Aircraft operation

Abstract

Gas turbine engine (10) for an aircraft. The gas turbine engine comprising: a rich-burning, fast-cooling, lean-burning (RQL) combustor device (16) having a number of fuel spray nozzles (403, 404) in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6.A depleted-PMD cruise nvPM emission index ratio is defined as: where: EIcruise (depleted) is defined as: EImaxTO is the loss-corrected system nvPM emission index in mg / kg of the gas turbine engine (10) when operating at approximately 100% of available thrust under given operating conditions; EIclimb is the loss-corrected system 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 depleted-PMD cruise nvPM emission index ratio is less than 0.2. The gas turbine engine (10) is designed to supply fuel including sustainable aviation fuel (SAF) to the fuel spray nozzles (403, 404).The invention also discloses a method of operating the gas turbine engine. Figure for the abstract: [Figure 6].

Description

Title of the invention: Operation of an aircraft

[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 one or more of the following characteristics:

[0008] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; and

[0009] in which:

[0010] A first ratio of the nvPM emission index between idle and PMD can be defined as:

[0011] [Math.l] FT ElmaxTO slow motion

[0012] where:

[0013] Δ / 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] Ê / maxio is the nvPM emission index corrected for losses in the system 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 ratio of the nvPM emission index between idle and PMD of the gas turbine engine is less than 0.8; and

[0016] The gas turbine engine is designed to supply fuel including sustainable aviation fuel (SAF) to the 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 by way of example, lower nvPMs under cruise conditions can significantly reduce the impact of radiative forcing from contrails. Strictly by way of further example, lower nvPMs under idle conditions can significantly improve local ground air quality in the region where the engine is operating. Strictly by way of further example, lower nvPMs under PMD conditions can significantly reduce the peak nvPM production rate during the flight cycle and / or improve ground air quality and / or air quality in the region where the engine is operating. 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, one or more of the following parameters can 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 the fuel must be distributed, ignited, and / or consumed within the gas turbine engine. These considerations may apply to all aspects of disclosure.

[0019] The first nvPM emission index ratio between idle and PMD of the gas turbine engine may be greater than zero.

[0020] The first nvPM emission index ratio between idle and PMD may be less than 0.708 and preferably less than 0.649 and more preferably less than 0.59.

[0021] The first nvPM emission index ratio between idle and PMD can be less than or equal to 0.5 and preferably less than or equal to 0.4 and more preferably less than or equal to 0.3.

[0022] The first nvPM emission index ratio between idle and PMD can be less than or equal to 0.105 and preferably less than or equal to 0.0962 and more preferably less than or equal to 0.0875.

[0023] The first nvPM emission index ratio between idle and PMD may be less than 0.01, 0.05, 0.07, 0.1, 0.15, 0.2, 0.25, 0.3, 0.35, 0.4, 0.45, 0.5, 0.55, 0.6, 0.65, 0.7, 0.75, 0.8, or within any defined range between any two of these values. For example, the first nvPM emission index ratio between idle and PMD may be within a range between 0.01 and 0.2, 0.01 and 0.15, 0.01 and 0.07, or 0.01 and 0.05.

[0024] The first nvPM emission index ratio between idle and PMD can be greater than or equal to 0.0103 and preferably greater than or equal to 0.0115 and more preferably greater than or equal to 0.0128.

[0025] The first nvPM emission index ratio between idle and PMD can be in the range of 0.0103 to 0.105 and preferably in the range of 0.0115 to 0.0962 and more preferably in the range of 0.0128 to 0.0875.

[0026] A second nvPM emission index ratio between idle and PMD can be defined as:

[0027] [Math.2] FF / Elmxsnl l,p

[0028] where:

[0029] £7raiti,sAF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine if it is operating at about 7% of available thrust under the given operating conditions, or under other different operating conditions, and if a fuel supplied to the combustor device includes a sustainable aviation fuel;

[0030] £7maxTo,sAF is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine if it is operating at approximately 100% of the available thrust under the same given operating conditions in which £7raiti,sAF is calculated, and if a fuel supplied to the combustor device includes a sustainable aviation fuel;

[0031] Elraienti,ff is the nvPM emission index corrected for losses in the system 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 £7raiti,sAF is calculated, and if a fuel supplied to the combustor device is a fossil-based hydrocarbon fuel;

[0032] Ê / maxTorF 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 £7raiti,sAF is calculated, and if a fuel supplied to the combustor device is a fossil-based hydrocarbon fuel;

[0033] and in which the second ratio of the nvPM emission index between idle and PMD of the gas turbine engine may be less than 1.

[0034] The second nvPM emission index ratio between idle and PMD may be greater than zero.

[0035] The second nvPM emission index ratio between idle and PMD 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.

[0036] The second nvPM emission index ratio between idle and PMD 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.

[0037] The second nvPM emission index ratio between idle and PMD 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.

[0038] The second nvPM emission index ratio between idle and PMD 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.

[0039] The second nvPM emission index ratio between idle and PMD 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.

[0040] The second NVPM emission index ratio between idle and PMD 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 emission index ratio of nvPM between idle and PMD can be in a range between 0.25 and 0.4 or 0.3 and 0.35.

[0041] According to a second aspect, a gas turbine engine for an aircraft is provided, comprising one or more of the following features:

[0042] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; and wherein:

[0043] A second ratio of the nvPM emission index between idle and PMD can be defined as:

[0044] [Math.3] Elralent-LSAF / P] _ / tùæxTO. SAF PF / EIm^TO [!F

[0045] where:

[0046] Elraienti,saf is the nvPM emissions index adjusted for losses in the system in mg / kg of the gas turbine engine if it operates at approximately 7% of the available thrust under given operating conditions and if a fuel supplied to the combustor device includes a sustainable aviation fuel;

[0047] £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 £7raiti,sAF is calculated, and if a fuel supplied to the combustor device includes a sustainable aviation fuel;

[0048] ^ / slow,ff is the nvPM emission index corrected for losses in the system 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 £7raiti,sAF is calculated, and if a fuel supplied to the combustor device is a fossil-based hydrocarbon fuel;

[0049] Ê / maxTorF 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 £7raiti,sAF is calculated, and if a fuel supplied to the combustor device is a fossil-based hydrocarbon fuel;

[0050] the second ratio of the nvPM emission index between idle and PMD of the gas turbine engine is less than 1; and

[0051] The gas turbine engine is designed to supply fuel including sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0052] The second nvPM emission index ratio between idle and PMD defined in the second aspect can be as defined above in relation to the first aspect.

[0053] 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 fuel spray nozzles.

[0054] According to a fourth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0055] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; and wherein:

[0056] A first ratio of the nvPM emission index between idle and PMD can be defined as:

[0057] [Math.4] 121 slow motion ^^maxTO

[0058] where:

[0059] Δ / 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

[0060] 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;

[0061] the first ratio of the nvPM emission index between idle and PMD of the gas turbine engine is less than 0.8; and

[0062] The process includes supplying fuel comprising a sustainable aviation fuel to the fuel spray nozzles.

[0063] The first nvPM emission index ratio between idle and PMD can be as defined above in relation to the first aspect.

[0064] A second nvPM emission index ratio between idle and PMD can be defined as:

[0065] [Math.5] EJ slowed down SAF / pr + / 1 maxTÜ, SAP The slow motion, FF / p ] / ^maxTO, FF

[0066] where:

[0067] £' / raienti,sAF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine if it is operating 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;

[0068] £'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 £7raiti,sAF is calculated, and if a fuel supplied to the combustor device includes a sustainable aviation fuel;

[0069] £7raiti,ff is the nvPM emissions index adjusted for losses in the system 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 £7raiti,sAF is calculated, and if a fuel supplied to the combustor device is a fossil-based hydrocarbon fuel;

[0070] £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 under which saf is calculated, and if a fuel supplied to the combustor device is a fossil-based hydrocarbon fuel; and

[0071] in which the second nvPM emission index ratio between idle and PMD of the gas turbine engine may be less than 1.

[0072] The second nvPM emission index ratio between idle and PMD can be as defined above in relation to the first aspect.

[0073] According to a fifth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0074] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; and wherein:

[0075] A second ratio of the nvPM emission index between idle and PMD can be defined as:

[0076] [Math.6] Saf / Elma!cfo, SAF EI™1^ Fe / EltnaiTQ l,p

[0077] where:

[0078] £ / raienti,sAF is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 7% of the thrust available under given operating conditions and if a fuel supplied to the combustor device includes a sustainable aviation fuel;

[0079] £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 saf is calculated, and if a fuel supplied to the combustor device includes a sustainable aviation fuel;

[0080] £7raienti,FF 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 ELaienti.sAF is calculated, and if a fuel supplied to the combustor device is a fossil-based hydrocarbon fuel;

[0081] £7maxTo,FF 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 £7raiti,sAF is calculated, and if a fuel supplied to the combustor device is a fossil-based hydrocarbon fuel;

[0082] the second ratio of the nvPM emission index between idle and PMD of the gas turbine engine is less than 1; and

[0083] The process includes supplying fuel comprising a sustainable aviation fuel to the fuel spray nozzles.

[0084] The second ratio of nvPM emission index between idle and PMD can be as defined above in relation to the first aspect.

[0085] According to a sixth aspect, a gas turbine engine for an aircraft is provided, comprising one or more of the following features:

[0086] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; and wherein:

[0087] A fuel flow nvPM emission index ratio can be defined as:

[0088] [Math.7] raiti'^'^ (.slowdown maiTO

[0089] where:

[0090] £ / raiti 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;

[0091] 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;

[0092] Wf Ralient is the fuel flow rate to the spray nozzles of fuel in kg / s at approximately 7% of the available thrust under given operating conditions; and

[0093] IVf^axTO 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;

[0094] the nvPM emission index ratio of the gas turbine engine fuel flow is less than 0.08; and

[0095] The gas turbine engine is designed to supply fuel including sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0096] The fuel flow nvPM emission index ratio may be less than 0.0798 and preferably less than 0.0731 and more preferably less than 0.0665.

[0097] The fuel flow nvPM emission index ratio may be less than or equal to 0.06 and preferably less than or equal to 0.04 and more preferably less than or equal to 0.02.

[0098] The fuel flow nvPM emission index ratio may be less than or equal to 0.0119 and preferably less than or equal to 0.0109 and more preferably less than or equal to 0.00986.

[0099] The fuel flow nvPM emission index ratio may be less than 0.01, 0.015, 0.02, 0.025, 0.03, 0.035, 0.04, 0.045, 0.05, 0.055, 0.06, 0.065, 0.07, 0.075, 0.08 or within any range defined between any two of these values.

[0100] The fuel flow nvPM emission index ratio may be greater than zero. The fuel flow nvPM emission index ratio may be greater than or equal to 0.000993 and preferably greater than or equal to 0.00111 and more preferably greater than or equal to 0.00124.

[0101] The fuel flow nvPM emission index ratio may be in the range of 0.000993 to 0.0119 and preferably in the range of 0.00111 to 0.0109 and even more preferably in the range of 0.00124 to 0.00986.

[0102] The fuel flow nvPM emission index ratio may be less than 0.0009, 0.003, 0.005, 0.007, 0.009, 0.011, 0.013, 0.015, 0.017, 0.019, 0.021, 0.03, 0.04, 0.05, 0.06, 0.07, or 0.08, or within any defined range between any two of these values. For example, the nvPM emission index ratio fuel flow 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.

[0103] Wf idle can be in the range of 0.0516 to 0.119 kg / s. Wf r-idle can be in the range of 0.0581 to 0.109 kg / s. Wf idle can be in the range of 0.0645 to 0.0990 kg / s. Wfr idle can be in the range of 0.0645 to 0.0850 kg / s. IVf idle can be in the range of 0.0645 to 0.0750 kg / s.

[0104] WfjnaxTo May be in the range of 0.441 to 1.23 kg / s. Wfj-^axTO May be in the range of 0.496 to 1.13 kg / s. W fjnaxTO Can be in the range of 0.551 to 1.03 kg / s. Wf^axTO Can be in the range of 0.551 to 0.850 kg / s. IVfjnaXTO Can be in the range of 0.551 to 0.750 kg / s.

[0105] 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 fuel spray nozzles.

[0106] According to an eighth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0107] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; and wherein:

[0108] A fuel flow nvPM emission index ratio can be defined as:

[0109] [Math.8] And slowed down^Wf, slowed down niaxTf^ Wf,. maxTO

[0110] where:

[0111] 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

[0112] 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;

[0113] Wf idle is 'c fuel flow rate to the spray nozzles of fuel in kg / s at approximately 7% of the available thrust under given operating conditions; and

[0114] WfjuaxTO is the fuel flow rate to the fuel spray nozzles in kg / s at approximately 100% of the available thrust under the given operating conditions;

[0115] the gas turbine engine fuel flow nvPM emission index ratio is less than 0.08; and

[0116] The process includes supplying fuel comprising a sustainable aviation fuel to the fuel spray nozzles.

[0117] Any one of the ratios of the nvPM emissions index of fuel flow, IVf raiti and W f maxTO may be as defined above in relation to the sixth aspect.

[0118] According to a ninth aspect, a gas turbine engine for an aircraft is provided, comprising one or more of the following features:

[0119] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; and wherein:

[0120] A thrust nvPM emission index ratio can be defined as:

[0121] [Math.9] E~To / f„to FIralenti / p 7 r slowed down

[0122] where:

[0123] 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;

[0124] Ê / maxio 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;

[0125] PmaxTO is 'a Thrust of the gas turbine engine at approximately 100% of the thrust available in kN under the given operating conditions; and

[0126] F idle is 'a Gas turbine engine thrust at about 7% of thrust available in kN under the given operating conditions;

[0127] the nvPM emission index ratio at thrust is greater than 0.09; and

[0128] The gas turbine engine is designed to supply fuel comprising a sustainable aviation fuel (SAF) at fuel spray nozzles.

[0129] The nvPM emission index ratio at thrust may be greater than 0.0949 and preferably greater than 0.106 and more preferably greater than 0.118.

[0130] The nvPM emission index ratio at thrust may be greater than or equal to 0.15 and preferably greater than or equal to 0.3 and more preferably greater than or equal to 0.45.

[0131] The nvPM emission index ratio at thrust may be greater than or equal to 0.64 and preferably greater than or equal to 0.72 and more preferably greater than or equal to 0.8.

[0132] The nvPM emission index ratio at thrust may be less than or equal to 6.53 and preferably less than or equal to 5.98 and more preferably less than or equal to 5.44.

[0133] The nvPM emission index ratio at thrust can be in the range of 0.640 to 6.53 and preferably in the range of 0.720 to 5.98 and even more preferably in the range of 0.800 to 5.44.

[0134] The nvPM emission index ratio at thrust may be greater than 0.094, 0.1, 0.2, 0.3, 0.4, 0.5, 0.6, 0.7, 0.8, 0.9, 1, 1.1, 1.2, 1.3, 1.4, 1.5, 1.6, 1.7, 1.8, 1.9, 2, 2.5, 3, 3.5, 4, 4.5, 5, 5.5, or 6, or within any defined range between any two of these values. For example, the nvPM emission index ratio at thrust may be in the range of 0.3 to 4.5, 0.4 to 4, 1 to 4.5, or 1.4 to 4.

[0135] FmaxTO May be in the range of 54.1 to 177 kN and preferably in the range of 60.8 kN to 163 kN and preferably in the range of 67.6 kN to 148 kN. Alternatively, FmaxfQ 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.

[0136] Fraienti May be in the range of 3.78 kN to 12.4 kN and preferably in the range of 4.26 kN to 11.4 kN. Fraieiln and preferably in the range of 4.73 kN to 10.4 kN. Alternatively, Fra}enn may be in the range of 3.5 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.

[0137] 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 fuel spray nozzles.

[0138] According to an eleventh aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0139] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device

[0140] having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; and wherein:

[0141] A thrust nvPM emission index ratio can be defined as:

[0142] [Math. 10] / iaay j Q Flraleuti / p , / In slow motion

[0143] where:

[0144] 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

[0145] 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;

[0146] FmaxTO is 'the thrust of the gas turbine engine at approximately 100% of the available thrust in kN under the given operating conditions,

[0147] F idle is 'a Gas turbine engine thrust at about 7% of thrust available in kN under the given operating conditions; and

[0148] the nvPM emission index ratio at thrust is greater than 0.09; and

[0149] The process includes supplying fuel comprising a sustainable aviation fuel to the fuel spray nozzles.

[0150] Any one of the nvPM emission index ratios at boost, FlnaxTO and F idle can be defined in relation to the ninth aspect.

[0151] According to a twelfth aspect, a gas turbine engine for an aircraft is provided, comprising one or more of the following features:

[0152] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; and wherein:

[0153] A ratio of the nvPM emission index between depleted cruise and PMD can be defined as:

[0154] [Math. 11] FI cruise (impoverished^ / PI — ' 7 '-'itiaxTO BPR

[0155] where:

[0156] ^ / cruise (impoverished) can be defined as:

[0157] [Math. 12] ^1 maxTO~^"^Mounted 2

[0158] 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;

[0159] Elmontée 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

[0160] BPR is the bypass ratio of the gas turbine engine;

[0161] the ratio of the nvPM emission index between the depleted cruise and the PMD is greater than 0.2; and

[0162] The gas turbine engine is designed to supply fuel including sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0163] The ratio of the nvPM emission index between the depleted cruise and the PMD may be less than 0.18, preferably less than 0.16, and more preferably less than 0.14.

[0164] The ratio of the nvPM emission index between the depleted cruise and the PMD may be less than 0.135, preferably less than 0.124 and more preferably less than 0.113.

[0165] The ratio of the nvPM emission index between the depleted cruise and the PMD may be less than or equal to 0.118, preferably less than or equal to 0.109, and more preferably less than or equal to 0.0983.

[0166] The ratio of the nvPM emission index between the depleted cruise and the PMD may be greater than zero. The ratio of the nvPM emission index between the depleted cruise and the PMD may be greater than or equal to 0.0426, preferably greater than or equal to 0.0479, and more preferably greater than or equal to 0.0533.

[0167] The nvPM emission index ratio between depleted cruise and PMD may be in the range of 0.0426 to 0.118, preferably in the range of 0.0479 to 0.109 and more preferably in the range of 0.0533 to 0.0983.

[0168] The nvPM emission index ratio between the depleted cruise and the 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 defined range between any two of these values.

[0169] The BPR may be in the range of 6.66 to 15.3 and more preferably in the range of 7.49 to 14.0 and even more preferably in the range of 8.33 to 12.8. Alternatively, the BPR may 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.

[0170] A ratio of the nvPM emissions index between rich cruising and PMD can be defined as:

[0171] [Math. 13] BPR

[0172] where:

[0173] Elcroisière (rich) Pcut be defined as:

[0174] [Math. 14] El - +FI, 1 mount 1 approach 2

[0175] £7 rise 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;

[0176] ^ / approach is the 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 in which £7climb is calculated; and

[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 the same operating conditions under which £7montée is calculated; and

[0178] in which the ratio of the nvPM emission index between rich cruise and PMD is less than 0.07.

[0179] The nvPM emission index ratio between rich cruise and PMD may be less than 0.065, preferably less than 0.06, and more preferably less than 0.055.

[0180] The nvPM emission index ratio between rich cruise and PMD may be less than 0.0635, preferably less than 0.0582 and more preferably less than 0.0529.

[0181] The nvPM emission index ratio between rich cruise and PMD may be less than or equal to 0.0462, preferably less than or equal to 0.0424, and more preferably less than or equal to 0.0385.

[0182] The NvPM emission index ratio between rich cruise and PMD may be greater than zero. The NvPM emission index ratio between rich cruise and PMD may be greater than or equal to 0.0128, preferably greater than or equal to 0.0144, and more preferably greater than or equal to 0.016.

[0183] The nvPM emission index ratio between rich cruise and PMD may be in the range of 0.0128 to 0.0462, preferably in the range of 0.0144 to 0.0424 and more preferably in the range of 0.0160 to 0.0385.

[0184] The nvPM emission index ratio between rich cruise and PMD may be less than 0.005, 0.01, 0.015, 0.02, 0.025, 0.03, 0.035, 0.04, 0.045, 0.05, 0.055, 0.06, 0.065, or 0.07, or within any defined range between any two of these values.

[0185] The nvPM emission index ratio between rich cruise and PMD may be less than or equal to 0.005, 0.007, 0.01, 0.013, 0.015, 0.017, 0.02, 0.023, 0.025, 0.027, 0.03, 0.033, 0.035, 0.037, 0.04, 0.043, 0.045, or 0.0462, or within any defined range between any two of these values.

[0186] According to a thirteenth aspect, a gas turbine engine for an aircraft is provided, comprising one or more of the following features:

[0187] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; and wherein:

[0188] A ratio of the nvPM emissions index between rich cruising and PMD can be defined as:

[0189] [Math. 15] croisièreiricho} / HJ ________ f '-''insxTt! BPR

[0190] where:

[0191] Elcroisière (rich) can be defined as:

[0192] [Math. 16] pj. +FT rises e^^^nppi-oche 2

[0193] 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;

[0194] ^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;

[0195] Elmmtrn 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 under the same operating conditions under which EImontée is calculated;

[0196] BPR is the bypass ratio of the gas turbine engine;

[0197] the ratio of the nvPM emissions index between the rich cruise and the PMD is greater than 0.07; and

[0198] The gas turbine engine is designed to supply fuel including sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0199] The nvPM emission index ratio between rich cruise and PMD and / or BPR may be as defined above in relation to the twelfth aspect.

[0200] 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 fuel spray nozzles.

[0201] According to a fifteenth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0202] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; and wherein:

[0203] A ratio of the nvPM emission index between depleted cruise and PMD can be defined as:

[0204] [Math. 17] (depleted) / BI„ BPR

[0205] where:

[0206] ^ / cruise (impoverished) can be defined as:

[0207] [Math. 18] PT_4-P7 ^-11 max i ' O' J-'1 monté e 2

[0208] 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;

[0209] 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

[0210] BPR is the bypass ratio of the gas turbine engine;

[0211] the ratio of the nvPM emission index between depleted cruise and PMD is greater than 0.2; and

[0212] The process includes supplying fuel comprising a sustainable aviation fuel to the fuel spray nozzles.

[0213] The nvPM emission index ratio between depleted cruise and PMD and / or BPR may be as defined above in relation to the twelfth aspect.

[0214] A ratio of the nvPM emissions index between rich cruising and PMD can be defined as:

[0215] [Math. 19] wasteland cruise} / FJ _ f jnaxTU BPR

[0216] where:

[0217] Elcroisière (rich) Pcut can be defined as:

[0218] [Math.20] El - +FI , 1 mon te 1 approach 2

[0219] £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, or under other different operating conditions;

[0220] ^ / 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 ^ / climb is calculated;

[0221] 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 under which EImOntée is calculated; and

[0222] in which the ratio of the nvPM emission index between rich cruise and PMD may be less than 0.07.

[0223] The nvPM emission index ratio between rich cruise and PMD and / or BPR may be as defined above in relation to the twelfth aspect.

[0224] According to a sixteenth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0225] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; and wherein:

[0226] A ratio of the nvPM emissions index between rich cruising and PMD can be defined as:

[0227] [Math.21] The MSXTO cruise BPR

[0228] where:

[0229] E / cruise (rich) can be defined as:

[0230] [Math.22] AND, +FI J nient ee1^2 approach 2

[0231] E / climb 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;

[0232] Ê / 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 EImontée is calculated;

[0233] 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 the same operating conditions in which EImontée is calculated;

[0234] BPR is the bypass ratio of the gas turbine engine;

[0235] the ratio of the nvPM emissions index between rich cruise and PMD may be less than 0.07; and

[0236] The process includes supplying fuel comprising a sustainable aviation fuel to the fuel spray nozzles.

[0237] The nvPM emission index ratio between rich cruise and PMD and / or BPR may be as defined above in relation to the twelfth aspect.

[0238] According to a seventeenth aspect, a gas turbine engine for an aircraft is provided, comprising one or more of the following features:

[0239] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; and wherein:

[0240] a ratio of nvPM emission index to PMD can be defined as:

[0241] [Math.23] ElmasTO, SAF EEuaxT O,FF

[0242] where:

[0243] E / maxTo.sAF is the 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 fuel spray nozzles includes sustainable aviation fuel (SAF); and

[0244] E / maxTorr 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 under given operating conditions if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0245] the ratio of the nvPM emission index to the PMD of the gas turbine engine is less than 1; and

[0246] The gas turbine engine is designed to supply fuel including a SAF to the fuel spray nozzles.

[0247] The ratio of the nvPM emission index to PMD may be greater than zero.

[0248] 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.

[0249] 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.

[0250] 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.

[0251] 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.

[0252] 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.

[0253] 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.

[0254] The ratio of the nvPM emission index to PMD may be 0.5, 0.51, 0.52, 0.53, 0.54, 0.55, 0.56, 0.57, 0.58, 0.59, 0.6, 0.61, 0.62, 0.63, 0.64, 0.65, 0.66, 0.67, 0.68, 0.69, 0.7, 0.71, 0.72, 0.73, 0.74, 0.75, 0.76, 0.77, 0.78, 0.79, 0.8, or within any range defined between any two of these values.

[0255] An upstream nvPM emission index ratio can be defined as:

[0256] [Math.24] The climb, SAF p J ■-1 ascent,FF

[0257] where:

[0258] £7,sAF is the 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, or under different given operating conditions, if a fuel supplied to the fuel spray nozzles includes sustainable aviation fuel (SAF); and

[0259] £ / climb,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 given in which £7 rise,sAF is calculated and if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0260] and in which the nvPM emission index ratio at the rise of the gas turbine engine may be less than 1.

[0261] The upstream nvPM emission index ratio may be greater than zero.

[0262] 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.

[0263] 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.

[0264] 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.

[0265] 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.

[0266] 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.

[0267] 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.

[0268] 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.

[0269] An nvPM emission index ratio approach can be defined as:

[0270] [Math.25] approach, SAP approach,FF

[0271] where:

[0272] £7apProche,sAF is the 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, or under different given operating conditions, if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel (SAF); and

[0273] £7approach,FF is the 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 in which £7approach,sAF is calculated and if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0274] in which the nvPM emission index ratio approaching the gas turbine engine may be less than 1.

[0275] The nvPM emission index ratio at the approach may be greater than zero.

[0276] The nvPM emission index ratio of 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.

[0277] 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.

[0278] 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.

[0279] 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.

[0280] 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.

[0281] 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.

[0282] The nvPM emission index ratio of 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.

[0283] An nvPM emission index ratio at idle can be defined as:

[0284] [Math.26] Elralenti, SAF Fl, 1 raiti.r F

[0285] where:

[0286] £7raiti,sAF is the 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, or under different operating conditions, and whether a fuel supplied to the fuel spray nozzles includes sustainable aviation fuel (SAF); and

[0287] £7raienti,FF 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 under which saf is calculated and if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0288] in which the nvPM emission index ratio at idle of the gas turbine engine may be less than 1.

[0289] The nvPM emission index ratio at idle may be greater than zero.

[0290] 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.

[0291] 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.

[0292] 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.

[0293] 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.

[0294] 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.

[0295] 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.

[0296] The nvPM emission index ratio at idle may be 0.06, 0.07, 0.08, 0.09, 0.1, 0.11, 0.12, 0.13, 0.14, 0.15, 0.16, 0.17, 0.18, 0.19, 0.2, 0.21, 0.22, 0.23, 0.24, 0.25, 0.26, 0.27, 0.28, 0.29, 0.3, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39, or 0.4, or within any defined range between any two of these values.

[0297] According to an eighteenth aspect, a gas turbine engine for an aircraft is provided, comprising one or more of the following features:

[0298] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; and wherein:

[0299] an upstream nvPM emission index ratio can be defined as:

[0300] [Math.27] SAF ^^inontée.FF

[0301] where:

[0302] £ / climb,sAF is the 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 fuel spray nozzles includes sustainable aviation fuel (SAF); and

[0303] £ / climb,FF 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 given operating conditions in which EImontéeetSAF is calculated and if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0304] the nvPM emission index ratio at the gas turbine engine's climb is less than 1; and

[0305] The gas turbine engine is designed to supply fuel including a SAF to the fuel spray nozzles.

[0306] The upstream nvPM emission index ratio can be as defined in relation to the seventeenth aspect.

[0307] According to a nineteenth aspect, a gas turbine engine for an aircraft is provided, comprising one or more of the following features:

[0308] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; and wherein:

[0309] an nvPM emission index ratio approaching can be defined as:

[0310] [Math.28] Approach, SAF Approach, FF

[0311] where:

[0312] ^ / approach,saf is the nvPM emissions index corrected for losses in the system in mg / kg of the gas turbine engine during operation at approximately 30% of available thrust under given operating conditions if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel (SAF); and

[0313] £ZiPProche,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 under the same given operating conditions in which £7approche,sAF is calculated and if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0314] the nvPM emission index ratio approaching the gas turbine engine is less than 1; and

[0315] The gas turbine engine is designed to supply fuel including a SAF to the fuel spray nozzles.

[0316] The nvPM emission index ratio to the approach may be as defined above in relation to the seventeenth aspect.

[0317] According to a twentieth aspect, a gas turbine engine for an aircraft is provided, comprising one or more of the following features:

[0318] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; and wherein:

[0319] An NVPM emission index ratio at idle can be defined as:

[0320] [Math.29] Elralenti, SAF El lentement,FF

[0321] where:

[0322] £7raiti,sAF is the 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 fuel spray nozzles includes sustainable aviation fuel (SAF); and

[0323] ^ / idle,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 under the same given operating conditions in which ELaienti.sAF is calculated and if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0324] the nvPM emission index ratio at idle of the gas turbine engine is less than 1; and

[0325] The gas turbine engine is designed to supply fuel including a SAF to the fuel spray nozzles.

[0326] The nvPM emission index ratio at idle can be as defined above in relation to the seventeenth aspect

[0327] According to a twenty-first aspect, a method of operating the gas turbine engine according to any one of the seventeenth, eighteenth, nineteenth or twentieth aspects is provided, the method comprising supplying fuel including a sustainable aviation fuel to the fuel spray nozzles.

[0328] According to a twenty-second aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0329] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; and wherein:

[0330] A ratio of the nvPM emission index to PMD can be defined as:

[0331] [Math.30] EIniaxTO, SAF EïmasTO.FF

[0332] where:

[0333] £7maxTo,sAF is the 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 fuel spray nozzles includes a sustainable aviation fuel (SAF); and

[0334] ^7maxTo,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 under given operating conditions if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0335] the ratio of the nvPM emission index to the PMD of the gas turbine engine is less than 1; and

[0336] The process includes supplying fuel comprising a sustainable aviation fuel to the fuel spray nozzles.

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

[0338] An upstream nvPM emission index ratio can be defined as:

[0339] [Math.31] El montée, SAF 1:1 montée,FF

[0340] where:

[0341] Æ' / climb.sAF is the 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, or under different given operating conditions, if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel (SAF); and

[0342] Ί / montéerF is the 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 given in which EImontéeetSAP is determined and if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and in which

[0343] the nvPM emission index ratio at the climb of the gas turbine engine may be less than 1.

[0344] The upstream nvPM emission index ratio can be as defined in relation to the seventeenth aspect.

[0345] An nvPM emission index ratio approach can be defined as:

[0346] [Math.32] EJ approach, SAP ET. "1 approach, FF

[0347] where:

[0348] ^ / approach,saf is the 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, or under different given operating conditions, if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel (SAF); and

[0349] £7approach,ff is the 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 given in which EZapproach.sAF is determined and if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and in which

[0350] the nvPM emission index ratio approaching the gas turbine engine may be less than 1.

[0351] The nvPM emission index ratio to the approach may be as defined in relation to the seventeenth aspect.

[0352] An nvPM emission index ratio at idle can be defined as:

[0353] [Math.33] EJraiti, SAP EpglenthFP

[0354] where:

[0355] £4aienti,sAF is the 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, or under different given operating conditions, and whether a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel (SAF); and

[0356] £7raiti,FF 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 under which £7raiti,sAF is determined and if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0357] in which the nvPM emission index ratio at idle of the gas turbine engine may be less than 1.

[0358] The nvPM emission index ratio at idle can be as defined in relation to the seventeenth aspect.

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

[0360] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; and wherein:

[0361] an upstream nvPM emission index ratio can be defined as:

[0362] [Math.34] The slowdown, SAF pr ^slowerErF

[0363] where:

[0364] ^ / climb,saf is the 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 fuel spray nozzles includes a sustainable aviation fuel (SAF); and

[0365] Ί / montéerF 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 given operating conditions in which EImontéeetSAP is determined and if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0366] the nvPM emission index ratio at the climb of the gas turbine engine may be less than 1; and

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

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

[0369] According to a twenty-fourth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0370] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; and wherein:

[0371] an nvPM emission index ratio approach can be defined as:

[0372] [Math.35] EI approaches, SAF El approaches, FF

[0373] where:

[0374] £7approach,sAF is the 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 fuel spray nozzles comprises a sustainable aviation fuel (SAF); and

[0375] £7apProche,FF is the 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 in which £7apProche,sAF is determined and if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0376] the nvPM emission index ratio approaching the gas turbine engine may be less than 1; and

[0377] The process includes supplying fuel comprising a sustainable aviation fuel to the fuel spray nozzles.

[0378] The nvPM emission index ratio to the approach may be as defined in relation to the seventeenth aspect.

[0379] According to a twenty-fifth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0380] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; and wherein:

[0381] An NVPM emission index ratio at idle can be defined as:

[0382] [Math.36] SAF EIralenti,FF

[0383] where:

[0384] £7raiti,sAF is the 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 fuel spray nozzles includes sustainable aviation fuel (SAF); and

[0385] £7raienti,FF 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 in which saf is determined and if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0386] the nvPM emission index ratio at idle of the gas turbine engine may be less than 1; and

[0387] The process includes supplying fuel comprising a sustainable aviation fuel to the fuel spray nozzles.

[0388] The nvPM emission index ratio at idle can be as defined in relation to the seventeenth aspect.

[0389] According to a twenty-sixth aspect, a gas turbine engine for an aircraft is provided, comprising one or more of the following features:

[0390] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; and wherein:

[0391] A fuel flow modified by the ratio of the emission index of nvPM to PMD can be defined as:

[0392] [Math.37] EIIUaxTO, SAF , , ta r EI^aff X WOnaxTO

[0393] where:

[0394] £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 under given operating conditions if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel (SAF);

[0395] £ / maxTo,FF is the 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 fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0396] TVf jnaxTO is the mass flow rate of fuel supplied to the spray nozzles of fuel in kg / s when the gas turbine engine is operating at approximately 100% of the available thrust under given operating conditions;

[0397] the fuel flow modified by the ratio of the emission index of nvPM to the PMD of the gas turbine engine in kg / s is less than 2; and

[0398] The gas turbine engine is designed to supply fuel including a SAF to the fuel spray nozzles.

[0399] The fuel flow modified by the ratio of the emission index of nvPM to PMD in kg / s may be greater than zero.

[0400] The fuel flow modified by the ratio of the emission index of nvPM to PMD in kg / s may be less than 1.23, more preferably less than 1.13 and even more preferably less than 1.03.

[0401] The fuel flow modified by the ratio of the emission index of nvPM to PMD in kg / s may be less than or equal to 0.793, more preferably less than or equal to 0.727 and more preferably less than or equal to 0.661.

[0402] The fuel flow modified by the ratio of the emission index of nvPM to PMD in kg / s may be greater than or equal to 0.284, preferably greater than or equal to 0.32 and more preferably greater than or equal to 0.356.

[0403] The fuel flow modified by the ratio of the emission index of nvPM to PMD in kg / s can be in the range of 0.284 to 0.793, preferably in the range of 0.320 to 0.727 and more preferably in the range of 0.356 to 0.661.

[0404] The fuel flow modified by the ratio of the emission index of nvPM to PMD in kg / s may be less than or equal to 0.2, 0.25, 0.3, 0.35, 0.4, 0.45, 0.5, 0.55, 0.6, 0.65, 0.7, 0.75, 0.8, 0.85, 0.9, 0.95, 1, 1.05, 1.1, 1.15, 1.2, 1.25, or 1.3, or within any defined range between any two of these values. For example, the fuel flow modified by the ratio of the emission index of nvPM to PMD in kg / s may be in a range between 0.45 and 0.65 or between 0.45 and 0.6.

[0405] Wf^axTO May be in the range of 0.441 to 1.23 kg / s, and preferably may be in the range of 0.496 to 1.13 kg / s, and more preferably can be in the range of 0.551 to 1.03 kg / s. Alternatively, Wf^axTO can be in the range of 0.551 to 0.850 kg / s. Wf^naxTO can be in the range of 0.551 to 0.750 kg / s.

[0406] A fuel flow modified by the nvPM emission index ratio on climb can be defined as:

[0407] [Math.38] inontSe, SAF T 4 Z EImantée,FF XW fJUOIltée

[0408] where:

[0409] £7,sAF is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at about 85% of available thrust under the given operating conditions, or under other different operating conditions, if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; and

[0410] £ / climb,FF is the 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 in which ^ / climb,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0411] Wf bridged is 'c mass flow rate of fuel supplied to the spray nozzles of fuel in kg / s when the gas turbine engine is operating at approximately 85% of available thrust under the same operating conditions in which ^ / climb.sAF and EImontéetPP are calculated; and in which

[0412] the fuel flow modified by the ratio of the emission index of nvPM at the climb of the gas turbine engine can be kg / s is less than 2.

[0413] The fuel flow modified by the ratio of the emission index of nvPM on climb in kg / s may be greater than zero.

[0414] The fuel flow modified by the ratio of the emission index of nvPM on climb in kg / s may be less than 1.01, more preferably less than 0.923 and even more preferably less than 0.839.

[0415] The fuel flow modified by the ratio of the emission index of nvPM on climb in kg / s may be less than or equal to 0.479, more preferably less than or equal to 0.439 and more preferably less than or equal to 0.399.

[0416] The fuel flow modified by the ratio of the emission index of nvPM on climb in kg / s may be greater than or equal to 0.175, preferably greater than or equal to 0.197 and more preferably greater than or equal to 0.219.

[0417] The fuel flow modified by the ratio of the emission index of nvPM on climb in kg / s can be in the range of 0.175 to 0.479, preferably in the range of 0.197 to 0.439 and more preferably in the range of 0.219 to 0.399.

[0418] The fuel flow modified by the nvPM emission index ratio on climb may be less than or equal to 0.15, 0.175, 0.2, 0.225, 0.25, 0.275, 0.3, 0.325, 0.35, 0.375, 0.4, 0.425, 0.45, 0.475, 0.5, 0.525, 0.55, 0.575, 0.6, 0.625, 0.65, 0.675, 0.7, 0.725, 0.75, 0.775, 0.8, 0.825, 0.85, 0.875, 0.9, 0.925, 0.95, 0.975, or 1, or within any defined range between any two of these values. For example, the fuel flow modified by the NvPM emissions index ratio on climb in kg / s could be in a range between 0.275 and 0.475 or 0.3 and 0.4.

[0419] Wf jnoDtée May be in the range of 0.369 to 1.01 kg / s, and preferably may be in the range of 0.415 to 0.923 kg / s, and more preferably in the range of 0.461 to 0.839 kg / s. Alternatively, IVf^lontée can be in the range of 0.461 to 0.650 kg / s. Wfpontée can be in the range of 0.461 to 0.600 kg / s.

[0420] A fuel flow modified by the nvPM emission index ratio on approach can be defined as:

[0421] [Math.39] approach, SAF EIapprOche,FF XW f,approche

[0422] where:

[0423] £7apProche,sAF is the nvPM emissions index adjusted for losses in the system in mg / kg of the gas turbine engine when operating at approximately 30% of available thrust under given operating conditions, or under other different operating conditions, if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; and

[0424] £7aPProche,FF is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at about 30% of available thrust under the same operating conditions in which ^4pprochejsAF is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0425] Wf approach is the mass flow rate of fuel supplied to the spray nozzles of fuel in kg / s when the gas turbine engine is operating at approximately 30% of available thrust under the same operating conditions in which Approach,saf and EZapproach,FF are calculated; and in which

[0426] the fuel flow modified by the nvPM emission index ratio approaching the gas turbine engine in kg / s may be less than 0.4.

[0427] The fuel flow modified by the nvPM emission index ratio to the kg / s approach may be greater than zero.

[0428] The fuel flow modified by the nvPM emission index ratio approach in kg / s may be less than 0.334, more preferably less than 0.306 and even more preferably less than 0.278.

[0429] The fuel flow modified by the nvPM emission index ratio approach in kg / s may be less than or equal to 0.0513, more preferably less than or equal to 0.047 and further preferably less than or equal to 0.0428.

[0430] The fuel flow modified by the nvPM emission index ratio approach in kg / s may be greater than or equal to 0.0204, preferably greater than or equal to 0.0229 and even more preferably greater than or equal to 0.0255.

[0431] The fuel flow modified by the emission index ratio of nvPM to the kg / s approach can be in the range of 0.0204 to 0.0513, preferably in the range of 0.0229 to 0.0470 and more preferably in the range of 0.0255 to 0.0428.

[0432] The fuel flow modified by the nvPM emission index ratio at the approach in kg / s may be less than or equal to 0.02, 0.05, 0.075, 0.1, 0.125, 0.15, 0.175, 0.2, 0.225, 0.25, 0.275, 0.3, 0.325, 0.35, 0.375, or 0.4, or within any defined range between any two of these values. For example, the fuel flow modified by the nvPM emission index ratio at the approach in kg / s may be in a range between 0.02 and 0.1 or 0.05 and 0.075.

[0433] Wf approach May be in the range of 0.133 to 0.334 kg / s, and preferably may be in the range of 0.149 to 0.306 kg / s, and more preferably in the range of 0.166 to 0.278 kg / s. Alternatively, Wf 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.

[0434] A fuel flow modified by the nvPM emission index ratio at idle can be defined as:

[0435] [Math.40] Elraienti, SAP EIralenti,FF X Wfsalent!

[0436] where:

[0437] £7raiti,sAF 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 given operating conditions, or under other different operating conditions, if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; and

[0438] £7aienti,FF 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 operating conditions in which £'4aienti,sAF is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0439] IVfralenti is the mass flow rate of fuel supplied to the spray nozzles of fuel in kg / s when the gas turbine engine is operating at approximately 7% of available thrust under the same operating conditions in which idle, safe and £7raiti, FF are calculated; and in which

[0440] the fuel flow modified by the emission index ratio of nvPM at idle of the gas turbine engine in kg / s may be less than 0.2.

[0441] The fuel flow modified by the nvPM emission index ratio at idle in kg / s may be greater than zero.

[0442] The fuel flow modified by the emission index ratio of nvPM at idle in kg / s may be less than 0.119, more preferably less than 0.109 and even more preferably less than 0.099.

[0443] The fuel flow modified by the emission index ratio of nvPM at idle in kg / s may be less than or equal to 0.0114, more preferably less than or equal to 0.0105 and still preferably less than or equal to 0.00949.

[0444] The fuel flow modified by the emission index ratio of nvPM at idle in kg / s may be greater than or equal to 0.00494, preferably greater than or equal to 0.00556 and even more preferably greater than or equal to 0.00618.

[0445] The fuel flow modified by the emission index ratio of nvPM at idle in kg / s can be in the range of 0.00494 to 0.0114, preferably in the range of 0.00556 to 0.0105 and more preferably in the range of 0.00618 to 0.00949.

[0446] The fuel flow modified by the emission index ratio of nvPM at idle in kg / s may be less than or equal to 0.004, 0.01, 0.015, 0.02, 0.025, 0.03, 0.035, 0.04, 0.045, 0.05, 0.055, 0.06, 0.065, 0.07, 0.075, 0.08, 0.085, 0.09, 0.095, 0.1, 0.11, or 0.12, or within any defined range between any two of these values. For example, the fuel flow modified by the nvPM emission index ratio at idle in kg / s can be in a range between 0.01 and 0.03 or 0.015 and 0.025.

[0447] IVf idle speed may be in the range of 0.0516 to 0.119 kg / s, and preferably may be in the range of 0.0581 to 0.109 kg / s, and more preferably may be in the range of 0.0645 to 0.0990 kg / s. Alternatively, IVf idle speed may be in the range of 0.0645 to 0.0850 kg / s. TVfr idle speed may be in the range of 0.0645 to 0.0750 kg / s.

[0448] According to a twenty-seventh aspect, a gas turbine engine for an aircraft is provided, comprising one or more of the following features:

[0449] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; and wherein:

[0450]

[0451]

[0452]

[0453]

[0454]

[0455]

[0456]

[0457]

[0458]

[0459]

[0460]

[0461]

[0462] A fuel flow modified by the nvPM emission index ratio during ascent can be defined as: [Math.41] The climb, SAF T 4 Z EImantée,FF X VV fpontée Or : £7montée,sAF is the nvPM emissions 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 fuel spray nozzles includes sustainable aviation fuel; and EImontée,FF is the 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 under which EImontée,sAF is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and IVf^oHtée is the mass flow rate of fuel supplied to the fuel spray nozzles in kg / s when the gas turbine engine is operating at approximately 85% of available thrust under the same operating conditions in which ^ / climb,SAF ^t ^ / climb,FF SOttt Calculates , the fuel flow modified by the nvPM emission index ratio at the gas turbine engine's climb rate is kg / s less than 2; and The gas turbine engine is designed to supply fuel including a SAF to the fuel spray nozzles. The fuel flow modified by an emission index ratio of nvPM on ascent and / or IVf^lontée Can be as defined above in relation to the twenty-sixth aspect. According to a twenty-eighth aspect, a gas turbine engine for an aircraft is provided, comprising one or more of the following features: a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; and wherein: A fuel flow modified by the nvPM emission index ratio at the approach can be defined as: [Math.42] '1 approach, SAF The approach, FF XW approaches

[0463] where:

[0464] ÆZipproche.sAF is the 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, or under other different operating conditions, if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; and

[0465] Elapproche,ff is the nvPM emissions index corrected for losses in the system in mg / kg of the gas turbine engine during operation at approximately 30% of available thrust under the same operating conditions in which ^ / approach,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0466] Wf approach is the mass flow rate of fuel supplied to the spray nozzles of fuel in kg / s when the gas turbine engine is operating at approximately 30% of available thrust under the same operating conditions in which ^ / approach,SAF and ^ / approach,FF SOttt Calculates , and daUS which

[0467] the fuel flow modified by the nvPM emission index ratio approaching the gas turbine engine in kg / s is less than 0.4; and

[0468] The gas turbine engine is designed to supply fuel including a SAF to the fuel spray nozzles.

[0469] The fuel flow modified by an nvPM emission index ratio at approach and / or W Approach may be as defined above in relation to the twenty-sixth aspect.

[0470] According to a twenty-ninth aspect, a gas turbine engine for an aircraft is provided, comprising one or more of the following features:

[0471] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; and wherein:

[0472] A fuel flow modified by the nvPM emission index ratio at idle can be defined as:

[0473] [Math.43] The slowdown, SAF Tjy "EÏr^ntnFF X slowed down

[0474] where:

[0475] ^ / slow,saf is the nvPM emissions index corrected for losses in the system in mg / kg of the gas turbine engine during operation at approximately 7% of available thrust under given operating conditions and if a fuel supplied to the fuel spray nozzles includes sustainable aviation fuel; and

[0476] £7raiti,FF 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 operating conditions in which Ralido,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0477] PVfralenti is the mass flow rate of fuel supplied to the spray nozzles of fuel in kg / s when the gas turbine engine is operating at about 7% of available thrust under the same operating conditions in which ^idle,SAF and £7idle,FF SOUt Calculated;

[0478] the fuel flow modified by the emission index ratio of nvPM at idle of the gas turbine engine in kg / s is less than 0.2; and

[0479] The gas turbine engine is designed to supply fuel including a SAF to the fuel spray nozzles.

[0480] The fuel flow modified by an emission index ratio of nvPM at idle and / or Wfralenti may be as defined above in relation to the twenty-sixth aspect.

[0481] 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 fuel spray nozzles.

[0482] According to a thirty-first aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0483] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; wherein:

[0484] A fuel flow modified by the ratio of the emission index of nvPM to PMD can be defined as:

[0485] [Math.44] ElmaxTO, SAF EI!liaxTo.FF X Wf,maxTO

[0486] where:

[0487] £'7maxTo,sAF is the 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 fuel spray nozzles includes sustainable aviation fuel (SAF);

[0488] £ / maxTo,FF is the 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 fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0489] WfjnaxTo is the mass flow rate of fuel supplied to the spray nozzles of fuel in kg / s when the gas turbine engine is operating at approximately 100% of the available thrust under given operating conditions;

[0490] the fuel flow modified by the ratio of the emission index of nvPM to PMD in kg / s is less than 2; and

[0491] The process includes supplying fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0492] The fuel flow modified by an emission index ratio of nvPM to PMD and / or WfjnaxTO may be as defined above in relation to the twenty-sixth aspect.

[0493] A fuel flow modified by the nvPM emission index ratio on climb can be defined as:

[0494] [Math.45] El mounted, SAF T47 EImanted, Fp X VVf bridged

[0495] where:

[0496] rise, saf is the nvPM emissions index adjusted for losses in the system in mg / kg of the gas turbine engine when operating at approximately 85% of available thrust under given operating conditions, or under other different operating conditions, if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; and

[0497] Ê / montéerF is the 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 in which £7montée,sAF is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0498] Wf bridged is 'c mass flow rate of fuel supplied to the spray nozzles of fuel in kg / s when the gas turbine engine is operating at approximately 85% of available thrust under the same operating conditions in which LAnomée.sAi and EImontéeSsp are calculated; and in which

[0499] the fuel flow modified by the ratio of the emission index of nvPM at the climb of the gas turbine engine in kg / s may be less than 2.

[0500] The fuel flow modified by an emission index ratio of nvPM on climb and / or bridged Wf may be as defined above in relation to the twenty-sixth aspect.

[0501] A fuel flow modified by the nvPM emission index ratio on approach can be defined as:

[0502] [Math.46] The approach, SAF The approach, FF VV f, approach

[0503] where:

[0504] ^^approach.sAF is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 30% of available thrust under the given operating conditions, or under other different operating conditions, if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; and

[0505] Elapproche,ff is the nvPM emissions index corrected for losses in the system in mg / kg of the gas turbine engine during operation at approximately 30% of available thrust under the same operating conditions in which E / approach,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0506] Wf approach is the mass flow rate of fuel supplied to the spray nozzles of fuel in kg / s when the gas turbine engine is operating at approximately 30% of available thrust under the same operating conditions in which ^ / approach,SAF and / ^approach,FF SOttt Calculates , and daUS which

[0507] The fuel flow modified by the nvPM emission index ratio approaching the gas turbine engine in kg / s may be less than 0.4.

[0508] The fuel flow modified by an approach nvPM emission index ratio and / or approach Wf can be as defined above in relation to the twenty-sixth aspect.

[0509] A fuel flow modified by the nvPM emission index ratio at idle can be defined as:

[0510] [Math.47] Elralenti, SAF jp EIra}e!M,FF X Wf slowed down

[0511] where:

[0512] £7raiti,sAF is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 7%

[0513]

[0514]

[0515]

[0516]

[0517]

[0518]

[0519]

[0520]

[0521]

[0522]

[0523] of the thrust available under the given operating conditions, or under other different operating conditions, if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; and Ralido,ff 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 operating conditions in which Ralido,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and IVfjajenti is the mass flow rate of fuel supplied to the fuel spray nozzles in kg / s when the gas turbine engine is operating at approximately 7% of available thrust under the same operating conditions in which E / raienti,saf and EZraientijF are calculated; and in which the fuel flow modified by the emission index ratio of nvPM at idle of the gas turbine engine in kg / s may be less than 0.2. The fuel flow modified by an NVPM emission index ratio at idle and / or Wfj-aientiTO may be as defined above in relation to the twenty-sixth appearance. According to a thirty-second aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising: a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; and wherein: A fuel flow modified by the nvPM emission index ratio during ascent can be defined as: [Math.48] The climb, SAP -rr Emmitée,FF X Wf,montée Or : Elmontée.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 fuel spray nozzles includes sustainable aviation fuel; and 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 in which ^ascent,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0524] Wf bridged is 'c mass flow rate of fuel supplied to the spray nozzles of fuel in kg / s when the gas turbine engine is operating at approximately 85% of available thrust under the same operating conditions in which ^ / climb,SAF and Elclimb,FF SOttt Calculates ,

[0525] the fuel flow modified by the ratio of the nvPM emission index to the climb of the gas turbine engine in kg / s is less than 2; and

[0526] The process includes supplying fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0527] The fuel flow modified by an emission index ratio of nvPM on ascent and / or WfpWntée May be as defined above in relation to the twenty-sixth aspect.

[0528] According to a thirty-third aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0529] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; and wherein:

[0530] A fuel flow modified by the nvPM emission index ratio on approach can be defined as:

[0531] [Math.49] He approached SAF EIappeoche,FF A vv Approach

[0532] where:

[0533] £7approach,sAF is the 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, or under other different operating conditions, if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; and

[0534] £7aPProche,FF is the 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 in which Approche,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0535] Wf approach is the mass flow rate of fuel supplied to the spray nozzles of fuel in kg / s when the gas turbine engine is operating at approximately 30% of the thrust available under the same operating conditions in which ^^approach.SAF ^t Elapproche,FF SOttt Calcules ,

[0536] the fuel flow modified by the nvPM emission index ratio approaching the gas turbine engine in kg / s is less than 0.4; and

[0537] The process includes supplying fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0538] The fuel flow modified by an nvPM emission index ratio at approach and / or W Approach may be as defined above in relation to the twenty-sixth aspect.

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

[0540] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; and wherein:

[0541] A fuel flow modified by the nvPM emission index ratio at idle can be defined as:

[0542] [Math.50] El SAF T47 EIraIenttFF X Wf,ralenti

[0543] where:

[0544] £7raiti,sAF is the 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 fuel spray nozzles includes a sustainable aviation fuel; and

[0545] £7raienti,FF 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 operating conditions in which Æ'A-aienti,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0546] Wfiaient! is the mass flow rate of fuel supplied to the spray nozzles of fuel in kg / s when the gas turbine engine is operating at approximately 7% of available thrust under the same operating conditions under which ^idle,SAF and Æ'Idle.FF are calculated;

[0547] the fuel flow modified by the emission index ratio of nvPM at idle of the gas turbine engine in kg / s is less than 0.2; and

[0548] The process includes supplying fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0549] The fuel flow modified by an emission index ratio of nvPM at idle and / or Wf at idle may be as defined above in relation to the twenty-sixth aspect.

[0550] According to a thirty-fifth aspect, a gas turbine engine for an aircraft is provided, comprising one or more of the following features:

[0551] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; wherein:

[0552] A depleted cruise nvPM emission index ratio can be defined as:

[0553] [Math.51] FI / crcasièr^appauvrie]. SAF PT croisier^appauvrie), FF

[0554] where:

[0555] ^ / cruise (impoverished),saf can be defined as:

[0556] [Math.52] ^nwTO, SaA~ El montée, SAF

[0557] E / cruise (impoverished e )^f can be defined as:

[0558] [Math.53] ElmàxTO, FA Elmontée, FF 2

[0559] £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 under given operating conditions if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel;

[0560] Æ' / 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 under given operating conditions if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel;

[0561] £7maXTo,FF is the 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 fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0562] Æ' / ascent.FF 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 if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0563] the nvPM emission index ratio in lean cruise of the gas turbine engine is less than 1; and

[0564] The gas turbine engine is designed to supply fuel including sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0565] The nvPM emission index ratio in depleted cruise can be greater than zero.

[0566] The nvPM emission index ratio in depleted cruise may be less than or equal to 0.9, preferably less than or equal to 0.8 and more preferably less than or equal to 0.75.

[0567] The nvPM emission index ratio in depleted cruise may be less than or equal to 0.709, preferably less than or equal to 0.65 and more preferably less than or equal to 0.591.

[0568] The nvPM emission index ratio in depleted cruise may be greater than or equal to 0.451, preferably greater than or equal to 0.507 and more preferably greater than or equal to 0.563.

[0569] The nvPM emission index ratio in depleted cruise may be in the range of 0.451 to 0.709, preferably in the range of 0.507 to 0.650 and more preferably in the range of 0.563 to 0.591.

[0570] The nvPM emission index ratio in depleted 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 within any defined range between any two of these values. For example, the second 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.

[0571] 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, or 0.71, or within any range defined between any two of these values.

[0572] A ratio of the nvPM emission index between idle and PMD can be defined as:

[0573] [Math.54] Elr^nSAF / El rnaxTOjSAF EIjAienti.Fr / EI maxTO.FF

[0574] where:

[0575] ^7raiti,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 under given operating conditions, or under other different operating conditions, if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel;

[0576] £7maxTo,sAF is the 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 in which £7raiti,sAF is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel;

[0577] ^ / slow,rr is the nvPM emissions index corrected for losses in the system in mg / kg of the gas turbine engine during operation at approximately 7% of available thrust under the same operating conditions in which ^- / idle,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0578] £7maxTo,FF is the 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 in which Æ'A-aienti,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0579] and in which the ratio of the nvPM emission index between idle and PMD of the gas turbine engine may be less than 1.

[0580] The nvPM emission index ratio between idle and PMD may be greater than zero.

[0581] The ratio of the nvPM emission index between idle and PMD may be less than or equal to 0.8 and preferably less than or equal to 0.6, even more preferably less than or equal to 0.4 and even more preferably less than or equal to 0.2.

[0582] The nvPM emission index ratio between idle and PMD may be less than or equal to 0.178 and preferably less than or equal to 0.164 and more preferably less than or equal to 0.149.

[0583] The nvPM emission index ratio between idle and PMD may be greater than or equal to 0.03 and preferably greater than or equal to 0.06 and more preferably greater than or equal to 0.09.

[0584] The nvPM emission index ratio between idle and PMD may be greater than or equal to 0.118 and preferably may be greater than or equal to 0.133 and more preferably may be greater than or equal to 0.148.

[0585] The nvPM emission index ratio between idle and PMD 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.

[0586] The nvPM emission index ratio between idle and 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 within any defined range between any two of these values. For example, the nvPM emission index ratio between idle and PMD may be within a range between 0.25 and 0.4 or 0.3 and 0.35.

[0587] The nvPM emission index ratio between idle and PMD 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.

[0588] A ratio of nvPM emissions index between depleted cruise and PMD can be defined as:

[0589] [Math.55] maxTO SAP El / nmsiéi iàppsurrKlFf / EJ maxT OP P

[0590] where:

[0591] E / cruise (impoverished e),saf can be defined as:

[0592] [Math.56] ^■huaxTO, SAf+ SIMontée, SAP 2

[0593] ^ / cruise (impoverished e )^f can be defined as:

[0594] [Math.57] EF ax F O. FF* E-Imontée. FF

[0595] £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 under given operating conditions, or under other different operating conditions, if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel;

[0596] £7montée,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 in which Ê / maxTo.sAF is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel;

[0597] £7maxTo,FF 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 in which £7maxT0>sAF is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0598] Ê / montéerF 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 in which £7maXTo,sAF is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0599] and in which the ratio of the nvPM emission index between the lean cruise and the PMD of the gas turbine engine may be less than 1.

[0600] The ratio of nvPM emission index between depleted cruise and PMD may be greater than zero.

[0601] The nvPM emission index ratio between depleted cruise and PMD may be less than or equal to 0.98, preferably less than or equal to 0.95 and more preferably less than or equal to 0.92. The nvPM emission index ratio between depleted cruise and PMD may be less than or equal to 0.914.

[0602] The ratio of the nvPM emission index between the depleted cruise and the PMD is greater than or equal to 0.698, preferably greater than or equal to 0.785 and more preferably greater than or equal to 0.873.

[0603] The ratio of the nvPM emission index between the depleted cruise and the PMD may be in the range of 0.698 to 0.980, preferably in the range of 0.785 to 0.950 and more preferably in the range of 0.873 to 0.914.

[0604] The nvPM emission index ratio between the depleted cruise and the 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.

[0605] The ratio of the nvPM emissions index between the depleted cruise and the PMD can be 0.69, 0.71, 0.73, 0.75, 0.77, 0.79, 0.81, 0.83, 0.85, 0.87, 0.89, 0.91, 0.93, 0.95, 0.97, or 0.99, or within any defined range between any two of these values. For example, the ratio of the nvPM emissions index between the depleted cruise and the PMD can be in a range between 0.91 and 0.99 or between 0.93 and 0.97.

[0606] A ratio of the nvPM emission index between idle and lean cruise can be defined as:

[0607] [Math.58] Eli ssai,. SAF / El crolsjèr^appau^ EJrpsan.FF / pI crais:are(appauvI-je\pp

[0608] where:

[0609] £7raiti,sAF is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at about 7% of available thrust under given operating conditions, or under other different operating conditions, if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel;

[0610] ^ / cruise (impoverished),saf can be defined as:

[0611] [Math.59] ElmaxTO, SAF+ ^^montée, SAP 2

[0612] £7maxTo,sAF is the 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 in which £7raiti,sAF is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel;

[0613] ^ / rise,saf is the nvPM emissions index corrected for losses in the system in mg / kg of the gas turbine engine when operating at about 85% of available thrust under the same operating conditions in which ^- / idle,saf is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel;

[0614] ^ / slow,ff is the nvPM emissions index corrected for losses in the system in mg / kg of the gas turbine engine when operating at about 7% of available thrust under the same operating conditions in which Æ'A-aienti,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0615] E / cm,,,^ (impoverished e ),ff can be defined as:

[0616] [Math.60] ElmàxTO, ascent, FF 2

[0617] £7maxTo,FF 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 in which ^ralid,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0618] £ / climb,FF 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 in which Ralido,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0619] and in which the ratio of the nvPM emission index between idle and lean cruise of the gas turbine engine may be less than 1.

[0620] The ratio of the nvPM emission index between idle and lean cruise can be greater than zero.

[0621] The ratio of the nvPM emission index between idle and 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 even more preferably less than or equal to 0.3.

[0622] The NvPM emission index ratio between idle and lean cruise may be less than or equal to 0.204, preferably less than or equal to 0.187 and more preferably less than or equal to 0.17.

[0623] The NvPM emission index ratio between idle and lean cruise may be greater than or equal to 0.129, preferably greater than or equal to 0.146 and more preferably greater than or equal to 0.162.

[0624] The NvPM emission index ratio between idle and lean cruise may be in the range of 0.129 to 0.204, preferably in the range of 0.146 to 0.187 and more preferably in the range of 0.162 to 0.170.

[0625] The NvPM emission index ratio between idle and 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 between idle and lean-fuel cruise may be in the range of 0.3 to 0.4 or 0.3 to 0.35.

[0626] The NVPM emission index ratio between idle and lean cruise can be 0.12, 0.125, 0.13, 0.135, 0.14, 0.145, 0.15, 0.155, 0.16, 0.165, 0.17, 0.175, 0.18, 0.185, 0.19, 0.195, 0.2, 0.205, or 0.21, or within any range defined between any two of these values.

[0627] A rich cruise nvPM emission index ratio can be defined as:

[0628] [Math.61] EIcroisièr^riche), SAP FT 1 cruise ship (luxury), FF

[0629] where:

[0630] ELroisière (rich), saf can be defined as:

[0631] [Math.62] EJ ascending, SAf^ EI approaching. SAF 2

[0632] £' / Cruise (rich),ff Pcut to be defined as:

[0633] [Math.63] EJ uphill, Ff+ EJ approach, FF

[0634] £ / climb,sAF is the nvPM emissions 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, if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel;

[0635] The approach,saf is the nvPM emissions index adjusted for losses in the system in mg / kg of the gas turbine engine when operating at approximately 30% of available thrust under the same operating conditions in which Ê / climb.sAF is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel;

[0636] ^ / rise,ff is the nvPM emissions index corrected for losses in the system in mg / kg of the gas turbine engine during operation at approximately 85% of available thrust under the same operating conditions in which ^ / climb,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0637] ^ / 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 under the same operating conditions in which ^ / climb,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0638] and in which the rich cruise nvPM emission index ratio of the gas turbine engine may be less than 1.

[0639] The nvPM emission index ratio in rich cruise can be greater than zero.

[0640] The rich cruise nvPM emission index ratio may be less than or equal to 0.9, preferably less than or equal to 0.8, still preferably less than or equal to 0.7.

[0641] The rich cruise nvPM emission index ratio may be less than or equal to 0.564, preferably less than or equal to 0.517 and more preferably less than or equal to 0.47.

[0642] The rich cruise nvPM emission index ratio may be greater than or equal to 0.3, preferably greater than or equal to 0.337 and more preferably greater than or equal to 0.375.

[0643] The rich cruise nvPM emission index ratio may be in the range of 0.300 to 0.564, preferably in the range of 0.337 to 0.517 and more preferably in the range of 0.375 to 0.470.

[0644] The rich cruise 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. For example, the rich cruise nvPM emission index ratio may be in the range of 0.45 to 0.7 or 0.5 to 0.65.

[0645] The rich cruise nvPM emission index ratio may be 0.3, 0.32, 0.34, 0.36, 0.38, 0.4, 0.42, 0.44, 0.46, 0.48, 0.5, 0.52, 0.54, 0.56, 0.58, or 0.6, or within any defined range between any two of these values.

[0646] A ratio of the nvPM emissions index between rich cruising and PMD can be defined as:

[0647] [Math.64] El croteièreÿich&FF / .EI nit3xTO FF

[0648] where:

[0649] ^ / cruise (rich), saf can be defined as:

[0650] [Math.65] Upgrading, SAP^approach, SAP 2

[0651] and Ê / cruise (rich)jf can be defined as:

[0652] [Math.66] The ascent, Fp+ Elappfoçhe. FF 2

[0653] and where:

[0654] £7 rise,sAF is the loss-corrected nvPM emission index 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, if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel;

[0655] £7apProche,sAF is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at about 30% of available thrust under the same operating conditions in which Ê / climb.sAF is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel;

[0656] £ / climb,FF is the 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 in which ^ / climb.sAF is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0657] £7approach,FF is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at about 30% of available thrust under the same operating conditions in which ^ / climb,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0658] £ / maxTo,sAF 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 in which Ê / climb.sAF is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel; and

[0659] £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 under the same operating conditions in which Ê / climb.sAF is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0660] and in which the ratio of the nvPM emission index between the rich cruise and the PMD of the gas turbine engine is less than 1.

[0661] The ratio of nvPM emission index between rich cruise and PMD may be greater than zero.

[0662] The nvPM emission index ratio between rich cruise and PMD may be less than or equal to 0.95, preferably less than or equal to 0.9 and even more preferably less than or equal to 0.875.

[0663] The nvPM emission index ratio between rich cruise and PMD may be less than or equal to 0.873, preferably less than or equal to 0.8 and more preferably less than or equal to 0.728.

[0664] The nvPM emission index ratio between rich cruise and PMD may be greater than or equal to 0.464, preferably greater than or equal to 0.523 and more preferably greater than or equal to 0.581.

[0665] The nvPM emission index ratio between rich cruise and PMD may be in the range of 0.464 to 0.873, preferably in the range of 0.523 to 0.800 and more preferably in the range of 0.581 to 0.728.

[0666] The NvPM emission index ratio between rich cruise and 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. For example, the NvPM emission index ratio between rich cruise and PMD may be in the range of 0.65 to 0.9 or 0.7 to 0.85 or 0.75 to 0.8.

[0667] The nvPM emission index ratio between rich cruise and PMD can be 0.45, 0.475, 0.5, 0.525, 0.55, 0.575, 0.6, 0.625, 0.65, 0.675, 0.7, 0.725, 0.75, 0.775, 0.8, 0.825, 0.85, or 0.875, or within any defined range between any two of these values.

[0668] A ratio of the nvPM emission index between idle and rich cruising can be defined as:

[0669] [Math.67] EIFont^AP / EI EIraIelMPP / EI Croisc^^

[0670] where:

[0671] £7raiti,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 under given operating conditions, or under other different operating conditions, and if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel (SAF);

[0672] £7raiti,FF is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at about 7% of available thrust under the same operating conditions in which £7raiti,sAF is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0673] ^ / cruise (rich), saf can be defined as:

[0674] [Math.68] SAf+ EIapproche. SAF 2

[0675] ^ / cruise (rich),ff can be defined as:

[0676] [Math.69] EJ iïiOatee,FF+ EJ approach, FF 2

[0677] ^ / climb,saf is the 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 in which ^- / idle,saf is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel;

[0678] £ / climb,FF 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 in which £7raiti,sAF is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0679] £7approach,sAF is the 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 in which £7raiti,sAF is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel; and

[0680] ^ / 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 under the same operating conditions in which Æ'A-aienti,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0681] and in which the ratio of the nvPM emission index between idle and rich cruise of the gas turbine engine may be less than 1.

[0682] The ratio of the nvPM emission index between idle and rich cruise can be greater than zero.

[0683] The NvPM emission index ratio between idle and rich cruise may be less than or equal to 0.8 and preferably less than or equal to 0.6 and more preferably less than or equal to 0.4.

[0684] The NvPM emission index ratio between idle and rich cruise may be less than or equal to 0.307, preferably less than or equal to 0.281 and more preferably less than or equal to 0.256.

[0685] The NvPM emission index ratio between idle and rich cruise may be greater than or equal to 0.163, preferably greater than or equal to 0.183 and more preferably greater than or equal to 0.203.

[0686] The NvPM emission index ratio between idle and rich cruise may be in the range of 0.163 to 0.307, preferably in the range of 0.183 to 0.281 and more preferably in the range of 0.203 to 0.256.

[0687] The NvPM emission index ratio between idle and rich cruising 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 between idle and rich cruising may be in the range of 0.3 to 0.5, or 0.35 to 0.45.

[0688] The ratio of the nvPM emission index between idle and rich cruising can be 0.16, 0.165, 0.17, 0.175, 0.18, 0.185, 0.19, 0.195, 0.2, 0.205, 0.21, 0.215, 0.22, 0.225, 0.23, 0.235, 0.24, 0.245, 0.25, 0.255, 0.26, 0.265, 0.27, 0.275, 0.28, 0.285, 0.29, 0.295, 0.3, 0.305, or 0.31, or within any defined range between any two of these values.

[0689] According to a thirty-sixth aspect, a gas turbine engine for an aircraft is provided, comprising one or more of the following features:

[0690] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; wherein:

[0691] A ratio of the nvPM emission index between idle and PMD can be defined as:

[0692] [Math.70] E^Imilsaf / EI jnaxTO SAp EI^mu-f / EI max TaFP

[0693] where:

[0694] £7raiti,sAF is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at about 7% of available thrust under given operating conditions if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel;

[0695] £7maxTo,sAF is the 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 in which £7raiti,sAF is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel;

[0696] ^ / slow,ff is the nvPM emissions index corrected for losses in the system in mg / kg of the gas turbine engine during operation at approximately 7% of available thrust under the same operating conditions in which ^- / idle,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0697] S / maxTorF 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 under the same operating conditions in which Æ'A-aienti,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0698] and the ratio of the nvPM emission index between idle and PMD of the gas turbine engine may be less than 1; and

[0699] The gas turbine engine is designed to supply fuel including sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0700] The ratio of the nvPM emission index between idle and PMD can be as defined above in relation to the thirty-fifth aspect.

[0701] According to a thirty-seventh aspect, a gas turbine engine for an aircraft is provided, comprising one or more of the following features:

[0702] a rich burning, rapid cooling, lean burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6;

[0703] in which:

[0704] A ratio of the nvPM emission index between depleted cruise and PMD can be defined as:

[0705] [Math.71] SAf / EI mâxTO SAP f / EÏ Ia[ixp Qp p

[0706] where:

[0707] ELroisière (impoverished),saf can be defined as:

[0708] [Math.72] F7mnxTO, SAfd" ^^inonÇée,, SAP 2

[0709] ^ / cruise (impoverished e )^f can be defined as:

[0710] [Math.73] ElmaxTO, Fp+ ^^niontée, FF 2

[0711] £7maxTo,sAF is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at about 100% of available thrust under given operating conditions if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel;

[0712] £7montée,sAF is the 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 in which ^ / maxTo.sAF is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel;

[0713] £7maXTo,FF is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at approximately 100% of the available thrust under the same operating conditions in which Ê / maxTo.sAF is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0714] Æ' / ascent.FF 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 in which £7maxT0>sAF is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0715] and the ratio of the nvPM emission index between the lean cruise and PMD of the gas turbine engine may be less than 1; and

[0716] The gas turbine engine is designed to supply fuel including sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0717] The ratio of nvPM emission index between depleted cruise and PMD can be as defined above in relation to the thirty-fifth aspect.

[0718] According to a thirty-eighth aspect, a gas turbine engine for an aircraft is provided, comprising one or more of the following features:

[0719] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; wherein:

[0720] A ratio of the nvPM emission index between idle and lean cruise can be defined as:

[0721] [Math.74] sat / EJ erofs^r^ppauyrfelsAF ElrsE,,ti °F / EI cruise^^

[0722] where:

[0723] ^7raiti,sAF is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at about 7% of available thrust under given operating conditions if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel;

[0724] EZcroisière (impoverished),saf can be defined as:

[0725] [Math.75] ^ImanTO, SAf+ mounted, SAF 2

[0726] Æ' / maxTo.sAF 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 in which E / raiti,saf is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel;

[0727] £'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 under the same operating conditions in which Ralide,saf is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel;

[0728] £7raiti,FF is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at about 7% of available thrust under the same operating conditions in which Ralido,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0729] ^ / cruise (impoverished e ),ff can be defined as:

[0730] [Math.76] ElmàxTO, FF+ El climb, FF 2

[0731] S / maxTorF 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 in which ^ralid,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0732] ^ / rise,ff is the nvPM emissions index corrected for losses in the system in mg / kg of the gas turbine engine when operating at about 85% of available thrust under the same operating conditions in which ^- / idle,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0733] ​​the ratio of the nvPM emission index between idle and lean cruise of the gas turbine engine is less than 1; and

[0734] The gas turbine engine is designed to supply fuel including sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0735] The ratio of the nvPM emission index between idle and lean cruise can be as defined above in relation to the thirty-fifth aspect.

[0736] According to a thirty-ninth aspect, a gas turbine engine for an aircraft is provided, comprising one or more of the following features:

[0737] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; wherein:

[0738] A rich cruise nvPM emission index ratio can be defined as:

[0739] [Math.77] DT rich cruise), S AF 1 rich cruise), FF

[0740] where:

[0741] E / cruise (rich), saf can be defined as:

[0742] [Math.78] Ei nlomée, SAF^~ Elapproach SAF 2 =

[0743] ^ / cruise (rich)^f can be defined as:

[0744] [Math.79] Ascent, FF^“ ^approached, FF

[0745] £'7montée,sAF is the nvPM emissions 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 if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel;

[0746] £7appropOche,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 under the same operating conditions in which EImontée,sAF is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel;

[0747] ^ / rise,ff is the nvPM emissions index corrected for losses in the system in mg / kg of the gas turbine engine during operation at approximately 85% of available thrust under the same operating conditions in which ^ / climb,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0748] ^ / 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 under the same operating conditions in which Ê / climb.sAF is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0749] the rich cruise nvPM emission index ratio of the gas turbine engine is less than 1; and

[0750] The gas turbine engine is designed to supply fuel including sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0751] The rich cruise nvPM emission index ratio can be as defined above in relation to the thirty-fifth aspect.

[0752] According to a fortieth aspect, a gas turbine engine for an aircraft is provided, comprising one or more of the following features:

[0753] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; wherein:

[0754] A ratio of the nvPM emission index between rich cruising and PMD can be defined as:

[0755] [Math.80] maxT O.SAF croinièr^flch^hp / .Fl njpxTO FF

[0756] where:

[0757] EZcroisière (rich), saf can be defined as:

[0758] [Math.81] rise, SAfE^approach, SAP 2

[0759] and Ê / cruise (rich)^f can be defined as:

[0760] [Math. 82] Elniontée, Fp+ approach. FF 2

[0761] and where:

[0762] £7montée,sAF is the nvPM emissions 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 and if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel;

[0763] £7aPProche,sAF is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at about 30% of available thrust under the same operating conditions in which Ê / climb.sAF is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel;

[0764] ^ / rise,ff is the nvPM emissions index corrected for losses in the system in mg / kg of the gas turbine engine when operating at about 85% of available thrust under the same operating conditions in which ^ / climb,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0765] ^ / approach,ff is the nvPM emissions index corrected for losses in the system in mg / kg of the gas turbine engine during operation at approximately 30% of available thrust under the same operating conditions in which ^ascent,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0766] £' / maxTo,sAF 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 in which ÊZmontée.sAF is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel; and

[0767] £7maxTo,FF is the 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 in which £7montée,sAF is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0768] the ratio of the nvPM emission index between the rich cruise and PMD of the gas turbine engine is less than 1; and

[0769] The gas turbine engine is designed to supply fuel including sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0770] The nvPM emission index ratio between rich cruise and PMD can be as defined above in relation to the thirty-fifth aspect.

[0771] According to a forty-first aspect, a gas turbine engine for an aircraft is provided, comprising one or more of the following features:

[0772] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; wherein:

[0773] A ratio of the nvPM emission index between idle and rich cruising can be defined as:

[0774] [Math. 83] ELalenti.Fp / EI crojsièr^rj^

[0775] where:

[0776] ^ / 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 under given operating conditions and if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel (SAF);

[0777] ^ / slow,ff is the nvPM emissions index corrected for losses in the system in mg / kg of the gas turbine engine during operation at approximately 7% of available thrust under the same operating conditions in which E / raiti,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0778] E / cruise (rich), saf can be defined as:

[0779] [Math. 84] The rise, SA approach, SAP 2

[0780] ^' / cruise (rich)jf can be defined as:

[0781] [Math.85] The climb, FT^~ EI approaches. FF 2

[0782] 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 under the same operating conditions in which £7raiti,sAF is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel;

[0783] £' / climb,FF 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 in which ^idle,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0784] ^ / approach,saf is the nvPM emissions index adjusted for losses in the system in mg / kg of the gas turbine engine during operation at approximately 30% of available thrust under the same operating conditions in which Æ'A-aienti,saf is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel; and

[0785] ^ / 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 under the same operating conditions in which ^- / idle,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0786] the ratio of the nvPM emission index between idle and rich cruise of the gas turbine engine is less than 1; and

[0787] The gas turbine engine is designed to supply fuel including sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0788] The ratio of the nvPM emission index between idle and rich cruising can be as defined above in relation to the thirty-fifth aspect.

[0789] 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 process including the supply of fuel comprising a durable aviation fuel to the fuel spray nozzles.

[0790] According to a forty-third aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0791] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; wherein:

[0792] A depleted cruise nvPM emission index ratio can be defined as:

[0793] [Math. 86] FT, , cruise^impoverished, SAF -°1 cruise (impoverished^ FF

[0794] where:

[0795] E / cruise (impoverished e),saf can be defined as:

[0796] [Math. 87] ^maxTO, SÂF+ Elmontée, SAF 2

[0797] (impoverished e )^f can be defined as:

[0798] [Math.88] ElmaxTO, :0 + El montée,FF

[0799] £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 under given operating conditions and if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel;

[0800] £7 rise,sAF is the nvPM emissions 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 and if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel;

[0801] £7maxTo,FF is the 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 fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0802] Æ' / ascent.FF 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 if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0803] the PMD nvPM emission index ratio in depleted cruise is greater than 1; and

[0804] The process includes supplying fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0805] The nvPM emission index ratio in depleted cruise can be as defined above in relation to the thirty-fifth aspect.

[0806] A ratio of the nvPM emission index between idle and PMD can be defined as:

[0807] [Math. 89] EJrslenll Saf / E1 maxpOiS^p EI^Pnti.FF / EI maxTO.FF

[0808] where:

[0809] £7raiti,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 under given operating conditions, or under other different operating conditions, and if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel;

[0810] £'7maxTo,sAF is the 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 in which Æ'A-aienti,saf is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel;

[0811] ^ / slow,ff is the nvPM emissions index corrected for losses in the system in mg / kg of the gas turbine engine during operation at approximately 7% of available thrust under the same operating conditions in which Æ'A-aienti,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0812] S / maxTorF 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 in which ^- / idle,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and in which

[0813] the ratio of the nvPM emission index between idle and PMD of the gas turbine engine is less than 1.

[0814] The ratio of the nvPM emission index between idle and PMD can be as defined above in relation to the thirty-fifth aspect.

[0815] A ratio of the nvPM emissions index between depleted cruise and PMD can be defined as:

[0816] [Math.90] ms;» rfappau™ saf / EI jjJax^Q,SAF 1 IcroisièrHappaavrtelFl' / E! yq pp

[0817] where:

[0818] E / cruise (impoverished e),saf can be defined as:

[0819] [Math.91] ^^maxTO. SAp+ EEnontée, SAF

[0820] ^ / cruise (impoverished e )^f can be defined as:

[0821] [Math.92] EIiuàxTO, Ff+ ascent, FF

[0822] £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 fuel spray nozzles includes a sustainable aviation fuel;

[0823] 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 under the same operating conditions in which £7maxTo,sAF is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel;

[0824] £7maxTo,FF 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 in which £7maxTo,sAF is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0825] EImontée,pp is the 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 in which Ê / maxTo.sAF is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and in which

[0826] the nvPM emission index ratio between lean cruise and PMD of the gas turbine engine is less than 1.

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

[0828] A ratio of the nvPM emission index between idle and lean cruise can be defined as:

[0829] [Math.93] SAF / EI orolsière(appauvrie\SA EI'«Fnt,,FF / EI l;roisière(appa Uvrie],FF

[0830] where:

[0831] £7raiti,sAF is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at about 7% of available thrust under given operating conditions, or under other different operating conditions, and if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel;

[0832] EZcroisière (impoverished),saf can be defined as:

[0833] [Math.94] El maxTO. montée, SAF 2

[0834] £7maxTo,sAF is the 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 in which ^- / idle,saf is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel;

[0835] ^ / climb,saf is the 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 in which ^- / idle,saf is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel;

[0836] ^ / slow,ff is the nvPM emissions index corrected for losses in the system in mg / kg of the gas turbine engine when operating at about 7% of available thrust under the same operating conditions in which ^ / idle,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0837] E / cruise (impoverished e),ff can be defined as:

[0838] [Math.95] ElmaxTO. Ff+ El montée. FF

[0839] £ / maxTo,FF 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 in which Ralide,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0840] £ / climb,FF is the 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 in which ^ / idle,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and in which

[0841] The NVPM emission index ratio between idle and lean cruise of the gas turbine engine may be less than 1.

[0842] The ratio of the nvPM emission index between idle and lean cruise can be as defined above in relation to the thirty-fifth aspect.

[0843] A rich cruise nvPM emission index ratio can be defined as:

[0844] [Math.96] rr Groisièr^Eiche), SAF 1 cruise (cheating), FF

[0845] where:

[0846] ^ / cruise (rich), saf can be defined as:

[0847] [Math.97] El climb, SAF^” E approach, SAP

[0848] E / cruise (rich),ff can be defined as:

[0849] [Math.98] Eljpontée. FF-*- EIapproach, FF 2

[0850] ^ / rise,saf is the nvPM emissions index corrected for losses in the system 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 fuel spray nozzles includes a sustainable aviation fuel;

[0851] ^ / approach,saf is the nvPM emissions index adjusted for losses in the system in mg / kg of the gas turbine engine when operating at approximately 30% of available thrust under the same operating conditions in which Ê / climb.sAF is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel;

[0852] ^ / rise,ff is the nvPM emissions index corrected for losses in the system in mg / kg of the gas turbine engine during operation at approximately 85% of the thrust available under the same operating conditions in which Ê / climb.sAF is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0853] Elapproche,ff is the nvPM emissions index corrected for losses in the system in mg / kg of the gas turbine engine during operation at approximately 30% of available thrust under the same operating conditions in which E / climb,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0854] The rich cruise nvPM emission index ratio of the gas turbine engine may be less than 1.

[0855] The rich cruise nvPM emission index ratio can be as defined above in relation to the thirty-fifth aspect.

[0856] A ratio of the nvPM emissions index between rich cruising and PMD can be defined as:

[0857] [Math.99] ÿjrjxTO SAF El ■cnx^^cb^'r / EI

[0858] where:

[0859] E / cruise (rich), saf can be defined as:

[0860] [Math. 100] The climb, SAF^ The approach, SAP 2

[0861] and Ê / cruise (rich),ff can be defined as:

[0862] [Math. 101] The climb, FpE The approach, FF 2

[0863] and where:

[0864] E Zmontée,saf is the nvPM emissions index corrected for losses in the system 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 fuel spray nozzles includes a sustainable aviation fuel;

[0865] E ZapprOche,saf is the nvPM emissions index adjusted for losses in the system in mg / kg of the gas turbine engine when operating at about 30% of available thrust under the same operating conditions in which £7mOntée,sAF is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel;

[0866] E / rise,ff is the nvPM emission index corrected for losses in the system in mg / kg of the gas turbine engine during operation at approximately 85% of the thrust available under the same operating conditions in which Ê / climb.sAF is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0867] E Zapproche,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 under the same operating conditions in which Ê / climb.sAF is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0868] E1 maxTo.sAF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine

[0869] during operation at approximately 100% of available thrust under the same operating conditions under which EI™^^ is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel; and

[0870] E ZmaxTo,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 under the same operating conditions in which ^ / climb,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0871] and wherein the ratio of the nvPM emission index between the rich cruise and the PMD of the gas turbine engine is less than 1.

[0872] The ratio of nvPM emission index between rich cruise and PMD can be as defined above in relation to the thirty-fifth aspect.

[0873] A ratio of the nvPM emission index between idle and rich cruising can be defined as:

[0874] [Math. 102] crmsler^ Ei miaiti.F^ / EI crojsjère(rich)EF

[0875] where:

[0876] E 4aienti,saf is the nvPM emissions index adjusted for losses in the system in mg / kg of the gas turbine engine when operating at about 7% of available thrust under given operating conditions, or under other different operating conditions, and if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel (SAF);

[0877] E / ralenti,ff is the nvPM emission index corrected for losses in the system in mg / kg of the gas turbine engine when operating at about 7% of available thrust under the same operating conditions in which / ^ / idle,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0878] E (rich), saf can be defined as:

[0879] [Math. 103] The rise, SAfE approach, SAP 2

[0880] E / cruise (rich)^f can be defined as:

[0881] [Math. 104] The rise,PP^ EIapprixjle, PP 2

[0882] 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 under the same operating conditions in which Ralide,saf is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel;

[0883] E ZmOntée,ff is the nvPM emissions index corrected for losses in the system in mg / kg of the gas turbine engine when operating at about 85% of available thrust under the same operating conditions in which £7raiti,sAF is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0884] E Zapproche,saf is the nvPM emissions index adjusted for losses in the system in mg / kg of the gas turbine engine during operation at approximately 30% of available thrust under the same operating conditions in which / ^ / idle,saf is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel; and

[0885] E / 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 under the same operating conditions in which / ^ / idle,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0886] and in which the ratio of the nvPM emission index between idle and rich cruise of the gas turbine engine may be less than 1.

[0887] The ratio of the nvPM emission index between idle and rich cruising can be as defined above in relation to the thirty-fifth aspect.

[0888] According to a forty-fourth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0889] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; wherein:

[0890] A ratio of the nvPM emission index between idle and PMD can be defined as:

[0891] [Math. 105] EIralé„il SAF / EI inax TO &^ Eh^uFF / EI Uax TO pp

[0892] where:

[0893] E / raiti,saf is the nvPM emissions index adjusted for losses in the system 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 fuel spray nozzles includes a sustainable aviation fuel;

[0894] E / maXTo,sAF is the 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 in which ^ / idle,saf is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel;

[0895] E / raienti,F? is the nvPM emission index corrected for losses in the system in mg / kg of the gas turbine engine during operation at approximately 7% of available thrust under the same operating conditions in which ^ / ■aienti,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0896] E / maXTo,FF is the 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 in which ^ / idle,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0897] and the ratio of the nvPM emission index between idle and PMD of the gas turbine engine may be less than 1; and

[0898] The process includes supplying fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0899] The ratio of the nvPM emission index between idle and PMD can be as defined above in relation to the thirty-fifth aspect.

[0900] According to a forty-fifth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0901] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; wherein:

[0902] A ratio of the nvPM emission index between depleted cruise and PMD can be defined as:

[0903] [Math. 106] maxTO SAP El / nmsiéi iàppsurrKlFf / EJ maxT OP P

[0904] where:

[0905] E / cruise (impoverished e),saf can be defined as:

[0906] [Math. 107] EEnaxTO, SAP^ SIMontée, SAP 2

[0907] E4r™„ere (impoverished e )^f can be defined as:

[0908] [Math. 108] EF ax F O. FF* E-Imontée. FF

[0909] £7maxTo,sAF is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at about 100% of available thrust under given operating conditions if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel;

[0910] E Zmontée,saf is the nvPM emissions index corrected for losses in the system in mg / kg of the gas turbine engine when operating at about 85% of available thrust under the same operating conditions in which ^ / maxTo.sAF is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel;

[0911] E ZmaxTorr is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine operating at approximately 100% of available thrust under the same operating conditions in which ^ / maxTo.sAF is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0912] E Zmontée,ff is the nvPM emissions index corrected for losses in the system in mg / kg of the gas turbine engine when operating at about 85% of available thrust under the same operating conditions in which £7maxTo,sAF is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0913] and the ratio of the nvPM emission index between the lean cruise and PMD of the gas turbine engine may be less than 1; and

[0914] The process includes supplying fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0915] The ratio of nvPM emission index between depleted cruise and PMD can be as defined above in relation to the thirty-fifth aspect.

[0916] According to a forty-sixth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0917] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; wherein:

[0918] A ratio of the nvPM emission index between idle and lean cruise can be defined as:

[0919] [Math. 109] Eli ssai,. SAF / El crolsjèr^appau^ I fear:are(I-I-poor\pp

[0920] where:

[0921] £7raiti,sAF is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine when operating at about 7% of available thrust under given operating conditions if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel;

[0922] EZcruise (impoverished), saf can be defined as:

[0923] [Math. 110] ElmaxTO, SAF+ ^^montée, SAP 2

[0924] £7maxTo,sAF is the 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 in which £7raiti,sAF is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel;

[0925] E Zmontée,saf is the nvPM emissions index corrected for losses in the system in mg / kg of the gas turbine engine when operating at about 85% of available thrust under the same operating conditions in which ^ / idle,saf is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel;

[0926] E 4aienti,F? is the nvPM emission index corrected for losses in the system in mg / kg of the gas turbine engine when operating at about 7% of available thrust under the same operating conditions in which E / raiti,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0927] E ZCTOisière (impoverished) ,ff can be defined as:

[0928] [Math. 111] EïmaxTO, El montée, FF 2

[0929] S / maxTOFF 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 in which Ralide,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0930] E Zmontée,ff 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 in which ^ / ralido,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0931] the ratio of the nvPM emission index between idle and lean cruise of the gas turbine engine is less than 1; and

[0932] The process includes supplying fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0933] The ratio of the nvPM emission index between idle and lean cruise can be as defined above in relation to the thirty-fifth aspect.

[0934] According to a forty-seventh aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0935] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; wherein:

[0936] A rich cruise nvPM emission index ratio can be defined as:

[0937] [Math. 112] c 1 cruise^richeX SAF rrrr^rrrrrrrrrrrrrrrrrrrrrrrrrrrrrrrrrrrrrrrr 1 cruise!era^ichel FF

[0938] where:

[0939] E / cruise (rich), saf can be defined as:

[0940] [Math. 113] The climb. SA The approach, SAF 2

[0941] E / cruise (rich),ff can be defined as:

[0942] [Math. 114] The climb, FA" Elappfoché, FF

[0943] £ / 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 under given operating conditions if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel;

[0944] E Zapproche,saf is the nvPM emissions index adjusted for losses in the system in mg / kg of the gas turbine engine when operating at approximately 30% of available thrust under the same operating conditions in which Ê / climb.sAF is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel;

[0945] E ZmOntée,ff is the nvPM emissions index corrected for losses in the system in mg / kg of the gas turbine engine during operation at approximately 85% of available thrust under the same operating conditions in which Ê / climb.sAF is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and

[0946] E / 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 under the same operating conditions in which ^ / climb,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0947] the rich cruise nvPM emission index ratio of the gas turbine engine is less than 1; and

[0948] The process includes supplying fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0949] The rich cruise nvPM emission index ratio can be as defined above in relation to the thirty-fifth aspect.

[0950] According to a forty-eighth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0951] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; wherein:

[0952] A ratio of the nvPM emissions index between rich cruising and PMD can be defined as:

[0953] [Math. 115] EIcrolslèrt(rkM,SAp / EI maxTO:SAp maxTO pp

[0954] where:

[0955] ^ / cruise (rich), saf can be defined as:

[0956] [Math. 116] The climb, SAF^^approach, SAP 2

[0957] and E / cruise (rich)jf Pcut can be defined as:

[0958] [Math. 117] Eïmonifif^pp~^ EIapprrx]lef pp 2

[0959] and where:

[0960] E / mOntée,saf is the nvPM emissions index corrected for losses in the system 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 fuel spray nozzles includes a sustainable aviation fuel;

[0961] E / approach,saf is the nvPM emissions index adjusted for losses in the system in mg / kg of the gas turbine engine when operating at about 30% of available thrust under the same operating conditions in which ^ / climb,saf is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel;

[0962] E / rise,ff is the nvPM emission index corrected for losses in the system in mg / kg of the gas turbine engine when operating at about 85% of available thrust under the same operating conditions in which Ê / climb.sAF is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0963] E / 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 under the same operating conditions in which Ê / climb.sAF is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0964] E ZmaxTo.sAF 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 in which ^ / climb,saf is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel; and

[0965] E / 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 in which ^ascent,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0966] the ratio of the nvPM emission index between the rich cruise and PMD of the gas turbine engine is less than 1; and

[0967] The process includes supplying fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0968] The ratio of nvPM emission index between rich cruise and PMD can be as defined above in relation to the thirty-fifth aspect.

[0969] According to a forty-ninth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising:

[0970] a rich-burning, rapid-cooling, lean-burning (RQL) combustor device having a number of fuel spray nozzles in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; wherein:

[0971] A ratio of the nvPM emission index between idle and rich cruising can be defined as:

[0972] [Math. 118] £I^OFr / 'PI crojsîére^icheï.FF

[0973] where:

[0974] E 4aienti,saf is the nvPM emissions index adjusted for losses in the system 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 fuel spray nozzles includes a sustainable aviation fuel (SAF);

[0975] E Laitsi,ff is the nvPM emissions index corrected for losses in the system in mg / kg of the gas turbine engine when operating at about 7% of available thrust under the same operating conditions in which / ^ / idle,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0976] E / cruise (rich), saf can be defined as:

[0977] [Math. 119] Ascent, SaF E 1 approach, SAF

[0978] E / cruise (rich),ff can be defined as:

[0979] [Math. 120] El + El , -mcniee.FF' 1 approach. FF

[0980] £ / climb,sAF is the nvPM emissions 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 in which idle,saf is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel;

[0981] E Zmontée,ff is the nvPM emissions index corrected for losses in the system in mg / kg of the gas turbine engine when operating at about 85% of available thrust under the same operating conditions in which £4aienti,sAF is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0982] E / approach,saf is the nvPM emissions index adjusted for losses in the system in mg / kg of the gas turbine engine during operation at approximately 30% of available thrust under the same operating conditions in which / ^ / idle,saf is calculated if a fuel supplied to the fuel spray nozzles is a sustainable aviation fuel; and

[0983] E / 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 under the same operating conditions in which / ^ / idle,saf is calculated if a fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel;

[0984] the ratio of the nvPM emission index between idle and rich cruise of the gas turbine engine is less than 1; and

[0985] The process includes supplying fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0986] The ratio of the nvPM emission index between idle and rich cruising can be as defined above in relation to the thirty-fifth aspect.

[0987] In any one of the aspects defined above, one or more of the emission indices of the gas turbine engine may be defined as follows:

[0988] S / maxTo can be in the range of 4.96 to 145 mg / kg and preferably in the range of 5.58 to 133 mg / kg and more preferably in the range of 6.21 to 121 mg / kg.

[0989] S / maxTo can be in the range of 4.96 to 93.3 mg / kg and preferably in the range of 5.58 to 85.5 mg / kg and more preferably in the range of 6.21 to 77.8 mg / kg.

[0990] S / maxTo can be in the range of 4.96 to 144 mg / kg and preferably in the range of 5.58 to 132 mg / kg and more preferably in the range of 6.21 to 120 mg / kg.

[0991] S / maxTo can be in the range of 7.69 to 145 mg / kg and preferably in the range of 8.65 to 133 mg / kg and more preferably in the range of 9.61 to 121 mg / kg.

[0992] The rise can be in the range of 1.82 to 124 mg / kg and preferably in the range of 2.05 to 114 mg / kg and more preferably in the range of 2.28 to 103 mg / kg.

[0993] ^ / mount,saf can be in the range of 1.82 to 58.6 mg / kg and preferably in the range from 2.05 to 53.7 mg / kg and more preferably in the range of 2.28 to 48.9 mg / kg.

[0994] ^ / mount,saf can be in the range of 1.82 to 123 mg / kg and preferably in the range from 2.05 to 113 mg / kg and more preferably in the range of 2.28 to 102 mg / kg.

[0995] EImontée,FF can be in the range of 3.84 to 124 mg / kg and preferably in the range of 4.32 to 114 mg / kg and more preferably in the range of 4.80 to 103 mg / kg.

[0996] ^ / approach may be in the range of 0.0328 to 17.6 mg / kg and preferably in the range of 0.0369 to 16.2 mg / kg and more preferably in the range of 0.0410 to 14.7 mg / kg.

[0997] ^ / approach,saf can be in the range of 0.0328 to 2.70 mg / kg and preferably in the range of 0.0369 to 2.48 mg / kg and more preferably in the range of 0.0410 to 2.25 mg / kg.

[0998] £7apProche,sAF may be in the range of 0.0328 to 17.5 mg / kg and preferably in the range of 0.0369 to 16.1 mg / kg and more preferably in the range of 0.0410 to 14.6 mg / kg.

[0999] ^ / approach,ff can be in the range of 0.213 to 17.6 mg / kg and preferably in the range of 0.240 to 16.2 mg / kg and preferably in the range of 0.267 to 14.7 mg / kg-

[1000] Ê / raienti may be in the range of 0.118 to 41.4 mg / kg and preferably in the range of 0.132 to 38.0 mg / kg and more preferably in the range of 0.147 to 34.5 mg / kg.

[1001] Æ'Zraiti.sAF may be in the range of 0.118 to 3.97 mg / kg and preferably in the range of 0.132 to 3.64 mg / kg and more preferably in the range of 0.147 to 3.31 mg / kg.

[1002] ^ / slow,saf can be in the range of 0.118 to 41.3 mg / kg and preferably in the range from 0.132 to 37.9 mg / kg and more preferably in the range of 0.147 to 34.4 mg / kg.

[1003] ^ / slow,ff can be in the range of 1.23 to 41.4 mg / kg and preferably in the range from 1.38 to 38.0 mg / kg and more preferably in the range of 1.54 to 34.5 mg / kg.

[1004] The following statements may apply to any one of the first to forty-ninth aspects defined above:

[1005] The fuel spray nozzles may include one or more duplex nozzles and one or more single-flow nozzles. The combustor device may include 10 to 14 duplex fuel spray nozzles. The combustor device may include 4 to 8 single-flow fuel spray nozzles.

[1006] The duplex fuel spray nozzles can be arranged in groups around the circumference of the combustor device. The groups of duplex fuel spray nozzles can comprise at least two groups arranged diametrically opposed to each other. Each group of duplex fuel spray nozzles can comprise 2 to 8 nozzles.

[1007] The combustor device may include one or more igniters. The igniter, or each igniter, may be arranged adjacent to one or more of the duplex fuel spray nozzles.

[1008] The number of fuel spray nozzles per unit size of engine core can be in the range of 2.5 to 4.5. The number of fuel spray nozzles per unit size of engine core can be in the range of 3 to 4.

[1009] 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%.

[1010] 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.

[1011] 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 in 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.

[1012] A gas turbine engine conforming to any aspect of this description may include an engine core comprising a turbine, a combustor device, 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).

[1013] 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, by means of 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 may be a first turbine, the core shaft may be a first core shaft, and the gas turbine engine may further include a second turbine and a second core shaft connecting the second turbine to the compressor. The second turbine, compressor, and second core shaft may be arranged to rotate at a higher speed than the first core shaft. In such an arrangement, the second turbine may be positioned axially upstream of the first turbine. The compressor may be a first compressor, and the gas turbine engine may further include a second compressor. The gas turbine engine may further include a third turbine and a third core shaft connecting the third turbine to the second compressor.The third turbine, the second compressor, and the third core shaft can be arranged to rotate at a higher speed than the second core shaft. In such an arrangement, the third turbine can be positioned axially upstream of the second turbine.

[1014] 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 compressor, such that the turbine and compressor rotate at the same speed (with the blower rotating at a lower speed).

[1015] 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.

[1016] In such an arrangement, the second compressor can be positioned axially downstream of the first compressor. The second compressor can be arranged to receive (for example receive directly, for example via a generally annular conduit) a flow from the first compressor.

[1017] The gearbox can be arranged to be driven by the core shaft that is designed to rotate (e.g., during operation) at the lowest rotational speed (e.g., the first core shaft in the example above). Alternatively, the gearbox can be arranged to be driven only by the core shaft that is designed to rotate (e.g., during operation) at the lowest rotational speed (e.g., only by the first core shaft, and not the second core shaft, in the example above). Alternatively, the gearbox can be arranged to be driven by any shaft or shafts, for example, the first and / or second shafts in the example above.

[1018] 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.

[1019] 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 in an inclusive range bounded by any two of the values ​​from the preceding sentence (i.e. the values ​​can form upper or lower bounds), for example in 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.

[1020] In any gas turbine engine as described and / or claimed herein, fuel of a given composition or mixture is supplied to a combustor device, which may be supplied downstream of the blower and compressor(s) with respect to the flow path (e.g., axially downstream). For example, the combustor device may 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 of the combustor device may be supplied at the inlet of the second turbine, when a second turbine is supplied. The combustor device may be supplied upstream of the turbine(s).

[1021] 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 11, 12, 13, 14, or 15 compressor stages (in addition to the blower). Such an engine may, for example, comprise 3 stages in the first compressor (or "low pressure") and either 8, 9, 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 either 8, 9, 10, or 11 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 with 4 stages in the first stage. compressor (or "low pressure") and 10 stages in the second compressor (or "high pressure").

[1022] 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, or vice versa, as required. 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, the gas turbine engine may be a "geared" gas turbine engine comprising 5, 6, or 7 turbine stages. Such a geared gas turbine engine may include a second ("high-pressure") turbine having 2 stages.Such a geared gas turbine engine may include a first turbine (or "low-pressure" turbine) having 3 or 4 stages. As a further example, the gas turbine engine may be a direct-drive gas turbine engine comprising a first turbine (or "low-pressure" turbine) having 3, 4, 5, 6, or 7 stages. Such a direct-drive gas turbine engine may include a second turbine (or "high-pressure" turbine) having 2 stages. Alternatively, such an engine may, for example, include a second turbine (or "intermediate-pressure" turbine) having 1, 2, or 3 stages. Such a direct-drive gas turbine engine may also include a third turbine (or "high-pressure" turbine) having 1, 2, or 3 stages. As a further example, the first turbine (or "low-pressure" turbine) may include 6 stages and the third turbine (or "high-pressure" turbine) may include 1 stage.

[1023] 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 furthest forward) 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.

[1024] 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), 280 cm (approximately 110 inches). 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 in 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 in 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 in the range of 170 cm to 180 cm, 190 cm to 200 cm, 200 cm to 210 cm, 210 cm to 220 cm, 210 cm to 230 cm, 220 cm to 230 cm, 260 cm to 270 cm, 280 cm to 290 cm, 290 cm to 300 cm, or 340 cm to 360 cm.

[1025] 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: 3,500 rpm, 3,450 rpm, 3,400 rpm, 3,350 rpm, 3,300 rpm, 3,250 rpm, 3,200 rpm, 3,150 rpm, 3,100 rpm, 3,050 rpm, 3,000 rpm, 2,950 rpm, 2,900 rpm, 2,850 rpm, 2,800 rpm, 2,750 rpm, 2,700 rpm, 2,650 rpm, 2,600 rpm, 2,550 rpm, 2 500 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, 1 350 rpm, 1300 rpm, 1250 rpm, 1200 rpm, 1150 rpm, 1100 rpm, 1050 rpm, 1000 rpm, or 950 rpm. The rotation speed. The fan speed 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). 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 2,750 to 2,900 rpm, 2,750 to 2,800 rpm, or 2,800 to 2,900 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 2,500 rpm to 2,800 rpm, or 2,500 rpm to 2,750 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 1,500 rpm to 1,800 rpm, or 1,500 rpm to 1,700 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 1,500 rpm, 950 rpm to 1,200 rpm, 950 rpm to 1,100 rpm, 950 rpm to 1,050 rpm, 950 rpm to 1,000 rpm, or 1,000 rpm to 1,050 rpm.Strictly by way of further non-limiting example, the fan speed under cruising conditions for a direct-drive engine with a fan diameter in the range of 170 cm to 200 cm can be in the range of 3,400 to 4,600 rpm, for example, 3,600 to 4,600 rpm, or 3,600 to 3,900 rpm. Strictly by way of further non-limiting example, the fan speed under cruising conditions for a direct-drive engine with a fan diameter in the range of 300 cm to 350 cm can be in the range of 1,800 to 3,200 rpm, or 1,950 to 2,900 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 cm to 290 cm can be in the range of 1,800 to 2,900 rpm, or 2,050 to 2,700 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 290 cm to 310 cm can be in the range of 1,800 to 2,950 rpm, or 2,100 to 2,700 rpm.

[1026] 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 average 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 blower load under cruise conditions may be greater than, or may be any 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 blower peak load 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 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).

[1027] Gas turbine engines in accordance with this description may have any desired bypass ratio (BPR), where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the 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 according to this disclosure may be within the range of 8 to 1, 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 . from 12.5 to 15. The bypass duct may be at least substantially annular. The bypass duct may be radially external to the reactor block. The radially external surface of the bypass duct may be defined by a nacelle and / or a fan housing.

[1028] 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 by way of non-limiting example, 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 non-limiting example, 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 non-limiting example, 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 under 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 under 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 under 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 as a non-limiting example, the overall pressure ratio 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 35 to 60, or 40 to 50. Strictly as an example. By way of non-limiting example, the overall pressure ratio under cruise conditions for 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 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 290 cm to 310 cm can be in the range of 35 to 60, or 37 to 47.

[1029] 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⁻¹s, 145 Nkg⁻¹s, 140 Nkg⁻¹s, 135 Nkg⁻¹s, 130 Nkg⁻¹s, 125 Nkg⁻¹s, 120 Nkg⁻¹s, 115 Nkg⁻¹s, 110 Nkg⁻¹ls, 105 Nkg⁻¹ls, 100 Nkg⁻¹ls, 99 Nkg⁻¹ls, 98 Nkg⁻¹ls, 97 Nkg⁻¹ls, 96 Nkg⁻¹ls, 95 Nkg⁻¹ls, 94 Nkg⁻¹ls, 93 Nkg⁻¹ls, 92 Nkg⁻¹ls, 91 Nkg⁻¹ls, 90 Nkg⁻¹ls, 89 Nkg⁻¹ls, 88 Nkg⁻¹ls, 87 Nkg⁻¹ls, 86 Nkg-ls, 85 Nkg-ls, 80 Nkg-ls, 75 Nkg's, 70 Nkg's, 65 Nkg's, 60 Nkg's, 55 Nkg's, 50 Nkg's, 45 Nkg's, 40 Nkg's, 35 Nkg's, 30 Nkg's, 25 Nkg's, 20 Nkg 's or 15 Nkg 's.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 bounds), for example, in the range of 80 Nkg-ls to 100 Nkg-ls, 85 Nkg-ls to 100 Nkg-ls, or 92 Nkg-ls to 100 Nkg-ls. Such engines can be particularly efficient compared to conventional gas turbine engines. Strictly by way of non-limiting example, the specific thrust 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 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 cm² to 230 cm² can be in the range of 90 Nkg-ls to 100 Nkg-ls, or 95 Nkg-ls to 100 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 340 cm² to 360 cm² can be in the range of... from 70 Nkg-ls to 95 Nkg-ls, from 80 Nkg-ls to 95 Nkg-ls, or from 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 having 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, from 20 Nkg-ls to 80 Nkg-ls, or from 25 Nkg-ls to 70 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 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 with 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 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 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.

[1030] 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 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). Strictly by way of non-limiting example, a gas turbine as described and / or claimed herein may be capable of producing 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. kN, or from 350 kN to 400 kN. Strictly by way of 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 from 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 by way of non-limiting example, the maximum thrust 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 100 kN to 200 kN, 110 kN to 180 kN, or 120 kN to 170 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 110 cm to 120 cm can be in the range of 60 kN to 70 kN, or 60 kN to 65 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 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 by way of non-limiting example, the maximum thrust 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 250 kN to 450 kN, 250 kN to 400 kN, or 250 kN to 350 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 270 cm to 290 cm can be in the range of 250 kN to 500 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 290 cm to 310 cm can be in the range of 250 kN to 550 kN, 300 kN to 525 kN, or 300 kN to 500 kN. The thrust mentioned above can 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.

[1031] 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.

[1032] 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 could 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 with a fan diameter in the range of 210 cm² to 230 cm² can be in the range of 1,525 K to 1,650 K. Strictly by way of non-limiting example, the TET under cruise conditions of a geared gas turbine engine with a fan diameter in the range of 340 cm² to 360 cm² can be in the range of 1,550 K to 1,660 K, or 1,550 K to 1,600 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 170 cm² to 200 cm² can be in the range of 1,400 K to 1,400 K. 650 K, or from 1,425 K to 1,625 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 having a fan diameter in the range of 270 cm to 290 cm can be in the range of 1400 K to 1600 K, from 1425 K to 1575 K, . or from 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 limits), for example 1530 K to 1600 K.

[1033] 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, 1825K, 1830K, 1835K, 1840K, 1845K, 1850K, 1855K, 1860K, 1865K, 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 by way of non-limiting example, the maximum TET of a geared gas turbine engine with a fan diameter in the range of 200 cm to 210 cm 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 with 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, or 1750 K to 1960 K. 850 K, or from 1750 K to 1800 K.Strictly by way of non-limiting example, the maximum TET of a geared gas turbine engine with a fan diameter in the range of 340 cm² to 360 cm² 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 cm 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 having a fan diameter in the range of 260 cm to 285 cm 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 an engine with . A direct-drive gas turbine with a fan diameter in the range of 270 cm to 290 cm can have a TET in the range of 1750 K to 1950 K, or 1800 K to 1900 K. Strictly as a non-limiting example, the maximum TET of a direct-drive gas turbine engine with a fan diameter in the range of 290 cm to 310 cm can have a TET 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.

[1034] 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:

[1035] [Math. 121] Heart size =

[1036] Where ^2 = the mass flow rate, in pounds per second, of the air at the inlet of the high-pressure compressor, T% = 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 square inch, of the air at the outlet of the high-pressure compressor. A unit of core size is therefore expressed as:

[1037] [Math. 122] i s* K 2 *po

[1038] 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 by way of 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 as a 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 as a non-limiting example, the core size under cruise conditions of a gas turbine engine. geared having a blower 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.

[1039] 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 with a fan diameter in the range of 260 cm² to 285 cm² 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 with a fan diameter in the range of 270 cm² to 290 cm² can be in the range of 9 to 13, 10.5 to 12.5, or 11 to 12. Strictly by way of non-limiting example, the size of The core under cruising 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 cruising conditions can be within an inclusive range bounded by any two of the core size values ​​above (i.e., the values ​​can form upper or lower bounds), e.g., 4.5 to 9.5.

[1040] 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 include at least two regions made using different materials. For example, The 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, the 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.

[1041] A blower as described and / or claimed herein may include a central portion from which the blower blades may extend, for example, in a radial direction. The blower blades may be connected to the central portion in any desired manner. For example, each blower blade may include a fastening element that may 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 that may notch into and / or engage with a corresponding notch in the hub / disc in order to secure the blower blade to the hub / disc. As a further example, the blower blades may be integrally formed with a central portion. Such an arrangement may be designated as a bladed disc or bladed ring.Any suitable process can be used to manufacture such a bladed disc or 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 linear friction welding.

[1042] 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 size of the bypass duct to be varied during operation. The general principles of this description may apply to engines with or without a VAN.

[1043] The fan of a gas turbine as described and / or claimed herein may have any desired number of fan blades, for example, 14, 16, 18, 20, 22, 24, or 26 fan blades. When the fan blades have a carbon fiber composite body, there may be 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.

[1044] 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.

[1045] 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.

[1046] 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.

[1047] 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 further non-limiting examples, the engine may produce between 5% and 8% of the available thrust. In further 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 further non-limiting examples, the engine may produce between 6% and 12% of the available thrust. In further 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 further non-limiting examples, the engine may produce between 95% and 100% of the available thrust. In other non-limiting examples, the engine may produce 100% of the available thrust.

[1048] Strictly by way of non-limiting example, climb may refer to a phase of engine operation during which 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 further non-limiting examples, the engine may produce between 80% and 95% of the available thrust. In further non-limiting examples, the engine may produce between 85% and 90% of the available thrust. For example, climb may refer to a phase of operation within an aircraft flight cycle between takeoff and arrival in cruise conditions, 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 may 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 for engine thrust.

[1049] 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).Cruising conditions can thus define an operating point of the gas turbine engine which provides a thrust which would ensure steady-state operation (i.e. the maintenance of a constant altitude). and / or a constant Mach number), or at least near-steady-state operation (i.e., maintaining at least a near-steady altitude and / or a near-steady Mach number) at mid-cruise of 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, or at least near-steady-state, operation of that aircraft at mid-cruise.

[1050] 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 gas turbine engine for an aircraft, the mid-cruise thrust, atmospheric conditions, and Mach number are known, and therefore the engine operating point at cruise conditions can be clearly defined.

[1051] 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 range of Mach 0.78, 0.79, or 0.8, in the range 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.

[1052] 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. The conditions cruising conditions can correspond to typical atmospheric conditions at any given altitude within these ranges.

[1053] 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 required level of net thrust. The known required level of net thrust is, of course, dependent on the engine and its intended application and may be, for example, a value in the range of 20 kN to 40 kN.

[1054] 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 required level of net thrust. The known required level of 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.

[1055] During operation, a gas turbine engine described and / or claimed herein may operate under the cruise conditions defined elsewhere in this document. Such cruise conditions may be determined by the cruise conditions (e.g., mid-cruise conditions) of an aircraft on which at least one (e.g., two or four) gas turbine engine may be mounted to provide propulsive thrust.

[1056] 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 further 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.

[1057] 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, conditions of Cruise according to this aspect may correspond to the mid-cruise of the aircraft, as defined elsewhere here.

[1058] 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).

[1059] 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.

[1060] During operation, under any one 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 one 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 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), for example, in the range from 0.4 to 2.5, from 0.42 to 2.08, or from 0.43 to 2.08.

[1061] 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 hydrogen mass fraction % of the fuel 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 13.41 to 15.3, or from 13.42 to 15.3.

[1062] During operation, under any one 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 one 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 (i.e., the values ​​may form upper or lower bounds), for example, within the range of 0.8 to 5, 0.88 to 4.75, or 0.89 to 4.75.

[1063] Strictly by way of non-limiting example, the percentage CO2 reduction per MJ of fuel energy may result from the fuel supplied to the combustor device having 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.

[1064] 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; comprising the bounding values ​​of X and Y.

[1065] A person skilled in the art would understand that, unless mutually exclusive, a feature or parameter described in relation to any of the above aspects may be applied to any other aspect. Furthermore, unless mutually exclusive, any feature or parameter contained or described herein may be applied to any aspect and / or combined with any other characteristic or parameter described herein.

[1066] Unless they are mutually exclusive, any parameter or value contained or described herein may be applied to and / or combined with any additional parameters and / or values ​​contained or described herein. For example, a first parameter or value contained or described herein (for example, parameter A) may be applied to and / or combined with any other parameters and / or values ​​contained or described herein (for example, one or more parameters B; a parameter C; and a parameter D, and so on) to express a product of their relationship. For example, a person skilled in the art will understand that where parameter A is described separately from parameter B, a product of their relationship may be expressed, for example, by A / B, B / a, B*A, or any other application, combination, or function of parameter A with respect to parameter B, as required.

[1067] 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 such additional parameters or values ​​related to those contained or described herein or that can be determined and / or deduced from them, to express their relationship with respect 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 parameter B; parameter C; and 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 where a parameter C can be considered separate from a parameter D, their relationship can be expressed, for example, as C+D, CD, CD, C / D, C / D, C*D, or any such additional product, application, combination, function, or expression of parameter C with respect to parameter D, or vice versa, as required. Brief description of the drawings

[1068] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[1069] [Fig.1] is a cross-sectional side view of a gas turbine engine;

[1070] [Fig.2] is a close-sectional side view of an upstream part of a geared gas turbine engine;

[1071] [Fig.3] is a partially cutaway view of a reducer for a gas turbine engine;

[1072] [Fig.4] is a close-up cross-sectional side view of a direct-drive gas turbine engine;

[1073] [Fig.5] is a schematic view of an aircraft having two gas turbine engines of the present application mounted on it;

[1074] [Fig.6] is a cross-sectional view through a combustor device of the engine of [Fig.1] in a plane perpendicular to the main axis of rotation of the engine;

[1075] [Fig.7] is a schematic cross-section of a duplex fuel spray nozzle of the combustor device of [Fig.6];

[1076] [Fig.8] is a schematic cross-section of a single-flow fuel spray nozzle of the combustor device of [Fig.6];

[1077] [Fig.9] is a schematic cross-sectional view of the engine in [Fig.1];

[1078] [Fig.10] is an additional partial cross-sectional view of the engine of [Fig.1];

[1079] [Fig. 1 1] is a schematic representation of a propulsion system for a aircraft including the engine of the [Fig. 1]; and

[1080] [Fig. 12] shows a method of operation of the gas turbine engine.

[1081] DETAILED DESCRIPTION OF THE DISCLOSURE

[1082] 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 that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 includes a core 11 that receives the core airflow A. The engine core 11 comprises, in axial flow series, a low-pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, a low-pressure turbine 19, and a core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass duct 22. The fan 23 is attached to and driven by the low-pressure turbine 19 via of a first low pressure shaft 26 and an epicycloidal reducer 30.

[1083] 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 combustion chamber 16, with the term "combustion equipment" being used interchangeably. 16” and “combustion chamber 16” are used interchangeably here. 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 high-pressure interconnecting 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 propulsive thrust. The epicyclic reducer 30 is a reduction gearbox.

[1084] 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 toothed ring or crown 38 which is coupled, via links 40, to a stationary support structure 24.

[1085] 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.

[1086] 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.

[1087] 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.

[1088] It should be borne in mind that the arrangement shown in Figures 2 and 3 is by way of example only, and that various alternatives are within the scope of this disclosure. Strictly by way of example, any suitable arrangement may be used for positioning the gearbox 30 in the motor 10 and / or for connecting 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) may have any desired degree of rigidity or flexibility.As a further example, any suitable arrangement of bearings between rotating and stationary parts of the motor (e.g., between the input and output shafts from the gearbox and fixed structures, such as the gearbox housing) may be used, and disclosure is not limited to the arrangement given as an example in [Fig. 2]. For example, where the gearbox 30 has a star arrangement (described above), those skilled in the art will readily understand that the arrangement of the output and support links and bearing locations will typically be different from that shown as an example in [Fig. 2].

[1089] 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.

[1090] Optionally, the reducer may drive additional and / or alternative components (for example, the intermediate pressure compressor and / or a booster).

[1091] 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.

[1092] 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].

[1093] With reference to [Fig.4], a gas turbine engine is usually indicated as 10, having a main axis of rotation 9. The engine 10 comprises, in axial flow series, an air inlet 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 inlet 12 and the exhaust nozzle 20.

[1094] 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.

[1095] 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 combustion chamber 16, with the terms "combustion equipment 16" and "combustion chamber 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.

[1096] Other gas turbine engines to which this description may be applied may have alternative configurations. By way of example, such engines may have an alternative number of interconnecting shafts (for example, two) and / or an alternative number of compressors and / or turbines. In addition, the engine may include a reduction gear provided in the drive train from a turbine to a compressor and / or a blower.

[1097] 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.

[1098] 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 principal axis of rotation 9), a radial direction (in the downward-upward direction in [Fig. 1]), and a circumferential direction (perpendicular to the page in the view of [Fig. 1]). The axial, radial, and circumferential directions are mutually perpendicular.

[1099] 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. Fuel is supplied to the fuel tanks 50a and 50b from a fuel inlet port 62. Other fuel systems can be used with other fuel tank arrangements.

[1100] The fuel F supplied to the combustion equipment 16 may comprise a fossil-based hydrocarbon fuel, such as kerosene. Thus, the fuel F may comprise molecules from one or more of the n- chemical families Alkanes, isoalkanes, cycloalkanes, and aromatics. There is an expectation within the aviation industry regarding a trend toward the use of fuels other than the traditional kerosene-based jet fuels currently in use. 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 provided, fuel F may include one or more trace elements, such as sulfur, nitrogen, oxygen, inorganic substances, and metals.

[1101] 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.

[1102] The functional performance of a given fuel composition, or of a fuel mixture F for use in a given mission, can be defined, at least in part, by the fuel's ability to complete the Brayton cycle of the gas turbine engine 10. The parameters defining the functional performance may include, for example, specific energy; energy density; thermal stability; and emissions comprising 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 that exist in an exit plane of the exhaust nozzle of the gas turbine engine and do not volatilize when heated to a temperature of 350 °C.Any reference to soot or smoke here may also apply to other types of particulate matter emissions known in the art. Gaseous emissions may include... 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.

[1103] 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 fuels with high aromatic content; and maximum surface tension (mN / m) can ensure sufficient spray break and fuel atomization.

[1104] 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.

[1105] A number of sustainable aviation fuel blends have been approved for use. For example, some approved blends include Blend ratios of up to 10% sustainable aviation fuel are permitted, while other approved blends include blend 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 containing) 100% sustainable aviation fuel (SAF) will eventually be approved for use.

[1106] 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 more of the following: higher iso-alkane and higher 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.

[1107] In some examples, the sustainable aviation fuel(s), or the mixture(s) supplied to the combustion equipment 16 may have a relatively lower aromatics and / or other non-paraffinic content than kerosene. The sustainable aviation fuel may have an aromatics content of 30%, 20%, 15%, 10%, 8%, 5%, or less than 5%, for example; 4%, 3%, 2%, 1%, or less than 1%; 0.75%, 0.5%, 0.25%, or less than 0.25%; 0.2%, 0.1%, or less than 0.1%; 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.

[1108] 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.

[1109] Figure 6 shows a cross-section through the combustor device 16 of the engine 10 of Figure 1 in a plane perpendicular to the main axis of rotation 9 of the engine 10. The combustor device 16 comprises an annular combustion chamber 401, defined by a sleeve 402. Alternative combustor device configurations can be used in other embodiments, for example, annular, canned, etc. The combustor device 16 comprises a plurality of fuel spray nozzles 403, 404 arranged around the circumference of the combustor device 16. Each fuel spray nozzle 403, 404 comprises one or more fuel injectors arranged to inject fuel into the combustion chamber 401. In this example, the combustor device 16 comprises 16 fuel spray nozzles 403, 404.In other examples, the combustor device 16 may include any appropriate number of fuel spray nozzles 403, 404, for example, a number of fuel spray nozzles in the range of 14 to 22. In some examples, the number of fuel spray nozzles 403, 404 may be between 16 and 20. In still other examples, the number of fuel spray nozzles may be 14, 15, 16, 17, 18, 19, 20, 21, 22 or a number in a defined range between any two of the values ​​in this sentence.

[1110] The number of fuel spray nozzles 403, 404 can also be quantified as a ratio between the number of fuel spray nozzles and the engine core size. The core size defines the size of the core 11 of the gas turbine engine 10. The engine core size can be defined as:

[1111] [Math. 123] Heart size = m2~pr

[1112] Where Ûl? = the mass flow rate, in pounds per second, of the air at the inlet of the high-pressure compressor 15, = the temperature, in Kelvin, of the air at the outlet of the compressor high pressure 15, and = 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:

[1113] [Math. 124] i s* K 2 • po

[1114] 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.

[1115] The number of fuel spray nozzles per unit engine core size (in the units given above) may be in the range of 2 to 6, for example, 2, 3, 4, 5, 6, or in a defined range between any two of these values. The number of fuel spray nozzles per unit engine core size may be in the range of 2.7 to 4, preferably between 3 and 3.6. In some preferred examples, the number of fuel spray nozzles per unit engine core size may be in the range of 2.5 to 4.5, for example, 2.5, 3, 3.5, 4, or 4.5, or any defined range between any two of these values. In yet other examples, the number of fuel spray nozzles per unit engine core size may be in the range of 3 to 4, for example 3.0, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, or 4.0, or any range defined between any two of these values.In still other examples, the number of fuel spray nozzles per unit engine core size may be 2.0, 2.1, 2.2, 2.3, 2.4, 2.5, 2.6, 2.7, 2.8, 2.9, 3.0, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4.0, 4.1, 4.2, 4.3, 4.4, 4.5, 4.6, 4.7, 4.8, 4.9, 5.0, 5.1, 5.2, 5.3, 5.4, 5.5, 5.6, 5.7, 5.8, 5.9, or 6.0, or within a defined range between any two of these values. .

[1116] The core size is defined here at an engine operating condition corresponding to a maximum value of the semi-dimensional flux at a high-pressure compressor inlet, defined as:

[1117] [Math. 125] “2^

[1118] where 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.

[1119] The operating condition corresponding to the maximum semi-dimensional flux 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 of the present application. In a specific example, the peak climb may correspond to operation at an altitude between 30,000 feet and 39,000 feet (more specifically 35,000 feet), a forward speed of Mach number 0.75 to 0.85, and an ambient air temperature (AAT) of ISA+10 K to ISA+15 K.

[1120] The combustor device 16 includes a number of duplex fuel spray nozzles 403 (also known as internally mounted nozzles) in which a primary fuel injector is integrated in the same fuel nozzle as a main fuel injector. The combustor device 16 also includes a number of single-flow fuel spray nozzles 404, each of which includes only a main fuel injector. In other examples, the combustor device 16 may include only duplex fuel spray nozzles or only single-flow fuel spray nozzles.

[1121] In this example, the combustor device 16 comprises 12 duplex fuel spray nozzles 403 and 4 single-flow fuel spray nozzles 404. The duplex fuel spray nozzles 403 are shown in [Fig. 6] by hatched circles. The duplex fuel spray nozzles 403 are arranged in groups of three around the circumference of the combustor device 16, with each group positioned diametrically opposite another group. In other examples, the combustor device 16 may comprise any appropriate number of duplex fuel spray nozzles, for example in the range of 10 to 14 nozzles, and any appropriate number of single-flow fuel spray nozzles, for example in the range of 4 to 8 nozzles.In some examples, the number of duplex fuel spray nozzles may be 10, 11, 12, 13 or 14, or within a defined range between any two of these values. In some examples, the number of single-flow fuel spray nozzles may be 4, 5, 6, 7, or 8, or within a defined range between any two of these values. Duplex fuel spray nozzles may be arranged in any appropriate number of groups, or they may not be arranged in groups. If so, each group of duplex fuel spray nozzles may include any appropriate number of nozzles, for example, in the range of 2 to 8 nozzles. In some examples, each group of duplex nozzles may include 2, 3, 4, 5, 6, 7, or 8 fuel spray nozzles, or a number within a defined range between any two of these values.

[1122] The combustor device 16 further includes four igniters 405 arranged to ignite an air-fuel mixture in the combustion chamber 401 during operation. Each igniter 405 is arranged adjacent to one of the groups of duplex fuel-spray nozzles 403. The duplex nozzles 403 are therefore each located closer to a respective igniter (for example, its nearest igniter) than to the single-flow nozzles 404. Each igniter 405 is arranged diametrically opposite another of the igniters 405. In other examples, the combustor device may include fewer or more igniters, for example, a number of igniters in the range of 1 to 8, and the igniters may be arranged differently. For example, one or more of the igniters may not be arranged adjacent to one of the duplex fuel spray nozzle groups and one or more of the igniters may not be arranged diametrically opposite another of the igniters.In some examples, the combustor device may include 1, 2, 3, 4, 5, 6, 7 or 8 igniters, or a number within a defined range between any two of the values ​​in this sentence.

[1123] In the example shown, when the engine 10 is operating at low power (below a stage point), for example during or shortly after starting, fuel is supplied only to the primary injectors of the duplex fuel spray nozzles 403 for distribution to the combustion chamber 401. A higher fuel flow rate is therefore supplied to the duplex nozzles 403 compared with the single-flow nozzles 404 below the stage point. As the power output of the engine 10 and the mass flow rate of air through the engine 10 increase, the stage point is reached at which additional fuel is supplied to the primary fuel injectors of one or more of the duplex fuel spray nozzles 403 and to the primary fuel injectors of one or more of the single-flow fuel spray nozzles 404 for distribution to the combustion chamber 401.In this example, at higher power levels, fuel is injected by all the main fuel injectors from both the 403 duplex fuel spray nozzles and the spray nozzles. single-flow fuel 404, in addition to the fuel injected by the primary injectors of the duplex fuel spray nozzles 403. In this example, the fuel flow rate supplied to the primary injectors of the single-flow fuel spray nozzles 404 is less than or equal to the fuel flow rate supplied to the primary injectors of the duplex fuel spray nozzles 403. Thus, because both the primary and primary injectors of the duplex fuel spray nozzles 403 receive fuel, the duplex fuel spray nozzles 403 receive more fuel than the single-flow fuel spray nozzles 404 at and above the stage point.In an alternative example, fuel is supplied only to the main fuel injectors of one or more of the duplex fuel spray nozzles 403 and to the main fuel injectors of one or more of the single-flow fuel spray nozzles 404 at and above the stage point; that is, fuel is not supplied to the primary injectors of the duplex fuel spray nozzles 403.

[1124] The fuel flow distributed to the plurality of fuel spray nozzles is thus oriented such that the fuel flow to a first subset of the plurality of fuel spray nozzles (the duplex fuel spray nozzles 403 in this example) is greater than that distributed to a second subset of the fuel spray nozzles (the single-flow fuel spray nozzles 404 in this example). This can provide a primary fuel flow to fuel spray nozzles that are located relatively closer to the igniters 405 to aid ignition and flame stability at low engine power, during engine starting, or during engine relighting.In some examples, the first subset (e.g., duplex nozzles) of fuel spray nozzles may comprise at least half, preferably at least two-thirds, of the total number of fuel spray nozzles.

[1125] In other examples, the fuel flow rate at each fuel spray nozzle supplied in the combustor device may be the same, and there may be no direction of the fuel flow to any subset of the nozzles. In such an example, all the fuel flow nozzles may be single-flow nozzles, or they may all be duplex nozzles. In still other examples, other arrangements of fuel spray nozzles may be supplied in which fuel is directed to those adjacent to, or closer to, the igniters. For example, two subsets (which may be independently controlled) of duplex nozzles or two subsets of single-flow nozzles may be supplied, which may be directed as described above.

[1126] Figure 7 shows one of the duplex fuel spray nozzles 403 of the combustor device 16. The duplex nozzle 403 includes a primary fuel injector 501, a main fuel injector 502, and an air duct 503. The primary injector 501 includes a primary inlet 504 arranged to receive a primary fuel flow P and a primary fuel circuit 505 arranged to distribute the primary fuel flow to the outlet 506 of the nozzle 403. The main injector 502 includes a main inlet 507 arranged to receive a main fuel flow M and a main fuel circuit 508 arranged to distribute the main fuel flow to the outlet 506 of the nozzle 403. The air duct 503 receives high-pressure air from the high-pressure compressor 15 and distributes the high-pressure air to the outlet 506 of the nozzle 403.

[1127] The duplex nozzle 403 is designed to produce, at the outlet 506 of the nozzle 403, a primary fuel cone from the primary injector 501 and a main fuel cone from the main injector 502 (illustrated in [Fig. 7] as dashed lines labeled P and M, respectively). When both the primary and main injectors 501 and 502 are active, the primary and main cones are arranged concentrically, with the main cone arranged annularly outside the primary cone. Those skilled in the art will be familiar with such fuel spray patterns.

[1128] It should be borne in mind that the duplex nozzle 403 of [Fig.7] is simply given as an example and that other examples may use an alternative configuration of the duplex nozzle.

[1129] Figure 8 shows one of the single-flow fuel spray nozzles 404 of the combustor device 16. The nozzle 404 includes a main fuel injector 601, comprising a main inlet 602 arranged to receive a main fuel flow M and a main fuel circuit 603 arranged to distribute the main fuel flow to the outlet 604 of the nozzle 404. The nozzle 404 is designed to produce a main fuel cone at the outlet 604 of the nozzle 404 (represented by the dashed lines referenced M). Similarly, air is supplied to the outlet 604 of the nozzle by an air duct 605.

[1130] It will be borne in mind that the single-flow fuel spray 404 of [Fig.8] is simply given as an example and that other examples may use an alternative configuration of single-flow fuel spray nozzle 404.

[1131] Figures 9 and 10 each show a section through the engine 10, viewed perpendicular to the main axis of rotation 9, comprising a portion of the combustor device 16 including one of the duplex fuel spray nozzles 403 and one of the igniters 405. A similar arrangement is provided at the location of the single-flow fuel spray nozzles 404 as for duplex nozzles 403. The combustor device 16 is mounted within a cavity 406 formed by an inner air casing 407 and an outer air casing 408. In operation, the high-pressure compressor 15 distributes high-pressure air D to the cavity 406 via a diffuser 409. At this point, a quantity of the air enters the combustor device 16 as combustion air E through the fuel nozzle 403 and / or mixing orifices at the inlet of the combustor device 16. The remaining air flows around the combustor device 16 as cooling air G, a quantity of which is admitted downstream of the fuel nozzle 403 as described below with reference to [Fig. 10].

[1132] One or more temperature and / or pressure probes (not shown) may be installed in the diffuser housing 409 and arranged to measure the temperature and / or pressure of the high-pressure air D delivered to the cavity 406 from the high-pressure compressor 15 via the diffuser 409 (i.e., the temperature and pressure at the outlet of the high-pressure compressor 15). Such a temperature probe may be designated as probe T3 and such a pressure probe may be designated as probe P3. It should be borne in mind that the motor 10 may include any suitable arrangement of pressure and temperature probes that may be positioned at any suitable location within the motor 10. As used here, T3 and P3, and any other numbered pressures and temperatures, may be defined using the station numbering listed in the SAE AS755 standard.

[1133] The combustor device 16 operates as a rich-burn, rapid-cooling, lean-burn (RQL) combustor device. In other examples, the combustor device 16 may be an alternative type of combustor device, such as a standard rich-burn combustor device (without fuel flow direction). With reference to [Fig. 10], the combustion chamber 401 of the combustor device 16 RQL is divided into three zones along the length of the combustor device 16: a rich zone 801, a rapid cooling zone 802, and a lean zone 803. In operation, a rich air-fuel mixture is introduced into the rich zone 801 from the fuel spray nozzle 403 where it is ignited by the igniter 405. Within the rich zone 801, fuel is burned at a fuel / air ratio greater than stoichiometry (for example, at an equivalence ratio of about 1.8).Air is then introduced into the combustion products, via primary ports 804 arranged in the jacket 402 of the combustor device 16, before the combustion products reach the rapid cooling zone 802. Additional air is added to the still-burning fuel via the primary ports 804 (which may be designated as cooling ports). Air is added through the primary ports 804 at a higher rate (e.g., higher). (within the rich zone), thus cooling the combustion to a fuel / air ratio significantly lower than the stoichiometric fuel / air ratio (for example, to an equivalence ratio between 0.5 and 0.7), while still allowing fuel to burn. Consequently, very little of the combustion process can be carried out at near the stoichiometric fuel / air ratios, and therefore relatively little nitrogen oxide (NOx) is produced. Air is then reintroduced into the combustion products via secondary ports 805 arranged in the jacket 402 of the combustor device 16, while the combustion products are in the lean zone 803 (or just before they reach the lean zone 803). Within the depleted zone 803, fuel is burned at a fuel / air ratio lower than stoichiometry (e.g., at an equivalence ratio of about 0.5 and 0.7).After passing through the depleted zone 803, the combustion products exit the combustor device 16. The secondary orifices 805 can be designated as dilution orifices, and can be arranged to gradually introduce dilution air into the depleted zone 803. Fuel added by the fuel spray nozzle is substantially burned completely by the time the air exits at an outlet of the combustor device, before flowing to the turbine.

[1134] Figure 11 shows a portion of a propulsion system 900 for an aircraft. The propulsion system 900 comprises the gas turbine engine 10 of Figure 1. The engine 10 further comprises a fuel system and an oil system. The fuel system comprises: a low-pressure fuel pump 902, a fuel-oil heat exchanger 903, a main (or high-pressure) fuel pump 904, a control device 908, and a fuel distribution valve 909. The propulsion system 900 further comprises a fuel tank 901. The oil system comprises an oil tank 905, an oil supply pump 906, and a main oil pump 907. In this example, the low-pressure fuel pump 902 is shown forming part of the gas turbine engine 10.In other examples, the low-pressure fuel pump, or auxiliary fuel pumps, may be supplied as part of the fuel system on board the aircraft to which the gas turbine engine is mounted.

[1135] The low-pressure fuel pump 902 is arranged to distribute fuel from the fuel tank 901 to the fuel-oil heat exchanger 903 via a suitable arrangement of pipes, conduits, etc. (not shown). The main fuel pump 904 is designed to distribute fuel from the fuel-oil heat exchanger 903 to the fuel spray nozzles of the combustor device 16 via the fuel distribution valve 909 and a suitable arrangement of pipes, conduits, etc. (not shown). The fuel distribution valve 909 is arranged to distribute fuel between a The system consists of a main manifold 909a and a primary manifold 909b. The main manifold is fluidly connected to the main injectors of each of the fuel spray nozzles 404, 403, as shown in [Fig. 11]. It therefore supplies fuel to all of the duplex fuel spray nozzles 403 and single-flow nozzles 404. The primary manifold 909b is fluidly connected to the primary injectors of each of the duplex fuel spray nozzles 403. The primary manifold 909b can therefore be used to provide a higher fuel flow rate to the first subset of fuel spray nozzles (e.g., the duplex fuel spray nozzles 403 in this example) compared to the fuel flow rate supplied to the second subset of fuel spray nozzles via the main manifold 909a.For example, below a certain engine power threshold, fuel may be delivered only to the first subset of fuel spray nozzles via the primary manifold 909b, or delivered to the first subset of fuel spray nozzles at a higher fuel flow rate compared to the second subset of fuel spray nozzles. This can limit the production of undesirable combustion products such as nitrogen oxides (NOx), unburned hydrocarbons (HC), and carbon monoxide (CO), and can direct fuel flow to the injectors closest to the igniters to aid flame stability and ignition at low engine power levels.

[1136] The oil supply pump 906 is arranged to distribute lubricating oil from the oil reservoir 905 to the fuel-oil heat exchanger 903 via a suitable arrangement of pipes, conduits, etc. (not shown). The main oil pump 907 is arranged to distribute oil from the fuel-oil heat exchanger 903 to engine components 10 as required via a suitable oil distribution arrangement (not shown). The fuel flow path from the fuel reservoir 901 to the combustor device 16, via the pumps 902, 904, and the fuel-oil heat exchanger 903, during operation, is illustrated in [Fig. 11] by dashed or dotted arrows.The oil flow path from the oil reservoir 905 to the fuel-oil heat exchanger 903, via the oil supply pump 906, and over the operating engine components 10 is illustrated in [Fig. 11] by solid arrows.

[1137] The control device 908 comprises a suitable arrangement of processors and electronic memory. The control device 908 is in communication with a fuel-oil heat exchanger 903, as illustrated by the dashed and dotted line in [Fig. 11], and is designed to control the operation of the fuel-oil heat exchanger 903. In some examples, the control device Control device 908 can be designed to control the flow of oil through the fuel-oil heat exchanger 903. Control device 908 is designed to control the operation of the fuel-oil heat exchanger 903 by providing control signals to the fuel-oil heat exchanger 903. Control device 908 is designed to control the operation of the fuel-oil heat exchanger 903 to adjust at least one property or parameter of the fuel at the inlet of the combustor device 16. In the illustrated example, control device 908 is designed to control the operation of the fuel-oil heat exchanger 903 to control the viscosity of the fuel at the inlet of the combustor device 16.In other examples, the control device 908 can be designed, either additionally or alternatively, to control the operation of the fuel-oil heat exchanger 903 to control the fuel temperature at the inlet of the combustor device. The control device 908 can be a separate control device as illustrated, or it can be part of an electronic engine control device (EEC) arranged to control other engine functions.

[1138] In the example shown, the fuel-oil heat exchanger 903 is arranged between the low-pressure fuel pump 902 and the main fuel pump 904, although the fuel-oil heat exchanger 903 may be arranged in any suitable location or position relative to the other components of the propulsion system 900. In other examples, the propulsion system 900 may include one or more additional heat exchangers arranged to receive oil from the oil system, or the propulsion system 900 may include one or more additional oil systems arranged to supply oil to one or more additional heat exchangers. It should be borne in mind that the propulsion system 900 as shown in [Fig. 11] is simply a schematic view of an illustrative propulsion system.

[1139] In one example, the control device 908 is designed to control the operation of the fuel-oil heat exchanger 903 to lower the fuel viscosity to 0.58 mm2 / s or less at the inlet of the combustor device 16 under cruising conditions. Alternatively, the control device 908 can be designed to control the operation of the fuel-oil heat exchanger 903 to lower the fuel viscosity to between 0.58 mm² / s and 0.30 mm² / s, for example 0.58, 0.57, 0.56, 0.55, 0.54, 0.53, 0.52, 0.51, 0.50, 0.49, 0.48, 0.47, 0.46, 0.45, 0.44, 0.43, 0.42, 0.41, 0.40, 0.39, 0.38, 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31 or 0.30 mm² / s at the device inlet combustor 16 under cruising conditions. Alternatively, the control device 908 can be designed to control the operation of the fuel-oil heat exchanger 903 to lower the fuel viscosity to 0.57, 0.56, 0.55, 0.54, 0.53, 0.52, 0.51, 0.50, 0.49, 0.48, 0.47, 0.46, 0.45, 0.44, 0.43, 0.42, 0.41, 0.40, 0.39, 0.38, 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31 or 0.30 mm2 / s or less, or within any defined range between any two of these values, at the inlet of the combustor device 16 under cruising conditions. The control device 908 can be designed to control the operation of the fuel-oil heat exchanger 903 to lower the fuel viscosity between 0.55 mm2 / s and 0.35 mm2 / s, 0.53 mm2 / s and 0.35 mm2 / s, 0.50 mm2 / s and 0.35 mm2 / s, 0.48 mm2 / s and 0.35 mm2 / s, 0.48 mm2 / s and 0.38 mm2 / s, 0.48 mm2 / s and 0.40 mm2 / s, 0.46 mm2 / s and 0.40 mm2 / s, 0.44 mm2 / s and 0.40 mm2 / s, or 0.44 mm2 / s and 0.42 mm2 / s at the inlet of the combustor device 16 under cruising conditions.

[1140] The control device 908 may, in addition or alternatively, be designed to control the operation of the engine 10 such that a reduction of 10 to 70% in the average of particles / kg of nvPM in the exhaust of the gas turbine engine 10 when the engine 10 is operating at an available thrust of 85% for given operating conditions, and in the average of particles / kg of nvPM in the exhaust of the gas turbine engine 10 when the engine 10 is operating at an available thrust of 30% for given operating conditions, is achieved when fuel supplied to the combustion chamber 16 is sustainable aviation fuel instead of a fossil-based hydrocarbon fuel. In other examples, the reduction in nvPM may be as otherwise defined herein.

[1141] In this example, or any other example described herein, the control device 908 is designed to control the fuel distribution valve 909 to control the distribution of fuel to the fuel spray nozzles of the combustor device 16. The control device 908 is designed to direct the fuel flow to the nozzles such that the first subset of a plurality of fuel spray nozzles receives more fuel than the second subset. The control device 908 is designed to control the fuel distribution valve 909 such that below a certain point, fuel is distributed only to the primary fuel injectors of the duplex fuel spray nozzles 403.Above a stage point, the control device 908 is designed to control the fuel distribution valve 909 so that additional fuel is distributed to the main fuel injectors of the duplex fuel spray nozzles 403 and the single-flow fuel spray nozzles 404. As such, the duplex fuel spray nozzles 403 receive more fuel than the single-flow fuel spray nozzles 404 (below and possibly above the stage point). The control device 908 can alternatively be designed to control the fuel distribution valve 909 to regulate fuel distribution such that any suitable subset of fuel spray nozzles 403, 404 receives more fuel than other fuel spray nozzles 403, 404. This advantageously allows for optimizing fuel distribution for engine performance, emissions, or any other suitable criterion. The fuel distribution system shown in the Figures should be understood as an example of how fuel is directed to the fuel spray nozzles; other configurations are possible. For example, two sets of independent single-flow nozzles may be provided.

[1142] The gas turbine engine 10 of this application is designed to supply fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles 403, 404. In other words, the gas turbine engine 10 is designed to inject fuel (F) comprising a sustainable aviation fuel (SAF) into the combustion chamber (401). During operation, therefore, the fuel supplied to the fuel spray nozzles 403, 404 includes SAF.

[1143] “Fuel including a SAF” may mean that the fuel supplied to the device The combustion chamber 401, via the fuel spray nozzles 403 and 404, contains a percentage of 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%. More generally, "fuel containing SAF" may mean a fuel comprising any mixture of SAF and fossil kerosene fuel, containing up to 100% SAF and no fossil kerosene fuel. Fuel containing SAF may be a fuel containing a percentage of SAF of 1%, 2%, 5%, 10%, 15%, 20%, 30%, 40%, 50%, 60%, 70%, 80%, 90%, 95%, 100%, or within any defined range between any two of these values.

[1144] “SAF” means a kerosene-type fuel whose hydrocarbon component is substantially full of paraffinic hydrocarbons. "SAF" means, in addition or alternatively, a kerosene-type fuel with a hydrogen mass fraction in the range of 13.7% to 16.9%, for example, 15.3%. "A fossil-based hydrocarbon fuel" or "fossil fuel" used anywhere here means a fossil-based kerosene with a hydrogen mass fraction in the range of 12.0% to 14.8%, for example, 13.4%.

[1145] Emissions of non-volatile particulate matter (nvPM)

[1146] An emission index (El) of nvPM is defined herein as the mass of nvPM produced per unit mass of fuel used by the combustor device 16 of the gas turbine engine 10 of any example described herein. In the present application, The nvPM emission index is the mass of nvPM produced by the gas turbine engine 10 in milligrams divided by the corresponding mass of fuel used by the engine 10 in kilograms, unless otherwise stated.

[1147] Examples of both uncorrected and loss-corrected (SL) exhaust emissions from series aircraft engines (using fossil-based JET Al kerosene-type aviation fuel), measured in accordance with the procedures in Volume II of ICAO Annex 16 and, where appropriate, certified by the States of engine design in accordance with their national regulations, are provided in the ICAO Aircraft Engine Emissions Database, hosted on behalf of the International Civil Aviation Organization (ICAO) by the European Union Aviation Safety Agency (EASA). The database covers engine types whose emissions are regulated, namely turbofan and turbofan engines with a static thrust exceeding 26.7 kilonewtons of nvPM emission indices.

[1148] Examples of empirical correction factors for the mass and number El of nvPM, together with definitions, symbols, SI units, acronyms, and procedures for estimating system losses are given in Annex 16 to the Convention on International Civil Aviation, entitled "Environmental Protection"; Volume II - Emissions from Aviation Engines Fourth Edition, July 2017; Appendix 8. Procedures for Estimating Corrections for Non-Volatile Particle nvPM Losses in the System; which is incorporated herein by reference.Within this, it is noted that the implementation of the nvPM sampling and measurement system can result in a significant loss of particles on the order of 50% for mass and 90% for the number of nvPM (for example due to the loss of particles on the walls of the sampling system caused by deposition mechanisms), and that particle losses depend on size and therefore depend on the operating condition of the engine, the combustion chamber technology, and other miscellaneous factors.As such, the loss-corrected nvPM emission indices refer to nvPM emissions at the engine outlet adjusted for particle size-dependent losses in the sampling and measurement system, excluding thermophoretic losses in the collection part, assuming that the nvPM in the exhaust outlet plane of an engine has a log-normal distribution, a fixed value of effective nvPM density, a fixed value of geometric standard deviation, limiting the nvPM mass concentration to the detection limit, a minimum particle size of 0.01 pm, and no coagulation.

[1149] For example, as specified in Annex 16 of the Convention on International Civil Aviation, entitled “Environmental Protection”, the factor of The correction of mass concentration (EImass e) as a function of losses in the system without correction of thermophoretic losses in the collection section is defined as the ratio between the estimated value of the mass concentration in the outlet plane of the engine exhaust nozzle, without correction as a function of thermophoretic losses in the collection section, and the measured mass concentration, and can be calculated as follows:

[1150] [Math. 126] "nvPMraasseEP ^SL mass— DF.xnvPM — 1 m&sse_ST P

[1151] Where:

[1152] EImass e represents the mass emission index of nvPM corrected for thermophoretic losses, in mg / kg of fuel;

[1153] KsL_maSS e represents the Elmasse correction factor as a function of losses in the system, without correction of thermophoretic losses in the Collection part, in pg / m3

[1154] nvPMmass e _ep represents the estimated value of the mass concentration of nvPM in the outlet plane of the engine exhaust nozzle, not corrected for thermophoretic losses in the Collection part;

[1155] DFi represents the first dilution factor; and,

[1156] nvPMmass e _ST represents the mass concentration of nvPM, after dilution, measured by an instrument under STP conditions, pg / m3.

[1157] Furthermore, the correction factor of EInum as a function of losses in the system is defined as the ratio between the estimated value of the number concentration in the outlet plane of the engine exhaust nozzle, without correction as a function of thermophoretic losses in the Collection part, and the measured number correction, and can be calculated as follows:

[1158] [Math. 127] _ nvPMDun} EP &SL num — DF.xDF? nvPM — rz num_STP

[1159] Where:

[1160] EInum represents the emission index in number of nvPM corrected according to thermophoretic losses, in number / kg of fuel;

[1161] KSL_num represents the correction factor of EInum as a function of losses in the system, without correction of thermophoretic losses in the Collection part, number / m3;

[1162] nvPMnum _Ep represents the estimated value of the number concentration of nvPM in the outlet plane of the engine exhaust nozzle, not corrected for thermophoretic losses in the Collection part;

[1163] DFi represents the first dilution factor;

[1164] DF2 represents the second dilution factor (VPR) according to the calibration; and,

[1165] nvPMnumSTRepresents the concentration in number of nvPM, after dilution, measured by an instrument under STP conditions, number / cm3.

[1166] For example, as specified in Annex 16 of the Convention on International Civil Aviation, entitled "Environmental Protection", the mass (nvPMmass e _Ep) and the number (nvPMnum Ep) in the exhaust nozzle outlet plane of the engine can be determined using the following procedure: a. For a measured nvPMnumSTP, start with an initial value of [Math. 128] nvPMnum EP = 3 x DF Y x DF2 x nvPM aumSTP

[1167] b) An initial value of 0.02 pm must be assumed for the mean geometric diameter, Dmg, of the log-normal distribution of particle size (pm).

[1168] c) Starting from the assumed initial values ​​of nvPMnum EP and Dmg obtained at points a) and b), estimate the mass (nvPMmass e_est) and number (nvPMnum_EST) concentrations of nvPM, where:

[1169] (nvPMmass e est) represents the estimated value of the mass concentration, after dilution, measured by an instrument (i.e., corrected for dilution), pg / m3; and,

[1170] (nvPMnum est) represents the estimated number concentration, after dilution, measured by an instrument (i.e., corrected for dilution), number / cm3; using the following equations:

[1171] [Math.129] nvP M massEST =2^.01 pmniasse (Dm) x xnvPMmun FPx

[1172] [Math. 130] nvPMnum_EST = (D^x nvPMDW1EP xx Ali^DJ

[1173] Where:

[1174] nvPMmi represents the instrument measuring the mass of non-volatile particles;

[1175] qmass e(Dm) represents the overall penetration fraction in the sampling and measurement system for the nvPMmi without thermophoretic losses in the Collection part at the electrical mobility size of the particles, Dm;

[1176] p represents the assumed effective density of nvPM, g / cm3;

[1177] Dm represents the electrical mobility diameter of nvPM, pm;

[1178] fign(Dm) represents the log-normal distribution function with geometric standard deviation parameters, ôg, and geometric mean diameter, Dmg;

[1179] and where:

[1180] nvPMni represents the instrument measuring the number of non-volatile particles;

[1181] i]num(Dm) represents the overall penetration fraction in the sampling and measurement system for nvPMni without thermophoretic losses in the Collection part at the electrical mobility size of the particles, Dm;

[1182] and where:

[1183] [Math.131] (. , 2 TJ =____1____ x 1

[1184] Ah / T) ) = — x 1 is 'a 'ar8cur of an interval in the natural logarithm of n log^ê) base ;

[1185] ôg is the assumed geometric standard deviation of the log-normal distribution;

[1186] e is Euler's number; and,

[1187] n is the number of particle intervals per decade.

[1188] d) Determine the difference, ô (defined as the sum of the squares of the relative differences between the measured and calculated mass and number concentrations, corrected for dilution) between nvPMnum_STP, nvPMmass e _STp and the estimated values ​​of the number concentration of nvPM (nvpMnum_EST) and the mass concentration of nvPM (nvPMmass e _est) from the initial values ​​of the exhaust nozzle outlet plane of the engine using the following equation:

[1189] [Math. 132] „ _ / DF^DF^nvPMn^^ \ 2 t DFpcnvPM^ STrnvPMm IS \ 2 $ ~ \ DF^DF^nvPMnum STP / ' (DF^ïvPMmasse STP /

[1190] e) Repeat steps c) to d) varying nvPMnum EP and Dmg until ô is reduced to less than 1 x 109.

[1191] f) Once ô is reduced to less than 1 x 109, the final values ​​of nvPMnum EP and Dmg are those associated with this minimized value of ô.

[1192] g) Using nvPMnum EP and Dmg from step f), nvPMnum EP must be determined using the following expression:

[1193] [Math. 133] nvPMmass EP = 01 puf 6 "

[1194] Alternative methods for determining and / or correcting losses in the system are also proposed, for example, by:

[1195] Durand et al., 2023 (Correction for particle loss in a regulatory aviation nvPM emissions System using measured particle size, Journal of Aerosol Science, Volume 169, 2023, 106140, ISSN 0021-8502);

[1196] Corbin et al., 2022 (Aircraft-engine particulate matter émissions from conventional and sustainable aviation fuel combustion : Comparison of measurement techniques for mass, number, and size, Atmospheric Measurement Techniques, 15 (10) (2022), pp. 3223-3242);

[1197] Durdina et al., 2021 (Réduction of nonvolatile particulate matter émissions of a commercial turbofan engine at the ground level from the use of a sustainable aviation fuel blend, Environmental Science and Technology, 55 (21) (2021), pp. 14576-14585);

[1198] Harper et al., 2022 (Influence of alternative fuel properties and combustor operating conditions on the nvPM and gaseous émissions produced by a small-scale RQL combustor, Fuel, 315 (2022), Article 123045) ; et,

[1199] Saffaripour et al., 2019 (A review on the morphologicalproperties of non-volatile particulate matter emissions from aircraft turbine engines, Journal of Aerosol Science, 105467 (2019)); each incorporated here for reference.

[1200] It is understood that data and indices from the nvPM sampling and measurement system that have not been corrected for system losses do not provide an accurate representation of actual exhaust emission levels. Conversely, it is understood that data corrected for system losses and nvPM emission indices corrected for system losses provide an accurate representation of actual exhaust emission levels, which can be expressed, for example, as one or more system loss-corrected (SL) indices, as a number of nvPMs corrected for system losses (SL) (i.e., in number / kg of fuel), or as a mass in number of nvPMs corrected for system losses (SL) (i.e., in mg / kg).Thus, it is understood that the mass of nvPM produced by the gas turbine engine 10 refers to the data corrected for losses in the system (SL) in milligrams (mg) divided by the corresponding mass of fuel used by the engine 10 in kilograms (kg) unless otherwise stated.

[1201] The nvPM emission index, corrected for system losses or otherwise, can be defined at various operating phases of the gas turbine engine 10, for example, idle, maximum takeoff, climb, and approach. An emission index can further be defined according to the type of fuel supplied to the combustor device 16.

[1202] The following emission index parameters are defined for the gas turbine engine 10:

[1203] i) LLdern is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine 10 if it is operating at 7% of thrust available under given operating conditions. Operation at 7% of available thrust may correspond to operation in an idle operating phase of the gas turbine engine 10;

[1204] ii) E7maxTO is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine 10 if it operates at 100% of available thrust under given operating conditions. Operation at 100% of available thrust may correspond to operation in a maximum takeoff operating phase of the gas turbine engine 10;

[1205] iiij £7montée is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine 10 during operation at 85% of available thrust under given operating conditions. Operation at 85% of available thrust may correspond to operation in a climb phase of the gas turbine engine 10;

[1206] iv) £7approach is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine 10 when operating at 30% of available thrust under given operating conditions. Operation at 30% of available thrust may correspond to operation in an approach phase of the gas turbine engine 10.

[1207] Available thrust under given operating conditions (i.e., engine power setting) is defined as a percentage of the maximum rated engine thrust (Foe) as defined in the article. In other words, a percentage of "available thrust" refers to a percentage of maximum thrust, where maximum thrust is "100% of available thrust," and "given operating conditions" refers to predetermined operating conditions under which the maximum rated engine thrust, i.e., 100% of available thrust, is measured. The predetermined operating conditions may be those of the ISA at sea level where the reference absolute humidity is 0.00634 kg water / kg dry air. The predetermined operating conditions may be static at sea level.The predetermined operating conditions may include no customer demand and / or no power draw. The predetermined operating conditions may be daytime conditions. The predetermined operating conditions may be at approximately 60% relative humidity.

[1208] The nvPM emission indices defined above can further be defined according to the fuel supplied to the combustor device. The fuel-specific nvPM emission index values ​​are defined as follows:

[1209] i) Æ' / raienti.FF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine 10 if it operates at 7% of available thrust under given operating conditions and if a fuel supplied to the combustor device 16 is a fossil-based hydrocarbon fuel;

[1210] ii) £7maxTo,FF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine 10 if it is operating at 100% of available thrust under given operating conditions and if a fuel supplied to the combustor device 16 is a fossil-based hydrocarbon fuel;

[1211] iii) £7montée,FF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine 10 when operating at 85% of available thrust under given operating conditions and if a fuel supplied to the combustor device 16 is a fossil-based hydrocarbon fuel;

[1212] iv) ^ / approach,ff is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine 10 when operating at 30% of available thrust under given operating conditions and if a fuel supplied to the combustor device 16 is a fossil-based hydrocarbon fuel;

[1213] (v) E / raienti,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine 10 if it is operating at 7% of available thrust under given operating conditions and if a fuel supplied to the combustor device 16 includes a sustainable aviation fuel (SAF);

[1214] vi) Ê / maxTo.sAF is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine 10 if it is operating at 100% of available thrust under given operating conditions and if a fuel supplied to the combustor device 16 is a sustainable aviation fuel (SAF);

[1215] vii) ^ / ascent,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine 10 when operating at 85% of available thrust under given operating conditions and if a fuel supplied to the combustor device 16 includes a sustainable aviation fuel (SAF); and

[1216] viii) E / approach,saf is the nvPM emission index corrected for system losses in mg / kg of the gas turbine engine 10 when operating at 30% of available thrust under given operating conditions and if a fuel supplied to the combustor device 16 includes a sustainable aviation fuel (SAF).

[1217] Fuel flow

[1218] A fuel flow rate Wf of the gas turbine engine 10 is defined as the fuel flow rate to the fuel spray nozzles 403, 404 of the combustor device 16 (i.e., when the engine 10 is in operation). The flow rate Fuel flow is defined for operation at different percentages of available thrust under given operating conditions as defined above. TVf raiti is the fuel flow rate to the fuel spray nozzles 403, 404 in kg / s at 7% of available thrust under given operating conditions and may correspond to operation in an idle phase of the gas turbine engine 10. IVfrnaxTO is the fuel flow rate to the fuel spray nozzles 403, 404 in kg / s at 100% of available thrust under given operating conditions and may correspond to operation in a maximum takeoff phase of the gas turbine engine 10.Wfjnontée is defined as the mass flow rate of fuel supplied to the fuel spray nozzles 403, 404 in kg / s when the gas turbine engine 10 is operating at 85% of the available thrust under given operating conditions and can correspond to operation in a climb phase of the gas turbine engine 10. IVRapproche is the mass flow rate of fuel supplied to the fuel spray nozzles 403, 404 in kg / s when the gas turbine engine 10 is operating at 30% of the available thrust under given operating conditions and can correspond to operation in an approach phase of the gas turbine engine 10.

[1219] In any example defined or claimed anywhere herein, W frnaxTO may be in the range of 0.441 to 1.23 kg / s, preferably 0.496 to 1.13 kg / s, more preferably 0.551 to 1.03 kg / s. TVfrnaxTO may be in the range of 0.551 to 0.850 kg / s. IVfrnaxTO may be in the range of 0.551 to 0.750 kg / s.

[1220] In any example defined or claimed anywhere herein, IVfrnotée may be in the range of 0.369 to 1.01 kg / s, preferably from 0.415 to 0.923 kg / s, more preferably from 0.461 to 0.839 kg / s. TVfrnontée may be in the range of 0.461 to 0.650 kg / s. IVfrnontée may be in the range of 0.461 to 0.600 kg / s.

[1221] In any example defined or claimed anywhere herein, TVf approach may be in the range of 0.133 to 0.334 kg / s, preferably 0.149 to 0.306 kg / s, more preferably 0.166 to 0.278 kg / s. Wf approach may be in the range of 0.166 to 0.300 kg / s. Wf approach may be in the range of 0.166 to 0.250 kg / s.

[1222] In any example defined or claimed anywhere herein, IV / idle may be in the range of 0.0516 to 0.119 kg / s, preferably from 0.0581 to 0.109 kg / s, more preferably from 0.0645 to 0.0990 kg / s. TVf idle Pcut may be in the range of 0.0645 to 0.0850 kg / s. TVf idle may be in the range of 0.0645 to 0.0750 kg / s.

[1223] Engine thrust

[1224] The thrust of the gas turbine engine 10 is given by the symbol F and is defined for operation at different percentages of the available thrust under given operating conditions as defined above. Fmax is defined as the thrust of the gas turbine engine 10 at 100% of the available thrust under given operating conditions in kN. Fmax is defined as the thrust of the gas turbine engine 10 at 7% of the available thrust under given operating conditions in kN.

[1225] In any of the examples defined or claimed anywhere herein, ^maxTO may be ^ans 'a P'a8c from 54.1 kN to 177 kN and preferably in the range of 60.8 kN to 163 kN and more preferably in the range of 67.6 kN to 148 kN. The value of F^axTO corresponds to the maximum nominal thrust Foo. Alternatively, FmaxTO 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 more preferably in the range of 60 kN to 70 kN.

[1226] In any of the examples defined or claimed anywhere herein, Fslowed may be in the range of 3.78 kN to 12.4 kN and preferably in the range of 4.26 kN to 11.4 kN and more preferably in the range of 4.73 kN to 10.4 kN. Alternatively, Fslowed may be in the range of 3.5 kN to 6 kN and preferably in the range of 4 kN to 5.5 kN and preferably in the range of 4.2 k...

Claims

Demands

1. Gas turbine engine (10) for an aircraft, comprising: a rich-burning, fast-cooling, lean-burning (RQL) combustor device (16) having a number of fuel spray nozzles (403, 404) in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; and wherein: a lean-cruise nvPM-PMD emission index ratio is defined as: [Math 160] BPR ^ / cruise (lean) CSt defini COmme .[Math 161] ElmaxTo+E1 climb 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); the depleted cruise-PMD nvPM emission index ratio is greater than 0.2; and the gas turbine engine (10) is designed to supply fuel including a sustainable aviation fuel (SAF) to the fuel spray nozzles (403, 404).

2. Gas turbine engine (10) according to claim 1, wherein the depleted cruise nvPM-PMD emission index ratio is less than 0.18, preferably less than 0.16, and more preferably less than 0.

14.

3. 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 0.0426, of preference greater than or equal to 0.0479, and even more preferably greater than or equal to 0.0533.

4. Gas turbine engine (10) according to any one of the preceding claims, wherein the depleted cruise nvPM-PMD emission index ratio is in the range of 0.0426 to 0.118, preferably in the range of 0.0479 to 0.109 and more preferably in the range of 0.0533 to 0.0983.

5. 5 Gas turbine engine (10) according to any one of the preceding claims, wherein a rich cruise nvPM-PMD emission index ratio is defined as: [Math 162] Elcrt,isIm^[lchfA / E-iiiiaxTo where ; Rich Cruise BPR is defined as: [Math 163] EIBKmtée+EIapprc^ is the loss-corrected nvPM emission index in mg / kg of the gas turbine engine (10) when operating at about 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 about 30% of available thrust under the same operating conditions under which EImontée is calculated;^ / maxTo 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 the same operating conditions under which EImontée is calculated; and the rich cruise nvPM-PMD emission index ratio is less than 0.

07.

6. Gas turbine engine (10) according to claim 5, wherein the rich cruise nvPM-PMD emission index ratio is less than 0.065, preferably less than 0.06, and more preferably less than 0.

055.

7. Gas turbine engine (10) according to any one of claims 5 to 6, wherein the emission index ratio of nvPM in rich-PMD cruising is greater than or equal to 0.0128, preferably greater than or equal to 0.0144, and more preferably greater than or equal to 0.

016.

8. Gas turbine engine (10) according to any one of claims 5 to 7, wherein the rich cruise nvPM-PMD emission index ratio is in the range of 0.0128 to 0.0462, preferably in the range of 0.0144 to 0.0424 and more preferably in the range of 0.0160 to 0.0385.

9. Gas turbine engine according to any one of the preceding claims, wherein the fuel spray nozzles comprise one or more duplex nozzles (403) and one or more single-flow nozzles (404), and preferably the combustor device (16) comprises 10 to 14 duplex fuel spray nozzles and 4 to 8 single-flow fuel spray nozzles.

10. 10 Gas turbine engine according to claim 9, wherein the duplex fuel spray nozzles (403) are arranged in groups around the circumference of the combustor device (16) and optionally wherein the groups of duplex fuel spray nozzles (403) comprise at least two groups arranged diametrically opposite each other.

11. Gas turbine engine according to claim 10 wherein each group of duplex fuel spray nozzles (403) comprises 2 to 8 nozzles (403).

12. Gas turbine engine according to claim 9 or claim 10 or claim 11, wherein the combustor device (16) comprises one or more igniters (405) and the, or each, igniter (405) is arranged adjacent to one or more of the duplex fuel spray nozzles (403).

13. Gas turbine engine according to any one of the preceding claims, wherein the number of fuel spray nozzles per unit engine core size is in the range of 2.5 to 4.5, and more preferably in the range of 3 to 4.

14. Gas turbine engine according to any one of the preceding claims, wherein the fuel supplied to the combustor device (16) 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%.

15. 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 fuel spray nozzles (403, 404).

16. Method (1000) of operating a gas turbine engine (10), the gas turbine engine comprising: a rich-burning, fast-cooling, lean-burning (RQL) combustor device (16) having a number of fuel spray nozzles (403, 404) in the range of 14 to 22 or a number of fuel spray nozzles per unit engine core size in the range of 2 to 6; and wherein: a lean-cruise nvPM-PMD emission index ratio is defined as: [Math 164] BPR ^ / cruise ( lean ) ^St defini COmme .[Math 165] maxTo+EImontée EIm^m 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; ^montée 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-PMD nvPM emission index ratio is greater than 0.2; and the gas turbine engine (10) is designed to supply fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles (403, 404); in which the process includes supplying (1002) fuel comprising a sustainable aviation fuel to the fuel spray nozzles (404, 403).

17. 17 Method (1000) according to claim 16, wherein a rich cruise nvPM-PMD emission index ratio is defined as: [Math 166] BPR Cruise (rich) is defined as: [Math 167] EIim-emission-index + EIapprox 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; Σapprox 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 EIim-emission-index is calculated;^ / maxTo 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 the same operating conditions under which EImontée is calculated; and the rich cruise nvPM-PMD emission index ratio is less than 0.07.

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