AN AIRPLANE

Abstract

A gas turbine engine (10) for an aircraft. The gas turbine engine comprises the following: a rich-burn, rapid-cool, lean-burn (RQL) combustion chamber (16) having a number of fuel spray nozzles (403, 404) in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6. A thrust-nvPM emission index ratio is defined as follows: EI max A b F max A b EIL eer FL eer where: EI Leer The nvPM emission index of the gas turbine engine (10), corrected for system losses, in mg / kg at operation with approximately 7% of the available thrust under given operating conditions, is; EI maxAb the nvPM emission index of the gas turbine engine (10) in mg / kg, corrected for system losses, at operation with approximately 100% of the available thrust under the given operating conditions; F maxAbThe thrust of the gas turbine engine (10) at approximately 100% of the available thrust in kN is F Leer The thrust of the gas turbine engine (10) is at approximately 7% of the available thrust in kN; the thrust-nvPM emission index ratio is greater than 0.09; and the gas turbine engine (10) is configured to supply the fuel spray nozzles (403, 404) with a fuel comprising sustainable aviation fuel (SAF). A method for operating the gas turbine engine is also disclosed.

Description

AREA

[0001] The present disclosure relates to emissions of non-volatile particulate matter (nvPM) from gas turbine engines, in particular gas turbine engines for an aircraft. The present disclosure provides various methods for operating a gas turbine engine and gas turbine engines. Aspects of the present application relate to methods for operating gas turbine engines using a fuel comprising a sustainable aviation fuel (NFF) and to gas turbine engines configured for operation with a fuel comprising an NFF. BACKGROUND

[0002] The aviation industry is expected to see a shift towards the use of fuels that differ from the currently widely used conventional kerosene-based jet fuels.

[0003] The inventors have determined that the emissions of a gas turbine engine are significantly dependent on the fuel used, in particular that the amount of nvPM generated by the engine varies depending on the operating parameters and the type of fuel used. Therefore, there is a need to consider the fuel properties of these other fuels and to adapt methods for operating gas turbine engines accordingly. SUMMARY

[0004] According to a first aspect, a gas turbine engine is intended for an aircraft, which includes one or more of the following elements: a combustion chamber for rich burning, rapid cooling, lean burning (RQL combustion chamber) having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; and where: A first nvPM emission index ratio of idle to MAB (maximum take-off) can be defined as follows: EILeerEImaxAb where: EGG Leer the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 7% of the available thrust under given operating conditions; and EGG maxAb The nvPM emission index of the gas turbine engine, corrected for system losses, in mg / kg, is used when operating at approximately 100% of the available thrust under the given operating conditions; the first nvPM emission index ratio idle to MAB of the gas turbine engine is less than 0.8; and the gas turbine engine is configured to provide the fuel spray nozzles with a sustainable aviation fuel (SAF)-based fuel.

[0005] A reduced nvPM content in the exhaust gas of a gas turbine engine advantageously contributes to a reduction in the engine's undesirable emissions. For example, depending on the operating conditions, such a reduction in nvPM can lead to a reduction in soot deposits within and / or downstream of the combustion chamber and / or to an improvement in local air quality. Furthermore, in certain phases of an aircraft flight (in which contrail formation would otherwise be expected), a reduction in the nvPM content in the exhaust gas can lead to a reduced intensity of contrails and / or a shorter time to contrail dissipation. It has also been recognized that specific parts of the flight cycle in which the nvPM value is reduced (or most significantly reduced) can be targeted to achieve a desired outcome, for example, with regard to environmental impact.By way of example, a lower nvPM value under cruise flight conditions can particularly reduce the impact of radiative forcing of contrails. Also by way of example, a lower nvPM value under idle conditions can particularly improve local air quality at ground level in the engine operating area. Again by way of example, a lower nvPM value under MAB conditions can particularly reduce the maximum nvPM production rate during the flight cycle and / or improve air quality at ground level and / or in the engine operating area. These considerations can apply to all aspects of the disclosure.

[0006] It has been established that when certain types of fuel, such as sustainable aviation fuel, are burned, a number of parameters related to the operation of a gas turbine engine affect, or are a significant factor in, the configuration and arrangement of the engine's combustion chamber. Accordingly, one or more parameters of the following aspects may be advantageously considered when determining, for example, operating settings, the combustion chamber arrangement, and / or the combustion chamber configuration, in order to influence and / or optimize how this fuel is distributed, ignited, and / or burned in the gas turbine engine. These considerations can apply to all aspects of the disclosure.

[0007] The first nvPM emission index ratio idle to MAB of the gas turbine engine can be greater than zero.

[0008] The first nvPM emission index ratio idle to MAB can be less than 0.708 and preferably less than 0.649 and more preferably less than 0.59.

[0009] The first nvPM emission index ratio idle to MAB 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.

[0010] The first nvPM emission index ratio idle to MAB 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.

[0011] The first nvPM emission index ratio (idle to MAB) can 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, or 0.8, or within any range defined by any two of these values. For example, the first nvPM emission index ratio (idle to MAB) can be in a range between 0.01 and 0.2, 0.01 and 0.15, 0.01 and 0.07, or 0.01 and 0.05.

[0012] The first nvPM emission index ratio idle to MAB 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.

[0013] The first nvPM emission index ratio idle to MAB 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.

[0014] A second nvPM emission index ratio, idle to MAB, can be defined as follows: EILeer,NFKEImaxAb,NFKEILeer,FKEImaxAb,FK where: EGG Leer,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 7% of the available thrust under the given operating conditions or under other operating conditions, and if a fuel supplied to the combustion chamber includes a sustainable aviation fuel; EGG maxAb,NFK The nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, during operation at approximately 100% of available thrust under the same given operating conditions as those under which EI Leer,NFK is calculated, and, if a fuel supplied to the combustion chamber includes a sustainable aviation fuel, is; EGG Leer,FKThe system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 7% of available thrust under the same given operating conditions as those under which EI Leer,NFK is calculated, and, if one of the fuels supplied to the combustion chamber is a fossil-based hydrocarbon fuel, is; EGG maxAb,FK The nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, during operation at approximately 100% of available thrust under the same given operating conditions as those under which EI Leer,NFK is calculated, and, if one of the fuels supplied to the combustion chamber is a fossil-based hydrocarbon fuel, is; and where the second nvPM emission index ratio idle to MAB of the gas turbine engine can be less than 1.

[0015] The second nvPM emission index ratio, idle to MAB, can be greater than zero.

[0016] The second nvPM emission index ratio, idle to MAB, can 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.

[0017] The second nvPM emission index ratio, idle to MAB, can 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.

[0018] The second nvPM emission index ratio, idle to MAB, can 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.

[0019] The second nvPM emission index ratio, idle to MAB, can be greater than or equal to 0.118, and preferably greater than or equal to 0.133, and more preferably greater than or equal to 0.148.

[0020] The second nvPM emission index ratio, idle to MAB, can be in the range of 0.118 to 0.178, and preferably in the range of 0.133 to 0.164, and more preferably in the range of 0.148 to 0.149.

[0021] The second nvPM emission index ratio, idle to MAB, can 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 range defined by any two of these values. For example, the second nvPM emission index ratio, idle to MAB, can be in a range between 0.25 and 0.4 or 0.3 and 0.35.

[0022] According to a second aspect, a gas turbine engine is intended for an aircraft that includes one or more of the following elements: a rich-burn, rapid-cooling, lean-burn (RQL) combustion chamber having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; and wherein: A second nvPM emission index ratio, idle to MAB, can be defined as follows: EILeer,NFKEImaxAb,NFKEILeer,FKEImaxAb,FK where: EGG Leer,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 7% of the available thrust under given operating conditions and if a fuel supplied to the combustion chamber includes a sustainable aviation fuel; EGG maxAb,NFKThe nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, during operation at approximately 100% of available thrust under the same given operating conditions as those under which EI Leer,NFK is calculated, and, if a fuel supplied to the combustion chamber includes a sustainable aviation fuel, is; EGG Leer,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 7% of available thrust under the same given operating conditions as those under which EI Leer,NFK is calculated, and, if one of the fuels supplied to the combustion chamber is a fossil-based hydrocarbon fuel, is; EGG maxAb,FKThe nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, during operation at approximately 100% of available thrust under the same given operating conditions as those under which EI Leer,NFK is calculated, and, if one of the fuels supplied to the combustion chamber is a fossil-based hydrocarbon fuel, is; the second nvPM emission index ratio idle to MAB of the gas turbine engine is less than 1; and the gas turbine engine is configured to provide the fuel spray nozzles with a sustainable aviation fuel (SAF)-based fuel.

[0023] The second nvPM emission index ratio idle to MAB defined in the second aspect can be defined in the context of the first aspect as above.

[0024] According to a third aspect, a method for operating the gas turbine engine of the first aspect or the second aspect is provided, wherein the method includes supplying fuel comprising a sustainable aviation fuel to the fuel spray nozzles.

[0025] According to a fourth aspect, a method for operating a gas turbine engine is provided, wherein the gas turbine engine comprises the following: a rich-burn, rapid-cooling, lean-burn (RQL) combustion chamber having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; and wherein: A first nvPM emission index ratio of idle to MAB (maximum take-off) can be defined as follows: EILeerEImaxAb where: EGG Leerthe nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 7% of the available thrust under given operating conditions; and EGG maxAb The nvPM emission index of the gas turbine engine, corrected for system losses, in mg / kg, is used when operating at approximately 100% of the available thrust under the given operating conditions; the first nvPM emission index ratio idle to MAB of the gas turbine engine is less than 0.8; and The process involves supplying a sustainable aviation fuel to the fuel spray nozzles.

[0026] The first nvPM emission index ratio idle to MAB can be defined as above in connection with the first aspect.

[0027] A second nvPM emission index ratio, idle to MAB, can be defined as follows: EILeer,NFKEImaxAb,NFKEILeer,FKEImaxAb,FK where: EGG Leer,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 7% of the available thrust under the given operating conditions or under other operating conditions, and if a fuel supplied to the combustion chamber includes a sustainable aviation fuel; EGG maxAb,NFK The nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, during operation at approximately 100% of available thrust under the same given operating conditions as those under which EI Leer,NFK is calculated, and, if a fuel supplied to the combustion chamber includes a sustainable aviation fuel, is; EGG Leer,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 7% of available thrust under the same given operating conditions as those under which EILeer,NFK is calculated, and, if one of the fuels supplied to the combustion chamber is a fossil-based hydrocarbon fuel, is; EGG maxAb,FK The nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, during operation at approximately 100% of available thrust under the same given operating conditions as those under which EI Leer,NFK is calculated, and, if one of the fuels supplied to the combustion chamber is a fossil-based hydrocarbon fuel, is; and where the second nvPM emission index ratio idle to MAB of the gas turbine engine may be less than 1.

[0028] The second nvPM emission index ratio, idle to MAB, can be defined as above in connection with the first aspect.

[0029] According to a fifth aspect, a method for operating a gas turbine engine is provided, wherein the gas turbine engine comprises: a rich-burn, rapid-cooling, lean-burn (RQL) combustion chamber having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; and wherein: A second nvPM emission index ratio, idle to MAB, can be defined as follows: EILeer,NFKEImaxAb,NFKEILeer,FKEImaxAb,FK where: EGG Leer,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 7% of the available thrust under given operating conditions and if a fuel supplied to the combustion chamber includes a sustainable aviation fuel; EGG maxAb,NFKThe nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, during operation at approximately 100% of available thrust under the same given operating conditions as those under which EI Leer,NFK is calculated, and, if a fuel supplied to the combustion chamber includes a sustainable aviation fuel, is; EGG Leer,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 7% of available thrust under the same given operating conditions as those under which EI Leer,NFK is calculated, and, if one of the fuels supplied to the combustion chamber is a fossil-based hydrocarbon fuel, is; EGG maxAb,FKThe nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, during operation at approximately 100% of available thrust under the same given operating conditions as those under which EI Leer,NFK is calculated, and, if one of the fuels supplied to the combustion chamber is a fossil-based hydrocarbon fuel, is; the second nvPM emission index ratio idle to MAB of the gas turbine engine is less than 1; and The process involves supplying a sustainable aviation fuel to the fuel spray nozzles.

[0030] The second nvPM emission index ratio, idle to MAB, can be defined as above in connection with the first aspect.

[0031] According to a sixth aspect, a gas turbine engine is intended for an aircraft that includes one or more of the following elements: a rich-burn, rapid-cooling, lean-burn (RQL) combustion chamber having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; and wherein: A fuel flow-nvPM emission index ratio can be defined as follows: EILempty×WK,EmptyEImaxAb×WK,maxAb where: EGG Leer The nvPM emission index of the gas turbine engine, corrected for system losses, in mg / kg, is used when operating at approximately 7% of the available thrust under given operating conditions; EGG maxAb The nvPM emission index of the gas turbine engine, corrected for system losses, in mg / kg, is used when operating at approximately 100% of the available thrust under the given operating conditions; W K,LeerThe fuel flow rate to the fuel spray nozzles in kg / s is approximately 7% of the available thrust under the given operating conditions; and W K,maxAb the fuel flow rate to the fuel spray nozzles in kg / s at approximately 100% of the available thrust under the given operating conditions; the fuel flow-nvPM emission index ratio of the gas turbine engine is less than 0.08; and the gas turbine engine is configured to provide the fuel spray nozzles with a sustainable aviation fuel (SAF)-based fuel.

[0032] The fuel flow nvPM emission index ratio can be less than 0.0798 and preferably less than 0.0731 and more preferably less than 0.0665.

[0033] The fuel flow nvPM emission index ratio can 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.

[0034] The fuel flow nvPM emission index ratio can 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.

[0035] The fuel flow nvPM emission index ratio can 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 by any two of these values.

[0036] The fuel flow-nvPM emission index ratio can be greater than zero. The fuel flow-nvPM emission index ratio can 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.

[0037] The fuel flow nvPM emission index ratio can be in the range of 0.000993 to 0.0119 and preferably in the range of 0.00111 to 0.0109 and more preferably in the range of 0.00124 to 0.00986.

[0038] The fuel flow-nvPM emission index ratio can 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 range defined by any two of these values. For example, the fuel flow-nvPM emission index ratio can be in a range between 0.0009 and 0.021, 0.009 and 0.019, 0.009 and 0.007, or 0.009 and 0.006.

[0039] W K,Leer It can range from 0.0516 to 0.119 kg / s. W K,Leer It can range from 0.0581 to 0.109 kg / s. W K,Leer It can range from 0.0645 to 0.0990 kg / s. W K,Leer It can range from 0.0645 to 0.0850 kg / s. W K,Leer can range from 0.0645 to 0.0750 kg / s.

[0040] W K,maxAb It can range from 0.441 to 1.23 kg / s. W K,maxAb It can range from 0.496 to 1.13 kg / s. W K,maxAbIt can range from 0.551 to 1.03 kg / s. W K,maxAb It can range from 0.551 to 0.850 kg / s. W K,maxAb can range from 0.551 to 0.750 kg / s.

[0041] According to a seventh aspect, a method for operating the gas turbine engine of the sixth aspect is provided, the method comprising the provision of fuel comprising a sustainable aviation fuel to the fuel spray nozzles.

[0042] According to an eighth aspect, a method for operating a gas turbine engine is provided, wherein the gas turbine engine comprises the following: a rich-burn, rapid-cooling, lean-burn (RQL) combustion chamber having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; and wherein: A fuel flow-nvPM emission index ratio can be defined as follows: EILempty×WK,EmptyEImaxAb×WK,maxAb where: EGG Leer the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 7% of the available thrust under given operating conditions; and EGG maxAb The nvPM emission index of the gas turbine engine, corrected for system losses, in mg / kg, is used when operating at approximately 100% of the available thrust under the given operating conditions; W K,Leer The fuel flow rate to the fuel spray nozzles in kg / s is approximately 7% of the available thrust under the given operating conditions; and W K,maxAb the fuel flow rate to the fuel spray nozzles in kg / s at approximately 100% of the available thrust under the given operating conditions; the fuel flow-nvPM emission index ratio of the gas turbine engine is less than 0.08; and The process involves supplying a sustainable aviation fuel to the fuel spray nozzles.

[0043] Each of the fuel flow nvPM emission index ratios, W K,Leer and W K,maxAb can be defined as above in connection with the sixth aspect. According to a ninth aspect, a gas turbine engine for an aircraft is intended to include one or more of the following elements: a rich-burn, rapid-cooling, lean-burn (RQL) combustion chamber having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; and wherein: A thrust-nvPM emission index ratio can be defined as follows: EImaxAbFmaxAbEILeerFLeer where: EGG Leer The nvPM emission index of the gas turbine engine, corrected for system losses, in mg / kg, is used when operating at approximately 7% of the available thrust under given operating conditions; EGG maxAb The nvPM emission index of the gas turbine engine, corrected for system losses, in mg / kg, is used when operating at approximately 100% of the available thrust under the given operating conditions; F maxAb the thrust of the gas turbine engine is approximately 100% of the available thrust in kN under the given operating conditions; and F Leer The thrust of the gas turbine engine is approximately 7% of the available thrust in kN under the given operating conditions; the thrust-nvPM emission index ratio is greater than 0.09; and the gas turbine engine is configured to provide the fuel spray nozzles with a sustainable aviation fuel (SAF)-based fuel.

[0044] The thrust-nvPM emission index ratio can be greater than 0.0949 and preferably greater than 0.106 and more preferably greater than 0.118.

[0045] The thrust-nvPM emission index ratio can 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.

[0046] The thrust-nvPM emission index ratio can 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.

[0047] The thrust-nvPM emission index ratio can 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.

[0048] The thrust-nvPM emission index ratio can be in the range of 0.640 to 6.53 and preferably in the range of 0.720 to 5.98 and more preferably in the range of 0.800 to 5.44.

[0049] The thrust-to-nvPM emission index ratio can 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 range defined by any two of these values. For example, the thrust-to-nvPM emission index ratio can be in a range between 0.3 and 4.5, 0.4 and 4, 1 and 4.5, or 1.4 and 4.

[0050] F maxAb can be in the range of 54.1 kN 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, F can be maxAb 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.

[0051] F Leer can 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. F Leerand preferably in the range of 4.73 kN to 10.4 kN. Alternatively, F Leer 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.

[0052] According to a tenth aspect, a method for operating the gas turbine engine of the ninth aspect is provided, the method comprising the provision of fuel comprising a sustainable aviation fuel to the fuel spray nozzles.

[0053] According to an eleventh aspect, a method for operating a gas turbine engine is provided, wherein the gas turbine engine comprises the following: a rich-burn, rapid-cooling, lean-burn (RQL) combustion chamber having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; and wherein: A thrust-nvPM emission index ratio can be defined as follows: EImaxAbFmaxAbEILeerFLeer where: EGG Leer the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 7% of the available thrust under given operating conditions; and EGG maxAb The nvPM emission index of the gas turbine engine, corrected for system losses, in mg / kg, is used when operating at approximately 100% of the available thrust under the given operating conditions; F maxAbThe thrust of the gas turbine engine at approximately 100% of the available thrust in kN under the given operating conditions is, F Leer The thrust of the gas turbine engine is approximately 7% of the available thrust in kN under the given operating conditions; and the thrust-nvPM emission index ratio is greater than 0.09; and The process involves supplying a sustainable aviation fuel to the fuel spray nozzles.

[0054] Each of the thrust-nvPM emission index ratios, F maxAb and F Leer can be defined as in connection with the ninth aspect.

[0055] According to a twelfth aspect, a gas turbine engine is intended for an aircraft that includes one or more of the following elements: a rich-burn, rapid-cooling, lean-burn (RQL) combustion chamber having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; and wherein: An nvPM emission index ratio for lean-burn cruise flight to MAB can be defined as follows: EIReise(lean)EImaxAbBPV where: EGG Reise(mager) which can be defined as follows: EImaxAb+EISteig2 EGG maxAb The nvPM emission index of the gas turbine engine, corrected for system losses, in mg / kg, is used when operating at approximately 100% of the available thrust under the given operating conditions; EGG Steig the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 85% of the available thrust under the given operating conditions; and BPV is the bypass ratio of the gas turbine engine; the nvPM emission index ratio of lean-burn cruise to MAB is less than 0.2; and the gas turbine engine is configured to provide the fuel spray nozzles with a sustainable aviation fuel (SAF)-based fuel.

[0056] The nvPM emission index ratio of lean-burn cruise to MAB can be less than 0.18, preferably less than 0.16 and more preferably less than 0.14.

[0057] The nvPM emission index ratio of lean-burn cruise to MAB can be less than 0.135, preferably less than 0.124 and more preferably less than 0.113.

[0058] The nvPM emission index ratio of lean-burn cruise to MAB can 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.

[0059] The nvPM emission index ratio for lean-burn cruise flight to MAB can be greater than zero. The nvPM emission index ratio for lean-burn cruise flight to MAB can 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.

[0060] The nvPM emission index ratio for lean-burn cruise to MAB can 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.

[0061] The nvPM emission index ratio of lean-burn cruise to MAB can 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 lie within any range defined by any two of these values.

[0062] The BPV can be in the range of 6.66 to 15.3 and higher, 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 BPV can be in the range of 3.5 to 6.5 and higher, preferably in the range of 4 to 6, and even more preferably in the range of either 4 to 5 or 5 to 6.

[0063] An nvPM emission index ratio for cruise flight with rich combustion to MAB can be defined as follows: EIReise(fett)EImaxAbBPV where: EGG Reise(fett) which can be defined as follows: ICE climb + ICE approach 2 EGG Steig the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 85% of the available thrust under the given operating conditions or under other operating conditions; EGG AnflugThe system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EI Steig is calculated; and EGG maxAb The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI Steig is calculated; and where the nvPM emission index ratio of cruise flight with rich combustion to MAB is less than 0.07.

[0064] The nvPM emission index ratio of cruise flight with rich combustion to MAB can be less than 0.065, preferably less than 0.06 and more preferably less than 0.055.

[0065] The nvPM emission index ratio of cruise flight with rich combustion to MAB can be less than 0.0635, preferably less than 0.0582 and more preferably less than 0.0529.

[0066] The nvPM emission index ratio of cruise flight with rich combustion to MAB can 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.

[0067] The nvPM emission index ratio for cruise flight with rich combustion to MAB can be greater than zero. The nvPM emission index ratio for cruise flight with rich combustion to MAB can 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.

[0068] The nvPM emission index ratio of cruise flight with rich combustion to MAB can 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.

[0069] The nvPM emission index ratio of rich-burn cruise flight to MAB can 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 lie within any range defined by any two of these values.

[0070] The nvPM emission index ratio of rich-burn cruise flight to MAB can 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 lie within any range defined by any two of these values.

[0071] According to a thirteenth aspect, a gas turbine engine is intended for an aircraft that includes one or more of the following elements: a rich-burn, rapid-cooling, lean-burn (RQL) combustion chamber having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; and wherein: An nvPM emission index ratio for cruise flight with rich combustion to MAB can be defined as follows: EIReise(fett)EImaxAbBPV where: EI Reise(fett) which can be defined as follows: ICE climb + ICE approach 2 EGG Steig The nvPM emission index of the gas turbine engine, corrected for system losses, in mg / kg, is used when operating at approximately 85% of the available thrust under the given operating conditions; EGG Anflug The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EI Steigis calculated; EGG maxAb The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI Steig is calculated; BPV is the bypass ratio of the gas turbine engine; the nvPM emission index ratio of cruise flight with rich combustion to MAB is less than 0.07; and the gas turbine engine is configured to provide the fuel spray nozzles with a sustainable aviation fuel (SAF)-based fuel.

[0072] The nvPM emission index ratio of cruise flight with rich combustion to MAB and / or the BPV can be defined as above in connection with the twelfth aspect.

[0073] According to a fourteenth aspect, a method for operating the gas turbine engine of the twelfth or thirteenth aspect is provided, the method comprising the provision of fuel comprising a sustainable aviation fuel to the fuel spray nozzles.

[0074] According to a fifteenth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising the following: a rich-burn, rapid-cooling, lean-burn (RQL) combustion chamber having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; and wherein: An nvPM emission index ratio for lean-burn cruise flight to MAB can be defined as follows: EIReise(lean)EImaxAbBPV where: EGG Reise(mager) which can be defined as follows: EImaxAb+EISteig2 EGG maxAb The nvPM emission index of the gas turbine engine, corrected for system losses, in mg / kg, is used when operating at approximately 100% of the available thrust under the given operating conditions; EGG Steig the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 85% of the available thrust under the given operating conditions; and BPV is the bypass ratio of the gas turbine engine; the nvPM emission index ratio of lean-burn cruise to MAB is less than 0.2; and The process involves supplying a sustainable aviation fuel to the fuel spray nozzles.

[0075] The nvPM emission index ratio of lean-burn cruise to MAB and / or the BPV can be defined as above in connection with the twelfth aspect.

[0076] An nvPM emission index ratio for cruise flight with rich combustion to MAB can be defined as follows: EIReise(fett)EImaxAbBPV where: EGG Reise(fett) which can be defined as follows: ICE climb + ICE approach 2 EGG Steig the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 85% of the available thrust under the given operating conditions or under other operating conditions; EGG Anflug The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EI Steig is calculated; EGG maxAbThe system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI Steig is calculated; and where the nvPM emission index ratio of cruise flight with rich combustion to MAB can be less than 0.07.

[0077] The nvPM emission index ratio of cruise flight with rich combustion to MAB and / or the BPV can be defined as above in connection with the twelfth aspect.

[0078] According to a sixteenth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising the following: a rich-burn, rapid-cooling, lean-burn (RQL) combustion chamber having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; and wherein: An nvPM emission index ratio for cruise flight with rich combustion to MAB can be defined as follows: EIReise(fett)EImaxAbBPV where: EGG Reise(fett) which can be defined as follows: ICE climb + ICE approach 2 EGG Steig The nvPM emission index of the gas turbine engine, corrected for system losses, in mg / kg, is used when operating at approximately 85% of the available thrust under the given operating conditions; EGG Anflug The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EISteig is calculated; EGG maxAb The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI Steig is calculated; BPV is the bypass ratio of the gas turbine engine; the nvPM emission index ratio of cruise flight with rich combustion to MAB can be less than 0.07; and The process involves supplying a sustainable aviation fuel to the fuel spray nozzles.

[0079] The nvPM emission index ratio of cruise flight with rich combustion to MAB and / or the BPV can be defined as above in connection with the twelfth aspect.

[0080] According to a seventeenth aspect, a gas turbine engine is intended for an aircraft, comprising one or more of the following elements: a rich-burn, rapid-cooling, lean-burn (RQL) combustion chamber having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; and wherein: A MAB-nvPM emission index ratio can be defined as follows: EImaxAb,NFKEImaxAb,FK where: EGG maxAb,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at approximately 100% of available thrust under given operating conditions, where the fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel (SAF); and EGG maxAb,FKthe nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at approximately 100% of available thrust under the given operating conditions, assuming the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; the MAB-nvPM emission index ratio of the gas turbine engine is less than 1; and the gas turbine engine is configured to provide the fuel spray nozzles with an NFK-rich fuel.

[0081] The MAB-nvPM emission index ratio can be greater than zero.

[0082] The MAB-nvPM emission index ratio can be less than or equal to 0.93, can preferably be less than or equal to 0.86, and can more preferably be less than or equal to 0.79.

[0083] The MAB-nvPM emission index ratio can be less than or equal to 0.776, can preferably be less than or equal to 0.711, and can more preferably be less than or equal to 0.646.

[0084] The MAB-nvPM emission index ratio can be greater than or equal to 0.15, can preferably be greater than or equal to 0.3, and can more preferably be greater than or equal to 0.45.

[0085] The MAB-nvPM emission index ratio can be greater than or equal to 0.516, can preferably be greater than or equal to 0.581, and can more preferably be greater than or equal to 0.645.

[0086] The MAB-nvPM emission index ratio can be in the range of 0.516 to 0.776, and preferably in the range of 0.581 to 0.711, and more preferably in the range of 0.645 to 0.646.

[0087] The MAB-nvPM emission index ratio can be less than 1, 0.95, 0.9, 0.85, 0.8, 0.75, 0.7, 0.65, 0.6, 0.55, 0.5, 0.45, 0.4, 0.35, 0.3, 0.25, 0.2, 0.15, 0.1 or 0.05, or within any range defined by any two of these values.

[0088] The MAB-nvPM emission index ratio can be 0.5, 0.51, 0.52, 0.53, 0.54, 0.55, 0.56, 0.57, 0.58, 0.59, 0.6, 0.61, 0.62, 0.63, 0.64, 0.65, 0.66, 0.67, 0.68, 0.69, 0.7, 0.71, 0.72, 0.73, 0.74, 0.75, 0.76, 0.77, 0.78, 0.79 or 0.8, or within any range defined by any two of these values.

[0089] A climb-nvPM emission index ratio can be defined as follows: Ice cream dough, NFK Ice cream dough, FK where: EGG Steig,NFKthe nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 85% of available thrust under the given operating conditions, or under other given operating conditions where a fuel supplied to the fuel spray nozzles includes sustainable aviation fuel (SAF); and EGG Steig,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same given operating conditions as those under which EI Steig,NFK is calculated, and, if the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel, is; and where the climb-nvPM emission index ratio of the gas turbine engine can be less than 1.

[0090] The climb-nvPM emission index ratio can be greater than zero.

[0091] The climb-nvPM emission index ratio can be less than or equal to 0.9, can preferably be less than or equal to 0.75, and can more preferably be less than or equal to 0.6.

[0092] The climb-nvPM emission index ratio can be less than or equal to 0.57, can preferably be less than or equal to 0.523, and can more preferably be less than or equal to 0.475.

[0093] The climb-nvPM emission index ratio can be greater than or equal to 0.1, can preferably be greater than or equal to 0.2, and can more preferably be greater than or equal to 0.3.

[0094] The climb-nvPM emission index ratio can be greater than or equal to 0.379, can preferably be greater than or equal to 0.427, and can more preferably be greater than or equal to 0.474.

[0095] The climb-nvPM emission index ratio can be in the range of 0.379 to 0.570, and preferably in the range of 0.427 to 0.523, and more preferably in the range of 0.474 to 0.475.

[0096] The climb-nvPM emission index ratio can be less than 1, 0.95, 0.9, 0.85, 0.8, 0.75, 0.7, 0.65, 0.6, 0.55, 0.5, 0.45, 0.4, 0.35, 0.3, 0.25, 0.2, 0.15, 0.1 or 0.05, or within any range defined by any two of these values.

[0097] The climb-nvPM emission index ratio can be 0.4, 0.41, 0.42, 0.43, 0.44, 0.45, 0.46, 0.47, 0.48, 0.49, 0.5, 0.51, 0.52, 0.53, 0.54, 0.55, 0.56, 0.57, 0.58, 0.59, 0.6 or within any range defined by any two of these values.

[0098] An approach-nvPM emission index ratio can be defined as follows: EI approach, NFKE approach, FK where: EGG Anflug,NFKthe nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 30% of available thrust under the given operating conditions, or under other given operating conditions where a fuel supplied to the fuel spray nozzles includes sustainable aviation fuel (SAF); and EGG Anflug,FK The nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, during operation at approximately 30% of available thrust under the same given operating conditions as those under which EI Anflug,NFK is calculated, and if the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel, then; and where the approach nvPM emission index ratio of the gas turbine engine can be less than 1.

[0099] The approach-nvPM emission index ratio can be greater than zero.

[0100] The approach nvPM emission index ratio can be less than or equal to 0.8, can preferably be less than or equal to 0.5, and can more preferably be less than or equal to 0.2.

[0101] The approach nvPM emission index ratio can be less than or equal to 0.185, can preferably be less than or equal to 0.169, and can more preferably be less than or equal to 0.154.

[0102] The approach-nvPM emission index ratio can be greater than or equal to 0.03, can preferably be greater than or equal to 0.06, and can more preferably be greater than or equal to 0.09.

[0103] The approach-nvPM emission index ratio can be greater than or equal to 0.122, can preferably be greater than or equal to 0.138, and can more preferably be greater than or equal to 0.153.

[0104] The approach-nvPM emission index ratio can be in the range of 0.122 to 0.185, and preferably in the range of 0.138 to 0.169, and more preferably in the range of 0.153 to 0.154.

[0105] The approach-nvPM emission index ratio can be less than 1, 0.95, 0.9, 0.85, 0.8, 0.75, 0.7, 0.65, 0.6, 0.55, 0.5, 0.45, 0.4, 0.35, 0.3, 0.25, 0.2, 0.15, 0.1 or 0.05, or within any range defined by any two of these values.

[0106] The approach-nvPM emission index ratio can 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 by any two of these values.

[0107] An idle-nvPM emission index ratio can be defined as follows: EILeer,NFKEILeer,FK where: EGG Leer,NFKthe nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 7% of available thrust under the given operating conditions or under other given operating conditions, and if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel (SAF); and EGG Leer,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 7% of available thrust under the same given operating conditions as those under which EI Leer,NFK is calculated, and, if the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel, is; and where the idle nvPM emission index ratio of the gas turbine engine may be less than 1.

[0108] The idle-nvPM emission index ratio can be greater than zero.

[0109] The idle nvPM emission index ratio can be less than or equal to 0.8, can preferably be less than or equal to 0.5, and can more preferably be less than or equal to 0.2.

[0110] The idle nvPM emission index ratio can be less than or equal to 0.115, can preferably be less than or equal to 0.106, and can more preferably be less than or equal to 0.0959.

[0111] The idle-nvPM emission index ratio can be greater than or equal to 0.02, can preferably be greater than or equal to 0.04, and can more preferably be greater than or equal to 0.06.

[0112] The idle nvPM emission index ratio can be greater than or equal to 0.0766, can preferably be greater than or equal to 0.0862, and can more preferably be greater than or equal to 0.0958.

[0113] The idle-nvPM emission index ratio can be in the range of 0.0766 to 0.115, and preferably in the range of 0.0862 to 0.106, and more preferably in the range of 0.0958 to 0.0959.

[0114] The idle-nvPM emission index ratio can be less than 1, 0.95, 0.9, 0.85, 0.8, 0.75, 0.7, 0.65, 0.6, 0.55, 0.5, 0.45, 0.4, 0.35, 0.3, 0.25, 0.2, 0.15, 0.1 or 0.05, or within any range defined by any two of these values.

[0115] The idle-nvPM emission index ratio can 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 range defined by any two of these values.

[0116] According to an eighteenth aspect, a gas turbine engine is intended for an aircraft that includes one or more of the following elements: a rich-burn, rapid-cooling, lean-burn (RQL) combustion chamber having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; and wherein: A climb-to-emissions index ratio can be defined as follows: Ice cream dough, NFK Ice cream dough, FK where: EGG Steig,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 85% of available thrust under given operating conditions, where the fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel (SAF); and EGG Steig,FKThe system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same given operating conditions as those under which EI Steig,NFK is calculated, and, if the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel, is; the climb-to-nvPM emission index ratio of the gas turbine engine is less than 1; and the gas turbine engine is configured to provide the fuel spray nozzles with an NFK-rich fuel.

[0117] The climb-nvPM emission index ratio can be defined as in connection with the seventeenth aspect.

[0118] According to a nineteenth aspect, a gas turbine engine is intended for an aircraft that includes one or more of the following elements: a rich-burn, rapid-cooling, lean-burn (RQL) combustion chamber having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; and wherein: An approach-nvPM emission index ratio can be defined as follows: EI approach, NFKE approach, FK where: EGG Anflug,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 30% of available thrust under given operating conditions, when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel (SAF); and EGG Anflug,FKThe nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, during operation at approximately 30% of available thrust under the same given operating conditions as those under which EI Anflug,NFK is calculated, and if the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel, it is; the approach-nvPM emission index ratio of the gas turbine engine is less than 1; and the gas turbine engine is configured to provide the fuel spray nozzles with an NFK-rich fuel.

[0119] The approach-nvPM emission index ratio can be defined as above in connection with the seventeenth aspect.

[0120] According to a twentieth aspect, a gas turbine engine is intended for an aircraft that includes one or more of the following elements: a rich-burn, rapid-cooling, lean-burn (RQL) combustion chamber having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; and wherein: An idle-nvPM emission index ratio can be defined as follows: EILeer,NFKEILeer,FK where: EGG Leer,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with 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); and EGG Leer,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 7% of available thrust under the same given operating conditions as those under which EILeer,NFK is calculated, and, if the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel, is; the idle nvPM emission index ratio of the gas turbine engine is less than 1; and the gas turbine engine is configured to provide the fuel spray nozzles with an NFK-rich fuel.

[0121] The idle-nvPM emission index ratio can be defined as above in connection with the seventeenth aspect.

[0122] According to a twenty-first aspect, a method for operating the gas turbine engine of any one of the seventeenth, eighteenth, nineteenth or twentieth aspects is provided, the method comprising supplying fuel comprising a sustainable aviation fuel to the fuel spray nozzles.

[0123] According to a twenty-second aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising the following: a rich-burn, rapid-cooling, lean-burn (RQL) combustion chamber having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; and wherein: A MAB-nvPM emission index ratio can be defined as follows: EImaxAb,NFKEImaxAb,FK where: EGG maxAb,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at 100% of available thrust under given operating conditions, where the fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel (SAF); and EGG maxAb,FKthe nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at 100% of available thrust under the given operating conditions, when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; the MAB-nvPM emission index ratio of the gas turbine engine is less than 1; and The process involves supplying a sustainable aviation fuel to the fuel spray nozzles.

[0124] The MAB-nvPM emission index ratio can be defined as in connection with the seventeenth aspect.

[0125] A climb-nvPM emission index ratio can be defined as follows: Ice cream dough, NFK Ice cream dough, FK where: EGG Steig,NFKthe nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 85% of available thrust under the given operating conditions, or under other given operating conditions where a fuel supplied to the fuel spray nozzles includes sustainable aviation fuel (SAF); and EGG Steig,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same given operating conditions as those under which EI Steig,NFK is determined, and if the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel, is; and wherein The climb-nvPM emission index ratio of the gas turbine engine may be less than 1.

[0126] The climb-nvPM emission index ratio can be defined as in connection with the seventeenth aspect.

[0127] An approach-nvPM emission index ratio can be defined as follows: EI approach, NFKE approach, FK where: EGG Anflug,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 30% of available thrust under the given operating conditions, or under other given operating conditions where a fuel supplied to the fuel spray nozzles includes sustainable aviation fuel (SAF); and EGG Anflug,FK The nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, during operation at approximately 30% of available thrust under the same given operating conditions as those under which EI Anflug,NFKis determined, and if the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel, is; and wherein The approach nvPM emission index ratio of the gas turbine engine may be less than 1.

[0128] The approach-nvPM emission index ratio can be defined as in connection with the seventeenth aspect.

[0129] An idle-nvPM emission index ratio can be defined as follows: EILeer,NFKEILeer,FK where: EGG Leer,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 7% of available thrust under the given operating conditions or under other given operating conditions, and if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel (SAF); and EGG Leer,FKThe system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 7% of available thrust under the same given operating conditions as those under which EI Leer,NFK is determined, and, if the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel, is; and where the idle nvPM emission index ratio of the gas turbine engine may be less than 1.

[0130] The idle-nvPM emission index ratio can be defined as in connection with the seventeenth aspect.

[0131] According to a twenty-third aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising the following: a rich-burn, rapid-cooling, lean-burn (RQL) combustion chamber having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; and wherein: A climb-to-emissions index ratio can be defined as follows: Ice cream dough, NFK Ice cream dough, FK where: EGG Steig,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 85% of available thrust under given operating conditions, where the fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel (SAF); and EGG Steig,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same given operating conditions as those under which EI Steig,NFKis determined, and if the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel, it is; the climb-to-nvPM emission index ratio of the gas turbine engine may be less than 1; and The process involves supplying a sustainable aviation fuel to the fuel spray nozzles.

[0132] The climb-nvPM emission index ratio can be defined as in connection with the seventeenth aspect.

[0133] According to a twenty-fourth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising the following: a rich-burn, rapid-cooling, lean-burn (RQL) combustion chamber having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; and wherein: An approach-nvPM emission index ratio can be defined as follows: EI approach, NFKE approach, FK where: EGG Anflug,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 30% of available thrust under given operating conditions, when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel (SAF); and EGG Anflug,FK The nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, during operation at approximately 30% of available thrust under the same given operating conditions as those under which EIAnflug,NFK is determined, and if the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel, then; and the approach-nvPM emission index ratio of the gas turbine engine may be less than 1; and The process involves supplying a sustainable aviation fuel to the fuel spray nozzles.

[0134] The approach-nvPM emission index ratio can be defined as in connection with the seventeenth aspect.

[0135] According to a twenty-fifth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising the following: a rich-burn, rapid-cooling, lean-burn (RQL) combustion chamber having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; and wherein: An idle-nvPM emission index ratio can be defined as follows: EILeer,NFKEILeer,FK where: EGG Leer,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with 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); and EGG Leer,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 7% of available thrust under the same given operating conditions as those under which EILeer,NFK is determined, and, if the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel, is; and the idle nvPM emission index ratio of the gas turbine engine can be less than 1; and The process involves supplying a sustainable aviation fuel to the fuel spray nozzles.

[0136] The idle-nvPM emission index ratio can be defined as in connection with the seventeenth aspect.

[0137] According to a twenty-sixth aspect, a gas turbine engine is intended for an aircraft, comprising one or more of the following elements: a rich-burn, rapid-cooling, lean-burn (RQL) combustion chamber having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; and wherein: A fuel stream modified with respect to the MAB-nvPM emission index ratio can be defined as follows: EImaxAb,NFKEImaxAb,FK×WK,maxAb where: EGG maxAb,NFK The nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at approximately 100% of available thrust under given operating conditions, where the fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel (SAF); EGG maxAb,FKthe nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at approximately 100% of available thrust under the given operating conditions, assuming the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and W K,maxAb the mass flow rate of the fuel supplied to the fuel spray nozzles in kg / s when the gas turbine engine is operated at approximately 100% of the available thrust under the given operating conditions; the fuel flow of the gas turbine engine, modified with respect to the MAB-nvPM emission index ratio, is less than 2 kg / s; and the gas turbine engine is configured to provide the fuel spray nozzles with an NFK-rich fuel.

[0138] The fuel flow in kg / s modified with respect to the MAB-nvPM emission index ratio can be greater than zero.

[0139] The fuel flow rate in kg / s modified with respect to the MAB-nvPM emission index ratio can be less than 1.23, more preferably less than 1.13 and even more preferably less than 1.03.

[0140] The fuel flow rate in kg / s modified with respect to the MAB-nvPM emission index ratio can be less than or equal to 0.793, more preferably less than or equal to 0.727 and even more preferably less than or equal to 0.661.

[0141] The fuel flow in kg / s modified with respect to the MAB-nvPM emission index ratio can be greater than or equal to 0.284, more preferably greater than or equal to 0.32 and even more preferably greater than or equal to 0.356.

[0142] The fuel flow rate in kg / s modified with respect to the MAB-nvPM emission index ratio 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.

[0143] The fuel flow rate modified with respect to the MAB-nvPM emission index ratio, in kg / s, can be less than or equal to 0.2, 0.25, 0.3, 0.35, 0.4, 0.45, 0.5, 0.55, 0.6, 0.65, 0.7, 0.75, 0.8, 0.85, 0.9, 0.95, 1, 1.05, 1.1, 1.15, 1.2, 1.25, or 1.3, or within any range defined by any two of these values. For example, the fuel flow rate modified with respect to the MAB-nvPM emission index ratio, in kg / s, can be within a range of 0.45 to 0.65 or 0.45 to 0.6.

[0144] W K,maxAbcan be in the range of 0.441 to 1.23 kg / s, and preferably in the range of 0.496 to 1.13 kg / s, and more preferably in the range of 0.551 to 1.03 kg / s. Alternatively, W K,maxAb in the range of 0.551 to 0.850 kg / s. W K,maxAb can range from 0.551 to 0.750 kg / s.

[0145] A fuel flow modified with respect to the climb-nvPM emission index ratio can be defined as follows: EISdough,NFKEISteig,FK×WK,Steig where: EGG Steig,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 85% of available thrust under the given operating conditions or under other operating conditions, if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; and EGG Steig,FKThe system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI Steig,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and W K,Steig the mass flow rate of the fuel supplied to the fuel spray nozzles in kg / s when the gas turbine engine is operated at approximately 85% of the available thrust under the same operating conditions as those in which EI Steig,NFK and egg Steig,FK to be calculated; and where The fuel flow rate of the gas turbine engine, modified with respect to the climb-nvPM emission index ratio, may be less than 2 kg / s.

[0146] The fuel flow in kg / s modified with respect to the climb-nvPM emission index ratio can be greater than zero.

[0147] The fuel flow rate in kg / s, modified with respect to the climb-nvPM emission index ratio, can be less than 1.01, more preferably less than 0.923 and even more preferably less than 0.839.

[0148] The fuel flow rate in kg / s modified with respect to the climb-nvPM emission index ratio can be less than or equal to 0.479, more preferably less than or equal to 0.439 and even more preferably less than or equal to 0.399.

[0149] The fuel flow rate in kg / s modified with respect to the climb-nvPM emission index ratio can be greater than or equal to 0.175, more preferably greater than or equal to 0.197 and even more preferably greater than or equal to 0.219.

[0150] The fuel flow rate in kg / s, modified with respect to the climb-nvPM emission index ratio, 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.

[0151] The fuel flow rate in kg / s modified with respect to the climb-nvPM emission index ratio can 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 in any range defined by any two of these values. For example, the fuel flow rate modified with respect to the climb-nvPM emission index ratio, in kg / s, can be in a range between 0.275 and 0.475 or 0.3 and 0.4.

[0152] W K,Steigcan be in the range of 0.369 to 1.01 kg / s, and preferably 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, W K,Steig in the range of 0.461 to 0.650 kg / s. W K,Steig It can range from 0.461 to 0.600 kg / s.

[0153] A fuel flow modified with respect to the approach-nvPM emission index ratio can be defined as follows: EI approach, NFKE approach, FK×WK, approach where: EGG Anflug,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 30% of available thrust under the given operating conditions or under other operating conditions, if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; and EGG Anflug,FKThe system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EI Anflug,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and W K,Antlug the mass flow rate of the fuel supplied to the fuel spray nozzles in kg / s when the gas turbine engine is operated at approximately 30% of the available thrust under the same operating conditions as those in which EI Anflug,NFK and egg Anflug,FK to be calculated; and where The fuel flow of the gas turbine engine, modified with respect to the approach nvPM emission index ratio, may be less than 0.4 kg / s.

[0154] The fuel flow in kg / s, modified with respect to the approach-nvPM emission index ratio, can be greater than zero.

[0155] The fuel flow in kg / s modified with respect to the approach-nvPM emission index ratio can be less than 0.334, more preferably less than 0.306 and even more preferably less than 0.278.

[0156] The fuel flow in kg / s modified with respect to the approach-nvPM emission index ratio can be less than or equal to 0.0513, more preferably less than or equal to 0.047 and even more preferably less than or equal to 0.0428.

[0157] The fuel flow in kg / s modified with respect to the approach nvPM emission index ratio can be greater than or equal to 0.0204, more preferably greater than or equal to 0.0229 and even more preferably greater than or equal to 0.0255.

[0158] The fuel flow rate modified with respect to the approach-nvPM emission index ratio in kg / s 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.

[0159] The fuel flow rate in kg / s modified for the approach-nvPM emission index ratio can 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 range defined by any two of these values. For example, the fuel flow rate in kg / s modified for the approach-nvPM emission index ratio can be in the range between 0.02 and 0.1 or 0.05 and 0.075.

[0160] W K,Anflugcan be in the range of 0.133 to 0.334 kg / s, and preferably 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, W K,Anflug in the range of 0.166 to 0.300 kg / s. W K,Anflug can range from 0.166 to 0.250 kg / s.

[0161] A fuel flow modified with respect to the idle-nvPM emission index ratio can be defined as follows: EILeer,NFKEILeer,FK×WK,Leer where: EGG Leer,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 7% of available thrust under the given operating conditions or under other operating conditions, if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; and EGG Leer,FKThe system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 7% of the available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and W K,Leer the mass flow rate of the fuel supplied to the fuel spray nozzles in kg / s when the gas turbine engine is operated at approximately 7% of the available thrust under the same operating conditions as those in which EI Leer,NFK and egg Leer,FK to be calculated; and where The fuel flow rate of the gas turbine engine, modified with respect to the idle-nvPM emission index ratio, can be less than 0.2 kg / s.

[0162] The fuel flow rate in kg / s, modified with respect to the idle-nvPM emission index ratio, can be greater than zero.

[0163] The fuel flow rate in kg / s, modified with respect to the idle-nvPM emission index ratio, can be less than 0.119, more preferably less than 0.109 and even more preferably less than 0.099.

[0164] The fuel flow rate in kg / s modified with respect to the idle-nvPM emission index ratio can be less than or equal to 0.0114, more preferably less than or equal to 0.0105 and even more preferably less than or equal to 0.00949.

[0165] The fuel flow rate in kg / s modified with respect to the idle-nvPM emission index ratio can be greater than or equal to 0.00494, more preferably greater than or equal to 0.00556 and even more preferably greater than or equal to 0.00618.

[0166] The fuel flow rate in kg / s modified with respect to the idle-nvPM emission index ratio 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.

[0167] The fuel flow rate in kg / s modified with respect to the idle-nvPM emission index ratio can be less than or equal to 0.004, 0.01, 0.015, 0.02, 0.025, 0.03, 0.035, 0.04, 0.045, 0.05, 0.055, 0.06, 0.065, 0.07, 0.075, 0.08, 0.085, 0.09, 0.095, 0.1, 0.11, or 0.12, or within any range defined by any two of these values. For example, the fuel flow rate in kg / s modified with respect to the idle-nvPM emission index ratio can be in a range between 0.01 and 0.03 or 0.015 and 0.025.

[0168] W K,Leercan be in the range of 0.0516 to 0.119 kg / s, and preferably in the range of 0.0581 to 0.109 kg / s, and more preferably in the range of 0.0645 to 0.0990 kg / s. Alternatively, W K,Leer in the range of 0.0645 to 0.0850 kg / s. W K,Leer can range from 0.0645 to 0.0750 kg / s.

[0169] According to a twenty-seventh aspect, a gas turbine engine is intended for an aircraft, comprising one or more of the following elements: a rich-burn, rapid-cooling, lean-burn (RQL) combustion chamber having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; and wherein: A fuel flow modified with respect to the climb-nvPM emission index ratio can be defined as follows: EISdough,NFKEISteig,FK×WK,Steig where: EGG Steig,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 85% of available thrust under given operating conditions and if the fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; and EGG Steig,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI Steig,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and W K,Steigthe mass flow rate of the fuel supplied to the fuel spray nozzles in kg / s when the gas turbine engine is operated at approximately 85% of the available thrust under the same operating conditions as those in which EI Steig,NFK and egg Steig,FK to be calculated; the fuel flow rate of the gas turbine engine, modified with respect to the climb-nvPM emission index ratio, is less than 2 in kg / s; and the gas turbine engine is configured to provide the fuel spray nozzles with an NFK-rich fuel.

[0170] The fuel flow and / or W modified with respect to the climb-nvPM emission index ratio K,Steig can be defined as above in connection with the twenty-sixth aspect.

[0171] According to a twenty-eighth aspect, a gas turbine engine is intended for an aircraft, comprising one or more of the following elements: a rich-burn, rapid-cooling, lean-burn (RQL) combustion chamber having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; and wherein: A fuel flow modified with respect to the approach nvPM emission index ratio can be defined as follows: EI approach, NFKE approach, FK×WK, approach where: EGG Anflug,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 30% of available thrust under the given operating conditions or under other operating conditions, if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; and EGG Anflug,FKThe system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EI Anflug,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and W K,Anflug the mass flow rate of the fuel supplied to the fuel spray nozzles in kg / s when the gas turbine engine is operated at approximately 30% of the available thrust under the same operating conditions as those in which EI Anflug,NFK and egg Anflug,FK calculated; and where the fuel flow of the gas turbine engine, modified with respect to the approach nvPM emission index ratio, is less than 0.4 kg / s; and the gas turbine engine is configured to provide the fuel spray nozzles with an NFK-rich fuel.

[0172] The fuel flow and / or W modified with respect to the approach-nvPM emission index ratio K,Anflug can be defined as above in connection with the twenty-sixth aspect.

[0173] According to a twenty-ninth aspect, a gas turbine engine is intended for an aircraft, comprising one or more of the following elements: a rich-burn, rapid-cooling, lean-burn (RQL) combustion chamber having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; and wherein: A fuel flow modified with respect to the idle-nvPM emission index ratio can be defined as follows: EILeer,NFKEILeer,FK×WK,Leer where: EGG Leer,NFKthe nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with 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 EGG Leer,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 7% of the available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and W K,Leer the mass flow rate of the fuel supplied to the fuel spray nozzles in kg / s when the gas turbine engine is operated at approximately 7% of the available thrust under the same operating conditions as those in which EI Leer,NFK and eggLeer,FK to be calculated; the fuel flow rate of the gas turbine engine, modified with respect to the idle-nvPM emission index ratio, is less than 0.2 kg / s; and the gas turbine engine is configured to provide the fuel spray nozzles with an NFK-rich fuel.

[0174] The fuel flow modified with respect to the idle-nvPM emission index ratio and / or W K,Leer can be defined as above in connection with the twenty-sixth aspect.

[0175] According to a thirtieth aspect, a method for operating the gas turbine engine of the twenty-sixth, twenty-seventh, twenty-eighth or twenty-ninth aspect is provided, the method comprising supplying fuel comprising a sustainable aviation fuel to the fuel spray nozzles.

[0176] According to a thirty-first aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising the following: a combustion chamber for rich burning, rapid cooling, lean burning (RQL combustion chamber) having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; wherein: A fuel stream modified with respect to the MAB-nvPM emission index ratio can be defined as follows: EImaxAb,NFKEImaxAb,FK×WK,maxAb where: EGG maxAb,NFK The nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at approximately 100% of available thrust under given operating conditions, where the fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel (SAF); EGG maxAb,FKthe nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at approximately 100% of available thrust under the given operating conditions, assuming the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and W K,maxAb the mass flow rate of the fuel supplied to the fuel spray nozzles in kg / s when the gas turbine engine is operated at approximately 100% of the available thrust under the given operating conditions; the fuel flow in kg / s modified with respect to the MAB-nvPM emission index ratio is less than 2; and The process involves supplying fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0177] The fuel flow modified with respect to the MAB-nvPM emission index ratio and / or W K,maxAbcan be defined as above in connection with the twenty-sixth aspect.

[0178] A fuel flow modified with respect to the climb-nvPM emission index ratio can be defined as follows: EISdough,NFKEISteig,FK×WK,Steig where: EGG Steig,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 85% of available thrust under the given operating conditions or under other operating conditions, if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; and EGG Steig,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI Steig,NFKis calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and W K,Steig the mass flow rate of the fuel supplied to the fuel spray nozzles in kg / s when the gas turbine engine is operated at approximately 85% of the available thrust under the same operating conditions as those in which EI Steig,NFK and egg Steig,FK to be calculated; and where The fuel flow rate of the gas turbine engine, modified with respect to the climb-nvPM emission index ratio, may be less than 2 kg / s.

[0179] The fuel flow and / or W modified with respect to the climb-nvPM emission index ratio K,Steig can be defined as above in connection with the twenty-sixth aspect.

[0180] A fuel flow modified with respect to the approach-nvPM emission index ratio can be defined as follows: EI approach, NFKE approach, FK×WK, approach where: EGG Anflug,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 30% of available thrust under the given operating conditions or under other operating conditions, if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; and EGG Anflug,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EI Anflug,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and WK,Anflug the mass flow rate of the fuel supplied to the fuel spray nozzles in kg / s when the gas turbine engine is operated at approximately 30% of the available thrust under the same operating conditions as those in which EI Anflug,NFK and egg Anflug,FK to be calculated; and where The fuel flow of the gas turbine engine, modified with respect to the approach nvPM emission index ratio, may be less than 0.4 kg / s.

[0181] The fuel flow and / or W modified with respect to the approach-nvPM emission index ratio K,Anflug can be defined as above in connection with the twenty-sixth aspect.

[0182] A fuel flow modified with respect to the idle-nvPM emission index ratio can be defined as follows: EILeer,NFKEILeer,FK×WK,Leer where: EGG Leer,NFKthe nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 7% of available thrust under the given operating conditions or under other operating conditions, if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; and EGG Leer,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 7% of the available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and W K,Leerthe mass flow rate of the fuel supplied to the fuel spray nozzles in kg / s when the gas turbine engine is operated at approximately 7% of the available thrust under the same operating conditions as those in which EI Leer,NFK and egg Leer,FK to be calculated; and where The fuel flow rate of the gas turbine engine, modified with respect to the idle-nvPM emission index ratio, can be less than 0.2 kg / s.

[0183] The fuel flow modified with respect to the idle-nvPM emission index ratio and / or W K,LeerAb can be defined as above in connection with the twenty-sixth aspect.

[0184] According to a thirty-second aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising the following: a rich-burn, rapid-cooling, lean-burn (RQL) combustion chamber having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; and wherein: A fuel flow modified with respect to the climb-nvPM emission index ratio can be defined as follows: EISdough,NFKEISteig,FK×WK,Steig where: EGG Steig,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 85% of available thrust under given operating conditions and if the fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; and EGG Steig,FKThe system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI Steig,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and W K,Steig the mass flow rate of the fuel supplied to the fuel spray nozzles in kg / s when the gas turbine engine is operated at approximately 85% of the available thrust under the same operating conditions as those in which EI Steig,NFK and egg Steig,FK to be calculated; the fuel flow rate of the gas turbine engine, modified with respect to the climb-nvPM emission index ratio, is less than 2 in kg / s; and The process involves supplying fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0185] The fuel flow and / or W modified with respect to the climb-nvPM emission index ratio K,Steig can be defined as above in connection with the twenty-sixth aspect.

[0186] According to a thirty-third aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising the following: a rich-burn, rapid-cooling, lean-burn (RQL) combustion chamber having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; and wherein: A fuel flow modified with respect to the approach nvPM emission index ratio can be defined as follows: EI approach, NFKE approach, FK×WK, approach where: EGG Anflug,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 30% of available thrust under given operating conditions or under other operating conditions, if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; and EGG Anflug,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EI Anflug,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and W K,Anflugthe mass flow rate of the fuel supplied to the fuel spray nozzles in kg / s when the gas turbine engine is operated at approximately 30% of the available thrust under the same operating conditions as those in which EI Anflug,NFK and egg Anflug,FK to be calculated; the fuel flow rate of the gas turbine engine, modified with respect to the approach-nvPM emission index ratio, is less than 0.4 kg / s; and The process involves supplying fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0187] The fuel flow and / or W modified with respect to the approach-nvPM emission index ratio K,Anflug can be defined as above in connection with the twenty-sixth aspect.

[0188] According to a thirty-fourth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising the following: a rich-burn, rapid-cooling, lean-burn (RQL) combustion chamber having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; and wherein: A fuel flow modified with respect to the idle-nvPM emission index ratio can be defined as follows: EILeer,NFKEILeer,FK×WK,Leer where: EGG Leer,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with 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 EGGLeer,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 7% of the available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and W K,Leer the mass flow rate of the fuel supplied to the fuel spray nozzles in kg / s when the gas turbine engine is operated at approximately 7% of the available thrust under the same operating conditions as those in which EI Leer,NFK and egg Leer,FK to be calculated; the fuel flow rate of the gas turbine engine, modified with respect to the idle-nvPM emission index ratio, is less than 0.2 kg / s; and The process involves supplying fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0189] The fuel flow modified with respect to the idle-nvPM emission index ratio and / or W K,Leer can be defined as above in connection with the twenty-sixth aspect.

[0190] According to a thirty-fifth aspect, a gas turbine engine is intended for an aircraft, comprising one or more of the following elements: a combustion chamber for rich burning, rapid cooling, lean burning (RQL combustion chamber) having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; wherein: An nvPM emission index ratio for lean-burn cruise flight can be defined as follows: EIJourney(lean),NFKEIJourney(lean),FK where: EGG Reise(mager),NFK which can be defined as follows: EImaxAb,NFK+EISteig,NFK2 EGG Reise(mager),FK which can be defined as follows: EImaxAb,FK+EISteig,FK2 EGG maxAb,NFK The nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at approximately 100% of available thrust under given operating conditions, assuming that the fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Steig,NFK The nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 85% of available thrust under given operating conditions, when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG maxAb,FKthe nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at approximately 100% of available thrust under the given operating conditions, assuming the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and EGG Steig,FK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 85% of the available thrust under the given operating conditions, when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; the nvPM emission index ratio for cruise flight with lean combustion of the gas turbine engine is less than 1; and the gas turbine engine is configured to provide the fuel spray nozzles with a sustainable aviation fuel (SAF)-based fuel.

[0191] The nvPM emission index ratio for lean-burn cruise flight can be greater than zero.

[0192] The nvPM emission index ratio for lean-burn cruise flight can 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.

[0193] The nvPM emission index ratio for lean-burn cruise flight can 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.

[0194] The nvPM emission index ratio for lean-burn cruise flight can 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.

[0195] The nvPM emission index ratio for lean-burn cruise flight can 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.

[0196] The nvPM emission index ratio for lean-burn cruise flight can be less than 1, 0.95, 0.9, 0.85, 0.8, 0.75, 0.7, 0.65, 0.6, 0.55, 0.5, 0.45, 0.4, 0.35, 0.3, 0.25, 0.2, 0.15, 0.1, or 0.05, or within any range defined by any two of these values. For example, the nvPM emission index ratio for lean-burn cruise flight can be in the range of 0.65 to 0.85 or 0.7 to 0.75.

[0197] The nvPM emission index ratio for lean-burn cruise flight 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 by any two of these values.

[0198] An nvPM emission index ratio of idle to MAB can be defined as follows: EILeer,NFKEImaxAb,NFKEILeer,FKEImaxAb,FK where: EGG Leer,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 7% of the available thrust under the given operating conditions or under other operating conditions, if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG maxAb,NFKThe system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI Leer,NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Leer,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 7% of the available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and EGG maxAb,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EILeer,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and where the nvPM emission index ratio idle to MAB of the gas turbine engine can be less than 1.

[0199] The nvPM emission index ratio of idle to MAB can be greater than zero.

[0200] The nvPM emission index ratio idle to MAB can be less than or equal to 0.8 and 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 0.2.

[0201] The nvPM emission index ratio idle to MAB can 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.

[0202] The nvPM emission index ratio idle to MAB can 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.

[0203] The nvPM emission index ratio idle to MAB can be greater than or equal to 0.118 and can preferably be greater than or equal to 0.133 and can more preferably be greater than or equal to 0.148.

[0204] The nvPM emission index ratio idle to MAB can be in the range of 0.118 to 0.178 and can preferably be in the range of 0.133 to 0.164 and can more preferably be in the range of 0.148 to 0.149.

[0205] The nvPM emission index ratio (idle to MAB) can be less than 1, 0.95, 0.9, 0.85, 0.8, 0.75, 0.7, 0.65, 0.6, 0.55, 0.5, 0.45, 0.4, 0.35, 0.3, 0.25, 0.2, 0.15, 0.1, or 0.05, or within any range defined by any two of these values. For example, the nvPM emission index ratio (idle to MAB) can be in the range of 0.25 to 0.4 or 0.3 to 0.35.

[0206] The nvPM emission index ratio idle to MAB 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 lie within any range defined by any two of these values.

[0207] An nvPM emission index ratio for lean-burn cruise flight to MAB can be defined as follows: EITravel(lean),NFKEImaxAb,NFKEITravel(lean),FKEImaxAb,FK where: EGG Reise(mager),NFK which can be defined as follows: EImaxAb,NFK+EISteig,NFK2 EGGReise(mager),FK which can be defined as follows: EImaxAb,FK+EISteig,FK2 EGG maxAb,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 100% of the available thrust under the given operating conditions or under other operating conditions, if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Steig,NFK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI maxAb,NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG maxAb,FKThe system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI maxAb,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and EGG Steig,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI maxAb,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and where the nvPM emission index ratio of lean-burn cruise to MAB of the gas turbine engine can be less than 1.

[0208] The nvPM emission index ratio of lean-burn cruise to MAB can be greater than zero.

[0209] The nvPM emission index ratio for lean-burn cruise flight to MAB can 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 for lean-burn cruise flight to MAB can be less than or equal to 0.914.

[0210] The nvPM emission index ratio of lean-burn cruise to MAB 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.

[0211] The nvPM emission index ratio for lean-burn cruise to MAB can 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.

[0212] The nvPM emission index ratio of lean-burn cruise to MAB can be less than 1, 0.95, 0.9, 0.85, 0.8, 0.75, 0.7, 0.65, 0.6, 0.55, 0.5, 0.45, 0.4, 0.35, 0.3, 0.25, 0.2, 0.15, 0.1 or 0.05, or lie within any range defined by any two of these values.

[0213] The nvPM emission index ratio for lean-burn cruise to MAB 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 range defined by any two of these values. For example, the nvPM emission index ratio for lean-burn cruise to MAB can be in a range between 0.91 and 0.99 or between 0.93 and 0.97.

[0214] An nvPM emission index ratio of idle to cruise flight with lean combustion can be defined as follows: EILeer,NFKEIReise(mager),NFKEILeer,FKEIReise(mager),FK where: EGG Leer,NFKthe nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 7% of the available thrust under the given operating conditions or under other operating conditions, if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Reise(mager),NFK which can be defined as follows: EImaxAb,NFK+EISteig,NFK2 EGG maxAb,NFK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI Leer,NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Steig,NFKThe system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Leer,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 7% of the available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; EGG Reise(mager),FK which can be defined as follows: EImaxAb,FK+EISteig,FK2 EGG maxAb,FKThe system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI Leer‚NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and EGG Steig,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI Leer‚NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and where the nvPM emission index ratio of idle to cruise flight with lean combustion of the gas turbine engine can be less than 1.

[0215] The nvPM emission index ratio of idle to cruise flight with lean combustion can be greater than zero.

[0216] The nvPM emission index ratio of idle to cruise flight with lean combustion 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.

[0217] The nvPM emission index ratio of idle to cruise flight with lean combustion can 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.

[0218] The nvPM emission index ratio of idle to cruise flight with lean combustion can 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.

[0219] The nvPM emission index ratio of idle to cruise flight with lean combustion can 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.

[0220] The nvPM emission index ratio at idle to cruise with lean combustion can be less than 1, 0.95, 0.9, 0.85, 0.8, 0.75, 0.7, 0.65, 0.6, 0.55, 0.5, 0.45, 0.4, 0.35, 0.3, 0.25, 0.2, 0.15, 0.1, or 0.05, or within any range defined by any two of these values. For example, the nvPM emission index ratio at idle to cruise with lean combustion can be in the range between 0.3 and 0.4 or 0.3 and 0.35.

[0221] The nvPM emission index ratio idle to cruise flight with lean combustion 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 lie in any range defined by any two of these values.

[0222] An nvPM emission index ratio for cruise flight with rich combustion can be defined as follows: EITravel(bold),NFKEITravel(bold),FK where: EGG Reise(fett),NFK which can be defined as follows: EI climb, NFK+EI approach, NFK2 EGG Reise(fett),FK which can be defined as follows: EI climb, FK+EI approach, FK2 EGG Steig,NFKthe nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 85% of the available thrust under the given operating conditions or under other operating conditions, if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Anflug,NFK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EI Steig,NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Steig,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI Steig,NFKis calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and EGG Anflug,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EI Steig,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and where the nvPM emission index ratio for cruise flight with rich combustion of the gas turbine engine can be less than 1.

[0223] The nvPM emission index ratio for cruise flight with rich combustion can be greater than zero.

[0224] The nvPM emission index ratio for cruise flight with rich combustion can be less than or equal to 0.9, preferably less than or equal to 0.8, more preferably less than or equal to 0.7.

[0225] The nvPM emission index ratio for cruise flight with rich combustion can 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.

[0226] The nvPM emission index ratio for cruise flight with rich combustion can 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.

[0227] The nvPM emission index ratio for cruise flight with rich combustion can 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.

[0228] The nvPM emission index ratio for rich-burn cruise flight can be less than 1, 0.95, 0.9, 0.85, 0.8, 0.75, 0.7, 0.65, 0.6, 0.55, 0.5, 0.45, 0.4, 0.35, 0.3, 0.25, 0.2, 0.15, 0.1, or 0.05, or within any range defined by any two of these values. For example, the nvPM emission index ratio for rich-burn cruise flight can be in a range between 0.45 and 0.7 or 0.5 and 0.65.

[0229] The nvPM emission index ratio for cruise flight with rich combustion can 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 lie within any range defined by any two of these values.

[0230] An nvPM emission index ratio for cruise flight with rich combustion to MAB can be defined as follows: EITravel(bold),NFKEImaxAb,NFKEITravel(bold),FKEImaxAb,FK where: EGG Reise(fett),NFKwhich can be defined as follows: EI climb, NFK+EI approach, NFK2 and egg Reise(fett),FK which can be defined as follows: EI climb, FK+EI approach, FK2 and whereby: EGG Steig,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 85% of the available thrust under the given operating conditions or under other operating conditions, if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Anflug,NFK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EI Steig,NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Steig,FKThe system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI Steig,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; EGG Anflug,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EI Steig,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; EGG maxAb,NFKThe system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI Steig,NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; and EGG maxAb,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI Steig,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and where the nvPM emission index ratio of cruise flight with rich combustion to MAB of the gas turbine engine is less than 1.

[0231] The nvPM emission index ratio of cruise flight with rich combustion to MAB can be greater than zero.

[0232] The nvPM emission index ratio of cruise flight with rich combustion to MAB can be less than or equal to 0.95, preferably less than or equal to 0.9 and more preferably less than or equal to 0.875.

[0233] The nvPM emission index ratio of cruise flight with rich combustion to MAB can 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.

[0234] The nvPM emission index ratio of cruise flight with rich combustion to MAB can 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.

[0235] The nvPM emission index ratio for cruise flight with rich combustion to MAB can 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.

[0236] The nvPM emission index ratio of rich-burn cruise to MAB can be less than 1, 0.95, 0.9, 0.85, 0.8, 0.75, 0.7, 0.65, 0.6, 0.55, 0.5, 0.45, 0.4, 0.35, 0.3, 0.25, 0.2, 0.15, 0.1, or 0.05, or within any range defined by any two of these values. For example, the nvPM emission index ratio of rich-burn cruise to MAB can be in the range of 0.65 to 0.9, 0.7 to 0.85, or 0.75 to 0.8.

[0237] The nvPM emission index ratio of cruise flight with rich combustion to MAB 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 lie in any range defined by any two of these values.

[0238] An nvPM emission index ratio of idle to cruise flight with rich combustion can be defined as follows: EILeer,NFKEIReise(fett),NFKEILeer,FKEIReise(fett),FK where: EGG Leer,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 7% of the available thrust under the given operating conditions or under other given operating conditions, and if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel (SAF); EGG Leer,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 7% of the available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; EGG Reise(fett),NFK which can be defined as follows: EI climb, NFK+EI approach, NFK2 EGG Reise(fett),FKwhich can be defined as follows: EI climb, FK+EI approach, FK2 EGG Steig,NFK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Steig,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; EGG Anflug,NFKThe system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EI Leer‚NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; and EGG Anflug,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EI Leer‚NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and where the nvPM emission index ratio of idle to cruise flight with rich combustion of the gas turbine engine can be less than 1.

[0239] The nvPM emission index ratio of idle to cruise flight with rich combustion can be greater than zero.

[0240] The nvPM emission index ratio of idle to cruise flight with rich combustion can 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.

[0241] The nvPM emission index ratio of idle to cruise flight with rich combustion can 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.

[0242] The nvPM emission index ratio of idle to cruise flight with rich combustion can 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.

[0243] The nvPM emission index ratio of idle to cruise flight with rich combustion can 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.

[0244] The nvPM emission index ratio at idle to cruise with rich combustion can be less than 1, 0.95, 0.9, 0.85, 0.8, 0.75, 0.7, 0.65, 0.6, 0.55, 0.5, 0.45, 0.4, 0.35, 0.3, 0.25, 0.2, 0.15, 0.1, or 0.05, or within any range defined by any two of these values. For example, the nvPM emission index ratio at idle to cruise with rich combustion can be in a range between 0.3 and 0.5 or 0.35 and 0.45.

[0245] The nvPM emission index ratio of idle to cruise flight with rich combustion 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 lie within any range defined by any two of these values.

[0246] According to a thirty-sixth aspect, a gas turbine engine is intended for an aircraft, comprising one or more of the following elements: a combustion chamber for rich burning, rapid cooling, lean burning (RQL combustion chamber) having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; wherein: An nvPM emission index ratio of idle to MAB can be defined as follows: EILeer,NFKEImaxAb,NFKEILeer,FKEImaxAb,FK where: EGG Leer,NFK The nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 7% of available thrust under given operating conditions, when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG maxAb,NFK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI Leer,NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Leer,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 7% of the available thrust under the same operating conditions as those under which EI Leer,NFKis calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and EGG maxAb,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI Leer,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; the nvPM emission index ratio at idle to MAB of the gas turbine engine may be less than 1; and the gas turbine engine is configured to provide the fuel spray nozzles with a sustainable aviation fuel (SAF)-based fuel.

[0247] The nvPM emission index ratio idle to MAB can be defined as above in connection with the thirty-fifth aspect.

[0248] According to a thirty-seventh aspect, a gas turbine engine is intended for an aircraft, comprising one or more of the following elements: a combustion chamber for rich burning, rapid cooling, lean burning (RQL combustion chamber) having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; where: An nvPM emission index ratio for lean-burn cruise flight to MAB can be defined as follows: EITravel(lean),NFKEImaxAb,NFKEITravel(lean),FKEImaxAb,FK where: EGG Reise(mager),NFK which can be defined as follows: EImaxAb,NFK+EISteig,NFK2 EGG Reise(mager),FK which can be defined as follows: EImaxAb,FK+EISteig,FK2 EGG maxAb,NFK The nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at approximately 100% of available thrust under given operating conditions, assuming that the fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Steig,NFK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI maxAb,NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG maxAb,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI maxAb,NFKis calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and EGG Steig,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI maxAb,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; The nvPM emission index ratio of lean-burn cruise flight to MAB of the gas turbine engine can be less than 1; and the gas turbine engine is configured to provide the fuel spray nozzles with a sustainable aviation fuel (SAF)-based fuel.

[0249] The nvPM emission index ratio of lean-burn cruise to MAB can be defined as above in connection with the thirty-fifth aspect.

[0250] According to a thirty-eighth aspect, a gas turbine engine is intended for an aircraft, comprising one or more of the following elements: a combustion chamber for rich burning, rapid cooling, lean burning (RQL combustion chamber) having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; wherein: An nvPM emission index ratio of idle to cruise flight with lean combustion can be defined as follows: EILeer,NFKEIReise(mager),NFKEILeer,FKEIReise(mager),FK where: EGG Leer,NFKThe nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 7% of available thrust under given operating conditions, when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Reise(mager),NFK which can be defined as follows: EImaxAb,NFK+EISteig,NFK2 EGG maxAb,NFK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI Leer,NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Steig,NFK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EILeer,NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Leer,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 7% of the available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; EGG Reise(mager),FK which can be defined as follows: EImaxAb,FK+EISteig,FK2 EGG maxAb,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI Leer,NFKis calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and EGG Steig,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; the nvPM emission index ratio of idle to cruise flight with lean combustion of the gas turbine engine is less than 1; and the gas turbine engine is configured to provide the fuel spray nozzles with a sustainable aviation fuel (SAF)-based fuel.

[0251] The nvPM emission index ratio of idle to cruise flight with lean combustion can be defined as above in connection with the thirty-fifth aspect.

[0252] According to a thirty-ninth aspect, a gas turbine engine is intended for an aircraft, comprising one or more of the following elements: a combustion chamber for rich burning, rapid cooling, lean burning (RQL combustion chamber) having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; wherein: An nvPM emission index ratio for cruise flight with rich combustion can be defined as follows: EITravel(bold),NFKEITravel(bold),FK where: EGG Reise(fett),NFK which can be defined as follows: EI climb, NFK+EI approach, NFK2 EGG Reise(fett),FK which can be defined as follows: EI climb, FK+EI approach, FK2 EGG Steig,NFK The nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 85% of available thrust under given operating conditions, when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Anflug,NFK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EI Steig,NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Steig,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI Steig,NFKis calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and EGG Anflug,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EI Steig,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; the nvPM emission index ratio for cruise flight with rich combustion of the gas turbine engine is less than 1; and the gas turbine engine is configured to provide the fuel spray nozzles with a sustainable aviation fuel (SAF)-based fuel.

[0253] The nvPM emission index ratio for cruise flight with rich combustion can be defined as above in relation to the thirty-fifth aspect.

[0254] According to a fortieth aspect, a gas turbine engine is intended for an aircraft that includes one or more of the following elements: a combustion chamber for rich burning, rapid cooling, lean burning (RQL combustion chamber) having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; wherein: An nvPM emission index ratio for cruise flight with rich combustion to MAB can be defined as follows: EITravel(bold),NFKEImaxAb,NFKEITravel(bold),FKEImaxAb,FK where: EGG Reise(fett),NFK which can be defined as follows: EI climb, NFK+EI approach, NFK2 and egg Reise(fett),FK which can be defined as follows: EI climb, FK+EI approach, FK2 and whereby: EGG Steig,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 85% of the available thrust under given operating conditions and if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Anflug,NFK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EI Steig,NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Steig,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EISteig,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; EGG Anflug,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EI Steig,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; EGG maxAb,NFK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI Steig,NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; and EGG maxAb,FKThe system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI Steig,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; the nvPM emission index ratio of cruise flight with rich combustion to MAB of the gas turbine engine is less than 1; and the gas turbine engine is configured to provide the fuel spray nozzles with a sustainable aviation fuel (SAF)-based fuel.

[0255] The nvPM emission index ratio of cruise flight with rich combustion to MAB can be defined as above in connection with the thirty-fifth aspect.

[0256] According to a forty-first aspect, a gas turbine engine is intended for an aircraft, comprising one or more of the following elements: a combustion chamber for rich burning, rapid cooling, lean burning (RQL combustion chamber) having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; wherein: An nvPM emission index ratio of idle to cruise flight with rich combustion can be defined as follows: EILeer,NFKEIReise(fett),NFKEILeer,FKEIReise(fett),FK where: EGG Leer,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 7% of the available thrust under given operating conditions, and if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel (SAF); EGG Leer,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 7% of the available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; EGG Reise(fett),NFK which can be defined as follows: EI climb, NFK+EI approach, NFK2 EGG Reise(fett),FK which can be defined as follows: EI climb, FK+EI approach, FK2 EGG Steig,NFK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI Leer‚NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Steig,FKThe system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI Leer‚NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; EGG Anflug,NFK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EI Leer‚NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; and EGG Anflug,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EILeer‚NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; the nvPM emission index ratio of idle to cruise flight with rich combustion of the gas turbine engine is less than 1; and the gas turbine engine is configured to provide the fuel spray nozzles with a sustainable aviation fuel (SAF)-based fuel.

[0257] The nvPM emission index ratio of idle to cruise flight with rich combustion can be defined as above in connection with the thirty-fifth aspect.

[0258] According to a forty-second aspect, a method for operating the gas turbine engine of any or any of the thirty-fifth, thirty-sixth, thirty-seventh, thirty-eighth, thirty-ninth, fortieth or forty-first aspects is provided, the method comprising supplying fuel comprising a sustainable aviation fuel to the fuel spray nozzles.

[0259] According to a forty-third aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising the following: a combustion chamber for rich burning, rapid cooling, lean burning (RQL combustion chamber) having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; wherein: An nvPM emission index ratio for lean-burn cruise flight can be defined as follows: EIJourney(lean),NFKEIJourney(lean),FK where: EGG Reise(mager),NFK which can be defined as follows: EImaxAb,NFK+EISteig,NFK2 EGG Reise(mager),FK which can be defined as follows: EImaxAb,FK+EISteig,FK2 EGG maxAb,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 100% of the available thrust under the given operating conditions and if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Steig,NFKthe nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 85% of the available thrust under the given operating conditions and if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG maxAb,FK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at approximately 100% of available thrust under the given operating conditions, assuming the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and EGG Steig,FKthe nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 85% of the available thrust under the given operating conditions, when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; the nvPM emission index ratio for lean-burn cruise flight is less than 1; and The process involves supplying fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0260] The nvPM emission index ratio for lean-burn cruise flight can be defined as above in relation to the thirty-fifth aspect.

[0261] An nvPM emission index ratio of idle to MAB can be defined as follows: EILeer,NFKEImaxAb,NFKEILeer,FKEImaxAb,FK where: EGG Leer,NFKthe nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 7% of the available thrust under the given operating conditions or under other operating conditions, and if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG maxAb,NFK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI Leer,NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Leer,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 7% of the available thrust under the same operating conditions as those under which EI Leer,NFKis calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and EGG maxAb,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI Leer,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and wherein The nvPM emission index ratio at idle to MAB of the gas turbine engine is less than 1.

[0262] The nvPM emission index ratio idle to MAB can be defined as above in connection with the thirty-fifth aspect.

[0263] An nvPM emission index ratio for lean-burn cruise flight to MAB can be defined as follows: EITravel(lean),NFKEImaxAb,NFKEITravel(lean),FKEImaxAb,FK where: EGG Reise(mager),NFK which can be defined as follows: EImaxAb,NFK+EISteig,NFK2 EGG Reise(mager),FK which can be defined as follows: EImaxAb,FK+EISteig,FK2 EGG maxAb,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 100% of the available thrust under the given operating conditions or under other operating conditions, and if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Steig,NFK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI maxAb,NFKis calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG maxAb,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI maxAb,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and EGG Steig,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI maxAb,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and wherein The nvPM emission index ratio of lean-burn cruise flight to MAB of the gas turbine engine is less than 1.

[0264] The nvPM emission index ratio of lean-burn cruise to MAB can be defined as above in connection with the thirty-fifth aspect.

[0265] An nvPM emission index ratio of idle to cruise flight with lean combustion can be defined as follows: EILeer,NFKEIReise(mager),NFKEILeer,FKEIReise(mager),FK where: EGG Leer,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 7% of the available thrust under the given operating conditions or under other operating conditions, and if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Reise(mager),NFK which can be defined as follows: EImaxAb,NFK+EISteig,NFK2 EGG maxAb,NFK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI Leer,NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Steig,NFK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Leer,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 7% of the available thrust under the same operating conditions as those under which EI Leer,NFKis calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; EGG Reise(mager),FK which can be defined as follows: EImaxAb,FK+EISteig,FK2 EGG maxAb,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI Leer,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and EGG Steig,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI Leer,NFKis calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and where and where the nvPM emission index ratio idle to cruise flight with lean combustion of the gas turbine engine may be less than 1.

[0266] The nvPM emission index ratio of idle to cruise flight with lean combustion can be defined as above in connection with the thirty-fifth aspect.

[0267] An nvPM emission index ratio for cruise flight with rich combustion can be defined as follows: EITravel(bold),NFKEITravel(bold),FK where: EGG Reise(fett),NFK which can be defined as follows: EI climb, NFK+EI approach, NFK2 EGG Reise(fett),FK which can be defined as follows: EI climb, FK+EI approach, FK2 EGG Steig,NFKthe nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 85% of the available thrust under the given operating conditions or under other operating conditions, and if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Anflug,NFK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EI Steig,NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Steig,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI Steig,NFKis calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and EGG Anflug,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EI Steig,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and where the nvPM emission index ratio for cruise flight with rich combustion of the gas turbine engine can be less than 1.

[0268] The nvPM emission index ratio for cruise flight with rich combustion can be defined as above in relation to the thirty-fifth aspect.

[0269] An nvPM emission index ratio for cruise flight with rich combustion to MAB can be defined as follows: EITravel(bold),NFKEImaxAb,NFKEITravel(bold),FKEImaxAb,FK where: EGG Reise(fett),NFK which can be defined as follows: EI climb, NFK+EI approach, NFK2 and egg Reise(fett),FK which can be defined as follows: EI climb, FK+EI approach, FK2 and whereby: EGG Steig,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 85% of the available thrust under the given operating conditions or under other operating conditions, and if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Anflug,NFKThe system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EI Steig,NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Steig,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI Steig,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; EGG Anflug,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EISteig,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; EGG maxAb,NFK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI Steig,NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; and EGG maxAb,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI Steig,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and where the nvPM emission index ratio of cruise flight with rich combustion to MAB of the gas turbine engine is less than 1.

[0270] The nvPM emission index ratio of cruise flight with rich combustion to MAB can be defined as above in connection with the thirty-fifth aspect.

[0271] An nvPM emission index ratio of idle to cruise flight with rich combustion can be defined as follows: EILeer,NFKEIReise(fett),NFKEILeer,FKEIReise(fett),FK where: EGG Leer,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 7% of the available thrust under the given operating conditions or under other given operating conditions, and if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel (SAF); EGG Leer,FKThe system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 7% of the available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; EGG Reise(fett),NFK which can be defined as follows: EI climb, NFK+EI approach, NFK2 EGG Reise(fett),FK which can be defined as follows: EI climb, FK+EI approach, FK2 EGG Steig,NFK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Steig,FKThe system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI Leer‚NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; EGG Anflug,NFK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EI Leer‚NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; and EGG Anflug,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EILeer‚NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and where the nvPM emission index ratio of idle to cruise flight with rich combustion of the gas turbine engine can be less than 1.

[0272] The nvPM emission index ratio of idle to cruise flight with rich combustion can be defined as above in connection with the thirty-fifth aspect.

[0273] According to a forty-fourth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising the following: a combustion chamber for rich burning, rapid cooling, lean burning (RQL combustion chamber) having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; wherein: An nvPM emission index ratio of idle to MAB can be defined as follows: EILeer,NFKEImaxAb,NFKEILeer,FKEImaxAb,FK where: EGG Leer,NFK The nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 7% of available thrust under given operating conditions, when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG maxAb,NFK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI Leer,NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Leer,FKThe system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 7% of the available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and EGG maxAb,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI Leer,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; the nvPM emission index ratio at idle to MAB of the gas turbine engine may be less than 1; and The process involves supplying fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0274] The nvPM emission index ratio idle to MAB can be defined as above in connection with the thirty-fifth aspect.

[0275] According to a forty-fifth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising the following: a combustion chamber for rich burning, rapid cooling, lean burning (RQL combustion chamber) having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; wherein: An nvPM emission index ratio for lean-burn cruise flight to MAB can be defined as follows: EITravel(lean),NFKEImaxAb,NFKEITravel(lean),FKEImaxAb,FK where: EGG Reise(mager),NFKwhich can be defined as follows: EImaxAb,NFK+EISteig,NFK2 EGG Reise(mager),FK which can be defined as follows: EImaxAb,FK+EISteig,FK2 EGG maxAb,NFK The nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at approximately 100% of available thrust under given operating conditions, assuming that the fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Steig,NFK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI maxAb,NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG maxAb,FKThe system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI maxAb,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and EGG Steig,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI maxAb,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; The nvPM emission index ratio of lean-burn cruise flight to MAB of the gas turbine engine can be less than 1; and The process involves supplying fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0276] The nvPM emission index ratio of lean-burn cruise to MAB can be defined as above in connection with the thirty-fifth aspect.

[0277] According to a forty-sixth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising the following: a combustion chamber for rich burning, rapid cooling, lean burning (RQL combustion chamber) having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; wherein: An nvPM emission index ratio of idle to cruise flight with lean combustion can be defined as follows: EILeer,NFKEIReise(mager),NFKEILeer,FKEIReise(mager),FK where: EGG Leer,NFK The nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 7% of available thrust under given operating conditions, when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Reise(mager),NFK which can be defined as follows: EImaxAb,NFK+EISteig,NFK2 EGG maxAb,NFK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI Leer,NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Steig,NFKThe system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Leer,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 7% of the available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; EGG Reise(mager),FK which can be defined as follows: EImaxAb,FK+EISteig,FK2 EGG maxAb,FKThe system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI Leer‚NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and EGG Steig,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI Leer‚NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; the nvPM emission index ratio of idle to cruise flight with lean combustion of the gas turbine engine is less than 1; and The process involves supplying fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0278] The nvPM emission index ratio of idle to cruise flight with lean combustion can be defined as above in connection with the thirty-fifth aspect.

[0279] According to a forty-seventh aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising the following: a combustion chamber for rich burning, rapid cooling, lean burning (RQL combustion chamber) having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; wherein: An nvPM emission index ratio for cruise flight with rich combustion can be defined as follows: EITravel(bold),NFKEITravel(bold),FK where: EGGReise(fett),NFK which can be defined as follows: EI climb, NFK+EI approach, NFK2 EGG Reise(fett),FK which can be defined as follows: EI climb, FK+EI approach, FK2 EGG Steig,NFK The nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 85% of available thrust under given operating conditions, when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Anflug,NFK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EI Steig,NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Steig,FKThe system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI Steig,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; and EGG Anflug,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EI Steig,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; the nvPM emission index ratio for cruise flight with rich combustion of the gas turbine engine is less than 1; and The process involves supplying fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0280] The nvPM emission index ratio for cruise flight with rich combustion can be defined as above in relation to the thirty-fifth aspect.

[0281] According to a forty-eighth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising the following: a combustion chamber for rich burning, rapid cooling, lean burning (RQL combustion chamber) having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; wherein: An nvPM emission index ratio for cruise flight with rich combustion to MAB can be defined as follows: EITravel(bold),NFKEImaxAb,NFKEITravel(bold),FKEImaxAb,FK where: EGG Reise(fett),NFK which can be defined as follows: EI climb, NFK+EI approach, NFK2 and egg Reise(fett),FK which can be defined as follows: EI climb, FK+EI approach, FK2 and whereby: EGG Steig,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 85% of the available thrust under given operating conditions and if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Anflug,NFK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EI Steig,NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Steig,FKThe system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI Steig,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; EGG Anflug,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EI Steig,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; EGG maxAb,NFKThe system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI Steig,NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; and EGG maxAb,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 100% of the available thrust under the same operating conditions as those in which EI Steig,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; the nvPM emission index ratio of cruise flight with rich combustion to MAB of the gas turbine engine is less than 1; and The process involves supplying fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0282] The nvPM emission index ratio of cruise flight with rich combustion to MAB can be defined as above in connection with the thirty-fifth aspect.

[0283] According to a forty-ninth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising the following: a combustion chamber for rich burning, rapid cooling, lean burning (RQL combustion chamber) having a number of fuel spray nozzles in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; wherein: An nvPM emission index ratio of idle to cruise flight with rich combustion can be defined as follows: EILeer,NFKEIReise(fett),NFKEILeer,FKEIReise(fett),FK where: EGG Leer,NFK the nvPM emission index of the gas turbine engine in mg / kg, corrected for system losses, at operation with approximately 7% of the available thrust under given operating conditions, and if a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel (SAF); EGG Leer,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 7% of the available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; EGG Reise(fett),NFK which can be defined as follows: EI climb, NFK+EI approach, NFK2 EGG Reise(fett),FK which can be defined as follows: EI climb, FK+EI approach, FK2 EGG Steig,NFKThe system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI Leer‚NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; EGG Steig,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 85% of available thrust under the same operating conditions as those under which EI Leer‚NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; EGG Anflug,NFK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EILeer‚NFK is calculated when a fuel supplied to the fuel spray nozzles includes a sustainable aviation fuel; and EGG Anflug,FK The system loss-corrected nvPM emission index of the gas turbine engine in mg / kg at operation with approximately 30% of the available thrust under the same operating conditions as those under which EI Leer‚NFK is calculated when the fuel supplied to the fuel spray nozzles is a fossil-based hydrocarbon fuel; the nvPM emission index ratio of idle to cruise flight with rich combustion of the gas turbine engine is less than 1; and The process involves supplying fuel comprising a sustainable aviation fuel (SAF) to the fuel spray nozzles.

[0284] The nvPM emission index ratio of idle to cruise flight with rich combustion can be defined as above in connection with the thirty-fifth aspect.

[0285] In each of the aspects defined above, one or more emission indices of the gas turbine engine can be defined as follows: EGG maxAb 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. EGG maxAb,NFK 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. EGG maxAb,NFK 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. EGG maxAb,FKcan 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. EGG Steig 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. EGG Steig,NFK can be in the range of 1.82 to 58.6 mg / kg and preferably in the range of 2.05 to 53.7 mg / kg and more, preferably in the range of 2.28 to 48.9 mg / kg. EGG Steig,NFK can be in the range of 1.82 to 123 mg / kg and preferably in the range of 2.05 to 113 mg / kg and more, preferably in the range of 2.28 to 102 mg / kg. EGG Steig,FK 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. EGG Anflugcan 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. EGG Anflug,NFK 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. EGG Anflug,NFK can 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. EGG Anflug,FK 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 more preferably in the range of 0.267 to 14.7 mg / kg. EGG Leer can 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. EGG Leer,MKcan 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. EGG Leer,MK can be in the range of 0.118 to 41.3 mg / kg and preferably in the range of 0.132 to 37.9 mg / kg and more preferably in the range of 0.147 to 34.4 mg / kg. EGG Leer,FK can be in the range of 1.23 to 41.4 mg / kg and preferably in the range of 1.38 to 38.0 mg / kg and more preferably in the range of 1.54 to 34.5 mg / kg.

[0286] The following statements can apply to each of the aspects defined above, from one to forty-nine: The fuel spray nozzles can include one or more duplex nozzles and one or more single-stream nozzles. The combustion chamber can contain 10-14 duplex fuel spray nozzles. The combustion chamber can contain 4-8 single-stream fuel spray nozzles.

[0287] The duplex fuel spray nozzles can be arranged in groups around the circumference of the combustion chamber. The groups of duplex fuel spray nozzles can comprise at least two groups arranged diametrically opposite each other. Each group of duplex fuel spray nozzles can contain 2-8 nozzles.

[0288] The combustion chamber can include one or more igniters. The igniter(s) can be located adjacent to one or more of the duplex fuel spray nozzles.

[0289] The number of fuel spray nozzles per unit of engine core size can range from 2.5 to 4.5. The number of fuel spray nozzles per unit of engine core size can range from 3 to 4.

[0290] The fuel supplied to the combustion chamber can contain NFK (non-fuel fuel) in the range of 50% to 100%. The fuel supplied to the combustion chamber can contain NFK in the range of 70% to 100%. The fuel supplied to the combustion chamber can contain NFK in the range of 90% to 100%.

[0291] All the features revealed above in connection with one aspect can be combined with the feature of another aspect, provided they do not mutually exclude each other.

[0292] As stated elsewhere herein, the present disclosure may be applicable to a relevant configuration of a gas turbine engine. Such a gas turbine engine may be, for example, a turbofan engine, an open-rotor engine (in which the propeller is not enclosed by a nacelle), a turboprop engine, or a turbojet engine. Each such engine may or may not be equipped with an afterburner. Such a gas turbine engine may, for example, be configured for power generation applications on land or at sea.

[0293] A gas turbine engine according to the present disclosure may comprise an engine core that includes a turbine, a combustion chamber, a compressor, and a core shaft connecting the turbine to the compressor. Such a gas turbine engine may include a fan (with fan blades). Such a fan may be arranged upstream of the engine core. Alternatively, in some examples, the gas turbine engine may include a fan arranged downstream of the engine core, for example, if the gas turbine engine has an open rotor or is a turboprop engine (in which case the fan may be referred to as a propeller).

[0294] An engine according to the present disclosure can be a gas turbine engine. Such an engine can be a direct-drive turbofan engine in which the fan is connected directly to the fan drive turbine via a core shaft, for example, without a gearbox. In such a direct-drive turbofan engine, the rotational speed of the fan can be expressed as being equal to that of the fan drive turbine. By way of example only, the fan drive turbine can be a first turbine, the core shaft can be a first core shaft, and the gas turbine engine can further comprise a second turbine and a second core shaft connecting the second turbine to the compressor. The second turbine, the compressor, and the second core shaft can be arranged to rotate at a higher speed than the first core shaft. In such an arrangement, the second turbine can be positioned axially upstream of the first turbine.The compressor can be a first compressor, and the gas turbine engine can further include a second compressor. The gas turbine engine can also 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.

[0295] An engine according to the present disclosure can be a geared gas turbine engine. In such an arrangement, the engine has a fan that is driven via a gearbox. Accordingly, such a gas turbine engine can include a gearbox that receives a drive from the core shaft and provides an output to the fan such that the fan is driven at a lower speed than the core shaft. The drive to the gearbox can be directly from the core shaft or indirectly from the core shaft, for example via a spur gear shaft and / or a gear. The core shaft can rigidly connect the turbine and the compressor so that the turbine and compressor rotate at the same speed (with the fan rotating at a lower speed).

[0296] The gas turbine engine, as described and / or claimed herein, may have any suitable general architecture. For example, the gas turbine engine may have any number of shafts connecting turbines and compressors, for example, one, two, or three shafts. 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, the second compressor, and the second core shaft may be arranged to rotate at a higher speed than the first core shaft.

[0297] In such an arrangement, the second compressor can be positioned axially downstream of the first compressor. The second compressor can be arranged to receive a flow from the first compressor (for example, by receiving it directly, or via a generally annular channel).

[0298] The gearbox can be arranged to be driven by the core shaft configured to rotate at the lowest speed (for example, in operation) (for example, the first core shaft in the example above). Alternatively, the gearbox can be arranged to be driven only by the core shaft configured to rotate at the lowest speed (for example, in operation) (for example, only the first core shaft and not the second core shaft in the example above). Finally, the gearbox can be arranged to be driven by any one or any number of shafts, for example, the first and / or second shaft in the example above.

[0299] The gearbox can be a reduction gearbox (by having a lower speed at the fan output than at the input from the main shaft). Any type of gearbox can be used. For example, the gearbox can be a planetary gearbox or a star gearbox, as described in more detail elsewhere herein. Such a gearbox can be single-stage. Alternatively, such a gearbox can be a summing gearbox, for example, a compound planetary gearbox (where the input is at the sun gear and the output at the ring gear, and which is therefore also called a compound star gearbox), e.g., with two reduction stages.

[0300] The transmission can have any desired gear ratio (defined as the speed of the input shaft divided by the speed of the output shaft). For example, the gear ratio can be greater than or one of the following values: 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 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, the gear ratio can be greater than 2.5, but it can also be in the range of 3.0 to 4.2 or 3.2 to 3.8. For example, the transmission could be a star gear with a gear ratio in the range of 3.1 or 3.2 to 3.8. Also for example, the transmission could be a star gear with a gear ratio in the range of 3.0 to 3.1.The transmission could also be, purely by way of example, a star gear with a gear ratio in the range of 3.3 to 3.6 or from 3.6 to 4.2. Also purely by way of example, the transmission could be a summing gear, for example a summing star gear, with 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 summing gear, for example a summing star gear, can lie within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​can form upper or lower limits), for example, in the range of 4.0 to 14.0, 6.0 to 12.0, or 8.0 to 10.0. In some arrangements, the gear ratio may lie outside these ranges.

[0301] In any gas turbine engine described and / or claimed herein, fuel of a given composition or mixture is supplied to a combustion chamber, which may be located downstream of the fan and compressor(s) with respect to the flow path (for example, axially downstream). For example, the combustion chamber may be located directly downstream of the second compressor (for example, at its outlet), provided a second compressor is provided. Also, purely by way of example, the flow may be supplied at the outlet of the combustion chamber to the inlet of the second turbine, provided a second turbine is provided. The combustion chamber may be located upstream of the turbine(s).

[0302] The compressor(s) (for example, the first compressor and the second compressor, as described above) can comprise any number of stages, for example, multiple stages. Each stage can include a set of rotor blades and a set of stator blades, which can be variable stator blades (since their angle of incidence can be variable). The set of rotor blades and the set of stator blades can be axially offset from each other. For example, the gas turbine engine can be a direct-drive turbofan gas turbine engine comprising 11, 12, 13, 14, or 15 compressor stages (in addition to the fan). Such an engine can, for example, have 3 stages in the first (or "low-pressure") compressor and either 8, 9, 10, or 11 stages in the second (or "high-pressure") compressor.Alternatively, such an engine could, for example, have 4 stages in the first (or "low-pressure") compressor and either 8, 9, 10, or 11 stages in the second (or "high-pressure") compressor. Alternatively, such an engine could, for example, have 7, 8, or 9 stages in a first (or "intermediate-pressure") compressor and 5, 6, or 7 stages in the second (or "high-pressure") compressor. Also purely by way of example, the gas turbine engine could be a "geared" gas turbine engine (in which the fan is driven by a first core shaft via a reduction gearbox) that includes 11, 12, 13, 14, or 15 compressor stages (in addition to the fan). Such an engine could have 3 or 4 stages in the first (or "low-pressure") compressor and 8, 9, or 10 stages in the second (or "high-pressure") compressor.As another example, the gas turbine engine can be a "geared" gas turbine engine with 4 stages in the first (or "low-pressure") compressor and 10 stages in the second (or "high-pressure") compressor.

[0303] The turbine, or each turbine (for example, the first turbine and the second turbine, as described above), can comprise any number of stages, for example, multiple stages. Each stage can comprise a set of rotor blades and a set of stator blades, or vice versa, as required. The respective sets of rotor blades and stator blades can be axially offset from one another. In each stage, the set of rotor blades can be located downstream of the respective set of stator blades. By way of example only, the gas turbine engine can comprise 5, 6, 7, 8, or 9 turbine stages. For example, the gas turbine engine can be a geared gas turbine engine comprising 5, 6, or 7 turbine stages. Such a geared gas turbine engine can include a second (high-pressure) turbine with 2 stages. Such a geared gas turbine engine can also include a first (or low-pressure) turbine with 3 or 4 stages.For example, a gas turbine engine could be a direct-drive gas turbine engine comprising a first (or "low-pressure") turbine with 3, 4, 5, 6, or 7 stages. Such a direct-drive gas turbine engine could include a second (or "high-pressure") turbine with 2 stages. Alternatively, such an engine could, for example, include a second (or "intermediate-pressure") turbine with 1, 2, or 3 stages. Such a direct-drive gas turbine engine could also include a third (or "high-pressure") turbine with 1, 2, or 3 stages. Again, for example, the first (or "low-pressure") turbine could have 6 stages, and the third (or "high-pressure") turbine could have 1 stage.

[0304] Each fan blade can be defined as having a radial span extending from a root (or hub) at a radially inner 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 can be smaller than or can be one of the following values: 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 fan blade at the hub to the radius of the fan blade at the tip can lie within an enclosing 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 of 0.29 to 0.30.These ratios can be generally referred to as the hub-to-tip ratio. As a non-restrictive example, the hub-to-tip ratio can range from 0.40 to 0.50, from 0.42 to 0.48, or from 0.43 to 0.47. The hub radius and the tip radius can both be measured at the leading edge (or axially foremost portion) of the fan blade. The hub-to-tip ratio refers, of course, to the gas-washed portion of the fan blade, i.e., the section radially outside any platform.

[0305] The radius of the fan can be measured between the centerline of the engine and the tip of a fan blade at its leading edge. The fan diameter (which can simply be twice the fan radius) can be larger than or any of the following values: 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). inch), 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 fan diameter 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 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. As a purely non-restrictive example, the fan diameter can range from 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.

[0306] The fan's rotational speed can vary during use. Generally, fans with a larger diameter have a lower rotational speed. As a purely non-restrictive example, the fan speed under cruise flight conditions can be lower than or any of the following values: 3500 rpm, 3450 rpm, 3400 rpm, 3350 rpm, 3300 rpm, 3250 rpm, 3200 rpm, 3150 rpm, 3100 rpm, 3050 rpm, 3000 rpm, 2950 rpm, 2900 rpm, 2850 rpm, 2800 rpm, 2750 rpm, 2700 rpm, 2650 rpm, 2600 rpm, 2550 rpm, 2500 rpm, 2450 rpm, 2400 rpm, 2350 rpm 2300 rpm, 2250 rpm, 2200 rpm, 2150 rpm, 2100 rpm, 2050 rpm, 2000 rpm, 1950 rpm, 1900 rpm, 1850 rpm, 1800 rpm, 1750 rpm, 1700 rpm, 1650 rpm, 1600 rpm, 1550 rpm, 1500 rpm, 1450 rpm, 1400 rpm, 1350 rpm, 1300 rpm, 1250 rpm, 1200 rpm, 1150 rpm, 1100 rpm, 1050 rpm, 1000 rpm or 950 rpm.The fan speed under cruise conditions can lie 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). Purely as another non-restrictive example, the fan speed under cruise conditions for a geared gas turbine engine with a fan diameter in the range of 200 cm to 210 cm could be in the range of 2750 to 2900 rpm, 2750 to 2800 rpm, or 2800 to 2900 rpm. Purely as another non-restrictive example, the fan speed under cruise conditions for a geared gas turbine engine with a fan diameter in the range of 210 cm to 230 cm could be in the range of 2500 to 2800 rpm or 2500 to 2750 rpm.As a further non-restrictive example, the fan speed under cruise conditions for a geared gas turbine engine with a fan diameter of 340 cm to 360 cm could be between 1500 rpm and 1800 rpm or between 1500 rpm and 1700 rpm. As a further non-restrictive example, the fan speed under cruise conditions for a geared gas turbine engine with a fan diameter of 380 cm to 420 cm could be between 950 rpm and 1500 rpm, between 950 rpm and 1200 rpm, between 950 rpm and 1100 rpm, between 950 rpm and 1050 rpm, between 950 rpm and 1000 rpm, or between 1000 rpm and 1050 rpm. As just another non-restrictive example, the fan speed under cruise flight 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 3400 to 4600 rpm, for example from 3600 to 4600 rpm or from 3600 to 3900 rpm.As a further non-restrictive example, the fan speed under cruise flight conditions for a direct-drive engine with a fan diameter in the range of 300 cm to 350 cm could be between 1800 and 3200 rpm or between 1950 and 2900 rpm. As a further non-restrictive example, the fan speed under cruise flight conditions for a direct-drive engine with a fan diameter in the range of 270 cm to 290 cm could be between 1800 and 2900 rpm or between 2050 and 2700 rpm. As a further non-restrictive example, the fan speed under cruise flight conditions for a direct-drive engine with a fan diameter in the range of 290 cm to 310 cm could be between 1800 and 2950 rpm or between 2100 and 2700 rpm.

[0307] In the operation of a gas turbine engine, the fan (with its associated fan blades) rotates around an axis of rotation. This rotation causes the tip of the fan blade to move at a velocity UStip. The work done by the fan blades on the stream results in a rise dH in the specific enthalpy of the stream. A fan tip load can be defined as dH / UStip², where dH is the rise in specific enthalpy (e.g., the average 1-D rise in specific enthalpy) across the fan, and UStip is the (translational) velocity of the fan tip, for example, at the leading edge of the tip (which can be defined as the fan tip radius at the leading edge multiplied by the angular velocity).The fan peak load under cruise flight conditions can be greater than or can be any of the following values: 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 (where all values ​​are dimensionless). The fan 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, in the range of 0.15 to 0.20, 0.28 to 0.35, 0.29 to 0.35, 0.29 to 0.30, or 0.30 to 0.35 (for example, for a geared gas turbine engine).

[0308] Gas turbine engines according to the present disclosure can have a desired bypass ratio (BPV), wherein the bypass ratio is defined as the ratio of the mass flow rate of the current through the bypass channel to the mass flow rate of the current through the core. In some configurations, the bypass ratio under cruise flight conditions may be greater than or any of the following values: 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 cruise flight conditions may be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e.,(The values ​​can form upper or lower limits), for example, in the range of 3.5 to 6.5, 4 to 6, 5 to 6, 12 to 16, 13 to 15, or 13 to 14. Purely as a non-limiting example, the bypass ratio under cruise flight conditions of a direct-drive gas turbine engine according to the present disclosure can be in 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. Purely as another non-limiting example, the bypass ratio under cruise flight conditions of a geared gas turbine engine according to the present disclosure can be in the range of 10 to 12, 12 to 15, or 12.5 to 15. The bypass channel can be at least substantially annular. The bypass duct can be located radially outside the core engine. The radial outer surface of the bypass duct can be defined by a nacelle and / or a fan housing.

[0309] The overall pressure ratio (OPR) of a gas turbine engine, as described and / or claimed herein, can be defined as the ratio of the dynamic pressure at the outlet of the high-pressure compressor (before entering the combustion chamber) to the dynamic pressure upstream of the fan. As a non-restrictive example, the total pressure ratio of a gas turbine engine, as described and / or claimed herein, under cruise flight conditions may be greater than or may be any of the following values: 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 total pressure ratio under cruise flight 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, in the range of 30 to 40 or 50 to 70.As a purely non-restrictive example, the total pressure ratio under cruise flight 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. As a purely non-restrictive example, the total pressure ratio under cruise flight 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. As a purely non-restrictive example, the total pressure ratio under cruise flight 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. As a purely non-restrictive example, the total pressure ratio under cruise flight 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.As a purely non-restrictive example, the overall pressure ratio under cruise flight 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. As a purely non-restrictive example, the overall pressure ratio under cruise flight 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. As a purely non-restrictive example, the overall pressure ratio under cruise flight conditions of a direct-drive gas turbine engine with a fan diameter in the range of 300 cm to 350 cm can be in the range of 35 to 60 or 40 to 50. As a purely non-restrictive example, the overall pressure ratio under cruise flight conditions of a direct-drive gas turbine engine with a fan diameter in the range of 260 cm to 285 cm can be in the range of 35 to 60 or from 35 to 50.As a further non-restrictive example, the overall pressure ratio under cruise flight conditions for 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 35 to 60 or from 37 to 47. As a further non-restrictive example, the overall pressure ratio under cruise flight conditions for 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 35 to 60 or from 37 to 47.

[0310] 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, the specific thrust for a given thrust condition may depend on the specific composition of the fuel supplied to the combustion chamber.

[0311] Under cruise flight conditions, the specific thrust of an engine described and / or claimed herein may be less than or may be one of the following values: 150 Nkg -1 s, 145 Nkg -1 s, 140 Nkg -1 s, 135 Nkg -1 s, 130 Nkg -1 s, 125 Nkg -1 s, 120 Nkg - 1 s, 115 Nkg -1 s, 110 Nkg-1s, 105 Nkg-1s, 100 Nkg-1s, 99 Nkg-1s, 98 Nkg-1s, 97 Nkg-1s, 96 Nkg-1s, 95 Nkg-1s, 94 Nkg-1s, 93 Nkg-1s, 92 Nkg-1s, 91 Nkg-1s, 90 Nkg-1s, 89 Nkg-1s, 88 Nkg-1s, 87 Nkg-1s, 86 Nkg-1s, 85 Nkg-1s, 80 Nkg-1s, 75 Nkg -1 s, 70 Nkg - 1 s, 65 Nkg -1 s, 60 Nkg -1 s, 55 Nkg -1 s, 50 Nkg -1 s, 45 Nkg -1 s, 40 Nkg -1 s, 35 Nkg -1 s, 30 Nkg - 1 s, 25 Nkg -1 s, 20 Nkg -1 s or 15 Nkg -1The specific thrust under cruise flight conditions can lie within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​can form upper or lower limits), for example, in the range of 80 Nkg-1s to 100 Nkg-1s, 85 Nkg-1s to 100 Nkg-1s, or 92 Nkg-1s to 100 Nkg-1s. Such engines can be particularly efficient compared to conventional gas turbine engines. Purely as a non-restrictive example, the specific thrust under cruise flight conditions of a geared gas turbine engine with a fan diameter in the range of 200 cm to 210 cm can be in the range of 90 Nkg-1s to 98 Nkg-1s or 92 Nkg-1s to 98 Nkg-1s. As a purely non-restrictive example, the specific thrust under cruise flight conditions of a geared gas turbine engine with a fan diameter in the range of 210 cm to 230 cm can be in the range of 90 Nkg-1s to 100 Nkg-1s or from 95 Nkg-1s to 100 Nkg-1s.As a purely non-restrictive example, the specific thrust under cruise flight conditions of a geared gas turbine engine with a fan diameter in the range of 340 cm to 360 cm could be in the range of 70 Nkg-1s to 95 Nkg-1s, 80 Nkg-1s to 95 Nkg-1s, or 85 Nkg-1s to 95 Nkg-1s. Similarly, as a purely non-restrictive example, the specific thrust under cruise flight conditions of a direct-drive gas turbine engine with a fan diameter in the range of 170 cm to 200 cm could be in the range of 20 Nkg-1s to 90 Nkg-1s, 20 Nkg-1s to 80 Nkg-1s, or 25 Nkg-1s to 70 Nkg-1s. As a purely non-restrictive example, the specific thrust under cruise flight conditions of a direct-drive gas turbine engine with a fan diameter in the range of 300 cm to 350 cm can be in the range of 90 Nkg-1s to 120 Nkg-1s or from 100 Nkg-1s to 115 Nkg-1s.As a purely non-restrictive example, the specific thrust under cruise flight 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-1s to 120 Nkg-1s, 30 Nkg-1s to 115 Nkg-1s, or 40 Nkg-1s to 115 Nkg-1s. As a purely non-restrictive example, the specific thrust under cruise flight 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-1s to 120 Nkg-1s or 95 Nkg-1s to 115 Nkg-1s. As a purely non-restrictive example, the specific thrust under cruise flight conditions 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 95 Nkg-1s to 130 Nkg-1s or from 105 Nkg-1s to 125 Nkg-1s.

[0312] A gas turbine engine, as described and / or claimed herein, can produce any desired maximum thrust. Purely as a non-restrictive example, a gas turbine such as described and / or claimed herein may be capable of generating a maximum thrust of at least or exactly 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, 130kN, 131kN, 132kN, 133kN, 134kN, 135kN, 136kN, 137kN, 138kN, 139kN, 140kN, 141kN, 142kN, 143kN, 144kN, 145kN, 146 kN, 147 kN, 148 kN, 149 kN, 150 kN, 151 kN, 152kN, 153 kN, 154 kN, 155 kN, 160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350kN, 375kN, to generate 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 lie 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). Purely as a non-restrictive example, a gas turbine such as described and / or claimed herein may be capable of generating a maximum thrust in the range of 50 kN to 85 kN, 57 kN to 78 kN, 60 kN to 73 kN, 60 kN to 70 kN, 65 kN to 150 kN, 105 kN to 150 kN, 155 kN to 170 kN, 330 kN to 420 kN, or 350 kN to 400 kN. As a purely non-restrictive example, the maximum thrust 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 130 kN to 160 kN or from 130 kN to 150 kN.As a purely non-restrictive example, the maximum thrust of a geared gas turbine engine with a fan diameter in the range of 210 cm to 230 cm could 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 to 145 kN. As a purely non-restrictive example, the maximum thrust of a geared gas turbine engine with a fan diameter in the range of 340 cm to 360 cm could be in the range of 370 kN to 500 kN. As a purely non-restrictive 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.As a purely non-restrictive example, the maximum thrust of a direct-drive gas turbine engine with a fan diameter in the range of 110 cm to 120 cm could be in the range of 60 kN to 70 kN or from 60 kN to 65 kN. As a purely non-restrictive example, the maximum thrust of a direct-drive gas turbine engine with a fan diameter in the range of 120 cm to 130 cm could be in the range of 65 kN to 75 kN or from 67 kN to 73 kN. As a purely non-restrictive 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.As a purely non-restrictive example, the maximum thrust of a direct-drive gas turbine engine with a fan diameter in the range of 260 cm to 285 cm could be in the range of 250 kN to 450 kN, 250 kN to 400 kN, or 250 kN to 350 kN. Similarly, as a purely non-restrictive example, the maximum thrust of a direct-drive gas turbine engine with a fan diameter in the range of 270 cm to 290 cm could be in the range of 250 kN to 500 kN, 275 kN to 400 kN, or 275 kN to 375 kN. As a purely non-restrictive example, the maximum thrust of a direct-drive gas turbine engine with a fan diameter of 290 cm to 310 cm can range from 250 kN to 550 kN, from 300 kN to 525 kN, or from 300 kN to 500 kN. The thrust mentioned above can be the maximum net thrust under normal atmospheric conditions at sea level at +15°C (ambient pressure 101.3 kPa, temperature 30°C) with the engine running statically.

[0313] In operation, the flow temperature at the inlet to the high-pressure turbine can be particularly high. This temperature, which can be referred to as TET, can be measured at the outlet to the combustion chamber, for example, immediately upstream of the first turbine guide vane, which itself can be called the nozzle guide vane. In some cases, the TET for a given thrust condition may depend on the specific composition of the fuel supplied to the combustion chamber.

[0314] Under cruise flight conditions, the TET can be at least one of 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 cruise conditions can lie 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). Thus, purely as a non-restrictive example, the TET under cruise conditions of a geared gas turbine engine with a fan diameter in the range of 200 cm to 210 cm can lie in the range of 1540 K to 1600 K or 1570 K to 1590 K.As a purely non-restrictive example, the TET under cruise flight conditions of a geared gas turbine engine with a fan diameter in the range of 210 cm to 230 cm could be in the range of 1525 K to 1650 K. As a purely non-restrictive example, the TET under cruise flight conditions of a geared gas turbine engine with a fan diameter in the range of 340 cm to 360 cm could be in the range of 1550 K to 1660 K or 1550 K to 1600 K. As a purely non-restrictive example, the TET under cruise flight conditions of a direct-drive gas turbine engine with a fan diameter in the range of 170 cm to 200 cm could be in the range of 1400 K to 1650 K or 1425 K to 1625 K. As a purely non-restrictive example, the TET under cruise flight conditions of a direct-drive gas turbine engine with a fan diameter in the range of 300 cm to 350 cm can be in the range of 1400 K to 1650 K, from 1425 K to 1575 K, from 1425 K to 1525 K or from 1475 K to 1550 K.As a purely non-restrictive example, the TET under cruise flight 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 1400 K to 1650 K or from 1425 K to 1625 K. As a purely non-restrictive example, the TET under cruise flight 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 1400 K to 1600 K, from 1425 K to 1575 K, or from 1450 K to 1550 K. As a purely non-restrictive example, the TET under cruise conditions 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 1400 K to 1650 K, 1450 K to 1600 K, or 1475 K to 1575 K. The TET under cruise conditions can be in an inclusive range bounded by any two of the values ​​in this paragraph (i.e.,, the values ​​can form upper or lower limits), for example 1530 K to 1600 K.

[0315] The maximum TET in engine use can be at least one of the following: 1700 K, 1750 K, 1755 K, 1760 K, 1765 K, 1770 K, 1775 K, 1780 K, 1785 K, 1790 K, 1795 K, 1800 K, 1805 K, 1810 K, 1815 K, 1820 K, 1825 K, 1830 K, 1835 K, 1840 K, 1845 K, 1850 K, 1855 K, 1860 K, 1865 K, 1870 K, 1875 K, 1880 K, 1885 K, 1890 K, 1895 K, 1900 K, 1905 K The maximum TET during engine use can be 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 use can be within an inclusive range bounded by any two of the TET values ​​in this paragraph (i.e., the values ​​can form upper or lower limits).Thus, purely as a non-restrictive example, the maximum TET of a geared gas turbine engine with a fan diameter in the range of 200 cm to 210 cm could 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. Similarly, as a non-restrictive example, the maximum TET of a geared gas turbine engine with a fan diameter in the range of 210 cm to 230 cm could be in the range of 1700 K to 1960 K, 1750 K to 1900 K, 1750 K to 1850 K, or 1750 K to 1800 K. As a purely non-restrictive example, the maximum TET 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 1800 K to 1960 K, 1800 K to 1900 K or 1850 K to 1900 K.As a purely non-restrictive example, the maximum TET of a direct-drive gas turbine engine with a fan diameter in the range of 170 cm to 200 cm could be in the range of 1700 K to 1950 K or from 1750 K to 1900 K. As a purely non-restrictive example, the maximum TET of a direct-drive gas turbine engine with a fan diameter in the range of 300 cm to 350 cm could be in the range of 1700 K to 1950 K, from 1750 K to 1900 K, or from 1775 K to 1900 K. As a purely non-restrictive example, the maximum TET of a direct-drive gas turbine engine with a fan diameter in the range of 260 cm to 285 cm could be in the range of 1700 K to 1950 K or from 1750 K to 1900 K. As a purely non-restrictive example, the maximum TET 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 1750 K to 1950 K or from 1800 K to 1900 K.As a purely non-restrictive 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 be in the range of 1750 K to 2000 K or from 1800 K to 1950 K. The maximum TET can occur, for example, under high thrust conditions, such as maximum takeoff conditions (MAB conditions).

[0316] A gas turbine engine, as described and / or claimed herein, can have any desired high-pressure compressor size, also known as the core size. The core size defines the size of the engine core. Engine core size can be defined as follows: Core size = m˙2T3P3 where ṁ2 = the mass flow rate of air entering the high-pressure compressor in pounds per second, T3 = the temperature of air exiting the high-pressure compressor in Kelvin, and P3 = the pressure of air exiting the high-pressure compressor in pounds per second squared per inch. One unit of the core quantity is therefore expressed as: s⋅K12⋅in Under cruise conditions, the core size can be at least or exactly 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 core size under cruise 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). Thus, purely as a non-restrictive example, the core size under cruise flight 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 3 to 8, 3.5 to 7, 4 to 6.5, 4.5 to 6.5, or 5 to 6. Purely as a non-restrictive example, the core size under cruise flight 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 7 to 15, 8 to 14, 9 to 13, 10 to 12.5, or 11 to 12.As a purely non-restrictive example, the core size under cruise flight 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 5 to 13, 6.5 to 12.5, 7.5 to 11.5, 8.5 to 10.5 or 9 to 10.

[0317] As a purely non-restrictive example, the core size under cruise flight conditions of a direct-drive gas turbine engine with a fan diameter in the range of 110 cm to 130 cm could 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. Similarly, as a purely non-restrictive example, the core size under cruise flight conditions of a direct-drive gas turbine engine with a fan diameter in the range of 170 cm to 200 cm could 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. As a purely non-restrictive example, the core size under cruise flight conditions of a direct-drive gas turbine engine with a fan diameter in the range of 300 cm to 350 cm can be in the range of 9 to 17, 10 to 16, 11 to 16, 13 to 14.5 or 13 to 14.As a purely non-restrictive example, the core size under cruise flight conditions of a direct-drive gas turbine engine with a fan diameter in the range of 260 cm to 285 cm could 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. As a purely non-restrictive example, the core size under cruise flight conditions of a direct-drive gas turbine engine with a fan diameter in the range of 270 cm to 290 cm could be in the range of 9 to 13, 10.5 to 12.5, or 11 to 12. As a purely non-restrictive example, the core size under cruise flight conditions 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 11 to 16, 12 to 15, 13 to 15, or 13 to 14.5. The core size under cruise flight conditions can be within an inclusive range bounded by any two of the above values ​​for the core size (i.e.,, the values ​​can form upper or lower limits), for example 4.5 to 9.5.

[0318] A fan blade and / or an air guide section of a fan blade, as described and / or claimed herein, may be made of any suitable material or combination of materials. For example, at least a portion of the fan blade and / or air guide section may be made at least partially of a composite material, such as a metal matrix composite and / or an organic matrix composite, such as a carbon fiber composite. As another example, at least a portion of the fan blade and / or air guide section may be made at least partially of a metal, such as a titanium-based metal, an aluminum-based material (such as an aluminum-lithium alloy), or a steel-based material. The fan blade may comprise at least two sections manufactured using different materials.For example, the fan blade can have a protective leading edge, which may be made using a material better able to withstand impacts (e.g., from birds, ice, or other materials) than the rest of the blade. Such a leading edge could, for example, be made using titanium or a titanium-based alloy. Thus, purely by way of 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 fan blade can weigh at least or exactly 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 lie within an inclusive range bounded by any two of the values ​​for the fan blade weight in the preceding sentence (i.e., the values ​​can form upper or lower bounds). For example, the weight of the fan blade for a geared gas turbine engine with 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, 4.0 kg to 6.0 kg, or 5.0 kg to 5.5 kg.

[0319] A fan as described and / or claimed herein may comprise a central section from which the fan blades may extend, for example, in a radial direction. The fan blades may be attached to the central section in any desired manner. For example, each fan blade may include a fastening that can engage in a corresponding slot in the hub (or disc). Such a fastening may, purely by way of example, be in the form of a dovetail that can be inserted and / or snapped into a corresponding slot in the hub / disc to secure the fan blade to the hub / disc. As another example, the fan blades may be formed in one piece with a central section. Such an arrangement may be referred to as a blade disc or blade ring. Any suitable method may be used to manufacture such a blade disc or blade ring.For example, at least part of the fan blades may be machined from a single block and / or at least part of the fan blades may be attached to the hub / disc by welding, such as linear friction welding.

[0320] 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 can allow the bypass channel outlet area to be varied during operation. The general principles of this disclosure can be applied to engines with or without a VAN.

[0321] The fan of a gas turbine, as described and / or claimed herein, can have any number of fan blades, for example, 14, 16, 18, 20, 22, 24, or 26 fan blades. If the fan blades have a carbon fiber composite body, there can be 14, 16, or 18 fan blades. If the fan blades have a metallic body (for example, an aluminum-lithium or titanium alloy), there can be 18, 20, or 22 fan blades.

[0322] If the gas turbine engine has an open rotor or is a turboprop engine, it can include two counter-rotating propeller stages connected via a shaft to, and driven by, a free-running turbine. The propellers can rotate in opposite directions, with one rotating clockwise and the other counterclockwise around the engine's axis of rotation. Alternatively, the gas turbine engine can include a propeller stage and a guide vane stage configured downstream of the propeller stage. The guide vane stage can have a variable pitch. Accordingly, high-pressure, intermediate-pressure, and free-running turbines can each drive high-pressure and intermediate-pressure compressors and propellers via suitable connecting shafts. The propellers can thus provide the majority of the propulsive thrust.

[0323] If the gas turbine engine has an open rotor or is a turboprop engine, one or more of the propeller stages may be driven by a gearbox. The gearbox may be of the type described herein.

[0324] The terms idle, taxiing, takeoff, climb, cruise, descent, approach, and landing (or one or more sections thereof) used herein have their usual meanings and are readily understandable to a person skilled in the art. For example, a person skilled in the art would immediately recognize that, in the case of a specific gas turbine engine for an aircraft, each term refers to the entirety or to one or more parts of an operating phase of the engine within a given use of an aircraft for which the gas turbine engine was designed to be mounted.

[0325] As a purely non-restrictive example, ground idle can refer to an operating phase of the engine in which the aircraft is stationary and in contact with the ground, but with the requirement that the engine is running. For example, during idle, the engine may produce between 3% and 9% of its available thrust. In other non-restrictive examples, the engine may produce between 5% and 8% of its available thrust. In further non-restrictive examples, the engine may produce between 6% and 7% of its available thrust. As a purely non-restrictive example, taxiing can refer to an operating phase of the engine in which 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 its available thrust.In further non-restrictive examples, the engine can generate between 6% and 12% of its available thrust. In further non-restrictive examples, the engine can generate between 7% and 10% of its available thrust. Purely as a non-restrictive example, "takeoff" can refer to an operational phase of the engine in which the aircraft is propelled by the thrust generated by the engine. In an initial stage within the takeoff phase, the aircraft can be propelled while in contact with the ground. In a later stage within the takeoff phase, the aircraft can be propelled while out of contact with the ground. For example, the engine can generate between 90% and 100% of its available thrust during takeoff. In further non-restrictive examples, the engine can generate between 95% and 100% of its available thrust.In other non-restrictive examples, the engine can generate 100% of the available thrust.

[0326] As a purely non-restrictive example, climb can refer to an operational phase of the engine in which the aircraft is propelled by the thrust generated by the engine. For example, during climb, the engine may generate between 75% and 100% of its available thrust. In other non-restrictive examples, the engine may generate between 80% and 95% of its available thrust. In still other non-restrictive examples, the engine may generate between 85% and 90% of its available thrust. For example, climb can refer to an operational phase within an aircraft's flight cycle between takeoff and reaching cruise conditions, with reaching cruise conditions defining the beginning of the cruise phase of the aircraft's flight, or a portion thereof.Additionally or alternatively, the term climb can refer, for example, to a nominal point or one or more nominal periods of time during the flight cycle of an aircraft between takeoff and landing, in which a relative increase in altitude is required, which may necessitate additional thrust from the engine.

[0327] As used herein, the term cruise conditions, which can define the cruise phase of flight, has the conventional meaning and is readily understood by experts. Thus, cruise conditions for a particular gas turbine engine for an aircraft can refer to the engine's operating point at the mid-cruise point of a given mission (which in the industry may be called the "economic mission") of an aircraft for which the gas turbine engine was designed to be attached. In this respect, mid-cruise can be considered the point in an aircraft's flight cycle at which 50% of the total fuel burned between the peak of climb and the start of descent has been consumed (which can be approximated by the midpoint—in terms of time and / or distance—between the peak of climb and the start of descent).Cruise flight conditions can thus define an operating point of the gas turbine engine that provides thrust sufficient to ensure stable operation (i.e., maintaining a constant altitude and / or Mach number) or at least substantially stable operation (i.e., maintaining at least a substantially constant altitude and / or Mach number) at mid-cruise flight of an aircraft for which the gas turbine engine is designed to be fitted, taking into account the number of engines provided for that aircraft.For example, if an engine is designed to be attached to an aircraft that has two engines of the same type, the engine can provide, under cruise conditions, half of the total thrust required for stable operation, or at least substantially stable operation, of the aircraft at mid-cruise.

[0328] In other words, the cruise flight conditions for a given gas turbine engine for an aircraft can be defined as the engine's operating point that provides a certain thrust (required to provide—in combination with any other engines on the aircraft—steady operation, or at least substantially stable operation, of the aircraft for which it is designed to be fitted, at a given mid-cruise Mach number) under atmospheric conditions at mid-cruise (defined by the International Standard Atmosphere according to ISO 2533 at the altitude at mid-cruise). For any given gas turbine engine for an aircraft, the mid-cruise thrust, the atmospheric conditions, and the Mach number are known, and thus the engine's operating point under cruise conditions can be clearly defined.

[0329] As a purely non-restrictive example, the forward speed under cruise conditions can be any point in the range of Mach 0.7 to 0.9, for example, 0.75 to 0.85, 0.76 to 0.84, 0.77 to 0.83, 0.78 to 0.82, 0.79 to 0.81, in the order of Mach 0.78, 0.79, or 0.8, in the order of Mach 0.85, or in the range of 0.8 to 0.85. Any single speed within these ranges can be part of the cruise conditions. For some aircraft, the cruise conditions may lie outside these ranges, for example, below Mach 0.7 or above Mach 0.9.

[0330] Purely as a non-restrictive example, cruise flight conditions can correspond to standard atmospheric conditions (according to the International Standard Atmosphere, ISA) at an altitude in the range of 10,000 m to 15,000 m, for example in the range of 10,000 m to 12,000 m, for example in the range of 10,400 m to 11,600 m (approximately 38,000 ft), for example in the range of 10,500 m to 11,500 m, for example in the range of 10,600 m to 11,400 m, for example in the range of 10,700 m (approximately 35,000 ft) to 11,300 m, for example in the range of 10,800 m to 11,200 m, for example in the range of 10,900 m to 11,100 m, for example in the order of 11,000 m. Cruise flight conditions can correspond to standard atmospheric conditions at any given altitude in these areas.

[0331] As a purely non-restrictive example, cruise flight conditions could correspond to a forward Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35,000 feet (10,668 m). Under such cruise flight conditions, the engine can provide a known required net thrust level. This known required net thrust level naturally depends on the engine and its intended application and could, for example, be a value in the range of 20 kN to 40 kN.

[0332] As a purely illustrative example, cruise flight conditions could correspond to a forward Mach number of 0.85 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38,000 feet (11,582 m). Under such cruise flight conditions, the engine can provide a known required net thrust level. This known required net thrust level naturally depends on the engine and its intended use and could, for example, be a value in the range of 35 kN to 65 kN.

[0333] In use, a gas turbine engine as described and / or claimed herein can operate under cruise flight conditions defined elsewhere herein. Such cruise flight conditions may be determined by the cruise flight conditions (for example, mid-cruise conditions) of an aircraft on which at least one (for example, two or four) gas turbine engine(s) can be mounted to provide thrust.

[0334] Furthermore, a person skilled in the art would immediately recognize that descent and approach, or both, refer to an operational phase within a flight cycle between cruise and landing of the aircraft, with the approach, in particular, forming part of the landing and takeoff (LTO) phase. As a purely non-restrictive example, the engine may generate less than 50% of its available thrust during one or both of the descent and approach phases. In other non-restrictive examples, the engine may generate between 25% and 40% of its available thrust. In further non-restrictive examples, the engine may generate between 30% and 35% of its available thrust. Additionally or alternatively, the term descent may refer to a nominal point in an aircraft's flight cycle between takeoff and landing where a relative decrease in altitude is required and which may necessitate a reduced thrust requirement from the engine.

[0335] According to one aspect, an aircraft comprising a gas turbine engine as described and / or claimed herein is provided. The aircraft according to this aspect is the aircraft for which the gas turbine engine has been designed to be attached. Accordingly, the cruise flight conditions according to this aspect may correspond to the cruise flight center of the aircraft as defined elsewhere herein.

[0336] According to one aspect, a method for operating a gas turbine engine as described and / or claimed herein is provided. Operation can take place under any suitable conditions, such as may be defined elsewhere herein (for example, with regard to thrust, atmospheric conditions, and Mach number).

[0337] According to one aspect, a method for operating an aircraft comprising a gas turbine engine as described and / or claimed herein is provided. Operation according to this aspect may include (or be) operation under any suitable condition, for example, at the aircraft's mid-cruise flight, as defined elsewhere herein.

[0338] In operation, under one or more of the operating conditions and / or thrust settings disclosed or described herein, the reduction in the CO2-related mass emission index (EI) provided by one or more of the gas turbine engine configurations disclosed or described herein may be greater than or equal to one of the following values: 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,if the fuel supplied to the combustion chamber includes sustainable aviation fuel or a mixture of sustainable aviation fuel comprising up to 100% sustainable aviation fuel. The CO2 percentage emission index (EI) can be expressed in kg CO2 per kg of fuel. The bypass ratio 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 0.4 to 2.5, 0.42 to 2.08, or 0.43 to 2.08.

[0339] As a purely non-restrictive example, the reduction of the CO2 percentage-related emissions index (EI) can result from the fact that the fuel supplied to the combustion chamber has a hydrogen mass fraction in percent greater than one of the following values: 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.

[0340] The hydrogen mass fraction in percent of the fuel can be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​can form upper or lower limits), for example, in the range of 13.41 to 15.3 or of 13.42 to 15.3.

[0341] In operation, under one or more of the operating conditions and / or thrust settings disclosed or described herein, the reduction in the CO2 percentage per MJ of fuel energy provided by one or more of the gas turbine engine configurations disclosed or described herein may be greater than or equal to one of the following values: 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, if the fuel supplied to the combustion chamber includes sustainable aviation fuel or a mixture of sustainable aviation fuel comprising up to 100% sustainable aviation fuel. The bypass ratio can lie 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 0.8 to 5, 0.88 to 4.75, or 0.89 to 4.75.

[0342] As a purely non-restrictive example, the reduction in the CO2 percentage per MJ of fuel energy can result from the fuel supplied to the combustion chamber having a specific fuel energy (in MJ per kg) greater than one of the following values: 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 of the fuel can lie within an inclusive range defined by any two the values ​​in the preceding sentence are limited (i.e., the values ​​can form upper or lower limits), for example in the range of 42.8 to 45, from 43 to 44.5 or from 43 to 44.

[0343] As used herein, a range “from value X to value Y” or “between value X and value Y” or the like, unless otherwise specified, denotes an enclosing range; including the limiting values ​​of X and Y.

[0344] The person skilled in the art will understand that, except in the case of mutual exclusion, a feature or parameter described in relation to one of the aspects mentioned above can be applied to any other aspect. Furthermore, except in the case of mutual exclusion, each feature or parameter contained or described herein can be applied to any aspect and / or combined with any other feature or parameter contained or described herein.

[0345] Provided they are not mutually exclusive, any parameter or value contained or described herein may be applied to and / or combined with one or more further parameters and / or values ​​contained or described herein. For example, a first parameter or value contained or described herein (e.g., parameter A) may be applied to and / or combined with one or more further parameters and / or values ​​contained or described herein (e.g., one or more parameters B, parameter C, and parameter D, etc.) to express a product of their relationship. A person skilled in the art would understand, for example, that in cases where parameter A is disclosed separately from parameter B, a product of their relationship may be expressed, for example, as A / B, B / A, B*A, or, as required, any other such application, combination, or function of parameter A in relation to parameter B.

[0346] Provided they are not mutually exclusive, any parameters or values ​​related to, or determinable and / or derivable from, those contained or described herein may be applied to and / or combined with one or more further parameters or values ​​contained or described herein and / or any further parameters or values ​​related to, or determinable and / or derivable from, those contained or described herein to express their relationship with respect to engine emissions and / or nvPM. For example, using associated determinable and / or derivable temperatures, pressures, operating parameters, rotational speeds, flow rates, or engine operating conditions, a first parameter or value (for example, parameter A) may be applied to any one or more further parameters or values ​​(for example, one or more of parameters B, parameter C, and parameter D, etc.).) applied and / or combined with these to express their relationship with respect to engine emissions and / or nvPM. For example, experts would understand that if parameter C can be considered separately from parameter D, their relationship can be expressed, for example, as C+D, CD, CD, C / D, C / D, C*D, or, as required, any further product, application, combination, function, or expression of parameter C relative to parameter D or vice versa. BRIEF DESCRIPTION OF THE DRAWINGS

[0347] The embodiments are now described only by way of example with reference to the figures in which: Fig. 1 is a cross-sectional side view of a gas turbine engine; Fig. 2 is a close-up cross-sectional side view of an upstream part of a geared gas turbine engine; Fig. 3 is a partially cutaway view of a gearbox for a gas turbine engine; Fig. 4 is a close-up cross-sectional side view of a direct-drive gas turbine engine; Fig. 5 is a schematic view of an aircraft on which two gas turbine engines are mounted in accordance with the present application; Fig. 6 a cross-sectional view through a combustion chamber of the engine Fig. 1 in a plane normal to the main axis of rotation of the engine; Fig. 7 a schematic cross-section of a duplex fuel spray nozzle of the combustion chamber Fig. 6 is; Fig. 8 a schematic cross-section of a single-flow fuel spray nozzle of the combustion chamber Fig. 6 is; Fig. 9 a partial sectional view of the engine from Fig. 1 is; Fig. 10 another partial sectional view of the engine from Fig. 1 is; Fig. 11 A schematic representation of a propulsion system for an aircraft, comprising the engine made of Fig. 1, is; and Fig. Figure 12 shows a method for operating the gas turbine engine. DETAILED DESCRIPTION OF THE PROOF

[0348] Fig. Figure 1 illustrates a gas turbine engine 10 with a main axis of rotation 9. The engine 10 includes an air inlet 12 and a propulsion fan 23, which generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 includes a core 11, which receives the core airflow A. The engine core 11 comprises, in axial flow sequence, 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 channel 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass channel 22. The fan 23 is attached to and driven by the low-pressure turbine 19 via a first low-pressure shaft 26 and a planetary gear set 30.

[0349] In operation, the core airflow A is accelerated and compressed by the low-pressure compressor 14 and directed into the high-pressure compressor 15, where it is further compressed. 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 combusted. The combustion equipment 16 can be referred to as the combustion chamber 16, the terms "combustion equipment 16" and "combustion chamber 16" being used synonymously herein. The resulting hot combustion products then expand through and drive the high-pressure and low-pressure turbines 17 and 19, respectively, before being discharged through the core outlet nozzle 20 to provide some thrust. The high-pressure turbine 17 drives the high-pressure compressor 15 via a suitable connecting second high-pressure shaft 27.The fan 23 generally acts by pressurizing the bypass airflow B, which flows through the bypass channel 22, so that the bypass airflow B is expelled through the bypass outlet nozzle 18 to generally provide the majority of the propulsion thrust. The planetary gear set 30 is a reduction gear set.

[0350] An exemplary arrangement for a geared turbofan gas turbine engine 10 is shown in Fig. 2 shown. The low-pressure turbine 19 (see Fig. 1) Drives the low-pressure shaft 26, which is coupled to a central pinion or sun gear 28 of the planetary gear set 30. The low-pressure shaft 26 can be referred to as the drive shaft for the planetary gear set 30. Radially outward from and engaging with the sun gear 28 is a plurality of planet gears 32, which are coupled to one another by a planet carrier 34. The planet carrier 34 restricts the planet gears 32 such that they rotate synchronously around the sun gear 28, while enabling each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled to the fan 23 via linkage 36 and a fan output shaft 42 to drive the fan 23 to rotate about the engine axis 9. Radially outwards from and engaging with the planet gears 32 is a ring gear or hollow gear 38, which is coupled via linkage 40 to a stationary support structure 24.

[0351] It should be noted that the terms “low-pressure turbine” and “low-pressure compressor” are used herein to refer to the lowest-pressure turbine stages and the lowest-pressure compressor stages (i.e., excluding fan 23), and / or the turbine and compressor stages connected by the lowest-rotating connecting shaft (26) in the engine (i.e., excluding the gearbox output shaft driving fan 23). In some literature, the “low-pressure turbine” and “low-pressure compressor” referred to herein may alternatively be known as the “intermediate-pressure turbine” and “intermediate-pressure compressor.” Where such alternative nomenclature is used, fan 23 may be described as a first or lowest-pressure compression stage.

[0352] The planetary gear set 30 is exemplified in Fig. 3 shown in more detail. Each of the sun gear 28, the planet gears 32 and the ring gear 38 has teeth on its circumference to mesh with the other gears. For clarity, in Fig. Figure 3, however, illustrates only exemplary sections of the teeth. Four planet gears 32 are illustrated, although it is obvious to the person skilled in the art that more or fewer planet gears 32 may be provided within the scope of protection of the claimed invention. Practical applications of a planetary gear drive 30 generally include at least three planet gears 32, for example, five planet gears 32.

[0353] This is exemplified in the Fig. 2 and Fig. The illustrated planetary gear set 30 is a planetary gear set in that the planet carrier 34 is coupled to the output shaft via linkage 36, with the ring gear 38 being fixed. However, any other suitable type of planetary gear set 30 can be used. As another example, the planetary gear set 30 can be a star arrangement in which the planet carrier 34 is held fixed and the ring gear (or toothed ring) 38 can rotate. In such an arrangement, the fan 23 is driven by the ring gear 38. As a further alternative example, the gear set 30 can be a differential gear set in which both the ring gear 38 and the planet carrier 34 can rotate.

[0354] This is the in Fig. 2 and Fig. The arrangement shown in Figure 3 is only exemplary, and various alternatives fall within the scope of protection of this disclosure. Purely by way of example, any suitable arrangement for housing the gearbox 30 in the engine 10 and / or for connecting the gearbox 30 to the engine 10 can be used. The connections (such as the linkage 36, 40 in the example of Figure 3) can also be used by way of example. Fig. 2) between the gearbox 30 and other parts of the engine 10 (such as the input shaft 26, the output shaft, and the stationary support structure 24) exhibit any desired degree of stiffness or flexibility. Likewise, any suitable arrangement of the bearings between rotating and stationary parts of the engine (for example, between the input and output shafts of the gearbox and the stationary structures, such as the gearbox housing) can be used as an example, and the disclosure is not limited to the exemplary arrangement of Fig. 2 limited. If, for example, the gearbox 30 has a star arrangement (described above), it would be readily apparent to experts that the arrangement of output and support linkages and bearing positions is usually determined by the one in Fig. would differ from the arrangement shown in example 2.

[0355] Accordingly, the present disclosure extends to a gas turbine engine with any arrangement of gear types (for example, star or planetary gears), support structures, drive and output shaft arrangement and bearing positions.

[0356] Optionally, the gearbox can drive additional and / or alternative components (e.g., the intermediate pressure compressor and / or a booster compressor).

[0357] Other gas turbine engines in which the present disclosure can be used may have alternative configurations. For example, such engines may have an alternative number of compressors and / or turbines and / or an alternative number of connecting shafts. Also by way of example, the [reference to be added] Fig. Figure 1 shows a gas turbine engine with a split flow nozzle 18, 20, meaning that the flow through the bypass channel 22 has its own nozzle (the bypass outlet nozzle) 18, which is separate from and radially outside the core outlet nozzle 20. However, this is not limiting, and any aspect of the present disclosure can also be used in engines in which the flow through the bypass channel 22 and the flow through the core 11 are mixed or combined upstream of (or before) a single nozzle, which may be called a mixing flow nozzle. One or both nozzles (whether mixing or split flow) may have a fixed or variable area.

[0358] As a further example, other gas turbine engines in which the present disclosure can be used may not have a gearbox for the main shaft(s) and may instead be direct-drive engines. A cross-sectional view of such an engine is shown in Fig. 4 shown.

[0359] As in Fig. As shown in Figure 4, a gas turbine engine is generally specified at 10 with a main axis of rotation 9. The engine 10 comprises, in axial flow sequence, an air inlet 12, a propulsion fan 23, an intermediate pressure compressor 14, a high pressure compressor 15, combustion equipment 16, a high pressure turbine 17, an intermediate pressure turbine 19a, a low pressure turbine 19, and an exhaust nozzle 20. A nacelle 21 surrounds the engine 10 and defines both the inlet 12 and the exhaust nozzle 20.

[0360] In operation, air entering inlet 12 is accelerated by fan 23 to generate two airflows: a core airflow A and a bypass airflow B. Core airflow A flows into intermediate-pressure compressor 14, and bypass airflow B passes through a bypass duct 22 to provide thrust. Intermediate-pressure compressor 14 compresses airflow A before delivering it to high-pressure compressor 15, where further compression takes place.

[0361] 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 combusted. The combustion equipment 16 can be referred to as the combustion chamber 16, the terms "combustion equipment 16" and "combustion chamber 16" being used synonymously herein. The resulting hot combustion products then expand through and drive the high-pressure, intermediate-pressure, and low-pressure turbines 17, 19a, and 19, respectively, before being discharged through the nozzle 20 to provide additional thrust. The high-pressure turbine 17, the intermediate-pressure turbine 19a, and the low-pressure turbine 19 each drive the high-pressure compressor 15, the intermediate-pressure compressor 14, and the fan 23, respectively, via suitable connecting shafts.

[0362] Other gas turbine engines in which the present disclosure can be used may have alternative configurations. For example, such engines may have an alternative number of connecting shafts (e.g., two) and / or an alternative number of compressors and / or turbines. Furthermore, the engine may include a gearbox provided in the drive train from a turbine to a compressor and / or a fan.

[0363] While the described example refers to a turbofan engine, the disclosure can be used, for example, for any type of gas turbine engine, such as an open rotor (where the fan stage is not enclosed by a nacelle) or a turboprop engine. In some arrangements, the gas turbine engine 10 may not include a gearbox 30.

[0364] The geometry of the gas turbine engine 10 and of its components is defined by a conventional axis system, which has an axial direction (aligned with the main axis of rotation 9), a radial direction (in the direction from bottom to top in Fig. 1) and a circumferential direction (perpendicular to the side in the view of Fig. 1) includes. The axial, radial and circumferential directions are perpendicular to each other.

[0365] Fig. Figure 5 shows an aircraft 1 on which two gas turbine engines 10 of the present disclosure are mounted, one on each wing. The aircraft 1 includes a fuel system 2 comprising a fuselage fuel tank 50a and two wing fuel tanks 50b. The fuel F is supplied to the gas turbine engines from the fuel system. The fuel tanks 50a and 50b are supplied with fuel via a fuel inlet port 62. Other fuel tank configurations may be used for other fuel systems.

[0366] The fuel F supplied to the combustion equipment 16 can comprise a fossil-based hydrocarbon fuel, such as kerosene. Thus, the fuel F can include molecules from one or more of the chemical families of n-alkanes, isoalkanes, cycloalkanes, and aromatics. Since a trend toward the use of fuels different from the currently widely used conventional kerosene-based jet fuels is anticipated in the aviation industry, the fuel F, when blended, mixed, or replaced with an alternative fuel, can include renewable hydrocarbons produced from biological or non-biological resources, also known as sustainable aviation fuel (SAF). In each of the examples provided, the fuel F can include one or more trace elements, including, for example, sulfur, nitrogen, oxygen, inorganic substances, and metals.

[0367] Experts understand renewable fuels (NFFs) to include, for example, biofuels, renewable aviation fuels, renewable jet fuels, alternative fuels, or biojet fuels produced from biological or non-biological resources. Thus, experts understand NFFs to be fuels produced from sustainable and / or renewable resources. For example, according to the general understanding, NFFs are 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 waste oil and fats; municipal solid waste; cellulose-containing waste (such as corn stalks); cover crops such as camelina, Carinata, and pennycress; non-biogenic alternative fuels; jatropha; halophytes and algae, and not from fossil hydrocarbons derived, for example, from fossil oil and / or natural gas.Accordingly, NFK is understood to be a substance consisting of renewable hydrocarbons. Furthermore, NFK is understood to mean that it does not include fossil fuels or fossil-based hydrocarbons.

[0368] The functional performance of a particular fuel composition or fuel mixture F for use in a specific application can be defined, at least in part, by the fuel's ability to support the Brayton cycle of the gas turbine engine 10. Parameters defining functional performance may include, for example, specific energy; energy density; thermal stability; and emissions, including gaseous and / or particulate matter. In this context, particulate emissions may include soot particles produced during the combustion of fuel F, also referred to as non-volatile particles (nvPM). Thus, nvPM can be defined as emitted particles present at the outlet plane of the exhaust nozzle of a gas turbine engine that do not volatilize when heated to a temperature of 350 °C.All references herein to soot or smoke apply equally to other types of particulate 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), including, for example, sulfur dioxide (SO2) and / or sulfur trioxide (SO3); and volatile organic compounds (VOCs) produced during the combustion of fuel F. All references herein to gaseous emissions apply equally to other types of gaseous emissions known in the art.

[0369] A relatively higher specific energy (i.e., energy per unit mass), expressed as 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 as MJ / l, can at least partially reduce the takeoff fuel volume, which can be particularly important for volume-constrained missions or military operations with a single refueling stop. Relatively higher thermal stability (i.e., inhibition of fuel degradation or coking under thermal stress) can allow the fuel to maintain elevated temperatures in the engine and fuel injectors, potentially providing relative improvements in combustion efficiency.Reduced emissions, including particulate matter, can enable reduced contrail formation, thus lowering the environmental impact of a given mission. Other fuel properties can also be crucial for functional performance. For example, a relatively lower freezing point (°C) can enable long-range missions to optimize flight profiles; minimal 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 a maximum surface tension (mN / m) can ensure adequate fuel spray pattern and atomization.

[0370] The ratio of hydrogen atoms to carbon atoms in a molecule can influence the specific energy of a given fuel composition or mixture. Fuels with higher ratios of hydrogen atoms to carbon atoms can exhibit higher specific energies in the absence of bond strain. In some examples, fossil-based hydrocarbon fuels can comprise molecules with approximately 7 to 18 carbon atoms, with a significant portion of a given composition consisting of molecules containing 9 to 15 carbon atoms, averaging 12 carbon atoms.

[0371] A number of sustainable aviation fuel blends have been approved for use. For example, some approved blends include up to 10% sustainable aviation fuel, while others include up to 50% sustainable aviation fuel (the remainder being one or more fossil-based hydrocarbon fuels, such as kerosene), with approval pending for further compositions. However, the aviation industry anticipates that sustainable aviation fuel blends containing up to (and including) 100% sustainable aviation fuel (SAF) will eventually be approved for use.

[0372] Sustainable aviation fuels can comprise one or more of n-alkanes, isoalkanes, cycloalkanes, and aromatics, and can be produced, for example, from one or more of synthesis gas (syngas); lipids (e.g., fats, oils, and greases); sugars; and alcohols. Thus, sustainable aviation fuels can have a lower content of aromatics and sulfur, or both, compared to fossil hydrocarbon fuels. Additionally or alternatively, sustainable aviation fuels can have a higher content of isoalkanes and cycloalkanes, or both, compared to fossil hydrocarbon fuels. In some examples, sustainable aviation fuels can have a density below 100% (e.g., between 90% and 98%) that of kerosene, and a specific energy above 100% (e.g., between 101% and 105%) that of kerosene.For example, the calorific value of sustainable aviation fuels can be between 101% and 105% of that of kerosene.

[0373] In some examples, the sustainable aviation fuel(s) or mixture(s) supplied to the combustion equipment 16 may have a relatively lower content of aromatic and / or other non-paraffinic components than kerosene. The sustainable aviation fuel may have an aromatic content of, for example, 30%, 20%, 15%, 10%, 8%, 5% or less than 5%; e.g., 4%, 3%, 2%, 1% or less than 1%; e.g., 0.75%, 0.5%, 0.25% or less than 0.25%; e.g., 0.2%, 0.1% or less than 0.1%; e.g., 0.01%, 0.001% or 0%. The aromatic content of the sustainable aviation fuel can be within an inclusive value or range that is limited by or lies within any two of the values ​​mentioned in the previous sentence (i.e., the values ​​can form upper or lower limits), e.g.13.5%, 8.5%, 2.5%, 0.35%, 0.15%, 0.05%, 0.005% or 0%; or 0% to 0.75%, 0% to 0.5% or 0.1% to 0.25%; or 0.15% to 0.65%, 0.35% to 0.55% or 0.035% to 0.055%; depending on one or more of preference, fuel stock or supplier and variations in the composition thereof.

[0374] At least in part due to their molecular structure, sustainable aviation fuels can offer advantages, including, for example, one or more of a higher specific energy (despite, in some examples, a lower energy density); a higher specific heat capacity; higher thermal stability; higher lubricity; lower viscosity; lower surface tension; a lower freezing point; lower soot emissions; less NOx; and lower CO2 emissions relative to fossil-based hydrocarbon fuels (e.g., when burned in combustion equipment). Accordingly, sustainable aviation fuels, compared to fossil hydrocarbon fuels such as kerosene, can lead to one or both of a relative decrease in specific fuel consumption and a relative decrease in maintenance costs.

[0375] Fig. Figure 6 shows a section through the combustion chamber 16 of engine 10. Fig. 1 in a plane normal to the main axis of rotation 9 of the engine 10. The combustion chamber 16 comprises an annular combustion space 401 defined by a lining 402. Alternative combustion chamber configurations can be used in other embodiments, for example, tubular, cup-shaped, etc. The combustion chamber 16 comprises a plurality of fuel spray nozzles 403, 404 arranged around the circumference of the combustion chamber 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 combustion chamber 16 comprises 16 fuel spray nozzles 403, 404. In other examples, the combustion chamber 16 can comprise any suitable number of fuel spray nozzles 403, 404, for example, a number of fuel spray nozzles in the range of 14-22. In some examples, the number of fuel spray nozzles 403, 404 can range between 16 and 20.In other examples, the number of fuel spray nozzles can be 14, 15, 16, 17, 18, 19, 20, 21, 22, or a number within a range defined between any two of the values ​​in this sentence.

[0376] The number of fuel spray nozzles 403, 404 can also be quantified as the ratio of the number of fuel spray nozzles to the engine core size. The core size defines the size of the core 11 of the gas turbine engine 10. Engine core size can be defined as follows: Core size = m˙2T3P3 where ṁ2 = the mass flow rate of air entering the high-pressure compressor 15 in pounds per second, T3 = the temperature of air exiting the high-pressure compressor 15 in Kelvin, and P3 = the pressure of air exiting the high-pressure compressor 15 in pounds per second per square inch. One unit of the core quantity is therefore expressed as: s⋅K12⋅in

[0377] The core size (in s·K) 1 / 2 The engine core size (in s·K) can be between 4 and 7, for example 4, 4.5, 5, 5.5, 6, 6.5, or 7, or any range defined between any two of these values. In some examples, the engine core size (in s·K) can be... 1 / 2 ·in) in the range of 5.0, 5.1, 5.2, 5.3, 5.4, 5.5, 5.6, 5.6, 5.7, 5.8, 5.9 or 6 or any range defined between any two of these values. In still further examples, the engine core size (in s·K) can be 1 / 2 ·in) 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.

[0378] The number of fuel spray nozzles per unit of engine core size (in the units specified above) can be in the range of 2 to 6, for example 2, 3, 4, 5, 6, or within any range defined between any two such values. The number of fuel spray nozzles per unit of engine core size can be in the range of 2.7 to 4, preferably in the range of 3 to 3.6. In some preferred examples, the number of fuel spray nozzles per unit of engine core size can be in the range of 2.5 to 4.5, for example 2.5, 3, 3.5, 4, or 4.5, or any range defined between any two such values. In yet other examples, the number of fuel spray nozzles per unit of engine core size can 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 further examples, the number of fuel spray nozzles per unit of engine core size can be 2.0, 2.1, 2.2, 2.3, 2.4, 2.5, 2.6, 2.7, 2.8, 2.9, 3.0, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4.0, 4.1, 4.2, 4.3, 4.4, 4.5, 4.6, 4.7, 4.8, 4.9, 5.0, 5.1, 5.2, 5.3, 5.4, 5.5, 5.6, 5.7, 5.8, 5.9 or 6.0, or within a range defined between any two such values.

[0379] The core parameter is defined herein in an engine operating condition that corresponds to a maximum value of the half-dimensionless flow at high-pressure compressor inlet, defined as: m˙2T2P2 where ṁ2 is the mass flow rate (in pounds per second) of air entering the high-pressure compressor, T2 is the temperature (in Kelvin) of air entering the high-pressure compressor, and P2 is the pressure (in pounds per second per square inch) of air entering the high-pressure compressor.

[0380] The operating condition corresponding to the maximum half-dimensionless flow at the high-pressure compressor inlet can be the peak of the climb operating condition. The core quantity referred to herein can therefore be defined at the peak of the climb operating condition. The peak of the climb can be defined as in the prior art and understood by a person skilled in the art for a specific implementation of a gas turbine of the present application. In a specific example, the peak of the climb can correspond to operation at an altitude between 30,000 feet and 39,000 feet (more precisely, 35,000 feet), a forward speed of Mach number 0.75 to 0.85, and an ambient air temperature (TAMB) of ISA+10 K to ISA+15 K.

[0381] The combustion chamber 16 comprises a number of duplex fuel spray nozzles 403 (also known as internally staged nozzles) in which a primary fuel injector is integrated into the same fuel nozzle as the main fuel injector. The combustion chamber 16 also comprises a number of single-stream fuel spray nozzles 404, each comprising only one main fuel injector. In other examples, the combustion chamber 16 may comprise only duplex fuel spray nozzles or only single-stream fuel spray nozzles.

[0382] In this example, the combustion chamber 16 comprises 12 duplex fuel spray nozzles 403 and 4 single-flow fuel spray nozzles 404. The duplex fuel spray nozzles 403 are in Fig. Figure 6 is illustrated by hatched circles. The duplex fuel spray nozzles 403 are arranged in groups of three around the circumference of the combustion chamber 16, each group being diametrically opposed to any other group. In other examples, the combustion chamber 16 may contain any suitable number of duplex fuel spray nozzles, for example, in the range of 10–14 nozzles, and any suitable number of single-stream fuel spray nozzles, for example, in the range of 4–8 nozzles. In some examples, the number of duplex fuel spray nozzles may be 10, 11, 12, 13, or 14, or within a range defined between any two values ​​in this sentence. In some examples, the number of single-stream fuel spray nozzles may be 4, 5, 6, 7, or 8, or within a range defined between any two values ​​in this sentence.The duplex fuel spray nozzles can be arranged in any suitable number of groups or they can be arranged without groups. Where applicable, each group of duplex fuel spray nozzles can comprise any suitable number of nozzles, for example, in the range of 2 to 8 nozzles. In some examples, each group of duplex nozzles can comprise 2, 3, 4, 5, 6, 7, or 8 fuel spray nozzles, or a number within a range defined between any two of these values.

[0383] The combustion chamber 16 further comprises four igniters 405 arranged to ignite an air-fuel mixture in the combustion chamber 401 during operation. Each igniter 405 is located 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 (e.g., its nearest igniter) compared to the single-flow nozzles 404. Each igniter 405 is arranged diametrically opposite another igniter 405. In other examples, the combustion chamber may include fewer or more igniters, for example, a number of igniters in the range 1-8, and the igniters may be arranged differently. For example, one or more of the igniters may not be located adjacent to one of the groups of duplex fuel spray nozzles, and one or more of the igniters may not be located diametrically opposite another igniter.In some examples, the combustion chamber can contain 1, 2, 3, 4, 5, 6, 7 or 8 igniters, or a number within a range defined between any two of the values ​​in this set.

[0384] In the example shown, when the engine 10 is operated at low power (below a step point), for example during or shortly thereafter during start-up, fuel is supplied only to the primary injectors of the duplex fuel spray nozzles 403 for delivery to the combustion chamber 401. The duplex nozzles 403 therefore receive a higher fuel flow rate below the step point compared to the single-flow nozzles 404. As the power output of the engine 10 and the mass airflow through the engine 10 increase, the step point is reached, at which fuel is supplied to the combustion chamber 401 in addition to the primary fuel injectors of one or more of the duplex fuel spray nozzles 403 and the primary fuel injectors of one or more of the single-flow fuel spray nozzles 404.In this example, at higher power levels, fuel is injected through all main fuel injectors of both the duplex fuel spray nozzles 403 and the single-flow fuel spray nozzles 404, in addition to fuel injected through the primary injectors of the duplex fuel spray nozzles 403. In this example, the fuel flow rate supplied to the main injectors of the single-flow fuel spray nozzles 404 is less than or equal to the fuel flow rate supplied to the main injectors of the duplex fuel spray nozzles 403. Therefore, since both the primary and main injectors of the duplex fuel spray nozzles 403 receive fuel, the duplex fuel spray nozzles 403 receive more fuel than the single-flow fuel spray nozzles 404 at and above the stage point.In an alternative example, fuel is supplied only to the main fuel injectors of one or more of the one or more duplex fuel spray nozzles 403 and to the main fuel injectors of one or more of the single-flow fuel spray nozzles 404 at and above the stage point, i.e., fuel is not supplied to the primary injectors of the duplex fuel spray nozzles 403.

[0385] The fuel flow supplied to the plurality of fuel spray nozzles is therefore unbalanced such that the fuel flow rate to a first subset of the plurality of fuel spray nozzles (in this example, the duplex fuel spray nozzles 403) is greater than that supplied to a second subset of the fuel spray nozzles (in this example, the single-flow fuel spray nozzles 404). This can enable a primary fuel flow to be supplied to fuel spray nozzles located relatively closer to the igniters 405 to aid ignition and flame stability at low engine power, engine start-up, or during engine restart. In some examples, the first subset (e.g., the duplex nozzles) of fuel spray nozzles can comprise at least half, and preferably at least two-thirds, of the total number of fuel spray nozzles.

[0386] In other examples, the fuel flow rate to each fuel spray nozzle provided in the combustion chamber may be the same, and there may be no fuel flow imbalance to any subset of nozzles. In such an example, all fuel flow nozzles may be single-stream nozzles, or they may all be duplex nozzles. In still other examples, other arrangements of fuel spray nozzles may be provided, in which fuel is unbalanced to those closer to or adjacent to the igniters. For example, two subsets (independently controllable) of duplex nozzles or two subsets of single-stream nozzles may be provided, which may be biased unbalanced as described above.

[0387] Fig. Figure 7 shows one of the duplex fuel spray nozzles 403 of the combustion chamber 16. The duplex nozzle 403 comprises a primary fuel injector 501, a main fuel injector 502, and an air duct 503. The primary injector 501 comprises a primary inlet 504, arranged to receive a primary fuel flow P, and a primary fuel circuit 505, arranged to deliver the primary fuel flow to the outlet 506 of the nozzle 403. The main injector 502 comprises a main inlet 507, arranged to receive a main fuel flow M, and a main fuel circuit 508, arranged to deliver 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 releases the high-pressure air to the outlet 506 of the nozzle 403.

[0388] The duplex nozzle 403 is configured to produce a primary fuel cone from the primary injector 501 and a main fuel cone from the main injector 502 at the outlet 506 of the nozzle 403 (in Fig. 7 illustrated by the dashed lines, each labeled P and M). When both the primary and main injectors 501, 502 are active, the primary and main cones are arranged concentrically, with the main cone arranged annularly outside the primary cone. Those skilled in the art are familiar with such fuel spray patterns.

[0389] The duplex nozzle 403 is from Fig. 7 is merely an example and other examples may use an alternative configuration of the duplex nozzle.

[0390] Fig. Figure 8 shows one of the single-flow fuel spray nozzles 404 of the combustion chamber 16. The nozzle 404 comprises a main fuel injector 601, including a main inlet 602 arranged to receive a main fuel flow M, and a main fuel circuit 603 arranged to deliver the main fuel flow to the outlet 604 of the nozzle 404. The nozzle 404 is configured to produce a main fuel cone at the outlet 604 of the nozzle 404 (shown by the dashed lines marked M). Air is similarly supplied to the outlet 604 of the nozzle through an air duct 605.

[0391] The single-flow fuel spray nozzle 404 is from Fig. 8 is merely an example and other examples may use an alternative configuration of the single-flow fuel spray nozzle 404.

[0392] Fig. 9 and Fig. Figures 10 each show a section through the engine 10, viewed perpendicular to the main axis of rotation 9, including a section of the combustion chamber 16 that comprises 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-stream fuel spray nozzles 404 as for the duplex nozzles 403. The combustion chamber 16 is mounted within a cavity 406 formed by an inner air housing 407 and an outer air housing 408. During operation, the high-pressure compressor 15 supplies high-pressure air D to the cavity 406 via a diffuser 409. At this point, a quantity of air enters the combustion chamber 16 as combustion air E through the fuel nozzle 403 and / or mixing ports at the inlet to the combustion chamber 16. The remaining air flows around the combustion chamber 16 as cooling air G, a portion of which is admitted downstream of the fuel nozzle 403, as described below with reference to Fig. 10 described.

[0393] One or more temperature and / or pressure probes (not shown) may be installed and arranged in the housing of the diffuser 409 to measure the temperature and / or pressure of the high-pressure air D discharged to the cavity 406 from the high-pressure compressor 15 via the diffuser 409 (i.e., temperature and pressure at the outlet of the high-pressure compressor 15). Such a temperature probe may be designated as a T3 probe, and such a pressure probe may be designated as a P3 probe. It is understood that the engine 10 may include any suitable arrangement of pressure and temperature probes, which may be positioned at any suitable location within the engine 10. As used herein, T3 and P3, and any other numbered pressures and temperatures, may be defined using the station numbering specified in standard SAE AS755.

[0394] Combustion chamber 16 operates as a fat-burning rapid quench lean-burn combustion chamber (RQL combustion chamber). In other examples, combustion chamber 16 may be an alternative type of combustion chamber, such as a standard fat-burning combustion chamber (no unbalanced fuel flow). How Fig. As can be seen from Figure 10, the combustion chamber 401 of the RQL combustion chamber 16 is divided into three zones along its length: a rich zone 801, a rapid quench zone 802, and a lean zone 803. During operation, a rich air-fuel mixture is introduced from the fuel spray nozzle 403 into the rich zone 801, where it is ignited by the igniter 405. Within the rich zone 801, fuel is burned at a fuel / air ratio higher than stoichiometric (for example, at an equivalence ratio of approximately 1.8). Air is then introduced into the combustion products via primary ports 804, which are located in the lining 402 of the combustion chamber 16, before the combustion products reach the rapid quench zone 802. Additional air is supplied to the still-burning fuel via the primary ports 804 (which can be referred to as quench ports). Air is drawn through the primary ports 804 at a higher rate (e.g.Air is then introduced into the combustion products via secondary ports 805 located in the lining 402 of the combustion chamber 16, which reduces the combustion to a significantly lower than stoichiometric fuel / air ratio (for example, at an equivalent ratio between 0.5 and 0.7), while still allowing the fuel to burn. Consequently, very little of the combustion process can be carried out at near-stoichiometric fuel / air ratios, and therefore relatively little nitrogen oxides (NOx) are produced. Air is then introduced into the combustion products again via secondary ports 805 located in the lining 402 of the combustion chamber 16, while the combustion products are in the lean zone 803 (or immediately before they reach the lean zone 803). Within the lean zone 803, fuel is burned at a fuel / air ratio lower than stoichiometric (for example, at an equivalent ratio between 0.5 and 0.7).After passing through the lean zone 803, the combustion products exit the combustion chamber 16. The secondary ports 805 can be referred to as dilution ports and can be arranged to gradually introduce dilution air into the lean zone 803. The fuel added by the fuel spray nozzle is substantially completely combusted until the air exits at a combustion chamber outlet before flowing to the turbine.

[0395] Fig. Figure 11 shows a section of a propulsion system 900 for an aircraft. The propulsion system 900 comprises the gas turbine engine 10 of Fig. 1. The engine 10 further comprises a fuel system and an oil system. The fuel system comprises: a low-pressure fuel pump 902, a fuel-oil heat exchanger 903, a main (or high-pressure) fuel pump 904, a control unit 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 feed pump 906, and a main oil pump 907. In the present example, the low-pressure fuel pump 902 is configured as part of the gas turbine engine 10. In other examples, the low-pressure fuel pump or additional fuel pumps may be provided as part of the fuel system on board the aircraft on which the gas turbine engine is mounted.

[0396] The low-pressure fuel pump 902 is arranged to deliver fuel from the fuel tank 901 to the fuel-oil heat exchanger 903 via a suitable arrangement of pipes, lines, etc. (not shown). The main fuel pump 904 is configured to deliver fuel from the fuel-oil heat exchanger 903 to the fuel distribution valve 909 and a suitable arrangement of pipes, lines, etc. (not shown) to the fuel spray nozzles of the combustion chamber 16. The fuel distribution valve 909 is arranged to distribute fuel between a main distributor 909a and a primary distributor 909b. The main distributor is fluidically connected to the main injectors of each of the fuel spray nozzles 404, 403, as shown in Fig. Figure 11 shows that it supplies fuel to all duplex 403 and single-flow 404 fuel spray nozzles. The primary distributor 909b is fluidically connected to the primary injectors of each of the duplex fuel spray nozzles 403. The primary distributor 909b can therefore be used to provide a higher fuel flow rate to the first subset of fuel spray nozzles (e.g., in this example, the duplex fuel spray nozzles 403) compared to the fuel flow rate supplied to the second subset of fuel spray nozzles via the primary distributor 909a. For example, below a threshold engine power, fuel can be supplied only to the first subset of fuel spray nozzles via the primary distributor 909b, or the first subset of fuel spray nozzles can be supplied at a higher fuel flow rate than the second subset of fuel spray nozzles.This can limit the production of unwanted combustion products, such as nitrogen oxides (NOx), unburned hydrocarbons (HC) and carbon monoxide (CO), and can create an imbalance in the fuel flow to injectors closest to the igniters to support flame stability and ignition at low engine power.

[0397] The oil transfer pump 906 is arranged to supply lubricating oil from the oil tank 905 to the fuel-oil heat exchanger 903 via a suitable arrangement of pipes, lines, etc. (not shown). The main oil pump 907 is arranged to supply oil from the fuel-oil heat exchanger 903 to components of the engine 10 via a suitable oil distribution arrangement (not shown), if required. The flow path of fuel from the fuel tank 901 via the pumps 902, 904, and the fuel-oil heat exchanger 903 to the combustion chamber 16 during operation is shown in Fig. Figure 11 is illustrated by dashed or dotted arrows. The flow path of oil from the oil tank 905 via the oil transfer pump 906 to the fuel-oil heat exchanger 903 and onward to components of the engine 10 in operation is shown in Figure 11. Fig. 11 illustrated by solid arrows.

[0398] The 908 controller comprises a suitable arrangement of processors and electronic memory. The 908 controller communicates as indicated by the dashed and dotted lines in Fig. Figure 11 illustrates the use of the fuel-oil heat exchanger 903 and is configured to control its operation. In some examples, the controller 908 may be configured to control the flow rate of oil through the fuel-oil heat exchanger 903. The controller 908 is configured to control the operation of the fuel-oil heat exchanger 903 by providing control signals to the fuel-oil heat exchanger 903. The controller 908 is configured to control the operation of the fuel-oil heat exchanger 903 to adjust at least one property or parameter of the fuel as it enters the combustion chamber 16. In the example shown, the controller 908 is configured to control the operation of the fuel-oil heat exchanger 903 to control the viscosity of the fuel as it enters the combustion chamber 16.In other examples, the control unit 908 may be additionally or alternatively configured to control the operation of the fuel-oil heat exchanger 903 to control the fuel temperature upon entry into the combustion chamber. The control unit 908 may, as illustrated, be a separate control unit or may form part of an electronic engine control (EEC) system arranged to control other engine functions.

[0399] 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 can be arranged at any suitable location or position relative to the other components of the drive system 900. In other examples, the drive system 900 may include one or more additional heat exchangers arranged to receive oil from the oil system, or the drive system 900 may include one or more additional oil systems arranged to supply oil to the one or more additional heat exchangers. The drive system 900 is, as shown in Fig. Figure 11 shows only a schematic view of an illustrative drive system.

[0400] In one example, the controller 908 is configured to control the operation of the fuel-oil heat exchanger 903 to maintain the fuel viscosity at 0.58 mm. 2 / s or lower upon entry into combustion chamber 16 under cruise flight conditions. Alternatively, the control unit 908 can be configured to control the operation of the fuel-oil heat exchanger 903 to maintain the fuel viscosity between 0.58 mm 2 / s and 0.30 mm 2 / s, for example 0.58, 0.57, 0.56, 0.55, 0.54, 0.53, 0.52, 0.51, 0.50, 0.49, 0.48, 0.47, 0.46, 0.45, 0.44, 0.43, 0.42, 0.41, 0.40, 0.39, 0.38, 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31 or 0.30 mm 2 / s upon entry into combustion chamber 16 under cruise flight conditions. Alternatively, the control unit 908 can be configured to control the operation of the fuel-oil heat exchanger 903 to reduce the fuel viscosity to 0.57, 0.56, 0.55, 0.54, 0.53, 0.52, 0.51, 0.50, 0.49, 0.48, 0.47, 0.46, 0.45, 0.44, 0.43, 0.42, 0.41, 0.40, 0.39, 0.38, 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, or 0.30 mm 2 / s or lower, or to within any range defined between any two of these values, upon entry into combustion chamber 16 under cruise flight conditions. The control unit 908 can be configured to control the operation of the fuel-oil heat exchanger 903 to reduce the fuel viscosity to between 0.55 mm 2 / s and 0.35 mm 2 / s, 0.53 mm 2 / s and 0.35 mm 2 / s, 0.50 mm 2 / s and 0.35 mm 2 / s, 0.48 mm 2 / s and 0.35 mm 2 / s, 0.48 mm 2 / s and 0.38 mm 2 / s, 0.48 mm 2 / s and 0.40 mm 2 / s, 0.46 mm 2 / s and 0.40 mm 2 / s, 0.44 mm 2 / s and 0.40 mm 2 / s or 0.44 mm 2 / s and 0.42 mm 2 to reduce / s upon entry into combustion chamber 16 under cruise flight conditions.

[0401] The control unit 908 can additionally or alternatively be configured to control the operation of engine 10 such that a reduction of 10–70% of the average number of particles / kg nvPM in the exhaust gas of the gas turbine engine 10 is achieved when the engine 10 is operated at 85% of the available thrust under given operating conditions, and of 10–70% of the average number of particles / kg nvPM in the exhaust gas of the gas turbine engine 10 when the engine 10 is operated at 30% of the available thrust under given operating conditions, provided that the fuel supplied to the combustion chamber 16 is a sustainable aviation fuel instead of a fossil-based hydrocarbon fuel. In other examples, the nvPM reduction may be defined differently than described herein.

[0402] In this example, or any other example described herein, the controller 908 is configured to control the fuel distribution valve 909 such that it controls the supply of fuel to the fuel spray nozzles of combustion chamber 16. The controller 908 is configured to create an imbalance in 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 controller 908 is configured to control the fuel distribution valve 909 such that, below a step point, fuel is delivered only to the primary fuel injectors of the duplex fuel spray nozzles 403. Above a step point, the controller 908 is configured to control the fuel distribution valve 909 such that fuel is additionally delivered to the main fuel injectors of the duplex fuel spray nozzles 403 and the single-flow fuel spray nozzles 404.Therefore, the duplex fuel spray nozzles 403 receive more fuel than the single-stream fuel spray nozzles 404 (below and optionally above the step point). The control 908 can alternatively be configured to control the fuel distribution valve 909 such that it controls the fuel supply so that any suitable subset of fuel spray nozzles 403, 404 receives more fuel than the other fuel spray nozzles 403, 404. This advantageously allows for optimization of the fuel supply with respect to engine performance, emissions, or any other suitable criteria. The fuel delivery system shown in the figures is to be understood as one example of how fuel can be supplied unbalanced to the fuel spray nozzles, although others are possible. For example, two sets of independent single-stream nozzles can be provided.

[0403] The gas turbine engine 10 of the present application is configured to supply fuel comprising sustainable aviation fuel (SAF) to the fuel spray nozzles 403, 404. In other words, the gas turbine engine 10 is configured to inject fuel (F) comprising sustainable aviation fuel (SAF) into the combustion chamber 401. In operation, the fuel supplied to the fuel spray nozzles 403, 404 therefore comprises SAF.

[0404] By “NFK-comprehensive fuel” we can mean that the fuel supplied to combustion chamber 16 (and combustion chamber 401) via the fuel spray nozzles 403, 404 comprises an NFK percentage 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, by “NFK-comprehensive fuel” we can mean a fuel comprising any mixture of NFK and fossil kerosene fuel, including up to 100% NFK and no fossil kerosene fuel. The NFK-containing fuel can be a fuel with an NFK percentage of 1%, 2%, 5%, 10%, 15%, 20%, 30%, 40%, 50%, 60%, 70%, 80%, 90%, 95%, 100%, or in any range defined by any two of these values.

[0405] By "NFK" we mean a kerosene fuel whose hydrocarbon component consists essentially exclusively of paraffinic hydrocarbons. "NFK" also or alternatively refers to a kerosene fuel with a hydrogen mass fraction in the range of 13.7% to 16.9%, for example, 15.3%. By "fossil-based hydrocarbon fuel" or "fossil fuel" we mean kerosene derived from fossil fuels with a hydrogen mass fraction in the range of 12.0% to 14.8%, for example, 13.4%. Emissions of non-volatile particulate matter (nvPM)

[0406] An nvPM emission index (EI) is defined herein as the mass of nvPM produced per unit mass of fuel used by the combustion chamber 16 of the gas turbine engine 10 of an example described herein. In the present application, unless otherwise specified, 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 consumed by the engine 10 in kilograms.

[0407] Examples of both uncorrected and system loss (SV) corrected exhaust emissions from production aircraft engines (using fossil JET A-1 kerosene aviation fuel), measured according to the procedures in ICAO Annex 16, Volume II, and, where indicated, certified by the countries of manufacture of the engines in accordance with their national regulations, are included in the "ICAO Aircraft Engine Emissions Databank," which is maintained by the European Union Aviation Safety Agency (EASA) on behalf of the International Civil Aviation Organization (ICAO). The database covers engine types whose emissions are regulated, namely turbojet and turbofan engines with a static thrust of more than 26.7 kilonewtons nvPM emission indices.

[0408] Examples of empirical correction factors for nvPM mass and number EIs, as well as definitions, symbols, SI units, abbreviations, and procedures for estimating system losses, can be found in Annex 16 of the Convention on International Civil Aviation “Environmental Protection,” Volume II – Aircraft Engine Emissions, Fourth Edition, July 2017; Annex 8 – Procedures for Estimating Non-Volatile Particulate Matter System Loss Corrections; which is incorporated herein by reference. It is noted therein that the implementation of the nvPM sampling and measurement system may result in significant particle loss on the order of 50% of the nvPM mass and 90% of the nvPM number (e.g., because particles are lost through deposition mechanisms on the walls of the sampling system) and that particle losses are size-dependent and therefore dependent on engine operating conditions, combustion chamber technology, and various other factors.Therefore, the nvPM emission indices corrected for system losses refer to nvPM emissions at the engine outlet, corrected for particle size-dependent losses in the sampling and measurement system, excluding thermophoretic losses in the collector section, assuming that the nvPM at the engine exhaust outlet plane have a log-normal distribution, a constant value for the effective nvPM density, a fixed value for the geometric standard deviation, a limitation of the nvPM mass concentration to the detection limit, a minimum particle size cutoff value of 0.01 µm, and no coagulation.

[0409] As specified, for example, in Annex 16 of the Convention on International Civil Aviation “Environmental Protection”, the EI Masse-Correction factor for system losses without correction for thermophoretic losses in the manifold is defined as the ratio between the estimated mass concentration at the outlet plane of the engine exhaust nozzle without correction for thermophoretic losses in the manifold and the measured mass concentration and can be calculated as follows: KSVMass=nvPMMassAESV1×nvPMMassSTP Where: EGG Masse for the nvPM mass emission index corrected for thermophoretic losses in mg / kg fuel; K SV_Masse for the EI Masse Correction factor for system losses without correction for thermophoretic losses in the collection section in µg / m³ 3 stands; nvPM Masse_AE for the estimated nvPM mass concentration at the exit plane of the engine's exhaust nozzle without correction for thermophoretic losses in the manifold; VF1 stands for the dilution factor of the first stage; and nvPMMasse_STP for the diluted nvPM mass concentration under STP conditions of the instrument in µg / m³ 3 stands.

[0410] Furthermore, the EI Anz -Correction factor for system losses defined as the ratio between the estimated number concentration at the exhaust plane of the engine nozzle without correction for thermophoretic losses in the manifold and the measured number correction, and can be calculated as follows: KSVAnz=nvPMAnzAESV1×SV2nvPMAnzSTP Where: EGG Anz for the nvPM number emission index corrected for thermophoretic losses in number / kg of fuel; K SV_Anz for the EI Anz -Correction factor for system losses without correction for thermophoretic losses in the manifold in number / cm² 3 stands; nvPM Anz_AEfor the estimated nvPM number concentration at the exit plane of the engine's exhaust nozzle without correction for thermophoretic losses in the manifold; VF1 stands for the dilution factor of the first stage; VF2 stands for the second-stage dilution factor (VPR) according to calibration; and nvPM Anz_STP for the diluted nvPM number concentration under STP conditions of the instrument in number / cm² 3 stands.

[0411] For example, as specified in Annex 16 of the Convention on International Civil Aviation “Environmental Protection”, the mass at the exit plane of the engine exhaust nozzle (nvPM) Masse_AE ) and number at the exit plane of the engine's exhaust nozzle (nvPM) Anz_AE ) can be determined using the following procedure: a) For a measured nvPM Anz_STP begin with an initial value of nvPMAnzAE=3×SV1×SV2×nvPMAnzSTP b) For the geometric mean diameter D mg For the lognormal particle size distribution (µm), an initial value of 0.02 µm should be assumed. c) Starting with initially assumed values ​​of nvPM Anz_AE and D mg Estimating the nvPM mass (nvPM) from a) and b) Masse_SCH ) and number concentration (nvPM) Anz_SCH ), where: (nvPM Masse_SCH ) for the estimated undiluted (i.e., corrected for dilution) mass concentration at the instrument in µg / m³ 3 stands; and (nvPM Anz_SCH ) for the estimated undiluted (i.e., corrected for dilution) number concentration at the instrument in numbers / cm² 3 stands; using the following equations: nvPMMasse_SCH=∑Dm=0.01μm1μmηMass(Dm)×ρπDm36×nvPMAnz_AE×flgn(Dm)×Δln(Dm) nvPMAnzSCH=∑Dm=0.01μm1μmηAnz(Dm)×nvPMAnzAE×flgn(Dm)×Δln(Dm) Where: nvPMmi stands for the instrument for the mass of non-volatile particles; η Masse (D m ) for the total penetration fraction of the sampling and measurement system for the nvPMmi without thermophoretic losses in the collection section at the electrical mobility particle size D m stands; ρ for the assumed effective nvPM density in g / cm³ 3 stands; D m for the electrical mobility diameter of the nvPM in µm; f lgn (D m ) for the lognormal distribution function with the parameters geometric standard deviation δ g and geometric mean diameter D mg stands; and whereby: nvPMni stands for the instrument for the number of non-volatile particles; η Anz (D m) for the total penetration fraction of the sampling and measurement system for the nvPMni without thermophoretic losses in the collection part at the electrical mobility particle size Dm; and whereby: flgn(Dm)=12πln(σg)×e−12{In(Dm)−ln(Dmg)ln(σg)}2 Δln(Dm)=1n×1log10(e) the width of a size class in the natural logarithm of the base; δ g the assumed geometric standard deviation of the log-normal distribution; e is Euler's number; and n is the number of particle size containers per decade. d) Determining the difference δ (defined as the sum of the squares of the relative differences between measured and calculated dilution-corrected mass and number concentrations) between nvPM Anz_STP , nvPM Masse_STP and the estimates of the nvPM number concentration (nvPM Anz_SCH ) and the nvPM mass concentration (nvPM Masse_SCH) from the initial values ​​of the exit plane of the engine's exhaust nozzle using the following equation: δ=(SV1×SV2×nvPMAnzSTP−nvPMAnzSCHSV1×SV2×nvPMAnzSTP)2 +(SV1×nvPMMasseSTP−nvPMMasseSCHSV1×nvPMMasseSTP)2 e) Repeat steps c) to d), where nvPM Anz_AE and D mg can be varied until δ is less than 1 × 10 -9 sinks. f) As soon as δ is less than 1 × 10 -9 The final values ​​of nvPM have dropped. Anz_AE and D mg those that are associated with this minimized value of δ. g) Using nvPM Anz_AE and D mg from step f) Determining nvPM Anz_AE by means of the following expression: nvPMMassAE=∑Dm=0.01μm1μmρπDm36×nvPMAnzAE׃ign(Dm)×Δln(Dm)

[0412] Alternative methodologies for determining and / or correcting system losses are also proposed, for example, by: Durand et al., 2023 (Correction for particle loss in a regulatory aviation nvPM emissions system using measured particle size, Journal of Aerosol Science, Band 169, 2023, 106140, ISSN 0021-8502); Corbin et al., 2022 (Aircraft-engine particulate matter emissions from conventional and sustainable aviation fuel combustion: Comparison of measurement techniques für mass, number, and size, Atmospheric Measurement Techniques, 15 (10) (2022), S. 3223-3242); Durdina et al., 2021 (Reduction of nonvolatile particulate matter emissions 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), S. 14576-14585); Harper et al., 2022 (Influence of alternative fuel properties and combustor operating conditions on the nvPM and gaseous emissions produced by a small-scale RQL combustor, Fuel, 315 (2022), Artikel 123045); und Saffaripour et al., 2019 (A review on the morphological properties of non-volatile particulate matter emissions from aircraft turbine engines, Journal of Aerosol Science, 105467 (2019)); which are hereby incorporated by reference.

[0413] It is understood that data and indices from the nvPM sampling and measurement system, which have not been corrected for system losses, do not provide an accurate representation of actual exhaust emission levels. Conversely, it is understood that data and nvPM emission indices corrected for sampling and measurement system losses provide an accurate representation of actual exhaust emission levels, which can be expressed, for example, as one or more of the following values: indices corrected for system loss (SV), nvPM count corrected for system loss (SV) (i.e., in count / kg of fuel), or nvPM mass corrected for system loss (SV) (i.e., in mg / kg).It is understood that, unless otherwise stated, the nvPM mass generated by the gas turbine engine 10 refers to the data in milligrams (mg) corrected for system loss (SV) divided by the corresponding mass of fuel consumed by the engine 10 in kilograms (kg).

[0414] The nvPM emission index, whether corrected for system losses or not, can be defined for different operating phases of the gas turbine engine 10, for example at idle, at maximum takeoff (startup), during climb, and during approach. An emission index can also be defined depending on the type of fuel supplied to the combustion chamber 16.

[0415] The following emission index parameters are defined for the gas turbine engine 10: i) EI LeerThe nvPM emission index of the gas turbine engine 10, corrected for system losses, is expressed in mg / kg at 7% of available thrust under given operating conditions. Operation at 7% of available thrust can be equivalent to operation in an idle phase of the gas turbine engine 10. ii) EI maxAb is the nvPM emission index of the gas turbine engine 10 in mg / kg, corrected for system losses, during operation at 100% of available thrust under given operating conditions. Operation at 100% of available thrust can correspond to operation during a maximum liftoff phase of the gas turbine engine 10; iii) EI SteigThe nvPM emission index of the gas turbine engine 10, corrected for system losses, is expressed in mg / kg at 85% of available thrust under given operating conditions. Operation at 85% of available thrust can correspond to operation during a climb phase of the gas turbine engine 10. iv) EI Anflug The nvPM emission index of the gas turbine engine 10, corrected for system losses, is given in mg / kg during operation at 30% of available thrust under given operating conditions. Operation at 30% of available thrust can correspond to operation during an approach phase of the gas turbine engine 10.

[0416] The available thrust under given operating conditions (i.e., engine power setting) is defined as a percentage of the engine's maximum rated thrust (F). 00), as defined in the prior art. 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 engine's maximum rated thrust, i.e., 100% of available thrust, is measured. The predetermined operating conditions may be ISA at sea level, where the absolute reference 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 bleed air and / or power offtakes. The predetermined operating conditions may be daytime conditions. The predetermined operating conditions may be at approximately 60% relative humidity.

[0417] The nvPM emission indices defined above can be further defined depending on the fuel supplied to the combustion chamber. The fuel-specific values ​​of the nvPM emission index are defined as follows: i) EI Leer,FK is the nvPM emission index of the gas turbine engine 10 in mg / kg, corrected for system losses, when operating at 7% of the available thrust under given operating conditions and when a fuel supplied to the combustion chamber 16 is a fossil-based hydrocarbon fuel; ii) EI maxAb,FK is the nvPM emission index of the gas turbine engine 10 in mg / kg, corrected for system losses, when operating at 100% of the available thrust under given operating conditions and when a fuel supplied to the combustion chamber 16 is a fossil-based hydrocarbon fuel; iii) EI Steig,FKis the nvPM emission index of the gas turbine engine 10 in mg / kg, corrected for system losses, when operating at 85% of the available thrust under given operating conditions and when a fuel supplied to the combustion chamber 16 is a fossil-based hydrocarbon fuel; iv) EI Anflug,FK is the nvPM emission index of the gas turbine engine 10 in mg / kg, corrected for system losses, when operating at 30% of the available thrust under given operating conditions and when a fuel supplied to the combustion chamber 16 is a fossil-based hydrocarbon fuel; v) EI Leer,MK is the nvPM emission index of the gas turbine engine 10 in mg / kg, corrected for system losses, when operating at 7% of the available thrust under given operating conditions and when a fuel supplied to the combustion chamber 16 includes a sustainable aviation fuel (SAF); vi) EI maxAb,NFKis the nvPM emission index of the gas turbine engine 10 in mg / kg, corrected for system losses, when operating at 100% of the available thrust under given operating conditions and when a fuel supplied to the combustion chamber 16 includes a sustainable aviation fuel (SAF); vii) EI Steig,NFK is the system loss-corrected nvPM emission index of the gas turbine engine 10 in mg / kg at operation with 85% of the available thrust under given operating conditions and when a fuel supplied to the combustion chamber 16 includes a sustainable aviation fuel (SAF); and viii) EI Anflug,NFK is the nvPM emission index of the gas turbine engine 10 in mg / kg, corrected for system losses, when operating at 30% of the available thrust under given operating conditions and when a fuel supplied to the combustion chamber 16 includes a sustainable aviation fuel (SAF). Fuel flow rate

[0418] A fuel flow rate W K The fuel flow rate of the gas turbine engine 10 is defined as the rate of fuel flow to the fuel spray nozzles 403, 404 of the combustion chamber 16 (i.e., when the engine 10 is in operation). The fuel flow rate is defined for operation at different percentages of the available thrust under given operating conditions, as defined above. K,Leer The fuel flow rate to the fuel spray nozzles 403, 404 in kg / s at 7% of the available thrust under given operating conditions is given and can correspond to operation in an idle operating phase of the gas turbine engine 10. K,maxAb The fuel flow rate to the fuel spray nozzles 403, 404 in kg / s at 100% of the available thrust under given operating conditions is given and can correspond to operation during a maximum liftoff phase of the gas turbine engine 10. K,Steigis defined as the mass flow rate of the fuel supplied to the fuel spray nozzles 403, 404 in kg / s when the gas turbine engine 10 is operated at 85% of the available thrust under given operating conditions, and can correspond to operation of the gas turbine engine 10 in a climb operation phase. K,Anflug is the mass flow rate of the fuel supplied to the fuel spray nozzles 403, 404 in kg / s when the gas turbine engine 10 is operated at 30% of the available thrust under given operating conditions, and can correspond to operation of the gas turbine engine 10 in an approach operation phase.

[0419] In each example defined or claimed herein, W K,maxAb 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. W K,maxAb It can range from 0.551 to 0.850 kg / s. W K,maxAbcan range from 0.551 to 0.750 kg / s.

[0420] In each example defined or claimed herein, W K,Steig in the range of 0.369 to 1.01 kg / s, preferably 0.415 to 0.923 kg / s, more preferably 0.461 to 0.839 kg / s. W K,Steig It can range from 0.461 to 0.650 kg / s. W K,Steig It can range from 0.461 to 0.600 kg / s.

[0421] In each example defined or claimed herein, W K,Anflug 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. W K,Anflug It can range from 0.166 to 0.300 kg / s. W K,Anflug can range from 0.166 to 0.250 kg / s.

[0422] In each example defined or claimed herein, W K,Leer in the range of 0.0516 to 0.119 kg / s, preferably 0.0581 to 0.109 kg / s, more preferably 0.0645 to 0.0990 kg / s. W K,LeerIt can range from 0.0645 to 0.0850 kg / s. W K,Leer can range from 0.0645 to 0.0750 kg / s. Engine thrust

[0423] The thrust of the gas turbine engine 10 is denoted by the symbol F and is defined for operation at different percentages of the available thrust under given operating conditions, as defined above. F maxAb is defined as the thrust of the gas turbine engine 10 at 100% of the available thrust under given operating conditions in kN. F Leer is defined as the thrust of the gas turbine engine 10 at 7% of the available thrust under given operating conditions in kN.

[0424] In each example defined or claimed herein, F maxAb in the range of 54.1 kN 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. The value of F maxAbcorresponds to the maximum nominal thrust F 00 Alternatively, F maxAb 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.

[0425] In each example defined or claimed herein, F Leer 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, F can be Leer 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.

[0426] The thrust at other operating points (e.g., approach and climb) can be calculated by determining the corresponding percentage of the maximum rated thrust F. 00 be defined. Bypass ratio

[0427] The bypass ratio (BPV) is defined as the ratio of the mass flow rate of the flow through the bypass channel to the mass flow rate of the flow through the engine core.

[0428] In each example defined or claimed herein, the BPV 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 BPV 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 either 4 to 5 or 5 to 6. First and second nvPM emission index ratio idle to MAB

[0429] A first nvPM emission index ratio idle to MAB is defined in the following equation (1): EILeerEImaxAb

[0430] EGG Leer and egg maxAbare as defined elsewhere herein. The first nvPM emission index ratio, idle to MAB, represents the ratio of the nvPM emission index corrected for system losses under idle conditions (e.g., at 7% of available thrust) to the nvPM emission index corrected for system losses at maximum takeoff (e.g., at 100% of available thrust). The first nvPM emission index ratio, idle to MAB, of the gas turbine engine 10 may be less than 0.8.

[0431] In other examples, the first nvPM emission index ratio idle to MAB of the gas turbine engine 10 can be less than 0.708 and preferably less than 0.649 and more preferably less than 0.59.

[0432] The first nvPM emission index ratio idle to MAB of the gas turbine engine 10 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.

[0433] The first nvPM emission index ratio idle to MAB of th...

Claims

[1] Gas turbine engine (10) for an aircraft, comprising: a combustion chamber (16) for rich combustion, rapid cooling, lean combustion (RQL combustion chamber) having a number of fuel spray nozzles (403, 404) in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; and where: A thrust-nvPM emission index ratio is defined as follows: EImaxAbFmaxAbEILeerFLeer where: EGG Leer the nvPM emission index of the gas turbine engine (10) in mg / kg corrected for system losses at operation with approximately 7% of the available thrust under given operating conditions; EGG maxAb the nvPM emission index of the gas turbine engine (10) in mg / kg corrected for system losses at operation with approximately 100% of the available thrust under the given operating conditions; F maxAbthe thrust of the gas turbine engine (10) at approximately 100% of the available thrust in kN is, F Leer the thrust of the gas turbine engine (10) is at about 7% of the available thrust in kN; the thrust-nvPM emission index ratio is greater than 0.09; and the gas turbine engine (10) is configured to supply the fuel spray nozzles (403, 404) with a sustainable aviation fuel (SAF). [2] Gas turbine engine (10) according to claim 1, wherein the thrust-nvPM emission index ratio is greater than 0.0949 and preferably greater than 0.106 and more preferably greater than 0.

118. [3] Gas turbine engine (10) according to claim 1 or claim 2, wherein the thrust-nvPM emission index ratio is 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. [4] Gas turbine engine (10) according to one of the preceding claims, wherein the thrust-nvPM emission index ratio is 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. [5] Gas turbine engine (10) according to one of the preceding claims, wherein the thrust-nvPM emission index ratio is 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. [6] Gas turbine engine (10) according to one of the preceding claims, wherein the thrust-nvPM emission index ratio is 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. [7] Gas turbine engine according to any one of the preceding claims, wherein: a) F maxAb in the range of 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; and / or b) F Leer 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. [8] Gas turbine engine according to 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 the combustion chamber (16) preferably comprises 10-14 duplex fuel spray nozzles and 4-8 single-flow fuel spray nozzles. [9] Gas turbine engine according to claim 8, wherein the duplex fuel spray nozzles (403) are arranged in groups around the circumference of the combustion chamber (16) and wherein optionally the groups of duplex fuel spray nozzles (403) comprise at least two groups arranged diametrically opposite each other. [10] Gas turbine engine according to claim 9, wherein each group of duplex fuel spray nozzles (403) comprises 2-8 nozzles (403). [11] Gas turbine engine according to claim 8 or claim 9 or claim 10, wherein the combustion chamber (16) comprises one or more igniters (405) and the igniter or each igniter (405) is arranged adjacent to one or more of the duplex fuel spray nozzles (403). [12] Gas turbine engine according to one of the preceding claims, wherein the number of fuel spray nozzles per unit of engine core size is in the range of 2.5 to 4.5 and more, preferably in the range of 3 to 4. [13] Gas turbine engine according to one of the preceding claims, wherein the fuel supplied to the combustion chamber (16) comprises a NFK percentage 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%. [14] Method (1000) for operating the gas turbine engine (10) according to one of the preceding claims, wherein the method comprises providing (1002) fuel comprising a sustainable aviation fuel to the fuel spray nozzles (403, 404). [15] Method (1000) for operating a gas turbine engine (10), wherein the gas turbine engine (10) comprises: a combustion chamber (16) for rich combustion, rapid cooling, lean combustion (RQL combustion chamber) having a number of fuel spray nozzles (403, 404) in the range of 14-22 or a number of fuel spray nozzles per unit of engine core size in the range of 2 to 6; and where: A thrust-nvPM emission index ratio is defined as follows: EImaxAbFmaxAbEILeerFLeer where: EGG Leerthe nvPM emission index of the gas turbine engine (10) in mg / kg corrected for system losses at operation with approximately 7% of the available thrust under given operating conditions; EGG maxAb the nvPM emission index of the gas turbine engine (10) in mg / kg corrected for system losses at operation with approximately 100% of the available thrust under the given operating conditions; F maxAb the thrust of the gas turbine engine (10) at approximately 100% of the available thrust in kN is, F Leer the thrust of the gas turbine engine (10) is at about 7% of the available thrust in kN; the thrust-nvPM emission index ratio is greater than 0.09; and the gas turbine engine (10) is configured to provide the fuel spray nozzles (403, 404) with a sustainable aviation fuel (SAF) comprising fuel; wherein the method comprises the provision (1002) of fuel comprising a sustainable aviation fuel to the fuel spray nozzles (403, 404).

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