Aircraft Propulsion System

Abstract

A propulsion system (2) for an aircraft (1) includes a gas turbine engine (10) and a fuel tank (50, 53) arranged to store fuel to power the gas turbine engine (10), in which the stored fuel comprises at least a proportion of a sustainable aviation fuel – SAF – having a density of between 90% and 98% of the density, ρK, of kerosene and a calorific value of between 101% and 105% of the calorific value PCK, of kerosene. The gas turbine engine 10 includes a combustion chamber 16; and a fuel pump 108 arranged to supply fuel to the combustion chamber 16 at an energy flow rate, C, the pump 108 being arranged to deliver fuel at the outlet at a volumetric flow rate, Q, the percentage of fuel passing through the pump that is not supplied to the combustion chamber 16 being called a spillage percentage.The fuel comprises X% of SAF, where X% is in the range of 5% to 100%, and has a density, ρF, and a calorific value PCF. The propulsion system 2 is arranged such that: the fuel change spillage ratio, Rd: is equal to: Figure for the abbreviation: 13.

Description

Description Title of the invention: Propulsion System This description relates to aircraft propulsion systems and aircraft operating methods involving the management of various types of fuel, including refueling control, and methods for modifying aircraft to enable the implementation of such operating methods. This description further relates to methods for, and systems for, determining one or more fuel characteristics of an aviation fuel. This description also relates to the operation of an aircraft according to the determined fuel characteristics. There is an expectation in the aviation industry of a trend towards the use of fuels other than the traditional kerosene-based jet fuels generally used today. According to a first aspect, a method is provided for refueling an aircraft comprising a gas turbine engine and a fuel tank arranged to supply fuel to the gas turbine engine, the method comprising: obtaining the amount of energy required for a planned flight profile; obtaining a fuel calorific value available to the aircraft for refueling; calculating the amount of available fuel needed to supply the required energy; and refueling with the calculated quantity of fuel available. Knowledge of the calorific value(s) of the fuel(s) available to an aircraft can enable more efficient, tailored control of the propulsion system. For example, switching to a fuel with a higher calorific value can allow for a smaller quantity (mass or volume) of fuel to meet an aircraft's energy requirements for a flight. Furthermore, since more power is needed to lift a greater mass of fuel, carrying enough fuel for the planned flight (including a safety margin above the expected energy demand), but not completely filling the tank(s), can provide a performance bonus by reducing the aircraft's takeoff weight. Knowledge of fuel calorific value can therefore be used as a tool to improve aircraft performance, e.g., by avoiding carrying excess fuel weight. Calculating the amount of available fuel needed to provide the required energy may include obtaining an energy content from fuel already in the fuel tank(s) — e.g., that remaining from a refueling event previous - and the subtraction of this from the determined amount of energy required for the planned flight profile. Obtaining the calorific value of the fuel available for aircraft refueling may include receiving calorific value data input for the available fuel, e.g., through a user interface. Obtaining the calorific value of the fuel available to the aircraft may include receiving calorific value data in an electronic communication, e.g., from scanning a barcode or QR code associated with the fuel supply, or receiving a message from the fuel supplier or refueling facility. Determining the calorific value of the fuel available to the aircraft may involve a chemical and / or physical determination of the calorific value of the available fuel, possibly by performing one or more of the following: 1. the identification of a tracer in the available fuel, such as a dye, or a trace substance used as a marker, and the consultation of a calorific value corresponding to this tracer (a trace substance intrinsically present in the fuel which may vary between fuels can be used to identify a fuel, and / or a substance may be added voluntarily to act as a tracer); ii. the deduction of the calorific value from one or more detected physical or chemical properties of the available fuel; and / or lil, the combustion of a sample of the available fuel to determine its calorific value directly, possibly using the gas turbine engine. The chemical and / or physical determination of the calorific value of the available fuel can be carried out on board the aircraft, and may possibly be carried out during aircraft operation, e.g. the supply of lighting, heating and / or air conditioning while stationary / before refueling is completed. Specifically (but not exclusively) in scenarios where calorific value is entered manually, the procedure may further include performing a check to verify the calorific value / data entered. This check may involve measuring the calorific value of the fuel being used in the gas turbine engine, possibly during pre-takeoff aircraft operations (e.g., taxiing or gate operations), and comparing this to the calorific value obtained via a different route. If the discrepancy exceeds a certain threshold, the aircraft may be returned to the gate / refueling may be restarted. The determination of the fuel's calorific value can be carried out by the over- monitoring engine parameters during an initial period of aircraft operation during which the gas turbine engine uses fuel; and determining the fuel calorific value based on the monitored engine parameters. This determination step can be used to determine the calorific value of fuel already in the aircraft's tanks—e.g., during a flight prior to the refueling event—so that the energy remaining in the tanks can be calculated and deduced from the amount required during refueling. This determination step can also be used for the new fuel to be supplied to the aircraft—e.g.the combustion of a small sample in the / a gas turbine engine while the aircraft is stationary and before refueling is completed, so as to determine the amount of additional fuel to request or accept, or as part of checks to confirm the calorific value obtained during refueling. The gas turbine engine can be a main propulsion gas turbine engine, or it can be a gas turbine engine in an auxiliary power unit (APU), which may or may not be configured to provide propulsive power to the aircraft. In some cases, fuel properties can be determined by burning a small sample in the APU before starting the main engine(s). In additional or alternative examples, the combustion of a fuel sample to determine fuel characteristics can be carried out in one or more of the main propulsion engine(s). According to a second aspect, a propulsion system is provided for an aircraft, the propulsion system comprising: a gas turbine engine; a fuel tank arranged to hold fuel to power the gas turbine engine; and a supply manager designed for: to obtain the amount of energy required for a planned flight profile; obtain a fuel calorific value available for the aircraft; calculate the mass or volume of available fuel needed to provide the required energy; and deliver at the output the mass or volume of available fuel necessary to allow the aircraft to be refueled accordingly. The amount of energy required for a planned flight profile may include a safety margin beyond the expected amount needed for the flight. The size of this safety margin can be determined based on a variety of flight parameters, environmental conditions, and aircraft characteristics. The propulsion system may further include one or more sensors arranged to physically and / or chemically detect one or more properties or characteristics of the fuel. One or more fuel properties / sensor data can then be used to deduce or calculate the fuel's calorific value—this can be used for either or both of the fuel already present in the aircraft's tank(s) before refueling, and the fuel being supplied to the aircraft during refueling. The type of sensor and its location can be chosen accordingly. The propulsion system can be arranged to implement the process of the first aspect. The gas turbine engine may include: an engine core comprising a turbine, a compressor, and a central shaft connecting the turbine to the compressor; and a blower located upstream of the engine core, the blower comprising a plurality of blower blades and being arranged to be driven by an output of the core shaft. In another respect, it provides a non-transient, computer-readable medium on which instructions are stored which, when executed by a processor, cause the processor to: to obtain the amount of energy required for an aircraft's intended flight profile; obtain a fuel calorific value available for the aircraft; calculate the amount of available fuel needed to provide the required energy; and deliver at the output the mass or volume of available fuel necessary to allow the aircraft to be refueled accordingly. The instructions can also be arranged to cause the processor to command a fuel input to the aircraft, so that the aircraft is fueled with the calculated amount of fuel available. The instructions can be arranged to cause the processor to perform the process of the first aspect. According to a third aspect, a method is provided for controlling the refueling of an aircraft comprising a gas turbine engine and a fuel tank arranged to supply fuel to the gas turbine engine, the method comprising: the reception of a calorific value data input for fuel supplied to the aircraft during refueling; the determination independently of at least one of the following: (i) the calorific value of the fuel supplied to the gas turbine engine during operation; and (ii) the calorific value of the fuel supplied to the aircraft during refueling; and providing an alert if the determined calorific value of fuel is inconsistent with the received calorific value data input. As used above, determining the calorific value "independently" means determining the calorific value without using, or referring to, input calorific value data—two values ​​for the calorific value can therefore be obtained separately and compared. The two values ​​can thus be obtained in completely different ways—e.g., one can be supplied by a fuel supplier, and the other can be calculated from sensor data. The inventors were aware that, since different fuels can have different properties while still meeting standards, knowledge of the fuel(s) available for an aircraft can enable more efficient, tailored control of the propulsion system. Verification systems, for example, implementing fuel property checks to determine if changes to aircraft control are required based on fuel characteristics, can therefore be implemented to increase confidence in, and improve the reliability of, such new propulsion system control techniques. The third aspect check may be performed for safety reasons — to ensure that the total energy in the fuel on board the aircraft is sufficient for the intended flight, including any foreseen safety margin (e.g., in case of bad weather or the need to divert to a different airport). The control of the third aspect can be carried out for reasons of aircraft performance optimization — confidence in the knowledge of the fuel(s) on board the aircraft can allow the engine operation to be matched to the available fuel(s). The process can be carried out by an aircraft system, and / or by an off-site unit. The gas turbine engine can be a main, propulsive gas turbine, or a gas turbine engine of an auxiliary power unit (APU), which may or may not be arranged to provide propulsive power to the aircraft. The process may include determining the calorific value of the fuel supplied to the aircraft during refueling in a fuel test unit. The fuel test unit may be provided off-site at a refueling location. The determination of the calorific value of the fuel supplied to the gas turbine engine in operation can be carried out during at least one of the following: (i) one or more ground operations prior to takeoff (e.g., engine warm-up and / or aircraft taxiing), and (ii) the climb. It should be borne in mind that performing the determination relatively early in flight can facilitate action. Appropriate corrective action should be taken if the determined value is inconsistent with the received calorific value data. Ideally, the check can be performed before the aircraft takes off, in cases where the fuel level is not as expected to such an extent that refueling is advised. The determination of the calorific value of the fuel supplied to the gas turbine engine in use can be carried out by burning fuel taken from the fuel tank arranged to be used to supply fuel to the gas turbine engine in an auxiliary power unit (APU) of the aircraft and determining the calorific value. The calorific value of the fuel supplied to the operating gas turbine engine can be determined by monitoring engine parameters (of the propulsion gas turbine engine and / or a gas turbine engine in an APU) during an initial period of aircraft operation in which the gas turbine engine is using the fuel; and determining the fuel calorific value based on the monitored engine parameters. Optionally, one or more engine parameters can be monitored during this initial period. The determination of the calorific value of the fuel supplied to the aircraft during refueling can be carried out by identifying a tracer in the fuel supplied to the aircraft (e.g., a dye or a trace marker element), and consulting a calorific value corresponding to that tracer. The determination of the calorific value of the fuel supplied to the aircraft during refueling can be carried out by deducing the calorific value of one or more detected physical or chemical properties of the available fuel. Receiving calorific value data input for fuel supplied to the aircraft during refueling may include receiving data input through an aircraft user interface, e.g., a value entered or a value selected from a menu. The receipt of calorific value data input for fuel supplied to the aircraft during refueling may include the receipt of data communicated electronically to the aircraft, e.g., sent by a supplier or obtained from scanning a barcode, QR code, or other code associated with the fuel supplied. According to a fourth aspect, a propulsion system is provided for an aircraft, the propulsion system comprising: a gas turbine engine; a fuel tank arranged to hold fuel to power the gas turbine engine; and a fuel tracking system designed to: receive a data input of calorific value for fuel supplied to the aircraft during refueling: determine at least one of the following: (i) the calorific value of the fuel supplied to the gas turbine engine during operation; and (ii) the calorific value of the fuel supplied to the aircraft during refueling; and provide an alert if the determined calorific value of the fuel is inconsistent with the received calorific value data input. The gas turbine engine may include: an engine core comprising a turbine, a compressor, and a central shaft connecting the turbine to the compressor; and a blower located upstream of the engine core, the blower comprising a plurality of blower blades and being arranged to be driven by an output of the core shaft. The propulsion system may further include one or more sensors arranged to physically and / or chemically detect one or more fuel properties, one or more fuel properties being used to deduce or calculate the fuel's calorific value. These sensors may be used to determine the calorific value of either the fuel supplied to the gas turbine engine during operation, or the fuel being delivered to the aircraft during refueling. The type of sensor and its location may be chosen accordingly. The propulsion system may further include one or more sensors arranged to detect at least one engine parameter during an initial period of aircraft operating time during which the gas turbine engine (a propulsion gas turbine engine and / or a gas turbine engine of an APU) uses fuel; the at least one engine parameter being used to determine the calorific value of the fuel. The propulsion system can be arranged to implement the process of the third aspect. According to a fifth aspect, a method is provided for refueling an aircraft comprising a gas turbine engine and a fuel tank arranged to supply fuel to the gas turbine engine, the method comprising: the determination of an amount of energy required for a planned flight profile (possibly including a safety margin as mentioned above); the receipt of a data input of calorific value for the available fuel; calculating the amount of available fuel needed to provide the required energy; the delivery at the exit of the calculated quantity of available fuel necessary to allow the aircraft to be refueled accordingly; the determination of at least one of the following: (i) the calorific value of fuel supplied to the gas turbine engine in use after refueling; and (ii) the calorific value of the fuel supplied to the aircraft during refueling; and providing an alert if the determined calorific value of fuel is inconsistent with the received calorific value data input. Calculating the amount of available fuel (e.g., from a tank or refueling pipeline) needed to provide the required energy may involve determining the calorific value of the fuel already in the fuel tank and using this value to calculate the total available energy of fuel already on board, possibly using one or more calorific values ​​of one or more fuels on board and the quantity of each. This value can then be subtracted from the determined amount of energy required for the planned flight profile. Therefore, calculating the amount of available fuel needed to provide the required energy may involve determining the energy value of the fuel already in the fuel tank and subtracting this value from the determined amount of energy required for the planned flight profile. The method may further include calculating the calorific value of the blended fuel after refueling. The calorific value of the fuel supplied to the operating gas turbine engine can therefore be determined to be inconsistent with the received calorific value data input only when the calculated calorific value for the blended fuel does not correspond to the determined calorific value of the fuel supplied to the operating gas turbine engine. According to a sixth aspect, a method is provided for determining the calorific value of fuel supplied to a gas turbine of an aircraft, the method comprising: the detection of one or more engine parameters during an initial period of aircraft operating time during which the gas turbine engine uses fuel; and based on one or more detected engine parameters, the determination of a fuel calorific value. Detection can be repeated so that at least one motor parameter is monitored over time, and / or to provide control of a first detected power. Because different fuels can have different properties, such as different calorific values, while still meeting standards, the same gas turbine engine may perform differently when supplied with a different fuel. The performance of the aircraft's own gas turbine engine can therefore be used to determine the calorific value of a fuel currently being supplied to that same engine. Knowledge of the fuel(s) available to an aircraft is beneficial because it can enable more efficient, tailored control of propulsion systems and / or customized refueling. The first period of aircraft operating time may include at least one of (i) a period of engine operation before takeoff, and (ii) a period of time spent climbing. The first period of aircraft operation time may include time spent warming up the engine before any movement, time spent taxiing before takeoff, and / or time spent climbing. In other examples, the first period of time may be during cruise. It should be kept in mind that making the determination relatively early in the flight can facilitate the use of the determined calorific value. The determination may be made before the aircraft leaves the ground, in cases where the fuel is not as expected to such an extent that refueling is advised, or to judge the quantity of fuel to load. In other applications, the calorific value may be used to influence propulsion system control during flight / for the remainder of the flight following the determination. The engine parameters detected / monitored may include one or more of the following: fuel flow, shaft speed, combustion chamber temperature rise, thrust generated, and exhaust temperature. The process may further include carrying out a check to verify the determined calorific value, the check including comparing the calorific value of the fuel determined from the engine parameters with a calorific value of the fuel which is: a) brought to the aircraft (e.g., by electronic communication, or using a graphical user interface); b) calculated from detected fuel properties (e.g., sensor data); and / or c) determined from one or more engine parameters detected in a different time period of aircraft operation. The calorific value of the fuel can be determined as a relative value, or change, by comparison with that of a different fuel used by the same gas turbine engine. A change in the engine parameter(s) monitored / detected during the change from one fuel to another can be used to determine the change in calorific value. One or more engine parameters can therefore be detected / monitored in two different time periods—one for each of the two different fuels. The fuel change can be the only change made to the engine control between the two time periods. The two time periods can also be chosen such that the altitude and / or other external parameters are at least substantially the same for both, and can therefore be chosen to be close to each other in time, if not immediately consecutive. An interval can be left between the two time periods to allow for any transient behavior around the fuel change. Engine parameters can instead be monitored over a single period of time during which the switch between the two different fuels is made. The fuel switch can be the only change made to the engine control during this period. Any transient behavior around the fuel switch can be used in determining the calorific value in certain examples. Engine parameters can be or include shaft speed and fuel flow. Engine parameters may be or include a temperature rise across the combustion chamber and fuel flow. The gas turbine engine that is monitored to determine the calorific value of the fuel being consumed may be an aircraft propulsion gas turbine engine. The gas turbine engine being monitored to determine the calorific value of the fuel being consumed may be a gas turbine engine from an aircraft auxiliary power unit. In such cases, the calorific value can be determined before the aircraft's propulsion gas turbine engine is started. According to a seventh aspect, a method is provided for operating an aircraft comprising a gas turbine engine and a fuel tank arranged to supply fuel to the gas turbine engine, the method comprising: the detection of one or more engine parameters during an initial period of aircraft operating time during which the gas turbine engine uses fuel; depending on the detected engine parameter(s), the determination of data relating to or including the fuel's calorific value; and the adjustment (e.g., optimization) of engine performance during a second period of aircraft operating time based on the determined data. The detection can be repeated in such a way that at least one engine parameter is monitored over time, and / or to provide control of an initially detected power. Knowledge of fuel calorific value can therefore be used as a tool to improve aircraft performance in flight, e.g., by gauging the likelihood of contrail formation behind an aircraft under given atmospheric conditions and changing the fuel source or altitude in flight as appropriate, or using a lower calorific value fuel for lower thrust operations. The first aircraft operating time period may include at least one of the following: (i) engine operating time before takeoff, and (ii) time spent climbing. The first aircraft operating time period may include at least one of the following: time spent warming up the engine before movement, time spent taxiing before takeoff, and time spent climbing. Alternatively, or in addition, the first operating time period may be early in a cruise period (e.g., within the first 1%, 5%, or 10% of cruise time), thus allowing for subsequent engine performance optimization. The second period of aircraft operation time may be a period of time spent in cruise, and possibly may be the entire cruise. Engine performance tuning may include at least one of the following: * the change in fuel flow rate; * the change in pump discharge; * the change in altitude; * the change of the guide vane stage; and * the fuel change. According to an eighth aspect, a power system is provided for an aircraft, the power system comprising: a gas turbine engine; a fuel tank arranged to hold fuel to power the gas turbine engine; and a fuel tracking system designed to: detect one or more engine parameters during an initial period of aircraft operating time during which the gas turbine engine uses fuel; and determine the calorific value of the fuel based on the detected engine parameter(s). The gas turbine engine may include an engine core comprising a turbine, a compressor, and a central shaft connecting the turbine to the compressor; and a blower located upstream of the engine core, the blower comprising a plurality of blower blades and arranged to be driven by an output from the central shaft. The gas turbine engine can be configured to provide propulsive power to the aircraft. The power system can therefore be described as a propulsion system. Alternatively, the gas turbine engine can be an auxiliary power unit (APU) of the aircraft and may not be arranged to provide propulsive power. The power supply system may further include at least one sensor arranged to detect at least one motor parameter. The fuel system may further include one or more sensors arranged to physically and / or chemically detect one or more properties or characteristics of the fuel, one or more fuel properties or characteristics being used to deduce or calculate the calorific value of the fuel. According to a ninth aspect, a propulsion system is provided for an aircraft, the propulsion system comprising a gas turbine engine and one or more fuel tanks arranged to contain fuel to supply the gas turbine engine, one or more of the tanks containing a fuel that is a sustainable aviation fuel — SAF — or is a fuel mixture comprising SAF, the SAF having a density of between 90% and 98% of the density, px, of kerosene and a calorific value of between 101% and 105% of the calorific value, PC, of ​​kerosene. The propulsion system therefore includes at least one fuel tank designed to store fuel to power the gas turbine engine, in which the stored fuel comprises at least a proportion of sustainable aviation fuel (SAF). If the stored fuel is an SAF blend as opposed to pure SAF, the calorific value and density of the blend will be between those of SAF and those of kerosene. The gas turbine engine includes: a combustion chamber; and a fuel pump arranged to supply fuel from one or more fuel tanks to the combustion chamber at an energy flow rate, C, the fuel pump having an inlet arranged to receive fuel from one or more fuel tanks and an outlet arranged to deliver fuel at a pump outlet volumetric flow rate, Q, a proportion of the fuel delivered at the outlet being supplied to the combustion chamber and the remainder being recirculated (i.e., returned from the pump outlet to the inlet, either directly or indirectly through one or more systems or components), the percentage of fuel passing through the pump that is recirculated / that is not supplied to the combustion chamber being called a spillage percentage, and in which the fuel supplied to the pump comprises X % of SAF, where X % is in the range of 5 % to 100 %, with any remaining fuel being kerosene, and has a density, pz, and a calorific value PC,. The propulsion system is arranged in such a way that: Fuel change spill report (dimensionless), R4, of: percentage of spillage during cruise using Te kerosene 2: The A AE AT AT EE A EE ZA RCE EE EE EE EE EL Ra © spillage percentage during cruise using fuel with X% of SAF is equal to: 2-(€ / vese) 2-(e7 {PC:xp,) ] X can be greater than 50, such that the fuel supplied to the combustion chamber is greater than 50% SAF. We will bear in mind that different tanks may contain fuels with a different %SAF. The gas turbine engine can be arranged so that, for an engine with a maximum takeoff thrust in the range of 400 kN to 500 kN, R; is in the following range during cruise: 2- (a-sa0 ee - (0-63 = z ie ds ALAN) <R< Z ACTA) ; = {1 ST O- 15. 1? @- (0-63 / (0014x PC, 0.1 5 @- (0-580) / foniax PC / x0e) and possibly on the beach: - (Q-580 Ÿ ; 2- (0-60 ft s 250 / (one Peso) <R< # 20 / onssx Pen) f 2 2 ES to @-(0-580) / tonraxPC,xo,] { @- 10-580) / (oisxPC,x6,) where Q (fuel flow rate) is measured in liters per second, PC (calorific value) in MJ / kg, and p (density) in kg per liter, with the subscripts K and F used for kerosene and SAF or SAF blend fuel, respectively. The gas turbine engine can be arranged so that, for an engine with a maximum takeoff thrust in the range of 300 kN to 350 kN, R is in the following range during cruise: @-(to-asn toorr«Pegxn) | <R< 2-(ro-san (sorsPe,x0) | 2- ( (Q-#37) / (0027 <PCHx5,) ) 2- ((O-487) / im027x PEp1o;)} d and possibly on the beach: 2-(0-a3m (0.027 « PCyX Q- (0.447) 2195 / our ren) « pc 200 / omsre re) ©-iQ-487) / 1 TT @-(0-437 Sreitxn - (0-43 / (0027 sPyx2p) 7 @- 40-437) / fop271PE,x0,) where, as above, Q is measured in liters per second, PC in MJ / kg, and p in kg per liter. The SAF proportion (X%) can be gravimetric. The gas turbine engine can be arranged in such a way that R; < 1.04. The gas turbine engine can be arranged in such a way that R, = 1.003. The gas turbine engine can be arranged so that A; > 1.014. It can be R; > 1.014 when the proportion SAF is greater than or equal to 75%. In addition to considering fuel properties, R; can vary depending on one or more of the ambient temperature, altitude, and cruising stage. The gas turbine engine can be arranged so that R4 decreases by less than 0.15% between the beginning and end of the cruise, for a constant temperature and altitude. The gas turbine engine can be arranged so that R4 decreases by at least 0.11% when the altitude increases by at least 600 m. The gas turbine engine may include an engine core comprising a turbine, a combustion chamber, a compressor, and a central shaft connecting the turbine to the compressor; and a blower situated upstream of the engine core, the blower comprising a plurality of blower blades and being arranged to be driven by an output of the central shaft. A fuel tank in one or more tanks may contain pure SAF having a density between 90% and 98% of the density, px, of kerosene and a calorific value between 101% and 105% of the calorific value, PCx, of kerosene. Alternatively or in addition, a fuel tank in one or more tanks may contain a blended fuel comprising a proportion of SAF, the SAF having a density between 90% and 98% of the density, px, of kerosene and a calorific value between 101% and 105% of the calorific value, PC, of ​​kerosene, and being blended with a kerosene-based fuel to form the mixture. According to a tenth aspect, a propulsion system for an aircraft is provided, the propulsion system comprising a gas turbine engine and at least one fuel tank arranged to store fuel to power the gas turbine engine, in which the stored fuel comprises at least a proportion of a sustainable aviation fuel — SAF, the SAF having a density of between 90% and 98% of the density, px, of kerosene and a calorific value of between 101% and 105% of the calorific value PC, of ​​kerosene, the gas turbine engine comprising: a combustion chamber; and a fuel pump arranged to supply fuel from one or more of the fuel reservoirs to the combustion chamber, the fuel pump having an inlet arranged to receive fuel from one or more of the tanks of fuel, the fuel supplied to the pump comprising X% SAF (with any remaining fuel being kerosene) and an outlet arranged to deliver the fuel, a proportion of the fuel delivered at the outlet being supplied to the combustion chamber and the remainder being recirculated, the percentage of fuel passing through the pump that is recirculated / that is not supplied to the combustion chamber being called a spillage percentage, and wherein the propulsion system is arranged such that: the spillage ratio of fuel change, Re: R, = percentage of spillage during cruise using kerosene d — “percentage of spillage during cruise using a fuel with X % of SAF> where X % is at least 30%, is greater than or equal to 1.003. Here, a gravimetric SAF percentage is used, as opposed to a volumetric one—the fuel is therefore X% SAF by weight. Due to varying fuel densities, the volumetric SAF percentage would be slightly different—a volume percentage can be used in some applications, and the given numbers can be adjusted accordingly. The propulsion system, and more specifically the gas turbine engine, can be arranged in such a way that: At R, > 1+#%- When X is 50 (i.e., the fuel is 50% SAF by weight), the fuel change discharge ratio can be at least 1.0066. When X is 100, such that the fuel is pure SAF, the fuel change pour ratio can be at least 1.0138. The gas turbine engine can be configured such that R varies according to one or more of the following: ambient temperature, altitude, and cruise stage. The gas turbine engine can be configured such that R4 decreases by less than 0.15% between the beginning and end of cruise, for constant temperature and altitude. The gas turbine engine can be arranged so that R4 decreases by at least 0.11% when the altitude increases by at least 600 m. The gas turbine engine can be arranged so that R4 is less than or equal to 1.04. The gas turbine engine can be arranged so that Rç is greater than or equal to 1.003. The gas turbine engine can be arranged in such a way that R; is greater than or equal to 1.014. According to an eleventh aspect, a method for operating an aircraft comprising a propulsion system is provided, the propulsion system comprising an enginea gas turbine comprising a combustion chamber, and at least one fuel tank arranged to store fuel to supply the gas turbine engine, in which the stored fuel comprises at least a proportion of a durable aviation fuel — SAF — having a density of between 90% and 98% of the density, Pr, of kerosene and a calorific value of between 101% and 105% of the calorific value PC, of ​​kerosene, the gas turbine engine comprising a fuel pump arranged to supply fuel from one or more of the fuel tanks to the combustion chamber at an energy flow rate, C, the fuel pump having an inlet arranged to receive fuel from one or more fuel tanks and an outlet arranged to deliver fuel at a pump outlet volumetric flow rate, Q, a proportion of the fuel delivered at the outlet being supplied to the combustion chamber and the remainder being recirculated,The percentage of fuel passing through the pump that is recirculated is called the spillage percentage; the process includes: the supply of fuel from one or more of the fuel tanks to the gas turbine engine, the fuel supplied to the gas turbine engine comprising X% of SAF, where X% is in the range of 5% to 100%, with any remaining fuel being kerosene, and wherein the fuel has a density, p, and a calorific value PC; and the control of the propulsion system such that: Fuel change spill report, Re: fé, = percentage of spillage during cruise using Kerosene d percentage of spillage during cruise using fuel with X% SAF is equal to: Q- { E / {PCyx0,) ) (Cinema) The method may include controlling the propulsion system in such a way that any one of the conditions relating to R, as described above for the ninth and tenth aspects, apply. X% can be in the range of 50% to 100%. According to a twelfth aspect, a method is provided for operating an aircraft comprising a propulsion system including a gas turbine engine and one or more fuel tanks arranged to contain fuel to power the gas turbine engine, one or more of the tanks containing at least a proportion of a sustainable aviation fuel — SAF — having a density of between 90% and 98% of the density, px, of kerosene and a calorific value of between 101% and 105% of the calorific value, PC, of ​​kerosene: the gas turbine engine comprising: a combustion chamber; and a fuel pump arranged to supply fuel from one or more of the fuel reservoirs to the combustion chamber, the fuel pump having an inlet arranged to receive fuel from one or more fuel tanks, the fuel supplied to the pump comprising X % of SAF, any remainder of the fuel being kerosene, and an outlet arranged to deliver the fuel, a proportion of the fuel delivered at the outlet being supplied to the combustion chamber and the remainder being recirculated, the percentage of fuel passing through the pump that is recirculated being called a pour percentage. The propulsion system is arranged in such a way that: The fuel change spill ratio, R, is defined as: percentage of spillage during cruise using kerosene R, = a AAA NA SE ATEN ALL CAD AA A AA EEE DNA A AS NET d'— "percentage of spillage during cruise using a fuel with X Ce of SAF > When X% is at least 30% (i.e., when the fuel is at least 30% SAF by weight / mass), R4 is greater than or equal to 1.003. The method may include controlling the propulsion system in such a way that any one of the conditions relating to R, as described above for the ninth and tenth aspects, apply. According to a thirteenth aspect, a method is proposed for determining one or more fuel characteristics of an aviation fuel suitable for powering a gas turbine engine of an aircraft, the method comprising: the determination of a mass of a fuel loaded, or which has been loaded, onto the aircraft; the determination of a corresponding volume of fuel; and the determination of one or more fuel characteristics of the fuel as a function of the determined mass and volume. The inventors have in mind that by measuring a mass and volume parameter of the fuel being loaded, or which has been loaded, onto an aircraft one or more characteristics of the fuel can be determined. Determining one or more fuel characteristics may involve calculating a fuel density based on the determined fuel mass and volume. One or more fuel characteristics may be a function of the density. One or more fuel characteristics may be obtained by comparing the density with a fuel lookup table containing known densities and corresponding fuel characteristics to be determined. Determining the mass of the fuel may involve measuring a flow rate mass at a point within a fuel supply line used to deliver fuel to one or more fuel tanks on board the aircraft. The determination of fuel volume may include measuring a volumetric flow rate at a point in the fuel supply line used to convey fuel to one or more fuel tanks on board the aircraft. The determination of the mass of fuel may include the measurement of the mass and / or a change in mass of any or more of: the aircraft; one or more fuel tanks on board the aircraft; a fuel tank vehicle from which fuel is brought; or a storage tank from which fuel is brought to the aircraft. The determination of fuel volume may include measuring the volume and / or a change in the volume of fuel: stored in one or more fuel tanks on board the aircraft; and / or stored in a fuel storage tank from which fuel is brought to the aircraft. One or more specified fuel characteristics may include any one or more of the following: (i) a distribution of hydrocarbons from fuel (ii) a percentage of sustainable aviation fuel in the fuel; and / or (iii) the aromatic hydrocarbon content of the fuel. One or more fuel characteristics may include an indication that the fuel is a fossil fuel, e.g., kerosene. One or more fuel characteristics can also be determined based on an indicative signal of fuel temperature. According to a fourteenth aspect, a fuel characteristics determination system is provided for determining one or more fuel characteristics of an aviation fuel suitable for powering a gas turbine engine of an aircraft, the system comprising: a fuel characteristic determination module designed to: receive a fuel mass signal indicating a mass of fuel being loaded, or which has been loaded, onto the aircraft; receive a fuel volume signal indicating a volume of fuel being loaded or loaded onto the aircraft; and determine one or more fuel characteristics of the fuel based on fuel volume and fuel mass signals. The fuel characteristic determination module can be configured to calculate fuel density based on fuel mass signal and fuel volume signal. The fuel characteristics determination system may include in addition to a mass sensor arranged to measure a mass of fuel, in which the fuel mass signal is received from the mass sensor. The fuel characteristic determination system may further include a volume sensor arranged to measure a volume of fuel, in which the fuel volume signal is received from the volume sensor. The mass sensor can be a mass flow meter. The volume sensor can be a volumetric flow meter. The mass flow meter can be arranged to measure the mass flow rate at a point within a fuel supply line used to deliver fuel to one or more fuel tanks on board the aircraft. The volumetric flow meter can be arranged to measure a volumetric flow rate at a point within the fuel supply line used to deliver fuel to one or more fuel tanks on board the aircraft. The fuel mass signal can be a function of a measured mass and / or a change in mass of any one or more of the following: the aircraft; one or more fuel tanks on board the aircraft; a fuel tanker from which fuel is supplied; or a storage tank from which fuel is supplied to the aircraft. The mass sensor can be configured to measure any of the parameters listed above. The fuel volume signal can be a function of a measured volume and / or a change in fuel volume: stored in one or more fuel tanks on board the aircraft; and / or stored in a fuel storage tank from which fuel is supplied to the aircraft. The volume sensor can be configured to measure either of these parameters. One or more specified fuel characteristics may include any one or more of the following: (i) a distribution of hydrocarbons from fuel; (ii) a percentage of sustainable aviation fuel in the fuel; and / or (iii) the aromatic hydrocarbon content of the fuel. One or more fuel characteristics may include an indication that the fuel is a fossil fuel, e.g., kerosene. The determination module can also be arranged to determine one or more fuel characteristics based on an indicative signal of the fuel temperature. According to a fifteenth aspect, a method for operating an aircraft having a gas turbine engine is provided, the method comprising: the determination of one or more fuel characteristics using the thirteenth aspect method; and the operation of the aircraft according to one or more fuel characteristics. The operation of the aircraft according to one or more fuel characteristics may include: a) the modification of an aircraft control parameter, preferably a gas turbine engine control parameter, in response to one or more fuel characteristics; and / or b) the supply of fuel having different fuel characteristics during aircraft refueling. According to a sixteenth aspect, an aircraft is provided comprising the fuel characteristic determination system of the fourteenth aspect, the aircraft further comprising a control system arranged to control the operation of the aircraft according to one or more fuel characteristics determined by the fuel characteristic determination system. According to a seventeenth aspect, a method is provided for determining one or more fuel characteristics of an aviation fuel suitable for powering a gas turbine engine of an aircraft, the gas turbine engine having a combustion chamber to which fuel is supplied from an aircraft fuel system, the method comprising: the determination of a mass of fuel being supplied to the combustion chamber; determining a corresponding volume of fuel being supplied to the combustion chamber; and the determination of one or more fuel characteristics as a function of the determined mass and volume. The inventors have in mind that one or more fuel characteristics can be determined during the operation of a gas turbine engine by measuring a mass and volume parameter of the fuel as it is supplied to a combustion chamber of the gas turbine engine. One or more fuel characteristics can be determined based on the measured mass and volume of the fuel. Determining one or more fuel characteristics may involve calculating a fuel density based on the determined fuel mass and volume. One or more fuel characteristics may be a function of the density. One or more fuel characteristics may be obtained by comparing the density with a fuel lookup table containing known densities and corresponding fuel characteristics to be determined. Determining the mass of the fuel may include determining a mass flow rate of fuel supplied to the combustion chamber. The fuel system may include an engine fuel system that is part of the gas turbine engine. The mass flow rate can be measured at a point in a fuel line of the engine fuel system. The mass flow rate can be measured immediately before the fuel reaches the combustion chamber. The mass flow rate can be determined based on an operating parameter of a fuel pump supplied within the fuel system. The mass flow rate can also be determined based on a fuel flow measurement using a mass flow meter. Determining the fuel volume may include determining the volumetric flow rate of fuel being supplied to the combustion chamber. Volumetric flow rate can be measured at a point in a fuel line of the engine fuel system. Volumetric flow rate can be measured immediately before the fuel reaches the combustion chamber. Volumetric flow rate can be measured at a position adjacent to the mass flow rate. The specific volume can be determined based on an operating parameter of a pump supplied within the fuel system. The volumetric flow rate can be determined based on a fuel flow measurement using a volumetric flow meter. One or more of the specified fuel characteristics include any one or more of the following: (i) a distribution of hydrocarbons from fuel; (li) a percentage of sustainable aviation fuel in the fuel; and / or (iii) the aromatic hydrocarbon content of the fuel. One or more fuel characteristics include an indication that the fuel is a fossil fuel, e.g., kerosene. According to an eighteenth aspect, a fuel characteristic determination system is provided for determining one or more fuel characteristics of an aviation fuel suitable for powering a gas turbine engine of an aircraft, the gas turbine engine having a combustion chamber to which fuel is supplied from an aircraft fuel system, the fuel characteristic determination system comprising: a fuel characteristic determination module designed to: the reception of a fuel mass signal indicating a mass of fuel being delivered to the combustion chamber; the reception of a fuel volume signal indicating a volume of fuel being supplied to the combustion chamber; and determine one or more fuel characteristics of the fuel based on fuel volume and fuel mass signals. The fuel characteristic determination module can be configured to calculate fuel density based on fuel mass signal and fuel volume signal. The fuel characteristic determination system may further include a mass flow meter arranged to measure the mass flow rate of fuel being supplied to the combustion chamber. The fuel mass signal can be received from the mass flow meter. The fuel characteristic determination system may further include a volumetric flow meter arranged to measure the volumetric flow rate of fuel being supplied to the combustion chamber. The fuel volume signal can be received from the volumetric flow meter. The mass flow meter and / or volumetric flow meter can be arranged to measure the fuel flow at a point in a fuel line of a gas turbine engine fuel system. The mass flow meter and / or volumetric flow meter can also be arranged to measure the fuel flow at a point immediately before the fuel reaches the combustion chamber. The fuel system may further include a fuel pump arranged to supply fuel to the combustion chamber. The fuel mass signal and / or the fuel volume signal may be a function of a fuel pump operating parameter. One or more of the specified fuel characteristics include any one or more of the following: (i) a distribution of hydrocarbons from fuel; (ii) a percentage of sustainable aviation fuel in the fuel; and / or (iii) the aromatic hydrocarbon content of the fuel. One or more fuel characteristics include an indication that the fuel is a fossil fuel, e.g., kerosene. According to a nineteenth aspect, a method for operating an aircraft having a gas turbine engine is provided, the method comprising: the determination of one or more fuel characteristics using the seventeenth aspect method; and the operation of 1° aircraft according to one or more fuel characteristics. The operation of the aircraft according to one or more fuel characteristics may include: a) the modification of an aircraft control parameter, preferably a gas turbine engine control parameter, in response to one or more fuel characteristics; and / or b) the supply of fuel having different fuel characteristics during aircraft refueling. According to a twentieth aspect, an aircraft is provided comprising the fuel characteristic determination system of the eighteenth aspect, the aircraft further comprising a control system arranged to control the operation of the aircraft according to one or more fuel characteristics determined by the fuel characteristic determination system. This description can be applied to any relevant configuration of gas turbine engine. Such a gas turbine engine can be, for example, a turbofan engine, an open-rotor gas turbine engine (in which the propeller is not enclosed in a nacelle), a turboprop, or a turbojet. Any such engine may or may not be equipped with an afterburner. A gas turbine engine, according to any aspect of this description, may include an engine core comprising a turbine, a combustion chamber, a compressor, and a central shaft connecting the turbine to the compressor. Such a gas turbine engine may include a fan (having fan blades). Such a fan may be located upstream of the engine core. Alternatively, in some examples, the gas turbine engine may include a fan located downstream of the engine core, for example, where the gas turbine engine is an open rotor or a turboprop (in which case the fan may be called a propeller). When the gas turbine engine is an open rotor or turboprop, it may consist of two stages of counter-rotating propellers connected to and driven by a free-spinning turbine via a shaft. The propellers may rotate in opposite directions, such that one rotates clockwise and the other counterclockwise around the engine's axis of rotation. Alternatively, the gas turbine engine may consist of one stage of propellers and a stage of guide vanes configured downstream of the propeller stage. The guide vane stage may have variable pitch. Consequently, high-pressure, intermediate-pressure, and free-spinning turbines can drive compressors and high-pressure and intermediate-pressure propellers, respectively, via suitable interconnecting shafts. Thus, the propellers can provide the majority of the propulsive thrust. When the gas turbine engine is an open rotor or a turboprop, one or more of the propeller stages may be driven by a gearbox. The gearbox may be of the type described here. An engine as described herein can be a turbofan engine. Such an engine can 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, it can be said that the fan rotates at the same speed as the fan drive turbine. An engine as described herein may be a geared turbofan. In such an arrangement, the engine has a fan that is driven via a gearbox. Thus, such a gas turbine engine may include a gearbox that receives an input from the central shaft and delivers an output drive to the fan so as to rotate the fan at a lower speed than the central shaft. The input to the gearbox may be directly from the central shaft, or indirectly from the central shaft, for example, via a shaft and / or spur gear. The central shaft may rigidly 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 connecting turbines and compressors, for example, one, two, or three shafts. Purely by way of example, the turbine connected to the central shaft may be a first turbine, the compressor connected to the central shaft may be a first compressor, and the central shaft may be a first central shaft. The engine core may further include a second turbine, a second compressor, and a second central shaft connecting the second turbine to the second compressor. The second turbine, second compressor, and second central shaft may be arranged to rotate at a higher speed than the first central shaft. In such an arrangement, the second compressor can be positioned axially downstream of the first compressor. The second compressor can be configured to receive (for example, directly, for example, via a generally annular pipe) a flow from the first compressor. The gearbox can be arranged to be driven by the center shaft that is configured to rotate (e.g., during operation) at the lowest rotational speed (e.g., the first center shaft in the example above). Alternatively, the gearbox can be arranged to be driven only by the center shaft that is configured to rotate (e.g., during operation) at the lowest rotational speed (e.g., being only the first center shaft, and not the second center shaft, in the example above). The gearbox can be arranged to be driven by one or more shafts, for example, the first and / or second shafts in the example above. The gearbox can be a reduction gearbox (in that the output to the blower has a lower rotational speed than the input to the central shaft). Any type of gearbox can be used. For example, the gearbox can be a "satellite" or "star" gearbox, as described in more detail elsewhere here. Such a gearbox can be a single-stage gearbox. Alternatively, such a gearbox can be a compound gearbox, for example, a compound satellite gearbox (which may have the input on the sun gear and the output on the ring gear, and thus be called a "star compound" gearbox), for example, having two reduction stages. The gearbox can have any desired reduction ratio (defined as the input shaft speed divided by the output shaft speed), for example, greater than 2.5, for example, in the range of 3 to 4.2, or 3.2 to 3.8, for example, in 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 preceding sentence. Purely as an example, the gearbox could be a star-type gearbox with a reduction ratio in the range of 3.1 or 3.2 to 3.8. Purely as a further example, the gearbox could be a "star" gearbox with a reduction ratio in the range of 3.0 to 3.1. Purely as a further example, the gearbox could be a "star" gearbox with a reduction ratio in the range of 3.6 to 4.2.In some arrangements, the gear ratio may be outside these ranges. In any gas turbine engine as described and / or claimed herein, fuel of a given composition or mixture is supplied to a combustion chamber, which may be supplied downstream of the fan and compressor(s) relative to the flow path (e.g., axially downstream). For example, the combustion chamber may be directly downstream of (e.g., at the outlet of) the second compressor, when a second compressor is supplied. As a further example, the flow at the outlet of the combustion chamber may be supplied at the inlet of the second turbine, when a second turbine is supplied. The combustion chamber may be supplied upstream of the turbine(s). The compressor or compressors (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 The stator blade arrangement can be axially offset from one another. For example, the gas turbine engine can be a direct-drive, turbofan-type gas turbine engine comprising 13 or 14 compressor stages (in addition to the fan). Such an engine might, for instance, have 3 stages in the first compressor (or "low pressure") and either 10 or 11 stages in the second compressor (or "high pressure"). As a further example, the gas turbine engine can be a geared gas turbine engine (in which the fan is driven by a first central shaft via a reduction gearbox) comprising 11, 12, or 13 compressor stages (in addition to the fan). Such an engine might have 3 or 4 stages in the first compressor (or "low pressure") and 8 or 9 stages in the second compressor (or "high pressure").As a further example, the gas turbine engine can be a "geared" gas turbine engine having 4 stages in the first compressor (or "low pressure") and 10 stages in the second compressor (or "high pressure"). The turbine or turbines (for example, the first and second turbines as described above) can comprise any number of stages, for example, multiple stages. Each stage can comprise one row of rotor blades and one row of stator blades. The rotor and stator blade rows can be axially offset from each other. The second (or "high-pressure") turbine can comprise two stages in any arrangement (for example, whether geared or direct-drive). The gas turbine engine can be a direct-drive gas turbine engine comprising a first (or "low-pressure") turbine having five, six, or seven stages. Alternatively, the gas turbine engine can be a geared gas turbine engine comprising a first (or "low-pressure") turbine having three or four stages. Each fan blade can be defined as having a radial extent extending from a root (or hub) to a radially internal gas-washed location, or from a 0% extent position to a tip at a 100% extent position. The ratio of the fan blade radius at the hub to the fan blade radius at the tip can be less than (or on the order of) any 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 can be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​can form upper or lower bounds), for example, in the range of 0.28 to 0.32 or 0.29 to 0.30. These ratios can commonly be referred to as the hub-to-tip ratio.The spoke at the hub and the spoke at the . Both tip ratios can be measured at the leading edge portion (or axially forward) of the blade. The hub-to-tip ratio, of course, refers to the gas-washed portion of the fan blade, namely, the portion radially outside any platform. The blower radius can be measured between the engine centerline and the tip of a blower blade at its leading edge. The blower diameter (which can simply be twice the blower radius) can be greater than (or on the order of) any of the following: 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 (approximately 125 inches), 330 cm (approximately 130 inches), 340 cm (approximately 135 inches), 350 cm, 360 cm (approximately 140 inches), 370 cm (approximately 145 inches), 380 cm (approximately 150 inches), 390 cm (approximately 155 inches), 400 cm, 410 cm (approximately 160 inches) or 420 cm (approximately 165 inches).The blower diameter can be within an inclusive range delimited by any two of the values ​​in the preceding sentence (i.e., the values ​​can form upper or lower limits), for example, within the range of 240 cm to 280 cm or 330 cm to 380 cm. Purely as a non-limiting example, the blower diameter can be within 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. The fan speed can vary during operation. Generally, the fan speed is lower for fans with a larger diameter. As a purely non-limiting example, the fan speed under cruising conditions can be less than 3500 rpm, for example, less than 2500 rpm, or less than 2300 rpm. As a further non-limiting example, the fan speed under cruising conditions for a geared gas turbine engine with a fan diameter in the range of 200 cm to 210 cm can be in the range of 2750 to 2900 rpm. Purely as a further non-limiting example, the fan speed under cruising conditions for a "geared" gas turbine engine having a fan diameter in the range of 210 cm to 230 cm can be in the range of 2500 to 2800 rpm.Purely as a further non-limiting example, the fan speed under cruising conditions for a geared gas turbine engine with a fan diameter in the range of 340 cm to 360 cm can be in the range of 1500 to 1800 rpm. Purely as a further non-limiting example, the fan speed under cruising conditions for a direct-drive engine with a fan diameter in the range of 190 cm to 200 cm can be in the range of 3600 to 1800 rpm. 3900 rpm. Purely as a further non-limiting example, the fan speed under cruising conditions for a direct drive engine having a fan diameter in the range of 300 cm to 340 cm can be in the range of 2000 to 2800 rpm. During operation of the gas turbine engine, the fan (with associated fan blades) rotates around an axis of rotation. This rotation leads to the displacement of the fan blade tip with a velocity Ugoime. The work done by the fan blades on the flow results in an enthalpy rise dH of the flow. A fan tip load can be defined as dH / Upoine* where dH is the enthalpy rise (e.g., the average enthalpy increase 1-D) across the fan and Ufoine is the (translational) velocity of the fan tip, e.g., at the leading edge of the tip (which can be defined as the radius of the fan tip at the leading edge multiplied by the angular velocity).The peak fan load under cruise conditions may be greater than (or on the order of) any one 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 ​​being dimensionless). The end fan load may be within an inclusive range bounded by any two of the values ​​from the preceding sentence (i.e., the values ​​may form upper or lower bounds), for example, in the range of 0.28 to 0.31, or 0.29 to 0.3 (for example, for a geared gas turbine engine). Gas turbine engines according to this description may have any desired dilution ratio, where the dilution ratio is defined as the ratio of the mass flow rate of the flow through the bypass pipe to the mass flow rate of the flow through the core at cruising conditions. In some arrangements the dilution ratio may be greater than (or on the order of) any of the following values: 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 may be in an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower bounds), for example in the range of 12 to 16, or 13 to 15, or 13 to 14.Purely by way of non-limiting example, the bypass ratio of a direct-drive gas turbine engine as described herein may be in the range of 9:1 to 11:1. Purely by way of further non-limiting example, the bypass ratio of a geared gas turbine engine as described herein may be in the range of 12:1 to 15:1. The bypass pipe may be substantially annular. The bypass pipe may be radially external to the central engine. The radially external surface of the bypass pipe may be defined by a nacelle and / or [missing information]. a blower housing. 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 outlet of the highest-pressure compressor (before the combustion chamber inlet) to 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 herein during cruise operation may be greater than (or on the order of) any of the following values: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio may be within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower bounds), for example, in the range of 50 to 70.Purely by way of non-limiting example, the overall pressure ratio at cruising conditions of a geared gas turbine engine with a fan diameter in the range of 200 cm to 210 cm can be in the range of 40 to 45. Purely by way of non-limiting example, the overall pressure ratio at cruising conditions of a geared gas turbine engine with a fan diameter in the range of 210 cm to 230 cm can be in the range of 45 to 55. Purely by way of non-limiting example, the overall pressure ratio at cruising conditions of a geared gas turbine engine with a fan diameter in the range of 340 cm to 360 cm can be in the range of 50 to 60. Purely by way of non-limiting example, the overall pressure ratio at cruising conditions of a direct-drive gas turbine engine with a fan diameter in the The range of 300 cm to 340 cm can 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 rate through the engine. In some examples, specific thrust may depend, for a given thrust condition, on the specific fuel composition supplied to the combustion chamber. Under cruise conditions, the specific thrust of an engine described and / or claimed herein may be less than (or on the order of) any of the following values: 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 within an inclusive range bounded by any two of the values ​​in the preceding sentence (i.e., the values ​​may form upper or lower bounds), for example, in the range of 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.Purely 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 cm can be in the range of 90 Nkg-s to 95 Nkg-s. Purely by way of non-limiting example, the specific thrust... The specific thrust of a geared gas turbine engine with a fan diameter in the range of 210 cm to 230 cm can be in the range of 80 Nkg's to 90 Nkg's. Purely by way of non-limiting example, the specific thrust of a geared gas turbine engine with a fan diameter in the range of 340 cm to 360 cm can be in the range of 70 Nkg's to 90 Nkg's. Purely by way of non-limiting example, the specific thrust of a direct-drive gas turbine engine with a fan diameter in the range of 300 cm to 340 cm can be in the range of 90 Nkg's to 120 Nkg's. A gas turbine engine as described and / or claimed herein may have any desired maximum thrust. Purely by way of non-limiting example, a gas turbine as described and / or claimed herein may be capable of producing a maximum thrust of at least (or in the order of) any one of the following values: 100 kN, 110 kN, 120 kN, 130 kN, 140 kN, 150 kN, 160 kN, 170 kN, 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 bounds). Purely as a non-limiting example, a gas turbine as described and / or claimed herein may be capable of producing a maximum thrust in the range of 330 kN to 420 kN, e.g., 350 kN to 400 kN.Purely by way of non-limiting example, the maximum thrust of a geared gas turbine engine with a fan diameter in the range of 200 cm to 210 cm can be in the range of 140 kN to 160 kN. Purely by way of non-limiting example, the maximum thrust of a geared gas turbine engine with a fan diameter in the range of 210 cm to 230 cm can be in the range of 150 kN to 200 kN. Purely by way of non-limiting example, the maximum thrust of a geared gas turbine engine with a fan diameter in the range of 340 cm to 360 cm can be in the range of 370 kN to 500 kN. Purely as a 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 can be in the range of 370 kN to 500 kN.The thrust mentioned 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 stationary. During operation, the flow temperature at the inlet of the high-pressure turbine can be particularly high. This temperature, which may be called TET, can be measured at the outlet of the combustion chamber, for example, immediately upstream of the first turbine blade, which may itself be called a nozzle guide vane. In some examples, TET may depend on, for a Under given thrust conditions and the specific fuel composition supplied to the combustion chamber, the TET (Thrust Energy Transfer) during cruise can be at least (or on the order of) any of the following values: 1400 K, 1450 K, 1500 K, 1550 K, 1600 K, or 1650 K. Thus, strictly as a non-limiting example, the cruise TET of a geared gas turbine engine with a fan diameter in the range of 200 cm to 210 cm² can be in the range of 1540 K to 1600 K. Again, purely as a non-limiting example, the cruise TET of a geared gas turbine engine with a fan diameter in the range of 210 cm to 230 cm² can be in the range of 1590 kN to 1650 kN. Purely as a 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 can be in the range of 1600 K to 1660 K.Purely as a non-limiting example, the cruise TET of a direct-drive gas turbine engine with a fan diameter in the range of 300 cm to 340 cm can be in the range of 1590 K to 1650 K. Purely as a non-limiting example, the cruise TET of a direct-drive gas turbine engine with a fan diameter in the range of 300 cm to 340 cm can be in the range of 1570 kN to 1630 kN. The TET in cruising can be in an inclusive range delimited by any two of the values ​​from the preceding sentence (i.e., the values ​​can form upper or lower bounds). The maximum TET during engine operation can be, for example, at least (or in the order of) any of the following values: 1700 K, 1750 K, 1800 K, 1850 K, 1900 K, 1950 K, 2000 K, 2050 K, or 2100 K. Thus, purely by way of non-limiting example, the maximum TET of a geared gas turbine engine with a fan diameter in the range of 200 cm to 210 cm can be in the range of 1890 K to 1960 K. Also purely by way of non-limiting example, the maximum TET of a geared gas turbine engine with a fan diameter in the range of 210 cm to 230 cm can be in the range of 1890 kN to 1960 kN.As a non-limiting example, the maximum TET of a geared gas turbine engine with a fan diameter in the range of 340 cm to 360 cm can be in the range of 1890 kN to 1960 kN. As a purely non-limiting example, the maximum TET of a direct-drive gas turbine engine with a fan diameter in the range of 300 cm to 340 cm can be in the range of 1935 K to 1995 K. Also as a non-limiting example, the maximum TET of a direct-drive gas turbine engine with a fan diameter in the range of 300 cm to 340 cm can be in the range of 1890 K to 1950 K. The maximum TET can be in an inclusive range delimited by any two of the values ​​from the preceding sentence (i.e., the values ​​can form upper or lower bounds), for example in the range from 1800 K to 1950 K. Maximum TET can occur, for example, at a high thrust condition, for example at a maximum takeoff condition (MTO). A portion of a fan blade and / or airfoil of a fan blade described and / or claimed herein may be made from any suitable material or combination of materials. For example, at least a portion of the fan blade and / or airfoil may be made at least partially from a composite, such as a metal matrix composite and / or an organic matrix composite, such as a carbon fiber composite. As a further example, at least a portion of the fan blade and / or airfoil may be made at least partially from a metal, such as a titanium-based metal, an aluminum-based material (such as an aluminum-lithium alloy), or a steel-based material. The fan blade may include at least two regions made from different materials.For example, a fan blade may have a protective leading edge, which can be made of a material better able to withstand impacts (e.g., from birds, ice, or other materials) than the rest of the blade. Such a leading edge could, for example, be made of titanium or a titanium-based alloy. Thus, purely as an example, a fan blade could have a body made of carbon fiber or aluminum (such as a lithium aluminum alloy) with a titanium leading edge. A blower as described and / or claimed herein may include a central portion from which the blower blades may extend, for example, in a radial direction. The blower blades may be connected to the central portion in any desired manner. For example, each blower blade may include a fastener that can engage with a corresponding slot in the hub (or disc). Purely by way of example, such a fastener may be in the form of a dovetail that can be inserted into and / or engage with a corresponding slot in the hub / disk to secure the blower blade to the hub / disk. As a further example, the blower blades may be formed as a single piece with a central portion. Such an arrangement may be called a bladed disc or bladed ring. Any suitable method may be used to manufacture such a bladed disc or bladed ring.For example, at least part of the blower blades can be machined from a block and / or at least part of the blower blades can be joined to the hub / disc by welding, such as linear friction welding. The gas turbine engines described and / or claimed herein may or may not be supplied with a variable area nozzle (VAN). Such a variable area nozzle allows the outlet cross-section of the pipe to be varied during operation. of bypass. The general principles of this description can be applied to motors 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 metallic body (for example, aluminum-lithium or titanium alloy), there may be 18, 20, or 22 fan blades. As used here, the terms idle, taxiing, takeoff, climb, cruise, descent, approach, and landing have their classic meanings and would be readily understood by a person skilled in the art. Thus, for a given gas turbine engine for an aircraft, a person skilled in the art will immediately recognize each term as referring to a phase of engine operation within a given mission of the aircraft to which the gas turbine engine is intended to be connected. In this regard, ground idle can refer to a phase of engine operation where the aircraft is stationary and in contact with the ground, but where there is a requirement for engine running. At idle, the engine can produce between 3% and 9% of its available thrust. In further non-limiting examples, the engine can produce between 5% and 8% of available thrust. In further non-limiting examples, the engine can produce between 6% and 7% of available thrust. Taxiing can refer to a phase of engine operation where the aircraft is propelled along the ground by the thrust produced by the engine. During taxiing, the engine can produce between 5% and 15% of available thrust. In further non-limiting examples, the engine can produce between 6% and 12% of available thrust. In further non-limiting examples, the engine can produce between 7% and 10% of available thrust.Takeoff can refer to a phase of engine operation where the aircraft is propelled by the thrust produced by the engine. Initially, during the takeoff phase, the aircraft may be propelled while in contact with the ground. Later, during takeoff, the aircraft may be propelled while in contact with the ground. During takeoff, the engine may produce between 90% and 100% of available thrust. In further, non-limiting examples, the engine may produce between 95% and 100% of available thrust. In further, non-limiting examples, the engine may produce 100% of available thrust. Climb can refer to a phase of engine operation where the aircraft is propelled by the thrust produced by the engine. During climb, the engine can produce between 75% and 100% of available thrust. In additional, non-limiting examples, the engine can produce between 80% and 95% of thrust. available. In additional, non-limiting examples, the engine can produce between 85% and 90% of available thrust. In this sense, climb can refer to an operational phase within an aircraft flight cycle between takeoff and arrival at cruise conditions. Alternatively, climb can refer to a nominal point in an aircraft flight cycle between takeoff and landing, when a relative increase in altitude is required, which may necessitate an additional thrust demand from the engine. As used here, cruise conditions have the classical meaning and would be readily understood by a person skilled in the art. Thus, for a given gas turbine engine for an aircraft, a person skilled in the art would immediately recognize cruise conditions to mean the engine's mid-cruise operating point for a given mission (which may be called in the industry an "economic mission") of an aircraft to which the gas turbine engine is designed to be connected. In this sense, mid-cruise is the point in an aircraft's flight cycle at which 50% of the total fuel burned between the end of the climb and the beginning of the descent has been burned (which can be approximated by mid-course—in terms of time and / or distance between the end of the climb and the beginning of the descent).Cruise conditions thus define an operating point for the gas turbine engine that provides sufficient thrust to ensure static operation (i.e., maintaining a constant altitude and Mach number) at mid-cruise of the aircraft to which it is designed to be connected, taking into account the number of engines supplied to that aircraft. For example, when an engine is designed to be connected to an aircraft that has two engines of the same type, at cruise conditions the engine provides half of the total thrust that would be required for static operation of that aircraft at mid-cruise. In other words, for a given gas turbine engine for an aircraft, cruise conditions are defined as the engine's operating point that provides a specified thrust (required to provide—in combination with any of the other engines on the aircraft—stationary operation of the aircraft to which it is designed to be connected at a given mid-cruise Mach number) under mid-cruise atmospheric conditions (defined by the International Standard Atmosphere according to ISO 2533 at mid-cruise altitude). For any given gas turbine engine for an aircraft, the mid-cruise thrust, atmospheric conditions, and Mach number are known, and therefore the engine's operating point under cruise conditions is clearly defined. Purely as an example, the forward speed under cruising conditions could be any point in the range of Mach 0.7 to 0.9, for example, 0.75 to 0.85, for example, 0.76 to 0.84, for example, 0.77 to 0.83, for example, 0.78 to 0.82, for example, 0.79 to 0.81, for example, on the order of Mach 0.8, on the order of Mach 0.85, or in the The range is from 0.8 to 0.85. Any speed within these ranges alone may constitute a cruising condition. For some aircraft, cruising conditions may be outside these ranges, for example, below Mach 0.7 or above Mach 0.9. Purely as an example, cruising conditions might correspond to standard atmospheric conditions (according to the International Standard Atmosphere, ISA) at an altitude that is in the range of 10,000 m to 15,000 m, 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 ft), for example in the range of 10,500 m to 11,500 m, for example in the range of 10,600 m to 11,400 m, for example in the range of 10,700 m to 11,300 m, for example in the range of 10,700 m (around 35,000 ft) to 11,300 m, for example in the range of 10,800 m at 11,200 m, for example, in the range from 10,900 m to 11,100 m, for example, in the order of 11,000 m. Cruising conditions can correspond to standard atmospheric conditions at any given altitude within these ranges. Purely as an example, cruise conditions might correspond to an engine operating point that provides a known required thrust level (e.g., a value in the range of 30 kN to 35 kN) at a Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38,000 feet (11,582 m). As a further example, cruise conditions might correspond to an engine operating point that provides a known required thrust level (e.g., a value in the range of 50 kN to 65 kN) at a Mach number of 0.85 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35,000 feet (10,668 m). During operation, a gas turbine engine described and / or claimed herein may be operated under the cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions (e.g., mid-cruise conditions) of an aircraft on which at least one (e.g., two or four) gas turbine engine may be mounted to provide propulsive thrust. Furthermore, a person skilled in the art would immediately recognize either or both of the descent and approach to refer to an operational phase in an aircraft flight cycle between cruise and landing. During either or both of the descent and approach, the engine may produce between 20% and 50% of available thrust. In additional, non-limiting examples, the engine may produce between 25% and 40% of available thrust. In further, non-limiting examples, the engine may produce between 30% and 35% of available thrust. Additionally, or alternatively, descent may refer to a nominal point in a aircraft flight cycle between takeoff and landing, when a relative increase in altitude is required, and which may require a reduced thrust demand from the engine. According to one aspect, an aircraft comprising a gas turbine engine as described and / or claimed herein is provided. The aircraft according to this aspect is the aircraft to which the gas turbine engine was designed to be connected. Consequently, the cruise conditions according to this aspect correspond to the aircraft's mid-cruise, as defined elsewhere herein. In one aspect, a method for operating a gas turbine engine as described and / or claimed herein is provided. The operation may be under any suitable conditions, which may be as defined elsewhere herein (for example, in terms of thrust, atmospheric conditions, and Mach number). According to one aspect, a method for operating an aircraft as described and / or claimed herein is provided. The operation according to this aspect may (or can be) an operation under any suitable condition, for example, mid-cruise of the aircraft, as defined elsewhere herein. A person skilled in the art will bear in mind, unless they are mutually exclusive, that a characteristic or parameter described in relation to any of the above aspects can be applied to any other aspect. Furthermore, unless they are mutually exclusive, any characteristic or parameter described can be applied to any aspect and / or combined with any other characteristic or parameter described herein. Embodiments will now be described by way of example only, with reference to the figures, in which: The [Fig.1] is a cross-sectional side view of a gas turbine engine; Fig. 2 is a close-up cross-sectional side view of an upstream portion of a gas turbine engine; The [Fig.3] is a partially cut-out view of a gearbox for a gas turbine engine; Fig. 4 is a schematic view of an aircraft with a refueling manager; The [Fig.5] is a schematic representation of a fuel supply management process; The [Fig.6] is a schematic view of an aircraft [featuring a fuel tracking system; Figure 7 is a schematic representation of a calorific value control process; Figure 8 is a schematic representation of a fuel tracking system; Figure 9 is a schematic representation of a management process. fuel supply including the calorific value control process of the [Fig.7]; [Fig.10] is a schematic view of an aircraft incorporating a fuel tracking system; The [Fig.11] is a schematic representation of a method for determining calorific value; The [Fig.12] is a schematic representation of a fuel tracking system; The [Fig.13] is a schematic view of an aircraft fuel supply system, illustrating the spillage around the fuel pump; Fig. 14 is a schematic view of an aircraft fuel pump; Figure 15 is a schematic representation of a discharge management process: and The [Fig.16] is a schematic view of an aircraft including a fuel characteristic determination system; The [Fig.17] is a schematic representation of a method for determining one or more characteristics of an aviation fuel; The [Fig.18] is a schematic view of a system for determining the characteristics of fuel supplied within a fuel system of a gas turbine engine; Figure 19 is another schematic representation of a method for determining one or more characteristics of an aviation fuel; and The [Fig.20] is a schematic representation of an aircraft operating procedure. Figure 1 illustrates a gas turbine engine 10 having a main axis of rotation 9. The engine 10 includes an air intake 12 and a propulsive blower 23 which generates two airflows: a central airflow A and a bypass airflow B. The gas turbine engine 10 includes a core 11 which receives the central airflow A. The engine core 11 includes, in axial flow series, a low-pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, a low-pressure turbine 19, and a core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass duct 22. The blower 23 is attached to and driven by the low-pressure turbine 19 via a shaft 26 and an epicyclic gearbox 30. During operation, the core air flow A is accelerated and compressed by the low-pressure compressor 14 and directed to the high-pressure compressor 15 where further compression takes place. The compressed air discharged from the high-pressure compressor Pressure 15 is directed into the combustion unit 16 where it is mixed with fuel F and the mixture is burned. The combustion unit 16 may be called the combustion chamber 16, with the terms "combustion unit 16" and "combustion chamber 16" used interchangeably here. The resulting hot combustion products then expand through, and thus drive, the high-pressure and low-pressure turbines 17, 19 before being discharged through the nozzle 20 to provide some propulsive thrust. The high-pressure turbine 17 drives the high-pressure compressor 15 via a matching interconnecting shaft 27. The blower 23 typically provides the majority of the propulsive thrust. The epicyclic gearbox 30 is a reduction gearbox. An exemplary arrangement for a geared blower gas turbine engine 10 is shown in [Fig. 2]. The low-pressure turbine 19 (see [Fig. 1]) drives the shaft 26, which is coupled to a sun wheel, or sun gear, 28 of the epicyclic gear arrangement 30. Radially outward from the sun gear 28 and meshing with it are a plurality of satellite gears 32 which are coupled together by a satellite carrier 34. The satellite carrier 34 constrains the satellite gears 32 to precess around the sun gear 28 in synchronism while allowing each satellite gear 32 to rotate about its own axis. The satellite carrier 34 is coupled via links 36 to the blower 23 in order to drive its rotation around the motor axis 9.Radially outwards from the satellite gears 32 and meshing with them, there is a ring or annular gear 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 here can be taken to mean respectively the lowest-pressure turbine stages and the lowest-pressure compressor stages (i.e., without including the blower 23) and / or the turbine and compressor stages which are connected together by the shaft interconnection 26 with the lowest rotational speed in the engine (i.e., excluding the gearbox output shaft that drives the blower 23). In some literature, the "low-pressure turbine" and the "low-pressure compressor" are referred to here as the "intermediate-pressure turbine" and "intermediate-pressure compressor." When such alternative nomenclature is used, the blower 23 may be called a first-stage compressor, or the lowest-pressure blower. The epicyclic gearbox 30 is shown in more detail as an example in [Fig. 3]. Each of the sun gear 28, the planet gears 32, and the ring gear 38 has teeth around its periphery for meshing with the other gears. However, for clarity only, representative portions of the teeth are illustrated in [Fig. 3]. There are four planet gears 32 shown, although it is obvious to the competent reader that more or fewer planet gears 32 may be provided within the scope of application of the claimed invention. Practical applications of a planetary epicyclic gearbox 30 generally include at least three planet gears 32. The epicyclic gearbox 30 illustrated by way of example in Figures 2 and 3 is of the satellite type, in that the planet carrier 34 is coupled to an output shaft via linkages 36, with the ring gear 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 gear (or ring) 38 permitted to rotate. In such an arrangement, the blower 23 is driven by the ring gear 38. As a further alternative example, the gearbox 30 may be a differential gearbox in which both the ring gear 38 and the planet carrier 34 are permitted to rotate. It should be borne in mind that the arrangement shown in Figures 2 and 3 is by way of example only, and various alternatives are within the scope of this description. Purely by way of example, any suitable arrangement can 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 links 36, 40 in the example in [Fig. 2]) between the gearbox 30 and other parts of the motor 10 (such as the input shaft 26, the output shaft, and the 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 fixed structures, such as the gearbox casing) can be used, and the description is not limited to the exemplary arrangement in [Fig. 2]. For example, where the gearbox 30 has a star arrangement (described above), those skilled in the art will readily understand that the arrangement of the output and support links and bearing locations would typically be different from that shown as an example in [Fig. 2]. Accordingly, this description extends to a gas turbine engine having any arrangement of gearbox styles (e.g., star or satellite), support structures, input and output shaft arrangement, and bearing locations. Optionally, the gearbox may drive additional and / or alternative components (e.g., the intermediate pressure compressor and / or a booster). Other gas turbine engines to which this description may be applied may have alternative configurations. For example, such engines may have an alternative number of compressors and / or turbines and / or an alternative 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 pipe 22 has its own nozzle 18 that is separate from and radially outside the central engine nozzle 20. However, this is not limiting, and any aspect of this description may also apply to engines in which the flow through the bypass pipe 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 for a mixed or split flow) can have a fixed or variable surface area. While the example described relates to a turbofan engine, the description can be applied, for example, to any type of gas turbine engine, such as an open rotor (in which the fan stage is not enclosed by a nacelle) or a turboprop, for example. In some arrangements, the gas turbine engine 10 may not include a gearbox 30. The geometry of the gas turbine engine 10, and its 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 bottom-up direction in [Fig. 1]), and a circumferential direction (perpendicular to the page in the view of [Fig. 1]). The axial, radial, and circumferential directions are perpendicular to each other. The fuel F supplied to the combustion equipment 16 may include a fossil hydrocarbon fuel, such as kerosene. Thus, fuel F may include molecules from one or more of the chemical families of n-alkanes, iso-alkanes, cycloalkanes, and aromatics. In addition, or alternatively, fuel F may include renewable hydrocarbons produced from biological or non-biological resources, otherwise known as sustainable aviation fuel (SAF). In each of the examples provided, fuel F may include one or more trace elements comprising, for example, sulfur, nitrogen, oxygen, inorganic matter, and metals. The functional performance of a given composition, or fuel mixture intended for use in a given mission, can be defined, at least in part, by the fuel's ability to serve the Brayton cycle of the gas turbine engine. Parameters defining functional performance may include, for example, specific energy; energy density; thermal stability; and particulate emissions. A relatively higher specific energy (i.e., energy per unit mass), expressed in MJ / kg, can at least partially reduce takeoff weight, thus potentially providing a relative improvement in fuel efficiency. A relatively higher energy density (i.e., energy per unit volume), expressed in MJ / L, can at least partially reduce the fuel volume at takeoff, which can be particularly important for volume-limited missions or military operations involving refueling. A relatively higher thermal stability (i.e., inhibition of fuel degradation or co-kefaction under thermal stress) can allow the fuel to withstand high temperatures in the engine and fuel injectors, thus potentially providing relative improvements in combustion efficiency.Reduced emissions, including particulate matter, can lead to less contrail formation, thus reducing the environmental impact of a given mission. Other fuel properties can also be key to functional performance. For example, a relatively lower freezing point (°C) can enable long-range missions to optimize flight profiles; minimal aromatic concentrations (%) can ensure sufficient swelling of certain materials used in the construction of O-rings and seals that have been previously exposed to fuels with high aromatic content; and maximum surface tension (mN / m) can ensure sufficient spray breaking and atomization of the fuel. The ratio of hydrogen to carbon atoms in a molecule can influence the specific energy of a given composition, or fuel mixture. Fuels with higher hydrogen-to-carbon ratios can have higher specific energies in the absence of bonding constraints. For example, fossil hydrocarbon fuels may contain molecules of approximately 7 to 18 carbon atoms, with a significant portion of a given composition consisting of molecules with 9 to 15 carbon atoms, averaging 12 carbon atoms. A number of sustainable aviation fuel blends have been approved for use, comprising between 10% and 50% sustainable aviation fuel (the remainder consisting of one or more fossil hydrocarbon fuels, such as kerosene), with additional compositions pending approval. However, it is expected within the aviation industry that sustainable aviation fuel blends comprising up to (and containing) 100% sustainable aviation fuel (SAF) will eventually be approved for use. Sustainable aviation fuels may include one or more of the following: Sustainable aviation fuels can contain alkanes, isoalkanes, cycloalkanes, and aromatics, and can be produced, for example, from one or more of the following: synthesis gas (syngas); lipids (e.g., fats, oils, and greases); sugars; and alcohols. Thus, sustainable aviation fuels may include one or both of these fuels with a lower aromatics and sulfur content compared to fossil hydrocarbon fuels. Alternatively, or in addition, sustainable aviation fuels may include one or both of these fuels with a higher isoalkanes and cycloalkanes content compared to fossil hydrocarbon fuels. Therefore, in some examples, sustainable aviation fuels may include one or both of these fuels with a density between 90% and 98% of that of kerosene and a calorific value between 101% and 105% of that of kerosene. Because at least in part of the molecular structure of sustainable aviation fuels, sustainable aviation fuels can provide beneficial effects including, for example, one or more of the following: higher specific energy (despite, in some examples, lower energy density); higher specific heat capacity; higher thermal stability; higher lubricity; lower viscosity; lower surface tension; lower freezing point; lower soot emissions; and lower CO- emissions, compared to fossil hydrocarbon fuels (e.g., when burned in combustion equipment 16).Consequently, compared to fossil hydrocarbon fuels, such as kerosene, sustainable aviation fuels can lead to one or both of a relative decrease in specific fuel consumption, and a relative decrease in maintenance costs. As shown in Figures 4, 6, and 10, an aircraft may include several fuel tanks 50, 53; for example, a larger primary fuel tank 50 located in the aircraft fuselage, and a smaller fuel tank 53a, 53b located in each wing. In other examples, an aircraft may have only one fuel tank 50, and / or the wing fuel tanks 53 may be larger than the central fuel tank 50, or no central fuel tank may be provided (with all the fuel being stored in the aircraft wings) — it should be borne in mind that many different tank arrangements are envisaged and that the illustrated examples are provided for ease of description and are not intended to be limiting. Figures 4, 6, and 10 show an aircraft 1 with a propulsion system 2 comprising two gas turbine engines 10. Fuel from a fuel supply system on board the aircraft 1 is supplied to the gas turbine engines 10. The fuel supply system in the illustrated examples comprises a single fuel source. For the purposes of this application, "fuel source" means either 1) a single fuel tank, or 2) a plurality of fuel tanks that are fluidly interconnected. Each fuel source is arranged to provide a distinct fuel source; that is, the first fuel source may contain a first fuel with a different fuel characteristic, or multiple different fuel characteristics, from a second fuel contained in the second fuel source. The first and second fuel sources are therefore not fluidly coupled to each other in such a way as to separate the different fuels (at least under normal operating conditions). As used here, the term "fuel characteristics" refers to intrinsic or inherent fuel properties such as fuel composition, not variable properties such as volume or temperature. Examples of fuel characteristics include one or more of the following: i. 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; ii. Parameters of hydrocarbon distribution in the fuel, such as: @ the aromatic hydrocarbon content of the fuel, and possibly also / alternatively the multi-aromatic hydrocarbon content of the fuel; @ the hydrogen to carbon (H / C) ratio of the fuel; @ % of composition information for some or all of the hydrocarbons present; lii, the presence or percentage of a particular species or element, such as: @ the percentage of nitrogen-containing species in the fuel; @ the presence or percentage of a tracer species or trace element in the fuel (e.g., a trace substance inherently present in the fuel that may vary between fuels and thus used to identify a fuel, and / or a substance intentionally added 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 being used in a gas turbine engine 10, such as: @ the level of non-volatile particulate matter (nvPM) or CO- emissions during combustion (a value can be provided for a specific combustion chamber being operated under particular conditions to fairly compare fuels; a measured value can be adjusted accordingly depending on the properties and conditions of the combustion chamber); @ a level of fuel coking; v. one or more properties of the fuel itself, independent of use in an engine or combustion, such as: @ the thermal stability of the fuel (e.g., the thermal degradation temperature); and @ one or more physical properties such as density, viscosity, calorific value, freezing point, and / or heat capacity. In this example, the first fuel source comprises a central fuel tank 50, located primarily in the fuselage of aircraft 1, and a plurality of wing fuel tanks 53a, 53b, 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 for balancing. All tanks 50, 53 are fluidly interconnected in the example shown, thus forming a single fuel source. Each of the central fuel tank 50 and wing fuel tanks 53 can comprise a plurality of fluidly interconnected fuel tanks. In another example, the wing fuel tanks 53a, 53b may not be fluidly connected to the central tank 50, thus forming a separate second fuel source. For balancing purposes, one or more fuel tanks in the left wing may be fluidly connected to one or more fuel tanks in the right wing. This can be done either via one or more of the central fuel tank 50 (if this tank is not part of a separate fuel source), or by bypassing the central fuel tank(s), or both (for maximum flexibility and safety). In another example, the first fuel source comprises wing fuel tanks 53 and a central fuel tank 50, while a second fuel source comprises an additional central fuel tank. Fluid interconnection between the wing fuel tanks 53 and the central fuel tank 50 of the first fuel source can be provided to balance the aircraft 1. In some examples, the allocation of fuel tanks 50, 53 available on aircraft 1 may be constrained such that a first fuel source and a second fuel source are each substantially symmetrical with respect to the aircraft's centerline. In cases where an asymmetrical fuel tank allocation is permitted, a suitable means of fuel transfer may be provided between fuel tanks of the first fuel source and / or between fuel tanks of the second fuel source such that the position of the The aircraft's center of mass can be maintained within acceptable lateral limits throughout the flight. An aircraft 1 can be refueled by connecting a fuel storage tank 60, such as that supplied by an airport fuel truck, or a permanent pipeline, to a connection port 62 of the aircraft fuel line, via a fuel line 61. A desired quantity of fuel can be transferred from the fuel storage tank 60 to one or more tanks 50, 53 of the aircraft 1. Particularly in examples with more than one fuel source, in which different tanks 50, 53 are to be filled with different fuels, multiple connection ports 62 of the fuel line can be provided instead of one, and / or valves can be used to appropriately direct fuel. Aircraft 1 is typically refueled at multiple different airports, for example, at the beginning and end of a long-haul flight. While there are standards that all aviation fuels must meet, different aviation fuels have different compositions, for example, depending on their source (e.g., different petroleum sources, biofuels or other synthetic aviation fuels (often described as sustainable aviation fuels — SAFs), and / or blends of petroleum-based fuels and other fuels) and any additives included (e.g., such as antioxidants and metal deactivators, biocides, static reducers, anti-icing agents, corrosion inhibitors) and any impurities.Just as varying between airports and fuel suppliers, even for a given airport or fuel supplier, the composition of available aviation fuel can vary between batches. Furthermore, the fuel tanks 50, 53 of aircraft 1 are not usually emptied before being refueled for a subsequent flight, resulting in mixtures of different fuels within the tanks—effectively a fuel with a different composition resulting from the mixing. Different fuels can have different calorific values ​​and / or different densities, among other fuel characteristics. As used here, the term "calorific value" refers to the lower heating value (also called net calorific value) of the fuel, unless otherwise specified. Net calorific value is defined as the amount of heat released by the combustion of a specified quantity of the fuel, assuming that the latent heat of vaporization of water in the reaction products is not recovered (i.e., the water produced remains as water vapor after combustion). The calorific values ​​(also called heating powers) of fuels can be directly determined - for example, by measuring the energy released when a certain volume or mass of fuel is consumed in the gas turbine engine 10 — or calculated from other fuel parameters; e.g., by observing the hydrocarbon distribution of the fuel and the calorific value of each type of constituent hydrocarbon. Alternatively, or as a supplement to provide verification, the calorific value can be determined using external data, such as a lookup table for a tracer substance in the fuel, or data encoded in a barcode associated with the fuel, or other stored data. The determination can therefore be carried out in many different ways. For example: * A barcode of a fuel to be added to a fuel tank 50, 53 of aircraft 1 can be scanned to read fuel data, or an identified tracer substance (e.g., a dye) and fuel properties consulted based on that tracer: * Data can be entered manually, or transmitted to the aircraft | ; "a fuel sample can be extracted for ground-side analysis before takeoff; "Fuel properties can be inferred from measurements of propulsion system 2 activity during one or more aircraft operating periods, e.g., engine warm-up, taxiing, takeoff, climb and / or cruise; and / or One or more fuel properties can be detected on board, possibly in flight, for example, using online sensors and / or other measurements. The calorific value of the fuel can therefore be consulted, physically determined from the results of combustion (either under the wing or off-site), chemically determined from detected chemical properties, calculated / inferred from other detected properties of the fuel (either under the wing or off-site), or otherwise determined or obtained. In some examples, combinations of these techniques can be used to determine and / or verify one or more fuel characteristics, including calorific value. Fuel characteristics can be detected in various ways, both directly (e.g., from sensor data corresponding to the fuel characteristic in question) and indirectly (e.g., by deduction or calculation from other characteristics or measurements). Characteristics can be determined as relative values ​​with respect to another fuel, or as absolute values. For example, one or more of the following detection methods can be used: "The aromatic or cycloparaffin content of the fuel can be determined by function of measuring the swelling of a sensor component made of a sealing material such as a nitrile sealing material. "Trace substances or species, either naturally present in the fuel or added to act as tracers, can be used to determine fuel characteristics such as the percentage of sustainable aviation fuel in the fuel or whether the fuel is kerosene. * 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 fuel thermal stability or coking level — for example by measuring the accumulation of surface deposits on the piezoelectric crystal which will result in a change in vibrational mode. "Various fuel characteristics can be determined by collecting performance parameters of the gas turbine engine 10 during a first period of operation (such as during takeoff) and then during a second period of operation (e.g., during cruise) comparing these collected parameters to expected values ​​if fuel with known properties is used. * Various fuel characteristics including the aromatic hydrocarbon content of the fuel can be determined based on sensor measurements of the presence, absence, or degree of formation of condensation trails by the gas turbine 10 during its operation. Fuel characteristics including aromatic hydrocarbon content can be determined based on a 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 the operation of the aircraft 1 based on measurements taken when the fuel is consumed — for example using the fuel flow rate and shaft speed or the temperature change across the combustion chamber 16. "Various fuel characteristics can be determined by making an operational change arranged to affect the operation of the gas turbine engine 10, detecting a response to the operational change; and determining one or more fuel characteristics of the fuel based on the response to the operational change. "Various fuel characteristics can be determined in relation to the fuel characteristics of a first fuel by changing the fuel supplied to the gas turbine engine 10 from the first fuel to the second fuel, and the determination of one or more fuel characteristics of the second fuel based on a change in the relationship between T30 and one of the T40 and T41 (the relationship being indicative of the temperature increase across the combustion chamber 16). The characteristics can be determined as relative values ​​by comparison to the first fuel, or as an absolute value as a reference to the known values ​​for the first fuel. As used here and discussed below, T30, T40 and T41, and any other numbered pressures and temperatures, are defined using the station numbering listed in SAE AS755, in particular: * T30 = High pressure compressor (HPC) discharge temperature; * T40 = Combustion outlet temperatures; and "T4] = High pressure turbine (HPT) rotor inlet temperature. In some examples, obtaining the calorific value of the fuel available for aircraft 1 and / or the fuel already on board aircraft 1 prior to refueling may involve a chemical and / or physical determination of the calorific value of the available fuel. Equivalent determinations may be made for the fuel(s) to be used for refueling—either by means of ground tests (e.g., prior to refueling) or on-board tests (e.g., using one or more sensors 204 on a fuel supply line within aircraft 1, leading to a tank 50, 53). Optionally, this determination may be or include performing one or more of the following: (i) the identification of a tracer in the available fuel, such as a dye or trace marker element, and the consultation of a calorific value corresponding to that tracer; (ii) the deduction of the calorific value from one or more detected physical and / or chemical properties of the available fuel, for example, using a ground test unit to analyze a fuel sample; and / or (iii) the combustion of a sample of the available fuel to determine its calorific value directly, possibly using the gas turbine engine 10. The step of chemical and / or physical determination of the calorific value of the available fuel can therefore be carried out on aircraft 1 in some examples. The inventors were aware that, since different fuels can have different properties while still conforming to standards, knowledge of the fuel(s) available for an aircraft 1 can allow for more efficient, tailored control of the propulsion system 2. For example, switching to a fuel with a higher calorific value can allow for a smaller quantity (mass and / or volume) of fuel to meet an aircraft's energy requirements for a flight. Since more power is needed to lift a greater mass of fuel, taking on enough fuel for the planned flight (including a safety margin above the expected energy demand), but not completely filling tank(s) 50, 53, can provide a performance bonus by reducing the aircraft's takeoff weight 1. Knowledge of the fuel's calorific value can therefore be used as a tool to improve aircraft performance, e.g., by avoiding carrying excess fuel weight. In the example shown in [Fig. 4], two sensors 204a, 204b are provided, each arranged to physically and / or chemically detect one or more characteristics of the composition of the fuel being added to the fuel tank 50, 53 during refueling. One or more of these fuel characteristics can be used to deduce or calculate the calorific value. The sensors can be described as forming part of a refueling management system 204, 206. In alternative examples, none of such sensors 204 can be provided (for example, a barcode associated with the fuel storage tank 60 can be read and the corresponding fuel data provided to the manager fueling 206 rather than detecting fuel properties), or more or less, and / or located differently, sensors can be provided. In the examples described here with reference to Figures 4 and 5, before refueling an aircraft 1, the amount of energy required for a planned flight profile of that aircraft is determined. As used here, the "amount of energy required for a planned flight profile" generally includes a safety margin above a calculated amount of fuel needed for aircraft 1, with its expected takeoff mass, to complete the planned flight. The safety margin may include authorization for a possible diversion to an alternative destination airport, e.g., in the event of weather conditions preventing a safe landing at the planned destination airport. As used here, the term "flight profile" refers to the operational characteristics (e.g., height / altitude, power setting, flight path angle, airspeed, and the like) of an aircraft when flying along a flight path, as well as the flight path itself / track (route). In determining the amount of energy required to complete a planned flight profile, external conditions may also be considered (e.g., weather, as prevailing headwinds may increase the energy required). In addition, or alternatively, the size of the safety margin may be chosen to account for variations in external conditions. When an aircraft 1 is about to be refueled, in various examples, a power The calorific value of the fuel available to the aircraft (i.e., the fuel with which the aircraft is to be refueled) is obtained, possibly using any of the off-site methods described above. The total specific energy of the available fuel (either per unit mass or per unit volume) can thus be established. A quantity of fuel required to meet the energy requirements of the flight profile can then be determined. In some cases, an underwing determination method could be used (e.g., combustion of a sample or the use of onboard sensors). An underwing determination may lead to a need to pause and restart refueling while the required quantity of fuel is calculated. In many cases, the tank(s) 50, 53 of aircraft 1 are not completely empty when aircraft 1 needs to be refueled. In such cases, the calorific value of the fuel already on board aircraft 1 is also obtained, possibly by retrieving the fuel characteristics (e.g., calorific value, or other properties from which calorific value can be calculated) from storage, and / or by consuming a sample of fuel (e.g., during taxiing) and directly measuring the calorific value. From a known quantity of fuel remaining (mass, volume, or % full of tank(s) 50, 53) and the known calorific value, the total energy available from the fuel on board can be determined. The total energy available from the fuel on board can therefore be subtracted from the determined amount of energy required for the planned flight profile to determine the amount of energy needed to be added to aircraft 1 during refueling. Using the specific energy (or volumetric energy density) of the available fuel, a desired quantity (i.e., mass or volume) of fuel available to aircraft 1 can thus be calculated. In some examples, an aircraft 1 may have multiple tanks 50, 53 containing different fuels, possibly with different calorific values. The total fuel energy content for fuel on board can therefore be added through the different tanks 50, 53. Similarly, if an aircraft 1 is to be refueled with several different fuels, the determined amount of energy required can be distributed among the different fuels as desired, and a mass or volume of each fuel to be added can be calculated. Aircraft 1 can then be fueled with the calculated quantity (mass or volume) of available fuel. In some examples (particularly those where data analysis was used to determine one or more calorific values), after refueling, a check is then carried out to verify the calorific value of the fuel. This check may include measuring the calorific value of the fuel. during use in the gas turbine engine 10, possibly during engine warm-up before taxiing, or during the taxiing of aircraft 1 before takeoff. Ideally, the fuel characteristics can be determined (and any checks, or based on the determined characteristics, carried out) before aircraft 1 leaves the ground, in case the fuel is not as expected. The 2020 process carried out in these examples is illustrated in [Fig.5]. In step 2022, the amount of energy required for a planned flight profile, including a safety margin, is obtained. In this context, the amount of energy being "obtained" means that a numerical value or measurement of the required energy is somehow made available for use in process 2020—whether by calculation or other determination (e.g., from a lookup table), by communication with another system, by user input (e.g., to a graphical user interface), by retrieval from memory, or by any other suitable method. For example, the fueling manager 206 may receive a message containing an indicative value of the energy requirement, or may calculate such a value based on knowledge of the intended flight profile, aircraft 1, and aircraft load. In step 2024, the available fuel calorific value for aircraft 1 is obtained. In this context, the calorific value being "obtained" again means that a numerical value or measurement of the calorific value is somehow made available for use in this process 2020. The value can be obtained in any suitable way, e.g., by being entered manually, for example, via a graphical user interface communicating with the fuel manager 206, by being communicated electronically to the fuel manager 206, e.g., by wired or wireless communication from a barcode reader following barcode scanning, and / or by being determined from sensor data. The value can be stored for future use, possibly in memory, or in communication with the fuel manager 206. The steps of obtaining the required amount of energy in 2022 and obtaining the calorific value of the fuel available for refueling in 2024 can be carried out in either order, or simultaneously. It should be borne in mind that, if tank 50, 53 is not empty before refueling, the energy content of the fuel already on board can also be obtained — e.g., by recovering a previously stored calorific value for this fuel, and by calculating the energy content of the remaining fuel as a reference to it and the quantity of fuel remaining in tank(s) 50, 53, or by calculation from data of sensor. In step 2026, the amount of available fuel needed to provide the required energy is calculated, based on the values ​​obtained. The calculated amount can be a volume and / or a mass of fuel—refueling can be performed volumetrically or gravimetrically. In some examples, the calculation can be performed by the fuel manager's own fueling system (step 206). In examples where fuel remains on board, the energy of that fuel can be subtracted from the energy required for the planned flight profile before calculating the amount of available fuel needed. In step 2028, the calculated quantity (mass or volume) of available fuel required is distributed so as to allow aircraft 1 to be refueled accordingly. For example, a refueling manager 206 can display the quantity to a graphical user interface, can transmit the quantity to another system for the attention of a refueling technician, or can provide the value to an automated system so that this system automatically completes the refueling process, or triggers an alert, once the calculated quantity of fuel has been supplied. Updated values ​​for the calorific value of the fuel in fuel tank 50, 53 after refueling can be calculated and stored for future use, possibly by the refueling manager 206. In examples where an aircraft 1 has multiple fuel tanks 50, 53 that are fluidly connected such that the fuels in the tanks 50, 53 are equivalent, a single calorific value for the fuel can be stored and updated. In examples where an aircraft 1 has multiple fuel tanks 50, 53 that are not fluidly connected, such that there may be differences between the fuels in the different tanks 50, 53, a separate calorific value can be stored and updated for each tank 50, 53. The 2020 process can then be iterated at each refueling event. The 2020 procedure may further include the control 2029 of an automated refueling system to refuel the aircraft 1 in line with the calculated quantity of fuel. Alternatively, refueling may be carried out and / or completed manually, e.g., by a person reading the calculated quantity delivered at the outlet (mass or volume) of the required available fuel, possibly from a display associated with the connection port 62 of the aircraft's fuel line. In some examples, refueling can be started before one or more of the preceding process steps 2022-2028 are performed, these steps can be performed during refueling, and the refueling process can then be completed when the calculated amount of fuel has been added. In certain examples, particularly in implementations where a fuel calorific value is manually entered in step 2024, a check may then be performed to verify the entered data. The check may include measuring the fuel calorific value while in use in the gas turbine engine 10, possibly during the aircraft 1's taxiing before takeoff. The fuel calorific value may be determined by monitoring engine parameters during an initial period of aircraft operation time during which the gas turbine engine 10 is using the fuel; and determining the fuel calorific value based on the monitored engine parameters (e.g., fuel burn-up rate, fuel flow rate, temperatures, pressures, thrust generated, shaft speed, etc.).In the event of a significant discrepancy between the values ​​entered and determined, aircraft 1 may return to the terminal for further investigation, as a safety measure. A propulsion system 2 for an aircraft 1 can therefore include a refueling manager 206 arranged to: obtain 2022 a quantity of energy required for a planned flight profile, possibly including a safety margin; obtain 2024 a fuel calorific value available for the aircraft; Calculate the quantity (mass or volume) of available fuel needed to provide the required energy; and deliver at output 2028 the quantity (mass or volume) of available fuel necessary to enable aircraft 1 to be refueled accordingly. In some applications, a refueling manager may be provided separately; not on board the aircraft. Optionally, a refueling manager may be provided as part of a ground refueling station. If the flight profile does not include a safety margin, or if any reasons for extending a safety margin are found (e.g., predicted adverse weather conditions or abnormally high load), an (additional) safety margin may be added at the calculation stage. The fueling manager 206 can further be arranged to control 2029 an automated fueling system so that it fuels aircraft 1 in line with the calculated amount of fuel (e.g., adding only the calculated amount of fuel, within tolerances). The fueling manager 206 can be supplied as a separate fueling management unit integrated into the propulsion system 2, and / or as software and / or hardware incorporated into pre-existing aircraft control systems (e.g., as part of an electronic engine controller). (CEM) 42). In some examples, the fueling manager 206 can store calorific value data for any fuel currently in the aircraft's tank(s) 50, 53, and can cause this data to be updated after refueling. The data can be stored separately from the circuits performing the calculations and retrieved as needed—regardless of where the data is stored, this storage can be considered part of the fueling manager 206, whether or not it is integrated or physically connected in any way. In examples where a calorific value is consulted or read (e.g., from scanning a barcode), this value can be automatically transferred to the fuel manager 206, or can be entered into a user interface, or associated with the fuel manager 206, e.g., by a pilot or technician. In some applications, a non-transient, computer-readable medium containing instructions to bring about the 2020 process described above may be provided for use under the wing or off-site. The instructions may cause the processor to: obtain 2022 an amount of energy required for a planned flight profile of an aircraft 1; obtain 2024 a fuel calorific value available for the aircraft; Calculate the amount of available fuel needed to supply the required energy in 2026; and deliver at output 2028 the mass or volume of available fuel necessary to enable the aircraft to be refueled accordingly. In automated refueling implementations, the instructions can be further arranged to cause the processor to command a fuel input to aircraft 1, such that the aircraft is refueled with the calculated amount of available fuel. In other implementations, the output can be read by a person who then refuels aircraft 1 accordingly, and / or used to provide an alert when a sufficient volume or mass of fuel has been added. The inventors were aware that, since different fuels can have different properties while still meeting standards, knowledge of the fuel(s) available to an aircraft 1 can enable more efficient, tailored control of the propulsion system 2. For example, switching to a fuel with a higher calorific value can allow a smaller quantity of fuel to meet an aircraft's energy requirements for a flight. Since more power is needed to lift a greater mass of fuel, carrying sufficient fuel for the planned flight (including a margin of (Safety above expected energy demand), but not completely filling the tank(s), can provide a performance bonus by reducing the aircraft's takeoff weight. Knowledge of fuel calorific value can therefore be used as a tool to improve aircraft performance, e.g., by avoiding carrying excess fuel weight and / or by gauging the probability of contrail formation behind an aircraft under given atmospheric conditions. The calorific value of one or more fuels can be controlled by a second method to improve reliability, particularly if calorific value knowledge is to be used in any way that could affect safety (e.g., ensuring sufficient fuel on board for a safe landing at a destination airport). In various examples such as those described with respect to Figures 6 to 9, an aircraft 1 receives calorific value data input for fuel supplied to the aircraft 1 during refueling. For example, a barcode of a fuel to be added to a fuel tank 50, 53 of the aircraft can be scanned to read fuel data, or a tracer substance (e.g., a dye) can be identified and fuel properties consulted based on that tracer. Alternatively, calorific value can be provided to a pilot or technician, e.g., on a written label or orally, for input into an aircraft system. The calorific value data can be entered manually or transmitted electronically to the aircraft 1. In examples where calorific value is consulted or read (e.g.(from scanning a barcode), this value can be automatically transferred to an aircraft fuel tracking system 203, or can be entered into a user interface of, or associated with, fuel tracking system 203, e.g., by a pilot or technician. It should be borne in mind that verifying the calorific value of fuels can be important, particularly in scenarios involving manual input of calorific value data. The received input of calorific value data for the fuel supplied to aircraft 1 during refueling can therefore be verified, and an alert can be provided if the determined calorific value of the fuel is inconsistent with the received calorific value data input. In the example shown in [Fig. 6], two sensors 204a, 204b are provided, each arranged to physically and / or chemically detect one or more characteristics of the fuel composition being added to the fuel tank 50, 53 during refueling. One or more of these fuel characteristics can be used to deduce or calculate the calorific value. The sensors 204 can be described as part of a fuel monitoring system 203, which can be described as a fuel composition monitoring system 203. In alternative examples, none of such sensors 204 can be supplied (for example, a barcode associated with a fuel storage tank 60 can instead be read and the corresponding fuel data supplied to the fueling manager 206), or more or fewer, and / or located differently, sensors may be supplied. In many cases, the tank(s) 50, 53 of aircraft 1 are not totally empty when aircraft 1 needs to be refueled. In such cases, the calorific value of the fuel already on board aircraft 1 is also obtained, possibly by retrieving the fuel characteristics (e.g., calorific value, or other properties from which calorific value can be calculated) from storage, and / or by consuming a sample of fuel (e.g., during taxiing) and directly measuring the calorific value. From a known quantity of fuel remaining (mass, volume, or % full of the tank(s)), a known quantity of fuel added, and the known calorific values ​​of each, the calorific value of the resulting mixture can be determined where appropriate (e.g., where the fuels are mixed in the same tank or linked tanks, or are mixed during use). In various examples, the calorific value of the fuel arranged to be supplied to the gas turbine engine 10 is determined. It should be borne in mind that, in applications with multiple fuel sources, the fuel arranged to be supplied to the gas turbine engine 10 may differ from the fuel in any given tank 50, 53. This fuel may include a fuel mixture with which the aircraft 1 was recently refueled, and any fuel remaining in the aircraft's tank(s) 50, 53 following a previous refueling event. In some examples, a sample may be taken, e.g., from a pipe approaching the gas turbine engine 10, for analysis, for example, using an analytical unit or ground laboratory. In alternative or additional examples, the determination may be made under the wing, using one or more sensors or other sensors on board the aircraft 1. In alternative or additional examples, the calorific value of the fuel supplied to aircraft 1 during refueling is determined, for example, by a ground test prior to the refueling event, or by underwing tests in a pipe or tank containing only the newly added fuel. In examples with an active fuel management system, it may be possible to deliver the new fuel alone to the gas turbine engine 10, regardless of whether the older fuel is still on board (e.g., by choice of fuel tank / source), and to calculate the calorific value of the new fuel directly from the engine performance. In examples where ground fuel testing is performed, the calorific value of the fuel supplied to aircraft 1 during refueling can be determined in a fuel test unit 6l1a. In some examples, the fuel test unit may form part of a fuel delivery apparatus 61. In other examples, a fuel test unit 6la may be supplied by an airline or airport, or a sample may be sent to a laboratory for analysis. The 6la test unit of such examples can be supplied off-site to a refueling site. In examples where underwing fuel testing is performed, the determination of the calorific value of the fuel supplied to the gas turbine engine 10 in operation may be carried out during at least one of the aircraft's pre-takeoff taxis and during the climb. If a sufficiently large discrepancy is identified, the flight may be aborted subject to re-verification of the fuel. In examples where underwing fuel testing is performed, the calorific value of the fuel supplied to the operating gas turbine engine 10 can be determined by burning fuel taken from the fuel tank 50, 53 in an auxiliary aircraft power unit 1 and measuring the calorific value. Alternatively or in addition, the main gas turbine engine 10 can be used to perform the combustion. After refueling, a check is then carried out to verify the calorific value of the fuel on board the aircraft 1. The check may include the measurement of the calorific value of the fuel being used in the gas turbine engine 10, possibly during the taxiing of the aircraft | before takeoff. If a discrepancy between the determined calorific value and the received calorific value data exceeds a threshold, an alert may be provided—e.g., an audible and / or visual alarm, and / or a message sent to a pilot or other responsible party. It should be borne in mind that, when the determined calorific value is that of a fuel mixture created on board aircraft 1, the determined value is not directly compared to the received calorific value data; instead, adjustments are made to one or the other to reflect the contribution of the older fuel. The 2030 process carried out is illustrated in [Fig.7]. In step 2032, a data input for the calorific value of a fuel supplied to aircraft 1 during refueling is received. This data may be entered manually, transmitted to the fuel tracking system 203, obtained by reading a barcode, or similar. The input may include a calorific value of the fuel, and / or may be arranged to allow the calorific value to be determined. fuel (e.g., calculated or consulted). The input can be obtained in any suitable manner, e.g., by being entered manually, for example, through a graphical user interface communicating with the fuel tracking system 203, and / or communicated electronically to the fuel tracking system 203, e.g., by wired or wireless communication from a barcode scanner after reading a barcode. The input may be or include sensor data in certain applications, e.g., by providing one or more chemical or physical fuel characteristics. A fuel calorific value in, or determined from, the input can be stored for future use, possibly in the memory of, or in communication with, the fuel tracking system 203. In step 2034, a calorific value is determined for at least one of the following: (i) the fuel supplied to the gas turbine engine 10 in use (this fuel comprising a portion of, or being, the fuel supplied to aircraft 1 during refueling): and (ii) the fuel supplied to aircraft 1 during refueling. The determination in step 2034 does not use the calorific value data provided in step 2032, so the calorific value determinations in steps 2032 and 2034 are independent of each other. The determination in step 2034 may otherwise use any of the approaches described above. The value may be stored for future use, possibly in memory or in communication with the fuel tracking system 203. It should be borne in mind that the fuel supplied to the gas turbine engine 10 in use may differ from the fuel supplied to the aircraft | during refueling, given that any other fuel already on board the aircraft 1 may be used (either by mixing in the same tank 50, 53 as the new fuel, or by being supplied in a mixture from different on-board fuel sources). Steps 2032 (receipt of calorific value input) and 2034 (determination of calorific value, independently of this input) can be carried out in either order, or simultaneously. At step 2036, an alert is provided if the determined calorific value of the fuel is inconsistent with the received calorific value data input. It should be borne in mind that, if tank 50, 53 is not empty before refueling, the calorific value of fuel already on board can also be obtained — e.g., by recovering a previously stored calorific value for that fuel — and can also be used in the determination 2034 of whether the determined calorific value of the fuel is inconsistent with the received calorific value data input. Updated values ​​for the calorific value of the fuel in fuel tank 50, 53 after refueling can be calculated and stored for future use, possibly by the fuel tracking system 203. In examples where an aircraft 1 has multiple fuel tanks 50, 53 that are fluidly connected such that the fuels in the tanks 50, 53 are equivalent, a single fuel calorific value can be stored and updated. In examples where an aircraft 1 has multiple fuel tanks 50, 53 that are not fluidly connected, such that there may be differences between the fuels in the different tanks 50, 53, a separate calorific value can be stored and updated for each tank. The 2030 process can then be iterated at each refueling event. Implementations of the 2030 method therefore allow for a check to be performed to verify the entered data. The check may include measuring the calorific value of the fuel being used in the gas turbine engine 10, possibly during the engine 10 warm-up or the aircraft 10 taxiing before takeoff. The determination 2034 of the fuel calorific value can be carried out by monitoring engine parameters during an initial period of aircraft operation during which the gas turbine engine 10 is using the fuel; and determining the fuel calorific value as a function of the monitored engine parameters (e.g., fuel combustion rate, temperatures, pressures, thrust generated by the shaft speed(s), etc.).In the event of a significant discrepancy between the entered and determined values, aircraft 1 may return to the terminal for further investigation as a safety measure, and / or a propulsion system command based on knowledge of the fuel calorific value may not be carried out. A propulsion system 2 for an aircraft may therefore include a fuel tracking system 203 arranged to: receive a data input of the calorific value of fuel supplied to aircraft 1 during refueling; determine at least one of the following: 1, the calorific value of the fuel supplied to the gas turbine engine 10 during operation; and li. the calorific value of the fuel supplied to aircraft 1 during refueling; and provide an alert if the determined calorific value is inconsistent with the received calorific value data input. The 203 fuel tracking system can be supplied as a separate fuel tracking unit integrated into the propulsion system 2, and / or as software and / or hardware incorporated into pre-existing aircraft control systems. In some examples, the fuel tracking system 203 can store calorific value data for any fuel currently in the aircraft's tank(s) 50, 53, and can have this data updated after refueling. The data can be stored separately from the circuits performing the calculations and retrieved as needed—regardless of where the data is stored, this storage can be considered part of the fuel tracking system 203, whether or not it is integral or physically connected in any way. The fuel tracking system 203 includes a fuel composition tracker 202. The fuel composition tracker 202 in the example described includes a memory 202a arranged to store current fuel characteristic data (in particular, calorific values), and processing circuits 202c arranged to calculate updated values ​​for the fuel characteristic(s) of the fuel in the fuel tank 50, 53 after refueling. The calculated values ​​can then replace the fuel characteristic data previously stored in the memory, and / or be time-stamped and / or date-added to the memory. A log of fuel characteristic data over time can thus be assembled. The fuel composition tracker 202 of the example shown also includes a receiver 202b arranged to receive data relating to fuel composition (e.g., calorific value, or data that can be used to determine calorific value) and / or requests for fuel composition information. The fuel composition tracker 202 of the example shown also includes an output module 202c arranged to provide an alert if the determined calorific value of fuel supplied to the gas turbine engine 10 and / or newly added fuel is inconsistent with the calorific value data input received, for example by sending a signal to an alarm or other system. In examples where a calorific value is consulted or read (e.g., from scanning a barcode), this value can be automatically transferred to the fuel composition tracker 202, or can be entered into a user interface of, or associated with, the fuel composition tracker 202, e.g., by a pilot or technician. In certain applications, the calorific power control process 2030 can be incorporated into a refueling process 2020 as described previously with respect to [Fig. 5], as illustrated in [Fig. 9]. In such applications, the fuel composition tracker 202 can also be called a refueling manager, or can be part of a manager fueling 206. In step 2022, the amount of energy required for a planned flight profile, optionally including a safety margin, is obtained. In this context, the amount of energy "obtained" means that a numerical value or measurement of the required energy is somehow made available for use in process 2020, 2030—whether by calculation or other determination (e.g., from a lookup table), by communication with another system, by user input (e.g., to a graphical user interface), by retrieval from memory, or by any other suitable method. For example, the fueling manager 202, 206 may receive a message containing an indicative value of the energy requirement, or may calculate such a value based on knowledge of the intended flight profile and the aircraft 1 and the aircraft's load. The term "flight profile" is used as defined above. In step 2032, a data input of the calorific value of a fuel supplied to aircraft 1 during refueling is provided, as described above. Steps 2022 and 2032 can be carried out in either order, or simultaneously. It should be borne in mind that, if tank 50, 53 is not currently empty, the energy content of the fuel already on board can also be obtained — e.g., by recovering a previously stored calorific value for that fuel, and by calculating the energy content of the remaining fuel in reference to it and the quantity of fuel remaining in tank(s) 50, 53. In step 2026, the amount of available fuel needed to provide the required energy is calculated, based on the values ​​obtained. The calculated amount can be a volume and / or a mass of fuel—refueling can be performed volumetrically or gravimetrically. In some examples, the calculation can be performed by the fuel manager's own fueling system (step 206). In examples where fuel remains on board, the energy of that fuel can be subtracted from the energy required for the planned flight profile before calculating the amount of available fuel needed. In step 2028, the calculated quantity (mass or volume) of available fuel required is distributed so that aircraft 1 can be refueled accordingly. For example, a refueling manager 206 can display the quantity to a graphical user interface, transmit the quantity to another system for a refueling technician, or provide the value to an automated system so that this system automatically completes the refueling process once the calculated quantity of fuel has been supplied. Updated values ​​for the calorific value of the fuel in the tank of Fuel values ​​50, 53 after refueling can be calculated and stored for future use, possibly by the refueling manager 202, 206. In examples where an aircraft has multiple fuel tanks 50, 53 that are fluidly connected such that the fuels in the tanks 50, 53 are equivalent, a single fuel calorific value can be stored and updated. In examples where an aircraft has multiple fuel tanks 50, 53 that are not fluidly connected, such that there may be differences between the fuels in the different tanks 50, 53, a separate calorific value can be stored and updated for each tank. The 2020 procedure may further include the 2029 command of an automated refueling system to refuel the aircraft 1 in line with the calculated quantity of fuel. Alternatively, the 2029 refueling may be carried out manually, e.g., by a person reading the calculated quantity dispensed (mass or volume) of the required available fuel, possibly from a display associated with the connection port 62 of the aircraft's fuel line. In some examples, refueling can be started before one or more of steps 2022 to 2028 are performed; these steps can be performed during refueling, and the refueling process can then be completed when the calculated amount of fuel has been added. A 2034, 2036 check is then performed to verify the entered data, as described above. The control may include measuring the calorific value of the fuel being used in the gas turbine engine 10, possibly during engine warm-up or aircraft taxiing 10 prior to takeoff. The determination of the fuel calorific value may be carried out by monitoring engine parameters during an initial period of aircraft operating time during which the gas turbine engine 10 is using the fuel; and determining the fuel calorific value based on the monitored engine parameters (e.g., fuel burn-up rate, temperatures, shaft speeds, pressures, thrust generated, etc.). In the event of a significant discrepancy between the input and determined values, an alert is provided, and the aircraft may return to the terminal for further investigation as a safety precaution. The check can be carried out during or after refueling. The 2020, 2030 process can then be iterated at each refueling event. In certain examples, the 203 fuel tracking system can therefore be further configured to: obtain 2022 a quantity of energy required for a planned flight profile, including the safety margin; obtain 2032 a fuel calorific value available to the aircraft; Calculate the quantity (mass or volume) of available fuel needed to provide the required energy; and deliver at output 2038 the quantity (mass or volume) of available fuel necessary to enable the aircraft to be refueled accordingly. In such examples, the fuel tracking system 203 can be called the fueling manager 206. It should be borne in mind that the examples described with respect to Figures 4 and 5 can therefore be combined with the examples described with respect to Figures 6 to 9. In such examples, the calculation of the mass or volume of available fuel needed to provide the required energy includes obtaining a fuel calorific value already in the fuel tank 50, 53, and subtracting this from the energy required for the intended flight profile. The method may further include calculating the calorific value of the blended fuel after refueling. The calorific value of the fuel supplied to the gas turbine engine 10 can only be determined to be inconsistent with the received calorific value data input when the calculated calorific value for the blended fuel does not correspond to the determined calorific value of the fuel supplied to the gas turbine engine 10 during operation. The inventors were aware that, since different fuels can have different properties while still meeting standards, knowledge of the fuel(s) available to an aircraft 1 can enable more efficient, tailored control of the propulsion system. For example, switching to a fuel with a higher calorific value can allow a lower fuel flow rate to meet an aircraft's energy requirements at a particular point in the flight envelope, thus potentially supplying more fuel to auxiliary systems (e.g., fuel-driven hydraulic actuators or fuel oil heat exchangers) if the total pumped flow rate is constant. Knowledge of the fuel's calorific value can therefore be used as a tool to improve aircraft performance in flight even if it is not known at the refueling point, e.g.by assessing the probability of contrail formation behind an aircraft 1 under given atmospheric conditions and by changing fuel source or altitude during flight. In addition, the in-flight determination of calorific value can be used as a check to verify the data provided or before refueling. As discussed above, the calorific values ​​(also called heating powers) of fuels can be directly determined – for example, by measuring the energy released when a certain volume or mass of the fuel is burned in the gas turbine engine 10 — or calculated from other fuel parameters. In the examples currently being described, with respect to Figures 10 to 12, the calorific value of the fuel is directly determined by measuring the energy released (or by taking measurements which allow the energy released to be deduced) when a certain volume or mass of the fuel is consumed within a gas turbine engine 10, 44 of the aircraft 1. The performance of the aircraft's own main / propulsive gas turbine engine 10 and / or a gas turbine engine of an auxiliary power unit (APU) 44 is therefore used to determine this fuel characteristic. In addition to the propulsion system 2 described in relation to Figures 4 and 6, a power system 4 of the application shown in [Fig. 10] includes an auxiliary power unit (APU) 44. The more general term "power system" 4 can be used instead of propulsion system 2 when system 2,4 does not solely provide propulsive power, or when system 4 provides no propulsive power at all. A propulsion system 2 is an example of a power system 4. The APU 44 is a smaller gas turbine engine than the 10 on the wings of aircraft 1, and is arranged to provide electrical power to aircraft 1's systems; for example, lighting, heating, air conditioning, and / or the like. The APU 44 may be, for example, an APU in the Honeywell 331 series, such as the HGT1700 Auxiliary Power Unit (APU). In some applications, the APU 44 may be certified for in-flight use; in other applications, it may be certified for ground use only. An aircraft APU 44 is typically arranged to be started using one or more aircraft batteries to provide electrical power and possibly bleed air for air conditioning and engine starting. The APU 44 of the application shown is located towards the rear of the fuselage, and is not arranged to provide any propulsive power to aircraft 1.In alternative applications, the APU may be located differently (e.g., within a nacelle 21 of the aircraft 1), and / or may provide some propulsive power. In the example shown in [Fig. 10], the central fuel tank 50 is arranged to supply fuel to the APU 44; refueling arrangements may vary in other examples. Engine parameters during an initial period of aircraft operation time during which the gas turbine engine 10, 44 uses fuel are detected and possibly monitored over time. These parameters may include one or more of the following: thrust / propulsion delivered, the volume (or mass) of fuel used in a given time (e.g., calculated from a fuel pumping rate / fuel flow rate, taking into account a ratio of de- pump discharge (if applicable), exhaust temperature, one or more shaft speeds, one or more temperature readings from other components / in other locations, and / or one or more pressure measurements. In some examples, it may be assumed that 100% of the fuel supplied to a combustion chamber 16 of the gas turbine engine 10 is completely burned. In other examples, different assumptions about combustion completeness may be made where applicable. Sensors 224, for example temperature sensors 224a and pressure sensors 224b, can be supplied in association with the gas turbine engine or each gas turbine engine 10 in order to monitor the performance of the gas turbine engine 10. Depending on the monitored engine parameters, a calorific value of the fuel can then be determined. As discussed above, it should be borne in mind that checking fuel calorific values ​​can be important, particularly in examples involving manual input of calorific value data. The determined calorific value can therefore be checked by a second determination (and possibly additional determinations), and / or used to check a manually entered (or otherwise provided) value. An alert can be provided if the determined calorific value of the fuel is inconsistent with the received calorific value data input. Examples such as those currently described with reference to Figures 10 to 12 can therefore be combined with examples previously described with reference to Figures 6 to 9. In various examples, an aircraft 1 receives calorific value data input for fuel supplied to the aircraft 1 during refueling, and a comparison of the entered value with the determined value can provide this verification. For example, a barcode of a fuel to be added to a fuel tank 50, 53 of aircraft 1 can be scanned to read fuel data, or a tracer substance (e.g., a dye) can be identified, and fuel properties can be consulted based on that tracer. Alternatively or in addition, calorific value can be provided to a pilot or technician, e.g., on a written label or orally, for input into an aircraft system. Calorific value data can therefore be entered manually or transmitted to aircraft 1. In examples where calorific value is consulted or read (e.g.e.g., from the scanning of a barcode), this value can be automatically transferred to a fuel tracking system 203 of the aircraft 1, or can be entered into a user interface of, or associated with, fuel tracking system 203 e.g., by a pilot or technician. In some applications, the calorific value is determined by measuring or by otherwise determining the energy released when a certain volume or mass of fuel is consumed can be verified against a value determined in a different way, such as any of the determination methods described above. The calorific value of the fuel can therefore be determined physically from the results of combustion under the wing, and possibly verified with values ​​supplied to aircraft 1, determined chemically from detected chemical properties, or calculated from other detected fuel properties (either under the wing or off-site), for example using any of the detection techniques described above. In alternative or additional examples, the process of determining the calorific value as a function of engine performance can be repeated, and the values ​​compared, thus providing verification using the same approach. In alternative or additional examples, a calorific value as a function of engine performance can be determined in two different engines, thus providing verification; e.g., two different 10 propulsion gas turbine engines, or one 10 propulsion gas turbine engine and one 44 APU gas turbine engine. . Repeating the detection and determination process to obtain a second calorific value can be described as performing the same procedure in a second time period. The first and second time periods may be at different stages of aircraft operation (e.g., ground operations versus cruise, climb versus cruise, or cruise at a first altitude versus cruise at a second altitude), or they may be at the same stage of aircraft operation—therefore, in some cases, no changes can be made to the propulsion system control and / or the ambient conditions may be the same between the two time periods, such that the same values ​​of the monitored parameters (within the errors / natural variation) would be expected. If a discrepancy between the determined calorific value and the received calorific value data exceeds a threshold, an alert may be provided—e.g., an audible and / or visual alarm, and / or a message sent to a pilot or other responsible party. It should be borne in mind that, when the determined calorific value is that of a fuel mixture created on board aircraft 1, the determined value is not directly compared to the received calorific value data; instead, adjustments are made to one or the other to reflect the contribution of the different fuel(s). If a sufficiently large discrepancy is identified, the flight may be interrupted subject to a recheck of the fuel, and / or no specific command to Fuel for propulsion system 2 cannot be provided. In some examples, an aircraft 1 may include multiple fuel tanks 50, 53, and two or more of the fuel tanks may contain different fuels, which may have different calorific values. In such cases, a determination of the calorific value for the fuel in each tank 50, 53 may be made—for example, by using a fuel management system to initially supply 100% of the fuel from one tank and make a first determination, and then using the fuel management system to initially supply 100% of the fuel from the other tank and make a second determination.Alternatively or in addition, calorific values ​​for two (or more) different mixtures of two (or more) different fuels can be determined from the engine's performance on the specific mixtures, and the calorific values ​​of the fuels in individual tanks 50, 53 can be calculated based on these determinations. In other cases of this type, the two or more different fuels can still be supplied in the same ratio in a mixture, and the calorific value of the mixture can therefore be the only value of interest. In the examples described here, the calorific value of fuel supplied to the gas turbine engine 10 in use is determined. This fuel may include a fuel mixture with which the aircraft 1 was recently refueled, and any fuel remaining in the aircraft's tank(s) 50, 53 following a previous refueling event, and may include a mixture of fuels from different tanks. The determination(s) are performed in an early stage of aircraft operation, and the results can then be used to influence the control of the propulsion system 2, or of the aircraft 1 more generally, in one or more subsequent stages of aircraft operation. For example, the determination of the calorific value of the fuel supplied to the gas turbine engine 10 in operation can be carried out during the taxiing of the aircraft 1 before takeoff, and / or during other ground operations, and the result can be used to influence the control during one or more of the following stages: takeoff, climb, and cruise. Particularly in applications where an auxiliary power unit 44 is used for the determination, the determination can even be carried out at an airport gate, before the main gas turbine engine(s) 10 are started.Alternatively, the determination can be made under cruise conditions and the results used to influence the control later in the flight, e.g., during cruise later in the same flight. We will bear in mind that cruising conditions generally represent a large proportion of most commercial flights, and that the optimization of the propulsion system control for cruise therefore provides optimization for the bulk of the flight envelope in most cases. The 2040 process carried out is illustrated in [Fig.11]. In step 2042, engine parameters are detected or monitored during an initial period of time in which a fuel of interest is consumed in a gas turbine engine 10, 44 of aircraft 1. This period of time can be described as an initial period of aircraft operation time and can occur at any aircraft operating point, including when aircraft 1 is stationary (e.g., at a gate). Notably, in examples where engine parameters are monitored while aircraft 1 is stationary, the gas turbine engine used may be an APU 44 of aircraft 1, rather than one of the engines 10 arranged primarily to provide propulsive power. The first time period can be an idling period during the initial engine start-up. This can allow the fuel calorific value to be determined before the flight begins, and the operation of engine 10 can subsequently vary according to the determined calorific value. The method 2040 described here can therefore be used as part of an active control scheme. As an example of monitored engine parameters, the mass fuel flow rate in the combustion chamber 16, the shaft speed of one or more shafts 26, 27 of the engine 10, and / or one or more pressures and temperatures can be detected—either instantaneously at a point in the first time period, or with monitoring over the first time period. The fuel flow rate in the combustion chamber 16 can be measured directly—many current aircraft 1 have a fuel flow meter at this location, and a meter could be added if not. Alternatively, the fuel flow rate at the combustion chamber inlet can be inferred from data collected elsewhere—e.g., from the position of a fuel metering valve (such valves commonly provide position feedback), or from one or more pumps or flow meters located elsewhere.In examples with a positive displacement pump rather than a gravity pump, the mass flow rate can be calculated from the volumetric flow rate and fuel density, or the calorific value calculation can be adjusted accordingly. Furthermore, it should be borne in mind that the current aircraft 1 routinely monitors one or more shaft speeds, and that this information, like other data, is often provided to an electronic engine controller (EMC) 42. Step 2042 can therefore be performed without requiring any new sensors. In step 2044, a calorific value is determined for the fuel consumed, using the monitored engine parameters. For example, the mass flow rate of fuel used to achieve a given speed of the low-pressure shaft 26, or the speed of any of the drive shafts 26, 27, can be used to calculate the fuel's calorific value (either directly or using a lookup table of known values). In some applications, a single measurement can be taken to determine the calorific value. In other cases, the parameters (e.g., fuel flow rate and shaft speed) can be monitored over a longer operating period with the same fuel, for example, to improve the reliability of the calculation. The change in the relationship between fuel mass flow and shaft speed can be at least about 1:1 for the change in fuel calorific value (assuming there is no shaft gearing). In some applications, measurements can be taken both in the first time period during which the first fuel of interest is consumed in a gas turbine engine 10 and in a second time period during which a second fuel of interest is consumed in the same gas turbine engine 10. Other than the fuel change, all other engine control options / engine operating characteristics can be held constant between the first and second time periods, such that the fuel change is the only variable, and a system response (in terms of a change in one or more of the monitored parameters) can be attributed to the fuel change alone. The change in the relationship between the fuel mass flow rate and the speed of the selected shaft can have a close 1:1 relationship to the change in calorific value.The calorific value of the second fuel can therefore be determined based on knowledge of the calorific value of the first fuel, possibly as a relative value compared to that of the first fuel. As an alternative to considering shaft speed, the mass fuel flow rate can be kept constant during fuel changes, and a change (if any) in the temperature rise across the combustion chamber 16 can be taken into account. The combustion chamber outlet temperature – T40 – can be compared to the compressor outlet temperature – T30 – to obtain a measure of this temperature change (with the compressor outlet temperature closely corresponding to the combustion chamber inlet temperature). As used here, T30 and T40, and any other numbered temperatures, are defined using the station numbering listed in SAE AS755, in particular: T30 = High-pressure compressor outlet temperature (HPC) * T40 = Combustion outlet temperatures In current engines, T40 is generally not measured directly using conventional measurement technology, such as thermocouples, due to the high temperature. A direct temperature measurement can be taken optically, but alternatively or in addition, a value for T40 can instead be modeled or deduced from other measurements (e.g., using readings from thermocouples used for temperature measurement at other stations and knowledge of the architecture and thermal properties of the gas turbine engine). For a fuel with a higher calorific value, an increase in the temperature rise across the combustion chamber (T40-T30) would be expected, and vice versa, if the mass fuel flow rate and the change in combustion chamber temperature can be used as an alternative to shaft speed, or as a control for a calculation based on shaft speed. For the examples listed, the calculation of a change in calorific value when switching fuels is described. It should be kept in mind that absolute values ​​can be calculated, but considering a change in shaft speed and / or combustion chamber temperature rise (or another parameter) for a fixed fuel mass flow rate can offer improved accuracy in cases where there are relevant accuracy limitations on how the fuel mass flow rate can be measured. When changes are being evaluated, as described above, it may be desirable to have the first and second time periods as close as reasonably possible—a small interval may be left to ensure a complete fuel change in the combustion chamber 16 and to allow any transient effects to pass. The size of the required interval (if any) may depend on the fuel flow rate under the operating condition. The gas turbine engine 10 typically reacts almost instantaneously (within one second) to fuel differences once the fuel reaches the combustion chamber 16, and speed probes used for shaft speed measurements generally have a small time constant.Under conditions of relatively low power and low fuel flow, an interval of approximately ten seconds from the moment fuel enters the pylon connecting engine 10 to the airframe of aircraft 1 can be used. At higher power, where the fuel flow can be four times greater or more, an interval of two to three seconds from the fuel change at the pylon inlet may be appropriate. It should be borne in mind that the travel time from a fuel tank to engine 10 can vary depending on the tank's location as well as the fuel flow, and can be accommodated accordingly with knowledge of the specific aircraft 1 – the pylon inlet is [missing information]. so mentioned here to facilitate generalization, although the change of time from the opening or closing of a valve at or near a fuel tank 50, 53, or the activation or deactivation of a fuel pump 108, can be used in various applications, with the interval calculated with reference to the fuel flow time between the point of interest and the engine 10. The interval can therefore be arranged to allow the time required to flush the first fuel from the fuel supply pipes and for the second fuel to reach the combustion chamber 16, as well as a new static state to be reached. Furthermore, measurements can be averaged over a period of time (e.g., 5 to 30 seconds) within each time period, or within the second time period only, and any trends examined, to verify that a new steady state has been reached and / or to improve reliability. In other examples, the transient behavior itself can be used in the determination—no gaps can be left in such cases, and parameters can be monitored for a single time period encompassing the change. In certain applications, the propulsion system 2 of aircraft 1 can then be controlled 2046 differently depending on the determined calorific value of the fuel. The 2040 method can also be used to obtain near-instantaneous measurements during flight, for example, controlling the fuel calorific value when fuel is drawn from a different tank 50, 53, or a combination of tanks. This 2040 method can then be used as part of an active control scheme to control the subsequent operation of the gas turbine engine 10. A change in calorific value can be detected, e.g., as a function of a change in the relationship between fuel flow and shaft speed, and engine performance can be controlled accordingly, for example, by changing one or more of the following: fuel flow; pump discharge; altitude; guide vane pitch (where variable-position vanes are provided); and fuel (where multiple different fuels are available on board the aircraft 1). Thus, once the calorific value(s) of one or more fuels on board aircraft 1 have been determined, the propulsion system 2 can be controlled according to the calorific value(s) determined, for example, by: * Changing an operating parameter of an aircraft thermal management system 1 (e.g., a fuel-oil heat exchanger), or changing the fuel temperature supplied to the combustion chamber 16 of the engine 10. * When more than one fuel is stored on board an aircraft 1, the choice of fuel to be used for which operations (e.g., for ground operations as opposed to flight, or for operations with different thrust requirements) can be performed depending on the fuel. A fuel distribution system can therefore be appropriately controlled according to the fuel characteristics. * The adjustment of one or more flight control surfaces of aircraft 1 so as to change course and / or altitude according to knowledge of the calorific value of the fuel(s). * The change in the discharge percentage of a fuel pump 108 (i.e., the proportion of pumped fuel recirculated instead of being passed to the combustion chamber 16, as discussed in more detail below). The pump 108 and / or one or more valves can therefore be appropriately controlled according to the fuel characteristics. "The change in the planning of variable inlet guide vanes (VIGV, Variable Inlet Guide Vane), the VIGVs can therefore be moved, or a movement of the VIGVs is canceled, as appropriate depending on the fuel characteristics. It should be borne in mind that the fuel flow rate at the inlet of the combustion chamber 16 is generally already measured in modern gas turbine engines 10, with a gravimetric fuel flow meter often being supplied. Conversions can be made, where applicable, where a volumetric flow meter is supplied. In addition, the fuel flow rate can be supplemented or alternatively deduced from the position of a fuel metering valve that controls the fluid flow in the combustion chamber 16 and / or other circuits—such fuel metering valves generally provide position feedback, but, especially at cruising speed as opposed to idle, improved accuracy can be provided by a flow meter downstream of the fuel metering function.Similarly, shaft speeds are already recorded in modern gas turbine engines 10, so no additional hardware / sensors may be required to implement the method 2040 described here. The method 2040 can therefore be implemented in software, possibly as part of CEM 42, without requiring any physical modifications to the gas turbine engine 10. A propulsion system 2, or other power system 4 as discussed above, for an aircraft 1 may therefore include a fuel follower 202 arranged to: monitor 2042 engine parameters during an initial aircraft operating time period during which a 10.44 gas turbine engine uses fuel; and determine 2044 a calorific value of the fuel based on the monitored engine parameters. The fuel tracker 202 can then provide the calorific value determined in as an output. The determined calorific power can be supplied to an aircraft control system 42, for example, to be used to influence the control 2046 of the propulsion system 2. The fuel tracker 202 can be supplied as a separate refueling management unit integrated into the propulsion system 2, and / or as software and / or hardware incorporated into pre-existing aircraft control systems. In some examples, the fuel tracker can store calorific value data for any fuel currently in the aircraft's tank(s) 50, 53, and can cause this data to be updated based on new determinations (e.g., triggered by refueling). The data can be stored separately from the circuits performing the calculations and retrieved as needed—regardless of where the data is stored, this storage can be considered part of the fuel tracker 202, whether or not it is integral or physically connected in any way. The 202 fuel tracker can be part of a 203 fuel tracking system. The fuel monitoring system 203 includes the fuel composition tracker 202. The fuel composition tracker 202 of the example described includes a memory 202a arranged to store received values ​​for monitored engine parameters and determined fuel calorific values, and a processing circuit 202c arranged to calculate calorific values ​​based on the received values ​​for the monitored engine parameters. The calculated calorific value can then replace fuel characteristic data previously stored in the memory, and / or be time-stamped and / or date-added to the memory. A log of fuel characteristic data (in particular, calorific value, although other characteristics can also be stored) over time can thus be assembled. The fuel monitoring system 203 includes, or is in communication with, one or more sensors 224. The sensors 224, for example, temperature sensors 224a and pressure sensors 224b, are associated with a gas turbine engine 10 so as to monitor the performance of the gas turbine engine 10. The data from these sensors is used, possibly in conjunction with other supplied data, to calculate the fuel calorific value as a function of the engine performance when that fuel is burned. The fuel composition tracker 202 of the example shown also includes a receiver 202b arranged to receive data relating to fuel composition (including monitored engine parameters, or values ​​calculated elsewhere from them) and / or requests for fuel composition information and / or requests for composition information fuel. The fuel composition tracker 202 of the example shown also includes an output module 202d. In some applications, the output module 202d can be arranged to provide an alert if the determined calorific value of fuel supplied to the gas turbine engine 10 is inconsistent with a received calorific value data input, for example by sending a signal to an alarm or other system. In certain examples, the 203 fuel tracking system can therefore be further configured to: receive a fuel calorific value data input in aircraft 1, supplied to aircraft 1 during refueling: compare the input calorific value data to the determined calorific value; and provide an alert if the determined calorific value of fuel supplied to the gas turbine engine 10 is inconsistent with the received calorific value data input. In alternative or additional implementations, the output module 202d can send a message — for example, including a fuel calorific value, or a control instruction based on the calorific value — to an aircraft control system 42, e.g., an electronic engine controller (EMC), so as to influence the control of the propulsion system 2 based on fuel characteristics. The inventors were aware that, since different fuels can have different properties while still conforming to standards, knowledge of the fuel or fuels available to an aircraft 1 can enable more efficient, tailored control of the propulsion system 2, and in particular it may be appropriate to change the discharge rate around a fuel pump 108 of the gas turbine engine 10. For example, switching to a fuel with a higher calorific value may allow a smaller fuel flow to the combustion chamber 16 to meet the energy requirements of an aircraft at a particular point in the flight envelope, thus potentially supplying more fuel to auxiliary systems (e.g., fuel-driven hydraulic actuators or fuel oil heat exchangers) if the total pumped flow is constant. As shown in Figures 4, 6, and 10 and described above, an aircraft may include several fuel tanks 50, 53; for example, a larger primary fuel tank 50 located in the aircraft fuselage, and a smaller fuel tank 53a, 53b located in each wing. In other examples, an aircraft may have only one fuel tank 50. Many arrangements different fuel tanks are possible as described above, and tanks 50, 53 can form a single fuel source, or multiple fuel sources. The fuel used in a gas turbine engine 10 of the aircraft can therefore vary during flight (where an aircraft 1 has multiple distinct fuel sources) as well as between flights (since an aircraft 1 can be refueled with a different fuel). Tanks 50 and 53 can contain different fuels—for example, with one tank 50 containing jet fuel and another tank 53 containing SAF, or a kerosene-SAF mixture. The different fuels can be mixed en route to the combustion chamber 16. The percentage of SAF in a fuel supplied to the engine 10 can therefore vary between 0% and 100% during the operation of the aircraft 1 in some examples. The SAF can have a density (p) of between 90% and 98% of the density of kerosene. SAF can have a calorific value, (PC) between 101% and 105%, the calorific value of kerosene (calorific value being as defined above).For example, the calorific value of kerosene can be 43.1 MJ / kg (with a current minimum calorific value permitted in the fuel specification of 42.8 MJ / kg), while the calorific value of SAF can be 44.2 MJ / kg. The calorific value and density of fuel mixtures can vary accordingly, with a density of 90 to 100% of that of kerosene and a calorific value of 100% to 105% of that of kerosene. The aircraft | examples currently described include a fuel pump 108 arranged to pump fuel from one or more reservoirs 50, 53 to the gas turbine engine 10. The fuel pump 108 has an inlet 108a arranged to receive fuel and an outlet 108b through which fuel leaves the pump 108. The fuel received may be from a single tank, or may be a mixture from a combination of tanks 50, 53. The fuel delivered to the gas turbine engine 10 may therefore have a different composition from any fuel stored in a tank in some examples. The fuel may pass through one or more elements of a heat management unit, or other engine components, en route to the inlet 108a of the pump 108.In addition, the fuel leaving pump 108 is not brought entirely to the combustion chamber 16; rather, some is recirculated ("spillover"), and the recirculated fuel generally forms a proportion of the fuel entering the intake of pump 108a. The proportion of fuel passing through pump 108 that is recirculated is called the spillage or spillage percentage, namely: Total fuel flow, Q, at the pump outlet - Fuel flow into the combustion chamber - Spill — —————— 0 raanmanaraanmnnn rare pineapple antenna Total fuel flow, Q, at the pump outlet The recirculated fuel may include fuel returned directly from the outlet of pump 108b to the inlet of pump 108a, for example via a fuel return line 46 ([Fig. 13], flow S). In other examples, no fuel return line 46 may be present. The recirculated fuel may include fuel diverted from the combustion chamber inlet to serve other roles in auxiliary systems or engine components ([Fig. 13], flow S), for example, with the fuel acting as a heat transfer medium in one or more heat exchangers, or as a working fluid in one or more hydraulic fuel actuators. In addition, some of the recirculated fuel may be returned to a fuel tank 50, 53 prior to further use. The spilled fuel ([Fig.13], flow S, and S>) can therefore be used to perform engine functions as well as to allow a pump 108 to be operated with a defined flow rate even when there are fluctuations in combustion chamber fuel demand. As used here, "spill" consequently includes fuel used for some purpose other than being brought into the combustion chamber 16, not just fuel sent directly from the outlet of pump 108b to the inlet of pump 108a or returned to a tank 50, 53. When an aircraft 1 is operated with kerosene in cruise, spillages of 70 to 85% are common, with spillage often reaching 98% at idle in flight. The precise spillage levels depend on one or more of the aircraft and engine designs, thrust demand, ambient temperature, altitude, and stage of cruise (e.g., due to a higher aircraft weight from the additional fuel, and often a lower altitude, at the beginning of cruise, and a lighter aircraft and often a higher altitude at the end of cruise). A minimum spillage of at least 5% or 10% may be defined to ensure that sufficient fuel is supplied to the auxiliary systems 45. The pump 108 may be sized for the maximum takeoff thrust (PMD) of an aircraft 1, possibly at low altitude. About 10% of the pump's total capacity can typically be allocated to supplying auxiliary systems 45 at PMD.In cruise relative to PMD, a pump 108 with a rotation speed linked to a drive shaft will always rotate very quickly, but the flow required by the combustion chamber 16 will be significantly lower than at PMD; the fraction of excess flow discharged is therefore generally much higher in cruise than at PMD. A new 108 pump on any 10 engine will generally have some extra capacity (effectively an excessive fuel supply) because its performance is expected to deteriorate over time – the % of total flow The discharged fuel can therefore be greater for newer pumps 108 than for older pumps. It should be borne in mind that the total fuel discharge requirements for auxiliary systems 45, etc., depend on the number of auxiliary systems with a fuel supply, the flow requirements of these systems 45, and when the flow is required (e.g., fuel-driven hydraulic actuators do not usually all move at the same time). A reserve capacity margin is also generally provided so that more discharged fuel is usually available than required. One or more sensors may be provided to detect spillage directly. For example, the fuel flow into the combustion chamber 16 may be measured directly — many current aircraft 1 have a fuel flow meter at this location, and a meter could be added if not — or deduced from measurements elsewhere, and subtracted from a pump outlet flow value, Q, known to provide a measure of the fuel flow returned to recirculation. The inventors were aware that, since different fuels can have different properties while still meeting standards, the use of SAF or kerosene-SAF mixtures can alter the desired discharge under given conditions. In particular, as the SAF content increases, a desired discharge can generally decrease. One or more valves associated with the pump 108, with the fuel return line 46, and / or with auxiliary systems or components using fuel 45 can be used to control the discharge. The pump 108 and the valves (not shown) and the fuel supply lines connecting them together form a fuel supply system 230, as illustrated in [Fig. 13] and [Fig. 14]. A fuel change spillage ratio, R4, is defined as: R = percentage spillage during cruise using kerosene / percentage spillage during cruise using a fuel with X% SAF. R4 is therefore equal to 1 when X = 0, i.e., when the fuel is pure kerosene, but varies as X increases. It should be kept in mind that, in the calculation of R4, the conditions are considered identical except for the fuel change—that is, the same engine, same stage of flight, same altitude, etc. In various examples, X% is at least 30% (X = 30), and R4 is greater than or equal to 1.003. In various examples, when the fuel mixture is 50% SAF by weight (X = 50), the fuel change spillage ratio is at least 1.0066, and when X=100 such that the fuel is pure, the fuel change spillage ratio is at least 1.0138. The following relationship can be applied in various examples where X% is at least 10%, and possibly at least 30%, illustrating the relationship between %SAF and the disbursement: x Ra > 1 +—jigp À > 10 The fuel supply system 230 is arranged to supply fuel to the combustion chamber 16 at an energy flow rate, C, optionally measured in Megawatts, mW. This energy flow rate, C, can be controlled to satisfy / equalize the energy demand of the combustion chamber to obtain a given power output / thrust under given conditions. By defining the kerosene density as pK, the kerosene calorific value as PC, and the density and calorific value of the fuel supplied to the combustion chamber 16 as pF and PC, respectively, it has proven advantageous to control the pump flow rate, Q, and the discharge ratio as a function of the fuel properties and the energy demand of the combustion chamber such that: , = percentage of spillage during cruise using kerosene d = percentage of spillage during cruise using 16 fuel with X 0 of SAP Qc 4 _ 2 / (PC, <a.) 2- ° / PUpxu,) In various examples, the gas turbine engine 10 is arranged such that, for an engine 10 with a maximum takeoff thrust in the range of 400 kN to 500 kN, in cruise: Q-(0.4595) 7 {ssgx108x PE xp) <R< Q-te 29961 / (5ssc105< PC 0) ui = 2 Q = (0-4905) / {sg0a 10" PCex0,) > 2- (595) / (ste 109 PCpXi,) where Q is measured in imperial gallons per hour, PC in CHU / lb and p in lb per imperial gallon. Converting the units to measure Q in liters / second, PC in MJ / kg, and p in kg per liter gives: Q-(0-550) ë 15 (OPLEXPCEX05) rer <R 2- (20-638 / (optexPCrxn,) L- 10-60 / (ooiaxPe,xnç) sc (G-580) / (og14x PC xA,) d In various examples, the gas turbine engine 10 is arranged such that, for an engine 10 with a maximum takeoff thrust in the range of 300 kN to 350 kN, the following relationship is maintained in cruise: @- (0.457) [sir CR < Q-(0-3857) / (e33x1054Pc.xe5 2- (Q-385T) / (3 42x 10%4PCr*0;) , = 2 C@- 3457) / {1131054 PCpX 0, where Q is measured in imperial gallons per hour, PC in CHU and p in pounds per imperial gallon. Unit conversions to measure Q in liters / second, PC in MJ / kg and p in kg per liter gives: 0027 x PCF 4, * 487) 22-47 / pormecean) gp to LU ETE / ipamec D -tt-4nn 7.501 1170 SN PRO 203 Meres A @-(0-4871 / (0p27xPO,x0,} _ @ 10-437 / fao2TxPCyx0p) The fuel change spillage ratio, R4, can therefore be ordered according to the flow properties of the pump outlet and fuel. As mentioned previously, a desired spill can vary depending on one or more of the following: ambient temperature, altitude, and cruise stage. Ra can therefore vary accordingly. In various examples, Ra is decreased by less than 0.15%, and possibly by more than 0.1%, between the beginning and end of the cruise, for a constant temperature and altitude. In various examples, Rç is decreased by at least 0.11% when the altitude increases by at least 600 m. A process 3000 carried out in certain embodiments is illustrated in [Fig.15]. In step 3002, fuel from one or more of the fuel tanks 50, 53 is supplied to the gas turbine engine 10. The fuel supplied to the gas turbine engine 10 comprises X% of SAF, where X% is in the range of 5% to 100%, with any remaining fuel being kerosene. The fuel density is denoted pr and its calorific value as PC. Correspondingly, px and PC are used for kerosene. In some implementations, at step 3004, the propulsion system 2 is controlled such that: Fuel change spill report, R4: R percentage of spillage during cruise using kerosene d — “percentage of spillage in chalk pits using fuel with X% SAF is equal to: Q- < / (PCEX D) QC / (Pecxa,) Where Q is the fuel flow rate at the pump outlet, and C is the fuel energy flow rate entering the combustion chamber. The fuel change spillage ratio Rç can therefore be controlled based on knowledge of the fuel and the current engine operation. The calorific value and / or density of the fuel can be determined or otherwise obtained using any or more of the approaches described above. Command 3004 can be repeated and updated iteratively if / when the fuel changes. In alternative or additional implementations, and when the fuel supplied to combustion chamber 16 in step 3002 comprises at least 30% SAF (i.e., X > 30) in step 3004, the propulsion system 2 is controlled such that: Rx percentage spillage during cruise using Kerosene ad 5 TESTS STE AND AND EE ETEET SRE IS CASES AND TEE EEE EEE ETES STE percentage of spillage during cruise using fuel with X% SAF is greater than or equal to 1.003. In various applications of this type, the discharge is controlled in such a way that: XR,>21+#%5; For example, when X is 50, the fuel change spill ratio can be at least 1.0066, and when X is 100, such that the fuel is pure SAF, the fuel change spill ratio can be at least 1.0138. In various implementations of process 3000, the gas turbine engine 10 can be arranged such that R; $ 1.04 and / or the gas turbine engine 10 is arranged such that Ry > 1.003, and possibly R; > 1.014. In other applications, particularly those where the calorific value of the fuel is unknown, the process 3000 can be reversed—instead of controlling R4 as a function of engine activity and fuel properties, R4 can be iteratively adjusted when switching to a new fuel until a desired energy flow rate, C4, to the combustion chamber 16 is achieved for a known (e.g., defined or sensed) pump outlet flow rate, Q4. The change in discharge, as captured in the ratio R4, can therefore be used to determine the calorific value of a new fuel, if the fuel density is known, or a calorific value multiplied by the fuel density if not. A change in R4 can thus be used to determine the fuel properties. A propulsion system 2 for an aircraft, according to the examples currently described, may therefore include a gas turbine engine 10 and one or more fuel tanks 50, 53 arranged to hold fuel to supply fuel to the gas turbine engine 10, one or more of the tanks 50, 53 containing sustainable aviation fuel (SAF) either alone or as part of a mixture. The SAF has a density between 90% and 98% of the density, px, of kerosene and a calorific value between 101% and 105% of the calorific value, PCx, of kerosene. The gas turbine engine 10, in many examples, comprises an engine core 11 including a turbine 19, a combustion chamber 16, a compressor 14, and a central shaft 26 connecting the turbine to the compressor; and a blower 23 located upstream of the engine core, the blower comprising a plurality of blower blades and being arranged to be driven by an output of the core shaft 26; in addition to the fuel pump 108. The fuel pump 108 is arranged to supply fuel from one or more of the fuel tanks 50, 53 to the gas turbine engine 10 and provides a pump outlet volumetric flow rate, Q. The fuel pump 108 has an inlet 108a arranged to receive fuel from one or more fuel tanks 50, 53 and a discharge 108b arranged to supply fuel to the gas turbine engine 10, and is arranged to recirculate (reverse) excess fuel from the discharge to the inlet (directly or indirectly), the percentage of fuel passing through the pump that is recirculated being called a recirculation percentage. The fuel supplied to the gas turbine engine 10 comprises X% of SAF, where X% is in the range of 5% to 100%, and is possibly at least 30%, with any remaining fuel being kerosene.The fuel supplied to the gas turbine engine 10 has a density, pr, and a calorific value PC;. . The propulsion system 2 is arranged so that the fuel change spill ratio, R4, is as described above. The inventors also found that a measurement of the mass and volume of the fuel F used by the aircraft can be used to determine the fuel's characteristics. For example, such a measurement of fuel mass and volume can be performed during a refueling process in which fuel is loaded onto the aircraft. Figure 16 illustrates an aircraft 1 connected to a fuel storage tank 60 for refueling as described above. The fuel storage tank 60 may be mounted on a fuel delivery vehicle (e.g., a tanker truck) or may be a fixed storage tank from which the aircraft 1 can be refueled. The aircraft 1 includes a fuel line connection port 62 that is coupled to a fuel loading line 61 during refueling. The fuel loading line 61 may include a fuel hose of known design. The fuel line connection port 62 is fluidly coupled to the fuel tanks 53, 55 of the aircraft 1 by one or more fuel transmission lines 63 on board the aircraft so that the fuel received through the fuel loading line 61 is transferred and stored in the fuel tanks 53, 55.The fuel loading line 61 and the fuel transmission line 63 can together form a fuel supply line used to supply fuel to the fuel tanks 53, 55 on board the aircraft 1 from the fuel storage tank 60. In some examples, the fuel transmission line(s) 63 may not be present, the . fuel is instead distributed directly from a fuel line connection port to each fuel tank (or set of interconnected fuel tanks). With further reference to [Fig.16], aircraft 1 further comprises a fuel characteristic determination system 102. The fuel characteristic determination system 102 is arranged to determine one or more fuel characteristics of the fuel loaded, or which has been loaded, onto aircraft 1, such characteristics being any one of those described or claimed herein. The fuel characteristic determination system 102 generally includes a mass sensor 103, a volume sensor 104, and a fuel characteristic determination module 105. The mass sensor 103 is arranged to measure a mass of fuel loaded onto the aircraft 1. In the example described here, the mass sensor 103 is arranged to measure a mass flow rate of fuel as it flows from the fuel line connection port 62 to the fuel tanks 55, 53 on board the aircraft 1. The mass sensor in this example may be a mass flow meter arranged to measure the mass of fluid passing a fixed point within the fuel supply line per unit of time.The point at which the mass flow rate is measured can be at any point upstream of the aircraft fuel tanks 53, 55 in which the fuel is stored, so that the mass flow rate of fuel being loaded onto the aircraft 1 can be measured. In this example, the mass flow sensor 103 is located in a fuel line (e.g., a portion of the fuel line 63) fluidly connecting the fuel line connection port 62 to one of the aircraft fuel tanks 53, 55. However, it can be located at other points within the aircraft fuel system, for example, at the fuel line connection port 62. The mass flow meter 103 can be a Coriolis flow meter of known design.However, any other suitable type of mass flow meter may be used, for example, any type in which mass determination is not measured indirectly based on knowledge of the fuel density. The volume sensor 104 in the example described here is arranged to measure the volumetric flow rate of fuel loaded onto aircraft 1. In the example described here, the volume sensor 104 is arranged to measure the volumetric flow rate of fuel as it flows from the fuel line connection port 62 to the fuel tanks 53, 55 on board aircraft 1. The volume sensor 104 in this example can be a volumetric flow meter arranged to measure the volume of fluid passing a fixed point within the aircraft fuel system per unit of time. Similar to the mass flow meter 103, the de- The volumetric flow meter 104 can be located at any point upstream of the aircraft fuel tank(s) 53, 55 in which the fuel is stored. It can therefore be supplied in the same fuel line as the mass flow meter 103, and can be downstream of the mass flow meter 103, as illustrated in the figures, or upstream of it. Any suitable volumetric flow meter can be used, such as a turbine or pressure flow meter. The volumetric flow meter can be of a type that does not deduce the volumetric flow rate from a measured mass flow rate. Similarly, the mass flow meter can be of a type that does not deduce the mass flow rate from a measured volumetric flow rate. The mass flow rate and volumetric flow rates are thus measured independently of each other (and without requiring knowledge of the fuel density). The fuel characteristic determination module 105 is configured to determine one or more fuel characteristics of the fuel loaded onto the aircraft 1 based on the fuel mass and fuel volume determined by the mass sensor 103 and the volume sensor 104. The fuel characteristic determination module 105 is therefore in communication with the mass and volume sensors 103 and 104, as illustrated in the figures, so that it can receive signals from them that are indicative of the fuel mass and fuel volume. In the present example, the fuel characteristic determination module 105 is a separate unit and can communicate with an electronic engine controller (EMC) 42 of each of the gas turbine engines 10 supplied to the aircraft 1.One or more specific fuel characteristics can be communicated to the CEM 42 so that the respective engine 10 can be controlled accordingly, as will be discussed below. In other examples, the determination module 105 may be part of the CEM 42 of the (or each) engine 10. The fuel characteristic determined by the Fuel Characteristics Determination Module 105 can be any of those described or claimed herein. To determine a fuel characteristic, the Fuel Characteristics Determination Module 105 is configured to compare the fuel mass and fuel volume and determine a corresponding fuel density (e.g., by dividing the mass flow rate by the volumetric flow rate). Since fuels with different fuel characteristics will exhibit a known variation in density, a characteristic of the fuel being loaded onto the aircraft 1 can be deduced based on the density. In some embodiments, one or more fuel characteristics can be determined by calculating a deviation from the density value that would be expected if the fuel were fossil kerosene.In other embodiments, the determination module 105. can be configured to access a lookup table defining a known dependence of fuel characteristics on fuel density. The measured fuel density can then be compared to the values ​​in the lookup tables so that a fuel characteristic can be determined. In one example, the fuel characteristic might be the percentage of SAF in the fuel. The inventors observed that SAF has a lower density compared to fossil kerosene, and this difference can be used to infer the percentage of SAF present in the fuel based on a measurement of the fuel's density when loaded onto the aircraft. Other fuel characteristics may also have an associated variation in fuel density. For example, the Determination Module 105 can determine that the fuel is fossil kerosene (e.g., substantially 100% fossil kerosene) if the measured fuel density is that associated with fossil kerosene fuel without SAF present. In other examples, the determined fuel characteristics might include the fuel's hydrocarbon distribution or the fuel's aromatic hydrocarbon content.In other examples, the fuel density determined from mass and volume measurements can be considered as a fuel characteristic determined by the fuel characteristic determination system 102. In the example shown in [Fig. 16], a mass flow rate and a volumetric flow rate of fuel are measured in the fuel flowing through a conduit leading to the fuel tanks on board the aircraft. In some examples, the mass sensor is arranged to measure the mass of fuel stored in the aircraft's fuel tanks during loading, or after it has been loaded. For example, the mass sensor 103 can be arranged to measure an increase in the weight of the aircraft 1 when fuel is loaded, or an increase in the weight of the aircraft's own fuel tanks 53, 55 when they are filled with fuel. The fuel characteristic determination module 105 can, in such examples, base the fuel characteristic on a total mass of fuel loaded for a flight, or the mass per unit time loaded onto the aircraft, associated with a measured mass change. The same applies to the volume sensor 104.In certain examples, therefore, the volume sensor 104 can be arranged to determine the volume of fuel stored in the aircraft's fuel tanks. For example, the volume sensor 104 may include one or more level sensors arranged to measure the fuel level within the aircraft's fuel tank(s). The volume of fuel stored can then be determined. The determination module 105 can, in such examples, base the fuel characteristic(s) on a total volume, or volume per unit time, of fuel loaded onto the aircraft. The module of... determination can therefore more generally receive an indicative signal of the mass and / or volume of fuel (e.g., a weight or level measurement), rather than a direct mass or volume measurement. In various examples, determining the mass of fuel may involve measuring the mass and / or a change in mass of any one or more of the following: the aircraft; one or more fuel tanks on board the aircraft; a fuel tanker from which fuel is supplied; or a storage tank from which fuel is supplied to the aircraft. Determining the volume of fuel may involve measuring the volume and / or a change in the volume of fuel: (a) stored in one or more fuel tanks on board the aircraft; and / or (b) stored in a fuel storage tank from which fuel is supplied to the aircraft. In the example illustrated in [Fig. 16], the determination system 102 is located entirely on board aircraft 1. In other examples, which may not be the case, one or more components of the determination system 102 are not located on aircraft 1. For example, the mass and volume sensors 103, 104 and the determination module 105 may be included in a dedicated unit that is separate from aircraft 1. In some examples, the fuel characteristic determination system 102 may be located entirely outside the aircraft. In such an example, the fuel characteristic determination module 105 may determine a fuel characteristic that is then communicated to aircraft 1 (e.g., to an engine control module or engines 10). In this example, a data transfer link may be provided (e.g.(a wireless or wired data connection) and can be used to communicate fuel characteristics to the aircraft from the Fuel Characteristics Determination System 102. In some examples, the data transfer can be done manually by a user, e.g., a technician or other system operator can obtain the fuel characteristics from the Fuel Characteristics Determination System 102 and provide them manually to a control module on board the aircraft. In some examples, some of the system components 102 may be located at the fuel supply tank 60 (e.g., on board a fuel tanker). For example, sensors 103 and 104 may be located within the fuel supply line 61, and the fuel characteristic determination module 105 may be located on board the aircraft 1. In such an example, mass and volume measurements may be communicated to the determination unit on board the aircraft using any suitable data connection (wired or wireless). In some examples, the mass and volume sensors 103 and 104 may be arranged to measure the mass and / or volume of fuel being extracted from the tank fuel storage tank 60 connected to aircraft 1. In one example, the mass sensor 103 can be arranged to determine the mass of fuel delivered to the aircraft by measuring a change in the weight of the fuel storage tank 60, or a vehicle on which the fuel storage tank 60 is transported. Similarly, the volume sensor 104 can include a level sensor arranged to measure the level of fuel contained in the fuel storage tank 60. Measurements taken on or associated with the fuel storage tank can be used to determine a total mass and / or a total volume, or a mass per unit time and / or a volume per unit time of fuel supplied to the aircraft. In some examples, the fuel characteristic determination module 105 is configured to determine one or more fuel characteristics based, at least in part, on a fuel temperature. This can help distinguish variations in fuel density caused by changes in fuel temperature from those associated with the fuel characteristic(s) being determined. In some examples, the determination module 105 is configured to obtain an indicative signal of a current fuel temperature, for example, from a fuel temperature sensor configured to directly measure fuel temperature, or from an ambient temperature sensor. In the examples described here, the 105 fuel characteristic determination module is configured to determine one or more fuel characteristics based solely on the determined fuel mass and fuel volume. In other examples, the 105 determination module can be configured to combine fuel mass and fuel volume information with inputs from other sensors or other fuel characteristic determination methods as described elsewhere here. This can allow for the inferral of a wider range of fuel characteristic types, or can improve the accuracy of fuel characteristic determination. Figure 17 illustrates a method 1002 for determining one or more fuel characteristics of an aviation fuel that can be performed by the fuel characteristic determination system 102 shown in Figure 16. The method 1002 includes determining 1003 the mass of fuel being loaded (or having been loaded) onto the aircraft 1; determining 1004 the volume of fuel of the fuel loaded (or having been loaded) onto the aircraft 1; and determining 1005 one or more fuel characteristics based on the determined mass and volume of the fuel. The mass and volume measurements can be performed during a refueling process, e.g., to measure the mass and / or volume of fuel as it is being loaded onto the aircraft, or the mass and / or the total volume of fuel that has been loaded onto the aircraft once refueling is complete (e.g., before aircraft operation). The determination 1003 of the fuel mass may include measuring a mass flow rate at a point within a fuel supply line used to convey fuel to the aircraft's fuel tanks 53, 55. As discussed above, the point at which the mass flow rate is measured may be any point upstream of one or more fuel tanks 53, 55 on board the aircraft 1. For example, the fuel flow rate may be measured at a point on board the aircraft through which the fuel loaded onto the aircraft travels to reach the fuel tank(s). In other embodiments, the fuel flow rate may be measured at a point in a fuel loading system (i.e., not on the aircraft) such as a point within a fuel loading line 61 connected to the aircraft. The determination of fuel volume 1004 may include measuring a volumetric flow rate at a point within a fuel supply line used to deliver fuel to the aircraft's fuel tanks 53, 55. As discussed above, the point at which the volumetric flow rate is measured may be any point upstream of one or more fuel tanks 53, 55 on board the aircraft 1, similarly to the measurement of mass flow rate. The volumetric flow rate may be measured at a similar or adjacent location to the mass flow rate. As discussed above, in some examples, the determination 1003 of the fuel mass may include measuring the mass or a change in mass of the aircraft 1, one or more fuel tanks 53, 55 on board the aircraft to which fuel is being supplied, a tanker vehicle from which fuel is being supplied, or a fuel storage tank 60 from which it is supplied. The change in mass of the aircraft 1, the fuel tanks 53, 55, the tanker vehicle, or the fuel storage tank 60 may be used to determine a mass flow rate per unit time, or total mass, of fuel loaded onto the aircraft 1 for the purpose of comparison with a corresponding volume measurement. The fuel volume determination step 1004 may include measuring the volume or a change in volume of fuel within one or more fuel tanks 53, 55 on board aircraft 1 or the fuel storage tank 60. The measurement of the change in volume may include measuring a fluid level within the respective tank / fuel container in which the fuel is held. The change in fuel level may be used to determine a volume per unit of time, or total volume, of fuel loaded onto aircraft 1 for comparison with the corresponding mass measurement. The determination 1005 of one or more fuel characteristics may include the comparison of the fuel mass and the fuel volume de- completed. As discussed previously, this may involve calculating 1006 a fuel density based on mass and volume. One or more fuel characteristics may be determined 1005 based on a comparison of the calculated fuel density with a known density value associated with fuel having known characteristics. One or more of the determined fuel characteristics may be any of those described here that are associated with a corresponding characteristic fuel density. In the embodiment shown in [Fig. 17], one or more fuel characteristics are further determined 1007 as a function of a fuel temperature indicator signal. As discussed previously, the fuel temperature indicator signal can be from a sensor arranged to directly measure fuel temperature, or a sensor arranged to measure ambient temperature, or otherwise input into the determination module. In this example, fuel characteristics can be determined while loaded onto the aircraft. In such examples, one or more of the determined fuel characteristics can be communicated directly to the CEM 42 if it is running during refueling, or they can otherwise be stored and communicated to the CEM when it is activated. If the CEM is not active when the fuel characteristics are determined, they can be communicated to another aircraft control system. When fuel characteristics are determined for fuel loading onto the aircraft, this fuel may be blended with fuel already present in the fuel tanks (e.g., from previous flights). The determined fuel characteristics can then be combined with those determined from previous refuelings to determine the characteristics of the fuel stored in the aircraft's fuel tanks. This can be done using a summation process in which the quantity of fuel loaded into the tanks, the quantity of fuel used during each flight, and the corresponding characteristics of the loaded fuel are recorded and combined to determine the characteristics of the fuel actually stored in the aircraft's tanks at a given time. The inventors also found that fuel characteristics can be determined during operation of the gas turbine engine, rather than during refueling. Figure 18 illustrates another example of a fuel characteristic determination system 106. In this example, the fuel characteristic determination system 106 is configured to determine one or more characteristics of the fuel being used by the gas turbine engine of the present application during the operation of the engine 10. The example in Figure 18 therefore differs from that of [Fig.16] in that the mass and volume of fuel which is distributed to a combustion chamber of the engine are determined, e.g. the mass and volume can be measured while the fuel is burned by the gas turbine engine 10. More specifically, this can be the case during the flight of the aircraft 1 on which the gas turbine engines 10 are mounted, or during the operation of the aircraft 1 while it is on the ground (e.g., during taxiing). Figure 18 illustrates a schematic view of part of the aircraft fuel system 65 and the combustion equipment 16 of the gas turbine engine 10. The combustion equipment 16 comprises a plurality of fuel nozzles (not shown in Figure 18) arranged to inject fuel into a combustion can. Fuel is supplied to the combustion equipment 16 by a fuel distribution regulator 107 under the control of the EMC 42. Fuel is delivered to the fuel distribution regulator 107 by a fuel pump 108 from a fuel source 109 on board the aircraft 1 (e.g., the fuel tanks 53, 55 described above). The fuel distribution regulator 107 and the combustion equipment 16 (which may be called simply combustion chamber) may be of known design, and may be arranged for staged (lean mixture) or rich combustion. The fuel characteristic determination system 106 illustrated in [Fig.18] generally comprises a fuel characteristic determination module 110, a mass sensor 111 and a volume sensor 112. The system 106 illustrated in [Fig.18] differs from that of [Fig.16] in that it is arranged to measure the mass and volume of fuel as it is supplied and burned by the combustion equipment 16 of the gas turbine engine 10 when the engine is operated, rather than when the fuel is loaded onto the aircraft 1. Aircraft 1 includes an aircraft fuel supply system located on board the aircraft, which is adapted to supply fuel F to each of the gas turbine engines 10 for combustion in the engine combustion equipment 16 as described above. The aircraft fuel supply system is arranged 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 supply system together form the (overall) fuel system of aircraft 1, in which fuel is stored, distributed to the engine, and burned. The aircraft fuel system includes any component that can store fuel, or through which fuel flows during operation or refueling. The mass sensor 111 is arranged to measure the mass flow rate of fluid supplied to the combustion equipment 16. In the example described here, the mass flow sensor 111 is arranged to measure the mass of fuel per unit of time. flowing between the fuel delivery regulator 107 and the combustion equipment 16. Any suitable mass flow meter can be used for the mass flow sensor 111, such as a Coriolis flow meter. The mass flow meter can be any mass flow meter that performs a mass measurement that is not a function of knowledge of the fuel density. The mass flow sensor 111 can be arranged to measure the mass flow of fuel at any point in the aircraft fuel system that is upstream of the combustion equipment 16 (e.g., upstream of the fuel nozzles of the combustion equipment 16) and downstream of the fuel source 109 from which fuel is brought on board the aircraft 1 (e.g., downstream of one or more fuel tanks 53, 55 forming the fuel source).In some examples, the mass flow rate is therefore measured at a point within the engine fuel system, such as in a fuel line within or forming part of the gas turbine engine 10 (rather than being measured by a sensor provided on the aircraft 1 to which the gas turbine engine 10 is mounted). In some examples, the mass flow rate is measured at a point immediately before the fuel is burned, e.g., immediately before it enters the combustion chamber. In still other embodiments, the mass flow rate is measured at a point within the aircraft fuel supply system, e.g., before it enters the gas turbine engine 10. The volumetric flow sensor 112 is arranged to measure the volumetric flow rate of fluid supplied to the combustion equipment 16. In the example described here, the volumetric flow sensor is arranged to measure the volume of fuel per unit of time flowing between the fuel delivery regulator 107 and the combustion equipment 16. Any suitable volumetric flow meter can be used for the volumetric flow sensor 112, such as a pressure or turbine type flow meter. The volumetric flow meter can be of a type arranged to measure the volumetric flow rate without considering a mass measurement. Similarly, the mass flow meter can be of a type arranged to measure the mass flow rate without considering a volume measurement. The volumetric and mass flow sensors can be of the same type as that described in the embodiment of [Fig. 16]. The volumetric flow sensor 112 can be arranged to measure the volumetric flow rate of fuel at any point in the aircraft fuel system that is upstream of the combustion equipment 16 (e.g., upstream of the fuel nozzles of the combustion equipment 16) and downstream of the fuel source 109 from which the fuel is drawn (e.g., downstream of one or more fuel tanks 53, 55 forming the fuel source). It can therefore be in a similar or adjacent position to the mass sensor 111 (e.g., upstream or downstream). In some examples, the flow rate Volumetric flow rate is therefore also measured at a point within a fuel line in or forming part of the gas turbine engine 10 (rather than being measured by a sensor supplied on the aircraft 1 to which the gas turbine engine 10 is mounted). In some examples, the volumetric flow rate is measured at a point immediately before the fuel is burned in a manner similar to the mass flow rate, e.g., immediately before entering the combustion chamber 16. In some examples, the volumetric flow rate is measured as part of the aircraft fuel supply system, before the fuel reaches the gas turbine engines. The fuel characteristic determination module 110 communicates with the mass and volume sensors 111, 112, and is configured to receive signals from them indicating the mass and volume of fuel per unit time being burned by the combustion equipment 16. The fuel characteristic determination module 110 is configured to determine one or more fuel characteristics of the fuel as a function of the fuel mass and volume in a manner similar to the determination module 105 described in connection with [Fig. 16]. For example, the determination module 110 can be configured to calculate a fuel density from the indicative signal of the fuel mass and volume, based on which a fuel characteristic can be found as described previously. As can be seen in [Fig.

[18] , the determination module 110 can transmit one or more fuel characteristics to the CEM 42. In other examples, it may be part of the CEM 42. The fuel characteristic(s) determined by the Fuel Characteristic Determination Module 110 can be any of those described or claimed herein. To determine a fuel characteristic, the Fuel Characteristic Determination Module 110 is arranged to compare the fuel mass and fuel volume and determine a corresponding fuel density (e.g., by dividing the mass flow rate by the volumetric flow rate). Since fuels with different fuel characteristics will exhibit a known variation in density, a characteristic of the fuel being loaded onto the aircraft 1 can be deduced. In some embodiments, the fuel characteristic can be determined by calculating a deviation from the density value that would be expected if the fuel were fossil kerosene.In other implementations, the fuel characteristic determination module 110 can be configured to access a lookup table defining a known dependence of fuel characteristics on fuel density. The measured fuel density can then be compared to the values ​​in the lookup tables so that a fuel characteristic can be determined. The fuel characteristic determination module can therefore be operated in a manner corresponding to that of the fuel characteristic determination module. of the [Fig.16]. In one example, the fuel characteristic might be the percentage of SAF in the fuel. The inventors observed that SAF has a lower density compared to fossil kerosene, and this difference can be used to infer the percentage of SAF present in the fuel based on a measurement of the fuel's density when loaded onto the aircraft. Other fuel characteristics may also have an associated variation in fuel density. For example, the Fuel Characteristic Determination Module 110 can determine that the fuel is fossil kerosene (e.g., substantially 100% fossil kerosene) if the measured fuel density is that associated with fossil kerosene fuel without SAF present. In other examples, the determined fuel characteristics might include the fuel's hydrocarbon distribution or the fuel's aromatic hydrocarbon content.In other examples, fuel density can be considered as a fuel characteristic determined by the fuel characteristic determination system 106. In the example illustrated in [Fig. 18], the mass sensor 111 and the volume sensor 112 are arranged to measure the fuel flow through the engine fuel system. In other embodiments, the determination module 110 can be arranged to receive an indicative signal of the mass and / or volumetric flow rate, which may not be a direct measurement of the fuel flow. Such a signal may be an operating parameter of the fuel pump 108, such as the pump speed or discharge rate, which has a known relationship with the mass and / or volumetric fuel flow rate. In some examples, the fuel characteristic determination module 110 is configured to determine one or more fuel characteristics as a function of fuel temperature, similarly to that described above. This can help distinguish variations in fuel density caused by changes in fuel temperature from those associated with the characteristic or characteristics being determined. This can be important in examples where the fuel is heated above ambient temperature by a heat exchanger that is part of an engine thermal management system 10. Therefore, in some examples, the fuel characteristic determination module 110 is configured to obtain an indicative signal of the current fuel temperature at the point where the volume and / or mass are measured.The temperature can be measured by a fuel temperature sensor arranged to directly measure the fuel temperature, or by an ambient temperature sensor. In other examples, the fuel characteristic determination module 110 can receive an indicative signal of the temperature of. fuel that is deduced from other engine operating parameters rather than being based on a direct measurement. In the examples described herein, the fuel characteristic determination module 110 is configured to determine one or more fuel characteristics based solely on the determined fuel mass and fuel volume obtained during the operation of the gas turbine engine 10. In other examples, the fuel characteristic determination module 110 can be configured to combine fuel mass and fuel volume information with inputs from other sensors or other fuel characteristic determination methods as described elsewhere here. This can allow for the inference of a wider range or types of fuel characteristics, or improve the accuracy of fuel characteristic determination. Figure 19 illustrates a method 1008 for determining a fuel characteristic of an aviation fuel that can be performed by the fuel characteristic determination system 106 shown in Figure 18. The method 1008 comprises: determining 1009 the mass of fuel supplied to the combustion chamber; determining 1010 the corresponding volume of fuel supplied to the combustion chamber; and determining 1011 one or more fuel characteristics as a function of the determined mass and volume. The mass and volume of fuel are determined during the operation of the gas turbine engine. As discussed previously, the determination 1009 of the fuel mass includes determining the mass flow rate of fuel being supplied to the combustion chamber 16. The mass flow rate can be determined based on a measurement taken by a mass flow meter 111 on the fuel flowing to the combustion chamber 16. The mass flow rate can be measured at any point within the aircraft fuel system that is upstream of the combustion equipment 16. For example, upstream of the fuel nozzles of the combustion equipment 16, and downstream of the fuel source 109 from which fuel is supplied to the aircraft 1 (e.g., downstream of one or more fuel tanks 53, 55 forming the fuel source). In some examples, the mass flow rate can be measured at a point in a fuel line of the engine fuel system.The mass flow rate can be measured immediately before the fuel reaches the combustion chamber 16. The determination 1010 of the fuel volume includes determining the volumetric flow rate of fuel supplied to the combustion chamber 16. The volumetric flow rate can be determined 1010 based on a fuel flow measurement using a volumetric flow meter 112 as described above. The volumetric flow meter of fuel 112 can be arranged to measure volumetric flow rate at positions equivalent to those described above with reference to the mass flow meter. In some examples, mass and volumetric flow rates may be measured at positions close to each other, or immediately upstream or downstream of each other, or at distinct locations in the fuel supply system. In any of the examples here, mass and volume are measured for the same fuel flow rate; that is, mass and volume are measured at positions on a flow path carrying the same fuel composition, and there are no additional fuel well sources in between. This allows for meaningful comparison of the measured flow rates. In examples where fuels of different compositions are stored in the aircraft's fuel tanks 53 and 55, the mass and volume sensors are positioned to measure the same fuel flow rate.For example, volume and mass sensors can be located in a pipe carrying fuel from tank 53, which may store fuel of a different composition than that of tank 55. In such an example, an additional pair of mass and volume sensors can be provided in a pipe carrying fuel from tank 55 so that the characteristics of each fuel can be measured independently. In some examples, the steps 1009, 1010 for determining the fuel mass and fuel volume may involve basing the determination on a signal from which the flow rate can be deduced rather than a direct measurement. In some examples, therefore, the determination of the fuel mass or volume is a function of an operating parameter of the fuel pump 108 or of another signal indicating the mass / volume flow rate. As discussed above, the determination 1011 of one or more fuel characteristics includes comparing the determined fuel mass and fuel volume. This may involve calculating 1012 a fuel density based on mass and volume. One or more fuel characteristics may be determined 1011 by comparing the calculated fuel density with a known value associated with fuel having known characteristics. One or more of the determined fuel characteristics may be any of those described herein that are associated with a corresponding characteristic fuel density. In the embodiment shown in [Fig. 19], one or more fuel characteristics are further determined 1013 as a function of a fuel temperature indicator signal. As discussed previously, the fuel temperature indicator signal can be from a sensor arranged to directly measure the fuel temperature, or a sensor arranged to measure the ambient temperature, or otherwise entered into the module for determining another source. The fuel characteristics determined using any of the fuel characteristic determination systems or methods in the examples herein can be used in aircraft operation, and more specifically in the operation of the aircraft's gas turbine engine(s). This can allow the aircraft's operation to be modified in response to the determined fuel characteristic. This application further provides a method 1065 for operating an aircraft 1 powered by one or more gas turbine engines 10 as illustrated in [Fig. 20]. The method 1065 may be a method for operating the aircraft 1 of any of the examples described herein. The method 1065 includes the determination 1066 of one or more fuel characteristics. This may include the use of any of the methods described herein. The method 1065 further includes the operation 1067 of the aircraft 1 according to one or more fuel characteristics. The operation of the aircraft 1067 may more specifically include the operation of the gas turbine engine(s) 10 mounted on the aircraft 1, but may also include the operation of other parts of the aircraft. Once one or more fuel characteristics are known, the gas turbine engine 10 or the aircraft can more generally be controlled or operated in various ways to take advantage of this knowledge. The operation step 1067 of the gas turbine engine or the aircraft may include the modification 1067a of an aircraft control parameter, and specifically a gas turbine engine control parameter, in response to one or more fuel characteristics. The modification of the control parameter may include any or more of the following: 1) Modifying a control parameter of a gas turbine engine thermal management system (e.g., a fuel heat exchanger) based on one or more fuel characteristics. By modifying the operation of the heat exchanger, the fuel temperature supplied to the combustion chamber 16 of the engine 10 is affected. In one example, modifying the operation of the thermal management system or changing the fuel temperature might include increasing the fuel temperature if the fuel characteristics indicate that the fuel can tolerate operation at a higher temperature without risk of coking or thermal breakdown. (ii) Where more than one fuel is stored on board an aircraft 1, the modification of a control parameter that directs the selection of which fuel to use for which operations (e.g., for ground operations as opposed to flight, for low-temperature starting, or for operations with thrust demands) may result in a change to a control parameter that directs the selection of which fuel to use for which operations (e.g., for ground operations as opposed to flight, for low-temperature starting, or for operations with thrust demands). different) depending on fuel characteristics such as %SAF, nVPM generation potential, viscosity, and calorific value. An aircraft fuel delivery system can therefore be appropriately controlled based on these fuel characteristics. The fuel delivery system can be controlled to supply the engine with fuel having a different fuel characteristic than that measured in step 1066. This may involve, for example, supplying fuel with a relatively lower aromatic content; supplying fuel with a lower SAF content; or supplying kerosene-type fossil fuel. Fuel delivery can be controlled by switching between fuel tanks or by changing a fuel mixture ratio. iii) The modification of a control parameter to adjust one or more flight control surfaces of aircraft 1, so as to change route and / or altitude based on knowledge of the fuel. (iv) Modifying a control parameter to change the spillage percentage of a fuel pump (i.e., the proportion of pumped fuel recirculated instead of being passed into the combustion chamber) of an aircraft fuel system according to one or more fuel characteristics, for example, based on the %SAF of the fuel. The pump and / or one or more valves can therefore be appropriately controlled according to the fuel characteristics. (v) Modifying a control parameter to change the scheduling of variable intake guide vanes (VIGVs) based on fuel characteristics. The VIGVs can be moved, or a VIGV movement can be canceled, as appropriate, depending on the fuel characteristics. In the examples above, the gas turbine engine or aircraft is operated according to one or more fuel characteristics by making changes to how the aircraft or gas turbine engine is controlled during operation. This can be done, for example, by an engine control system (such as the CEM 42) by making changes to various engine control parameters. Similar changes can be implemented by other aircraft control systems during operation (e.g., during flight). The CEM can more generally be called an example of a control system 42 arranged to control aircraft operation (e.g., it can be a control module of a control system). This application further provides an aircraft 1 having a fuel characteristic determination system 102, 106 according to any one or more of the examples described or claimed herein. Aircraft 1 further comprises a system for A control system is arranged to direct aircraft operation according to one or more fuel characteristics determined by the fuel characteristic determination system. The control system may include the engine ECM 42, with which the fuel characteristic determination system may be in communication or partially integrated. In other examples, other aircraft control systems may be provided with fuel characteristics, and the aircraft controlled accordingly. Operational step 1067 of the gas turbine engine or aircraft, depending on one or more fuel characteristics, can be performed automatically in response to the determination of fuel properties without any pilot intervention. In some examples, it may be performed after approval by a pilot, following notification to the pilot of a proposed change. In some examples, step 1067a may involve the automatic execution of some changes and the request for others, depending on the nature of the change.In particular, changes that are "transparent" to the pilot — such as internal changes within engine flows that do not affect engine power output and would not be noticed by a pilot — can be made automatically, whereas any changes that the pilot would notice can be notified to the pilot (i.e., a notification indicating that the change will occur unless the pilot decides otherwise) or suggested to the pilot (i.e., the change will not occur without positive input from the pilot).In applications where a notification or suggestion is provided to a pilot, this may be provided on an aircraft cockpit display, and / or sent to a separate device such as a handheld tablet or other computing device, and / or announced via an audible sound such as synthesized speech or a recorded message or a particular tone indicating the proposed / notified change. In other examples, the gas turbine engine operation step 1067 according to one or more fuel characteristics may include the supply 1067b to the gas turbine engine of fuel having different characteristics than the fuel for which one or more fuel characteristics were measured in step 1066. This provision of a different fuel may include the loading of fuel having different fuel characteristics into the aircraft's fuel tanks during aircraft refueling. In some embodiments, one or more specified fuel characteristics may include the fuel density calculated from mass and volume measurements. In such an example, the aircraft can be operated according to the fuel density. It will be understood that the invention is not limited to the embodiments described above, and various modifications and improvements may be made without departing from the concepts described herein. Except in cases of mutual exclusion, any one of the features may be used separately or in combination with any other features, and the description extends to and includes all combinations and subcombinations of one or more features described herein.

Claims

Demands

1. DC propulsion system (2) for an aircraft (1) comprising an engine gas turbine (10) and at least one fuel tank (50, 53) arranged to store fuel to power the gas turbine engine (10), in which the stored fuel comprises at least a proportion of sustainable aviation fuel — SAF, having a density between 90% and 98% of the density, px, of kerosene and a calorific value between 101% and 105% of the calorific value PCx, of kerosene, the gas turbine engine (10) comprising: a combustion chamber (16); and a fuel pump (108) arranged to supply fuel to the combustion chamber (16) at an energy flow rate, C, the pump to fuel (108) having an inlet (108a) arranged to receive fuel from at least one fuel tank (50, 53) and a evacuation (108b) arranged to deliver fuel at the outlet to a pump outlet volumetric flow rate, Q, a proportion of the fuel delivered at the outlet being supplied to the combustion chamber (16) and the the remainder being put back into circulation, the percentage of fuel passing to through the pump (108) which is not supplied to the chamber of combustion (16) being called a spillage percentage, and in which the fuel supplied to the pump (108) comprises X % of SAF, where X% is in the range of 5% to 100%, with any remainder fuel being kerosene, and has a density, pr, and a ca- lorifique PC; and in which the propulsion system (2) is arranged such that: Fuel change spill report, Re: R, = percentage of spillage during cruise using kerosene d 7 "percentage of spillage during cruise using kerosene with SAF 0 is equal to: 9-(€ / this) 2-(€ 7 pcs) |

2. Propulsion system (2) according to claim 1, wherein the gas turbine engine (10) has a maximum thrust at takeoff in the range from 400 kN to 500 kN, and is arranged such that Rd is in the following cruising range: @-(1e-sa0 fnmmercce) | p « D VE fosicpc,ny | HQE î months = e- [ (2-63) / (ovidx POEX pp ) _ e- ( (@-580) / (omaxPC,xp;}) d measured in liters per second, PC in MJ / kg, and p in kg per liter.

3. Propulsion system (2) according to claim 1, wherein the gas turbine engine (10) has a maximum thrust at takeoff in the range from 300 kN to 350 kN, and is arranged such that Rd is in the following cruising range: ; | aù 2- ( (2-37) / (0pax pe xe,) ] o-| LO-487) / (DOZTAPE XP.) | orQ: such < RE 50 The 2- (10-487) / (00274 PCpxpp) ) O-((0-423) / iourarcrxo,) ) d is measured in liters per second, PC in MJ / kg, and p in kg per liter.

4. Propulsion system (2) according to claim 1, wherein the gas turbine engine (10) is arranged such that R < 1.04,

5. Propulsion system (2) according to claim 1, wherein the Gas turbine engine (10) is arranged such that A > 1.

003.

6. Propulsion system (2) according to claim 1, wherein the A gas turbine engine (10) is arranged such that Rd > 1.

014.

7. A propulsion system (2) according to claim 1, wherein Rd varies depending on at least one of the following: ambient temperature, altitude, and the cruising stage.

8. Propulsion system (2) for an aircraft (1) comprising an engine gas turbine (10) and at least one fuel tank (50, 53) arranged to store fuel to power the gas turbine engine (10), in which the stored fuel comprises at least a proportion of sustainable aviation fuel — SAF, having a density between 90% and 98% of the density, px, of kerosene and a calorific value between 101% and 105% of the calorific value (CV) of kerosene, the gas turbine engine (10) comprising: a combustion chamber (16); and a fuel pump (108) arranged to supply fuel to the combustion chamber (16), the fuel pump (108) having a intake (108a) arranged to receive fuel from at least one fuel tank (50, 53), fuel delivered to the pump (108) including X% SAF, with any remaining fuel being kerosene, and an evacuation (108b) arranged to deliver at the outlet the fuel, a proportion of the fuel delivered at the outlet being supplied to the combustion chamber (16) and the rest being put back into circulation, the percentage of fuel passing through the pump (108) that is not supplied to the combustion chamber (16) being called a percentage of spill, and in which the propulsion system (2) is arranged such that: Fuel change spill report, Rç: percentage of spillage during cruise using kerosene — A ere da oeane à dar à ie tan nr en ne angina Se epe ina KR, — mb EE D UE EE where X% is at least 30%, is greater than or equal to 1.

003.

9. Propulsion system (2) according to claim 10, wherein R, to L'é+eies

10. Propulsion system (2) according to claim 8, wherein X% is equal to 50%, and in which the spillage ratio of change fuel is at least 1.0066°.

11. Propulsion system (2) according to claim 8, wherein X% is equal to 100%, such that the fuel is pure SAF, and in which the fuel change spill ratio is at minus 1.0138.

12. Propulsion system (2) according to claim 8, wherein the Gas turbine engine (10) is arranged such that A; "1.

04.

13. Propulsion system (2) according to claim 8, wherein the Gas turbine engine (10) is configured such that Ry > 1.

003.

14. | Propulsion system (2) according to claim 8, wherein the gas turbine engine (10) is arranged such that R > 1.

014.

15. Method (3000) of operating an aircraft (1) comprising a system of propulsion (2), the propulsion system (2) comprising an engine gas turbine (10), and at least one fuel tank (50, 53) arranged to store fuel to power the gas turbine engine (10), in which the stored fuel comprises at least a proportion of sustainable aviation fuel — SAF — having a density between 90% and 98% of the density, px, of kerosene and a calorific value between 101% and 105% of the calorific value (CV) of kerosene, the gas turbine engine (10) including a fuel pump (108) arranged to supply fuel from at least one tank of fuel (50, 53) to a combustion chamber (16) of the engine at gas turbine (10) at an energy flow rate, C, fuel pump (108) having an intake (108a) arranged to receive fuel from the minus a fuel tank (50, 53) and an outlet (108b) arranged to deliver fuel at the outlet at a volumetric flow rate of pump outlet, Q, a proportion of the fuel delivered at the outlet being supplied to the combustion chamber (16) and the remainder being returned to circulation, the percentage of fuel passing through the pump (108) that is not supplied to the combustion chamber (16) being called a spillage percentage, the process (3000) comprising: supplying (3002) fuel from at least one fuel tank (50, 53) to the gas turbine engine (10), the fuel supplied to the gas turbine engine (10) comprising X% of SAF, where X% is in the range of 5% to 100%, with any remainder of the fuel being kerosene, and wherein the fuel has a density, Pr, and a calorific value PC; and the control (3004) of the propulsion system (2) such that: the fuel change spillage ratio, Ra: R, = percentage spillage in cruise using kerosene d = "percentage spillage in cruise using fuel with SAF A 5 is equal to: 2- (e, (PES 0)} 0-{e pren)

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