Gas turbine engine for an aircraft

Abstract

This application relates to a gas turbine engine (10) for an aircraft (1), comprising: a staged combustion system (64) having pilot fuel injectors (313) and main fuel injectors (314), the staged combustion system (64) being operable in a pilot-only injection operating range in which fuel is distributed only to the pilot fuel injectors (313), and a pilot and main injection operating range in which fuel is distributed at least to the main fuel injectors (314).The gas turbine further includes a fuel distribution regulator (306) designed to control fuel distribution to the pilot and main fuel injectors (313, 314), the fuel distribution regulator (306) being designed to receive fuel from a first fuel source (302) containing a first fuel having a first fuel characteristic and from a second fuel source (304) containing a second fuel having a second fuel characteristic, the second fuel characteristic being different from the first.The staged combustion system (64) is designed to switch between pilot-only injection and pilot- and main-injection operating ranges at a switching point that corresponds to a steady-state engine cruising mode. This switching point defines a boundary between a first engine cruising range (320a) and a second engine cruising range (320b). The fuel distribution regulator (306) is designed to deliver fuel to the pilot fuel injectors (313) during at least a portion of the first engine cruising range (320a) with a fuel characteristic different from that delivered to the pilot and / or main fuel injectors (313, 314) during the second engine cruising range (320b).The invention also describes a method (4014) for operating a gas turbine engine (10). Figure for the abstract: Fig. 5.

Description

Description Title of the invention: Gas turbine operation The present invention relates to a gas turbine engine for an aircraft, and a method of operating a gas turbine engine for an aircraft. The present disclosure further relates to a computer-implemented method for determining one or more fuel loading parameters for an aircraft, a system for determining fuel loading parameters, a method for determining a fuel allocation to the fleet-wide for a plurality of missions, and a fleet-wide fuel allocation determination system. The present invention further relates to a method of loading fuel onto the aircraft, and onto a plurality of aircraft performing the plurality of missions. In the aviation sector, a trend is expected towards the use of fuels different from the traditional kerosene-based jet fuels generally used today. According to a first aspect, a gas turbine engine for an aircraft is provided, comprising: a staged combustion system having pilot fuel injectors and main fuel injectors, the staged combustion system being operable in a pilot injection only operating range and a pilot and main injection operating range; And a fuel delivery regulator configured to control fuel delivery to the pilot and main fuel injectors, the fuel delivery regulator configured to receive fuel from a first fuel source containing a first fuel having a first characteristic fuel and a second fuel source containing a second fuel having a second fuel characteristic, the second fuel characteristic being different from the first, wherein the fuel delivery regulator is configured to distribute fuel to the pilot fuel injectors during at least a portion of the pilot injection only operating range having a fuel characteristic different from the fuel delivered to the pilot fuel injectors and / or main during at least part of the pilot and main injection operating range. The inventors have determined that it may be advantageous to supply fuel from different fuel sources having different characteristics to the fuel delivery regulator, and to distribute fuel to the pilot and main fuel injectors such that fuel of a characteristic different is provided in different ranges of combustion chamber operation. This more flexible fuel supply to the combustion chamber may allow fuel that exhibits favorable combustion properties, but is feed limited, to be used when those favorable properties will have a greater effect. The first fuel characteristic may be associated with a lower nvPM production level than the second fuel characteristic. Fuel delivered to the pilot fuel injectors for at least a portion of the pilot-only operating range may be associated with a lower nvPM production level than fuel delivered to the pilot and / or main fuel injectors for at least part of the pilot and main injection operating range. The fuel delivery regulator may be configured to distribute fuel from the first fuel source to the pilot fuel injectors during operation in both the pilot injection only and pilot and main injection operating ranges, and the fuel from the second fuel source to the main fuel injectors during operation in the pilot and main injection operating range. The fuel delivery regulator may be configured to distribute fuel from the first fuel source to the pilot fuel injectors during operation in at least a portion of the pilot injection only operating range, and distribute fuel from from the second fuel source to the pilot fuel injectors during the pilot and main injection operating range. The fuel delivery regulator may be configured to switch fuel delivery to the pilot fuel injectors between fuel from the first fuel source and fuel from the second fuel source at one or more operating points at within the pilot injection only operating range or at a limit thereof. The fuel delivery regulator may be configured to switch fuel delivery to the pilot fuel injectors between fuel from the first fuel source and fuel from the second fuel source in accordance with a mode signal indicating a change. of the operating range of the staged combustion system. Switching can occur at the staging point between the pilot injection only operating range and the pilot and main injection operating range. The fuel delivery regulator may be configured to switch fuel delivery back to the pilot fuel injectors between the fuel from the first fuel source and the fuel from the second fuel source at a threshold point within the pilot injection only operating range. Fuel from the second source may be delivered to the pilot fuel injectors at fuel flow rates below the threshold point, and fuel from the first source is delivered to the pilot fuel injectors at fuel flow rates between the threshold and the limit of the operating range for pilot injection only. The threshold point may be a threshold fuel flow rate at which the production of nvPM by the gas turbine engine passes a threshold quantity of the nvPM produced by the gas turbine engine during operation in which the pilot fuel injectors receive fuel having the second fuel characteristic Additionally or alternatively, the threshold point may be a predefined threshold fuel flow rate that is less than the fuel flow rate at the boundary between pilot injection only operation and pilot and main injection operation by a predefined amount. The predefined threshold fuel flow can be either a percentage of the fuel flow at the staging point, or an absolute value of fuel flow lower than that at the staging point. The fuel delivery regulator may include a fuel mixer configured to receive a supply of fuel from both the first and second fuel sources and output fuel from the first fuel source, fuel from the second fuel source , or a mixture of these. The fuel mixer may be configured to deliver fuel to the pilot fuel injectors. The fuel mixer may be configured to distribute a mixture of fuel from the first fuel source and fuel from the second fuel source to the pilot fuel injectors during at least a portion of the pilot injection operating range alone. The fuel mixer may be configured to deliver a fuel mixture to pilot fuel injectors having a variable mixing ratio of fuel from the first fuel source and fuel from the second fuel source, wherein the mixing ratio may be modified in the pilot injection only operating range depending on fuel flow. The proportion of fuel from the first fuel source to that of the second fuel source can be reduced with the reduction of a fuel flow rate within the pilot injection only operating range. The dependence of the proportion of fuel coming from the first fuel source compared to that coming from the second fuel source on the fuel flow rate is determined according to a desired obtained level of nyPM at a particular fuel flow rate, and can be determined such that the nyPM does not exceed a predetermined threshold or such that the production of nvPM is minimized over a period of operation of the gas turbine engine. The production of nvPM can be minimized on the LTO cycle. The fuel mixer may be configured to mix fuel to achieve a desired level of nvPM under one or more operating conditions of the gas turbine engine within the pilot injection only operating range. The fuel regulator may be configured to distribute fuel to the pilot fuel injectors to minimize a cost function dependent on one or more nvPM impact parameters. The nvPM impact parameter(s) may include any one or more of: 1) the height above ground level at which nvPM production takes place; ii) the location (e.g. longitude and latitude) of nVPM production: iii) weather and / or atmospheric conditions at an nvPM production location; iv) climate impacts associated with the location of nvPM production; v) the mass and / or size of individual nvPM particles produced; vi) potential contrail generation and / or contrail characteristics; vil) the local air quality (LAQ) impact of nvPM production; and or vili) the mass and / or number of nvPM. The mixing ratio provided by the fuel mixer can be determined at least partially according to one or more elements among: a) the quantity of fuel having the first fuel composition available for a proposed flight (e.g., estimating or measuring the quantity of fuel in the fuel tanks comprising the first and second fuel sources); b) the total fuel requirement quantity for the pilot fuel injectors during pilot-only operation for the entire flight within an operating range in which fuel is supplied from the first fuel source; and or c) a fuel composition limit parameter (e.g., certification limit, fuel composition available for refueling, aircraft / engine limits). According to a second aspect, the invention relates to a method of operating a gas turbine engine for an aircraft, the gas turbine engine comprising a staged combustion system having pilot fuel injectors and in- Main fuel jets, the staged combustion system capable of operating in a pilot injection only operating range and a pilot and main injection operating range, the method comprising: regulating fuel delivery to the pilot and main fuel injectors from a first fuel source containing a first fuel having a first fuel characteristic and a second fuel source containing a second fuel having a second fuel characteristic of fuel, the second fuel characteristic being different from the first, wherein regulating fuel delivery comprises delivering fuel to the pilot fuel injectors during at least a portion of the pilot injection only operating range having a fuel characteristic different from the fuel delivered to the pilot fuel injectors and / or main during at least part of the pilot and main injection operating range. The first fuel characteristic may be associated with a lower nvPM production level than the second fuel characteristic. Fuel delivered to the pilot fuel injectors for at least a portion of the pilot-only operating range may be associated with a lower nvPM production level than fuel delivered to the pilot and / or main fuel injectors for at least part of the pilot and main injection operating range. Controlling fuel delivery may include distributing fuel from the first fuel source to the pilot fuel injectors during operation in both the pilot injection only and pilot and main injection operating ranges, and fuel from the second fuel source to the main fuel injectors during operation in the pilot and main injection operating range. Controlling fuel delivery may include distributing fuel from the first fuel source to the pilot fuel injectors during operation in at least a portion of the pilot injection only operating range, and distributing fuel to from the second fuel source to the pilot fuel injectors during the pilot and main injection operating range. Controlling fuel delivery may include switching fuel delivery to the pilot fuel injectors between fuel from the first fuel source and fuel from the second fuel source at one or more operating points within of the operating range with pilot injection only or at a limit thereof. Controlling fuel delivery may include switching fuel delivery to the pilot fuel injectors between fuel from the first fuel source and fuel from the second fuel source according to a mode signal indicative of a change in fuel delivery. the operating range of the staged combustion system. Controlling fuel delivery may further include delivering fuel to the pilot fuel injectors between fuel from the first fuel source and fuel from the second fuel source at a threshold point within the fuel delivery range. pilot injection only operation. The threshold point may be a threshold fuel flow rate at which the production of nvPM by the gas turbine engine passes a threshold quantity of the nvPM produced by the gas turbine engine during operation in which the pilot fuel injectors receive fuel having the second fuel characteristic. Additionally or alternatively, the threshold point may be a predefined threshold fuel flow rate that is less than the fuel flow rate at the boundary between pilot injection only operation and pilot and main injection operation by a predefined amount. Fuel supply regulation may include: mixing a fuel supply from both the first and second fuel sources to form a blended fuel consisting of fuel from the first fuel source, fuel from the second fuel source, or a mixture thereof ; And delivering the mixed fuel to the pilot fuel injectors. Dispensing the mixed fuel may include delivering a mixture of fuel from the first fuel source and fuel from the second fuel source to the pilot fuel injectors during at least a portion of the pilot injection operating range. alone. Dispensing the mixed fuel may include delivering a fuel mixture to the pilot fuel injectors having a variable mixture ratio of fuel from the first fuel source and fuel from the second fuel source, wherein the ratio mixture rate is varied in the pilot injection only operating range depending on the fuel flow. The proportion of fuel from the first fuel source to that of the second fuel source can be reduced with the reduction of a fuel flow rate within the pilot injection only operating range. The dependence of the proportion of fuel coming from the first fuel source compared to that coming from the second fuel source by ratio to fuel flow may be determined according to a desired achieved level of nvPM at a particular fuel flow, and may be determined such that the nvPM does not exceed a predetermined threshold or such that the production of nvPM is reduced to the minimum over a period of operation of the gas turbine engine. The blending of the fuel may include blending the fuel to achieve a desired level of nvPM under one or more operating conditions of the gas turbine engine within the pilot injection only operating range. Controlling fuel delivery may include delivering fuel to pilot fuel injectors to minimize a cost function dependent on one or more nvPM impact parameters. The nvPM impact parameter(s) may include one or more elements from: 1) the height above ground level at which nvPM production takes place; 11) location of nvPM production; iii) weather and / or atmospheric conditions at the location of nvPM production; iv) climate impacts associated with the location of nvPM production; v) the mass and / or size of individual nvPM particles produced; vi) potential contrail generation and / or contrail characteristics; vil) the local air quality (LAQ) impact of nvPM production; and or vili) the mass and / or number of nvPM. The mixture of the fuel may comprise the mixture of the fuel at a mixing ratio determined at least partially according to one or more elements among: a) the quantity of fuel having the first fuel composition available for a proposed flight; b) the total fuel requirement quantity for the pilot fuel injectors during pilot-only operation for the entire flight within an operating range in which fuel is supplied from the first fuel source; and or c) a fuel composition limit parameter. According to another aspect, the invention relates to an aircraft comprising one or more gas turbine engines according to the first aspect, and optionally one or more of the associated statements above. According to a third aspect, the invention relates to a gas turbine engine for an aircraft, comprising: a staged combustion system having pilot fuel injectors and main fuel injectors, the staged combustion system capable of operating in a pilot injection only operating range in which fuel is delivered only to the pilot fuel injectors, and a pilot and main injection operating range in which fuel is delivered to at least the main fuel injectors; And a fuel delivery regulator configured to control fuel delivery to the pilot and main fuel injectors, the fuel delivery regulator configured to receive fuel from a first fuel source containing a first fuel having a first characteristic fuel and a second fuel source containing a second fuel having a second fuel characteristic, the second fuel characteristic being different from the first, wherein: the staged combustion system is designed to switch between the pilot injection only operating range and the pilot and main injection operating range at a staging point which corresponds to a cruise mode of operation at stabilized state of the engine, the staging point defining a limit between a first operating range at engine cruise speed and a second operating range at engine cruise speed; And the fuel delivery regulator is configured to distribute fuel to the pilot fuel injectors during at least a portion of the first engine cruise operating range having a fuel characteristic different from the fuel distributed to the pilot fuel injectors and / or or main during the second operating range at engine cruise speed. The inventors have determined that it may be advantageous to operate a staged combustion system such that it is in pilot injection only mode during at least a portion of its cruise operation, while selectively providing fuel from two different sources to the combustion chamber during cruise operation. The inventors have determined that by adjusting the staging point so that lower cruise operation can take place in the pilot injection only mode, certain engine emissions can be reduced and efficiency improved combustion. In combination with the selective use of fuels with different characteristics, the inventors have determined that disadvantageous effects on emissions that would otherwise result in a shift in the staging point can be mitigated. This therefore gives an overall improvement in combustion efficiency and a reduction in emissions through a combination of these factors. The first fuel characteristic may be associated with a lower nvPM production level than the second fuel characteristic. The fuel delivered to the pilot fuel injectors during the first range of engine cruise operation may be associated with an nvPM production level that is lower than that of the fuel delivered to the pilot and / or main fuel injectors during at least part of the second operating range at engine cruise speed. The first fuel characteristic may correspond to a greater proportion of SAF within the respective fuel compared to the second fuel characteristic, and the fuel dispensed during the first operating range at engine cruise speed may have a higher proportion of SAF compared to the fuel distributed during the second operating range at engine cruise speed. The first engine cruise operating range may correspond to operation of the aircraft in a later part of a cruise segment of a flight, and the second engine cruising operating range may correspond to operation of the aircraft in a relatively earlier part of the cruise segment. The first operating range at engine cruising speed may correspond to operation at subsonic cruising speed in the stabilized state of the engine and the second operating range at engine cruising speed may correspond to operation at supersonic cruise in the stabilized state of the engine. The fuel delivery regulator may include a fuel mixer configured to receive a supply of fuel from both the first and second fuel sources and output fuel from the first fuel source, fuel from the second fuel source , or a mixture of these. The fuel mixer may be configured to deliver fuel to the pilot fuel injectors, and optionally to the main fuel injectors. The proportion of fuel delivered from the first fuel source compared to that from the second fuel source may be determined according to a desired achieved level of nvPM production at a particular fuel flow rate within the first range operating at engine cruise speed, and can be determined such that the production of nvPM does not exceed a predetermined threshold or such that the production of nvPM is reduced to a minimum over a period of operation of the turbine engine at gas. The proportion of fuel dispensed from the first fuel source compared to that from the second fuel source during the first operating range at engine cruise speed can be determined at least partially according to one or more elements among: a) the quantity of fuel having the first fuel characteristic and the second fuel characteristic available for a proposed flight; b) the total fuel requirement quantity for the pilot fuel injectors during pilot-only operation for the entire flight within an operating range in which fuel is supplied from the first fuel source: and / Or c) a fuel composition limit parameter. According to a fourth aspect, the invention relates to a method of operating a gas turbine engine for an aircraft, the gas turbine engine comprising a staged combustion system having pilot fuel injectors and fuel injectors main, the staged combustion system being operable in a pilot injection only operating range in which the fuel is distributed only to the pilot fuel injectors, and a pilot and main injection operating range in which the fuel is distributed to the less to the main fuel injectors, the method comprising: regulating fuel delivery to the pilot and main fuel injectors from a first fuel source containing a first fuel having a first fuel characteristic and a second fuel source containing a second fuel having a second fuel characteristic fuel, the second fuel characteristic being different from the first; switching between the pilot injection only operating range and the pilot and main injection operating range at a staging point during a steady state cruise mode of operation of the engine to define a first operating range in engine cruise speed and a second operating range in engine cruise speed; And delivering fuel to the pilot fuel injectors during at least a portion of the first operating range at engine cruise speed having a fuel characteristic different from the fuel delivered to the pilot and / or main fuel injectors during the second operating range at engine cruising speed. The first fuel characteristic may be associated with an nvPM production level that is lower than that of the second fuel characteristic. The fuel delivered to the pilot fuel injectors during the first range of engine cruise operation may be associated with an nvPM production level that is lower than that of the fuel delivered to the pilot injectors. pilot and / or main fuel during at least part of the second operating range at engine cruise speed. The first fuel characteristic may correspond to a higher proportion of SAF within the respective fuel compared to the second fuel characteristic. The fuel distributed during the first operating range at engine cruise speed may have a higher proportion of SAF compared to the fuel distributed during the second operating range at engine cruise speed. The first engine cruise operating range may correspond to operation of the aircraft in a later part of a cruise segment of a flight, and the second engine cruising operating range may correspond to operation of the aircraft in a relatively earlier part of the cruise segment. The first operating range at engine cruising speed may correspond to operation at subsonic cruising speed in the stabilized state of the engine and the second operating range at engine cruising speed may correspond to operation at supersonic cruise in the stabilized state of the engine. Controlling fuel delivery may include dispensing fuel from the first fuel source, fuel from the second fuel source, or a mixture thereof using a mixer. fuel. The fuel mixer may be configured to deliver fuel to the pilot fuel injectors, and optionally to the main fuel injectors. The proportion of fuel delivered from the first fuel source compared to that from the second fuel source may be determined according to a desired achieved level of nvPM production at a particular fuel flow rate within the first range operating at engine cruise speed, and can be determined such that the nvPM does not exceed a predetermined threshold or such that the production of nvPM is reduced to a minimum over a period of operation of the gas turbine engine. The proportion of fuel distributed from the first fuel source compared to that from the second fuel source during the first operating range at engine cruise speed can be determined at least partially according to one or more elements among: a) the quantity of fuel having the first fuel characteristic and the second fuel characteristic available for a proposed flight; b) the total fuel requirement quantity for the pilot fuel injectors during pilot injection only operation for the entire flight within an operating range in which fuel is supplied from the first fuel source: and / or c) a fuel composition limit parameter. According to another aspect, the invention relates to an aircraft comprising one or more gas turbine engines according to the third aspect, and optionally one or more of the associated statements above. According to a fifth aspect, the invention relates to a gas turbine engine for an aircraft, comprising: a staged combustion system having pilot fuel injectors and main fuel injectors, the staged combustion system capable of operating in a pilot injection only operating range in which fuel is delivered only to the pilot fuel injectors, and a pilot and main injection operating range in which fuel is delivered to at least the main fuel injectors at a relative speed defined by a pilot and main injection staging ratio; And a fuel delivery regulator configured to control fuel delivery to the pilot and main fuel injectors, the fuel delivery regulator configured to receive fuel from a first fuel source containing a first fuel having a first characteristic fuel and a second fuel source containing a second fuel having a second fuel characteristic, the second fuel characteristic being different from the first, wherein: the staged combustion system can also be operated in a transition operating range between the operating ranges with pilot injection only and with pilot and main injection; in the transition operating range fuel is delivered to the pilot and main fuel injectors at a transition staging ratio that is different from the pilot and main injection staging ratio; And the fuel delivery regulator is configured to distribute fuel to the pilot and / or main fuel injectors during the transition operating range having a fuel characteristic different from the fuel delivered to the pilot and / or main fuel injectors for at least one part of the pilot and main injection operating range. The inventors have determined that a transition region that can be included allows a transition between the pilot-only injection staging ratio (100:0) and the selected pilot and main injection staging ratio. The inventors have determined that by configuring the staged combustion system to operate within a range of transition between pilot injection operation and pilot and main injection operation, the amount of CO and HC emissions implemented in this range of engine power settings can be reduced. The inventors have also determined that any detrimental change in engine emissions obtained from the transition region can be mitigated at least in part by using fuel having a different fuel characteristic in the transition operating range compared to that used in the transition region. at least one pilot and main injection operation. The first fuel characteristic may be associated with a lower nvPM production level than the second fuel characteristic. Fuel delivered to at least the pilot fuel injectors during the transition operating range may be associated with a lower nvPM production level than fuel delivered to the pilot and / or main fuel injectors during at least a portion of the pilot and main injection operating range. The first fuel characteristic may correspond to a higher proportion of SAF within the respective fuel compared to the second fuel characteristic. Fuel delivered during the transition operating range may have a higher SAF proportion compared to fuel delivered during at least a portion of the pilot and main injection operating range. The transition staging ratio may vary continuously with changing motor power within at least a portion of the transition operating range. The continuous variation may be such that, within the transition operating range, the proportion of the total fuel flow to the fuel injectors that can be attributed to fuel flow to the pilot fuel injectors decreases with increasing of engine power and the proportion of the total fuel flow to the fuel injectors that can be attributed to fuel flow to the main fuel injectors increases with increasing engine power. The transition staging ratio may have a constant intermediate value within at least a portion of the transition operating range that is different from the pilot and main injection staging ratio. The constant intermediate value can be between that of the pilot injection range alone and that of the pilot and main injection range. The transition staging ratio can vary between a series of constant intermediate values ​​(i.e. different from each other), each being different from the pilot and main injection staging ratio. Each constant intermediate value can be included between that of the pilot injection range alone and that of the pilot and main injection range. The fuel delivery regulator may include a fuel mixer configured to receive a supply of fuel from both the first and second fuel sources and output fuel from the first fuel source, fuel from the second fuel source , or a mixture of these. The fuel mixer may be configured to deliver fuel to the pilot fuel injectors, and optionally to the main fuel injectors. The proportion of fuel delivered from the first fuel source compared to that from the second fuel source may be determined according to a desired achieved level of nvPM production at a particular fuel flow rate within the range. transition operation, and can be determined such that the production of nvPM does not exceed a predetermined threshold or such that the production of nvPM is reduced to a minimum over a period of operation of the gas turbine engine. The proportion of fuel distributed from the first fuel source compared to that from the second fuel source during the transition operating range may be determined at least partially according to any one or more of: a) the quantity of fuel having the first fuel characteristic and the second fuel characteristic available for a proposed flight; b) the total fuel requirement quantity for the fuel injectors during pilot injection only operation for the entire flight within an operating range in which fuel is supplied from the first fuel source; and or c) a fuel composition limit parameter. According to a sixth aspect, the invention relates to a method of operating a gas turbine engine for an aircraft, the gas turbine engine comprising a staged combustion system having pilot fuel injectors and fuel injectors main, the staged combustion system being operable in a pilot injection only operating range in which the fuel is distributed only to the pilot fuel injectors, and a pilot and main injection operating range in which the fuel is distributed to the less to the main fuel injectors at a relative speed defined by a pilot and main injection staging ratio, the method comprising: regulating fuel delivery to the pilot and main fuel injectors from a first fuel source containing a first fuel having a first fuel characteristic and a second fuel source containing a second fuel having a second fuel characteristic, the second fuel characteristic being different from the first, operation of the staged combustion system in a transitional operating range between pilot injection only and pilot and main injection operating ranges in which fuel is supplied to the pilot and main fuel injectors at a staging ratio transition which is different from the pilot and main injection staging ratio, wherein regulating fuel delivery includes delivering fuel to the pilot and / or main fuel injectors during the transition operating range having a fuel characteristic different from the fuel delivered to the pilot and / or main fuel injectors for at least part of the pilot and main injection operating range. The first fuel characteristic may be associated with a lower nvPM production level than the second fuel characteristic. Fuel delivered to at least the pilot fuel injectors during the transition operating range may be associated with a lower nvPM production level than fuel delivered to the pilot and / or main fuel injectors during at least a portion of the pilot and main injection operating range. The first fuel characteristic may correspond to a higher proportion of SAF within the respective fuel compared to the second fuel characteristic, and the fuel dispensed during the transition operating range may have a higher proportion of SAF compared to the second fuel characteristic. to fuel delivered during at least part of the pilot and main injection operating range. The transition staging ratio may vary continuously with changing motor power within at least a portion of the transition operating range. The continuous variation may be such that, within the transition operating range, the proportion of the total fuel flow to the fuel injectors that can be attributed to fuel flow to the pilot fuel injectors decreases with increasing of engine power and the proportion of the total fuel flow to the fuel injectors that can be attributed to fuel flow to the main fuel injectors increases with increasing engine power. The transition staging ratio may have a constant intermediate value within at least a portion of the transition operating range that is different of the pilot and main injection staging ratio, and possibly lies between that of the pilot injection range alone and that of the pilot and main injection range. The transition staging ratio may vary between a series of constant intermediate values, each different from the pilot and main injection staging ratio, and optionally each lying between that of the pilot injection range alone and that of the pilot and main injection range. Controlling fuel delivery may include dispensing fuel from the first fuel source, fuel from the second fuel source, or a mixture thereof, using a fuel mixer. The fuel mixer may be configured to deliver fuel to the pilot fuel injectors, and optionally to the main fuel injectors. The proportion of fuel delivered from the first fuel source compared to that from the second fuel source may be determined according to a desired achieved level of nvPM production at a particular fuel flow rate within the operating range of transition. The fuel proportion may be determined such that nvPM production does not exceed a predetermined threshold or such that nvPM production is minimized over a period of operation of the gas turbine engine. The proportion of fuel distributed from the first fuel source compared to that from the second fuel source during the transition operating range may be determined at least partially according to any one or more of: a) the quantity of fuel having the first fuel characteristic and the second fuel characteristic available for a proposed flight; b) the total fuel requirement quantity for the fuel injectors during pilot injection only operation for the entire flight within an operating range in which fuel is supplied from the first fuel source; and or c) a fuel composition limit parameter. According to another aspect, the invention relates to an aircraft comprising one or more gas turbine engines according to the fifth aspect, and optionally one or more of the associated statements above. According to a seventh aspect, the invention relates to a gas turbine engine for an aircraft, comprising: a staged combustion system having pilot fuel injectors and main fuel injectors, the staged combustion system operable in a pilot injection only operating range in which a fuel is distributed only to the pilot fuel injectors, and a pilot and main injection operating range in which a fuel is distributed at least to the main fuel injectors; And a fuel delivery regulator configured to control fuel delivery to the pilot and main fuel injectors, the fuel delivery regulator configured to receive fuel from a first fuel source containing a first fuel having a first characteristic fuel and a second fuel source containing a second fuel having a second fuel characteristic, the second fuel characteristic being different from the first, wherein: the staged combustion system is designed to operate in an acceleration mode in which acceleration of the engine from a steady state operating mode is caused; And The fuel delivery regulator is configured to distribute fuel to the pilot and / or main fuel injectors, during operation in at least a portion of the throttle mode, having a fuel characteristic different from the fuel delivered to the pilot fuel injectors and / or main during at least part of the operating mode in the stabilized state. To reduce the production of excessive amounts of nvPM during acceleration, it is known to switch to a "throttle" mode of operation of a gas turbine engine in which the staging point occurs at a power setting weaker engine. The inventors have observed that switching to such a known acceleration mode can, however, present a certain number of disadvantages. For example, an increase in HC and CO emissions may be caused. In an acceleration mode of the present application the fuel distribution regulator is designed to distribute fuel to the fuel injectors (i.e. the pilot and / or main fuel injectors) having a characteristic of fuel different from the fuel distributed to the fuel injectors (i.e. the pilot and / or main fuel injectors) during at least part of the steady state operating mode. The inventors have determined that increased nvPM emissions when the engine is operating in an acceleration mode can be avoided or reduced by using a fuel with different characteristics than that used during steady-state operation. This may allow the staging point during the acceleration mode to remain the same or similar to that of the steady state operating mode, avoiding or reducing / limiting a disadvantageous increase in HC or CO emissions. The first fuel characteristic can be associated with a production level lower nvPM than the second fuel characteristic. The fuel delivered to at least the pilot fuel injectors during the boost mode may be associated with an nvPM production level that is lower than that of the fuel delivered to the pilot and / or main fuel injectors during operation in the boost mode. the stabilized state. The first fuel characteristic may correspond to a higher proportion of SAF within the respective fuel compared to the second fuel characteristic, and the fuel dispensed during the acceleration mode may have a higher proportion of SAF. The staged combustion system may be configured to switch between operation in the pilot injection only and pilot and main injection operating ranges at a stage point. The staging point may be at the same or greater engine power in the acceleration mode compared to the steady state mode. The staging point may be at a lower engine power in the acceleration mode compared to the steady state mode and is at a higher power than a default staging point at which the staged combustion system is controlled when fuel of a different characteristic cannot be supplied to the combustion system. The fuel delivery regulator may be configured to distribute fuel to the pilot fuel injectors during pilot injection only operation in the throttle mode that has a different fuel characteristic than the fuel distributed to the main fuel injectors during pilot operation. pilot and main injection in the steady state operating mode of the engine. The fuel delivery regulator may be configured to distribute fuel during the pilot injection only operating range in the acceleration mode having a fuel characteristic determined based on a control parameter upon which the production of nvPM by the fuel depends. engine. The control parameter may be a fuel-air ratio in a combustion chamber of the staged combustion system. As the fuel-air ratio decreases, the proportion of fuel associated with low nvPM production delivered to the pilot fuel injectors can also be decreased. The fuel delivery regulator may be configured to switch fuel delivery to the main and / or pilot fuel injectors to that having a different fuel characteristic at a starting point of a period of operation in the mode 'acceleration. The fuel delivery regulator may be configured to revert to delivering fuel having the same fuel characteristic as that delivered in the steady state mode after a transition to pilot and main injection operation. The fuel delivery regulator may be configured to deliver fuel to the pilot and / or main fuel injectors during the acceleration mode at a rate greater than that sufficient to maintain steady-state operation of the engine. According to an eighth aspect, the invention relates to a method of operating a gas turbine engine for an aircraft, the gas turbine engine comprising a staged combustion system having pilot fuel injectors and fuel injectors main, the staged combustion system being operable in a pilot injection only operating range in which the fuel is distributed only to the pilot fuel injectors, and a pilot and main injection operating range in which the fuel is distributed to the less to the main fuel injectors, the method comprising: regulating fuel delivery to the pilot and main fuel injectors from a first fuel source containing a first fuel having a first fuel characteristic and a second fuel source containing a second fuel having a second fuel characteristic fuel, the second fuel characteristic being different from the first; operating the staged combustion system in an acceleration mode in which acceleration of the engine from a steady-state operating mode is caused; And delivering fuel to the pilot and / or main fuel injectors, during operation in at least a portion of the acceleration mode, having a fuel characteristic different from the fuel delivered to the pilot and / or main fuel injectors for at least part of the operating mode in the stabilized state. The first fuel characteristic may be associated with an nvPM production level that is lower than that of the second fuel characteristic, and the fuel delivered to at least the pilot fuel injectors during the acceleration mode may be associated with a level of nvPM production that is less than that of the fuel delivered to the pilot and / or main fuel injectors during operation in the steady state mode. The first fuel characteristic may correspond to a higher proportion of SAF within the respective fuel compared to the second fuel characteristic, and the fuel dispensed during the acceleration mode may have a higher proportion of FAS. The staged combustion system may be configured to switch between operation in the pilot injection only and pilot and main injection operating ranges at a stage point. The staging point may be at the same or greater engine power in the acceleration mode compared to the steady state mode. The staging point may be at a lower engine power in the boost mode compared to the steady state mode, and may be at a higher power than a default staging point at which the staged combustion system is controlled when fuel of a different characteristic cannot be supplied to the combustion system. Fuel may be delivered to the pilot fuel injectors during pilot-only operation in the throttle mode that has a different fuel characteristic than fuel delivered to the main fuel injectors during pilot-and-main-injection operation in the steady state operating mode of the engine. Delivery of fuel to the fuel injectors may include delivery of fuel during the pilot injection only operating range in the throttle mode having fuel characteristics based on a control parameter upon which nvPM production by the fuel injector depends. engine. The control parameter may be a fuel-air ratio in a combustion chamber of the staged combustion system. As the fuel-air ratio decreases, the proportion of fuel associated with low nvPM production delivered to the pilot fuel injectors can also be decreased. Fuel delivery to the fuel injectors may include switching fuel delivery to the main and / or pilot fuel injectors to one having a different fuel characteristic at a starting point of an operating period in the fuel injector. acceleration mode. Delivering fuel to the fuel injectors may include returning to delivery of fuel having the same fuel characteristic as that delivered in the steady state mode after a transition to pilot and main injection operation. During the acceleration mode fuel may be delivered to the fuel injectors at a rate greater than sufficient to maintain steady-state operation. According to another aspect, the invention relates to an aircraft comprising one or more gas turbine engines according to the seventh aspect, and optionally one or more of the associated statements above. According to a ninth aspect, the invention relates to a gas turbine engine for an aircraft, comprising: a staged combustion system having pilot fuel injectors and main fuel injectors, the staged combustion system being operable in a pilot injection only operating range and a pilot and main injection operating range; And a fuel delivery regulator configured to control fuel delivery to the pilot and main fuel injectors; a fuel characteristic determination module configured to determine one or more fuel characteristics of the fuel being dispensed into the staged combustion system; And a control device configured to determine a staging point defining the point at which the staged combustion system is switched between pilot injection only operation and pilot and main injection operation, the staging point being determined based on the determined fuel characteristic(s), and the controller being configured to control the staged combustion system based on the determined stage point. The inventors have determined that the staging point at which a staged combustion system is controlled can be determined based on the characteristics of the fuel supplied to the combustion chamber. As noted above in connection with the third, fourth, seventh and eighth aspects, the staging point may be controlled in cases where fuel of different characteristics is available to reduce certain engine emissions. The fuel characteristic(s) may indicate that the fuel is associated with a lower nvPM production level compared to fossil kerosene. The fuel characteristic(s) may include any one or more of: (i) a percentage of sustainable aviation fuel in the fuel; (ii) aromatic hydrocarbon content of the fuel; and or (iii) naphthalene content of the fuel. The controller may be configured to determine the staging point such that a staging point associated with one or more fuel characteristics that indicate that the fuel is associated with low nvPM production corresponds to a setting of higher engine power compared to a staging point associated with one or more fuel characteristics that indicate that the fuel is associated with relatively higher nvPM production. The determined staging point may be a cruise staging point, and the controller may be configured to control the system combustion using the staging point determined during an engine cruise operating condition. The controller may be configured to determine the staging point such that the staged combustion system is configured to switch between the pilot injection only operating range and the pilot and main injection operating range at a point d staging which corresponds to an operating mode at cruising speed in the stabilized state of the engine, the staging point defining a limit between a first operating range at engine cruising speed and a second range of operation at engine cruise speed. The first operating range in cruise mode may correspond to operation of the aircraft in a later part of a cruise segment of a flight, and the second operating range in cruising mode may correspond to operation of the the aircraft in a relatively earlier part of the cruise segment. The first operating range in cruising speed may correspond to operation in subsonic cruising speed in the stabilized state of the engine and the second operating range in engine cruising speed corresponds to operation in supersonic cruising speed at stabilized state of the engine. The determined staging point may be an engine acceleration staging point, and the controller may be configured to control the staged combustion system using the determined staging point during an acceleration operating condition. of the motor. The engine throttle staging point may be determined to be the same (e.g., set to be the same) as a cruise RPM staging point in response to the one or more fuel characteristics. According to a tenth aspect, the invention relates to a method of operating a gas turbine engine for an aircraft, the gas turbine engine comprising a staged combustion system having pilot fuel injectors and fuel injectors main, the staged combustion system being able to operate in a pilot injection only operating range and a pilot and main injection operating range, the method comprising: determining one or more fuel characteristics of a fuel being dispensed into the staged combustion system; determining a staging point, defining the point at which the staged combustion system is switched between pilot injection only operation and pilot and main injection operation, depending on the determined fuel characteristic(s); And the control of the staged combustion system according to the determined stage point. The fuel characteristic(s) may indicate that the fuel is associated with a lower nvPM production level compared to fossil kerosene. The fuel characteristic(s) includes any one or more of: (i) a percentage of sustainable aviation fuel in the fuel; (ii) aromatic hydrocarbon content of the fuel; and or (iii) naphthalene content of the fuel. Determining the staging point may include determining the staging point such that a staging point associated with one or more fuel characteristics that indicate that the fuel is associated with low nvPM production corresponds to at a higher engine power setting relative to a staging point associated with one or more fuel characteristics that indicate that the fuel is associated with relatively higher nvPM production. The determined staging point may be a cruise staging point, and the combustion system may be controlled using the determined staging point during a cruise operating condition of the engine. The staging point may be determined such that the staged combustion system is configured to switch between the pilot injection only operating range and the pilot and main injection operating range at a staging point that corresponds to a operating mode in cruise mode in the stabilized state of the engine. The staging point may define a limit between a first operating range at engine cruise speed and a second operating range at engine cruise speed. The first operating range in cruise mode may correspond to operation of the aircraft in a later part of a cruise segment of a flight, and the second operating range in cruising mode may correspond to operation of the the aircraft in a relatively earlier part of the cruise segment. The first operating range in cruising speed may correspond to operation in subsonic cruising speed in the stabilized state of the engine and the second operating range in cruising speed may correspond to operation in supersonic cruising speed at stabilized state of the engine. The determined staging point may be an engine acceleration staging point, and the staged combustion system may be controlled using the determined staging point during an engine acceleration operating condition. The engine throttle staging point may be determined to be the same as a cruise RPM staging point in response to the one or more fuel characteristics. According to an eleventh aspect, the invention relates to an aircraft comprising one or more gas turbine engines according to the ninth aspect, and optionally one or more of the associated statements above. According to a twelfth aspect, the invention relates to a gas turbine engine for an aircraft, comprising: a staged combustion system having pilot fuel injectors and main fuel injectors, the staged combustion system being operable in a pilot injection only operating range and a pilot and main injection operating range; a fuel delivery regulator configured to control fuel delivery to the pilot and main fuel injectors; a fuel characteristic determination module configured to determine one or more fuel characteristics of the fuel being dispensed into the staged combustion system; And a controller configured to determine a staging ratio defining the ratio between the fuel flow of the pilot fuel injector and the fuel flow of the main fuel injector, the staging ratio being determined by function of the fuel characteristic(s), and the control device being configured to control the staged combustion system as a function of the determined stage ratio, The inventors have determined that the staging ratio can be determined based on the fuel characteristic(s) to more effectively manage engine emissions. This may, for example, allow the staging ratio to be adjusted to reduce CO and HC production in a manner that would otherwise lead to high levels of nvPM production as discussed in connection with the fifth and sixth aspect above. The fuel characteristic(s) may indicate that the fuel is associated with a lower nvPM production level compared to fossil kerosene. The fuel characteristic(s) may include any one or more of: (i) a percentage of sustainable aviation fuel in the fuel; (ii) aromatic hydrocarbon content of the fuel; and or (iii) naphthalene content of the fuel. The control device can be configured to: control the staged combustion system during the pilot and main injection operating range according to a pilot and main injection stage ratio, and in which the stage ratio determined according to the fuel characteristic(s) is a ratio of transition staging; And controlling the staged combustion system such that it is operated in a transitional operating range between the pilot injection only operating range and the pilot and main injection operating range, wherein, within the range transition operation, the staged combustion system is controlled according to the transition stage ratio, the transition stage ratio being different from the pilot and main injection stage ratio. The transition staging ratio may exhibit continuous variation with changing motor power within the transition operating range. The continuous variation may be such that the proportion of the total fuel flow to the fuel injectors that can be attributed to a fuel flow to the pilot fuel injectors decreases with increasing engine power. The proportion of the total fuel flow to the fuel injectors that can be attributed to fuel flow to the main fuel injectors may decrease with increasing engine power in the transition operating range. The transition staging ratio can vary between a series of constant intermediate values, each different from the pilot and main injection staging ratio. Each intermediate value can be between that of the pilot injection range alone and that of the pilot and main injection range. According to a thirteenth aspect, the invention relates to a gas turbine engine comprising a staged combustion system having pilot fuel injectors and main fuel injectors, the staged combustion system capable of operating in an injection operating range pilot only and an operating range with pilot and main injection, the method comprising: determining one or more fuel characteristics of a fuel being dispensed into the staged combustion system; determining a staging ratio defining the ratio between the fuel flow of the pilot fuel injector and the fuel flow of the main fuel injector, the staging ratio being determined according to the characteristic(s) fuel risks; And controlling the staged combustion system according to the determined stage ratio. The fuel characteristic(s) may indicate that the fuel is associated with a lower nvPM production level compared to fossil kerosene. The fuel characteristic(s) may include any one or more of: (i) a percentage of sustainable aviation fuel in the fuel; (ii) aromatic hydrocarbon content of the fuel; and or (iii) naphthalene content of the fuel. Determining the staging ratio may include determining a transition staging ratio; And controlling the staged combustion system may include controlling the staged combustion system so that it is operated in a transitional operating range between the pilot injection only operating range and the pilot injection operating range and main. In the transition operating range the staged combustion system can be controlled according to the transition stage ratio, the transition stage ratio may be different from a pilot and main injection stage ratio according to which the staged combustion chamber is controlled during the pilot and main injection operating range. The transition staging ratio may exhibit continuous variation with changing motor power within the transition operating range. The continuous variation may be such that the proportion of the total fuel flow to the fuel injectors that can be attributed to a fuel flow to the pilot fuel injectors decreases with increasing engine power, and the proportion of the total fuel flow to the fuel injectors that can be attributed to fuel flow to the main fuel injectors increases with increasing engine power in the transition operating range. The transition staging ratio can vary between a series of constant intermediate values, each different from the pilot and main injection staging ratio. Each intermediate value can be between that of the pilot injection range alone and that of the pilot and main injection range. According to a fourteenth aspect, the invention relates to an aircraft comprising one or more gas turbine engines according to the thirteenth aspect, and optionally one or more of the associated statements above. According to a fifteenth aspect, the invention relates to a computer-implemented method for determining a fuel allocation for an aircraft, in which: the aircraft includes a first fuel source configured to contain a first fuel having a first fuel characteristic and a second fuel source configured to contain a second fuel having a second fuel characteristic, the second fuel characteristic being different from the first one ; the aircraft includes one or more gas turbine engines powered by fuel from the first and second fuel sources; the one or more gas turbine engines each include a combustion system staged having pilot fuel injectors and main fuel injectors, the staged combustion system being operable in a pilot injection only operating range and a pilot and main injection operating range; the one or more gas turbine engines each include a fuel delivery regulator configured to control fuel delivery to the pilot and main fuel injectors from the first fuel source and the second fuel source, the process comprising: obtaining a proposed mission description including a list of operating points for the gas turbine engine(s) during the mission; obtaining nvPM (non-volatile particulate matter) impact parameters for the gas turbine engine(s), the impact parameters being associated with each operating point of the proposed mission using fuel from the first fuel source, fuel from the second fuel source, or a mixture thereof; calculating an optimized set of one or more fuel characteristics for each operating point of the proposed mission defined in the mission description based on the nvPM impact parameters; And determining a fuel allocation based on the optimized set of one or more fuel characteristics. The inventors have determined that by calculating the fuel allocation in this manner, fuel can be allocated to a mission such that fuel having the required characteristics can be provided to the aircraft to carry out the proposed mission. while reducing the impact of nvPM. This can allow better use of the characteristics of the available fuel in reducing nvPM compared to loading a defined quantity of different types of available fuel, whatever the mission to be carried out using this fuel. The first fuel characteristic may be associated with a lower nvPM production level than the second fuel characteristic. Additionally or alternatively, the first fuel characteristic and the second fuel characteristic may be a percentage of SAF present in the respective fuel. Each operating point in the mission description may include any one or more of: one or more operating conditions in which the gas turbine engines must operate, one or more fuel flow values ​​corresponding to an operating point, and / or an operating time at a corresponding operating point. The nvPM impact parameters may include a nyPM impact parameter defining an amount of nvPM produced by the respective gas turbine engine for different respective fuel characteristics including the first fuel, the second fuel, or a mixture thereof. -this at each operating point of the flight description. The fuel allocation may include any one or more of: 1) a quantity of fuel allocated to each of the first and second fuel sources; 11) the first fuel characteristic; iii) the second fuel characteristic; and or iv) a fuel mixture ratio. The method may further include determining one or more fuel usage parameters corresponding to the fuel allocation. Fuel usage settings can define how fuel is used during the mission defined by the mission description. The fuel usage parameter(s) may include any one or more of: i) a mixing schedule according to which fuel from the first fuel source and the second fuel source is mixed by the fuel distribution regulator: ii) a switchover schedule in which the fuel delivery regulator is configured to switch between dispensing fuel from the first fuel source and the second fuel source; iii) an allocation of fuel tanks provided in the aircraft to form the first fuel source and the second fuel source; and or iv) an isolation valve setting for fuel tanks forming the first fuel source and the second fuel source. The optimized set of fuel characteristic(s) may further be determined based on any one or more of: 1) the achievable range of fuel characteristics that can be provided by the fuel delivery regulator; (ii) a total quantity of a non-default fuel allocated to the mission; lii) a total fuel requirement for the mission; iv) the capacities of the aircraft's fuel tanks; eVou (v) restrictions on how aircraft fuel tanks may be allocated to the first or second fuel source. Calculating the optimized set of one or more fuel characteristics may include minimizing a cost function dependent on the nvPM impact parameter(s). The nvPM impact parameter(s) may include any one or more of: 1) the height above ground level at which nvPM production takes place; 11) location of nvPM production; iii) weather and / or atmospheric conditions at an nvPM production location; iv) climate impacts associated with the location of nvPM production; v) mass / size of individual nvPM particles produced; vi) potential contrail generation and / or contrail characteristics; vii) local air quality (LAQ) impact of nvPM production; and or viii) the quantity of nvPM produced (e.g. mass and / or number). According to a sixteenth aspect, the invention relates to a method of loading fuel onto an aircraft, comprising: determining a fuel allocation using the method according to the fifteenth aspect and optionally any one or more of the associated statements above; And loading fuel onto the aircraft based on fuel allocation. According to a seventeenth aspect, the invention relates to a non-transitory computer-readable medium on which are stored instructions which, when executed by a processor, cause the processor to implement the method of the fifteenth aspect and optionally a or any of the associated statements above. According to an eighteenth aspect, the invention relates to a fuel allocation determination system for determining a fuel allocation for an aircraft, the fuel allocation determination system comprising a computer device configured to implement the method of the fifteenth aspect and possibly one or more of any associated statements above. According to a nineteenth aspect, the invention relates to a fuel allocation determination system for determining a fuel allocation for an aircraft, in which: the aircraft includes a first fuel source configured to contain a first fuel having a first fuel characteristic and a second fuel source configured to contain a second fuel having a second fuel characteristic, the second fuel characteristic being different from the first one ; the aircraft includes one or more gas turbine engines powered by fuel from the first and second fuel sources; the gas turbine engine(s) each comprise a staged combustion system having pilot fuel injectors and main fuel injectors, the staged combustion system being operable in a pilot injection only operating range and a range of pilot and main injection operation; the one or more gas turbine engines each include a fuel delivery regulator configured to control fuel delivery to the pilot and main fuel injectors from the first fuel source and the second fuel source, the system for determining fuel loading parameters comprising: a mission description obtaining module configured to obtain a proposed mission description including a list of operating conditions for the gas turbine engine(s) during the mission: an impact parameter obtaining module configured to obtain impact parameters of the nvPMs for the gas turbine engine(s), the impact parameters being associated with each operating point of the proposed mission using compositions fuels that include fuel from the first fuel source, fuel from the second fuel source, or a mixture thereof; a fuel characteristics calculation module configured to calculate an optimized set of one or more fuel characteristics for each operating point of the proposed mission defined in the mission description based on the nvPM impact parameters; And a fuel allocation determination module configured to determine a fuel allocation based on the optimized set of one or more fuel characteristics. The first fuel characteristic may be associated with a lower nvPM production level than the second fuel characteristic. Additionally or alternatively, the first fuel characteristic and the second fuel characteristic may be a percentage of SAF present in the respective fuel. Each of the operating points of the mission description obtained by the mission description obtaining module may include any one or more of: one or more operating conditions in which the gas turbine engines must operate, one or more fuel flow values ​​corresponding to an operating point, and / or an operating time at a corresponding operating point. The nvPM impact parameters obtained by the parameter obtaining module impact parameters may include an nvPM impact parameter defining an amount of nvPM produced by the respective gas turbine engine for different respective fuel characteristics including the first fuel, the second fuel, or a mixture thereof at each operating point of the mission description. The fuel allocation determined by the fuel allocation determination module may include any one or more of: 1) a quantity of fuel allocated to each of the first and second fuel sources; 11) the first fuel characteristic; iii) the second fuel characteristic; and or iv) a fuel mixture ratio. The fuel allocation determination system may further comprise a fuel usage parameter determination module configured to determine one or more fuel usage parameters corresponding to the fuel allocation, the fuel usage parameters fuel use defining how the fuel is to be used during the mission defined by the mission description. The fuel usage parameter(s) may optionally include any one or more of: 1) a mixing schedule in which fuel from the first fuel source and the second fuel source is mixed by the fuel distribution regulator; ii) a switchover schedule in which the fuel delivery regulator is configured to switch between dispensing fuel from the first fuel source and the second fuel source; iii) an allocation of fuel tanks provided in the aircraft to form the first fuel source and the second fuel source; and or iv) an isolation valve setting for fuel tanks forming the first fuel source and the second fuel source. The optimized set of one or more fuel characteristics determined by the fuel characteristics calculation module can also be determined based on one or more elements among: 1) the achievable range of fuel characteristics that can be provided by the fuel delivery regulator; (ii) a total quantity of a non-default fuel allocated to the mission; iii) a total fuel requirement for the mission; iv) aircraft fuel tank capacities; and or (v) restrictions on how aircraft fuel tanks may be allocated to the first or second fuel source. The fuel characteristics calculation module may be configured to calculate the optimized set of one or more fuel characteristics by minimizing a cost function dependent on the nvPM impact parameter(s). The nvPM impact parameter(s) may include any one or more of: 1) the height above ground level at which nvPM production takes place; 11) location of nvPM production; iii) weather and / or atmospheric conditions at an nvPM production location; iv) climate impacts associated with the location of nvPM production; v) the mass of nvPM particles produced; vi) potential contrail generation and / or contrail characteristics; vii) local air quality (LAQ) impact of nvPM production; and or viii) the quantity of nvPM produced (e.g. mass / number). According to a twentieth aspect, the invention relates to an aircraft comprising: a first fuel source configured to contain a first fuel having a first fuel characteristic and a second fuel source configured to contain a second fuel having a second fuel characteristic, the second fuel characteristic being different from the first; one or more gas turbine engines powered by fuel from the first and second fuel sources, wherein: the gas turbine engine(s) each comprise a staged combustion system having pilot fuel injectors and main fuel injectors, the staged combustion system being operable in a pilot injection only operating range and a range of pilot and main injection operation; the one or more gas turbine engines each include a fuel delivery regulator configured to control fuel delivery to the pilot and main fuel injectors from the first fuel source, the second fuel source, or a mixture of these; And a fuel allocation determination system according to the eighteenth or nineteenth aspect and optionally any one or more of the associated statements above. In a twenty-first aspect, the invention relates to a computer-implemented method for determining a fleet-wide fuel allocation for a a plurality of missions flown by a plurality of aircraft, the plurality of missions being fueled from a fuel source comprising a quantity of a default fuel and a quantity of a non-default fuel, allocating fuel indicating the quantity of the non-default fuel and the default fuel to be assigned to each mission of the plurality of missions, the default fuel and the non-default fuel having one or more fuel characteristics different from each other , the process comprising: obtaining an initial proposed fuel allocation for each mission of the plurality of missions; performing a fleet-wide optimization in which the proposed fuel allocation of each mission of the plurality of missions is modified within the constraints of the default and / or non-default total available fuel of the fuel source to minimize a sum of nvPM impact parameters per mission across the entire plurality of missions, each mission of the plurality of missions being associated with a respective nvPM impact parameter per mission determined according to a fuel usage for that mission, the fuel usage defining how the fuel allocation for the respective mission is to be used during that mission; And determining the fleet-wide fuel allocation for the plurality of missions based on the fleet-wide optimization. The inventors have determined that the fuel available for a plurality of missions can be shared intelligently between these missions in order to advantageously use different types of fuel available. This allows less available fuel to be shared between missions in order to use it more efficiently, for example so that the overall impact of mission nvPMs is reduced. Non-default fuel may have a lower nvPM production level than the default fuel. The non-default fuel may be formed from a mixture of a first fuel having a first fuel characteristic and a second fuel having a second fuel characteristic, different from the first. The first and second fuel characteristics may be a percentage of SAF within the respective fuel. The non-default fuel may be a SAF-rich fuel and the default fuel may be a relatively low-SAF fuel (i.e., having a lower SAF content compared to the SAF-rich fuel). Achieving fleet-wide optimization may include: i) performing an outer loop optimization in which the fuel allocation of one or more of the plurality of missions is modified to reduce the sum of the nvPM impact parameters per mission of the plurality of missions; And ii) performing an inner loop optimization in which the fuel usage for each mission of the plurality of missions is obtained based on the constraints of the modified fuel allocation to determine a new proposed fuel usage for each mission of the plurality of missions. Steps 1) and ii) may be repeated until an optimized fuel usage for each mission of the plurality of missions is determined which corresponds to a minimized sum of the nvPM impact parameters per mission. The inner loop optimization may include obtaining a pre-prepared solution for fuel usage for a respective mission. The proposed fuel allocation for each of the plurality of missions may be achieved by obtaining an optimized fuel usage for the respective mission defining how the fuel should be used to minimize the nvPM impact parameter per mission for this mission. Optimized fuel usage for each mission can be achieved by performing per-mission optimization. Optimization per mission may include, for each respective mission: determining a type and / or operational capabilities of a combustion chamber used by the respective aircraft used for the mission; determining a total fuel requirement for the respective mission; determining a quantity of fuel required for each type of fuel injector provided in the combustion chamber for the respective mission where more than one type of injector is provided; determining the nvPM emissions dependence for each engine operating point of the mission using fuel having the characteristics of the default fuel, the non-default fuel, or a mixture thereof; And determining an optimized fuel usage that minimizes total nvPM emissions for the respective mission. Determining a type of combustion chamber used by the aircraft may include determining whether the aircraft includes a lean-burn staged combustion chamber or a rich-burn combustion chamber. If the combustion chamber is a lean burn staged combustion chamber having pilot and main fuel injectors, determining a quantity of fuel required for each type of fuel injector may include: a) determining a quantity of fuel required for the pilot injectors during pilot and main injection operation; and or b) determining a quantity of fuel required for the pilot fuel injectors during pilot injection only operation; and or c) determining a quantity of fuel required for the pilot fuel injectors operating within a threshold range of operation at fuel flow rates below that of the staging point. Fleet-wide optimization can be based on: a percentage of a first fuel having a first fuel characteristic in the default fuel defining the lowest possible percentage of fuel having the first fuel characteristic that can be used for combustion; and or a percentage of the first fuel having the first fuel characteristic in the non-default fuel defining the highest possible percentage of fuel having the first fuel characteristic that can be used for combustion; and or the amount of non-default fuel available for the plurality of missions. According to a twenty-second aspect, the invention relates to a method of loading fuel onto a plurality of aircraft carrying out a plurality of missions, the plurality of missions being supplied with fuel from a fuel source comprising a quantity of a default fuel and a quantity of a non-default fuel, the method comprising: determining a fuel allocation for the plurality of missions using the method of the twenty-first aspect and optionally any one or more of the associated statements above; And loading fuel onto the plurality of aircraft based on the fuel allocation. According to a twenty-third aspect, the invention relates to a non-transitory computer-readable medium on which are stored instructions which, when executed by a processor, cause the processor to implement the method of the twenty-first aspect and possibly one or more of any associated statements above. According to a twenty-fourth aspect, the invention relates to a Motte-scale fuel allocation determination system for determining a fuel allocation for a fleet for a plurality of missions, the fuel allocation determination system fleet-wide fuel comprising a computing device configured to implement the method of the twenty-first aspect and optionally one or more of any associated statements above. According to a twenty-fifth aspect, the invention relates to a system for determining fleet-wide fuel allocation to determine a fuel allocation for a plurality of missions flown by a plurality of aircraft, the plurality of missions being fueled from a fuel source comprising a quantity of a default fuel and a quantity of a non-default fuel, the fuel allocation indicating the quantity of the non-default fuel and the default fuel to be allocated to each mission of the plurality of missions, the default fuel and the non-default fuel having one or more fuel characteristics different from each other, the method comprising: an initial proposed fuel allocation obtaining module configured to obtain an initial proposed fuel allocation for each mission of the plurality of missions; a fleet-wide optimization module configured to perform a fleet-wide optimization in which the proposed fuel allocation of each mission of the plurality of missions is modified within the constraints of the default and / or non-default fuel available total of the fuel source to minimize a sum of nvPM impact parameters per mission across the entire plurality of missions, each mission of the plurality of missions being associated with a respective per-mission nvPM impact parameter determined based on a proposed fuel usage for that mission, the fuel usage defining how the fuel allocation for the respective mission is to be used during that mission; And a fleet-wide fuel allocation determination module configured to determine the fleet-wide fuel allocation for the plurality of missions based on the fleet-wide optimization. Non-default fuel may have a lower nvPM production level than the default fuel. The non-default fuel may be formed from a mixture of a first fuel having a first fuel characteristic and a second fuel having a second fuel characteristic, different from the first. The first and second fuel characteristics may be a percentage of SAF within the respective fuel, and wherein the non-default fuel is a high SAF fuel and the default fuel may be a relatively low SAF fuel. The Fleet Wide Optimization module can be configured to perform the following steps: i) performing an outer loop optimization in which the fuel allocation of one or more of the plurality of missions is modified to reduce the sum of the nvPM impact parameters per mission of the plurality of missions; And ii) performing an inner loop optimization in which the fuel usage for each mission of the plurality of missions is obtained based on the constraints of the modified fuel allocation to determine a new proposed fuel usage for each mission of the plurality of missions. The fleet-wide optimization module may be configured to repeat steps i) and ii) until an optimized fuel usage for each of the plurality of missions is determined which corresponds to a minimized sum of nvPM impact parameters per mission. The fleet-wide optimization module can be configured to perform the inner-loop optimization by obtaining a pre-prepared solution for fuel usage for a respective mission. The fleet-wide optimization module may be configured to obtain the proposed fuel allocation for each of the plurality of missions by obtaining an optimized fuel usage for the respective mission defining how the fuel should be used to minimize the per-mission nvPM impact parameter for this mission. The fleet-wide optimization module may be configured to obtain the optimized fuel usage for each mission by performing per-mission optimization, where the per-mission optimization optionally includes, for each mission: determining a type and / or operational capabilities of a combustion chamber used by the respective aircraft used for the mission; determining a total fuel requirement for the respective mission; determining a quantity of fuel required for each type of fuel injector provided in the combustion chamber for the respective mission where more than one type of injector is provided; determining the dependence of nvPM emissions for each mission engine operating point of the mission using fuel having the characteristics of default fuel, non-default fuel, or a mixture of these; And determining an optimized fuel usage that minimizes total nvPM emissions for the respective mission. Determining a type of combustion chamber used by the aircraft may include determining whether the aircraft includes a lean-burn staged combustion chamber or a rich-burn combustion chamber. If the combustion chamber is a lean burn staged combustion chamber having pilot and main fuel injectors, determining a quantity of fuel required for each type of fuel injector can to understand : determining a quantity of fuel required for the pilot fuel injectors during pilot and main injection operation; and or determining a quantity of fuel required for the pilot fuel injectors during pilot injection only operation; and or determining a quantity of fuel required for the pilot fuel injectors operating within a threshold range of operation at fuel flow rates below that of the staging point. The fleet-wide optimization module can be configured to base fleet-wide optimization on: a percentage of a first fuel having a first fuel characteristic in the default fuel defining the lowest possible percentage of fuel having the first fuel characteristic that can be used for combustion; and or a percentage of the first fuel having the first fuel characteristic in the non-default fuel defining the highest possible percentage of fuel having the first fuel characteristic that can be used for combustion; and or the amount of non-default fuel available for the plurality of missions. The present description may apply to any relevant gas turbine engine configuration. Such a gas turbine engine may be, for example, a turbojet gas turbine engine, an open rotor gas turbine engine (in which the propeller is not surrounded by a nacelle), a turboprop or turbojet. Any such engine may or may not be provided with a post-combustion chamber. A gas turbine engine according to any aspect of the present invention may include an engine core including 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 (comprising fan blades). Such a fan can be located upstream of the engine core. Alternatively, in some examples, the gas turbine engine may include a fan located downstream of the engine core, for example when the gas turbine engine is an open rotor or turboprop (in which case the fan may be referred to as a propeller ). When the gas turbine engine is an open rotor or turboprop engine, the gas turbine engine may include two stages of counter-rotating propellers attached to and driven by a free power turbine via a shaft . The propellers can rotate in opposite directions so that one rotates clockwise and the other counterclockwise around the axis of rotation of the motor. Alternatively, the gas turbine engine may include a propeller stage and a guide vane stage configured downstream of the propeller stage. The guide vane stage can be of variable pitch. Thus, the high pressure, intermediate pressure and free power turbines respectively can drive high pressure and intermediate pressure compressors and propellers by suitable interconnecting shafts. Thus, the propellers can provide most of the propulsive thrust. When the gas turbine engine is an open rotor or turboprop engine, one or more of the propeller stages may be driven by a gearbox. The gearbox may be of the type described here. An engine according to the present invention may be a turbofan engine. Such an engine may be a direct-drive turbofan engine in which the fan is directly connected to the fan drive turbine, for example without a gearbox. In such a direct-drive turbofan engine, the fan can be said to rotate at the same rotational speed as the fan drive turbine. An engine according to the present invention may be a gearbox turbofan engine. In such an arrangement, the engine has a fan which is driven through a gearbox. Thus, such a gas turbine engine may include a gearbox which receives an input from the core shaft and outputs a drive to the fan so as to drive the fan at a rotational speed lower than that of the core tree. Input to the gearbox may be directly from the core shaft, or indirectly from the core shaft, for example via a spur shaft and / or gear. The core shaft can connect the turbine and the compressor, such that the turbine and the compressor rotate at the same speed (with the fan rotating at a lower speed). The gas turbine engine as described and / or claimed herein may have any suitable general architecture. For example, the gas turbine engine may have any desired number of shafts which connect turbines and compressors, for example, one, two or three shafts. Strictly by way of example, 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 include 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 configured to rotate at a higher rotational speed than the first core shaft. In such an arrangement, the second compressor can be positioned axially downstream of the first compressor. The second compressor may be designed to receive (e.g. receive directly, for example via a generally annular conduit) a flow from the first compressor. The gearbox may be designed to be driven by the core shaft which is configured to rotate (e.g. in use) at the lowest rotational speed (e.g. the first core shaft in the example above). For example, the gearbox may be designed to be driven only by the core shaft which is configured to rotate (e.g. in use) at the lowest rotational speed (e.g. only the first shaft kernel tree, not the second kernel tree, in the example above). Alternatively, the gearbox may be designed to be driven by one or more shafts, for example the first and / or second shafts in the example above. The gearbox may be a reduction gearbox (in that the output to the blower is one rotational speed lower than the input from the core shaft). Any type of gearbox can be used. For example, the gearbox may be a "planetary" or "epicyclic planetary" gearbox, as described in more detail elsewhere herein. Such a gearbox can be single stage. Alternatively, such a gearbox may be a compound gearbox, for example a compound planetary gearbox (which may have the input on the sun gear and the output on the ring gear, and thus be called “compound planetary epicyclic” gearbox), for example having two reduction stages. The gearbox can have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example greater than 2.5, for example in the range of 3 to 4.2, or 3.2 to 3.8, for example of the order of or at least 3, 3.1, 3.2, 3.3, 3.4 , 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. The gear ratio can be, for example, between any two of the values ​​in the previous sentence. Strictly by way of example, the gearbox may be an "epicyclic planetary" gearbox having a reduction ratio in the range of 3.1 or 3.2 to 3.8. Strictly by way of further example, the gearbox may be an "epicyclic planetary" gearbox having a reduction ratio in the range of 3.0 to 3.1. Strictly by way of further example, the gearbox may be a "planetary" gearbox having a reduction ratio in the range of 3.6 to 4.2. In some arrangements, the gear ratio may be outside of these ranges. In any gas turbine engine as described and / or claimed herein, fuel of a given composition or mixture is supplied to a chamber combustion, which can be arranged downstream of the blower and the compressor(s) relative to the flow path (for example axially downstream). For example, the combustion chamber may be directly downstream of (e.g. at the outlet of) the second compressor, where a second compressor is arranged. As a further example, the flow leaving the combustion chamber can be supplied to the inlet of the second turbine, where a second turbine is disposed. The combustion chamber can be placed upstream of the turbine(s). The or each compressor (for example the first compressor and the second compressor as described above) may comprise any number of stages, for example multiple stages. Each stage may comprise a row of rotor blades and a row of stator blades, which may be variable stator blades (in that their angle of incidence may be variable). The row of rotor blades and the row of stator blades can be offset axially from one another. For example, the gas turbine engine may be a direct drive turbofan gas turbine engine comprising 13 or 14 compressor stages (in addition to the fan). Such a motor can for example comprise 3 stages in the first compressor (or “low pressure” compressor) and either 10 or 11 stages in the second compressor (or “high pressure” compressor). As a further example, the gas turbine engine may be a "gear" gas turbine engine (in which the fan is driven by a first core shaft through a gearbox) comprising 11, 12 or 13 compressor stages (in addition to the blower). Such an engine can include 3 or 4 stages in the first compressor (or “low pressure” compressor) and 8 or 9 stages in the second compressor (or “high pressure” compressor). As a further example, the gas turbine engine may be a "gear" gas turbine engine having 4 stages in the first compressor (or "low pressure" compressor) and 10 stages in the second compressor (or “high pressure” compressor). The or each turbine (e.g. the first turbine and the second turbine as described above) may comprise any number of stages, e.g. multiple stages. Each stage may include a row of rotor blades and a row of stator blades. The row of rotor blades and the row of stator blades can be offset axially from one another. The second turbine (or “high pressure” turbine) may comprise 2 stages in any arrangement (e.g. whether a gear or direct drive motor). The gas turbine engine may be a direct drive gas turbine engine comprising a first turbine (or "low pressure" turbine) having 5, 6 or 7 stages. Alternatively, the gas turbine engine may be a "high speed" gas turbine engine. gears” comprising a first turbine (or “low pressure” turbine) having 3 or 4 stages. Each fan blade may be defined as having a radial span extending from a root (or hub) at a radially internal gas scrubbing location, or a 0% span position, to one end at a 100% span position. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be less than (or of the order of) any number of: 0.4, 0 .39, 0.38, 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27 , 0.26, or 0.25. The ratio of the fan blade radius at the hub to the fan blade radius at the tip may be in an inclusive range delimited by any two of the values ​​in the preceding sentence (i.e. (say that the values ​​can form upper or lower limits), for example in the range 0.28 to 0.32 or 0.29 to 0.30. These ratios may commonly be referred to as hub-to-end ratio. The radius at the hub and the radius at the tip can both be measured at the leading edge (or axially most forward) portion of the blade. The hub-to-tip ratio refers, of course, to the gas-washed portion of the fan blade, that is, the portion radially outboard of any platform. The fan radius can be measured from the engine centerline to the tip of a fan blade at its leading edge. The fan diameter (which can simply be twice the fan radius) can be greater than (or of the order of) any number among: 140 cm, 170 cm, 180 cm, 190 cm, 200 cm, 210 cm, 220 cm, 230 cm, 240 cm, 250 cm (approximately 100 inches), 260 cm, 270 cm (approximately 105 inches), 280 cm (approximately 110 inches), 290 cm (approximately 115 inches), 300 cm (approximately 120 inches), 310 cm, 320 cm (about 125 inches), 330 cm (about 130 inches), 340 cm (about 135 inches), 350 cm, 360 cm (about 140 inches), 370 cm (about 145 inches), 380 (approximately 150 inches) cm, 390 cm (approximately 155 inches), 400 cm, 410 cm (approximately 160 inches) or 420 cm (approximately 165 inches). The blower diameter may be within an inclusive range delimited by any two of the values ​​in the preceding sentence (i.e. the values ​​may form upper or lower limits), for example in the range 240 cm to 280 cm or 330 cm to 380 cm. Strictly by way of non-limiting example, the blower diameter may be in the range of 170 cm to 180 cm, 190 cm to 200 cm, 200 cm to 210 cm, 210 cm to 230 cm, 290 cm to 300 cm or 340 cm to 360 cm. Blower rotation speed may vary during use. Generally, the rotation speed is lower for larger diameter blowers. Strictly by way of non-limiting example, the rotation speed of the fan in cruising speed conditions may be less than 3,500 rpm, for example less than 2,500 rpm, for example less than 2,300 rpm. Strictly by way of further non-limiting example, the fan rotation speed under cruise conditions for a "gear" gas turbine engine having a fan diameter in the range of 200 cm to 210 em may be in the range of 2750 to 2900 rpm. Strictly by way of further non-limiting example, the fan rotation speed under cruise conditions for a "gear" gas turbine engine having a fan diameter in the range of 210 cm to 230 cm cm can be in the range of 2500 to 2800 rpm. Strictly by way of further non-limiting example, the fan rotation speed under cruise conditions for a "gear" gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 1500 to 1800 rpm. Strictly by way of further non-limiting example, the fan rotation speed under cruising conditions for a direct drive engine having a fan diameter in the range of 190 cm to 200 cm may be in the range of 3,600 to 3,900 rpm. Strictly by way of further non-limiting example, the fan rotation speed under cruising conditions for a direct drive engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 2000 to 2800 rpm. When operating the gas turbine engine, the fan (with associated fan blades) rotates about an axis of rotation. This rotation causes the end of the fan blade to move with a speed U,,. The work carried out by the fan blades 23 on the flow leads to an increase in enthalpy dH of the flow. The loading of a fan tip can be defined as dH / Usi,> , where dH is the enthalpy rise (e.g. 1-D average enthalpy rise) on the blower and U,, is the (translational) speed of the fan tip, e.g. at the leading edge of the tip (which can be defined as the fan tip radius at the leading edge times the speed angular). Fan tip loading at cruise conditions may be greater than (or on the order of) any of: 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 0.4 (all values ​​dimensionless). The blower end loading may be in an inclusive range delimited by any two of the values ​​in the preceding sentence (i.e. the values ​​may form upper or lower limits), for example in the range 0 .28 to 0.31, or 0.29 to 0.3 (e.g. for a geared gas turbine engine). Gas turbine engines according to the present invention can have any desired dilution rate, where the dilution rate is defined as the ratio of the mass flow rate of the flow through the bypass conduit to the mass flow rate of the flow through the core at cruise conditions. In certain arrangements the dilution rate may be greater than (or of the order of) any number among: 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 dilution ratio can be in one 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 12 to 16, 13 to 15, or 13 to 14 Strictly by way of non-limiting example, the bypass ratio of a direct drive gas turbine engine according to the present description may be in the range of 9:1 to 11:1. Strictly by way of non-limiting example, the bypass ratio of a geared gas turbine engine according to the present disclosure may be in the range of 12:1 to 15:1. The bypass conduit may be substantially annular. The bypass conduit may be radially outboard of the base engine. The radially external surface of the bypass duct may be defined by a nacelle and / or a fan casing. The overall pressure ratio of a gas turbine engine as described and / or claimed herein can be defined as the ratio of the stagnation pressure at the compressor outlet to the highest pressure (before entering the combustion chamber) at the stagnation pressure upstream of the fan. By way of non-limiting example, the overall pressure ratio of a gas turbine engine as described and / or claimed here at cruising speed may be greater than (or of the order of) any number among: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio may be in an inclusive range delimited by any two of the values ​​in the preceding sentence (i.e., the values may form upper or lower limits), for example in the range 50 to 70. Strictly by way of non-limiting example, the overall pressure ratio under cruising conditions of a gas turbine engine gear having a fan diameter in the range of 200 cm to 210 cm may be in the range of 40 to 45. Strictly by way of non-limiting example, the overall pressure ratio at cruise speed conditions of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 45 to 55. Strictly by way of non-limiting example, the overall pressure ratio under operating conditions cruising speed of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 50 to 60. Strictly by way of non-limiting example, the overall pressure ratio under cruising conditions of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 50 to 60. 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, specific thrust may depend, for a given thrust condition, on the specific composition of the fuel supplied to the combustion chamber. Under cruising conditions, the specific thrust of an engine described and / or claimed here may be less than (or of the order of) any number among: 110 Nkg" s, 105 Nkg"! s, 100 Nkg-" s, 95 Nkg"! s, 90 Nkg"! s, 85 Nkg" s or 80 Nkg" s. The specific thrust may be in an inclusive range delimited by any two of the values ​​in the preceding sentence (i.e. the values ​​may form upper or lower limits), for example in the range of 80 Nkg! s to 100 Nkg" s or 85 Nkg- s to 95 Nkg"! s. Such engines can be particularly efficient compared to conventional gas turbine engines. Strictly by way of non-limiting example, the specific thrust of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 em may be in the range of 90 Nkg" s to 95 Nkg" s. Strictly by way of non-limiting example, the specific thrust of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 80 Nkg" s to 90 Nkg" s. Strictly as a non-limiting example, the specific thrust of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 70 Nkg" s at 90 Nkg"s. Strictly as a non-limiting example, the specific thrust of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 90 Nkg-! s to 120 Nkg"! s. A gas turbine engine as described and / or claimed herein can have any desired maximum thrust. Strictly by way of non-limiting example, a gas turbine as described and / or claimed here may be capable of producing a maximum thrust of at least (or of the order of) any following number: 100 kKN, 110 kN, 120 kN, 130 kN, 140 kN, 150 kN, 160 kN, 170 kKN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN, 450 kN, 500 kN, or 550 kN . The maximum thrust 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 limits). Strictly by way of non-limiting example, a gas turbine as described and / or claimed herein may be capable of producing maximum thrust in the range of 330 kN to 420 KN, for example 350 kN to 400 kN. Strictly by way of non-limiting example, the maximum thrust of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 140 kN to 160 kN. Strictly by way of non-limiting example, the maximum thrust of a geared gas turbine engine having a diameter of blower in the range of 210 cm to 230 em can be in the range of 150 kN to 200 kN. Strictly by way of non-limiting example, the maximum thrust of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 370 kN to 500 kN. Strictly by way of non-limiting example, the maximum thrust of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 370 kN to 500 kN. The thrust discussed above may be the maximum net thrust under standard atmospheric conditions at sea level plus 15 degrees C (ambient pressure 101.3 kPa, temperature 30 degrees C), with the engine static. During use, the flow temperature at the inlet of the high pressure turbine can be particularly high. This temperature, which can be called TET, can be measured at the outlet of the combustion chamber, for example immediately upstream of the first turbine blade, which can itself be called nozzle guide vane. In some examples, the TET may depend, for a given thrust condition, on the specific composition of the fuel supplied to the combustion chamber. At cruising speed, the TET can be at least (or of the order of) any number among: 1,400 K, 1,450 K, 1,500 K, 1,550 K, 1,600 K or 1,650 K. Thus, Strictly by way of non-limiting example, the cruise TET of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 cm may be in the range of 1540 K to 1 600 K. Strictly by way of non-limiting example, the cruising TET of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 cm may be in the range of 1590 K to 1650 K. Strictly by way of non-limiting example, the cruise TET of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range from 1600 K to 1660 K. Strictly by way of non-limiting example, the TET at cruise speed of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm can be in the range of 1590 K to 1650 K. Strictly by way of non-limiting example, the cruising TET of a direct drive gas turbine engine having a fan diameter in the range of 300 cm at 340 cm can be in the range of 1570 K to 1630 K. The cruise TET may be within an inclusive range delimited by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower limits). The maximum TET when using the motor can be, for example, at least (or of the order of) any number among: 1700 K, 1750 K, 1800 K, 1850 K, 1900 K , 1950 K, 2000 K, 2050 K or 2100 K. Thus, strictly by way of non-limiting example, the maximum TET of a geared gas turbine engine having a fan diameter in the range of 200 cm to 210 em may be in the range of 1890 K to 1960 K. Strictly by way of non-limiting example, the maximum TET of a geared gas turbine engine having a fan diameter in the range of 210 cm to 230 em may be in the range of 1890 K to 1960 K. Strictly by way of non-limiting example, the maximum TET of a geared gas turbine engine having a fan diameter in the range of 340 cm to 360 cm may be in the range of 1890 K to 1960 K. Strictly by way of non-limiting example, the maximum TET of a direct drive gas turbine engine having a fan diameter in the range from 300 cm to 340 cm may be in the range of 1935 K to 1995 K. Strictly by way of non-limiting example, the maximum TET of a direct drive gas turbine engine having a fan diameter in the range of 300 cm to 340 cm may be in the range of 1890 K to 1950 K. The maximum TET may be in an inclusive range delimited by any two of the values ​​in the preceding sentence (i.e. the values ​​may form upper or lower limits), for example in the range 1800 K to 1950 K. The maximum TET can occur, for example, in a high thrust condition, for example in a maximum take-off (MTO) condition. A fan blade and / or airfoil portion of a fan blade described and / or claimed herein may be made from any suitable material or combination of materials. For example at least part of the fan blade and / or the aerodynamic profile may be manufactured at least in part from a composite, for example a metal matrix composite and / or an organic matrix composite, such as 'a carbon fiber composite. As a further example at least part of the fan blade and / or the aerodynamic profile may be manufactured at least in part from a metal, such as a titanium-based metal or a titanium-based material. aluminum (such as an aluminum-lithium alloy) or a steel-based material. The fan blade may include at least two regions manufactured using different materials. For example, the fan blade may have a protective leading edge, which may be made using a material better able to withstand impact (e.g. from birds, ice or other material). than the rest of the dawn. Such a leading edge can, for example, be manufactured using titanium or a titanium-based alloy. Thus, by way of example only, the fan blade may have a body based on carbon fibers or aluminum (such as a lithium aluminum alloy) with a leading edge made of titanium. A fan as described and / or claimed herein may comprise a central portion, from which the fan blades may extend, for example in a radial direction. The fan blades can be attached to the part central in any way desired. For example, each fan blade may include an accessory that can engage a corresponding slot in the hub (or disk). Strictly by way of example, such an accessory may be in the form of a dovetail which may fit into and / or engage a corresponding slot in the hub to secure the fan blade to the hub / disc. For example, the fan blades may be an integral part of a central part. Such an arrangement may be called a vaned disk or vaned ring. Any suitable method can be used to make such a bladed disk or bladed ring. For example, at least a portion of the fan blades may be machined from a block and / or at least a portion of the fan blades may be attached to the hub / disc by welding, including linear friction welding. The gas turbine engines described and / or claimed herein may or may not have a variable area nozzle (VAN). Such a nozzle with a variable section can make it possible to vary the outlet area of ​​the diversion conduit during use. The general principles of the present invention can be applied to engines with or without VAN. The fan of a gas turbine as described and / or claimed herein may have any desired number of fan blades, for example, 14, 16, 18, 20, 22, 24 or 26 fan blades. When the fan blades have a carbon fiber composite body, there may be 16 or 18 fan blades. When the fan blades have a metal body (for example aluminum-lithium or titanium alloy), there may be 18, 20 or 22 fan blades. As used herein, the terms idle, taxi, takeoff, climb, cruise, descent, approach and landing have the conventional meaning and are readily understood by those skilled in the art. Thus, for a given gas turbine engine for an aircraft, those skilled in the art will immediately recognize each term to designate a phase of operation of the engine within a given mission of an aircraft to which the gas turbine engine is intended to be fixed. In this regard, ground idle may refer to a phase of engine operation where the aircraft is stabilized and in contact with the ground, but where the engine must operate. At idle, the engine may be producing between 3% and 9% of the engine's available thrust. In additional non-limiting examples, the engine may be producing between 5% and 8% of the available thrust. In additional non-limiting examples, the engine may be producing between 6% and 7% of the available thrust. Taxiing may refer to a phase of engine operation where the aircraft is propelled along the ground by the thrust produced by the engine. When driving, the engine may be producing between 5% and 15% of available thrust. In additional non-limiting examples, the engine may be producing between 6% and 12% of the available thrust. In additional non-limiting examples, the engine may be producing between 7% and 10% of the available thrust. Takeoff may refer to a phase of engine operation where the aircraft is propelled by the thrust produced by the engine. At an initial stage in the takeoff phase, the aircraft may be propelled while the aircraft is in contact with the ground. At a later stage in the takeoff phase, the aircraft may be propelled while the aircraft is not in contact with the ground. During takeoff, the engine may be producing between 90% and 100% of the available thrust. In additional non-limiting examples, the engine may be producing between 95% and 100% of the available thrust. In additional non-limiting examples, the engine may be producing 100% of the available thrust. Climb may refer to a phase of engine operation where the aircraft is propelled by the thrust produced by the engine. During climb, the engine may be producing between 75% and 100% of the available thrust. In additional non-limiting examples, the engine may be producing between 80% and 95% of the available thrust. In additional non-limiting examples, the engine may be producing between 85% and 90% of the available thrust. In this regard, climb may refer to a phase of operation within an aircraft flight cycle between takeoff and arrival at cruise conditions. Additionally or alternatively, climb may refer to a nominal point in an aircraft flight cycle between takeoff and landing, when a relative increase in altitude is required, which may require a request for thrust additional engine load. As used herein, cruise conditions have the conventional meaning and are readily understood by those skilled in the art. Thus, for a given gas turbine engine for an aircraft, those skilled in the art will immediately recognize the cruise regime conditions as being the operating point of the engine in the middle of the cruise regime of a given mission (which may be referred to in the industry as the "economic mission") of an aircraft to which the gas turbine engine is intended to be attached. In this regard, mid-cruise is the point in an aircraft flight cycle at which 50% of the total fuel that is burned between the top of the climb and the start of the descent has been burned (which can be approximated by the midpoint in terms of time and / or distance) between the top of the climb and the start of the descent. The cruise conditions thus define an operating point of the gas turbine engine which provides a thrust which would ensure operation in the stabilized state (i.e. maintaining an altitude constant and a constant Mach number) in the middle of cruise of an aircraft to which it is intended to be attached, taking into account the number of engines supplied to this aircraft. For example when an engine is designed to be attached to an aircraft which has two engines of the same type, under cruise conditions the engine provides half the total thrust that would be required for steady state operation of this aircraft in mid-cruise. In other words, for a given gas turbine engine for an aircraft, cruise conditions are defined as the engine operating point that provides a specified thrust (required to provide (in combination with any which other engines on the aircraft) operation in the stabilized state of the aircraft to which it is intended to be fixed at a given mid-cruise regime Mach number) in the atmospheric conditions of mid-cruise regime ( defined by the International Standard Atmosphere according to ISO 2533 at the mid-cruise altitude). For any given gas turbine engine for an aircraft, the mid-cruise thrust, atmospheric conditions and Mach number are known, and hence the operating point of the engine at cruise RPM conditions. is clearly defined. Strictly by way of example, the forward speed at the cruise condition may be any point in the range Mach 0.7 to 0.9, e.g. 0.75 to 0.85, e.g. example 0.76 to 0.84, for example 0.77 to 0.83, for example 0.78 to 0.82, for example 0.79 to 0.81, for example of the order of Mach 0.8 , of the order of Mach 0.85 or in the range of 0.8 to 0.85. Any single speed within these ranges may be part of the cruise RPM condition. For some aircraft, cruise conditions may be outside of these ranges, for example below Mach 0.7 or above Mach 0.9. Strictly by way of example, cruise conditions may correspond to standard atmospheric conditions (according to the International Standard Atmosphere, ISA) at an altitude which is 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 (around 38,000 feet), 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 (about 35,000 feet) to 11,300 m, for example in the range of 10,800 m to 11,200 m, for example in the range from 10,900 m to 11,100 m, for example of the order of 11,000 m. Cruise conditions may correspond to standard atmospheric conditions at any given altitude within these ranges. Strictly by way of example, the cruise speed conditions may correspond to an engine operating point which provides a level of thrust known required value (e.g. a value in the range 30 kN to 35 kN) at a forward Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38,000 feet (11,582m). Strictly by way of further example, cruise conditions may correspond to an engine operating point which provides a known required level of thrust (e.g. a value in the range 50 kN to 65 kN) at a number of Forward Mach of 0.85 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35,000 feet (10,668 m). In use, a gas turbine engine described and / or claimed herein may operate at the cruise rpm conditions defined elsewhere herein. Such cruise speed conditions may be determined by the cruise speed conditions (e.g. mid-cruise conditions) of an aircraft on which at least one (e.g. 2 or 4) gas turbine engine can be mounted to provide propulsion thrust. Furthermore, those skilled in the art will immediately recognize descent and / or approach to refer to a phase of operation within an aircraft flight cycle between cruise and landing of the aircraft. . During descent and / or approach, the engine may be producing between 20% and 50% of the available thrust. In additional non-limiting examples, the engine may be producing between 25% and 40% of the available thrust. In additional non-limiting examples, the engine may be producing between 30% and 35% of the available thrust. Additionally or alternatively, descent may refer to a nominal point in an aircraft flight cycle between takeoff and landing, when a relative decrease in altitude is required, which may require a request for thrust reduced from the engine. According to one aspect, the invention relates to an aircraft comprising a gas turbine engine as described and / or claimed here. The aircraft according to this aspect is the aircraft to which the gas turbine engine is intended to be attached. Thus, the cruise regime conditions in this aspect correspond to the midpoint of the aircraft cruise regime, as defined elsewhere herein. According to one aspect, the invention relates to a method of operating a gas turbine engine as described and / or claimed here. Operation may be under any suitable condition, which may be as defined elsewhere herein (e.g. in terms of thrust, atmospheric conditions and Mach number). According to one aspect, the invention relates to a method of operating an aircraft comprising a gas turbine engine as described and / or claimed here. Operation in this aspect may include (or may be) operation in any suitable condition, e.g. mid-cruise. the aircraft, as defined elsewhere in the description. Those skilled in the art will have in mind that, except where mutually exclusive, a characteristic or parameter described in connection with any of the above aspects can be applied to any other appearance. Additionally, except where mutually exclusive, any feature or parameter described herein may be applied to any aspect and / or combined with any other feature or parameter described herein. here. Embodiments will now be described by way of example only, with reference to the figures, in which: [Fig.1] is a side sectional view of a gas turbine engine; [Fig.2] is a side view in close section of an upstream part of a gas turbine engine; [Fig.3] is a partially cut away view of a gearbox for a gas turbine engine; [Fig.4] is a schematic view of an aircraft comprising two fuel sources; [Fig.5] is a schematic view of a combustion system of a gas turbine engine connected to two fuel sources; [Fig.6] illustrates a relationship between fuel flow and nvPM number; [Fig.7] illustrates dependencies of nvPM number on fuel flow for various fuel compositions; [Fig.8] illustrates another dependence of nvPM number on fuel flow for various fuel compositions; [Fig.9] represents a schematic view of another combustion system; [Fig.10] illustrates the dependence of the number of nvPM on the fuel flow for the arrangement of [Fig.9]; [Fig.11] illustrates a method of operating a gas turbine engine; [Fig.12] represents a schematic view of another combustion system; [Fig.13] illustrates the dependence of the number of nvPM on the fuel flow for the arrangement of [Fig.12]; [Fig.14] illustrates another method of operating a gas turbine engine; [Fig.15] illustrates another example of the dependence of the number of nvPM on the fuel flow for the arrangement of [Fig.12]; [Fig.16] illustrates another method of operating a gas turbine engine; [Fig.17] represents a schematic view of another combustion system; [Fig.18] illustrates the dependence of the number of nvPM on the fuel flow for the arrangement shown in [Fig.17]; [Fig.19] illustrates a region of achievable nvPM numbers depending on the flow rate fuel for the arrangement shown in [Fig.17]; [Fig.20] illustrates another dependence of the number of nvPM on the fuel flow for the arrangement shown in [Fig.17]; [Fig.21] illustrates another method of operating a gas turbine engine; [Fig.22] illustrates the dependence of the emission index (ED) of carbon monoxide (CO) and unburned hydrocarbons (HC) on the engine power setting; [Fig.23] illustrates the dependence of the emission index (EI) of carbon monoxide (CO) and unburned hydrocarbons (HC) on the engine power setting with a staging point (SP) set to define the first and second operating ranges at cruise speed; [Fig.24] illustrates the dependence of the nvPM emission index on the engine power setting; [Fig.25] illustrates the dependence of the nvPM emission index on the engine power setting to illustrate the result of the selective use of a fuel of a different characteristic during a first operation in engine cruise speed; [Fig.26] illustrates another method of operating a gas turbine engine; [Fig.27] illustrates the dependence of the CO and HC emission index (EI) on the engine power setting to illustrate the effect of a transition operating range between a operating range with pilot injection only and pilot and main injection; [Fig.28] illustrates the dependence of the nvPM emission index on the engine power setting to further illustrate the effect of the transition operating range; [Fig.29] illustrates another method of operating a gas turbine engine; [Fig.30] illustrates another method of operating a gas turbine engine; [Fig.31] represents a schematic view of another combustion system; [Fig.32] illustrates another method of operating a gas turbine engine; [Fig.33] illustrates another method of operating a gas turbine engine; [Fig.34] illustrates a method for determining one or more fuel loading parameters; [Fig.35] illustrates a method of loading fuel onto an aircraft; [Fig.36] shows a schematic view of a system for determining fuel loading parameters; [Fig.37] illustrates a method for determining an optimized fleet-wide fuel allocation for a plurality of flights; [Fig.38] illustrates a flight optimization to determine optimized fuel usage; [Fig.39] illustrates a method of loading fuel onto a plurality of aircraft: [Fig.40] shows a schematic view of a system for determining fuel allocation parameters on a fleet scale; And [Fig.41] shows a schematic view of a computing device. [Fig.1] illustrates a gas turbine engine 10 having a main axis of rotation 9. The engine 10 comprises an air inlet 12 and a propulsion fan 23 which generates two air flows: a flow of central air À and a bypass air flow B. The gas turbine engine 10 comprises a central part 11 which receives the central air flow À. The central part 11 comprises, in axial flow series, a low pressure compressor 14, a high pressure compressor 15, combustion equipment 16, a high pressure turbine 17, a low pressure turbine 19 and a nozzle. central exhaust 20. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass air flow B flows through the bypass duct 22 The fan 23 is attached to and driven by the low pressure turbine 19 via a shaft 26 and an epicyclic gearbox 30. In use, the central airflow λ is accelerated and compressed by the low pressure compressor 14 and directed into the high pressure compressor 15 where further compression occurs. The compressed air discharged from the high pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel F and the mixture is burned. Combustion equipment 16 may be referred to as combustion chamber 16, the terms "combustion equipment 16" and "combustion chamber 16" being used interchangeably herein. The hot combustion products obtained expand through the high pressure and low pressure turbines 17, 19 and thereby drive them, before being evacuated through the nozzle 20 to provide a certain propulsive thrust. The high pressure turbine 17 drives the high pressure compressor 15 through a suitable interconnecting shaft 27. The fan 23 generally provides the majority of the propulsive thrust. The epicyclic gear box 30 is a reduction box. An example arrangement for a geared fan gas turbine engine 10 is shown in [Fig.2]. The low pressure turbine 19 (see [Fig.1]) drives the shaft 26, which is coupled to a sun wheel, or sun gear, 28 of the epicyclic gear arrangement 30. Radially outwardly from the The sun gear 28 and meshing therewith are a plurality of planet gears 32 which are coupled together by a planet carrier 34. The planet carrier 34 forces the planet gears 32 to precede each other around the sun gear 28. solar gear 28 in synchronism while allowing each satellite gear 32 to rotate around its own axis. The planet carrier 34 is coupled via links 36 to the fan 23 in order to cause its rotation around the motor axis 9. Radially towards the outside of the planet gears 32 and meshing with them There is a ring or crown 38 which is coupled, via links 40, to a stationary support structure 24. It should be noted that the terms "low pressure turbine" and "low pressure compressor", as used herein, may be taken to mean the lowest pressure turbine stages and the lowest pressure compressor stages ( that is to say not including the fan 23) respectively and / or the turbine and compressor stages which are connected together by the interconnection shaft 26 with the lowest rotation speed in the engine (c that is to say not including the gearbox output shaft which drives the fan 23). In some literature, the "low pressure turbine" and "low pressure compressor" mentioned herein may alternatively be referred to as "intermediate pressure turbine" and "intermediate pressure compressor". When such alternative nomenclature is used, the blower 23 may be referred to as the first compression stage, or lowest compression stage. The epicyclic gearbox 30 is illustrated by way of example in more detail in [Fig.3]. Each of the sun gear 28, planet gears 32 and ring gear 38 includes teeth around their periphery for meshing with the other gears. However, for clarity only, exemplary portions of the teeth are shown in |Fig.3]. Four planet gears 32 are illustrated, although it will be apparent to those skilled in the art that more or fewer planet gears 32 may be provided within the scope of the claimed invention. Practical applications of a gearbox planetary planetary gears 30 generally comprise at least three planetary gears 32. The epicyclic gearbox 30 illustrated by way of example in Figures 2 and 3 is of the planetary type, in that the planet carrier 34 is coupled to an output shaft via connections 36, with the crown 38 fixed. However, any other suitable type of epicyclic gearbox 30 may be used. As a further example, the epicyclic gearbox 30 may be a star arrangement, in which the planet carrier 34 is held fixed, with the crown (or ring) 38 allowed to rotate. In such an arrangement the fan 23 is driven by the ring gear 38. As another additional example, the gearbox 30 may be a differential gearbox in which the ring gear 38 and the planet carrier 34 are both allowed to turn. It will be understood that the arrangement illustrated in Figures 2 and 3 is given for example only, and various variations are within the scope of this description. Strictly by way of example, any suitable arrangement may be used to locate the gearbox 30 in the motor 10 and / or to connect the gearbox 30 to the motor 10. As a further example, the connections (such as the connections 36, 40 in the example of [Fig.2]) between the gearbox 30 and other parts of the motor 10 (such as the input shaft 26, the output shaft and the fixed structure 24) can have any desired degree of rigidity or flexibility. As a further example, any suitable arrangement of bearings between rotating and stationary parts of the engine (e.g. between the input and output shafts of the gearbox and the fixed structures, notably the crankcase gearbox) can be used, and the invention is not limited to the arrangement given as an example in [Fig.2]. For example, when the gearbox 30 has a star arrangement (described above), those skilled in the art will easily understand that the arrangement of the output and support links and the bearing locations would generally be different from that shown as an example in [Fig.2]. Thus, the present description extends to a gas turbine engine having any arrangement of gearbox styles (e.g. star or planetary), support structures, shaft arrangement entry and exit, and bearing locations. Optionally, the gearbox can drive additional and / or different elements (e.g. the intermediate pressure compressor and / or a booster). Other gas turbine engines to which this description may be applied may have different configurations. For example, such engines may have a different number of compressors and / or turbines and / or a different number of interconnecting shafts. As a further example, the gas turbine engine shown in [Fig.1] has a split flow nozzle 18, 20, meaning that the flow through the bypass conduit 22 has its own nozzle 18 which is distinct from the central engine nozzle 20 and radially outside thereof. However, this is not limiting, and any aspect of the present description may also apply to engines in which the flow through the bypass conduit 22 and the flow through the core 11 are mixed, or combined, before (or upstream) of a single nozzle, which may be called a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable surface area. While the example described relates to a turbofan engine, the description can apply, for example, to any type of gas turbine engine, such as an open rotor (in which the fan stage n 'is not surrounded by a nacelle) or a turbo- thruster, for example. In some arrangements, the gas turbine engine 10 may not include a gearbox 30. The geometry of the gas turbine engine 10, and associated components, is defined by a conventional axis system, comprising an axial direction (which is aligned with the axis of rotation 9), a radial direction (in the direction of bottom to top in [Fig.1]), and a circumferential direction (perpendicular to the page in the view of [Fig.1]). The axial, radial and circumferential directions are perpendicular to each other. The fuel F supplied to the combustion equipment 16 may comprise a fossil hydrocarbon fuel, such as kerosene. Thus, the fuel F may include molecules from one or more of the chemical families of n-alkanes, isoalkanes, cycloalkanes and aromatic compounds. Additionally or alternatively, F fuel may include renewable hydrocarbons produced from biological or non-biological resources, otherwise known as sustainable aviation fuel (SAF). In each of the examples provided, the fuel F may comprise one or more trace elements including, for example, sulfur, nitrogen, oxygen, inorganic materials and metals. Functional performance of a given composition, or fuel mixture to be used in a given mission, may be defined, at least in part, by the ability of the fuel to sustain 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 particulate matter. Relatively higher specific energy (i.e. energy per unit mass), expressed in MJ / kg, can at least partially reduce takeoff weight, potentially allowing a relative improvement in fuel efficiency. A relatively higher energy density (i.e. energy per unit volume), expressed in MJ / L, can at least partially reduce the volume of fuel at takeoff, which may be particularly important for long-range missions. limited volume or military operations involving resupply. Relatively higher thermal stability (i.e., inhibition of fuel to degrade or coke under thermal stress) can allow the fuel to sustain high temperatures in the engine and fuel injectors, potentially allowing make improvements in combustion efficiency. Reducing emissions, including particulate matter, can help reduce the formation of contrails, while reducing the environmental impact of a given mission. Other fuel properties may also be critical to functional performance. For example, a relatively lower freezing point (°C) may enable long-range missions action to optimize flight profiles; Minimum aromatic concentrations (%) may ensure sufficient swelling of certain materials used in the construction of O-rings and gaskets that have been previously exposed to fuels with high aromatic content; and, maximum surface tension (mN / m) can ensure sufficient spray fragmentation and fuel atomization. The ratio of the number of hydrogen atoms to the number of carbon atoms in a molecule can influence the specific energy of a given composition, or fuel mixture. Fuels with higher ratios of hydrogen atoms to carbon atoms can have higher specific energies in the absence of bonding constraints. For example, fossil hydrocarbon fuels may include molecules of approximately 7 to 18 carbon atoms, with a significant portion of a given composition coming from molecules having 9 to 15 carbon atoms, with an average of 12 carbon atoms. carbon. A number of sustainable aviation fuel blends have been approved for use, comprising between 10% and 50% sustainable aviation fuel (the remainder comprising one or more fossil hydrocarbon fuels, such as kerosene), with compositions additional documents awaiting approval. However, the aviation industry expects the use of sustainable aviation fuel blends up to (and including) 100% sustainable aviation fuel (SAF) to be approved for use. Sustainable aviation fuels may include one or more of n-alkanes, isoalkanes, cycloalkanes and aromatics, and may be produced, for example, from one or more synthetic gases (syngas); lipids (e.g. fats, oils and greases); sugars; and alcohols. Thus, sustainable aviation fuels may include a lower content of aromatic compounds and / or sulfur, compared to fossil hydrocarbon fuels. Additionally or alternatively, sustainable aviation fuels may include a higher level of isoalkane and / or cycloalkane, compared to fossil hydrocarbon fuels. Thus, in some examples, sustainable aviation fuels may include a density of between 90% and 98% of that of kerosene and / or a calorific value of between 101% and 105% of that of kerosene. Due at least in part to the molecular structure of sustainable aviation fuels, sustainable aviation fuels may have beneficial effects including, for example, one or more of a higher specific energy (despite, in some examples, a higher density lower energy); higher specific heat capacity; higher thermal stability; a power read- higher gloss; lower viscosity; lower surface tension; a lower freezing point; lower soot emissions; and, CO emissions; lower, compared to fossil hydrocarbon fuels (for example when burned in combustion equipment 16). Thus, compared to fossil hydrocarbon fuels, such as kerosene, sustainable aviation fuels can lead to a relative reduction in specific fuel consumption, and / or a relative reduction in maintenance costs. Aircraft fuel system and fuel characteristics An aircraft 1 comprising two gas turbine engines 10 according to any of the examples described here is illustrated in [Fig.4]. In this example, the aircraft 1 comprises two gas turbine engines 10, but in other examples it may comprise one or more gas turbine engines. The aircraft 1 further comprises an aircraft fuel system located on board the aircraft suitable for supplying fuel F to each of the gas turbine engines 10 to be burned in the engine combustion equipment 16 such as described above. The aircraft fuel system is configured to supply fuel to an engine fuel system provided on each of the gas turbine engines 10. The engine fuel system and the aircraft fuel system together form the engine fuel system. fuel supply of aircraft 1. The aircraft fuel system includes two fuel sources, a first fuel source 302 and a second fuel source 304 (having different shading in [Fig.4]). For the purposes of this application the expression "fuel source" means either 1) a single fuel tank, or 2) a plurality of fuel tanks which may or may not be fluidly interconnected. Each of the fuel sources is configured to provide distinct fuel sources, i.e. the first fuel source may contain a first fuel having one or more characteristics different from a second fuel contained in the second fuel source. fuel. The first and second fuel sources are therefore not fluidly coupled to each other so as to separate the different fuels (at least under normal operating conditions as indicated elsewhere here). The aircraft fuel system may include a plurality of fuel tanks that may be in fluid communication selectively in different arrangements to form the first and second fuel sources. The fuel tanks may include associated shut-off valves so that different combinations of tanks can be grouped in different configurations. In such examples, there may be several ways of assigning individual fuel tanks to the first fuel source and the second fuel source. fuel source. In other examples the allocation of fuel tanks to each of the fuel sources may be fixed. In the present example, the total volume of fuel tanks forming the first fuel source 302 is less than or equal to the volume of the fuel tanks forming the second fuel source 304. This may however not be the case for other examples . In the present example, the first and second fuel sources 302, 304 include a plurality of wing fuel tanks 53, where at least one wing fuel tank is located in the left wing and at least one wing fuel tank is located in the right wing, and a central fuel tank 55 located primarily in the fuselage of the aircraft 1. Each of the central fuel tank 55 and the wing fuel tanks 53 may include a plurality of fluidly interconnected fuel tanks not shown in the figures. As shown by the shading in [Fig.4], the wing fuel tanks 53 form the first fuel source 302, the second fuel source 304 being formed by the central fuel tank 55, in the present example . For balancing purposes, one or more fuel tanks in the left wing may be in fluid communication with one or more fuel tanks in the right wing. This is achievable either through a central fuel tank (if that tank is not part of the other fuel source), or by bypassing the central fuel tank(s), or both (for flexibility and maximum security). In another example, the second fuel source 304 includes wing fuel tanks and a center fuel tank, while the first fuel source 302 includes an additional center fuel tank. The fluid interconnection between the wing fuel tanks and the central fuel tank of the second fuel source may be provided to balance the aircraft. In some examples, the allocation of available fuel tanks on the aircraft may be constrained such that the first fuel source and the second fuel source 302, 304 are each substantially symmetrical with respect to the centerline of the aircraft. aircraft. In cases where an asymmetrical allocation of fuel tanks is permitted, suitable fuel transfer means are provided between fuel tanks of the first fuel source and / or between fuel tanks of the second fuel source of such so that the position of the center of mass of the aircraft can be maintained within acceptable lateral limits throughout the flight. A surge fuel tank could in some examples be part of the first fuel source 302, or in other examples be part of the second fuel source 304. The allocation of fuel tanks to the first fuel source 302 and the second fuel source 304 may be constrained such that neither the first fuel source 302 nor the second fuel source 304 Only include the trim fuel tank. While the examples described herein include only a first and a second fuel source 302, 304, in other examples additional fuel sources may be provided such that fuel from any number of sources, each having a different fuel characteristic or characteristics can be supplied to the combustion chamber through the fuel delivery regulator. The first and second fuel sources 302, 304 may be used to store fuel having different characteristics. This can make it possible to supply fuel with different characteristics to the combustion equipment 16 as will be described in more detail in the various examples below. As used herein, the term "fuel characteristics" means intrinsic or inherent fuel properties such as fuel composition, invariable properties such as volume or temperature. Examples of fuel characteristics include one or more of: 1. the percentage of sustainable aviation fuel (SAF) in the fuel, or an indication that the fuel is a fossil fuel, for example fossil kerosene, or pure SAF; li. parameters of a fuel hydrocarbon distribution, such as: @ the aromatic hydrocarbon content of the fuel, and possibly also / as a variant the multi-aromatic hydrocarbon content of the fuel; @ the hydrogen to carbon (H / C) ratio of the fuel; @ composition information in % for some or all of the hydrocarbons present; iii, the presence or percentage of a particular element or of a particular species, in particular: @ the percentage of nitrogen species in the fuel; @ the presence or percentage of a species of tracer or trace element in the fuel (e.g. a trace substance inherently present in the fuel which may vary between fuels and thus be used to identify a fuel, and / or a substance added voluntarily to act as a tracer); @ the naphthalene content of the fuel; @ the sulfur content of the fuel; @ the cycloparaffin content of the fuel; @ the oxygen content of the fuel; iv, one or more properties of the fuel when used in a gas turbine engine 10, such as: @ the rate of non-volatile particulate matter (nvPM) or CO emissions; to combustion (a value may be provided for a specific combustion chamber operating under particular conditions to fairly compare fuels; a measured value may be adjusted accordingly depending on the properties and conditions of the combustion chamber) ; @ the level of fuel scaling; v. one or more properties of the fuel itself, independently of use in an engine or combustion, such as: @ the thermal stability of the fuel (for example the thermal breakdown temperature); And @ one or more physical properties such as density, viscosity, calorific value, freezing temperature and / or heat capacity. The aircraft 1 may be refueled by connecting a refueling source 60, such as that provided by an airport fuel truck, a fixed fuel storage, or a permanent pipeline, to a connection port 62 of the fuel line of the aircraft. aircraft, via a fuel line 61. A desired quantity of fuel can be transferred from the fuel source 60 to one or more tanks 53, 55 of the aircraft 1. In the present example in which different tanks 53, 55 must be filled with different fuels, multiple ports 62 for connecting the fuel line are provided. The refueling source 60 therefore presents in this example two fuels, having different fuel characteristics, in separate containers 60a, 60b (for example a fuel by default and not by default as described below) in separate tanks. In other examples, valves may be used to direct fuel appropriately if it is received from a single connection port. Fuel characteristics can be obtained by either: (ÿ) the physical and / or chemical detection of one or more characteristics of a fuel, either during operation of the aircraft (for example on the wing), or when the aircraft! is being refueled; (iii) receiving data, for example from input provided at a user interface, or data transmitted to the aircraft 1. Fuel characteristics can be detected in different ways, both direct (e.g. from sensor data corresponding to the fuel characteristic in question) and indirect (e.g. by inference or calculation to from other characteristics or measurements). The characteristics can be determined as relative values ​​compared to another fuel, or as absolute values. For example, one or more of the following detection methods may be used: * The content of aromatic compounds or cycloparaffin in the fuel can be determined based on measurements of the swelling of a sensor component made from a sealing material such as a nitrile sealing material. “Trace substances or species, either naturally present in the fuel or added to act as a tracer, can be used to determine fuel characteristics such as the percentage of sustainable aviation fuel in the fuel. the fuel or whether the fuel is kerosene or not. * Measurements of the vibrational mode of a piezoelectric crystal exposed to fuel can be used as a basis for determining various fuel characteristics including fuel aromatic content, fuel oxygen content, and thermal stability or level fuel scaling (for example by measuring the accumulation of surface deposits on the piezoelectric crystal which will lead to a change in vibration mode). “Various fuel characteristics may be determined by collecting performance parameters of the gas turbine engine 10 during a first period of operation (e.g. during takeoff) and then during a second period of operation (e.g. during cruise) comparing these parameters collected at expected values ​​if fuel with known properties is used. Different fuel characteristics including aromatic hydrocarbon content of the fuel can be determined based on sensor measurements of the presence, absence or degree of contrail formation by the gas turbine 10 during operation . * Fuel characteristics including aromatic hydrocarbon content can be determined based on UV-Vis spectroscopy measurement performed on the fuel. Various fuel characteristics including sulfur content, naphthalene content, aromatic hydrogen content and hydrogen to carbon ratio can be determined by measuring substances present in the exhaust gases emitted by the gas turbine engine 10 during its use. * The calorific value of the fuel can be determined during operation of the aircraft 1 based on measurements taken when the fuel is burned, for example using fuel flow and shaft speed or temperature change on the combustion chamber assembly 16. “Various fuel characteristics may be determined by making an operational change designed to affect the operation of the gas turbine engine 10, detecting a response to the operational change; and determining the fuel characteristic(s) of the fuel based on the response to the operational change. “Various fuel characteristics may be determined with respect to the fuel characteristics of a first fuel by changing a fuel introduced into the gas turbine engine 10 from the first fuel to a second fuel, and determining the characteristic(s) of fuel of the second fuel as a function of a change in a relationship between T30 and T40 or T41 (the relationship being indicative of the increase in temperature across the entire combustion chamber 16). The characteristics can be determined as relative values ​​compared to the first fuel, or as absolute values ​​by reference to known values ​​for the first fuel. As used herein and noted below, T30, T40 and T41, and any other numbered pressures and temperatures, are defined using the station numbering listed in SAE AS755, specifically: * T30 = high pressure compressor (HPC) outlet temperature; “T40 = combustion outlet temperatures; And * T41 = high pressure rotor inlet temperature (HPT). Combusti system [Fig.5] schematically illustrates additional details of the combustion equipment 16 (which may be simply referred to as "the combustion chamber") of the gas turbine engine 10. In this example, the combustion equipment is part of a staged combustion chamber system 64 into which fuel is injected via pilot and main fuel injectors. Fuel is supplied to the fuel injectors by means of a fuel distribution regulator 306 under the control of an electronic engine controller (EEC) 42. The fuel is distributed to the fuel distribution regulator by fuel pumps 308a, 308b. In the present example, fuel is supplied to the fuel delivery regulator 306 from two different fuel sources (the first fuel source 302 and the second fuel source 304) as described above. Fuel from each of the first and second fuel sources 302, 304 is supplied by a respective first fuel pump 308a and second fuel pump 308b. Each of these fuel pumps may be mechanically driven by an accessory gearbox. In alternative configurations, for example in a more electric engine (MEE) configuration, the fuel pump may be electric. The professional will understand that any known suitable configuration of a fuel pump or combination of fuel pumps can be used to distribute fuel from the fuel tanks 53, 55 to the combustion chamber 16. The high pressure fuel is distributed by the fuel delivery regulator 306 into a pilot manifold 309 and a main manifold 310. In the present example, the staged combustion system includes a plurality of fuel nozzles 311 which are designed as duplex fuel injectors (also known as internally staged). In the present example, 16 circumferentially spaced fuel nozzles are arranged around an annular combustion chamber 312. As will be understood, more or fewer fuel injectors may be provided depending on size and design requirements. of the physics engine. Additionally, different combustion chamber configurations can be used, e.g. shaped like a cannula, tube, etc. Fuel from pilot manifold 309 and main manifold 310 is distributed to all fuel nozzles 311 for injection into tubular combustion chamber 312. In the present example, a central pilot fuel injector 313 produces a jet of pilot fuel in a primary zone of the tubular combustion chamber 312, while a concentric main fuel injector 314 produces a main jet of fuel. Fuel delivery through the pilot and main fuel injectors 313, 314 is staged, so at low power levels (and therefore low air mass flows), fuel is distributed by the central pilot fuel injector 313 at a rich fuel-air ratio (i.e., at an equivalence ratio greater than unity) for improved flame stability. In the present example, as power and mass flow increase, a staging point (SP) at which fuel is delivered through all or part of the main fuel injectors of the fuel nozzles is reached, supplementing the fuel flow pilot fuel injectors. The main fuel injectors 314 are designed to inject fuel at a lean fuel-to-air ratio (i.e., at an equivalence ratio less than unity). At this point, the airflow is such that the equivalence ratio immediately downstream of the pilot fuel injectors is also fuel lean. In this example, at higher power levels, fuel is injected through all main fuel injectors. The staged combustion system is therefore characterized by a “pilot injection only” operating zone and a “pilot and main injection” operating zone. The two regions are separated by the staging point. In “pilot injection only” operation, any flow of fuel supplied within the combustion chamber is supplied only by one or more injectors. fuel pilots. In “pilot and main injection” operation, fuel is supplied to the combustion chamber by the pilot fuel injector(s) and the main fuel injector(s), or only the main fuel injectors. The fuel delivery regulator 306 is therefore designed to distribute fuel to the main fuel injectors, or to both the main and pilot fuel injectors, during pilot and main injection operation. The fuel delivery regulator may be configured to distribute fuel to both the pilot and main fuel injectors during pilot and main injection operation rather than turning off the pilot fuel injectors during pilot and main injection operation. rapid emergency to keep the combustion chamber lit. During pilot injection only operation, the fuel delivery regulator supplies fuel only to the pilot fuel injectors. For a particular operating condition such as altitude or ambient air pressure, the staging point SP is generally defined by a threshold value of an engine operating parameter representative of the engine power setting, such as T30 (total temperature at the compressor outlet) or calculated T40 (total temperature at the combustion chamber outlet) or the combustion chamber air-fuel ratio (FAR). It will be understood that different threshold values ​​may apply for different operating conditions. Those skilled in the art are familiar with such operation of staged combustion systems with the aim of influencing lean combustion at high powers while respecting the limits of flame stability at lower powers. Additionally, it also recognizes other staged combustion configurations, for example those with separate pilot and main fuel injectors (as opposed to duplex or internal staged), which may be arranged in parallel (radially separated, aligned axially) or in series (axially separated, radially aligned). It will be understood that the principles described herein can be applied to any staged combustion system including main and pilot fuel injectors. The balance of fuel injection by the pilot fuel injectors 313 and the main fuel injectors 314 is controlled by the electronic engine governor 42, which provides control signals to the fuel delivery regulator 306 reporting the total fuel which is to be injected in the form of a fuel flow rate (WF) and the ratio of the pilot fuel injector fuel flow to the main fuel injector fuel flow (staging ratio). The fuel delivery regulator 306 is configured to use these control signals to distribute the requested fuel flow based on the overall fuel flow. requested fuel and the ratio between the pilot fuel flow and the main fuel flow. In alternative examples, the electronic engine governor 42 may instead be configured to provide control signals to the fuel delivery regulator 306 that report the pilot fuel flow (Wrpiæ:) and the main fuel flow (Wrman). . It goes without saying that the information transmitted is equivalent. State-of-the-art systems can simply provide fuel from a single source to a fuel delivery regulator such that fuel with the same fuel characteristics (i.e., same com - fuel position) is supplied to both the pilot fuel injectors and the main fuel injectors at all operating points of the combustion chamber. In the present application however, the fuel delivery regulator 306 is configured to selectively supply fuel to the main and pilot manifold from two different fuel sources as will be described later. Combusti ill L plus d' In state-of-the-art staged combustion systems, there is access to fuel only from a single source, which is used in both pilot-only and pilot-injection operating ranges. pilot and main. Fuel having the same characteristics is therefore supplied to the combustion chamber fuel injectors by prior art fuel delivery regulators independently of the fuel flow rate. The inventors have determined that it may be advantageous to supply fuel from different fuel sources having different characteristics to the fuel delivery regulator, and to distribute fuel to the pilot and main fuel injectors such that fuel of a different characteristic is supplied in different ranges of the combustion chamber operation. Referring again to the staged combustion system shown in [Fig.5], the fuel distribution regulator 306 is therefore designed to distribute fuel to the pilot fuel injectors 313 during at least part of the operating range. (i.e., a mode of operation) with pilot injection only having a fuel characteristic different from that of the fuel supplied to the pilot fuel injectors 313 and / or main fuel injectors 314 for at least part of the range pilot and main injection operation. The different fuel characteristic may be provided using fuel from first fuel source 302, fuel from second fuel source 304, or a mixture thereof. An advantage of this more flexible fuel supply to the fuel chamber combustion 16 is that fuel which has favorable combustion properties, but whose supply is limited, can be used when these favorable properties will have a greater effect. nvPM emission from staged combustion [Fig.6] illustrates a relationship between fuel flow W; and the soot particle emission rate (represented as "non-volatile particulate matter (nvPM) number" and referred to as "nvPM#" herein) for a typical lean burn combustion chamber (e.g. a combustion chamber system staged combustion). Any reference here to soot may also apply to other types of nvPM. [Fig.6] shows a first region, to the left of the staging point SP (at a fuel flow rate lower than that of the staging point), in which only the pilot fuel injectors of the fuel chamber combustion are operational, and in which nvPM# rises rapidly with increasing fuel flow Wz. Operation in this first region will be referred to as operation in the “pilot injection only” operating range. [Fig.6] shows a second region, to the right of the stage point SP (at a higher fuel flow than at the stage point), in which the pilot fuel injectors and the main fuel injectors operate , and in which nvPM# is much lower and rises only slowly (if at all) with an increase in W,. Operation in this second region will be called operation in the “pilot and main injection” operating range. Although a non-zero value of nvPM# in the "pilot and main injection" operating range is shown in [Fig.6], it will be borne in mind that in certain lean mixture combustion chambers nvPM# in the “pilot and main injection” range could be significantly zero. It should be understood that the shape of the curve shown in [Fig.6] is only an example given for illustration purposes. Generally the nvPM# in pilot injection only operation is significantly higher than that in pilot and main injection operation, and there is a clear boundary between the two operating ranges. Some examples shown below take further advantage of the positive gradient of the curve shown in pilot injection only operation, but once again the precise shape of the curve shown in [Fig.6] is only one example of the dependence between W; and nvPM#. The position of the stage point SP may be chosen such that the fuel flow during most or all cruise conditions falls to the right of the stage point SP i.e. in a region operating conditions with pilot and main injection corresponding to very low values ​​(or perhaps sens- obviously zero) of nvPM#. Conversely, operation at many other points in a flight, such as taxiing, approach, descent, may correspond to "pilot injection only" operation in which nvPM (soot) emissions are comparatively high. It should be understood that the staging point SP can be at different values ​​of W, for different flight conditions such as altitude. For example, at cruise regime where the air density is low and therefore the mass flow rate of the working fluid through the gas turbine core is also low, the staging point SP may correspond to a significantly lower fuel flow rate. to what it would be on the ground where the density of the ambient air is much higher. The staging point SP applying to a particular flight condition may be defined as an absolute value of fuel flow, or a predetermined percentage of any W£ which applies to the flight condition. Other definitions of staging point may also be used as will be understood by those skilled in the art. The present application refers to the number of nvPM (or nvPM#) and its dependence on W;, as illustrated in [Fig.6]. However, it will be understood that a corresponding relationship for the mass of nvPM and its dependence on Wr can also be considered, and that the apparatus and methods defined herein can be adapted in this sense. nvPM emissions control The level of nvPM emissions from a gas turbine engine depends on the characteristics of the fuel used. For example, some aircraft fuels may be characterized by a lower proportion of certain constituents known to cause soot emissions compared to a typical fossil kerosene, and thus produce a lower amount of soot at the same chamber fuel flow rate. combustion. While nvPM emission from a gas turbine engine can be effectively reduced by using fuel associated with a low nvPM production index, this may not always be possible. For example, the availability of such fuels may not be sufficient for use for the entire duration of a flight. There may also be other technical, regulatory or financial constraints on the use of such low nvPM fuels in large fuel volumes or in significant concentrations within a blended fuel. The inventors have determined that by favoring the use of fuel associated with a relatively low level of nvPM emission (for example compared to fossil kerosene) during the pilot injection only operating range of the combustion chamber compared to the injection operating range pilot and main, nvPM emissions can still be significantly reduced, while not requiring the use of a large quantity of the low nvPM fuel over the entire operating range of the combustion chamber. In the present example therefore, the first fuel maintained in the first fuel source 302 is associated with an nvPM production level which is lower than that of the second fuel maintained in the second fuel source 304 (e.g. when it is used under corresponding conditions). Fuel delivered to the pilot fuel injectors during at least a portion of the pilot injection only operating range is also associated with an nvPM production level that is lower than that of the fuel delivered to the pilot fuel injectors and / or main during at least part of the pilot and main injection operating range. As indicated below, fuel may be supplied to the pilot and main fuel injectors 313, 314 exclusively from one of the two available fuel sources 302, 304, or as a mixture of fuel from the first and second fuel sources 302, 304. In some examples, the fuel system may be configured such that the main fuel injectors 314 may be supplied from either the first fuel source 302 or the second fuel source 304, and also such that the Pilot fuel injectors 313 may be powered from either the first fuel source 302 or the second fuel source 304. Although some of the examples presented below do not include such a flexible fuel system, ensure that a Any fuel tank that can supply fuel to any fuel injector of any engine has safety advantages. In some of the examples described here, an individual fuel injector 313, 314 is supplied with either fuel only from the first fuel source 302, or fuel from the second fuel source 304, i.e. -say the fuel system switches between the two fuel sources. In other examples, the fuel system also includes the equipment necessary to achieve fuel mixing from the two onboard fuel sources (for example, high soot fuels and low soot fuels), the mixture ratio being varied according to various decision-making criteria to produce a fuel composition whose characteristics are equal to or equal to those of the low soot producing fuel composition (100:0 mixture ratio) to the fuel composition producing a high level of soot (mix ratio 0:100) or somewhere between both (mixing ratio x:100-x where OÔ < x < 100). Various examples may involve the ability to switch between predetermined fuel compositions and / or produce a mixed fuel composition for only the pilot fuel injectors 313. The main fuel injectors 314 may in such examples be supplied at any time with fuel from one of the fuel sources (e.g. fuel associated with high nvPM level). However, for added flexibility, some examples may allow the main fuel injectors 314 to be switched to fuel from the other source (e.g. low nvPM fuel composition) and / or a fuel composition mixed during certain abnormal operating periods, for example in the event of loss of fuel associated with high nvPM level due to, for example, a fuel leak. In some examples, the fuel characteristic by which the fuel from the first fuel source 302 differs from the fuel from the second fuel source 304 may be a percentage of sustainable aviation fuel (SAF) present in the respective fuel. Fuel distributed to pilot fuel injectors 313 during at least a portion of the pilot injection only operating range would similarly have a different SAF percentage compared to fuel distributed to pilot fuel injectors 313 and / or main fuel injectors 314 during at least a portion of the pilot injection-only operating range. minus part of the pilot and main injection operating range. Compared to fossil kerosene, SAF offers significantly lower soot emissions, or more generally nvPM. When SAF is used as part of a fuel composition blended with fossil kerosene, roughly the higher the percentage of SAF in the mixture (and therefore the lower the percentage of fossil kerosene), the greater the reduction in nvPM emissions are high, at least in certain SAF percentage ranges. This effect is illustrated in [Fig.7] which shows nvPM# emissions dependencies at a fuel flow rate Wr for a default fuel composition (solid line) and for three other fuel compositions A , B and C (marked) for a particular flight condition defined for example by an altitude and a forward speed. The default fuel composition and fuel compositions A, B and C are characterized by progressively higher SAF percentages and correspondingly lower fossil kerosene percentages, respectively. The default fuel composition may be 100% fossil kerosene, or a blend comprising predominantly fossil kerosene with a small percentage of SAF, as may be available by default at some airports. In [Fig.7], the emission reduction factor of nvPM# relative to the default fuel composition achieved by each blended fuel composition is shown as unchanged with W; during operation with pilot injection only and also without change with W, during operation with pilot and main injection, although the reduction factors in these two operating regions are shown to be different from each other. It will be understood that other dependencies on Wz may be observed in various implementations and are also contemplated by the present application. [Fig.8] shows such a further dependence on W;, characterized in pilot injection only operation by an increase in the reduction factor of nvPM# for each blended fuel composition as W, is reduced from the staging point SP towards a lower level of fuel flow. Other variations may be possible; in particular the dependence on W; of the nvPM# reduction factor cannot be the same for each blended fuel composition, and / or need not be a function of W; monotonically increasing or decreasing. Although the following examples in the present application are based on the shape illustrated in [Fig.7], it will be understood that more general shapes are also contemplated. The inventors have determined that when a gas turbine engine 10 operates in the pilot and main injection operating range soot emissions are inherently low and substitution of the default fuel composition (e.g. fossil kerosene) with SAF (or a high percentage SAF blend) or other low soot fuel will produce a small reduction in soot emissions. Conversely when the engine 10 operates at a fuel flow which is below the staging point SP, but nevertheless close thereto, soot emissions can be significantly reduced by using a fuel composition which includes a higher percentage of low soot producing fuel (e.g. SAF) and a lower percentage of fossil kerosene (or other high soot producing fuel) compared to a default fuel composition. By using fuel with a different characteristic in different combustion chamber operating ranges the percentage of SAF within the fuel composition burned in the pilot injection only mode can be increased (e.g., to the maximum), while reducing (e.g. to a minimum) the percentage of SAF within the fuel composition burned in the pilot and main injection mode. This may allow the nvPM reducing effect of the available SAF to have a greater impact compared to using a constant percentage composition of SAF throughout the range. complete operation of the combustion chamber. In various other examples one or more other characteristics of the fuel may be modified alternatively to or in addition to the SAF percentage. Changes in other characteristics between the fuels of the first and second fuel sources may additionally or alternatively be associated with different levels of the nvPM#. For example, the first fuel source 302 may be characterized by fuel with a lower proportion of certain constituents that cause soot or other nvPM emissions compared to that of the second fuel source 304. In some examples, the fuel of the first fuel source 302 can be characterized by a lower aromatic hydrocarbon content, or more particularly a lower naphthalene content compared to the second fuel source 304. The lower content of such compounds producing Soot may be associated with SAF when comparing to fossil kerosene. This may not always be the case, however. Some TAS may be associated with a higher level of soot production compared to others (e.g. if they have a higher amount of added synthetic aromatic content), or may be associated with a higher level higher soot production compared to a fossil fuel such as fossil kerosene from which the content of aromatic compounds has been eliminated in order to leave mainly a paraffinic content. Various examples of fuel delivery to the main and pilot fuel injectors 313, 314 from either of the first and second fuel sources 302, 304 (exclusively, at least under normal operating conditions) or d A mixture of these at various points during the operating ranges of the combustion chamber are described below. Figures 9, 10 and 11 [Fig.9] illustrates an example in which the fuel delivery regulator 306 is configured to distribute fuel from the first fuel source 302 to the pilot fuel injectors 313 during operation in both operating ranges pilot-only and pilot-and-main-injection, and fuel from the second fuel source 304 to the main fuel injectors 314 during operation in the pilot-and-main-injection operating range. In this example, the fuel delivery regulator 306 includes a pilot regulator 306a in fluid communication with the first fuel source 302 via the first fuel pump 308a. The fuel delivery regulator 306 further includes a separate main regulator 306b in fluid communication with the second fuel source 304 via the second fuel pump 308b. THE Pilot regulator 306a is configured to distribute fuel to the pilot manifold 309 and pilot fuel injectors 313. The main regulator 306b is configured to distribute fuel to the main manifold 310 and the main fuel injectors 314. The fuel delivery regulator 306 therefore comprises two separate flow paths through which fuel from each fuel source 302, 304 is supplied to the combustion chamber 16. The speed of fuel flow through each of the pilot regulator 306a and that of the main regulator 306b can be regulated independently of each other using control signals received by the fuel distribution regulator 306 from the EEC 42. The pilot fuel injectors 313 are therefore at all times supplied with fuel from the first fuel source 302 and the main fuel injectors 314 are at all times supplied with fuel from the second fuel source 304. The fuel flow to the main fuel injectors 314 may be substantially zero in the range pilot injection only operation. This means that the composition of fuel passing through an individual fuel injector does not change throughout the flight and is predetermined before flight (at least during normal operating conditions). In the example of [Fig.9], the first fuel source 302 contains a fuel which is associated with low nvPM production, for example a fuel having a relatively high SAF content (for example a fuel rich in SAF ). The second fuel source 304 contains a fuel associated with high nvPM production, for example a fuel having a relatively low SAF content (i.e. lower than that of the first fuel), for example a fuel low in SAF. The term "high SAF" may be used here to indicate a fuel with a higher SAF content compared to a "low SAF" fuel. The term "high SAF" may include fuel that is 100% SAF. “Low SAF” fuel may include fuel that is 0% SAF, for example fossil kerosene. In some examples, the high SAF fuel may include up to 50% SAF, and the low SAF fuel substantially zero% SAF; for example, it may be a standard fossil kerosene fuel. In the example of [Fig.9], the composition of fuel rich in SAF can be determined by one or more of the following factors: a) the quantity of SAF available for, or allocated to, a proposed flight; b) the total fuel requirement for the pilot fuel injectors for the entire flight (calculated according to methods known to those skilled in the art); And c) any limits on the maximum allowable percentage of SAF, e.g. certification limits, or e.g. technical limits related to the specific aircraft and / or the pilot fuel injectors themselves, or the Maximum mixture percentage at which SAF is available at the fuel loading point. The desired percentage of SAF in the composition of the SAF-rich fuel can be calculated by 100% times factor a), divided by factor b), subject to a maximum allowable value which is the minimum of the various potential limits identified in factor c). An adjustment may be necessary to take into account the different volumetric energy densities of SAF and fossil kerosene, according to methods known to those skilled in the art. In this example, the low-SAF fuel composition may have a default SAF percentage of zero (or the minimum possible given the default fuel supply at the airport where fuel refueling occurs). aircraft), but any SAF assigned to the proposed flight that was not incorporated into the SAF-rich fuel composition for the pilot fuel injectors 313 will be used as part of the low-SAF fuel composition for the pilot fuel injectors 313. main fuels 314. The percentage of SAF thus obtained in the low-SAF fuel composition will be capped by any certification limits or by any technical limits linked to the specific aircraft and / or main fuel injectors, or the maximum mixture percentage in which SAF is available at the fuel loading point. Operation of the combustion chamber in [Fig.9] may not be possible for all flights because the size of the fuel tanks available for the high-SAF fuel composition and for the low-SAF fuel composition are not possible. cannot accommodate the fuel volume requirements for pilot fuel injectors 313 and / or main fuel injectors 314. This may be the case for long-haul flights in which all fuel tanks must be fully filled before departure and for which the respective fuel volume requirements for different fuel injector types may not exactly match the source volumes. [Fig.10] illustrates the dependence of the number of nvPM on the fuel flow W; for the fuel regulator arrangement of [Fig.9] (dashed line) compared to the corresponding dependence for a default fuel composition such as fossil kerosene (solid line). In this illustration, it is assumed that the SAF-rich fuel composition corresponds to the fuel composition À in [Fig.7]. As can be seen in [Fig.10], an advantageous reduction in the number of nvPM is provided within the pilot injection only operating range, whereas in the pilot and main injection operating range it does not. there is little or no change in the number of nvPM. In this example, a limited amount of available SAF has therefore been more effectively targeted to a part of the operating range (i.e., pilot injection only) where it can provide the greatest benefit in terms of nvPM reduction. [Fig.11] illustrates a method 4000 of operating a gas turbine engine which can be implemented using the staged combustion chamber system of [Fig.9]. The method 4000 includes regulating 4002 the delivery of fuel to the pilot and main fuel injectors 313, 314 from the first fuel source 302 and the second fuel source 304. As noted above, regulating the fuel delivery generally includes providing fuel to the pilot fuel injectors 313 during at least a portion of the pilot injection only operating range having a fuel characteristic different from that of the fuel delivered to the pilot fuel injectors 313 and / or to the pilot fuel injectors 313 and / or to the pilot fuel injectors 313. main fuel injectors 314 for at least part of the pilot and main injection operating range. In the example of [Fig.11], the regulation 4002 of the fuel delivery includes the distribution 4004 of fuel from the first fuel source 302 to the pilot fuel injectors 313 during operation in both the pilot injection only and pilot and main injection operating ranges, and fuel from the second fuel source 304 to the main fuel injectors 314 during operation in the pilot and main injection operating range. Any of the other features described above in connection with [Fig.9] may be incorporated into the method of [Fig.11], although they are not repeated here. Figures 12, 13 and 14 In the example shown in Figures 9, 10 and 11, the advantageous reduction in the number of nvPM (compared to the default fuel composition) observed particularly in pilot injection only operation (i.e. to the left of point SP) is made possible by prioritizing the SAF to the pilot fuel injectors, thereby allowing a higher percentage of SAF in the fuel composition supplied to these fuel injectors. However, in pilot and main injection operation (in which soot production is minimal even when using fossil kerosene) nevertheless an SAF-rich fuel composition is still supplied to the pilot fuel injectors. Even more efficient use of SAF-rich fuel can be achieved by providing additional flexibility of fuel delivery to the combustion chamber. [Fig.12] illustrates an example in which the fuel delivery regulator 306 is designed to switch the delivery of fuel to the pilot fuel injectors 313 between the first fuel source 302 and the second fuel source. fuel 304, In the example described the switchover occurs at the border of the operating range with pilot injection only (for example at the staging point SP). In other examples, there may be multiple tipping points including at the boundary and / or within the pilot injection only operating range. More generally therefore, the fuel delivery regulator 306 is designed to switch the delivery of fuel to the pilot fuel injectors 313 between the first and second fuel sources 302, 304 at one or more operating points within the operating range with pilot injection alone or at a border thereof. With reference to [Fig.12], the pilot regulator 306a comprises two remote regulators that can be controlled independently, a first regulator 315a in fluid communication with the first fuel source 302, and a second regulator 315b in fluid communication with the second fuel source 304. The first and second regulators 31Sa, 315b of the pilot regulator 306a can be controlled by signals received from the EEC 42. The first and second regulators 315a, 315b are in fluid communication with the pilot manifold 309 so that they can control the delivery of fuel to the pilot fuel injectors 313. The fuel delivery regulator 306 is configured to switch the delivery of fuel to the pilot fuel injectors 313 between fuel from the first fuel source 302 and fuel from the second fuel source 304 according to a mode signal indicating a change of the operating range of the staged combustion system. The mode signal can be obtained by the EEC 42 from the combustion system 64 (or combustion chamber 16) and a corresponding control signal sent to the regulators 315a, 315b so that switching can occur at the level of the staging point between the operating range with pilot injection only and the operating range with pilot and main injection. The fuel distribution regulator 306 is thus designed to switch the pilot fuel injectors 313 between the first fuel and the second fuel (or vice versa) each time the staging point SP is crossed. The EEC can be configured to receive from the combustion system 64 (or the combustion chamber 16) a signal indicating its mode of operation (pilot injection only or pilot and main injection). Alternatively, the EEC 42 may command the combustion system 64 to switch from one operating mode to another. In this example, the mode signal may already be available at EEC 42, and may be used to switch fuel delivery by fuel delivery controller 306. Switching fuel delivery according to a mode signal mode indicating a change of mode of the combustion chamber 16 can guarantee a acceptable level of synchronization between the operating mode of the combustion device and the fuel composition of the pilot fuel injectors. This can help ensure rapid switching between fuel sources when the staging point is crossed. In the example described herein, the pilot regulator 306a is configured to switch between supplying fuel exclusively from the first fuel source 302 and exclusively from the second fuel source 304. The first and second fuel regulators 315a, 315b can therefore be arranged to switch between operation where: i) the first is completely closed (so that no fuel from the corresponding source is supplied to the combustion chamber 16), and the second is used to regulate the flow speed of the other fuel source; and ii) the second is completely closed (so that no fuel from the corresponding source is supplied to the combustion chamber 16), and the first is used to regulate the flow speed of the other fuel source . In other examples, any other suitable arrangement of pilot fuel regulator 306a may be provided to allow switching between fuel sources. The fuel delivery regulator 306 illustrated in [Fig. 12] may be designed to distribute fuel such that when the engine 10 is operating in the pilot injection only mode, the pilot fuel injectors 313 are supplied with a composition fuel rich in SAF, and when the engine is operating in the pilot and main injection mode, the pilot fuel injectors 313 and the main fuel injectors 314 are all supplied with a fuel composition low in SAF. In this example, the composition of SAF-rich fuel will be determined by the following factors: a) the quantity of SAF available for, or allocated to, a proposed flight; b) the total fuel requirement for the pilot fuel injectors during pilot injection only operation for the entire flight (calculated according to methods known to those skilled in the art); And c) any limits on the maximum allowable percentage of SAF, e.g. certification limits, or e.g. technical limits related to the specific aircraft and / or the pilot fuel injectors themselves, or the maximum mixture percentage at which SAF is available at the fuel loading point. The desired percentage of SAF in the composition of the SAF-rich fuel is then calculated simply by 100% times factor a) divided by factor b), subject to a maximum allowable value which is the minimum of the various limits po- potential identified in factor c). An adjustment may be necessary to take into account the different volumetric energy densities of SAF and fossil kerosene, according to methods known to those skilled in the art. The composition of fuel low in SAF can be determined using the same method and the same constraints as for the example described in conjunction with Figures 9, 10 and 11, further constrained by any practical limits linked to the fuel injectors. fuel pilots 313 if necessary. [Fig.13] illustrates the dependence of the number of nvPM on the fuel flow W, for the fuel delivery regulator of [Fig.12] (dashed line) compared to the corresponding dependence for a fuel composition by defect such as fossil kerosene (solid line). In this illustration it is assumed that the SAF-rich fuel composition corresponds to fuel composition B in [Fig.7] and that the SAF-poor fuel composition corresponds to the default fuel composition. As can be seen in [Fig.13], for a fixed amount of SAF assigned to an individual flight, the fuel delivery regulator 316 of [Fig.12] allows the SAF-rich fuel composition to have a higher SAF content percentage compared to the example in [Fig.9], due to a more restricted use of the SAF-rich fuel composition compared to pilot injection only operation. Therefore, the SAF is more effectively prioritized toward an operating region in which the beneficial nyPM reduction is greater. As with the example in Figures 9, 10 and 11, fuel tank capacities may prevent operation according to the example in [Fig.12] for some flights, as it may be difficult to match the volumes of the different fuel tanks with the required volumes of high-SAF and low-SAF fuel compositions. [Fig.14] illustrates a method of operating a gas turbine engine 10 which can be implemented using the staged combustion chamber system of [Fig.12]. Process steps common to the process of [Fig.11] are marked in this way. In this example, the regulation 4002 of the fuel distribution includes the switching 4006 of the distribution of fuel to the pilot fuel injectors 313 between the first fuel source 302 and the second fuel source 304 at the staging point SP. Any of the features described above in connection with [Fig.12] may be incorporated into the method of [Fig.14], although they are not repeated here. Figures 15 and 16 The fuel distribution regulator of [Fig.12] can in certain examples Be designed to switch fuel delivery to the pilot fuel injectors 313 again between the first fuel source 302 and the second fuel source 304 at a threshold point TP within the pilot injection only operating range. Fuel from the second fuel source 304 may be delivered to the pilot fuel injectors 313 at fuel flow rates below the threshold point, with fuel from the first fuel source 302 delivered to the pilot fuel injectors 313 at flow rates of fuel between the threshold and the boundary of the pilot injection only operating range (staging point), and fuel from the second fuel source 304 can be delivered to the pilot fuel injectors 313 at fuel flow rates above the border. The main fuel injectors 314 may be supplied with fuel from the second fuel source 304 at any time. Similar to the examples described previously, the first fuel contained in the first fuel source 302 can be a fuel associated with a low level of nvPM such as a fuel rich in SAF, while the second fuel contained in the second source of 304 fuel may be associated with a high level of nvPM such as low SAF fuel. When the combustion chamber 16 operates in pilot injection only mode near the staging point SP (as determined by the threshold position), the pilot fuel injectors 313 are supplied with the SAF-rich fuel composition. In all other cases the pilot fuel injectors 313 are supplied with a fuel composition low in SAF. This example may provide even more efficient use of SAF and may be further advantageous over examples where a switchover occurs only at the SP staging point in cases where the percentage of SAF in the high-carbon fuel composition SAF in the second example is limited by SAF availability rather than engine certification limits. The present example in which two tipping points are provided can also overcome difficulties with potential mismatches between the capacities of individual fuel tanks and the required volumes of high-SAF and low-SAF fuel compositions. The position of the switching threshold TP can be determined according to one or more different factors so that the switching occurs "near" the staging point SP. For example, the threshold can be: a) a first threshold fuel flow rate at which the production of nvPM by the gas turbine engine 10 passes a threshold quantity of the nvPM produced by the gas turbine engine during operation in which the pilot fuel injectors 313 receive fuel having the second fuel characteristic. For example, the Threshold can be defined as operating in pilot injection only mode at a fuel flow rate which, for a default fuel composition such as fossil kerosene, would correspond to soot emissions exceeding a threshold. Soot emissions or nvPM produced by the gas turbine engine can be defined either as the number of soot particles emitted per unit mass of fuel (i.e., a number emission index), or as as the number of soot particles emitted per unit time (i.e., also taking into account fuel flow), or as the number of soot particles emitted per unit flight distance (this is i.e. also taking into account the speed of the aircraft). b) a second threshold defined as a predefined threshold fuel flow less than the fuel flow at the staging point SP (as defined for current flight conditions such as current altitude). The predefined threshold can be either a percentage of the fuel flow at the staging point, or an absolute flow value lower than that at the staging point. The first threshold and / or the second threshold may be defined with reference to the quantity of SAF available for a proposed flight, under the assumption that no more than the available SAF will be incorporated into the SAF-rich fuel composition, subject to previously identified constraints, and taking into account the fuel volume requirement for a proposed flight for the pilot fuel injectors during operation “near the SP point” according to a candidate value of the first threshold and / or the second threshold, the fuel volume requirement being determined using methods well known to those skilled in the art. The second threshold may be defined with reference to the fuel tank volume available on board the aircraft 1 for the SAF-rich fuel composition, again taking into account the fuel volume requirement for a proposed flight for the pilot fuel injectors when operating "near point SP", the fuel volume requirement being determined using methods well known to those skilled in the art. For certain flights, in particular long-haul flights for which all available fuel capacity must be used, the size of the available fuel tanks 53, 55 may limit the options available regarding the second threshold, to ensure that the capacity of the fuel tank used for SAF-rich fuel composition is completely used. This would impose corresponding constraints on the percentage of SAF in the composition of SAF-rich fuel In cases where the first threshold and / or the second threshold have not been defined with reference to the available quantity of SAF, the composition of SAF-rich fuel may be determined by one or more of the following factors: a) the quantity of SAF available for, or allocated to, a proposed flight; b) the total fuel requirement for the pilot fuel injectors during pilot-only operation near the staging point for the entire flight (calculated according to methods known to those skilled in the art and taking into account the volumes of the relevant fuel tanks as described above); And c) any limits on the maximum allowable percentage of SAF, e.g. certification limits, or e.g. practical limits related to the specific aircraft and / or the pilot fuel injectors themselves, or the maximum mixture percentage at which the SAF is available at the fuel loading point. The percentage of SAF in the composition of the SAF-rich fuel is then calculated simply as 100% times factor a) divided by factor b), subject to a maximum allowable value which is the minimum of the various potential limits identified in the factor c). An adjustment may be necessary to take into account the different volumetric energy densities of SAF and fossil kerosene, according to methods known to those skilled in the art. The composition of low-SAF fuel can be determined using the same method and the same constraints as described for example in connection with [Fig.12]. [Fig.15] illustrates the dependence of the number of nvPM on the fuel flow W£ (dashed line) for an example in which two switching points are planned, one at a threshold TP within the pilot injection only operating range and a second TP2 at staging point SP, compared to the corresponding dependence for a default fuel composition such as fossil kerosene (solid line). In this illustration, it is assumed that the SAF-rich fuel composition corresponds to fuel composition C in [Fig.7] and the low-SAF fuel composition corresponds to the default fuel composition. The tipping points TP, TP2 are visible by the rapid changes in the number of nvPM at the corresponding fuel flow. [Fig.16] illustrates a method 4000 of operating a gas turbine engine 10 which can be implemented using the staged combustion chamber system of [Fig.12] in which two switching points TP , TP2 are provided. Process steps common to the process of [Fig.14] are marked in this sense. In this example, the regulation 4002 of the fuel distribution further comprises a step 4008 of switching the fuel distribution to the pilot fuel injectors between the first fuel source 302 and the second fuel source 304 at a threshold point in the operating range with pilot injection only in addition to the 4006 tilting at the staging point. Any of the features described above in connection with [Fig.15] can be incorporated into the method of [Fig.16]. Figures 17, 18 and 19 In the examples described above, the fuel supplied to the pilot and main fuel injectors 313, 314 is limited to the characteristics of the two predefined fuels (the first and second fuels) contained in the first and second fuel sources 302, 304. The inventors have determined that further benefits can be achieved by supplying fuel containing a mixture of fuel from the first and second fuel sources 302, 304 to the combustion chamber 16 to give greater flexibility of the fuel composition supplied to the fuel injectors 313, 314. [Fig.17] illustrates an example in which the fuel delivery regulator 306 includes a fuel mixer 318. The fuel mixer 318 is configured to receive a fuel supply from the first and second fuel sources 302, 304 and to output fuel from the first fuel source 302, fuel from the second fuel source 304, or a mixture thereof (e.g. a mixture ratio which may vary between 100% of the first fuel and 0% of the second fuel, up to 00% of the first fuel and 100% of the second fuel and any ratio thereof). The fuel mixer 318 is in fluid communication with the pilot manifold 309, and is configured to distribute fuel to the pilot fuel injectors 313. The fuel delivery regulator 306 includes a main regulator 306b connected to the second fuel source 304 and designed to feed the main manifold 310 and main fuel injectors 314 in a manner similar to other examples. In other examples, the mixer may be configured to supply both the main and pilot fuel injectors 313, 314. Using the fuel mixer 318 a fuel mixture may be delivered to the pilot fuel injectors 313 for at least part of the operation in the pilot injection only range. For other parts of pilot injection only operation, and during pilot and main injection operation, the pilot fuel injectors 313 may receive fuel from only one of the fuel sources through the fuel mixer 313. A mixture Fuel from both sources may also be supplied during the entire pilot injection only operating range and / or during the pilot and main injection operating range. Similar to the examples described previously, the first fuel contained in the first fuel source 302 can be a fuel associated with a low level of nvPM such as a fuel rich in SAF, while the second fuel contained in the second source of 304 fuel may be associated with a high level of nvPM emissions such as low SAF fuel. When the room of combustion 16 operates in a pilot injection mode only a fuel mixture may be supplied to the pilot fuel injectors 313 such that the supplied mixed fuel is associated with a lower nvPM level compared to the fuel supplied to the main fuel injectors and / or pilots 313, 314 during pilot and main injection operation where the nvPM level is intrinsically lower. This could enable even more efficient use of low nvPM producing fuels. The fuel supplied to the pilot fuel injectors 313 during at least a portion of pilot injection only operation may therefore contain more of the first fuel compared to the fuel supplied to the main fuel injectors 314 during pilot and main injection operation. In the example of [Fig.17], the fuel distribution regulator 306 allows fuels from the first and second sources 302, 304 to be mixed by the fuel mixer 318 at a desired mixing ratio and d 'be distributed to the pilot manifold 309. This contrasts with the example of Figures 12, 13 and 14, where a fuel mixture formed by mixing a quantity of fuel coming from the first source 302 and a quantity of fuel coming from the second source 304 cannot be provided. A number of advantages are associated with providing blended fuel in this way. For example, it is possible to avoid a hard limit or tipping point (e.g., the TP threshold point described above) in the pilot-only region, below which a low-SAF fuel composition is supplied to the pilot fuel injectors 313 and above which an SAF-rich fuel composition is supplied to the pilot fuel injectors 313. This may reduce the risk of a sudden increase in soot emissions as the fuel flow is reduced, which could otherwise result in unnecessarily high soot emissions during, for example, approach and / or final approach flight phases. Fuel blending may also be even more advantageous because a fuel blend further reducing nyvPM can be used close to the SP staging point (i.e., at fuel flow rates just below the staging point ) when the number of nvPM is at its maximum. This can provide greater overall control of nvPM emission, and use low nvPM fuels more efficiently. The percentage of SAF within the first fuel rich in SAF can be as high as possible, subject to any limits on the maximum admissible percentage of SAF, for example certification limits, or for example technical limits linked to the specific aircraft and / or to the pilot fuel injectors themselves, or the maximum mixture percentage at which the SAF is available at the point loading fuel. The percentage of SAF within the first SAF-rich fuel may also be constrained by the required fill factor of the fuel tank(s) used for the SAF-rich fuel composition, in conjunction with the amount of SAF assigned to the proposed flight. For example, on long-haul flights requiring a complete filling of all fuel tanks 53, 55, the volume of SAF-rich fuel composition may not be less than the capacity of the smallest individual fuel tank that can be used. understand the first fuel source. The composition of the fuel low in SAF, that is to say the second fuel, can be determined using the same method and the same constraints as for the examples described previously. In some examples, the fuel delivery regulator 306 may be designed to provide a constant ratio mixture of fuel from the first fuel source 302 and fuel from the second fuel source 304. This may provide the ability to provide the combustion chamber a fuel having different characteristics than the fuels available to power the aircraft 1. This can provide greater flexibility and better control of nvPM. For example, the mixture ratio may be determined and set for a particular flight once the quantity of fuel in the first and second fuel sources 302, 304 assigned to the flight is known. In other examples, the fuel mixer 318 is configured to deliver a fuel mixture to the pilot fuel injectors 313 having a variable mixing ratio of fuel from the first fuel source 302 and fuel from the second fuel source. fuel 304. The mixture ratio can be varied in the pilot injection only operating range based on the fuel flow rate, or based on the fuel flow rate divided by the fuel flow rate at the staging point SP. In some examples, the fuel mixture may be varied such that the proportion of fuel coming from the first fuel source 302 compared to that coming from the second fuel source 304 is reduced with decreasing fuel flow in the range. pilot injection only operation. This can help reduce the amount of low nvPM producing fuel (e.g. SAF) as the flow rate decreases. In other examples, opposite dependencies of the first fuel content with a fuel flow rate may be anticipated. For example, in some cases the percentage reduction in nvPM due to the use of SAF may be greater at lower power settings (e.g. low fuel flow) than at higher power settings (e.g. example a high fuel flow). The fuel distribution regulator 306 of the example in [Fig.17] can be designed to: a) dispensing fuel from the second fuel source 304, for example, fuel producing high nvPM levels such as lean SAF fuel to both the pilot fuel injectors 313 and the main fuel injectors 314 at fuel flow rates above the staging point; b) delivering fuel to the pilot fuel injectors 313 from the first fuel source 302, for example, fuel producing low levels of nvPM such as SAF-rich fuel at and / or immediately below the flow rate of fuel at staging point SP; And c) delivering fuel to the pilot fuel injectors 313 having a mixture comprising progressively less of the first fuel from the first fuel source 302 and correspondingly more of the second fuel from the second fuel source 304 as the flow rate of overall fuel is reduced below the staging point. The ratio of first fuel to second fuel being supplied to pilot fuel injectors 313 during pilot injection only operation may be varied according to a fuel mixing schedule. The fuel mixture schedule may be determined by the EEC 42 and used to send control signals to the fuel delivery regulator 306 to control the mixture ratio. In some examples, the dependence of the proportion of fuel coming from the first fuel source 302 relative to that coming from the second fuel source 304 on the fuel flow rate can be determined based on a desired obtained level of nvPM at a particular fuel flow. For example, the mixture ratio may be determined such that the nvPM number for any specific fuel flow rate does not exceed a predetermined threshold. For a given flight condition (such as altitude and forward speed), a look-up table can be used to determine the mixture ratio (e.g. SAF percentage) needed to achieve a particular level of number of nvPM to a particular Wr. Given that we know the characteristic values ​​of the first and second fuels (i.e. the percentage of SAF within each of the two predetermined fuel compositions, rich in SAF rich and poor in SAF) the proportion of the composition of SAF-rich fuel to be delivered to the pilot fuel injectors can be determined. The proportion of low SAF fuel mixed with the high SAF fuel to produce an instantaneous fuel composition to be delivered to the pilot fuel injectors 313 can then also be determined in order to keep nvPM production within a limit. threshold throughout pilot injection only operation. The determination of the proportion of each fuel constituting the mixture can be made by the EEC 42 based on information from the look-up table, information on the characteristics of the fuel contained in the first and second fuel sources and current fuel flow. Once a mixture ratio is determined by the EEC 42 the mixer 318 can be controlled to do so by control signals sent from the EEC 42 to the fuel delivery regulator 306. In some examples, the EEC 42 can calculate the mixture ratio in real time in response to changes in the current flight condition or current atmospheric conditions. In another example, the fuel blend schedule (taking into account the variation with W; of the reduction factor of nvPM# (or mass of nvPM) for a given fuel composition compared to that for the default or lean composition in SAF) can be determined so as to reduce to a minimum (or keep within a predefined threshold) the total number (or mass) of nvPM emitted during a period of operation of the gas turbine engine 10 such as a cycle of landing and takeoff (LTO). Before a flight, using knowledge of the desired dependence of the nvPM number on Wz and flight conditions, the characteristics of each of the first and second fuels, and knowledge of the amount of fuel that will be used at each value from W; at each flight condition, the total quantity of each of the first and second fuels required for a proposed flight may be determined as described later. The aircraft can therefore be loaded with a suitable quantity of fuel before the flight. [Fig.18] illustrates an example of the dependence of the number of nvPM on the fuel flow W; for the example shown in [Fig.17] (dashed line) compared to the corresponding dependence for a default fuel composition such as fossil kerosene (solid line). In this example, the first fuel is a fuel composition rich in SAF which corresponds to the fuel composition C of [Fig.7] and the second fuel is a fuel composition poor in SAF which corresponds to the composition of default fuel in [Fig.7]. In the example of [Fig.18], the characteristic of the fuel supplied to the pilot fuel injectors 313 during operation with pilot injection only is determined such that the number of nvPM does not exceed a predetermined threshold which in the Example shown corresponds to the nvPM number of the SAF-rich fuel composition at the staging point SP. In this example there is an operating zone at low values ​​of Wr in which the pilot fuel injectors 313 are powered with the fuel composition low in SÂF. As W; increases and the nvPM number increases, at some point it is necessary to begin blending in a certain SAF-rich fuel composition in order to prevent the nvPM number from rising above the predetermined threshold. As W, increases further, the proportion of SAF-rich fuel in the mixture continues to increase until it reaches 100% at the staging point SP. At even higher values ​​of W,, corresponding to pilot and main injection operation, the pilot fuel injectors 313 are again fed with a fuel composition low in SAF. Although [Fig.18] shows a capped nvPM number as a function of Wz, it will be borne in mind that through the appropriate determination of the dependence of the mixing ratio on W; over the entire pilot injection only operating region, it is possible to obtain any desired dependence of the number of nvPM with respect to Wz, subject to an upper limit corresponding to the dependence of the second fuel (this is i.e. fuel associated with high nvPM level, such as low SAF fuel composition) and a lower limit corresponding to the dependence of the first fuel (i.e. fuel associated with low level of nvPM, such as SAF-rich fuel composition). If the SAF-rich fuel composition matches the fuel composition C from 7, the treatable region is represented by the hatched area in [Fig.19]. Within the processable region, any single-valued function of the number of nvPM versus Wz can in principle be obtained by an appropriate mixing program as a function of W; in pilot injection only operation. [Fig.20] shows another example of how the number of nvPM can depend on fuel flow. By supplying a fuel mixture to the pilot fuel injectors in this manner in pilot injection only operation (or at least within one or more regions within pilot injection only operation) the intensity of the fuel rate variation of the number of nvPM with W; can be made significantly lower than for the default fuel composition such as fossil kerosene, and also less lower than for other examples in which switching between fuel sources 302, 304 occurs. , in pilot injection only operation, using the example in [Fig.17], Wr can be modified for reasons other than soot emissions without causing very significant changes in emissions of soot. For example, in US2022042465, fuel flow rates of individual engines are varied during final approach to limit purge noise. If we assume that The thrust requirement during final approach corresponds to a pilot injection only operating region, then the examples herein in which fuel is mixed by the fuel mixer 318 of the present invention can significantly reduce the change. soot emissions that would otherwise accompany such noise-driven fuel flow changes. The examples shown in Figures 13 and 15 can be considered as special cases of the blended fuel example in which the SAF-rich fuel composition is used not only at the staging point SP but also throughout an entire region. extending in some way to the left of the staging point (i.e. at lower fuel flow rates). In the example in [Fig.13] this region extends to the left side of the graph. In the example in [Fig.15] there is a tipping point at which the fuel composition is strongly tilted towards the low SAF fuel composition in the pilot injection only operating range. It will be borne in mind that switching from one type of fuel to another is equivalent to changing the mixture ratio from 0:100 to 100:0, or vice versa. [Fig.21] illustrates a method 4000 of operating a gas turbine engine 10 which can be implemented using the system of [Fig.17]. Steps common to other processes described previously are marked in this sense. In this example, fuel delivery control 4002 includes: mixing 4010 a fuel supply from both the first and second fuel sources 302, 304 to form a blended fuel consisting of fuel from the first fuel source. fuel 302, fuel from the second fuel source 304, or a mixture thereof; and the distribution 4012 of the mixed fuel to the pilot fuel injectors 313. Any of the features described above in connection with [Fig.17] can be incorporated into the method of [Fig.21], even if they are not repeated here. nvPM cost function In the examples described above the control of the fuel characteristics supplied to the fuel injectors is based on a desired form of the dependence of the number of nvPM on W; from which the fuel characteristic(s) to be used at each value of Wr are determined to enable the desired shape to be obtained. The fuel delivery regulator in these examples can therefore be controlled to minimize engine nvPM production. The nvPM production can be in terms of the nvPM content of the engine exhaust e.g. the mass or number of nyPM particles produced. In order to provide even further improved control of nvPM emission, additional factors may be considered as the inventors have identified that one emitted soot particle is not as disadvantageous as another emitted soot particle. Therefore in some examples one can take into account the amount of cost or harm caused by each soot particle emitted. For example, we can consider that soot particles emitted close to the ground have a greater impact on human health than soot particles emitted several thousand feet above the ground (indeed soot emissions at less altitude of 3,000 feet above the runway are regulated, while soot emissions at higher altitudes are not). The fuel regulator of any of the examples described herein may be configured to distribute fuel to the pilot fuel injectors to regulate (e.g., optimize or reduce) the impact of engine exhaust nvPMs. The fuel delivery regulator 306 may therefore be configured to distribute fuel to the pilot fuel injectors to minimize a cost function based on one or more nvPM impact parameters. nvPM impact parameters may relate to the cost or harm of emitting nvPM (e.g., soot) of a particular type or in a certain situation. This may allow the gas turbine engine 10 to be operated in a manner that reduces nvPM emissions that might otherwise cause the greatest harm compared to those that have less effect on the environment and / or human health. . The nvPM impact parameter(s) on which the cost function is based may include one or more elements of: 1) the height above the ground at which nvPM production takes place; li) the position (e.g. location i.e. longitude and latitude) of the nvPM production. The cost of nvPM emission can be set based on the 3D position of soot emissions (e.g. altitude, longitude, latitude), and can take into account proximity to population centers or other important locations; iii) weather / atmospheric conditions at an nvPM production location. For example the cost function can take into account prevailing winds, weather patterns and atmospheric behavior such as removal processes in order to assess the future location of emitted soot particles relative to population centers or other important places; iv) climate impacts associated with the location of nvPM production. For example, the climatic impacts of soot deposition at certain locations can be taken into account. Such locations may otherwise include high albedo surfaces, for example; v) mass / size of individual nvPM particles produced. The cost function can be defined to prioritize reductions in the number and / or mass of nvPMs within one or more specific particle size ranges, e.g. relative to reductions in the number and / or mass of nvPMs in one or more other size ranges. The cost function may, for example, include a sum weighted over the number (or mass) of particles in different size ranges, with the weighting set to reflect the notion that certain size ranges may be more harmful to health, or result in a more unfavorable environmental outcome, than others (and therefore be more important than the cost function), but without completely ignoring other size ranges; (vi) potential contrail generation and / or contrail characteristics. For example, the cost function may take into account the probability that certain nvPM emissions will cause the aircraft to produce a contrail, or affect the properties or characteristics of a contrail produced by the aircraft. . This may apply more particularly to contrails formed at the top of the aircraft descent phase where combustion chamber operation would likely be significantly below the staging point and fuel usage at low level of nvPM could have a greater effect; vii) local air quality (LAQ) impact of nvPM production; and or vili) the quantity of nvPM produced. The cost function may take into account the quantity of nvPM emissions produced by the engine, for example in terms of total mass and / or total number of nvPM particles in a given time, or an emissions index of the mass. (or number) of one nvPM per unit mass of fuel consumed. Any number of the above impact parameters may be set to define a cost function which is to be minimized and the fuel delivery regulator 306 controlled to supply fuel of the required characteristics to the combustion chamber. 16. In some examples, the nvPM production cost function may be incorporated into a broader cost function including other costs. Control of the fuel delivery regulator 306 of any example described herein may be based on the cost function. For example, the onboard mixture of the first and second fuels (rich in SAF and lean in SAF) can be used to achieve a desired number of nyvPM at each operating condition within pilot injection only operation. The fuel characteristics provided to the pilot fuel injectors 313 in pilot injection only operation (and therefore the corresponding mixing ratio of the amount of low-SAF composition to the amount of high-SAF fuel composition) can no longer be a simple function of Wr (in fact, it may not even be a single-valued function of Wr) but may vary with other parameters of so as to minimize an overall cost function. Switching between pilot injection only operation and pilot and main injection operation during cruise In known staged combustion systems, the stage point SP is usually chosen such that operation at cruise speed takes place within the pilot and main injection operating region. Operation at low power settings, such as taxi, descent, and approach, often takes place within the pilot-injection-only operating region. The inventors have determined that it is further advantageous to operate a staged combustion system such that it is in pilot injection only mode during at least a portion of its cruise operation, while also selectively providing fuel from two different sources to the combustion chamber during cruise operation. In another example of the present application, the staged combustion system 64 illustrated in [Fig.5] is additionally or alternatively designed to switch between the pilot injection only operating range and the pilot injection operating range and main at the staging point selected so that it corresponds to a cruising mode of operation in the stabilized state of the engine 10. The staging point in this example is designed such that it defines a limit between a first operating range at engine cruise speed and a second operating range at engine cruise speed. In other words, the staging point is selected to occur at an engine power setting (or other motor operating parameter indicative of the engine power setting) that is above the setting minimum engine power at which the engine operates at cruising speed in the stabilized state. This defines a range of relatively lower cruise engine power conditions, in which the staged combustion system operates in the pilot injection only mode of operation, compared to pilot and main injection operation at engine power settings in relatively higher cruising speed. This effectively moves the staging point to a higher power setting so that the combustion system can still operate in pilot injection only operation at higher power settings than the minimum cruise power setting. The pilot injection only and pilot and main injection ranges are therefore redefined such that the staged combustion system 64 switches between operating modes at one W; different, that is to say that the border separating the ranges is changed compared to the systems of the state of the art. In this example, the fuel delivery regulator 306 is designed to distribute fuel to the pilot fuel injectors 313 during at least part or preferably all of the first cruise operating range having a fuel characteristic different from that of the fuel distributed to the pilot fuel injectors 313 and / or to the main fuel injectors 314 during the second operating range at cruise speed. More specifically, during the first range of cruise operation, the fuel regulator 306 supplies fuel to the pilot fuel injectors from available fuel sources which is different from that supplied to the fuel injectors during the second range. operating at cruising speed. This may be fuel supplied to both the main and pilot fuel injectors during main and pilot injection operation. The inventors have determined that by adjusting the staging point so that lower cruise operation can take place in the pilot injection only mode, certain engine emissions can be reduced and combustion efficiency can be reduced. improved. In combination with the selective use of fuels with different characteristics, the inventors have determined that disadvantageous effects on emissions that would otherwise result in a shift in the staging point can be mitigated. This therefore gives an overall improvement in combustion efficiency and a reduction in emissions through a combination of these factors. Pollutant emissions can be characterized by an emission index (ED), detailing the mass (or number) of a particular pollutant per unit mass of fuel consumed. The inventors have observed that in operation with pilot injection only, at very low engine powers, the mass emission index of carbon monoxide (CO) and unburned hydrocarbons (HC) can be relatively high. Not only does this represent a release of pollutants into the atmosphere, but it also reduces fuel efficiency due to incomplete combustion of the fuel, i.e. not all of the fuel is completely burned. . This may result in increased mission fuel burn and / or reduced aircraft payload range capability. As engine power increases (still in pilot injection only operation), both EI(CO) and EI(HC) decrease and then remain low. However, when moving to even higher engine power settings and transitioning to pilot and main injection operation, the EI(CO) and EI(HC) become high again before dropping back to low again. even higher engine power settings. This is illustrated in [Fig.22], which shows the dependence of EI(CO) and EI(HC) on the motor power setting, with the vertical dotted line representing the SP staging point. During the first range of cruise operation, the engine operates at relatively low power settings, such that if the staging point were selected to cause the combustion system to operate in pilot and main injection mode, the emission of CO and HC would be relatively high. This is due to the fact that even if the staging point was chosen such that cruise operation takes place in pilot and main injection mode, that is to say above the SP staging point , some parts of the cruise operation could be quite close to the staging point (e.g. just above it). Disadvantageously, this would mean that HC and CO emissions may be high during such operation, and combustion efficiency may be disadvantageously and materially reduced (resulting in lower fuel consumption). higher fuel and / or a reduction in aircraft payload range capability). By setting the staging point so that low power cruise operation occurs in the pilot injection only mode, this disadvantageous CO and HC emission can be reduced or avoided. This is illustrated in [Fig.23]. which shows the effect of adjusting the staging point SP' to define a first cruising speed operating range 320a and a second cruising speed operating range 320b. As can be seen in [Fig.23], the first cruising speed operating range 320a corresponds to a low power cruising speed, which is below the staging point SP', the second operating range in cruising speed 320b corresponding to operation in cruising speed of higher power above the staging point SP', In the first operating range in cruising speed 320a, the staged combustion system 64 is configured to operate in pilot injection only mode, thus providing low CO and HC emissions. Due to the shifting of the staging point to a higher power setting (e.g., from thin to thick dotted lines in [Fig.23] marked SP and SP respectively”), CO and HC emissions remain weak over a larger range of motor power settings (e.g., thick solid line versus thin dashed line in the dependence of EI(CO) and EI(HC) on power setting). The inventors have further found that the nvPM emission index increases rapidly with engine power in the pilot injection only mode but is usually uniformly low in the pilot and main injection mode. For nvPM, the emissions index can be described with reference to the mass of nvPM or the number of nvPM particles (per unit mass of fuel). An example of the dependence of nvPM on the engine power setting is illustrated in Figure [Fig.24]. As a result, by effectively moving the staging point to a higher power setting, the default El(nvPM) value is advantageously increased. As can be inferred from [Fig.24], operating in pilot injection only mode at higher power ratings would result in increased nvPM production, as there would be less of the engine power band in fuel injection operation pilot and main when nvPM production is low, and nyPM would increase rapidly at higher powers in pilot-only injection mode of operation. The inventors have determined that this otherwise detrimental increase in nvPM can be mitigated by the selective use of a fuel of a different fuel characteristic during at least that portion of the pilot injection only cruise operating range at which nvPM production would otherwise be high. For example, the fuel delivered to the pilot fuel injectors 313 during the first engine cruise operating range 320a may be chosen to be fuel associated with an nvPM production level that is lower than that of the fuel delivered to the injectors. pilot and / or main fuel 313, 314 during at least part of the second operating range at engine cruise speed 320b. This can be achieved by the fuel contained in the first fuel source 302 being associated with low nvPM production compared to the second fuel source 304 (under corresponding combustion conditions). The effect of selectively using a fuel of a different characteristic during the first range of engine cruise operation 320a is illustrated in [Fig.25]. The thick solid line in the dependence of El(nvPM) on power can be compared to the thick dashed line in the first cruise operating range 320a which is produced using a low nvPM producing fuel. In some examples, the fuel characteristic by which the fuel from the first fuel source 302 differs from the fuel from the second fuel source 304 may be the percentage of sustainable aviation fuel (SAF) present in the respective fuel as indicated above. The fuel distributed to the pilot fuel injectors 313 during at least a portion of the pilot injection-only cruise operating range would similarly have a different SAF percentage compared to the fuel distributed to the pilot fuel injectors 313 and / or main fuel injectors 314 for at least part of the pilot and main injection operating range. As noted above, compared to fossil kerosene, SAF gives significantly less nvPM, and therefore can be used to mitigate the increase in nvPM that would otherwise be obtained from additional pilot-only cruise operation. For example, the fuel supplied to the pilot fuel injectors during the first cruise operating range 320a may have a greater proportion of SAF compared to that supplied to the fuel injectors during the second cruising operating range 320b. This allows a limited amount of SAF (or other low nvPM fuel) to be selectively used to reduce overall nvPM, CO and HC emissions. This may not always be the case, however. In some examples, fossil kerosene can be processed to remove aromatic components, particularly naphthalenes, to produce a largely paraffinic fuel of fossil origin that would be a fuel producing low levels of nvPM. Other fuel characteristics may therefore be associated with a low nvPM level, such as percent aromatic content or naphthalene content. The inventors have therefore noted that through a combination of the selective use of two types of fuel and operation in pilot injection mode only at a low cruising speed, the advantageous reduction in CO and HC emissions is accompanied a lower disadvantageous increase in nvPM than would occur without using a high-SAF fuel composition (or another fuel associated with a low level of nvPM) within the relevant range of engine power settings. The selective pilot injection only cruise operation and selective fuel usage of the present application are advantageous over other known methods for reducing overall nvPM, HC and CO emissions. For example, alternatives include a "pilot and semi-main injection" operating zone which is between the pilot injection only region and the pilot and main injection region. When the engine power setting corresponds to the pilot and semi-main injection region, instead of lighting all the main fuel burners, only a proportion (e.g. half) of them are ignited, with e.g. an alternating on-off-on-off distribution around the combustion chamber ring (other arrangements could be considered such as lighting main burners on one half of the ring and not on the other half ). However, this results in uneven combustion properties around the combustion chamber ring and can also be disadvantageous to turbine operation and turbine life. Another approach is to have a staged pilot system in which the "pilot injection only" operating region is divided into subregions, in which progressively larger numbers of pilot burners are switched to progressively higher engine power settings in the overall “pilot injection only” region. Such a configuration disadvantageously adds weight and complexity which can be avoided by the methods of the present application. In one example, the first cruise operating range corresponds to operation of the aircraft in a later part of a cruise segment of a flight, and the second cruising operating range corresponds to operation of the aircraft in a relatively anterior part of the cruise segment. For example, during a flight cycle of the aircraft, the operation in cruise mode can be divided into one or more cruise segments. These may correspond to steady-state operation at different altitudes. Toward the end of a cruise segment, the engine power setting required to maintain steady-state cruise at the specified Mach number and specified altitude will decrease as the aircraft burns fuel and lightens up. As less thrust is therefore required at a later part of the cruise segment, the engine power setting is reduced. The combustion system of the present application can therefore be designed to switch to pilot injection only operation at low engine powers towards the end of a cruise segment, rather than remaining in pilot injection operation and main throughout the cruise segment (or segments if there is more than one) of the flight. This helps reduce HC and CO production that might otherwise occur towards the end of the cruise segment. In another example, the first cruising speed operating range 320a corresponds to subsonic cruising speed operation in the stabilized state of the engine and the second cruising speed operating range 320b corresponds to cruising speed operation supersonic in the stabilized state of the engine. In this example, the gas turbine engine 10 is designed to provide both subsonic and supersonic cruise operation of the aircraft on which it is mounted. The staging point can be determined such that operation in cruise mode with pilot injection only corresponds to subsonic operation at relatively low engine power, while operation in cruise mode with pilot and main injection corresponds to operation supersonic at higher engine power. For a supersonic aircraft, supersonic cruise is likely to be possible over oceans, but over land the cruise may be constrained to be subsonic for reasons of compliance with noise regulations. . In such an example, the supersonic cruise regime can cor- respond to a high engine power setting, well above the staging point. Subsonic cruise may however in this example correspond to much lower engine power settings which could be near a default or state of the art staging point, resulting in low combustion efficiency and high CO and HC emissions during subsonic cruise. A supersonic aircraft can consume a material proportion of its fuel at subsonic cruise, and thus the combustion efficiency in this operating condition is very important. By operating in pilot injection only mode during subsonic cruise, CO and HC emissions can be reduced, while fuel of a different specification can be supplied to minimize any increase in production of nvPM. Steady-state supersonic cruise operation may occur before steady-state subsonic cruise operation, or vice versa. In some examples, steady-state subsonic cruise operation may be subsonic cruise operation over land, while steady-state supersonic cruise operation may be in-state operation. supersonic cruise regime over water (e.g. over the sea). In order to supply fuel to the pilot fuel injectors 313 during the first cruise operation which is different from that supplied to the combustion chamber during the second cruise operation the fuel distribution regulator 306 is designed to supply fuel selectively from the first and second fuel sources 302, 304 as described above (for example fuel coming only from the first fuel source 302, fuel coming only from the second fuel source 304, or a mixture of these). In one example, the fuel distribution regulator 306 may include a fuel mixer 318 as illustrated in [Fig.17]. Similar to that described above, the fuel mixer is configured to receive a supply of fuel from the first and second fuel sources 302, 304 and to output fuel consisting of the fuel from the first fuel source, fuel from the second fuel source, or a mixture thereof. This mixed fuel is supplied to the pilot fuel injectors 313, with fuel from a single fuel source (the second fuel source 304 in the example above) being supplied to the main fuel injectors 314. In others Examples, the mixer can be designed to feed both the main and pilot fuel injectors with an appropriate fuel mixture. In yet other examples, the fuel delivery regulator 306 according to any examples described herein may be used to provide fuel in the first and second ranges of cruise operation (e.g., those using independent regulators for the first and second fuel sources or a pilot regulator designed to switch between sources fuel). As indicated above in connection with the example illustrated in [Fig.17], the dependence of the proportion of fuel coming from the first fuel source 302 compared to that coming from the second fuel source 304 compared at the fuel flow rate can be determined based on a desired achieved level of nvPM at a particular fuel flow rate. For example, the proportion of fuel dispensed from the first fuel source 304 compared to that of the second fuel source 304 may be determined according to a desired obtained level of nvPM at a particular fuel flow rate in the first range of operation at engine cruise speed. The quantity of fuel supplied from each fuel source may be determined such that the nvPM value does not exceed a predetermined threshold or such that the production of nvPM is minimized over a period of operation of the turbine engine. gas as described above. In any of the examples described here this can be achieved by appropriate control of the mixer 318, or any of the other fuel delivery regulators 306 described herein. In certain examples, the proportion of fuel distributed from the first fuel source 302 compared to that distributed from the second fuel source 304 during the first cruising range of operation can be determined at least partially according to one or any number of elements: a) the quantity of fuel having the first fuel characteristic and the quantity of fuel having the second fuel characteristic available for a proposed flight. This can for example be provided by an estimation or measurement of the quantity of fuel in the fuel tanks 53, 55 making up the first and second fuel sources; b) the total fuel requirement amount for the fuel injectors 313 during pilot injection only operation for the entire flight within an operating range in which fuel is supplied from the first fuel source 302; elou c) a fuel composition limit parameter (e.g. a certification limit, fuel composition available for refueling, or aircraft / engine limits). [Fig.26] illustrates a method 4014 of operating a gas turbine engine according to the example above in which fuel of a different characteristic is distributed to the pilot fuel injectors during the first operating range at engine cruise speed. The method 4014 includes regulating 4016 the distribution of fuel to the pilot and / or main fuel injectors 313, 314 from the first fuel source 302 containing a first fuel having a first fuel characteristic and from the second source of fuel 304 containing a second fuel having a second fuel characteristic, as described above using the fuel delivery regulator 306 of any example herein (i.e., fuel coming only from the first fuel source 302, fuel from only the second fuel source 304, or a mixture thereof). The second fuel characteristic is different from the first as described above. The method further includes switching 4018 between the pilot injection only operating range and the pilot and main injection operating range at a staging point during a steady state cruise mode of operation of the engine to define a first engine cruise speed operating range 320a and a second engine cruise speed operating range 320b. The method further comprises distributing fuel 4020 to the pilot fuel injectors 313 during at least a portion of the first engine cruise operating range 320a having a fuel characteristic different from the fuel distributed to the pilot fuel injectors and / or main 313, 314 during the second operating range at engine cruise speed 320b. Any of the features described above in connection with the examples in which a different fuel is provided in the first and second engine cruise operating ranges may be incorporated into the method 4014, even if they are not repeated here. ariation of the staging ratio in a cruise regime in the stabilized state p In another example of the present application, the staged combustion system 64 illustrated in [Fig.5] is additionally or alternatively designed to operate in a transitional operating range between the pilot injection only operating range and the pilot and main injection operating range. In the present examples, the transition operating range is provided during cruise operation in the stabilized state of the engine. In the transition operating range, fuel is delivered to both the pilot fuel injectors and the main fuel injectors 313, 314 at a transition staging ratio. As defined elsewhere herein, the staging ratio defines the relative fuel mass flow rates attributable to the pilot fuel injectors 313 Compared to main fuel injectors 314, In the pilot injection only operating range, the staging ratio is, by definition, 100:0. In the pilot and main injection operating range, the combustion chamber is configured to operate with a pilot and main injection staging ratio which may be 20:80 or 30:40, etc. Within the transition operating range the combustion chamber is configured to operate using a transition staging ratio which is different from the pilot and main injection staging ratio (and different from the staging ratio pilot injection only, since fuel is delivered to both the main and pilot fuel injectors during the transition operating range). More specifically, in the transition operating range, the proportion of fuel flow attributable to the pilot fuel injectors compared to the main fuel injectors is greater than for the ratio of pilot and main injection. In other words, a smooth or intermediate transition is provided, in which a greater proportion of fuel is supplied to the pilot fuel injectors compared to a direct switch between all fuel being supplied to the pilot fuel injectors in the pilot injection only mode, and the pilot and main injection staging ratio in which a smaller proportion of fuel is supplied to the pilot fuel injectors. In other words, rather than moving the precise staging point to a higher power setting (as described elsewhere herein), the inventors have determined that a transition region that can be included is characterized by a transition slower (compared to the engine power setting) from the pilot-only injection staging ratio (100:0) to the chosen pilot and main injection staging ratio (for example 20:80). The inventors have determined that by configuring the staged combustion system 64 to operate in a transition range (i.e., a transition operating range) between pilot injection and pilot and main injection operation, the amount of emissions of CO and HC implemented in this range of engine power settings can be reduced. This is illustrated in [Fig.27], which shows the dependence of the emission index (El) of CO and HC for a staged combustion chamber (such as that described here) on the engine power setting . The transition range is marked 322, and is at an engine power range between pilot injection only operation and pilot and main injection operation. At engine powers within the transition range a high level of HC and CO would be emitted if the combustion chamber 16 were intended to operate at the pilot and main injection staging ratio, as illustrated by the dotted line in the HC and CO curve in [Fig.27]. By operating at transition staging ratio (which in this case gradually decreases as engine power increases), the CO and HC emission index is reduced, as shown by the thick solid line in the curve of HC and CO in the transition range. In this example, the fuel delivery regulator 306 is configured to distribute fuel to the pilot 313 and / or main 314 fuel injectors during the transition operating range having a different fuel characteristic than the fuel delivered to the fuel injectors pilot and / or main 313, 314 for at least part of the pilot and main injection operating range. The inventors have further determined that, disadvantageously, the default EI(nvPM) value in transition range 322 is increased for fuel having identical characteristics. This can be seen in [Fig.28]. which illustrates the dependence of the nvPM emission index on engine power for a staged combustion chamber. The thin dotted line in [Fig.28] represents nvPM emissions if transition range 322 was not provided. The thick solid line shows the effect of introducing transition range 322 on nvPM emissions. In the present example, the inventors have determined that the EK(nvPM) value in the transition range 322 can be attenuated by the use of a fuel of a different characteristic (e.g. fuel rich in SAF) as represented by the thick dashed line in the nvPM curve within the transition range. Thus, the nvPM penalty (necessary to achieve lower CO / HC emissions in transition range 322) is reduced by using a SAF-rich fuel composition in transition range 322 rather than using a fuel composition defaults to transition range 322. In some examples the fuel distributed to at least the pilot fuel injectors 313 by the fuel delivery regulator 306 during the transition operating range 322 may be chosen to be fuel associated with an nvPM production level which is lower than that of the fuel supplied to the pilot and / or main fuel injectors 313, 314 during at least a portion of the pilot and main injection operating range. This can be achieved by the fuel contained in the first fuel source 302 being associated with low nvPM production compared to the second fuel source (under corresponding combustion conditions). In some examples, the fuel delivered to both the pilot and main fuel injectors 313, 314 by the fuel delivery regulator 306 during the transition operating range 322 is low nvPM producing fuel. In some other examples, only fuel delivered to the pilot fuel injectors 313 during the Transition operating range 322 is chosen to be associated with a low nvPM level (i.e. fuel delivered to the main fuel injectors remains the same). This may allow better utilization of the amount of low nvPM fuel available, since supplying fuel to the main fuel injectors (i.e., the lean portion of the fuel system) may have less effect on reduction nvPMs. In some examples, the fuel characteristic by which the fuel from the first fuel source 302 differs from the fuel from the second fuel source 304 may be the percentage of sustainable aviation fuel (SAF) present in the respective fuel as indicated above. Fuel distributed to pilot fuel injectors 313 during transition operating range 322 would similarly have a different SAF percentage compared to fuel distributed to pilot fuel injectors 313 and / or main fuel injectors 314 during at least a portion of the operating range. with pilot and main injection. As noted above, compared to fossil kerosene, SAF gives significantly less nvPM, and can therefore be used to mitigate the increase in nvPM that would otherwise be obtained from additional cruise operation. For example, fuel supplied to the pilot fuel injectors during the transition operating range may have a greater proportion of SAF compared to that supplied to the fuel injectors during the pilot and main injection operating range. This allows a limited available amount of SAF (or other low nvPM fuel) to be selectively used to reduce overall nvPM, CO and HC emissions. This may not always be the case, however. In some examples, fossil kerosene can be processed to remove aromatic components, particularly naphthalenes, to produce a largely paraffinic fuel of fossil origin that would be a fuel producing low levels of nvPM. Other fuel characteristics may therefore be associated with a low nvPM level, such as percent aromatic content or naphthalene content. The first and second fuels can therefore differ in characteristics other than the SAF content while still obtaining the desired effect on nvPM emissions. During the range of transition operation 322 the transition staging ratio may vary with a changing motor power setting. In one example, the transition staging ratio exhibits a continuous variation with changing motor power within the transition operating range 322. This can provide a smooth transition between the staging ratio within the operating range with pilot injection only and that with pilot injection and main. The continuous variation may be such that the proportion of the total fuel flow to the fuel injectors (i.e. the total being distributed to the pilot and main fuel injectors) that can be attributed to the flow of fuel to the pilot fuel injectors 313 decreases with increasing engine power during the transition operating range 322. The proportion of the total fuel flow to the fuel injectors that can be attributed to a fuel flow to the main fuel injectors 314 instead increases with increasing engine power in the transition operating range 322. In other examples, the transition staging ratio may have a constant intermediate value which is different from the pilot and main injection staging ratio. The transition staging ratio can be between the pilot injection ratio only and the pilot and main injection ratio. This therefore allows a more gradual transition between the pilot injection ratio alone and the pilot and main injection ratio. For example, the transition staging ratio may be 70:30, which is between 100:0 in the pilot injection only range, and a pilot and main injection staging ratio which is e.g. from 8:80 p.m. or 30:70 p.m. In yet other examples, the transition staging ratio varies between a series of constant intermediate values, each different from the pilot and main injection staging ratio. Furthermore, each of the intermediate ratios can be included between that of the pilot injection range alone and that of the pilot and main injection range. For example, the transition staging ratio can be a series of values ​​of 80:20, 60:40, and 40:60. The ratio of pilot and main injection can in this example be 20:80. The intermediate staging ratios can thus decrease towards the pilot and main injection ratio (i.e. a progressively smaller proportion of the total fuel is supplied to the pilot fuel injectors, and a progressively smaller proportion greater of the total fuel is supplied to the main fuel injectors). This is, however, only an example, and any other number and values ​​of intermediate transition staging ratios may be used. In some examples, the transition staging ratio may have a continuous variation with motor power in one part of the transition operating range and may have one or more constant values ​​in another part of the transition operating range . The examples above can therefore be combined. In other examples, the staging ratio may have a continuous variation over the entire transition operating range, or one or more constant values ​​over the entire transition operating range. In order to supply fuel to the pilot fuel injectors 313 during the transition operating range 322 which is different from that supplied to the combustion chamber 16 during the pilot and main injection operating range the fuel delivery regulator 306 is designed to supply fuel selectively from the first and second fuel sources 302, 304 as described above (for example from only the first fuel source 302, only the second fuel source 304, or a mixture of these). In one example, the fuel delivery regulator 306 may include a fuel mixer 318 as shown in [Fig.17]. Similar to that described above, the fuel mixer is configured to receive a fuel supply from both the first and second fuel sources 302, 304 and to output fuel from the first source. fuel, fuel from the second fuel source, or a mixture thereof. This mixed fuel is supplied to the pilot fuel injectors 313, with fuel from a single fuel source (the second fuel source 304 in the example above) being supplied to the main fuel injectors 314. In other cases Examples, the mixer can be designed to feed both the main and pilot fuel injectors with an appropriate fuel mixture. In yet other examples, the fuel delivery regulator 306 of any of the examples described herein may be used to provide fuel in the pilot-only, transition, and pilot-and-main-injection ranges (e.g., those using independent regulators for the first and second fuel sources or a pilot regulator designed to switch between fuel sources). As indicated above in connection with the example illustrated in [Fig.17], the dependence of the proportion of fuel coming from the first fuel source 302 compared to that coming from the second fuel source 304 compared at the fuel flow rate can be determined based on a desired achieved level of nvPM at a particular fuel flow rate. For example, the proportion of fuel delivered from the first fuel source 302 compared to that from the second fuel source 304 may be determined according to a desired obtained level of nvPM at a particular fuel flow rate in the transition range. 322. The quantity of fuel supplied from each source may be determined such that the nvPM value does not exceed a predetermined threshold or such that the production of nvPM is minimized over a period of operation of the gas turbine engine as described above. In any of the examples described herein this may be achieved by appropriate control of the mixer 318, or any of the other fuel delivery regulators 306 described herein. In some examples, the proportion of fuel dispensed from the first fuel source 302 compared to that dispensed from the second fuel source 304 during the transition operating range 322 may be determined at least partially based on one or more elements any of: a) the quantity of fuel having the first fuel characteristic and the quantity of fuel having the second fuel characteristic available for a proposed flight. This can for example be provided by an estimation or measurement of the quantity of fuel in the fuel tanks 53, 55 making up the first and second fuel sources; b) the total fuel requirement amount for the fuel injectors 313 during pilot-only operation for the entire flight within an operating range in which fuel is supplied from the first fuel source 302; and or c) a fuel composition limit parameter (e.g. a certification limit, fuel composition available for refueling, or aircraft / engine limits). [Fig.29] illustrates a method 4022 of operating a gas turbine engine according to the example above in which fuel of a different characteristic is distributed to the pilot fuel injectors during the transition range 322 compa - relative to the pilot and main injection operating range. The method 4022 includes regulating 4024 the distribution of fuel to the pilot and / or main fuel injectors 313, 314 from the first fuel source 302 containing a first fuel having a first fuel characteristic and from the second source of fuel 304 containing a second fuel having a second fuel characteristic as described above using the fuel delivery regulator 306 of any example herein (e.g. such that the fuel dispensed is fuel originating only from the first fuel source, fuel from only the second fuel source, or a mixture thereof). The second fuel characteristic is different from the first as described above. The method further includes operating 4026 of the staged combustion system in a transition operating range 322 between the pilot injection only and pilot and main injection operating ranges in which fuel is delivered to the pilot fuel injectors and main to a transition staging ratio which is different from the pilot and main injection staging ratio. Controlling 4024 the delivery of fuel to the pilot and / or main fuel injectors includes the delivery 4028 of fuel to the pilot and / or main fuel injectors during the operating range of transition 322 having a fuel characteristic different from that of the fuel supplied to the pilot and / or main fuel injectors during at least a portion of the pilot and main injection operating range. Any of the features described above in connection with the examples in which a different fuel is delivered in the transition range 322 may be incorporated into the method 4022, although they are not repeated here. Staged combustion chamber control during acceleration An acceleration of the engine is caused by deliberately increasing the fuel flow to a value greater than that necessary to maintain operation in the stabilized state. The inventors have observed that this "supercharging" causes an initial increase in the fuel-air ratio in the combustion chamber and can lead to increased production of nvPM (soot or smoke), particularly at engine power settings corresponding to injection operation. pilot alone. To reduce the production of excessive amounts of nvPM during acceleration, it is known to switch to a "throttle" mode of operation of a gas turbine engine in which the staging point occurs at a power setting weaker engine. Thus, the transition to pilot and main injection operation takes place at a lower engine power setting than it would during steady state operation. This helps reduce the overall amount of nyPM produced during acceleration. The inventors have observed that switching to such a known acceleration mode can, however, present a certain number of disadvantages. For example, an increase in HC and CO emissions may be caused. In some known solutions, maintaining an acceptable level of CO and HC results in an increase in NOx emissions because the poorest flames which produce the least NOx in this region produce too much CO and HC , and conversely design changes to reduce CO and HC can lead to increased NOx emissions across the entire power curve. For example a known solution could be to physically change and enrich the entire main fuel injector to reduce CO and HC emissions, which would then lead to higher NOx levels at all power levels. In another example of the present application, the staged combustion system 64 illustrated in [Fig.5] is additionally or alternatively designed to operate in an acceleration mode of operation. In the acceleration mode the fuel delivery regulator 306 is designed to distribute fuel to the fuel injectors (i.e. the pilot and / or main fuel injectors 313, 314) having a characteristic of fuel different from that of the fuel distributed to the fuel injectors (i.e. to the pilot and / or main fuel injectors 313, 314) during at least part of the steady state operating mode. Engine throttle mode is a mode in which the fuel delivery regulator 306 is designed to deliver fuel to the fuel injectors 313, 314 at a rate greater than sufficient to maintain operation at the same time. stabilized state of the engine. This causes the acceleration of the motor, by which we designate the increase in the angular speed of one or more bodies of the motor. In the steady state operating mode the fuel injectors are not supercharged so that no engine acceleration occurs. The inventors have determined that increased nvPM emissions when the engine is operating in an acceleration mode can be avoided or reduced by using a fuel with different characteristics than that used during steady-state operation. This allows the staging point during the acceleration mode to remain the same or similar to that of the steady state operating mode, thereby avoiding or reducing / limiting the disadvantageous increase in HC or CO emissions. In some examples therefore the fuel delivered to the fuel injectors (e.g., pilot and / or main fuel injectors 313, 314) during the boost mode of operation may be chosen to be fuel associated with a production level. of nvPM that is lower than that of the fuel delivered to the fuel injectors (e.g., pilot and / or main fuel injectors 313, 314) during at least a portion of the steady state mode of operation. This can be achieved by the fuel contained in the first fuel source 302 being associated with a low production of nyPM compared to the fuel contained in the second fuel source 304 (under corresponding combustion conditions). Fuel associated with lower nvPM production may be supplied to at least the pilot fuel injectors 313 in throttle mode operation since this will have the most significant effect on nvPM production. Preferably, fuel associated with lower nvPM production may be supplied only to the pilot fuel injectors during the acceleration mode to better utilize a limited supply of this fuel. In some examples, the fuel characteristic by which the fuel from the first fuel source 302 differs from the fuel from the second fuel source 304 may be the percentage of sustainable aviation fuel (SAF) present in the respective fuel as indicated above. Fuel delivered to the fuel injectors during the throttle mode of operation would have to even a different SAF percentage compared to the fuel delivered to the pilot 313 and / or main 314 fuel injectors during at least a portion of the steady state mode of operation. As noted above, compared to fossil kerosene, SAF gives significantly less nvPM, and can therefore be used to mitigate the increase in nvPM that would otherwise be obtained from supercharging the combustion chamber during the combustion mode. 'acceleration. For example, the fuel supplied to the fuel injectors 313, 314 during the acceleration mode may have a greater proportion of SAF compared to that supplied to the fuel injectors during the steady state mode. This allows a limited available amount of SAF (or other low nvPM fuel) to be selectively used to reduce overall nvPM, CO and HC emissions. As noted above in conjunction with other examples, however, this may not always be the case. In some examples, fossil kerosene can be processed to remove aromatic components, particularly naphthalenes, to produce a largely paraffinic fuel of fossil origin which would be a fuel producing low levels of nvPM. Other fuel characteristics may therefore be associated with a low nvPM level, such as percent aromatic content or naphthalene content. The first and second fuels can therefore differ in characteristics other than the SAF content while still obtaining the desired effect on nvPM emissions. In some examples, the fuel delivery regulator 306 is configured to dispense fuel having a different fuel characteristic during an acceleration mode of operation that follows a steady state mode of operation in which the combustion chamber system 64 operates in pilot injection only mode. In some examples therefore, fuel characteristics are changed only for engine accelerations that begin from an engine power setting that is below the steady state staging point. Since the power setting is below the steady state staging point, the combustion chamber will operate in the pilot injection only mode, and acceleration may result in an increase in nvPM emissions without a change in power. characteristic of the fuel provided by the methods of the present application. In other examples however, the acceleration may be from any mode of steady state operation, whether operation in operating ranges with pilot injection only or pilot and main injection. The staged combustion system 64 may be configured to switch between the pilot injection only operating range and the pilot and main injection operating ranges at the same or greater engine power in the mode. acceleration compared to the steady state mode. In other words, the staged combustion system 64 is controlled (e.g. by the EEC) using a stage point which is at the same or higher engine power setting in the acceleration mode compared to the at-speed mode. the stabilized state. This means that the staging point is not reduced to a lower engine power setting during acceleration mode, thereby avoiding or reducing any increase in HC or CO emissions, or requiring all injectors to be enriched with fuel. main fuel for all engine powers. More particularly, in certain examples, the staging point can be the same both in the acceleration operating mode and that in the stabilized state. The staging point used in the acceleration mode may be referred to as the "acceleration staging point", while the staging point used during steady state operation may be referred to as the "acceleration staging point". cruising regime”. In this example therefore, the throttle stage point is at a power setting greater than or equal to the cruise stage stage point. In other examples, the staged combustion system may be designed to switch between the pilot injection only operating range and the pilot and main injection operating ranges at lower engine power in the acceleration mode compared to the steady state mode. The staged combustion system 64 is controlled (e.g. by the EEC) in this example using a stage point which is at a lower engine power setting in the throttle mode compared to the steady state mode. In this example, the throttle staging point is greater than a default throttle staging point at which the combustion system would be controlled if fuel of a different characteristic could not be supplied to the combustion chamber. combustion. The throttle staging point is therefore reduced (relative to the cruise staging point) by an amount less than the reduction to the default throttle staging point that would be found in systems known. The fuel delivery regulator 306 may in some examples be designed to distribute fuel to the pilot fuel injectors 313 during pilot injection only operation in the acceleration mode which has a different fuel characteristic than that of the pilot fuel injector 313. fuel delivered to the main fuel injectors 314 during pilot and main injection operation in the engine's steady-state operating mode. In this example, fuel of a different characteristic (e.g., SAF-rich fuel) is delivered to the pilot fuel injectors 313 when operating in the pilot injection only mode during engine acceleration. Fuel that is associated with a high nvPM level (e.g. lean SAF fuel) is then supplied to the main fuel injectors 314 during pilot and main injection steady state operation where nvPM emission is inherently lower. This allows a limited amount of fuel within the first fuel source (eg, SAF-rich fuel) to be used more efficiently by using it selectively during engine acceleration with pilot injection only. In this example, fuel of a different fuel characteristic is delivered to the pilot fuel injectors 313 in only part of the throttle mode of operation. Once the staging point is reached in the throttle mode, fuel delivery may revert to the same as that used during steady-state operation (e.g. it may revert to using a fuel low in SAF). In other examples, the SAF-rich fuel (or a fuel of another different characteristic) may be used at all engine powers in the acceleration mode (e.g. in pilot injection only and pilot injection operation). and main). In one example, the fuel delivery regulator 306 is configured to distribute fuel during the pilot injection only operating range in the throttle mode having a fuel characteristic determined based on a fuel-to-air ratio in the chamber. combustion 16. In this example, as the fuel-air ratio decreases, the proportion of fuel associated with low nvPM production delivered to the pilot fuel injectors can also be decreased. Since the fuel-air ratio reduces (e.g. because as the engine body speed increases the mass flow rate of air inside the combustion chamber also increases) the amount of a low-producing fuel of nvPM (e.g., SAF-rich fuel) can also be reduced. This allows the low nvPM producing fuel to be used during a portion of the acceleration mode of operation in which its benefits are greater, thereby better utilizing a limited availability of this fuel. In other examples, the fuel characteristics may be determined based on another appropriate control parameter that is related to the formation of combustion chamber nvPM other than the fuel-air ratio. In other examples, the fuel delivery regulator 306 is configured to switch fuel delivery to the fuel injectors (the eVou pilot main fuel injectors 313, 314) to that having a different fuel characteristic at a starting point. of one period of operation in the acceleration mode. In this example therefore, the fuel composition may be switched to another composition (e.g., switched to a SAF-rich fuel) at the start of an acceleration period. In some examples, fuel regulator 306 may continue to dispense the same fuel throughout the fuel mode. acceleration operation. In other examples, fuel delivery regulator 306 is configured to revert to delivering fuel having the same fuel characteristic as that delivered in the steady state mode after a transition to pilot injection operation. and main. The fuel composition is therefore switched back to the default steady state fuel characteristics once the staging point has passed. This can also help restrict the use of fuel associated with low nvPM emissions to only part of the acceleration mode, allowing a limited supply of this fuel to be used more efficiently. To distribute fuel to the pilot and / or main fuel injectors 313, 314 during the acceleration mode of the combustion system which is different from that distributed to the combustion chamber during the steady state mode of operation the fuel delivery regulator 306 is configured to selectively distribute fuel from the first and second fuel sources 302, 304 as described above. In one example, the fuel distribution regulator 306 may include a fuel mixer 318 as illustrated in [Fig.17]. Similar to that described above, the fuel mixer is configured to receive a supply of fuel from both the first and second fuel sources 302, 304 and to output fuel from the first fuel source 302 , fuel from the second fuel source 304, or a mixture thereof. This mixed fuel is supplied to the pilot fuel injectors 313, with fuel from a single fuel source (the second fuel source 304 in the example above) being supplied to the main fuel injectors 314, in others Examples, the mixer may be designed to supply both the main and pilot fuel injectors 313, 314 with an appropriate fuel mixture. In yet other examples, the fuel delivery regulator 306 of any of the examples described herein may be used to distribute fuel during at least a portion of the acceleration mode (to the pilot fuel injectors and / or main) that is different from that distributed during at least part of the steady-state mode of operation (e.g., those using independent regulators for the first and second fuel sources or a pilot regulator designed to switch between fuel sources). As indicated above in connection with the example illustrated in [Fig.17], the dependence of the proportion of fuel coming from the first fuel source 302 compared to that coming from the second fuel source 304 compared at the fuel flow rate can be determined based on a desired achieved level of nvPM at a particular fuel flow rate. For example, the The proportion of fuel delivered from the first fuel source 302 compared to that from the second fuel source 304 can be determined according to a desired obtained level of nvPM at a particular fuel flow rate during the fuel mode. acceleration. The quantity of fuel supplied from each source may be determined such that the nvPM value does not exceed a predetermined threshold or such that the production of nvPM is minimized over a period of operation of the gas turbine engine. as described above. In any of the examples described here this can be achieved by appropriate control of the mixer 318, or any of the other fuel delivery regulators 306 described herein. In some examples, the proportion of fuel distributed from the first fuel source 302 compared to that distributed from the second fuel source 304 during the acceleration mode of the combustion system can be determined at least partially according to one or any number of elements: a) the quantity of fuel having the first fuel characteristic and the quantity of fuel having the second fuel characteristic available for a proposed flight. This can for example be provided by an estimation or measurement of the quantity of fuel in the fuel tanks 53, 55 composing the first and second fuel sources 302, 304: b) the total fuel requirement amount for the pilot fuel injectors 313 during pilot-only operation for the entire flight within an operating range in which fuel is supplied from the first fuel source 302; and or c) a fuel composition limit parameter (e.g. a certification limit, fuel composition available for refueling, or aircraft / engine limits). [Fig.30] illustrates a method 4030 of operating a gas turbine engine according to the example above in which fuel of a different characteristic is distributed to the fuel injectors (e.g. pilot fuel injectors and / or main 313, 314) during at least part of the acceleration mode compared to during at least part of the mode in the stabilized state. The method 4030 includes regulating 4032 the delivery of fuel to the pilot and / or main fuel injectors 313, 314 from the first fuel source 302 containing a first fuel having a first fuel characteristic and the second fuel source 304 containing a second fuel having a second fuel characteristic, as described above using the fuel delivery regulator 306 of any example herein. The second characteristic of fuel is different from the first as described above. The method further includes operating 4034 the staged combustion system 64 in an acceleration mode in which acceleration of the engine from a steady-state operating mode is caused. The method 4030 further includes delivering fuel 4036 to the fuel injectors (e.g. pilot and / or main fuel injectors 313, 314), during at least a portion of operation in the acceleration mode, having a fuel characteristic different from that of the fuel delivered to the fuel injectors (for example the pilot and / or main fuel injectors 313, 314) during at least part of the engine's steady state operating mode. Any of the features described above in connection with the examples in which a different fuel is provided in the acceleration mode compared to a steady state cruise mode of operation may be incorporated into the process 4030 [Fig.30], even if they are not repeated here. D inati int di lat istics d In other examples of the present application, the staging point at which the staged combustion system 64 is operated may be determined based on a determination of the characteristics of the fuel with which it is supplied. The inventors have determined that the staging point can be based on fuel characteristics to advantageously utilize the particular fuel being supplied to the engine 10. In some examples, the aircraft shown in [Fig.4] may be designed to have only one fuel source. In such an example, the fuel tanks 53, 55 can be fluidly connected to form a single fuel source on board the aircraft. The aircraft can therefore carry fuel having the same fuel characteristics, rather than fuels having different characteristics. [Fig.31] illustrates an example of a staged combustion system 64 having characteristics corresponding to those of [Fig.5], but which is supplied with fuel from a single source on board the aircraft 1. In this example, fuel is received by the fuel distribution regulator 306 via a fuel pump 308, from a single fuel source, for example fuel tanks 53, 55 shown in [Fig.4]. The characteristics common with the example of [Fig.5] are marked in this sense, and will therefore not be described again. In the example illustrated in [Fig.31], the gas turbine engine 10 further comprises a fuel characteristic determination module 330. The fuel characteristic determination module 330 is configured to determine one or more several characteristics of the fuel being distributed in the fuel distribution regulator 306. In the example described, the fuel characteristics determination module 330 is in communication with a sensor device 332, which is configured to perform a measurement on the fuel flowing to the fuel delivery regulator 306 to determine the characteristics of the fuel. Sensor device 332 may take a number of different forms, and may operate in accordance with any of the examples of determining fuel characteristics described herein. In other examples, the fuel characteristic determination module 330 may receive signals from a sensor device located elsewhere on board the aircraft that is configured to perform a fuel characteristic measurement. In yet other examples, the fuel characteristics determination module 330 may obtain fuel characteristics from sources other than a sensor device as described elsewhere herein, for example the fuel characteristics may be received by via a data communications channel, or user input. The EEC 42 (which may be more generally referred to as a "controller") is in communication with the fuel characteristic determination module 330 such that it can receive the fuel characteristics of the fuel being dispensed into the fuel delivery regulator 306. In the present example, the fuel characteristics determination module 330 is shown separately from the EEC, but in other examples they may be combined. The controller 42 is configured to determine the staging point at which the staged combustion system is switched between its pilot injection only operation and pilot and main injection operation. The staging point is determined based on the fuel characteristic(s). Once the staging point is determined in this way it is used by the controller 42 to control the operation of the staged combustion system 64, for example it is used to control the operation of the fuel distribution regulator so that a suitable fuel flow is provided to pilot manifold 309, or pilot manifold 309 and main manifold 310 for operation in pilot injection only and pilot and main injection modes respectively. The inventors have determined that staged combustion system 64 can be advantageously controlled based on the characteristics of the fuel it is being fed. In particular, the staging point can be chosen to advantageously use the characteristics of the fuel being distributed into the engine. In one example, the staging point is determined based on the fuel characteristic(s) indicating that the fuel is associated with a low level of nvPM production (e.g. low compared to fossil kerosene, under corresponding combustion conditions). This may allow the staging point to be adjusted so that it matches an engine operating condition that would otherwise lead to high levels of nvPM production. For example, as shown in Figures 23 and 25, if a low nvPM producing fuel is determined to be being distributed into the combustion chamber the staging point can be adjusted to reduce CO and HC production, without result in a disadvantageous increase in nvPM production that would otherwise occur if a fuel producing a relatively high level of nvPM had been used. In some examples, the fuel characteristic on which the staging point determination is based may be the percentage of sustainable aviation fuel (SAF) present in the respective fuel. As noted above, compared to fossil kerosene, SAF yields significantly less nvPM, and therefore can be used to mitigate changes in staging point that would otherwise increase nvPM production. In some examples, fossil kerosene can be processed to remove aromatic components, particularly naphthalenes, to produce a largely paraffinic fuel of fossil origin that would be a fuel producing low levels of nvPM. Other fuel characteristics may therefore be associated with a low nvPM level, such as percent aromatic content or naphthalene content. In other examples, the fuel characteristic(s) by which the staging point is determined may include an aromatic hydrocarbon content of the fuel, and / or a naphthalene content of the fuel. These characteristics can also indicate the level of nvPM that will be produced by the fuel, and help determine the staging point in this sense. The controller 42 may be configured to determine the staging point such that the staging point associated with the fuel that is a low nvPM producing fuel corresponds to a higher engine power setting compared to the staging point. associated with one or more fuel characteristics that indicate that the fuel is associated with relatively higher nvPM production. In other words, the staging point can be increased at higher engine power for a fuel that is associated with lower nvPM production compared to a fuel associated with higher nvPM production. The engine power at which the staging point occurs can therefore be increased with decreasing fuel nvPM production. As indicated above in conjunction with Figures 23 and 25, this increase in the point staging can help reduce CO and HC production, without significantly increasing nvPM. The low nvPM fuel may be a fuel that produces a lower nvPM emission compared to fossil kerosene under corresponding combustion conditions. The fuel with a low nvPM level may be a fuel rich in SAF, which has at least one SAF content, and preferably an SAF content greater than 10%, or more preferably equal to or greater than 50%. In some examples, the staging point determined based on fuel characteristics may be a cruise-rate staging point with which the combustion system 64 is controlled during steady-state cruise-rate operation. of the motor. The control device 42 can be configured to determine the staging point so that it corresponds to an engine power setting which causes a switch between operation with pilot injection only and with pilot and main injection during the cruise speed at the stabilized state. The staging point determined by the control device 42 as a function of the fuel characteristics can therefore create a limit between a first operating range at engine cruise speed and a second operating range at engine cruise speed. The staging point can be selected to define these two operating points at cruise speed if it is established that a fuel with a relatively low nvPM level is used (for example, a fuel with a high SAF content and / or with a low content of aromatic compounds and / or naphthalene). This may allow pilot injection only operation at cruise to advantageously utilize its low HO and CO output, while avoiding the increase in nyvPM that would otherwise occur if a high nvPM fuel was supplied by the engine. (see figures 27 and 28, and the associated description above). The first operating range in cruise mode may correspond to operation of the aircraft in a later part of a cruise segment of a flight, and the second operating range in cruising mode may correspond to operation of the the aircraft in a relatively earlier part of the cruise segment. As noted above, this may allow the staged combustion system 64 to switch to pilot injection only operation during a later portion of a cruise segment (e.g. a cruise segment at a constant altitude). In another example, the first operating range in cruising speed may correspond to operation in subsonic cruising speed in the stabilized state of the engine and the second operating range in cruising speed may correspond to operation in cruising speed. supersonic in the stabilized state of the engine. In these two examples, the staging point is selected such that pilot injection only operation occurs at low engine power cruise operation, for example either in a later part of a cruise segment or during subsonic cruise speed. Steady-state supersonic cruise operation may occur before steady-state subsonic cruise operation, or vice versa. In some examples, steady-state subsonic cruise operation may be subsonic cruise operation over land, while steady-state supersonic cruise operation may be in-state operation. supersonic cruise regime above water. According to another example, the staging point determined according to the fuel characteristics may be an engine acceleration staging point according to which the staged combustion system is controlled during an engine acceleration condition. In such an example, the controller may be configured to select the engine throttle staging point such that it is the same as that used at cruise speed. Indeed, this is achievable if it has been established that a fuel with a low nvPM level is in use (for example, a fuel with a high SAF content and / or low aromatic content and / or low in naphthalene). As noted above, this may allow acceleration to occur (at least in part) while remaining in the pilot-injection-only mode despite increasing fuel flow to provide the acceleration. This can help avoid an increase in HC and CO production that would otherwise occur by increasing the staging point during engine acceleration. In certain examples, the control device 42 can be designed to determine the staging point in cruising and / or acceleration mode, and can be designed to determine staging points both for operation on - personal / subsonic and for earlier / later parts of a cruise segment. [Fig.32] illustrates a method 4038 of operating a gas turbine engine for an aircraft. The method can be implemented by the apparatus of [Fig.31]. The method 4038 includes determining 4040 one or more fuel characteristics of a fuel being dispensed into the combustion system; determining 4042 a staging point defining the point at which the staged combustion system is switched between pilot injection only operation and pilot and main injection operation depending on the determined fuel characteristic(s); and the control 4044 of the staged combustion system according to the determined stage point. The fuel characteristic(s) may indicate that the fuel is associated with a low level of nvPM production compared to fossil kerosene as indicated above. The fuel characteristic(s) includes any one or more of: (i) a percentage of sustainable aviation fuel in the fuel; (ii) aromatic hydrocarbon content of the fuel; and / or (iii) naphthalene content of the fuel. Other fuel specifications may be used. The staging point determination 4042 may include the staging point determination 4046 such that a staging point associated with one or more fuel characteristics that indicate that the fuel is associated with low nvPM production corresponds to at a higher engine power setting relative to the staging point associated with one or more fuel characteristics that indicate that the fuel is associated with relatively higher nvPM production. The stage point determined 4042 by the method of [Fig.32] may be a cruise stage stage point, and may be used during a later part of a cruise segment of a flight, or during operation in subsonic cruise mode of an aircraft with supersonic capacity as indicated in the various examples above. Any of the features described above in connection with the examples in which the staging point is determined according to one or more fuel characteristics may be incorporated into method 4038 of [Fig.32]. In the examples above, the staging point is determined based on the nvPM production characteristics of the fuel. This is, however, only an example, and other fuel characteristics may be taken into account to determine a suitable staging point to take advantage of the characteristics of the fuel being delivered into the combustion chamber. In other examples, the controller 42 illustrated in [Fig.31] is additionally, or alternatively, designed to determine a staging ratio according to the fuel characteristic(s). As noted elsewhere herein, the controller 42 is configured to control the fuel delivery regulator 306 (and therefore the staged combustion system 64) according to a stage ratio. The staging ratio defines the ratio between the fuel flow of the pilot fuel injector 313 and the fuel flow of the main fuel injector 314. The inventors have determined that the staging ratio can be chosen to advantageously exploit certain fuel characteristics of the fuel which is supplied to the fuel delivery regulator by intelligently selecting the ratio based on the fuel characteristics. For example, the staging ratio may be determined based on the fuel characteristic(s) indicating that the fuel is associated with a low level of nvPM production (e.g., low compared to fossil kerosene, under conditions of corresponding combustion). This may allow the staging ratio to be adjusted to reduce CO and HC production in a manner that would otherwise lead to high levels of nvPM production as shown in various examples above. In some examples, the fuel characteristic(s) upon which the determination of the staging ratio is based may be the percentage of sustainable aviation fuel (SAF) present in the respective fuel. As noted above, compared to fossil kerosene, SAF yields significantly less nvPM, and therefore can be used to mitigate changes in staging ratio that would otherwise increase nvPM production. In some examples, fossil kerosene can be processed to remove aromatic components, particularly naphthalenes, to produce a largely paraffinic fuel of fossil origin that would be a fuel producing low levels of nvPM. Other fuel characteristics may therefore be associated with a low nvPM level, such as percent aromatic content or naphthalene content. In other examples, the fuel characteristic(s) by which the staging ratio is determined may include an aromatic hydrocarbon content of the fuel, and / or a naphthalene content of the fuel. These characteristics can also indicate the level of nvPM that will be produced by the fuel, and help determine the staging ratio in this sense. The controller 42 may be configured to determine a transition staging ratio that allows the staged combustion system 64 to be operated in a transition operating range between the pilot injection only operating range and the of pilot and main injection operation as shown above in conjunction with Figures 27 and 28. The transition staging ratio can be determined depending on the fuel characteristic(s) indicating that the fuel being dispensed into the fuel regulator fuel distribution 306 is associated with a low nvPM production level (for example depending on SAF content, aromatic content or naphthalene content; low nvPM production being relatively less important than for kerosene fuel fossil). The transition staging ratio may be different from a pilot and main injection staging ratio according to which the control device 42 controls the staged combustion system during pilot and main injection operation. The pilot injection staging ratio and main can be a default report, and can be determined according to known techniques. As indicated above in conjunction with Figures 27 and 28, by configuring the staged combustion system 64 to operate in the transition range between pilot injection only operation and pilot and main injection operation the emission quantity of CO and HC implementation in this range of engine power settings may be reduced. By detecting that the combustion chamber is being supplied with a fuel associated with low nvPM production, the controller can determine that the transition staging ratio can be used to reduce CO and HC emissions, without causing a excessive increase in nvPM production. Any of the transition staging ratio characteristics described above may be incorporated into the examples described herein in which the transition staging ratio is selected based on the fuel characteristic(s). For example, during the transition phase the transition staging ratio may vary with a changing motor power setting. In one example, the transition staging ratio exhibits continuous variation with changing motor power within the transition operating range. This can ensure a smooth transition between the staging ratio in the pilot injection only and the pilot and main injection operating range. The continuous variation may be such that the proportion of the total fuel flow to the fuel injectors (i.e. the total being distributed to the pilot and main fuel injectors) can be attributed to the fuel flow to pilot fuel injectors 313 decreases with increasing engine power during the transition operating range. The proportion of the total fuel flow to the fuel injectors that can be attributed to fuel flow to the main fuel injectors 314 instead increases with increasing engine power in the operating range of transition. In other examples, the transition staging ratio has a constant intermediate value which is different from the pilot and main injection staging ratio. The transition staging ratio can be between the pilot injection ratio only and the pilot and main injection ratio. This therefore allows a more gradual transition between the pilot injection ratio alone and the pilot and main injection ratio. For example, the transition staging ratio may be 70:30, which is between 100:0 in the pilot injection only range, and a pilot and main injection staging ratio which may be 8:80 p.m. or 30:70 p.m. In yet other examples, the transition staging ratio varies between a series of constant intermediate values, each different from the ratio pilot and main injection staging. Furthermore, each of the intermediate ratios can be included between that of the pilot injection range alone and that of the pilot and main injection range. For example, the transition staging ratio can vary between a series of values ​​of 80:20, 60:40 and 40:60. The ratio of pilot and main injection can be 20:80 in this example. The intermediate staging ratios can thus decrease towards the pilot and main injection ratio (i.e. a progressively smaller proportion of the total fuel is supplied to the pilot fuel injectors 313, and a proportion progressively more of the total fuel is supplied to the main fuel injectors 314). This is, however, only an example, and any other number and values ​​of intermediate transition staging ratios may be used. In some examples, the transition staging ratio may have a continuous variation with motor power in one part of the transition operating range and may have one or more constant values ​​in another part of the transition operating range . The examples above can therefore be combined. In other examples, the staging ratio may vary with motor power over the entire transition operating range, or may have one or more constant values ​​with motor power over the entire operating range. of transition. [Fig.33] illustrates an example of a 4050 method of operating a gas turbine engine. The method 4050 includes: determining 4052 one or more fuel characteristics of a fuel being dispensed into the staged combustion system 64; determining 4054 a staging ratio defining the ratio between the fuel flow of the pilot fuel injector and the fuel flow of the main fuel injector; and the control 4056 of the staged combustion system 64 according to the determined stage ratio. Method 4050 may be carried out by the apparatus shown in [Fig.31] (in addition to, or alternatively to, method 4038 in which a staging point is determined). The fuel characteristic(s) may indicate that the fuel is associated with a low nvPM production level compared to fossil kerosene as noted above. The fuel characteristic(s) includes any one or more of: (i) a percentage of sustainable aviation fuel in the fuel; (ii) aromatic hydrocarbon content of the fuel; eV / or (iii) a naphthalene content of the fuel. Other fuel specifications may be used. Control 4056 of staged combustion system 64 may include control 4058 of staged combustion system 64 during the pilot and main injection operating range according to an injection stage ratio pilot and main as indicated above, The determination 4054 of the staging ratio may include the determination 4060 of a transition staging ratio. The control 4056 of the staged combustion system 64 may then include the control 4062 thereof so that it is actuated in a transitional operating range between the pilot injection only operating range and the pilot injection operating range and main as indicated above. The transition staging ratio can be determined according to any of the examples given above. Any of the features described above in connection with the examples in which the staging ratio is determined according to one or more fuel characteristics may be incorporated into the method 4050 of [Fig.33]. In the examples above, the staging ratio is determined based on the nvPM production characteristics of the fuel. This is, however, only an example, and other fuel characteristics may be taken into account to determine a suitable staging point to take advantage of the characteristics of the fuel being delivered into the combustion chamber. Operating conditions Regulation of fuel delivery according to any of the examples described herein is understood to be suitable for operation under normal operating conditions of the gas turbine engine 10 or aircraft 1. During operating conditions abnormal fuel distribution to the combustion chamber 16 as described here (for example to regulate the nvPM level) can be overridden. The regulation of fuel distribution described here is therefore applicable during at least part of the operation of the associated gas turbine engine, for example when the availability of fuel does not prevent it, either due to tank capacity limits, or in unexpected abnormal operating conditions. For example, in abnormal operation, particularly after an engine failure, a requirement to transfer fuel from one wing tank to another wing tank (in order to maintain a lateral position of the aircraft center of mass acceptable) can override the regulation of fuel delivery according to the present application, particularly in cases requiring a transfer of fuel from the first fuel source 302 to the second fuel source 304 or vice versa. However, if both wing fuel tanks are part of the same fuel source (i.e., both are part of the first fuel source 302 or both are part of the second fuel source 302 304) and are fluidly interconnected either directly or through one or more additional fuel tanks which are also part of the same fuel source (respectively the first or second fuel source) then the fuel regulation described herein can continue to operate despite such abnormal conditions. When the fuel delivery regulator 306 is configured to switch between sources or mix fuel from the first and second fuel sources 302, 304, the switching or mixing may be overridden if one of the fuel sources is depleted due to leakage or other unexpected fuel reasons (e.g. incorrect fuel loading). In some or all of the examples, particularly that shown in [Fig.9], a fuel bypass may be provided in which fuel can be distributed between the first and second fuel sources in the event of an abnormal operating condition. For example, in the example of [Fig.9] a fuel bypass can be arranged upstream of the fuel distribution regulator 306, or between fuel tanks of the first and second fuel sources 302, 304, which provides emergency interconnection between the first and second fuel sources. The pilot regulator 306a and the main regulator 306b are therefore capable of being supplied with fuel coming from either of the first and second fuel sources 302, 304 in the event of failure of one of the fuel sources during operation. Fuel supply regulation may further be overridden during a portion of an aircraft mission if insufficient quantities of the first and second fuels may be contained in the aircraft fuel tanks. For some missions, the total fuel loading requirement proposed for a flight may dictate the minimum quantity with which each fuel tank must be filled, and this may override some other means of fuel delivery control in the various examples. described here. Carb allocation calculation The present application further relates to a method for determining a fuel allocation for an aircraft. The method makes it possible to determine a fuel allocation according to which fuel is loaded on the aircraft 1 to carry out a proposed flight or a proposed mission. The aircraft for which the method is used may be that illustrated in [Fig.4], which includes a first fuel source 302 configured to contain a first fuel having a first fuel characteristic and a second fuel source 304 configured to containing a second fuel having a second fuel characteristic, the second fuel characteristic being different from the first. As indicated above, the aircraft 1 includes one or more gas turbine engines 10 powered by fuel from the first and second fuel sources 302, 304. The gas turbine engines 10 each include a fuel distribution regulator. fuel 306 designed to provide fuel from each fuel source, or a mixture thereof, and a staged combustion system 64, as illustrated in [Fig.5], or as described elsewhere in the present invention. A method 4070 of determining a fuel allocation is illustrated in [Fig.34]. The method 4070 includes obtaining 4072 a proposed mission description including a list of operating points for the gas turbine engine(s) 10 of the aircraft 1 during an operating mission. The operating point list includes information on the operation of the gas turbine engines 10 of the aircraft 1 which is expected for a particular planned period of operation for which fuel is to be loaded on the aircraft. The operating point list may include a series of information from which the expected impact of gas turbine engine nvPMs during each part of the operating mission may be determined for fuels of different characteristics used. The operating points of the mission description may include any one or more of: one or more conditions under which the gas turbine engines 10 must operate (e.g., location and / or ambient air conditions expected for the specific mission), one or more fuel flow values ​​corresponding to an operating point, and an operating time at a corresponding operating point. The operating points may therefore indicate that the mission includes, for example, a period of engine operation in a cruise operating condition, under certain ambient conditions, and in which a specifi...

Claims

Demands

1. Gas turbine engine (10) for an aircraft (1), comprising: a staged combustion system (64) having fuel injectors pilots (313) and main fuel injectors (314), the system staged combustion (64) capable of operating within a range of operating pilot-only injection operation in which a fuel is distributed only to pilot fuel injectors (313), and a pilot and main injection operating range in which a fuel is distributed at least to the fuel injectors principals (314); and a fuel distribution regulator (306) designed to control fuel distribution to the pilot fuel injectors and main components (313, 314), the fuel distribution regulator (306) being designed to receive fuel from a primary source of fuel (302) containing a first fuel having a first ca- fuel characteristic and a second fuel source (304) containing a second fuel having a second characteristic of fuel, the second fuel characteristic being different from the first, in which: the staged combustion system (64) is designed to switch between the operating range with pilot injection only and the operating range pilot and main injection operation at a single stage point which corresponds to a cruising operating mode in the state The engine is stabilized, the gearing point defining a boundary between a first operating range in cruising mode engine (320a) and a second operating range at engine speed of motor cruising (320b); and The fuel distribution regulator (306) is designed to distribute fuel to the pilot fuel injectors (313) for at a portion of the first operating range in steady state motor cruise (320a) having a fuel characteristic different from the fuel distributed to the pilot fuel injectors and / or main (313, 314) during the second operating range in engine cruising mode (320b).

2. Gas turbine engine (10) according to claim 1, wherein the The first fuel characteristic is associated with a level of nvPM production which is lower than that of the second characteristic fuel type, and the fuel distributed to the injectors of pilot fuel (313) during the first operating range in engine cruising speed (320a) is associated with a level of nvPM production which is lower than that of the fuel distributed to pilot and / or main fuel injectors (313, 314) during at minus a portion of the second operating range in steady state motor cruise (320b).

3. Gas turbine engine (10) according to claim 2, wherein the The first fuel characteristic corresponds to a proportion of SAF peri-fuel within the respective fuel compared to the second fuel characteristic, and the fuel distributed during the first operating range in cruising mode engine (320a) has a higher proportion of SAF compared to fuel distributed during the second operating range in engine cruising speed (320b).

4. Gas turbine engine (10) according to any one of claims previous, in which the first operating range in regime engine cruising (320a) corresponds to an operation of the aircraft (1) in a later part of a cruise segment of a flight, and the second operating range in cruise mode of engine (320b) corresponds to an operation of the aircraft (1) in a relatively anterior part of the cruising segment.

5. Gas turbine engine (10) according to any one of claims 1 at 3, in which the first operating range in steady state of engine cruising (320a) corresponds to operation in steady state subsonic cruising in the stabilized state of the engine (10) and the second engine cruising speed operating range (320b) corresponds to supersonic cruising operation at the stabilized state of the engine (10).

6. Gas turbine engine (10) according to any one of claims previous, in which: The fuel distribution regulator (306) includes a mixer fuel (318) designed to receive a fuel supply at both from the first and second fuel sources (302, 304) and to extract fuel from the first fuel source (302), from the fuel from the second fuel source (304), or a mixture of these; and the fuel mixer (318) is designed to distribute fuel to the pilot fuel injectors (313), and possibly to the in- main fuel injectors (314).

7. Gas turbine engine (10) according to any one of claims previous, in which the proportion of fuel distributed from the first fuel source (302) compared to that from the second fuel source (304) is determined according to a level desired production of nvPM at a fuel flow rate per- particular within the first operating range in steady state motor cruise (320a), and can be determined such that the nvPM production does not exceed a predetermined threshold or such so that nvPM production is reduced to a minimum on a operating period of the gas turbine engine (10).

8. Gas turbine engine (10) according to any one of claims previous, in which the proportion of fuel distributed from the first fuel source (302) compared to that from from the second fuel source (304) during the first range of engine cruising speed operation (320a) is determined at least partially according to one or more of the following elements among: a) the quantity of fuel having the first fuel characteristic and the second characteristic of fuel available for a flight propose ; b) the total fuel requirement quantity for the injectors pilot fuel during pilot-only injection operation for the entire flight within an operating range in which fuel is supplied from the first source of fuel (302); and / or c) a fuel composition limit parameter.

9. Method (4014) of operating a gas turbine engine (10) for an aircraft (1), the gas turbine engine (10) comprising a system of staged combustion (64) having pilot fuel injectors (313) and main fuel injectors (314), the combustion system stepped (64) capable of operating within an operating range at Pilot injection only, in which fuel is distributed only to the pilot fuel injectors (313), and a range of operation pilot and main injection operation in which the fuel is distributed at least to the main fuel injectors (314), the method (4014) comprising: the regulation (4016) of a fuel distribution to the injectors of pilot and main injection fuel (313, 314) from a first fuel source (302) containing a first fuel having a first fuel characteristic and a second source of fuel (304) containing a second fuel having a second ca- fuel characteristic, the second fuel characteristic being different from the first one; the switch (4018) between the injection operating range pilot only and the operating range with pilot and main injection at a staging point during a mode of operation in regime cruising in the engine's stabilized state to define a first range engine cruising speed (320a) and a second operating range in cruising mode motor (320b); and fuel distribution (4020) to fuel injectors pilots (313) during at least part of the first operating range engine cruising speed operation (320a) having a characteristic fuel type different from the fuel distributed to the injectors pilot and / or main fuel (313, 314) during the second range of engine cruising speed operation (320b).

10. 0 Method (4014) according to claim 9, wherein the first ca- fuel characteristic is associated with a production level of nvPM which is lower than that of the second fuel characteristic, and the fuel distributed to the pilot fuel injectors (313) during the first operating range in cruise mode of motor (320a) is associated with an nvPM production level which is lower than that of the fuel distributed to the fuel injectors pilots and / or principals (313, 314) for at least part of the second operating range in cruising mode motor (320b).

11. 1 Method (4014) according to claim 10, wherein the first ca- fuel characteristic corresponds to a higher proportion of SAF within the respective fuel compared to the second characteristic fuel teristics, and the fuel distributed during the first engine cruising operating range (320a) has a higher proportion of SAF compared to the fuel distributed during the second operating range in cruise mode of motor (320b).

12. 2 Method (4014) according to any one of claims 9 to 11, in which the first operating range in cruising mode of engine (320a) corresponds to an operation of the aircraft (1) in a later part of a cruise segment of a flight, and the second beach engine cruising speed (320b) corresponds to an aircraft operation (1) in a relatively an- exterior of the cruise segment.

13. 3 Method (4014) according to any one of claims 9 to 12, in which : the regulation (4016) of a fuel distribution includes the dis- fuel contribution from the first fuel source (302), of fuel from the second fuel source (304), or from a mixing these using a fuel mixer (318); and the fuel mixer (318) is designed to distribute fuel to the pilot fuel injectors (313), and possibly to the in- main fuel injectors (314).

14. 4 Method (4014) according to any one of claims 9 to 13, in which the proportion of fuel distributed from the first fuel source (302) compared to that from the second fuel source (304) is determined according to a level obtained desired production of nvPM at a particular fuel flow rate within the first operating range in cruising mode of motor (320a), and can be determined such that the level of nvPM does not exceed a predetermined threshold or in such a way that the nvPM production should be reduced to a minimum over a period of operation. operation of the gas turbine engine (10).

15. 5 Method (4014) according to any one of claims 9 to 14, in which the proportion of fuel distributed from the first fuel source (302) compared to that from the second fuel source (304) during the first operating range in engine cruising mode (320a) is determined at least by- tiellement according to one or more of the following elements: a) the quantity of fuel having the first fuel characteristic and the second characteristic of fuel available for a flight propose ; b) the total fuel requirement quantity for the injectors pilot fuel during pilot-only injection operation for the entire flight within an operating range in which the fuel is supplied from the first fuel source; € c) a fuel composition limit parameter.

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