Gas turbine emissions

Abstract

Gas turbine engine (10) for an aircraft. The gas turbine engine includes a combustor device (16), comprising a combustion chamber (120) and a plurality of fuel spray nozzles (124) designed to inject fuel into the combustion chamber (120), wherein the plurality of fuel spray nozzles (124) comprises a first subset (124A) of fuel spray nozzles (124) and a second subset (124B) of fuel spray nozzles (124), wherein the combustor device (16) can operate under a condition in which each of the fuel spray nozzles of the first subset (124A) of fuel spray nozzles (124) is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset (124B) of fuel spray nozzles (124), wherein a ratio of the number of nozzlesof fuel spray nozzles (124) in the first subset (124A) of fuel spray nozzles (124) to the number of fuel spray nozzles (124) in the second subset (124B) of fuel spray nozzles (124) is in the range of 1:2 to 1:5. A depleted cruise nvPM-PMD emission index ratio is defined as: where: EIcruise (depleted) is defined as: EImaxTO is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine (10) when operating at approximately 100% of available thrust under given operating conditions; EImontée is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine (10) when operating at approximately 85% of available thrust under given operating conditions; and BPR is the bypass ratio of the gas turbine engine (10). The index ratioThe nvPM emissions during depleted-PMD cruise are less than 0.2. The gas turbine engine (10) is designed to supply fuel comprising sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles (124). The invention also discloses a method of operating the gas turbine engine. Figure for the abstract: Figure 6

Description

Title of the invention: Gas turbine emissions

[0001] SCOPE OF APPLICATION

[0002] This disclosure relates to emissions of non-volatile particulate matter (nvPM) from gas turbine engines, specifically aircraft gas turbine engines. This disclosure provides various methods of operating a gas turbine engine and gas turbine engines. Certain aspects of this application relate to methods of operating gas turbine engines using a fuel that includes a sustainable aviation fuel (SAF) and gas turbine engines designed to operate using a fuel that includes an SAF.

[0003] CONTEXT

[0004] There is an expectation in the aeronautical industry regarding a trend towards the use of fuels other than the traditional kerosene-based jet fuels generally used at present.

[0005] The inventors have observed that the emissions of a gas turbine engine are sensitive to the fuel used; in particular, the amount of nvPM produced by the engine varies depending on the operating parameters and the type of fuel used. Therefore, there is a need to take into account the fuel properties of these different fuels and to adjust the operating processes of gas turbine engines accordingly.

[0006] SUMMARY

[0007] According to a first aspect, a gas turbine engine for an aircraft is provided, comprising any one or more of the following features:

[0008] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of spray nozzles of fuel in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and

[0009] in which:

[0010] A first nvPM emission index ratio at idle-PMD can be defined as:

[0011] [Math.l] pj 1 Elmax TO slow motion

[0012] where:

[0013] Æ' / raienti is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine if it operates at approximately 7% of the available thrust under given operating conditions; and

[0014] S / maxTo is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine if it operates at approximately 100% of the available thrust under the given operating conditions;

[0015] the first NVPM-PMD emission index ratio of the gas turbine engine at idle is less than 60; and

[0016] The gas turbine engine is designed to supply fuel including sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles.

[0017] Advantageously, reduced nvPMs in the exhaust of a gas turbine engine contribute to a reduction in undesirable engine emissions. For example, depending on operating conditions, reducing nvPMs in this way can 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 aircraft flight (where contrails are expected to form), reduced nvPMs in the exhaust can lead to reduced condensation drag force and / or a longer time for a condensation trail to disperse. Moreover, it has been recognized that certain parts of the flight cycle during which nvPMs are reduced (or most reduced) can be targeted to achieve a desired outcome, for example, in terms of environmental impact.Strictly as an example, lower nvPMs under cruise conditions can significantly reduce the impact of radiative forcing from contrails. Strictly as a further example, lower nvPMs under idle conditions can significantly improve local ground air quality in the region where the engine is operating. Strictly as a further example, lower nvPMs under PMD conditions can significantly reduce the peak nvPM production rate during the [period of time missing]. flight cycle and / or improve air quality on the ground and / or in the area where the engine operates. These considerations may apply to all aspects of disclosure.

[0018] It has been determined that a number of parameters related to the operation of a gas turbine engine influence, or are an important factor in, the configuration and arrangement of the engine's combustor device when certain types of fuel, such as sustainable aviation fuel, are consumed. Accordingly, any one or more parameters of the following aspects may be advantageously taken into account when determining, for example, operating settings, combustor device arrangement, and / or combustor device configuration, to influence and / or optimize how fuel is delivered, ignited, and / or consumed within the gas turbine engine. These considerations may apply to all aspects of the disclosure.

[0019] The first NvPM-PMD emission index ratio of the gas turbine engine at idle may be greater than zero.

[0020] The first nvPM-PMD emission index ratio at idle may be less than 58.4 and preferably may be less than 53.5 and more preferably may be less than 48.6.

[0021] The first nvPM-PMD emission index ratio at idle may be less than or equal to 45 and preferably may be less than or equal to 30 and more preferably may be less than or equal to 15.

[0022] The first nvPM-PMD emission index ratio at idle may be less than or equal to 8.65 and preferably may be less than or equal to 7.93 and more preferably may be less than or equal to 7.21.

[0023] The first nvPM-PMD emission index ratio at idle may be less than or equal to 0.155 and preferably may be less than or equal to 0.142 and more preferably may be less than or equal to 0.129.

[0024] The first nvPM-PMD emission index ratio can be greater than or equal to 0.038 and preferably can be greater than or equal to 0.0428 and more preferably can be greater than or equal to 0.0475.

[0025] The first nvPM-PMD emission index ratio can be greater than or equal to 1.21 and preferably can be greater than or equal to 1.36 and more preferably can be greater than or equal to 1.52.

[0026] The first NvPM-PMD emission index ratio can be in the range of 0.0380 to 8.65 and preferably can be in the range of 0.0428 to 7.93 and more preferably can be in the range of 0.0475 to 7.21.

[0027] The first NvPM-PMD emission index ratio can be in the range of 0.0380 to 0.155 and preferably can be in the range of 0.0428 to 0.142 and more preferably can be in the range of 0.0475 to 0.129.

[0028] The first NvPM-PMD emission index ratio can be in the range of 1.21 to 8.65 and preferably can be in the range of 1.36 to 7.93 and more preferably can be in the range of 1.52 to 7.21.

[0029] The first NVPM emission index ratio at idle-PMD 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, 0.2, 2, 4, 6, 8, 10, 12, 14, 16, 18, 20, 22, 24, 26, 28, 30, 32, 34, 36, 38, 40, 42, 44, 46, 48, 50, 52, 54, 56, 58, 60, or any defined range between any two of these values. For example, the first NvPM-PMD emission index ratio could be in a range between 0.01 and 0.2, 0.01 and 0.15, 0.01 and 0.07, or 0.01 and 0.05.

[0030] A second nvPM emission index ratio at idle-PMD can be defined as:

[0031] [Math.2] EIraieati SAP / El^rc, SAF *-1 slow motion, FF / PI / CiniaxTO, FF

[0032] where:

[0033] Ê / raienti ,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine if it operates at about 7% of available thrust under given operating conditions, or under other different operating conditions, and if a fuel supplied to the combustor device includes a sustainable aviation fuel;

[0034] £7maxTo,sAF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine if it operates at about 100% of available thrust under the same given operating conditions in which EIràleaü jSaf is calculated, and if a fuel supplied to the combustor device includes a sustainable aviation fuel;

[0035] Ê / raienti jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine if it operates at about 7% of available thrust under the same given operating conditions in which EIalenü >Saf is calculated, and if a fuel supplied to the combustor device is a fossil-based hydrocarbon fuel;

[0036] £7maxTojF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine if it operates at about 100% of available thrust under the same given operating conditions in which EIalenü >Saf is calculated, and if a fuel supplied to the combustor device is a fossil-based hydrocarbon fuel;

[0037] and in which the second ratio of the nvPM idle-PMD emission index of the gas turbine engine may be less than 1.

[0038] The second nvPM-PMD emission index ratio at idle may be greater than zero.

[0039] The second nvPM-PMD emission index ratio can be less than or equal to 0.8 and preferably can be less than or equal to 0.6 and more preferably can be less than or equal to 0.4 and more preferably can be less than or equal to 0.2.

[0040] The second nvPM-PMD emission index ratio can be less than or equal to 0.178 and preferably can be less than or equal to 0.164 and more preferably can be less than or equal to 0.149.

[0041] The second nvPM-PMD emission index ratio can be greater than or equal to 0.03 and preferably can be greater than or equal to 0.06 and more preferably can be greater than or equal to 0.09.

[0042] The second nvPM-PMD emission index ratio can be greater than or equal to 0.118 and preferably can be greater than or equal to 0.133 and more preferably can be greater than or equal to 0.148.

[0043] The second nvPM-PMD emission index ratio can be in the range of 0.118 to 0.178 and preferably can be in the range of 0.133 to 0.164 and more preferably can be in the range of 0.148 to 0.149.

[0044] The second nvPM-PMD emission index ratio at idle 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 within any defined range between any two of these values. For example, the second nvPM-PMD emission index ratio at idle may be within a range between 0.25 and 0.4 or between 0.3 and 0.35.

[0045] According to a second aspect, a gas turbine engine for an aircraft is provided, comprising one or more of the following features:

[0046] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles,in which the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and in which: ,

[0047] A second nvPM emission index ratio at idle-PMD can be defined as:

[0048] [Math.3] The slowdown. SAF / Fl / 1 tùaxTO. SAF FF / EIiiiaxTa [!F

[0049] where:

[0050] Ê / raienti ,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine if it operates at about 7% of available thrust under given operating conditions and if a fuel supplied to the combustor device includes a sustainable aviation fuel;

[0051] £7maxTo,sAF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine if it operates at about 100% of available thrust under the same given operating conditions in which EIràleaü jSaf is calculated, and if a fuel supplied to the combustor device includes a sustainable aviation fuel;

[0052] Ê / raienti jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine if it operates at about 7% of available thrust under the same given operating conditions in which EIalenü >Saf is calculated, and if a fuel supplied to the combustor device is a fossil-based hydrocarbon fuel;

[0053] £7maxTojF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine if it operates at about 100% of available thrust under the same given operating conditions in which EIalenü >Saf is calculated, and if a fuel supplied to the combustor device is a fossil-based hydrocarbon fuel;

[0054] the second ratio of the nvPM emission index at idle to PMD of the gas turbine engine is less than 1; and

[0055] The gas turbine engine is designed to supply fuel including sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles.

[0056] The second nvPM idle-PMD emission index ratio defined in the second aspect can be as defined above in relation to the first aspect.

[0057] According to a third aspect, a method of operating the gas turbine engine of the first aspect or of the second aspect is provided, the method comprising supplying fuel including a sustainable aviation fuel to the plurality of fuel spray nozzles.

[0058] According to a fourth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0059] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles,in which the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and in which: ,

[0060] A first nvPM emission index ratio at idle-PMD can be defined as:

[0061] [Math.4] pr 1 slow motion ^^maxTO

[0062] where:

[0063] Æ' / raienti is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine if it operates at approximately 7% of the available thrust under given operating conditions; and

[0064] S / maxTo is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine if it operates at approximately 100% of the available thrust under the given operating conditions;

[0065] the first NVPM-PMD emission index ratio of the gas turbine engine at idle is less than 60; and

[0066] The process includes supplying fuel comprising a sustainable aviation fuel to the plurality of fuel spray nozzles.

[0067] The first nvPM emission index ratio at idle-PMD can be as defined above in relation to the first aspect.

[0068] A second nvPM emission index ratio at idle-PMD can be defined as:

[0069] [Math.5] El idle, SAP / pi / J-1maxTO,SAb The slow motion, FF / pi „ / 1 snaxTQ. FF

[0070] where:

[0071] Ê / raienti ,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine if it operates at about 7% of available thrust under given operating conditions, or under other different operating conditions, and if a fuel supplied to the combustor device includes a sustainable aviation fuel;

[0072] £'7maxTo,sAF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine if it operates at about 100% of available thrust under the same given operating conditions in which EIalenü >Saf is calculated, and if a fuel supplied to the combustor device includes a sustainable aviation fuel;

[0073] Ê / raienti jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine if it operates at about 7% of available thrust under the same given operating conditions in which EIalenü >Saf is calculated, and if a fuel supplied to the combustor device is a fossil-based hydrocarbon fuel;

[0074] £7maxTo,FF is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine if it operates at approximately 100% of available thrust under the same given operating conditions in which EIràleaü jSaf is calculated, and if a fuel supplied to the combustor device is a fossil-based hydrocarbon fuel; and

[0075] in which the second ratio of the nvPM idle-PMD emission index of the gas turbine engine may be less than 1.

[0076] The second nvPM emission index ratio at idle-PMD can be as defined above in relation to the first aspect.

[0077] According to a fifth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0078] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and in which:

[0079] A second nvPM emission index ratio at idle-PMD can be defined as:

[0080] [Math.6] The slowed-down SAF / p I. / 1 maxTu, S AF

[0081] where:

[0082] Ê / raienti ,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine if it operates at about 7% of available thrust under given operating conditions and if a fuel supplied to the combustor device includes a sustainable aviation fuel;

[0083] £7maxTo,sAF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine if it operates at about 100% of available thrust under the same given operating conditions in which EIaienü ,saf is calculated, and if a fuel supplied to the combustor device includes a sustainable aviation fuel;

[0084] Æ' / raienti jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine if it operates at about 7% of available thrust under the same given operating conditions in which EIalenü >Saf is calculated, and if a fuel supplied to the combustor device is a fossil-based hydrocarbon fuel;

[0085] £7maxTojF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine if it operates at about 100% of available thrust under the same given operating conditions in which EIalenü ,Sai is calculated, and if a fuel supplied to the combustor device is a fossil-based hydrocarbon fuel;

[0086] the second ratio of the nvPM emission index at idle to PMD of the gas turbine engine is less than 1; and

[0087] The process includes supplying fuel comprising a sustainable aviation fuel to the plurality of fuel spray nozzles.

[0088] The second nvPM emission index ratio at idle-PMD can be as defined above in relation to the first aspect.

[0089] According to a sixth aspect, a gas turbine engine for an aircraft is provided, comprising any one or more of the following features:

[0090] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a fuel flow rate higher than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and wherein:

[0091] A fuel flow rate nvPM emission index ratio can be defined as:

[0092] [Math.7] EIra]eatiXW fraientj ^masTCi^ W. mas'TO

[0093] where:

[0094] Æ' / raienti is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine if it operates at approximately 7% of available thrust under given operating conditions;

[0095] S / maxTo is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine if it operates at approximately 100% of the available thrust under the given operating conditions;

[0096] Wç idle is the fuel flow rate to the fuel spray nozzles in kg / s at approximately 7% of the available thrust under given operating conditions; and

[0097] W f maxTO is the fuel flow rate to the spray nozzles of fuel in kg / s at approximately 100% of the available thrust under given operating conditions;

[0098] the gas turbine engine fuel flow rate nvPM emission index ratio is less than 6; and

[0099] The gas turbine engine is designed to supply fuel including sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles.

[0100] The nvPM emission index ratio of the fuel flow may be less than 5.93 and preferably may be less than 5.44 and more preferably may be less than 4.94.

[0101] The nvPM emission index ratio of the fuel flow may be less than or equal to 4.5 and preferably may be less than or equal to 3 and more preferably may be less than or equal to 1.5.

[0102] The nvPM emission index ratio of the fuel flow may be less than or equal to 0.879 and preferably may be less than or equal to 0.806 and more preferably may be less than or equal to 0.733.

[0103] The nvPM emission index ratio of the fuel flow may be less than or equal to 0.0181 and preferably may be less than or equal to 0.0166 and more preferably may be less than or equal to 0.0151.

[0104] The ratio of the nvPM emission index of the fuel flow rate may be greater than or equal to 0.00351 and preferably may be greater than or equal to 0.00395 and more preferably may be less than or equal to 0.00439.

[0105] The ratio of the nvPM emission index of the fuel flow rate may be greater than or equal to 0.109 and preferably may be greater than or equal to 0.123 and more preferably may be less than or equal to 0.137.

[0106] The nvPM emission index ratio of the fuel flow can be in the range of 0.00351 to 0.879 and preferably in the range of 0.00395 to 0.806 and more preferably in the range of 0.00439 to 0.733.

[0107] The ratio of the nvPM emission index of the fuel flow can be in the range of 0.00351 to 0.0181 and preferably in the range of 0.00395 to 0.0166 and more preferably in the range of 0.00439 to 0.0151.

[0108] The ratio of the nvPM emission index of the fuel flow rate may be in the range of 0.109 to 0.879 and preferably in the range of 0.123 to 0.806 and more preferably in the range of 0.137 to 0.733.

[0109] The nvPM emission index ratio of the fuel flow can be 0.003, 0.004, 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, 0.1, 0.2, 0.3, 0.4, 0.5, 0.6, 0.7, 0.8, 0.9, 1, 1.5, 2, 2.5, 3, 3.5, 4, 4.5, 5, 5.5, 6 or any range defined between any two of these values. Alternatively, the fuel flow nvPM emission index ratio can be in a range between 0.0009 and 0.021, 0.009 and 0.019, 0.009 and 0.007, or 0.009 and 0.006.

[0110] WfjnaxTO May be in the range of 0.595 to 1.29 kg / s and preferably may be in the range of 0.670 to 1.19 kg / s and more preferably may be in the range of 0.744 to 1.08 kg / s. [YES] Wf raiti can be in the range of 0.0695 to 0.118 kg / s and preferably can be in the range of 0.0782 to 0.108 kg / s and more preferably can be in the range of 0.0869 to 0.0981 kg / s.

[0112] Wf idle May be in the range of 0.0712 to 0.117 kg / s and preferably may be in the range of 0.0801 to 0.107 kg / s and more preferably may be in the range of 0.0890 to 0.0970 kg / s. Wf maxTO May be in the range of 0.595 to 1.28 kg / s and preferably may be in the range of 0.670 to 1.17 kg / s and more preferably may be in the range of 0.744 to 1.07 kg / s.

[0113] Wf^maxTo May be in the range of 0.701 to 1.29 kg / s and preferably may be in the range of 0.788 to 1.19 kg / s and more preferably may be in the range of 0.876 to 1.08 kg / s.

[0114] Wfj-afgnfj can be in the range of 0.0645 to 0.0850 kg / s. Wf rafenfj can be in the range of 0.0645 to 0.0750 kg / s.

[0115] Wf maxTo can be in the range of 0.551 to 0.850 kg / s. can be in the range of 0.551 to 0.750 kg / s.

[0116] According to a seventh aspect, a method of operating the gas turbine engine of the sixth aspect is provided, the method comprising supplying fuel including a sustainable aviation fuel to the plurality of fuel spray nozzles.

[0117] According to an eighth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0118] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles,in which the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and in which: ,

[0119] A fuel flow rate nvPM emission index ratio can be defined as:

[0120] [Math.8] ■^slow motion* slowed down maxTO

[0121] where:

[0122] Δ / raienti is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine if it operates at approximately 7% of available thrust under given operating conditions; and

[0123] S / maxTo is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine if it operates at approximately 100% of the available thrust under the given operating conditions;

[0124] Wf raiti is the fuel flow velocity to the fuel spray nozzles in kg / s at approximately 7% of the available thrust under the given operating conditions; and

[0125] is the fuel flow rate to the spray nozzles of fuel in kg / s at approximately 100% of the available thrust under given operating conditions;

[0126] the gas turbine engine fuel flow rate nvPM emission index ratio is less than 6; and

[0127] The process includes supplying fuel comprising a sustainable aviation fuel to the plurality of fuel spray nozzles.

[0128] Any one of the ratio of the nvPM emissions index of fuel flow, of W{raiti and of WfjnaxTO May be as defined above in relation to the sixth aspect.

[0129] According to a ninth aspect, a gas turbine engine for an aircraft is provided, comprising any one or more of the following features:

[0130] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles,in which the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and in which: ,

[0131] A thrust nvPM emission index ratio can be defined as:

[0132] [Math.9] / iaax j O / To raltinti

[0133] where:

[0134] Ê / raienti is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine if it operates at about 7% of available thrust under given operating conditions;

[0135] S / maxTo is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine if it operates at approximately 100% of the available thrust under the given operating conditions;

[0136] maxTO is 'a Thrust of the gas turbine engine at approximately 100% of the available thrust in kN under the given operating conditions; and

[0137] -^idle is 'a Thrust of the gas turbine engine at approximately 7% of the available thrust in kN under the given operating conditions;

[0138] the nvPM emission index ratio at thrust is greater than 0.001; and

[0139] The gas turbine engine is designed to supply fuel comprising a sustainable aviation fuel (SAF) with a plurality of fuel spray nozzles.

[0140] The nvPM emission index ratio at thrust may be greater than 0.00115 and preferably may be greater than 0.00129 and more preferably may be greater than 0.00144.

[0141] The nvPM emission index ratio at thrust may be greater than 0.0644 and preferably may be greater than 0.0724 and more preferably may be greater than 0.0805.

[0142] The nvPM emission index ratio at thrust may be greater than or equal to 0.003 and preferably may be greater than or equal to 0.005 and more preferably may be greater than or equal to 0.007.

[0143] The nvPM emission index ratio at thrust may be greater than or equal to 0.00776 and preferably may be greater than or equal to 0.00874 and more preferably may be greater than or equal to 0.00971.

[0144] The nvPM emission index ratio at thrust may be greater than or equal to 0.434 and preferably may be greater than or equal to 0.488 and more preferably may be greater than or equal to 0.542.

[0145] The nvPM emission index ratio at thrust may be less than or equal to 1.77 and preferably may be less than or equal to 1.62 and more preferably may be less than or equal to 1.48.

[0146] The nvPM emission index ratio at thrust may be less than or equal to 0.0553 and preferably may be less than or equal to 0.0507 and more preferably may be less than or equal to 0.0461.

[0147] The nvPM emission index ratio at thrust may be in the range of 0.00776 to 1.77 and preferably may be in the range of 0.00874 to 1.62 and more preferably may be in the range of 0.00971 to 1.48.

[0148] The nvPM emission index ratio at thrust may be greater than 0.001, 0.0015, 0.002, 0.0025, 0.003, 0.0035, 0.004, 0.0045, 0.005, 0.01, 0.02, 0.03, 0.04, 0.05, 0.06, 0.07, 0.08, 0.09, 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, 1, 1.05, 1.1, 1.15, 1.2, 1.25, 1.3, 1.35, 1.4, 1.45, 1.5, 1.55, 1.6, 1.65, 1.7, 1.75, 1.8, 1.85, 1.9, 1.95, 2, 2.5, 3, 3.5, 4, 4.5, 5, 5.5, or 6, or any defined range between any two of these values. For example, the nvPM to thrust emission index ratio can be in the range of 0.3 to 4.5, 0.4 to 4, 1 to 4.5, or 1.4 to 4.

[0149] The nvPM emission index ratio at thrust may be in the range of 0.434 to 1.77 and preferably may be in the range of 0.488 to 1.62 and more preferably may be in the range of 0.542 to 1.48.

[0150] The nvPM emission index ratio at thrust may be in the range of 0.00776 to 0.0553 and preferably may be in the range of 0.00874 to 0.0507 and more preferably may be in the range of 0.00971 to 0.0461.

[0151] -PmaxTO May be ^ans 'a P'a8c 85.4 kN to 172 kN and preferably may be in the range of 96.1 kN to 158 kN and preferably may be in the range of 106 kN to 144 kN.

[0152] -^slow Can be ^ans 'a P'a8c 5.98 kN to 12.1 kN and preferably can be in the range of 6.72 kN to 11.1 kN and more preferably can be in the range of 7.47 kN to 10.1 kN.

[0153] FmaxTO May be ^ans 'a P'a8c 89.0 kN to 157 kN and preferably may be in the range of 100 kN to 144 kN and more preferably may be in the range of 111 kN to 131 kN.

[0154] -^slow speed Can be ^in the P'a8c 6.23 kN to 11.0 kN and preferably can be in the range of 7.00 kN to 10.1 kN and more preferably can be in the range of 7.78 kN to 9.13 kN.

[0155] -PmaxTO May be in the range of 50 kN to 85 kN and preferably in the range of 57 kN to 78 kN and preferably in the range of 60 kN to 73 kN, and preferably in the range of 60 kN to 70 kN.

[0156] -^slowdown Can be ^ans 'a P'a8c ^,5 kN to 6 kN and preferably in the range of 4 kN to 5,5 kN and preferably in the range of 4,2 kN to 5.2 kN, and preferably in the range of 4,2 kN to 5 kN.

[0157] According to a tenth aspect, a method of operating the gas turbine engine of the ninth aspect is provided, the method comprising supplying fuel including a sustainable aviation fuel to the plurality of fuel spray nozzles.

[0158] According to an eleventh aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0159] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles,in which the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and in which: ,

[0160] a nvPM emission index ratio to thrust can be defined as:

[0161] [Math. 10]

[0162] where:

[0163] Æ' / raienti is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine if it operates at approximately 7% of available thrust under given operating conditions; and

[0164] S / maxTo is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine if it operates at approximately 100% of the available thrust under the given operating conditions;

[0165] is 'the gas turbine engine thrust at approximately 100% of the thrust available in kN under given operating conditions,

[0166] -^raient! is 'a Thrust of the gas turbine engine at approximately 7% of the available thrust in kN under the given operating conditions; and

[0167] the nvPM emission index ratio at thrust is greater than 0.001; and

[0168] The method includes supplying fuel comprising a sustainable aviation fuel to the plurality of fuel spray nozzles.

[0169] Any one of the ratio of the emission index of nvPM at thrust, -FmaxTo and ■^idle May be as defined in relation to the ninth aspect.

[0170] According to a twelfth aspect, a gas turbine engine for an aircraft is provided, comprising any one or more of the following features:

[0171] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles,in which the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and in which: ,

[0172] A depleted cruise nvPM-PMD emission index ratio can be defined as:

[0173] [Math. 11] cruise [impoverished] / El __________ / niaxTQ BPR

[0174] where:

[0175] E / cruise (impoverished) can be defined as:

[0176] [Math. 12] pi +FT 2 max TO 1 * m on ta a 2

[0177] Ê / maxio is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 100% of available thrust under given operating conditions;

[0178] EImontée is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 85% of available thrust under given operating conditions; and

[0179] BPR is the bypass ratio of the gas turbine engine;

[0180] the ratio of the nvPM emissions index in depleted cruise-PMD is greater than 0.2; and

[0181] The gas turbine engine is designed to supply fuel including sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles.

[0182] The ratio of the nvPM emission index in depleted cruise-PMD may be less than 0.119 and preferably may be less than 0.109 and more preferably may be less than 0.0989.

[0183] The ratio of the nvPM emissions index in depleted cruise-PMD may be less than 0.101 and preferably may be less than 0.0922 and more preferably may be less than 0.0838.

[0184] The nvPM emission index ratio in depleted cruise-PMD may be less than or equal to 0.095 and preferably may be less than or equal to 0.092 and preferably further may be less than or equal to 0.089.

[0185] The ratio of the nvPM emissions index in depleted cruise-PMD may be less than or equal to 0.106 and preferably may be less than or equal to 0.0972 and preferably further may be less than or equal to 0.0883.

[0186] The ratio of the nvPM emissions index in depleted cruise-PMD may be less than or equal to 0.0887 and preferably may be less than or equal to 0.0813, and preferably further may be less than or equal to 0.0739.

[0187] The ratio of the nvPM emissions index in depleted cruise-PMD may be greater than or equal to 0.0519 and preferably may be greater than or equal to 0.0584, and preferably further may be greater than or equal to 0.0649.

[0188] The ratio of the nvPM emissions index in depleted cruise-PMD may be greater than or equal to 0.0578 and preferably may be greater than or equal to 0.0651, and preferably further may be greater than or equal to 0.0723.

[0189] The nvPM emission index ratio in depleted cruise-PMD may be in the range of 0.0519 to 0.106 and preferably may be in the range of 0.0584 to 0.0972 and preferably further may be in the range of 0.0649 to 0.0883.

[0190] The nvPM emission index ratio in depleted cruise-PMD may be in the range of 0.0519 to 0.0887 and preferably may be in the range of 0.0584 to 0.0813 and preferably further may be in the range of 0.0649 to 0.0739.

[0191] The nvPM emission index ratio in depleted cruise-PMD can be in the range of 0.0578 to 0.106 and preferably can be in the range of 0.0651 to 0.0972 and more preferably can be in the range of 0.0723 to 0.0883.

[0192] The ratio of the nvPM emission index in depleted cruise-PMD 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.

[0193] BPR may be in the range of 6.63 to 13.4 and preferably may be in the range of 7.46 to 12.3 and preferably further may be in the range of 8.29 to 11.1.

[0194] BPR may be in the range of 8.36 to 13.4 and preferably may be in the range of 9.40 to 12.3 and preferably further may be in the range of 10.4 to 11.1.

[0195] BPR may be in the range of 6.63 to 10.3 and preferably may be in the range of 7.46 to 9.38 and preferably further may be in the range of 8.29 to 8.53.

[0196] BPR can be in the range of 3.5 to 6.5 and more preferably in the range of 4 to 6 and even more preferably in the range of 4 to 5 or 5 to 6.

[0197] A rich cruise nvPM-PMD emission index ratio can be defined as:

[0198] [Math. 13] cruise (luxury) / H 7 _____________ / ■'-■'maxTO BPR

[0199] where:

[0200] El^iète (rich) can be defined as:

[0201] [Math. 14] pi . +pr . ■^-'moiitee^-^1 approach 2

[0202] EImontée is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust under given operating conditions, or under other different operating conditions;

[0203] ^fapproach is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 30% of available thrust under the same operating conditions under which EImontée is calculated; and

[0204] Elmmtrn is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 100% of available thrust under the same operating conditions under which EImontée is calculated; and

[0205] wherein the rich cruise nvPM-PMD emission index ratio is less than 20.

[0206] The rich cruise nvPM-PMD emission index ratio may be less than 19 and preferably may be less than 17.5 and more preferably may be less than 15.9.

[0207] The rich cruise nvPM-PMD emission index ratio may be less than or equal to 12 and preferably may be less than or equal to 9 and more preferably may be less than or equal to 6.

[0208] The rich cruise nvPM-PMD emission index ratio may be less than or equal to 4.54 and preferably may be less than or equal to 4.17 and more preferably may be less than or equal to 3.79.

[0209] The rich cruise nvPM-PMD emission index ratio may be less than or equal to 0.066 and preferably may be less than or equal to 0.0605 and more preferably may be less than or equal to 0.055.

[0210] The rich-cruising nvPM-PMD emission index ratio can be greater than or equal to 0.0374 and preferably greater than or equal to 0.0421 and more preferably greater than or equal to 0.0468.

[0211] The rich cruise nvPM-PMD emission index ratio may be greater than or equal to 1.41 and preferably may be greater than or equal to 1.58 and more preferably may be greater than or equal to 1.76.

[0212] The rich-PMD nvPM emission index ratio may be in the range of 0.0374 to 4.54, preferably may be in the range of 0.0421 to 4.17 and preferably further may be in the range of 0.0468 to 3.79.

[0213] The rich-PMD nvPM emission index ratio may be in the range of 0.0374 to 0.0660, preferably may be in the range of 0.0421 to 0.0605 and preferably further may be in the range of 0.0468 to 0.0550.

[0214] The rich-PMD nvPM emission index ratio may be in the range of 1.41 to 4.54, preferably may be in the range of 1.58 to 4.17 and preferably furthermore may be in the range of 1.76 to 3.79.

[0215] The rich cruise nvPM-PMD emission index ratio may be less than 1, 2, 3, 4, 5, 6, 7, 8, 9, 10, 11, 12, 13, 14, 15, 16, 17, 18, 19, or 20, or within any defined range between any two of these values.

[0216] The rich cruise nvPM-PMD emission index ratio may be less than 0.03, 0.33, 0.63, 0.93, 1.23, 1.53, 1.83, 2.13, 2.43, 2.73, 3.03, 3.33, 3.63, 3.93, 4.23, 4.53, 4.83, 5.13, 5.43, 5.73, or 6.03, or within any defined range between any two of these values.

[0217] According to a thirteenth aspect, a gas turbine engine for an aircraft is provided, comprising any one or more of the following features:

[0218] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, in which The combustor device can operate under a condition in which each of the fuel spray nozzles in the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles in the second subset of fuel spray nozzles, in which the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and in which:

[0219] A rich cruise nvPM-PMD emission index ratio can be defined as:

[0220] [Math. 15] ^A:croisière(riche) / pf ___________ / ■^ HriaxTu BPR

[0221] where:

[0222] £' / Cruise (rich) can be defined as:

[0223] [Math. 16] PT ±PT J ascent ■ 1 approach ' 2

[0224] EImontée is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 85% of available thrust under given operating conditions;

[0225] ^fapproach is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 30% of available thrust under the same operating conditions under which EImontée is calculated;

[0226] Elmmtrn is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 100% of available thrust under the same operating conditions under which EImontée is calculated;

[0227] BPR is the bypass ratio of the gas turbine engine;

[0228] the rich cruise nvPM-PMD emission index ratio is greater than 20; and

[0229] The gas turbine engine is designed to supply fuel including sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles.

[0230] The rich cruise nvPM-PMD emission index ratio and / or the BPR may be as defined above in relation to the twelfth aspect.

[0231] According to a fourteenth aspect, a method of operating the gas turbine engine of the twelfth or thirteenth aspect is provided, the method comprising supplying fuel including a sustainable aviation fuel to the plurality of fuel spray nozzles.

[0232] According to a fifteenth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0233] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles,in which the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and in which: ,

[0234] A depleted cruise nvPM-PMD emission index ratio can be defined as:

[0235] [Math. 17] fi cruise (impoverished.) / EI BPR

[0236] where:

[0237] EZcruise (impoverished) can be defined as:

[0238] [Math. 18] AnaxTCd”^Cancelled

[0239] S / maxTo is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 100% of available thrust under given operating conditions;

[0240] EImontée is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 85% of available thrust under given operating conditions; and

[0241] BPR is the bypass ratio of the gas turbine engine;

[0242] the ratio of the nvPM emissions index in depleted cruise-PMD is greater than 0.2; and

[0243] The process includes supplying fuel comprising a sustainable aviation fuel to the plurality of fuel spray nozzles.

[0244] The ratio of nvPM emissions index in depleted cruise-PMD and / or the BPR may be as defined above in relation to the twelfth aspect.

[0245] A rich cruise nvPM-PMD emission index ratio can be defined as:

[0246] [Math. 19] cruise (rich) / F / —, ________ f ^-^inaxTO BPR

[0247] where:

[0248] EIaoisiè[e (rich) can be defined as:

[0249] [Math.20] pi +pr, 2-1J ascent 'approach

[0250] F / climb is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust under given operating conditions, or under other different operating conditions;

[0251] F / approach is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 30% of available thrust under the same operating conditions under which Elclimb is calculated;

[0252] F / maxio is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 100% of available thrust under the same operating conditions under which Elmontée is calculated; and

[0253] in which the rich cruise nvPM-PMD emission index ratio may be less than 20.

[0254] The rich cruise nvPM-PMD emission index ratio and / or the BPR may be as defined above in relation to the twelfth aspect.

[0255] According to a sixteenth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0256] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of nozzles fuel spray is supplied with fuel at a higher fuel flow rate to each of the fuel spray nozzles in the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and wherein:

[0257] A rich cruise nvPM-PMD emission index ratio can be defined as:

[0258] [Math.21] Elcroisière (rich) / H] __ _______ / -'-•'maxTO BPR

[0259] where:

[0260] ^' / cruise (rich) can be defined as:

[0261] [Math.22] ^Anoiitée+^Lpproche 2

[0262] £7mOntée is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust under given operating conditions;

[0263] ^fapproach is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 30% of available thrust under the same operating conditions under which ^ / climb is calculated;

[0264] Ê / maxio is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 100% of available thrust under the same operating conditions in which EImontée is calculated;

[0265] BPR is the bypass ratio of the gas turbine engine;

[0266] the rich cruise nvPM-PMD emission index ratio may be less than 20; and

[0267] The process includes supplying fuel comprising a sustainable aviation fuel to the plurality of fuel spray nozzles.

[0268] The rich cruise nvPM-PMD emission index ratio and / or the BPR may be as defined above in relation to the twelfth aspect.

[0269] According to a seventeenth aspect, a gas turbine engine for an aircraft is provided, comprising any one or more of the following features:

[0270] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a fuel flow rate higher than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and wherein:

[0271] a ratio of nvPM emissions index to PMD can be defined as:

[0272] [Math.23] ^■^maxTO, SAP ^LiaxTO.FF

[0273] where:

[0274] £'7maxTo,sAF is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 100% of available thrust under given operating conditions if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel (SAF); and

[0275] £7maxTo,FF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 100% of available thrust for given operating conditions if a fuel supplied to the plurality of fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0276] the ratio of the nvPM emission index to the PMD of the gas turbine engine is less than 1; and

[0277] The gas turbine engine is designed to supply fuel including a SAF to the plurality of fuel spray nozzles.

[0278] The ratio of the nvPM emission index to PMD may be greater than zero.

[0279] The ratio of the nvPM emission index to PMD may be less than or equal to 0.93, and preferably may be less than or equal to 0.86, and more preferably may be less than or equal to 0.79.

[0280] The ratio of the nvPM emission index to PMD may be less than or equal to 0.776, and preferably may be less than or equal to 0.711, and more preferably may be less than or equal to 0.646.

[0281] The ratio of the nvPM emission index to PMD may be greater than or equal to 0.15, and preferably may be greater than or equal to 0.3, and more preferably may be greater than or equal to 0.45.

[0282] The ratio of the nvPM emission index to PMD may be greater than or equal to 0.516, and preferably may be greater than or equal to 0.581, and more preferably may be greater than or equal to 0.645.

[0283] The ratio of the nvPM to PMD emission index may be in the range of 0.516 to 0.776, and preferably may be in the range of 0.581 to 0.711, and more preferably may be in the range of 0.645 to 0.646.

[0284] The ratio of the nvPM emission index to PMD 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.

[0285] The ratio of the nvPM emission index to PMD can 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.

[0286] An upstream nvPM emission index ratio can be defined as:

[0287] [Math.24] ^7montêe, SAP pr Jmontœ,FF

[0288] where:

[0289] F / climb,saf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 85% of available thrust for given operating conditions, or for different given operating conditions, if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel (SAF); and

[0290] F / climb jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust for the same given operating conditions at which EImontée saf is calculated and if a fuel supplied to the plurality of fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0291] and in which the nvPM emission index ratio at the rise of the gas turbine engine may be less than 1.

[0292] The upstream nvPM emission index ratio may be greater than zero.

[0293] The upstream nvPM emission index ratio may be less than or equal to 0.9, and preferably may be less than or equal to 0.75, and more preferably may be less than or equal to 0.6.

[0294] The upstroke nvPM emission index ratio may be less than or equal to 0.57, and preferably may be less than or equal to 0.523, and more preferably may be less than or equal to 0.475.

[0295] The upstroke nvPM emission index ratio may be greater than or equal to 0.1, and preferably may be greater than or equal to 0.2, and more preferably may be greater than or equal to 0.3.

[0296] The upstroke nvPM emission index ratio may be greater than or equal to 0.379, and preferably may be greater than or equal to 0.427, and more preferably may be greater than or equal to 0.474.

[0297] The upstream nvPM emission index ratio may be in the range of 0.379 to 0.570, and preferably may be in the range of 0.427 to 0.523, and more preferably may be in the range of 0.474 to 0.475.

[0298] The upstream nvPM emission 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.

[0299] The upstream nvPM emission index ratio may be less than 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.

[0300] An nvPM emission index ratio approach can be defined as:

[0301] [Math.25] approach, SAP jp i Cm 1 approach,FF

[0302] where:

[0303] ^ / approach,saf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 30% of available thrust for given operating conditions, or for different given operating conditions, if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel (SAF); and

[0304] £7approach jf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 30% of available thrust under the same given operating conditions at which £7approach ,saf is calculated and if a fuel supplied to the plurality of fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0305] in which the nvPM emission index ratio approaching the gas turbine engine may be less than 1.

[0306] The nvPM emission index ratio at the approach may be greater than zero.

[0307] The nvPM emission index ratio at the approach may be less than or equal to 0.8, and preferably may be less than or equal to 0.5, and more preferably may be less than or equal to 0.2.

[0308] The nvPM emission index ratio at the approach may be less than or equal to 0.185, and preferably may be less than or equal to 0.169, and more preferably may be less than or equal to 0.154.

[0309] The nvPM emission index ratio at the approach may be greater than or equal to 0.03, and preferably may be greater than or equal to 0.06, and more preferably may be greater than or equal to 0.09.

[0310] The nvPM emission index ratio at the approach may be greater than or equal to 0.122, and preferably may be greater than or equal to 0.138, and more preferably may be greater than or equal to 0.153.

[0311] The nvPM emission index ratio at the approach may be in the range of 0.122 to 0.185, and preferably may be in the range of 0.138 to 0.169, and more preferably may be in the range of 0.153 to 0.154.

[0312] The nvPM emission index ratio at the approach 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.

[0313] The nvPM emission index ratio at the approach 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.

[0314] An nvPM emission index ratio at idle can be defined as:

[0315] [Math.26] ■Ë'Laienti, SAP ^LentiFF

[0316] where:

[0317] Ê / raienti ,saf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 7% of available thrust for given operating conditions, or for different given operating conditions, and if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel (SAF); and

[0318] Ê / raienti jf is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 7% of available thrust under the same given operating conditions at which EIialeaü saf is calculated and if a fuel supplied to the plurality of fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0319] in which the nvPM emission index ratio at idle of the gas turbine engine may be less than 1.

[0320] The nvPM emission index ratio at idle may be greater than zero.

[0321] The nvPM emission index ratio at idle may be less than or equal to 0.8, and preferably may be less than or equal to 0.5, and more preferably may be less than or equal to 0.2.

[0322] The nvPM emission index ratio at idle may be less than or equal to 0.115, and preferably may be less than or equal to 0.106, and more preferably may be less than or equal to 0.0959.

[0323] The nvPM emission index ratio at idle may be greater than or equal to 0.02, and preferably may be greater than or equal to 0.04, and more preferably may be greater than or equal to 0.06.

[0324] The nvPM emission index ratio at idle may be greater than or equal to 0.0766, and preferably may be greater than or equal to 0.0862, and more preferably may be greater than or equal to 0.0958.

[0325] The nvPM emission index ratio at idle may be in the range of 0.0766 to 0.115, and preferably may be in the range of 0.0862 to 0.106, and more preferably may be in the range of 0.0958 to 0.0959.

[0326] The nvPM emission index ratio at idle 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.

[0327] The nvPM emission index ratio at idle can be 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, or within any defined range between any two of these values.

[0328] According to an eighteenth aspect, a gas turbine engine for an aircraft is provided, comprising any one or more of the following features:

[0329] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles,in which the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and in which: ,

[0330] an upstream nvPM emission index ratio can be defined as:

[0331] [Math.27] ^■^montême, SAP pr Jmontœ,FF

[0332] where:

[0333] F / climb,saf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 85% of available thrust under given operating conditions, if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel (SAF); and

[0334] F / climb jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust for the same given operating conditions at which EImontée saf is calculated and if a fuel supplied to the plurality of fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0335] the nvPM emission index ratio at the gas turbine engine's climb is less than 1; and

[0336] The gas turbine engine is designed to supply fuel including a SAF to the plurality of fuel spray nozzles.

[0337] The upstream nvPM emission index ratio can be as defined in relation to the seventeenth aspect.

[0338] According to a nineteenth aspect, a gas turbine engine for an aircraft is provided, comprising any one or more of the following features:

[0339] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles,in which the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and in which: ,

[0340] A ratio of nvPM emissions index to the approach can be defined as:

[0341] [Math.28] approach, SAF PT l PC 1 approach, FF

[0342] where:

[0343] ^ / approach,saf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 30% of available thrust under given operating conditions if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel (SAF); and

[0344] £7approach jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 30% of available thrust for the same given operating conditions at which £7approach ,saf is calculated and if a fuel supplied to the plurality of fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0345] the nvPM emission index ratio approaching the gas turbine engine is less than 1; and

[0346] The gas turbine engine is designed to supply fuel including a SAF to the plurality of fuel spray nozzles.

[0347] The nvPM emission index ratio to the approach may be as defined above in relation to the seventeenth aspect.

[0348] According to a twentieth aspect, a gas turbine engine for an aircraft is provided, comprising any one or more of the following features:

[0349] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles,in which the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and in which: ,

[0350] An NVPM emission index ratio at idle can be defined as:

[0351] [Math.29] ®A-slow, SAF ^■^slowFF

[0352] where:

[0353] EZraiti,saf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 7% of available thrust under given operating conditions if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel (SAF); and

[0354] Æ' / raienti jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 7% of available thrust for the same given operating conditions at which EIialeaü saf is calculated and if a fuel supplied to the plurality of fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0355] the nvPM emission index ratio at idle of the gas turbine engine is less than 1; and

[0356] The gas turbine engine is designed to supply fuel including a SAF to the plurality of fuel spray nozzles.

[0357] The nvPM emission index ratio at idle may be as defined above in relation to the seventeenth aspect.

[0358] According to a twenty-first aspect, a method of operating the gas turbine engine of any of the seventeenth, eighteenth, nineteenth or twentieth aspects is provided, the method comprising supplying fuel comprising a sustainable aviation fuel to the plurality of fuel spray nozzles.

[0359] According to a twenty-second aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0360] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of spray nozzles of fuel in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and in which:

[0361] A ratio of the nvPM emission index to PMD can be defined as:

[0362] [Math.30] E / maxTO, SAP ^maxTO.LT

[0363] where:

[0364] £ / maxTo,sAF is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at 100% of available thrust under given operating conditions if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel (SAF); and

[0365] £ / maxTo,FF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at 100% of available thrust for given operating conditions if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0366] the ratio of the nvPM emission index to the PMD of the gas turbine engine is less than 1; and

[0367] The process includes supplying fuel comprising a sustainable aviation fuel to the plurality of fuel spray nozzles.

[0368] The ratio of the nvPM emission index to PMD can be as defined in relation to the seventeenth aspect.

[0369] An upstream nvPM emission index ratio can be defined as:

[0370] [Math.31] ■^up, SAP ■^fmontée.FF

[0371] where:

[0372] £7,saf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 85% of available thrust for given operating conditions, or for different given operating conditions, if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel (SAF); and

[0373] £7montée ,ff is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 85% of available thrust for the same given operating conditions to which EImontée saf is determined and whether a fuel supplied to the plurality of fuel spray nozzles is a fossil-based hydrocarbon fuel; and in which

[0374] the nvPM emission index ratio at the climb of the gas turbine engine may be less than 1.

[0375] The upstream nvPM emission index ratio can be as defined in relation to the seventeenth aspect.

[0376] An nvPM emission index ratio approach can be defined as:

[0377] [Math.32] The approach, SAF The approach, FF

[0378] where:

[0379] ^ / approach,saf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 30% of available thrust for given operating conditions, or for different given operating conditions, if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel (SAF); and

[0380] ^fapproach jf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 30% of available thrust under the same given operating conditions at which £7approach ,saf is determined and if a fuel supplied to the plurality of fuel spray nozzles is a fossil-based hydrocarbon fuel; and in which

[0381] the nvPM emission index ratio approaching the gas turbine engine may be less than 1.

[0382] The nvPM emission index ratio to the approach may be as defined in relation to the seventeenth aspect.

[0383] An nvPM emission index ratio at idle can be defined as:

[0384] [Math.33] 77ralenti, SAF ^77raieI1ti pp

[0385] where:

[0386] Æ' / raienti ,saf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 7% of available thrust for given operating conditions, or for different given operating conditions, and if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel (SAF); and

[0387] Ê / raienti jf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 7% of available thrust under the same given operating conditions at which EIialeaü saf is determined and if a fuel supplied to the plurality of fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0388] in which the nvPM emission index ratio at idle of the gas turbine engine may be less than 1.

[0389] The nvPM emission index ratio at idle can be as defined in relation to the seventeenth aspect.

[0390] According to a twenty-third aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0391] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles,in which the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and in which: ,

[0392] an upstream nvPM emission index ratio can be defined as:

[0393] [Math.34] PT 1 ascent, SAF ■^^ascent^FF

[0394] where:

[0395] F / climb,saf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 85% of available thrust under given operating conditions if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel (SAF); and

[0396] F / climb jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 85% of available thrust for the same given operating conditions to which EImontée saf is determined and whether a fuel supplied to the plurality of fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0397] the nvPM emission index ratio at the climb of the gas turbine engine may be less than 1; and

[0398] The process includes supplying fuel comprising a sustainable aviation fuel to the plurality of fuel spray nozzles.

[0399] The upstream nvPM emission index ratio can be as defined in relation to the seventeenth aspect.

[0400] According to a twenty-fourth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0401] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles,in which the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and in which: ,

[0402] an nvPM emission index ratio approaching can be defined as:

[0403] [Math.35] F7 approach, P r, approach,. FF

[0404] where:

[0405] ^ / approach,saf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 30% of available thrust under given operating conditions if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel (SAF); and

[0406] £7approach jf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 30% of available thrust for the same given operating conditions to which EI^^ ,Sai is determined and whether a fuel supplied to the plurality of fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0407] the nvPM emission index ratio approaching the gas turbine engine may be less than 1; and

[0408] The process includes supplying fuel comprising a sustainable aviation fuel to the plurality of fuel spray nozzles.

[0409] The nvPM emission index ratio to the approach may be as defined in relation to the seventeenth aspect.

[0410] According to a twenty-fifth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0411] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles,in which the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and in which: ,

[0412] an nvPM emission index ratio at idle can be defined as:

[0413] [Math.36] The slowdown,^ SAF ■^fraleiltiFF

[0414] where:

[0415] Æ' / raienti ,saf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 7% of available thrust under given operating conditions if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel (SAF); and

[0416] Æ' / raienti jf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 7% of available thrust for the same given operating conditions to which EIialenti saf is determined and whether a fuel supplied to the plurality of fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0417] the nvPM emission index ratio at idle of the gas turbine engine may be less than 1; and

[0418] The process includes supplying fuel comprising a sustainable aviation fuel to the plurality of fuel spray nozzles.

[0419] The nvPM emission index ratio at idle can be as defined in relation to the seventeenth aspect.

[0420] According to a twenty-sixth aspect, a gas turbine engine for an aircraft is provided, comprising any one or more of the following features:

[0421] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles,in which the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and in which: ,

[0422] A fuel flow rate modified by a ratio of the nvPM to PMD emission index can be defined as:

[0423] [Math.37] EfmàxTO, SAF , Tj) T ^maxTOFF XW f^TO

[0424] where:

[0425] £7maxTo,sAF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 100% of available thrust for given operating conditions if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel (SAF);

[0426] £7maxTo,FF is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 100% of available thrust under given operating conditions if a fuel supplied to the A plurality of fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0427] ^fjnaxTO is 'c mass flow rate of fuel supplied to the plurality of fuel spray nozzles in kg / s when the gas turbine engine is operating at about 100% of the thrust available for the given operating conditions;

[0428] the fuel flow rate modified by a ratio of the emission index of nvPM to the PMD of the gas turbine engine in kg / s is less than 2; and

[0429] The gas turbine engine is designed to supply fuel including a SAF to the plurality of fuel spray nozzles.

[0430] The fuel flow rate modified by an emission index ratio of nvPM to PMD in kg / s may be greater than zero.

[0431] The fuel flow rate modified by an emission index ratio of nvPM to PMD in kg / s may be less than 1.29, more preferably may be less than 1.19 and even more preferably may be less than 1.08.

[0432] The fuel flow rate modified by an emission index ratio of nvPM to PMD in kg / s may be greater than or equal to 0.384 and preferably may be greater than or equal to 0.432 and preferably further may be greater than or equal to 0.481.

[0433] The fuel flow rate modified by an emission index ratio of nvPM to PMD in kg / s may be less than or equal to 0.834, more preferably may be less than or equal to 0.764 and preferably further may be less than or equal to 0.695.

[0434] The fuel flow modified by an emission index ratio of nvPM to PMD in kg / s can be in the range of 0.384 to 0.834 and preferably can be in the range of 0.432 to 0.764 and preferably further can be in the range of 0.481 to 0.695.

[0435] The fuel flow rate modified by an emission index ratio of nvPM to PMD in kg / s may be greater than or equal to 0.453 and preferably may be greater than or equal to 0.509 and preferably further may be greater than or equal to 0.566.

[0436] The fuel flow modified by an emission index ratio of nvPM to PMD in kg / s can be in the range of 0.453 to 0.834 and preferably can be in the range of 0.509 to 0.764 and preferably further can be in the range of 0.566 to 0.695.

[0437] The fuel flow rate modified by an emission index ratio of nvPM to PMD in kg / s may be 0.38, 0.384, 0.4, 0.42, 0.44, 0.46, 0.48, 0.5, 0.52, 0.54, 0.56, 0.58, 0.6, 0.62, 0.64, 0.66, 0.68, 0.7, 0.72, 0.74, 0.76, 0.78, 0.8, 0.82, 0.83, 0.834 or within any defined range between any two of these values. Alternatively, the fuel flow rate modified by an emission index ratio of nvPM at PMD in kg / s can be in a range between 0.45 and 0.65 or 0.45 and 0.6.

[0438] ^fjnaxTO May be in the range of 0.595 to 1.29 kg / s and preferably may be in the range of 0.670 to 1.19 kg / s, and more preferably may be in the range of 0.744 to 1.08 kg / s.

[0439] Wf niaxTO May be in the range of 0.595 to 1.28 kg / s and preferably may be in the range of 0.670 to 1.17 kg / s, and more preferably can be in the range of 0.744 to 1.07 kg / s.

[0440] WfjnaxTO May be in the range of 0.701 to 1.29 kg / s and preferably may be in the range of 0.788 to 1.19 kg / s, and more preferably may be in the range of 0.876 to 1.08 kg / s.

[0441] ^fjnaxTO May be in the range of 0.551 to 0.850 kg / s. VVfmaïyQ may be in the range of 0.551 to 0.750 kg / s.

[0442] A fuel flow rate modified by an nvPM emission index ratio on climb can be defined as:

[0443] [Math.38] •Ë'-tmontée, SAP y , £7montfeFF X ^fjnotée

[0444] where:

[0445] £7 rise,saf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 85% of the available thrust for the given operating conditions, or under other different operating conditions, if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel; and

[0446] £7montée jf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 85% of available thrust under the same operating conditions at which EImontée ,Sai is calculated, if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and

[0447] ^fjiiontée is the mass flow rate of fuel supplied to the plurality of fuel spray nozzles in kg / s when the gas turbine engine is operating at approximately 85% of the available thrust under the same operating conditions under which EImontée jSaf and EImontée |F are calculated; and in which

[0448] the fuel flow modified by an emission index ratio of nvPM at the climb of the gas turbine engine in kg / s may be less than 2.

[0449] The fuel flow rate modified by an emission index ratio of nvPM on climb in kg / s may be greater than zero.

[0450] The fuel flow rate modified by a climb nvPM emission index ratio in kg / s may be less than 1.05, more preferably may be less than 0.96 and even more preferably may be less than 0.873.

[0451] The fuel flow rate modified by an emission index ratio of nvPM on climb in kg / s may be greater than or equal to 0.234 and preferably may be greater than or equal to 0.263 and preferably further may be greater than or equal to 0.292.

[0452] The fuel flow rate modified by an emission index ratio of nvPM on climb in kg / s may be less than or equal to 0.498, more preferably may be less than or equal to 0.456 and more preferably furthermore may be less than or equal to 0.415.

[0453] The fuel flow rate modified by an emission index ratio of nvPM on climb in kg / s can be in the range of 0.234 to 0.498 and preferably can be in the range of 0.263 to 0.456 and preferably further can be in the range of 0.292 to 0.415.

[0454] The fuel flow rate modified by a climb nvPM emission index ratio in kg / s may be less than or equal to 0.496 and preferably may be less than or equal to 0.455 and preferably further may be less than or equal to 0.413.

[0455] The fuel flow rate modified by an emission index ratio of nvPM on climb in kg / s can be in the range of 0.234 to 0.496 and preferably can be in the range of 0.263 to 0.455 and preferably further can be in the range of 0.292 to 0.413.

[0456] The fuel flow rate modified by an emission index ratio of nvPM on climb in kg / s may be greater than or equal to 0.274 and preferably may be greater than or equal to 0.308 and preferably further may be greater than or equal to 0.342.

[0457] The fuel flow rate modified by a climb nvPM emission index ratio in kg / s may be less than or equal to 0.498 and preferably may be less than or equal to 0.456 and preferably further may be less than or equal to 0.415.

[0458] The fuel flow rate modified by an emission index ratio of nvPM on climb in kg / s can be in the range of 0.274 to 0.498 and preferably can be in the range of 0.308 to 0.456 and preferably further can be in the range of 0.342 to 0.415.

[0459] The fuel flow rate modified by an NvPM emission index ratio during climb may be less than or equal to 0.23, 0.234, 0.24, 0.26, 0.28, 0.3, 0.32, 0.34, 0.36, 0.38, 0.4, 0.42, 0.44, 0.46, 0.48, 0.498, 0.5, or any range defined between any two of these values. Alternatively, the fuel flow rate modified by an NvPM emission index ratio during climb in kg / s may be in a range between 0.275 and 0.475 or 0.3 and 0.4

[0460] rise May be in the range of 0.492 to 1.05 kg / s and preferably may be in the range of 0.554 to 0.960 kg / s, and more preferably may be in the range of 0.616 to 0.873 kg / s.

[0461] Wfpontée May be in the range of 0.492 to 1.05 kg / s and preferably may be in the range of 0.554 to 0.957 kg / s, and more preferably may be in the range of 0.616 to 0.870 kg / s.

[0462] ^fjnotée May be in the range of 0.577 to 1.05 kg / s and preferably may be in the range of 0.649 to 0.960 kg / s, and more preferably may be in the range of 0.721 to 0.873 kg / s.

[0463] Ascent May be in the range of 0.461 to 0.650 kg / s. Wfjnotée May be in the range of 0.461 to 0.600 kg / s.

[0464] A fuel flow rate modified by an approach nvPM emission index ratio can be defined as:

[0465] [Math.39] ^approach, SAF y. r LApproche.FF UpprOClie

[0466] where:

[0467] ^7appropOChe ,saf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 30% of the available thrust for the given operating conditions, or under other different operating conditions, if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel; and

[0468] ^fapproach jf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 30% of available thrust under the same operating conditions to which £7approach saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and

[0469] approach is 'c mass flow rate of fuel supplied to the plurality of fuel spray nozzles in kg / s when the gas turbine engine is operating at about 30% of the available thrust for the same operating conditions at which EI.^^ >Saf and EI.^^ >FF are calculated; and in which

[0470] the fuel flow modified by an emission index ratio of nvPM approaching the gas turbine engine in kg / s may be less than 0.4.

[0471] The fuel flow rate modified by an nvPM emission index ratio in the kg / s approach may be greater than zero.

[0472] The fuel flow rate modified by an emission index ratio of nvPM to the kg / s approach may be less than 0.343, more preferably may be less than 0.314 and even more preferably may be less than 0.286.

[0473] The fuel flow rate modified by an nvPM emission index ratio approach in kg / s may be greater than or equal to 0.0269 and preferably may be greater than or equal to 0.0302 and preferably further may be greater than or equal to 0.0336.

[0474] The fuel flow rate modified by an emission index ratio of nvPM to the kg / s approach may be less than or equal to 0.0526, more preferably may be less than or equal to 0.0482 and preferably further may be less than or equal to 0.0439.

[0475] The fuel flow rate modified by an emission index ratio of nvPM to the approach in kg / s may be in the range of 0.0269 to 0.0526 and preferably may be in the range of 0.0302 to 0.0482 and preferably further may be in the range of 0.0336 to 0.0439.

[0476] The fuel flow rate modified by an emission index ratio of nvPM to the approach in kg / s may be less than or equal to 0.0524 and preferably may be less than or equal to 0.0480 and preferably further may be less than or equal to 0.0437.

[0477] The fuel flow rate modified by an emission index ratio of nvPM to the approach in kg / s may be in the range of 0.0269 to 0.0524 and preferably may be in the range of 0.0302 to 0.0480 and preferably further may be in the range of 0.0336 to 0.0437.

[0478] The fuel flow rate modified by an nvPM emission index ratio approach in kg / s may be greater than or equal to 0.0301 and preferably may be greater than or equal to 0.0339 and preferably further may be greater than or equal to 0.0376.

[0479] The fuel flow rate modified by an nvPM emission index ratio approach in kg / s may be less than or equal to 0.0526 and preferably may be less than or equal to 0.0482 and preferably further may be less than or equal to 0.0439.

[0480] The fuel flow rate modified by an emission index ratio of nvPM to the approach in kg / s may be in the range of 0.0301 to 0.0526 and preferably may be in the range of 0.0339 to 0.0482 and preferably further may be in the range of 0.0376 to 0.0439.

[0481] The fuel flow rate modified by an approach nvPM emission index ratio in kg / s may be 0.026, 0.0269, 0.028, 0.03, 0.032, 0.034, 0.036, 0.038, 0.04, 0.042, 0.044, 0.046, 0.048, 0.05, 0.052, 0.0526, 0.053, or any defined range between any two of these values. Alternatively, the fuel flow rate modified by an approach nvPM emission index ratio in kg / s may be in a range between 0.02 and 0.1 or 0.05 and 0.075.

[0482] ^tapproche may be in the range of 0.175 to 0.343 kg / s and preferably may be in the range of 0.197 to 0.314 kg / s, and more preferably may be in the range of 0.219 to 0.286 kg / s.

[0483] ^f, approach may be in the range of 0.175 to 0.341 kg / s and preferably may be in the range of 0.197 to 0.313 kg / s, and more preferably may be in the range of 0.219 to 0.284 kg / s.

[0484] ^fyapproche may be in the range of 0.196 to 0.343 kg / s, and preferably may be in the range of 0.220 to 0.314 kg / s, and more preferably in the range of 0.245 to 0.286 kg / s.

[0485] approach can be in the range of 0.166 to 0.300 kg / s. Wf approach can be in the range of 0.166 to 0.250 kg / s.

[0486] A fuel flow rate modified by an nvPM emission index ratio at idle can be defined as:

[0487] [Math.40] ^slow motion, SAF r .r ^slow motion FF Slow motion

[0488] where:

[0489] Æ' / raienti ,saf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 7% of the available thrust for the given operating conditions, or under other different operating conditions, if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel; and

[0490] Δ / raienti jf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 7% of available thrust under the same operating conditions at which EIialenti ,Sai is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and

[0491] ^fj'alient is the mass flow rate of fuel supplied to the plurality of fuel spray nozzles in kg / s when the gas turbine engine is operating at approximately 7% of the available thrust under the same operating conditions at which ^ / ralient ,saf and E / raiti ,ff are calculated; and in which

[0492] the fuel flow modified by an emission index ratio of nvPM at idle of the gas turbine engine in kg / s may be less than 0.2.

[0493] The fuel flow rate modified by an emission index ratio of nvPM at idle in kg / s may be greater than zero.

[0494] The fuel flow modified by an emission index ratio of nvPM at idle in kg / s may be less than 0.118, and more preferably may be less than 0.108 and even more preferably may be less than 0.0981.

[0495] The fuel flow modified by an emission index ratio of nvPM at idle in kg / s may be greater than or equal to 0.00666, and preferably may be greater than or equal to 0.00749 and preferably further may be greater than or equal to 0.00833.

[0496] The fuel flow modified by an emission index ratio of nvPM at idle in kg / s may be less than or equal to 0.0113, and more preferably may be less than or equal to 0.0104 and preferably further may be less than or equal to 0.0094.

[0497] The fuel flow modified by an emission index ratio of nvPM at idle in kg / s may be in the range of 0.00666 to 0.0113, and preferably may be in the range of 0.00749 to 0.0104 and preferably further may be in the range of 0.00833 to 0.0094.

[0498] The fuel flow modified by an emission index ratio of nvPM at idle in kg / s may be greater than or equal to 0.00682, and preferably may be greater than or equal to 0.00767 and preferably further may be greater than or equal to 0.00853.

[0499] The fuel flow modified by an emission index ratio of nvPM at idle in kg / s may be less than or equal to 0.0112, and preferably may be less than or equal to 0.0103 and preferably further may be less than or equal to 0.00929.

[0500] The fuel flow modified by an emission index ratio of nvPM at idle in kg / s may be in the range of 0.00682 to 0.0112, and preferably may be in the range of 0.00767 to 0.0103 and preferably further may be in the range of 0.00853 to 0.00929.

[0501] The fuel flow rate modified by an emission index ratio of nvPM at idle in kg / s may be 0.0065, 0.00666, 0.007, 0.0075, 0.008, 0.0085, 0.009, 0.0095, 0.01, 0.0113, 0.015, or any defined range between any two of these values. Alternatively, the fuel flow rate modified by an emission index ratio of nvPM at idle in kg / s may be in a range between 0.01 and 0.03 or 0.015 and 0.025.

[0502] Wf idle May be in the range of 0.0695 to 0.118 kg / s, and preferably may be in the range of 0.0782 to 0.108 kg / s, and more preferably may be in the range of 0.0869 to 0.0981 kg / s.

[0503] Wf idle May be in the range of 0.0712 to 0.117 kg / s, and preferably may be in the range of 0.0801 to 0.107 kg / s, and more preferably can be in the range of 0.0890 to 0.0970 kg / s.

[0504] FWf idle speed can be in the range of 0.0645 to 0.0850 kg / s. Wç idle speed can be in the range of 0.0645 to 0.0750 kg / s.

[0505]

[0506]

[0507]

[0508]

[0509]

[0510]

[0511]

[0512]

[0513] According to a twenty-seventh aspect, a gas turbine engine for an aircraft is provided, comprising any one or more of the following features: a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles,in which the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and in which: , A fuel flow rate modified by an NvPM emission index ratio during ascent can be defined as: [Math.41] ^^up, SAP X climb Or : El^ontée ,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 85% of available thrust under given operating conditions and if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel; and EImontée FF is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 85% of available thrust under the same operating conditions at which EImontée jSaf is calculated, if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and ^fmontée is 'c mass flow rate of fuel supplied to the plurality of fuel spray nozzles in kg / s when the gas turbine engine is operating at about 85% of the available thrust for the same operating conditions to which EImontée Sai and EImontée >Ff are calculated; the fuel flow rate modified by an NvPM emission index ratio at the gas turbine engine's climb is kg / s is less than 2; and

[0514] The gas turbine engine is designed to supply fuel including a SAF to the plurality of fuel spray nozzles.

[0515] The fuel flow modified by an emission index ratio of nvPM on ascent and / or ^(ascent May be as defined above in relation to the twenty-sixth aspect.

[0516] According to a twenty-eighth aspect, a gas turbine engine for an aircraft is provided, comprising any one or more of the following features:

[0517] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles,in which the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and in which: ,

[0518] A fuel flow rate modified by an approach nvPM emission index ratio can be defined as:

[0519] [Math.42] ■F7 approach, SAF T 47 Approach,FF X Approach

[0520] where:

[0521] £7 approach,saf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 30% of the available thrust for the given operating conditions, or under other different operating conditions, if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel; and

[0522] ^fapproach jf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 30% of available thrust under the same operating conditions to which £7aPPreach saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and

[0523] ^tapproche is 'c mass flow rate of fuel supplied to the plurality of fuel spray nozzles in kg / s when the gas turbine engine is operating at about 30% of the available thrust for the same operating conditions at which £7approche ,saf and £7approche jf are calculated; and in which

[0524] the fuel flow rate modified by an emission index ratio of nvPM approaching the gas turbine engine in kg / s is less than 0.4; and

[0525] The gas turbine engine is designed to supply fuel including a SAF to the plurality of fuel spray nozzles.

[0526] The fuel flow modified by an NVPM emission index ratio at approach and / or Wf, approach Pcut be as defined above in relation to the twenty-sixth aspect.

[0527] According to a twenty-ninth aspect, a gas turbine engine for an aircraft is provided, comprising any one or more of the following features:

[0528] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles,in which the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and in which: ,

[0529] A fuel flow rate modified by an NVPM emission index ratio at idle can be defined as:

[0530] [Math.43] Slow motion, SAP ■^slow motion,FF slow motion

[0531] where:

[0532] Ê / nienti ,saf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 7% of available thrust under given operating conditions and if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel; and

[0533] Δ / raienti jf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 7% of available thrust under the same operating conditions at which Δ / raienti jSaf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and

[0534] Wf raiti is the mass flow rate of fuel supplied to the plurality of fuel spray nozzles in kg / s when the gas turbine engine is operating at about 7% of the available thrust for the same operating conditions at which ^ralid ,saf and EIaienti ,ff are calculated;

[0535] the fuel flow rate modified by an emission index ratio of nvPM at idle of the gas turbine engine in kg / s is less than 0.2; and

[0536] The gas turbine engine is designed to supply fuel including a SAF to the plurality of fuel spray nozzles.

[0537] The fuel flow modified by an emission index ratio of nvPM at idle and / or W'fjaienti may be as defined above in relation to the twenty-sixth aspect.

[0538] According to a thirtieth aspect, a method of operating the gas turbine engine of the twenty-sixth, twenty-seventh, twenty-eighth or twenty-ninth aspect is provided, the method comprising supplying fuel including a sustainable aviation fuel to the plurality of fuel spray nozzles.

[0539] According to a thirty-first aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0540] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles,in which the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; where: ,

[0541]

[0542]

[0543]

[0544]

[0545]

[0546]

[0547]

[0548]

[0549]

[0550]

[0551]

[0552]

[0553] A fuel flow rate modified by a ratio of the nvPM to PMD emission index can be defined as: [Math.44] gfmaxTO, SAP ■FAnaxTO.FF XO^TO Or : Ê / maxTo.sAF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 100% of available thrust for given operating conditions if a fuel supplied to the plurality of fuel spray nozzles includes a sustainable aviation fuel (SAF); E / maxTorF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 100% of available thrust under given operating conditions if a fuel supplied to the plurality of fuel spray nozzles is a fossil-based hydrocarbon fuel; and WfjnaxTo is 'c mass flow rate of fuel supplied to the plurality of fuel spray nozzles in kg / s when the gas turbine engine is operating at about 100% of the thrust available for the given operating conditions; the fuel flow rate modified by a ratio of the emission index of nvPM to PMD in kg / s is less than 2; and the process includes supplying fuel comprising a sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles. The fuel flow rate modified by a ratio of the emissions index from nvPM to PMD and / or Wf,maxTQ can be as defined above in relation to the twenty- sixth aspect. A fuel flow rate modified by a ratio of nvPM emissions index to the Ascent can be defined as: [Math.45] ^7 ascent, SAF pr mounts,FF X Wf fruitee Or : ^ / ascent,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 85% of the available thrust for the given operating conditions, or under other different operating conditions, if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel; and

[0554] Elmontée FF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 85% of available thrust under the same operating conditions at which EImontée jSaf is calculated, if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and

[0555] Wf niont.ée is 'c mass flow rate of fuel supplied to the plurality of nozzles of fuel spray in kg / s when the gas turbine engine is operating at approximately 85% of available thrust under the same operating conditions under which EImontée, Sai and EImontée >FF are calculated; and in which

[0556] the fuel flow modified by an emission index ratio of nvPM at the climb of the gas turbine engine in kg / s may be less than 2.

[0557] The fuel flow modified by an emission index ratio of nvPM on climb and / or Wf climb May be as defined above in relation to the twenty-sixth aspect.

[0558] A fuel flow rate modified by an approach nvPM emission index ratio can be defined as:

[0559] [Math.46] The approach. SAP r,, ^approach,FF * Approach

[0560] where:

[0561] ^ / approach,saf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 30% of the available thrust for the given operating conditions, or under other different operating conditions, if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel; and

[0562] £7approach jf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 30% of available thrust under the same operating conditions at which £7approach saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and

[0563] ^tapproche is 'c mass flow rate of fuel supplied to the plurality of fuel spray nozzles in kg / s when the gas turbine engine is operating at about 30% of the available thrust for the same operating conditions at which EI.^^ >Saf and EI.^^ FF are calculated; and in which

[0564] the fuel flow modified by an emission index ratio of nvPM approaching the gas turbine engine in kg / s may be less than 0.4.

[0565] The fuel flow modified by an emission index ratio of nvPM to the approach and / or W^fapproc]ie can be as defined above in relation to the twenty-sixth aspect.

[0566] A fuel flow rate modified by an nvPM emission index ratio at idle can be defined as:

[0567] [Math.47] Slow motion. SAP 7^7 Slow Motion, FF XW Slow Motion

[0568] where:

[0569] Æ' / raienti ,saf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 7% of the available thrust for the given operating conditions, or under other different operating conditions, if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel; and

[0570] Æ' / raienti jf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 7% of available thrust under the same operating conditions at which EIaienti jSaf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and

[0571] Wfjalentiest 'c mass flow rate of fuel supplied to the plurality of fuel spray nozzles in kg / s when the gas turbine engine is operating at about 7% of the available thrust for the same operating conditions at which ^ / ralentiest ,saf and EIialenti n are calculated; and in which

[0572] the fuel flow modified by an emission index ratio of nvPM at idle of the gas turbine engine in kg / s may be less than 0.2.

[0573] The fuel flow modified by an emission index ratio of nvPM at idle and / or Wf raitiTO May be as defined above in relation to the twenty-sixth aspect.

[0574] According to a thirty-second aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0575] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a flow rate of fuel higher than each of the fuel spray nozzles of the second subset of fuel spray nozzles, in which a ratio of number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of spray nozzles fuel in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and in which:

[0576]

[0577]

[0578]

[0579]

[0580]

[0581]

[0582]

[0583]

[0584]

[0585]

[0586] a fuel flow rate modified by a ratio of nvPM emissions index to the Ascent can be defined as: [Math.48] ^upgrade, SAP pr ascent,FF Wf, uphill Or : E / ascent, saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 85% of available thrust under given operating conditions and if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel; and E / rise jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 85% of available thrust under the same operating conditions at which EIrise jSaf is calculated, if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and ^fjnotée is 'c mass flow rate of fuel supplied to the plurality of fuel spray nozzles in kg / s when the gas turbine engine is operating at about 85% of the available thrust for the same operating conditions to which EImontée Sai and EImontée >FF are calculated; the fuel flow rate modified by a ratio of the nvPM emission index at the gas turbine engine's climb rate in kg / s is less than 2; and the process includes supplying fuel comprising a sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles. The fuel flow modified by a ratio of nvPM emission index on ascent and / or Wf^noted May be as defined above in relation to the twenty-sixth aspect. According to a thirty-third aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising: a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of spray nozzles fuel comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a fuel flow rate greater than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and wherein:

[0587] A fuel flow rate modified by an approach nvPM emission index ratio can be defined as:

[0588] [Math.49] The approach, SA F xr ^approach,FF X approach

[0589] where:

[0590] ^ / approach,saf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 30% of available thrust for given operating conditions, or under other different operating conditions, if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel; and

[0591] ^ / approach jf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 30% of available thrust under the same operating conditions to which £7aPPreach saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and

[0592] Wf,approach is the mass flow rate of fuel supplied to the plurality of fuel spray nozzles in kg / s when the gas turbine engine is operating at about 30% of the available thrust for the same operating conditions at which ^ / approach,saf and £7approach,ff are calculated,

[0593] the fuel flow rate modified by an emission index ratio of nvPM approaching the gas turbine engine in kg / s is less than 0.4; and

[0594] The process includes supplying fuel comprising a sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles.

[0595] The fuel flow modified by an emission index ratio of nvPM to the approach and / or W^fapproc]ie can be as defined above in relation to the twenty-sixth aspect.

[0596] According to a thirty-fourth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0597] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles,in which the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and in which: ,

[0598] A fuel flow rate modified by an NvPM emission index ratio at idle can be defined as:

[0599] [Math.50] Slow motion. SAP 7^7 Slow Motion, FF XW Slow Motion

[0600] where:

[0601] Æ' / raienti ,saf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 7% of available thrust under given operating conditions and if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel; and

[0602] ΊLaienti jf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 7% of available thrust under the same operating conditions at which Ίaienti jSaf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and

[0603] Wfjalenti is the mass flow rate of fuel supplied to the plurality of fuel spray nozzles in kg / s when the gas turbine engine is operating at approximately 7% of the available thrust for the same operating conditions under which ^idle, saf and EIaienti are calculated;

[0604] the fuel flow rate modified by an emission index ratio of nvPM at idle of the gas turbine engine in kg / s is less than 0.2; and

[0605] The process includes supplying fuel comprising a sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles.

[0606] The fuel flow modified by an emission index ratio of nvPM at idle and / or Wfj'at idle may be as defined above in relation to the twenty-sixth aspect.

[0607] According to a thirty-fifth aspect, a gas turbine engine for an aircraft is provided, comprising any one or more of the following features:

[0608] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles,in which the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; where: ,

[0609] A depleted cruise nvPM emission index ratio can be defined as:

[0610] [Math.51] P 7 1 impoverished cruise SAF 1 Groisiere(impoverishment^ FF

[0611] where:

[0612] Elcroisière (impoverished), saf can be defined as:

[0613] [Math.52] ^•fmaxTO, SAf+ ^Anontée, SAF 2

[0614] E / eroisière (impoverished) ,ff can be defined as:

[0615] [Math.53] -F-^maxTO, II-*- FF

[0616] £ / maxTo,sAF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 100% of available thrust for given operating conditions if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0617] £7 rise,saf is the system loss corrected nvPM emission index in mg / kg of the gas turbine engine when operating at about 85% of available thrust for given operating conditions if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0618] £7maxTo,FF is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 100% of available thrust under given operating conditions if a fuel supplied to the plurality of fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0619] £7jf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at about 85% of available thrust for given operating conditions if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0620] the nvPM emission index ratio in lean cruise of the gas turbine engine is less than 1; and

[0621] The gas turbine engine is designed to supply fuel including sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles.

[0622] The nvPM emission index ratio in depleted cruise can be greater than zero.

[0623] The nvPM emission index ratio in depleted cruise may be less than or equal to 0.9 and preferably less than or equal to 0.8.

[0624] The nvPM emission index ratio in depleted cruise may be less than or equal to 0.732, and preferably may be less than or equal to 0.671 and preferably further may be less than or equal to 0.61.

[0625] The nvPM emission index ratio in depleted cruise may be greater than or equal to 0.455, and preferably may be greater than or equal to 0.512 and preferably further may be greater than or equal to 0.569.

[0626] The nvPM emission index ratio in depleted cruise can be in the range of 0.455 to 0.732, and preferably can be in the range of 0.512 to 0.671 and preferably further can be in the range of 0.569 to 0.610.

[0627] The nvPM emission index ratio in depleted cruise may be less than 1, 0.95, 0.9, 0.85, 0.8, 0.75, 0.70, 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 defined range between any two of these values. For example, the nvPM emission index ratio in depleted cruise may be within a range between 0.65 and 0.85 or between 0.7 and 0.75.

[0628] The nvPM emission index ratio in depleted cruise can 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, 0.71, 0.72, 0.73, 0.732, or within any range defined between any two of these values.

[0629] A ratio of nvPM emissions index at idle-PMD can be defined as:

[0630] [Math.54] The slowdown, S^ / E1 lnaXTO,SAF

[0631] where:

[0632] Æ' / raienti ,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 7% of available thrust for given operating conditions, or under other different operating conditions, if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0633] £7maxTo,sAF is the nvPM emissions index corrected for system losses in mg / kg of the gas turbine engine when operating at about 100% of available thrust for the same operating conditions at which EIialeaü jSaf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0634] Δ / raienti jf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 7% of available thrust under the same operating conditions at which Δ / raienti jSaf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and

[0635] S / maxTojF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 100% of available thrust for the same operating conditions at which EIialenti ,Sai is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0636] and in which the ratio of the nvPM at idle-PMD emission index of the gas turbine engine may be less than 1.

[0637] The nvPM-PMD emission index ratio at idle may be greater than zero.

[0638] The ratio of the nvPM emission index at idle to PMD may be less than or equal to to 0.8 and preferably may be less than or equal to 0.6, and preferably again may be less than or equal to 0.4 and preferably again may be less than or equal to 0.2.

[0639] The nvPM-PMD emission index ratio may be less than or equal to 0.178, and preferably may be less than or equal to 0.164 and preferably further may be less than or equal to 0.149.

[0640] The nvPM-PMD emission index ratio may be greater than or equal to 0.118 and preferably may be greater than or equal to 0.133 and preferably further may be greater than or equal to 0.148.

[0641] The nvPM-PMD emission index ratio may be in the range of 0.118 to 0.178 and preferably may be in the range of 0.133 to 0.164 and preferably further may be in the range of 0.148 to 0.149.

[0642] The nvPM-PMD emission index ratio at idle 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 defined range between any two of these values. For example, the nvPM-PMD emission index ratio at idle may be within a range between 0.25 and 0.4 or between 0.3 and 0.35.

[0643] The nvPM idle-PMD emission index ratio can 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 defined range between any two of these values.

[0644] A depleted cruise nvPM / PMD emission index ratio can be defined as:

[0645] [Math.55] '-f impoverished cruise!, maxTO.SAF maxTO FF

[0646] where:

[0647] E / cruise (impoverished), saf can be defined as:

[0648] [Math.56] ^7maxTO, SAF+ ^finontée. FAS 2

[0649] ^ / cruise (impoverished) ,ff can be defined as:

[0650] [Math.57] FîmaxTO, Fp+ FF 2

[0651] £7maxTo,sAF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 100% of the thrust available for the given operating conditions, or under other different operating conditions, if a fuel supplied to the plurality of fuel spray nozzles includes a durable aviation fuel;

[0652] Ê / montée ,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust for the same operating conditions at which £7maxTo,sAF is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0653] £ / maxTo,FF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 100% of available thrust under the same operating conditions at which £7maxTo,sAF is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and

[0654] Ê / montée jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust for the same operating conditions at which £7maxTo,sAF is calculated, if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0655] and in which the ratio of the nvPM emissions index in lean cruise / PMD of the gas turbine engine may be less than 1.

[0656] The ratio of nvPM emissions index in depleted cruise / PMD may be greater than zero.

[0657] The ratio of the nvPM emissions index in depleted cruise / PMD may be less than or equal to 0.98 and preferably may be less than or equal to 0.96.

[0658] The ratio of the nvPM emissions index in depleted cruise / PMD may be less than or equal to 0.95, and preferably may be less than or equal to 0.944.

[0659] The ratio of the nvPM emissions index in depleted cruise / PMD may be greater than or equal to 0.704, preferably greater than or equal to 0.792 and preferably further greater than or equal to 0.88.

[0660] The nvPM emission index ratio in depleted cruise / PMD may be in the range of 0.704 to 0.960, and preferably may be in the range of 0.792 to 0.950 and preferably further may be in the range of 0.880 to 0.944.

[0661] The depleted cruise nvPM / PMD emission 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. For example, the depleted cruise nvPM / PMD emission index ratio may be in the range of 0.91 to 0.99 or 0.93 to 0.97.

[0662] The depleted cruise nvPM / PMD emission index ratio may be 0.7, 0.704, 0.725, 0.75, 0.775, 0.8, 0.825, 0.85, 0.875, 0.9, 0.925, 0.95, 0.96, or within any defined range between any two of these values.

[0663] A nvPM emission index ratio at idle / lean cruise can be defined as:

[0664] [Math.58] ^ / EI cruise (impoverished).SAF CTOisière (impoverished);FF

[0665] where:

[0666] Ê / raienti ,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 7% of available thrust for given operating conditions, or under other different operating conditions, if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0667] EZcroisière (impoverished),saf can be defined as:

[0668] [Math.59] / TtmaxTO. SAf+ ■^ / rise, SAP 2

[0669] £7maxTo,sAF is the nvPM emissions index corrected for system losses in mg / kg of the gas turbine engine when operating at about 100% of available thrust for the same operating conditions at which ELaienti ,saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0670] EImontée saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust for the same operating conditions at which EIialeaü jSaf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0671] EIialenti >FF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 7% of available thrust for the same operating conditions at which EIialenti ,Sai is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0672] E / cruise (impoverished),ff can be defined as:

[0673] [Math.60] / L A. . FF+ f 7mont.ée, FF 2

[0674] £7maxTo,FF is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 100% of thrust available under the same operating conditions under which EIalenti >Saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and

[0675] £7montée jf is the system loss corrected nvPM emission index in mg / kg of the gas turbine engine when operating at about 85% of available thrust under the same operating conditions at which EIialenti ,Sai is calculated, if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0676] and in which the nvPM emission index ratio at idle / lean cruise of the gas turbine engine may be less than 1.

[0677] The nvPM emission index ratio at idle / lean cruise can be greater than zero.

[0678] The nvPM emission index ratio at idle / lean cruise is less than or equal to 0.8, preferably less than or equal to 0.6, more preferably less than or equal to 0.4, and more preferably further less than or equal to 0.3.

[0679] The nvPM emission index ratio at idle / lean cruise may be less than or equal to 0.203, and preferably may be less than or equal to 0.186 and preferably further may be less than or equal to 0.169.

[0680] The nvPM emission index ratio at idle / lean cruise may be greater than or equal to 0.125, and preferably may be greater than or equal to 0.141 and preferably further may be greater than or equal to 0.157.

[0681] The nvPM emission index ratio at idle / lean cruise may be in the range of 0.125 to 0.203, and preferably may be in the range of 0.141 to 0.186, and preferably further may be in the range of 0.157 to 0.169.

[0682] The nvPM emission index ratio at idle / lean-fuel cruise 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. For example, the nvPM emission index ratio at idle / lean-fuel cruise may be in the range of 0.3 to 0.4 or 0.3 to 0.35.

[0683] The nvPM emission index ratio at idle / lean cruise can be 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, 0.203, or within any defined range between any two of these values.

[0684] A rich cruise nvPM emission index ratio can be defined as:

[0685] [Math.61] ^^(.Toisiètetriclie), SAP El croLsière(ridie), FF

[0686] where:

[0687] E / cruise (rich), saf can be defined as:

[0688] [Math.62] ■^Ascent, SAIElapproche. §AF 2

[0689] E / cruise (rich) jf can be defined as:

[0690] [Math.63] The climb, Ffd” The approach, FF 2

[0691] E / climb,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust for given operating conditions, or under other different operating conditions, if a fuel supplied to the plurality of fuel spray nozzles includes a sustainable aviation fuel;

[0692] E / approach saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 30% of available thrust for the same operating conditions at which E / climb ,saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0693] E / climb jf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 85% of available thrust under the same operating conditions at which E / climb >Saf is calculated, if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and

[0694] EEPprOche jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 30% of available thrust for the same operating conditions at which E / climb,saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0695] and in which the rich cruise nvPM emission index ratio of the gas turbine engine may be less than 1.

[0696] The nvPM emission index ratio in rich cruise can be greater than zero.

[0697] The rich cruise nvPM emission index ratio may be less than or equal to 0.8 and preferably may be less than or equal to 0.6 and even more preferably may be less than or equal to 0.4.

[0698] The rich cruise nvPM emission index ratio may be less than or equal to 0.303, and preferably may be less than or equal to 0.278 and preferably further may be less than or equal to 0.252.

[0699]

[0700]

[0701]

[0702]

[0703]

[0704]

[0705]

[0706]

[0707]

[0708]

[0709]

[0710]

[0711]

[0712] The rich cruise nvPM emission index ratio may be greater than or equal to 0.123, and preferably may be greater than or equal to 0.138 and preferably further may be greater than or equal to 0.154. The rich cruise nvPM emission index ratio can be in the range of 0.123 to 0.303, and preferably can be in the range of 0.138 to 0.278 and preferably further can be in the range of 0.154 to 0.252. The rich cruise nvPM emission index ratio can 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. For example, the rich cruise nvPM emission index ratio can be in the range between 0.45 and 0.7 or between 0.5 and 0.65. The upstream nvPM emission index ratio may be less than 0.12, 0.123, 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, 0.303, or within any defined range between any two of these values. A rich cruise nvPM / PMD emission index ratio can be defined as: [Math.64] A^cruisierhrichejSAF / F / mæcpQ F\F croisièreyâche'j.FF / El maXTO FF Or : Æ' / cruise (rich) ,saf can be defined as: [Math.65] ^Lricaitée, SaU" ^ / approach, SAF 2 and E / cruise (rich) jf can be defined as: [Math.66] ■^TmoiltëejFF^' approach, FF 2 and where: EImontée saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust for given operating conditions, or under other different operating conditions, if a fuel supplied to the plurality of fuel spray nozzles includes a sustainable aviation fuel; ^approach, saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 30% of available thrust under the same operating conditions as ^ascent,saf is calculated if a fuel supplied to the plurality of fuel spray nozzles includes a durable aviation fuel;

[0713] £7montée jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust under the same operating conditions at which EImontée ,Sai is calculated, if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0714] E / approach,ff is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 30% of available thrust for the same operating conditions at which E / climb,saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0715] £7maxTo,sAF is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 100% of available thrust under the same operating conditions at which ^ / climb,saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel; and

[0716] S / maxTojF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 100% of available thrust for the same operating conditions at which EImontée Saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0717] and in which the rich cruise nvPM / PMD emission index ratio of the gas turbine engine may be less than 1.

[0718] The rich cruise nvPM / PMD emission index ratio may be greater than zero.

[0719] The rich cruise nvPM / PMD emission index ratio may be less than or equal to 0.8, and preferably may be less than or equal to 0.6, and even more preferably may be less than or equal to 0.5.

[0720] The rich cruise nvPM / PMD emission index ratio may be less than or equal to 0.469, and preferably may be less than or equal to 0.43 and preferably further may be less than or equal to 0.391.

[0721] The rich cruise nvPM / PMD emission index ratio may be greater than or equal to 0.191, and preferably may be greater than or equal to 0.214 and preferably further may be greater than or equal to 0.238.

[0722] The rich cruise nvPM / PMD emission index ratio may be in the range of 0.191 to 0.469, preferably may be in the range of 0.214 to 0.430 and preferably furthermore may be in the range of 0.238 to 0.391.

[0723] The rich cruise nvPM / PMD emission 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. For example, the rich cruise nvPM / PMD emission index ratio may be in the range of 0.65 to 0.9 or 0.7 to 0.85 or 0.75 to 0.8.

[0724] The ratio of the nvPM emission index to PMD may be 0.19, 0.191, 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, 0.41, 0.42, 0.43, 0.44, 0.45, 0.46, 0.469, 0.47, or within any range defined between any two of these values.

[0725] A rich idle / cruise nvPM emission index ratio can be defined as:

[0726] [Math.67] ■E-îraiti,SAF / E7 cruise (rich),SAF E / raloai.FF / E / ^j^^.j^^

[0727] where:

[0728] Ê / raienti ,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 7% of available thrust for given operating conditions, or under other different operating conditions, and if a fuel supplied to the plurality of fuel spray nozzles includes a sustainable aviation fuel (SAF);

[0729] Ί / raienti jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 7% of available thrust for the same operating conditions at which EIaienti jSaf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0730] ^ / cruise (rich) ,saf can be defined as:

[0731] [Math.68] ■^uphill, SaU" ■^■fapprocilej saf 2

[0732] E / cruise (rich) jF can be defined as:

[0733] ​​[Math.69] ^Uioiitée.Fld" EJapproach, FF 2

[0734] EImontée saf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 85% of available thrust under the same operating conditions as ELaienti ,saf is calculated if a fuel supplied to the plurality of fuel spray nozzles includes a sustainable aviation fuel;

[0735] £7montée jf is the system loss corrected nvPM emission index in mg / kg of the gas turbine engine when operating at about 85% of available thrust under the same operating conditions at which EIaienti jSaf is calculated, if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0736] ^ / approach ,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 30% of available thrust under the same operating conditions at which EIialeaü jSaf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel; and

[0737] ^ / approach ,ff is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 30% of available thrust for the same operating conditions at which EIiaienti ,Sai is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0738] and in which the nvPM emission index ratio at idle / rich cruise of the gas turbine engine may be less than 1.

[0739] The nvPM emission index ratio at idle / rich cruise can be greater than zero.

[0740] The nvPM emission index ratio at idle / rich cruise may be less than or equal to 0.9 and preferably may be less than or equal to 0.8.

[0741] The nvPM emission index ratio at idle / rich cruise may be less than or equal to 0.746, and preferably may be less than or equal to 0.683 and preferably further may be less than or equal to 0.621.

[0742] The nvPM emission index ratio at idle / rich cruise may be greater than or equal to 0.304, and preferably may be greater than or equal to 0.342 and preferably further may be greater than or equal to 0.38.

[0743] The nvPM emission index ratio at idle / rich cruise may be in the range of 0.304 to 0.746, and preferably may be in the range of 0.342 to 0.683 and preferably further may be in the range of 0.380 to 0.621.

[0744] The nvPM emission index ratio at idle / rich cruise 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. For example, the nvPM emission index ratio at idle / rich cruise may be in the range of 0.3 to 0.5, or 0.35 to 0.45.

[0745] The ratio of the nvPM emission index at idle / rich cruise can be 0.3, 0.304, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39, 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, 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, 0.746, 0.75 or within any defined range between any two of these values

[0746] According to a thirty-sixth aspect, a gas turbine engine for an aircraft is provided, comprising any one or more of the following features:

[0747] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles,in which the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; where: ,

[0748] an nvPM idling-PMD emission index ratio can be defined as:

[0749] [Math.70] £ / rait., S^ / EI maxTO|SAF Elrala^ / E! maxTaPF

[0750] where:

[0751] Æ' / raienti ,saf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at about 7% of available thrust for given operating conditions if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0752] £7maxTo,sAF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 100% of available thrust for the same operating conditions at which EIialeaü jSaf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0753] Æ' / raienti jf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 7% of thrust available under the same operating conditions under which EIalenti >Saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and

[0754] £7maxTo,FF is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at about 100% of available thrust for the same operating conditions at which EIialenti ,Sai is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0755] and the ratio of the nvPM at idle to PMD emission index of the gas turbine engine may be less than 1; and

[0756] The gas turbine engine is designed to supply fuel including sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles.

[0757] The nvPM idle-PMD emission index ratio can be as defined above in relation to the thirty-fifth aspect.

[0758] According to a thirty-seventh aspect, a gas turbine engine for an aircraft is provided, comprising any one or more of the following features:

[0759] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles,in which the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5;

[0760] in which:

[0761] A depleted cruise nvPM / PMD emission index ratio can be defined as:

[0762] [Math.71] ®™elB&ppal.vri4SAI^ maxTO FF

[0763] where:

[0764] E / cruise (impoverished), SAF can be defined as:

[0765] [Math.72] EïiiiaxTO, SAf+ Elmontée, SAF 2

[0766] E / cruise (impoverished),ff can be defined as:

[0767] [Math.73] F7maxTO, FF+ E'^montee, FF

[0768] f / maxTo.sAF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 100% of available thrust for given operating conditions if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0769] f / ascent,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust for the same operating conditions at which £7maxTo,sAF is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0770] f / maxTo.FF is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 100% of available thrust under the same operating conditions at which £7maxTo,sAF is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and

[0771] f / ascent jf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at about 85% of available thrust under the same operating conditions at which £7maxTo,sAF is calculated, if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0772] and the NVPM / PMD emission index ratio of the gas turbine engine in lean cruise may be less than 1; and

[0773] The gas turbine engine is designed to supply fuel including sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles.

[0774] The ratio of nvPM emissions index in depleted cruise / PMD can be as defined above in relation to the thirty-fifth aspect.

[0775] According to a thirty-eighth aspect, a gas turbine engine for an aircraft is provided, comprising any one or more of the following features:

[0776] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles,in which the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; where: ,

[0777] A ratio of the nvPM emission index at idle / lean cruise can be defined as:

[0778] [Math.74] cruisefappauTO^

[0779] where:

[0780] Æ' / raienti ,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 7% of available thrust for given operating conditions if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0781] ^ / cruise (impoverished) ,saf can be defined as:

[0782] [Math.75] ^■fmàxTO, SAld” montage, SAP 2

[0783] S / maxTo.sAF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 100% of available thrust for the same operating conditions at which EIialeaü jSaf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0784] EImontée saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 85% of available thrust under the same operating conditions as EIialeaü Saf is calculated if a fuel supplied to the plurality of fuel spray nozzles includes a sustainable aviation fuel;

[0785] Ί / raienti jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 7% of available thrust for the same operating conditions at which EIaienti jSaf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0786] ^ / cruise (impoverished) ,ff can be defined as:

[0787] [Math.76] ^TfliaxTO, Fp+ El montée, FF 2

[0788] £7maxTo,FF is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 100% of available thrust under the same operating conditions at which EIaienti Saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and

[0789] EImontée jFF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust under the same operating conditions at which EIaienti Saf is calculated, if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0790] the nvPM emission index ratio at idle / lean cruise of the gas turbine engine is less than 1; and

[0791] The gas turbine engine is designed to supply fuel including sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles.

[0792] The nvPM emission index ratio at idle / lean cruise can be as defined above in relation to the thirty-fifth aspect.

[0793] According to a thirty-ninth aspect, a gas turbine engine for an aircraft is provided, comprising any one or more of the following features:

[0794] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a flow rate of fuel greater than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; where:

[0795] A rich cruise nvPM emission index ratio can be defined as:

[0796] [Math.77] ^^tToisiètetriclie), SAP ^fcroLsière(riclie), FF

[0797] where:

[0798] E / cruise (rich), saf can be defined as:

[0799] [Math.78] ^ / ascent, SAF^- EJapproach, SAF 2

[0800] ^' / cruise (rich) jf can be defined as:

[0801] [Math.79] Elmontée, 11 +^Japproche, FF 2

[0802] EImontée saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust for given operating conditions if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0803] ^ / approach ,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 30% of available thrust for the same operating conditions at which ^ / climb ,saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0804] EImontée jFF is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 85% of available thrust under the same operating conditions at which EImontée jSaf is calculated, if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and

[0805] ^ / approach jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 30% of available thrust for the same operating conditions at which ^ / climb ,saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0806] the rich cruise nvPM emission index ratio of the gas turbine engine is less than 1; and

[0807] The gas turbine engine is designed to supply fuel including sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles.

[0808] The rich cruise nvPM emission index ratio can be as defined above in relation to the thirty-fifth aspect.

[0809] According to a fortieth aspect, a gas turbine engine for an aircraft is provided, comprising any one or more of the following features:

[0810] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles,in which the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; where: ,

[0811] A rich cruise nvPM / PMD emission index ratio can be defined as:

[0812] [Math.80] ^cruise0-iche$SAB / -Er F\F cruiseirichetFF / El rna^TO FF

[0813] where:

[0814] E / cruise (rich), saf can be defined as:

[0815] [Math.81] ^'■hnontée. ^ / approach, SAF 2

[0816] and E / cruise (rich) jf can be defined as:

[0817] [Math.82] ^7 ascent, pd~ approach, FF 2

[0818] and where:

[0819] EImontée >Saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust for given operating conditions and if a fuel supplied to the plurality of fuel spray nozzles includes a sustainable aviation fuel;

[0820] The approach, saf is the nvPM emissions index adjusted for losses of the system in mg / kg of the gas turbine engine when operating at about 30% of available thrust for the same operating conditions at which ^ / climb ,saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0821] Ί / rise jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust for the same operating conditions at which EIrise ,Sai is calculated, if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0822] ^ / approach ,ff is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 30% of available thrust for the same operating conditions at which EImontée ,saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0823] Elm^w.sAF is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 100% of available thrust under the same operating conditions at which EImontée ,saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel; and

[0824] £7maXTojF is the system loss corrected nvPM emission index in mg / kg of the gas turbine engine when operating at about 100% of available thrust under the same operating conditions at which EImontée ,saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0825] the rich cruise nvPM / PMD emission index ratio of the gas turbine engine is less than 1; and

[0826] The gas turbine engine is designed to supply fuel including sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles.

[0827] The rich cruise nvPM / PMD emission index ratio can be as defined above in relation to the thirty-fifth aspect.

[0828] According to a forty-first aspect, a gas turbine engine for an aircraft is provided, comprising any one or more of the following features:

[0829] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles,in which the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; where: ,

[0830] A rich idle / cruise nvPM emission index ratio can be defined as:

[0831] [Math. 83] £ Aa|mti.SAt / E7 £ J raia.ti.Fi / E7 cruise(riclie),FF

[0832] where:

[0833] Æ' / raienti ,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 7% of available thrust for given operating conditions and if a fuel supplied to the plurality of fuel spray nozzles includes a sustainable aviation fuel (SAF);

[0834] Æ' / raienti jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 7% of available thrust for the same operating conditions at which EIaienti jSaf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0835] ^ / cruise (rich) ,saf can be defined as:

[0836] [Math. 84] ^Ariontée, Saf+ El approche, SAF 2

[0837] £' / cruise (rich) ,ff can be defined as:

[0838] [Math.85] -S'-Gnoritée.FF^ S^approdie, FF 2

[0839] £7mOntée ,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust for the same operating conditions at which £7raiti ,saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0840] £7mOntée jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust for the same operating conditions at which EIialenti ,Sai is calculated, if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0841] ^ / approach ,saf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 30% of available thrust under the same operating conditions to which EIiaienti >Saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel; and

[0842] ^ / approach ,ff is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 30% of available thrust for the same operating conditions to which EIiaienti >Saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0843] the nvPM emission index ratio at idle / rich cruise of the gas turbine engine is less than 1; and

[0844] The gas turbine engine is designed to supply fuel including sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles.

[0845] The nvPM emission index ratio at idle / rich cruise can be as defined above in relation to the thirty-fifth aspect.

[0846] According to a forty-second aspect, a method of operating the gas turbine engine of any one or more of the thirty-fifth, thirty-sixth, thirty-seventh, thirty-eighth, thirty-ninth, fortieth, or forty-first aspect is provided, the method comprising supplying fuel comprising a sustainable aviation fuel to the plurality of fuel spray nozzles.

[0847] According to a forty-third aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0848] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles,in which the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; where: ,

[0849] A depleted cruise nvPM emission index ratio can be defined as:

[0850] [Math. 86] FI t cruise (impoverished), SAF J cruise (impoverished), FF

[0851] where:

[0852] (impoverished), saf can be defined as:

[0853] [Math.87] ■^^maxTO. SaA" mounted, SAF 2

[0854] E / eroisière (impoverished) ,ff can be defined as:

[0855] [Math.88] ■^■^maxTQ, II + -Çf montée, FF

[0856] £ / maxTo,sAF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 100% of available thrust for the given operating conditions and if a fuel supplied to the plurality of fuel spray nozzles includes a sustainable aviation fuel;

[0857] EImontée saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust for given operating conditions and if a fuel supplied to the plurality of fuel spray nozzles includes a sustainable aviation fuel;

[0858] £ / maxTo,FF is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 100% of available thrust under given operating conditions if a fuel supplied to the plurality of fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0859] S / ascent jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust for given operating conditions if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0860] the nvPM emission index ratio in depleted cruise is greater than 1; and

[0861] The method includes supplying fuel comprising a sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles.

[0862] The nvPM emission index ratio in depleted cruise can be as defined above in relation to the thirty-fifth aspect.

[0863] A ratio of nvPM emissions index at idle-PMD can be defined as:

[0864] [Math. 89] maxTO

[0865] where:

[0866] ^ / idle,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 7% of available thrust for given operating conditions, or under other different operating conditions, and if a fuel supplied to the plurality of fuel spray nozzles includes a sustainable aviation fuel;

[0867] £ / maxTo,sAF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 100% of available thrust for the same operating conditions at which EIialeaü Saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0868] EIialenti >FF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 7% of available thrust under the same operating conditions at which EIialenti >Saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and

[0869] £7maxTo,FF is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 100% of available thrust under the same operating conditions to which EIialenti >Saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and in which

[0870] The NVPM-PMD emission index ratio of the gas turbine engine may be less than 1.

[0871] The nvPM idle-PMD emission index ratio can be as defined above in relation to the thirty-fifth aspect.

[0872] A depleted cruise nvPM / PMD emission index ratio can be defined as:

[0873] [Math.90] cruise and impoverished}. Saf / ff m3XT O.SAF ■^-croisiêr^.impoverished^F / ^ IIlâxTO FF

[0874] where:

[0875] E / cruise (impoverished), SAF can be defined as:

[0876] [Math.91] ■FJnraxTO. SAld” -FLnontée, SAP 2

[0877] E / eroisière (impoverished) ,ff can be defined as:

[0878] [Math.92] ■E-EnaxTQ, pp+ EI^jppp

[0879] £7maxTo,sAF is the nvPM emissions index corrected for system losses in mg / kg of the gas turbine engine when operating at about 100% of available thrust under given operating conditions, or under other different operating conditions, and if a fuel supplied to the plurality of fuel spray nozzles includes a sustainable aviation fuel;

[0880] EImontée saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust for the same operating conditions at which £7maxTo,sAF is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0881] £7maxTo,FF is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 100% of available thrust under the same operating conditions at which £7maxTo,sAF is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and

[0882] -EImontée jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 85% of available thrust for the same operating conditions to which £' / maXTo,sAF is calculated, if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and in which

[0883] The NVPM / PMD emission index ratio of the gas turbine engine in lean cruise may be less than 1.

[0884] The ratio of nvPM emissions index in depleted cruise / PMD can be as defined above in relation to the thirty-fifth aspect.

[0885] A nvPM emission index ratio at idle / lean cruise can be defined as:

[0886] [Math.93] cruisefappauTO^

[0887] where:

[0888] Ê / raienti ,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 7% of available thrust for given operating conditions, or under other different operating conditions, and if a fuel supplied to the plurality of fuel spray nozzles includes a sustainable aviation fuel;

[0889] ^ / cruise (impoverished) ,saf can be defined as:

[0890] [Math.94] ^■fmàxTO, Sàf+ montée, SAP 2

[0891] S / maxTo.sAF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 100% of available thrust for the same operating conditions at which ELaienti ,saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0892] EImontée saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust for the same operating conditions at which EIialeaü jSaf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0893] EIialenti >FF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 7% of available thrust for the same operating conditions at which EIialenti ,Sai is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0894] (depleted) ,ff can be defined as:

[0895] [Math.95] ^rnàxTO. II + ■^ / rise, FF 2

[0896] ÆZmaxTojF is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 100% of available thrust under the same operating conditions at which EIaienti jSaf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and

[0897] £7montée jf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 85% of available thrust under the same operating conditions at which EIialenti ,Sai is calculated, if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and in which

[0898] The NVPM emission index ratio at idle / lean cruise of the gas turbine engine may be less than 1.

[0899] The nvPM emission index ratio at idle / lean cruise can be as defined above in relation to the thirty-fifth aspect.

[0900] A rich cruise nvPM emission index ratio can be defined as:

[0901] [Math.96] Efcroisière(rich), SAF El croLsièrefiidie), FF

[0902] where:

[0903] ^ / cruise (rich) ,saf can be defined as:

[0904] [Math.97] ^^uphill. SAF^” ^Uppraclie, SAF 2

[0905] Elcroisière (rich) ,ff can be defined as:

[0906] [Math.98] -^ascent, fU"^approach, FF 2

[0907] EImontée saf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at about 85% of available thrust for given operating conditions, or under other different operating conditions, and if a fuel supplied to the plurality of fuel spray nozzles includes a sustainable aviation fuel;

[0908] ^ / approach ,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 30% of available thrust for the same operating conditions to which ^ascent,saf is calculated if a fuel supplied to the plurality of fuel spray nozzles includes a durable aviation fuel;

[0909] S / montée jf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 85% of available thrust under the same operating conditions at which EImontée >Saf is calculated, if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and

[0910] Ê / approach ,ff is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 30% of available thrust under the same operating conditions at which ^ / climb ,saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and

[0911] in which the rich cruise nvPM emission index ratio of the gas turbine engine may be less than 1.

[0912] The rich cruise nvPM emission index ratio can be as defined above in relation to the thirty-fifth aspect.

[0913] A rich cruise nvPM / PMD emission index ratio can be defined as:

[0914] [Math.99] E^ctoisièrichel.SAv / EI maxTO F SLoisieraîridieAFF^^

[0915] where:

[0916] E / cruise (rich), saf can be defined as:

[0917] [Math. 100] Elmontée, SaU- -^Uproche, SAF 2

[0918] and Elcroisière (rich) jf can be defined as:

[0919] [Math. 101] ^7 ascent, Fld- ^The approach, FF 2

[0920] and where:

[0921] Æ' / ascent ,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust under given operating conditions, or under other different operating conditions, and if a fuel supplied to the plurality of fuel spray nozzles includes a sustainable aviation fuel;

[0922] ^ / approach ,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 30% of available thrust for the same operating conditions to which ^ascent,saf is calculated if a fuel supplied to the plurality of fuel spray nozzles includes a durable aviation fuel;

[0923] S / rise jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust for the same operating conditions at which EIrise ,Sai is calculated, if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0924] The approach, ff is the nvPM emissions index corrected for losses in the system in mg / kg of the gas turbine engine when operating at about 30% of available thrust for the same operating conditions at which E / climb,saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0925] £ / maxTo,sAF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine

[0926] when operating at approximately 100% of the available thrust under the same operating conditions at which the Saf rise is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel; and

[0927] ÆZmaxTojF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 100% of available thrust for the same operating conditions at which EImontée Saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0928] and in which the rich cruise nvPM / PMD emission index ratio of the gas turbine engine may be less than 1.

[0929] The rich cruise nvPM / PMD emission index ratio can be as defined above in relation to the thirty-fifth aspect.

[0930] A rich idle / cruise nvPM emission index ratio can be defined as:

[0931] [Math. 102]

[0932] where:

[0933] EIialenti saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 7% of available thrust for given operating conditions, or under other different operating conditions, and if a fuel supplied to the plurality of fuel spray nozzles includes a sustainable aviation fuel (SAF);

[0934] Ί / raienti jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 7% of available thrust for the same operating conditions at which EIaienti jSaf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0935] ^ / cruise (rich) ,saf can be defined as:

[0936] [Math. 103] ^Uphill. SAF^ ^ / approach, SAF 2

[0937] E / cruise (rich) jF can be defined as:

[0938] [Math. 104] ^^JF climb^- approach, FF 2

[0939] EImontée saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust for the same operating conditions at which EIialeaü Saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0940] EImontée jFF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust for the same operating conditions at which EIaienti Saf is calculated, if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0941] £7approach,saf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 30% of available thrust under the same operating conditions at which £7raiti,saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel; and

[0942] Ê / approach jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 30% of available thrust for the same operating conditions at which EIiaienti ,Sai is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0943] and in which the nvPM emission index ratio at idle / rich cruise of the gas turbine engine may be less than 1.

[0944] The nvPM emission index ratio at idle / rich cruise can be as defined above in relation to the thirty-fifth aspect.

[0945] According to a forty-fourth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0946] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles,in which the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; where: ,

[0947] an nvPM-PMD emission index ratio can be defined as:

[0948] [Math. 105] ^ / El maxTO.SAF Ë'Iraiœti,FF / El maxTOpF

[0949] where:

[0950] Æ' / raienti ,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 7% of available thrust for given operating conditions if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0951] £7maxTo,sAF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 100% of available thrust for the same operating conditions at which jSaf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0952] Æ' / raienti jf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 7% of available thrust under the same operating conditions at which £7raiti ,saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and

[0953] S / maxTojF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 100% of thrust available for the same operating conditions under which EIalenti >Saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0954] and the ratio of the nvPM at idle to PMD emission index of the gas turbine engine may be less than 1; and

[0955] The method includes supplying fuel comprising a sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles.

[0956] The nvPM idle-PMD emission index ratio can be as defined above in relation to the thirty-fifth aspect.

[0957] According to a forty-fifth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0958] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles,in which the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; where: ,

[0959] A depleted cruise nvPM / PMD emission index ratio can be defined as:

[0960] [Math. 106] ^^croîMèrefappauyri^. SAf / ^ rnaxTO SAF ® cruise <appauvrie>.FF / Ê / maxfO pp

[0961] where:

[0962] EZcroisière (impoverished),saf can be defined as:

[0963] [Math. 107] ^■fmaxTO, SAEH” ■ELriontée. gAF 2

[0964] E / eroisière (impoverished) ,ff can be defined as:

[0965] [Math. 108] ■F^maxTO, Fp+ ^montée, FF 2

[0966] £ / maxTo,sAF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 100% of available thrust for given operating conditions if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0967] £7 rise,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust under the same operating conditions at which is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0968] £7maxTo,FF is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 100% of available thrust under the same operating conditions at which £7maxTo,sAF is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and

[0969] £7montée jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust for the same operating conditions at which £7maxTo,sAF is calculated, if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0970] and the NVPM / PMD emission index ratio of the gas turbine engine in lean cruise may be less than 1; and

[0971] The process includes supplying fuel comprising a sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles.

[0972] The ratio of nvPM emissions index in depleted cruise / PMD can be as defined above in relation to the thirty-fifth aspect.

[0973] According to a forty-sixth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0974] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; where:

[0975] A ratio of the nvPM emission index at idle / lean cruise can be defined as:

[0976] [Math. 109] ÊUonl,.FF / Elcroigiè

[0977] where:

[0978] Æ' / raienti ,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 7% of available thrust for given operating conditions if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0979] ^ / cruise (impoverished) ,saf can be defined as:

[0980] [Math. 110] ElmaxTO, SAP+ ^Aiiantée, SAP 2

[0981] £7maxTo,sAF is the nvPM emissions index adjusted for losses of the system in mg / kg of the gas turbine engine when operating at about 100% of available thrust for the same operating conditions at which EIialeaü jSaf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0982] EImontée saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust for the same operating conditions at which EIialeaü Saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[0983] Æ' / raienti jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 7% of available thrust for the same operating conditions at which EIialenti ,Sai is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0984] E / cruise (impoverished),ff can be defined as:

[0985] [Math. 111] ^fmaxTO, II-*- FF

[0986] £ / maxTo,FF is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 100% of available thrust under the same operating conditions at which EIaienti Saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and

[0987] £7montée jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust under the same operating conditions at which EIaienti Saf is calculated, if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[0988] the nvPM emission index ratio at idle / lean cruise of the gas turbine engine is less than 1; and

[0989] The process includes supplying fuel comprising a sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles.

[0990] The nvPM emission index ratio at idle / lean cruise can be as defined above in relation to the thirty-fifth aspect.

[0991] According to a forty-seventh aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0992] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles,in which the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; where: ,

[0993] A rich cruise nvPM emission index ratio can be defined as:

[0994] [Math. 112] PT ■ 1 cruise (luxury), SAP pr 1 croLsière(richel FF

[0995] where:

[0996] ^ / cruise (rich) ,saf can be defined as:

[0997] [Math. 113] ^Lintée, SAp4 E ^approach, SAF 2

[0998] ^ / cruise (rich) jf can be defined as:

[0999] [Math. 114] ^^Ascent. FF-^Approach. FF

[1000] EImontée saf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at about 85% of available thrust for given operating conditions if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[1001] ^fapproach ,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 30% of available thrust for the same operating conditions at which ^ / climb ,saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[1002] Ί / rise jf is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 85% of available thrust under the same operating conditions at which EIrise Saf is calculated, if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel; and

[1003] ^ / approach jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 30% of available thrust for the same operating conditions at which ^ / climb ,saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[1004] the rich cruise nvPM emission index ratio of the gas turbine engine is less than 1; and

[1005] The process includes supplying fuel comprising a sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles.

[1006] The rich cruise nvPM emission index ratio can be as defined above in relation to the thirty-fifth aspect.

[1007] According to a forty-eighth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[1008] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of spray nozzles fuel comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a fuel flow rate greater than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; wherein:

[1009] A rich cruise nvPM / PMD emission index ratio can be defined as:

[1010] [Math. 115] B / maxTO SAF B^ctoisierefrichéJFF / B / pp

[1011] where:

[1012] E / cruise (rich), saf can be defined as:

[1013] [Math. 116] f^montée, SAF^- E / approche. SAF 2

[1014] and Elcroisière (rich) ff can be defined as:

[1015] [Math. 117] ^fhlOlltée.FEd” The approach, FF 2

[1016] and where:

[1017] EImontée saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust for given operating conditions and if a fuel supplied to the plurality of fuel spray nozzles includes a sustainable aviation fuel;

[1018] ^ / approach ,saf is the nvPM emissions index corrected for system losses in mg / kg of the gas turbine engine when operating at about 30% of available thrust for the same operating conditions at which EImontée ,saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[1019] EImontée ff is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 85% of available thrust under the same operating conditions to which EImontée ,saf is calculated, if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[1020] ^ / approach jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 30% of available thrust for the same operating conditions at which ^climb,saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[1021] £ / maxTo,sAF is the system loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 100% of available thrust under the same operating conditions at which Æ / climb,saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel; and

[1022] £7maxTo,FF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 100% of available thrust for the same operating conditions at which EImontée ,Sai is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[1023] the rich cruise nvPM / PMD emission index ratio of the gas turbine engine is less than 1; and

[1024] The process includes supplying fuel comprising a sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles.

[1025] The rich cruise nvPM / PMD emission index ratio may be as defined above in relation to the thirty-fifth aspect.

[1026] According to a forty-ninth aspect, a method of operating a gas turbine engine is provided, the gas turbine engine comprising:

[1027] a combustor device, comprising a combustion chamber and a plurality of fuel spray nozzles designed to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustor device can operate under a condition in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with fuel at a higher fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein a ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of spray nozzles of fuel in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; where:

[1028] A ratio of the nvPM emissions index at idle / rich cruise can be defined as:

[1029] [Math. 118]

[1030] where:

[1031] Æ' / raienti ,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 7% of available thrust for given operating conditions and if a fuel supplied to the plurality of fuel spray nozzles includes sustainable aviation fuel (SAF);

[1032] Æ' / raienti jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 7% of available thrust for the same operating conditions at which EIaienti jSaf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[1033] ^ / cruise (rich) ,saf can be defined as:

[1034] [Math. 119] Annotated; SAfF ^approach, SAF 2

[1035] Elcroisière (rich) ,ff can be defined as:

[1036] [Math. 120] ^7 JfF El ascent approach. FF 2

[1037] EImontée saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust for the same operating conditions at which E / raienti ,saf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a sustainable aviation fuel;

[1038] EImontée jFF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 85% of available thrust under the same operating conditions at which EIialenti ,Sai is calculated, if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[1039] Ê / approach ,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at approximately 30% of available thrust under the same operating conditions to which EIiaienti ,Sai is calculated if a fuel supplied to the plurality of fuel spray nozzles includes a sustainable aviation fuel; and

[1040] ^ / approach jf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine when operating at about 30% of available thrust for the same operating conditions at which EIialeaü jSaf is calculated if a fuel supplied to the plurality of fuel spray nozzles comprises a fossil-based hydrocarbon fuel;

[1041] the nvPM emission index ratio at idle / rich cruise of the gas turbine engine is less than 1; and

[1042] The process includes supplying fuel comprising a sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles.

[1043] The ratio of the nvPM emission index at idle / rich cruise can be as defined above in relation to the thirty-fifth aspect.

[1044] In any one of the aspects defined above, any one or more of the emission indices of the gas turbine engine may be defined as follows:

[1045] S / maxTo may be in the range of 0.00893 to 4.72 mg / kg and preferably in the range of 0.0100 to 4.33 mg / kg and more preferably in the range of 0.0111 to 3.94 mg / kg.

[1046] S / maxTo may be in the range of 0.767 to 4.72 mg / kg and preferably in the range of 0.863 to 4.33 mg / kg and more preferably in the range of 0.959 to 3.94 mg / kg.

[1047] S / maxTo may be in the range of 0.00893 to 0.0809 mg / kg and preferably in the range of 0.0100 to 0.0741 mg / kg and more preferably in the range of 0.0111 to 0.0674 mg / kg.

[1048] £7maxTo,sAF may be in the range of 0.00893 to 3.05 mg / kg and preferably in the range of 0.0100 to 2.80 mg / kg and more preferably in the range of 0.0111 to 2.54 mg / kg.

[1049] £'7maxTo,sAF may be in the range of 0.767 to 3.05 mg / kg and preferably in the range of 0.863 to 2.80 mg / kg and more preferably in the range of 0.959 to 2.54 mg / kg-

[1050] £7maxTo,sAF may be in the range of 0.00893 to 0.0523 mg / kg and preferably in the range of 0.0100 to 0.0479 mg / kg and more preferably in the range of 0.0111 to 0.0436 mg / kg.

[1051] £'7maxTo,sAF may be in the range of 0.00893 to 4.71 mg / kg and preferably in the range of 0.0100 to 4.32 mg / kg and more preferably in the range of 0.0111 to 3.93 mg / kg.

[1052] £'7maxTo,sAF may be in the range of 0.767 to 4.71 mg / kg and preferably in the range of 0.863 to 4.32 mg / kg and more preferably in the range of 0.959 to 3.93 mg / kg-

[1053] £ / maxTo,sAF may be in the range of 0.00893 to 0.0808 mg / kg and preferably in the range of 0.0100 to 0.0740 mg / kg and more preferably in the range of 0.0111 to 0.0673 mg / kg.

[1054] £7maxTo,FF may be in the range of 0.0138 to 4.72 mg / kg and preferably in the range of 0.0155 to 4.33 mg / kg and more preferably in the range of 0.0172 to 3.94 mg / kg.

[1055] £7maxTo,FF may be in the range of 1.18 to 4.72 mg / kg and preferably in the range of 1.33 to 4.33 mg / kg and more preferably in the range of 1.48 to 3.94 mg / kg.

[1056] £7maxTo,FF may be in the range of 0.0138 to 0.0809 mg / kg and preferably in the range of 0.0155 to 0.0741 mg / kg and more preferably in the range of 0.0172 to 0.0674 mg / kg.

[1057] £7 rise may be in the range of 0.00438 to 2.30 mg / kg and preferably in the range of 0.00493 to 2.11 mg / kg and more preferably in the range of 0.00548 to 1.92 mg / kg-

[1058] £7 rise may be in the range of 0.460 to 2.30 mg / kg and preferably in the range of 0.517 to 2.11 mg / kg and more preferably in the range of 0.575 to 1.92 mg / kg.

[1059] £7 rise may be in the range of 0.00438 to 0.0221 mg / kg and preferably in the range of 0.00493 to 0.0202 mg / kg and more preferably in the range of 0.00548 to 0.0184 mg / kg.

[1060] £7montée ,saf may be in the range of 0.00438 to 1.09 mg / kg and preferably in the range of 0.00493 to 0.999 mg / kg and more preferably in the range of 0.00548 to 0.909 mg / kg.

[1061] £7montée ,saf may be in the range of 0.460 to 1.09 mg / kg and preferably in the range of 0.517 to 0.999 mg / kg and more preferably in the range of 0.575 to 0.909 mg / kg.

[1062] ElmOI1tée ,saf may be in the range of 0.00438 to 0.0105 mg / kg and preferably in the range of 0.00493 to 0.00959 mg / kg and more preferably in the range of 0.00548 to 0.00872 mg / kg.

[1063] Ê / montée ,saf may be in the range of 0.00438 to 2.29 mg / kg and preferably in the range of 0.00493 to 2.10 mg / kg and more preferably in the range of 0.00548 to 1.91 mg / kg.

[1064] E / mount, saf can be in the range of 0.460 to 2.29 mg / kg and preferably in the range of 0.517 to 2.10 mg / kg and more preferably in the range of 0.575 to 1.91 mg / kg-

[1065] S / montée ,saf may be in the range of 0.00438 to 0.0220 mg / kg and preferably in the range of 0.00493 to 0.0201 mg / kg and more preferably in the range of 0.00548 to 0.0183 mg / kg.

[1066] Elmontée jj may be in the range of 0.00923 to 2.30 mg / kg and preferably in the range of 0.0103 to 2.11 mg / kg and more preferably in the range of 0.0115 to 1.92 mg / kg.

[1067] Ê / montée jf may be in the range of 0.969 to 2.30 mg / kg and preferably in the range of 1.09 to 2.11 mg / kg and more preferably in the range of 1.21 to 1.92 mg / kg.

[1068] £7montée ,ff may be in the range of 0.00923 to 0.0221 mg / kg and preferably in the range of 0.0103 to 0.0202 mg / kg and more preferably in the range of 0.0115 to 0.0184 mg / kg.

[1069] ^ / approach may be in the range of 0.337 to 12.6 mg / kg and preferably in the range of 0.379 to 11.6 mg / kg and more preferably in the range of 0.421 to 10.5 mg / kg.

[1070] ^ / approach may be in the range of 0.571 to 9.89 mg / kg and preferably in the range of 0.643 to 9.07 mg / kg and more preferably in the range of 0.714 to 8.25 mg / kg.

[1071] ^ / approach, saf can be in the range of 0.337 to 1.94 mg / kg and preferably in the range of 0.379 to 1.78 mg / kg and more preferably in the range of 0.421 to 1.62 mg / kg-

[1072] ^ / approach, saf can be in the range of 0.571 to 1.52 mg / kg and preferably in the range of 0.643 to 1.40 mg / kg and more preferably in the range of 0.714 to 1.27 mg / kg-

[1073] ^ / approach, saf can be in the range of 0.337 to 12.5 mg / kg and preferably in the range of 0.379 to 11.5 mg / kg and more preferably in the range of 0.421 to 10.4 mg / kg-

[1074] ^ / approach, saf can be in the range of 0.571 to 9.88 mg / kg and preferably in the range of 0.643 to 9.06 mg / kg and more preferably in the range of 0.714 to 8.24 mg / kg-

[1075] ^ / approach jf can be in the range of 2.19 to 12.6 mg / kg and preferably in the range of 2.47 to 11.6 mg / kg and more preferably in the range of 2.74 to 10.5 mg / kg.

[1076] ^ / approach jf can be in the range of 3.72 to 9.89 mg / kg and preferably in the range of 4.18 to 9.07 mg / kg and more preferably in the range of 4.65 to 8.25 mg / kg.

[1077] EZraienü may be in the range of 0.0525 to 1.55 mg / kg and preferably in the range of 0.0591 to 1.43 mg / kg and more preferably in the range of 0.0657 to 1.30 mg / kg.

[1078] Ê / raienti may be in the range of 0.0858 to 1.55 mg / kg and preferably in the range of 0.0966 to 1.43 mg / kg and more preferably in the range of 0.107 to 1.30 mg / kg.

[1079] ^ / slowed down can be in the range of 0.0525 to 1.01 mg / kg and preferably in the range of 0.0591 to 0.925 mg / kg and more preferably in the range of 0.0657 to 0.841 mg / kg-

[1080] Ê / raienti ,saf may be in the range of 0.0525 to 0.149 mg / kg and preferably in the range of 0.0591 to 0.137 mg / kg and more preferably in the range of 0.0657 to 0.124 mg / kg.

[1081] Ê / raienti ,saf may be in the range of 0.0858 to 0.149 mg / kg and preferably in the range of 0.0966 to 0.137 mg / kg and more preferably in the range of 0.107 to 0.124 mg / kg.

[1082] Ê / raienti ,saf may be in the range of 0.0525 to 0.0967 mg / kg and preferably in the range of 0.0591 to 0.0886 mg / kg and more preferably in the range of 0.0657 to 0.0806 mg / kg.

[1083] Ê / raienti ,saf may be in the range of 0.0525 to 1.54 mg / kg and preferably in the range of 0.0591 to 1.42 mg / kg and more preferably in the range of 0.0657 to 1.29 mg / kg.

[1084] Ê / raienti ,saf may be in the range of 0.0858 to 1.54 mg / kg and preferably in the range of 0.0966 to 1.42 mg / kg and more preferably in the range of 0.107 to 1.29 mg / kg-

[1085] Ê / raienti ,saf may be in the range of 0.0525 to 1.00 mg / kg and preferably in the range of 0.0591 to 0.924 mg / kg and more preferably in the range of 0.0657 to 0.840 mg / kg.

[1086] Ê / raienti jf may be in the range of 0.548 to 1.55 mg / kg and preferably in the range of 0.617 to 1.43 mg / kg and more preferably in the range of 0.686 to 1.30 mg / kg.

[1087] Ê / raienti ,ff may be in the range of 0.896 to 1.55 mg / kg and preferably in the range of 1.00 to 1.43 mg / kg and more preferably in the range of 1.12 to 1.30 mg / kg.

[1088] Ê / raienti jf may be in the range of 0.548 to 1.01 mg / kg and preferably in the range of 0.617 to 0.925 mg / kg and more preferably in the range of 0.686 to 0.841 mg / kg.

[1089] The following statements may apply to any one of the first to forty-ninth aspects defined above:

[1090] The ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles may be in the range of 1:3 to 1:4, or preferably in the range of 1:3.5 to 1:4.

[1091] The first subset of fuel spray nozzles may comprise between 1 and 10 fuel spray nozzles. The first subset of fuel spray nozzles may comprise between 3 and 5 fuel spray nozzles.

[1092] The second subset of fuel spray nozzles may comprise between 10 and 25 fuel spray nozzles. The second subset The fuel spray nozzle assembly can include between 13 and 20 fuel spray nozzles. The second subset of fuel spray nozzles can include between 13 and 17 fuel spray nozzles.

[1093] The combustor device may include one or more igniters.

[1094] Each of the first subset of fuel spray nozzles may be located closer to one or more of the respective igniters than the second subset. In addition or alternatively, one or more of the igniters may be arranged diametrically opposite another or more of the igniters.

[1095] The fuel supplied to the combustor device may include a %SAF in the range of 50% to 100%. The fuel supplied to the combustor device may include a %SAF in the range of 70% to 100%. The fuel supplied to the combustor device may include a %SAF in the range of 90% to 100%.

[1096] Any one of the above-disclosed features in relation to one aspect may be combined with the feature of another aspect unless they are mutually exclusive.

[1097] As stated elsewhere herein, this description may apply to any relevant configuration of a gas turbine engine. Such a gas turbine engine may be, for example, a turbofan gas turbine engine, an open-rotor gas turbine engine (in which the propeller is not enclosed by a nacelle), a turboprop engine, or a turbojet engine. Any such engine may or may not be equipped with an afterburner. Such a gas turbine engine may be, for example, designed for land-based or marine power generation applications.

[1098] A gas turbine engine conforming to any aspect of this description may include an engine core comprising a turbine, a combustion chamber, a compressor, and a core shaft connecting the turbine to the compressor. Such a gas turbine engine may include a fan (having fan blades). Such a fan may be located upstream of the engine core. Alternatively, in some examples, the gas turbine engine may include a fan located downstream of the engine core, for example, in the case where the gas turbine engine is an open-rotor or turboprop engine (in which case the fan may be called a propeller).

[1099] An engine according to this disclosure may be a gas turbine engine. Such an engine may be a direct-drive, turbofan engine in which the blower is directly connected to the blower drive turbine, for example without a reduction gear, via a core shaft. In such a direct-drive, turbofan engine, the blower may be said to rotate at the same rotational speed as the blower drive turbine. Strictly by way of example, the The blower drive turbine can be a first turbine, the core shaft can be a first core shaft, and the gas turbine engine can further include a second turbine and a second core shaft connecting the second turbine to the compressor. The second turbine, compressor, and second core shaft can be arranged to rotate at a higher speed than the first core shaft. In such an arrangement, the second turbine can be positioned axially upstream of the first turbine.

[1100] An engine according to this disclosure may be a geared gas turbine engine. In such an arrangement, the engine has a blower that is driven via a reduction gear. Accordingly, such a gas turbine engine may include a reduction gear that receives an input from the core shaft and delivers a drive to the blower so as to drive the blower at a rotational speed lower than that of the core shaft. The input to the reduction gear may be directly from the core shaft, or indirectly from the core shaft, for example, via a straight shaft and / or gear. The core shaft may connect the turbine and the compressor, such that the turbine and the compressor rotate at the same speed (with the blower rotating at a lower speed).

[1101] The gas turbine engine as described and / or claimed herein may have any suitable general architecture. For example, the gas turbine engine may have any desired number of shafts connecting turbines and compressors, for example, one, two, or three shafts. By way of example only, the turbine connected to the core shaft may be a first turbine, the compressor connected to the core shaft may be a first compressor, and the core shaft may be a first core shaft. The engine core may further comprise a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, second compressor, and second core shaft may be arranged to rotate at a higher speed than the first core shaft.

[1102] In such an arrangement, the second compressor can be positioned axially downstream of the first compressor. The second compressor can be arranged to receive (for example, directly, for example via a generally annular conduit) a flow from the first compressor.

[1103] The gearbox can be arranged to be driven by the core shaft that is designed to rotate (for example, during operation) at the lowest rotational speed (for example, the first core shaft in the example above). For example, the gearbox can be arranged to be driven only by the core shaft that is designed to rotate (for example, during operation) at the lowest rotational speed. low (for example, only by the first core tree, and not the second core tree, in the example above). Alternatively, the reducer can be arranged to be driven by any one or more trees, for example the first and / or second trees in the example above.

[1104] The gearbox may be a reduction box (in that the output to the blower has a lower rotational speed than the input from the core shaft). Any type of gearbox may be used. For example, the gearbox may be a 'planetary' or 'star' gearbox, as described in more detail elsewhere in this document. Such a gearbox may be a single stage. Alternatively, such a gearbox may be a compound gearbox, for example a compound planetary gearbox (which may have the input on the sun gear and the output on the ring gear, and thus be called a 'compound star' gearbox), for example with two reduction stages.

[1105] The gearbox can have any desired gear ratio (defined as the speed of rotation of the input shaft divided by the speed of rotation of the output shaft). For example, the gear ratio can be greater than, or can be any of the following: 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, or 4.2. The gear ratio can be an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​can form upper or lower bounds). The gear ratio can be, for example, greater than 2.5, for example in the range of 3.0 to 4.2, or from 3.2 to 3.8. Strictly as an example, the reducer can be a "star" reducer having a gear ratio in the range of 3.1 or 3.2 to 3.8.Strictly as a further example, the reducer may be a "star" reducer with a reduction ratio in the range of 3.0 to 3.1. Strictly as a further example, the reducer may be a "planetary" reducer with a reduction ratio in the range of 3.3 to 3.6, or from 3.6 to 4.2. Strictly as a further example, the reducer may be a compound reducer, for example a compound star reducer, having a gear ratio of, or at least: 4.0, 4.5, 5.0, 5.5, 6.0, 6.5, 7.0, 7.5, 8.0, 8.5, 9.0, 9.5, 10.0, 10.5, 11.0, 11.5, 12.0, 12.5, 13.0, 13.5, 14.0.The gear ratio of a compound reducer, for example a compound star reducer, can be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​can form upper or lower bounds), for example, within the range from 4.0 to 14.0, or from 6.0 to 12.0, or from 8.0 to 10.0. In some arrangements, the gear ratio can be outside these ranges.

[1106] In any gas turbine engine as described and / or claimed herein, fuel of a given composition or mixture is supplied to a chamber of Combustion can be supplied downstream of the blower and compressor(s) relative to the flow path (e.g., axially downstream). For example, the combustor can be directly downstream of (e.g., at the outlet of) the second compressor when a second compressor is supplied. As a further example, the flow at the outlet to the combustor can be supplied at the inlet of the second turbine when a second turbine is supplied. The combustor can also be supplied upstream of the turbine(s).

[1107] The compressor or compressors (for example, the first and second compressors as described above) may comprise any number of stages, for example, multiple stages. Each stage may comprise a row of rotor blades and a row of stator blades, which may be variable stator blades (in that their angle of incidence may be variable). The row of rotor blades and the row of stator blades may be axially offset from each other. For example, the gas turbine engine may be a direct-drive, double-flow gas turbine engine comprising 13 or 14 compressor stages (in addition to the fan). Such an engine may, for example, comprise 3 stages in the first compressor (or "low pressure") and either 10 or 11 stages in the second compressor (or "high pressure").Alternatively, such an engine may, for example, comprise 4 stages in the first compressor (or "low pressure") and 10 stages in the second compressor (or "high pressure"). Alternatively, such an engine may, for example, comprise 7, 8, or 9 stages in a first compressor (or "intermediate pressure") and 5, 6, or 7 stages in the second compressor (or "high pressure"). As a further example, the gas turbine engine may be a "gear-driven" gas turbine engine (in which the fan is driven by a first core shaft via a reduction gearbox) comprising 11, 12, 13, 14, or 15 compressor stages (in addition to the fan). Such an engine may comprise 3 or 4 stages in the first compressor (or "low pressure") and 8, 9, or 10 stages in the second compressor (or "high pressure").As a further example, the gas turbine engine can be a "geared" gas turbine engine having 4 stages in the first compressor (or "low pressure") and 10 stages in the second compressor (or "high pressure").

[1108] The turbine or turbines (for example, the first and second turbines as described above) may comprise any number of stages, for example, multiple stages. Each stage may comprise one row of rotor blades and one row of stator blades. The respective rows of rotor blades and stator blades may be axially offset from each other. In each stage, the row of rotor blades may be arranged downstream of the respective row of stator blades. By way of example, the gas turbine engine may comprise 5, 6, 7, 8, or 9 turbine stages. For example, a gas turbine engine can be a geared gas turbine engine with 5, 6, or 7 turbine stages. Such a geared gas turbine engine can include a second turbine (high-pressure) with 2 stages. Such a geared gas turbine engine can include a first turbine (low-pressure) with 3 or 4 stages. As a further example, a gas turbine engine can be a direct-drive gas turbine engine with a first turbine (low-pressure) with 5, 6, or 7 stages. Such a direct-drive gas turbine engine can include a second turbine (high-pressure) with 2 stages. Alternatively, such an engine can, for example, include a second turbine (intermediate-pressure) with 1, 2, or 3 stages.Such a direct-drive gas turbine engine may also include a third turbine (or "high pressure") having 1, 2 or 3 stages.

[1109] Each fan blade can be defined as having a radial span extending from a foot (or hub) at a radially internal gas-washed location, or a 0% span position, to a tip at a 100% span position. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be less than, or may be any of: 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, 0.30, 0.29, 0.28, 0.27, 0.26, or 0.25. The ratio of the radius of the blower blade at the hub to the radius of the blower blade at the tip can be in an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​can form upper or lower bounds), for example in the range of 0.28 to 0.32, or 0.29 to 0.30.These ratios can commonly be referred to as the hub-to-tip ratio. Strictly as a non-limiting example, the hub-to-tip ratio can be in the range of 0.40 to 0.50, 0.42 to 0.48, or 0.43 to 0.47. Both the hub radius and the tip radius can be measured at the leading edge (or axially forward) portion of the fan blade. The hub-to-tip ratio refers, of course, to the gas-washed portion of the fan blade—that is, the portion radially outside any platform.

[1110] The radius of the fan can be measured between the midline of the engine and the tip of a fan blade at its leading edge. The blower diameter (which can simply be twice the blower radius) can be greater than, and can be any of: 110 cm, 120 cm, 130 cm, 140 cm, 150 cm, 160 cm, 170 cm, 180 cm, 190 cm, 200 cm, 205 cm, 206 cm, 207 cm, 208 cm, 209 cm, 210 cm, 211 cm, 212 cm, 213 cm, 214 cm, 215 cm, 216 cm, 217 cm, 218 cm, 219 cm, 220 cm, 230 cm, 240 cm, 250 cm (approximately 100 inches), 260 cm, 270 cm (approximately 105 inches). inches), 280 cm (about 110 inches), 290 cm (about 115 inches), 300 cm (about 120 inches), 310 cm, 320 cm (about 125 inches), 330 cm (about 130 inches), 340 cm (about 135 inches), 345 cm, 350 cm, 355 cm, 360 cm (about 140 inches), 370 cm (about 145 inches), 380 cm (about 150 inches), 390 cm (about 155 inches), 400 cm, 410 cm (about 160 inches) or 420 cm (about 165 inches). The blower diameter can be within an inclusive range delimited by any two of the values ​​in the preceding sentence (i.e., the values ​​can form upper or lower limits), for example, within the range from 110 cm to 120 cm, 120 cm to 130 cm, 210 cm to 240 cm, 250 cm to 280 cm, 320 cm to 380 cm, or 380 cm to 420 cm. Strictly as a non-limiting example, the blower diameter can be within the range from 170 cm to 180 cm, 190 cm to 200 cm, 200 cm to 210 cm, 210 cm to 220 cm, 220 cm to 230 cm, 290 cm to 300 cm, or 340 cm to 360 cm.

[1111] The rotational speed of the blower may vary during use. Generally, the rotational speed is lower for blowers with a larger diameter.Strictly by way of non-limiting example, the fan speed under cruising conditions may be less than, or may be any of the following: 3500 rpm, 3450 rpm, 3400 rpm, 3350 rpm, 3300 rpm, 3250 rpm, 3200 rpm, 3150 rpm, 3100 rpm, 3050 rpm, 3000 rpm, 2950 rpm, 2900 rpm, 2850 rpm, 2800 rpm, 2750 rpm, 2700 rpm, 2650 rpm, 2600 rpm, 2550 rpm, 2500 rpm, 2450 rpm, 2400 rpm, 2350 rpm, 2300 rpm, 2250 rpm, 2200 rpm, 2150 rpm, 2100 rpm, 2050 rpm, 2000 rpm, 1950 rpm, 1900 rpm, 1850 rpm, 1800 rpm, 1750 rpm, 1700 rpm, 1650 rpm, 1600 rpm, 1550 rpm, 1500 rpm, 1450 rpm, 1400 rpm, 1350 rpm, 1300 rpm, 1250 rpm, 1200 rpm, 1150 rpm, 1 100 rpm, 1050 rpm, 1000 rpm, or 950 rpm.The fan speed under cruising conditions can be within an inclusive range delimited by any two of the values ​​in the preceding sentence (i.e., the values ​​can form upper or lower limits). Strictly as a further non-limiting example, the fan speed under cruising conditions for a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm can be in the range of 2750 to 2900 rpm, 2750 to 2800 rpm, or 2800 to 2900 rpm. Strictly as a further non-limiting example, the fan speed under cruising conditions for a "geared" gas turbine engine having a fan diameter in the range of 210 cm to 230 cm can be in the range of 2500 rpm to 2800 rpm, or 2500 rpm to 2750 rpm.Strictly as a further non-limiting example, the fan speed under cruising conditions for a "geared" gas turbine engine having . A fan diameter in the range of 340 cm to 360 cm can be in the range of 1500 rpm to 1800 rpm, or 1500 rpm to 1700 rpm. Strictly as a further non-limiting example, the fan speed under cruising conditions for a geared gas turbine engine having a fan diameter in the range of 380 cm to 420 cm can be in the range of 950 rpm to 1500 rpm, 950 rpm to 1200 rpm, 950 rpm to 1100 rpm, 950 rpm to 1050 rpm, 950 rpm to 1000 rpm, or 1000 rpm to 1050 rpm. Strictly as a further non-limiting example, the fan speed under cruising conditions for a direct drive engine having a fan diameter in the range of 170 cm to 200 cm can be in the range of 3400 to 4600 rpm, for example 3600 to 4600 rpm, or 3600 to 3900 rpm.Strictly as a further non-limiting example, the fan speed under cruising conditions for a direct-drive engine with a fan diameter in the range of 300 cm² to 350 cm² can be in the range of 1800 to 2800 rpm, or 1950 to 2550 rpm. Strictly as a further non-limiting example, the fan speed under cruising conditions for a direct-drive engine with a fan diameter in the range of 270 cm² to 290 cm² can be in the range of 1800 to 2800 rpm, or 2050 to 2450 rpm. Strictly as a further non-limiting example, the fan speed under cruising conditions for a direct drive engine having a fan diameter in the range of 290 cm to 310 cm can be in the range of 1800 to 2800 rpm, or 2100 to 2500 rpm.

[1112] During operation of the gas turbine engine, the fan (with the associated fan blades) rotates about an axis of rotation. This rotation causes the tip of the fan blade to move at a velocity Utip. The work done by the fan blades on the flow results in a specific enthalpy increase dH of the flow. A fan tip load can be defined by dH / Utip², where dH is the specific enthalpy increase (e.g., the mean specific enthalpy increase 1-D) across the fan and Utip is the (translational) velocity of the fan tip, e.g., at the leading edge of the tip (which can be defined as a fan tip radius at the leading edge multiplied by the angular velocity).The peak fan load under cruise conditions can be greater than, or can be any one of the following: 0.15, 0.16, 0.17, 0.18, 0.19, 0.20, 0.21, 0.22, 0.23, 0.24, 0.25, 0.26, 0.27, 0.28, 0.29, 0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39, or 0.40 (all values ​​being dimensionless). The peak fan load can be within an inclusive range bounded by any two of the values ​​in the preceding sentence. (that is, the values ​​can form upper or lower limits), for example in the range of 0.15 to 0.20, 0.28 to 0.35, 0.29 to 0.35, 0.29 to 0.30, or 0.30 to 0.35 (for example for a geared gas turbine engine).

[1113] Gas turbine engines in accordance with this description may have any desired bypass ratio (BPR), where the bypass ratio may be defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core. In certain arrangements, the derivative ratio under cruising conditions may be greater than, or may be any of the following: 3, 3.5, 4, 4.5, 5, 5.5, 6, 6.5, 7, 7.5, 8, 8.5, 9, 9.5, 10, 10.1, 10.2, 10.3, 10.4, 10.5, 10.6, 10.7, 10.8, 10.9, 11, 11.1, 11.2, 11.3, 11.4, 11.5, 11.6, 11.7, 11.8, 11.9, 12, 12.1, 12.2, 12.3, 12.4, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20.The bypass ratio under cruising conditions may be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower bounds), for example, within the range of 3.5 to 6.5, 4 to 6, 4 to 5, 5 to 6, 12 to 16, 13 to 15, or 13 to 14. Strictly by way of non-limiting example, the bypass ratio under cruising conditions of a direct-drive gas turbine engine as described herein may be within the range of 8 to 11, 8 to 10, 9 to 11, 9 to 10, 10 to 16, 12 to 16, 13 to 15, or 13 to 14. Strictly by way of further non-limiting example, the bypass ratio under cruising conditions of a geared gas turbine engine according to this disclosure may be in the range of 10 to 12, 12 to 15, or 12.5 to 15. The bypass conduit may be at least substantially annular.The bypass duct can be radially external to the reactor block. The radially external surface of the bypass duct can be defined by a nacelle and / or a fan housing.

[1114] The overall pressure ratio (OPR) of a gas turbine engine as described and / or claimed herein may be defined as the ratio of the stagnation pressure at the outlet of the highest pressure compressor (before an inlet to the combustor device) to the stagnation pressure upstream of the blower. By way of non-limiting example, the overall pressure ratio of a gas turbine engine as described and / or claimed herein under cruising conditions may be greater than, or may be any of the following: 25, 30, 31, 32, 33, 34, 35, 36, 37, 38, 39, 40, 41, 42, 43, 44, 45, 46, 47, 48, 49, 50, 51, 52, 53, 54, 55, 56, 57, 58, 59, 60, 65, 70, 75. The overall pressure ratio under cruising conditions may be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower bounds), for example, within the range from 30 to 40 or from 50 to 70.Strictly as an example. By way of example only, the overall pressure ratio under cruising conditions of a geared gas turbine engine with a fan diameter in the range of 110 cm to 120 cm can be in the range of 30 to 40, or 31 to 36. Strictly by way of example only, the overall pressure ratio under cruising conditions of a geared gas turbine engine with a fan diameter in the range of 120 cm to 130 cm can be in the range of 30 to 40, or 31 to 36. Strictly by way of example only, the overall pressure ratio under cruising conditions of a geared gas turbine engine with a fan diameter in the range of 200 cm to 210 cm can be in the range of 40 to 44.Strictly by way of non-limiting example, the overall pressure ratio at cruising conditions of a geared gas turbine engine with a fan diameter in the range of 210 cm to 230 cm can be in the range of 44 to 55. Strictly by way of non-limiting example, the overall pressure ratio at cruising conditions of a geared gas turbine engine with a fan diameter in the range of 340 cm to 360 cm can be in the range of 50 to 60. Strictly by way of non-limiting example, the overall pressure ratio at cruising conditions of a direct-drive gas turbine engine with a fan diameter in the range of 170 cm to 200 cm can be in the range of 35 to 60, or 40 to 50.Strictly by way of non-limiting example, the overall pressure ratio under cruise conditions of a direct-drive gas turbine engine with a fan diameter in the range of 300 cm to 350 cm can be in the range of 35 to 60, or 40 to 50. Strictly by way of non-limiting example, the overall pressure ratio under cruise conditions of a direct-drive gas turbine engine with a fan diameter in the range of 260 cm to 285 cm can be in the range of 35 to 60, or 35 to 50. Strictly by way of further non-limiting example, the overall pressure ratio under cruise conditions for a direct-drive engine with a fan diameter in the range of 270 cm to 290 cm can be in the range of 35 to 60, or 37 to 47.Strictly as a further non-limiting example, the overall pressure ratio under cruising conditions for a direct drive engine having a fan diameter in the range of 290 cm to 310 cm can be in the range of 35 to 60, or 37 to 47.

[1115] The specific thrust of an engine can be defined as the net thrust of the engine divided by the total mass flow rate through the engine. In some examples, specific thrust may depend, for a given thrust condition, on the specific composition of fuel supplied to the combustor device. Under cruise conditions, the specific thrust of an engine described and / or claimed herein may be less than, or may be any of the following: 150 Nkg, 145 Nkg's, 140 Nkg's, 135 Nkg's, 130 Nkg's, 125 Nkg's, 120 Nkg's, 115 Nkg's, 110 Nkg-ls, 105 Nkg-ls, 100 Nkg-ls, 99 Nkg-ls, 98 Nkg-ls, 97 Nkg-ls, 96 Nkg-ls, 95 Nkg-ls, 94 Nkg-ls, 93 Nkg-ls, 92 Nkg-ls, 91 Nkg-ls, 90 Nkg-ls, 89 Nkg-ls, 88 Nkg-ls, 87 Nkg-ls, 86 Nkg-ls, 85 Nkg-ls, 80 Nkg-ls, 75 Nkg's, 70 Nkg's, 65 Nkg's, 60 Nkg⁴s, 55 Nkg⁴s, 50 Nkg⁴s, 45 Nkg⁴s, 40 Nkg⁴s, 35 Nkg⁴s, 30 Nkg⁴s, 25 Nkg⁴s, 20 Nkg⁴s, or 15 Nkg⁴s. The specific thrust under cruising conditions can be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​can form upper or lower limits), for example, in the range from 80 Nkg⁻¹ to 100 Nkg⁻¹s, from 85 Nkg⁻¹ to 100 Nkg⁻¹s, or from 92 Nkg⁻¹s to 100 Nkg⁻¹s. Such engines can be particularly efficient compared to conventional gas turbine engines.Strictly by way of non-limiting example, the specific thrust under cruise conditions of a geared gas turbine engine with a fan diameter in the range of 200 cm to 210 cm can be in the range of 90 Nkg-ls to 98 Nkg-ls, or 92 Nkg-ls to 98 Nkg-ls. Strictly by way of non-limiting example, the specific thrust under cruise conditions of a geared gas turbine engine with a fan diameter in the range of 210 cm to 230 cm can be in the range of 90 Nkg-ls to 100 Nkg-ls, or 95 Nkg-ls to 100 Nkg-ls. Strictly as a non-limiting example, the specific thrust under cruise conditions of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm can be in the range of 70 Nkg-ls to 95 Nkg-ls, 80 Nkg-ls to 95 Nkg-ls, or 85 Nkg-ls to 95 Nkg-ls.Strictly by way of non-limiting example, the specific thrust under cruise conditions of a direct-drive gas turbine engine with a fan diameter in the range of 170 cm to 200 cm can be in the range of 20 Nkg-ls to 90 Nkg-ls, 20 Nkg-ls to 80 Nkg-ls, or 25 Nkg-ls to 70 Nkg-ls. Strictly by way of non-limiting example, the specific thrust under cruise conditions of a direct-drive gas turbine engine with a fan diameter in the range of 300 cm to 350 cm can be in the range of 90 Nkg-ls to 120 Nkg-ls, or 100 Nkg-ls to 115 Nkg-ls. Strictly by way of non-limiting example, the specific thrust under cruise conditions of a direct drive gas turbine engine having a fan diameter in the range of 260 cm to 285 cm can be in the range of 20 Nkg-ls to 120 Nkg-ls, 30 Nkg-ls to 115 Nkg-ls, or 40 Nkg-ls to 115 Nkg-ls.Strictly as a non-limiting example, the specific thrust under cruise conditions of a direct-drive gas turbine engine having a fan diameter in the range of 270 cm to 290 cm can be in the range of 90 Nkg-ls to 120 Nkg-ls, or 95. Nkg-ls to 115 Nkg-ls. Strictly by way of non-limiting example, the specific thrust under cruise conditions of a direct drive gas turbine engine having a fan diameter in the range of 290 cm to 310 cm can be in the range of 95 Nkg-ls to 130 Nkg-ls, or 105 Nkg-ls to 125 Nkg-ls.

[1116] A gas turbine engine as described and / or claimed herein can have any desired maximum thrust. Strictly by way of non-limiting example, a gas turbine as described and / or claimed herein may be capable of producing a maximum thrust of at least, may be any one of the following: 50 kN, 55 kN, 56 kN, 57 kN, 58 kN, 59 kN, 60 kN, 61 kN, 62 kN, 63 kN, 64 kN, 65 kN, 66 kN, 67 kN, 68 kN, 69 kN, 70 kN, 71 kN, 72 kN, 73 kN, 74 kN, 75 kN, 76 kN, 77 kN, 78 kN, 79 kN, 80 kN, 90 kN, 100 kN, 105 kN, 110 kN, 115 kN, 120 kN, 125 kN, 130 kN, 131 kN, 132 kN, 133 kN, 134 kN, 135 kN, 136 kN, 137 kN, 138 kN, 139 kN, 140 kN, 141 kN, 142 kN, 143 kN, 144 kN, 145 kN, 146 kN, 147 kN, 148 kN, 149 kN, 150 kN, 151 kN, 152 kN, 153 kN, 154 kN, 155 kN, 160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250kN, 300kN, 350 kN, 375 kN, 400 kN, 425 kN, 450 kN, 475 kN, 500 kN, 525 kN, 550 kN, 600 kN, 650 kN, or 700 kN.The maximum thrust may be within an inclusive range delimited by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower bounds). Strictly by way of non-limiting example, a gas turbine as described and / or claimed herein may produce a maximum thrust in the range of 50 kN to 85 kN, 57 kN to 78 kN, 60 kN to 73 kN, 60 kN to 70 kN, 65 kN to 150 kN, 105 kN to 150 kN, 155 kN to 170 kN, 330 kN to 420 kN, or 350 kN to 400 kN. Strictly as a non-limiting example, the maximum thrust of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm can be in the range of 130 kN to 160 kN, or 130 kN to 150 kN.Strictly by way of non-limiting example, the maximum thrust of a geared gas turbine engine with a fan diameter in the range of 210 cm to 230 cm can be in the range of 130 kN to 200 kN, or 135 kN to 180 kN, or 135 kN to 170 kN, or 135 kN to 160 kN, or 135 kN to 150 kN, or 135 kN or 145 kN. Strictly by way of non-limiting example, the maximum thrust of a geared gas turbine engine with a fan diameter in the range of 340 cm to 360 cm can be in the range of 370 kN to 500 kN. Strictly as a non-limiting example, the maximum thrust of a direct drive gas turbine engine having a fan diameter in the range of 170 cm to 200 cm can be in the range of 100 kN to 200 kN, 110 kN to 180 kN, or 120 kN to 170 kN.Strictly as a non-limiting example, the maximum thrust of a direct drive gas turbine engine having a fan diameter in the range of 110 cm to 120 cm can be in the range of . 60 kN to 70 kN, or 60 kN to 65 kN. Strictly by way of non-limiting example, the maximum thrust of a direct-drive gas turbine engine with a fan diameter in the range of 120 cm to 130 cm can be in the range of 65 kN to 75 kN, or 67 kN to 73 kN. Strictly by way of non-limiting example, the maximum thrust of a direct-drive gas turbine engine with a fan diameter in the range of 300 cm to 350 cm can be in the range of 250 kN to 600 kN, 270 kN to 550 kN, 300 kN to 500 kN, 300 kN to 400 kN, or 425 kN to 525 kN. Strictly as a non-limiting example, the maximum thrust of a direct drive gas turbine engine having a fan diameter in the range of 260 cm to 285 cm can be in the range of 250 kN to 450 kN, 275 kN to 400 kN, or 275 kN to 375 kN.Strictly by way of non-limiting example, the maximum thrust of a direct-drive gas turbine engine with a fan diameter in the range of 270 cm to 290 cm can be in the range of 250 kN to 500 kN, 275 kN to 400 kN, or 300 kN to 375 kN. Strictly by way of non-limiting example, the maximum thrust of a direct-drive gas turbine engine with a fan diameter in the range of 290 cm to 310 cm can be in the range of 250 kN to 550 kN, 300 kN to 525 kN, or 325 kN to 500 kN. The thrust mentioned above may be the maximum net thrust under typical atmospheric conditions at sea level plus 15 degrees C (ambient pressure of 101.3 kPa, temperature of 30 degrees C), with the engine stationary.

[1117] During operation, the flow temperature at the inlet of the high-pressure turbine can be particularly high. This temperature, which may be called TET, can be measured at the outlet of the combustor device, for example, immediately upstream of the first turbine blade, which itself may be called the nozzle guide blade. In some examples, the TET may depend, for a given thrust condition, on the specific composition of the fuel supplied to the combustor device.

[1118] Under cruising conditions, the TET may be at least, or may be any of the following: 1400 K, 1450 K, 1455 K, 1460 K, 1465 K, 1470 K, 1475 K, 1480 K, 1490 K, 1495 K, 1500 K, 1505 K, 1510 K, 1515 K, 1520 K, 1525 K, 1530 K, 1535 K, 1540 K, 1545 K, 1550 K, 1555 K, 1560 K, 1565 K, 1570 K, 1575 K, 1580 K, 1585 K, 1590 K, 1595 K, 1600 K, 1650 K, 1700 K, or 1750 K. The TET under cruising conditions can be within an inclusive range delimited by any two of the values ​​in the preceding sentence (i.e., the values ​​can form upper or lower bounds). Thus, strictly by way of non-limiting example, the TET under cruising conditions of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 1540 K to 1600 K, or 1570 K to 1590 K. Strictly by way of non-limiting example, the TET under cruise conditions of a geared gas turbine engine having a fan diameter in the range of 210 cm² to 230 cm² may be in the range of 1525 K to 1650 K. Strictly by way of non-limiting example, the TET under cruise conditions of a geared gas turbine engine having a fan diameter in the range of 340 cm² to 360 cm² may be in the range of 1550 K to 1660 K, or 1550 K to 1600 K. Strictly by way of non-limiting example, the TET under cruise conditions of a direct-drive gas turbine engine having a fan diameter in the range of 170 cm to 200 cm can be in the range of 1400 K to 1650 K, or 1425 K to 1625 K.Strictly by way of non-limiting example, the TET under cruise conditions of a direct-drive gas turbine engine with a fan diameter in the range of 300 cm² to 350 cm² can be in the range of 1400 K to 1650 K, 1425 K to 1575 K, 1425 K to 1525 K, or 1475 K to 1550 K. Strictly by way of non-limiting example, the TET under cruise conditions of a direct-drive gas turbine engine with a fan diameter in the range of 260 cm² to 285 cm² can be in the range of 1400 K to 1650 K, or 1425 K to 1625 K. Strictly by way of non-limiting example, the TET under cruise conditions of a direct-drive gas turbine engine with a diameter blower in the range of 270 cm to 290 cm can be in the range of 1400 K to 1600 K, 1425 K to 1575 K, or 1450 K to 1550 K.Strictly by way of non-limiting example, the TET under cruise conditions of a direct-drive gas turbine engine having a fan diameter in the range of 290 cm to 310 cm may be in the range of 1400 K to 1650 K, 1450 K to 1600 K, or 1475 K to 1575 K. The TET under cruise conditions may be in an inclusive range bounded by any two of the TET values ​​in this paragraph (i.e., the values ​​may form upper or lower bounds), for example 1530 K to 1600 K.

[1119] The maximum TET during engine operation may be at least, or may be any one of the following: 1700 K, 1750 K, 1755 K, 1760 K, 1765 K, 1770 K, 1775 K, 1780 K, 1785 K, 1790 K, 1795 K, 1800 K, 1805 K, 1810 K, 1815 K, 1820 K, 1825 K, 1830 K, 1835 K, 1840 K, 1845 K, 1850 K, 1855 K, 1860 K, 1865 K, 1870 K, 1875 K, 1880 K, 1885 K, 1890 K, 1895 K, 1900 K, 1905 K, 1910 K, 1915 K, 1920 K, 1925 K, 1930 K, 1935 K, 1940 K, 1945 K, 1950 K, 1955 K, 1960 K, 1965 K, 1970 K, 1975 K, 1980 K, 1985 K, 1990 K, 1995 K, 2000 K, 2050 K, or 2100 K. The maximum TET during engine operation may be within an inclusive range delimited by any two of the TET values ​​in this paragraph (i.e., the values ​​may form upper or lower limits). Thus, strictly as By way of non-limiting example, the maximum TET of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm can be in the range of 1800 K to 1960 K, 1850 K to 1960 K, 1875 K to 1960 K, 1900 K to 1960 K, or 1900 K to 1950 K. Strictly by way of non-limiting example, the maximum TET of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm can be in the range of 1700 K to 1960 K, 1750 K to 1900 K, 1750 K to 1850 K, or 1750 K to 1800 K. Strictly as a non-limiting example, the maximum TET of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm can be in the range of 1800 K to 1960 K, 1800 K to 1900 K, or 1850 K to 1900 K.Strictly by way of non-limiting example, the maximum TET of a direct-drive gas turbine engine with a fan diameter in the range of 170 cm² to 200 cm² can be in the range of 1700 K to 1950 K, or 1750 K to 1900 K. Strictly by way of non-limiting example, the maximum TET of a direct-drive gas turbine engine with a fan diameter in the range of 300 cm² to 350 cm² can be in the range of 1700 K to 1950 K, 1750 K to 1900 K, or 1775 K to 1900 K. Strictly by way of non-limiting example, the maximum TET of a direct-drive gas turbine engine with a fan diameter in the range of 260 cm² to 285 cm² can be in the range of 1700 K to 1950 K, or from 1750 K to 1900 K.Strictly by way of non-limiting example, the maximum TET of a direct-drive gas turbine engine with a fan diameter in the range of 270 cm² to 290 cm² can be in the range of 1750 K to 1950 K, or 1800 K to 1900 K. Strictly by way of non-limiting example, the maximum TET of a direct-drive gas turbine engine with a fan diameter in the range of 290 cm² to 310 cm² can be in the range of 1750 K to 2000 K, or 1800 K to 1950 K. The maximum TET can occur, for example, under a high thrust condition, such as a maximum takeoff thrust (PMD) condition.

[1120] A gas turbine engine as described and / or claimed herein may have any desired high-pressure compressor size, also known as core size. The core size defines the size of the engine core. The engine core size may be defined as:

[1121] [Math. 121] Heart size = ùi2~p-

[1122] Where = the mass flow rate, in pounds per second, of the air at the inlet of the high-pressure compressor, = the temperature, in Kelvin, of the air at the outlet of the high-pressure compressor, and = the pressure, in pounds inches per second squared per inch square, of the air at the outlet of the high-pressure compressor. A unit of core size is therefore expressed as:

[1123] [Math. 122] i s*K 2 • po

[1124] Under cruising conditions, the heart size may be at least, or may be any one of the following: 2, 2.5, 3, 3.5, 4, 4.5, 5, 5.5, 6, 6.5, 7, 7.5, 8, 8.5, 9, 9.5, 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5, or 20. The heart size under cruising conditions may be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower bounds). lower). Thus, strictly as a non-limiting example, the core size under cruising conditions of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm can be in the range of 3 to 8, 3.5 to 7, 4 to 6.5, 4.5 to 6.5, or 5 to 6.Strictly by way of non-limiting example, the core size under cruise conditions of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm can be in the range of 7 to 15, 8 to 14, 9 to 13, 10 to 12.5 or 11 to 12. Strictly by way of non-limiting example, the core size under cruise conditions of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm can be in the range of 5 to 13, 6.5 to 12.5, 7.5 to 11.5, 8.5 to 10.5, or 9 to 10.

[1125] Strictly by way of non-limiting example, the core size under cruise conditions of a direct-drive gas turbine engine having a fan diameter in the range of 110 cm to 130 cm may be in the range of 3 to 8, 3.5 to 7, 4 to 6.5, 4.5 to 6.5, or 5 to 6. Strictly by way of non-limiting example, the core size under cruise conditions of a direct-drive gas turbine engine having a fan diameter in the range of 170 cm to 200 cm may be in the range of 3 to 12, 3.5 to 11, 3.5 to 10, 3.5 to 9, 3.5 to 8, 4 to 7, 4 to 6.5, 4.5 to 6.5, or 5 to 6. Strictly by way of non-limiting example As a non-limiting example, the core size under cruising conditions of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 350 cm can be in the range of 9 to 17, 10 to 16, 11 to 16, 13 to 14.5 or 13 to 14.Strictly by way of non-limiting example, the core size under cruise conditions of a direct drive gas turbine engine having a fan diameter in the range of 260 cm to 285 cm can be in the range of 3 to 12, 3.5 to 11, 3.5 to 10, 3.5 to 9, 3.5 to 8, 4 to 7, 4 to 6.5, 4.5 to 6.5, or 5 to 6. Strictly by way of non-limiting example, the core size under cruise conditions of a direct drive gas turbine engine having a fan diameter in the range of 270 cm. The core size at 290 cm can be in the range of 9 to 13, 10.5 to 12.5, or 11 to 12. Strictly as a non-limiting example, the core size under cruise conditions of a direct-drive gas turbine engine having a fan diameter in the range of 290 cm to 310 cm can be in the range of 11 to 16, 12 to 15, 13 to 15, or 13 to 14.5. The core size under cruise conditions can be in an inclusive range bounded by any two of the core size values ​​above (i.e., the values ​​can form upper or lower bounds), for example, 4.5 to 9.5.

[1126] A portion of a fan blade and / or airfoil of a fan blade described and / or claimed herein may be made from any suitable material or combination of materials. For example, at least a portion of the fan blade and / or airfoil may be made at least in part from a composite, for example, a metal matrix composite and / or an organic matrix composite, such as a carbon fiber composite. As a further example, at least a portion of the fan blade and / or airfoil may be made at least in part from a metal, such as a titanium-based metal or an aluminum-based material (such as an aluminum-lithium alloy) or a steel-based material. The fan blade may comprise at least two regions made using different materials.For example, a fan blade may have a protective leading edge, which can be made of a material more resistant to impact (e.g., from birds, ice, or other materials) than the rest of the blade. Such a leading edge could, for example, be made of titanium or a titanium-based alloy. Thus, strictly as an example, a fan blade could have a carbon fiber or aluminum-based body (such as an aluminum-lithium alloy) with a titanium leading edge. The blower blade may have a weight of at least, or may be any of the following: 3.0 kg, 3.5 kg, 4.0 kg, 4.5 kg, 5.0 kg, 5.5 kg, 6.0 kg, 6.5 kg, 7.0 kg, 7.5 kg, 8.0 kg, 8.5 kg, 9.0 kg, 9.5 kg, 10.0 kg, 10.5 kg, 11.0 kg, 11.5 kg, 12.0 kg, 12.5 kg, 13.0 kg, 13.5 kg, 14.0 kg, 14.5 kg, 15.0 kg, 15.5 kg, 16.0 kg, 16.5 kg, 17.0 kg, 17.5 kg, 18.0 kg, 18.5 kg, 19.0 kg, 19.5 kg, or 20.0 kg.The weight of the fan blade can be within an inclusive range bounded by any two of the fan blade weight values ​​from the preceding sentence (i.e., the values ​​can form upper or lower bounds). For example, the weight of the fan blade of a "geared" gas turbine having a fan diameter in the range of 200 cm to 230 cm can be in the range of 3.0 kg to 6.0 kg, or 4.0 kg to 6.0 kg, or 5.0 kg to 5.5 kg.

[1127] A blower as described and / or claimed herein may include a central part, from which the blower blades may extend, for example in a radial direction. The fan blades can be connected to the central portion in any desired manner. For example, each fan blade may include a fastening element that can engage with a corresponding notch in the hub (or disc). Strictly by way of example, such a fastening element may be in the form of a dovetail joint that can be notched into and / or engage with a corresponding notch in the hub / disc to secure the fan blade to the hub / disc. As a further example, the fan blades may be formed integrally with a central portion. Such an arrangement may be designated as a bladed disc or a bladed ring. Any suitable process may be used to manufacture such a bladed disc or a bladed ring.For example, at least part of the fan blades can be machined from a block and / or at least part of the fan blades can be joined to the hub / disc by welding, such as a linear friction weld.

[1128] The gas turbine engines described and / or claimed herein may or may not be equipped with a variable area nozzle (VAN). Such a variable area nozzle may allow the outlet area of ​​the bypass duct to be varied during operation. The general principles of this description may apply to engines with or without a VAN.

[1129] The fan of a gas turbine as described and / or claimed herein may have any desired number of fan blades, for example, 14, 16, 18, 20, 22, 24, or 26 fan blades. When the fan blades have a carbon fibre composite body, there may be 14, 16, or 18 fan blades. When the fan blades have a metallic body (for example, aluminum-lithium alloy or titanium), there may be 18, 20, or 22 fan blades.

[1130] When the gas turbine engine is an open-rotor or turboprop engine, the gas turbine engine may comprise two counter-rotating propeller stages fixed to and driven by a free-power turbine via a shaft. The propellers may rotate in opposite directions so that one rotates clockwise and the other counterclockwise about the engine's axis of rotation. Alternatively, the gas turbine engine may comprise a propeller stage and a guide vane stage designed downstream of the propeller stage. The guide vane stage may have variable pitch. Thus, high-pressure, intermediate-pressure, and free-power turbines may drive high-pressure and intermediate-pressure propellers and compressors, respectively, via suitable interconnecting shafts. In this way, the propellers may provide the majority of the propulsion thrust.

[1131] In the case where the gas turbine engine is an open rotor or turboprop engine, one or more of the propeller stages may be driven by a reduction gear. The reduction gear may be of the type described here.

[1132] As used herein, the terms idle, taxiing, takeoff, climb, cruise, descent, approach, and landing (or any part thereof) have the classical meaning and would be readily understood by a person skilled in the art. Thus, for a given gas turbine engine for an aircraft, a person skilled in the art would immediately recognize that each term refers to all, or any part, of a phase of engine operation within a given mission of an aircraft to which the gas turbine engine is designed to be fitted.

[1133] Strictly by way of non-limiting example, ground idle may refer to a phase of engine operation where the aircraft is stationary and in contact with the ground, but where there is a need for the engine to run. For example, at idle, the engine may produce between 3% and 9% of the available engine thrust. In other non-limiting examples, the engine may produce between 5% and 8% of the available thrust. In other non-limiting examples, the engine may produce between 6% and 7% of the available thrust. Strictly by way of non-limiting example, taxiing may refer to a phase of engine operation where the aircraft is propelled along the ground by the thrust produced by the engine. For example, during taxiing, the engine may produce between 5% and 15% of the available thrust. In other non-limiting examples, the engine may produce between 6% and 12% of the available thrust.In other non-limiting examples, the engine may produce between 7% and 10% of the available thrust. Strictly as a non-limiting example, takeoff may refer to a phase of engine operation where the aircraft is propelled by the thrust produced by the engine. At an initial stage in the takeoff phase, the aircraft may be propelled while in contact with the ground. At a later stage in the takeoff phase, the aircraft may be propelled while not in contact with the ground. For example, during takeoff, the engine may produce between 90% and 100% of the available thrust. In other non-limiting examples, the engine may produce between 95% and 100% of the available thrust. In still other non-limiting examples, the engine may produce 100% of the available thrust.

[1134] Strictly by way of non-limiting example, climb may refer to a phase of engine operation where the aircraft is propelled by the thrust produced by the engine. For example, during climb, the engine may produce between 75% and 100% of the available thrust. In other non-limiting examples, the engine may produce between 80% and 95% of the available thrust. In still other non-limiting examples, the engine may produce between 85% and 90% of the thrust. available. For example, climb can refer to an operating phase within an aircraft flight cycle between takeoff and arrival in cruise conditions, with arrival in cruise conditions thus defining the beginning of the cruise phase, or a portion thereof, of the aircraft flight. In addition, or alternatively, climb can refer to, for example, a nominal point in, or one or more nominal periods during, an aircraft flight cycle between takeoff and landing, where a relative increase in altitude is required, which may necessitate an additional demand on engine thrust.

[1135] As used here, cruise conditions, which may define the cruise phase, have a classical meaning and will be readily understood by those skilled in the art. Thus, for a given gas turbine engine for an aircraft, cruise conditions may refer to the engine's operating point at mid-cruise of a given mission (which may be called in the industry an "economic mission") of an aircraft to which the gas turbine engine is designed to be attached. In this sense, mid-cruise may be considered as the point in an aircraft's flight cycle at which 50% of the total fuel burned between the end of the climb and the beginning of the descent has been burned (which may be approximated by the midpoint—in terms of time and / or distance—between the end of the climb and the beginning of the descent).Cruise conditions can thus define an operating point for a gas turbine engine that provides sufficient thrust to ensure steady-state operation (i.e., maintaining a constant altitude and / or Mach number), or at least substantially steady-state operation (i.e., maintaining at least a substantially constant altitude and / or Mach number), at mid-cruise speed for an aircraft to which it is designed to be attached, taking into account the number of engines supplied to that aircraft. For example, when an engine is designed to be attached to an aircraft that has two engines of the same type, under cruise conditions the engine can provide half the total thrust that would be required for steady-state operation, or at least substantially steady-state operation, of that aircraft at mid-cruise speed.

[1136] In other words, for a given gas turbine engine for an aircraft, cruise conditions can be defined as the engine operating point that provides a specified thrust (required to provide—in combination with any other engines on the aircraft—steady-state, or at least substantially steady-state, operation of the aircraft for which it is designed to be fixed at a given mid-cruise Mach number) at mid-cruise atmospheric conditions (defined by the International Standard Atmosphere according to ISO 2533 at the mid-cruise altitude). For any Given a gas turbine engine for an aircraft, the mid-cruise thrust, atmospheric conditions, and Mach number are known, and therefore the engine's operating point at cruise conditions can be clearly defined.

[1137] Strictly by way of non-limiting example, forward speed in cruise conditions may be any point in the range from Mach 0.7 to 0.9, for example 0.75 to 0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example 0.78 to 0.82, for example 0.79 to 0.81, for example in the order of Mach 0.78, 0.79, or 0.8, in the order of Mach 0.85, or in the range from 0.8 to 0.85. Any single speed within these ranges may be part of the cruise condition. For a particular aircraft, the cruise conditions may be outside these ranges, for example, below Mach 0.7 or above Mach 0.9.

[1138] Strictly by way of non-limiting example, cruising conditions may correspond to typical atmospheric conditions (according to the International Standard Atmosphere, ISA) at an altitude that is in the range of 10,000 m to 15,000 m, e.g. in the range of 10,000 m to 12,000 m, e.g. in the range of 10,400 m to 11,600 m (about 38,000 ft), e.g. in the range of 10,500 m to 11,500 m, e.g. in the range of 10,600 m to 11,400 m, e.g. in the range of 10,700 m (about 35,000 ft) to 11,300 m, e.g. in the range of 10,800 m to 11,200 m, e.g. in the range from 10,900 m to 11,100 m, for example around 11,000 m. Cruising conditions can correspond to typical atmospheric conditions at any given altitude within these ranges.

[1139] Strictly by way of non-limiting example, cruise conditions may correspond to a forward Mach number of 0.8 and typical atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35,000 feet (10,668 m). Under such cruise conditions, the engine can provide a known level of required net thrust. The known level of required net thrust is, of course, dependent on the engine and its intended application and may be, for example, a value ranging from 20 kN to 40 kN.

[1140] Strictly by way of further example, cruise conditions may correspond to a forward Mach number of 0.85 and typical atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38,000 feet (11,582 m). Under such cruise conditions, the engine can provide a known level of required net thrust. The known level of required net thrust is, of course, dependent on the engine and its intended application and may be, for example, a value ranging from 35 kN to 65 kN.

[1141] In operation, a gas turbine engine described and / or claimed herein can operate under the cruising conditions defined elsewhere in this document. Such cruising conditions can be determined by the cruising conditions (e.g., mid-cruise conditions) of an aircraft to which at least one (e.g., 2 or 4) gas turbine engine can be fitted to provide propulsive thrust.

[1142] Furthermore, a person skilled in the art would immediately recognize that either or both of a descent and an approach refer to an operating phase within an aircraft flight cycle between cruise and landing, the approach in particular being part of the landing and takeoff (LTO) phase. Strictly by way of non-limiting example, during either or both of the descent and approach, the engine may produce less than 50% of available thrust. In other non-limiting examples, the engine may produce between 25% and 40% of available thrust. In still other non-limiting examples, the engine may produce between 30% and 35% of available thrust.In addition or as an alternative, descent can refer to a nominal point in an aircraft flight cycle between takeoff and landing, where a relative decrease in altitude is required, and which may necessitate a reduced thrust demand from the engine.

[1143] According to one aspect, an aircraft is supplied comprising a gas turbine engine as described and / or claimed herein. The aircraft according to this aspect is the aircraft to which the gas turbine engine was designed to be fitted. Accordingly, cruise conditions according to this aspect may correspond to mid-cruise of the aircraft, as defined elsewhere herein.

[1144] According to one aspect, a method of operating a gas turbine engine as described and / or claimed herein is provided. The operation may be under any suitable conditions, which may be as defined elsewhere herein (for example, in terms of thrust, atmospheric conditions, and Mach number).

[1145] According to one aspect, a method of operating an aircraft comprising a gas turbine engine as described and / or claimed herein is provided. Operation according to this aspect may (or may be) operation under any appropriate condition, for example, mid-cruise of the aircraft, as defined elsewhere herein.

[1146] During operation, under any or more of the operating conditions and / or thrust settings disclosed or described herein, a reduction in the mass CO2 emission index (El) in percentage, provided by any or more of the gas turbine engine configurations disclosed or described herein, may be greater than or equal to any of the following: 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, 0.2, 0.22, 0.24, 0.26, 0.28, 0.3, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39, 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, 0.65, 0.7, 0.75, 0.8, 0.85, 0.9, 0.95, 1, 1.1, 1.2, 1.3 1.4, 1.5, 1.6, 1.7, 1.8, 1.9, 2, 2.1, 2.2, 2.3, 2.4, 2.5, 2.6, 2.7, 2.8, 2.9, 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.5, 5, 5.5, or 6, when the fuel supplied to the combustor device comprises sustainable aviation fuel, or a blend of sustainable aviation fuel comprising up to 100% sustainable aviation fuel. The CO2 emission index (El) as a percentage may be expressed in kg of CO2 per kg of fuel. The percentage reduction provided can be in an inclusive range delimited by any two of the values ​​in the preceding sentence (i.e., the values ​​can form upper or lower bounds), for example in the range from 0.4 to 2.5, from 0.42 to 2.08, or from 0.43 to 2.08.

[1147] Strictly by way of non-limiting example, the reduction in the CO2 emission index (El) in percentage may result from the fact that the fuel supplied to the combustor device comprises a percentage mass fraction of hydrogen greater than any of the following: 13.4, 13.41, 13.42, 13.43, 13.44, 13.45, 13.46, 13.47, 13.48, 13.49, 13.5, 13.51, 13.52, 13.53, 13.54, 13.55, 13.56, 13.57, 13.58, 13.59, 13.6, 13.65, 13.7, 13.75, 13.8, 13.85, 13.9, 13.95, 14, 14.1, 14.2, 14.3, 14.4, 14.5, 14.6, 14.7, 14.8, 14.9, 15, 15.1, 15.2, 15.3, 15.4, 15.5, 15.6, 15.7, 15.8, 15.9, 16, 16.5, 17, or 17.5. The % mass fraction of hydrogen in the fuel can be in an inclusive range delimited by any two of the values ​​in the preceding sentence (i.e., the values ​​can form upper or lower bounds), for example in the range from 13.41 to 15.3, or from 13.42 to 15.3.

[1148] During operation, under any or more of the operating conditions and / or thrust settings disclosed or described herein, a CO2 reduction in percentage per MJ of fuel energy, provided by any or more of the gas turbine engine configurations disclosed or described herein, may be greater than or equal to any of the following: 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, 0.2, 0.22, 0.24, 0.26, 0.28, 0.3, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39, 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, 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, 0.81, 0.82, 0.83, 0.84, 0.85, 0.86, 0.87, 0.88, 0.89, 0.9, 0.91, 0.92, 0.93, 0.94, 0.95, 0.96, 0.97, 0.98, 0.99, 1, 1.1, 1.2, 1.3, 1.4, 1.5, 1.6, 1.7, 1.8, 1.9, 2, 2.1, 2.2, 2.3, 2.4, 2.5, 2.6, 2.7, 2.8, 2.9, 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1, 4.2, 4.3, 4.4, 4.5 4.6, 4.7, 4.8, 4.9, 5, 5.5, 6.5, 7, or 7.5, when the fuel supplied to the combustor device comprises sustainable aviation fuel, or a blend of sustainable aviation fuel comprising up to 100% sustainable aviation fuel. The percentage reduction supplied may be within an inclusive range delimited by any two of the values ​​in the preceding sentence. (that is, the values ​​can form upper or lower limits), for example in the range from 0.8 to 5, from 0.88 to 4.75, or from 0.89 to 4.75.

[1149] Strictly by way of non-limiting example, the percentage CO2 reduction per MJ of fuel energy may result from the fact that the fuel supplied to the combustor device comprises a specific fuel energy (in MJ per kg) greater than any of the following: 41.5, 41.6, 41.7, 41.8, 41.9, 42, 42.1, 42.2, 42.3, 42.4, 42.5, 42.6, 42.7, 42.8, 42.9, 43, 43.1, 43.2, 43.3, 43.4, 43.5, 43.6, 43.7, 43.8, 43.9, 44, 44.5, 45, 45.5, or 46. The specific fuel energy fuel can be in an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​can form upper or lower bounds), for example in the range from 42.8 to 45, from 43 to 44.5, or from 43 to 44.

[1150] As used herein, unless otherwise indicated, a range "from value X to value Y" or "between value X and value Y", or similar, means an inclusive range; including the bounding values ​​of X and Y.

[1151] A person skilled in the art would understand that, except for mutual exclusivity, a feature or parameter described in relation to any of the above aspects may be applied to any other aspect. Furthermore, except for mutual exclusivity, any feature or parameter contained or described herein may be applied to any aspect and / or combined with any other feature or parameter described herein.

[1152] Furthermore, except for mutual exclusivity, any parameter or value contained or described herein may be applied to and / or combined with any one or more other parameters and / or values ​​described herein. For example, a first parameter or value contained or described herein (e.g., parameter A) may be applied to and / or combined with any one or more other parameters and / or values ​​contained or described herein (e.g., one or more parameters B; a parameter C; and a parameter D, and so on) to express their relationship. For example, a person skilled in the art will understand that where parameter A is described separately from parameter B, their relationship may be expressed, for example, as A+B, BA, AB, A / B, B / A, B*A, or any other product, application, combination, function, or expression of parameter A with respect to parameter B, or vice versa, as required.

[1153] Unless they are mutually exclusive, any parameters or values ​​related to those contained or described herein, or that can be determined and / or deduced from them, may be applied and / or combined with any additional parameters or values ​​contained or described herein and / or any other parameters or values ​​related to those contained or described herein, or that can be determined and / or deduced from them, to express their relationship in relation to engine emissions and / or nvPM. For example, using temperatures, pressures, operating parameters, rotational speeds, flow rates, or related engine operating conditions that can be determined and / or deduced, a first parameter or value (e.g., parameter A) can be applied and / or combined with any one or more additional parameters or values ​​(e.g., any one or more of a parameter B; a parameter C; and a parameter D, etc.), to express their relationship with respect to engine emissions and / or nvPM.For example, a person skilled in the art will understand that when a parameter C can be considered separate from a parameter D, their relationship can be expressed, for example, by C+D, CD, CD, C / D, C / D, C*D, or any such product, application, combination, function, or additional expression of parameter C with respect to parameter D, or vice versa, as required. Brief description of the drawings

[1154] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[1155] [Fig.1] is a cross-sectional side view of a gas turbine engine;

[1156] [Fig.2] is a close-up cross-sectional side view of an upstream part of an engine geared gas turbine;

[1157] [Fig. 3] is a partially cutaway view of a gearbox for a turbine motor gas;

[1158] [Fig.4] is a close-up cross-sectional side view of a direct-drive gas turbine engine;

[1159] [Fig. 5] is a schematic view of an aircraft having two gas turbine engines of the present application mounted on it;

[1160] [Fig. 6] is a schematic representation of a distribution system of fuel and combustor device of a gas turbine engine;

[1161] [Fig.7] is a cross-sectional view through the combustor device of a gas turbine engine along the main axis of rotation of the engine;

[1162] [Fig. 8] is another schematic representation of a fuel distribution system and combustor device of a gas turbine engine; and

[1163] [Fig.9] shows a method of operation of the gas turbine engine.

[1164] DETAILED DESCRIPTION OF THE DISCLOSURE

[1165] Figure [1] illustrates a gas turbine engine 10 having a main axis of rotation 9. The engine 10 includes an air intake 12 and a propulsion fan 23 which generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 includes a core 11 which receives the core airflow A. The engine core 11 comprises, in axial flow series, a low-pressure compressor 14, a high-pressure compressor 15, a combustion unit 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 blower 23 is attached to and driven by the low-pressure turbine 19 via a first low-pressure shaft 26 and an epicyclic reduction gear 30.

[1166] During operation, 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 discharged from the high-pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel F and the mixture is burned. The combustion equipment 16 may be referred to as the combustor device 16, with the terms "combustion equipment 16" and "combustor device 16" used interchangeably herein. The resulting hot combustion products then expand through, and thereby drive, the high-pressure and low-pressure turbines 17, 19 before being discharged through the core exhaust nozzle 20 to provide some propulsive thrust. The high-pressure turbine 17 drives the high-pressure compressor 15 via a suitable second interconnecting high-pressure shaft 27.The blower 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 discharged through the bypass exhaust nozzle 18 to generally provide the majority of the propulsion thrust. The epicyclic reducer 30 is a reduction gearbox.

[1167] An exemplary arrangement for a geared blower 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 reducer 30. The low-pressure shaft 26 can be called the input shaft for the epicyclic reducer 30. Radially outward from the sun gear 28 and meshing with it are a plurality of planet gears 32, which are coupled together by a planet carrier 34. The planet carrier 34 forces the planet gears 32 to change orientation about the sun gear 28 synchronously while allowing each planet gear 32 to rotate about its own axis.The planet carrier 34 is coupled via links 36 to the blower 23 through an output blower shaft 42 in order to drive the blower 23 in rotation about the motor shaft 9. Radially outward from the planet gears 32 and meshing with them, there is a ring or . toothed crown 38 which is coupled, via links 40, to a stationary support structure 24.

[1168] It should be noted that the terms 'low-pressure turbine' and 'low-pressure compressor' as used herein may be taken to mean the lowest-pressure turbine stages and the lowest-pressure compressor stages (i.e., not including the blower 23) respectively and / or the turbine and compressor stages which are connected together by the interconnecting shaft (26) with the lowest rotational speed in the motor (i.e., not including the reduction output shaft which drives the blower 23). In some literature, the 'low-pressure turbine' and 'low-pressure compressor' referred to herein may alternatively be known as the 'intermediate-pressure turbine' and 'intermediate-pressure compressor'.When such alternative nomenclature is used, the blower 23 may be designated as the first compression stage or the lowest pressure compression stage.

[1169] The epicyclic reducer 30 is shown in more detail by way of example in [Fig. 3]. Each of the sun gear 28, the planet gears 32, and the ring gear 38 has teeth around its periphery for meshing with the other gears. However, for clarity only, only exemplary portions of the teeth are shown in [Fig. 3]. There are four planet gears 32 shown, although it will be apparent to the specialist reader that more or fewer planet gears 32 may be provided within the scope of the claimed invention. Practical applications of a planetary epicyclic reducer 30 generally include at least three planet gears 32, for example, five planet gears 32.

[1170] The epicyclic reducer 30 illustrated by way of example in Figures 2 and 3 is of the planetary type, in that the planet carrier 34 is coupled to the output shaft via links 36, with the ring gear 38 fixed. However, any other suitable type of epicyclic reducer 30 may be used. As a further example, the epicyclic reducer 30 may be a star arrangement, in which the planet carrier 34 is held fixed, with the ring gear (or ring) 38 permitted to rotate. In such an arrangement, the blower 23 is driven by the ring gear 38. As another alternative example, the reducer 30 may be a differential reducer in which both the ring gear 38 and the planet carrier 34 are permitted to rotate.

[1171] It should be borne in mind that the arrangement shown in [Fig. 2] and [Fig. 3] is by way of example only, and that various alternatives are within the scope of this disclosure. Strictly by way of example, any suitable arrangement may be used to position the reducer 30 in the motor 10 and / or to connect the gearbox 30 to the motor 10. By way of further example, the connections (such as links 36, 40 in the example in [Fig. 2]) between the gearbox 30 and other parts of the motor 10 (such as the input shaft 26, the output shaft, and the stationary support structure 24) can have any desired degree of rigidity or flexibility. By way of further example, any suitable arrangement of bearings between rotating and stationary parts of the motor (e.g., between the input and output shafts from the gearbox and fixed structures, such as the gearbox housing) can be used, and disclosure is not limited to the arrangement given as an example in [Fig. 2].For example, when the reducer 30 has a star arrangement (described above), the person skilled in the art will readily understand that the arrangement of the output and support links and the bearing locations will typically be different from that shown as an example in [Fig.2].

[1172] Thus, the present description extends to a gas turbine engine having any arrangement of gearbox styles (for example, star or planetary), support structures, input and output shaft arrangements, and bearing locations.

[1173] Optionally, the reducer may drive additional and / or alternative components (for example, the intermediate pressure compressor and / or a booster).

[1174] Other gas turbine engines to which this description may be applied may have alternative configurations. For example, such engines may have a different number of compressors and / or turbines and / or a different number of interconnecting shafts. As a further example, the gas turbine engine shown in [Fig. 1] has a split-flow nozzle 18, 20, which means that the flow through the bypass duct 22 has its own nozzle (the bypass exhaust nozzle 18) which is separated from and radially outside the core exhaust nozzle 20. However, this is not limiting, and any aspect of the present description can also be applied to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be called a mixed-flow nozzle.The nozzle(s) (whether mixed or split flow) can have a fixed or variable area.

[1175] By way of further example, other gas turbine engines to which the present description may be applied may not have a reduction gear for the main shaft(s), being instead direct-drive engines. A cross-sectional view of such an engine is shown in [Fig. 4].

[1176] With reference to [Fig. 4], a gas turbine engine is usually denoted as 10, having a main axis of rotation 9. The engine 10 comprises, in flow series axial, an air intake 12, a propulsion fan 23, an intermediate pressure compressor 14, a high pressure compressor 15, a combustion unit 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.

[1177] During operation, the air entering the intake 12 is accelerated by the blower 23 to produce two air streams: a core air stream A and a bypass air stream B. The core air stream A flows into the intermediate pressure compressor 14, and the bypass air stream B passes through a bypass duct 22 to provide propulsion thrust. The intermediate pressure compressor 14 compresses the air stream A before delivering this air to the high-pressure compressor 15 where further compression takes place.

[1178] The compressed air discharged from the high-pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel F and the mixture is burned. The combustion equipment 16 may be referred to as the combustor device 16, with the terms "combustion equipment 16" and "combustor device 16" used interchangeably herein. The resulting hot combustion products then expand, thereby driving the high-pressure, intermediate-pressure, and low-pressure turbines 17, 19a, 19 before being discharged through the nozzle 20 to provide additional propulsion thrust. The high-pressure turbine 17, intermediate-pressure turbine 19a, and low-pressure turbine 19 drive the high-pressure compressor 15, the intermediate-pressure compressor 14, and the blower 23, respectively, each via a suitable interconnecting shaft.

[1179] Other gas turbine engines to which this description may be applied may have alternative configurations. For example, such engines may have an alternative number of interconnecting shafts (e.g., two) and / or an alternative number of compressors and / or turbines. In addition, the engine may include a reduction gear supplied in the drive train from a turbine to a compressor and / or a blower.

[1180] While the example described relates to a turbofan engine, the description may apply, for example, to any type of gas turbine engine, such as an open rotor (in which the fan stage is not enclosed by a nacelle) or a turboprop, for example. In some arrangements, the gas turbine engine 10 may not include a reduction gear 30.

[1181] The geometry of the gas turbine engine 10, and of its components, is defined by a conventional axis system, comprising an axial direction (which is aligned with the main axis of rotation 9), a radial direction (in the downward upward direction on the [Fig. 1]), and a circumferential direction (perpendicular to the page on the (view of [Fig.1]). The axial, radial and circumferential directions are mutually perpendicular.

[1182] Figure 5 shows an aircraft 1 on which two gas turbine engines 10 of this disclosure are mounted, one on each wing. The aircraft 1 includes a fuel system 2 comprising a fuselage fuel tank 50a and two wing fuel tanks 50b. Fuel F is supplied by the fuel system to the gas turbine engines. The fuel tanks 50a and 50b are supplied with fuel from a fuel inlet port 62. Other fuel systems can be used with other fuel tank arrangements.

[1183] The fuel F supplied to the combustion equipment 16 may include a fossil-based hydrocarbon fuel, such as kerosene. Thus, fuel F may include molecules of one or more of the chemical families of n-alkanes, iso-alkanes, cycloalkanes, and aromatics. Given the expectation in the aviation industry of a trend toward the use of fuels other than the traditional kerosene-based jet fuels generally used today, fuel F may include renewable hydrocarbons produced from biological or non-biological resources, also known as sustainable aviation fuel (SAF), when blended or mixed with, or substituted for, an alternative fuel. In each of the examples given, fuel F may include one or more trace elements comprising, for example, sulfur, nitrogen, oxygen, inorganic substances, and metals.

[1184] For those skilled in the art, SAF means, for example, a biofuel, a renewable aviation fuel, a renewable jet fuel, or an alternative jet fuel or biofuel, produced from biological or non-biological resources. Thus, for those skilled in the art, SAF includes, for example, a fuel produced from sustainable and / or renewable resources.For example, it is understood that a SAF is commonly synthesized from carbon-containing gases extracted from the atmosphere and / or captured during industrial processes; or from a wide range of sustainable feedstocks such as, for example, used oils and greases; municipal solid waste; cellulosic waste (such as corn stalks); cover crops such as camelina, carinata, and moneywort; non-biogenic alternative fuels; jatropha; halophytes, and algae, rather than from fossil-based hydrocarbons derived, for example, from fossil-based petroleum and / or natural gas. Accordingly, it is understood that an SAF includes renewable hydrocarbons. In addition, it is understood that an SAF does not include fossil fuels or fossil-based hydrocarbons.

[1185] The functional performance of a given fuel composition, or of a fuel mixture F for use in a given mission, may These parameters can be defined, at least in part, by the fuel's ability to complete the Brayton cycle of the gas turbine engine. The 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). Thus, nvPM can be defined as emitted particles exiting at the nozzle outlet plane of a gas turbine engine that do not volatilize when heated to a temperature of 350 °C. Any reference here to soot or smoke can also be applied equivalently to other types of particulate matter emissions known in the art.The gaseous emissions may include any one or more of the following: nitrogen oxides (NOx); carbon monoxide (CO); carbon dioxide (CO2); unburned hydrocarbons (UHC); sulfur oxides (SOx) comprising, for example, sulfur dioxide (SO2) and / or sulfur trioxide (SO3); and volatile organic compounds (VOCs) created by the combustion of said fuel F. Any reference to gaseous emissions herein may also apply to other types of gaseous emissions known in the art.

[1186] A relatively higher specific energy (i.e., energy per unit mass), expressed in MJ / kg, can at least partially reduce the takeoff weight, potentially providing a relative improvement in fuel efficiency. A relatively higher energy density (i.e., energy per unit volume), expressed in MJ / L, can at least partially reduce the fuel volume at takeoff, which may be particularly important for volume-limited missions or military operations involving refueling. A relatively higher thermal stability (i.e., inhibiting fuel degradation or coking under thermal stress) can allow the fuel to withstand elevated temperatures in the engine and fuel injectors, potentially providing relative improvements in combustion efficiency.Reduced emissions, including particulate matter, can lead to less condensation trail formation, thus reducing the environmental impact of a given mission. Other fuel properties can also be critical to functional performance. For example, a relatively lower freezing point (°C) can allow for optimized flight profiles on long-range missions; minimum aromatic concentrations (%) can ensure sufficient swelling of certain materials used in the construction of O-rings and seals previously exposed to high-temperature fuels. aromatic content; and, a maximum surface tension (mN / m) can ensure sufficient spray break and atomization of the fuel.

[1187] The ratio of hydrogen to carbon atoms in a molecule can influence the specific energy of a given composition, or fuel mixture. Fuels with higher hydrogen-to-carbon ratios may have higher specific energies in the absence of bond strains. 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 consisting of molecules with 9 to 15 carbons, and an average of 12 carbons.

[1188] A number of sustainable aviation fuel blends have been approved for use. For example, some approved blends include blending ratios of up to 10% sustainable aviation fuel, while other approved blends include blending ratios of up to 50% sustainable aviation fuel (the remainder comprising one or more fossil-based hydrocarbon fuels, such as kerosene), with additional compositions pending approval. However, there is an expectation within the aviation industry that sustainable aviation fuel blends comprising up to (and including) 100% sustainable aviation fuel (SAF) will eventually be approved for use.

[1189] Sustainable aviation fuels may comprise one or more of n-alkanes, iso-alkanes, cycloalkanes, and aromatics, and may be produced, for example, from one or more of a synthesis gas (synthesis gas); lipids (e.g., fats, oils, and greases); sugars; and alcohols. Thus, sustainable aviation fuels may comprise lower aromatic and / or sulfur contents compared to fossil-based hydrocarbon fuels. In addition, or alternatively, sustainable aviation fuels may comprise one or both of higher iso-alkane and cycloalkane contents compared to fossil-based hydrocarbon fuels.In some examples, sustainable aviation fuels may have a density below 100%, for example between 90% and 98%, of that of kerosene, and / or a specific energy above 100%, for example between 101% and 105%, of that of kerosene. For example, the calorific value of sustainable aviation fuels may be between 101% and 105% of that of kerosene.

[1190] In some examples, the sustainable aviation fuel(s), or the mixture(s) supplied to the combustion equipment 16 may have a content of aromatics and / or other non-paraffinic elements relatively lower than that of kerosene. Sustainable aviation fuel may have an aromatics content of, for example, 30%, 20%, 15%, 10%, 8%, 5%, or less than 5%; for example, 4%, 3%, 2%, 1%, or less than 1%; for example, 0.75%, 0.5%, 0.25%, or less than 0.25%; for example, 0.2%, 0.1%, or less than 0.1%; for example, 0.01%, 0.001%, or 0%.The aromatics content of sustainable aviation fuel may be within one inclusive figure or range delimited by one or two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower bounds), for example 13.5%, 8.5%, 2.5%, 0.35%, 0.15%, 0.05%, 0.005% or 0%; or from 0% to 0.75%, from 0% to 0.5%, or from 0.1% to 0.25%; or from 0.15% to 0.65%, 0.35% to 0.55%, or 0.035% to 0.055%; depending on one or more of a preference, stock or supplier of fuel, and variations in composition therein.

[1191] Thanks at least in part to the molecular structure of sustainable aviation fuels, sustainable aviation fuels can provide beneficial effects including, for example, one or more of the following: higher specific energy (despite, in some examples, a lower energy density); higher specific heat capacity; higher thermal stability; superior lubricating power; lower viscosity; lower surface tension; lower freezing point; lower soot emissions; lower NOx levels; and lower CO2 emissions, compared to fossil-based hydrocarbon fuels (e.g., when burned in combustion equipment 16).Thus, compared to fossil-based hydrocarbon fuels, such as kerosene, sustainable aviation fuels can lead to one or both of a relative decrease in specific fuel consumption and a relative decrease in maintenance costs.

[1192] Figure 6 shows a schematic representation of a fuel distribution system 102 and the combustor device 16 of a gas turbine engine 10 of any example described herein. The combustor device 16 is designed to use a staged lean-burn combustion. Fuel is divided among pilot fuel injectors and main fuel injectors by means of a fuel system control device, which in the example shown is supplied by a fuel metering unit (FMU) 104 under the control of an electronic engine control device (EEC) 106. Fuel is delivered to the fuel metering unit 104 by a fuel pump 108. In the example shown, the fuel pump 108 is mechanically driven by an accessory drive gearbox (AGB) 110, although the fuel pump 108 can alternatively be electrically driven. The fuel pump 108 shown in [Fig. 6] can be one of multiple fuel pumps supplied within the fuel distribution system 102. For example, fuel pump 108 may be a high-pressure fuel pump supplied on the gas turbine engine 10, with one or more additional lower-pressure fuel pumps also supplied, optionally on board the aircraft rather than being part of the gas turbine engine 10.

[1193] High-pressure fuel is delivered by the fuel metering unit 104 into one or more fuel manifolds for distribution to pilot fuel injectors 116A and main fuel injectors 116B. Fuel delivery via pilot fuel injectors 116A and main fuel injectors 116B is staged, so at low power levels (and therefore low mass airflows) it is primarily or entirely delivered by the pilot fuel injectors 116A at a rich fuel-to-air ratio (i.e., an equivalence ratio greater than one) for improved flame stability. As power and mass flow increase, a stage is reached where fuel is delivered by some or all of the main fuel injectors 116B, supplementing the fuel flow from the pilot fuel injectors 116A.The 116B main fuel injectors are designed to inject fuel at a lean fuel-to-air ratio (i.e., an equivalence ratio less than one). At this point, the airflow is such that the equivalence ratio immediately downstream of the 116A pilot fuel injectors is also lean. In the example shown, at higher power levels, fuel is injected by all 116B main fuel injectors.

[1194] Men skilled in the art will be familiarized with such operation of staged combustion systems in order to effect a lean burn at high power levels while respecting flammability limits at lower power levels.

[1195] The fuel injection balance by the pilot fuel injectors 116A and the main fuel injectors 116B is controlled by the electronic engine control device 106, which provides control signals to the fuel metering unit 104. The control signals may be indicative directly or indirectly of the total fuel to be injected, for example in the form of a fuel flow rate and the ratio of the pilot fuel injector fuel flow rate to the main injector fuel flow rate.

[1196] Figure 7 shows a section through the combustor device 16 in a plane perpendicular to the main axis of rotation 9 of the engine 10. The combustor device 16 comprises an annular combustion chamber 120, defined by a liner 122. Other combustor device configurations may alternatively be used, for example, cannular combustor devices, canned combustor devices, etc.

[1197] The combustor device 16 comprises a plurality of fuel spray nozzles 124 arranged around a circumference of the combustor device 16 and designed to inject fuel into the combustion chamber 120. In the example shown, the combustor device 16 comprises sixteen (16) fuel spray nozzles 124. The combustor device 16 may alternatively comprise any appropriate number of fuel spray nozzles, for example, 8, 9, 10, 11, 12, 13, 14, 15, 17, 18, 19, 20, 21, 22, 23, 24, 25, 26, 27, 28, 29, 30 fuel spray nozzles, and so on. The combustor device may include a number of fuel spray nozzles within an inclusive range defined between any two of the values ​​in the preceding sentence, the two values ​​forming the upper and lower limits of the range and being included within the range.For example, the combustor device may include between 14 and 27 fuel spray nozzles, or between 16 and 25 fuel spray nozzles, or between 18 and 23 fuel spray nozzles.

[1198] A core size of a gas turbine engine is defined as (with reference to the arrangement shown in [Fig. 1]):

[1199] [Math. 123] Jt- heart size = iù2 • -pf

[1200] where 11½ is the mass flow rate, in pounds per second, of the air at the inlet of the high-pressure compressor 15, T3 is the temperature, in Kelvin, of the air at the outlet of the high-pressure compressor 15, and is the pressure, in pounds inches per second squared per square inch, of the air at the outlet of the high-pressure compressor 15. A unit of core size is therefore expressed as:

[1201] [Math. 124] i s* K 2 *po

[1202] The core size (in s.K1 / 2 .in) of the motor may be between 4 and 7, for example 4, 4.5, 5, 5.5, 6, 6.5 or 7, or any defined range between any two of these values. In some examples, the motor core size (in s.K1 / 2 .in) may be in the range of 5.0, 5.1, 5.2, 5.3, 5.4, 5.5, 5.6, 5.6, 5.7, 5.8, 5.9, or 6, or any defined range between any two of these values. In yet other examples, the motor core size (in s.K1 / 2 .in) may be in the range of 5.25, 5.26, 5.27, 5.28, 5.29, 5.30, 5.31, 5.32, 5.33, 5.34, 5.35, 5.36, 5.37, 5.38, 5.39, 5.40, 5.41, 5.42, 5.43, 5.44, or 5.45, or any range defined between any two of these values.

[1203] The number of fuel spray nozzles 124 per unit engine core size (in s.K1 / 2 .in) may be between 2 and 6. The number may be, for example, 2, 3, 4, 5, or 6, or any defined range between any two of these values. In some examples, the number may be between 3 and 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 defined range between any two of these values.

[1204] The number of fuel spray nozzles per unit engine core size may be between 2 and 7, or more preferably between 2.1 and 6.5, or more preferably between 2.4 and 3.4.

[1205] In still other examples, the number of fuel spray nozzles per unit engine core size may be 2.0, 2.1, 2.2, 2.3, 2.4, 2.5, 2.6, 2.7, 2.8, 2.9, 3.0, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4.0, 4.1, 4.2, 4.3, 4.4, 4.5, 4.6, 4.7, 4.8, 4.9, 5.0, 5.1, 5.2, 5.3, 5.4, 5.5, 5.6, 5.7, 5.8, 5.9, 6.0, 6.1, 6,2, 6,3, 6,4, 6,5, 6,6, 6,7, 6,8, 6,9, 7,0, or within a defined range between any two of these values.

[1206] The core size is defined here at an engine operating condition corresponding to a maximum value of the semi-dimensional flow rate at a high-pressure compressor inlet, defined as:

[1207] [Math. 125]

[1208] where UI2 is the mass flow rate (in pounds per second) of the air at the inlet of the high-pressure compressor, T2 is the temperature (in Kelvin) of the air at the inlet of the high-pressure compressor, and P2 is the pressure (in pounds inches per second squared per square inch) of the air at the inlet of the high-pressure compressor.

[1209] The operating condition corresponding to the maximum semi-dimensional flow rate at the high-pressure compressor inlet can be the peak climb operating condition. The core size designated herein can therefore be defined at the peak climb operating condition. The peak climb can be as defined in the art and as understood by those skilled in the art for a specific implementation of a gas turbine engine of the present application. In a specific example, the peak climb can correspond to operation at an altitude between 30,000 feet and 39,000 feet (more specifically 35,000 feet), a forward speed of Mach number 0.75 to 0.85, and an ambient air temperature (AAT) of ISA+10 K to ISA+15 K.

[1210] In the example shown, each fuel spray nozzle 124 includes a duplex fuel spray nozzle (also called an internal stage nozzle) where a pilot fuel injector 116A is integrated into the same fuel spray nozzle 124 as a primary fuel injector 116B. However, it is envisaged that other types of staged combustion configurations could be used, for example, those with pilot and main fuel injectors in separate fuel spray nozzles rather than both contained in duplex or internally staged fuel spray nozzles. Indeed, it will be understood that the principles described here can be applied to any staged combustion system comprising pilot and main fuel injectors.

[1211] Returning to [Fig. 6], the fuel distribution system 102 includes a separation valve (SV) 112 designed to separate the fuel flow between the fuel spray nozzles 124 of the combustor device 16 such that the pilot inject...

Claims

Demands

1. A gas turbine engine (10) for an aircraft, comprising: a combustor device (16), including a combustion chamber (120) and a plurality of fuel spray nozzles (124) designed to inject fuel into the combustion chamber (120), wherein the plurality of fuel spray nozzles (124) comprises a first subset (124A) of fuel spray nozzles (124) and a second subset (124B) of fuel spray nozzles (124), wherein the combustor device (16) can operate under a condition in which each of the fuel spray nozzles of the first subset (124A) of fuel spray nozzles (124) is supplied with fuel at a fuel flow rate greater than each of the fuel spray nozzles of the second subset (124B) of fuel spray nozzles (124), wherein a ratio of the number nozzlesfuel spray (124) in the first subset (124A) of fuel spray nozzles (124) to the number of fuel spray nozzles (124) in the second subset (124B) of fuel spray nozzles (124) is in the range of 1:2 to 1:5; and wherein: a depleted cruise nvPM-PMD emission index ratio is defined as: [Math. 160] WHERE; BPR ^ / cruise (depleted) CSt defini COmme . [Math. 161] EImay^o+EImontée is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine (10) when operating at approximately 100% of available thrust under given operating conditions; ^ / climb is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine (10) when operating at approximately 85% of available thrust under given operating conditions; and BPR is the bypass ratio of the gas turbine engine (10); the depleted cruise nvPM-PMD emission index ratio is less than 0.2; and the gas turbine engine (10) is designed to supply fuel comprising a sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles (124).

2. Gas turbine engine (10) according to claim 1, wherein the depleted cruise nvPM-PMD emission index ratio is less than 0.119, preferably less than 0.109 and more preferably less than 0.0989.

3. Gas turbine engine (10) according to any one of the preceding claims, wherein the depleted cruise nvPM-PMD emission index ratio is less than or equal to 0.106, preferably less than or equal to 0.0972, and more preferably less than or equal to 0.0883.

4. Gas turbine engine (10) according to any one of the preceding claims, wherein the depleted cruise nvPM-PMD emission index ratio is greater than or equal to 0.0519, preferably greater than or equal to 0.0584, and more preferably greater than or equal to 0.0649.

5. Gas turbine engine (10) according to any one of the preceding claims, wherein: a rich-PMD cruise nvPM emission index ratio is defined as: [Math. 162] ra™sl«8(rlcl,e / EImaxTO Where • BPR Elaoisière (rich) is defined as: [Math. 163] ^climb+^approach is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine (10) when operating at approximately 85% of available thrust under the given operating conditions, or under other different operating conditions; £7approach is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine (10) when operating at approximately 30% of available thrust in the same operating conditions under which EI^^ is calculated; and £7maxTo is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine (10) when operating at about 100% of available thrust under the same operating conditions under which EZmontée is calculated; and the rich cruise nvPM-PMD emission index ratio is less than 20.

6. Gas turbine engine (10) according to claim 5, wherein the rich cruise nvPM-PMD emission index ratio is less than 19, preferably less than 17.5 and more preferably less than 15.

9.

7. Gas turbine engine (10) according to claim 5 or claim 6, wherein the rich cruise nvPM-PMD emission index ratio is less than or equal to 12, preferably less than or equal to 9 and more preferably less than or equal to 6.

8. Gas turbine engine (10) according to any one of claims 5 to 7, wherein the rich cruise nvPM-PMD emission index ratio is greater than or equal to 0.0374, preferably greater than or equal to 0.0421, and more preferably greater than or equal to 0.0468.

9. A gas turbine engine (10) according to any one of the preceding claims, wherein the ratio of the number of fuel spray nozzles (124) in the first subset (124A) of fuel spray nozzles (124) to the number of fuel spray nozzles (124) in the second subset (124B) of fuel spray nozzles (124) is in the range of 1:3 to 1:4, or preferably in the range of 1:3.5 to

10. 1.H-. Gas turbine engine (10) according to any one of the preceding claims, wherein the combustor device (16) comprises one or more igniters (126).

11. Gas turbine engine (10) according to claim 10, wherein each of the first subassembly (124A) of fuel spray nozzles (124) is located closer to one or more of the respective igniters (126) than the second subassembly (124B), and / or wherein one or more igniters (126) are arranged diametrically opposite one or more other igniters among the igniters (126).

12. Gas turbine engine (10) according to any one of the preceding claims, wherein the fuel supplied to the plurality of fuel spray nozzles (124) comprises a %SAF in the range of 50% to 100%, preferably in the range of 70% to 100%, and more preferably in the range of 90% to 100%.

13. Method (1000) of operating the gas turbine engine (10) according to any one of the preceding claims, the method comprising supplying (1002) fuel comprising a sustainable aviation fuel to the plurality of fuel spray nozzles (124).

14. Method (1000) of operating a gas turbine engine (10), the gas turbine engine (10) comprising: a combustor device (16), comprising a combustion chamber (120) and a plurality of fuel spray nozzles (124) designed to inject fuel into the combustion chamber (120), in which the plurality of fuel spray nozzles (124) comprises a first subset (124A) of fuel spray nozzles (124) and a second subset (124B) of fuel spray nozzles (124),in which the combustor device (16) can operate under a condition in which each of the fuel spray nozzles of the first subset (124A) of fuel spray nozzles (124) is supplied with fuel at a fuel flow rate greater than each of the fuel spray nozzles of the second subset (124B) of fuel spray nozzles (124), in which a ratio of the number of fuel spray nozzles (124) in the first subset (124A) of fuel spray nozzles (124) to the number of fuel spray nozzles (124) in the second subset (124B) of fuel spray nozzles (124) is in the range of 1:2 to 1:5; and in which: a depleted cruise nvPM-PMD emission index ratio is defined as: [Math. 164], BPR ^ / cruise (depleted) ^St defini COnHHC . [Math. 165] EEwmï^EPniontée EIm^^ is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine (10) when operating at approximately 100% of available thrust under given operating conditions; ^climb is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine (10) when operating at approximately 85% of available thrust under given operating conditions; and BPR is the bypass ratio of the gas turbine engine (10); and the ratio of the nvPM emission index in depleted cruise-PMD is less than 0.2; and the method (1000) includes supplying (1002) fuel comprising a sustainable aviation fuel to the plurality of fuel spray nozzles (124).

15. Method (1000) according to claim 14, wherein: a rich-PMD cruise nvPM emission index ratio is defined as: [Math. 166] WHERE I BPR Æ / cruise (rich) is defined as: [Math. 167] ■E'^mantae+^'^approach EImontée is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine (10) when operating at approximately 85% of available thrust under the given operating conditions, or under other different operating conditions; Rapproche is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine (10) when operating at approximately 30% of available thrust under the same operating conditions under which EImontée is calculated; £7maxTo is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine (10) during of operation at approximately 100% of available thrust under the same operating conditions in which EZmontée is calculated; and The ratio of the nvPM emissions index in rich cruising to PMD is less than 20.

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