FUEL FLOW

Abstract

A gas turbine engine (10) for an aircraft is disclosed. The gas turbine engine comprises: a combustion chamber (16) comprising a combustion chamber (120) and a plurality of fuel spray nozzles (124) configured to inject fuel into the combustion chamber (120), wherein the plurality of fuel spray nozzles (124) comprises a first subset (124A) of fuel spray nozzles (124) and a second subset (124B) of fuel spray nozzles (124), wherein the combustion chamber (16) is operable in a state in which each of the fuel spray nozzles of the first subset (124A) of the fuel spray nozzles (124) is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset (124B) of the fuel spray nozzles (124), wherein the number of fuel spray nozzles (124) in the first subset (124A) of the fuel spray nozzles (124) to the number of fuel spray nozzles (124) in the second subset (124B) the fuel spray nozzles (124) are in the range of 1:2 to 1:5.A fuel flow-nvPM emission index ratio is defined as follows: . EIL eer × WK , L eer EI max A b × WK , max A b 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; 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,maxAbThe fuel flow rate to the fuel spray nozzles is 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 (10) is less than 6. The gas turbine engine (10) is configured to supply a sustainable aviation fuel (SAF) to the plurality of fuel spray nozzles (124). A method for operating a 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 for 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 (SAF), as well as to gas turbine engines configured for operation with an SAF fuel. 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 one aspect, a gas turbine engine is intended for an aircraft, comprising one or more of the following elements: a combustion chamber comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and where: A first nvPM emission index ratio idle to MAB (maximum takeoff) 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 60; and the gas turbine engine is configured to provide a sustainable aviation fuel (SAF) to the multitude of fuel spray nozzles.

[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 emission 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 58.4 and can preferably be less than 53.5 and can more preferably be less than 48.6.

[0009] The first nvPM emission index ratio idle to MAB can be less than or equal to 45, and can preferably be less than or equal to 30, and can more preferably be less than or equal to 15.

[0010] The first nvPM emission index ratio idle to MAB can be less than or equal to 8.65 and can preferably be less than or equal to 7.93 and can more preferably be less than or equal to 7.21.

[0011] The first nvPM emission index ratio idle to MAB can be less than or equal to 0.155 and can preferably be less than or equal to 0.142 and can more preferably be less than or equal to 0.129.

[0012] The first nvPM emission index ratio idle to MAB can be greater than or equal to 0.038 and can preferably be greater than or equal to 0.0428 and can more preferably be greater than or equal to 0.0475.

[0013] The first nvPM emission index ratio idle to MAB can be greater than or equal to 1.21 and can preferably be greater than or equal to 1.36 and can more preferably be greater than or equal to 1.52.

[0014] The first nvPM emission index ratio idle to MAB can be in the range of 0.0380 to 8.65 and can preferably be in the range of 0.0428 to 7.93 and can more preferably be in the range of 0.0475 to 7.21.

[0015] The first nvPM emission index ratio idle to MAB can be in the range of 0.0380 to 0.155 and can preferably be in the range of 0.0428 to 0.142 and can more preferably be in the range of 0.0475 to 0.129.

[0016] The first nvPM emission index ratio idle to MAB can be in the range of 1.21 to 8.65 and can preferably be in the range of 1.36 to 7.93 and can more preferably be in the range of 1.52 to 7.21.

[0017] The first nvPM emission index ratio (idle 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, 0.2, 2, 4, 6, 8, 10, 12, 14, 16, 18, 20, 22, 24, 26, 28, 30, 32, 34, 36, 38, 40, 42, 44, 46, 48, 50, 52, 54, 56. 58, 60, or 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 of 0.01 to 0.2, 0.01 to 0.15, 0.01 to 0.07, or 0.01 to 0.05.

[0018] 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 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 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 the available thrust under the same given operating conditions as those under which EI Leer,NFKis calculated, and, if one of the fuels supplied to the combustion chamber is a fossil-based hydrocarbon fuel, is; 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 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.

[0019] The second nvPM emission index ratio, idle to MAB, can be greater than zero.

[0020] 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, and more preferably less than or equal to 0.2.

[0021] 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.

[0022] 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.

[0023] 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.

[0024] 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.

[0025] 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.

[0026] According to a second aspect, a gas turbine engine is intended for an aircraft that includes one or more of the following elements: a combustion chamber comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and where: 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,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 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 the available thrust under the same given operating conditions as those under which EI Leer,NFKis calculated, and, if one of the fuels supplied to the combustion chamber is a fossil-based hydrocarbon fuel, is; 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 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 a sustainable aviation fuel (SAF) to the multitude of fuel spray nozzles.

[0027] 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.

[0028] 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 plurality of fuel spray nozzles.

[0029] According to a fourth aspect, a method for operating a gas turbine engine is provided, which includes the gas turbine engine: a combustion chamber comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and where: A first nvPM emission index ratio idle to MAB (maximum takeoff) 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 60; and The process involves supplying a sustainable aviation fuel to the multitude of fuel spray nozzles.

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

[0031] 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 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 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 the 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 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 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.

[0032] The second nvPM emission index ratio, idle to MAB, can be defined as above in connection with the first aspect.

[0033] According to a fifth aspect, a method for operating a gas turbine engine is provided, wherein the gas turbine engine comprises: a combustion chamber comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; 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,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 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 the available thrust under the same given operating conditions as those under which EI Leer,NFKis calculated, and, if one of the fuels supplied to the combustion chamber is a fossil-based hydrocarbon fuel, is; 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 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 multitude of fuel spray nozzles.

[0034] The second nvPM emission index ratio, idle to MAB, can be defined as above in connection with the first aspect.

[0035] According to a sixth aspect, a gas turbine engine is intended for an aircraft that includes one or more of the following elements: a combustion chamber comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and where: 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,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 6; and the gas turbine engine is configured to provide a sustainable aviation fuel (SAF) to the multitude of fuel spray nozzles.

[0036] The fuel flow-nvPM emission index ratio can be less than 5.93 and preferably less than 5.44 and more preferably less than 4.94.

[0037] The fuel flow nvPM emission index ratio can be less than or equal to 4.5, and preferably less than or equal to 3, and more preferably less than or equal to 1.5.

[0038] The fuel flow nvPM emission index ratio can be less than or equal to 0.879, and preferably less than or equal to 0.806, and more preferably less than or equal to 0.733.

[0039] The fuel flow nvPM emission index ratio can be less than or equal to 0.0181, and preferably less than or equal to 0.0166, and more preferably less than or equal to 0.0151.

[0040] The fuel flow nvPM emission index ratio can be greater than or equal to 0.00351, and preferably greater than or equal to 0.00395, and more preferably greater than or equal to 0.00439.

[0041] The fuel flow nvPM emission index ratio can be greater than or equal to 0.109, and preferably greater than or equal to 0.123, and more preferably greater than or equal to 0.137.

[0042] The fuel flow nvPM emission index ratio can be in the range of 0.00351 to 0.879, and preferably in the range of 0.00395 to 0.806, and more preferably in the range of 0.00439 to 0.733.

[0043] The fuel flow nvPM emission index ratio can be in the range of 0.00351 to 0.0181 and can preferably be in the range of 0.00395 to 0.0166 and can more preferably be in the range of 0.00439 to 0.0151.

[0044] The fuel flow nvPM emission index ratio can be in the range of 0.109 to 0.879, and preferably in the range of 0.123 to 0.806, and more preferably in the range of 0.137 to 0.733.

[0045] The fuel flow nvPM emission index ratio can be 0.003, 0.004, 0.005, 0.01, 0.015, 0.02, 0.025, 0.03, 0.035, 0.04, 0.045, 0.05, 0.055, 0.06, 0.065, 0.07, 0.1, 0.2, 0.3, 0.4, 0.5, 0.6, 0.7, 0.8, 0.9, 1, 1.5, 2, 2.5, 3, 3.5, 4, 4.5, 5, 5.5, 6 or any range defined by any two of these values. Alternatively, the fuel flow nvPM emission index ratio can be in a range between 0.0009 and 0.021, 0.009 and 0.019, 0.009 and 0.007 or 0.009 and 0.006.

[0046] W K,maxAb can be in the range of 0.595 to 1.29 kg / s and can preferably be in the range of 0.670 to 1.19 kg / s and can more preferably be in the range of 0.744 to 1.08 kg / s.

[0047] W K,Leer can be in the range of 0.0695 to 0.118 kg / s and can preferably be in the range of 0.0782 to 0.108 kg / s and can more preferably be in the range of 0.0869 to 0.0981 kg / s.

[0048] W K,Leer can be in the range of 0.0712 to 0.117 kg / s and preferably in the range of 0.0801 to 0.107 kg / s and more, preferably in the range of 0.0890 to 0.0970 kg / s. K,maxAb can be in the range of 0.595 to 1.28 kg / s and preferably in the range of 0.670 to 1.17 kg / s and more preferably in the range of 0.744 to 1.07 kg / s.

[0049] W K,maxAbcan be in the range of 0.701 to 1.29 kg / s and can preferably be in the range of 0.788 to 1.19 kg / s and can more preferably be in the range of 0.876 to 1.08 kg / s.

[0050] 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.

[0051] 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.

[0052] 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 multitude of fuel spray nozzles.

[0053] According to an eighth aspect, a method for operating a gas turbine engine is provided, wherein the gas turbine engine comprises the following: a combustion chamber comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and where: 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 6; and The process involves supplying a sustainable aviation fuel to the multitude of fuel spray nozzles.

[0054] Each of the fuel flow nvPM emission index ratios, W K,Leer and W K,maxaAb can be defined as above in connection with the sixth aspect.

[0055] According to a ninth aspect, a gas turbine engine is intended for an aircraft that includes one or more of the following elements: a combustion chamber comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and where: A thrust-nvPM emission index ratio can be defined as follows: EImaxAbFmaxAbEILerFLeer where: EGG LeerThe 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 0.001; and the gas turbine engine is configured to provide a sustainable aviation fuel (SAF) to the multitude of fuel spray nozzles.

[0056] The thrust-nvPM emission index ratio can be greater than 0.00115 and can preferably be greater than 0.00129 and can more preferably be greater than 0.00144.

[0057] The thrust-nvPM emission index ratio can be greater than 0.0644 and can preferably be greater than 0.0724 and can more preferably be greater than 0.0805.

[0058] The thrust-nvPM emission index ratio can be greater than or equal to 0.003, can preferably be greater than or equal to 0.005, and can more preferably be greater than or equal to 0.007.

[0059] The thrust-nvPM emission index ratio can be greater than or equal to 0.00776, can preferably be greater than or equal to 0.00874, and can more preferably be greater than or equal to 0.00971.

[0060] The thrust-nvPM emission index ratio can be greater than or equal to 0.434, can preferably be greater than or equal to 0.488, and can more preferably be greater than or equal to 0.542.

[0061] The thrust-nvPM emission index ratio can be less than or equal to 1.77, can preferably be less than or equal to 1.62, and can more preferably be less than or equal to 1.48.

[0062] The thrust-nvPM emission index ratio can be less than or equal to 0.0553, can preferably be less than or equal to 0.0507, and can more preferably be less than or equal to 0.0461.

[0063] The thrust-nvPM emission index ratio can be in the range of 0.00776 to 1.77, and preferably in the range of 0.00874 to 1.62, and more preferably in the range of 0.00971 to 1.48.

[0064] The thrust-nvPM emission index ratio can be greater than 0.001, 0.0015, 0.002, 0.0025, 0.003, 0.0035, 0.004, 0.0045, 0.005, 0.01, 0.02, 0.03, 0.04, 0.05, 0.06, 0.07, 0.08, 0.09, 0.1, 0.15, 0.2, 0.25, 0.3, 0.35, 0.4, 0.45, 0.5, 0.55, 0.6, 0.65, 0.7, 0.75, 0.8, 0.85, 0.9, 0.95, 1, 1.05, The value can be 1.1, 1.15, 1.2, 1.25, 1.3, 1.35, 1.4, 1.45, 1.5, 1.55, 1.6, 1.65, 1.7, 1.75, 1.8, 1.85, 1.9, 1.95, 2, 2.5, 3, 3.5, 4, 4.5, 5, 5.5, or 6, or lie within any range defined by any two of these values. For example, the thrust-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.

[0065] The thrust-nvPM emission index ratio can be in the range of 0.434 to 1.77, and preferably in the range of 0.488 to 1.62, and more preferably in the range of 0.542 to 1.48.

[0066] The thrust-nvPM emission index ratio can be in the range of 0.00776 to 0.0553, and preferably in the range of 0.00874 to 0.0507, and more preferably in the range of 0.00971 to 0.0461.

[0067] F maxAb can be in the range of 85.4 kN to 172 kN and can preferably be in the range of 96.1 kN to 158 kN and can preferably be in the range of 106 kN to 144 kN.

[0068] F Leer can be in the range of 5.98 kN to 12.1 kN and can preferably be in the range of 6.72 kN to 11.1 kN and can more preferably be in the range of 7.47 kN to 10.1 kN.

[0069] F maxAb can be in the range of 89.0 kN to 157 kN and can preferably be in the range of 100 kN to 144 kN and can more preferably be in the range of 111 kN to 131 kN.

[0070] F Leercan be in the range of 6.23 kN to 11.0 kN and can preferably be in the range of 7.00 kN to 10.1 kN and can more preferably be in the range of 7.78 kN to 9.13 kN.

[0071] F maxAb Can be in the range of 50 kN to 85 kN and preferably in the range of 57 kN to 78 kN and preferably in the range of 60 kN to 73 kN and more, preferably in the range of 60 kN to 70 kN.

[0072] F Leer Can be in the range of 3.5 kN to 6 kN and preferably in the range of 4 kN to 5.5 kN and preferably in the range of 4.2 kN to 5.2 kN and more, preferably in the range of 4.2 kN to 5 kN.

[0073] 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 multitude of fuel spray nozzles.

[0074] According to an eleventh aspect, a method for operating a gas turbine engine is provided, wherein the gas turbine engine comprises the following: a combustion chamber comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and where: 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 maxAb The 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 0.001; and The process involves supplying a sustainable aviation fuel to the multitude of fuel spray nozzles.

[0075] Each of the thrust-nvPM emission index ratios, F maxAb and F Leer can be defined as in connection with the ninth aspect.

[0076] According to a twelfth aspect, a gas turbine engine is intended for an aircraft that includes one or more of the following elements: a combustion chamber comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and where: An nvPM emission index ratio for lean-burn cruise flight to MAB can be defined as follows: EIReise(lean)EImaxAbBVP 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 nvPM is the system loss-corrected emission index of the gas turbine engine in mg / kg 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 a sustainable aviation fuel (SAF) to the multitude of fuel spray nozzles.

[0077] The nvPM emission index ratio of lean-burn cruise to MAB can be less than 0.119, and preferably less than 0.109, and more preferably less than 0.0989.

[0078] The nvPM emission index ratio of lean-burn cruise to MAB can be less than 0.101 and can preferably be less than 0.0922 and can more preferably be less than 0.0838.

[0079] The nvPM emission index ratio of lean-burn cruise to MAB can be less than or equal to 0.095, and preferably less than or equal to 0.092, and further preferably less than or equal to 0.089.

[0080] The nvPM emission index ratio of lean-burn cruise to MAB can be less than or equal to 0.106, and preferably less than or equal to 0.0972, and further preferably less than or equal to 0.0883.

[0081] The nvPM emission index ratio of lean-burn cruise to MAB can be less than or equal to 0.0887, and preferably less than or equal to 0.0813, and further preferably less than or equal to 0.0739.

[0082] The nvPM emission index ratio of lean-burn cruise to MAB can be greater than or equal to 0.0519, and preferably greater than or equal to 0.0584, and more preferably greater than or equal to 0.0649.

[0083] The nvPM emission index ratio of lean-burn cruise to MAB can be greater than or equal to 0.0578, and preferably greater than or equal to 0.0651, and more preferably greater than or equal to 0.0723.

[0084] The nvPM emission index ratio for lean-burn cruise flight to MAB can be in the range of 0.0519 to 0.106, and preferably in the range of 0.0584 to 0.0972, and more preferably in the range of 0.0649 to 0.0883.

[0085] The nvPM emission index ratio for lean-burn cruise flight to MAB can be in the range of 0.0519 to 0.0887, and preferably in the range of 0.0584 to 0.0813, and more preferably in the range of 0.0649 to 0.0739.

[0086] The nvPM emission index ratio for lean-burn cruise to MAB can be in the range of 0.0578 to 0.106, and preferably in the range of 0.0651 to 0.0972, and more preferably in the range of 0.0723 to 0.0883.

[0087] 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.

[0088] The BPV can be in the range of 6.63 to 13.4 and preferably in the range of 7.46 to 12.3 and more preferably in the range of 8.29 to 11.1.

[0089] The BPV can be in the range of 8.36 to 13.4 and preferably in the range of 9.40 to 12.3 and further preferably in the range of 10.4 to 11.1.

[0090] The BPV can be in the range of 6.63 to 10.3 and preferably in the range of 7.46 to 9.38 and further preferably in the range of 8.29 to 8.53.

[0091] The BPV can be in the range of 3.5 to 6.5 and more, preferably in the range of 4 to 6 and even more preferably in the range of either 4 to 5 or 5 to 6.

[0092] An nvPM emission index ratio for cruise flight with rich combustion to MAB can be defined as follows: EIRaise(fett)EImaxAbBPV where: EGG Reise(fett) which can be defined as follows: ICE climb + ICE approach 2 EGG Steigthe 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; and EGG maxAb 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 operating conditions as those under which EI Steig is calculated; and where the nvPM emission index ratio of cruise flight with rich combustion to MAB is less than 20.

[0093] The nvPM emission index ratio of cruise flight with rich combustion to MAB can be less than 19 and can preferably be less than 17.5 and can more preferably be less than 15.9.

[0094] The nvPM emission index ratio of cruise flight with rich combustion to MAB can be less than or equal to 12, and can preferably be less than or equal to 9, and can more preferably be less than or equal to 6.

[0095] The nvPM emission index ratio of cruise flight with rich combustion to MAB can be less than or equal to 4.54 and can preferably be less than or equal to 4.17 and can further preferably be less than or equal to 3.79.

[0096] The nvPM emission index ratio of cruise flight with rich combustion to MAB can be less than or equal to 0.066 and can preferably be less than or equal to 0.0605 and can further preferably be less than or equal to 0.055.

[0097] The nvPM emission index ratio of cruise flight with rich combustion to MAB can be greater than or equal to 0.0374 and can preferably be greater than or equal to 0.0421 and can more preferably be greater than or equal to 0.0468.

[0098] The nvPM emission index ratio of cruise flight with rich combustion to MAB can be greater than or equal to 1.41 and can preferably be greater than or equal to 1.58 and can further preferably be greater than or equal to 1.76.

[0099] The nvPM emission index ratio of cruise flight with rich combustion to MAB can be in the range of 0.0374 to 4.54 and can preferably be in the range of 0.0421 to 4.17 and can further preferably be in the range of 0.0468 to 3.79.

[0100] The nvPM emission index ratio of cruise flight with rich combustion to MAB can be in the range of 0.0374 to 0.0660 and can preferably be in the range of 0.0421 to 0.0605 and can further preferably be in the range of 0.0468 to 0.0550.

[0101] The nvPM emission index ratio for cruise flight with rich combustion to MAB can be in the range of 1.41 to 4.54 and can preferably be in the range of 1.58 to 4.17 and can further preferably be in the range of 1.76 to 3.79.

[0102] The nvPM emission index ratio of rich combustion cruise flight to MAB can be less than 1, 2, 3, 4, 5, 6, 7, 8, 9, 10, 11, 12, 13, 14, 15, 16, 17, 18, 19 or 20, or lie within any range defined by any two of these values.

[0103] The nvPM emission index ratio of rich combustion cruise flight to MAB can be less than 0.03, 0.33, 0.63, 0.93, 1.23, 1.53, 1.83, 2.13, 2.43, 2.73, 3.03, 3.33, 3.63, 3.93, 4.23, 4.53, 4.83, 5.13, 5.43, 5.73 or 6.03, or lie within any range defined by any two of these values.

[0104] According to a thirteenth aspect, a gas turbine engine is intended for an aircraft that includes one or more of the following elements: a combustion chamber comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and where: An nvPM emission index ratio for cruise flight with rich combustion to MAB can be defined as follows: EIRaise(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 EI Steig is calculated; EGG maxAb 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 operating conditions as those under 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 20; and the gas turbine engine is configured to provide a sustainable aviation fuel (SAF) to the multitude of fuel spray nozzles.

[0105] 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.

[0106] 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 plurality of fuel spray nozzles.

[0107] According to a fifteenth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising the following: a combustion chamber comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and where: 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 nvPM is the system loss-corrected emission index of the gas turbine engine in mg / kg 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 multitude of fuel spray nozzles.

[0108] The nvPM emission index ratio of lean-burn cruise flight to MAB and to or the BPV can be defined as above in connection with the twelfth aspect.

[0109] 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; EGGmaxAb 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 operating conditions as those under which EI Steig is calculated; and where the nvPM emission index ratio of cruise flight with rich combustion to MAB can be less than 20.

[0110] 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.

[0111] According to a sixteenth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising the following: a combustion chamber comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; 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 EI Steig is calculated; EGG maxAb 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 operating conditions as those under 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 20; and The process involves supplying a sustainable aviation fuel to the multitude of fuel spray nozzles.

[0112] 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.

[0113] According to a seventeenth aspect, a gas turbine engine is intended for an aircraft, comprising one or more of the following elements: a combustion chamber comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and where: A MAB-nvPM emission index ratio can be defined as follows: EImaxAb,NFKEImaxAb,FK where: EGG maxAb,NFKthe 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 one of the fuels supplied to the multiple fuel spray nozzles includes a sustainable aviation fuel (SAF); and EGG maxAb,FK 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, when one of the multiple fuel spray nozzles supplies 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 a fuel comprehensive to the multitude of fuel spray nozzles.

[0114] The MAB-nvPM emission index ratio can be greater than zero.

[0115] 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.

[0116] 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.

[0117] 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.

[0118] 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.

[0119] 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.

[0120] 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.

[0121] 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.

[0122] A climb-nvPM emission 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 the given operating conditions, or under other given operating conditions where a fuel supplied to the multiple 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,NFK is calculated, and, if one of the multiple fuel spray nozzles supplied 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.

[0123] The climb-nvPM emission index ratio can be greater than zero.

[0124] 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.

[0125] 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.

[0126] 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.

[0127] 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.

[0128] 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.

[0129] 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.

[0130] 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.

[0131] 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 multiple 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 EI Anflug,NFK is calculated and, if one of the multiple fuel spray nozzles is a fossil-based hydrocarbon fuel, is; and where the approach nvPM emission index ratio of the gas turbine engine may be less than 1.

[0132] The approach-nvPM emission index ratio can be greater than zero.

[0133] 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.

[0134] 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.

[0135] 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.

[0136] 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.

[0137] 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.

[0138] 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.

[0139] 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.

[0140] 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 the available thrust under the given operating conditions or under other given operating conditions, and if a fuel supplied to the multitude of 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 the available thrust under the same given operating conditions as those under which EI Leer,NFK is calculated, and, if one of the multiple fuel spray nozzles supplied is a fossil-based hydrocarbon fuel, it is; and where the idle nvPM emission index ratio of the gas turbine engine can be less than 1.

[0141] The idle-nvPM emission index ratio can be greater than zero.

[0142] 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.

[0143] 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.

[0144] 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.

[0145] 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.

[0146] 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.

[0147] 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.

[0148] The idle nvPM emission index ratio can be 0.08, 0.09, 0.1, 0.11, 0.12, 0.13, 0.14, 0.15, 0.16, 0.17, 0.18, 0.19, 0.2 or within any range defined by any two of these values.

[0149] According to an eighteenth aspect, a gas turbine engine is intended for an aircraft that includes one or more of the following elements: a combustion chamber comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and where: A climb-to-emissions index ratio can be defined as follows: Ice cream dough, NFK Ice cream dough, FK where: EGGSteig,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 one of the fuels supplied to the multiple 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,NFK is calculated, and, if one of the multiple fuel spray nozzles supplied 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 a fuel comprehensive to the multitude of fuel spray nozzles.

[0150] The climb-nvPM emission index ratio can be defined as in connection with the seventeenth aspect.

[0151] According to a nineteenth aspect, a gas turbine engine is intended for an aircraft that includes one or more of the following elements: a combustion chamber comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and where: 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 given operating conditions, where one of the fuels supplied to the multiple 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 EI Anflug,NFK is calculated and, if one of the multiple fuel spray nozzles supplied is a fossil-based hydrocarbon fuel, 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 a fuel comprehensive to the multitude of fuel spray nozzles.

[0152] The approach-nvPM emission index ratio can be defined as above in connection with the seventeenth aspect.

[0153] According to a twentieth aspect, a gas turbine engine is intended for an aircraft that includes one or more of the following elements: a combustion chamber comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and where: 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 given operating conditions, and if a fuel supplied to the multiple 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 the available thrust under the same given operating conditions as those under which EI Leer,NFK is calculated, and, if one of the multiple fuel spray nozzles supplied 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 a fuel comprehensive to the multitude of fuel spray nozzles.

[0154] The idle-nvPM emission index ratio can be defined as above in connection with the seventeenth aspect.

[0155] 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 plurality of fuel spray nozzles.

[0156] According to a twenty-second aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising the following: a combustion chamber comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and where: A MAB-nvPM emission index ratio can be defined as follows: EImaxAb,NFKEImaxAb,FK where: EGGmaxAb,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 one of the fuels supplied to the multiple fuel spray nozzles includes a sustainable aviation fuel (SAF); and EGG maxAb,FK 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, when one of the multiple fuel spray nozzles supplies 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 multitude of fuel spray nozzles.

[0157] The MAB-nvPM emission index ratio can be defined as in connection with the seventeenth aspect.

[0158] A climb-nvPM emission 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 the given operating conditions, or under other given operating conditions where a fuel supplied to the multiple 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 one of the multiple fuel spray nozzles supplied 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.

[0159] The climb-nvPM emission index ratio can be defined as in connection with the seventeenth aspect.

[0160] 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 multiple 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 EI Anflug,NFK is determined, and, if one of the multiple fuel spray nozzles supplied 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.

[0161] The approach-nvPM emission index ratio can be defined as in connection with the seventeenth aspect.

[0162] 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 the available thrust under the given operating conditions or under other given operating conditions, and if a fuel supplied to the multitude of 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 the available thrust under the same given operating conditions as those under which EI Leer,NFK is determined, and, if one of the multiple fuel spray nozzles supplied is a fossil-based hydrocarbon fuel, is; and where the idle nvPM emission index ratio of the gas turbine engine can be less than 1.

[0163] The idle-nvPM emission index ratio can be defined as in connection with the seventeenth aspect.

[0164] According to a twenty-third aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising the following: a combustion chamber comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and where: A climb-to-emissions 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 given operating conditions, where one of the fuels supplied to the multiple 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,NFK is determined, and, if one of the multiple fuel spray nozzles supplied is a fossil-based hydrocarbon fuel, 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 multitude of fuel spray nozzles.

[0165] The climb-nvPM emission index ratio can be defined as in connection with the seventeenth aspect.

[0166] According to a twenty-fourth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising the following: a combustion chamber comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and where: 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 given operating conditions, where one of the fuels supplied to the multiple 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 EI Anflug,NFK is determined, and, if one of the multiple fuel spray nozzles supplied is a fossil-based hydrocarbon fuel, is; 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 multitude of fuel spray nozzles.

[0167] The approach-nvPM emission index ratio can be defined as in connection with the seventeenth aspect.

[0168] According to a twenty-fifth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising the following: a combustion chamber comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and where: 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 given operating conditions, and if a fuel supplied to the multiple 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 the available thrust under the same given operating conditions as those under which EI Leer,NFK is determined, and, if one of the multiple fuel spray nozzles supplied 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 multitude of fuel spray nozzles.

[0169] The idle-nvPM emission index ratio can be defined as in connection with the seventeenth aspect.

[0170] According to a twenty-sixth aspect, a gas turbine engine is intended for an aircraft, comprising one or more of the following elements: a combustion chamber comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and where: 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 one of the fuels supplied to the multiple fuel spray nozzles includes a sustainable aviation fuel (SAF). 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, where one of the fuels supplied to the multiple fuel spray nozzles is a fossil-based hydrocarbon fuel; and W K,maxAbthe mass flow rate of the fuel supplied to the multitude of 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 a fuel comprehensive to the multitude of fuel spray nozzles.

[0171] The fuel flow in kg / s modified with respect to the MAB-nvPM emission index ratio can be greater than zero.

[0172] The fuel flow in kg / s modified with respect to the MAB-nvPM emission index ratio can be less than 1.29, can more preferably be less than 1.19 and can even more preferably be less than 1.08.

[0173] The fuel flow rate in kg / s modified with respect to the MAB-nvPM emission index ratio can be greater than or equal to 0.384 and can preferably be greater than or equal to 0.432 and can further preferably be greater than or equal to 0.481.

[0174] The fuel flow in kg / s modified with respect to the MAB-nvPM emission index ratio can be less than or equal to 0.834, can more preferably be less than or equal to 0.764 and can even more preferably be less than or equal to 0.695.

[0175] The fuel flow rate in kg / s modified with respect to the MAB-nvPM emission index ratio can be in the range of 0.384 to 0.834 and can preferably be in the range of 0.432 to 0.764 and can more preferably be in the range of 0.481 to 0.695.

[0176] The fuel flow in kg / s modified with respect to the MAB-nvPM emission index ratio can be greater than or equal to 0.453 and can preferably be greater than or equal to 0.509 and can further preferably be greater than or equal to 0.566.

[0177] The fuel flow rate in kg / s modified with respect to the MAB-nvPM emission index ratio can be in the range of 0.453 to 0.834 and can preferably be in the range of 0.509 to 0.764 and can more preferably be in the range of 0.566 to 0.695.

[0178] The fuel flow in kg / s modified with respect to the MAB-nvPM emission index ratio can be 0.38, 0.384, 0.4, 0.42, 0.44, 0.46, 0.48, 0.5, 0.52, 0.54, 0.56, 0.58, 0.6, 0.62, 0.64, 0.66, 0.68, 0.7, 0.72, 0.74, 0.76, 0.78, 0.8, 0.82, 0.83, 0.834 or lie within any range defined by any two of these values. Alternatively, the fuel flow in kg / s modified with respect to the MAB-nvPM emission index ratio can be in a range between 0.45 and 0.65 or 0.45 and 0.6.

[0179] W K,maxAb can be in the range of 0.595 to 1.29 kg / s and can preferably be in the range of 0.670 to 1.19 kg / s and can more preferably be in the range of 0.744 to 1.08 kg / s.

[0180] W K,maxAb can be in the range of 0.595 to 1.28 kg / s and can preferably be in the range of 0.670 to 1.17 kg / s and can more preferably be in the range of 0.744 to 1.07 kg / s.

[0181] W K,maxAbcan be in the range of 0.701 to 1.29 kg / s and can preferably be in the range of 0.788 to 1.19 kg / s and can more preferably be in the range of 0.876 to 1.08 kg / s.

[0182] 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.

[0183] 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 multiple 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 by any of the multiple fuel spray nozzles is a fossil-based hydrocarbon fuel; and W K,Steig the mass flow rate of the fuel supplied to the multitude of 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 calculated; and where the fuel flow of the gas turbine engine, modified with respect to the climb-nvPM emission index ratio, can be less than 2 kg / s.

[0184] The fuel flow in kg / s modified with respect to the climb-nvPM emission index ratio can be greater than zero.

[0185] The fuel flow rate in kg / s modified with respect to the climb-nvPM emission index ratio can be less than 1.05, can more preferably be less than 0.96 and can even more preferably be less than 0.873.

[0186] The fuel flow rate in kg / s modified with respect to the climb-nvPM emission index ratio can be greater than or equal to 0.234 and can preferably be greater than or equal to 0.263 and can further preferably be greater than or equal to 0.292.

[0187] The fuel flow rate in kg / s modified with respect to the climb-nvPM emission index ratio can be less than or equal to 0.498, can more preferably be less than or equal to 0.456 and can even more preferably be less than or equal to 0.415.

[0188] The fuel flow rate in kg / s modified with respect to the climb-nvPM emission index ratio can be in the range of 0.234 to 0.498 and can preferably be in the range of 0.263 to 0.456 and can more preferably be in the range of 0.292 to 0.415.

[0189] The fuel flow rate in kg / s modified with respect to the climb-nvPM emission index ratio can be less than or equal to 0.496, and preferably less than or equal to 0.455, and even more preferably less than or equal to 0.413.

[0190] The fuel flow rate in kg / s modified with respect to the climb-nvPM emission index ratio can be in the range of 0.234 to 0.496 and can preferably be in the range of 0.263 to 0.455 and can more preferably be in the range of 0.292 to 0.413.

[0191] The fuel flow rate in kg / s modified with respect to the climb-nvPM emission index ratio can be greater than or equal to 0.274 and can preferably be greater than or equal to 0.308 and can further preferably be greater than or equal to 0.342.

[0192] The fuel flow rate in kg / s modified with respect to the climb-nvPM emission index ratio can be less than or equal to 0.498, and preferably less than or equal to 0.456, and even more preferably less than or equal to 0.415.

[0193] The fuel flow rate in kg / s modified with respect to the climb-nvPM emission index ratio can be in the range of 0.274 to 0.498 and can preferably be in the range of 0.308 to 0.456 and can more preferably be in the range of 0.342 to 0.415.

[0194] The fuel flow modified with respect to the climb-nvPM emission index ratio can be less than or equal to 0.23, 0.234, 0.24, 0.26, 0.28, 0.3, 0.32, 0.34, 0.36, 0.38, 0.4, 0.42, 0.44, 0.46, 0.48, 0.498, 0.5, or within any range defined by any two of these values. Alternatively, the fuel flow modified with respect to the climb-nvPM emission index ratio, expressed in kg / s, can be within a range of 0.275 to 0.475 or 0.3 to 0.4.

[0195] W K,Steig can be in the range of 0.492 to 1.05 kg / s and can preferably be in the range of 0.554 to 0.960 kg / s and can more preferably be in the range of 0.616 to 0.873 kg / s.

[0196] W K,Steig can be in the range of 0.492 to 1.05 kg / s and can preferably be in the range of 0.554 to 0.957 kg / s and can more preferably be in the range of 0.616 to 0.870 kg / s.

[0197] W K,Steigcan be in the range of 0.577 to 1.05 kg / s and can preferably be in the range of 0.649 to 0.960 kg / s and can more preferably be in the range of 0.721 to 0.873 kg / s.

[0198] 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.

[0199] 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 multiple 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 by any of the multiple fuel spray nozzles is a fossil-based hydrocarbon fuel; and W K,Anflug the mass flow rate of the fuel supplied to the multitude of 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.

[0200] The fuel flow in kg / s, modified with respect to the approach-nvPM emission index ratio, can be greater than zero.

[0201] The fuel flow in kg / s modified with respect to the approach-nvPM emission index ratio can be less than 0.343, can more preferably be less than 0.314 and can even more preferably be less than 0.286.

[0202] The fuel flow rate in kg / s modified with respect to the approach nvPM emission index ratio can be greater than or equal to 0.0269 and can preferably be greater than or equal to 0.0302 and can further preferably be greater than or equal to 0.0336.

[0203] The fuel flow in kg / s modified with respect to the approach-nvPM emission index ratio can be less than or equal to 0.0526, can more preferably be less than or equal to 0.0482 and can further preferably be less than or equal to 0.0439.

[0204] The fuel flow rate in kg / s modified with respect to the approach-nvPM emission index ratio can be in the range of 0.0269 to 0.0526 and can preferably be in the range of 0.0302 to 0.0482 and can further preferably be in the range of 0.0336 to 0.0439.

[0205] The fuel flow in kg / s modified with respect to the approach nvPM emission index ratio can be less than or equal to 0.0524 and can preferably be less than or equal to 0.0480 and can further preferably be less than or equal to 0.0437.

[0206] The fuel flow rate in kg / s modified with respect to the approach-nvPM emission index ratio can be in the range of 0.0269 to 0.0524 and can preferably be in the range of 0.0302 to 0.0480 and can further preferably be in the range of 0.0336 to 0.0437.

[0207] The fuel flow in kg / s modified with respect to the approach-nvPM emission index ratio can be greater than or equal to 0.0301 and can preferably be greater than or equal to 0.0339 and can further preferably be greater than or equal to 0.0376.

[0208] The fuel flow rate in kg / s modified with respect to the approach nvPM emission index ratio can be less than or equal to 0.0526 and can preferably be less than or equal to 0.0482 and can further preferably be less than or equal to 0.0439.

[0209] The fuel flow rate in kg / s, modified with respect to the approach-nvPM emission index ratio, can be in the range of 0.0301 to 0.0526 and can preferably be in the range of 0.0339 to 0.0482 and can further preferably be in the range of 0.0376 to 0.0439.

[0210] The fuel flow rate in kg / s modified with respect to the approach-nvPM emission index ratio can be 0.026, 0.0269, 0.028, 0.03, 0.032, 0.034, 0.036, 0.038, 0.04, 0.042, 0.044, 0.046, 0.048, 0.05, 0.052, 0.0526, 0.053, or within any range defined by any two of these values. Alternatively, the fuel flow rate in kg / s modified with respect to the approach-nvPM emission index ratio can be within a range of 0.02 to 0.1 or 0.05 to 0.075.

[0211] W K,Anflug can be in the range of 0.175 to 0.343 kg / s and can preferably be in the range of 0.197 to 0.314 kg / s and can more preferably be in the range of 0.219 to 0.286 kg / s.

[0212] W K,Anflug can be in the range of 0.175 to 0.341 kg / s and can preferably be in the range of 0.197 to 0.313 kg / s and can more preferably be in the range of 0.219 to 0.284 kg / s.

[0213] W K,Anflugcan be in the range of 0.196 to 0.343 kg / s and can preferably be in the range of 0.220 to 0.314 kg / s and can more preferably be in the range of 0.245 to 0.286 kg / s.

[0214] 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.

[0215] 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 multiple 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 available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied by any of the multiple fuel spray nozzles is a fossil-based hydrocarbon fuel; and W K,Leer the mass flow rate of the fuel supplied to the multitude of 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.

[0216] The fuel flow rate in kg / s, modified with respect to the idle-nvPM emission index ratio, can be greater than zero.

[0217] The fuel flow rate in kg / s modified with respect to the idle-nvPM emission index ratio can be less than 0.118, and more preferably less than 0.108, and even more preferably less than 0.0981.

[0218] The fuel flow rate in kg / s modified with respect to the idle-nvPM emission index ratio can be greater than or equal to 0.00666 and can preferably be greater than or equal to 0.00749 and can further preferably be greater than or equal to 0.00833.

[0219] The fuel flow rate in kg / s modified with respect to the idle-nvPM emission index ratio can be less than or equal to 0.0113, and more preferably less than or equal to 0.0104, and further preferably less than or equal to 0.0094.

[0220] The fuel flow rate in kg / s modified with respect to the idle-nvPM emission index ratio can be in the range of 0.00666 to 0.0113 and can preferably be in the range of 0.00749 to 0.0104 and can further preferably be in the range of 0.00833 to 0.0094.

[0221] The fuel flow rate in kg / s modified with respect to the idle-nvPM emission index ratio can be greater than or equal to 0.00682, and preferably greater than or equal to 0.00767, and further preferably greater than or equal to 0.00853. The fuel flow rate in kg / s modified with respect to the idle-nvPM emission index ratio can be less than or equal to 0.0112, and preferably less than or equal to 0.0103, and further preferably less than or equal to 0.00929.

[0222] The fuel flow rate in kg / s modified with respect to the idle-nvPM emission index ratio can be in the range of 0.00682 to 0.0112 and can preferably be in the range of 0.00767 to 0.0103 and can further preferably be in the range of 0.00853 to 0.00929.

[0223] The fuel flow rate modified with respect to the idle-nvPM emission index ratio, expressed in kg / s, can be 0.0065, 0.00666, 0.007, 0.0075, 0.008, 0.0085, 0.009, 0.0095, 0.01, 0.0113, 0.015, or within any range defined by any two of these values. Alternatively, the fuel flow rate modified with respect to the idle-nvPM emission index ratio, expressed in kg / s, can be within a range between 0.01 and 0.03 or 0.015 and 0.025.

[0224] W K,Leer can be in the range of 0.0695 to 0.118 kg / s and can preferably be in the range of 0.0782 to 0.108 kg / s and can more preferably be in the range of 0.0869 to 0.0981 kg / s.

[0225] W K,Leer can be in the range of 0.0712 to 0.117 kg / s and can preferably be in the range of 0.0801 to 0.107 kg / s and can more preferably be in the range of 0.0890 to 0.0970 kg / s.

[0226] 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.

[0227] According to a twenty-seventh aspect, a gas turbine engine is intended for an aircraft, comprising one or more of the following elements: a combustion chamber comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and where: 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 one of the fuels supplied to the multiple 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 by any of the multiple fuel spray nozzles is a fossil-based hydrocarbon fuel; and W K,Steigthe mass flow rate of the fuel supplied to the multitude of 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 a fuel comprehensive to the multitude of fuel spray nozzles.

[0228] 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.

[0229] According to a twenty-eighth aspect, a gas turbine engine is intended for an aircraft, comprising one or more of the following elements: a combustion chamber comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; 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 multiple 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,NFKis calculated when the fuel supplied by any of the multiple fuel spray nozzles is a fossil-based hydrocarbon fuel; and W K,Anflug the mass flow rate of the fuel supplied to the multitude of 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 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 gas turbine engine is configured to provide a fuel comprehensive to the multitude of fuel spray nozzles.

[0230] The fuel flow and / or W modified with respect to the approach-nvPM emission index ratio K,Anflugcan be defined as above in connection with the twenty-sixth aspect.

[0231] According to a twenty-ninth aspect, a gas turbine engine is intended for an aircraft, comprising one or more of the following elements: a combustion chamber comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and where: 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 multiple 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 available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied by any of the multiple fuel spray nozzles is a fossil-based hydrocarbon fuel; and W K,Leerthe mass flow rate of the fuel supplied to the multitude of 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 gas turbine engine is configured to provide a fuel comprehensive to the multitude of fuel spray nozzles.

[0232] 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.

[0233] 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 the provision of fuel comprising a sustainable aviation fuel to the plurality of fuel spray nozzles.

[0234] 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 comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; where: 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 one of the fuels supplied to the multiple fuel spray nozzles includes a sustainable aviation fuel (SAF). 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, where one of the fuels supplied to the multiple fuel spray nozzles is a fossil-based hydrocarbon fuel; and W K,maxAbthe mass flow rate of the fuel supplied to the multitude of 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 a sustainable aviation fuel (SAF) to the multitude of fuel spray nozzles.

[0235] The fuel flow modified with respect to the MAB-nvPM emission index ratio and / or W K,maxAb can be defined as above in connection with the twenty-sixth aspect.

[0236] 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,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 operating conditions, if a fuel supplied to the multiple 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 by any of the multiple fuel spray nozzles is a fossil-based hydrocarbon fuel; and W K,Steigthe mass flow rate of the fuel supplied to the multitude of 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 calculated; and where the fuel flow of the gas turbine engine, modified with respect to the climb-nvPM emission index ratio, can be less than 2 kg / s.

[0237] 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.

[0238] 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,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 operating conditions, if a fuel supplied to the multiple 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 by any of the multiple fuel spray nozzles is a fossil-based hydrocarbon fuel; and W K,Anflugthe mass flow rate of the fuel supplied to the multitude of 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.

[0239] 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.

[0240] 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 multiple 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 available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied by any of the multiple fuel spray nozzles is a fossil-based hydrocarbon fuel; and W K,Leerthe mass flow rate of the fuel supplied to the multitude of 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.

[0241] 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.

[0242] According to a thirty-second aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising the following: a combustion chamber comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and where: 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 one of the fuels supplied to the multiple 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 by any of the multiple fuel spray nozzles is a fossil-based hydrocarbon fuel; and W K,Steigthe mass flow rate of the fuel supplied to the multitude of 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 a sustainable aviation fuel (SAF) to the multitude of fuel spray nozzles.

[0243] 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.

[0244] According to a thirty-third aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising the following: a combustion chamber comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and where: 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, where a fuel supplied to the multiple 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,NFKis calculated when the fuel supplied by any of the multiple fuel spray nozzles is a fossil-based hydrocarbon fuel; and W K,Anflug the mass flow rate of the fuel supplied to the multitude of 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 a sustainable aviation fuel (SAF) to the multitude of fuel spray nozzles.

[0245] The fuel flow and / or W modified with respect to the approach-nvPM emission index ratio K,Anflugcan be defined as above in connection with the twenty-sixth aspect.

[0246] According to a thirty-fourth aspect, a method for operating a gas turbine engine is provided, the gas turbine engine comprising the following: a combustion chamber comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; and where: 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 multiple 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 available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied by any of the multiple fuel spray nozzles is a fossil-based hydrocarbon fuel; and W K,Leerthe mass flow rate of the fuel supplied to the multitude of 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 a sustainable aviation fuel (SAF) to the multitude of fuel spray nozzles.

[0247] 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.

[0248] 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 comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; where: 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, where the fuel supplied to the multiple 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 given operating conditions, where one of the fuels supplied to the multiple 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, where one of the fuels supplied to the multiple 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 one of the fuels supplied to the multiple 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 a sustainable aviation fuel (SAF) to the multitude of fuel spray nozzles.

[0249] The nvPM emission index ratio for lean-burn cruise flight can be greater than zero.

[0250] The nvPM emission index ratio for lean-burn cruise flight can be less than or equal to 0.9 and preferably less than or equal to 0.8.

[0251] The nvPM emission index ratio for lean-burn cruise flight can be less than or equal to 0.732 and can preferably be less than or equal to 0.671 and can further preferably be less than or equal to 0.61.

[0252] The nvPM emission index ratio for lean-burn cruise flight can be greater than or equal to 0.455, and preferably greater than or equal to 0.512, and more preferably greater than or equal to 0.569.

[0253] The nvPM emission index ratio for lean-burn cruise flight can be in the range of 0.455 to 0.732, and preferably in the range of 0.512 to 0.671, and more preferably in the range of 0.569 to 0.610.

[0254] 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, 0.70, 0.65, 0.6, 0.55, 0.5, 0.45, 0.4, 0.35, 0.3, 0.25, 0.2, 0.15, 0.1, or 0.05, or within any 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.

[0255] 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, 0.71, 0.72, 0.73 or 0.732, or within any range defined by any two of these values.

[0256] 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 multitude of fuel spray nozzles 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 operating conditions as those under which EI Leer,NFK is calculated when a fuel supplied to the multitude of 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 available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied by any of the multiple fuel spray nozzles is a fossil-based hydrocarbon fuel; and 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 operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied to one of the multiple 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.

[0257] The nvPM emission index ratio of idle to MAB can be greater than zero.

[0258] 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, and even more preferably less than or equal to 0.4, and even more preferably less than or equal to 0.2.

[0259] The nvPM emission index ratio idle to MAB can be less than or equal to 0.178 and can preferably be less than or equal to 0.164 and can further preferably be less than or equal to 0.149.

[0260] 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.

[0261] 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.

[0262] 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.

[0263] 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.

[0264] 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, if a fuel supplied to the multitude of 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 maxAb,NFK is calculated when a fuel supplied to the multitude of 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, during operation at approximately 100% of available thrust under the same operating conditions as those under which EI maxAb,NFK is calculated when the fuel supplied by any of the multiple fuel spray nozzles is a fossil-based hydrocarbon 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 maxAb,NFK is calculated when the fuel supplied to one of the multiple 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.

[0265] The nvPM emission index ratio of lean-burn cruise to MAB can be greater than zero.

[0266] The nvPM emission index ratio of lean-burn cruise to MAB can be less than or equal to 0.98 and can preferably be less than or equal to 0.96.

[0267] The nvPM emission index ratio of lean-burn cruise to MAB can be less than or equal to 0.95 and can preferably be less than or equal to 0.944.

[0268] The nvPM emission index ratio of lean-burn cruise to MAB can be greater than or equal to 0.704, preferably greater than or equal to 0.792 and more preferably greater than or equal to 0.88.

[0269] The nvPM emission index ratio for lean-burn cruise to MAB can be in the range of 0.704 to 0.960, and preferably in the range of 0.792 to 0.950, and more preferably in the range of 0.880 to 0.944.

[0270] The nvPM emission index ratio for 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 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 the range of 0.91 to 0.99 or 0.93 to 0.97.

[0271] The nvPM emission index ratio of lean-burn cruise to MAB can be 0.7, 0.704, 0.725, 0.75, 0.775, 0.8, 0.825, 0.85, 0.875, 0.9, 0.925, 0.95, 0.96 or lie within any range defined by any two of these values.

[0272] 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 multitude of 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 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 operating conditions as those under which EI Leer,NFK is calculated when a fuel supplied to the multitude of 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 multitude of 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 available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied to one of the multiple 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 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 operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied by any of the multiple 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 one of the multiple 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.

[0273] The nvPM emission index ratio of idle to cruise flight with lean combustion can be greater than zero.

[0274] The nvPM emission index ratio of idle to cruise flight with lean combustion can be 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.

[0275] The nvPM emission index ratio of idle to cruise flight with lean combustion can be less than or equal to 0.203 and can preferably be less than or equal to 0.186 and can further preferably be less than or equal to 0.169.

[0276] The nvPM emission index ratio of idle to cruise flight with lean combustion can be greater than or equal to 0.125 and can preferably be greater than or equal to 0.141 and can more preferably be greater than or equal to 0.157.

[0277] The nvPM emission index ratio of idle to cruise flight with lean combustion can be in the range of 0.125 to 0.203 and can preferably be in the range of 0.141 to 0.186 and can more preferably be in the range of 0.157 to 0.169.

[0278] 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.

[0279] The nvPM emission index ratio idle to cruise flight with lean combustion can be 0.125, 0.13, 0.135, 0.14, 0.145, 0.15, 0.155, 0.16, 0.165, 0.17, 0.175, 0.18, 0.185, 0.19, 0.195, 0.2, 0.203 or lie in any range defined by any two of these values.

[0280] 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 multitude of 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 multitude of 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 by any of the multiple 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 one of the multiple 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.

[0281] The nvPM emission index ratio for cruise flight with rich combustion can be greater than zero.

[0282] The nvPM emission index ratio for cruise flight with rich combustion can be less than or equal to 0.8, and preferably less than or equal to 0.6, and even more preferably less than or equal to 0.4.

[0283] The nvPM emission index ratio for cruise flight with rich combustion can be less than or equal to 0.303 and can preferably be less than or equal to 0.278 and can further preferably be less than or equal to 0.252.

[0284] The nvPM emission index ratio for cruise flight with rich combustion can be greater than or equal to 0.123 and can preferably be greater than or equal to 0.138 and can more preferably be greater than or equal to 0.154.

[0285] The nvPM emission index ratio for cruise flight with rich combustion can be in the range of 0.123 to 0.303 and can preferably be in the range of 0.138 to 0.278 and can further preferably be in the range of 0.154 to 0.252.

[0286] 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.

[0287] The nvPM emission index ratio for cruise flight with rich combustion can be 0.12, 0.123, 0.13, 0.14, 0.15, 0.16, 0.17, 0.18, 0.19, 0.2, 0.21, 0.22, 0.23, 0.24, 0.25, 0.26, 0.27, 0.28, 0.29, 0.3, 0.303 or lie within any range defined by any two of these values.

[0288] 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: EGGReise(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, if a fuel supplied to the multitude of 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 multitude of 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 one of the multiple 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 one of the multiple fuel spray nozzles is a fossil-based hydrocarbon 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 operating conditions as those under which EI Steig,NFK is calculated when a fuel supplied to the multitude of fuel spray nozzles includes a sustainable aviation fuel; and 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 operating conditions as those under which EI Steig,NFK is calculated when the fuel supplied to one of the multiple 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 can be less than 1.

[0289] The nvPM emission index ratio of cruise flight with rich combustion to MAB can be greater than zero.

[0290] The nvPM emission index ratio of cruise flight with rich combustion 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 0.5.

[0291] The nvPM emission index ratio of cruise flight with rich combustion to MAB can be less than or equal to 0.469 and can preferably be less than or equal to 0.43 and further preferably less than or equal to 0.391.

[0292] The nvPM emission index ratio of cruise flight with rich combustion to MAB can be greater than or equal to 0.191 and can preferably be greater than or equal to 0.214 and can further preferably be greater than or equal to 0.238.

[0293] The nvPM emission index ratio for cruise flight with rich combustion to MAB can be in the range of 0.191 to 0.469 and can preferably be in the range of 0.214 to 0.430 and can further preferably be in the range of 0.238 to 0.391.

[0294] 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.

[0295] The nvPM emission index ratio of cruise flight with rich combustion to MAB can be 0.19, 0.191, 0.2, 0.21, 0.22, 0.23, 0.24, 0.25, 0.26, 0.27, 0.28, 0.29, 0.3, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39, 0.4, 0.41, 0.42, 0.43, 0.44, 0.45, 0.46, 0.469 or 0.47 or lie in any range defined by any two of these values.

[0296] 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 multitude of 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 available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied to one of the multiple 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 multitude of 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 one of the multiple 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 multitude of 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 Leer,NFK is calculated when the fuel supplied to one of the multiple 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.

[0297] The nvPM emission index ratio of idle to cruise flight with rich combustion can be greater than zero.

[0298] The nvPM emission index ratio of idle to cruise flight with rich combustion can be less than or equal to 0.9 and can preferably be less than or equal to 0.8.

[0299] The nvPM emission index ratio of idle to cruise flight with rich combustion can be less than or equal to 0.746 and can preferably be less than or equal to 0.683 and can further preferably be less than or equal to 0.621.

[0300] The nvPM emission index ratio of idle to cruise flight with rich combustion can be greater than or equal to 0.304 and can preferably be greater than or equal to 0.342 and can more preferably be greater than or equal to 0.38.

[0301] The nvPM emission index ratio of idle to cruise flight with rich combustion can be in the range of 0.304 to 0.746 and can preferably be in the range of 0.342 to 0.683 and can further preferably be in the range of 0.380 to 0.621.

[0302] 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.

[0303] The nvPM emission index ratio of idle to cruise flight with rich combustion can be 0.3, 0.304, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39, 0.4, 0.41, 0.42, 0.43, 0.44, 0.45, 0.46, 0.47, 0.48, 0.49, 0.5, 0.51, 0.52, 0.53, 0.54, 0.55, 0.56, 0.57, 0.58, 0.59, 0.6, 0.61, 0.62, 0.63, 0.64, 0.65, 0.66, 0.67, 0.68, 0.69, 0.7, 0.71, 0.72, 0.73, 0.74, 0.746, 0.75 or lie within any range defined by any two of these values.

[0304] 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 comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; where: 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 given operating conditions, where one of the fuels supplied to the multiple fuel spray nozzles 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 operating conditions as those under which EI Leer,NFK is calculated when a fuel supplied to the multitude of 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 available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied by any of the multiple fuel spray nozzles is a fossil-based hydrocarbon fuel; and 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 operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied to one of the multiple 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 a sustainable aviation fuel (SAF) to the multitude of fuel spray nozzles.

[0305] The nvPM emission index ratio idle to MAB can be defined as above in connection with the thirty-fifth aspect.

[0306] 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 comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; where: 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, where the fuel supplied to the multiple 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 multitude of fuel spray nozzles includes a sustainable aviation fuel; EGGmaxAb,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 operating conditions as those under which EI maxAb,NFK is calculated when the fuel supplied by any of the multiple 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 one of the multiple 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 a sustainable aviation fuel (SAF) to the multitude of fuel spray nozzles.

[0307] The nvPM emission index ratio of lean-burn cruise to MAB can be defined as above in connection with the thirty-fifth aspect.

[0308] 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 comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; where: 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 given operating conditions, where one of the fuels supplied to the multiple 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 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 operating conditions as those under which EI Leer,NFK is calculated when a fuel supplied to the multitude of 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 multitude of 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 available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied to one of the multiple 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 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 operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied by any of the multiple 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 one of the multiple 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 a sustainable aviation fuel (SAF) to the multitude of fuel spray nozzles.

[0309] 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.

[0310] 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 comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; where: 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 the available thrust under given operating conditions, where one of the fuels supplied to the multiple 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 multitude of fuel spray nozzles includes a sustainable aviation fuel; EI 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 by any of the multiple 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 one of the multiple 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 a sustainable aviation fuel (SAF) to the multitude of fuel spray nozzles.

[0311] The nvPM emission index ratio for cruise flight with rich combustion can be defined as above in relation to the thirty-fifth aspect.

[0312] 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 comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; where: 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 multiple 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,NFKis calculated when a fuel supplied to the multitude of 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 one of the multiple 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 one of the multiple fuel spray nozzles is a fossil-based hydrocarbon fuel; EGGmaxAb,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 operating conditions as those under which EI Steig,NFK is calculated when a fuel supplied to the multitude of fuel spray nozzles includes a sustainable aviation fuel; and 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 operating conditions as those under which EI Steig,NFK is calculated when the fuel supplied to one of the multiple 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 a sustainable aviation fuel (SAF) to the multitude of fuel spray nozzles.

[0313] 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.

[0314] 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 comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; where: 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 multitude of 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 available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied to one of the multiple 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 multitude of 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 one of the multiple 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 multitude of 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 one of the multiple 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 a sustainable aviation fuel (SAF) to the multitude of fuel spray nozzles.

[0315] 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.

[0316] 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 plurality of fuel spray nozzles.

[0317] 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 comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; where: 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 multiple 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 multiple 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, where one of the fuels supplied to the multiple 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 one of the fuels supplied to the multiple 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 a sustainable aviation fuel (SAF) to the multitude of fuel spray nozzles.

[0318] The nvPM emission index ratio for lean-burn cruise flight can be defined as above in relation to the thirty-fifth aspect.

[0319] 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 multiple fuel spray nozzles 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 operating conditions as those under which EI Leer,NFK is calculated when a fuel supplied to the multitude of fuel spray nozzles includes a sustainable aviation fuel; EI 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 operating conditions as those under which EI Leer,NFKis calculated when the fuel supplied by any of the multiple fuel spray nozzles is a fossil-based hydrocarbon fuel; and 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 operating conditions as those under which EI Leer,NFK is calculated when one of the multiple fuel spray nozzles supplies is a fossil-based hydrocarbon fuel; and wherein The nvPM emission index ratio at idle to MAB of the gas turbine engine can be less than 1.

[0320] The nvPM emission index ratio idle to MAB can be defined as above in connection with the thirty-fifth aspect.

[0321] 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 multiple 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 maxAb,NFK is calculated when a fuel supplied to the multitude of 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, during operation at approximately 100% of available thrust under the same operating conditions as those under which EI maxAb,NFK is calculated when the fuel supplied by any of the multiple fuel spray nozzles is a fossil-based hydrocarbon 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 maxAb,NFK is calculated when one of the multiple fuel spray nozzles supplies 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 can be less than 1.

[0322] The nvPM emission index ratio of lean-burn cruise to MAB can be defined as above in connection with the thirty-fifth aspect.

[0323] 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, and if a fuel supplied to the multiple 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 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 operating conditions as those under which EI Leer,NFK is calculated when a fuel supplied to the multitude of 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 multitude of fuel spray nozzles includes a sustainable aviation fuel; EI 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 operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied to one of the multiple 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 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 operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied by any of the multiple 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 one of the multiple fuel spray nozzles supplies is a fossil-based hydrocarbon fuel; and wherein 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.

[0324] 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.

[0325] 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 the available thrust under the given operating conditions or under other operating conditions, and if a fuel supplied to the multiple 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 multitude of 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 by any of the multiple fuel spray nozzles is a fossil-based hydrocarbon 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 Steig,NFK is calculated when the fuel supplied by any of the multiple 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.

[0326] The nvPM emission index ratio for cruise flight with rich combustion can be defined as above in relation to the thirty-fifth aspect.

[0327] 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 multiple 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 multitude of 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 one of the multiple 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 one of the multiple fuel spray nozzles is a fossil-based hydrocarbon fuel; EGG maxAb,NFK is the nvPM emission index of the gas turbine engine, corrected for system losses, in mg / kg in an 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 multitude of fuel spray nozzles includes a sustainable aviation fuel; and 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 operating conditions as those under which EI Steig,NFK is calculated when the fuel supplied to one of the multiple 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 can be less than 1.

[0328] 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.

[0329] 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 multitude of 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 available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied to one of the multiple 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 multitude of 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 one of the multiple 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 multitude of 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 Leer,NFK is calculated when the fuel supplied to one of the multiple 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.

[0330] 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.

[0331] 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 comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; where: 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 given operating conditions, where one of the fuels supplied to the multiple fuel spray nozzles 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 operating conditions as those under which EI Leer,NFK is calculated when a fuel supplied to the multitude of 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 available thrust under the same operating conditions as those under which EI Leer,NFKis calculated when the fuel supplied by any of the multiple fuel spray nozzles is a fossil-based hydrocarbon fuel; and 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 operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied to one of the multiple 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 a sustainable aviation fuel (SAF) to the multitude of fuel spray nozzles.

[0332] The nvPM emission index ratio idle to MAB can be defined as above in connection with the thirty-fifth aspect.

[0333] 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 comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; where: 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, where the fuel supplied to the multiple 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 multitude of fuel spray nozzles includes a sustainable aviation fuel; EGGmaxAb,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 operating conditions as those under which EI maxAb,NFK is calculated when the fuel supplied by any of the multiple 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 one of the multiple 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 a sustainable aviation fuel (SAF) to the multitude of fuel spray nozzles.

[0334] The nvPM emission index ratio of lean-burn cruise to MAB can be defined as above in connection with the thirty-fifth aspect.

[0335] 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 comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; where: 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 given operating conditions, where one of the fuels supplied to the multiple 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 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 operating conditions as those under which EI Leer,NFK is calculated when a fuel supplied to the multitude of 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 multitude of 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 available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied to one of the multiple 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 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 operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied by any of the multiple 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 one of the multiple 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 a sustainable aviation fuel (SAF) to the multitude of fuel spray nozzles.

[0336] 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.

[0337] 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 comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; where: 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 the available thrust under given operating conditions, where one of the fuels supplied to the multiple 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 multitude of 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 by any of the multiple 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 one of the multiple 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 a sustainable aviation fuel (SAF) to the multitude of fuel spray nozzles.

[0338] The nvPM emission index ratio for cruise flight with rich combustion can be defined as above in relation to the thirty-fifth aspect.

[0339] 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 comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; 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, NFK2 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 multiple 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 multitude of 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 one of the multiple 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 one of the multiple fuel spray nozzles is a fossil-based hydrocarbon 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 operating conditions as those under which EI Steig,NFK is calculated when a fuel supplied to the multitude of fuel spray nozzles includes a sustainable aviation fuel; and 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 operating conditions as those under which EI Steig,NFK is calculated when the fuel supplied to one of the multiple 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 a sustainable aviation fuel (SAF) to the multitude of fuel spray nozzles.

[0340] 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.

[0341] 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 comprising a combustion chamber and a plurality of fuel spray nozzles configured to inject fuel into the combustion chamber, wherein the plurality of fuel spray nozzles comprises a first subset of fuel spray nozzles and a second subset of fuel spray nozzles, wherein the combustion chamber is operable in a state in which each of the fuel spray nozzles of the first subset of fuel spray nozzles is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset of fuel spray nozzles, wherein the ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles is in the range of 1:2 to 1:5; where: 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 multitude of 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 available thrust under the same operating conditions as those under which EI Leer,NFK is calculated when the fuel supplied to one of the multiple fuel spray nozzles is a fossil-based hydrocarbon fuel; EGG Reise(fett),NFKwhich 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 multitude of 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 one of the multiple 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 multitude of 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 one of the multiple 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 a sustainable aviation fuel (SAF) to the multitude of fuel spray nozzles.

[0342] 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.

[0343] 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 0.00893 to 4.72 mg / kg and preferably in the range of 0.0100 to 4.33 mg / kg and more preferably in the range of 0.0111 to 3.94 mg / kg. EGG maxAb can be in the range of 0.767 to 4.72 mg / kg and preferably in the range of 0.863 to 4.33 mg / kg and more preferably in the range of 0.959 to 3.94 mg / kg. EGG maxAbcan be in the range of 0.00893 to 0.0809 mg / kg and preferably in the range of 0.0100 to 0.0741 mg / kg and more preferably in the range of 0.0111 to 0.0674 mg / kg. EGG maxAb,NFK can be in the range of 0.00893 to 3.05 mg / kg and preferably in the range of 0.0100 to 2.80 mg / kg and more, preferably in the range of 0.0111 to 2.54 mg / kg. EGG maxAb,NFK can be in the range of 0.767 to 3.05 mg / kg and preferably in the range of 0.863 to 2.80 mg / kg and more preferably in the range of 0.959 to 2.54 mg / kg. EGG maxAb,NFK can be in the range of 0.00893 to 0.0523 mg / kg and preferably in the range of 0.0100 to 0.0479 mg / kg and more preferably in the range of 0.0111 to 0.0436 mg / kg. EGG maxAb,NFK can be in the range of 0.00893 to 4.71 mg / kg and preferably in the range of 0.0100 to 4.32 mg / kg and more, preferably in the range of 0.0111 to 3.93 mg / kg. EGG maxAb,NFKcan be in the range of 0.767 to 4.71 mg / kg and preferably in the range of 0.863 to 4.32 mg / kg and more preferably in the range of 0.959 to 3.93 mg / kg. EGG maxAb,NFK can be in the range of 0.00893 to 0.0808 mg / kg and preferably in the range of 0.0100 to 0.0740 mg / kg and more preferably in the range of 0.0111 to 0.0673 mg / kg. EGG maxAb,FK can be in the range of 0.0138 to 4.72 mg / kg and preferably in the range of 0.0155 to 4.33 mg / kg and more, preferably in the range of 0.0172 to 3.94 mg / kg. EGG maxAb,FK can be in the range of 1.18 to 4.72 mg / kg and preferably in the range of 1.33 to 4.33 mg / kg and more preferably in the range of 1.48 to 3.94 mg / kg. EGG maxAb,FK can be in the range of 0.0138 to 0.0809 mg / kg and preferably in the range of 0.0155 to 0.0741 mg / kg and more preferably in the range of 0.0172 to 0.0674 mg / kg. EGG maxAb,NFKcan be in the range of 0.00438 to 2.30 mg / kg and preferably in the range of 0.00493 to 2.11 mg / kg and more, preferably in the range of 0.00548 to 1.92 mg / kg. EGG maxAb,NFK can be in the range of 0.460 to 2.30 mg / kg and preferably in the range of 0.517 to 2.11 mg / kg and more preferably in the range of 0.575 to 1.92 mg / kg. EGG maxAb,NFK can be in the range of 0.00438 to 0.0221 mg / kg and preferably in the range of 0.00493 to 0.0202 mg / kg and more preferably in the range of 0.00548 to 0.0184 mg / kg. EGG Steig,NFK can be in the range of 0.00438 to 1.09 mg / kg and preferably in the range of 0.00493 to 0.999 mg / kg and more preferably in the range of 0.00548 to 0.909 mg / kg. EGG Steig,NFK can be in the range of 0.460 to 1.09 mg / kg and preferably in the range of 0.517 to 0.999 mg / kg and more preferably in the range of 0.575 to 0.909 mg / kg. EGG Steig,NFKcan be in the range of 0.00438 to 0.0105 mg / kg and preferably in the range of 0.00493 to 0.00959 mg / kg and more preferably in the range of 0.00548 to 0.00872 mg / kg. EGG Steig,NFK can be in the range of 0.00438 to 2.29 mg / kg and preferably in the range of 0.00493 to 2.10 mg / kg and more, preferably in the range of 0.00548 to 1.91 mg / kg. EGG Steig,NFK can be in the range of 0.460 to 2.29 mg / kg and preferably in the range of 0.517 to 2.10 mg / kg and more preferably in the range of 0.575 to 1.91 mg / kg. EGG Steig,NFK can be in the range of 0.00438 to 0.0220 mg / kg and preferably in the range of 0.00493 to 0.0201 mg / kg and more preferably in the range of 0.00548 to 0.0183 mg / kg. EGG Steig,FK can be in the range of 0.00923 to 2.30 mg / kg and preferably in the range of 0.0103 to 2.11 mg / kg and more, preferably in the range of 0.0115 to 1.92 mg / kg. EGG Steig,FKcan be in the range of 0.969 to 2.30 mg / kg and preferably in the range of 1.09 to 2.11 mg / kg and more preferably in the range of 1.21 to 1.92 mg / kg. EGG Steig,FK can be in the range of 0.00923 to 0.0221 mg / kg and preferably in the range of 0.0103 to 0.0202 mg / kg and more preferably in the range of 0.0115 to 0.0184 mg / kg. EGG Anflug can be in the range of 0.337 to 12.6 mg / kg and preferably in the range of 0.379 to 11.6 mg / kg and more, preferably in the range of 0.421 to 10.5 mg / kg. EGG Anflug can be in the range of 0.571 to 9.89 mg / kg and preferably in the range of 0.643 to 9.07 mg / kg and more preferably in the range of 0.714 to 8.25 mg / kg. EGG Anflug,NFK can be in the range of 0.337 to 1.94 mg / kg, and preferably in the range of 0.379 to 1.78 mg / kg, and more preferably in the range of 0.421 to 1.62 mg / kg. EI Anflug,NFKcan be in the range of 0.571 to 1.52 mg / kg and preferably in the range of 0.643 to 1.40 mg / kg and more preferably in the range of 0.714 to 1.27 mg / kg. EGG Anflug,NFK can be in the range of 0.337 to 12.5 mg / kg and preferably in the range of 0.379 to 11.5 mg / kg and more, preferably in the range of 0.421 to 10.4 mg / kg. EGG Anflug,NFK can be in the range of 0.571 to 9.88 mg / kg and preferably in the range of 0.643 to 9.06 mg / kg and more preferably in the range of 0.714 to 8.24 mg / kg. EGG Anflug,FK can be in the range of 2.19 to 12.6 mg / kg and preferably in the range of 2.47 to 11.6 mg / kg and more preferably in the range of 2.74 to 10.5 mg / kg. EGG Anflug,FK can be in the range of 3.72 to 9.89 mg / kg and preferably in the range of 4.18 to 9.07 mg / kg and more, preferably in the range of 4.65 to 8.25 mg / kg. EGG Leercan be in the range of 0.0525 to 1.55 mg / kg and preferably in the range of 0.0591 to 1.43 mg / kg and more preferably in the range of 0.0657 to 1.30 mg / kg. EGG Leer can be in the range of 0.0858 to 1.55 mg / kg and preferably in the range of 0.0966 to 1.43 mg / kg and more preferably in the range of 0.107 to 1.30 mg / kg. EGG Leer can be in the range of 0.0525 to 1.01 mg / kg and preferably in the range of 0.0591 to 0.925 mg / kg and more, preferably in the range of 0.0657 to 0.841 mg / kg. EGG Leer,NFK can be in the range of 0.0525 to 0.149 mg / kg and preferably in the range of 0.0591 to 0.137 mg / kg and more preferably in the range of 0.0657 to 0.124 mg / kg. EGG Leer,NFK can be in the range of 0.0858 to 0.149 mg / kg and preferably in the range of 0.0966 to 0.137 mg / kg and more preferably in the range of 0.107 to 0.124 mg / kg. EGG Leer,NFKcan be in the range of 0.0525 to 0.0967 mg / kg and preferably in the range of 0.0591 to 0.0886 mg / kg and more preferably in the range of 0.0657 to 0.0806 mg / kg. EGG Leer,NFK can be in the range of 0.0525 to 1.54 mg / kg, and preferably in the range of 0.0591 to 1.42 mg / kg, and more preferably in the range of 0.0657 to 1.29 mg / kg. EI Leer,NFK can be in the range of 0.0858 to 1.54 mg / kg and preferably in the range of 0.0966 to 1.42 mg / kg and more preferably in the range of 0.107 to 1.29 mg / kg. EGG Leer,NFK can be in the range of 0.0525 to 1.00 mg / kg and preferably in the range of 0.0591 to 0.924 mg / kg and more, preferably in the range of 0.0657 to 0.840 mg / kg. EGG Leer,FK can be in the range of 0.548 to 1.55 mg / kg and preferably in the range of 0.617 to 1.43 mg / kg and more preferably in the range of 0.686 to 1.30 mg / kg. EGG Leer,FKcan be in the range of 0.896 to 1.55 mg / kg and preferably in the range of 1.00 to 1.43 mg / kg and more, preferably in the range of 1.12 to 1.30 mg / kg. EGG Leer,FK can be in the range of 0.548 to 1.01 mg / kg and preferably in the range of 0.617 to 0.925 mg / kg and more preferably in the range of 0.686 to 0.841 mg / kg.

[0344] The following statements can apply to each of the aspects defined above, from one to forty-nine: The ratio of the number of fuel spray nozzles in the first subset of fuel spray nozzles to the number of fuel spray nozzles in the second subset of fuel spray nozzles can be in the range of 1:3 to 1:4 or preferably in the range of 1:3.5 to 1:4.

[0345] The first subset of fuel spray nozzles can include between 1 and 10 fuel spray nozzles. The first subset of fuel spray nozzles can include between 3 and 5 fuel spray nozzles.

[0346] The second subset of fuel spray nozzles can include between 10 and 25 fuel spray nozzles. The second subset of fuel spray nozzles can include between 13 and 20 fuel spray nozzles. The second subset of fuel spray nozzles can include between 13 and 17 fuel spray nozzles.

[0347] The combustion chamber can include one or more igniters.

[0348] Each of the first subset of fuel spray nozzles can be positioned closer to one or more of the igniters than the second subset. Alternatively, or in addition, one or more of the igniters can be positioned diametrically opposite one or more of the other igniters.

[0349] 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%.

[0350] 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.

[0351] 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.

[0352] 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).

[0353] 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.

[0354] 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).

[0355] 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.

[0356] 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, directly, or via a generally annular channel).

[0357] The gearbox can be arranged to be driven by the core shaft configured to rotate at the lowest speed (for example, in operation) – 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) – 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 more shafts, for example, the first and / or second shaft in the example above.

[0358] 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 compound gearbox, for example, a compound planetary gearbox (where the input is on the sun gear and the output is on the ring gear, and which is therefore also called a compound star gearbox), e.g., with two reduction stages.

[0359] 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. Alternatively, and also purely for illustrative purposes, 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.

[0360] 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), where a second compressor is provided. As another example, the flow at the outlet to the combustion chamber may be supplied to the inlet of the second turbine, where a second turbine is provided. The combustion chamber may be located upstream of the turbine(s).

[0361] The compressor or compressors (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 engine comprising 13 or 14 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 10 or 11 stages in the second (or "high-pressure") compressor. Alternatively, such an engine can, for example, have 4 stages in the first (or "low-pressure") compressor and 10 stages in the second (or "high-pressure") compressor.Alternatively, such an engine could, for example, comprise 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) comprising 11, 12, 13, 14, or 15 compressor stages (in addition to the fan). Such an engine could comprise 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 a further example, the gas turbine engine could 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.

[0362] 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. 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.As a purely illustrative example, the gas turbine engine could be a direct-drive gas turbine engine comprising a first (or "low-pressure") turbine with 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.

[0363] 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.

[0364] 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 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, 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 could lie within the range of 170 cm to 180 cm, 190 cm to 200 cm, 200 cm to 210 cm, 210 cm to 220 cm, 220 cm to 230 cm, 290 cm to 300 cm, or 340 cm to 360 cm.

[0365] 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 of 300 cm to 350 cm could be between 1800 and 2800 rpm or between 1950 and 2550 rpm. As a further non-restrictive example, the fan speed under cruise flight conditions for a direct-drive engine with a fan diameter of 270 cm to 290 cm could be between 1800 and 2800 rpm or between 2050 and 2450 rpm. As a further non-restrictive example, the fan speed under cruise flight conditions for a direct-drive engine with a fan diameter of 290 cm to 310 cm could be between 1800 and 2800 rpm or between 2100 and 2500 rpm.

[0366] 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 of 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).

[0367] Gas turbine engines according to the present disclosure can have a desired bypass ratio (BPV), wherein the bypass ratio can be 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 11, 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.

[0368] 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.

[0369] The specific thrust of an engine can be defined as the net thrust of the engine divided by the total mass flow through the engine. In some examples, the specific thrust for a given thrust condition may depend on the specific composition of fuel supplied to the combustion chamber. 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- 1 s, 95 Nkg- 1 s, 94 Nkg- 1 s, 93 Nkg- 1 s, 92 Nkg- 1 s, 91 Nkg- 1 s, 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 order 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 can be in the range of 70 Nkg-1s to 95 Nkg-. 1 s, from 80 Nkg-1s to 95 Nkg-1s or from 85 Nkg-1s to 95 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 170 cm to 200 cm can be in the range of 20 Nkg-1s to 90 Nkg- 1The specific thrust under cruise flight conditions of a direct-drive gas turbine engine with a fan diameter of 300 cm to 350 cm could be in the range of 90 Nkg to 120 Nkg or 100 Nkg to 115 Nkg. Similarly, the specific thrust under cruise flight conditions of a direct-drive gas turbine engine with a fan diameter of 260 cm to 285 cm could be in the range of 20 Nkg to 120 Nkg, 30 Nkg to 115 Nkg, or 40 Nkg to 115 Nkg. 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 from 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.

[0370] 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, in the range of 100 kN to 200 kN, from 110 kN to 180 kN or from 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, 275 kN to 400 kN, or 275 kN to 375 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 300 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 325 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.

[0371] 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.

[0372] 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.

[0373] 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 operation may 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 operation may be within an inclusive range bounded by any two of the TET values ​​in this paragraph (i.e., the values ​​may 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 a maximum takeoff condition (MAB condition).

[0374] 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 square inch. One unit of the core quantity is therefore expressed as: s⋅K12⋅in

[0375] 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.

[0376] 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 core size values ​​(i.e.,, the values ​​can form upper or lower limits), for example 4.5 to 9.5.

[0377] 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.

[0378] 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.

[0379] 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.

[0380] 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.

[0381] 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.

[0382] 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.

[0383] 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.

[0384] 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.

[0385] 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.

[0386] 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 midpoint of cruise flight during 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 fitted. In this context, 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 commencement 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 commencement 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 flight conditions, half of the total thrust required for stable operation, or at least substantially stable operation, of the aircraft during cruise flight.

[0387] 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 amount of 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 attached, at a given Mach number during cruise flight) under atmospheric conditions during cruise flight (defined by the International Standard Atmosphere according to ISO 2533 at a given altitude during cruise flight). For any given gas turbine engine for an aircraft, the cruise flight thrust, atmospheric conditions, and Mach number are known, and thus the engine's operating point under cruise flight conditions can be clearly defined.

[0388] 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.

[0389] 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.

[0390] 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 use and could, for example, be a value in the range of 20 kN to 40 kN.

[0391] 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.

[0392] 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, cruise flight conditions) of an aircraft on which at least one (for example, two or four) gas turbine engines can be mounted to provide thrust.

[0393] 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.

[0394] 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.

[0395] 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).

[0396] 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, during cruise flight of the aircraft, as defined elsewhere herein.

[0397] 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.

[0398] 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 that is 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. The hydrogen mass fraction as a percentage 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 bounds), for example, in the range of 13.41 to 15.3 or from 13.42 to 15.3.

[0399] 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, from 0.88 to 4.75, or from 0.89 to 4.75.

[0400] 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 can be 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.

[0401] 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.

[0402] 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.

[0403] 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 other 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 other parameters and / or values ​​contained or described herein (e.g., one or more of parameter B, parameter C, and parameter D, etc.) to express their relationship.The person skilled in the art would understand, for example, that if parameter A is disclosed separately from parameter B, their relationship can be expressed, for example, as follows: A+B, BA, AB, A / B, B / A, B*A or any further product, application, combination, function or, as required, any expression of parameter A relative to parameter B or vice versa.

[0404] 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 in order 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) can be applied to and / or combined with any or any further parameters or values ​​(for example, one or more of parameter B; parameter C; and parameter D, etc.) to express their relationship with respect to engine emissions and / or nvPM. For example, 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 any further product, application, combination, function, or expression of parameter C relative to parameter D or vice versa, as required. BRIEF DESCRIPTION OF THE DRAWINGS

[0405] 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 is a schematic representation of a fuel distribution system and the combustion chamber of a gas turbine engine; Fig. 7 is a cross-sectional view through the combustion chamber of a gas turbine engine along the main axis of rotation of the engine; Fig. 8 is another schematic representation of a fuel distribution system and the combustion chamber of a gas turbine engine; and Fig. Figure 9 shows a method for operating the gas turbine engine. DETAILED DESCRIPTION OF THE PROOF

[0406] 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.

[0407] 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.

[0408] 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.

[0409] 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.

[0410] The planetary gear set 30 is exemplified in Fig. Figure 3 shows 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.

[0411] 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.

[0412] 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. 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.

[0413] Accordingly, the present disclosure extends to a gas turbine engine with any arrangement of gear types (for example, star or planetary gears), support structures, input and output shaft arrangements, and bearing positions.

[0414] Optionally, the gearbox can drive additional and / or alternative components (e.g., the intermediate pressure compressor and / or a booster compressor).

[0415] 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.

[0416] 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.

[0417] How Fig. As can be seen from Figure 4, a gas turbine engine is generally specified at 10 and has 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.

[0418] 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 flows through a bypass duct 22 to provide thrust. Intermediate-pressure compressor 14 compresses airflow A before delivering the air to high-pressure compressor 15, where further compression takes place.

[0419] 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.

[0420] Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of 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.

[0421] 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.

[0422] The geometry of the gas turbine engine 10 and of its components is defined by a conventional axis system, which has an axial direction (the axial direction (aligned with the main axis of rotation 9) and a radial direction (in the direction from bottom to top). 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.

[0423] 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.

[0424] 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.

[0425] For example, experts understand NFK to mean a biofuel, a renewable aviation fuel, a renewable jet fuel, an alternative fuel, or a biojet fuel produced from biological or non-biological resources. Experts thus understand NFK to mean, for example, a fuel produced from sustainable and / or renewable resources. For instance, according to the general understanding, NFK is 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.

[0426] 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 exhaust nozzle outlet of a gas turbine engine that do not volatilize upon heating to a temperature of 350 °C.All references herein to soot or smoke may equally apply 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.

[0427] 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 missions 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.

[0428] 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.

[0429] 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.

[0430] 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.

[0431] 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 may 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 ​​may 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.

[0432] 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.

[0433] Fig. Figure 6 shows a schematic representation of a fuel distribution system 102 and the combustion chamber 16 of the gas turbine engine 10 of any example described herein. The combustion chamber 16 is configured to utilize staged lean combustion. Fuel is distributed to pilot fuel injectors and main fuel injectors by means of a fuel system control unit, which in the example shown is provided by a fuel metering unit (FMU) 104 under the control of an electronic engine control unit (EEC) 106. Fuel is supplied to the fuel metering unit 104 via a fuel pump 108. In the example shown, the fuel pump 108 is mechanically driven by an accessory gearbox (AGB) 110, although the fuel pump 108 can alternatively be electrically driven. The in Fig. The fuel pump 108 shown in Figure 6 can be one of several fuel pumps provided within the fuel distribution system 102. For example, the fuel pump 108 can be a high-pressure fuel pump provided at the gas turbine engine 10; optionally, instead of forming part of the gas turbine engine 10, one or more additional low-pressure fuel pumps may also be provided on board the aircraft.

[0434] The high-pressure fuel is supplied by the fuel metering unit 104 to one or more fuel distributors for distribution to pilot fuel injectors 116A and main fuel injectors 116B. The fuel supply via the pilot fuel injectors 116A and the main fuel injectors 116B is staged. Therefore, at low power outputs (and thus low air mass flow rates), fuel is supplied mainly or entirely through the pilot fuel injectors 116A at a rich fuel-air ratio (i.e., at an equivalence ratio greater than one) for improved flame stability. As power output and mass flow rate increase, a stage point is reached at which fuel is supplied through some or all of the main fuel injectors 116B, thus supplementing the fuel flow from the pilot fuel injectors 116A.The main fuel injectors 116B are configured to inject fuel at a lean fuel-air ratio (i.e., at an equivalence ratio of less than one). At this point, the airflow is such that the equivalence ratio immediately downstream of the pilot fuel injectors 116A is also fuel-lean. In the example shown, at higher power levels, fuel is injected through all main fuel injectors 116B.

[0435] The person skilled in the art will be familiar with such operation of staged combustion systems to achieve lean combustion at high power outputs while simultaneously observing flammability limits at lower power outputs.

[0436] The balance of fuel injection through the pilot fuel injectors 116A and the main fuel injectors 116B is controlled by the electronic engine control unit 106, which provides control signals to the fuel metering unit 104. These control signals can specify, directly or indirectly, the total fuel to be injected, for example, in the form of a fuel flow rate, and the ratio of the fuel flow from the pilot fuel injector to the fuel flow from the main fuel injector.

[0437] Fig. Figure 7 shows a section through the combustion chamber 16 in a plane perpendicular to the main axis of rotation 9 of the engine 10. The combustion chamber 16 comprises an annular combustion chamber 120, which is defined by a lining 122. Other combustion chamber configurations can be used alternatively, for example, tubular combustion chambers, cup-shaped combustion chambers, etc.

[0438] The combustion chamber 16 comprises a plurality of fuel spray nozzles 124 arranged around a circumference of the combustion chamber 16 and configured to inject fuel into the combustion chamber 120. In the example shown, the combustion chamber 16 comprises sixteen (16) fuel spray nozzles 124. Alternatively, the combustion chamber 16 may comprise any suitable number of fuel spray nozzles, for example, 8, 9, 10, 11, 12, 13, 14, 15, 17, 18, 19, 20, 21, 22, 23, 24, 25, 26, 27, 28, 29, 30 fuel spray nozzles, and so on. The combustion chamber may comprise a number of fuel spray nozzles within an enclosing region defined between any two values ​​in the preceding sentence, the two values ​​forming the upper and lower bounds of the region and enclosed within it.

[0439] For example, the combustion chamber can include between 14 and 27 fuel spray nozzles, or between 16 and 25 fuel spray nozzles, or between 18 and 23 fuel spray nozzles.

[0440] A key parameter of a gas turbine engine is defined as (with reference to the in Fig. (1 arrangement shown): Core size = m˙2⋅T3P3 where ṁ2 = the mass flow rate of air in pounds per second entering the high-pressure compressor 15, T3 = the temperature of air in Kelvin exiting the high-pressure compressor 15, and P3 = the pressure of air in pounds per second squared per square inch exiting the high-pressure compressor 15. One unit of the core quantity is therefore expressed as: s⋅K12⋅in

[0441] The core size (in sK) 1 / 2The engine core size (in sK) 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 sK) 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 yet other examples, the engine core size (in sK) 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.

[0442] A number of fuel spray nozzles: 124 per unit engine core size (in sK). 1 / 2The number (.in) can be between 2 and 6. For example, the number can be 2, 3, 4, 5, or 6, or any range defined between any two of these values. In some examples, the number can be between 3 and 4, for example, 3.0, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, or 4.0, or any range defined between any two of these values.

[0443] The number of fuel spray nozzles per unit engine core size can be between 2 and 7 or more, preferably between 2.1 and 6.5 or more, preferably between 2.4 and 3.4.

[0444] In further examples, the number of fuel spray nozzles per unit 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, 6.0, 6.1, 6.2, 6.3, 6.4, 6.5, 6.6, 6.7, 6.8, 6.9, 7.0 or within a range defined between any two of these values.

[0445] 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˙2=T2P2 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.

[0446] 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 engine 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.

[0447] In the example shown, each fuel spray nozzle 124 comprises a duplex fuel spray nozzle (also known as an internally staged nozzle) in which a pilot fuel injector 116A is integrated into the same fuel spray nozzle 124 as a main fuel injector 116B. However, it is intended that other types of staged combustion configurations can be used, for example, those with pilot fuel injectors and main fuel injectors in separate fuel spray nozzles, instead of both being contained in duplex or internally staged fuel spray nozzles. Indeed, it is understood that the principles disclosed herein can be applied to any staged combustion system that includes pilot fuel injectors and main fuel injectors.

[0448] How Fig. As further detailed in Section 6, the fuel distribution system 102 comprises a divider valve (SV) 112 configured to divide the fuel flow between the fuel spray nozzles 124 of the combustion chamber 16 such that each pilot injector 116A of a first subset 124A of the fuel spray nozzles 124 receives a higher fuel flow rate than each pilot injector 116A of a second subset 124B of the spray nozzles 124 below a stage point, for example, down to a threshold fuel flow rate or engine power. Below the stage point, the pilot injectors of the second subset 124B may receive no fuel or they may receive fuel at a lower fuel flow rate compared to those of the first subset 124A. Below the stage point, no fuel is supplied to the main injectors 116B.Above the step point, all pilot injectors 116A can be supplied with fuel at the same fuel flow rate. In the present example, close as shown in . Fig. Figure 7 shows the first and second subsets 124A and 124B of fuel spray nozzles, all of which are provided in the combustion chamber. The electronic engine control unit 106 is configured to control the divider valve 112, although the divider valve 112 may alternatively be mechanically controlled or have a fixed configuration.

[0449] In the Fig. In the example shown, the first subset 124A of the fuel spray nozzles 124 comprises 2 fuel spray nozzles 124 (shown hatched). Alternatively, the first subset 124A of the fuel spray noz...

Claims

[1] Gas turbine engine (10) for an aircraft, comprising: a combustion chamber (16) comprising a combustion chamber (120) and a plurality of fuel spray nozzles (124) configured to inject fuel into the combustion chamber (120), wherein the plurality of fuel spray nozzles (124) comprises a first subset (124A) of fuel spray nozzles (124) and a second subset (124B) of fuel spray nozzles (124), wherein the combustion chamber (16) is operable in a state in which each of the fuel spray nozzles of the first subset (124A) of the fuel spray nozzles (124) is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset (124B) of the fuel spray nozzles (124), wherein the number of fuel spray nozzles (124) in the first subset (124A) of the fuel spray nozzles (124) to the number of fuel spray nozzles (124) in the second subset (124B) the fuel spray nozzles (124) in the range of 1:2 to 1:5; and where: a fuel flow-nvPM emission index ratio is defined as follows: EILempty×WK,EmptyEImaxAb×WK,maxAb 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; 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 (10) is less than 6; and the gas turbine engine (10) is configured to provide a sustainable aviation fuel (SAF) to the multiple fuel spray nozzles (124). [2] Gas turbine engine (10) according to claim 1, wherein the fuel flow-nvPM emission index ratio is less than 5.93 and preferably less than 5.44 and more preferably less than 4.

94. [3] Gas turbine engine (10) according to claim 1 or claim 2, wherein the fuel flow-nvPM emission index ratio is less than or equal to 4.5 and preferably less than or equal to 3 and more preferably less than or equal to 1.

5. [4] Gas turbine engine (10) according to one of the preceding claims, wherein the fuel flow-nvPM emission index ratio is less than or equal to 0.879 and preferably less than or equal to 0.806 and more preferably less than or equal to 0.

733. [5] Gas turbine engine (10) according to one of the preceding claims, wherein the fuel flow-nvPM emission index ratio is less than or equal to 0.0181 and preferably less than or equal to 0.0166 and more preferably less than or equal to 0.0151. [6] Gas turbine engine (10) according to one of the preceding claims, wherein the fuel flow-nvPM emission index ratio is greater than or equal to 0.00351 and preferably greater than or equal to 0.00395 and more preferably greater than or equal to 0.00439. [7] Gas turbine engine (10) according to one of the preceding claims, wherein the fuel flow-nvPM emission index ratio is greater than or equal to 0.109 and preferably greater than or equal to 0.123 and more preferably greater than or equal to 0.

137. [8] Gas turbine engine (10) according to one of the preceding claims, wherein the fuel flow-nvPM emission index ratio is in the range of 0.00351 to 0.879 and preferably in the range of 0.00395 to 0.806 and more preferably in the range of 0.00439 to 0.

733. [9] Gas turbine engine (10) according to any one of the preceding claims, wherein: a) W K,Leer in the range of 0.0695 to 0.118 kg / s, and preferably in the range of 0.0782 to 0.108 kg / s, and more preferably in the range of 0.0869 to 0.0981 kg / s; and / or b) W K,maxAb in the range of 0.595 to 1.29 kg / s and preferably in the range of 0.670 to 1.19 kg / s and more preferably in the range of 0.744 to 1.08 kg / s. [10] Gas turbine engine (10) according to one of the preceding claims, wherein the ratio of the number of fuel spray nozzles (124) in the first subset (124A) of the fuel spray nozzles (124) to the number of fuel spray nozzles (124) in the second subset (124B) of the fuel spray nozzles (124) is in the range of 1:3 to 1:4 or preferably in the range of 1:3.5 to 1:

4. [11] Gas turbine engine (10) according to one of the preceding claims, wherein the combustion chamber (16) comprises one or more igniters (126). [12] Gas turbine engine (10) according to claim 11, wherein of the first subset (124A) of the fuel spray nozzles (124) each is arranged closer to a respective one or more of the igniters (126) than the second subset (124B) and / or wherein one or more of the igniters (126) are arranged diametrically opposite one or more other igniters (126). [13] Gas turbine engine (10) according to one of the preceding claims, wherein the fuel supplied to the combustion chamber (16) 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%. [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 plurality of fuel spray nozzles (124). [15] Method (1000) for operating a gas turbine engine (10), comprising the gas turbine engine (10): a combustion chamber (16) comprising a combustion chamber (120) and a plurality of fuel spray nozzles (124) configured to inject fuel into the combustion chamber (120), wherein the plurality of fuel spray nozzles (124) comprises a first subset (124A) of fuel spray nozzles (124) and a second subset (124B) of fuel spray nozzles (124), wherein the combustion chamber (16) is operable in a state in which each of the fuel spray nozzles of the first subset (124A) of the fuel spray nozzles (124) is supplied with a greater fuel flow rate than each of the fuel spray nozzles of the second subset (124B) of the fuel spray nozzles (124), wherein the number of fuel spray nozzles (124) in the first subset (124A) of the fuel spray nozzles (124) to the number of fuel spray nozzles (124) in the second subset (124B) the fuel spray nozzles (124) in the range of 1:2 to 1:5; and where: a fuel flow-nvPM emission index ratio is defined as follows: EILempty×WK,EmptyEImaxAb×WK,maxAb 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; 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 conditions. Operating conditions; and the fuel flow-nvPM emission index ratio of the gas turbine engine (10) is less than 6; and wherein the method comprises the provision (1002) of fuel comprising a sustainable aviation fuel to the plurality of fuel spray nozzles (124).

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