Emissions of non-volatile particulate matter from gas turbine engines combusting sustainable aviation fuel and fossil-based hydrocarbon fuel

Abstract

A gas turbine engine includes: a rich burn, quick quench, lean burn combustor having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit engine core size in the range 2 to 6. An MTO nvPM emissions index ratio is defined as: $\frac{{EI}_{\max TO, SAF}}{{EI}_{\max TO, FF}}$
where: EI.sub.maxTO,SAF and EI.sub.maxTO,FF are respectively the nvPM emissions index in mg/kg of the gas turbine engine when operating at around 100% available thrust for the given operating conditions if a fuel provided to the fuel spray nozzles includes a sustainable aviation fuel (SAF) or is a fossil-based hydrocarbon fuel. The MTO nvPM emissions index ratio of the gas turbine engine is less than 1. The gas turbine engine is configured to provide fuel including a SAF to the fuel spray nozzles. Also disclosed is a method of operating a gas turbine engine.

Claims

1. A gas turbine engine for an aircraft, comprising: a rich burn, quick quench, lean burn (RQL) combustor having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit engine core size in the range 2 to 6 where engine core size is defined, at a top of climb operating condition, as: $Core size = {\dot{m}}_{2} \frac{\sqrt{T_{3}}}{P_{3}},$ where {dot over (m)}.sub.2=mass flow rate, in lbs per second, of air on entry to a high-pressure compressor, T.sub.3=temperature, in Kelvin, of air on exit from the high-pressure compressor, and P.sub.3=pressure, in lb inches per second squared per inch squared, of air on exit from the high-pressure compressor such that a unit of core size is expressed as: $s .Math. K^{\frac{1}{2}} .Math. in;$ and wherein: an MTO nvPM emissions index ratio is defined as: $\frac{{EI}_{maxTO, SAF}}{{EI}_{maxTO, FF}}$ where: EI.sub.maxTO,SAF is the nvPM emissions index in mg/kg of the gas turbine engine when operating at around 100% available thrust for given operating conditions if a fuel provided to the fuel spray nozzles comprises a sustainable aviation fuel (SAF); and EI.sub.maxTO,FF is the nvPM emissions index in mg/kg of the gas turbine engine when operating at around 100% available thrust for the given operating conditions if a fuel provided to the fuel spray nozzles is a fossil-based hydrocarbon fuel; the MTO nvPM emissions index ratio of the gas turbine engine is less than 1; and the gas turbine engine is configured to provide fuel comprising a SAF to the fuel spray nozzles, wherein the fuel spray nozzles comprises one or more duplex nozzles and one or more single flow nozzles; and the combustor comprises one or more ignitors each arranged adjacent to one or more of the duplex nozzles, and/or wherein the number of fuel spray nozzles per unit engine core size is in the range 2.5 to 4.5.

2. The gas turbine engine of claim 1, wherein the MTO nvPM emissions index ratio is less than or equal to 0.93.

3. The gas turbine engine of claim 1, wherein the MTO nvPM emissions index ratio is greater than or equal to 0.15.

4. The gas turbine engine of claim 1, wherein a climb nvPM emissions index ratio is defined as: $\frac{{EI}_{climb, SAF}}{{EI}_{climb, FF}}$ where: EI.sub.climb,SAF is the nvPM emissions index in mg/kg of the gas turbine engine when operating at around 85% available thrust for the given operating conditions, or for different given operating conditions, if a fuel provided to the fuel spray nozzles comprises a sustainable aviation fuel (SAF); and EI.sub.climb,FF is the nvPM emissions index in mg/kg of the gas turbine engine when operating at around 85% available thrust for the same given operating conditions at which EI.sub.climb,SAF is calculated and if a fuel provided to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and wherein the climb nvPM emissions index ratio of the gas turbine engine is less than 1.

5. The gas turbine engine of claim 4, wherein the climb nvPM emissions index ratio is less than or equal to 0.9.

6. The gas turbine engine of claim 4, wherein the climb nvPM emissions index ratio is greater than or equal to 0.1.

7. The gas turbine engine of claim 1, wherein an approach nvPM emissions index ratio is defined as: $\frac{{EI}_{approach, SAF}}{{EI}_{approach, FF}}$ where: EI.sub.approach,SAF is the nvPM emissions index in mg/kg of the gas turbine engine when operating at around 30% available thrust for the given operating conditions, or for different given operating conditions, if a fuel provided to the fuel spray nozzles comprises a sustainable aviation fuel (SAF); and EI.sub.approach,FF is the nvPM emissions index in mg/kg of the gas turbine engine when operating at around 30% available thrust for the same given operating conditions at which EI.sub.approach,SAF is calculated and if a fuel provided to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and wherein the approach nvPM emissions index ratio of the gas turbine engine is less than 1.

8. The gas turbine engine of claim 7, wherein the approach nvPM emissions index ratio is less than or equal to 0.8.

9. The gas turbine engine of claim 7, wherein the approach nvPM emissions index ratio is greater than or equal to 0.03.

10. The gas turbine engine of claim 1, wherein an idle nvPM emissions index ratio is defined as: $\frac{{EI}_{idle, SAF}}{{EI}_{idle, FF}}$ where: EI.sub.idle,SAF is the nvPM emissions index in mg/kg of the gas turbine engine when operating at around 7% available thrust for the given operating conditions, or for different given operating conditions, and if a fuel provided to the fuel spray nozzles comprises a sustainable aviation fuel (SAF); and EI.sub.idle,FF is the nvPM emissions index in mg/kg of the gas turbine engine when operating at around 7% available thrust for the same given operating conditions at which EI.sub.idle,SAF is calculated and if a fuel provided to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and wherein the idle nvPM emissions index ratio of the gas turbine engine is less than 1.

11. The gas turbine engine of claim 10, wherein the idle nvPM emissions index ratio is less than or equal to 0.8.

12. The gas turbine engine of claim 10, wherein the idle nvPM emissions index ratio is greater than or equal to 0.02.

13. The gas turbine engine of claim 1, wherein the fuel provided to the fuel spray nozzles comprises a % SAF in the range of 50% to 100%.

14. A method of operating the gas turbine engine of claim 1, the method comprising providing fuel comprising a sustainable aviation fuel to the fuel spray nozzles.

15. The gas turbine engine of claim 1, wherein the top of climb occurs at 30,000 to 39,000 feet, a forward speed of Mach Number 0.75 to 0.85, and ambient air temperature of ISA+10K to ISA+15K.

16. The gas turbine engine of claim 15, wherein the top of climb occurs at 35,000 feet.

17. A method of operating a gas turbine engine, the gas turbine engine comprising: a rich burn, quick quench, lean burn (RQL) combustor having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit engine core size in the range 2 to 6 where engine core size is defined, at a top of climb operating condition, as: $Core size = {\dot{m}}_{2} \frac{\sqrt{T_{3}}}{P_{3}},$ where {dot over (m)}.sub.2=mass flow rate, in lbs per second, of air on entry to a high-pressure compressor, T.sub.3=temperature, in Kelvin, of air on exit from the high-pressure compressor, and P.sub.3=pressure, in lb inches per second squared per inch squared, of air on exit from the high-pressure compressor such that a unit of core size is expressed as: $s .Math. K^{\frac{1}{2}} .Math. in;$ and wherein an MTO nvPM emissions index ratio is defined as: $\frac{{EI}_{maxTO, SAF}}{{EI}_{maxTO, FF}}$ where: EI.sub.maxTO,SAF is the nvPM emissions index in mg/kg of the gas turbine engine when operating at around 100% available thrust for given operating conditions if a fuel provided to the fuel spray nozzles comprises a sustainable aviation fuel (SAF); and EI.sub.maxTO,FF is the nvPM emissions index in mg/kg of the gas turbine engine when operating at around 100% available thrust for the given operating conditions if a fuel provided to the fuel spray nozzles is a fossil-based hydrocarbon fuel; the MTO nvPM emissions index ratio of the gas turbine engine is less than 1; the gas turbine engine is configured to provide fuel comprising a SAF to the fuel spray nozzles; the fuel spray nozzles comprises one or more duplex nozzles and one or more single flow nozzles; and the combustor comprises one or more ignitors each arranged adjacent to one or more of the duplex nozzles, and/or wherein the number of fuel spray nozzles per unit engine core size is in the range 2.5 to 4.5; wherein the method comprises providing fuel comprising a sustainable aviation fuel to the fuel spray nozzles.

18. The method of claim 17, wherein a climb nvPM emissions index ratio is defined as: $\frac{{EI}_{climb, SAF}}{{EI}_{climb, FF}}$ where: EI.sub.climb,SAF is the nvPM emissions index in mg/kg of the gas turbine engine when operating at around 85% available thrust for the given operating conditions, or for different given operating conditions, if a fuel provided to the fuel spray nozzles comprises a sustainable aviation fuel (SAF); and EI.sub.climb,FF is the nvPM emissions index in mg/kg of the gas turbine engine when operating at around 85% available thrust for the same given operating conditions at which EI.sub.climb,SAF is calculated and if a fuel provided to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and wherein the climb nvPM emissions index ratio of the gas turbine engine is less than 1.

19. The method of claim 17, wherein an approach nvPM emissions index ratio is defined as: $\frac{{EI}_{approach, SAF}}{{EI}_{approach, FF}}$ where: EI.sub.approach,SAF is the nvPM emissions index in mg/kg of the gas turbine engine when operating at around 30% available thrust for the given operating conditions, or for different given operating conditions, if a fuel provided to the fuel spray nozzles comprises a sustainable aviation fuel (SAF); and EI.sub.approach,FF is the nvPM emissions index in mg/kg of the gas turbine engine when operating at around 30% available thrust for the same given operating conditions at which EI.sub.approach,SAF is calculated and if a fuel provided to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and wherein the approach nvPM emissions index ratio of the gas turbine engine is less than 1.

20. The method of claim 17, wherein an idle nvPM emissions index ratio is defined as: $\frac{{EI}_{idle, SAF}}{{EI}_{idle, FF}}$ where: EI.sub.idle,SAF is the nvPM emissions index in mg/kg of the gas turbine engine when operating at around 7% available thrust for the given operating conditions, or for different given operating conditions, and if a fuel provided to the fuel spray nozzles comprises a sustainable aviation fuel (SAF); and EI.sub.idle,FF is the nvPM emissions index in mg/kg of the gas turbine engine when operating at around 7% available thrust for the same given operating conditions at which EI.sub.idle,SAF is calculated and if a fuel provided to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and wherein the idle nvPM emissions index ratio of the gas turbine engine is less than 1.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

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

(2) FIG. 1 is a sectional side view of a gas turbine engine;

(3) FIG. 2 is a close up sectional side view of an upstream portion of a geared gas turbine engine;

(4) FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine;

(5) FIG. 4 is a close up sectional side view of a direct drive gas turbine engine;

(6) FIG. 5 is a schematic view of an aircraft having two gas turbine engines of the present application mounted thereon;

(7) FIG. 6 is a sectional view through a combustor of the engine of FIG. 1 in a plane normal to the principal rotational axis of the engine;

(8) FIG. 7 is a schematic cross-section of a duplex fuel spray nozzle of the combustor of FIG. 6;

(9) FIG. 8 is a schematic cross-section of a single flow fuel spray nozzle of the combustor of FIG. 6;

(10) FIG. 9 is a partial sectional view of the engine of FIG. 1;

(11) FIG. 10 is a further partial sectional view of the engine of FIG. 1;

(12) FIG. 11 is a schematic representation of a propulsion system for an aircraft comprising the engine of FIG. 1; and

(13) FIG. 12 shows a method of operating the gas turbine engine.

DETAILED DESCRIPTION OF THE DISCLOSURE

(14) FIG. 1 illustrates a gas turbine engine 10 having a principal rotational axis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23 that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 comprises a core 11 that receives the core airflow A. The engine core 11 comprises, in axial flow series, a low pressure compressor 14, a high pressure compressor 15, combustion equipment 16, a high pressure turbine 17, a low pressure turbine 19 and a core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass duct 22. The fan 23 is attached to and driven by the low pressure turbine 19 via a first, low pressure shaft 26 and an epicyclic gearbox 30.

(15) In use, the core airflow 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 exhausted from the high pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel F and the mixture is combusted. The combustion equipment 16 may be referred to as the combustor 16, with the terms combustion equipment 16 and combustor 16 used interchangeably herein. The resultant hot combustion products then expand through, and thereby drive, the high pressure and low pressure turbines 17, 19 before being exhausted through the core exhaust nozzle 20 to provide some propulsive thrust. The high pressure turbine 17 drives the high pressure compressor 15 by a suitable interconnecting second, high pressure shaft 27. The fan 23 generally acts to impart increased pressure to the bypass airflow 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 propulsive thrust. The epicyclic gearbox 30 is a reduction gearbox.

(16) An exemplary arrangement for a geared fan gas turbine engine 10 is shown in FIG. 2. The low pressure turbine 19 (see FIG. 1) drives the low pressure shaft 26, which is coupled to a sun wheel, or sun gear, 28 of the epicyclic gearbox 30. The low pressure shaft 26 may be referred to as an input shaft for the epicyclic gearbox 30. Radially outward of the sun gear 28 and intermeshing therewith is a plurality of planet gears 32 that are coupled together by a planet carrier 34. The planet carrier 34 constrains the planet gears 32 to precess around the sun gear 28 in synchronicity whilst enabling each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled via linkages 36 to the fan 23 through an output fan shaft 42 in order to drive the fan 23 in rotation about the engine axis 9. Radially outward of the planet gears 32 and intermeshing therewith is an annulus or ring gear 38 that is coupled, via linkages 40, to a stationary supporting structure 24.

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

(18) The epicyclic gearbox 30 is shown by way of example in greater detail in FIG. 3. Each of the sun gear 28, planet gears 32 and ring gear 38 comprise teeth about their periphery to intermesh with the other gears. However, for clarity only exemplary portions of the teeth are illustrated in FIG. 3. There are four planet gears 32 illustrated, although it will be apparent to the skilled reader that more or fewer planet gears 32 may be provided within the scope of the claimed invention. Practical applications of a planetary epicyclic gearbox 30 generally comprise at least three planet gears 32, for example five planet gears 32.

(19) The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3 is of the planetary type, in that the planet carrier 34 is coupled to the output shaft via linkages 36, with the ring gear 38 fixed. However, any other suitable type of epicyclic gearbox 30 may be used. By way of further example, the epicyclic gearbox 30 may be a star arrangement, in which the planet carrier 34 is held fixed, with the ring (or annulus) gear 38 allowed to rotate. In such an arrangement the fan 23 is driven by the ring gear 38. By way of further alternative example, the gearbox 30 may be a differential gearbox in which the ring gear 38 and the planet carrier 34 are both allowed to rotate.

(20) It will be appreciated that the arrangement shown in FIG. 2 and FIG. 3 is by way of example only, and various alternatives are within the scope of the present disclosure. Purely by way of example, any suitable arrangement may be used for locating the gearbox 30 in the engine 10 and/or for connecting the gearbox 30 to the engine 10. By way of further example, the connections (such as the linkages 36, 40 in the FIG. 2 example) between the gearbox 30 and other parts of the engine 10 (such as the input shaft 26, the output shaft and the stationary supporting structure 24) may have any desired degree of stiffness or flexibility. By way of further example, any suitable arrangement of the bearings between rotating and stationary parts of the engine (for example between the input and output shafts from the gearbox and the fixed structures, such as the gearbox casing) may be used, and the disclosure is not limited to the exemplary arrangement of FIG. 2. For example, where the gearbox 30 has a star arrangement (described above), the skilled person would readily understand that the arrangement of output and support linkages and bearing locations would typically be different to that shown by way of example in FIG. 2.

(21) Accordingly, the present disclosure extends to a gas turbine engine having any arrangement of gearbox styles (for example star or planetary), support structures, input and output shaft arrangement, and bearing locations.

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

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

(24) By way of further example, other gas turbine engines to which the present disclosure may be applied may have no gearbox for the main shaft(s), instead being direct drive engines. A cross-sectional view of one such engine is shown in FIG. 4.

(25) With reference to FIG. 4, a gas turbine engine is generally indicated at 10, having a principal rotational axis 9. The engine 10 comprises, in axial flow series, an air intake 12, a propulsive fan 23, an intermediate pressure compressor 14, a high pressure compressor 15, combustion equipment 16, a high pressure turbine 17, an intermediate pressure turbine 19a, a low-pressure turbine 19 and an exhaust nozzle 20. A nacelle 21 surrounds the engine 10 and defines both the intake 12 and the exhaust nozzle 20.

(26) In use, air entering the intake 12 is accelerated by the fan 23 to produce two air flows: a core airflow A and a bypass airflow B. The core airflow A flows into the intermediate pressure compressor 14, and the bypass air flow B passes through a bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 14 compresses the airflow A before delivering that air to the high pressure compressor 15 where further compression takes place.

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

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

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

(30) The geometry of the gas turbine engine 10, and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned with the principal rotational axis 9), a radial direction (in the bottom-to-top direction in FIG. 1), and a circumferential direction (perpendicular to the page in the FIG. 1 view). The axial, radial and circumferential directions are mutually perpendicular.

(31) FIG. 5 shows an aircraft 1 on which two gas turbine engines 10 of the present disclosure are mounted, one on each wing. The aircraft 1 comprises a fuel system 2 which comprises a fuselage fuel tank 50a and two wing fuel tanks 50b. Fuel F is provided from the fuel system to the gas turbine engines. The fuel tanks 50a, 50b are supplied with fuel from a fuel input port 62. Other fuel systems may be used with other layouts of fuel tank.

(32) The fuel F provided to the combustion equipment 16 may comprise a fossil-based hydrocarbon fuel, such as Kerosene. Thus, the fuel F may comprise molecules from one or more of the chemical families of n-alkanes, iso-alkanes, cycloalkanes, and aromatics. As there is an expectation in the aviation industry of a trend towards the use of fuels different from the traditional kerosene-based jet fuels generally used at present, when blended with, mixed with, or replaced by an alternative fuel, the fuel F may comprise renewable hydrocarbons produced from biological or non-biological resources, otherwise known as sustainable aviation fuel (SAF). In each of the provided examples, the fuel F may comprise one or more trace elements including, for example, sulphur, nitrogen, oxygen, inorganics, and metals.

(33) SAF is understood by the Skilled Person to refer to, for example, a biofuel, renewable aviation fuel, renewable jet fuel, alternative fuel or biojet fuel, produced from biological or non-biological resources. SAF is understood to be commonly synthesised from carbon-containing gasses drawn out of the atmosphere and/or captured from industrial processes; or from a wide range of sustainable feedstocks such as, for example, waste oil and fats; municipal solid waste; cellulosic waste (such as corn stalks); cover crops such as camelina, carinata, and pennycress; non-biogenic alternative fuels; jatropha; halophytes and algae, rather than from fossil-based hydrocarbons. SAF is understood as not encompassing fossil fuels.

(34) Functional performance of a given fuel composition, or blend of fuel F for use in a given mission, may be defined, at least in part, by the ability of the fuel to service 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 gaseous and/or particulate matter. In this regard, particulate matter emissions may include soot particles created by the combustion of said fuel F, also known as non-volatile particulate matter (nvPM). Any reference herein to soot or smoke may apply equally to other types of particulate matter emissions known within the art. Gaseous emissions may include any one or more of nitrogen oxides (NOX); carbon monoxide (CO); carbon dioxide (CO2); unburned hydrocarbons (UHC); sulphur oxides (SOx) including, for example, sulphur dioxide (SO2) and/or sulphur trioxide (SO3); and, volatile organic compounds (VOC) created by the combustion of said fuel F. Any reference herein to gaseous emissions may apply equally to other types of gaseous emissions known within the art.

(35) A relatively higher specific energy (i.e. energy per unit mass), expressed as MJ/kg, may at least partially reduce take-off weight, thus potentially providing a relative improvement in fuel efficiency. A relatively higher energy density (i.e. energy per unit volume), expressed as MJ/L, may at least partially reduce take-off fuel volume, which may be particularly important for volume-limited missions or military operations involving refuelling. A relatively higher thermal stability (i.e. inhibition of fuel to degrade or coke under thermal stress) may permit the fuel to sustain elevated temperatures in the engine and fuel injectors, thus potentially providing relative improvements in combustion efficiency. Reduced emissions, including particulate matter, may permit reduced contrail formation, whilst reducing the environmental impact of a given mission. Other properties of the fuel may also be key to functional performance. For example, a relatively lower freeze point (C) may allow long-range missions to optimise flight profiles; minimum aromatic concentrations (%) may ensure sufficient swelling of certain materials used in the construction of o-rings and seals that have been previously exposed to fuels with high aromatic contents; and, a maximum surface tension (mN/m) may ensure sufficient spray break-up and atomisation of the fuel.

(36) The ratio of the number of hydrogen atoms to the number of carbon atoms in a molecule may influence the specific energy of a given composition, or blend of fuel. Fuels with higher ratios of hydrogen atoms to carbon atoms may have higher specific energies in the absence of bond strain. In some examples, fossil-based hydrocarbon fuels may comprise molecules with approximately 7 to 18 carbon atoms, with a significant portion of a given composition stemming from molecules with 9 to 15 carbons, with an average of 12 carbons.

(37) A number of sustainable aviation fuel blends have been approved for use. For example, some approved blends comprise blend ratios of up to 10% sustainable aviation fuel, whilst other approved blends comprise blend ratios of up to 50% sustainable aviation fuel (the remainder comprising one or more fossil-based hydrocarbon fuels, such as Kerosene), with further compositions awaiting approval. However, there is an anticipation in the aviation industry that sustainable aviation fuel blends comprising up to (and including) 100% sustainable aviation fuel (SAF) will be eventually approved for use.

(38) Sustainable aviation fuels may comprise one or more of n-alkanes, iso-alkanes, cyclo-alkanes, and aromatics, and may be produced, for example, from one or more of synthesis gas (syngas); lipids (e.g. fats, oils, and greases); sugars; and alcohols. Thus, sustainable aviation fuels may comprise either or both of a lower aromatic and sulphur content, relative to fossil-based hydrocarbon fuels. Additionally or alternatively, sustainable aviation fuels may comprise either or both of a higher iso-alkane and cyclo-alkane content, relative to fossil-based hydrocarbon fuels. In some examples, sustainable aviation fuels may comprise either or both of a density of below 100%, for example between 90% and 98%, that of kerosene and a specific energy of above 100%, for example between 101% and 105%, that of kerosene. For example, the calorific value of sustainable aviation fuels may be between 101% and 105% that of kerosene.

(39) In some examples, the sustainable aviation fuel(s), or blend(s) provided to the combustion equipment 16 may be relatively lower in aromatic and/or other non-paraffinic content than that of kerosene. The sustainable aviation fuel may comprise an aromatic content of e.g. 30%, 20%, 15%, 10%, 8%, 5%, or less than 5%; e.g. 4%, 3%, 2%, 1%, or less than 1%; e.g. 0.75%, 0.5%, 0.25%, or less than 0.25%; e.g. 0.2%, 0.1%, or less than 0.1%; e.g. 0.01%, 0.001%, or 0%. The aromatic content of the sustainable aviation fuel may be in an inclusive figure or range bounded by or within any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), e.g. 13.5%, 8.5%, 2.5%, 0.35%, 0.15%, 0.05%, 0.005%, or 0%; or 0% to 0.75%, 0% to 0.5%, or 0.1% to 0.25%; or 0.15% to 0.65%, 0.35% to 0.55%, or 0.035% to 0.055%; according to one or more of preference, fuel stock or supplier, and compositional variation therein.

(40) Owing 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 a higher specific energy (despite, in some examples, a lower energy density); higher specific heat capacity; higher thermal stability; higher lubricity; lower viscosity; lower surface tension; lower freeze point; lower soot emissions; lower NOx; and, lower CO2 emissions, relative to fossil-based hydrocarbon fuels (e.g. when combusted in the combustion equipment 16). Accordingly, relative to fossil-based hydrocarbon fuels, such as Kerosene, sustainable aviation fuels may lead to either or both of a relative decrease in specific fuel consumption, and a relative decrease in maintenance costs.

(41) FIG. 6 shows a section through the combustor 16 of the engine 10 of FIG. 1 in a plane normal to the principal rotational axis 9 of the engine 10. The combustor 16 comprises an annular combustion chamber 401, defined by a liner 402. Alternative combustor configurations may be used in other embodiments, for example cannular, canned, etc. The combustor 16 comprises a plurality of fuel spray nozzles 403, 404 arranged about the circumference of the combustor 16. Each fuel spray nozzle 403, 404 comprises one or more fuel injectors arranged to inject fuel into the combustion chamber 401. In this example, the combustor 16 comprises 16 fuel spray nozzles 403, 404. In other examples, the combustor 16 may comprise any suitable number of fuel spray nozzles 403, 404, for example a number of fuel spray nozzles in the range 14-22. In some examples, the number of fuel spray nozzles 403, 404 may be between 16 and 20. In yet other examples, the number of fuel spray nozzles may be 14, 15, 16, 17, 18, 19, 20, 21, 22, or a number within a range defined between any two of the values in this sentence.

(42) The number of fuel spray nozzles 403, 404 can also be quantified as a ratio of a number of fuel spray nozzles to engine core size. The core size defines the size of the core 11 of the gas turbine engine 10. Engine core size can be defined as:

(43) $Core size = {\overset{}{m}}_{2} \frac{\sqrt{T_{3}}}{P_{3}}$

(44) Where {dot over (m)}.sub.2=the mass flow rate, in lbs per second, of air on entry to the high-pressure compressor 15, T.sub.3=the temperature, in Kelvin, of air on exit from the high-pressure compressor 15, and P.sub.3=the pressure, in lb inches per second squared per inch squared, of air on exit from the high-pressure compressor 15. A unit of core size is therefore expressed as:

(45) $s .Math. K^{\frac{1}{2}} .Math. i n$

(46) The core size (in s.Math.K.sup.1/2.Math.in) of the engine may 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 engine core size (in s.Math.K.sup.1/2.Math.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 range defined between any two of these values. In yet further examples, the engine core size (in s.Math.K.sup.1/2.Math.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 range defined between any two of these values.

(47) The number of fuel spray nozzles per unit engine core size (in the units given above) may be in the range of 2 to 6, for example, 2, 3, 4, 5, 6, or within a range defined between any two of those values. The number of fuel spray nozzles per unit engine core size may be in the range 2.7 to 4, preferably in the range 3 to 3.6. 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 range defined between any two of these values. In yet further examples, the number of fuel spray nozzles per unit engine core size may be in the range of 3 to 4, for example 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 range defined between any two of these values. In yet further 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 range defined between any two of those values.

(48) The core size is defined herein at an engine operation condition corresponding to a maximum value of the semi-non-dimensional flow at high pressure compressor entry, defined as:

(49) ${\dot{m}}_{2} \frac{\sqrt{T_{2}}}{P_{2}}$ where {dot over (m)}.sub.2 is the mass flow rate (in lbs per second) of air on entry to the high-pressure compressor, T.sub.2 is the temperature (in Kelvin) of air on entry to the high pressure compressor, and P.sub.2 is the pressure (in lb inches per second squared per inch squared) of air on entry to the high-pressure compressor.

(50) The operating condition corresponding to the maximum semi-non-dimensional flow at High Pressure compressor entry may be the top of climb operating condition. The core size referred to herein may therefore be defined at the top of climb operating condition. The top of climb may be as defined in the art and as understood by the skilled person for a specific implementation of a gas turbine of the present application. In one specific example, the top of climb may correspond to operating at an altitude of between 30,000 ft to 39,000 ft (more specifically 35,000 ft), a forward speed of Mach Number 0.75 to 0.85, and ambient air temperature (TAMB) of ISA+10K to ISA+15K.

(51) The combustor 16 comprises a number of duplex fuel spray nozzles 403 (also known as internally-staged nozzles) in which a primary fuel injector is integrated in the same fuel nozzle as a main fuel injector. The combustor 16 also comprises a number of single flow fuel spray nozzles 404 which each comprise a main fuel injector only. In other examples, the combustor 16 may comprise only duplex fuel spray nozzles or only single flow fuel spray nozzles.

(52) In this example, the combustor 16 comprises 12 duplex fuel spray nozzles 403 and 4 single flow fuel spray nozzles 404. The duplex fuel spray nozzles 403 are illustrated in FIG. 6 by shaded circles. The duplex fuel spray nozzles 403 are arranged in groups of three about the circumference of the combustor 16, with each group being arranged diametrically opposite another of the groups. In other examples, the combustor 16 may comprise any suitable number of duplex fuel spray nozzles, for example in the range of 10-14 nozzles, and any suitable number of single flow fuel spray nozzles, for example in the range of 4-8 nozzles. In some examples, the number of duplex fuel spray nozzles may be 10, 11, 12, 13 or 14, or within a range defined 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 range defined 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 comprise any suitable number of nozzles, for example in the range of 2 to 8 nozzles. In some examples, each group of duplex nozzles may comprise 2, 3, 4, 5, 6, 7 or 8 fuel spray nozzles, or a number within a range defined between any two of those values.

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

(54) In the example shown, when the engine 10 is operating at low power (below a staging point), for example during or shortly after start-up, fuel is supplied only to the primary injectors of the duplex fuel spray nozzles 403 for delivery to the combustion chamber 401. A greater fuel flow rate 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 of air through engine 10 increases, the staging point is reached at which fuel is additionally supplied to the main fuel injectors of one or more of the 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 main fuel injectors of both the duplex fuel spray nozzles 403 and the single flow fuel spray nozzles 404, in addition to fuel injected by the primary injectors of the duplex fuel spray nozzles 403. In this example, the flow rate of fuel supplied to the main injectors of the single flow fuel spray nozzles 404 is less than or equal to the flow rate of fuel supplied to the main injectors of the duplex fuel spray nozzles 403. Therefore, because both the primary and main injectors of the duplex fuel spray nozzles 403 are receiving 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 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 at and above the staging point, i.e., fuel is not supplied to the primary injectors of the duplex fuel spray nozzles 403.

(55) Fuel flow delivered to the plurality of fuel spray nozzles is therefore biased such that the fuel flow rate 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 delivered 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 provided to fuel spray nozzles which are located relatively closer to the ignitors 405 to aid ignition and flame stability at low engine powers, engine start-up, or during an engine re-light. In some examples, the first subset (e.g. the duplex nozzles) of fuel spray nozzles may comprise at least one half, preferably at least two thirds, of the total number of fuel spray nozzles.

(56) In other examples, the rate of fuel flow to each fuel spray nozzle provided in the combustor may be the same and there may be no biasing of the fuel flow to a subset of the nozzles. In such an example, all of the fuel flow nozzles may be single flow nozzles or they may all be duplex nozzles. In yet other examples, other arrangements of fuel spray nozzles may be provided in which fuel is biased to those adjacent, or closer, to the ignitors. For example, two subsets (that are independently controllable) of duplex nozzles or two subsets of single flow nozzles may be provided which can be biased as described above.

(57) FIG. 7 shows one of the duplex fuel spray nozzles 403 of the combustor 16. The duplex nozzle 403 comprises a primary fuel injector 501, a main fuel injector 502 and an air duct 503. The primary injector 501 comprises a primary inlet 504 arranged to receive a primary flow of fuel P and a primary fuel circuit 505 arranged to deliver 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 deliver the main flow of fuel to the outlet 506 of the nozzle 403. The air duct 503 receives high-pressure air from the high-pressure compressor 15 and delivers the high-pressure air to the outlet 506 of the nozzle 403.

(58) The duplex nozzle 403 is configured to produce, at the outlet 506 of the nozzle 403, a primary cone of fuel from the primary injector 501 and a main cone of fuel from the main injector 502 (illustrated in FIG. 7 by the dashed lines labelled P and M respectively). When both the primary and main injectors 501, 502 are active, the primary and main cones are arranged concentrically, with the main cone arranged annularly outside of the primary cone. Those skilled in the art will be familiar with such fuel spray patterns.

(59) It will be appreciated that the duplex nozzle 403 of FIG. 7 is merely exemplary and that other examples may utilise an alternative configuration of duplex nozzle.

(60) FIG. 8 shows one of the single flow fuel spray nozzles 404 of the combustor 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 main fuel circuit 603 arranged to deliver 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 dashed lines labelled M). Air is similarly suppled to the outlet 604 of the nozzle by an air duct 605.

(61) It will be appreciated that the single flow fuel spray 404 of FIG. 8 is merely exemplary and that other examples may utilise an alternative configuration of single flow fuel spray nozzle 404.

(62) FIGS. 9 and 10 each show a section through the engine 10, viewed perpendicularly to the principal rotational axis 9, including a portion of the combustor 16 comprising one of the duplex fuel spray nozzles 403 and one of the ignitors 405. A similar arrangement is provided at the location of the single flow fuel spray nozzles 404 as for the duplex nozzles 403. The combustor 16 is mounted within a cavity 406 formed by an inner air casing 407 and outer air casing 408. In operation, the high-pressure compressor 15 delivers high-pressure air D to the cavity 406 via a diffuser 409. At this point, a quantity of the air enters the combustor 16 as combustion air E through the fuel nozzle 403 and/or mixing ports at the entrance to the combustor 16. The remaining air flows around the combustor 16 as cooling air G, some of which is admitted downstream of the fuel nozzle 403 as described below with reference to FIG. 10.

(63) One or more temperature and/or pressure probes (not shown) may be installed in the casing of the diffuser 409 and arranged to measure the temperature and/or pressure of the high-pressure air D delivered to the cavity 406 from the high-pressure compressor 15 via the diffuser 409 (i.e. the temperature and pressure at the high-pressor compressor 15 exit). Such a temperature probe may be referred to as a T3 probe and such a pressure probe may be referred to as a P3 probe. It will be appreciated that the engine 10 may comprise any suitable arrangement of pressure and temperature probes which may be positioned at any suitable location within the engine 10. As used herein, T3 and P3, and any other numbered pressures and temperatures, may be defined using the station numbering listed in standard SAE AS755.

(64) The combustor 16 operates as a rich burn, quick quench, lean burn (RQL) combustor. In other examples, the combustor 16 may be an alternative type of combustor, such as a standard rich-burn combustor (with no fuel flow biasing). Referring to FIG. 10, the combustion chamber 401 of the RQL combustor 16 is divided into three zones along the length of the combustor 16: a rich zone 801, a quick quench 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 ignitor 405. Within the rich zone 801 fuel is burnt at a fuel/air ratio higher than stoichiometric (for example, at an equivalence ratio of about 1.8). Air is then introduced to the combustion products, via primary ports 804 arranged in the liner 402 of the combustor 16, before the combustion products reach the quick quench zone 802. Further air is added to the still burning fuel via the primary ports 804 (which may be referred to as quench ports). Air is added by the primary ports 804 at a higher rate (e.g. higher than within the rich zone) thereby quenching the combustion to a significantly lower than stoichiometric fuel/air ratio (for example, at an equivalence ratio of between 0.5 and 0.7), while continuing to allow the fuel to burn. Consequently, very little of the combustion process may be carried out at close to stoichiometric fuel/air ratios, and so relatively little nitrogen oxides (NOx) is produced. Air is then again introduced to the combustion products, via secondary ports 805 arranged in the liner 402 of the combustor 16, while the combustion products are in the lean zone 803 (or just before they reach the lean zone 803). Within the lean zone 803, fuel is burnt at a fuel/air ratio lower than stoichiometric (for example, at an equivalence ratio of between 0.5 and 0.7). After passing through the lean zone 803, the combustion products exit the combustor 16. The secondary ports 805 may be referred to as dilution ports, and may be arranged to gradually introduce dilution air into the lean zone 803. Fuel added by the fuel spray nozzle is substantially completely burnt by the time the air exits at an outlet of the combustor, prior to flowing to the turbine.

(65) FIG. 11 shows a portion of a propulsion system 900 for an aircraft. The propulsion system 900 comprises the gas turbine engine 10 of FIG. 1. The engine 10 further comprises a fuel system and an oil system. The fuel system comprises: 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 distributing valve 909. The propulsion system 900 further comprises a fuel tank 901. The oil system comprises 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 forming 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 on board the aircraft to which the gas turbine engine is mounted.

(66) The low pressure fuel pump 902 is arranged to deliver fuel from the fuel tank 901 to the fuel-oil heat exchanger 903 via a suitable arrangement of pipes, conduits etc. (not shown). The main fuel pump 904 is configured to deliver fuel from the fuel-oil heat exchanger 903 to the fuel spray nozzles of the combustor 16 via the fuel distributing valve 909 and a suitable arrangement of pipes, conduits etc. (not shown). The fuel distributing valve 909 is arranged to distribute 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 FIG. 11. It therefore provides fuel to all of the duplex 403 and single flow 404 fuel spray nozzles. 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 therefore be used to provide a greater flow rate of fuel to the first subset of fuel spray nozzles (e.g. the duplex fuel spray nozzles 403 in the present example) compared to the flow rate of fuel provided to the second subset of fuel spray nozzles via the main manifold 909a. For example, below a threshold engine power, fuel may be supplied only to the first subset of fuel spray nozzles via the primary manifold 909b, or supplied to the first subset of fuel spray nozzles at a greater fuel flow rate compared to the second subset of fuel spray nozzles. This may limit the production of undesirable combustion products such as oxides of nitrogen (NOx), unburned hydrocarbons (HC), and carbon monoxide (CO), and may bias fuel flow to injectors nearest to the ignitors to aid flame stability and ignition at low engine powers.

(67) The oil feed pump 906 is arranged to deliver lubricating oil from the oil tank 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 deliver 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 flow path of fuel from the fuel tank 901 to the combustor 16, via the pumps 902, 904 and the fuel-oil heat exchanger 903, in operation is illustrated in FIG. 11 by dashed or dotted arrows. The flow path of oil from the oil tank 905 to the fuel-oil heat exchanger 903, via the oil feed pump 906, and on to components of the engine 10 in operation is illustrated in FIG. 11 by solid arrows.

(68) The controller 908 comprises a suitable arrangement of processors and electronic memory. The controller 908 is in communication with fuel-oil heat exchanger 903, as illustrated by the dashed and dotted line in FIG. 11, and is configured to control operation of the fuel-oil heat exchanger 903. In some examples, the controller 908 may be configured to control a flow rate of oil through the fuel-oil heat exchanger 903. The controller 908 is configured to control operation of the fuel-oil heat exchanger 903 by providing control signals to the fuel-oil heat exchanger 903. The controller 908 is configured to control operation of the fuel-oil heat exchanger 903 to adjust at least one property or parameter of the fuel on entry to the combustor 16. In the example shown, the controller 908 is configured to control operation of the fuel-oil heat exchanger 903 to control a viscosity of the fuel on entry to the combustor 16. In other examples, the controller 908 may additionally or alternatively be configured to control operation of the fuel-oil heat exchanger 903 to control a temperature of the fuel on entry to the combustor. The controller 908 may be a separate controller as illustrated, or may form part of an Engine Electronic Controller (EEC) arranged to control other engine functions.

(69) 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 at any suitable location or position relative to the other components of the propulsion system 900. In other examples, the propulsion system 900 may comprise one or more further heat exchangers arranged to receive oil from the oil system, or the propulsion system 900 may comprise one or more further oil systems arranged to supply oil to the one or more further heat exchangers. It will be appreciated that the propulsion system 900 as shown in FIG. 11 is merely a schematic view of an illustrative propulsion system.

(70) In one example, the controller 908 is configured to control operation of the fuel-oil heat exchanger 903 to lower the fuel viscosity to 0.58 mm.sup.2/s or lower on entry to the combustor 16 at cruise conditions. Alternatively, the controller 908 may be configured to control operation of the fuel-oil heat exchanger 903 to lower the fuel viscosity to between 0.58 mm.sup.2/s and 0.30 mm.sup.2/s, for 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.sup.2/s on entry to the combustor 16 at cruise conditions. Alternatively, the controller 908 may be configured to control operation of the fuel-oil heat exchanger 903 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.sup.2/s or lower, or to within any range defined between any two of these values, on entry to the combustor 16 at cruise conditions. The controller 908 may be configured to control operation of the fuel-oil heat exchanger 903 to lower the fuel viscosity to between 0.55 mm.sup.2/s and 0.35 mm.sup.2/s, 0.53 mm.sup.2/s and 0.35 mm.sup.2/s, 0.50 mm.sup.2/s and 0.35 mm.sup.2/s, 0.48 mm.sup.2/s and 0.35 mm.sup.2/s, 0.48 mm.sup.2/s and 0.38 mm.sup.2/s, 0.48 mm.sup.2/s and 0.40 mm.sup.2/s, 0.46 mm.sup.2/s and 0.40 mm.sup.2/s, 0.44 mm.sup.2/s and 0.40 mm.sup.2/s, or 0.44 mm.sup.2/s and 0.42 mm.sup.2/s on entry to the combustor 16 at cruise conditions.

(71) The controller 908 may additionally or alternatively be configured to control operation of the engine 10 such that a reduction of 10-70% in an average of particles/kg of nvPM in the exhaust of the gas turbine engine 10 when the engine 10 is operating at 85% available thrust for given operating conditions and particles/kg of nvPM in the exhaust of the gas turbine engine 10 when the engine 10 is operating at 30% available thrust for the given operating conditions is obtained when a fuel provided to the combustor 16 is a sustainable aviation fuel instead of a fossil-based hydrocarbon fuel. In other examples, the nvPM reduction may be as otherwise defined herein.

(72) In this example, or any other example described herein, the controller 908 is configured to control the fuel distribution valve 909 to control delivery of the fuel to the fuel spray nozzles of the combustor 16. The controller 908 is configured to bias fuel flow to the nozzles such that the first subset of plurality of fuel spray nozzles 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 delivered only to the primary fuel injectors of the duplex fuel spray nozzles 403. Above a staging point, the controller 908 is configured to control the fuel distribution valve 909 such that fuel is additionally delivered to the main fuel injectors of the duplex fuel spray nozzles 403 and the single flow fuel spray nozzles 404. As such, the duplex fuel spray nozzles 403 receive more fuel than the single flow fuel spray nozzles 404 (below and optionally above the staging point). The controller 908 may alternatively be configured to control the fuel distribution valve 909 to control delivery of the fuel such that any suitable subset of fuel spray nozzles 403, 404 receive more fuel than the other fuel spray nozzles 403, 404. This advantageously enables fuel delivery to be optimised for engine performance, emissions, or any other suitable criteria. The fuel delivery system shown in the Figures is to be understood as one example of how fuel is biased to the fuel spray nozzles, with others being possible. For example, two sets of independent single flow nozzles may be provided.

(73) The gas turbine engine 10 of the present application is configured to provide fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles 403, 404. In other words, the gas turbine engine 10 is configured to inject fuel (F) comprising a sustainable aviation fuel (SAF) into the combustion chamber 401. In use, therefore, fuel provided to the fuel spray nozzles 403, 404 comprises SAF.

(74) By fuel comprising SAF we may mean that the fuel provided to the combustor 16 (and to the combustion chamber 401), via the fuel spray nozzles 403, 404, comprises a % SAF in the range of 50% to 100%, preferably in the range 70% to 100%, and more preferably in the range 90% to 100%. More generally, by fuel comprising SAF we may mean a fuel comprising any blend of SAF and fossil kerosene fuel, including up to 100% SAF and no fossil kerosene fuel. The fuel comprising SAF may be a fuel comprising a percentage SAF of 1%, 2%, 5%, 10%, 15%, 20%, 30%, 40%, 50%, 60%, 70%, 80%, 90%, 95%, 100%, or within any range defined between any two of those values.

(75) By SAF we mean a kerosene-type fuel whose hydrocarbon component is substantially all paraffinic hydrocarbons. By SAF we additionally or alternatively mean a kerosene-type fuel with a hydrogen mass fraction in the range of 13.7% to 16.9%, for example 15.3%. By a fossil-based hydrocarbon fuel or fossil fuel used anywhere herein we mean a fossil derived kerosene with a hydrogen mass fraction in the range of 12.0% to 14.8%, for example 13.4%.

(76) Non-Volatile Particulate Matter (nvPM) Emissions

(77) An nvPM emissions index (EI) is defined herein as the mass of nvPM produced per unit mass of fuel used by the combustor 16 of the gas turbine engine 10 of any example described herein. In the present application, the nvPM emissions index is the mass of nvPM produced by the gas turbine engine 10 in milligrams divided by the corresponding mass of the fuel used by the engine 10 in kilograms.

(78) The nvPM emissions index can be defined at various operating phases of the gas turbine engine 10, for example at idle, max take off, climb and approach. An emissions index may be further defined depending on the type of fuel being provided to the combustor 16.

(79) The following emissions index parameters are defined for the gas turbine engine 10: i) EI.sub.idle is the nvPM emissions index in mg/kg of the gas turbine engine 10 if operating at 7% available thrust for given operating conditions. Operation at 7% available thrust may correspond to operating at an idle operating phase of the gas turbine engine 10; ii) EI.sub.maxTO is the nvPM emissions index in mg/kg of the gas turbine engine 10 if operating at 100% available thrust for given operating conditions. Operation at 100% available thrust may correspond to operating at a max take off operating phase of the gas turbine engine 10; iii) EI.sub.climb is the nvPM emissions index in mg/kg of the gas turbine engine 10 when operating at 85% available thrust for given operating conditions. Operation at 85% available thrust may correspond to operating at a climb operating phase of the gas turbine engine 10; iv) EI.sub.approach is the nvPM emissions index in mg/kg of the gas turbine engine 10 when operating at 30% available thrust for given operating conditions. Operation at 30% available thrust may correspond to operating at an approach operating phase of the gas turbine engine 10.

(80) The available thrust for given operating conditions (i.e. engine power setting) is defined as a percentage of the engine maximum rated thrust (Foo) as defined in the art. In other words, a percentage available thrust refers to a percentage of a maximum thrust, where the maximum thrust is 100% available thrust, and given operating conditions refers to predetermined operating conditions at which engine maximum rated thrust, i.e., 100% available thrust, is measured. The predetermined operating conditions may be ISA at sea level where the reference absolute humidity is 0.00634 kg water/kg dry air. The predetermined operating conditions may be at sea level static. The predetermined operating conditions may include no customer bleeds and/or no power offtakes. The predetermined operating conditions may be at day conditions. The predetermined operating conditions may be at around 60% relative humidity.

(81) The nvPM emissions indexes defined above may be further defined according to the fuel being supplied to the combustor. Fuel specific values of the nvPM emissions index are defined as follows: i) EI.sub.idle,FF is the nvPM emissions index in mg/kg of the gas turbine engine 10 if operating at 7% available thrust for given operating conditions and if a fuel provided to the combustor 16 is a fossil-based hydrocarbon fuel; ii) EI.sub.maxTO,FF is the nvPM emissions index in mg/kg of the gas turbine engine 10 if operating at 100% available thrust for given operating conditions and if a fuel provided to the combustor 16 is a fossil-based hydrocarbon fuel; iii) EI.sub.climb,FF is the nvPM emissions index in mg/kg of the gas turbine engine 10 when operating at 85% available thrust for given operating conditions and if a fuel provided to the combustor 16 is a fossil-based hydrocarbon fuel; iv) EI.sub.approach,FF is the nvPM emissions index in mg/kg of the gas turbine engine 10 when operating at 30% available thrust for given operating conditions and if a fuel provided to the combustor 16 is a fossil-based hydrocarbon fuel; v) EI.sub.idle,SAF is the nvPM emissions index in mg/kg of the gas turbine engine 10 if operating at 7% available thrust for given operating conditions and if a fuel provided to the combustor 16 comprises a sustainable aviation fuel (SAF); vi) EI.sub.maxTO,SAF is the nvPM emissions index in mg/kg of the gas turbine engine 10 if operating at 100% available thrust for given operating conditions and if a fuel provided to the combustor 16 comprises a sustainable aviation fuel (SAF); vii) EI.sub.climb,SAF is the nvPM emissions index in mg/kg of the gas turbine engine 10 when operating at 85% available thrust for given operating conditions and if a fuel provided to the combustor 16 comprises a sustainable aviation fuel (SAF); and viii) EI.sub.approach,SAF is the nvPM emissions index in mg/kg of the gas turbine engine 10 when operating at 30% available thrust for given operating conditions and if a fuel provided to the combustor 16 comprises a sustainable aviation fuel (SAF).
Fuel Flow Rate

(82) A fuel flow rate W.sub.f, of the gas turbine engine 10 is defined as the rate of fuel flow to the fuel spray nozzles 403, 404 of the combustor 16 (i.e. when the engine 10 is in use). The fuel flow rate is defined for operation at different percentages of available thrust for given operating conditions as defined above. W.sub.f,idle is the rate of fuel flow to the fuel spray nozzles 403, 404 in kg/s at 7% available thrust for given operating conditions and may correspond to operating at an idle operating phase of the gas turbine engine 10. W.sub.f,maxTO is the rate of fuel flow to the fuel spray nozzles 403, 404 in kg/s at 100% available thrust for given operating conditions and may correspond to operating at a max take off operating phase of the gas turbine engine 10. W.sub.f,climb is defined as the mass flow rate of fuel provided to the fuel spray nozzles 403, 404 in kg/s when the gas turbine engine 10 is operating at 85% available thrust for given operating conditions and may correspond to operating at a climb operating phase of the gas turbine engine 10. W.sub.f,approach is the mass flow rate of fuel provided to the fuel spray nozzles 403, 404 in kg/s when the gas turbine engine 10 is operating at 30% available thrust for given operating conditions and may correspond to operating at an approach operating phase of the gas turbine engine 10.

(83) In any example defined or claimed anywhere herein, W.sub.f,maxTO may be in the range 0.441 to 1.23 kg/s, preferably 0.496 to 1.13 kg/s, more preferably 0.551 to 1.03 kg/s.

(84) In any example defined or claimed anywhere herein, W.sub.f,climb may be in the range 0.369 to 1.01 kg/s, preferably 0.415 to 0.923 kg/s, more preferably 0.461 to 0.839 kg/s.

(85) In any example defined or claimed anywhere herein, W.sub.f,approach may be in the range 0.133 to 0.334 kg/s, preferably 0.149 to 0.306 kg/s, more preferably 0.166 to 0.278 kg/s.

(86) In any example defined or claimed anywhere herein, W.sub.f,idle may be in the range 0.0516 to 0.119 kg/s, preferably 0.0581 to 0.109 kg/s, more preferably 0.0645 to 0.0990 kg/s.

(87) Engine Thrust

(88) The thrust of the gas turbine engine 10 is given the symbol F and is defined for operation at different percentages of available thrust for given operating conditions as defined above. F.sub.maxTO is defined as the thrust of the gas turbine engine 10 at 100% available thrust for given operating conditions in kN. F.sub.idle is defined as the thrust of the gas turbine engine 10 at 7% available thrust for given operating conditions in kN.

(89) In any of the examples defined or claimed anywhere herein, F.sub.maxTO may be in the range 54.1 kN to 177 kN and preferably in the range 60.8 kN to 163 kN and more preferably in the range 67.6 kN to 148 kN. The value of F.sub.maxTO corresponds to the maximum rated thrust Foo.

(90) In any of the examples defined or claimed anywhere herein, F.sub.idle may be in the range 3.78 kN to 12.4 kN and preferably in the range 4.26 kN to 11.4 kN and more preferably in the range 4.73 kN to 10.4 kN.

(91) The thrust at other operating points (e.g. approach and climb) may be defined by taking the relevant percentage value of the maximum rated thrust, Foo.

(92) Bypass Ratio

(93) The bypass ratio (BPR) is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core.

(94) In any of the examples defined or claimed anywhere herein, the BPR may be in the range of 6.66 to 15.3 and more preferably in the range of 7.49 to 14.0 and even more preferably in the range of 8.33 to 12.8.

(95) First and Second Idle-MTO nvPM Emissions Index Ratios

(96) A first idle-MTO nvPM emissions index ratio is defined in equation (1) below:

(97) $\begin{matrix} \frac{{EI}_{idle}}{{EI}_{ma xTO}} & (1) \end{matrix}$ EI.sub.idle and EI.sub.maxTO are as defined elsewhere herein. The first idle-MTO nvPM emissions index ratio represents the ratio of the nvPM emissions index at idle (e.g. at 7% available thrust) to the nvPM emissions index at max take off (e.g. at 100% available thrust). The first idle-MTO nvPM emissions index ratio of the gas turbine engine 10 may be less than 0.8.

(98) In other examples, the first idle-MTO nvPM emissions index ratio of the gas turbine engine 10 may be less than 0.708 and preferably less than 0.649 and more preferably less than 0.59.

(99) The first idle-MTO nvPM emissions index ratio of the gas turbine engine 10 may be less than or equal to 0.5 and preferably less than or equal to 0.4 and more preferably less than or equal to 0.3.

(100) The first idle-MTO nvPM emissions index ratio of the gas turbine engine 10 may be less than or equal to 0.105 and preferably less than or equal to 0.0962 and more preferably less than or equal to 0.0875.

(101) More generally, the first idle-MTO nvPM emissions index ratio may be less than 0.01, 0.05, 0.07, 0.1, 0.15, 0.2, 0.25, 0.3, 0.35, 0.4, 0.45, 0.5, 0.55, 0.6, 0.65, 0.7, 0.75, 0.8, or in any range defined between any two of these values.

(102) In any of the examples above where an upper bound of the first idle-MTO nvPM emissions index is defined, the first idle-MTO nvPM emissions index may have a lower bound of greater than zero.

(103) In any of the examples above, the first idle-MTO nvPM emissions index ratio of the gas turbine engine 10 may be greater than or equal to 0.0103 and preferably greater than or equal to 0.0115 and more preferably greater than or equal to 0.0128.

(104) In one example, the first idle-MTO nvPM emissions index ratio of the gas turbine engine 10 may be in the range of 0.0103 to 0.105 and preferably in the range of 0.0115 to 0.0962 and more preferably in the range of 0.0128 to 0.0875.

(105) As the values in the previous paragraphs correspond to where the gas turbine engine 10 is operated using fuel comprising a SAF, the first idle-MTO nvPM emissions index ratio equation (1) above is equivalent to EI.sub.idle,SAF/EI.sub.maxTO,SAF for the values in the previous paragraphs.

(106) A second idle-MTO nvPM emissions index ratio (this is also referred to as a an idle-MTO nvPM emissions index ratio elsewhere herein) is defined in equation (2) below:

(107) $\begin{matrix} \frac{{EI}_{idle, SAF} / {EI}_{maxTO, SAF}}{{EI}_{idle, FF} / {EI}_{maxTO, FF}} & (2) \end{matrix}$ EI.sub.idle,SAF, EI.sub.maxTO,SAF EI.sub.idle,FF and EI.sub.maxTO,FF are as defined elsewhere herein. The second idle-MTO nvPM emissions index ratio represents a ratio of the first idle-MTO nvPM emissions index ratio when the gas turbine engine is operated using fuel comprising SAF compared to if it were operated using a fossil-based hydrocarbon fuel.

(108) Additionally, or alternatively, in any example defined or claimed herein, the second idle-MTO nvPM emissions index ratio of the gas turbine engine 10 may be less than 1.

(109) In some examples, the second idle-MTO nvPM emissions index ratio of the gas turbine engine 10 may be less than or equal to 0.8 and preferably less than or equal to 0.6 and more preferably less than or equal to 0.4.

(110) The second idle-MTO nvPM emissions index ratio of the gas turbine engine 10 may be less than or equal to 0.178 and preferably less than or equal to 0.164 and more preferably less than or equal to 0.149.

(111) More generally, the second idle-MTO nvPM emissions index ratio may be less than 0.05, 0.1, 0.15, 0.2, 0.25, 0.3, 0.35, 0.4, 0.45, 0.5, 0.55, 0.6, 0.65, 0.7, 0.75, 0.8, 0.85, 0.9, 0.95, or 1, or in any range defined between any two of these values.

(112) In any of the examples above in which an upper bound of the second idle-MTO nvPM emissions index is defined the lower bound may be greater than zero.

(113) In any of the examples above, the second idle-MTO nvPM emissions index ratio of the gas turbine engine 10 may be greater than or equal to 0.03 and preferably greater than or equal to 0.06 and more preferably greater than or equal to 0.09.

(114) The second idle-MTO nvPM emissions index ratio of the gas turbine engine 10 may be greater than or equal to 0.118 and preferably greater than or equal to 0.133 and more preferably greater than or equal to 0.148.

(115) In one example, the second idle-MTO nvPM emissions index ratio of the gas turbine engine 10 may be in the range 0.118 to 0.178 and preferably in the range 0.133 to 0.164 and more preferably in the range 0.148 to 0.149.

(116) In some examples, the second idle-MTO nvPM emissions index ratio of the gas turbine engine 10 may be 0.118, 0.12, 0.125, 0.13, 0.135, 0.14, 0.145, 0.15, 0.155, 0.16, 0.165, 0.17, 0.175, 0.178, or within any range defined between any two of these values.

(117) Fuel-Flow nvPM Emissions Index Ratio

(118) A fuel-flow nvPM emissions index ratio is defined as:

(119) $\begin{matrix} \frac{{EI}_{idle} W_{f, idle}}{{EI}_{ma xTO} W_{f, ma xTO}} & (3) \end{matrix}$ where W.sub.f,idle is as defined above i.e. is the mass flow rate of fuel provided to the fuel spray nozzles in kg/s at 7% available thrust for given operating conditions; and W.sub.f,maxTO is the mass flow rate of fuel provided to the fuel spray nozzles in kg/s at 100% available thrust for the same given operating conditions. EI.sub.idle and EI.sub.maxTO are as defined elsewhere herein. The fuel flow nvPM emissions index ratio represents the ratio of the nvPM emissions index at idle (e.g. at 7% available thrust) multiplied by the respective fuel flow rate to the nvPM emissions index at max take off (e.g. at 100% available thrust) multiplied by the respective fuel flow rate. Additionally, or alternatively, in any example defined or claimed herein, the fuel-flow nvPM emissions index ratio of the gas turbine engine 10 may be less than 0.08.

(120) In other examples, the fuel-flow nvPM emissions index ratio of the gas turbine engine 10 may be less than 0.0798 and preferably less than 0.0731 and more preferably less than 0.0665.

(121) The fuel-flow nvPM emissions index ratio of the gas turbine engine 10 may be less than or equal to 0.06 and preferably less than or equal to 0.04 and more preferably less than or equal to 0.02.

(122) The fuel-flow nvPM emissions index ratio of the gas turbine engine 10 may less than or equal to 0.0119 and preferably less than or equal to 0.0109 and more preferably less than or equal to 0.00986.

(123) More generally, the fuel-flow nvPM emissions index ratio may be less than 0.01, 0.015, 0.02, 0.025, 0.03, 0.035, 0.04, 0.045, 0.05, 0.055, 0.06, 0.065, 0.07, 0.075, 0.08 or any range defined between any two of these values.

(124) In any of the examples above in which only an upper bound for the fuel-flow nvPM emissions index ratio is defined, the lower bound may be greater than zero.

(125) In any of the above examples, the fuel-flow nvPM emissions index ratio of the gas turbine engine 10 may be greater than or equal to 0.000993 and preferably greater than or equal to 0.00111 and more preferably greater than or equal to 0.00124.

(126) In one example, the fuel-flow nvPM emissions index ratio of the gas turbine engine 10 may be in the range of 0.000993 to 0.0119 and preferably in the range of 0.00111 to 0.0109 and even more preferably in the range of 0.00124 to 0.00986.

(127) Thrust nvPM Emissions Index Ratio

(128) A thrust nvPM emissions index ratio is defined as:

(129) 0 $\begin{matrix} \frac{{EI}_{maxTO} / F_{maxTO}}{{EI}_{idle} / F_{idle}} & (4) \end{matrix}$ where F.sub.maxTO is as defined above i.e. is the thrust of the gas turbine engine 10 at 100% available thrust in kN (i.e. the maximum rated thrust, Foo) and F.sub.idle is the thrust of the gas turbine engine 10 at 7% available thrust in kN (i.e. F.sub.000.07). EI.sub.idle and EI.sub.maxTO are as defined elsewhere herein. The thrust nvPM emissions index ratio represents the ratio of the nvPM emissions index at max take off divided by respective thrust to the nvPM emissions index at idle divided by the respective thrust. Additionally, or alternatively, in any example defined or claimed herein, the thrust nvPM emissions index ratio of the gas turbine engine 10 may be greater than 0.09.

(130) In some examples, the thrust nvPM emissions index ratio of the gas turbine engine 10 may be greater than 0.0949 and preferably greater than 0.106 and more preferably greater than 0.118.

(131) The thrust nvPM emissions index ratio of the gas turbine engine 10 may be greater than or equal to 0.15 and preferably greater than or equal to 0.3 and more preferably greater than or equal to 0.45.

(132) The thrust nvPM emissions index ratio of the gas turbine engine 10 may be greater than or equal to 0.64 and preferably greater than or equal to 0.72 and more preferably greater than or equal to 0.8.

(133) More generally, the thrust nvPM emissions index ratio may be greater than 0.094, 0.1, 0.2, 0.3, 0.4, 0.5, 0.6, 0.7, 0.8, 0.9, 1, 1.1, 1.2, 1.3, 1.4, 1.5, 1.6, 1.7, 1.8, 1.9, 2 or within any range defined between any two of these values.

(134) In any of the examples above in which only a lower bound of the thrust nvPM emissions index ration is defined, the upper bound may be as defined in the following paragraphs.

(135) The thrust nvPM emissions index ratio of the gas turbine engine 10 may be less than or equal to 6.53 and preferably less than or equal to 5.98 and more preferably less than or equal to 5.44.

(136) In one example, the thrust nvPM emissions index ratio of the gas turbine engine 10 may be in the range of 0.640 to 6.53 and preferably in the range of 0.720 to 5.98 and even more preferably in the range 0.800 to 5.44.

(137) Lean and Rich Cruise-MTO nvPM Emissions Index Ratio

(138) A lean cruise-MTO nvPM emissions index ratio is defined as:

(139) $\begin{matrix} \frac{{EI}_{cruise (lean)} / {EI}_{\max TO}}{BPR} & (5) \end{matrix}$ where EI.sub.cruise (lean) is defined as:

(140) $\begin{matrix} \frac{{EI}_{maxTO} + {EI}_{climb}}{2} & (6) \end{matrix}$ EI.sub.cruise (lean) represents the nvPM emissions index when the gas turbine engine 10 is operating in a lean cruise operating phase. EI.sub.cruise (lean) is determined by finding the average (mean) of the nvPM emissions index corresponding to when the gas turbine engine 10 is operating in a max take off operating phase (i.e. at 100% available thrust) and when it is operating in a climb operating phase (i.e. operating at 85% available thrust). In equations (5) and (6) above, EI.sub.maxTO and EI.sub.climb are as defined elsewhere herein. BPR is the bypass ratio of the gas turbine engine 10 as defined above. The bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core. The lean cruise-MTO nvPM emissions index ratio represents a ratio of the emissions index at lean cruise to the emissions index at max take off, divided by the BPR.

(141) Additionally, or alternatively, in any example defined or claimed herein, the lean cruise-MTO nvPM emissions index ratio of the gas turbine engine 10 may be less than 0.2.

(142) The lean cruise-MTO nvPM emissions index ratio of the gas turbine engine 10 may be less than 0.18, preferably less than 0.16, and further preferably less than 0.14.

(143) The lean cruise-MTO nvPM emissions index ratio of the gas turbine engine 10 may be less than 0.135, preferably less than 0.124 and further preferably less than 0.113.

(144) The lean cruise-MTO nvPM emissions index ratio of the gas turbine engine 10 may be less than or equal to 0.118, preferably less than or equal to 0.109, and further preferably less than or equal to 0.0983.

(145) More generally, the lean cruise-MTO nvPM emissions index ratio may be less than 0.01, 0.02, 0.03, 0.04, 0.05, 0.06, 0.07, 0.08, 0.09, 0.1, 0.11, 0.12, 0.13, 0.14, 0.15, 0.16, 0.17, 0.18, 0.19, or 0.2, or within any range defined between any two of these values.

(146) In any of the examples defined above, where only an upper bound for the lean-cruise-MTO nvPM emissions index ratio is defined, the lower bound may be greater than zero.

(147) In any of the examples above, the lean cruise-MTO nvPM emissions index ratio of the gas turbine engine 10 may be greater than or equal to 0.0426, preferably greater than or equal to 0.0479, and further preferably greater than or equal to 0.0533.

(148) In one example, the lean cruise-MTO nvPM emissions index ratio of the gas turbine engine 10 may be in the range of 0.0426 to 0.118, preferably in the range of 0.0479 to 0.109 and further preferably in the range of 0.0533 to 0.0983.

(149) A rich cruise-MTO nvPM emissions index ratio is defined as:

(150) $\begin{matrix} \frac{{EI}_{cruise (rich)} / {EI}_{ma xTO}}{BPR} & (7) \end{matrix}$ where EI.sub.cruise (rich) is defined as:

(151) $\begin{matrix} \frac{{EI}_{climb} + {EI}_{approach}}{2} & (8) \end{matrix}$ EI.sub.cruise (rich) represents the nvPM emissions index when the gas turbine engine 10 is operating in a rich cruise operating phase. EI.sub.cruise (rich) is determined by finding the average (mean) of the nvPM emissions index corresponding to when the gas turbine engine 10 is operating in a climb operating phase (i.e. at 85% available thrust) and when it is operating in an approach operating phase (i.e. operating at 30% available thrust). In equations (7) and (8) above, EI.sub.maxTO, EI.sub.climb, EI.sub.approach, and BPR are as defined above. The rich cruise-MTO nvPM emissions index ratio represents a ratio of the emissions index at rich cruise to the emissions index at max take off, divided by the BPR.

(152) Additionally, or alternatively, in any example defined or claimed herein, the rich cruise-MTO nvPM emissions index ratio of the gas turbine engine 10 may be less than 0.07.

(153) The rich cruise-MTO nvPM emissions index ratio of the gas turbine engine 10 may be less than 0.0635, preferably less than 0.0582 and further preferably less than 0.0529.

(154) The rich cruise-MTO nvPM emissions index ratio of the gas turbine engine 10 may be less than 0.065, preferably less than 0.06, and further preferably less than 0.055.

(155) The rich cruise-MTO nvPM emissions index ratio of the gas turbine engine 10 may be less than or equal to 0.0462, preferably less than or equal to 0.0424, and further preferably less than or equal to 0.0385.

(156) More generally, the rich cruise-MTO nvPM emissions index ratio may be less than 0.005, 0.01, 0.015, 0.02, 0.025, 0.03, 0.035, 0.04, 0.045, 0.05, 0.055, 0.06, 0.065, 0.07, or within any range defined between any two of these values. Alternatively, the rich cruise-MTO nvPM emissions index ratio may be less than or equal to 0.005, 0.007, 0.01, 0.013, 0.015, 0.017, 0.02, 0.023, 0.025, 0.027, 0.03, 0.033, 0.035, 0.037, 0.04, 0.043, 0.045, 0.0462, or within any range defined between any two of these values.

(157) In any of the examples defined above, where only an upper bound for the rich-cruise-MTO nvPM emissions index ratio is defined, the lower bound may be greater than zero.

(158) In any of the examples defined above, the rich cruise-MTO nvPM emissions index ratio of the gas turbine engine 10 may be greater than or equal to 0.0128, preferably greater than or equal to 0.0144, and further preferably greater than or equal to 0.016.

(159) In one example, the rich cruise-MTO nvPM emissions index ratio of the gas turbine engine 10 may be in the range of 0.0128 to 0.0462, preferably in the range of 0.0144 to 0.0424 and further preferably in the range of 0.0160 to 0.0385.

(160) The bypass ratio of the gas turbine engine 10 may be as defined anywhere herein. More specifically, the bypass ratio may be in the range of 6.66 to 15.3 and more preferably in the range of 7.49 to 14.0 and even more preferably in the range of 8.33 to 12.8.

(161) MTO, Climb, Approach and Idle nvPM Emissions Index Ratio

(162) An MTO nvPM emissions index ratio is defined as:

(163) $\begin{matrix} \frac{{EI}_{maxTO, SAF}}{{EI}_{maxTO, FF}} & (9) \end{matrix}$

(164) Where EI.sub.maxTO,SAF and EI.sub.maxTO,FF are as defined above. The MTO nvPM emissions index ratio represents a ratio of the nvPM emissions index corresponding to operating the gas turbine engine 10 during the max take off operating phase using a fuel comprising SAF to if the gas turbine engine 10 were instead operated using a fossil-based hydrocarbon fuel.

(165) Additionally, or alternatively, in any example defined or claimed herein, the MTO nvPM emissions index ratio of the gas turbine engine 10 may be less than 1.

(166) The MTO nvPM emissions index ratio of the gas turbine engine 10 may be less than or equal to 0.93, and preferably less than or equal to 0.86, and more preferably less than or equal to 0.79.

(167) The MTO nvPM emissions index ratio of the gas turbine engine 10 may be less than or equal to 0.776, and preferably less than or equal to 0.711, and more preferably less than or equal to 0.646.

(168) More generally, the MTO nvPM emissions index ratio of the gas turbine engine 10 may be less than 1, 0.95, 0.9, 0.85, 0.8, 0.75, 0.7, 0.65, 0.6, 0.55, 0.5, 0.45, 0.4, 0.35, 0.3, 0.25, 0.2, 0.15, 0.1, or 0.05, or any range defined between any two of these values.

(169) In any of the examples above in which only an upper bound of the MTO nvPM emissions index ratio is defined, the lower bound may be greater than zero.

(170) In any of the examples above, the MTO nvPM emissions index ratio of the gas turbine engine 10 may be greater than or equal to 0.15, and preferably greater than or equal to 0.3, and more preferably greater than or equal to 0.45.

(171) The MTO nvPM emissions index ratio of the gas turbine engine 10 may be greater than or equal to 0.516, and preferably greater than or equal to 0.581, and more preferably greater than or equal to 0.645.

(172) In one example, the MTO nvPM emissions index ratio of the gas turbine engine 10 may be in the range of 0.516 to 0.776, and preferably in the range of 0.581 to 0.711, and more preferably in the range of 0.645 to 0.646.

(173) In other examples, the MTO nvPM emissions index ratio may be 0.5, 0.51, 0.52, 0.53, 0.54, 0.55, 0.56, 0.57, 0.58, 0.59, 0.6, 0.61, 0.62, 0.63, 0.64, 0.65, 0.66, 0.67, 0.68, 0.69, 0.7, 0.71, 0.72, 0.73, 0.74, 0.75, 0.76, 0.77, 0.78, 0.79, 0.8, or within any range defined between any two of these values.

(174) To evaluate the value of the MTO nvPM emissions index ratio a value for EI.sub.maxTO,SAF may be determined for operation at 100% available thrust using a fuel comprising SAF. The value of EI.sub.maxTO,FF may be determined for corresponding operation at the same given operating conditions except if the gas turbine engine were instead operated using a fossil-based hydrocarbon fuel.

(175) A climb nvPM emissions index ratio is defined as:

(176) $\begin{matrix} \frac{{EI}_{climb, SAF}}{{EI}_{climb, FF}} & (10) \end{matrix}$

(177) Where EI.sub.climb,SAF and EI.sub.climb,FF are as defined above. The climb nvPM emissions index ratio represents a ratio of the nvPM emissions index corresponding to operating the gas turbine engine 10 during the climb operating phase using a fuel comprising SAF to if the gas turbine engine 10 was instead operated using a fossil-based hydrocarbon fuel.

(178) Additionally, or alternatively, in any example defined or claimed herein, the climb nvPM emissions index ratio of the gas turbine engine 10 may be less than 1.

(179) The climb nvPM emissions index ratio of the gas turbine engine 10 may be less than or equal to 0.9, and preferably less than or equal to 0.75, and more preferably less than or equal to 0.6.

(180) The climb nvPM emissions index ratio of the gas turbine engine 10 may be less than or equal to 0.57, and preferably less than or equal to 0.523, and more preferably less than or equal to 0.475.

(181) More generally, the climb nvPM emissions index ratio of the gas turbine engine 10 may be less than 1, 0.95, 0.9, 0.85, 0.8, 0.75, 0.7, 0.65, 0.6, 0.55, 0.5, 0.45, 0.4, 0.35, 0.3, 0.25, 0.2, 0.15, 0.1, or 0.05, or any range defined between any two of these values.

(182) In any of the examples defined above in which only an upper bound for the climb nvPM emissions index ratio is defined, the lower bound may be greater than zero.

(183) In any of the examples above, the climb nvPM emissions index ratio of the gas turbine engine 10 may be greater than or equal to 0.1, and preferably greater than or equal to 0.2, and more preferably greater than or equal to 0.3.

(184) The climb nvPM emissions index ratio of the gas turbine engine 10 may be greater than or equal to 0.379, and preferably greater than or equal to 0.427, and more preferably greater than or equal to 0.474.

(185) In one example, the climb nvPM emissions index ratio of the gas turbine engine 10 may be in the range of 0.379 to 0.570, and preferably in the range of 0.427 to 0.523, and more preferably in the range of 0.474 to 0.475.

(186) In other examples, the climb nvPM emissions index ratio may be 0.4, 0.41, 0.42, 0.43, 0.44, 0.45, 0.46, 0.47, 0.48, 0.49, 0.5, 0.51, 0.52, 0.53, 0.54, 0.55, 0.56, 0.57, 0.58, 0.59, 0.6, or within any range defined between any two of these values.

(187) An approach nvPM emissions index ratio is defined as:

(188) $\begin{matrix} \frac{{EI}_{approach, SAF}}{{EI}_{approach, FF}} & (11) \end{matrix}$

(189) Where EI.sub.approach,SAF and EI.sub.approach,FF are as defined above. The approach nvPM emissions index ratio represents a ratio of the nvPM emissions index corresponding to operating the gas turbine engine 10 during the approach operating phase using a fuel comprising SAF to if the gas turbine engine 10 was instead operated using a fossil-based hydrocarbon fuel.

(190) Additionally, or alternatively, in any example defined or claimed herein, the approach nvPM emissions index ratio of the gas turbine engine 10 may be less than 1.

(191) The approach nvPM emissions index ratio of the gas turbine engine 10 may be less than or equal to 0.8, and preferably less than or equal to 0.5, and more preferably less than or equal to 0.2.

(192) The approach nvPM emissions index ratio of the gas turbine engine 10 may be less than or equal to 0.185, and preferably less than or equal to 0.169, and more preferably less than or equal to 0.154.

(193) More generally, the approach nvPM emissions index ratio of the gas turbine engine 10 may be less than 1, 0.95, 0.9, 0.85, 0.8, 0.75, 0.7, 0.65, 0.6, 0.55, 0.5, 0.45, 0.4, 0.35, 0.3, 0.25, 0.2, 0.15, 0.1, or 0.05, or any range defined between any two of these values.

(194) In any example defined above in which only an upper bound for the approach nvPM emissions index ratio is defined, the lower bound may be greater than zero.

(195) In any example above, the approach nvPM emissions index ratio of the gas turbine engine 10 may be greater than or equal to 0.03, and preferably greater than or equal to 0.06, and more preferably greater than or equal to 0.09.

(196) The approach nvPM emissions index ratio of the gas turbine engine 10 may be greater than or equal to 0.122, and preferably greater than or equal to 0.138, and more preferably greater than or equal to 0.153.

(197) In one example, the approach nvPM emissions index ratio of the gas turbine engine 10 may be in the range of 0.122 to 0.185, and preferably in the range of 0.138 to 0.169, and more preferably in the range of 0.153 to 0.154.

(198) In other examples, the approach nvPM emissions index ratio may be 0.1, 0.11, 0.12, 0.13, 0.14, 0.15, 0.16, 0.17, 0.17, 0.18, 0.19, 0.2, or within any range defined between any two of these values.

(199) An idle nvPM emissions index ratio is defined as:

(200) $\begin{matrix} \frac{{EI}_{idle, SAF}}{{EI}_{idle, FF}} & (12) \end{matrix}$

(201) Where EI.sub.idle,SAF and EI.sub.idle,FF are as defined above. The idle nvPM emissions index ratio represents a ratio of the nvPM emissions index corresponding to operating the gas turbine engine 10 during the idle operating phase using a fuel comprising SAF to if the gas turbine engine 10 was instead operated using a fossil-based hydrocarbon fuel.

(202) Additionally, or alternatively, in any example defined or claimed herein, the idle nvPM emissions index ratio of the gas turbine engine 10 may be less than 1.

(203) The idle nvPM emissions index ratio of the gas turbine engine 10 may be less than or equal to 0.8, and preferably less than or equal to 0.5, and more preferably less than or equal to 0.2.

(204) The idle nvPM emissions index ratio of the gas turbine engine 10 may be less than or equal to 0.115, and preferably less than or equal to 0.106, and more preferably less than or equal to 0.0959.

(205) More generally, the idle nvPM emissions index ratio may be less than 1, 0.95, 0.9, 0.85, 0.8, 0.75, 0.7, 0.65, 0.6, 0.55, 0.5, 0.45, 0.4, 0.35, 0.3, 0.25, 0.2, 0.15, 0.1, or 0.05, or any range defined between any two of these values.

(206) In any example defined above in which only an upper bound is given for the idle nvPM emissions index ratio, the lower bound may be greater than zero.

(207) The idle nvPM emissions index ratio of the gas turbine engine 10 may be greater than or equal to 0.02, and preferably greater than or equal to 0.04, and more preferably greater than or equal to 0.06.

(208) The idle nvPM emissions index ratio of the gas turbine engine 10 may be greater than or equal to 0.0766, and preferably greater than or equal to 0.0862, and more preferably greater than or equal to 0.0958.

(209) In one example, the idle nvPM emissions index ratio of the gas turbine engine 10 may be in the range of 0.0766 to 0.115, and preferably in the range of 0.0862 to 0.106, and more preferably in the range of 0.0958 to 0.0959.

(210) In some examples, the idle nvPM emissions index ratio of the gas turbine engine 10 may be 0.06, 0.07, 0.08, 0.09, 0.1, 0.11, 0.12, 0.13, 0.14, 0.15, 0.16, 0.17, 0.18, 0.19, 0.2, 0.21, 0.22, 0.23, 0.24, 0.25, 0.26, 0.27, 0.28, 0.29, 0.3, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39, 0.4, or within any range defined between any two of these values.

(211) MTO, Climb, Approach and Idle nvPM Emissions Index Ratio-Modified Fuel Flow

(212) An MTO nvPM emissions index ratio-modified fuel flow is defined as:

(213) $\begin{matrix} \frac{{EI}_{maxTO, SAF}}{{EI}_{maxTO, FF}} W_{f, maxTO} & (13) \end{matrix}$

(214) Where EI.sub.maxTO,SAF and EI.sub.maxTO,FF are as defined above. W.sub.f,maxTO is as defined elsewhere herein i.e. is the mass flow rate of fuel provided to the fuel spray nozzles 403, 404 in kg/s when the gas turbine engine 10 is operating at 100% available thrust for given operating conditions (e.g. during an MTO operating phase). The MTO nvPM emissions index ratio-modified fuel flow represents the fuel flow at MTO operation scaled by the respective nvPM emissions index ratio.

(215) Additionally, or alternatively, in any example defined or claimed herein, the MTO nvPM emissions index ratio-modified fuel flow of the gas turbine engine 10 in kg/s may be less than 2. More specifically, the MTO nvPM emissions index ratio-modified fuel flow of the gas turbine engine 10 in kg/s may be less than 1.23, more preferably less than 1.13 and yet even more preferably less than 1.03.

(216) The MTO nvPM emissions index ratio-modified fuel flow of the gas turbine engine 10 in kg/s may be less than or equal to 0.793, more preferably less than or equal to 0.727 and further preferably less than or equal to 0.661.

(217) In any of the examples in the previous paragraphs where only an upper bound is defined, the MTO nvPM emissions index ratio-modified fuel flow of the gas turbine engine 10 in kg/s may additionally be greater than zero.

(218) In any of the examples above the MTO nvPM emissions index ratio-modified fuel flow of the gas turbine engine 10 in kg/s may be greater than or equal to 0.284, preferably greater than or equal to 0.32 and further preferably greater than or equal to 0.356.

(219) In some examples, the MTO nvPM emissions index ratio-modified fuel flow of the gas turbine engine 10 in kg/s may be in the range in the range 0.284 to 0.793, preferably in the range 0.320 to 0.727 and further preferably in the range 0.356 to 0.661.

(220) In some examples, the MTO nvPM emissions index ratio-modified fuel flow of the gas turbine engine 10 in kg/s may be less than or equal to 0.2, 0.25, 0.3, 0.35, 0.4, 0.45, 0.5, 0.55, 0.6, 0.65, 0.7, 0.75, 0.8, 0.85, 0.9, 0.95, 1, 1.05, 1.1, 1.15, 1.2, 1.25, or 1.3 or within any range defined between any two of these values.

(221) W.sub.f,maxTO may be as defined anywhere else herein. In any of the examples above, W.sub.f,maxTO may be in the range 0.441 to 1.23 kg/s, preferably in the range 0.496 to 1.13 kg/s, more preferably in the range 0.551 to 1.03 kg/s.

(222) A climb nvPM emissions index ratio-modified fuel flow is defined as:

(223) 0 $\begin{matrix} \frac{{EI}_{climb, SAF}}{{EI}_{climb, FF}} W_{f, climb} & (14) \end{matrix}$

(224) Where EI.sub.climb,SAF and EI.sub.climb,FF are as defined above. W.sub.f,climb is as defined elsewhere herein i.e. is the mass flow rate of fuel provided to the fuel spray nozzles 403, 404 in kg/s when the gas turbine engine 10 is operating at 85% available thrust for given operating conditions (e.g. during a climb operating phase). The climb nvPM emissions index ratio-modified fuel flow represents the fuel flow at climb operation scaled by the respective nvPM emissions index ratio.

(225) Additionally, or alternatively, in any example defined or claimed herein, the climb nvPM emissions index ratio-modified fuel flow of the gas turbine engine 10 in kg/s may be less than 2. More specifically, the climb nvPM emissions index ratio-modified fuel flow in kg/s may be less than 1.01, more preferably less than 0.923 and yet even more preferably less than 0.839.

(226) The climb nvPM emissions index ratio-modified fuel flow of the gas turbine engine 10 in kg/s may be less than or equal to 0.479, more preferably less than or equal to 0.439 and further preferably less than or equal to 0.399.

(227) In any of the examples in the previous paragraphs where only an upper bound of the climb nvPM emissions index ratio-modified fuel flow is defined, the lower bound may be greater than zero.

(228) In any example above, the climb nvPM emissions index ratio-modified fuel flow of the gas turbine engine 10 in kg/s may be greater than or equal to 0.175, preferably greater than or equal to 0.197 and further preferably greater than or equal to 0.219.

(229) In some examples, the climb nvPM emissions index ratio-modified fuel flow of the gas turbine engine 10 in kg/s may be in the range 0.175 to 0.479, preferably in the range 0.197 to 0.439 and further preferably in the range 0.219 to 0.399.

(230) In some examples, the climb nvPM emissions index ratio-modified fuel flow of the gas turbine engine 10 in kg/s may be less than or equal to 0.15, 0.175, 0.2, 0.225, 0.25, 0.275, 0.3, 0.325, 0.35, 0.375, 0.4, 0.425, 0.45, 0.475, 0.5, 0.525, 0.55, 0.575, 0.6, 0.625, 0.65, 0.675, 0.7, 0.725, 0.75, 0.775, 0.8, 0.825, 0.85, 0.875, 0.9, 0.925, 0.95, 0.975 or 1, or within any range defined between any two of these values.

(231) The W.sub.f,climb of the gas turbine engine 10 may be as defined anywhere else herein. In any of the examples above, W.sub.f,climb may be in the range 0.369 to 1.01 kg/s, preferably in the range 0.415 to 0.923 kg/s, more preferably in the range 0.461 to 0.839 kg/s.

(232) An approach nvPM emissions index ratio-modified fuel flow is defined as:

(233) $\begin{matrix} \frac{{EI}_{approach, SAF}}{{EI}_{approach, FF}} W_{f, approach} & (15) \end{matrix}$

(234) Where EI.sub.approach,SAF and EI.sub.approach,FF are as defined above. W.sub.f,approach is as defined elsewhere herein i.e. the mass flow rate of fuel provided to the fuel spray nozzles 403, 404 in kg/s when the gas turbine engine 10 is operating at 30% available thrust for given operating conditions (e.g. during an approach operating phase). The approach nvPM emissions index ratio-modified fuel flow represents the fuel flow at approach operation scaled by the respective nvPM emissions index ratio.

(235) Additionally, or alternatively, in any example defined or claimed herein, the approach nvPM emissions index ratio-modified fuel flow of the gas turbine engine 10 in kg/s may be less than 0.4. More specifically, the approach nvPM emissions index ratio-modified fuel flow in kg/s may be less than 0.334, more preferably less than 0.306 and yet even more preferably less than 0.278.

(236) The approach nvPM emissions index ratio-modified fuel flow of the gas turbine engine 10 in kg/s may be less than or equal to 0.0513, more preferably less than or equal to 0.047 and further preferably less than or equal to 0.0428.

(237) In any of the examples in the previous paragraphs where only an upper bound is defined, the approach nvPM emissions index ratio-modified fuel flow of the gas turbine engine 10 in kg/s may additionally be greater than zero.

(238) In any example above, the approach nvPM emissions index ratio-modified fuel flow of the gas turbine engine 10 in kg/s may be greater than or equal to 0.0204, preferably greater than or equal to 0.0229 and further preferably greater than or equal to 0.0255.

(239) In some examples, the approach nvPM emissions index ratio-modified fuel flow of the gas turbine engine 10 in kg/s may be in the range 0.0204 to 0.0513, preferably in the range 0.0229 to 0.0470 and further preferably in the range 0.0255 to 0.0428.

(240) In some examples, the approach nvPM emissions index ratio-modified fuel flow of the gas turbine engine 10 in kg/s may be less than or equal to 0.02, 0.05, 0.075, 0.1, 0.125, 0.15, 0.175, 0.2, 0.225, 0.25, 0.275, 0.3, 0.325, 0.35, 0.375, 0.4, or within any range defined between any two of these values.

(241) The W.sub.f,approach of the gas turbine engine 10 may be as defined anywhere else herein. In any of the examples above, W.sub.f,approach may be in the range 0.133 to 0.334 kg/s, preferably in the range 0.149 to 0.306 kg/s, more preferably in the range 0.166 to 0.278 kg/s.

(242) An idle nvPM emissions index ratio-modified fuel flow is defined as:

(243) $\begin{matrix} \frac{{EI}_{idle, SAF}}{{EI}_{idle, FF}} W_{f, idle} & (16) \end{matrix}$

(244) Where EI.sub.idle,SAF and EI.sub.idle,FF are as defined above. W.sub.f,idle is defined as the mass flow rate of fuel provided to the fuel spray nozzles 403, 404 in kg/s when the gas turbine engine 10 is operating at 7% available thrust for given operating conditions (e.g. during an idle operating phase). The idle nvPM emissions index ratio-modified fuel flow represents the fuel flow at idle operation scaled by the respective nvPM emissions index ratio.

(245) Additionally, or alternatively, in any example defined or claimed herein, the idle nvPM emissions index ratio-modified fuel flow of the gas turbine engine 10 in kg/s may be less than 0.2. More specifically, the idle nvPM emissions index ratio-modified fuel flow in kg/s may be less than 0.119, more preferably less than 0.109 and yet even more preferably less than 0.099.

(246) In some examples, the idle nvPM emissions index ratio-modified fuel flow of the gas turbine engine 10 in kg/s may be less than or equal to 0.0114, more preferably less than or equal to 0.0105 and further preferably less than or equal to 0.00949.

(247) In any of the examples in the previous paragraph, where only an upper bound is defined, the idle nvPM emissions index ratio-modified fuel flow of the gas turbine engine 10 in kg/s may additionally be greater than zero.

(248) In any example above, the idle nvPM emissions index ratio-modified fuel flow of the gas turbine engine 10 in kg/s may be greater than or equal to 0.00494, preferably greater than or equal to 0.00556 and further preferably greater than or equal to 0.00618.

(249) In some examples, the idle nvPM emissions index ratio-modified fuel flow of the gas turbine engine 10 in kg/s may be in the range 0.00494 to 0.0114, preferably in the range 0.00556 to 0.0105 and further preferably in the range 0.00618 to 0.00949.

(250) In some examples, the idle nvPM emissions index ratio-modified fuel flow of the gas turbine engine 10 in kg/s may be less than or equal to 0.004, 0.01, 0.015, 0.02, 0.025, 0.03, 0.035, 0.04, 0.045, 0.05, 0.055, 0.06, 0.065, 0.07, 0.075, 0.08, 0.085, 0.09, 0.095, 0.1, 0.11, or 0.12, or within any range defined between any two of these values.

(251) The W.sub.f,idle of the gas turbine engine 10 may be as defined anywhere else herein. In any of the examples above, W.sub.f,idle may be in the range 0.0516 to 0.119 kg/s, preferably in the range 0.0581 to 0.109 kg/s, more preferably in the range 0.0645 to 0.0990 kg/s.

(252) Rich Cruise, Rich Cruise/MTO, Idle/Rich Cruise, Lean Cruise, Lean Cruise/MTO, and Idle/Lean Cruise nvPM Emissions Index Ratios

(253) A rich cruise nvPM emissions index ratio is defined as:

(254) $\begin{matrix} \frac{{EI}_{cruise (rich), SAF}}{{EI}_{cruise (rich), FF}} & (17) \end{matrix}$ Where EI.sub.cruise(rich),SAF is defined as:

(255) $\begin{matrix} \frac{{EI}_{climb, SAF} + {EI}_{approach, SAF}}{2} & (18) \end{matrix}$ and EI.sub.cruise(rich),FF is defined as:

(256) $\begin{matrix} \frac{{EI}_{climb, FF} + {EI}_{approach, FF}}{2} & (19) \end{matrix}$ EI.sub.cruise(rich),SAF and EI.sub.cruise(rich),FF represent the nvPM emissions index when the gas turbine engine 10 is operating in a rich cruise operating phase and is running on a fuel comprising SAF and a fossil-based hydrocarbon fuel respectively. EI.sub.cruise(rich),SAF is determined by finding the average (mean) of the nvPM emissions indices when the gas turbine engine 10 is operating using a fuel comprising SAF in a climb operating phase (i.e. at 85% available thrust) and when it is operating using a fuel comprising SAF in an approach operating phase (i.e. operating at 30% available thrust). In equations (18) and (19) above, EI.sub.climb,SAF, EI.sub.approach,SAF, EI.sub.climb,FF and EI.sub.approach,FF are as defined elsewhere herein. The rich cruise nvPM emissions index ratio represents a ratio of the emissions index at rich cruise when using a fuel comprising SAF to the emissions index at rich cruise when using a fossil-based hydrocarbon fuel.

(257) Additionally or alternatively, in any example defined or claimed herein, the rich cruise nvPM emissions index ratio of the gas turbine engine 10 may be less than 1. More specifically, the rich cruise nvPM emissions index ratio of the gas turbine engine 10 may be less than or equal to 0.9, preferably less than or equal to 0.8, even preferably less than or equal to 0.7.

(258) In some examples, the rich cruise nvPM emissions index ratio of the gas turbine engine 10 may be less than or equal to 0.564, preferably less than or equal to 0.517 and further preferably less than or equal to 0.47.

(259) More generally, in some examples, the rich cruise nvPM emissions index ratio of the gas turbine engine 10 may be less than 1, 0.95, 0.9, 0.85, 0.8, 0.75, 0.7, 0.65, 0.6, 0.55, 0.5, 0.45, 0.4, 0.35, 0.3, 0.25, 0.2, 0.15, 0.1, or 0.05, or within any range defined between any two of these values.

(260) In any of the examples of the previous paragraphs, where only an upper bound is defined, the rich cruise nvPM emissions index ratio may be greater than zero.

(261) In any example above, the rich cruise nvPM emissions index ratio of the gas turbine engine 10 may be greater than or equal to 0.3, preferably greater than or equal to 0.337 and further preferably greater than or equal to 0.375.

(262) In some examples, the rich cruise nvPM emissions index ratio of the gas turbine engine 10 may be in the range 0.300 to 0.564, preferably in the range 0.337 to 0.517 and further preferably in the range 0.375 to 0.470.

(263) In some examples, the rich cruise nvPM emissions index ratio of the gas turbine engine 10 may be 0.3, 0.32, 0.34, 0.36, 0.38, 0.4, 0.42, 0.44, 0.46, 0.48, 0.5, 0.52, 0.54, 0.56, 0.58, 0.6, or within any range defined between any two of these values.

(264) A rich cruise/MTO nvPM emissions index ratio is defined as:

(265) $\begin{matrix} \frac{{EI}_{cruise (rich) . SAF} / {EI}_{maxTO, SAF}}{{EI}_{cruise (rich), FF} / {EI}_{maxTO, FF}} & (20) \end{matrix}$ where EI.sub.cruise(rich),SAF and EI.sub.cruise(rich),FF are as defined earlier in this section, and EI.sub.maxTO,SAF and EI.sub.maxTO,FF are as defined elsewhere herein. The rich cruise/MTO nvPM emissions index ratio represents a ratio of the nvPM emissions index using a fuel comprising SAF at rich cruise divided by that at MTO to the nvPM emissions index using fossil-based fuel at rich cruise divided by that at MTO.

(266) Additionally, or alternatively, in any example defined or claimed herein, the rich cruise/MTO nvPM emissions index ratio of the gas turbine engine 10 may be less than 1. More specifically, the rich cruise/MTO nvPM emissions index ratio of the gas turbine engine 10 may be less than or equal to 0.95, preferably less than or equal to 0.9 and even more preferably less than or equal to 0.875.

(267) In some examples, the rich cruise/MTO nvPM emissions index ratio of the gas turbine engine 10 may be less than or equal to 0.873, preferably less than or equal to 0.8 and further preferably less than or equal to 0.728.

(268) More generally, the rich cruise/MTO nvPM emissions index ratio of the gas turbine engine 10 may be less than 1, 0.95, 0.9, 0.85, 0.8, 0.75, 0.7, 0.65, 0.6, 0.55, 0.5, 0.45, 0.4, 0.35, 0.3, 0.25, 0.2, 0.15, 0.1, or 0.05, or any range defined between any two of these values.

(269) In any of the examples in the previous paragraphs in which only an upper bound is defined, the rich cruise/MTO nvPM emissions index ratio may be greater than zero.

(270) In any example above, the rich cruise/MTO nvPM emissions index ratio of the gas turbine engine 10 may be greater than or equal to 0.464, preferably greater than or equal to 0.523 and further preferably greater than or equal to 0.581.

(271) In some examples, the rich cruise/MTO nvPM emissions index ratio of the gas turbine engine 10 may be in the range 0.464 to 0.873, preferably in the range 0.523 to 0.800 and further preferably in the range 0.581 to 0.728.

(272) In some examples, the rich cruise/MTO nvPM emissions index ratio of the gas turbine engine 10 may be 0.45, 0.475, 0.5, 0.525, 0.55, 0.575, 0.6, 0.625, 0.65, 0.675, 0.7, 0.725, 0.75, 0.775, 0.8, 0.825, 0.85, or 0.875, or within any range defined between any two of these values.

(273) An idle/rich cruise nvPM emissions index ratio is defined as:

(274) $\begin{matrix} \frac{{EI}_{idle, SAF} / {EI}_{cruise (rich), SAF}}{{EI}_{idle, FF} / {EI}_{cruise (rich), FF}} & (21) \end{matrix}$ where EI.sub.cruise(rich),SAF and EI.sub.cruise(rich),FF are as defined earlier in this section, and EI.sub.idle,SAF and EI.sub.idle,FF are as defined elsewhere herein. The idle/rich cruise nvPM emissions index ratio represents a ratio of the nvPM emissions index using a fuel comprising SAF at idle divided by that at rich cruise to the nvPM emissions index using fossil-based fuel at idle divided by that at rich cruise.

(275) Additionally, or alternatively, in any example defined or claimed herein, the idle/rich cruise nvPM emissions index ratio of the gas turbine engine 10 may be less than 1. More specifically, the idle/rich cruise nvPM emissions index ratio of the gas turbine engine 10 may be less than or equal to 0.8 and preferably less than or equal to 0.6 and more preferably less than or equal to 0.4.

(276) In some examples, the idle/rich cruise nvPM emissions index ratio of the gas turbine engine 10 may be less than or equal to 0.307, preferably less than or equal to 0.281 and further preferably less than or equal to 0.256.

(277) More generally, the idle/rich cruise nvPM emissions index ratio of the gas turbine engine 10 may be less than 1, 0.95, 0.9, 0.85, 0.8, 0.75, 0.7, 0.65, 0.6, 0.55, 0.5, 0.45, 0.4, 0.35, 0.3, 0.25, 0.2, 0.15, 0.1, or 0.05, or any range defined between any two of these values.

(278) In any of the examples in the previous paragraphs in which only an upper bound is defined, the idle/rich cruise nvPM emissions index ratio of the gas turbine engine 10 may be greater than zero.

(279) In any example above, the idle/rich cruise nvPM emissions index ratio of the gas turbine engine 10 may be greater than or equal to 0.163, preferably greater than or equal to 0.183 and further preferably greater than or equal to 0.203.

(280) In some examples, the idle/rich cruise nvPM emissions index ratio of the gas turbine engine 10 may be in the range 0.163 to 0.307, preferably in the range 0.183 to 0.281 and further preferably in the range 0.203 to 0.256.

(281) In some examples, the idle/rich cruise nvPM emissions index ratio of the gas turbine engine 10 may be 0.16, 0.165, 0.17, 0.175, 0.18, 0.185, 0.19, 0.195, 0.2, 0.205, 0.21, 0.215, 0.22, 0.225, 0.23, 0.235, 0.24, 0.245, 0.25, 0.255, 0.26, 0.265, 0.27, 0.275, 0.28, 0.285, 0.29, 0.295, 0.3, 0.305, or 0.31, or within any range defined between any two of these values.

(282) A lean cruise nvPM emissions index ratio is defined as:

(283) $\begin{matrix} \frac{{EI}_{cruise (lean), SAF}}{{EI}_{cruise (lean), FF}} & (22) \end{matrix}$ Where EI.sub.cruise(lean),SAF is defined as:

(284) $\begin{matrix} \frac{{EI}_{maxTO, SAF} + {EI}_{climb, SAF}}{2} & (23) \end{matrix}$ and EI.sub.cruise(lean),FF is defined as:

(285) 0 $\begin{matrix} \frac{{EI}_{maxTO, FF} + {EI}_{climb, FF}}{2} & (24) \end{matrix}$ EI.sub.cruise(lean),SAF and EI.sub.cruise(lean),FF represent the nvPM emissions index when the gas turbine engine 10 is operating in a lean cruise operating phase and is running on a fuel comprising SAF and a fossil-based hydrocarbon fuel respectively. EI.sub.cruise(lean),SAF is determined by finding the average (mean) of the nvPM emissions indices when the gas turbine engine 10 is operating using a fuel comprising SAF in a max take off operating phase (i.e. at 100% available thrust) and when it is operating using a fuel comprising SAF in a climb operating phase (i.e. operating at 85% available thrust). In equations (23) and (24) above, EI.sub.maxTO,SAF, EI.sub.climb,SAF, EI.sub.maxTO,FF and EI.sub.climb,FF are as defined elsewhere herein. The lean cruise nvPM emissions index ratio represents a ratio of the emissions index at lean cruise when using a fuel comprising SAF to the emissions index at lean cruise when using a fossil-based hydrocarbon fuel.

(286) Additionally, or alternatively, in any example defined or claimed herein, the lean cruise nvPM emissions index ratio of the gas turbine engine 10 may be less than 1. More specifically the lean cruise nvPM emissions index ratio of the gas turbine engine 10 may be less than or equal to 0.9, preferably less than or equal to 0.8 and further preferably less than or equal to 0.75.

(287) In some examples, the lean cruise nvPM emissions index ratio of the gas turbine engine 10 may be less than or equal to 0.709, preferably less than or equal to 0.65 and further preferably less than or equal to 0.591.

(288) More generally, in some examples, the lean cruise nvPM emissions index ratio may be less than 1, 0.95, 0.9, 0.85, 0.8, 0.75, 0.7, 0.65, 0.6, 0.55, 0.5, 0.45, 0.4, 0.35, 0.3, 0.25, 0.2, 0.15, 0.1, or 0.05, or in any range defined between any two of these values.

(289) In any of the examples in the previous paragraphs where only an upper bound is defined, the lean cruise nvPM emissions index ratio of the gas turbine engine 10 may be greater than zero.

(290) In any example above the lean cruise nvPM emissions index ratio of the gas turbine engine 10 may be greater than or equal to 0.451, preferably greater than or equal to 0.507 and further preferably greater than or equal to 0.563.

(291) In some examples, the lean cruise nvPM emissions index ratio of the gas turbine engine 10 may be in the range 0.451 to 0.709, preferably in the range 0.507 to 0.650 and further preferably in the range 0.563 to 0.591.

(292) In some examples, the lean cruise nvPM emissions index ratio of the gas turbine engine 10 may be 0.45, 0.455, 0.46, 0.47, 0.48, 0.49, 0.5, 0.51, 0.52, 0.53, 0.54, 0.55, 0.56, 0.57, 0.58, 0.59, 0.6, 0.61, 0.62, 0.63, 0.64, 0.65, 0.66, 0.67, 0.68, 0.69, 0.7, or 0.71, or within any range defined between any two of these values.

(293) A lean cruise/MTO nvPM emissions index ratio is defined as:

(294) $\begin{matrix} \frac{{EI}_{cruise (lean), SAF} / {EI}_{maxTO, SAF}}{{EI}_{cruise (lean), FF} / {EI}_{maxTO, FF}} & (25) \end{matrix}$ where EI.sub.cruise(lean),SAF and EI.sub.cruise(lean),FF are as defined earlier in this section, and EI.sub.maxTO,SAF and EI.sub.maxTO,FF are as defined elsewhere herein. The lean cruise/MTO nvPM emissions index ratio represents a ratio of the nvPM emissions index ratio using a fuel comprising SAF at lean cruise divided by that at MTO to the nvPM emissions index ratio using fossil-based fuel at lean cruise divided by that at MTO.

(295) Additionally, or alternatively, in any example defined or claimed herein, the lean cruise/MTO nvPM emissions index ratio of the gas turbine engine 10 may be less than 1. More specifically, the lean cruise/MTO nvPM emissions index ratio of the gas turbine engine 10 may be less than or equal to 0.98, preferably less than or equal to 0.95 and further preferably less than or equal to 0.92. In some examples, the lean cruise/MTO nvPM emissions index ratio of the gas turbine engine 10 may be less than or equal to 0.914.

(296) More generally, in some examples, the lean cruise/MTO nvPM emissions index ratio may be less than 1, 0.95, 0.9, 0.85, 0.8, 0.75, 0.7, 0.65, 0.6, 0.55, 0.5, 0.45, 0.4, 0.35, 0.3, 0.25, 0.2, 0.15, 0.1, or 0.05, or any range defined between any two of these values.

(297) In any of the examples in the previous paragraphs, where only an upper bound is defined, the lean cruise/MTO nvPM emissions index ratio may be greater than zero.

(298) In any example above, the lean cruise/MTO nvPM emissions index ratio of the gas turbine engine 10 may be greater than or equal to 0.698, preferably greater than or equal to 0.785 and further preferably greater than or equal to 0.873.

(299) In some examples, the lean cruise/MTO nvPM emissions index ratio of the gas turbine engine 10 may be in the range 0.698 to 0.980, preferably in the range 0.785 to 0.950 and further preferably in the range 0.873 to 0.914.

(300) In some examples, the lean cruise/MTO nvPM emissions index ratio of the gas turbine engine 10 may be 0.69, 0.71, 0.73, 0.75, 0.77, 0.79, 0.81, 0.83, 0.85, 0.87, 0.89, 0.91, 0.93, 0.95, 0.97, or 0.99, or within any range defined between any two of these values.

(301) An idle/lean cruise nvPM emissions index ratio is defined as:

(302) $\begin{matrix} \frac{{EI}_{idle, SAF} / {EI}_{cruise (lean), SAF}}{{EI}_{idle, FF} / {EI}_{cruise (lean), FF}} & (26) \end{matrix}$ where EI.sub.cruise(lean),SAF and EI.sub.cruise(lean),FF are as defined earlier in this section, and EI.sub.idle,SAF and EI.sub.idle,FF are as defined elsewhere herein. The idle/lean cruise nvPM emissions index ratio represents a ratio of the nvPM emissions index ratio using a fuel comprising SAF at idle divided by that at lean cruise to the nvPM emissions index ratio using fossil-based fuel at idle divided by that at lean cruise.

(303) Additionally, or alternatively, in any example defined or claimed herein, the idle/lean cruise nvPM emissions index ratio of the gas turbine engine 10 may be less than 1. More specifically, the idle/lean cruise nvPM emissions index ratio of the gas turbine engine 10 may be less than or equal to 0.8, preferably less than or equal to 0.6, even preferably less than or equal to 0.4, and even further preferably less than or equal to 0.3.

(304) In some examples, the idle/lean cruise nvPM emissions index ratio of the gas turbine engine 10 may be less than or equal to 0.204, preferably less than or equal to 0.187 and further preferably less than or equal to 0.17.

(305) More generally, the idle/lean cruise nvPM emissions index ratio may be less than 1, 0.95, 0.9, 0.85, 0.8, 0.75, 0.7, 0.65, 0.6, 0.55, 0.5, 0.45, 0.4, 0.35, 0.3, 0.25, 0.2, 0.15, 0.1, or 0.05, or any range defined between any two of these values.

(306) In any of the examples in the paragraphs above, where only an upper bound is defined, the idle/lean cruise nvPM emissions index ratio may be greater than zero.

(307) In any example above, the idle/lean cruise nvPM emissions index ratio of the gas turbine engine 10 may be greater than or equal to 0.129, preferably greater than or equal to 0.146 and further preferably greater than or equal to 0.162.

(308) In some examples, the idle/lean cruise nvPM emissions index ratio of the gas turbine engine 10 may be in the range in the range 0.129 to 0.204, preferably in the range 0.146 to 0.187 and further preferably in the range 0.162 to 0.170.

(309) In some examples, the idle/lean cruise nvPM emissions index ratio may be 0.12, 0.125, 0.13, 0.135, 0.14, 0.145, 0.15, 0.155, 0.16, 0.165, 0.17, 0.175, 0.18, 0.185, 0.19, 0.195, 0.2, 0.205, or 0.21, or within any range defined between any two of these values.

(310) Emission Indices

(311) In any of the examples defined or claimed anywhere herein, any one or more of the emission indices of the gas turbine engine 10 may be as follows: EI.sub.maxTO may be in the range 4.96 to 145 mg/kg and preferably in the range 5.58 to 133 mg/kg and more preferably in the range 6.21 to 121 mg/kg. EI.sub.maxTO,SAF may be in the range of 4.96 to 93.3 mg/kg and preferably in the range of 5.58 to 85.5 mg/kg and more preferably in the range of 6.21 to 77.8 mg/kg. EI.sub.maxTO,SAF may be in the range of 4.96 to 144 mg/kg and preferably in the range of 5.58 to 132 mg/kg and more preferably in the range of 6.21 to 120 mg/kg. EI.sub.maxTO,FF may be in the range of 7.69 to 145 mg/kg and preferably in the range of 8.65 to 133 mg/kg and more preferably in the range of 9.61 to 121 mg/kg. EI.sub.climb may be in the range 1.82 to 124 mg/kg and preferably in the range 2.05 to 114 mg/kg and more preferably in the range 2.28 to 103 mg/kg. EI.sub.climb,SAF may be in the range of 1.82 to 58.6 mg/kg and preferably in the range of 2.05 to 53.7 mg/kg and more preferably in the range of 2.28 to 48.9 mg/kg. EI.sub.climb,SAF may be in the range of 1.82 to 123 mg/kg and preferably in the range of 2.05 to 113 mg/kg and more preferably in the range of 2.28 to 102 mg/kg. EI.sub.climb,FF may be in the range of 3.84 to 124 mg/kg and preferably in the range of 4.32 to 114 mg/kg and more preferably in the range of 4.80 to 103 mg/kg. EI.sub.approach may be in the range of 0.0328 to 17.6 mg/kg and preferably in the range of 0.0369 to 16.2 mg/kg and more preferably in the range of 0.0410 to 14.7 mg/kg. EI.sub.approach,SAF may be in the range of 0.0328 to 2.70 mg/kg and preferably in the range of 0.0369 to 2.48 mg/kg and more preferably in the range of 0.0410 to 2.25 mg/kg. EI.sub.approach,SAF may be in the range of 0.0328 to 17.5 mg/kg and preferably in the range of 0.0369 to 16.1 mg/kg and more preferably in the range of 0.0410 to 14.6 mg/kg. EI.sub.approach,FF may be in the range of 0.213 to 17.6 mg/kg and preferably in the range of 0.240 to 16.2 mg/kg and more preferably in the range of 0.267 to 14.7 mg/kg. EI.sub.idle may be in the range 0.118 to 41.4 mg/kg and preferably in the range 0.132 to 38.0 mg/kg and more preferably in the range 0.147 to 34.5 mg/kg. EI.sub.idle,SAF may be in the range of 0.118 to 3.97 mg/kg and preferably in the range of 0.132 to 3.64 mg/kg and more preferably in the range of 0.147 to 3.31 mg/kg. EI.sub.idle,SAF may be in the range of 0.118 to 41.3 mg/kg and preferably in the range of 0.132 to 37.9 mg/kg and more preferably in the range of 0.147 to 34.4 mg/kg. EI.sub.idle,FF may be in the range of 1.23 to 41.4 mg/kg and preferably in the range of 1.38 to 38.0 mg/kg and more preferably in the range of 1.54 to 34.5 mg/kg.
Method of Operating a Gas Turbine Engine

(312) FIG. 12 illustrates a method 1000 of operating the gas turbine engine 10 of any example or aspect defined or claimed herein. The method comprises providing 1002 fuel comprising a sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles 403, 404.

CONCLUSION

(313) Anything described in this section may apply to any aspect or example described or claimed anywhere herein.

(314) For any example gas turbine engine defined herein, the ratio defined in any one or more of expressions 1 to 26 may be as defined or claimed anywhere herein. In other words, the gas turbine engine 10 may be configured such than one or more of the ratios defined herein are within the ranges defined herein.

(315) Any reference to a ratio (or other parameter) of the gas turbine engine being within a specific range should be understood to mean that the gas turbine is configured such that or configured such that, in use the respective ratio or parameter is within the range. In other words, a reference to a ratio or parameter of the gas turbine engine being within a specified range should be understood to mean that the gas turbine engine is arranged such that the respective parameter or ratio is within that range when the gas turbine engine is in use.

(316) Any of the parameters defined herein may be determined at suitable given operating conditions. For example, the given operating conditions at which the emissions indices defined herein are determined may be ISA at sea level except that the reference absolute humidity shall be 0.00634 kg water/kg dry air. The predetermined operating conditions may be at sea level static. The predetermined operating conditions may include no customer bleeds and/or no power offtakes. The predetermined operating conditions may be at day conditions. The predetermined operating conditions may be at around 60% relative humidity. The same given operating conditions may be used to evaluate any other parameter defined herein, such as the BPR.

(317) The emissions indices may however be evaluated at other operating conditions. For example, other different operating conditions may be used so long as the same operating conditions are used for all the parameters within a respective ratio.

(318) The emission indices defined herein may be determined using any suitable method as would be known to the skilled person. For example, the procedure for calculating the emission indices defined herein may comprise plotting curves of the nvPM emissions index and NOx emissions index against T3. The known T3 at the 4 LTO reference points (7% thrust, 100% thrust, 85% thrust and 30% thrust) is then used to find the respective emissions index. T3 is defined using the station numbering listed in standard SAE AS755, i.e. T3=high pressure compressor outlet total temperature.

(319) Any reference herein to operation at 7% thrust may more generally be considered to be operation at idle. Any reference herein to operation at 100% thrust may more generally be considered to be operation at max take off. Any reference herein to operation at 85% thrust may more generally be considered to be operation at a climb operating phase. Any reference herein to operation at 30% thrust may more generally be considered to be operation at an approach operating phase. These operating phases may be as defined elsewhere herein.

(320) Any reference to a percentage available thrust given herein should be taken to mean at approximately or at around the specified thrust. For example, by 7% available thrust used anywhere herein we mean around 7% available thrust. Similarly, by 100% available thrust used anywhere herein we mean around 100% available thrust. By 85% available thrust used anywhere herein we mean around 85% available thrust. By 30% available thrust used anywhere herein we mean around 30% available thrust. By around used when specifying a thrust of XX we may mean XX15%, XX10%, XX5%, or XX2%. For example, by around 7% available thrust we may mean 2% to 12% available thrust or 5% to 9% available thrust. For example, by around 100% available thrust we may mean 90% to 100% available thrust or 95% to 100% available thrust or 98% to 100% available thrust. For example, by around 30% available thrust we may mean 20% to 40% available thrust or 25% to 35% available thrust or 28% to 32% available thrust. For example, by around 85% available thrust we may mean 70% to 100% available thrust or 75% to 95% available thrust or 80% to 90% available thrust or 83% to 87% available thrust.

(321) Any of the ranges defined herein should be understood as an inclusive range i.e. in the range A and B or in the range A to B should include the upper and lower boundaries A and B. Where a ratio is given as a single number, e.g., 0.5, this refers to the ratio of the given single number to 1, i.e., 0.5 is to be read as 0.5:1.

(322) Where a number is quoted to one significant figure, e.g., 0.5, this may refer to the same number when quoted to two significant figures where the second significant figure is 0, e.g., 0.50, or when quoted to three significant figures where the third significant figure is 0, e.g., 0.500. Where a number is quoted to two significant figures, e.g., 0.50, this may refer to the same number when quoted to three significant figures where the third significant figure is 0, e.g., 0.500. Where a number is quoted to three significant figures, e.g., 0.500, this may refer to the same number when quoted to two significant figures, e.g., 0.50, or one significant figure, e.g., 0.5. Where a number is quoted to two significant figures, e.g., 0.50, this may refer to the same number when quoted to one significant figure, e.g., 0.5.

(323) Advantageously, reduced nvPM in the exhaust of a gas turbine engine contributes to a reduction in undesirable emissions of the engine. For example, according to operational conditions, reducing nvPM in such a manner may lead to a reduced degree of soot deposits within the engine within and/or downstream of the combustor, and/or an improvement in local air quality. Furthermore, at certain stages of an aircraft flight (where contrails are otherwise expected to form) reduced nvPM in the exhaust may lead to reduced contrail strength and/or time taken for a contrail to disperse. Still further, it has been recognised that certain parts of the flight cycle at which the nvPM is reduced (or most reduced) can be targeted in order to achieve a desired outcome, for example in terms of environmental impact. Purely by way of example, lower nvPM at cruise conditions may particularly reduce the radiative forcing impact of contrails. Purely by way of further example, lower nvPM at idle conditions may particularly improve local air quality on the ground in the region of engine operation. Purely by way of further example, lower nvPM at MTO conditions may particularly reduce the maximum rate of nvPM production during the flight cycle and/or improve air quality on the ground and/or in the region of engine operation. These considerations may apply to all aspects of the disclosure.

(324) A number of parameters related to gas turbine engine operation have been determined to have an influence on, or are an important factor in, the configuration and arrangement of the combustor of the engine when certain types of fuel, such as a sustainable aviation fuel, are being combusted. Accordingly, any one or more parameters of the aspects disclosed or described above may be advantageously taken into account when determining, for example, operational settings, combustor arrangement and/or combustor configuration, to influence and/or optimise how that fuel is to be distributed, ignited, and/or combusted within the gas turbine engine. These considerations may apply to all aspects of the disclosure.

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

Emissions of non-volatile particulate matter from gas turbine engines combusting sustainable aviation fuel and fossil-based hydrocarbon fuel

Assignee

Inventors

Cpc classification

Classification Explorer

F05D2240/35

MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING

Classification Explorer

F02C9/40

MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING

Classification Explorer

F02C7/228

MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING

Classification Explorer

F05D2220/75

MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING

Classification Explorer

F02C7/22

MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING

International classification

Classification Explorer

F02C7/22

MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING

Abstract

Claims

Description