm ² /s or less when entering the combustion chamber at cruise conditions. The controller may be configured to control operation of the fuel-oil heat exchanger to transfer heat from the oil to the fuel before the fuel enters the combustion chamber so as to lower the fuel viscosity so that it lies within a defined range between any two of the values in the previous sentence.

The fuel supplied to the combustion chamber may have a calorific value of at least 43.5 MJ/kg, preferably between 43.5 MJ/kg and 44 MJ/kg, or at least 43.8 MJ/kg and preferably between 43.8 MJ/kg and 44 MJ/kg.

The fuel supplied to the combustion chamber may have an aromatic content of 10% or less by volume, preferably 5% or less, and more preferably 1% or less.

The characteristics of the following statements may apply to any of the above:

In any of the aforementioned aspects, the number of fuel spray nozzles may be between 14 and 22 and/or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6.

The (total) number of fuel spray nozzles can be between 16 and 20.

The number of fuel spray nozzles may be 14, 15, 16, 17, 18, 19, 20, 21, 22, or a number within a defined range between any two of the values in this sentence.

The number of fuel spray nozzles per unit engine core size may be in the range of 2.5 to 4.5, and more preferably in the range of 3 to 4.

The number of fuel spray nozzles per unit engine core size may be 2, 3, 4, 5, 6, or within a defined range between any two of these values, and more preferably 2.5, 3 , 3.5, 4 or 4.5, or within a defined range between any two of these values, and even more preferably 3.0, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, or 4.0, or within a defined range between any two of these values.

The number of fuel spray nozzles per unit size of engine core can be 2.0, 2.1, 2.2, 2.3, 2.4, 2.5, 2.6, 2.7, 2 .8, 2.9, 3.0, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4.0 , 4.1, 4.2, 4.3, 4.4, 4.5, 4.6, 4.7, 4.8, 4.9, 5.0, 5.1, 5.2, 5 ,3, 5.4, 5.5, 5.6, 5.7, 5.8, 5.9 or 6.0, or within a defined range between any two of these values.

The motor core size may be in the range of 4 to 7, and more preferably in the range of 5 to 6, and even more preferably in the range of 5.2 to 5.5.

Unless otherwise noted, all references to "core size" herein are in sK ^1/2 .in. units and all references to "number of fuel spray nozzles per unit engine core size" are to number of nozzles per unit size of motor core also in sK ^1/2 .in.

In any of the aforementioned aspects, the fuel spray nozzle subset may include at least half of the total number of fuel spray nozzles.

In any of the aforementioned aspects, the fuel spray nozzle subset may comprise at least two thirds of the total number of fuel spray nozzles.

In any of the aforementioned aspects, the fuel spray nozzle arrangement may include duplex nozzles and single flow nozzles.

In any of the aforementioned aspects, the fuel spray nozzle subset may include the duplex nozzles and the remaining fuel spray nozzles may include the single flow nozzles. The fuel spray nozzle sub-assembly may be a first fuel spray nozzle sub-assembly and includes the duplex nozzles and the fuel spray nozzle arrangement may include a second fuel spray nozzle sub-assembly. fuel including single flow nozzles.

The combustion chamber may have 10 to 14 duplex fuel spray nozzles and 4 to 8 single flow fuel spray nozzles.

Duplex fuel spray nozzles can be arranged in groups around the circumference of the combustion chamber.

The duplex fuel spray nozzle groups may include two or more groups arranged diametrically opposite each other.

Each group of duplex fuel spray nozzles can include 2-8 nozzles.

In any of the aforementioned aspects, the combustion chamber may include at least two igniters and the fuel spray nozzle subassembly may include at least two groups of nozzles. Each group of nozzles may be adjacent to one of the igniters.

In any of the aforementioned aspects, the gas turbine engine may include a plurality of igniters. The plurality of igniters may be arranged symmetrically around a circumference of the combustion chamber. Pairs of igniters may be arranged diametrically opposed to each other around a circumference of the combustion chamber.

Each igniter may be arranged adjacent one or more of the duplex fuel spray nozzles.

Each igniter may be arranged adjacent to one of the groups of duplex fuel spray nozzles.

The combustion chamber can include 1 to 8 igniters.

As noted elsewhere in this document, this description may apply to any relevant gas turbine engine configuration. Such a gas turbine engine may be, for example, a turbofan gas turbine engine, an open rotor gas turbine engine (in which the propeller is not surrounded by a nacelle), a turboprop engine. or a turbojet. Any such engine may or may not be provided with an afterburner. Such a gas turbine engine can be, for example, designed for power generation applications on land or at sea.

A gas turbine engine in accordance with any aspect of the present disclosure may include an engine core including a turbine, a combustion chamber, 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 can be located upstream of the engine core. Alternatively, in certain examples, the gas turbine engine may include a fan located downstream of the engine core, for example when the gas turbine engine is an open rotor or a turboprop (in which case the fan may be called a propeller). ).

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

When the gas turbine engine is an open rotor or turboprop engine, one or more of the propeller stages may be driven by a reduction gear. The reducer may be of the type described here.

An engine according to the present description may be a turbofan engine. Such an engine may be a direct-drive turbofan engine in which the fan is directly connected to the fan drive turbine, for example without a reduction gear, via a core shaft. In such a direct-drive turbofan engine, the fan can be said to rotate at the same rotational speed as the fan drive turbine. Strictly by way of example, the fan drive turbine may be a first turbine, the core shaft may be a first core shaft, and the gas turbine engine may further include a second turbine and a second core shaft connecting the second turbine to the compressor. The second turbine, the compressor and the second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft. In such an arrangement, the second turbine can be positioned axially upstream of the first turbine.

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

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

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

The reduction gear may be arranged to be driven by the core shaft which is configured to rotate (e.g. in use) at the lowest rotational speed (e.g. the first core shaft in the example below). above). For example, the gearbox may be arranged to be driven only by the core shaft which is configured to rotate (e.g. in use) at the lowest rotational speed (e.g. only by the first core shaft , and not the second heart tree, in the example above). Alternatively, the gearbox can be arranged to be driven by any shaft or shafts, for example the first and/or second shafts in the example above.

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

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

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

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

The or each turbine (e.g. the first turbine and the second turbine as described above) may comprise any number of stages, e.g. multiple stages. Each stage may include 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 can be axially offset from each other. The second turbine (or “high pressure” turbine) can comprise 2 stages in any arrangement (e.g. whether it is a gear or direct drive motor). The gas turbine engine may be a direct drive gas turbine engine comprising a first turbine (or "low pressure" turbine) having 5, 6 or 7 stages. Alternatively, the gas turbine engine may be a "geared" gas turbine engine comprising a first turbine (or "low pressure" turbine) having 3 or 4 stages.

Each fan blade may be defined as having a radial span extending from one foot (or hub) at a gas-washed radially internal location, or 0% span position, to one end at a 100% reach 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 of the order of) any one 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 can be included in a range bounded by any two of the values in the preceding sentence (i.e. (say that the values can form upper or lower limits), for example in the range 0.28 to 0.32 or 0.29 to 0.30. These ratios may commonly be referred to as the hub to end ratio. Both the hub radius and the tip radius can be measured at the leading edge (or axially most forward) portion of the blade. The hub-to-tip ratio refers, of course, to the gas-washed portion of the fan blade, that is, the portion radially outboard of any platform.

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

Blower rotation speed may vary during use. Generally, the rotation speed is lower for blowers with a larger diameter. Strictly by way of non-limiting example, the rotation speed of the fan at cruising conditions may be less than 3500 rpm, for example less than 2600 rpm, or less than 2500 rpm, or less than 2300 rpm. Strictly by way of further non-limiting example, the fan rotation speed at cruise conditions for a "gear" gas turbine engine having a fan diameter in the range from 200 cm to 210 cm may be within the range from 2750 to 2900 rpm. Strictly by way of further non-limiting example, the fan rotation speed at cruise conditions for a "gear" gas turbine engine having a fan diameter in the range from 210 cm to 230 cm may be within the range going from 2500 to 2800 rpm. Strictly by way of further non-limiting example, the fan rotation speed at cruise conditions for a "gear" gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be within the range from 1500 to 1800 rpm. Strictly as a further non-limiting example, the fan rotation speed at cruising conditions for a direct drive motor having a fan diameter in the range of 190 cm to 200 cm may be in the range of 3600 at 3900 rpm. Strictly by way of further non-limiting example, the fan rotation speed at cruising conditions for a direct drive motor having a fan diameter in the range of 300 cm to 340 cm may be in the range of 2000 at 2800 rpm.

In use of the gas turbine engine, the fan (with associated fan blades) rotates about an axis of rotation. This rotation results in a movement of the end of the fan blade with a speed U _tip . The work accomplished by the fan blades on the flow results in an increase in enthalpy dH of the flow. A fan tip load can be defined by dH/U _tip ² , where dH is the enthalpy increase (e.g. 1-D average enthalpy increase) across the fan and U _tip is the speed (translation) 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 load at cruise conditions may be greater than (or of the order of) any of: 0.28, 0.29, 0.30, 0.31, 0.32, 0 .33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39, or 0.4 (all values dimensionless). The blower tip load may be in an inclusive range delimited 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 0.31, or 0.29 to 0.3 (e.g. for a geared gas turbine engine).

Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio (BPR), where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the heart. In certain arrangements the bypass ratio at cruising conditions may be greater than (or of the order of) any of the following: 9, 9.5, 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The ratio bypass at cruise conditions may be in an inclusive range delimited by any two of the values in the preceding sentence (i.e. the values may form upper or lower limits), for example in the range 12 to 16, 13 to 15, or 13 to 14. Strictly by way of non-limiting example, the bypass ratio at cruise conditions of a direct drive gas turbine engine according to the present description may be in the range from 9:1 to 11:1. Strictly by way of further non-limiting example, the bypass ratio at cruise conditions of a geared gas turbine engine according to the present disclosure may be in the range of 12:1 to 15:1. The bypass conduit may be substantially annular. The bypass conduit may be radially outboard of the core motor. The radially external 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 stagnation pressure at the outlet of the highest pressure compressor (before entering the chamber combustion) at the stagnation pressure upstream of the blower. By way of non-limiting example, the overall pressure ratio of a gas turbine engine as described and/or claimed here at cruise conditions may be greater than (or of the order of) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio can be in an inclusive range delimited 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 50 to 70. Strictly by way of non-limiting example, the overall pressure ratio at cruising conditions of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 40 to 45. Strictly by way of non-limiting example, the overall pressure ratio at cruising conditions of a turbine engine at gear gas having a fan diameter in the range of 210 cm to 230 cm may be in the range of 45 to 55. Strictly by way of non-limiting example, the overall pressure ratio at cruise conditions of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 50 to 60. Strictly by way of non-limiting example, the overall pressure ratio at operating conditions cruise 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 can be defined as the net thrust of the engine divided by the total mass flow through the engine. In some examples, the specific thrust may depend, for a given thrust condition, on the specific composition of the fuel supplied to the combustion chamber. At cruise conditions, the specific thrust of an engine described and/or claimed here may be less than (or of the order of) any of the following: 110 Nkg ^-1 s, 105 Nkg ^-1 s, 100 Nkg ^-1 s, 95 Nkg ^-1 s, 90 Nkg ^-1 s, 85 Nkg ^-1 s or 80 Nkg ^-1 s. The specific thrust may be in an inclusive range delimited by any two of the values in the previous sentence (i.e. the values may form upper or lower limits), for example in the range 80 Nkg ^-1 s at 100 Nkg ^-1 s, or from 85 Nkg ^-1 s to 95 Nkg ^-1 s. Such engines can be particularly efficient compared to conventional gas turbine engines. Strictly by way of non-limiting example, the specific thrust of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 90 Nkg ^-1 s to 95 Nkg ^-1 s. Strictly by way of non-limiting example, the specific thrust of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 80 Nkg ^-1 s to 90 Nkg ^-1 s. Strictly by way of non-limiting example, the specific thrust of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 70 Nkg ^-1 s to 90 Nkg ^-1 s. Strictly as a non-limiting example, the specific thrust of a direct drive geared gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 90 Nkg ^{- 1} s at 120 Nkg ^-1 s.

A gas turbine engine as described and/or claimed herein can have any desired maximum thrust. Strictly by way of non-limiting example, a gas turbine as described and/or claimed here may be capable of producing a maximum thrust of at least (or of the order of) any of the following: 100 kN , 110 kN, 120 kN, 130 kN, 135 kN, 140 kN, 145 kN, 150 kN, 155 kN, 160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN, 450 kN, 500 kN or 550 kN. The maximum thrust may be an inclusive range bounded by any two of the values in the preceding sentence (i.e., the values may form upper or lower limits). Strictly by way of non-limiting example, a gas turbine as described and/or claimed herein may be capable of producing maximum thrust in the range from 155 kN to 170 kN, 330 kN to 420 kN, or 350 kN to 400 kN. Strictly 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. Strictly as a non-limiting example, the maximum thrust of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 150 kN to 200 kN. Strictly as a non-limiting example, the maximum thrust of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 370 kN to 500 kN. Strictly by way of non-limiting example, the maximum thrust of a direct drive geared gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 370 kN to 500kN. The thrust mentioned above may be the maximum net thrust under typical atmospheric conditions at sea level plus 15 degrees C (ambient pressure 101.3 kPa, temperature 30 degrees C), with the engine static.

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

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

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

A fan as described and/or claimed herein may comprise a central portion, from which the fan blades may extend, for example in a radial direction. The fan blades can be connected to the central part in any desired way. For example, each fan blade may include a fastener that may engage a corresponding notch in the hub (or disk). Strictly by way of example, such a fixing element may be in the form of a dovetail which can notch into and/or engage a corresponding notch in the hub/disc in order to secure the fan blade to 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 bladed disk or a bladed crown. Any suitable method may be used to manufacture such a bladed disk or bladed crown. 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 connected to the hub/disc by welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may not be provided with a variable area nozzle (VAN). Such a nozzle with a variable section can make it possible to vary the outlet area of the bypass duct during use. The general principles of this description can apply to engines with or without VAN.

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

As used herein, the terms idle, taxi, takeoff, climb, cruise, descent, approach, and landing have the conventional meaning and would be easily understood by those skilled in the art. Thus, for a given gas turbine engine intended for an aircraft, those skilled in the art will immediately recognize each term as referring to a phase of operation of the engine within the framework of a given mission of an aircraft to which the engine gas turbine is designed to be fixed.

In this sense, ground idling can refer to a phase of engine operation where the aircraft is stationary and in contact with the ground, but where there is a requirement to rev the engine. During idle, the engine can produce between 3% and 9% of the available engine thrust. In additional non-limiting examples, the engine may produce between 5% and 8% of the available thrust. In additional non-limiting examples, the engine may produce between 6% and 7% of the available thrust. Taxiing may refer to a phase of engine operation during which the aircraft is propelled to the ground by the thrust produced by the engine. While taxiing, the engine can produce between 5% and 15% of the available thrust. In additional non-limiting examples, the engine may produce between 6% and 12% of the available thrust. In additional non-limiting examples, the engine may produce between 7% and 10% of the available thrust. Takeoff may refer to a phase of engine operation during which the aircraft is propelled by the thrust produced by the engine. During an initial stage of the take-off phase, the aircraft can be propelled while it is in contact with the ground. During a later stage of the take-off phase, the aircraft can be propelled while it is not in contact with the ground. During takeoff, the engine can produce between 90% and 100% of the available thrust. In additional non-limiting examples, the engine may produce between 95% and 100% of the available thrust. In additional non-limiting examples, the engine may produce 100% of the available thrust.

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 can produce between 75% and 100% of the available thrust. In additional non-limiting examples, the engine may produce between 80% and 95% of the available thrust. In additional non-limiting examples, the engine may produce between 85% and 90% of the available thrust. In this sense, climb can refer to a phase of operation during an aircraft flight cycle between takeoff and arrival at cruise conditions. Additionally or alternatively, climb may refer to a nominal point in an aircraft flight cycle between takeoff and landing, where a relative increase in altitude is required, which may require an additional thrust request from the engine.

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

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

Strictly by way of example, the forward speed at cruise condition may be any point in the range from Mach 0.7 to 0.9, e.g. 0.75 to 0.85, e.g. 0, 76 to 0.84, for example 0.77 to 0.83, for example 0.78 to 0.82, for example 0.79 to 0.81, for example of the order of Mach 0.8, of l order of Mach 0.85 or in the range from 0.8 to 0.85. Any single speed within these ranges can be part of the cruise condition. For a certain aircraft, cruise conditions may be outside of these ranges, for example below Mach 0.7 or above Mach 0.9.

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

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

Strictly as an additional example, cruise conditions may correspond to a forward Mach number of 0.85 and typical atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38,000 feet (11 582m). At such cruise conditions, the engine can provide a known required level of net thrust. The known required net thrust level is, of course, dependent on the motor and its intended application and may be, for example, a value in the range 35 kN to 65 kN.

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

In addition, those skilled in the art would immediately recognize one and/or the other of the terms “descent” and “approach” to designate a phase of operation within a flight cycle of an aircraft, between cruise and the landing of the aircraft. During either descent and approach, the engine can produce between 20% and 50% of the available thrust. In additional non-limiting examples, the engine may produce between 25% and 40% of the available thrust. In additional non-limiting examples, the engine may produce between 30% and 35% of the available thrust. Additionally or alternatively, descent may refer to a nominal point in an aircraft flight cycle between takeoff and landing, where a relative decrease in altitude is required, and which may require a reduced thrust demand of the motor.

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

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

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

Those skilled in the art would understand that, unless mutually exclusive, a characteristic or parameter described in relation to any of the above aspects can be applied to any other aspect. Furthermore, unless mutually exclusive, any characteristic or any parameter described here may be applied to any aspect and/or associated with any other characteristic or any other parameter described here.

As used herein, a range "from value X to value Y" or "between value X and value Y", or the like, means an inclusive range; including the bounding values of X and Y.

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

is a side sectional view of a gas turbine engine;

is a side view in close section of an upstream part of a gas turbine engine;

is a partially cut away view of a reduction gear for a gas turbine engine;

is a cross section through an engine combustion chamber of the in a plane normal to the main axis of rotation of the engine;

is a schematic cross section of a duplex fuel spray nozzle of the combustion chamber of the ;

is a schematic cross section of a single flow fuel spray nozzle of the combustion chamber of the ;

is a partial cross-sectional view of the engine of the ;

is another partial cross-sectional view of the engine of the ;

is a schematic representation of a propulsion system for an aircraft comprising the engine of the ;

, And show methods of operating a gas turbine engine with respect to fuel parameters including calorific value, aromatic content and viscosity upon entering the combustion chamber at cruise conditions.

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

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

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

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

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

The epicyclic gearbox 30 illustrated by way of example in Figures 2 and 3 is of the planetary type, in that the planet carrier 34 is coupled to an output shaft via connections 36, with the toothed ring 38 fixed . However, any other suitable type of epicyclic gearbox 30 may be used. As a further example, the epicyclic gearbox 30 may be a star arrangement, in which the planet carrier 34 is held fixed, with the toothed ring (or ring) 38 allowed to rotate. In such an arrangement, the fan 23 is driven by the ring gear 38. As another alternative example, the gearbox 30 can be a differential gearbox in which the ring gear 38 and the planet carrier 34 are one and the same. 'other allowed to turn.

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

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

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

Other gas turbine engines to which the present description may be applied may have alternative configurations. For example, such engines may have another number of compressors and/or turbines and/or another number of interconnecting shafts. As an additional example, the gas turbine engine shown in has a split flow nozzle 18, 20 which means that the flow through the bypass conduit 22 has its own nozzle 18 which is independent of, and radially outside of, the core motor nozzle 20. However, this is not limiting, and any aspect of the present description may also apply to engines in which the flow through the bypass conduit 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be called a mixed flow nozzle. One and/or the other of the nozzles (whether mixed or divided flow) can have a fixed or variable area.

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

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

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

The functional performance of a given composition, or blend, of fuel for use in a given mission may be defined, at least in part, by the ability of the fuel to power 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 particles. Relatively higher specific energy (i.e. energy per unit mass), expressed in MJ/kg, can at least partially reduce takeoff weight, potentially providing a relative improvement in fuel efficiency. A relatively higher energy density (i.e. energy per unit volume), expressed in MJ/L, can at least partially reduce the volume of fuel at takeoff, which may be particularly important for long-range missions. limited volume or military operations involving refueling. Relatively higher thermal stability (i.e., inhibition of fuel degradation or scale formation under thermal stress) can allow the fuel to withstand high temperatures in the engine and fuel injectors, thus potentially providing relative improvements in combustion efficiency. Reduced emissions, particularly of particulate matter, can help reduce the formation of contrails, while reducing the environmental impact of a given mission. Other fuel properties may also be critical to functional performance. For example, a relatively lower freezing point (°C) may allow long-distance missions to optimize flight profiles; Minimum aromatic compound concentrations (%) may ensure sufficient swelling of certain materials used in the construction of O-rings and gaskets that have been previously exposed to fuels with high aromatic content; and, maximum surface tension (mN/m) can ensure sufficient spray breakup and fuel atomization.

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

A number of sustainable aviation fuel blends have been approved for use. For example, some approved blends include blend ratios of up to 10% sustainable aviation fuel, while other approved blends include blend ratios of between 10% and 50% sustainable aviation fuel (the remainder comprising a or several hydrocarbon fuels of fossil origin, such as kerosene), other compositions awaiting approval. However, it is anticipated in 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, iso-alkanes, cycloalkanes, and aromatics, and may be produced, for example, from one or more of syngases (syngases). ; lipids (e.g. fats, oils and fats); sugars; and alcohols. Thus, sustainable aviation fuels may include one or both of a lower content of aromatic compounds and sulfur, compared to hydrocarbon fuels of fossil origin. Additionally or alternatively, sustainable aviation fuels may include either/or a higher iso-alkane and cyclo-alkane content, relative to fossil-based hydrocarbon fuels. Thus, in certain examples, sustainable aeronautical fuels may comprise 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. kerosene.

Due at least in part to the molecular structure of sustainable aviation fuels, sustainable aviation 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; a lower freezing point; lower soot emissions; and, lower CO ₂ emissions, compared to hydrocarbon fuels of fossil origin (for example when burned in combustion equipment 16). Thus, compared to hydrocarbon fuels of fossil origin, such as kerosene, sustainable aeronautical fuels can result in one or/and the other of a relative reduction in specific fuel consumption, and a relative reduction in fuel costs. 'interview.

There shows a section through the combustion chamber 16 of the engine 10 of the in a plane normal to the main axis of rotation 9 of the engine 10. The combustion chamber 16 comprises an annular combustion chamber 401, defined by a liner 402. Alternative combustion chamber configurations can be used in other examples , for example cannular, canned, etc. The combustion chamber 16 includes a plurality of fuel spray nozzles 403, 404 arranged around the circumference of the combustion chamber 16. Each fuel spray nozzle 403, 404 includes one or more fuel injectors arranged to inject fuel in combustion chamber 401. In this example, combustion chamber 16 includes 16 fuel spray nozzles 403, 404. In other examples, combustion chamber 16 may include any suitable number of spray nozzles fuel spray nozzles 403, 404, for example a number of fuel spray nozzles in the range of 14 to 22. In some examples the number of fuel spray nozzles 403, 404 may be between 16 and 20. In others In further examples, the number of fuel spray nozzles may be 14, 15, 16, 17, 18, 19, 20, 21, 22, or a number within a defined range between any two of the values in this sentence .

The core size defines the size of core 11 of engine 10. The engine core size can be defined as:

Or = the mass flow rate, in pounds per second, of the air entering the high pressure compressor 15, = the temperature, in Kelvin, of the air at the outlet of the high pressure compressor 15, and = the pressure, in pounds inches per second squared per inch squared, of the air at the outlet of the high pressure compressor 15. A unit of core size is therefore expressed as:

The core size (in s.K ^1/2 .in ) of the motor can be between 4 and 7, for example 4, 4.5, 5, 5.5, 6, 6.5, or 7, or any range defined between any two of these values. In some examples, the motor core size (in s.K ^1/2 .in ) may be in the range of 5.0, 5.1, 5.2, 5.3, 5.4, 5.5, 5 ,6, 5.6, 5.7, 5.8, 5.9, or 6, or any defined range between any two of these values. In still other examples, the motor core size (in s.K ^1/2 .in ) may be in the range of 5.25, 5.26, 5.27, 5.28, 5.29, 5, 30, 5.31, 5.32, 5.33, 5.34, 5.35, 5.36, 5.37, 5.38, 5.39, 5.40, 5.41, 5.42, 5.43, 5.44, or 5.45, or any defined range between any two of these values.

The number of fuel spray nozzles per unit engine core size (in the units given above) may be in the range 2 to 6, for example, 2, 3, 4, 5, 6, or within of a defined range between any two of these values. In some preferred examples, the number of fuel spray nozzles per unit engine core size may be in the range of 2.5 to 4.5, for example 2.5, 3, 3.5, 4 or 4, 5, or any defined range between any two of these values. In still other examples, the number of fuel spray nozzles per unit engine core size may be in the range of 3 to 4, e.g. 3.0, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, or 4.0, or any defined range between any two of these values. In still other examples, the number of fuel spray nozzles per unit engine core size may be 2.0, 2.1, 2.2, 2.3, 2.4, 2.5, 2, 6, 2.7, 2.8, 2.9, 3.0, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4.0, 4.1, 4.2, 4.3, 4.4, 4.5, 4.6, 4.7, 4.8, 4.9, 5.0, 5, 1, 5.2, 5.3, 5.4, 5.5, 5.6, 5.7, 5.8, 5.9, or 6.0, or within a defined range between any two of these values.

The core size is defined here at an engine operating condition corresponding to a maximum value of the semi-non-dimensional flow at the inlet of the high pressure compressor, defined as:

Or = the mass flow rate, (in pounds per second) of the air entering the high pressure compressor, = the temperature (in Kelvin) of the air at the inlet of the high pressure compressor, and is the pressure (in pounds inches per second squared per inch squared) of the air at the inlet of the high pressure compressor.

The operating condition corresponding to the maximum semi-non-dimensional flow at the high pressure compressor inlet may be the end of rise operating condition. The core size referred to here can therefore be defined at the end of rise operating condition. The end of rise may be as defined in the art and as understood by those skilled in the art for a specific implementation of a gas turbine of the present application. In a specific example, the end of climb may correspond to operation at an altitude between 30,000 feet and 39,000 feet (more specifically 35,000 feet), a forward speed of a Mach number 0.75 to 0.85 , and an ambient air temperature (DTAMB) of ISA+10 K to ISA+15 K.

The combustion chamber 16 includes a number of duplex fuel spray nozzles 403 (also known as internally staged nozzles) in which a primary fuel injector is integrated into the same fuel nozzle as a fuel injector main. Combustion chamber 16 also includes a number of single flow fuel spray nozzles 404 which each include a main fuel injector only. In other examples, combustion chamber 16 may include only duplex fuel spray nozzles or only single flow fuel spray nozzles.

In this example, the combustion chamber 16 includes 12 duplex fuel spray nozzles 403 and 4 single flow fuel spray nozzles 404. The duplex fuel spray nozzles 403 are illustrated in the by shaded circles. The duplex fuel spray nozzles 403 are arranged in groups of three around the circumference of the combustion chamber 16, each group being arranged diametrically opposite another of the groups. In other examples, the combustion chamber 16 may include any suitable number of duplex fuel spray nozzles, for example in the range of 10 to 14 nozzles, and any suitable number of duplex fuel spray nozzles. single flow fuel, for example in the range of 4 to 8 nozzles. In some examples, the number of duplex fuel spray nozzles may be 10, 11, 12, 13 or 14, or within a defined range between any two of the values in this sentence. In some examples, the number of single flow fuel spray nozzles may be 4, 5, 6, 7 or 8, or within a defined range between any two of the values in this sentence. The duplex fuel spray nozzles may be arranged in any suitable number of groups, or may not be arranged in groups. Where applicable, each group of duplex fuel spray nozzles may include any suitable number of nozzles, for example in the range of 2 to 8 nozzles. In some examples, each group of duplex nozzles may include 2, 3, 4, 5, 6, 7 or 8 fuel spray nozzles, or a number within a defined range between any two of these values. The nozzle arrangement shown on the may be all of the fuel spray nozzles provided in the combustion chamber 16. In other examples, other fuel spray nozzles may be provided in the combustion chamber 16 in addition to those described.

The combustion chamber 16 further comprises four igniters 405 arranged to ignite an air-fuel mixture in the combustion chamber 401 in operation. Each igniter 405 is arranged adjacent to one of the groups of duplex fuel spray nozzles 403. The duplex nozzles 403 are for this reason each located closer to a respective igniter (e.g. its nearest igniter) compared to single flow nozzles 404. Each igniter 405 is arranged diametrically opposite another of the igniters 405. In other examples, the combustion chamber may include fewer or more igniters, for example a number of igniters in the range from 1 to 8, and the igniters can be arranged differently. For example, one or more of the igniters may not be arranged adjacent to one of the groups of duplex fuel spray nozzles and one or more of the igniters may not be arranged diametrically opposite another of the igniters. In some examples, the combustion chamber may include 1, 2, 3, 4, 5, 6, 7 or 8 igniters, or a number within a defined range between any two of the values in this phrase.

In the example shown, when the engine 10 is operating at low power (below a staging point), for example during or shortly after re-ignition or start-up, fuel is supplied only to the primary injectors of the spray nozzles of duplex fuel 403 for distribution to the combustion chamber 401. A greater flow of fuel is therefore provided to the duplex nozzles 403 compared to the single flow nozzles 404 below the staging point. As the power output of the engine 10 and the mass flow rate of air through the engine 10 increase, the staging point is reached at which additional fuel is supplied to the main fuel injectors of one or more duplex fuel spray nozzles 403 and to the main fuel injectors of one or more of the single flow fuel spray nozzles 404 for delivery to the combustion chamber 401. In the present example, at higher power levels , fuel is injected by all the primary fuel injectors of both the duplex fuel spray nozzles 403 and the single flow fuel spray nozzles 404, in addition to the fuel injected by the primary injectors of the fuel spray nozzles duplex 403. In this example, the fuel flow rate supplied to the main injectors of the single flow fuel spray nozzles 404 is less than or equal to the fuel flow rate supplied to the main injectors of the duplex fuel spray nozzles 403. For this reason, Because both the primary and main injectors of the duplex fuel spray nozzles 403 receive fuel, the duplex fuel spray nozzles 403 receive more fuel than the single flow fuel spray nozzles 404 at and above the staging point. In an alternative example, fuel is supplied only to the main fuel injectors of one or more of the duplex fuel spray nozzle(s) 403 and to the main fuel injectors of one or more of the flow fuel spray nozzles single 404 at and above the staging point, i.e., fuel is not supplied to the primary injectors of the duplex fuel spray nozzles 403.

The flow of fuel distributed to the plurality of fuel spray nozzles is therefore directed to a subset of the fuel spray nozzles adjacent the igniters. For example, fuel flow is biased such that the flow of fuel to a first subset of the plurality of fuel spray nozzles (the duplex fuel spray nozzles 403 in the present example) is greater than that distributed to a second subset of the fuel spray nozzles (the single flow fuel spray nozzles 404 in the present example). This may allow a primary fuel flow to be supplied to fuel spray nozzles which are located relatively closer to the igniters 405 to facilitate ignition and flame stability at low engine powers, upon engine startup, or during engine restart. In some examples, the first subset (e.g. duplex nozzles) of fuel spray nozzles may comprise at least half, preferably at least two thirds, of the total number of fuel spray nozzles.

In other examples, the rate of fuel flow to each fuel spray nozzle provided into the combustion chamber may be the same and there may be no bias on fuel flow to a subassembly. nozzles. In such an example, all of the fuel flow nozzles may be single flow nozzles or they may all be duplex nozzles. In still other examples, other arrangements of fuel spray nozzles may be provided in which fuel is directed to those adjacent to, or closer to, the igniters. For example, one can provide two subassemblies (which are independently controllable) of duplex nozzles or two subassemblies of single flow nozzles which can be actuated in the manner described above.

There shows one of the duplex fuel spray nozzles 403 of the combustion chamber 16.

The duplex nozzle 403 comprises a primary fuel injector 501, a main fuel injector 502 and an air conduit 503. The primary injector 501 comprises a primary inlet 504 arranged to receive a primary flow of fuel P and a fuel circuit primary 505 arranged to distribute the primary flow of fuel to the outlet 506 of the nozzle 403. The main injector 502 comprises a main inlet 507 arranged to receive a main flow of fuel M and a main fuel circuit 508 arranged to distribute the main fuel flow to outlet 506 of nozzle 403. Air conduit 503 receives high pressure air from high pressure compressor 15 and distributes high pressure air to outlet 506 of nozzle 403.

Duplex nozzle 403 is designed to produce, at outlet 506 of nozzle 403, a primary cone of fuel from primary injector 501 and a primary cone of fuel from primary injector 502 (shown in there by the dotted lines marked P and M respectively). When both of the primary and main injectors 501, 502 are active, the primary and main cones are arranged concentrically, the main cone being arranged annularly outside the primary cone. The skilled person will be familiar with such fuel spray patterns.

We will have in mind that the duplex nozzle 403 of the is given as an example only and that other examples may use an alternative duplex nozzle configuration.

There shows one of the single flow fuel spray nozzles 404 of the combustion chamber 16. The nozzle 404 comprises a main fuel injector 601, comprising a main inlet 602 arranged to receive a main flow of fuel M and a circuit of main fuel 603 arranged to distribute the main flow of fuel to the outlet 604 of the nozzle 404. The nozzle 404 is configured to produce a main cone of fuel at the outlet 604 of the nozzle 404 (shown by the dotted lines marked M). Air is similarly supplied to the outlet 604 of the nozzle via an air conduit 605.

It will be borne in mind that the single-flow fuel spray 404 of the is given as an example only and that other examples may use an alternative single flow fuel spray nozzle configuration 404.

Figures 7 and 8 each show a section through the engine 10, perpendicular to the main axis of rotation 9, comprising a part of the combustion chamber 16 comprising one of the duplex fuel spray nozzles 403 and one igniters 405. A similar arrangement is provided at the location of the single flow fuel spray nozzles 404 as for the duplex nozzles 403. The combustion chamber 16 is mounted within a cavity 406 formed by a housing of internal air 407 and an external air housing 408. In operation, the high pressure compressor 15 distributes high pressure air D to the cavity 406 via a diffuser 409. At this point, a quantity air enters the combustion chamber 16 as combustion air E via the fuel nozzle 403 and/or mixing ports at the inlet to the combustion chamber 16. remaining air flows around the combustion chamber 16 as cooling air G, part of which is admitted downstream of the fuel nozzle 403 as described below with reference to the .

One or more temperature and/or pressure probes (not shown) may be installed in the diffuser housing 409 and arranged to measure the temperature and/or pressure of the high pressure air D distributed to the cavity 406 from of the high pressure compressor 15 via the diffuser 409 (i.e. the temperature and pressure at the outlet of the high pressure compressor 15). Such a temperature probe may be called a T3 probe and such a pressure probe may be called a P3 probe. It will be borne in mind that the motor 10 may include any suitable arrangement of pressure and temperature probes which may be positioned at any suitable location within the motor 10. As used here, T3 and P3, and any other numbered pressures and temperatures, can be defined using the position numbering given in SAE AS755.

Combustion chamber 16 operates as a rich, rapid quench, lean (RQL) combustion chamber. In other examples, combustion chamber 16 may be an alternative type of combustion chamber, such as a standard rich combustion combustion chamber (without fuel flow bias). In reference to the , the combustion chamber 401 of the RQL combustion chamber 16 is divided into three zones along the length of the combustion chamber 16: a rich zone 801, a rapid extinguishing zone 802 and a lean zone 803. In operation , a rich air-fuel mixture is introduced into the rich zone 801 from the fuel spray nozzle 403 where it is ignited by the igniter 405. Within the rich zone 801 the fuel is burned at a fuel/fuel ratio. air above the stoichiometric ratio (for example, at an equivalence ratio of approximately 1.8). Air is then introduced into the combustion products, via primary orifices 804 arranged in the liner 402 of the combustion chamber 16, before the combustion products reach the rapid extinguishing zone 802 Additional air is added to the still burning fuel through primary ports 804 (which may be referred to as quenching ports). Air is added through the primary ports 804 at a higher rate (e.g., higher than in the rich zone) which quenches combustion at a ratio significantly lower than the stoichiometric fuel/air ratio (e.g., at a ratio equivalence between 0.5 and 0.7), while continuing to let the fuel burn. Therefore, a very small part of the combustion process can be carried out at near stoichiometric fuel/air ratios, and therefore relatively little nitrogen oxides (NOx) are produced. Air is then reintroduced into the combustion products, via secondary orifices 805 arranged in the liner 402 of the combustion chamber 16, while the combustion products are in the lean zone 803 (or just before until they reach the poor zone 803). Within lean zone 803, fuel is burned at an air/fuel ratio less than the stoichiometric ratio (e.g., at an equivalence ratio between 0.5 and 0.7). After passing through the lean zone 803, the combustion products leave the combustion chamber 16. The secondary orifices 805 can be called dilution orifices, and can be arranged to gradually introduce dilution air into the lean zone 803. The fuel added by the fuel spray nozzle is almost completely burned as the air exits at an outlet of the combustion chamber, before flowing to the turbine.

There shows part of a propulsion system 900 for an aircraft. The propulsion system 900 includes the gas turbine engine 10 of the . The engine 10 further includes a fuel system and an oil system. The fuel system includes: a low pressure fuel pump 902, a fuel-oil heat exchanger 903, a main (or high pressure) fuel pump 904, a controller 908, and a fuel distribution valve 909. The propulsion system 900 further includes a fuel tank 901. The oil system includes an oil tank 905, an oil feed pump 906, and a main oil pump 907. In the present example, the low pressure fuel pump 902 is shown as part of the gas turbine engine 10. In other examples, the low pressure fuel pump, or additional fuel pumps, may be provided as part of the fuel system to on board the aircraft to which the gas turbine engine is mounted.

The low pressure fuel pump 902 is arranged to distribute fuel from the fuel tank 901 to the fuel-oil heat exchanger 903 through a suitable arrangement of pipes, conduits, etc. (not shown). The main fuel pump 904 is configured to distribute fuel from the fuel-oil heat exchanger 903 to the fuel spray nozzles of the combustion chamber 16 via the fuel distribution valve 909 and of a suitable arrangement of pipes, conduits, etc. (not shown). The fuel distribution valve 909 is arranged to distribute the fuel between a main manifold 909a and a primary manifold 909b. The main manifold is fluidly connected to the main injectors of each of the fuel spray nozzles 404, 403 as shown in the . It therefore supplies fuel to all duplex fuel spray nozzles 403 and single flow 404. The primary manifold 909b is fluidly connected to the primary injectors of each of the duplex fuel spray nozzles 403. The primary manifold 909b can for this reason be used to provide a greater flow of fuel to the first subset of fuel spray nozzles (e.g. duplex fuel spray nozzles 403 in the present example) compared to the flow of fuel supplied to the second flow spray nozzle subassembly via main manifold 909a. For example, below a threshold engine power, fuel may be fed only (or at a greater fuel flow rate) to the first subset of fuel spray nozzles via the primary manifold 909b by comparison with the second subset of fuel spray nozzles.

The oil feed pump 906 is arranged to distribute lubricating oil from the oil reservoir 905 to the fuel-oil heat exchanger 903 via a suitable arrangement of pipes, conduits, etc. (not shown). The main oil pump 907 is arranged to distribute oil from the fuel-oil heat exchanger 903 to components of the engine 10 as required via a suitable oil distribution arrangement ( not shown). The fuel flow path from the fuel tank 901 to the combustion chamber 16, through the pumps 902, 904 and the fuel-oil heat exchanger 903, in operation is illustrated in the by dotted arrows. The oil flow path from the oil tank 905 to the fuel-oil heat exchanger 903, through the oil feed 906, and onto the engine components 10 in operation is shown on the by solid arrows.

The controller 908 includes an appropriate arrangement of processors and electronic memory. The controller 908 is in communication with the fuel-oil heat exchanger 903, as illustrated by the dotted and dashed line on the , and is configured to control the operation of the fuel-oil heat exchanger 903. The controller 908 is configured to control the operation of the fuel-oil heat exchanger 903, by providing control signals to the fuel-oil heat exchanger 903. Controller 908 is configured to control operation of fuel-oil heat exchanger 903 to adjust at least one property or parameter of the fuel upon entry into the combustion chamber 16. In the example shown, the controller 908 is configured to control the operation of the fuel-oil heat exchanger 903 to control a viscosity of the fuel upon entry into the combustion chamber 16. In other examples, the controller 908 may additionally or alternatively be configured to control the operation of the fuel-oil heat exchanger 903 to control a temperature of the fuel upon entry into the chamber. combustion. Controller 908 may be a separate controller as illustrated, or may be part of an electronic engine controller (EEC) arranged to control other engine functions.

In the example shown, the fuel-oil heat exchanger 903 is disposed between the low pressure fuel pump 902 and the main fuel pump 904, although the fuel-oil heat exchanger 903 may be disposed neither regardless of any appropriate location or position relative to other components of the propulsion system 900. In other examples, the propulsion system 900 may include one or more additional heat exchangers arranged to receive oil from the oil system , or the propulsion system 900 may include one or more additional oil systems arranged to supply oil to the additional heat exchanger(s). We will have in mind that the propulsion system 900 as shown on the is only a schematic view of a propulsion system serving as an illustration.

In one example, the controller 908 is configured to control the operation of the fuel-oil heat exchanger 903 to lower the fuel viscosity to 0.58 mm ² /s or less upon entering the chamber. combustion 16 at cruising conditions. Controller 908 may alternatively be configured to control operation of fuel-oil heat exchanger 903 to lower fuel viscosity to 0.48 mm ² /s or less upon entry into the fuel chamber. combustion at cruise conditions.

Alternatively, controller 908 may be configured to control operation of fuel-oil heat exchanger 903 to lower fuel viscosity between 0.58 mm ² /s and 0.30 mm ² /s, e.g. example 0.58, 0.57, 0.56, 0.55, 0.54, 0.53, 0.52, 0.51, 0.50, 0.49, 0.48, 0.47, 0 .46, 0.45, 0.44, 0.43, 0.42, 0.41, 0.40, 0.39, 0.38, 0.37, 0.36, 0.35, 0.34 , 0.33, 0.32, 0.31 or 0.30 mm ² /s upon entry into the combustion chamber 16 at cruising conditions. Controller 908 may be configured to control operation of fuel-oil heat exchanger 903 to lower fuel viscosity between 0.54 mm ² /s and 0.34 mm ² /s, 0.50 mm ² /s and 0.38 mm ² /s, 0.50 mm ² /s and 0.35 mm ² /s, or preferably between 0.48 mm ² /s and 0.40 mm ² /s, or between 0.44 mm ² /s and 0.42 mm ² /s when entering the combustion chamber at cruise conditions. Alternatively, controller 908 may be configured to control operation of fuel-oil heat exchanger 903 to lower fuel viscosity to 0.57, 0.56, 0.55, 0.54, 0 .53, 0.52, 0.51, 0.50, 0.49, 0.48, 0.47, 0.46, 0.45, 0.44, 0.43, 0.42, 0.41 , 0.40, 0.39, 0.38, 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31 or 0.30 mm ² /s or less upon entering the combustion chamber 16 at cruising conditions. The controller 908 may be configured to control the operation of the fuel-oil heat exchanger 903 to lower the fuel viscosity so that it is within a defined range between any two of the values in the preceding sentence.

The fuel viscosities described above may be appropriate for operating the gas turbine engine 10 with sustainable aviation fuel. The fuel circuit can for this reason be arranged to supply the combustion chamber 16 with a sustainable aeronautical fuel, for example by storing a sustainable aeronautical fuel within the fuel tank 901. The sustainable aeronautical fuel (SAF) can be such that defined anywhere in this document. According to any aspect of the present disclosure, the fuel supplied to the plurality of fuel spray nozzles may comprise a sustainable aviation fuel, for example having a SAF % in the range between 10% and 50%, or between 50% and 100%, or including 100% SAF. The present application is not, however, limited to operation using sustainable aviation fuel, and any of the parameters described herein may be appropriate for use with other fuels.

Sustainable aviation fuel can have a calorific value of at least 43.5 MJ/kg, for example between 43.5 MJ/kg and 44 MJ/kg. The sustainable aviation fuel may alternatively have a calorific value of at least 43.8 MJ/kg, for example between 43.8 MJ/kg and 44 MJ/kg. In some examples, the sustainable aviation fuel may have a calorific value of 43.5, 43.6, 43.7, 43.8, 43.9 or 44.0 MJ/kg, or within a defined range between any two of the values in this sentence.

Sustainable aviation fuel may have an aromatic content of 10% or less by volume. In some examples, the sustainable aviation fuel may have an aromatic content of 5% or less by volume, or more preferably 1% or less by volume. In some examples, the sustainable aviation fuel may have an aromatic content by volume of 1, 2, 3, 4, 5, 6, 7, 8 or 9% or less, or may be within a defined range between any two of the values in this sentence.

Although the fuel properties given in the previous two paragraphs (and elsewhere in this document) relate to sustainable aviation fuel, other types of fuel with the same properties can be supplied to the combustion chamber by the fuel system. fuel.

The controller 908 is also configured to control the distribution of fuel to the fuel spray nozzles 403 of the combustion chamber 16 to direct a flow of fuel to a subset of the fuel spray nozzles 403 adjacent to the igniter or igniters 405 during a relighting procedure. In the example shown, the controller 908 is configured to direct a flow of fuel to the groups of duplex fuel spray nozzles 403 adjacent to the igniters 405, by distributing the fuel entirely or mainly to the primary injectors of these nozzles fuel spray 403 during a relighting procedure. Controller 908 may alternatively be configured to direct fuel flow to any suitable fuel spray nozzle subset 403 (or fuel injector subset of the fuel spray nozzles 403) during a relighting procedure, to increase the probability of successful ignition. The fuel spray nozzle subassembly 403 to which fuel flow is directed during a relighting procedure may be chosen to take into account one or more properties of the fuel, e.g. calorific value, compound content aromatics or fuel viscosity (e.g., fuel viscosity upon entering the combustion chamber at cruise conditions), to increase a probability of successful ignition.

In the present example, the controller 908 is configured to direct a flow of fuel towards the nozzles such that the first subset of the plurality of fuel spray nozzles described above receives more fuel than the second subset. The controller 908 is configured to control the fuel distribution valve 909 such that below a staging point fuel is distributed only to the primary fuel injectors of the duplex fuel spray nozzles 403. Above a staging point, controller 908 is configured to control fuel distribution valve 909 such that fuel is further distributed to the main fuel injectors of the duplex fuel spray nozzles 403 and single-flow fuel spray nozzles 404. The combustion chamber 16 is therefore operated as a staged combustion chamber. As such, duplex fuel spray nozzles 403 receive more fuel than single flow fuel spray nozzles 404 (below and optionally above the staging point). The fuel delivery system shown in the Figures should be considered as an example of how fuel is directed to the spray nozzles, others being possible. For example, two sets of independent single flow nozzles may be provided. In some examples, fuel flow may be requested at other engine operating conditions such as during starting or at low power operation such as during idling.

There illustrates a method 1000 of operating a gas turbine engine 10 as described above. The method 1000 includes supplying 1001 of fuel to the combustion chamber 16 of the gas turbine engine. The fuel may be or may include sustainable aviation fuel. In the example shown, the fuel has a calorific value of at least 43.5 MJ/kg. In other examples, the fuel may have a calorific value between 43.5 MJ/kg and 44 MJ/kg. In still other examples, the fuel may have a calorific value of at least 43.8 MJ/kg. In yet other examples, the fuel may have a calorific value between 43.8 MJ/kg and 44 MJ/kg. In some examples, the fuel may have a calorific value of 43.5, 43.6, 43.7, 43.8, 43.9 or 44.0 MJ/kg, or within a defined range between any two values of this sentence.

There illustrates a method 1100 of operating a gas turbine engine 10 as described above. The method 1100 includes supplying 1101 of fuel to the combustion chamber 16 of the gas turbine engine. The fuel may be or may include sustainable aviation fuel. In the example shown, the fuel has an aromatic content of 10% or less by volume. In some examples, the fuel may have an aromatic content of 5% or less by volume, or more preferably 1% or less by volume. In some examples, the fuel may have an aromatic compound content by volume of 1, 2, 3, 4, 5, 6, 7, 8 or 9% or less, or may be within a defined range between any two values of this sentence.

There illustrates a method 1200 of operating a gas turbine engine 10 as described above. The method 1200 includes transferring 1201 heat from the oil to the fuel before the fuel enters the combustion chamber 16 so as to lower a viscosity of the fuel to 0.58 mm ² /s or less when entering the combustion chamber 16 at cruise conditions, or to lower the fuel viscosity to 0.48 mm ² /s or less when entering the combustion chamber at cruise conditions. The method 1200 includes providing 1202 fuel to the combustion chamber 16 of the gas turbine engine 10. In the example shown, the method 1200 includes transferring heat 1201 from the oil to the fuel by controlling the operation of a 903 fuel-oil heat exchanger.

The method 1200 may include transferring heat 1201 from the oil to the fuel so as to lower a viscosity of the fuel between 0.58 mm ² /s and 0.30 mm ² /s upon entry into the fuel chamber. combustion 16 at cruising conditions. The method 1200 may include transferring heat 1201 from the oil to the fuel so as to lower a fuel viscosity between 0.54 mm ² /s and 0.34 mm ² /s, 0.50 mm ² /s and 0.38 mm ² /s, 0.50 mm ² /s and 0.35 mm ² /s, or preferably between 0.48 mm ² /s and 0.40 mm ² /s, or between 0.44 mm ² /s and 0.42 mm ² /s when entering the combustion chamber at cruise conditions.

The method 1200 may include transferring heat 1201 from the oil to the fuel so as to lower a viscosity of the fuel to 0.58, 0.57, 0.56, 0.55, 0.54, 0.53, 0.52, 0.51, 0.50, 0.49, 0.48, 0.47, 0.46, 0.45, 0.44, 0.43, 0.42, 0.41, 0, 40, 0.39, 0.38, 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31 or 0.30 mm ² /s or less when entry into the combustion chamber 16 at cruising conditions. The method 1200 may include transferring heat 1201 from the oil to the fuel so as to lower a viscosity of the fuel so that it is within a defined range between any two of the values in the preceding sentence.

The method 1200 may include transferring heat 1201 from the oil to the fuel so as to lower a fuel viscosity between 0.55 mm ² /s and 0.35 mm ² /s, 0.50 mm ² /s and 0.35 mm ² /s, 0.48 mm ² /s and 0.35 mm ² /s, 0.48 mm ² /s and 0.38 mm ² /s, 0.48 mm ² /s and 0, 40 mm ² /s, 0.46 mm ² /s and 0.40 mm ² /s, 0.44 mm ² /s and 0.40 mm ² /s, or 0.44 mm ² /s and 0.42 mm ² /s upon entry into the combustion chamber 16 at cruising conditions.

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

Claims (15)

1 Method (1100) of operating a gas turbine engine (10), the gas turbine engine (10) comprising a staged combustion chamber (16) comprising an arrangement of fuel spray nozzles (403, 404 ) wherein a flow of fuel is directed to a subset of the nozzles (403, 404) adjacent one or more igniters (405) during a relighting procedure, wherein the method (1100) comprises providing fuel (1101 ) to the combustion chamber (16) having an aromatic compound content of 10% or less by volume. 2 Method (1100) according to claim 1, or any one of the other preceding claims, wherein the number of fuel spray nozzles (403, 404) is between 14 and 22 and/or a number of spray nozzles of fuel per unit engine core size in the range 2 to 6. 3 Method (1100) according to claim 1, or any one of the other preceding claims, wherein the subset of fuel spray nozzles (403, 404) comprises at least half of the total number of fuel spray nozzles. fuel (403, 404). 4 Method (1100) according to claim 3, wherein the subset of fuel spray nozzles (403, 404) comprises at least two thirds of the total number of fuel spray nozzles (403, 404). A method (1100) according to claim 1, or any of the other preceding claims, wherein the arrangement of fuel spray nozzles (403, 404) comprises duplex nozzles (403) and single flow nozzles ( 404). 6 The method (1100) of claim 5, wherein the fuel spray nozzle subset (403, 404) comprises the duplex nozzles (403) and the remaining fuel spray nozzles (403, 404) comprise the nozzles single flow (404). 7 Method (1100) according to claim 1, or any one of the other preceding claims, wherein the combustion chamber (16) comprises at least two igniters (405) and the sub-assembly of fuel spray nozzles (403 , 404) comprises at least two groups of nozzles (403, 404), each group of nozzles (403, 404) being adjacent to one of the igniters (405). 8 Method (1100) according to claim 1 or any one of the other preceding claims, in which:
a) the fuel supplied to the combustion chamber (16) has a calorific value of at least 43.5 MJ/kg, preferably between 43.5 MJ/kg and 44 MJ/kg, or at least 43, 8 MJ/kg and preferably between 43.8 MJ/kg and 44 MJ/kg; and or
b) the gas turbine engine (10) includes a fuel-oil heat exchanger (903), and the method (1100) includes transferring heat (1201) from the oil to the fuel before the fuel enters the combustion chamber (16) so as to lower the fuel viscosity to 0.58 mm ² /s or less upon entering the combustion chamber (16) at cruise conditions, preferably 0.48 mm ² /s or less, preferably between 0.40 mm ² /s and 0.48 mm ² /s and preferably between 0.42 mm ² /s and 0.44 mm ² /s. 9 Gas turbine engine (10) for an aircraft, comprising:
a staged combustion chamber (16) comprising an arrangement of fuel spray nozzles (403, 404) and one or more igniters (405); And
a controller (908) configured to direct a flow of fuel to a subset of the fuel spray nozzles (403, 404) adjacent the igniter or igniters (405) during a relighting procedure, the fuel having an aromatic compound content of 10% or less by volume. 10 Gas turbine engine (10) according to claim 9, wherein the number of fuel spray nozzles (403, 404) is between 14 and 22 and/or a number of fuel spray nozzles per unit size of engine core in the range of 2 to 6. 11 gas turbine engine (10) according to any one of claims 9 to 10 wherein the subset of fuel spray nozzles (403, 404) comprises at least half of the total number of fuel spray nozzles (403, 404), and optionally
wherein the fuel spray nozzle subassembly (403, 404) comprises at least two thirds of the total number of fuel spray nozzles (403, 404). A gas turbine engine (10) according to any one of claims 9 to 11, wherein the arrangement of fuel spray nozzles (403, 404) comprises duplex nozzles (403) and single flow nozzles ( 404), and optionally
wherein the fuel spray nozzle subset (403, 404) comprises the duplex nozzles (403) and the remaining fuel spray nozzles (403, 404) comprise the single flow nozzles (404). 13 Gas turbine engine (10) according to any one of claims 9 to 12, wherein the combustion chamber (16) comprises at least two igniters (405) and the fuel spray nozzle sub-assembly (403 , 404) comprises at least two groups of nozzles (403, 404), each group of nozzles (403, 404) being adjacent to one of the igniters (405). 14 gas turbine engine (10) according to claim 13, wherein the gas turbine engine (10) comprises a plurality of igniters (405) arranged symmetrically around a circumference of the combustion chamber (16), and optionally wherein pairs of igniters (405) are arranged diametrically opposed to each other around a circumference of the combustion chamber (16). 15 gas turbine engine (10) according to any one of claims 9 to 14, in which:
a) the fuel supplied to the combustion chamber (16) has a calorific value of at least 43.5 MJ/kg, preferably between 43.5 MJ/kg and 44 MJ/kg, or at least 43, 8 MJ/kg and preferably between 43.8 MJ/kg and 44 MJ/kg; and or
b) the gas turbine engine (10) comprises a fuel-oil heat exchanger (903), and the controller (908) is arranged to transfer heat from the oil to the fuel before the fuel enters the combustion chamber (16) so as to lower the fuel viscosity to 0.58 mm ² /s or less upon entering the combustion chamber (16) at cruise conditions, preferably 0 .48 mm ² /s or less, preferably between 0.40 mm ² /s and 0.48 mm ² /s and preferably between 0.42 mm ² /s and 0.44 mm ² /s.